Download - Project Gemini Familiarization Manual Vol2
PROJECT GEMINI
familiarizationmanual
SEDR300 COPYNO.
RENDEZVOUS and DOCKINGCONFIGURA TIONS
,,__ _____..____ O'_I_C'E LATEST CHANGED PAGES SUPERSEDE
__ THE SAME PAGES OF PREVIOUS DATEInsert changed pages into basic
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SECTION 8 IS CONTAINED IN ACONFIDENTIAL SUPPLEMENT TOTHIS MANUAL
MCDONNELLI JULY 1966CHANGED
22 AUGUST 1966
PREMI
INSERTLATESTCHANGED PAGES.DESTROYSUPERSEDEDPAGES.
LISTOFEFFECTIVEPAGES NOTE: The portion of the text affected by the changes is indicatedby a vertical line in the outer margins of the page.
TOTAL NUMBER OF PAGES IN THIS PUBLICATION IS1057 , CONSISTING OF THE FOLLOWING:
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FOEEWORD
/
Initiated by the NASA and implemented by McDonnell Aircraft Corporation, Project
Gemini is the second major step in the field of manned space exploration.
Closely allied to Project Mercury in concept and utilizing the knowledge gained
from the Mercury flights, Project Gemini utilizes a two man spacecraft considerably
more sophisticated than its predecessor. The Gemini spacecraft is maneuverable
_rithinits orbit and is capable of rendezvous and docking _Ith a second orbiting
vehicle•
PREPARED BY MCDONN_T_.T_TECHNICAL DATA DEPT.
Reviewedbyf_'
i Sr.MaintenanceEngineer
_ ReviewedbyI_ ' Sulmrz£so'r' '_e'ehniealDal_
I.
Reviewedby . /_. _--_'_._
_ _ -NASA- ResidentManager
I I
f-.
• NPROJECT GEMINI
familiarizationmanual
SEDR300 COPYNO.
_ RENDEZVOUS and DOCKINGCONFIGURATIONS
THIS DOCUMENT SUPERSEDESDOCUMENT DATED 31 MAY 1965
SECTION 8 IS CONTAINED IN ACONFIDENTIAL SUPPLEMENT TOTHIS MANUAL
MCDONNELL
I JULY 1966
f
A
SEDR300 _._
PROMINI
FOEEWORD
Initiated by the NASA and implemented by McDonnell Aircraft Corporation, Project
Gemini is the second major step in the field of manned space exploration.
Closely allied to Project Mercury in concept and utilizing the knowledge gained
from the Mercury flights, Project Gemini utilizes a two mau spacecraft considerably
more sophisticated than its predecessor. The Gemini spacecraft is maneuverable
within its orbit and is capable of rendezvous and docking with a second orbiting
vehicle.
PREPARED BY MCDONN_.T,TECHNICAL DATA DEPT.
Reviewed by S_.
Sr. Maintenance Engineer
Supez_lsor - "Technical I_ta
Revlewedby _ _. _
-NASA - Resident Manager
B
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----------PROJECTGEMINI
1lfI_ODUC'I_ON
The purpose Of this _.uuel is to describe the Gemini spacecraft systems and
_Jor components. The manual is intended as a femiliarization-indoctrination
aid and as a ready reference for detailed information on a specific system or
component. The manual is sectionalized by spacecraft systess or major assemblies.
Each section is as ccnplete as is practical to minimize the need for cro6s
referencing.
The infora_tion co_tained in this _anual (SEDR 300, VOL XI) is applicable to
rendezvous missions only and is accurate as of i April 1966.
For information pertaini_ to long range or modified (non-rendezvous) configura-
tions of the spacecraft, refer to HEDR 300, VCL. I.
C
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PROJECT GEMINI
SEC'TIOE_T_rPAGE
SECTIONI
SPAC_RAIeT MISSION ..... i-i
SECTIONlI
m.m_ s_ _smmr.xm ......................................2-I
SECTIONIZT
ZIi'_RXOR_ .............................................. 3-i
SECTIONIV
SErE SI_ .............................. _-i.--.--.--.-- .... -- . ----... . .. . ....
SECTIONV
ELECTRICALPOWER SYS_ ..............................._................ 5-1
SECTIONVI
C01T_0___L SYSTEM .............................................6-1
SECTIONVII
COOLINGSTS_ ..........................................................7-i
SECTIONVIII
GUII)AE_ AU _ SYST_ ............................................ 8-1
SECTIONIX
C0)a4_I_J_TI0_SSTSTml ...................................................9-1
S_J_20N X
INS_ATION AND RECORDING SYSTB( --'--" ....................- "- - - - - -" - I0-i
SEt.ON XI
PI_0_IICS A_D _E R0C_T ........................... " .......... ii-i
D
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PAGE
LANDING _ ......................................................... 12-1
SE@I'ZON_I
IX)CKING_ ......................................................... 13-i
SEC'I_ONXIV
TARGET DOCKING ADAF_R ................................................. 14-i
_6"1_ON XV
A[X_ _ DOgEING ADAPTER ....................................... 15-i
SECTION XVI
E_-_EHICULAR ACTIVITY ............................................... 16-I
E
SPACECRAFT MISSION
Section
TABLE OF CONTENTS
TITLE PAGE
MISSION DESCRIPTION ................................... 1-3MISSION OBJECTIVES ................................. ]-3SPACECRAFT DESCRIPTION ....................... 1-4LAUNCH VEHICLE DESCRIPTION ............... !-5CREW REQUIREMENTS ............................... 1-5 :_:...:_-_._SPACECRAFT RECOVERY 1-6 :..:.:.-:::::_:-._-.===......... ..o.o... .... ......o _
00tltl_OttQ_ _* _
:::::_:::::::::._.-=..-::_.,"_H6t_.Q_QOOQQ_° q
'°°°°°°°°°°0.°°°°0.._.°_°_''°°°O°°°°°*_H'°_°.._...
'_°°°°.°°°0°0.0.°°°°.°°_°°.°_°...°°°°°°0.0°°°°°°°°°°,.°°0.0.°°..°...°°°°.0°°°0°°,°o00°00°°0 ...... °°°..°°0°•'°°°°°°°°°0o00°°000.°o°°....°00._.°°°°.°°°°0°0°°°0°°.•,0.oo..0_._°..°°0°.°.°°°°_0.°_°.°°.°.0.._°°°°0°°°°0°°•.....°°°°°0.°°°°0°°°.°°°00°.°°.0.._°0.0°°°°..°°°°°.°°.0o._°°°o°..°°°°°°_°°°°°°°0..o.0.°o°°...°0....°°0°.°°°.°°°0.o0°0°.0°o..._0°0...°•.°..0°°°.o°.°°o...°°°0.0°°,.oo..._°_.o.°°0°..°°0°.°°°,..°...°.°°..°°°.°0°°0°°0°°..°°.o.°°°...°0°°°°°°°°°°°°°_°°.°..0°0..°00°0.°0.°°°°0,.°°°..°°°°°°°°°°.°°°°°°°°°,.... ..°°°°0.°00°0°°0°°°°°°,.... °°°°.0.°0°.°.°0°0°°°°°°.°..°°°°0°..0°00..°0°..°°°•,.°.°°°..0°°0°°°°.°°°°.°°°,.... °°0°°0.°0°°0.°°°0°.°0°°,0°°°°°.°.°°°°0°°00°o°0°°°•.°0°°°0°°°°°°0°°°°°0°°..0°•.°°°0°00°°°°°0°°..°°0°o0°.,_.°°°°0°°°°.00.000...°o°°.•.°°°°00°°0°°°00°.°..°o.0..._°.°°.°.°oo.°00°0°.o.°°..0••.. ................ .°°.0.°°
................. °.°.°°.°00•/ ............ °.°°°°..0°00°.°,..°.°...0°...°°°°.°°°°°°°•,.°°°00°°°°°°.°°.°0°°°°°°0°
...... 0.°°..°.°°.°°°°°°°°°°,°o°°.°.°°°°°°°°0°.°°0°..°°..... 00.°00..0°.°°°0°°°°°°.,.°.°0.•°°°°°.°°°°°°°.°00°°,..0°°°0°°0°.°°°°0..0°.o.o°,...0°°.0°.°°o°°.°°°o°°..oo.................... °..°.o0.................... °0...°,................ °°..°°. ........ °....... ° ......... ...°,.0..°°..°°.°0°0°°°00°°°°.0•..°0°00..0oo°..°°.°°..°°o°°...................... ..°°............... °.°°0°0°°°°°,..o°°0.°o°o.o°0°o0..00°°..,
1-1 __
--- SEDR300
RENDEZVOUS
RECO_/ERYSECTION
_ RE-ENTRY
CONTROLSYSTEM
RE-ENTRY MODULE < SECTION
LANDINGMODULE
EQUIPMENTSECTION
SPACECRAFT
SECTION -
ADAPTER
> EQUIPMENTSECTION
LAUNCH VEHICLE
/
I TITAN n LAUNCH VEHICLE
iFigure 1-1 Spacecraft Pre-Launch Configuration
1-2
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PROJEMINI
SECTION I SPACECRAFT MISSION
MISSI_ DESCRIPTION
Fundas_ntally_ the mission of Project Gemini is the insertion of a two man space-
craft into a semi-permanent orbit about the earth, the study of man's ability to
rendezvous and dock with another orbiting vehicle, and the subsequent safe return
of the spacecraft and its occupants to the earths surface. Previous missions
included manned and umaanned flights to study hmnan capabilities during extended
missions in space. Rendezvous and docking with an orbiting Agena Target Vehicle
or Au_nented Target Docking Adapter and Extra-Vehicular Activities are planned
for most missions.
MISSIONOBJECTIVES
Specifically, the project will seek to:
1. Demonstrate the ability of the spacecraft to perform in manual and/or auto-
matic modes of operation.
2. Evaluate the adequacy of major systems in the spacecraft.
3- Verify the functional relationships of the major systems and their integra-
tion into the spacecraft.
4. Determine man's requirements and performance capabilities in a space environ-
ment •
5. Determine man's interface problems, and develop operational techniques for
the most efficient use of on-board capabilities.
6. Evaluate system performance during rendezvous and docking.
7. Demonstrate the ability of the pilots to perform Extra-Vehicular Activities.
i
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$|DIt300
8. Develop operational techniques required for rendezvousing and docking with
another orbiting vehicle.
9- Develop controlled re-entry techniques required for landing in a predicted
touchdown area.
lO. Develop operational recovery techniques of both spacecraft and pilots.
SPACECRAFT DESCRIPTI_
C_EPAL
The Gemini Spacecraft (Figure I-i) is a conical structure 19 feet long and weighs
approximately 7000 lbs. Basically it consists of a re-entry module and an adapter.
RE-ENTRY MODU_E
The re-entry module consists of the heat shield, the crew and equipment section,
Re-entry Control System section and the rendezvous and recovery section. The
crew and equilznent section contains a pressurized area suitable for human occu-
pation, and a number of non-pressurized compartments for housing equipment.
External access doors are provided for equipment compartments. The Re-entry
Control System section contains the major Re-entry Control System components.
The rendezvous and recovery section contains the rendezvous radar equil_nent,
the drogue parachute and pilot parachute assemblies, and the main parachute
assembly. The rendezvous and recovery section is Jettisoned after re-entry
along with the drogue parachute.
ADAPTER
The adapter consists of the launch vehicle mating ring_ the equipment section and
the retrograde section. The launch vehicle mating ring is bolted to the launch
vehicle. A portion of the ring remains with the launch vehicle at spacecraft-
SEDR 300
PROJECT GEMINI
launch vehicle separation. The equipment section contains major c_nponents of the
Electrical, Propulsion, and Cooling Systems. The primary oxygen supply for the
Environmental Control System is also located in the equipment section. The retro-
grade section contains the retrograde rockets and some components of the Cooling
System.
LAUNCH VEHICLE DESCRIPTION
The vehicle used to launch the Gemini Spacecraft is the Gemini - Titan II, built
by the Martin C_npany. The Titan II is modified structurally and functionally to
accept the Gemini adapter and to provide for the interchange of electrical signals.
The Titan II is a two stage launch vehicle 90 feet long and i0 feet in diameter
from the thrust chamber to the spacecraft adapter. The first stage is 70 feet
long and develops approximately 430,000 pounds of thrust. The second stage is 20
feet long and develops about 100,000 pounds of thrust.
Titan II uses hypergolic (self-lgniting when mixed) propellants. Nitrogen
_troxide is the oxidizer and uns_,.-.etricaldlmethylhydrazine is the fuel. The
propellants can he stored within the launch vehicle indefinitely and ignite auto-
matieally when they are mixed in the propulsion chamber. The hypergolic propel-
lants will burn (although at a very rapid rate) rather than explode, which is
a significant safety advsntage.
CREW REQUIRemEnTS
The Gemini Spacecraft utilizes a two-man crew seated side by side. The crew
member on the left is referred to as the c_mAnd pilot and functions as space-
craft c,_._:_nder. The crew member on the right is referred to as the pilot. Crew
members are selected from the NASA astronaut group.
1-7
SEDR 300
SPACE_ _ECOVERY
The Gemini !_nding module will make a water landing in a pre-determined area. A
task force of ships, planes, and personnel will be standing by for locating and
retrieving the spacecraft and crew. In the event an abort or other abnormal
occurence results in the spacecraft landing in a remote location, electronic
and visual recovery aids and survival kits are provided in the spacecraft to
facilitate spacecraft retrieval and crew survival, respectively.
i-5
MAJOR STRUCTURALASSEMBLIES
TABLE OF CONTENTS
TITLE PAGE
GENERAL INFORMATION .......................... 2-3
RE-ENTRY MODULE .................................... 2-3 :".._.*:.__'--::'mL:-iiiiii
RENDEZVOUS AND RECOVERY SECTION ......... 2-3 :_'Z..-i'._i=:_-_i._._i-_-)N
RE-ENTRYCONTROLSYSTEMSECTION............ 2-8 iiii_N'_"_r'ii_'_'$_CABIN ........................................................ 2- 8 ::_!_iiii!ii_i_:"-_iiii_,..,........o...........°..
_.°.°. °°.°..°_°°°.°._.°°.
A DAPT ER .................................................... 2-17 iiiiiiiiiiiiiiiiiiii}iiiiii
RETROGRADE SECTION ................................. 2-17 i!iii!iHilHi!ii!iiiiiiiii,..°..........°o......°...,
EQUIPMENT SECTION .................................... 2-17 ::iiiiiii_i_iiiii!i_!!!!ii!::,,,°°...°.°°°°°.°....°..,..
SPACECRAFT LAUNCH VEHICLE MATING..2-19 iiiii_}iii_HHiiiii!i!!iiiiiiiiiiiiiiiiiiiiii!!iiiii!i_iiii!iiiii!i!iiiii!iiiil::!iiiiiiiiiiiiiiiiiiiii_i_:::::::::::::::::::::::::::::..... .o,°°° ........ °.° .......... ..,......o°.........°................... o.°o,...,::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::
.................... °°°°.°........................ °°°,
.......... ° ................
........... ° ......... °°°°°.
................ .°°°o°°°°°............. °,°°°° .........
................. ° .........
............ °.°° ....................... .°°°o°o, .......
............ °°.°°°°°,°°°°o.
................. °°°°.°.°.,
............ °..°.°°°..° ...................... ° ........
2-1 ...........................
s oR30oPROJECT GEMINI
\
Figure 2-1 Interior Arrangement(Typical)
2-2
SEDR 300
PROJ GEMINI
SECTION II MAJOR STRUCTURAL ASSemBLIES
GENERAL INFORMATION
The Gemini Spacecraft is basically of a conical configuration (Figure 2-i) consist-
ing of a re-entry module and an adapter as the two major assemblies. Spacecraft
construction is semimonocoque, utilizing titanium for the primary structure. It
is designed to shield the cabin pressure vessel from exessive temperature vari-
ations, noise and meteorite penetration (Figure 2-2). See Figures 2-3 and 2-4
for spacecraft orientation.
RE-ENTRY MODULE
The re-entry module (Figure 2-5) is separated into three primary sections which
include the Rendezvous and Recovery section (R and R), Re-entry Control System
section (RCS) and the cabin section. Also incorporated in the re-entry module is
the heat shield which is attached to the cabin, and a nose fairing which is at-
tached to the forward end of the R and R section. The nose fairing is ejected
during launch.
RENDEZVOUS AND RECOVERY SECTION
The (R and R) section (Figure 2-5), the forward section of the spacecraft, is
semiconical in shape and is attached to the Re-entry Control System section with
twenty-four bolts. Incorporated in this joint is a pyrotechnlc device which
severs all bolts causing the rendezvous section to separate from the RCS section
on signal for parachute deployment. A drogue parachute will assist in the removal
of this section. The R and R section utilizes rings, stringers and bulkheads of
titanium for its primary structure. The external surface is composed of beryllium
shingles, except for the nose fairing. The nose fairing is composed of fiberglass
reinforced plastic laminate.
2-3
.__ SEDR300 -___
PROJECT GEMINI
_-- SPACECRAFT
___ ADAPI"ER _ • RE-ENTRY MODULE
LANDING MODULE
ADAPTER
MATING ADAPTER ADAPIE?. RENDEZVOUSSECTION _ CABIN RECOVERY - •
SECTION SECTION SECTION
NOSE
FAIRII
/-- NOSE FAIRINGMATING LINE
RENDEZVOUS ANDRECOVERY SECTIONMATING LINE
RE- ENTRY CONTROL
SYSTEM SECTION/CABIN MATING LINE
RE-ENTRY MODULE/ADAPTERMATING LINE
//SPACECRAFT/LAUNCHVEHICLE MATING LINE
Figure2-2 SpacecraftGeneralNomenclature
2-4
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i
I_ 90.00 70.53
l,I0 °
20 °
// 7°
88°30DIA
Y0
219.03 (ORBIT C ONFIGURATION)
226.84 (LAUNCH CONFIGURATION)
RX LX
By
Figure 2-3 Spacecraft Dimensions
2-5
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(TOP)
xO.O0
yo.O0
XO.O0
Figure 2-4 Stations Diagram
2-6
f_ SEDR 300
"_ PROJECT GEMINI
INGRESS-EGRESS
LANDING MODULE
CABI
RE-ENTRy CONTROLSYSTEM SECTION
CABIN/ADAPTER _/_ _RENDEZVOUS ANDRETAINING STRAP __O_VERY SECTIONFAIRING(TYPICAL 3 PLACES)
OBSERVATION WINDOWS
I I _ NOSE
\ I /2 f ) FAIRING
LARGE PRESSURE
EQUIPMENT BAy
RCS THRUSTCHAMBER ASSEh_
SCANNER
-.....
-ECS EQUIPMENT DOOR
f--\/
Figure 2-5 Re-entry Module Structure
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RE-ENTRY CONTR_ SYST_ SECTION
The RCS section is located between, and mated to, the R and R and cabin sections
of the spacecraft (Fi_a_re 2-5). This section is cylindrical in shape and is
constructed of an inner titanium alloy cylinder, eight stringers, two rings and
eight beryllium shingles for its outer skin. The RCS section is designed to house
the fuel and oxidizer tanks, valves, tube assemblies, and thrust chamber assemblies
for the RCS.
A parachute adapter assembly is installed on the forward face of the RCS section
for attachment of the main parachute.
CABIN
The cabin (Figure 2-5)2 similar in shape to a truncated cone, is mated to the RCS
section and the adapter. The cabin has an internal pressure vessel (Figure 2-6)
shaped to provide am adequate crew station with a proper water flotation attitude.
The shape of the pressure vessel also allows space between it and the outer
conical shell for the installation of equipment.
The basic cabin stracture consists of a fusion welded titanium frame assembly to
which the side panels, small and large pressure bulkheads and hatch sill are seam
welded. The side panels, small and large pressure bulkheads are of double skin
construction amd reinforced by s_iffeners spotwelded in place. Two hatches are
hinged to the hatch sill for pilot ingress and egress. For heat protection, the
outer comical surface is covered with Rene' _I shingles and an ablative heat shield
is attached to the large end of the cabin section.
A spring loaded hoist loop, located near the heat shield between the hatch open- ._
ings, is erreeted after landing to facilitate engagement of a hoisting hook for
spacecraft retrieval.
2-8
,_._-_._ SEDR 300
"_'I=__' PROJ SC T G S M,N, ______,,!
Figure 2-6 Cabin Pressure Vessel
2-9
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PROJ EC'" GEMINIj-.
E_ui_nent Ba_,s
The equipment bays are located outside the cabin pressure vessel (Figure2-7). Two
bays are located outboard o_ the side panels and one bay beneath the pressure
vessel floor. The bays are structurally designed for mounting of the equipment.
Doors
To enclose the side equipment bays, two structural doors are provided on each side
of the cabin (Figure 2-7). These doors provide access to the components installed
in the equipment bays. The main landing gear bays, located below the left and
right equipment bays, are each enclosed by one door. The landing gear is not
installed but fittings are provided for the attachment of the gear for future
spacecraft. 0m the bottom of the cabin, between the landing gear doors, two
additional doors are installed. The forward door allows access to the lower
equipment compartment and the aft door provides access to the Environmental
Control System compartment which is a portion of the pressure vessel.
Hatches
Two large structural hatches (Figure 2-8) are incorporated for sealing the cabin
ingress or egress openings. The hatches are sy_etrieally spaced on the top side
of the cabin section. Each hatch is manually operated by means of a handle and
mechanical latching mechanism. Each is hinged on the outboard side. In an
emergency, the hatches are opened in a three sequence operation employing pyro-
technic actuators. When initiated, the actuators simultaneously unlock and open
the mechanical latches, open the hatches and supply hot gases to ignite the ejec-
tion seat rocket catapults. An external hatch linkage fitting is incorporated to
allow a recovery hatch handle to be inserted for opening the hatches from the
outside. The recovery hatch handle is stowed on the main parachute adapter
2-10
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/_ 53 63 640_>_90 65
[_[_103[_>_86 _8<_]
0_[_91 96<@8<!_C_>_95
[!_[Z>100 0_>ff[>[i!>99 5469
8479A 79
40A 40 \ 7027B 27,6,27-1 \
78A 78
27-2 6683
19
26 \ 7111
18A10
2A1
55
37IO5
NOTE?.8A
DOOR B:FECTIVIIY - S/C D
NO. DESCRIPTION NO. DESCRIPTION NO. DESCRIPEION
1 DROGUE CHUTE DOOR 28 SHINGLE 79 RECOVERY UGHT DOOR
2 DOCKING BAR CARTRIDGE ACCESS 28A z 16o.2o EQUIPMENT ACCESS 79A RECOVERY LIGHT DOOR RELEASEMECHANISM
2A PYRO ELECTRICAL DISCONNECT ACCESS 32 FORWARD EQUIPMENT BAY DOOR - LEFT 80 HOIST LOOP DOOR
3 SHINGLE 37 AFT EQUIPMENT BAY DOOR - LEFT 82 SHAPED CHARGE DETONATOR ACCESS
EMERGENCY DOCKING RELEASECARTRIDGE4 AND GUILLOTINE CARTRIDGE ACCESS 40 SHINGLE 83 COVER ASS'Y - PARACHUTE CONTROL CABLES
J RECOVERY LIGHT AND HOIST LOOP4A PILOT CHUTE DEPLOY SENSOR SWITCH ACCESS 4OA RIGGING AND CARTRIDGE ACCESS 84 COVER ASS'Y. - PARACHUTE CONTROL CABLES
5 i SHINGLE 46 SEPARATION SENSING SWITCH ACCESS 85 RADIOMETER
5A RADAR ACCESS 53 DAMS LINE GUILLOTINE ACCESS 86 CRYO SPECTROMETER/INTERFEROMETER
10 SHINGLE 54 F. LS.C. TUBING CUTTER ACCESS 89 MICROMETEOPJTE EXPERIMENT
11 INTERFACE ACCESS 55 FORWARD MANEUVERING ENGINE ACCESS 90 UFIF - VHF POLORIZATION
12 INTERFACE ACCESS 56 FUEL CELL SERVICE ACCESS 91 UHF - VHF FOLORIZATION ACCESS
13 INTERFACE ACCESS 62 QAMS OXIDIZER PURGE ACCESS 95 FITCH ION SENSOR ACCESS
13A GUILLOTINE CARTRIDGE ACCESS 63 DAMS LINE GUILLOTINE ACCESS 96 YAW ION SENSOR ACCESS
18 INTERFACE ACCESS 64 OAMS OXIDIZER PURGE ACCESS 98 YAW SENSOR SYSTEM
PYROTECHNIC SWITCH CARTRIDGE AND BRIDLE18A DISCONNECT CARTRIDGE ACCESS 65 OAMS MODULE SERVICE ACCESS 99 NUCLEAR EMULSION
19 RE-ENTRY CONTROL SYSTEM ACCESS 66 ECS SERVICE ACCESS 100 CRYO SPECTROMETER/INFEI_:EROMETER ACCESS
20 RE-ENTRY CONTROL SYSTEM ACCESS 69 ECS PUMP MODULE SERVICE ACCESS 102 RADIOMETE_ ACCESS
21 RE-ENTRY CONTROL SYSTEM ACCESS 70 ECS PUMP MODULE SERVICE ACCESS 103 SPACE POWER TOOL ACCESS
26 RE-ENTRY CONTROL SYSTEM ACCESS 71 SEPARATION SENSING SWITCH ACCESS 105 8ETA SPECTROMETER
27-I SHINGLE 75 ELECTRICAL DISCONNECT ACCESS
27-2 FRESH AIR DOOR 76 ELECTRICAL DISCONNECT ACCESS
27A z_6o.2o EQUIPMENT ACCESS 78 SHINGLE
27B Z160,20 EQUIPMENT ACCESS 78A ZI60.20 EQUIPMENT ACCESS
Figure 2-7 Access Doors Spacecraft 5,6,8 and Up (Sheet ] of 2)2-11
SEoR3oo
16A17A16 gA
31A 8A
35 31 7
74 15A
14A22
232g 29,6,
'30 30A32
"3334
3837
41
NOTE
48 _:)oeRFPET_Vm,- s/cD
101_<_ ._57 87_<_67 3'3' 49
NO. DESCRIF_rlON NO. DESCRIFTIO}_ NO. DESCRIPTION
6 EMERGENCY DOCKING RELEASECARTRIDGE _9_, Z160.20 EQUIPMENT ACCESS 51 B.I.A. RELAy PANEL ACCESSAND GUI.LLOT NE CARTRIDGE ACCESS
7 SHINGLE 30 SHINGLE 52 FORWARD MANEUVERING ENGINE ACCESS
8 SHINGLE 3OA Z160.20 EQUIPMENT ACCESS 5'7 SHAPED CHARGE DETONATOR ACCESS
8_ RADAR ACCESS 3_ SHINGLE 58 FUEL CELL SERVICE ACCESS
9 EMERGEt_Cy DOCKING RELEASECARTRIDGEAND GUILLOTINE CARTRIDGE ACCESS 31A Z160.20 EQUIPMENT ACCESS 59 GUILLOTINE CARTRIDGE ACCESS
9A DROGUE CHUTE DEPLOY SENSOR $W_TCH ACCESS _2 FORWARD EQUIPMENT BAY DOOR - LEFT 60 GUILLOTINE CARTRIDGE ACCESS
14 INTERFACE ACCESS 33 MAIN LANDING GEAR DOOR - LEFT 6 | OAMS FUEL PURGE ACCESS
1"4A GUILLOTINE ANWL ACCESS 34 CENTER EQUIPMENT BAY DOOR - FORWARD 6'7 ENGINE TO SCUPPER INTERFACE ACCESS
_ INTERFACE ACCESS 35 MAIN LANDING GEAR DOOR - RIGHT 68 ELECTRONIC MODULE TEST ACCESS
GUILLO'f'rNE CARTRIDGE AND LAUNCH15A GUILLOTINE ANVIL ACCESS 36 FORWARD EQUIPMENT BAY DOOR - RIGHT '72 VEHICLE ELEC CONN ACCESS, .
i 166A INTERFACE..... ACCESS 37 AFT EQUIPMENT BAY DOOR - L_ET '73 SRFARAT_O_ SENSING SWITCH ACCESSGUILLOTINE CARTRIDGE ACCESS 38 E.C.S. BAY DGOR 74 SHAPEO CHARGE DEFONATOR ACCESS
1'7 INTERFACE ACCESS 39 AFT EQUIPMENT BAY 0OOR - RIGHT 7'7 FUEL CELL PURGE ACCESS
17A PARAG LIDER ELECT. CONTROL gOX ACCESS 4_ PURGE FITTING ACCESS 8'7 SPECTROMETER/I NT ERFEROMETER
22 RE-ENTRY CONTROL SYSTEM, ACCESS 47 RELAY PANEL ACCESS 88 ELECTROSTATIC CHARGE SENSOR
23 RE-ENTRY CONTROL SYSTEM ACCESS 48 RELAY PANEL ACCESS 92 LLTV ACCESS
24 RE-ENTRY COI_n'ROL SYSTEM ACCESS 49 SEPARATION SENSING SWITCH ACCESS 93 ELECTROSTATIC CHARGE ACCESS
25 RE-ENTRY CONTROL SYSTEM ACCESS 50 GUILLOTINE CARTRIDGE ACCESS 94 LLTV FAIRING
29 SHINGLE 9'7 PITCH _NSOR SYSTEM
101 SPECTROMETER/IN FERFERQMETERACCESS
104 BETA SPECTROMETER ACCESS
|06 MAGNETOMETER ACCESS
Figure 2-7 Access Doors Spacecraft 5, 6, 8 and Up (Sheet 2 of 2)
2-12
._- SEDR300
HATCHLATCHSHOWN IN LATCHEDPOSITION
.............. ...................................................................................... <................
HATCH INTERIOR VIEW
.__. HATCHHANDLESHOW.,_STOWE_PO_mO_ !i_ii!! HA.CHHANDLE_.OWNINUNSTOWED.OSmON
Figure 2-8 Spacecraft Ingress/Egress Hatches
2-13
SEDR 300
assembly located on the forward face of the RCS section. A hatch curtain (Figure
2-9) is stowed along the hinge of each hatch. After water landing, when the
hatches are open, the curtains are installed to help prevent water from entering
the cabin.
Windows
Each of the ingress/egress hatches incorporates a visual observation window (Figure
2-10). Each window consists of an inner and outer glass assembly. The outer
assembly is a single flat paineand the inner panel assembly consists of two flat
panes. The panes consist of Vycor (96% silica). The panes in the right window
are optically ground for better resolution. Each surface of each pane, with the
exception of the outer surface of the outer pane_ is coated to lessen reflection
and glare from cabin lights and to aid in impeding ultraviolet radiation into the _
cabin eompaz-_ent.
Heat Shield
The heat shield is a dish-shaped structure composed of silicone elastomer filled,
phenolic impregnated, fiberglass honeycomb. It is an ablative device, 90 inches
in diameter with a spherical radius of 144 inches. The shield is designed to
protect the re-entry module from extreme thermal conditions during re-entry into
the atmosphere. The device is attached to the large diameter end of the cabin
structure by 1/4 inch bolts.
Shin_les
The external surface of the cabin is made up of beaded shingles of Rene' 41. The
R and R and RCS section surfaces are made up of umbeaded shingles of berylliu_.
The shingles protect the re-entry module structure from excessive heat and provide
additional rigidity for the cabin. The shingles are black on the outer surface
I _ PROJECT GEMINI
INBOARD HOOK ATTACHMENT
OUTBOARDHOOKATTACHMENT
HATCH CURTAIN SHOWN IN EXTENDED POSITION _'C_._ \\
(TYPICAL IN LEFTAND RIGHT SIDE) '_._. \\(ROTATED 180 ° } _._ _. \
_. STRAPSNAP(TYPICAL 5 PLACES
EACHSIDE)
' HATCH CURTAIN SHOWN IN STOWED POSITION
Figure 2-9 Hatch Curtain2-15
sEo,300P-R-OJ ECT'-G E M I N I
OUTERGLASSPANE
GASKET ASKET
O-RING
SHINGLEATTACHINGERACKET IVETS
_ DETAIL A-A
% _ HATCHOUTER
• / ML (REF)
AM_
OUTERWINDOW ASSY
z,3,._o // ////x_/_-.// ,d._ •ii"/ //// OB.,,VATIONWINDOWASS,MBLy
NUT-MIDDLEGLASSPANI
STAT,.O-SEALWASHER
ERGLASSPANE
KET
FRAME
INNER WINDOW ASSEMBLYSTAT-O-SEALWASHER
BOLT
Figure2-10 ObservationWindow
2-16
SEDR 300
EMINI
to control thermal radiation. The inner surface of the beryllium shingles are
coated with gold to provide a low emissivity surface.
ADAPTER
The adapter functions to mate the spacecraft to the launch vehicle, to provide
for mounting equipment and retrograde rockets , and to serve as a radiator for
the spacecraft coolant system. The adapter (Figure 2-2) is a truncated cone-shape,
simimonocoque structure consisting of circmuferential aluminum rings, extruded
magnesium alloy stringers, and magnesium skin. The extruded stringers are designed
in a bulb-tee shape to provide a flow path for the liquid coolant which transfers
heat to the adapter skin for radiation to space. The outer surface of the skin
is coated with white ceramic type paint and the inner surface is covered with
aluminum foil. The _nner adapter surfaces of spacecraft 9 through 12 are gold
plated. The forward end of the adapter is coupled to the aft end of the re-entry
module by utilizing three titanium tension straps (Figure 2-11).
RETROGRADE SECTION
The retrograde section, the smaller end of the adapter, provides for installation
of four retrograde rockets and six Orbital Attitude Maneuvering System thrust
chamber assemblies. To provide for the installation of the retrograde rockets,
the retrograde section employs an al1_nun I beam support assembly. The I beams
are assembled in the form of a cruciform with one retrograde rocket mounted in
each quadrant.
EQUIPMENT SECTION
The equipment section is the larger diameter end of the adapter. The section
provides hard points for the attachment of structural modules for the OAMS tanks,
Enviromnental Control System primary oxygen supply, fuel cell (batteries on
2-17
SEDR 300
// _'_AINJNGSTRA_
/;/ _'N
,_ BOLT
S'HER,CALWASHER I// WASHOR
N% AD_ _i _1/ N N _FARNG
(TI TAN' UM) ...... ___._%
WASHER /- "'_"RE-ENTRY MODULE(_F)
(REF)
. STRAP ASSEMBLY
_ (VIEW ROTATED FOR CLARITY)
ADAPTER(P_F)
SHAPED CHARGE '_;lASSEMBLY (REF) _"'_,
• t-..(REF)
RE-ENTRy MODULE j
STRAP ASSEMBLY (REF)
(TYP 2 PLACES)
RETAINING STRAP ASSY
SHAPED CHARGE J_ (TITANIUM)
ASSEMBLY (REF) _HEAT SHIELD (RER)_ ! I
ISECTION A-A
Figure2-IIRe-EntryModule-AdapterRetainingStraps
2-18
SEDR 300
PROJECT GEMINI
spacecraft 6), coolant, electrical and electronic components, Extra Vehicular
Activity (EVA) equipment on spacecraft 9 through 12, and Rendezvous Evaluation
Pod on spacecraft 5 only. A honeycomb blast shield is provided above the modules
to shield the equipment section and booster dome from excessive heat during retro-
grade rocket firing under abort conditions. Ten OAMS thrust chamber assemblies
are mounted on the large diameter end of the equipment section. A gold deposited
fiberglass temperature control cover protects the equipment fred solar radiation
through the open end of the adapter after separation from the launch vehicle.
SPACECRAFT/LAUNCH VEHICLE MATING
The spacecraft is mated to the Titan II Launch Vehicle with a machined aluminm_
alloy ring (Figure 2-12). This ring, 120 inches in diameter, mates with the launch
f_ vehicle mating ring. Twenty bolts secure the rings together. To provide for
alignment, the launch vehicle incorporates one steel 3/26 inch disBeter align-
ment pin located at TY and four index marks. To separate the spacecraft from the
launch vehicle, a pyrotechnic charge is fired, severing the adapter section
approximately 1½ inches above the launch vehicle/spacecraft mating point.
-i9
___ SEDR 300 =__
PROJECT GEMINI
(REF) /I
IOXIDIZER TANK(REF)
QUAD3
QUAD4 QUAD2
VEHICLE(REF)
SPACECRAFT TOLAUNCH VEHICLEATTACHMENT BOLTHOLES
MATING SECTIONSEPARATION ASSEMBLY
(REF) QUAD4 QUAD3
LX
MACHINED MATINGRING
QUAD1 QUAD2
SPACECRAFT
MATING LINE
--_ Zl3,44LAUNCHVEHICLE
RING ATTACH --BOLT
BY
SECTION A-A(TYPICAL 20 PLACES)
Figure 2-12 Spacecraft/Launch Vehicle Mating Ring
2-20
CABIN INTERIORARRANGEMENT
SectionT__._o__o_T____ III
TITLE PAGE
GENERAL .................................................... 3-3_-_ CREW SEATING .......................................... 3-3
SEAT DESCRIPTION .................................... 3-3SEAT EJECTION SYSTEM ............................ 3-5RESTRAINT SYSTEM .................................... 3-10EGRESS KIT ................................................. 3-14
BACKBOARD ASSEMBLY ........................... 3-15 !.._..=.:..=._PELV IC BL0 CK ........................................... 3-15 ii!iiii_i_;:.:.iiii_-_,............._.__..,
,..oo...oo...o.._.._..o...
BALLUTE SYSTEM ....................................... 3-16 !iiii!ii!i!iiiiiii':_!ii:'_iiPERSONELLPARACHUTE .............................. 3-16 iiiiiiiiiiiiiiiiiiiiiiiiiiiPARACHUTEDROGUEMORTAR................3-16 !!i!!iiiiiiiiiiiiiiiii!iiiiPERSONAL HARNESS ASSEMBLY ................ 3-17 iiiiiiiiii!iiii!ii!iiiiiiil_.°°°°°°°...o°o..°_°o.o°,..
SURVIVAL KIT ............................................. 3-17 i!i!iiii!i!i!!ii!i!!!iii!!!.°°°°o°_oo°°°°oo,°°°°..**°,
PYROTECHNIC DEVICES ............................. 3-19 !iiiiiiiiiiiiiiiilH!H!!!!INSTRUMENT PANELS ................................ 3-20 iliiiiiiiiiiiiilHiiiii!iii
.............. °°°°o.°°°o°°,
CABININTERIORUGHTING.......................3-20 !iiiiiiiiiiiiiiHiiiiii!iiiSTATIC SYSTEM .......................................... 3-24 _iiiHiiiii!!ii!i!i!i!iiiiiFOOD, WATER and EQUIPMENT iiiilH!i!!iiiiiiiiiliiiill
STOWAGE ................................................ 3-24 i!iiiiiiiiiiiiiiiiiiiiiiiii
WASTE DISPOSAL ...................................... 3-29 iiiiiiiiiiiii!ii!Hiiii!iiiSTOWAGE PROVISIONS ............................ 3-29 i_.:-:-::._i_':._!!":._iii:._!!_ii_iiiiiiii_iiiiiiiiiii!ii!
_-]. :::::::::::::::::::::::::::
___ SEDR300 _____
PROJECT GEMINI
EGRESS
ERVATION WINDOW
Ilia
CIRCU
CIRCUIT BREAKER PANEL
INSTRUMENT PANEL
CONTROL VESSEL (REF)
MANAGEMENT
EJECTION SEAT
WASTE STOWAGE
SECONDARY C CONTROL HANDLECONTROL HANDLE
Figure 3-1 Cabin Equipment (Typical)
3-2
__ $EDR300
PRO J EC---"C'T-"GEM IN I
SECTION III CABIN INTERIOR ARRANG_ERT
GENERAL
The equipment within the cabin is arranged to permit the comand pilot, seated
to the left, and the pilot, seated to the right, to operate the controls and
observe displays and instruments in full pressure suits in the restrained or
unrestrained position. The cabin air outflow is regulated during launch to
establish and maintain a 5.5 psi differential pressure between the cabin and
outside ambient condition. The cabin is maintained at a nominal 5.1 psia
throughout the flight by a cabin pressure regulator. The cabin equipment
(Figure 3-1) basically consists of crew ejection seats, instrument panels and
controls, lighting, food, water, waste collection, and miscellaneous equipment.
SmT G
The crew members are seated in the typical command pilot and pilot fashion, faced
toward the small end of the re-entry module. The seats are canted 12° out-
board and 8° forward to assure separation and to provide required elevation
in the event an off the pad ejection is necessitated.
Crew seating provisions include scats, restraint mechanisms, seat ejection
devices, seat man separator, survival gear, and an egress kit assembly effective
spacecraft 5 and 6 only.
SEATI) n:TON,
The crew seats (Figure 3-2) are all metal built-up assemblies consist_ns of a
torque box framed seat bucket, channeled backs and arm rests. The seat has
lateral and vertical stiffeners, designed for a single moment of thrust. The
3-3
SEO.30o..oJ.cTGEM,.,RAFT CONTAINER
PARACHUTE RISERS ISERANDAND SHOULDER BALLUTE RISER STORAGE
RESTRAINT SERAPS_
PERSONNEL PAR.ACHUTE_
BACKBOARD _ R BOARD
DROGUEMORTAR _
_ BLOCK
_1 ITIA REEL _,,
CONTROL
ELBOWRESTRAINT
SUIT ELECTRIC/d. ; STRAp
SUIT OXYGEN HOSES
NOTE
_!_ COMMAND PILOT EJECTION SEAT ILLUSTRATED.HARNESS RELEASEACTUATOR IS lOCATED ONOUTBOARD SIDE OF SEAT. "_-'_
S/C 5 AND 6 _
Ieigu_e 3-2 Gemini Ejection Seat Assembly
SEDR300
PROJ MINI
seat is supported at a single point at the top of the seat back. At this
point, the seat bolts to the rocket/catapult. Each seat is supported against
fore, aft, and side movement by slide blocks mounted on the seats and retained
in tee type rail assemblies attached to the large pressure bulkhead. The
seats incorporate a padded contoured headrest to support the pilots helmet.
Each seat slso incorporates a restraint system, harness release system and a
seat/man separator.
SEAT EJECTION SYSteM
The seat ejection system (Figure 3-3) provides the crew with a means of escaping
from the vicinity of the spacecraft in the event of an abort or in an emergency
condition during launch or re-entry. Crew member seats are ejected by means of
,_ rocket/catapults. Hot gas from each of the hatch actuators is routed to the
appropriate seat catapult where dual firing pins strike dual percussion primers,
thereby igniting the seat rocket/catapult main charge and ejecting the seats from
the spacecraft. Hot gas from the rocket/catapult main charge ignites the sustainer
rocket and the rocket provides additional separation from the spacecraft. In the
event ejection becomes necessary, after deployment of main landing system parachute
and while descending in the two point suspension, it is mandatory that the main
landing system parachute be Jettisoned before ejecting from the spacecraft.
The ejection sequence is initiated by manually pulling either ejection control
(D-rlug) located on the front of the seat buckets. During the launch phase of
flight each pilot erects and holds on the D-rlng. This action aids in stabilizing
the pilots arms and at the same time places them in a position for instant response.
The D-rlngs are normally stowed at the front of the seat and are pinned in a
3-5
___ SEDR 300 _3
PROJECT GEMINI
300 FEET
=PEET r _,,
,,i= =- 'tI NOTETHIS PLOT ILLUSTRATESTHETRAJECTORY OF A PILOTWHEN EJECTED OFF THE PAD.
100 FEET
0 FEET 100 FEET 200 FEET 300 FEET 400 FEET 500 FEET 600 FEET 700 FEET
EJECTION SEAT TRAJECTORY PLOT
EITHER PILOT CAN EJECT SEATS.
NOTE
_i_ _DF_i_ LINES
1 Jrili"ill _j MDFFIREsMANUALFIRING MECHANISM" ... --
ii!i!ii IL_MDF BURNS AT AFPROXJ/_TELY 24'000 FT/SEC
_IBOTH PILOTS HOLD EJECTION CONTROL LOOpTHE FJRMLY.
NOTEiii!ii MDFCROSSOVERNE_ORKI.ITATESSECONDS_TNT'ATOB!_iili _PULSE TRAVELS TO HATCH ACTUATORS IN 4 SEPAP-A,TE L[NES.
IEVENT TIME
.01
_:_:_::_:_:_:BI_:BREECH ASSEMBLY IGNITES.
iiiiil HATCH IS UNLATCHED,
___,N mATE EJECTION iiiiil HATCH IS O PENED.
Figure g-a Ejection Seat Sequence Of Operation (Sheet 1 of 4)
3-6
SEDR 300
:x:.:
_:_i_!i!_![]ROCKET EURN ouT.
i::i::iii::i EVENT TIME I
CATAPULT ::i::i!::::EGRESS iiiiii::
aNYARDS ii_ii:_!!i
i! BALLUTE ACTUATOR
BALLISTIC
_.. ,ATCHACTUATORGASIMPULSEO,RECTED_OROCKEE/CATA,ULTRY_LL,ST,Ciii:_iminiHOSE
PILOT] DROGUE
CATAPULT TIRES MORTARREEL
I11_ HARNESS RELEASEACTUATORS INITIATED. EGRESS LANYARDS PULLED,COM-MuNICATIONIS SEVERED.I i
EVENT TIME l HAR NES:•r_ [] RELEASE
.32 DISCONNECT(FROM SEAT)
]EJECTION SEAT MOVING UP, IGNITES EJECTION ROCKET APPROX]MATELY
4 INCHES FROM END OF RAIL TRAVEL. mwm_ SEAT/_N RECOVERY BEACONi SEPERATOREVENT TIME
3R iiiiiiim [] RAORY................................................................................... .:_i:i:I EJECTION SEAT CONTINUE3 ON T J T
:!::::::1:1::::i :_::i::::_:_::_::::!::!:!:_:::::: :__:k:!:!:!:!:::!:!!::i:i:::: :1:1:::_:__:::_:_:_:::!:!::::!:!:!:!::!:!:!:::::::::::::::::::::::::::::::::::::_:::::::::::::::::::::::::::::::_:_:::::::V:::::::::::::::::::::::::::::::::::_:::::::::::::::_:ii_i::
i_iii II _"]J HARNESS RELEASEACTUATOR TIRES BY LANYARD PULLI _-_::_
_ LAP BELT RELEASEASSEMBLY ACTIVATED. BACKBOARD AND SURVIVAL GEAR
_ _ ASSEMBLY RELEASED FROM SEAT.
::.:_.:- EVENT TIME
:.x.::.:.:+x
:_::::::::
::.:: :.::.:::::.::::.:
i_!ii_111HARNESSRELEASEACTOATORGAS,MPULSEOEL,VEREDTOSRA_/MANSE-I iii::iiii PARATOR BY BALLISTIC HOSE.
_ _ i_:_ii'iiiii:_ SEAT/MAN SEPARATOR SHOE EXTENDS AND REMOVES SLACK FROM STRAP
::::ii:'_i_-- ASSEMBLY.
/'x :iiiiiii_I_ PILOT WITH BACKBOARD AND SURVIVAL GEAR SEPARATE FROM SEAT
;::::::;_i_i__ PILOT DROGUE MORTAR, BALLUTE SYSTEM AND RECOVERY BEACON INITIATEDBY LANYARDS CONNECTED TO SEAT STRUCTURE.
EVENT1.50TIMEIEGRESS CONTINUES UNDER ROCKET POWER. PILOTS LEAVE SPACECRAFT
BAT 12° OUTBOARD OF "X" AXIS AND 8 ° 20' FORWARD OF "Z" AXIS.
Figure 3-3 Ejection Seat Sequence Of Operation (Sheet 2 of 4)
3-7
SEDR 300
N
NOTE
DROGUE MORTAR BAROSTAT IS ACTIVATED DURING SF_.T/MAN(4 PLACES) SEPARATION TO DEPLOY THE PARACHUTE AT 5700 FEET OR BELOW.
BALLUTE [] DROGUE MORTAR FIRES.
RIS_ _ 3.80iiiiiii!i
NOTE
iiiil _ENTT"_ESARE_ORE,EC.O"BELOW5=PEETO.LY
[]MORTAR SLUG DEPLOYS AND INFLATES PILOT AND MAIN PARACHUTES.
iiili!
ABOVE SEQUENCE ILLUSTRATION IS TYPICAL OF EJECTION
BETWEEN 7S00 AND 40,000 FEEl ONt_. ,_--_
I SALLUTE DEPLOYS AFTER A S SECOND DELAY.
NOTEI . BALLUTE BAROSTAT HAS BEEN ACl"EqATED TO JEI'TISON Tiff
BALLUTE AT 7500 FEET,2. TIME CHART APPLICABLE TO EJECTIO_ ABOVE 7500 FEET
ONLY.
] PARACHUTE FULLY INFLATED.
_ BACKBOARD AND SEAT SEPARATED FROM PILOT.
iiiFigure 3-3 Ejection Seat Sequence Of Operation (Sheet 3 of 4)
3-8
_@ SEDR300 _'_
PROJECT GEMINI
NOTESURVIVAL EQUIPMENT DEPLOYED AS BACKBOARD FALLS FROMPILOT. THE LiFE RAFT IS INFLATED MANUALLY DURING PARA-CHUTE DESCENT ALL SURVIVAL EQUIPMENT IS SECURED TOPILOT BY A LANYARD.
NOTE
_ PILOT DISCONNECTS OXYGEN INLET AND OUTLET HOSES• OXYGEN_i_ CONNECTION IN PRESSURESUIT IS SEALED CLOSED WHEN OXYGEN
_ii HOSE IS REMOVED.
_" _i!i
ii
Figure 3-3 Ejection Seat Sequence Of Operation (Sheet 4 of 4)
3-9
__. SEOR300
downward position at the front of the seat structure. The safety pin is re-
moved during launch and re-entry and during orbit.
Each pilot is restrained in his ejection seat by a restraint system (Figure 3-4)
consisting of personal harness, lap belt assembly, shoulder restraint, inertia
reel and leg restraint. Other portions of the restraint system are part of the
ejection seat (Figure 3-2). These seat restraints are the arm restraint loops,
elbow restraint and foot stirrups. The restraint system provides adequate support
and restraint during conditions of maximum acceleration and deceleration.
INERTIA _L
The inertia reel (Figure 3-_) is a two position locking device, located on the
rear of the backboard. Two straps connect the inertia reel and the personal
harness to restrain the pilots forward movement. The inertia reel control handle
is located on the front of the left arm rest and has two positions, manual lock
and automatic lock. Orbital flight is accomplished with the inertia reel in the
automatic lock position. Manual lock position is used during launch and re-entry.
The manual lock position prevents the pilots shoulders from moving forward.
To release his shoulders, when the inertia reel is in the manual lock position,
the pilot must position the control handle to the automatic position. The auto-
matic lock a11ows the astronaut to move forward slowly a maximum of 18 inches but
will lock with a 3 g deceleration. When the automatic lock has engaged, the lock
will ratchet and permit movement back into the seat, but will not permit forward
movement. The release of the automatic lock is accomplished by cycling the
control handle to manual and back to automatic lock.
3 -I0
SEDR 300
! s/!
A//
%.\j/_ _
- _
i
\..
. i \JJ PERSONALHARNESS
it...s
J
\ ..s'Dt )".....__........ I KITaN'<,_O
_.:._f............II
.I , LEG RESTRAINT
ISHOULDER I_]l LEGRESTRAINTAND SURVIVALKITLANYARDRESTRAINT
Figure 3-4 Restraint System (Sheet 1 of 2)
3-11
sEoR300PROJECT GEMINI
NOTE
LEFT HAND SEAT IS SHOWN(PILOT) RIGHT HAND SEAT HASINERTIA REEL CONTROL ONRIGHT HAND ARM REST.
BACKBOARD jASSEMBLY
_._%
ASSEMBLy
PERSONNELAREA (REF)
iNERTIA REEl.ASSEMBLY
iNERTiA REELCONTROL SHAFT
Figure 3-4 Restraint System (Sheet 2 of 2)
3-12
SEDR300 ___PROJECT GEMINI
ARMRESTRAINT
The arm restraint (Figure 3-4) is a welded, 1/2 inch diameter tube assembly
made up in the form of a loop. A loop is installed on each arm rest to retain
the pilots arms within the ejection envelope. When the arm restraint loop is
not required, it ,my be swung to the rear and down.
An elbow restraint is provided for the command pilot only. It is used to
stabilize his forearm during manual re-entry.
LRG RESTRAINT 8_RAP
The leg restraint (Figure 3-4) consists of two straps of dacron webbing with a
connecting slide buckle. One end of each strap is secured to the seat by round
metal eyelets. The left strap of each leg restraint has a metal end assembly
that permits the right strap to fold back on itself. Velcro tape on the right
strap is used to secure the strap end in position when the strap is drswn tight
over the pilots legs. During seat/,_n separation, the restraint strap eyelets
are automatically released from the base of the seat, freeing the restraint strap.
EJECTION SEAT FOOT STIRRUP
The ejection seat foot stirrups (Figure 3-2) consist of two welded frames attached
to the front of the ejection seat. Each stirrup has a short protruding platform
with small vertical edges rising along the outboard side. The stirrup is so
constructed that the pilots shoe heel will lock in place and prevent forward
movement of the foot while the small vertical edges will prevent side movement.
During seat ejection, the pilots feet will stay in place,
3-13
[_ SEDR300
LAP _LT
The lap belt (Figure 3-k) is an arrsngem_nt of dacron and nylon straps, designed
to restrain the pilot in the seat structure. Load carrying straps from the lap
belt are fastened to the backboard and seat. The lap belt has a manual quick
disconnect and a pyrotechnic release fitting near the center of the pilots lap.
The manual quick disconnect can be released with one finger. Lap belt tension is
adjusted by sliding excess strap through the pyrotechnic release. During ejection,
the lap belt ends attached to the seat structure are released just prior to seat/
ma, separation. During separation, the lap belt _Ins with the pilot. Five
seconds after the backboard drogue mortar fires, the pyrotechnic lap belt release
activates and allows the lap belt, backboard and seat to fall free.
A second -_-,,_i release for the lap belt is also available to the pilot. It is
located forward on the right arm rest and is referred to as the ditch control.
Releasing the lap belt with the ditch control allows the pilot to egress from
the _Anding module with the backboard and seat.
EGI_gSS KIT (Effective Spacecraft 5 and 6)
The egress kit assembly contains the bail out oxygen for an ejected pilot. The
egress kit rests in the ejection seat bucket and forms a mounting surface for the
egress kit cushion. The egress kit contains an oxygen supply, for breathing and
suit pressurization; a composite disconnect, which when separated closes the port
and prevents escepe of egress oxygen; a relief valve, to prevent pressure build
up in the pressure suit; a regulator, to reduce high pressure to a controlled flow
of low pressure oxygen, a pressure gage, for visually checking egress oxygen pres-
sure; and connecting lines. Three lanyards are attached between the egress kit and
SEDR300
PMINI
the spacecraft. These lanyards pull release plns to allow the composite d/s-
connect to separate, allow the oxygen to flow through the pressure regulator and
allow the relief valve to control the pilots suit pressure. When the drogue
mortar deploys the pilot parachute, a 5-second pyrotechnic time delay is initiated
and at burn out_ the egress kit with the backboard is separated from the pilot.
EGRESS KIT CUSHION (Effective Spacecraft 5 and 6)
The egress kit cushion (Figure 3-2) has a universal type of contour and is
attached to the top of the egress kit. The cushion is positioned forward of the
pelvic block and up to the ejection control handle access door.
The backboard assembly (Figure 3-2) is machined aluminum, designed and stressed to
retain the inertia reel, ballute, ballute release and deploy mechanism, drogue
mortar, parachute and survival kit. A cushion, contoured to the ind/vidual pilots
body requirements, is positioned on the forward surface of the backboard. The
cushion is provided to supply support and comfort to the pilots back. The inertia
reel straps and lap belt secures the pilot to the backboard. The backboard
accompanies the pilot through seat ejection to parachute deployment. Five seconds
after parachute deployment, the backboard with the seat is separated from the
pilot.
PELVIC BL0C[
The pelvic block (Figure 3-2), contoured to the lower torso of each pilot, is
positioned between the backboard assembly and the seat. The block supports
the pilots lower vertebra and pelvic structure. It remains with the seat
structure upon seat/man separation.
3-15
_@ SEDR300
PROJECT'---GEMINI
B_r.T._ESYSTD_
The ballute system (Figure 3-2) consists of a barostat controlled pyrotech-lc
initiator, combined wi_h a pyrotechnic gas generator, cutters and a packaged
ballute. The ballute, located on the back and lower left side of the pilots
backboard, is an aluminized nylon fabric enclosed cone. It is inflated by ram
air passing through four inlets located s_-._trically around the upper periphery.
The ballute is connected to the backboard through an 8 inch rlser, a 5 foot dual
bridle, and by a one inch wide dacron webbing passing through a pyrotec_,_c
actuated cutter. The ballute provldes the pilot with a stabilized, feet into the
wind, attitude for all ejections over 7,500 feet. The system is fully autcmatic
and is actuated at seat/man separation. At altitudes below 7,500 feet, the
barostat prevents deployment of the ballute ....
PERSONN_. PARASITE
The personnel parachute (Figure 3-2) is a standard 28 ft dia nylon parachute.
The parachute is located on the right rear of the pilots backboard. It is
deployed by the drogue mortar slug and pilot chute. The parachute risers are
attached to the pilots personal harness.
PARAC_UfE DROGUE MORTAR
The parachute drogue mortar (Figure 3-2) is a pyrotechnic device designed to eject
a i0 oz drogue slug with sufficient velocity to deploy the pilot chute of the
personnel parachute. The drogue mortar is a barostat operated firing mechanism,
but can be fired manually. It will fire and deploy the parachute at or below
5,700 feet plus a 2.3 seconds time delay from seat/man separation. An MDF chain
is initiated by the drogue mortar and separates the backboard and seat from the
pilot •
3-16
SEDR300
PRMINI
PERS0_L HARNESS ASSDmLY
The personal harness ass_ubly (Figure 3-_) provides a light, strong, and comfort-
able arr-ngement to attach the personnel parachute to the pilot. The harness is
constructed from nylon webbing formed into a double fignre-8. The two figure-
8's are joined by two cross straps, the waist strap, and the chest strap. Only
the chest strap is adjustable. A quick disconnect is placed forward and below
each shoulder for connection of the parachute risers and shoulder restraint
straps. Below the left quick disconnect, a small ring is incorporated to attach
the survival kit lanyard.
mVITAnm
The survival kit (Figure 3-2) is a packaged group of specially designed equipment
for the use of a downed pilot. Articles in this kit are intended to aid in
preserving life under varying environmental conditions. Deployment of the
survival kit is automatic if the pilot ejects and is also available to the pilot
if he lands with the spacecraft.
Deployment of the survival kit during the ejection cycle takes place as the
backboard and seat falls away from the parachuting pilot. As the backboard
falls, the survival kit lanyard, connected to the pilots harness, pulls a
pin on the life raft container. When the pin is removed, the daisy chain loops
are disengaged and the llfe raft and rucksack are extracted from the container.
The survival kit lanyard repeats the extraction process in removing the machete
and water bottle from the second container. The machete and water bottle are
stowed in a survival equipment container on the left front side of the backboard.
3-17
•PROJECT GEMINI
During seat/man separation, a lanyard between the seat structure and the rucksack
activates the radio beacon. As the pilot descends on his parachute, the survival
equilzmentis suspended below and the radio beacon transmits on an emergency
frequency. Direction finding equipment on aircraft and aboard ship can plot the
pilots position taking navigational fixes on the radio/beacon.
Survival equipment is divided into two major stowage containers. The life raft
container mounted on the left rear of the backboard has the following items:
Life raft container (Typical)
I Life Raft
I Sea anchor
1 4 inch x _ inch Foam rubber pad
1 C02 cylinder
1 Sea dye marker
I Sun bonnet
Rucksack (Typical)
1 Survival light
1 Strobe light
1 Flash light
4 Fish hooks
Fish line
2 Sewing needles and thread
1 Magnetic compass
1 Fire starter
4 Fire fuel
3 -18
___ SEDR 300
PROJECT GEMINI
I Whistle
i Signal mirror
14 Water purification tablets
I De-salter kit (less can)
8 De-salter tablets
1 Water bag
1 Repair kit
1 Medication kit (Typical)
6 Tablet packets
1 Small injector (1 CC)
i Large injector (2 CC)
_- i 3 inch x 3 inch compress
1 12 inch x 12 inch alu_m_uumfoil/
i Tube zinc oxide
1 pr Sun @lasses
1 Radio beacon
The forward survival kit, mounted on the forward surface of the backboard to the
left of the pilots shoulder, contains the following;
1 Water container with 3 lb of water
1 Machete with sheath
PYROTECHNIC DEVICES
There are 18 pyrotechnic devices incorporated in the cabin all of which pertain
to seat ejection, restraint release and parachute deployment. The pyrotechnic
devices are 2 hatch actuators, 2 seat rocket/catapults, 2 ballute deployment and
release mechanisms, 2 backboard and seat Jettison, 2 drogue mortars, 2 harness
3-]9
_@ SEDR300 __
PROJECT GEMINI
release actuators, 2 seat/man separator actuators, 2 hatch actuator initiators
and 2 hatch MDF (Mild Detonating Fuse) b6vnesses. The pyrotechnic devices, except
the drogue mortar, are safetied by stowing the ejection control handle (D-ring)
with a safety pin through the handle into the ejection control assembly. On
spacecraft 8 only, a second ejection control pyrotechnic safety pin is also
inserted in the side of the ejection control assembly to completely safety the
MDF manual firing mechanism.
INS_ PANELS
Instrument panels, switch and circuit breaker panels and pedestal panels (Figure
3-5) are arranged to place controls and indicators within reach and convenient
view of each crew member while in a full pressure suit. A swizzle stick, stowed
by the overhead switch and circuit breaker panel, enables a pilot to position r _
switches and rotate selectors on the opposite side of the cabin. With this
arrangement, one pilot can control the co,,_lete spacecraft and temporarily free
the second pilot of all duties.
CABIN IN_RIOR LIG_
Cabin interior lighting is provided by three types of lights located in five
separate locations, described as follows: Cabin flood lights are located aft and
above the center-line stowage area. A DIM control is located under the light to
control light intensity. Instrument flood lights are located at the forward
inner edge of the hatches. Each instrument flood light installation contains two
lamps, one lamp having a rod filter and the other a white filter projecting down-
ward. A DIM control and a RED-WHITE-OFF switch are provided at each of the lights.
Two utility lights attached to the ends of spiral extension cords are located on
the left and right side walls of the spacecraft interior. The lights stow in
3-20
_ . SEDR 300
RAD EVAF EVAP _l
1 _vpJss _e N_
-- (S/C 10 & 11)DETAILL
(S/C 10 & UP)©
f_
DETAIL K(S/C 9 & UP)
c_f 02 ¢Rvo H2
DETAIL M
(S/C 10 & UP)
Figure 3-5 Instrument Panels and Displays (Sheet 3 of 3)
3-23
__ SEDR300 _--'-_
PROJECT GEMINI
clips mounted on the side walls. An ON-OFF switch is located adjacent to the
AUX RECEP panel on each of the spacecraft side walls. The CTR LIGHTS, BRIGHT-
0FF-DIM switch and the CABIN LIGHTS switch-circuit breaker are located on the
overhead switch and circuit breaker panel.
ELECTRICAL OUTLETS
The two receptacles, powered by the spacecraft electrical system, are installed on
brackets _mmediately aft of the left and right switch/circult breaker panels.
These receptacles are controlled by adjacent ON-OFF switches and are used for
powering the utility light or other electrical equipment.
STATIC SYSTEM
The static pressure system is employed to operate the rate of descent indicator,
altimeter, and to supply pressure to the static pressure transducer for instrumen-
tation. The static system is also utilized to provide a differential pressure
for the cabin pressure transducer. The static ports (Figure 3-6), used for
atmospheric pressure pick-up, are located in the small end of the spacecraft
conical section. The static port (Figure 3-6), used for differential pressure
pick-up, is located on the forward surface of the small pressure bulkhead.
FOOD WA_ER AND EQUIPMENT STOWAGE,,, , ,
Containers to left, right and aft of pilots (Figure 3-7) are provided for equipment
and food storage. Although minor changes in storage containers are dictated by
mission requirements, the main containers are as follows: Center-llne stowage
box, used for larger size camera containers and EVA (Extra-Vehicular Activity)
chest pack; right aft pressurized stowage box, used to stow food initially and
later, body waste materials; left aft stowage box, used to stow food packages;
3-24
._,_ SEDR300 _j__ __
PROJECT GEMINI
rf (REB_SWITCH
$THATIAMCBEPRLEN(REFU_AO_
DETAIL A-A
\
2'gMALL PRESSU P_ "_BULKHEAD
\\
DETAIL B-B
Figure 3-6 Static System
3-25
_ SEDR 300
RIGHT BIO-MED RECORDER
_TOWAG E AREA
- NUCLEAR EMULSION EXPERIMENT
SIDEWALL STOWAGE CONTAINER EXTENSION
SIDEWALL STOWAGE CONTAINER
-URINE SUPPLIES FOUCH
PILOT EJECTION SEA1REMOVED FOR CLARITY
....:::::::::::::::::VOICE TAPE RECORDER
LEFT SIDE DRY STOWAG E BAG S ::_::::_:;:::::::ili__::::::::'::::::::::::
OPTI
VIEW LOOKING INTO COMMAND PILOTSSIDE RIGHT PEDESTAL
TV MONITOR STOWAGE AREA, ._
Figure 3-7 Spacecraft Interior Stowage Areas (Typical) (Sheet 1 of 2)
3-26
I_. SEDR 300
STOWAGE AREA
-LEFT AFT STOWAGE CONTAINER
LEFT BIO_/,ED RECORDERSTOWAGE ARE
RIGHT HAND POUCH STOWAGE PROVISION--
RIGHT SIDEWALL CONTAINER_
IN-FLIGHT MEDICAL KIT-
:: _i_ii!_:'_.... COMMAND PILOT EJECTIONSEAT REMOVED
::::::::::::::::::::::::::::::::::::::::::::::::
I SIDE DRY STOWAGE BAGS VIEW LOOKING INTO PILOTS SIDE
POUCH
CENTER CONSOLE STOWAGE PROVISION
Figure 3-7 Spacecraft Interior Stowage Areas (Typical) (Sheet 2 of 2)
3-27
__ SEDR 300 __
PROJECT GEMINI
right and left sidewall stowage boxes, used to stow small pieces of equipment;
left and right fabric covered sidewall stowage boxes, used to stow lightweight head
sets; hatch food pouches used to stow large quantities of food; and sidewall stow-
age box extensions used to stow penlight, spotmeter, exposure dial and tape recorder
cartridges. Equipment stowed in the above boxes may change with each mission.
Larger pieces of equipment, emergency equipment or equipment used on every flight,
have special stowage brackets or fabric pouches positioned throughout the interior
of the spacecraft. Examples of specific stowage brackets are as follows :
inflight medical kit, stowed aft of abort control handle; and the optical sight,
stowed under co_and pilots instrument panel. Without counting the food packages,
stowage facilities are furnished for more than 125 pieces of equipment.
During flight, various pieces of frequently used equipment are removed from
launch stowage areas and are stowed, with Velcro tape, on the spacecraft sidewalls,
and on the inside surfaces of the hatch. As debris accumulates during flight, it
is placed in the left aft debris area, located aft of the pilots seat. Prior to
descent, the equipment is re-stowed. Only a general rule can be applied to
stowage descriptions. Exposed film is placed in insulated containers, previously
occupied by cameras and lens, in the center line stowage box. The left aft
stowage box is filled and the remainder of the loose equipment is divided among
the sidewall stowage boxes on a planned basis. The pressurized stowage box is
used to store urine samples and waste containers.
A water storage container, with a 16-pound capacity, is located forward of the
aft pressure bulkhead, between the seats. As the water is used from the main
storage container, it is replenished by the water stowed in the adapter section.
___ SEDR300 -___
PROJECT GEMINI
Drinking is accomplished by means of a tube and manual valve system. Food and
water will be sufficient for the mission and a postlanding period of 48 hours.
WASTE DISPOSAL
Feces will be collected in a glove-like plastic bag. Urine samples are taken,
and the remainder disposed of by overboard dumping. The urine samples and feces
waste containers are stowed in the right aft pressurized container which allows
cabin depressurization without possible boiling off of the waste materials
moisture content.
STOWAGE PROVISIONS
Personal stowage facilities are provided by retaining removed portions of the
F_ pressure suit and other equipment as required. These provisions consist of
floor pouches, Velcro covered areas on the walls of the pressure vessel, adjacent
to the pilots and attached to the structure in usable areas. Items to be stowed
utilize the hook and pile principle of mating Velcro patches.
3-29/30
SEQUENCE SYSTEM
Secfionll/
TABLE OF CONTENTS IIV_i_iiiil :::::ill
TITLE PAGE _!!!!i iiiii:i!!!!
SYSTEM DESCRIPTION 4-3 _ _
SYSTEM OPERATION ..................................... 4-5 ::=_;_'--_;_;_;;_4_PRE-LAUNCH 4-5 _" "_ _
................................................LIFT -OFF ....................................................... 4-6 _E_: ,_BOOST AND STAGING 4-7 "'-'_'E_"_._
..................................SEPARATION AND INSERTION ........................ 4-9 _ji_--_-_iii_ _PREPARE-TO-GO TO RETROGRADE .................. 4-12 _ii_-_:--;.'-=::_ii_._
TIME TO RETROGRADE MINUS 30 SECONDS .... 4-18 _..:-...-_..:..--:_.:.:-..E.J:,_oo._.oo°°°.°.*_°oo_°_.
RETROGRADE SEQUENCE ............................... 4-19 ...........................::-:'"'".:."'-:'-:'""-:'_=,°o°,.°°...,.°°o.°o°._.,_,°°°o.._.°°°°°..°._°oo°°_
RE-ENTRY ..................................................... 4-23 iii!iiiiiiiiiii _ii_ _iiiiiii
,°,°,°.**°°,,.°°°°****_***,
ABORT MODES 4 24 ...........................•, •.. ,°, ...,,,., ............................. .°**,*°.°..°*.,,o°**°***°°,
A BO RT SEQ UEN CE......................................... 4- 30 iiiiiiiiii!iiiii_iiiiiii_ii
SYSTEM UNITS ............................................. 4-38 HiiiiiiiiHiiH!iiiiiiiiii*,,,°..o.°._°°o,..°°,°._,,:::::::::::::::::::::::::::
LEFTSWITCH/CIRCUIT BREAKER PANEL ............ 4-39 _iiiiiiiiiii_i!iiiiiii_i!!!...... ..,°...°.._,°°,.°oo°,
BO OST-I NSERT-ABORT CO NTROLS AN D iiiiii_iiiiiii!i_iiiiiiiii!........ .o....°°...,°°°..°.°. ........ .°.,,......°_..°_• .°° ....... .°....o°**°.°..,• ..° ...... ..o.°o°.°°o...°°,
INDICATORS ............................................... 4-44 !!ii_iiiiiiiiiiii_ii_iiiiii,°°...**..°°...°.....o°..,.°o. ........ .°°°°...**°...,.
SEQUENCE CONTROLS AND INDICATORS. ........ 4 -47 .....""........................................-'"'-"--'-.....'"'"'"'"'""................• oo. ....... .o.°o°.°.°,°,.....o°..o.°...o .,°.°.°.°°.°..
RE ENTRY VEHICLE RELAY PANELS 4 51 ..................................... ..,°...°...o.°°o,........... . ......... .°°,°.,
_o°, .°.o°o°°.°.°.. _ ° ......... °°°°°o ........ °.,......... ..o.o° ....... °,,......... .o°... ........ °.°..,.... .....°. ......... ......,
ADAPTER RELAY PANELS ................................ 4-52 iiiii!ii!iiiiiiiiiii!!iiiii..°,.°o ......... °° .........
SEPARATION SENSORS ................................... 4-53 iiiiiiiiiiiiiiiiiiiiiiiiiil,.o°, ........... ° ..........
::::::::::::::::::::::::::::::::::::::::::::::::::::::
_-1 :::::::::::::::::::::::::::
SEDR 300 ______ PROJECT GEMINI ,PACECRAPTSEPARAT,ON_---
SENSORS (3)
EQUIPMENT ADAPTERSEPARATION SENSORS I
__i EPARATING
SECTION
DAMS SQUIB FIRERELAY PA
ACECRAFT SEPARATIONTYPICAL SEPARATION SENSOR RELAY PANEL
INSTRUMENTATION SEQUENCE
MON'T:RA% RP2---RELAY
R & RSECTION SEPARATIONCONTROL RELA'
COMMUNICATIONS
RELAYPANEL_
NOSE FAIRING JEITISONRELAY
MAIN iNSTRUMENT PANEL(UPPER ASSEMBLY)
RELAY PANEL _/ _""_,RETRO SEQUENCE ADAPTERSEPARATE RELAY PANEL
ER POWER SUPPLY
RELAY PANEL
ACS-RCS _!
DROGUE CHUTE RELA_
RELAY PANELINSTRUMENTATION POWER ETTISON
DOCKING RELAY CONTROL RELAy PANEL RELAY PANEL
PANEL (S/C 6, 8 & UP) ECS RELAY STRIBUTIONPANEL RELAY PANEL
UMBILICAL PYRO SWITCH COMMAND PILOT'S PANEL
RELAY PANEL PANEL _ _t
LEFT SWITCH/CIRCUIT BREAKER PANEL _........,,, I
@ © _ :11i An_PT l
I II II I .....,[NGINE I ENGLNE I EI,_IN_ZI
_ i| i ,E,_e i@[IANrlNOCNILigOOSTINSERT I REIROSEQ ILANOINGSEQ SEQILGHTS ATTIND '_/!
Io1__ _'1 _ I _ I _ I_" I',,o_,,.L_ ,_,o,0,_tc,"
.,. PWR ,J_ 'JETT .... ,!. 2 ,_. 3 ,_. ' /
| i ' I--- I / _*'_ / I L I / _ -_-"xj_ BOOST._NSERTI R_T_O / LANDING t RETRO ROCkEI SQUIB5 I
OL-- _ \_ HANDLE
Figure 4-1 Sequential System
4-2
__ SEDR 300
PROJECT GEMINI
SECTION IV SEQUENCE SYSTEM
SYSTem4 DESCRIPTION
The Sequence System of Gemini Spacecraft 5, 6, and 8 through 12 comprises those
controls, indicators, relays, sensors and timing devices which provide semiauto-
matic control of the spacecraft and/or launch vehicle during the critical control
times, but which are not part of other systems. (See Figure t-l). The critical
times are: the time from booster engine ignition through insertion into orbit;
the time to prepare to go to retrograde through post-landing; and the time to abort.
The Gemini crew does not control the spacecraft during boost through Second Stage
Engine Cutoff (SSEC0). The spacecraft is controlled by Radio Guidance System (RGS)
and the Digital Command System (DCS), or by the Inertial Guidance System (IGS) and
_ the on-board computer. The crew does however, monitor certain indicators to keep
informed of the operation of the launch vehicle, to anticipate a crisis if one
should develops, and to know if and when mission abort is mandatory. After SSEC0
the command pilot takes necessary action to separate the spacecraft from the launch
vehicle and applies final thrust to place the spacecraft in the desired orbit.
During orbit, the Sequence System is in standby. The electronic timer, however,
which is part of the Time Reference System, is counting down the time-to-go to
retrograde.
At _ minutes and 16 seconds before retrograde, (TR-256 seconds), a Sequence System
relay is actuated, and several Sequence System indicators ill_uate amber. These
Xndlcators provide the crew with cues for necessary operations. Again at 30 seconds
before retrograde, the crew is reminded to separate the adapter equipment and arm
the automatic retrograde rocket firing circuits. The Sequence System, if properly
4-3
_i_. SEDR 300
_ .... RgROFIRE g_:_"_g_Oa-_ORBIT JETTADAPTFF T = TR
TR ","43 DEC.
/.
PREPARE =TO-GO TO RETROGRADE RETROGRADE ! DROGUE J'
_ SB_ARATION ARM AUTO RETROIND/SW: GREEN CHUTE J,
Y
AUTO RETROFIREINITIATED DEPLOYEDT=356 SEC. TR -30 MINUTES MANUAL FIRERETROSWITCH:DEP_ESSED SOK'
C-BAND BEACON SWITCH= ON # I RETROTIREPLATFORMALIGNED F3 RETROFIRE:S.5-SECOND TIME DELAY
: ORBIT TR -356 SECONDS RELAy ENERGIZED '2 RETROFIRE=11-SECOND TIME DELAy _!_
JF4RETROFIRE:16*5-3ECOND TIME DELAyINDICATE RETROATTITUDESWITCH
MAtN BATtERy SWITCHEDON 45-3ECOND TIME DELAy RELAYRENDEZVOUS j /v {BATTERYPOWERINDICATOR: GREEN) JETTISONRETROIND/SW: AMBER
RCS INDIC./SW: AMBER/GREEN RETROJETT SQUIB SW: ARMJ TR -30 SECONDS RELAy ENERGIZED JETT RETRO/SW: F_ESS-AMBERLIGHT OUT
I RETROROC RETSQUiBSI ARMED PYRO SWITCHES,GUILLOTINES ANDARM AUTO RETROIND./SW= AMBER SHADEDCHARGESIGNITEDSEPADAETIND./SWI AMEER/GREEN RSTROIND= OFFELECIND./SW: AMBER/GREEN _ R& RSECTION
BIA SQUIB 8US & MAIN BATTFFYS: OFF DAMS ]ND./SW: AMBER/GREEN / SEPARATION
EXTEND HFANT. & SELECTHF T/R _f I
RErROEOC/_T SQUIBSSWITCH TO SAFENULL INCREMENTAL VELOCIIY INDICATORSPACECRAFTSEPARATION: SENSE& INDICATEAFT THRUSIFFSARE FIRED
RE-ENTRY 1 1
_! SETSPCFTSW: PRESSED ROLLED 180°:,_c_ SECG= T +336 SEC. MAiN CHUTEGAME: POWERON DEPLOYED_<_ FAIRINGS JETTISONED RADIO GUIDANCE ]0.6K'_ ENGINE U THRUSTDECAy ( --_20 SEC) TLM MONITORBACKUP GUIDANCE SELECTOR 80K FT. - COMPUTERCOMMANDS HDS. DOW
: SECONDARY AUTO PILOT 3OMPUTER-OFFSECONDARY GUIDANCE INDICATOR
SEPARATION AND INSERTION GUIDANCE SWITCH OVER RELAY :i_ATTITUDERATEINDICATOR
_ RGS-IGS SWITCH "_
/ MDP ELECTRONIC SWITCHES
MALFUNCTION DETECTION PACKAGE LANDINGPRIMARy HYDRAULICPRESSURESENSORPRIMARY HYDRAULICSYSTEM ALTIMETFF-60K FT.PRIMARy HYDRAULICACTUATORS LANDING SQUIB BUS;ARM
ENGINE il SHUTDOWN PRiMARy AUTO PILOT BAROSTATSARMRELAYENGINE [I FUEL & OXIDANT METERS- GUIDANCE (TARS) 50R FT ALTITUDEREQ=D. PRESS, DROGUE CHUTE DEPLOY SWITCHENGINE II THRUSTCHAMFERPRESSURE>5,.,_ TELEMETRY AND GUIDANCE RCSRINGS SWITCH;OFFSTAGE I1 IGNITION 40R FT. BAROSTAT& INDICATOR
STAGES I & II SEPARATION DROGUE CHUTE; DEPLOY, DISREEF
STAGE I SHUTDOWN PARA DEPLOY SWITCHETAO,NG_NDICAT,ONSMON'TOREO R_RSECT,ONSEPARAT,ONCHUTESTAG'NGSW'TCHRS_LOS" F=OSW,TCHEEGU,LOT,NEE D,SREEEEDSTAGING CONlf.OL RELAYSENERGIZED MAIN CHUTE DEPLOYED
ENGINE I FUEL& OXIDANT METERSREQ'D. DESCENTANTENNA; EXTENDSTAGING PRESS. UHF RESCUEBEACON ANTENNA: EXTEND
_ T_l_ ENGINE I UNDE_FFESSUREINDiCATOR EXT WATERSEALVALVE; CLOSE
_ SECONDS STAGE I ENGINECH_RFF FFESSUREE;'_"tlJ 522K FT & UP: NORMAL INEFFTION
ii _ AND RE-ENTRY (_
II 70K-522KFT: MODIFIEDRE-ENTRYI-II 15K-70K FT: RIDE-IT-OUTABORT
_ I I O-15K _FT: SLATEJECTION SUSPENSION2-POINT
PROGRAM INITIATE SIGNAL COMMUNICATIONS SELEC] IMPACTFIRSTMOTION SENSING ABORT HANDLE POSITION: NORMAL PARA JETTSWITCHHOLD DOWN ]BOLTFIRECOMMAND SEQUENCE LIGHTS CHECK MAIN CHUTE JETTISON RELAY& IGN.NO-SECOND TIME DELAY SWITCH POSITIONS SELECT FLASHING LIGHT; EXTENDTH_.USTCffA,_BFF PRESSUREJE_JlLDUE CIRCUIT BREAKFFPOSITIONS SELECT HOIST LOOPSTAGE I ENGINE IGNITION CREWINGRESS HF WHIP: EXTEND
AC POWER& INSTRUMENTATION_ OFFLIFT-OFF _.T =O SEC LIFT-OFF PRE-LAU NCH
Figure 4-2 Sequential System Simplified Block Diagram
4-4
•. . S|DR300
PROJECT GEMINI
armed, will initiate retrograde automatically. The crew redundantly initiates retro-
grade manually as a safety precaution. During descent, altitude indicators illumi-
nate as cues to deploy parachutes. After splash down, the main parachute is
Jettisoned, and all systems are shutdown.
Four abort modes comprise the abort sequence. They are: seat ejection (mode I);
ride-it-out abort (mode I-If); modified re-entry (mode If); and normal re-entry
(mode Ill). The mode selected for abort is related to the spacecraft altitude at
the time the abort command is given.
SYST_40P_TION
To simplify explanation, the Sequence System is divided into eight stages. The
_. eight stages are; pre-launch, lift-off, boost and staging, separation and In-
sertion, prepare-to-go to retrograde, retrograde, re-entry, and abort. Figure _-2
shows these sequence stages, and the detailed function of each stage. Telemetry
guidance, landing and post-landing are related to but not part of the Sequence
System. The simplified block diagram is explained in the following paragraphs.
Pre-launch, lift-off, boost and staging, and separation and insertion are explained
first. Prepara-to-go to retrograde, retrograde, and re-entry are discussed next.
Abort is discussed last.
PRE-LAUNCH
The command pilot and the pilot ingress the Gemini cabin and take their assigned
crew stations. The hatches are closed and locked. The crew checks that both D-
rings are unstowed. The c_..and pilot makes sure that the abort control handle
is in the NORMAL position; the maneuver controller is stowed; the altimeter is
set; and the Incremental Velocity Indicator (M) is zeroed. He verifies that
4-5
SEDR300
the nine sequence indicators, the two ABORT indicator lights, the ATT RATE indi-
cator light, the SEC _IDANCE indicator light, both ENGINE I indicator lights, and
the ENGINE II indicator light are extinguished. He places the top three rows of
circuit breakers on the left switch/circult breaker panel to the closed (up)
position. He places the BOOST-INSERT and RETRO ROCE_T SQUIB switches in the
bottom bow to ABM, and the RETRO and LANDING switches to SAFE. He tests the nine
sequence indicators with the SEQ LIGHTS _ST switch. He selects switches for
gyro run-up and platform alignment, and performs on-board computer checkout.
The pilot places the four MAIN BATTERIES switches and the three SQUIB BA_EI_IES
switches to ON. Both pilots select and check their intercom and uhf communica-
tions. The remaining controls and indicators are also monitored or positioned as
required. The crew verifies and reports all systems ready for launch.
LIFT-OFF
When the pre-launch countdown reaches zero, the first stage engine ignition signal
is given from the blockhouse. Both first stage engines begin thrust chamber pres-
sure buildup. Both ENGINE I indicators illuminate red but extinguish in about one
second. When the thrust chamber pressure of these two engines exceeds 77 percent of
rated pressure, a two-second time delay is initiated in the blockhouse. If all
systems remain go during this delay, the hold-down-bolt fire cc-_ud is given and
the launch vehicle is commltted to flight. First motion sensors detect vehicle
ascent one and one-half inches off the pad, and energize time-zero relays in the
blockhouse and in the spacecraft. A l_5-second shutdown arm time delay is initiated
to prevent accidental booster engine shutdown prior to the scheduled staging time.
The umbilical release command is given, disconnecting the adapter, and re-entry
umbilicals. The on-hoard computer is switched from the guidance inhibit mode to
SEDR300
EMINI
the guidance initiate mode and enabled to accept acceleration data. The lift-off
signal is also applied to the electronic T.4m_rand the event tlmer. The electronlc
tlm_T beduinsto count down the tlme-to-go to retrograde. The event timer begins to
count up the time from lif_-ofT.
BOOST ARD STAGING
As the missile continues to climb, the crew monitor the boost sequence and ABORT
Ind/cators. The two ENGINE I under pressure Indleators, the ATT _ Ind£cator
and both ABORT indicators _st remain extinsulshed. The EIglRR II indlcator
illuminates amber. The STAGE I FUEL and OXIDT_R_ needles must indicate pressures
within the required limits, and the L0_G1TUDINAL ACC_RR_ n_st indicate an
increasing acceleration within prescribed limits for the flight t_me indicated by
_ the event timer. The pilots monitor their indicators and report via uhf l_nk to
the ground. Abort mode I prevails during the first 50 seconds of flight. Ground
stations notify the pilot_ when abort mode I is no longer appllcablet and when
abort mode I - II becomes applicable. Abort mode I-XI is in effect during approx-
imately the next _5 seconds of flight. At T+95 seconds, the crew receives and
acknowledges changeover to abort mode II.
At T+145 seconds, when the acceleration has climbed to nearly 6g's, the first
stage engine shutdown arm relays are energized. At appro_4m-tely T+153 seconds,
the thrust chamber pressure drops to less than 68 percent. The two ENG]I_EI indi-
cators illuminate red, and the staging control relays are energized. The staging
switches are closed. The stage I shutdown solenoids energize and both engines are
shutdown. Acceleration drops sharply to appro_dmately l.Sg's. The booster se-
/-_ quential system immed/ately ignites the second stage engine. The explosive bolts
which unite stage i and stage 2 are detonated, and the stages separate. Both
4-7
SEDR 300
STAGE I
STAGE I FUEL & OXIDIZER OC SIGNALS llJ. FUEL & OXIDIZERPRESSURE
INDICATORS
BACK-UP GUIDANCE "ON" SECONDARY_- GUIDANCE
LIGHT (AMBER)
ENGINE I
STAGE I SUBASSEMBLY I ENGINE UNDERPRESSURE SIGNAL _- UNDERPRESSURELIGHT (RED)
(SAI RT. HD. LT. _,
ATTITUDEATTITUDE OVERRATES
OVERRATELIGHT (RED)
ENGINE I I
STAGE I SUBASSEMBLY 2 ENGINE UNDERPRESSURE SIGNAL UNDERPRESSURETITAN _ LIGHT (RED)
LAUNCH (SA2LF.HD.LT.]VEHICLE
GROUND COMMAND SHUTDOWN (ABORT) ABORT LIGHT(RED) (COMMANDPILOT'SPANEL)
STAGE IISTAGE II FUEL & OXIDIZER DC SIGNALS __ FUEL & OXIDIZEI
-- PRESSURE
INDICATORS
GROUND COMMAND SHUTDOWN (ABORT) ABORT LIGHTlip (RED) (PILOT'S
PANEL) _
ENGINE II I
STAGE II FUEL INJECTOR UNDERPRESSURESIGNAL _ UNDERPRESSURELIGHT (AMBER)
LIFT'OFF SIGNAL J*='*_i
MAIN
I ABORTICOMMAND RELAYELECTRONIC
TIMER T,MERPOWER J I• _ _ I ELECTRONICR-_ A T,MERSTART T,MEB
TIMER TIMER POWER l l
- O_O " _ EVENT TIMER8-,, K3-,_IIT,MERSTART
NOTE
RELAy K3-11, THE LiFT-OFF RELAY, IS LOCATEDON THE COMMUNICATIONS RELAy PANEL.
Figure 4-3 Boost and Staging Sequence
4-8
SEDR300
PRO GEMINI
ENGINE I indicators are extinguished. Fuel in_ector pressure of the second stage
engine rapidly increases above 55 percent, ext!_tlshing the ENGINE II under-
pressure indicator. The IDNGITUDINAL ACCETRROMETER begins to c!_mb slowly. The
crew reports the results of the staging sequence to the ground station. (Refer to
The E_GINE II underpressure indicator, the Attitude Overrate (ATT RATE) indicator,
and the two ABORT indicators must remain extinguished. The STAGE 2 FUEL and
OXIDT_ needles must indicate the required pressures, and the LONGITUDINAL ACCEL-
EROME_R must show the required increase.
At approximately T+310 seconds, the spacecraft has climbed above 522,000 feet and
its velocity exceeds 80 percent of orbital velocity. The ground station notifies
the crew that abort mode IIl now replaces abort mode II. Both pilots acknowledge
the change of abort modes.
SEPARATIOR AND INSERTION
At T+330 seconds, the acceleration has climbed to almost 7g's, and the spacecraft
has nearly reached orbital velocity and altitude. Appro_ately 337 seconds after
lift-off, the blockhouse computer transmits the SSECO c_nd tones via the Digital
Command System to the launch vehicle. The SSECO solenoids energize SSECO occurs,
thrust decays, and acceleration falls rapidly. The on-board computer begins to
compute the delta-V required for insertion.
The c_nd pilot waits 20 seconds for launch vehicle thrust to decay. Near the
end of the thrust decay period, the c_nd pilot depresses and releases the J_A'A"
/_ FAIRING switch on the main instrument panel. This switch energizes nose fair'n5
Jettison relays K3-13 and K3-17 and scanner cover _ettlson relays K3-18 and K3-19.
4-9
SEDR 300
BIA SQUIBBUS NO. 1
BOOST SEP -22 - -26INSERT SPCFT
CONT. 1 SWITCH -42
S/C SEPARATION "_= "J"SQUIB BUS NO. I m
" K3-2i c LV/SC J____'_ I _.oSW,TCHG-|
¢-_ j s/cSEP',_T,ONS_PEDC_ROE_GN.ERNO.2 2-_
K_-_ J LV/SCW,BE_ I Io G0,LLOT.NEIGNITERi 1-1
COMMONCONTROL
BUS
I K3-4_ J UHFWHIPANTENNAJ____% IIA SO.NOlOWH,P ACTUATOR
ANTENNASUHF
NOTE
r==J _ | (_) LATCHING RELAY
NO.I NO. 1 O NO.2 O
J SPACECRAFT SEPARATION SENSOR SWITCHESJ SEPMAIN
BUS z,J _ SPCFT
IIASEQ LIGHTS SEQ K3-28'
PWR LIGHTS OSWITCH DIM
O AMBER
BIA SQUIB _ _ OFFBUS #l IND. LTS SEQ.
=q TEST LIGHTS OSWITCH RED & GREEN
l D.K3-,B- , II I i X| I SCANNER COVERI
K3 38 SQUIB 1-1BOOST-INSERT / Cl,,Ir -. I I
CONTROL,_ _ . _ I IJETT IFA,R,NG _ _ -- 1 I I
K3-|3 JETTISON
." _ 2!!13 . - IGNITER 1-1
RELAY REDUNDANT NOMENCLATURE RELAy PANELRELAY
K3-22 K3-23 SPACECRAFT SHAPED CHARGE IGNITION BIA S/C SEPARATION CONTROLK3-24 K3-25 LAUNCH VEHICLE/SPACECRAFT GUILLOTINE BIA S/C SEPARATION CONTROLK3-26 K3-27 LAUNCH VEHICLE/SPACECRAFT PYRO SWITCH BIA CONTROLK3-28 K3-29 SPACECRAFT SEPARATION SENSOR BIA CONTROLK3-42 K3"_13 UHF WHIP ANTENNA ACTUATOR COMMUNICATIONSK3-86 K3-87 NOSE FAIRING JETTISON LATCH DOCKING
K3-13 K3-17 NOSE FAIRING JETTISON BIA NOSE FAIRING JETTISONK3-18 K3-19 SCANNER COVER JETTISON ACS-RCS
K3-38 K3-39 SQU_BBUSABORT POWERD_SrR_BUT_ON
Figure 4-4 Spacecraft Separation Sequence
4-10
SI:DR 300
INI
These _ettison relays arm the nose fairing squibs and scanner cover squibs. The
squibs detonate explosive charges, which _ettison the fairing and cover.
When thrust decay is complete, the corn=and pilot, depresses and releases the ._P
SPCFT switch-indicator on the mean instrmnent panel shown on Figure _-1. When
the contacts of the SEP SPC_T switch-indicator closes, squib bus number 1 power
is applied through the closed BOOST-INSERT CONT1 circuit breaker to relays K3-22,
K3-2_, and K3-_2. (Refer to F_Eure _-_). K3-22 is the spacecraft shaped charge
ignition relay. K3-2_ is the launch vehicle/spacecraft wire guillotine relay.
K3-_2 is the uhf whip antenna extend relay. Redundant contacts of the SEP SPCFT
switch-indicator energize redundant relays with power from squib bus number 2.
For simplicity, redundant elements are no_ shown.
Time delays in the relays and pyrotechnics cause the separation events to occur
in the following sequence. K3-2_, contacts C energize the launch vehicle/spacecraft
pyrotechnic switch relay K3-26. K3-26, contacts C i_aediately fire the pyro-
tec_-ic switch, open-clrcuiting the wires on the battery side of the guillotine.
Next the wire guillotines are fired, severing the launch vehicle spacecraft wires
at the interface. Finally the spacecraft shaped charges are ignited, breaking
the structural bond between the launch vehicle and the spacecraft. The operation
of all pyrotechnics mentioned in this section is explained in Section XI.
The launch vehicle may now separate from the spacecraft, or thrust from the Orbit
Attitude and Maneuver System (OAMS) may be required to effect separation. When
two inches of separation exist at the interface, the spacecraft separation sensors
close. The spacecraft separation sensor relay K3-28 is energized when two of the
three sensor switches are actuated. Contacts A of K3-28 appl_ main bus power
4-11
s oR30o,..oJ.cTG..,,,.,!RELAy _EDUNDANI_ NOMENCLATURE RELAY PANEL
RELAy
K1-29 MAIN BATTERY POWERPOWER INDICATOR
EQUIPMENTK7-3 DISCONNECT ECS
K8-16 TR-256 SECONDS COMMU NICATIO N.C
POWER SYSTEM
ELECTRONIC K8-19 RCSAMBER LIGHT ACS & RCSTIMER TR-256 SEC SIGNAL
INDICATERETROK8-29 ATTITUDE IGS
T -25_ __
COMMON K1 I-5 RCSAUXILIARy ACS & RCSCONTROLBUS
K8-16 K11-7 KI I-9 RCS SQUIB FIRE ACS & RCS
Kll-8 Kl1-10 RCS SQUIB FiRE ACS & RCS
RCS RiNG "B"
Kll-34 Kll-35 SQUIB FIRE ACS & RCSIi
K12-5 RETRO BIAS IGS
LTS EQ(_ O BRIGHT
DiM ETROMAIN
BUS SLIGHTS
A'rr IND K12-5 _1
RETRO _ CONTROL
CONTROL O"% O_.BUS _ ,,_\
I I I ISWITCH IGS POWER 26V AC D BIAS FLIGHT
WIPER SUPPLY VOLTAGE DIRECTOR
BEF POSITION
LTS BRIGHT IbJ KI-29 K8-17
MAIN _ O_ DIMSEQ OLIGHTS
K1-29SEQ
LTS MAIN BATTERIES K ]-2r_ 9
CONTROL JBUS
NO. 1 NO.2 NO. 3 NO. 4
02 BATE VALVE
CONTROL O CABIN FAN •COMMON
(_ " _ 0 OFFCONTROL •
Bus 02.,PATET ..L.
°_."°RMiLo _ I --ABSOLUTE
P_ssw I .J-02 HI PATE I ,_. m
RECOCK _ -_,
Figure 4-5 Time to Retrograde Minus 256 Seconds Sequence (Sheet I of 2)
4-12
_.>. SEDR300
SEQ II l,v
LT_ BR,GHT_ ] _"-,RKI,-,BUS O Dj
SEQ LIGHTS
Kll-5i
RCS
sQu,B1 RCB [BUSll I ^I K11-7
RCS I
I K11-9RETRO
BUS12 A
KII-10
_ll K11-7_c_$ Ic,,K11_81
,coMMoN _'_ _,_11-R_CONTROL
BUS 'L_ I KII-S-- 11_40 K11-7 I Kl1-I0 T
D Xll-7 KII 34 PAGKAGE ABIA SQUIB 11-79 K11-34 - _; lBUS1 BII PRESSURANTISOLATION SQUIB
KTI-711-39
E Kll-7 PACKAGE C
.w_'_-_ BUS I Cll c, .11-.,, KII-7 - ,,I .. , ,so,.,TioNsQu,,R_RO _ EllBUS El BIA SQUIB T_ 11-77 Kll-34 -
F K11-7KI, 34 PAG'<AGE D
BUS1 _O__ 1 FUELL.LII-37 K11-8
r IKII_ PACKAGE AFlEe , , PRESSURANT
Kli-8 ,_rJ ' " I ISOLATION SQUIB11-36
KII-8 PACKAGE CElY , OXIDIZER
,_l " I ISOLATION SQUIB11-38 K11-8
=_= DI_ lImB FUEL41 I ISOLATION SQUIB
HI RATENNUNC1ATOR
SEQ PANEL)_'_ ' LTS BRIGHT _ K7-3
O_o l"l ,k C IIo,__ "CABIN
FAN CABIN FAN K7-3
_,N o_ 0 , E_4. I _'_F_IPOWER SUPPLyBUS SUIT FAN •
NO. I SUIT FAN 1 K7-3
POWER SUPPLySUIT PAN I O
NO.2_ O_ _rlK7-3 I
DIJ _ I SUIT FAN NO. 2
SUiT FAN 20 l I POWER SUPPLy
Figure 4-5 Time to Retrograde Minus 256 Seconds Sequence (Sheet 2 of 2)
4-13
$EDR 300 _____PROJECT"GEMINI
through the closed SF_ LIG_S PWR circuit bresd_erand the SEQ LIGHTS _IG_T-DIM
switch to the switch-indlcators. The SEP SPC_T switch-indicator illum_nates
green.
The co_And pilot observes the delta-V required for insertion which is now dis-
played on the M. He fires the aft thrusters until the IVI is nu!led. The
spacecraft is in the required orbit. The crew places the following switches to
these positions: RETRO ROCKET SQUI_ to SAFE, BOOST-INSERT SQUIB to SAFE, and
MAIN PAT_ERXES i, 2, 3 and _ to OFF. For the co,mmlcation switches positioned
at this time, refer to Section IX.
I_P_-TO-GO '10 I_'I'ROGP,A_
Approximatel_ 30 m_nutes before retrofire time, the crew places the C-band beacon _
switch to CONT and performs platform alignment procedures. Then maneuver the
spacecraft to the Blunt End Forward (_F) position.
At _R-256 second_ (_ minutes and 16 seconds before retrofire time), the electronic
timer energizes the TR-256 second relay K8-16. (See Figure _-5). The A contacts
of K8-16 close and energize K8-17, K8-19 and KS-_. K8-17 is the Electrical Power
System TR-256 relay, and its A contacts now close to illuminate the BTRY PWR in-
dicator amber. K8-19 is the Re-entry Control System (RCS) amber light relay, and
illuminates the RCS indicator amber. K8-29 is the indicate retrograde attitude
relay, and illuminates the IND RETRO ATT indicator amber.
The amber BTRY PWR indicator reminds the pilot to turn on the main batteries by
ixlacingthe four MA3_ BATTERIES switches to the ON position. Relay KI-29 is
energized through the ON position of the four battery switches. The _RY PWR _
indicator illuminates green.
$EDR 300
PROJEC'T--'G-EM IN I
Depressing the amber IND RETRO ATT switch-ind/cator energizes the retrograde bias
relay K12-5. K12-5 extinguishes the amber lamp and illuminates the green lamp of
the indicator. K12-5 also applies the retrograde attitude bias voltage to the
Flight Director Indicator (FDI), and electrically places the inertial platform in
the _EF mode. The FDI needles can now be used to orient the spacecraft in this
attitude.
DepresslnE the RCS switch-indicator energizes the four RCS squib fire relays KII-7,
F11-8, KII-9, and KII-IO. Relays KII-7 and KII-8 are energized from retrograde
bus number i while KII-9, and F1_-i0 are energized from retrograde bus number 2.
When any of the four RCS squib fire relays energize, the RCS auxiliary relay
KII-5 is latchedt changing the RCS indicator from an amber to a green ind/cation.
Relays KII-7 and KII-9 both fire the package A_ C, D, pressure isolation, oxidizer
isolation, and fuel isolation squibs of ring B. Relays KII-8 and KII-IO fire the
package A, C_ D, pressure isolation, oxidizer isolation, and fuel isolation squibs
of ring A. The RCS RING A and RING B switches are now placed to ACME, and the
attitude controller is operated to fire and test the RCS thrusters.
02 high rate flow is initiated after the TR-256 second sequences at the option
of the crew. When the CABIN FAN switch is placed to the 02 HI RATE position, the
disconnect relay K7-3 is energized. K7-3 removes power from the cabin fan power
supply and the two suit power supplies, and illuminates the amber 02 HI RATE
_ndicator.
After the TR-256 sequence, re-entry communications are selected, as discussed in
S _ Section IX.
4-15
_.--_ SEDR 300
I I---ELECTRONIC, SIGNAL
SEPOAMS
BRIGHT _ B B F LINE
PWR. SEQ. LIGHTS
_N BUS I11(4-24 q)
1(4-23 ]_
RETRO 01"Lt"_OSQUIBBUS #I RETRO SEQ.
CNTL I (_)
E11_ A I
_'h T iI
K4-64L___,._1(4_, _.-I.B 'T '1_-II 1(4-2111(4-21_ t
RETRO oi"0 1SQUIB RETRO SEQBUS #1 CNTL, I
REDUNDANT NOMENCLATURE RELAY PANELRELAY RELAY
K4-2 1(4-22 WIRE GUILLOTINE ADAPTER SEPARATE 6 K4-46
K4-3 K4-B ADAPTER SHAPED CHARGE ADAPTER SEPARATE A
1(4-15 I(4.'16 ADAPTER SEPARATE SENSOR RETRO FIRE J J K4-16
K4-74 1(4-21 WIRE GUILLOTINE ADAPTER SEPARATE : | | f---,
K4-25 K4-27 PYRO SWITCH RETRO SEPARATESQUIB 0
K4-26 K4-2B I_RO SWITCH RETRO SEPARATE BUS #I RETRO SEQ1(4-30 K4-48 RETRO ABORT PYRO SWITCH BIA CONTROL CNTL I ARM AU
RETRO J "_ IK4-36 TR ARM RETRO SEPARATE
_ RET_TR_B_=EOSEP_TE:' ;" B 4
K4-64 K4-65 SEPARATE E].ECIRECAL LATCH RETRO FIRE : | I XI
K4-66 1(4-67 ABORT DISCREI_ IGS K4-46 K4-36D
--II -"K4-36
RETRO _]
coMMoN o---"--o_r_o._'o__r____i_T_T3.CONTROL RETRO AUTO C= _°I,__L.===o=,,:=i NOTE
(_ LATCHING COIL ON K4-36 TO ARM TR SIGNALLATCHING RELAy RELAY
Y,=_ ;' _ =---_IK4-74 I _
r-L_----" TX RX 11
o--.--o ,' c>o__T,ou,..,_., "BUS #1 RETRO J
CNTL 1 SEPARATION SENSOR
II _ .,_K4-2
Figure 4-6 Time to Retrograde Minus 30 Seconds Sequence (Sheet 1 of 2)
4-16
SEDR 300
'____ PRO,_c-,-G_,N,f-.
'_ _ ID t K4-23 IGNITERLINESGUILLOTINEI_I
K4-23F
IFI l _'-_ _cW,_BGU,,L.- IG NITER C-1
K4-2
i ]IE! I _-_ BU,LLOTINBIGNITER 0-1
RETRO K4-2
SQUIB ;l' I
BUS #I _ ADAPTER EQUIP.
I_l I K4-2 SECTION WIREGUILL IGNITER E-I
K4-2
I°i • r ADA_ER• IOd_ I_l I _ SHA'EDC_'GEIGNITER 3-1
K4-3
I ' I ADAPTERIEI I K4-3 SHAPED CHARGE
IGNITER 2-1
K4-3
IFI i LAUNCH VEHICLEI_l I K_-2, _,cw,REGU,LLIGNITER CL2
K4-21
I POWBRW,,EE I GUILLOTINEf_ RBTRO K4-21
SQUIB CL/'_ I I- IGNITER DII2
BUS #2 K4-2_D
I A[_APTER EQUIP.
II I SECTION WIRE
IDI I K4-2i GUILL IGNITEREI-2
K4-21T
lit i AOAPTEREQUIPIFt - 1 WIRING PYROK4-25 SWITCHIGNITER D-!
K4-25E
E PI I ADAPTER EQUIP.: WIRING PYROI I K4-25 SWITCHIGNITER E-1
K4-25
,_ . I ,_,,B,QU,P.I_ I WIRINGPYROK4_ SWTCHGNTERg-I
RETRO K4-25
ou " i 1BUS #l L FUEL CELL
IFI _ PVRO SWITCH_ :_.........:_ K4-26 IGNITER
B-I
K4-26
E / PYROSWITCHI I .k K4-26_ IGNITER C-1
K4-26
ADAPTER EQUIPMENTWIRING"ROSWITCHIGNITER J-I
K4-26
Figure4-6 Time to Retrograde Minus 30 Seconds Sequence (Sheet 2of2)
4-17
SEDR 300
RETROGRADE MINUS 3O SECONDS
Thirty seconds prior to retrograde (_-30 seconds), the electronic timer initiates
a contact closure. This closure energizes the retrograde TR-30 seconds relay
K4-46, which illuminates the SEP OAMS LIRE_ SEP ELEC_ SEP ADAPT, and ARM AUTO
RETRO indicators amber. Figure 4-6 shows a logic presentation of the TR-30
second sequence. Some of the sequences shown in Figure 4-6 such as SEP OAM_
T.I_ES,SEP ELECT, and SEP ADAPT are performed redundantly. However, for sim-
plicity only the sequences powered from retrograde squib bus number I is shown.
As soon as the connand pilot observes that the four indicators have illuminated
amber, he depresses and releases the SEP 0AM_ LINE switch-indicator. This switch
closure energizes the OAMS propellant line guillotine relay K4-23 and the retro-
grade abort pyrotechnle squib relay K_-30. K4-23 changes the SEP OAM_ LINE indica-
tion from amber to green, fires the OAM$ propellant lines guillotine igniter i-i,
and then energizes pyrotechnic switch relays K4-25 and K4-26. Relay K4-25 and
K4-26 energize pyrotechnic switches B, C, D, E, F and J.
Next, the c@._ud pilot depresses and releases the SEP ELEC switch-indicator which
energizes wire guillotine relay K4-2. K4-2 ignites wire guillotine C_ D and E
and energizes the separate electrical latch relay K4-64. When K4-6_ energizes,
the SEP ELEC switch-indicator chemges from amber to green. Then, the c_o_-_nd
pilot initiates the equipment adapter separation sequence by depressing and re-
leasing the SEP ADAPT switch-indicator. Closure of the SEP ADAPT switch energizes
the adapter shaped charge relay K_-3 and abort discrete relay K4-66. K4-3 deton-
ates shaped charge igniter 2-1 and 3-1. The adapter equipment section separates,
and separation is sensed by three toggle sensor switches. The switches close when _"
the physical separation is one and one half inches. The closure of au_ two switches
_-18
SEDR 300
s PROJECT GEMINI
energizes the ad_pter separate sensor relay K4-15. K4-15 changes the _ ADAPT
switch-indicator from amber to green. The green SEP ADAPT light l-forms the crew
that the adapter equipment section has been Jettisoned from the spacecraft. K_-66
sends the abort transfer discrete to the on-board c_nputer.
Lastly, the co_and pilot depresses and releases the AP_ ADTO ]IE_O switch-lndl-
cator. The ARM AUTO RETRO switch latches the TR arm relay KM-36. This relay
changes the indication fr_n amber to green and arms the electronic timer for the
TR relay contact closure. The four _0 ROCKET SQUIB switches are now moved to
the ARM position.
RETROGRADESEQUENCE
_ A logic diagram of the retrograde sequence is shown in Figure _-7. As discussed
previously, whenever a sequence is initiated from retrograde squib 1_s _,m_er i,
there is an identical redundant sequence initiated from retrograde squib bus
number 2. The retrograde sequence is initiated by the TR signal __om the elec-
tronic timer. The redundant sequence is initiated manually by the crew.
At Retrograde (_), the electronic timer latches the TR signal relay K_-3_. The
TR signal relay in the latched condition energizes the retrorocket automatic fire
relay K_-7. K_-3_ also energizes the 45-second time delay relay K_-_, initiates
a 5.5-seconds, ll.O-seconds, and a 16.5-second time delay, and deactivates the
I_ platform free mode. The retrorocket automatic fire relay red,,ndemtly fires
retrorocket number I from retrograde squib bus number i end numbur 2. At the
end of the 5.5-second time delay, the retrorocket automatic fire relay K_-9 is
:4 energized. K_-9 ignites retrorocket number 3 from retrograde squib bus number l
a_ nwnber 2. Retrorocket number 2 is redundantly ignited from retrosrade squib
SEDR 300
T,_ER]S';G_,LF_'_q ..L © ,ND,C_TEBLATCHCO,LO__,CH,NGRE_
ELECTRONIC
(TRS)i (_ INDICATES RESETCOIL OF LATCHING RELAY
COMMON RETRO MANUAL RETRO MANUALCONTROL BUS FIRE SEQUENCE RETRO
T i
K4"40 CONTROL m
5 .S SEC K3-7I BU S K4-43
JjJ_ 5.5 SEC E K3i7j_._
T.o it K3-" II TO ,E ° ° I-
E
11.0 SEC K3_l K4-44
; K3-71 v 11 SEC
t.D III 41 D" t.°. t
T.D K3-71 , 16.5 SEC
rETRO :: ,4"I - cl I reTrO T.D.m SQUIBSQUIB
BUS _ -- BUSII 12 i
ROCKET K 7 ROCKETE #I
,, ,_N,TER ½E _-_ U'- I ,ON,tOtql - 41 T i
K_-8 K,-81ROCKET K_!_ ROCKET
E
,3 ,GN,tER LjEI "__e ly j. j '3 IGNITER71 - 41 T i
K4-18 K4-18
ROCKET 11 ROCKETR2 IGNITER 11 H2 iGNITER
• T RETROROCKET K 13 ROCKET
',,GN,TER sou,o I,E,?.,_,_E1 ',,ON,TERII - 4"I T 'BUS
,1 K,%1, - K4-_'/
"°'z°"I !'¢: !iH°"z°"JSCANNER _ K 18 SCANNER
HEAD SQUIB 18 COVERj SQU_S' I
JETT RETRO I t,0 0
RETRO SEQ. J -J-" K4-69
RETRO SQUIB CONTROL' i_ A 'q _ !C_,us,, 8R,o.t LtSEQ. LIGHTS O _V_ K4-19 K4-28MAIN PWR DIM _
Bus SEQ.UGHtS
Figure 4-7 Retrograde Sequence (Sheet 1 of 2)
4-20
___ SEOR 3O0 _.____
PROJECT GEMINIF
RETRO _ 1
REDUNDANTSQUIB RETRO RELAy RELAy NOMENCLATURE RELAY PANEL
BUS ` I F K4-17 ADAPTER RELAy
1_I SHAPED K3-1R K3-19 SCANNER COVER JETTISON ACS & RCS
K4-17 CHARGE I-1l K3-71 K3-72 SALVO RETROS RETROFIRE
K4-4 K4-6 45-SECOND TIME DELAY RETRO SEPARATE
K4-7 RETRO ROCKET/1 AUTO FIRE RETRO FiRE
K4__17 I RETRO K4-S REEROROCKET"IMANUALEtRE RETROFIRE
J __ K4-17 ADAPTER K4-9 RETRO ROCKET #3 AUTO FIRE RETRO FIRE
SHAPED K4-I0 RETRO ROCKET E3 MANUAL FIRE RETRO FIREK4-17 CHARGE 2-1
I. K4-11 RETRO ROCKET ``2 AUTO FIRE RETRO FIRE
K4-]2 RETRO ROCKET ``2 MANUAL FIRE RETRO FIRE
D _ 1 K4-13 RETRO ROCKET ``4 AUTO FIRE RETRO FIRE
RETRO K4-14 RETRO ROCKET ``4 MANUAL FIRE RETRO FIRE
l _)_I ADAPTER K4-|7 K4-18 RETRO SEPARATE SHAPED CHARGE RETRO SEPARATE
SHAPED K4-34 TR SIGNAL RETRO SEPARATE
K4-17_ C ! CHARGE 3-1
K4-37 MANUAL RETRO LATCH RETRO SEPARATE
K4-38 K4-39 SCANNER HEADS JETTISON ACS & RCS
K4-40 K4-43 5.5 - SECOND TIME DELAY RETROFIRE
--I K4"-41 K4-'44 11 - SECOND TIME DELAY RETROFIRElPYRO
_-_i K4-17 SWITCH K4-42 K4-45 16.5 - SECOND TIME DELAY RETROFIRE
K4 17 H-1 K4--62 K4-_3 RETRO BIAS OFF lOSK8-29 INDICATE RETRO ATTITUDE lOS
/ K|2-S RETRO BIAS lOS
RET CMN KI2-6 RE-ENTRY ROLL DISPLAy lOS
CTL BUS ATT IND K3-86 K3-87 NOSE FAIRING JETTISON LATCH DOCKING
CTL RETRO __ K4-69 K4-68 LATCH RELEASE DOCKING
_'_'_ I B IJ S _ (_ K4-73 K4-72 INDEX BAR JETT & LATCH DOOR RELEASE DOCKING
12-7 T II dl T ' K13-2 K13-1 EMERGENCYLATCHRELEASE DOCKINGj K4-63 ,L K13-3 K13-4 INDEX BAR EXTEND DOCKING
' A'tf-K4-63
K12-6 K12-6
ROLLMIX | L/H FLIGHT ROLLM[X R/H FLIGHT !i_ER_TIAL
DIRECTOR DIRECTOR lOS POWER _ _
INTERLOCK CONTROL I CONTROL SUPPLY MEASUREMENTELECTRONICS ELECTRONICS ELECTRONICS
MAINBUS IND RETRO
I II K,2_5ATTLT
SEQ LTS B _ _IV r---_lK8-29
BOOCK,NGSOO,B T .L.
"*BUS #1 _* _ _" K4-73 K12-.5 _"
LATCH # 1 RELEASE R B COVER #3
RETR( iGNITER//1 RETRO RELEASESQUIB l -73 IGNI_R1BUSt1 J _ LBUS#1
I
LATCH R2 RELEASE C I C COVER E2
IGNITER RI J -73 RELEASEIGNITER El
I i.I
II j LATCHOOLATCH ``3 RELEASE D I IlK4 D COVER /1IGNITER ``1 l -73 RELEASE
i IGNITER/1I.
I
I IINDEX BAR _ E T_ E INDEX BAR JETTEXTEND IGNITER ``1
IGNITER//I -73 (1 SEC PYRO T.D.)I
K4-73 _ "_ %-Figure 4-7 Retrograde Sequence (Sheet 2 of 2)
4-21
___ SEDR300 __
PROJECT GEMINI
_s number 1 and _mber 2 when the retroroaket aut_tic fire relay K4-31
energizes at the end of the 3_.0-second time de_y. Retroro_et automatic fire
relay K4-13 is energized at the end of the 16.5-second ttme delay. K_-23 redun-
dantly fires retrorocket number 4 from retrograde squib bus number i and number 2.
In order to assure retrograde rocket ignition, the commRud pilot initiates manual
retrograde ignition by depressing and releasing the MAN FIRE RETRO switch-indicator
approximately one second after automatic retrofire initiation. The MAN FIRE
RETR0 switch latches the manual retrograde latch relay K_-S7, energizes retrorocket
manual fire relay K4-8, and initiates the _5-second time delay relay K_-6. This
switch also initiates the 5.5-second, ll-second and 16.5-second time delays.
The 5.5, 11 and 16.5-second time delays energize retrorocket manual fire relays
K_-IO, K_-12 and K_-I4 respectively, which in turn fire retrorockets number 3, _
number 2, and number _ respectively. Retrorocket number i is fired by K_-8. As
in automatic retrorocket fire, each retrorocket is fired from retrograde squib
bus number i and number 2. Twenty-two seconds after retrofire is initiated, the
last retrorocket ceases firing. The coamand pilot moves the JETT RETRO SQUIB ARM
switch on the left switch circuit breaker panel from SAFE to ARM. Forty-five
seconds after retrograde ignition, K_-4 or K_-6 energizes and illuminates the
JETT RETRO lamp on the main instrument panel.
As soon as the command pilot observes the JETT RETRO indicator is amber, he
dspresses and releases this switch-indicator. The switch energizes the retro-
grade separate shaped charge relay K4-17, the retrograde bias off relay K_-62,
an_ the horizon scanner heads Jettison relay K_-38. Relay K_-I7 fires retrograde
adapter shaped charge igniter I-i, 2-1, and 3-1 and pyrotechnic switch H-I.
Relay K_-62 latches the re-entry roll dlsplay relay K12-6 removing roll m_
4-22
SEDR 300
PROJEC"T- GEMINI
interlock from the flight director controller. Kh-62 also resets two latch
relays: the retrograde bias relay _I__-5and the indicate retrograde attitude
relay K8-29. Relay K8-29 extlnsuishes the IRD RETRO ATT indicator. K_-I8 fires
horizon scanner cover squib i-i if it was not fired previous],7_urin8 the boost
phase. K_-38 ignites the horizon scanner head s_ulh i-i through an 80_aillisecond
pyrotech-lc time delay and Jettisons the scanner head. The firing of pyrotechnic
switch H-1 ext!_,n_ishes the SEP ELEC, SEP ADAPT, S_P OAMS, ARM ADTO RETRO and
JETT RLTRO indicators.
On spacecraft 6 and 8 through 12, the JETT RETRO switch also energizes latch release
relay K_-69 through the B contacts of thenose fairing Jettison latch relay K3-86.
K_-69 fires the release igniters of docking latches i, 2 and 3 to Jettison them.
K4-69 also energizes the index bar Jettison and latch door release relay K_-73.
KM-73 fires three latch door cover release igniters. These igniters release the
latch doors which cover the ports left by the Jettlsone_ docking latches. K4-73
also Jettisons the docking index bar. If the bar was not extended previously, it
is first extended and then Jettisoned. These functions are not a part of the
retrograde sequence dur_r_ an abort if the abort occurs prior to nose fairing
Jettison.
RE-EE_RY
After the retrograde adapter and horizon scanner heads have been Jettisoned, the
cosm_nd pilot places the RETR0 PWR and RETRO JETT squib switches to SAFE. Using
the attitude controller and the FDI needles, he rolls the spacecraft 180 degrees
so that the horizon is visible in the upper portion of his cabin windc_o He
F _ cban_es the ATTITUDE CONTROL mode select switch on the main instrument panel from
PULSE to RATE CMD (RE-ER_). The co_and pilot uses attitude control and
4-23
__ $EDR 300
PROJEC-T GEMINI
_a-euverlng eleetrmktes an_ the attitude controller to control the roll attitude
d_rin8 approxt_e]_r the next 1D minutes in which the alti___e a4mt-tshes to
_00,000 feet. As this altitude the FDI roll aeedles start to move, the computer
_ t11_ml--m_eB_ a_ the computer begins to calculate the point of impact.
The cnmmM pilot eha_es the ATT]_I)E C01_ROL mode selec_ switch _ RATE CMD
(I_-ENT) to HE-w_. The @om_uter n_ c_utes the roll attitude for optimmn re-
entry ]A_ an_ also mxtomatteal3_ controls the roll attitude. Du_ng approximately
the next 10 _tmates, the altitude decreases to 100,000 feet. At this altltude,
the altimeter tnd£eator be_ to come off the peg. At 80,000 feet, the ec_puter
Cc_sanam the spaoeorsi't to assize the best attitude for drogue parachute deploy-
ment. Then the ecmmnd p_et places all guidance and electronic switches to OFF.
ABORTMODES _-_
An abort ls an unschea, l_ed termination of the spacecraft mission. An abort may be
initiated at any time during the spacecraft mission. In all cases the actual abort
sequence has to be in4tiated by the crew after an abort cc_mnd has been received.
An abort indication consists of i11,m_natlon of the ABORT indicators located on
the coemand pilot and pilot's panels. The ABORT indicator may be illuminated hy
three different methods. During pre-launch prior to umhilic_l disconnect, the
ABORT indicator may be _11,-.1-atedfrom the blockhouse via hardline through the
launch vehicle tail plug connector. After umbilical release, the ABORT indicator
may be illu_d by ground co.w--_ to the spacecraft vla a channel of the DOS
or by ground conmand to the l_unch vehicle to shutdown the booster.
The abort sequence is part of the Sequence System. The abort sequence co_rises
the abort indicators, controls, relays, and pyrotecb-lcs. The part of the abort
sequence which the crew _-_e use of is determined by the abort mode in effect at
the time when the abort e_--.andis received or the decision to abort is made.
SEDR 300
1INITIATE NORMAL LANDING & RECOVERYDEPLOY EMERGENCY CHUTE AT 10.6K FT.DEPLOY DROGUE CHUTE AT 4OK FT.INITIATE NORMAL RE-ENTRYMANEUVER S/C TO RE'ENTRY ATTITUDEJETTISON RETRO ADAPTERRETRO ROCKETS SALVO FIREDSEPARATION FROM LAUNCH VEHICLE
ABOBT CONTROL HANDLE: ABORT NORMAL RE-ENTRY & LANDING iNITIATED5 SECONDS WAIT FOR THRUST DECAY JETT RETRO SW/LT: PRESSED/OFFABORT CONTROL HANDLE: SHUTDOWN JETT RETRO LT.:AMBERRETRO ROCKET SQUIB SWITCHES: ARMED (PRE-LAUNCH) 45 SEC. TIME DELAy FOR RETRO JETTISON
RETRO ROCKETS: RIPPLE FIRED MANUALLYAUTO RETRO SW/LTS: PRESSED/GREEN
ABORT MODE T . 1/ RCS, SEPOAMS LINES, SEP ELEC, SEPADAPT,(15,000 TO 75_000 FEET) RETRO ATTITUDE ASSUMED
BTRY FOWER LIGHT: GREEN
l MAIN BATTERIES (4): ON
IND. RETROATT SW: PRESSEDSC MANEUVERED AWAY FROM LVSEPSPCET INDICATOR: GREENSEP SPCET SWITCH PRESSEDDAMS PROP: ONDAMS PWR SW: MANUVR & ATTABORT HANDLE: SHUTDOWN
LANDING SITE CHOSEN & APPROACHED PILOT EVALUATION OF DISPLAYLIFE RAFT iNFLATED & HUNG FROM SPACESUIT ABORT iNDICATORS: REDSURVIVAL KIT LANYARD PULLED ABORT SITUATION ANALYZEDPERSONNEL CHUTE OPENS (BELOW 10s000 FT.)
BALLUPE DITCHED: 10,CCOFT10,000 FT. BAROSTAT ARMED ABORT MODE TITRALLUTE OpENS (ABOVE 17,000 ET) (ABOVE 522,000 FEET)
/_, BALLUTE LANYARD PULLED
SEAT-MAN SEPARATED T
SEPARATION SUSTA]NER FIREDSEATS GONE SENSED & TELEMETEREDSEATS EJECTEDEJECTION HATCHES ACTUATED & OPENEDD-RING PULLEDPILOT EVALUATION OF DISPLAYABORT INDICATORS: REDDESTRUCT SWITCHES ARMEDENGINE SHUTDOWN TONES SENT
FLIGHT DYNAMICS OFFICER NORMAL RE-ENTRY & LANDING PROCEDURESFLIGHT DIRECTOR CONIROL S,/C ATTITUDE TO BEF°BOOSTER SYSTEMS ENGINEER JETTISON RETRO SECTION: IND. OFFRANGE SAFETY OFFICER 45-SECOND TIME DELAY RELAY
GROUND STATION ABORT COMMANDS RETRO ROCKETS (4): FIRED SIMULTANEOUSLYABORT SITUATION ANALYZED SEP ELEC, SEPADAPT, ARM AUTO RETRO: GREENBOOST INDICATORS MONITORED RCS, SEP OAMS LINES INDICATORS: GREEN
Z7O TUBING CUTTER IGNITER
ABORT MODE I SHAPED CHARGE IGNITION RELAYSEQUIPMENT ADAPTER GUILLOTINE RELAYS
(LAUNCH TO 75,000 FEET) PYRO SWITCH RELAYS
T GUILLOTINE RELAYS
RETRO ABORT INTERLOCK RELAYSRETRO ABORT RELAYSABORT HANDLE: ABORTSTAGE I (OR II) ENGINE CUT-OFFABORT HANDLE: SHUTDOWNPILOT EVALUATION OF DISPLAYABORT INDICATORS REDGROUND STATION : ABORT COMMAND
MAIN CHUTE OPENS 5.0 SEC. ABORT SITUATION ANALYZEDSEAT-MAN SEPARATION 3.0 SEC. BOOST INDICATORS MONITOREDSUSTAINER FIRED 2.25 SEC. RETRO ROCKET SQUIB SWITCHES: ARMED (PRE-LAUNCH)SEATS GONE SENSORS (TELEMETERED) STOW D-RINGSSEATS EJECTED 2.0 SECHATCHES OPEN 1.5 SECEJECTION SEAT D-RING PULLED I SECPILOT EVALUATION OF DISPLAY ABORT MODE TrABORT INDICATORS (2) (75,000 TO 522,000 FEET)LV TAIL PLUG
kV PAD ABORT COMMAND
/ABORT MODE Z
f-_ i (PRE-LAUNCH)
Figure 4-8 Abort Modes Simplified Block Diagram
4-25
SEDR300 ____PROJECT GEMINI
The abort mode to be used at any time during the mission is determined by calcu-
lations made on the ground and depends on the altitude and velocity attained by
the spacecraft. The critical abort altitudes are 15,000 feet, 75,000 feet, and
522,000 feet. The spacecraft reaches 15,000 feet approximate],750 seconds after
llft-ofT, 75,000 feet approximately I00 seconds after lift-off, and 52,000 feet
approximately 310 seconds after lift-ofT. Below 15,000 feet, seat ejection (mode I)
is used. Between 15,000 and 75,000 feet, ride-lt-out abort (mode I-II) is used.
Between 75,000 and 522,000 feet, modified re-entry (mode II) is used. Above 522,000
feet normal re-entry (mode III) is used, except that the spacecraft electronic
timer does not illuminate the sequential indicators amber when the time to press
them occurs, ,,_less the timer is updated by ground c_and. Figure _-8 presents a
simplified block diagram of the abort sequences in each of the three modes.
Abort Mode I
When an abort becomes necessary during pre-launch, it is accomplished by usi_
abort mode I. The abort c_ is given from the blockhouse by hardline through
the launch velLicletall plug connector. The command lights both ABORT indicators
on the counand pilot and pLlot's panels. When the pilots see this display, they
4_.._4ately pull the D-rlngs attached to their ejection seats. When one D-ring
is pulled, both ejection systems are energized. One-half seconds later, the
hatches are open, and one-half second after that the seats have been e_ected.
Sensors detect the ejection of the seats and notify the blockhouse that the pilots
are out of the spacecraft. One-quarter second after the seats are e_ec_ed, a
sustalner rocket under each seat is fired, which extende the distance between the
pilots add the launch vehicle. Then a pyrotechnic ignites and separates the
ejection seat from the pilots. Two seconds after sustatner ignition, the main
-26
SEDR 300
PROd E-CT GEMINI
parachutes have opened and the pilots are lowered safely to the ground. For
illustrations and fuller descriptions of the equtl_nent used for seat ejection
abort, refer to Section III of this manual.
After normal llft-off, and before the Gemini-Titan reaches an alitutde of 15,000
feet, an abort condition could develop. The crew monitor their booster indicators
so that they are aware at all times of the manner in which the flight is proceed-
Ing. Booster operation data is telemetered to the ground for analysis and inter-
pretation. The range safety officer, the booster systems engineer, the flight
director, or the flight dynamics officer, who are on the ground, any decide that
danger is "1mm4nent and an abort mandatory.
f_ A clm_nel of the DCS is used to send the abort co_and to the spacecraft and
ground c_mmsndS are sent to the launch vehicle to shutdown the booster engines.
Then the engine shutdown tones are received, the destruct switches of the launch
vehicle are armed. The two ENGINE I indicators and both ABORT ind/cators illlm_nate
red. The command pilot and pilot evaluate these displays and pull the D-rlngs.
The hatches open and the pilots in their seats are ejected, Refer to Section III
for a description of the remainder of this sequence.
Abort Mode I - II
Abort mode I - II is the ride-it-out abort mode. It is effective at altitudes
between 15,000 and 75,000 feet approximately 50 seconds to I00 seconds after
lift-off. Abort mode I - II is used when a mode I abort is inadvisable and when a
delay to permit entry into the mode II conditions is impractical. The crew how-
_ ever has the option to eject or to ride-it-out depend_ upon their assessment of
the abort conditions. Therefore the D-rings are not stowed during the I - II mode.
4-27
SEDR300 j__._EMINI
Abort mode I - lI begins duri_ stage I boost approximately 50 seconds after lift-
off. If an abort cond_tion develops, and the crew elect to ride it out, the
co_d pilot moves the abort control handle from NORMALto _. He waits 5
seconds for booster thrust to decay, then moves the handle from SHUTD0_ to ABORT.
The retrograde abort relays and the retrograde a_ort interlock relays are ener-
gized. These relays arm the _uses needed for abort action. The retrograde co_on
control _s is arme_ from the co_on control _us. Retrograde squib buses n-m_er
1 and number 2 are arme_ from 0_ squib buses number 1 and number 2. On space-
craft 5 only, spacecraft separation squib buses n_nber 1 and number 2 are armed
fr_ Boost Tnaert Abort (BYe) squib buses number 1 and mmher 2. Two parallel
circuits are used for redundancy. This arming of buses by means of relays elimin-
ates the motion of t_ the switch ordinarily required to arm the buses. Then,
in rapid sucesston_ wire 8-tllqtine relays, pyrotechnic switch relays, and shaped
charge _ter relays are energized. The relays ignite the pyrotechnics at the
equi_nent adapter/retrogx_te adapter mati_ line, and the vehicles separate. Then,
the four retrorockets are salvo fired and the 81_cecraCt thrusts away from the
launch vehicle.
If the abort altitude is between 15,000 and 25_000 fee% the retrograde adapter
is Jettisoned 7 seconds after retrorocket salvo fire is initiated. If the abort
altitude is _etween _,000 and 75,000 feet, the retrograde adapter is Jettisoned
45 seconds after salvo fire.
After retrograde adapter Jettison, the spacecraft is maneuvered to the re-entry
attlt-_a. If the abort altltude is above _0,000 feet, the drogue parachute is
4eployed at _0,000 feet, an_ the main parachute at 10,600 feet. If the drogue
SEDR300
PROJ EC-T GEMINI
parachute fails or has not been deployed before the spacecraft descents to 10,600
feet, the emergency main parachute switch is used to deploy the main parachute.
If one of the two first stage engines should fall and the launch vehicle is above
_O,000 feet, the pilots may elect to remain with the spacecraft until the operating
engine has boosted them to 75_000 feet. At this altitude, abort mode I - II
bee_re._es inapplicable.
Abort Mod$ II
Abort mode II becomes effective above 75,000 feet. At approximately i00 seconds
after lift-off on a normal mission, the launch vehicle has hoostod the spacecraft
to an altitude of 75,000 feet. Ground station cc_uters calculate the t_e for
changeover from abert mode I - II to abort mode II. The ground station notifies
the crew via the uhf cu_=unicatic_s I_-_ of the change to abort mode II. Both
the c._,_dandpilot and pilot acknowledge the change via the same l_-k, and stow
the ejection seat handles (D-ring). Yn_tiation of abort mode I above 75,000
feet could be disastrous.
Abort mode IX begines during stage i boost before booster engine cutoff and ends
during stage 2 boost before second stage engine cutoff. The crew continues to
monitor the booster indicators. If they should notice an abort situation de-
veloping, they analyze it. The decision to abort may be theirs or it may come
fr_ the ground. If a 8round station sends the co_nand to abort, both ABC_
indicators illuminate red. In abort mode II, the command pilot must act. He
moves the abort handle to the SHOTDOWN position. The operatln6 engine is cutoff.
Since launch vehicle destruct is _-ent and escape from the fireball is urgent,
he moves the ABORT handle ¢o ABORT. The spacecraft is separated _ the launch
vehicle at the equii_ent adapter/retrograde adapter mating line. The retroroekets,
4-29
armed by four RETRO ROCKET SQUIB switches during pre-launch checkoff, are salvo
fired, prope11_ng the spacecraft away from the launch vehicle.
Since orbital velocity could not have been reached below 522,000 feet, the space-
craft i_ediately begins a re-entry trajectory. The spacecraft is maneuvered to
t_e retrograde blunt end forward attitude, the retrograde section is Jettisoned,
and normal landing procedures are initiated.
Abort Mode III
At approximately 310 seconds after lift-off, the launch vehicle reaches the alti-
tude of 522,000 feet and a velocity of approximately 21,000 feet per second. The
ground station c_.,-,._ndsa change from abort mode II to abort mode IIl via the uhf
link. _
If an abort after this time should become necessary, the ABORT indicators would
be illuminated red. The command pilot responds and moves the ABORT handle to the
SHUTDOWN position. The shutdown command is thus given to the second stage engine.
The ABORT handle remains in the S_0WN position. The command pilot then presses
the SEP SPCFT swltch-indicator on the main instrument panel. This switch fires
the shaped charges and severs the wiring at the launch vehicle/spacecraft mating
line as described earlier. 0A_8 thrust is applied to put distance between the
second stage and the spacecraft. The crew perform the TR_256 seconds and the
TR-30 seconds procedures, using the main instrument panel switch-indicators.
After retrofire has been initiated manually, normal re-entry, landing, and post-
landing procedures are followed.
ABORT SEQUENCE '
The abort sequence described herein occurs during abort modes II and I - II. The
4-30
SEDR300
PROJEC-T'-GEMI NI
description covers the series of events which the abort control handle and Figure
4-9 shows the electrical circuits which cause the abort sequence to occur. Figure
4-9 includes the switches, circuit breakers, buses, relays, and p_notechnic
igniters. A table of Figure 4-9 gives the names, reference designations and
relay panel locations of the relays and redundant relays of the abort sequence.
The redundant relays, their buses, fuses, and squibs (with a few exceptions) are
not shown, since the circuitry and end results are identical with those shown.
The o,,.Ission is made to maintain clarity and simplicity.
Abort mode I, the seat ejection mode, is not covered here. The events of this
mode are discussed in Section III of this Manual.
Abort mode llI is executed by performing a launch vehicle engine shutdown, a
spacecraft separation sequence and a retrograde sequence. Separation and retro-
grade in abort mode III differs from normal separation and retrograde in that
the abort sequence is performed without cues from the indicators on the main
instrument panel. The electrical circuits however are identical with those shown
in the shutdown sequence (Figure 4-9), the spacecraft separation (Figure 4-4), the
TR-256 seconds sequence (Figure 4-5), the TR-30 seconds sequence (Figure 4-6).
Shutdown
When the command pilot moves the abort control handle to SHUTDOWN, the SHUTDOWN
switch is closed. BIA common control bus power is applied to the launch vehicle
engine shutdown signal relays K3-28 and K3-49. This power is also applied to the
engine shutdown relays in the Tital Launch Vehicle. The operating engine(s) are
cut off. As K3-48 and K3-49 energize, common control bus power is applied through
their B contacts to the spacecraft instrumentation programmer. The programmer
4-31
sEoR3oo ......RIIK3-48"_ B, K3-4BCOMMON _ }1",._
CONTROL
BUS BII K3-49 K3-49 --
]_ BOI_NSER T _ BOOSTER CUT-OFF
COMMAND _BORT) INSTRUMENTATION
B 1J K3-93_l _ PILOT ACTUATED PROGRAMMERB K3_ ABORTSHUTDOWN
INITIATE BI r K3-92BOOST B K3-92 -- ,,rt _ : _IK3-_ :
i_ CUT-OFF 1 ABORT CONTROL mm
HANDLE SHUTDOWN ENGINE ,"
_MMONHCUT-OFE2J O' .... '_ T,TANVV : _ :
BUSCONTROL L_ j O'_ _ 11 "ENGINE
I - I
CUT-OFFCOMMAND
ABORT BOOn-INSERT ©l ..... _ _ : _ :'N"'ATE1-1CONTROL'o_ABORTB,A _-_ : _SQUIB _ RETRO PWR
BUS I ABORT CONTROL SWITCH i
]::I HANDLE ERZOBQ01BBusOAMB_SQUIB _ _ 1,1=O_ O- AlSl K3-38OAMBBQUIBC[ ; _ _ RETROSQU,B_BUS 2BUS 2
ECOMMON
CONTROL RETRO COMMONB I Kll'8
Kll-7
DII K11 34D_11-7 RCS RING A Kll-8
PACKAOEAFK"EBIA - -< PRESSURE
Rcs SQU_B ,BoLSQB ,i'[
ACTIVATE Ell K,I_4_KI,_7_ _ K'I-_- Kll-3_ _A_i_ZQGEB_ C I I _:_ C ! Kll-8 ..L
B,A ,I"-,..,__,:_ c,,. Ell
ql PACKAOE0 RAC_GE0 KII-BI
RETRO.A E _1-34 -34__ .--" DllI ".,L I I* , FUELB_0,B °"_:_1_- _,_ , ,BOLSQB, ,SOLBQB: _t" T
BUS 1 BUS I RETI_O SQUIBBUS 1
TABLE OF RELAYS
RElAy REDUNDANT RELAy NOMENCLATURE RELAY PANELRELAY
3-36 3_q7 RETROABORT RETROSEQUENCE3-38 3-39 SQUIB BUS ABORT POWER DISTRIBUTION
3'-48 3-49 L/V ENGINE SHUTDOWN SIGNAL INSTRUMENTATIONSEQUENTIAL MONITOR
3-59 3-60 ACS ABORT ACS-RCS3-71 3-72 SALVO RETRO RETROFIRE3-92 3-93 INSTRUMENTATION ABORT INSTRUMENTATION
SEQUENTIAL MONITOR4-2 4-22 WIRE GUILLOTINE RETRO SEQUENCE4-3 4-5 ADAPTER SHAPED CHARGE RETROSEQUENCE4-4 4-6 45-SECOND TiME DElAY RETRO SEPARATION4-7 4-8 RETRO ROCKET 1 AUTO FIRE RETROFIRE4-9 4-10 RETRO ROCKET 3 AUTO FIRE RETROFIRE4-11 4-12 RETRO ROCKET 2 AUTO FIRE RETROFIRE4-13 4-14 RETRO ROCKET 4 AUTO FIRE RETROFIRE4-23 4-24 DAMS LINES GUILLOTINE LATCH RETRO SEQUENCE4-25 4-27 PYRO SWITCH RETRO SEPARATION4-26 4-28 PYRO SWITCH RETRO SEPARATION4-30 4-48 L/V PYRO SWITCH ABORT BIA CONTROL4-40 4-43 5.5 SECOND TIME DELAY RETRO FIRE4-41 4-44 11.0-SECON D TIME DELAY RETROFI RE4-42 4-45 16,5-SECOND TIME DELAY RETROFIRE4-64 SEPARATE ELECTRICAL LATCH RETROFIRE4-66 4-67 ABORT DI SCRETE IO S
4-74 4-21 WIRE GUILLOTINE lATCH RETRO SEQUENCE11-7 11-9 RCS SQ UI B FIRE ACS-RCSll-B 11-10 RCS SQUIB FIRE ACS-RCS
11-25 11-26 RETRO ABORT INTERLOCK RETRO SEQUENCE11-34 11_5 RCS RING B SQUIB FIRE ACS-RCS
Figure 4-9 Abort Sequence (Sheet 1 of 2)
4-32
---":-"_. SEDR 300
I 'qL_ PROJECT GEMINI
(_) REI_O SQUIB
@ Bus, I IRETRO SQU_B _ CONTROL #1 _ Z70 TUBING
I 1 -3 CUTTER IGNITERBos, _ ; ,IK,,-25liK3-_DAMS LINE ._ TUBING CUTTER & "_ tGUILLOTINE SHAPED CHARGE
IGNITER K4-3 !EQUIPMENT I
- ADAPTER SHAPED
zli2_ ,,,K _ OILLOT,NEOPELLANTONE : IC_RGE'GNITERI_ NITER 1-II
K4 3 i EQUIPMENT
- ADAPTER SHAPEDPYROSW,TCN0_ B,, " _ Dl_'-3T_r,..C_RGE'GNTER
• I I K4-74 _ _ gS/C SEPARATION '_1 r_4-zo I-_ i.SQUIB BUS 1
C K4-26 g
PYRO _ITC_ RETROK3-71 i"G" IGNITER COMMON
- _rsEc_._.. AUTO
RETRO K3-71 K4-2
-_1 K4-25 RELAYS
IY PYRO :_NITCH D-ID K3-71
; _ K3-71--
_E_[ ADAPTER WIRING
_1 K4-25 C
; ; PYRO SWITCH E-1C K -71
I I - -ADAPTER WI_NG
/_". PYRO SWITCHES
D, E, F, B, C & J i RETRO ROCKET II K4-7
F K4-26 K4-26 _ IFUELCELLWIRINGI _ --_OA'SQUIB SWITCH F , _ i'_ _-, R_RoROCK_qF F,G..TEB - K_'-_-__-R''--. ',,GN,TERR_ROROCKETRBK_ K.-_
-'L I K4-26 PYRO 9_'ITCH "C = _oAR I ::_ K4-? RETRO ROCKET• F F
SALVO FIREROCKETS RETRO ROCKET 12 K4-10
! I_l_t" m FUEL CELL WIRING .__Oj_O_LR_ _tl j II ._ K4-' 1 RETRO ROCKETK4-26 PYRO SWITCH "J"F F
' .,--CL_I_ :: IGNITER -- K4L_F I__E_L '2 IGNITERRETRO ROCKET 14 K4-12
A K4-64 = SQUIB SWITCH p IK4-13 mI
I F F 14 IGNITER
' [_I"D"I KI"_J4_26 " RETRO SQUIB -- K4-1L_ _
K,-R I LV/SCW,RE _OB' --_1 _-_ K4-2 I GUILLOTINE, L_,_ _; ,GN,TERC
WIREOU,LLOT,NE K_-_ I POWOR_,REIGNITERS C, D & _I_E IVE K4-2 ;; I IGNITERDGUILLOTINE
NOTE
L_TCH COIL OF LATCHING RELAy
"_iK_D KI_ ! EQU'P ADAPTER I
2 GUILLOTINE
j IGNITER E
_" K4-2 (_)
K3-36K4-66
/_ _ 8A_DERTcTo_NsNSELE;EXCITAT, oB_ t__]
Figure 4-9 Abort Sequence (Sheet 2 of 2)
4-33
SEDR300
GEMINI
encodes the voltage from this bus as the booster cutoff command signal for tele-
metry transmission to the ground tracking station.
Abort _Itiate
_hen the c_mand pilots moves the abort control handle to ABORT, numerous relays
are energized, as shown of Figure _-9. However five of these relays are key
relays in that they control the principal abort operations. These operations are:
(i) telemetry of the abort action to the ground; (2) arming of the retrograde buses;
(S) activation of the RCS; (k) separation of the spacecraft from the launch vehicle;
and (5) salvo firing of the retro rockets.
The relays which control those operations are: (i) the instrumentation abort
relay, KS-_2; (2) the squib bus abort relay K3-38; (3) the Attitude Control System
abort relay K3-59; (4) the retrograde abort relay K3-36; and (5) the salvo retro-
grade relay K3-TI.
Abort Telemet:_
When the instrumentation abort relay K3-92 is energized by the abort switch, its
B contacts connect common control bus power to the spacecraft instrumentation
progra_,er. The progrs_mer encodes this signal as the pilot actuated abort signal
for tel_netery transmission to the ground.
Abort Squib _s Armin_
Abort, if it occurs, requires that power for the circuits used in the retrograde
phase of the m_ssion _come 1-_,ed.iatelyavailable. When the abort switch is
closed, squib bus power is applie_ to K3-38. K3-38 arms the retrograde squib
buses i and 2 and the retrograde co_on control bus. _"
SEDR 300
PROJE GEMINIf_
Re-entry Control System (RC8) Activation
Re-entry immediately and automatically follows an abort. Re-entry requires the
use of the RCS for control of the spacecraft during this phase. Hence the RCS
is activated. Activation involves opening and pressurizing the RCS fuel and
oxidizer lines. This is done by firing the squibs of the fuel, oxidizer, and
pressurant packages.
In operation, the abort switch applies BIA squib bus power to the Attitude Control
System abort relay K3-59. K3-59 applies retrograde squib bus power to RCS (ring A)
squib fire relay KII-8 and the RCS (ring B) squib i_re relay Kll-7. KII-8 applies
retrograde squib bus power to package A, C, and D igniters of RCS ring A. The
squibs thus fired open the ring A fuel and oxidizer lines and pressurize them.
KII-7 applies retrograde squib bus power to similar igniter of RCS ring B with
similar results.
The B contact of F13-7 and KII-8 energize the retrograde abort interlock relay
KII-22. KII-25, contact A initiates the station 7.70 separation sequence.
OAMS Lines and Lower Wires Guillotine
Since the retrorockets are to be fired in the abort modes controlled by the abort
switch, the spacecraft must separate from the launch vehicle at station ZTO.
ZTO is on the mating line between the spacecraft and the equipment adapter section.
To -_ke separation complete, the OAMS propellant lines which cross this station
must be sealed and guillotined.
The abort switch energizes the retrograde abort relay K3-36 which arms K%-23, the
OAMg lines guillotine latch relay; K4-30, the retrograde abort pyrotechnlc switch
relay; and K4-74, the wire guillotine relay. When _I-25 is energized, it energizes
4-35
SEDR300
K_-23, K4-30, and K_-7_. The D contacts of K_-23 apply power to the OA_ propel-
lant lines guiS.lotine ig-lter. The guillotine now seals and cuts the lines.
Pyrotecb-lc switch G fires, opening the launch vehicle/spacecraft interface clrcuita
The lower wire gundles are guillotin_l. The first step toward launch vehicle/
spacecraft separation has been taken.
pyrotec_nlc Switch Ignition
The second step in launch vehicle/spacecraft separation is the removal of power
from the hot wires crossing station ZTO. These wires like the propellant lines,
must also be gu_11otine_, an& the guillotine blade could cause a short circuit
of the spacecraft power. Pyrotechnic switches B, C, D, E, F, G and J must be
operated to remove power from the wires to be guillotined.
K3-36 and _1-25 apply power to launch vehicle/spacecraft pyrotechnic switch abort
relay K_-30 and to wire guillotine latch relay K_-TM_ In_tlatlng pyrotec_nlc switch
ignition. K_-SO applies Ixwer to launch ve1_icle/spacecraftwlr_ng pyrotecBnlc switch
G igniter, ope-4_g pyrotechnic switch G. KM-7_ energizes pyrotechnic switch relays
K_-25 and K_-26. K4-25 ignites equipment adapter pyrotechnic switches D, E end F.
Kk-26 ignites fuel cell wiring pyrotechnic switch B, C and S. With the operation
of the pyrotec_nlc switches, the second step in launch vehicle/spacecraft separation
has been taken.
U_per Wire Guillotine Y_nitlon
The thlr_ step in launch vehicle/spacecraft separation is the cutting of the upper
wires that cross station Z70. This is accomplished by actuatln_ the wire guill-
otines. Three wire guillotines igniters must be fired: the launch vehicle/space-
craft wlre guillotine igniter C_ the power wire guillotine igniter D, and equipment
adapter wlre guillotine Igniter E.
4-B6
SEDR 300
F PROJ E'C-'T GEMINI
When K_-25 and K_-26 energize_ they app]_ power through the A contacts of K3-TI to
wire guillotine relay K4-2. K_-2 fires the wire guillotine i_ntters C, D and Z
cutting the station ZTO wires°
K_-2, contact C energizes the separate electrical latch relay K_/o_, the adapter
shaped charge relay K_-3 and the abort discrete relay K_-66. K_-6_, contact A
latches K_-2 in the energized position. K_-66 changes the computer from the
ascent mede to the re-entry mode and enables the computer to accept re-entry
data and solve the re-entry problem. K4-3 prepares the way for the fourth step
in the separation of the launch vehicle from the spacecraft.
Tubing and Structural Bond Cutt_,
_ The fourth and final step is to sever the ad_pter skin at station Z70 an_ break
the launch vehicle to spacecraft structural bond.
When K4-2 causes K4-3, the adapter shaped charge relay_ to energize, K_-3 fires the
ZT0 tubing cutter igniter an& the equipment adapter shaped charge igniters. The
pyrotechnics co_plete the task of launch vehlcle/spacecraft separation.
Retrorocket Salvo Fire
The retrorockets are salvo fired at the same time that the tuh_n_ and structural
bond is cut. To salvo fire the retrorockets, power must be applied s_w.ltaneously
to the retrorocket automatic fire relays and thus to the retrorockets. Therefore
the 5.5, ii.0, and 16._second time delay relays must be bypassed. Contacts C, D
and E of K3-71 bypass the time delay relays. When K_-2 energizes, retrograde
co...._onbus power simultaneously energizes the retrorocket automatic fire relays
/-_ K4-7, K_-9, K_-II and K_-I3. As these relays energize, retrograde squib bus power
is applied to the igniters of retrorockets i, 3, 2 and _. Salvo burn lasts
4-37
SEDR300
PREMINI
approx_tely 5.5 seconds.
_etrograde Section Jettiso_
When the retrorocket automatic fire relays are energized by K4-2, the 45-second
time delay relay K_-_ is also energized. When K_-4 energizes after 45 seconds, it
_11_4nates the ZETT RE_110indicator as shown on Figure _-7. The ZETT P_TRO
switch-indicator is then pressed, and the retrograde section is Jettisoned in a
mode II abort. However, in a mode I - IT abort when the altitude is _etween
15,000 and 25,000 feet, the swltch-indicator is pressed seven seconds after the
retrorockets beg_n firing. After the retrograde section has been Jettisoned,
normal re-entry and landing procedures are initiated.
The Sequence System as shown in Figure 4-1 comprises the following units;
Left switch/circuit breaker panel, consisting of three rows of circuit breakers
and one row of switches.
Boost and staging indicators, consisting of seven lights and three meters on the
top of the co,_and pilot and pilot's panels.
Sequence controls, consisting of two pushbutton switches, eight switch-indicators,
and one indicator are located on the left side of the main instrument panel.
Re-entry switches and indicators, consisting of four switches on the main instru-
ment panel center console and one switch, two lights, and two meters on the
c_d pilot 's panel.
_h
Abort controls, consisting of two D-rings on the ejection seats and one abort
control handle on the left side of the cabin.
4-38
SEDR300
PROJECGEMINI
Relay panels, consisting of four relay panels in the re-entry module and four in
the equlllnentadapter an_ retrograde sections, and two In the rendezvous and
recovery section.
Separation sensing devices, consisting of three each in the equil3nentadapter
section and the retrograde section.
The c(_nents of the Sequence System are described below:
The switches amd circuit breakers on the left switch and circuit breaker panel
perform Important fumcttons in the operation of the Sequence System. The top tow
of circuit breakers however pertain largely to c_,_-.,,n_cattons. The second row
of clrcult bre_ ers perform functions related to the operation of the Sequence
System. Their functions are as follows:
E_CSRC__IC TIMER Circuit B_eaker
The electronic timer circuit breaker CB8-15 applies main bus power through contact
A of lift-off relay K3-11 to start the electronic timer when the lift-off signal
energizes the K3-11. The timer begins counting the t_ne-to-go to retrograde.
E_ T_ER Circuit Breaker_
The evont timer circuit breaker cBS-I_ applies main bus power through contact B of
lift-off relay K3-ll to start the event timer when the lift-off signal energizes
KS-11. The event counter counts the time since lift-off occured.
BOOST CUTOFF i Ci_,,cuit I_.'eakel"
The boost cutoff I circuit breaker CB3-8 applies BIA c..,..oncontrol bus power to
the booster shutdown switch on the abort control and to the secondary guidance
4-39
SEDR300
PROJGEMINI
(RGS-IGS) switch. This circuit breaker arms the booster shutdown circuit and
the secondary guidance manual swltch-over circuit.
BOOST CUTOFF 2 Circuit Breaker
The boost cutoff 2 circuit breaker CB3-21 applies BIA common control bus power
redundantly to the booster shutdown switch, and supplies power for the second
stage engine cutoff signal input to the computer.
RETRO AUTO Circuit Breaker
The retrograde fire automatic circuit bres_er CB_-l applies retrograde common
control bus power to the ARM AUTO EETRO switch. It provides power to salvo fire
the retrorockets during the abort sequence. If CB4-1 is not closed, the elec-
tronic timer TR contact closure will not automatically fire the retrorockets.
RETR0 MAN Circuit Breaker
The retrograde manual circuit breaker CB4-2 provides retrograde common control bus
power for manually firing the retrorockets, and salvo firing the retrorockets with
the abort control handle.
TR-256 Circuit Breaker
The retrograde minus 256 seconds circuit breaker CB8-16 applies common control bus
power to relay contacts in the electronic timer and contacts of the TR-256 second
relay. CB8-16 enables the TR-256 second signal to illuminate amber the IND RETRO
ATT, BTRT PWR, and RCS indicators on the main instrument panel.
SEQ LIGHTS POWER Circuit Breaker
The sequence lights power circuit breaker CB6-1 applies main bus power to the
sequence Ifght BRIGHT-DIM switch and to open contacts on the barostat switch
arm relay and the message acceptance pulse relay.
4-40
SEDR 300
PROJE(JT GEMINI
LIG_S C0_0L Circuit, ]_eaker
The sequence lights control circuit breaker CBI-13 applies c_n control bus
l_wer through the four MAIN BATTERIES switches to relay K1-29. When the main
battery power indicator relay KI-29 is energize_, the _fRY PWR indicator on the
main instrument panel is illuminated green.
The third row of circuit breakers on the left switch/circuit breaker panel perform
functions related to the Sequentiml System. The functions are the following:
ATT IND _ RETRO Circ_titBreake r
The attitude indicate control retrograds circuit breaker CB12-7 applies retrograde
c(......_n control bus power to the IND RETRO ATT switch-indicator and to contacts of
retrograde bias off relays K4-62 and K_-63. Power fram CB12-7 energizes retrograde
bias relay I(12-5when the JETT EETRO indicator is pressed.
BOOST-INSERT CONTROL i Circuit Breaker
The boost-insert control 1 circuit breaker CB3-1 provides BYA squib bus number i
power to 4-_tiate the abort sequence with the abort control handle, Jettison the
nose fairing and scanner cover, separate the spacecraft from the launch vehicle,
sense launch vehicle/spacecraft separation, extend the ,,_ and diplexer whip
antennas, and initiate several experiments.
B00_T-I_SERT _CC_ROL 2 Circuit Breaker
The boost-insert control 2 circuit breaker CB3-11 connects BIA SQUIB BUS number 2
power redundantly to the same switches to which CB3-1 connects power.
1_:_,0 _ _ i Circuit _reaker
'_ The retrograde sequence control 1 circuit breaker C_-S connects the retrograde
SqUib bus number i to the SEP OAMS T.'r_s switch-indlcator, the SEP ADAPT switch-
4-4l
SEDR300 ____PROJECT GEMINI
indicator, the _ ELEC switch-indicator, and the JETT RETR0 switch-indicator on
the main instrument panel. It also arms the abort discrete relays and the
equil_ent adapter separation sensor switches and relays.
RETR0 SE_ CRTL 2 Circuit Breaker
The retrograde sequence control 2 circuit breaker C_-28 connects the retrograde
squib bus number 2 redundantly to the same switches to which the retrograde
sequence control i circuit breaker co--ects power and arms the same cir_its.
The sequence lishts test switch connects main bus power to all amber-colored se-
quence lights and to all lights on the annunciator panel in the AMRRR positionst
and to all red or green sequence lights in the _.n & _ position.
SE_ LIGHTS (mIGHT-DIM) Switch
The sequence light brlght-dlm switch is a single-pole, double-throw toggle switch.
It connects the main bus through a diode to all sequence light circuits in the
_IGHT position. It connects the bus through a resistor to the same circuits
in the DI}_position.
The fourth row on the left switch/clrcult breaker panel contains eight switches.
These switches arm or safety the various squib buses used by the Sequential System.
Their functions are as follows.
swit
The boost-insert squib bus arm-safe switch iS a four pole, do_ble throw togsle
switch. In the ARM posltlon, thls switch arms the BIA squib buses I and 2 and the
BIA e.+-.-..oncontrol bus. These buses arm the SEP SPCFT switch-indicator, the
BOOST C0TOFF 1 and 2 circuit breakers_ the BOOST _d_ CBTL I and CsT,.2 circut
_42
SEDR 300
_-___ PROJECT GEMINI
breakers, and the relay contacts which fire the nose fairing Jettison, scanner cover
Jettison, OAMS activate, RCS activate, spacecraft separate, guillotine and
pyrotechnlcs.
Rm_Q _, (A_U-SA_)S_it,ch
The retrograde power squib bus arm-safe switch is a fcctr-pole, double-throv switch.
In the ARM position, it arms retrograde squib bus 1 and 2 and the retrograde
common control bus. Thru these buses it arms the RETRO _T ARM-SA_ s_ltch_ the
RETR0 ROCKETS0_J_S AI_-SA_ 1_ 2, 3_ and _ switches, the ATT ]_D CNTL B_R0,
RETR0 SE_. 1 and 2, and RETR0 AUTO and MAN circuit breakers on the left switch/
circuit breaker panel, and the RCS _ 1 and 2 circuit breakers on the overhead
mritch/circuit breaker panel.
REg_O _TETT (ARM,-SA_) Switch
The retrograde Jettison squib bus arm-safe switch is a two-pole double-throw toggle
switch. In the ARM position, it arms retrograde Jettison squib buses number 1 and
number 2. From these Imses, the retrograde Jettison relays Eat the power to fire
the retrograde adapter shaped charges and retrograde pyrotechnic s_teh H.
_0 _0CKZTS_U]3l, z, 3,_4 (Am_-SA_)SwAtches
The four retrograde rocket squib arm switches apply the voltages which _Ite the
four retrofire rockets to open contacts of the retro rocket automatic and manual
fire relays. In the safe position of these four switches, the ignition voltage is
removed from the relays. When both the I_TRO POWER squib arm switch and the four
RETRO R_ SQUIB arm switches ere placed %o the ARM position, the OAMS squib
buses i and 2 are connected redundantly to the retrorocket fire relays.
4-43
SEDR 300 _ ,,_ _j
PROJECT GEMINI
B00ST-IN_-_ CORTROLSAND INDICATORS
Seven indicators, three meters and four controls are provided for the boost-
insert-abort phase of the spacecraft mission.
ndicatora
The two ENG_ I indicators are provided on the co_ pilot's panel to indicate
thrust ch_nber underpressure of the first stage booster engines. Each indicator
illuminates red when the thrust chamber pressure of the engine is 68 percent of
rated pressure or less. Both indicators ill,-,_uate red at stage 1 ignition but
extin@uished 0.91%o 1.25 seconds later as the pressure increases above 68 percent.
Both indicators li!!-m_-atered at booster engine cut-off and extinguish quickly at
staging.
_GI_ II Indicator
The ENGINE II indicator on the comnand pilot's panel illuminates amher %o indicate
the fuel injector underpressure (or off) condition of the second stage engine.
The critical pressure for engine 2 is 55 percent of rated value. The indicator
ill,m_nates when the first stage engine is ignited and stays amber through first
stage boost. ApprOXimAtely one second after both ENGINE I indicators extinguish,
the ENGINE II indicator also extinguishes, indicating normal staging and engine 2
fuel injector pressure build up.
RATE Indicator
The attitude rate indicator on the command pilot's panel indicates an evaluation
of the launch vehicle attitude rates during the boost phase. The indicator is
extinguishes if the attitude rates remain within acceptable l_mttsp but ill_m4nates
red if the rates exceed these 1_m_ts.
SEDR300
f-.
S_C_ v_dicetor
The secondary guidance indicator on the c_,_A_& pilot's panel in.cares vhich
guidance system is in operation. The in_cator is extinguished to indAeate that
pr_mry guidance is being used. The indicator 43_,_-4natesamber tO indicate that
secondary guidance has been selected.
/_O_T _dAcators
Two ABORT indicators are provided, one for each pilot. Both indicators _11,_uate
red when the abort c_-.-_ndis transmitted. When the ABORT indicator is i11,,_uated,
4_4ate an_ appropriate action is imperative. The indicator signals the crew
to initiate _._,.ediatelythe abort mode appropriate for the altitude and velocity
of the spacecraft. These modes are described under Sequence System Operation.
During the l_ost phases the crew has been retainedvia the uhf com_m!cations link
of the abort mode in effect.
S_AG_ 1 I_/0XIDY_ _eters
The stage 1 fuel end oxidizer meters on the com_ud pilot's panel enable the crew
to monitor the current status and progress of the boost phase, and to anticipate
an abort condition if one should develop. These meters indicate the gas pressures
in psis of the stage 1 fuel and oxidizer tanks. Dual indicat_ needles are pro-
vid_l for redundancy. The range of the stage 1 meters is 3_ to _ psia. A time-
versus-pressure scale near the bott_ of the meter shows the minimum required
pressure at 20, 40, and 60 seconds after lift-off. Critical fuel tank pressure
is indicated by a shaded colvm_ at the low end of the scle. After stag_ with
no signals applie_, the meters indicate maximum psla.
!
4-_5
$EDR 300
PROGEMINI
STAGE 2, FIT_T./O,XID_ Meters
The stage 2 fuel and oxidizer meters on the co_and pilot's panel indicate stage
2 fuel and oxidizer tan_ pressure over a 70 to i0 psla range. Redundant pointers
are used. Critical fuel tank l_ressuresare indicated by a shaded column at the
low end of the scale. The S-flag at the 30-psia mark indicates the minimum
acceptable stored pressure in the tank before pressurization. After spacecraft
separation, the meters indicate waT1_,, psia.
Longltudlnal Accelerometer
The acceler_neter on the ccmnand pilot's panel indicates the rate in g's at which
the launch vehicle engines are changing the velocity of the spacecraft. The range
of the accelercmeter is m_-ns _'s tO l_'s. The meter has positive and negatlve
memor_ pointers. The aeceleraneter enables the crew to monitor the effectiveness
of the en_r_-es. It is a seco_aa_y indicator of staging.
_S-I_ g_danee Switch
The guidance switch above the abort control handle permits the c_ pilot to
manually change fr_n prlm-ry guidance to secondary backup guidance. When back-/
up guidance has been selected either manually or automatically during stage I
boos% and the groun_ station determines that primary guidance is feasible during
stage 2 boost, primary guidance can be selected again by m_aentarily placing the
guidance switch to the R_ position.
D-]_4-,-,s
A D-ring Is provided on the e_ectlon seat of each pilot. These rings are pulled
tO initiate mode I abort at altitude below 70,000 feet. Refer to Section III of
thls vol_ for the location and operation of these devices.
SEDR300
PROJECT GEMINI
Abort Control Handle
The abort control handle is located on the c_nd pilot's side of the cabin. It
is use_ for spacecraft re-entry in abort modes I-II, II an_ III. These modes
are effective above 25,000 feet. The three positions of this handle are NOI_4AL,
S_, and ABORT. In NOFdqAL, the handle is inoperative. When the handle is
moved to _, the e_ine cutoff c_u_and is sent _o the operating launch
vehicle englne. When the abort handle is moved to ABORT, an i_te spacecraft
separation and retrograde sequence is performed. These sequences differs from
the normal sequences in that they are performed without Ques fr_ the indicators
on the gin instrument panel.
SEQUENCE CORTROL8 ARD IRDICATORS
The switches_ indicators, and switches-lndicators on the .rain in_x-n._nt panel
center console have the fol_ring nomenclature, place in the mission sequence,
and functions.
FAIRING _shtnttton Switch
The Jettison fairing switch is used at the end of secon_ erase engine thrust decay,
by the ohm.and pilot to Jettison the nose fairing, an_ the horizon scanner head
cove r.
SEP SpC_ Switch-ludlcator
The separate spacecraft switch-indicator is used in the separation-insertion phase
of the sequence. The c_.._nd pilot presses the switch-indlcator approxlmate_v 20
seconds after second stage engine cutoff when the M displays the delta-V required
for _nser_ion. Pressing the switch-indicator causes several things to happen.
Pr4-_rily, it detonates pyrotechnic devices which separate the spacecraft from the
4-47
. $EDR 300 _ ._._[
PROJECT GEMINI
launch vehicle. Secondarily, it extends the uhf and diplexer antennas and readies
the acquisition aid beacon for use. As the spacecraf_ moves away from the launch
vehicle, separation sensors close and energize the spacecraft separation relays.
The relays illum_nate the In_cator green.
IND RE_IO ATT Switch-Ind/eator
The Indlcate retro_'aae attitude switch-lndicator is il/_ated ember when the
electronics timer energizes the TR-256 second relay. The amber light cues the
crew Co press the uwltch-indicator at this time. When pressured, a bias voltage
is placed on the pitch needle of the FDI, and the inertial platform is electrically
placed in the _ mode. When released the ember light is extinguished an_ a green
llght is _11umlna_.
The _attery power indleator _11_-In_ted amber by the _R-256 second relay. The
amber light cues the pilot to place the MAIN BAttERIES switch to ON, and the fuel
cell switch or _ ___P_IES switch to OFF. This change mus_ be made because
the ad_pter section will be Jettisoned at retrograde. When all of the _in _attery
switches are on, the Indicator changes from amber to green.
RCS Switch-Indicator
The RCS _n_4cator Is lllumlnate_ amber by the TR-256 second relay. The amber light
cues the command pilot to activate the 11CSby firing the fuel, oxidizer, and
pressurant isolation squibs. Pressing the switch-indicator energizes relays which
fire the squibs. The indicator changes from ember to green, indlcating that the
RCS has been ac_tva_e_. ._
SEDR 300
PROJ E-E'C'T'-GEMINI
SEP OAM_ LINES Switch-Indicator__ , ,,,
The separate OA_ lines indicator is illumln_ted amber by the TR-256 second relay
is the prepare-to-go to retrograde phase. The amber light cues the crew to seal
and sever the OA_ lines before jettisoning the adapter. Pressing the switch-
indicator energizes relays which ignite the pyrotechnics used to seal and sever
the lines. The relays also fire pyrotechnic switches and wire guillotines
severing some of the adapter-retrograde mating line wiring. The indicator
changes from amber to green.
E,LEC,,Switch-Indlcator
The separate electrical indicator is also ill_nated amber by the T2-256 second
relay. The amber light cues the crew to sever all the wiring at the retrograde/
F _ adapter mating line. Pressing the switch-indicator energizes the wire guillotine
relay. The pyrotechnics are detonated and the wiring is cut. The indicator
changes from amber to green to indicate that electrical separation has been
accomplished.
SEP ADAPT S_-itch-Indicator, , J
The separate adapter indicator is illuminated amber by the _R-256 second relay.
The amber light cues the crew to Jettison the adapter equipment section. Pressing
the s_itch-indicator causes the adapter shaped charge and the ZTO tubing cutter
pyrotechnic to be detonated, and the adapter section severed. Separation of the
adapter section is sensed by the equipment adapter separation sensors. Two
closed sensors energize the sensor relay and change the indicator fram amber to
green.
_ AUTO EETR0 Switch-lndicator
The arm automatic retrofire indicator is illuminated amber by the TR-SO second
relay, l_neamber light cues the crew to arm the automatic retrofire circuits so
4-49
SEDR 300
PRO,JECMINI
that when the electronic timer closes the TR contacts at TR time, the retro-
rockets will fire automatically. Pressing the switch-indicator completes the
patch from the retrograde co_on control bus to the timer TR contact, and also
energizes the TR arm relay. The relay changes the light from amber to green.
Contact closure at _R time energizes the TR signal relay. The signal relay
energizes the 45-second time delay relay, fires the retro rockets at 5.5 second
intervals, and puts the platform in the free mode.
MAN FIRE HETR0 Pushhutton Switch
The manual fire retrorockets switch connects the retrograde common control bus to
the manual retrograde latch relay. Contacts of this relay energizes the 45-second
time delay relay, fire the retrorockets at 5.5-second intervals, and place the
platform in the free mode operation.
JETT HETRO Switch-lndicator
The Jettison retrograde adapter indicator is illuminated amber by the 45-second
time delay relay 45 seconds after retrofire begins. The amber light cues the
crew to jettison the retrograde adapter. Pressing the indicator ignites pyro-
technic switch H and other pyrotechnic devices which disconnect and guillotine
the wires at the retrograde adapter section/re-entry vehicle mating line. It
fires the shaped charges which sever the retrograde adapter section from the
re-entry vehicle. It energizes the Horizon Sensor System scanner head jettison
relays which fire the jettison squibs and Jettison the scanner head. It removes
the retrograde attitude signals applied to the flight director needles at TR-256
seconds. It switches the FDI roll channel to the mix mode for re-entry. Finally
by igniting pyrotecbnl c switch H it extinguishes the IND RE_O ATT_ SEP OAMS T.T_E,
SEP ELEC, HEP ADAPT and ARM AUTO I_TRO green indicators and the JETT RETRO amber
4-5o
indicator.
RE-ENVY VEHICLERELAYPANE_
Ten Sequence System relay panels are installed in Gemini Spacecraft 5, 6, _nd 8
through 12. Four relay panels are located in the re-entry vehicle, three in the
retrograde section, one in the equipment section, and two in the rendezvous and
recovery section. See Figure 4-1. The following Sequence System relay panels
are in the re-entry module.
BIA Control Rela_ Panel
The boost-insert-abort control relay panel contains six relays to perform space-
craft separation indicator control and launch vehicle/spacecraft pyrotec_nl c
switch firing.
Retrograde Separation Relay Panel
The necessary functions required for adapter retrograde section separation are
performed by the fourteen relays of the retrograde separation relay p_el. The
relays perform such functions as pyrotechnic switch and shaped charge ignition,
TR-BO second indication, automatic IGS free mode selection, and arming of the
contacts of the Time Reference System.
ACS Scanner and RCS Squib Fire Relay Panel
Re-entry Control System squib firing, scanner cover and scanner heads jettison,
abort interlock RCS amber light actuation, and RCS ring B squib firing test prior
to launch are provided by the sixteen relays of the attitude control system scan-
ner and RCS squib fire relay panel.
Umbilical Pyrotechnic Switch Relay Panelf_
The umbilical pyrotechnic switch relay panel contains two relays which apply landing
squib bus I and 2 pc_¢er to re-entry umbilical wiring pyrotechnic switch.
4-51
_. SI:DR300
PROJ E-E'CT-'G'EM IN I._m_mmm
ADAP_ RELAY PANELS
The retrograde section contains the following three relay panels which control
spacecraft separation, retrofire, and equipment section separation. The equip-
ment section contains the Orbit Attitude Maneuver System squib fire relay
panel.
S_acecraft Separation Control Relay Panel
The spacecraft separation control relay panel contains six relays to perform the
following functions: shaped charge ignition, and launch vehicle/spacecraft
guillotine firing.
Retrograde Fire Rela_ P_nel
The retrofire relay panel has twenty relays. These relays control the automatic, --_
manual and salvo firing of the retro rockets, and time the 5.5-second firing
sequence.
Retrograde Sequence Adapter Separate Relay Panel
The retrograde sequence adapter separate relay panel contains twelve relays. The
relays are used for equipment adapter shaped charge ignition, propellant line
guillotine, electrical wire guillotine, and retrograde abort.
Orbit Attitude and Maneuver System Squib Fire Relay Panel
The 0A_¢3 squib fire relay panel contains six relays for firing the OA_ squibs
and contro111ug the regulator valves.
Rendezvous and Recovery SECTION RELAY PANELS
The Rendezvous and Recovery section contains two Sequence System relay panels :
the nose fairing Jettison relay panel, and the docking relay panel.
4-52
SEDR300 _.._
PROJ ECT-"G"EM IN I
Nose Fair_ Jettison Relay Panel
The nose fairing Jettison relay panel contains two relays which control the
Jettlsonfng of the nose fiairng.
DockinR Relay Panel
The docking relay panel has eleven relays which extend the docking index bar,
illuminate the MSG ACPT light, effect emergency release of the docking latches,
release and Jettison the locking latches at retros_ade, Jettison the index bar,
and cover the docking latch ports.
SEPARATION SENSORS
The Sequence System con+_ins two sets of separation sensors. These are the launch
_ vehlcle/spacecraf¢ separatio_ sensors and the equipment adapter/re-entry vehicle
separation sensors. Figure _-i shows their eonflguration and location. Figure
4-3 shows how the launch vehicle/spacecraft separation sensors operate. Figure
4-6 shows how the equipment section/re-entry vehicle separation sensors
operate. Separation sensors are toggle switches which are normally open before
separation is initiated. The separating structure will close the sensors as it
moves away from the spacecraft re-entry modnle. The closure of any two of a set of
three sensors is sufficient to sense and indicate separation.
ELECTRICAL POWERSYSTEM
SecfionV
TABLE OF CONTENTS/f--.
TITLE PAGE
SYSTEM DESCRIPTION ................................ 5-3
SYSTEM OPERATION ................................... 5-7 _-_=_E_ERE-LAUNCH ............................................... 5-7 iiii_.-':-:__ ":'-_ii---_ORBIT ......................................................... 5-16 ii!iiiiiiii._-_"..-_-j_j_
RE-ENTRY .................................................... 5-18 iiiiiiii!!i!iiiiiii!i!_
MONITOR AND DISPLAY ............................... 5-19 iiiiiiiiiiiiii!i!iiiiiiiiii
SYSTEMS UNITS ........................................... 5-22 !!!iHilH!!iiiiiHHililil• °..,.°, .°,.°,..°,,,°°.,t ,.
SILVER-ZINC BATTERIES ................................. 5-22 !_i_iiiiii_H_ii!!i!iiiiiil
POWER SYSTEM RELAY PANEL ...................... 5-23 iilHii!iiHi!i!HHiHHil
POWER DISTRIBUTION RELAY PANEL .............. 5-23 iiiiiiiiiiiiiiiiiiiiiiiiHi
ADAPTER POWERSUPPLY RELAY PANEL ......... 5-23 iiiiiiiiiiiiiiiiiiiiiiiiiii
AMMETERS .................................................. 5-23 iiiiiiiiiiiiiiiiiiiiHiiii!i,,,,,o,,,,,,,,,,, ........ ,_
VOLT METER ................................................ 5- 24 iHi!!H!iiii!iHiiiiiiiiillPOWER SYSTEM MONITOR ...................... ; .... 5-24 !!iiii_ii_iiiiii_i_i_iii_!i
_' FUEL CELL BATTERIES .................................... 5-25 iiHiiiiiiii!iiiiiiiiiiiiii::::::::::::::::::::::::::::::::::::::::::::::::::::::
REACTANT SUPPLY SYSTEM .......................... 5-33 ii!!!ii!ii!ilH!iiHiiiii!il,,,., .................. ,,,,,,,, ........... ,,,.., .................... ,,,,,.,..,
5-1 "_i'"'-_'.:--i_'i"_'.:i:
SEDR 300
s/c
,I _,c 2A_ s/c++&9
+° °"°+"" ++ ' IB ''"_+_-+++-'+°-+°_Pw_ CNIL pWR CNTL _ XOVE_ SECTIONI S_CIION 2
. J l i_+++i....... DETA".A| @ @
-T 7-=7, ,
j'+'++"L.!+,++, ..+?_,:+:,:,_'<,°;, ,,:,,_ ,.._r'_:! DETeU=I
:'++'" '_++I...... _+':'lj__,+, +..............+, .....+ ;;I I
.... ;,+,,,,i_. +-+,,,+.4,:+,,,+.: J_+%_I + +i3
• • i 2+: '°F_' I' + : .'o......,_o+.,,m,+/,,_,_,__SQUIBSILVER :'-+_;'--; ......................... ::........... ".--'| ..L i i,
_4AIN BATTERIES SQU_B BATTE_I+S ....._.]_..:_
:::][_,£.+:.+-, ......,,......++....I "I-] I-] +,.,,,+ .............., +,@o!,@+@',@@+@oi,@NIl:'8";, I ............. I+ +," "+""It£_'-++;I _"
MAIN SILVER JZINC BATTERIES I
IDETAIL E J i DETAIL Gii in iRenE Im im _ n ii/mm ERe
SEE DETAIL D-
SEE DETAIL AFUEL CELL
[s/c a & _ (SHO_TJ /r--FUEl CELL
__KI ,A.B+BEC.ON_UMBILICALCONNECTION
-ADAPTER pOWERSUPPLYRELAYPANEL _ i_/_
RANGE CONFIG •
0 2 TAqK
CONNECTION
Figure 5-1 Electrical Power System Installation (S/C 5,Sand 9)
5-2 . +:
$|DIt 300
PROJECT_GEMINI
SECTION V ELECTRICAL POWER SYSTEM
SYSTEM DESCRIPTION
The Electrical Power System for the Gemini Spacecraft basically consists of two
fuel cell battery sections, four silver-zlnc maln batteries and three silver-
zlnc squib batteries (spacecraft 6 uses three 400 ampere/hour sliver-zinc bat-
teries in lleu of the fuel cell batteries). No primary ac electrical power
system Is provided for the spacecraft. Devices requiring ac power obtain thls
power from self-contalned inverters within the individual systems. See Figure
5-1 for spacecraft 5, 8 and 9 configuration, Figure 5-2 for spacecraft lO, ll
and 12 configuration, and Figure 5-3 for spacecraft 6 configuration.
F _ The Electrical Power System includes switches, circuit breakers, relay panels,
ammeters, a voltmeter and telelights which provide control, distribution and moni-
toring for the system. Also included as an Electrical Power System subsystem Is
the Reactant Supply System (RSS) which provides storage and control of the reac-
tants (hydrogen and Oxygen) used for fuel cell battery operation (not applicable
to spacecraft 6). Provisions are made for utilizing external power and remote
monitoring of the spacecraft power buses during ground tests and pre-launch
operations.
The two fuel cell battery sections and four main batteries provide dc power to
the spacecraft main power bus (on spacecraft 6, the three adapter module bat-
teries and the four main batteries provide dc power to the main bus). The squib
batteries provide dc power to the common control bus and the two Orbital Attitude
Maneuvering System (O_MS) squib buses. The OAN_ squib buses in turn distribute
dc power to the Boost-Insert-Abort (BIA), retrograde, landing and agena squib
5-3
.f-_-_ SEDR 300
__ PROJECT GEMINI
5-4
SEDR 3OO ___@ PROJECT GEMINI
,.............._............... I,,I,.I._12Ai2°12cl,::_; " 2;2-UR-TW, i' ' i_ : ;_-'_;:£T ] ";__T _,
@ @ I_ l<_+!,,@J.@J,,@[I I-I-i-i-i
_...i!!ii.i_
5-8
SEDR 300
MINI
buses via the individual squib bus arming switches. See Figure 5-4 for
spacecraft 5 and 8 thru 12 configuration and Figure 5-5 for spacecraft 6
configuration Electrical Power System schematics.
The fuel cell battery sections, along with the required RSS components, are in-
stalled in the RSS/fuel cell module (Figure 5-6, 5-7 and 5-8) which is located in
the spacecraft equipment adapter section (on spacecraft 6, three silver-zinc
batteries are installed in the adapter battery module (Figure 5-9) which is
located in the spacecraft equipment adapter section.) The main and squib bat-
teries are installed in the right cabin equipment bay.
The fuel cell SECTION i and SECTION 2 POWER (PWR and CNTL switches on spacecraft
5 and 6), PURGE, X-OVER and stack control switches (IA through 2C) are located
on the right instrument panel. The fuel cell _ and CNTL switches are used to
control the battery module power on spacecraft 6. The PURGE and X-OVER switches
are inoperative on spacecraft 6.
On spacecraft 5 and 6, a dual-vertical-readout ammeter is located on the right
instrument panel. On spacecraft 5 and 8 through 12, a power system monitor consis-
ing of: a delta pressure indicator, three dual-vertical-readout ammeters and an
ac/dc voltmeter, with associated selector switches, is located on the right instru-
ment panel.
On spacecraft 5, two FCAP indicator lamps are located on the right instrument
panel. On spacecraft 6, a conventional voltmeter and _mmeter with associated
selector switches are located on the right instrument panel.
The MAINBAT_ERIES switches, SQUIB BATS_RIES switches, BUS TIE switches, agena
5-6
• _ : SEDR300 -
' : :],i
BUS ARM switch, FUEL CET.T,CONT, FCAP (FC PANE5 on spacecraft 5 and 6) and
FC 02 and H2 (CRYO 02 and H2 on spacecraft 1% Ii and 12) regulator and heater
circuit breakers are !locatedon the right switch/circuit breaker panel. The FC
02 and H2 regulator emd heater circuit breakers are inoperative on spacecraft 6.
The squib bus arming switches are located on the left switch/circuit breaker panel.
The BTRY P_R sequenc_ light, FCAP telelights, 02 and _ heater switches, 02/H2
quantity indicator (integral with Environmental Control system (ECS) 02 indicator)
and selector switch are located on the center instrument panel. The 02 and H2
heater switches, quantity indicator and selector switch are identified as CRYO
switches and indicator on spacecraft lO, Ii, and 12.
_" On spacecraft 8 and 9, an 02 CROSS-FEED switch is also located on the center
instrument panel. On spacecraft 1% ll and 12, this switch is identified as
H2 TANK VAC. The power system relay panel, power distribution relay panel and
adapter power supply relay panel are located in the left equipment area of the
cabin.
SYSteM OPR"RATION
PRE-LAUNCH
In order to conserve spacecraft battery power, external electrical power is uti-
lized during the pre-launch phase of the mission. External power is supplied to
the spacecraft common control, main and squib power buses through umbilical cables
2onnected to the re-entry module and equipment adapter section umbilical recep-
tacles. The external power source is provided by Aerospace Ground Equipment
(4).
5-7
s _--, SEDR 300
NO rE
CO==OL FC_?T,__. I_%lL¥_I
MAIN
I J _ _ _ BUS
,ASTACKoONCOST SW Clil _ l
_:_:_ OFF J SR II
AP LIGHTS
iiii_ oON _II -- -- Jl I STACKpWRAPToFUELCELLTORLYswITCHEsFUELIBCELL (S/C 51_...._i:_i:iiiiiiiiiiiii_::(s/c5) I APSW,TCHESAP LIGHTS
=========================I (s/c 8 &UP)I
' t,rO ON O_ STACK l C
SHUNTS ARE SHUNT "_"_
_sR J ..--1
_OwE' _1_,_ '_,'O%%R_-ll,AMP METERS _SR I
2A STACK OFF I
o _ I_J._STACK2AI •ON - I • _ PWRRLYI
i , _ -7 IT'-,L-_ ,-)1¼1i_ I 1_;_"_(,7=./ I I STACK2AI _ I
_STACK_ OFF ! JI I _ I I I_'_S_RL¥I =" ICONTSW_ 1.4-.4.,I,_ bl_l I L
.s/cs&uP.:::_ii::!!!iii!i!LN i ,r I'- - • , q"--.J,, ,,_-_ii_iiiiili_i_!_ -- --- J I,...............................r" "-1
SUPPLY .:_PURGE =:=:....
AMMETER(S/C 5)
]A 2B2C
: COST ICB _z ....._(
_,,,_, iiii_ SECT_ (S/C8 & uP)L__Ji ::" POWERSW_ ilSECT 2 SRRLy i I
Figure 5-4 Schematic-Electrical Power System S/C 5and 8 thru 12 (Sheet 1 of 2)
5-8
,--,._T_- SEO.3oo
k P.oJ.croE.,.,v
f_" [_. d_2 d. _o_ __'l_I_I_MA_._ATI_R_ES_' o_, u.-- BLOCKHOUSEAGE MAIN 8US
,'_' "': I TO VOLTMETER
MAIN BATTERIESSWITCHES O (TO SEQ SYSTEM J
_;_!_i_#_ 1 --_ -MAIN BATTER SEC RELAY u
PWR,NDRLY " " AGESQU,BAGE_O,. _O_.V_;O.
_ii BUS'_BUS',__ .._iiiiiii!i!_!Ni_i:;if
AGE |SQUIB l
FROM FUEL ARMING
OFF DIS- DIS-_TEST ARM ( ARM,
2C
_.:.x_.>._ VOLTMETER IAS SELSW _ , REFERTO IGS SWITCH FUNCTION LOGIC
TEST TEST SW _ BT CONTROL TO THE COMMON CONTROLSI C BUS AND OAMS SQUIB BUSES.
BUS NOTEEXTERI_[AL POWER APPLIED TOSQUIB BUSESWITH SQUIBBATTERY SWITCHES NO. 1
#I SQUIB BUS AND NO. 2 IN UMB POSITION.
: MAIN BATTERIES SWITCHES ON:
i ._0_111__ MAIN BATTERY POWER APPLIED" IRETRO JETT #2 TO MAIN BUS AND BTRY. PWR.================================
PWR _ _ dEl_ TELELIGHT ILLUMINATED GREEN,
DO e_l R_TRO _ SECTION I AND SECTION 2
BIA #1 POWER SWITCHES (SECT 1 ANDSAFE_ SECT 2 PWRAND CNTL.
_1_10 _ SWITCHES ON S/C 5)ON:
#I J J H2LO 2 SUPPLY AND H20 SHUTOFF
VALVES ENERGIZED TO OPENPOSITION.
WITH SECTION POWER SWITCHES
(S/C 8 & UP) IN WARMUP/_2 SQUIB BUS POSITION:ARMING RELAY
" _lllmO THE H2/O 2 SUPPLY AND H20SHUTOFF VALVES ARE ENERGIZED
1 _._1o4 TOOPENPOSmON.(STACK
POWER RELAYS AND STACK H2
SHUTOFF VALVES CANNOT BE3 J ENERGIZED.)
,o"_o_.,STACKCONTROLSW.C.ES" o,,_j2o,,,_!C_!B (IA THRU2C)oN:
SQU,, BAT,EmE$ _. OAMS , I I I )._._::_._-. STACK It 2 SHUTOFF VALVESSQUIB _ _
_" TO AGENA BUS #2 i ] __ j ENERGIZED TO OPEN POSITION,STACK POWER RELAYS ENERGIZED
' CIRCUITS J S_Fe SI |ii_ _ S_ TO LATCH POSITION,RETRO ROCKET SQUIB JE_i_ i::_A CONNECTING FUEL CELL STACKSTO MAIN BUS.
Figure 5-4 Schematic-Electrical Power System S/C 5 and 8 thru 12 (Sheet 2 of 2)
5-9
s y_ SEDR 300
SECt1 SECtl SECI2 SECT_ _INa_TTF_I_
/
COMMON _ TO LOAOS ":_::ii!i;_iii_ili)M:::i:i_::!ii_i?:ii?:i_i!i:ii_::ii::_....
CONTROL
B_S ''_ j 0 AG-ZN
OFF + MI
S_!!ER g" _ I + AC_v_ZNI '._/_ "--" TESTADAPTER J
BATTERY AG-ZN
I MODULE -- ÷ M3I
,'ON
BUS TIE
El UMBO"
.-- -UMBO
R :::::::::::::::::::::::x::::::::::
SECTION
AMMETERS BUS TIE --OFF 0 _I#2 UMB0 JOFF . SWITCHES
M1 SHUNT M2 SHUNT ::::::::::::::::::*:::::::::::::::::::::::::::::::'::
ON ....| ::::::::@ @o,,_OFF :::::::
SQUI_ BA'_TERIES
,\
Figure 5-5 Sehematie-Nleetrieal Power System (Spaeeeraf_ 6) (Sheet 1 of 2)
5-10
_ SEDR 300
,f_ i!!_ BATTERY I'EST AMMETER VOLTMETER
iiiii
BLOCKHOUSE ... ill
1 i',iAGE_,_BUB ill
l 2B2C
, iiliii iiiI _2A I IcI
-- B'II'--_, I TOCO_.- | CIRCUITS
ON OFF S
iii _gT BT
iiiii _ SWITCHES (TEST
AGE SQUIB AGE SQUIB AGENCROoML_OuN iiiiiill
AGE iiiii T TSQUIB i!ii TO T -5 | _lO MAIN BAITERY
A_,NG ill R_-_Y POWER,ND,_'TORRELAY#2 SWITCH #l ii
illillii!
i l _ SWITCHFUNCTIONLOGIC
ON ON DIS- ARM DIS- ( iiiiiARM ARM ii::ii I. SQUIB BATTERIES SWITCHES ON:
iili SQUIB BATTERY I_)WER APPLIED TO THE COMMON CONTROL AND DAMS SQUIBiiiii BUSES
_" ii:_i 2. MAIN BATTERIES SWITCHES ON:iiiii MAIN BATTERY POWER APPLIED TO MAIN BUS AND BTRY. PWRTELELIGHTiiiiii!iii ILLUMINATED GREEN.
iiili 3. FUEL CELL CNTL. } AND CNTL. 2 CIRCUIT BRF_KERS ENGAGED, SECTION II AND!!i SECTION R2 PWR AND CNTL. SWITCHES ON, AND STACK CONTROL SWITCHES
iii 'ATHRU2CDE=ESSBDTOONPOB,T'ON:I iiii STACK POWER RELAYS ENERGIZED TO LATCH POSITION AND ADAPTER MODULEI iii BATTERIES (At B, C) POWER APPLIED TO THE MAIN BUS.
iiiiiiiiliiiii iiiiiiiiiiiliiii iiiiiiiiiiiiiiiiiiiiii_ililii ii iiili iii-iJi iiliiiiiiii iiiiiiiil iiiiiill iiiiiii iiiilii i
f O_, A #1 SQUIB BUS "_==iiiiiiiiiiiiiiiii ' C_SAFE RETRO JETT '1
| L_ ARMINGRELAY. ,--_,_S 'Ai_'i'iiHiiii'iHi'iiii'i!=" A_.1+....... RE%FR0:!_ + _ ...........i_iii'iiRETRO JETTi_O_"_ RETRO JETT #2
BIA C_O SAFE !1
I..1" J I I
w.-_. I I #1 I
I#2 I
#2 SQUIB BUS _ TO RETRO
il)- RCW_KEI SQUIBS
I I
J -- 13 I
p, OAMBSQU,Bk _Y BUS _2 #4 I
Figure 5- 5 Schematic-Electrical Power System (Spacecraft 6) (Sheet 2 of 2)
5-11
-_-_ SEDR 300
FUELCELLBATTERYSECTIONS-
/FUEL CELL MODULESTRUCTURAL ASSEMBLY (REF)
_LATCH_YPE
SHUTOFF VALVES
TANK
TANK
PRESSURETRANSD
PRESSURE
TEMP SENSOR
QUANTITY SENSOR
ESSURESWITCH
QUANTITY INDICATORCONTROL
DC-AC iNVERTERS
QUANTITY _'_"
CONTROL QUANTITY SEN SOR PRESSURETRANSDUCER
Figure 5-6 RSS/Fuel Cell Module (S/C 5)
5-12
_" j/ "
!• iNVERTER
i02 FILL VENT
i iNDICATOR CONTROLi
_ PRESSUREREGULATORAND RELIEF VALVE
TANK
REGULATORAND RELIEF VALVE
TANK
CONTROL
SWITCH
pFILLVENT VALVE
CONNECTIONS
CONNECTOR
TEMPERATURECONNECTION
.... CONNECTION
PRESSURECONTROL SWIT( -L ......
HEAT
,_" DUAL SO LENO ID -_2.2- 2 CO N NE CTIO N
' SENSORCONNECTIONS
Figure 5-7 RSS/Fuel Cell Module (S/C 8 & 9)
5-13
__-_ SEDR 300
FUEL CELL BATTERY
AND RELIEF VALVES
STRUCTURAL ASSEMBLY (REF)
TANK
SHUTOFF VALVE
0 2 FILL AND
VENT , TANK
0 2 TANK HEATERCONNECTION
PRESSURETRANSDUCERCONNECTION
42O TANK0 TEMP SENSOR (REF)C(_NNEC TION
QUANTITY SENSOR HEAT EXCHANGERSCONNECTIONS
DC-AC INVERTER
H 2 FILL ANDQUANTITY INDICATOR VENT VALVESCONTROL
CONNECTION
QUANTITY SENSOR ESSURESWITCHCONNECTIONS
TEMP SENSOR :_
CONNECTIONPRESSURETRANSDUCER'
Figure 5-8 RSS/Fuel CellModule (S/C I0,Ii& 12)
5-14
_._. SEDR 300 ___j
©
JLESTRUCTURE ASSEMBLY (REF)
SILVER ZINC
(A, B, AND C)
Figure 5-9 Adapter Battery Module (S/C 6)
5-15
_. SEOR300
•\
SQUIB _S switches i and 2 must be placed in the umbilical (UMB) position
in order to apply external power to the spacecraft squib buses. Remote control
of the spacecraft squib bus arming relays, and remote monitoring of the space-
craft power buses is also accomplished through the re-entry _nd adapter umbili-
ca/ cables.
Prior to launch, a_1 MAIN BATJ_RIES and SQUIB BATTERIES switches, SECTION POWER
switches (SECT PWR and CNTL switches on spacecraft 5 and 6) and stack control
switches (IA through 2C) are set to the ON position to insure _v_w,lm redun-
dancy of the Electrical Power System during the launch phase of the mission.
On spacecraft 5 and 8 through 12, the fuel cell batteries are activated in
sufficient time prior to launch, to insure launch readiness of the fuel cell
batteries and RSS.
The common control bus and OA_S squib buses are switched from external power to
the squib batteries in sufficient time prior to launch, to verify the squib
battery circuits. The BIA buses are armed prior to launch by setting the BOOST-
INSERT ARM/SAFE switch to ARM position.
The re-entry and equipment adapter section umbilicals are disconnected from the
spacecraft Just prior to lift-off. Normally, umbilical separation is accomplished
by an electrical solenoid device. A backup method of separation is also provided
by a lanyard initiated mechanism which is actuated by movement of the launch
vehlcle.
ORBIT
From launch time until booster separation and insertion into orbit, both the fuel _,
5-16
DO. SEDR300
PROJEC-'G'EMINI
cell battery sections (module batteries on spacecraft 6) and the four main
batteries are connected in parallel to the main power bus. After booster
separation is accomplished, the MAIN BAT_EPS_S switches are placed in the OFF
position to conserve the main battery power. Also the pilots will disarm the
BIA squib buses by setting the BOOST-INSERT ARM/SAFE switch to SAFE.
In the event of an abort, all squib buses required for the abort function t which
are not armed prior to launch, are armed via the abort relays controlled by the
Sequential System. These relays effectively bypass the applicable squib bus arm-
ing switches which normally arm these buses.
The SQUIB BATSERIES switches remain in the ON position throughout the entire
mission until landing is accomplished. All three squib batteries are connected
to the common control bus through diodes for individual fault protection. Squib
batteries i and 2 are connected to the two 0_S squib buses via the de-energized
squib bus arming relays.
The 02 CROSS-FEED switch (spacecraft 8 and 9) re_ins in the CLOSED position
except in the event of a loss of RSS 02 tank pressure. This switch controls the
02 crossfeed valve, which provides the capability of connecting the ECS 02
supply to the fuel cell battery sections. (Refer to Figure 5-14).
The H2 TANK VAC switch (spacecraft lO, II and 12) provides the capability of
venting the area between the inner and outer wall of the RSS H2 supply tank to
outside vacuum in space. This switch, when in VENT position, initiates a pyro-
technic cutter device to perform this function. The H2 TANK VAC switch remains
in the SAFE position until the pilots elect to perform this function.
5-17
_. SEDR300
PROJECT GEMINI
The BUS TIE switches remain in the OFF position unless the necessity arises where
the pilots must use main bus power to fire the squibs. The BUS TIE switches pro-
vide a method of connecting the main bus to the OA_ squib and common control
buses. The agena BUS ARM switch will be used according to the mission require-
ments.
On spacecraft 5 and 8 through 12, a small percentage of the reactant gases must
be purged from the fuel cell batteries periodically to insure that the impurities
contained in the feed gases do not restrict reactant flow to the cells and to
remove any accumulation of product water in the gas lines. This purging function
is performed by the pilots manually actuating the 02 and H2 PURGE switches. The
pilots m_y increase the flow of gases to the fuel cell sections for more effective
purging by setting the X-0VER switch to ON position. _'"
HE-ENTRY
At 256 seconds before retrograde time (TR-256) the pilots will arm the retrograde
squib buses by setting the RETRO PWR ARM/SAFE switch and the individual RETRO
ROCKET SQUIBS No. l, 2, B or 4 ARM/SAFE switches to ARM position. The retro-
grade rockets are used according to the mission requirements.
The MAIN BATTERIES switches must be returned to the ON position at _R-256 seconds
to insure continuity of main bus power at the time of separation of the equipment
adapter section# containing the RSS/fuel cell module (battery module on space-
craft 6), fr_ the spacecraft. There is no automatic switching provided for this
function.
The stack control switches (IA through 2C) and the SECTION 1 and SECTION 2 POWER --_
switches (SECT i and SECT 2 PWR and CNTL switches on spacecraft 5 and 6) are
5-18
SEDR300
_ PROJEC--T GEMINI
set to OFF position after the main batteries are properly connected to the
main bus.
After retrograde rocket firing has been accomplished, the pilots will set the
RETRO JETT ARM/SAFE switch to ARM. This switch provides a method of arming the
JETT RETRO switch (center instrument panel), and is effectively an interlock to
prevent inadvertent Jettisoning of the retrograde section prior to firing of the
retrograde rockets.
After the equipment adapter and retrograde sections are separated from the space-
craft, the pilots will disarm the retrograde squib buses by setting the RETRO
PWR ARM/SAFE switch, HEEO JETT AR_SAFE switch and HETRO ROCEET SQUIBS AR_SAFE
switches to SAFE position. At this time, the landing squib buses are armed by
setting the LANDING ARM/SAFE switch to ARM position.
After landing is accomplished, the pilots will disarm the landing squib buses
by returning the LANDING ARM/SAFE switch to SAFE. At this time, power will be
removed from the OAM3 squib buses by setting SQUIB _IES switches 1 and 2
to OFF position. All unnecessary electrical equipment will be deactivated to
conserve the remaining spacecraft batteries for recovery equipment operation.
SQUIB BATTERIES switch 1 and the MAIN BATTERIES switches will remain in the
ON position to power the main and control buses throughout the recovery phase
of the mission.
MONITOR AND DISPLAY
Throughout the mission, visual displays of bus voltage and cub=rentare provided
by the system voltmeter and _mmeters. On spacecraft 5 and 8 through 12, a
5-19
SEDR 300
INI
power system monitor_ which consists of a delta pressure indicator, three dual
_-.,eters and an ac/dc voltmeter is utilized.
The a_eters monitor individual fuel cell stack current (IA through 2C). The volt-
meter, used in conjunction with a selector switch, displays individual fuel cell
stack voltage, co_on control bus voltage, GAMH squib bus i and 2 voltage, Inertial
Guidance System (IGS) inverter output voltage (spacecraft 8 through 12 only),
main bus voltage and individual main battery voltage with the selector switch
in Battery Test (BT) position and a particular MAIN _IES switch in _ST
position.
The delta pressure indicator, in conjunction with a selector switch, provides a
visual display of 02 versus H2 and 02 versus H20 differential pressure in the ___
fuel Cell battery sections. In the event that the differential pressure exceeds
the prescribed limits, the pilots must evaluate the fuel cell battery performance,
and if a _Ifunction exists, shut down the malfunctioning fuel cell battery sec-
tion by setting the applicable fuel cell power and stack control switches to OFF
position. The delta pressure indicator is inoperative on spacecraft 5.
An out of tolerance delta pressure indication is also provided by the fuel cell
delta pressure (FCAP) telelights on the center inatrument panel. The lights are
illumir_ted red when a malfunction exists. On spacecraft 5 only, two FCAP indi-
cator _ps on the right instrument panel are illuminated red when a malfunction
exists.
The reactant (02 and _) supply quantities are displayed on the ECS 02 quantity
indicator(centerinstrumentpanel) when the associatedselectorswitchis set _"
to FC 02 or FC H2 positions. On spacecraft lO, II and 12, this indicator
5-2o
SEDR 300
MINI
and switch are identified as CRY0 02 and H2.
The BTRY I_R sequence light, located on the center instrument panel, is illum-
inated amber at TR-256 seconds during the mission by action of the TR- 5 relay
in the power system relay panel. This informs the pilots that they must return
the MAIN BATTERIES switches to ON position to insure continuity of main bus power
due to the impending separation of the equipment adapter section containing the
adapter power supply (fuel cell battery sections on spacecraft 5 and 8 through
12 and silver-zinc batteries on spacecraft 6). With all main batteries properly
connected to the main bus, the BTRY PWR sequence light is illuminated green.
On spacecraft 5 and 6 the dual-vertical-readout section ammeter provides a
,_ display of section i and 2 main bus current. Section i includes 50 percent of
the adapter power supply current (fuel cell batteries or silver-zlnc batteries
as applicable) plus main batteries 1 and 2 current. Section 2 includes 50 percent
of the adapter power supply current plus main batteries 3 and 4 current.
The stack ammeter (used for battery test ammeter on spacecraft 6) with selector
switch in 1A, IB, 1C or 2A, 2B, 2C positions, displays 50 percent of adapter module
battery current. With the selector switch in BT position, the ammeter displays
individual main battery test current as the appropriate MAIN BATIERIES switch is
set to TEST position.
Displays of common control bus voltage, main bus voltage, OAMS squib bus voltage,
and adapter module battery voltage is provided by the system voltmeter and
selector switch. Individual main battery voltage (with the particular battery
f-_ removed from the main bus) is monitored with the voltmeter selector switch in
BT position and the applicable MAIN BATTERIES switch in TEST position.
5-21
SEDR300 _.___
PRO,JECT GEMINI
The FCAP telellghts and reactant quantity indications are not operative on
spacecraft 6.
SYSTEMS UNITS
SILVER-ZINC BATTERIES
The four main batteries are 45 ampere/hour, 16 cell, silver-zinc batteries. The
three squib batteries are 15 ampere/hour, 16 cell, silver-zinc batteries. The
squib batteries are special high-discharge-rate batteries which will maintain a
terminal voltage of 18 volts for one second under a 75 ampere load.
On spacecraft 6, three 400 ampere/hour, 16 cell, silver-zlnc batteries are
installed in the adapter battery module. These batteries are used in lleu of
fuel cell batteries. _11 of the silver-zlnc batteries have an open circuit
terminal voltage of 28.8 to 29.9 volts DC.
The main and squib battery cases are made of titanium. The approximate activated
(wet) weight for each squib battery is 8 lbs and each main battery 17 lbs. The
adapter module battery cases (spacecraft 6) are constructed of magnesium and the
approximate wet weight of each battery is 118 lbs.
The battery electrolyte consists of a 40 percent solution of reagent grade potas-
sium hydroxide and distilled water. The main and squib batteries have a vent
valve in each cell designed to prevent electrolyte loss and will vent the cell to
atmospheric pressure in the event a pressure in excess of 40 psig builds up within
the cell.
All of the silver-zlnc batteries are equipped with relief valves which maintain a
tolerable interior to exterior differential pressure in the battery cases. The
batteries are capable of operating in any attitude in a weightless state. Prior
5-22
____. SEDR 300
PROJECT GEMINI
to installation into the spacecraft, the batteries are activated and sealed at sea
level pressure. All of the batteries are coldplate mounted to control battery
temperature.
P_4_R SYS_M RELAY PANEL
The power system relay panel contains relays necessary for controlling and sequenc-
ing power system functions. The panel contains the control relays for the fuel cell
battery system and RSS, main battery power sequence light relay, TR- 5 relay and the
squib bus arming relays.
POWER DISTRIBUTION RELAY PANEL /
The power distribution relay panel contains the relays required for arming the
retrograde squib buses in the event of an abort. These relays are controlled by
the Sequential System.
ADAPTER POWER SUPPLY RELAY PANEL
The adapter power supply relay panel contains relays necessary for control11 ng
adapter module power to the main power bus. The relay panel contains the stack
power relays which connect the individual fuel cell stacks tothe main bus. (On
spacecraft 6 the stack power relays connect the adapter module batteries to the
main bus. ) The panel also contains diodes used for reverse current protection
between the adapter power supply and the spacecraft main power bus.
AMMEteRS
The main bus section _mmeter (spacecraft 5 and 6) is a dual-edge-readout vertical
reading meter having a 0-50 ampere range with a total accuracy of two percent.
/4 The NO. i scale displays main batteries i and 2 and 50 percent of the adapter
power supply current. The NO. 2 scale displays main batteries 3 and 4 _-d 50
5-23
k sEo3oo,. PROJECT GEMINI
percent of the adapter power supply current. The ammeter is sh_t co_ected
between the main power bus and spacecraft ground.
The fuel cell stack _eter (used as a battery ammeter on spacecraft 6) with
associated selector switch, provides a display of individual main battery test
current with the selector in BT position and a particular MAIN BATIERIES switch
in _EST position. With the selector switch in IA, IB, IC or 2A, 2B, 2C positions,
the ammeter displays 50 percent of the applicable adapter module battery current.
The meter has a 0-20 ampere scale.
VOLTME_IER
On spacecraft 6 the voltmeter, used in conjunction with a selector switch, dis-
plays main bus, common control bus and squib bus voltage. Individual main battery
voltage may be monitored with the voltmeter selector switch set to BT position and
a particular MAIN BATTERIES switch set to _EST position. The voltmeter displays
applicable adapter module batteries (A, B and C) voltage when the selector switch
is set to 1A, IB, 1C or 2A, 2B, 2C positions. The voltmeter has a 0-50 vdc range.
POWER SYSteM MONITOR
The power system monitor (not applicable on spacecraft 6) consists of five vertical
reading indicators; a delta pressure indicator, three dual-readout smmeters and an
ac/dc voltmeter. The delta pressure indicator (not operative on spacecraft 5) and
voltmeter are used in conjunction with selector switches located just below on the
instrument panel.
The _mmeters provide a display of individual fuel cell stack (1A through 2C)
current (reading must be multiplied by 0.8 on spacecraft 5). The voltmeter, with ....
the selector switch in appropriate position, displays individual fuel cell stack
5-24
SEDR300
voltage, main bus, squib bus and common control bus voltage, individual main
battery voltage (with a particular MAIN BATTERIES switch in _EST position) and IGS
inverter output voltage with the selector switch in AC position. The AC position
on the selector switch Is inoperative on spacecraft 5. The voltmeter has a 20-33
ac volt range and an 18-33 dc volt range.
The delta pressure indicator has a 0-1.5 psi range with the selector switch in \
either H2 position and a 0-6 psi range with the selector switch in either H20
position. This indicator displays 02 versus H2 differential pressure and 02
versus H20 differential pressure for each fuel cell battery section.
FUEL CELL BAT_"RTR.S
Construction
The fuel cell batteries used in the Gemini Spacecraft are of the solid ion-
exchange membrane type using hydrogen (H2) for fuel and oxygen (02) for an oxi-
dizer. The fuel cell battery system is comprised of two separate sections which
are sealed in air tight pressure containers. Each section is made up of three
interconnected fuel cell stacks with plumbing for transferring hydrogen, oxygen
and product water. (See Figure 5-10).
Each fuel cell stack consists of 32 individual fuel cells. Each basic fuel cell
is made up of two catalytic electrodes separated by a solid type electrolyte in
laminated form. (See Figure 5-11).
The electrolyte is composed of a sulfonated styrene polymer (plastic) approximately
0.i0 inches thick. Thin films of platinum catalyst, applied to both sides of the
_ electrolyte, act as electrodes and support ionization of hydrogen on the anode
side of the cell and oxidation on the cathode side of the cell.
5-25
Figure 5-10 Fuel Cell Battery Section
5-26
_. SEDR300
PROJECT GEMINI
/_
A thin titanium screen, imbedded into the platinum catalytic electrode, reduces
the internal resistance along the current flow path from the electrode to the
current collector and adds strength to the solid electrolyte.
On the hydrogen side of the fuel cell, a current collector is attached by means of
a glass-cloth-reinforced epoxy frame which assures a tight seal around the edges
of the cell, forming a closed chamber. Ribs in the collector are in contact with
the catalytic electrode on the fuel cell, providing a path for current flow.
The hydrogen fuel is admitted through an inlet tube in the frame of the current
collector and enters each gas channel between the collector ribs by way of a
series of slots in the tube. Another tube provides a purge outlet, making it
possible to flush accumulated inert gases from the cell. The collector plate
is made of approximately 0.003 inch thick titanium.
On the oxygen side of the cell, a current collector of the same configuration and
material as the hydrogen side collector is attached. Its ribs, located at right
angles to those of the other collector, provide structural support to the electro-
lyte-electrode structure.
A Dacron cloth wick, attached between the ribs, carries away the product water
through capillary action, by way of a termination bar on one side of the assembly.
Oxygen is admitted freely to this side of the fuel cell from the oxygen filled
area of the section container.
The cell cooling system consists of two separate tubes bonded in the cavity formed
by the construction of the oxygen side current collector and the back side of the
hydrogen current collector. Each tube passes through six of the collector ribs
and has the cooling capacity to maintain operating temperature. The cooling of
5-27
SEDR 300
the oxygen current collector, which holds the product water transport wicks
provides the coldplate for water condensation from the warmer oxygen electrode.
The individual fUel cell assemblies are arranged in series to form a stack as
shown in Figure 5-12. When assembling the cells into a stack, the ribs of the
oxygen side current collector contact the solid electrolyte of the fuel cell
assembly. Titanium terminal plates are installed on the ends of the two outside
cells to which connections are made for the external circuit. End plates, which
are honey-comb structures of epoxy-glass %_minate 0.5 inch thick, are installed
on the outside of the terminal plates.
Stainless steel insulated tie rods hold the stack together and maintain a compres-
sion load across the area of each cell assembly. This assures proper contact of
the solid electrolyte with the ribs of each current collector. The fuel cell
stacks are packaged in a pressure tight container, together with the necessary
reactant and coolant ducts and manifolds, water separator for each stack, and
required electrical power and instrumentation wiring.
The hydrogen inlet line, hydrogen purge llne, and the two coolant lines for each
cell lead from their respective common manifolds running the length of the stack.
The manifolds are made of an insulating plastic material and the individual cell
connections are potted in place after assembly to provide a leak-tight seal. The
oxygen sides of the cells are open to the oxygen environment surrounding the fuel
cell assemblies within the container.
An accessory pad is mounted on the outside of the fuel cell section container. It
includes the gas inlet and outlet fittings, purge and shutoff valves, water valve ...._
and electrical connectors. Structurally, the container is a titanium pressure
5-28
/-_ SEDR 300% -
MEMBRANE H2 ELECTROOE
(ELECTROLYTE) -- ;_[
0 2 ELECTRODE H2 FEED TUBE
FRAME
0 2 CURRENT' H2 CURRENT
COLLECTOR _ COLII!CTOR
(
cOOLANT,. _TUBES
¢PROD. H2O
REMOVA_/_WICKS
GETUBE
Figure 5-11 Basic Fuel Cell Assembly
OOLANT OUT
/_ - COOLANT IN
MANIFOLD INLET
0 2
CELL Wl,
I TERMINAL PLATE
•H2 FEED TUBES
EXCHANGE MEMBib_NL/ELECTRODE ASSEMBLY
JTACK TIE ROD
(-) TERMINAL PLATE
F _WATER SEPARATOR BASIN
HYDROGEN PURGE MANIFOLD
Figure 5-12 Fuel Cell Stack Assembly
5-29
SEDR 300
PROJECT GEMINI
vessel consisting of a central cylinder with two end covers and two mounting
brackets. Within the container_ the fuel cell stacks are mounted on fiberglass-
impregnated epoxy rails by bolts which pass through the stack plates. These rails
are in turn bolted to the mounting rings sandwiched between the two flanges on the
section container.
The hydrogen _nlfolds on each stack within a section are parallel fed with a
hydrogen shutoff valve and check valve in the feed line to each stack. Oxygen
is fed into the section container so that the entire free volume of the container
contains oxygen at approxlmately 22.5 psia. The coolant reaches the fuel cell
battery sections by two separate isolated lines. Any malfunction in the coolant
line in one section will not affect the cooling function of the coolant line in
the other section. "_'_,
Each stack in the section has its own water-oxygen separators which are manifolded
into a single line coming out of the section container. All hydrogen, oxygen
coolant, electrical and water storage pressure line connections at the section
container are fastened to standard bulkhead fittings on the accessory pad.
After the stacks are completely assembled within the container, all void spaces
are filled with unicellular foam. The purpose of this foaming is for vibration
dampening, accoustical noise deadening and minimizing free gas volume to prevent
possible fire propagation. Thin plastic covers are placed over the top and bottom
of each stack to manifold oxygen to the stack and to keep the foam material from
entering areas around the coo1_nt manifolds and oxygen water separator.
.Operation
The basic principle by which the fuel cell operates to produce electrical energy
5-30
SEDR300 _.__
PROd EC'T GEMINI
and water, is the controlled oxidation of hy&rogen. This is accomplished through
the use of the solid electrolyte ion-exchange membrane. On the hydrogen side of
the fuel cell, hydrogen gas disassociates on the cata]_vtic electrode to provide
hydrogen ions and electrons. The electrons are provided a conducting path of low
resistance by the current collector, either to an external load or to the next
series-connected fuel cell.
When a flow of electrons is allowed to do work and move to the oxygen side of the
fuel cell, the reaction will proceed. By use and replacement, hydrogen ions flow
through the solid electrolyte to the catalytic electrode on the oxygen side of
the fuel cell. When electrons are available on this surface, oxygen disassociates
and combines with the available hydrogen ions to fOrm water. (See Figure 5-13).
The co_gen current collector provides the means of distributing electrons and
condensing the product water on a surface to be transported away by the wick
system through capillary action. The individual cell wicks are integrated into
one large wick which routes the water to an absorbent material that separates
the water from the gas.
]By using the oxygen outlet pressure as a reference, a small pressure differential
is obtained over the length of the water removal system. This pressure is suffi-
cient to push the gas-free water toward the storage reservoir.
Waste heat, generated during the fuel cell battery ogeration, is dissipated by
means of the recirculating coolant provided by the spacecraft cooling system.
In addition, the total coolant flow provides the function of preheating the
_--" incoming reactant gases. In the spacecraft, the reactant gases are suFplied to the
fuel cell battery sections by the RSS. This system contains the reactant supply
5-31
sEoR300__..... PROJECT GEMINI
ELECTRODES
-_, ,IF
4_ + 4H++ O2._,,-- 2H202H2_ 4H+ ÷ 4_ OVERALL
2H2 + O2 _2H20
CHEMICAL REACTIONS
Figure 5-13 Principal of Operation.
5-32
,_ S|DR300
PMIN!
tanks, control valves, heat exchangers2 temperature sensors and heaters required
for management of the fuel cell reactants. (See Figure 5-14).
REACTANT SUPPLY SYSTEM
The RSS is essentially a subsystem for the fuel cell battery sections. The system
provides storage for the cryogenic hydrogen and oxygen, converts the reactants to
gaseous form and controls the flow of the gases to the fuel cell battery sections.
The RSS components are installed in the RSS/fUel cell module. On spacecraft i0,
ii and 12, the RSS 02 requirements are supplied from a central cryogenic 02 tank
located in the RSS/fuel cell module. This vessel also supplies 02 to the ECS.
See Figure 5-6, 5-7 and 5-8 for component installation and Figure 5-14 for a
functional diagram.
Components
Reactant Supply Tanks
Two tanks are utilized to separately contain the cryogenic hydrogen and oxygen
required for the operation of the fuel cell battery sections. The tanks are therm-
ally insulated to minimize heat conduction to the stored elements which would cause
the homogeneous solution to revert to a mixture of gas and liquid. The tanks are
capable of maintaining the stored liquids at supercritical pressures and cryogenic
temperatures.
The approximate total amount of liquid stored in the hydrogen vessel is 22.25 ibs
for long mission configurations and 5.80 ibs for short mission configurations.
The approximate total amount of liquid stored in the oxygen vessel is 180 ibs on
spacecraft 5, 106 Ibs on spacecraft i0, ii and 12 (for ECS and RSS) and 46.0 Ibs
_ on spacecraft8 and 9.
5-33
SEO3o0PROJECT GEMINI
J LEGEND r- CRYO PRESSURE]
_-,_ HYDROGEN J AND QUANTITY J•:.:.:-:<.:.:.:.:-:-:-:-:.OXYGEN J INDICATOR
mJ_u_o:x-,_g_VENT l TO T/M
VENT VALVE _ o21ooPP I
.4 I,_ ITCH SENSOR _ --"
' AUTO _ r PRESSU'R_AND__ HEATER J QUANTITYI ; . SWITCH
_ I ,ND_CATOR
ON _ _TEMP. rlYDROGEN Ic_NTORLI_ _
• _ A°G.E.
HEA.TER ECS
SECTION 1POWER SWITCH
!P,LLE TOECS, [_OPP
RESERVOIR TEMP. SENSOR
I • TO A.G.E.
VALV __.p, _HEAT EXCHANGER
CHECK VALVESLATCH-TYPE
HIGH PRESSURE: ml. sEc.I COOL COOLI
LOOP LOOP
H 2 STANDBY FROM FUEL
VENT _ CELL SECTIONS
TOSUPPLY ' CROSSOVER
FROM FUEL VALVE
CELL SECTIONS LATCH-. /_ • 0 2 HIGH TYPE
P'RI, S_C_ PRESSURECOOL COOL RELIEF VALVE
LOOPv LOOPREF ECS / NOT APPLICABLE
S/C 10, 11 & 12 TO S/C 10, II& 12FILL VALV_ _ CHECK VALVES
HEAT EXCHANGER
LATCH-TYPESHUTOFF
' TEMP. _ _ TE,.,P._NSOR VALVE 1J OXYGEN TO ECS TO A.G .E.
CONTAINER RESERVOIR
TO TELEM., _N
& A.G .E. -- _O OFF ,.LHEATER
02 SECTION 2QUANTITY POWER SWITCH
HEATERSWITCH
I oN tVENT VALVE AUTO
-- TOT/M
IFigure 5-14 RSS/Fuel Cell System Functional Diagram (S/C 5 & 8 thru 12) (Sheet 1 of 2)
5-34
- SEDR300
_._ H20 OUTLET PRESSUREREFERENCE
STACK CONTROL SWITCHES IAz IB & ICA
DIFFERENTIAL
H2 STACK
(LATCH-TYPE) OVERBOARDVENT
H20
SECTION 1 SHUTOFFVALVE PRESSURE(LATCH-TYPE) REGULATOR
WITH
COOLANT IN FILTERS _-
PRODUCT L'_ ....
H20 BLEED
02 PURGE _
(NC)
_x:_xxx xxxxxxxxxx × x:ooooo_
DIFFERENTIAL
XOVERSWITCH
STACK CONTROL SWITCHES 2A s 2B & 2C0 42 LBS TO
ON /_ \ DRINK CABIN
"TTT_TI TT TT ._o
H2 STACK SHUTOFFVALVES (LATCH-TYPE)
H20SHUTOFF
t
CONTROL
H2O OUTLET PRESSURE REFERENCE
Figure 5-14 RSS/Fuel Cell System Functional Diagram (S/C 5 & 8 thru 12) (Sheet 2 of 2)
5-35
SEDR 300 _._GEMINI
The hydrogen vessel is composed of titanium alloy and the oxygen vessel Is made
of a high strength nickel base alloy. Both vessels are spherical in shape and
double w_led. A vacuum space between the inner and outer wall (approximately one
inch) provides thermal insulation from ambient heat conduction. The inner wsl!
is supported in relation to the outer wall by an insulating material supplemented
by compression loading devices.
Each storage tank contains a fluid q,_ntlty sensor, a pressure sensor, a tempera-
ture sensor and an electrical heater (the 02 tank on spacecraft 1% ii and 12 has
two heaters) installed in the inner vessel in intimate contact with the stored
reactants. The fluid quantity sensor is an integral capacitance unit which oper-
ates in conjunction with an indicator control unit containing a null bridge
amplifier.
The sensor varies the capacitance (in proportion to fluid level) in a circuit
connected to the null bridge amplifier. The amplified signal is then used to
drive a servo motor, which in turn operates a visual indicator for quantity
indication. Power inverters supply 400 cycle, 26 vac power to the fluid quantity
circuits.
The temperature sensor is a platinum resistance device capable of transmitting a
source signal to a balanced bridge circuit. The sensor provides cryogenic fluid
temperature monitoring for telemetry and AGE.
The pressure sensor is a dual resistive element, diaphragm type transducer. The
sensor provides signals for cryogenic fluid pressure monitoring on a spacecraft
meter. The electrical heaters provide a method of accelerating pressure build-up
in the reactant supply tanks. The heaters may be operated either in a manual or
5-Sg
__ SEDR300
PROJEC--'T GEMINI
automatic mode. In the automatic mode a pressure switch removes power from the
heater element when the tank pressure builds up to a ncmir_1 900 psi8 in the
oxygen tank and a nominal 250 psig in the hydrogen tank. In the man,-_l mode a
spacecraft pressure meter indicates proper switch operation.
Fill and Vent Valves
The fill and vent valves provide a dual function in permitting simultaneous fill
and vent operations. Quick disconnect fittings are provided for rapid ground
service connection to both the storage t-n_ fill check valve and the vent check
valve. When fill connections are made, the pressure of the ground se_vlee connec-
tion against the fill and vent valve poppet shaft simultaneously opens both the
fill and vent ports. When ground service equipment is removed, the valve poppet
automatically returns to its normally spring-loaded-closed position. The vent
check valve is a single poppet type, spring-loaded-closed valve which opens when
system pressure exceeds 20 psig to relieve through the fill and vent valve vent
ports.
Heat Exchangers
The supply temperature control heat exchangers are finned heat exchangers in which
the supply fluid temperature is automatically controlled by absorbing heat from the
recirculating coolant fluid of the spacecraft cooling system. The special double-
pass design precludes freezing of the coolant and assures a reactant fluid supply
at 50°F minimum and 140 ° m_ximum.
Dual Pressure Regulator and Relief Valves
f_ The dual pressure regulator and relief valves are normally open poppet type regu-
lators which control downstream pressure to the fUel cell battery sections. The
5-37
___ SEDR 300 ...__ .__
PROJECT GEMINI
regulators maintain the hydrogen pressure at approximately 21.7 psia and the
oxygen pressure at approximately 2.2 psia. The oxygen side of the regulators
is referenced to hydrogen pressure. The hydrogen side of the regulators is
referenced to produce H20 pressure.
The relief valves provide overpressurization protection for the regulated pressure
to the fuel cell battery sections. This valve is precalibrated to operate at a
pressure of sppro_mately i0 psia above the normal supply level.
High Pressure Relief Valves
The high pressure relief valves are single poppet type, spring-loaded-closed
valves which provide system overpressurization protection. The valves vent
system gas to ambient when pressure exceeds the system limlts.
Solenoid Shutoff Valves
The solenoid shutoff valves are solenoid operated latching type valves which
eliminate fluid loss during the nonoperating standby periods. The valves are
normally open and are closed only during fill and standby periods by applying
power to the solenoids.
Crossover Valve
The crossover valve is a solenoid operated latching type valve which provides the
capability of selecting both dual pressure regulators to supply hydrogen and
oxygen to a fuel cell battery section for the purpose of increasing flow rate
for more effective purging. The crossover valve is contro_ed by the X-OVER
switch on the right instrument panel.
02 Crossfeed Reactant Valve (spacecraft 8 and 9) -_
The 02 cressfeed reactant valve is a solenoid operated, latching type valve Which
SEDR300
MINI
provides the capability of pressurizing the RSS with 02 from the ECS oxygen
supply. This provides a redundant method of supplying the proper reactant 02
pressure to the fuel cell battery sections in the event of a malfunction in the
RSS oxygen supply. The crossfeed valve is controlled by the 02 CROSS-FEED switch
located on the center instrument panel.
Operation
During pre-launch, the two separate reactant supply tanks are serviced (using AGE
equipment) with liquid hydrogen and oxygen. After the tanks are filled, in order
to accelerate pressure buildup within the tanks, the internal tank heaters are
operated, utilizing external electrical power. In approximately one hour the
liquid is converted into a high density, homogeneous fluid at a constant pressure.
During the fill operation, the solenoid shutoff valves between the storage tanks
and the dual pressure regulators are closed. Once operating pressure is obtained,
the solenoid shutoff valves may be opened by applying power to the coil of the
valves. The high density, homogeneous fluid will then flow upon demand.
The fluid flows from the supply tanks to the heat exchangers. The fluid tempera-
ture, when entering the heat exchangers is approximately -279°F for the oxygen
and approximately -423°F for the hydrogen. The heat exchangers absorb heat from
the recirculating coolant fluid of the spacecraft cooling system. This heat_
applied to the high density fluid, raises the temperature of the reactants to
approximately 50°F to 140°F.
The reactants, now in gaseous form, flow through the heat exchangers, past the
/_ high pressure relief valves and AGE temperature sensors, to the supply solenoid
shutoff valves. During fuel cell battery operation, if the demand on the fluid
_-B9
SEDR300 _--7PRO--O--j-EC' GEMINI
flow is inadequate to keep tank pressures within l_m_ts, the high pressure relief
valves will vent, externally, the excess fluid. The AGE temperature sensors on
the heat exchangers are used for pre-launch checkout only.
The reactants flow through the supply solenoid shutoff valves to the dual
pressure regulator and relief valves. The dual pressure regulators reduce the
pressure of the reactants to approximately 21.7 psia for the hydrogen and approx-
imately 20.5 psia for the oxygen. The gas now flows through the manual shutoff
valves and is then directed to the fuel cell battery sections at a flow rate that
is determined by both the electrical load applied and the frequency of purging.
The flow rate of the gases may be increased for more effective purging by opening
the crossover valve.
After launch, the supply tank heaters are operated by spacecraft power. The
heaters operate as required to maintain proper system pressures.
5-_o
ENVIRONMENTALCONTROL SYSTEM
TITLE PAGEIll
SYSTEM DESCRIPTION ................................ 6-3 Vl
OXYGEN SUPPLY SYSTEM ............................. 6-3
CABIN LOOP. ............................................... 6-8SUIT LOOP. .................................................. 6-11
WATER MANAGEMENT SYSTEM ..................... 6-18SYSTEM DISPLAYS AND CONTROLS................. 6-26
SYSTEM OPERATIONS ................................ 6-32SERVICE AND CHECKOUT ............................ 6-33
PRELA UN CH. 6- 33 _-_:_._............................................... _-:_._--.._LAUNCH ..................................................... 6 ....... ._---
0 RBIT......................................................... 6- 36 ::iiiii_i:'_".".-_-".._"=_"_
RE-ENTRY.................................................... 6-42 iii_!ii_iiii._ii_:-_-._::
P0 STLAN DIN G ............................................ 6-43 iiiiiiiiiiiiiiiiiiiii_i!;'_i.°.°°oo°ooooo°o°o°.**°°.o°.....o.°.°ooo°.,°°°°**°°°.,_
EMERG ENCY. ............................................... 6-44 ii_ii_!iiiiiiiiiiiiiiiii!i!
SYSTEM UNITS ............................................ 6-45 --..!----_..---_--:...-.:_--.:.:!_DUAL SECONDARY RATE AND SUIT iiiii_iiiiii!!i!!!iii_i_ii
SYSTEM SHUTOFF VALVE ............................ 6-45 i_i_iiiiiiii_i_i!!!_!!!!i--°*.o.....o....o.H.oo..._.°..°*....*......o....o..°.
SUIT OXYGEN DEMAND REGULATOR ............. 6-48 i!i!iiiiiiiiiiH_iiii!!!i!i................ o.°.,o.°o.°:::::::::::::::::::::::::::
CABIN PRESSURERELIEFVALVE .......... . ........... 6-50 _-'_----_!-_-_-.:--_.:_!'..::::::::::::::::::::::::::::::::::::::::::::::::::::::
SUIT CIRCUIT COMPRESSOR ........................... 6-53 :-..---_--_!_---.:---_--.:---!:::::::::::::::::::::::::::
SOLIDS TRAP............................................... 6-53 iiiiiiiiiiiiiiiii!_iiiiiiil.°.*.°°.°°...°°°...oo..°.H:::::::::::::::::::::::::::
_ DUAL CABIN PRESSUREREGULATOR ............... 6-53 iiiiiii_iii_iiiiiiiiiiiiiiiPRIMARY SUPERCRITICAL OXYGEN .................. ..°....................... ..........
........................ ...
CONTAINER .............................................. 6-57 iiiiiiiiiiiiiii_iiiiiiiiiii
SECONDARY OXYGEN CONTAINER ................ 6-57 ::iiiiii!iii_iiiiiiiiiiiiiii6-1 i'_'""_.-"_'":!'_":"_"_
PROJECT GEMINI
Figure 6-1 Environmental Control System
6-2
__. SEDR300
P ROJ ECE-CT'-'G-E-M I N I
SECTIONVZ E_VmOm_rAL COmmOT SYS_
S_S_m_ D_scan_IoN
The Environmental Control System (ECS) (Figure 6-1, 6-2) may be defined as a
system which provides a safe and comfortable gaseous atmosphere for the pilots.
The system must perform such tasks as providing fresh oxygen, pressurization,
temperature control, water removal and toxic gas removal. In addition to
providing atmospheric control for the pilots, the system provides equipment
cooling and regulated temperatures for certain pieces of equipment.
For ease of understanding, the Environmental Control System may be separated into
four systems or loops which operate somewhat independent of each other. These
loops are:
(i) The oxygen supply system.
(2) The cabin loop.
(S) Thesuitloop.
(_) The water management system.
OXYGEN SUPPLY SYST_
There are three oxygen systems: Primary, Secondary and Egress.
Primary Oxygen (Figure 6-3, 6-_)
This system stores and dispenses oxygen for breathing and for suit and cabin
pressurizatlon.
6-S
_--. SEDR 300
6-4
__ $EDR300 ._._
PROJECT GEMINI
This system provides oxygen during the period commencing two hours prior to
launch and terminating with jettison of the adapter section at retrograde.
The primary oxygen supply is stored at supercrltical pressure in a cryogenic
spherical container in the adapter section of the spacecraft. This container is
filled with liquid oxygen at atmospheric pressure. Heat is supplied by thermal
leakage through the container insulation and by activation of an electric heater
in order to build pressure to the critical point of 736 psia. Above this point
liquid oxygen becomes a homogeneous mixture, described for simplicity as a dense
supercritical fluid. This fluid is warmed, regulated and filtered before it
enters the suit or cabin loop.
On spacecraft L0 through 12, a long mission ECS oxygen tank will replace the short
mission Reactant Supply System (RSS) oxygen tank and the ECS oxygen tank, allowing
ECS breathing and the RSS oxygen to be stored in the same container. A tee fitting
on the cryogenic line allows both the ECS and RSS systems to receive oxygen from
the cc_i_on container.
The primary loop consists of the following components: primary oxygen container,
pressure control switch, pressure transducer, fill and vent valves, temperature
discharge sensor, pressure relief valve, pressure regulator, shutoff valve, filter,
check valves, and heat exchanger.
Secondary Oxygen (Figure 6-3, 6-4)
The secondary oxygen system is capable of performing the same functions as the
primary oxygen system and operates when pressure in the primary system falls
/_ below 75 + i0 psi. At retrograde, when the primary oxygen container is Jettisoned
6-5
SEDR 300 _]
_____ PROJECT GEMINI
Figure 6-3 Primary and Secondary Oxygen System
6-6
6-7
_. $EDR300 __
PROJECT GEMINI
with the equipment adapter, the secondary oxygen system ass_nes the duties of
the primary oxygen system.
The gaseous secondary oxygen supply is stored in two cylinders located in the
re-entry module. Each tank contains 6.5 pounds of usable oxygen pressurized to
5000 psig maxim_n at 70°F. This oxygen supply is then regulated before it enters
the suit or cabin loop.
The secondary system consists of two: tanks, fill valves, transducers, pressure
regulators, shutoff valves and check valves.
Egress System (Figure 6-5)
This system provides each pilot with oxygen for breathing and for suit pressuriza-
tion in the event that they initiate ejection procedures at 7%000 feet or below,
during launch or re-entry. The egress oxygen is provided on spacecraft 5 and 6
only.
The egress gaseous oxygen supply is stored in a tank located in each seat-mounted
egress kit. Each tank contains 0.31 pound of usable pressurized oxygen.
Each egress system consists of a tank, pressure regulator, pressure gage,
restrictor, check valve, shutoff valve, and composite disconnects.
CABIN LOOP (Figure 6-6)
Design cabin leakage at ground test conditions is 670 standard cubic centimeters
per minute, (scc/min) of nitrogen at 5.0 psig. Makeup oxygen, to maintain
cabin pressure at nominal 5.1 psia level, is called for by the cabin pressure ......
regulator. In order to obtain maximum utilization of oxygen, it first passes
6-8
./ _ SEDR300
f_
SUIT(
Figure 6-5 Egress Oxygen System (S/C 5 & 6)
6-9
SEOR3oo_'_ PROJECT GEMINI
M.AIN BUS TEMPERATURE[m_ PRESSURE
iNDICATOR l_J] INDICATOR
EQUIPMENTDISCONNECT
CABIN FAN CABIN _
CIRCUIT FAN _ TO _---
----- BREAKE_R OH__ INSTR,
COMMON _ C_CONTROL e
R CABIN AIRTEMPERATURE
w TRANSDUCER
TO 0 2 HI• RATEVALVE I
0 2 HI RATECIRCUIT BREAKER
( S/C $&6 - 8 & UP COOLANT
CABINcoNTROL_i 1" I !_
TEMP = : --
A BVALVE
_LI_I_ CABIN _.
PRESSURETRANSDUCER
CABIN FAN J
J CO 2 PRESSURE
TO AMPLIFIERJ_ INSTR.
CABIN HEAT 'EXCHANGER CABIN PRESSURE
REGULATOR
CABIN FAN FROM
PWR SUPPLY TO __ J OXYGENCABIN SUPPLY
CABIN
(S/C 5&6 ) REPRESSURIZATJONVALVE
J CABINj TO SUIT RELIEF
CABIN ) CIRCUITTO SUIT FAN TO SUIT FAN/ NO. 1 CIRCUIT NO. 2 CIRCUIT
_ CAB,N , r c o )CABIN AIRCIRCULATING VALVE -- -- --VALVE "q_
CABIN
OUTFLOW_
VALVE _
.... r _ .!SMALL PRESSURE
SNORKEl, _ BULKHEADINLET _ WATER SHUTOFFVALVE VALVE
Figure 6-6 Cabin Environmental Control
6-10
__ SEDR300 _____
PROJECT GEMINIf---
through the suit loop before it is dumped into the cabin through the suit
pressure relief valves.
primary cabin components for spacecraft 5 and 6 are a cabin heat exchanger and a
fan. These parts have been removed from spacecraft 8 through 12. This loop also
contains a relief valve for both positive and negative pressure relief, a pres-
sure regulator and manual valves to either due_ cabin pressure or repressurize.
In the latter operation, oxygen is supplied directly to the cabin.
LOOP(Fimxe 6-?, 6-8, 6-9)
The pilots are provided with redundant atmospheres by having a closed pressure
suit circuit within the pressurized cabin. This suit circuit provides for
cooling, pressurization, purification and water removal.
The suit loop is a closed system with two pressure suits operating in parallel.
Circulation of oxygen through the suit is provided by a centrifugal compressor.
Carbon dioxide and odors are removed by an absorber bed containing lithium
hydroxide and activated charcoal. The gases are cooled in a heat exchanger by
a liquid coolant, Monsanto MCS 198, to a temperature below the dew point. Water
condensing within the heat exchanger is dumped overboard or routed to the water
evaporator. The reconditioned Oxygen is mixed with fresh makeup oxygen.
The suit circuit has two modes of operation, the nox_al recirculation Node which
was discussed in the previous paragraph _-d the high rate mode which shuts off
the recirculation system and _umps oxygen directly into the suit.
6-11
...-__ -. SEDR 300
;_ PROJECT GEMINI
/ i
,.!.._" _._:_,; :ii
PEDESTAL I : i
i,i !
CENTER CONSOLE
\
',\
_'_\\. SMALL PRESSURE
. _ULKHEAD _E_
SUIT TEMPERATURE VALVE (2 REQD)(MECHANICALLY LINKEDTO CONTROL HANDLES "_ /ON PEDESTAL)
SOLID /(2 REQD) / ,"
/I"
i ,/]
/
_"_X_ /SECONDARY C /
PRESSUREREGULATOR(2 REQD)
x _ j
/,.# ,
\
TRANSDUCER DISCONNECT ..f ''''S'f
SUIT INLET (2 REQD)DISCONNECTDUCT (2 REQD)
SYSTEM SHUT-OFF VALVE
Figure 6-7 Suit Loop (ECS Package) (Sheet 1 of 2)
6-12
F__{\ SEDR 300 _]
PROJECT GEMINIF.
r---.-! ...... ]
PEDESTAL : :
iZ ZL:t............ .'_
CENTERCONSOLE
SMALL PRESSURE _
BULKHEAD (REF)1 _,_,
)NDENSATEWATER HOSE
TEE-SECONDARY
02 SYSTEM (2 ..............._.
OXYGEN TEST
i /i
z _'" //
//
HOSE
i //
FROM PRIMARY / _ /
0 2 SUPPLY _ ;' ;
//
CHECK /VALVE /
// /
COOLANT INLET ,/ ..'/
HOSE (SYSTEM NO. I) I "'"'_ CABIN
/02
\. / / REPRESSURIZATION\ i -'_ LINE
COOLANT OUTLET ..................... _ _/_"SECONDARy/,\ 02 ,/"
HOSE (SYSTEM NO. \ SUPPLY LiNE j/_"./
SECONDARY 0 2 fOXYGENHOSE ........_--_._._SUPPLY LINE SUCTION H20
OXYGEN HOSE VALVE
PRESSURE H20 (3 REQD)OXYGEN BULKHEAD
HOSE FITTING(SUIT CIRCUIT SUPPLY)
Figure 6-7 Suit Loop (ECS Package) (Sheet 2 of 2)
6-13
_-_ SEDR300 -_IFT_ _
l ! J PROJECT GEMINI
0 2 HIGH RATE _--_ SNORKEL _'_
CK SNORKE'[ SNORKEL
INLET
VALVE
02HIGHRATE
L A,_ER I PRESSURE
MANUAL I B_'K-H_D ' --
0 2 HIGH RATE J
_: _- _ -- --I -- INFLOW
II VALVE
"_ i L G I '
_ CHECK
VALVE
SUPPLYFROM OXYGENDUAL HIGH _p _X'_'_"_x_"_"v"_'_'_
FLOW FSUIT SYSTEM
SHUTOFF VALVE I
RIGHT
LO ABSOLUTE
NORMAL PRESSURELEFT ABSOLUTE NORMAL SWITCHPRESSURESWITCH "
CHECK FROMVALVE NO. I SUIT -- CHECK ._-_
PRESSURE PRESSUREDEMANDREGULATOR LEGULATOR
SOLIDS
TRAP SUIT CIRCUITAND PRESSUREBYPASS INDICATOR
SU,TC,RC0. POWERSUPPLYTEMPERATUREiNDICATOR
NO. 2 SUIT
I_ _ COMPRESSOR
CHECK
VALVE;UIT CIRCUIT
TEMPERATURE
___ SUIT CIRCUIT
TEMPERATURE;ENSOR SENSOR
CO,_,AND ODORNO. I SUIT FLOW RATE NO. 2 SUIT FLOW RATE ABS'_RBERAND SHUTOFF CONTROL AND SHUTOFF CONTROL
CHECK CHECK
VALVE VALVE
TO CABINPRESSURE _REGULATOR SUIT HEAT
TO WATER EXCHANGERDISCHARGE WATER
SEPARATOR
Figure 6-8 Suit Loop Recirculation Mode Schematic (S/C5, 6, 8 & Up) (Sheet 1 of 2)
6-14
_ ,;-:_. SEDR 300
-- PROJECT GEMINI
FAN POWER .9-- I DISCONNECT
CABIN KECIRC SUPPLY FQ,_.._,_j RELAY SUIT FAN NO. t
__ CIRCUIT BREAKER 24V DC
SUIT FAN MAIN BUSkl SWITCH
c,"_
J w;CABIN AIR J D,_M_ _ S/C S& 6 CABIN FAN
CIRCULATING I _ILI CABIN CIRCUIT _ SUIT FAVALVE I FAN
PURGE E,_ OFF
AI.VE _& i
•=_ 0 2 HIGH RATE
¢H
=_ SEQ LIGHTS
(SWITCHS)
COMMON
0 2 RATE CONTROLCONTROL _USCIRCUITBREAKER
_/C 8 & UP
TO 0 2 TO 0 2 CONTROLHI RATE CIRCUIT BREAKERVALVE
NO. I SUIT FANPOWER SUPPLY
NO. I SUIT
COMPRESSOR LEGEND
__ _ HIGH _TE 0 2
CONTAMINATED 0 2
CHECKVALVE
NOTE
RELAYS AND SWITCHES ARE SHOWNIN THE NORMAL POSITION.
t
Figure 6-8 Suit Loop Recirculation Mode Schematic (S/C5, 6, 8 & Up) (Sheet 2 of 2)
6-15
.f _. SEDR 300
L_!_ _ PROJECT GEMINI0 2 HIGH RATE I
SNORKELiNLETVALVE
MANUAL 0 2 I --I BULKHEAD 'HIGH RATE i
. i___ I VALVE/pc G )
_, ..
CHECKVALVE
SUPPLY
FLOW RATEAND
SHUTOFF VALVE
VALVEPRESSUREDEMANDREGULATOR REGULATOR
SOLIDSTRAP
AND AND
I I_ ,___' ( "_'NO. 2sU,tFAN
SUIT CIRCUIT __ POWER SUPPLY
TE_APERATURE_iM _F_ O
INDICATOR ..-:::.. ,,NO. 2 SUIT
MPRESSOR
CHECK
SUITCIRCUIT
TEMPERATURE ____
SENSOR SENSOR
_ _.. _ U_" F_W'P,_T_" ...... CO AND
AND SHUTOFF CONTROL ANO SHUTOFF CONTROL ABSORBER _-_ _,_
lCHECK
VALVE ,- LOO P_L VALVE C._K "W" _C®LANF__FRESSURE _REGULATOR
TO WAT_
DISCHARGE :__ EXCHANGERwATERSEPARATOR
Figure 6-9 Suit Loop High Rate Oxygen Mode Schematic (S/C 5, 6, 8 & Up) (Sheet 1 of 2)
6-16
.I: _-_. SEDR300
S/C 5AND 6
TO CABIN " EQUIPMENT
FAN POWERS._ _ I DISCONNECT SUITFAN NO. l
CABIN RECIRC SUPPLY F_ I RELAY CIRCUITBREAKER 24V DC
SWITCH MAIN BUS
I Vj l s/c 5AND6
CABIN AIR J ._ _ _,_ CABIN FAN
CIRCULATING J I"JC A-BINVALVE } __ FAN
PURGE E I _,VALVE OF
O2 HIGH RATE
SEQ LIGHTS(SWITCHS)
COMMON0 RATE CONTROLC_NTROL BUSCIRCUITBREAKi_
_CSAN0te
TRAOTEO_ITCOBRN'i'EA_ aOL
F_ NO. I SUITFAN
NO. 1SUIT 1COMPRESSOR
LEGEND
HIGH RATEO2
_C" _ CONTAMINATED0 2
1 '11-NOTE
RELAYSARESHOWN IN THELATCHEDAND ENERGIZEDPOSITION.
Figure 6-9 Suit Loop High Rate Oxygen Mode Schematic (S/C5, 6, 8 & Up) (Sheet 2 of 2)
6-17
__ SEDR300
PROJECT GEMINI
The suit loop consists of two suit pressure demand regulator valves, four check
valves, two throttle valves, two solid traps, a system shutoff and high flow
rate valve, two ccmpressors, one carbon dioxide and odor absorber, and a suit
heat exchanger.
WATER MANAG_NT SYSTEM (Figure 6-10)
The purpose of the water management system is to store and dispense drinking
water, collect and route unwanted water to the evaporator or dump overboard.
Drinking water is stored in a tank or tanks in the adapter. Each tank contains
a bladder and is pressurized to supply water to the transparant tank in the
re-entry module.
Spacecraft 5 utilizes two drinking water storage tanks which store both the
drinking water and the fuel cell by-product water. One tank uses a combination
of oxygen and fuel cell water as the pressurant while the other tank is
pressurized with oxygen.
Spacecraft 6 has only one storage tank and uses oxygen for the pressurant.
Spacecraft 8 through 12 utilizes two storage tanks. Fuel cell by-product water is
used as the pressurant for the drinking storage tank. The other tank is used to
store fuel cell water.
Urine and condensated water from the suit circuit heat exchanger is absorbed by
the wick in the water boiler or dumped overboard.
6 -18
sEo 3oo,:.o..,,:c:-,-G,:,v,,,,,,
ADAPTER WATER STORAGETANK "[_" (S/C 5 ONLY)(SEEDETAIL "C" SHEET5)-_ WATER MANAGEMENT PANEL WATER MANAGEMENT PANEL
(S/C 6, 8, AND UP) (S/C $ ONLY)
ADAPTER WATER STORAGE._f--_" TANKS (SEE SHEET 6 FOR
S/C 8 & 9) (SEE SHEET 7
FOR S/C 10& UP)
//ADAPTER WATER STORAGETANK "A" (S/C 5 ONLY)(SEEDETAIL "C" SHEET ._
iiiiiiiiiiiiiiiiiiiiiiiiii!iiiiiiiiii
•ADAPTER WATER STORAGE
TANK (S/C 6 ONLY)(SEE DETAIL "B" SHEET 3)
CABIN WATER STOP,AGE TANK
(SEE DETAIL "A" SHEET 2 FOR S/C(SEE DETAIL '_." SHEET 4 FOR S/C 6,8AND UP)
Figure 6-10Water Management System (Sheet 1 of 7)
6-19
___ SEDR 300 __
PROJECT GEMINI
COLL£CTOR
LIGHT
NTROLVALVE
PORT D
PORT A PORT E PORT F
"BELLOWS
ASSEMBLY
PORT C
PORT B
\ WATER CONTROL VALVE\ WATER TANK (AFT LOOKING FORWARD)
DETAIL "A" S/C 5 ONLY _-"',
Figure 6-10 Water Management System (Sheet 2 of 7)
6-20
._. _ SEDR 300 J__ _-_
PROJECT GEMINI
©©
?,
TANK
'_DETAIL B" S/C 6 ONLY
_$HUTOFF VALVE
PRESSURANTTANK
if "_,
Figure 6-10 Water Management System(Sheet 3of7)
6-21
.--_-_ SEDR 300
CHEMICAL VOLUMEURIN_
MEASURI NG SYST
I:'_T7
WATER DISPENSER_
.. \ 4J,4X_...
i : STORAGE
\\ DUMP SWITCH
l_r_><
\
SHUTOFF V_
DETAIL 'IA" S/C 6, 8, AND UP
Figure 6-10 Water Management System (Sheet 4 of 7)
6-22
__ SEDR 300 ___
PROJECT GEMINIf_
WAI
\\
\ ,,
\\
OAMS MODULE STRUCTURALASSEMBLY (REF)
DETAIL"C" (S/C 5 ONLY)
WATER TANKREGULATOR
/
Figure 6-10 Water Management System (Sheet 5 of 7)"
6-23
/ __'_. SEDR 300
WATER STORAGETANK
ADAPTER RETRO
_c .
EL CELL WATERSTORAGE TANK
t_;L_"-. SEDR 300
_" ___ PROJECT GEMINI
WATER STORAGETANK
ADAPTER RETROSECTION TANKS
WATER STORAGETANK
DETAIL"E" ($/C 10, 11& 12)
Figure 6-10 Water Management System (Sheet 7 of 7)6-25
SEDR 300 _._ .._PROJECT GEMINI
Con_onents of the water manasement system, in addition to the water tanks, are
a water control valve, condensate valve, water evaporator and two m-nual shutoff
valves.
Urine S_stem (Figure 6-11)
The urine system is designed for total management of the crewmens urinary output.
It samples and determines total volume of every urination and provides for
chemical ._lysis. Volume is determined by sampling every urination and using
a tracer dilution technique. Three tenths (0.3) ml of tracer chemical is added
to the urine and samples are taken. Durln8 postflight operations, the amount
that the tracer has been diluted in each sample determines the q-_ntity of urine
voided by the crewmen. The urine system consists of the following components;
a Chemical Urine Volume Measuring System (CUVI_) with selector valve, tracer
storage accm,_-tor and collection/mlxin8 bag, a urine receiver assembly with
collection bag, a urine qui_k-disconnect hose and urine solids trap filter, urine
sampling bags and a roll-on cuff receiver assembly.
SYS_ DISPLAYS AmD COF_OLS
The displays _.d controls for the Enviromnental Control System are provided in
the cabin and function as specified.
SECO_ O_ S_TOFF Handle
A tuRin,A1 secondary caqvgen shutoff handle is provided for each member of the flight
crew for complete and positive shutoff of each secondary oxygen container. The
handles are located aft of the right and left switch/circuit breaker panels.
The position (_EN or _ is noted.
r_(o 0_-=-._ SEOR3oo
NOTES
Figure 6-11 Urine System (Typical)
6-27
SEDR300
PROJECTGEMINI
O2 In'oH _ Telelight
The 02 HI RATE telelight located on the instr1_ent panel illuminates when the
high oxygen rate valve is opened manually or automatically.
The 02 HI RA_ switch on spacecraft 5 a,d 6 also activates the cabin fan. The
switch has three positions; CABIN FAN_ 02 HI _, and OFF. Spacecraft 8 through
12 do not have the cabin fan, and that position of the switch is not used.
C_ AIR RECIRC Handle
The cabin air recircnlation handle controls the recirculation valve which permits
entry of cabin air into the suit circuit for removal of odors and carbon dioxide.
This procedure will renovate cabin air without cabin decowpression and reduces _
the possibility of carbon dioxide pockets by increasing circulation of the cabin
atmosphere.
This handle will control the cabin air inlet valve which provides for ventilation
during landing and post!_nding pha_es of the mission.
CABIN VENT Handle
This handle controls the operation of the cabin outflow valve to permit emergency
decompression in orbit and cabin ventilation during the landing phase.
WA_R SEAL Handle
This handle provides for watertight closure of the cabin pressure relief valve
durir_ a water landing.
6 -28
__ SEDR 300
PROJECT GEMINI
0 2 HI RA'I_ RECOC_ Handlei i ii i
This handle provides for the manual return of the oxygen high rate valve to the
closed position, thereby restoring normal oxygen flow rate. Actuation of this
handle also reestablishes the capability of initiating high rate oxygen flow
when necessary.
_, A_d. SUIT _ _icator
A dual indicator provides for monitoring temperatures in the suit an& cabin
circuits. Range markings are calibrated in degrees Fahren_elt.
CABIN And P C02 PRESS Indicator
A dual indicator provides for monitoring cabin atmospheric pressure and the
amount of carbon dioxide at the suit inlet. Cabin atmospheric pressure is
calibrated in pounds per square inch. Carbon dioxide partial pressure is
calibrated in millimeters of mercury.
SEC 02 Indicator
A dual indicator is provide_ for monitoring pressure in the ind/vidual gaseous
oxygen containers in the secondary oxygen subsyst_n. The indicator range is
from O to 6000 psia t divided into 500-pound increnents and nmmbered at each
lO00-poun_ interval. Readings must be multiplied by i00 to obtain correct
valttes.
6-z9
__ SEDR300 _.___
PR-OJ E C-T GEMINI
ECS 02 QUANT % and PSIA Meter
This indicator provides for monitoring quantity and pressure of cryogenic oxygen
in the primary oxygen container. The quantity scale displays from 0 to 100
per cent in 2 per cent increments, numbered at 20 per cent intervals. The
pressure scale ranges from 0 to lO00 psia in 20-pound increments, numbered at
200-pound intervals. Red undermarkihgs are incorporated on the oxygen meter
to indicate the point at which thermal pressurization may be discontinued by
de-energizing the heaters. Spacecraft l0 through 12 reidentifies the indicator
as the CRYO meter, but it performs the same function as before.
Cryogenic Quantity Switch
This switch allows the same indicator to be used when monitoring the pressure and "-_.
quantity of cryogen in amy of the three cryogenic containers. The three contain-
ers are: the ECS primary oxygen supply, the RSS or fuel cell (FC) oxygen supply,
and the RSS or FC hydrogen supply. The switch is located below the indicator on
the center panel and has the following positions: ECS 02, FC 02, FC H2 (PX35)
and OFF. Spacecraft iO through 12 have a three position switch. The positions
are; 02, H2 (PX35) and OFF. The 02 position allows the indicator to monitor the
ECS-RSS oxygen supply and the H2 (PX35) position allows monitoring the RSS or FC
hydrogen supply.
ECS O2 HEATER Switch
This switch is connected to the heaters in the ECS primary oxygen container.
The switch has three positions; AUTO, OFF, and ON. It is located below the
flight plan display on the center panel. Spacecraft lO through 12 reidentifies
the ECS 02 HEATER switch as 02 HEATER switch and connects to heaters in the
6-30
___ SEDR300 _.._
PROJEC'T"-GEM IN I
ECS-RSS oxygen container.
SUIT FAN Switch
The switch has three positions; NO i, OFF, a_ NO 1 & 2. The switch is located
in the upper left hand corner of the center panel. This allows suit fans number
i end number 2 to operate together or independently. Suit fan NO 2 may be operated
by placing switch in NO i & 2 position _nd placing suit fan NO i circuit breaker
switch to OFF.
02 CROSS FEED Switch, ,, ,
The 02 CROSS FEED Switch, spacecraft 6, 8 and 9 when in the OPE_ position, per-
_. mits oxygen from the pri-_ry oxygen supply module for the ECS to be used in the
RSS in the event of RSS oxygen module failure. The reverse is also true.
Water Management Panel
A three knob panel is provided for managing, replenishing, and dumping waste water
and urine overboard.
SUIT and CABIN _ Controls
Dual concentric knobs are mounted between the ejection seats for suit and cabin
temperature control. These knobs control the operation of valves regulating
the flow rate of primary and secondary coolant through the suit and cabin heat
excb-ngers. Clockwise rotation results in increased temperatures.
MA_AL 0 2 NTGH RA't_ Handle
This b_n_Lle is located on the console between the members of the flight crew
and provides for manual control of the dual high oxygen rate and suit system
6-31
PROJECT GEMINI
shutoff valve. Actuation of the handle shall initiate the oxygen high flow
rate and de-energize the suit compressor. Resue_tion of normal system operation
shall be effected by actuation of the oxygen high rate recock handle.
SUIT FLOW Control Levers
An individual lever is provided for each member of the flight crew for regulation
of circulatory oxygen flow through the suit circuits. The levers are located on
the lower section of the pedestal and shall provide any selected flow valve
setting from fully open to fully closed. A detent provides an intermediate
position to prevent inadvertent shutoff of suit flow. This detent may be by-
passed by moving the lever outboard.
Cabin REPRESS Control
A rotary handle control is provided for cabin repressurization after a decompres-
sion has occurred and for ELSS oxygen supply. The control rotates approximately
90° between fully OPEN (repressurize) and fully CLOSED (off) positions. This
control is located on the center console between the suit flow control panels.
ECS RTR Telelight
This telelight, located on the annunciator panel of the center instrument panel,
illuminates when the heater in the primary oxygen container has been m_nually
activated.
SYS_ OPERATION
The Environmental Control System (Figure 6-1, 6-2) is semi-automatic in operation
and provides positive control in all modes of operation. There are six r_
6-32
__. SEDR 300 ___.]
PROJECT GEMINIf--
operational modes :
I• Pre-Launch
2. Launch
3. Orbit
4. Re-Entry
5• Postlandlng
6. Emergency
Prior to the pre-launch mode, it is necessary to service and to check the system
functionally.
SERVICE AND CHECEDUT
For this operationj it is assumed that the spacecraft has been mated with the
booster on the launch pad and in the unservieed condition. The primary,
secondary and egress oxygen storage tanks are filled. The water boiler and
drinking water supply tank are supplied with water. The cartridge in the suit
loop cannister is then replaced.
PRE-LAUNCH
The pre-launch phase is defined as the period after the servle_-_ has been
completed and prior to launch.
Suit Loop
The pilots in their suits_ with face plates open_ are connected to the suit
6-33
_. SEDR300
PROJECT GEMINIr
circuit. The suit circuit compressor is actuated and the suit temperature
control valve is adjusted to satisfy the pilot desiring the cooler temperature.
The other pilot becomes comfortable by adjusting his suit flow rate control
valve toward the closed position to obtain a warmer setting. A ground supply
of pure oxygen is connecte_ to the pressure suit circuit purge fitting. Flow
is In_tiated with the face plates closed. The suit circuit gas is sampled
periodically until an acceptable oxygen content is attained. A suit circuit
leakage test is conducted. After satisfactorily completing the suit circuit
leakage test, the primary and secondary oxygen manual shutoff valves are opened
and the suit circuit purge system is disconnected and removed.
Cab
The cabin hatches are closed. A ground supply of pure oxygen is connected to
the cabin purge fltting2 flow is initiated and the cabin is purged. The cabin
fan is actuated and the recirculation valve is opened. A cabin leakage test is
conducted. After satisfactorily completing the cabin purge and leakage test,
the cabin purge system is disconnected and removed and the cabin purge fitting
is capped.
Oxygen Loop (Figure 6-3, 6-4)
The primary and secondary oxygen manual shutoff valves are opened.
The liquid oxygen inside the primary supercritical container has been changing
from a liquid to a supercritical fluid by thermal leakage and heater activation.
6-$4
SEDR300
A pressure control switch provides for aut_tic or manual activation of these
heaters. The manual control switch is located on the center control panel.
An indicator also on the center control panel indicates both pressure and
quantity from a transducer and control unit that are attached to the container.
oThe oxygen gas flows from the container and is warmed to approximately 50 F
in a heat exchanger. This heat exchanger also contains a relief valve that
limits m_ximum pressure to i000 psig. This valve opens, permitting full flow
and reseats within the range of 945-1000 psig.
A discharge temperature sensor provides an indication, for telemetering onl_,
of the temperature in the primary oxygen line downstream of the heat exchanger.
The oxygen gas is regulated from i000 psia maximum to ii0 + i0 psig. Flow
capacity of 0.35 ib/min with an inlet pressure from 800 to i000 psia and an
oinlet temperature of 60 F. This regulator also contains a relief feature that
limits downstream pressure to 215 psig in the event of a failed-open condition.
A lO-micron filter provides filtration of the primary oxygen supply before it
enters the suit or cabin loop.
LAUNCH
Cabin Loop
The cabin pressure relief valve opens to l_t the pressure differential between
cabin and ambient to 5.5 + .0 psi..0
6-35
__ SEDR 300
PROJECINIf
suit (Fure 6-8)
Oxygen is supplied to the suit loop through the suit pressure regulator. The
suit pressure is controlle_ to between 2 and 9 inches of water above cabin
pressure by the suit pressure regulator.
Suit circuit oxygen from the suit circuit demand regulator recirculates through
the suit compressor, the carbon dioxide and odor absorber, the suit heat
exchanger and water separator, the pressure suits, and the suit circuit solids
traps. There are two compressors in the circuit. One is an alternate to be used
if a compressor failure occurs. The alternate compressor is activated by posi-
tloni_ the SUIT FAN switch on the center panel. The cartridge of lithium
hydroxide and activated charcoal remove carbon dioxide and odors of an organic
nature that could have any ill effects on the pilots. As suit circuit oxygen
flows through the suit heat exchanger, the temperature is controlled as selected
by the pilots.
Solid traps, located in the oxygen outlet ducts of both pilots' suits, remove
particulate solids, preventing contamination of the suit circuit system. An
integral by-pass opens if the traps become choked with collected solids permitting
continuous oxygen flow through the suit circuit.
ORBIT
Nonual cabin le_age allows the cabin pressure to decay to a nominal value of
5.1 psia. The cabin pressure control valve maintains this value automatically. 4
6-36
__. SEDR300 _____
PROJECT GEMINI
A dual cabin pressure regulator supplies makeup oxygen through the pilots'
pressure suits to the cabin on demand, as sensed by two aneroid elements within
the regulator. The regulator supplies the makeup oxygen at a controlled pressure
between 5.0 to 5.3 psia.
The cabin fan on spacecraft 5 and 6 circulates cabin air through the cabin heat
exchanger. The cabin fan has been removed on spacecraft 8 through 12. One or both
of the pilots may open their faceplates. The cabin air circulating valve is in
the open position to provide for recirculation of the cabin oxygen through the
suit circuit.
In the event of spacecraft depressurization, whether intentionally or by space-
_ craft puncture, the dual cabin pressure regulator closes when cabin pressure
decreases to 4.1 +0.2-0.I psia, preventing excessive loss of oxygen.
Loop(Figure6-8, 6-9)
The suit circuit demand regulator senses cabin pressure and maintains suit
circuit pressure at 2.5 to 3.5 inches of water below to 2 to 9 inches of water
above cabin pressure. Should cabin pressure decrease below 3.5 psia, the suit
+0._circuit demand regulators maintain the suit circuit pressure at 3.5 -0.0 psia
by constant bleed orifices and sensing aneroids within the regulator. When
cabin pressure is restored to 5.1 +0.2-0.i psla, the suit circuit demand regulators
return to normal operation.
In the event of cabin and suit circuit malfunction, the suit circuit will
automatically revert to the high rate of operation when suit circuit pressureS_
6- 37
SEDR300
PRINIm
+0.1decreases below 3.0 -0.0 psla. A suit circuit pressure sensing switch energizes
the solenoid of the dual high flow rate and system shutoff valve. This initiates
a highoxygenflowrateof0.08 ibf nperman(totalflow:0.16Ib/m).
This high flow rate flows directlY into the suits by-passing the suit demand
regulators. The suit recirculating system is shut off and the suit compressors
are de-activated when the solenoid of the dual high flow rate and system shutoff
valve has been energized. The 02 HI RATE light on the center panel illuminates
when the suit circuit is on the high flow rate. There is also a manual control
for the high flow rate and system shutoff valve located on the center console.
When the suit circuit pressure is restored to a level above 3.0 _:_ psia, the
high rate and system shutoff valve is reset manually by using the control marked
02 HIGH RATE RECOCK located on the center panel. This returns the suit circuit _
to normal operation by opening the system shutoff valve and closing the high
rate valve. The suit compressor is also reactivated.
Water Management System (Figure 6-12)
The drinking water system is pressurized and manually controlled by the pilots.
Water from the adapter supplY is used to replenish the cabin tank water supply.
The w_ter tank drink selector valve is set in the NORM position.
The pilots manually operate the drinking dispenser to provide drinking water
from the cabin storage tank.
The water separator removes metabolic moisture through a wicklng material
positioned between the plates of the suit heat exchanger.
6-9
J- :_-. SEDR 300
__ PROJECT GEMINI
_:::::::_::;::::::::::::::::::::::::::::::::::::::::::::_:::E]E:::::::::::_:::::::::::::::::::::::::;::::_:::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::_::::::_::_:::::::::::::::::::::::::_:::_::_::::::::::::::::::::::::_::::::::::::::
i!ii ............................................................................................... ;:P:A:C:E:C:R:A:F T:::;:: C:O:N:F I:G U:R:A:T I:O:N ................................................................................................ iiii
H20 CONDENSATEVALVE VALVE
J METABOLIC MOISTUREFROM SUIT HEATEXCHANGER
",11_
DRINKING
SUITCOMPRE SSOR
URINE CONTROLSOLENOID VALVEVALVE
OVERBOARD _'" __""_,,__DUMF _|_
DISCONNECT
LEGEND
UNIT NOMENCLATURE WATEREVAPORATOR
J _ _ DRINKING WATER
02
f-_, _ FUEL CELL WATER/
WASTE WATER
GAS
REClRCULATINGO _,'-'/11_\;"
Figure 6-12 Water Management Schematic (Sheet 1 of 2)
6-39
j-:_. SEDR 300
"_' PROJECT GEMINI
WATERPRESSUREREGULATOR SHUTOFF
VALVE
J t cT---_CAF
AOAFTERWATER | I _STORAGE TANK
TO CABINWATER TANK CAP
CHECK VALVE
SPACECRAFT 6 CONFIGURATION
r_
FUEL CELL WATER :_:_:_:_:_:`:_:_:_:_:___:_:_:_:.:_:_:_:_:_:_:`:.:.:_;___:_:.:_:_:_:_:_:_:_:_:_:_:_:`:_:`:___:_:_:_:_:_:_:`:_:_:_:`:_:_:_:.:_:_:_:_:_:_:.:_:`:
ABSOLUTE PRESSURE ABSOLUTE PRESSURE TO RSS !REGUI_ATOR
REGULATOR REGULATOR ..-..::;.:.:..
VENT_ f._ ................_ ..................
+,,,, dk_,_2_K..... _- ....._CAF
DISCONNECT CHECK VALVE
SPACECRAFT 8 AND UP CONFIGURATION
Figure 6-12 Water Management Schematic (Sheet 2 of 2)6-40
. SEDR 300
PROJECT GEMINI
Urine Disposal S_stem (Figure 6-Ii)
The Chemical Urine Volume Measuring System (CUVMS) has a four position selector
valve labeled URINATE, SAMPLE, Dt_4Pand BY-PASS. Positic_ selection is made by
rotating the selector valve which is attached to a multi-ported center plug. The
selector valve includes a positive displacement tracer metering pump which is
activated by the handle as it passes over a plunger between the BY-PASS an_
URINATE positions. This supplies a quantity of tracer solution to the passages
to mix with the urine in the collection/missing bag. Therefore, volume measurement
can only be accomplished when the selector is in the URINATE position and the tracer
chemical is added. The urine receiver assembly and collection bag provides for
collecting and sampling urine or overboard dump provisions but does not provide
_ for volume measurement. The urine quick-disconnect hose and filter assembly con-
sists of a section of flexible hose with quick disconnect couplers on each end.
This assembly connects the CUVMS or the urine receiver assembly to the water
management panel through an in-line urine solids trap filter for dumping urine
overboard. The urine sampling bags are plastic laminate with a valve that
connects to the sampler port on the CUVMS or the urine receiver assembly for
taking urine samples. The urine receiver roll-on cuff is the interface between
the crewmen and the CUVMS or the urine receiver assembly. It provides an air
and liquid tight seal for direct urine transfer. Selector valve positions on
the CUVMS direct urine flow as follows. The SAMPLE position directs urine and
tracer chemical mixture frc_ collection/mixing bag to the sample port and into the
sampling bag. The D_P position directs the urine/tracer chemical mixture over-
board. The BY-PASS position (normal purge position) directs the urine flow to
the water management panel for dumping overboard.
6 _41
SEDR 300 ._j______jPROJECT GEMINI
The dump selector valve on the water management panel is positioned to route
the urine either to the water boiler or dumped overboard. The normal procedure
is to dmnp. Before it is dumped the urine dump system is preheated by position-
ing its heater switch located on the water management panel. A urine dump
heater light is also provided and located on the water management panel. This
light illumluates when the heater is activated.
RE-EFf_Y
O_en S_stem
The primary oxygen system is disconnected when the adapter section is separated
from the re-entry module. This re_m_vesthe primly oxygen supply pressure which
automatical_ activates the secondary oxygen supply.
The system shutoff and high rate valve is positioned to the high rate position
before the adapter section is Jettisoned.
Cabin
The pressure in the suit and cabin remains constant at 5 psia (nominal) until
an altitude of approximately 27,000 feet is reached.
As ambient pressure increases during descent, the cabin pressure relief valve
edm_ts ambient air into the cabin, preventing high differential pressures. The
6-h2
PROJECT GEMINI
cabin pressure relief valve begins to open when the ambient pressure is 15.0
inches of water greater than cabin pressure and opens to maximum flow when the
pressure differential is 20 inches of water.
At an altitude of 25,600 feet, or below, the pilots manually open the cabin
inflow and outflow valves to circulate external air through the cabin and suit
circuit.
Maximum negative pressure on the cabin should not exceed 2 psi as controlled
by the cabin relief valve.
s,u t
Prior to re-entry the face plates should be closed. The hlgh flow rate of oxygenf-_
is flowing directly into the suit circuit.
When the cabin inflow valve is opened It activates the suit compressor and
external air Is circulated through the suit circuit.
POSTIANDING
Ventilation is provided by the suit compressor as long as electrical power is
available (12 hours m_,m_m).
Ambient air is drawn into the vehicle through the snorkel inflow valve, by the
suit compressor, circulated through the suit circuit into the cabin, then
discharged overboard through the outflow vent valve.
6-43
SEDR300
PROJEC'CT--G-EMINI
The snorkel inlet valve functions as a water check valve. When the snorkel inlet
valve is above water level t the ball check is retained freely in a wire mesh
cage_ permitting ambient air to enter the suit circuit. Normal oscillations
of the spacecraft in the sea may result in the snorkel valve being momentarily
submerged. This will cause the ball check to seat and is held there by suction
from the suit compressor. Opening the cabin alr circulating valve allows the
ball to drop from its seat.
To prevent water from entering the cabin through the cabin pressure relief valve,
the manual shutoff section of the valve is closed.
_ERGENCY
Cabin Loop
If cabin depressurization becomes necessary due to toxic contaminants or fire,
the cabin outflow valve is opened to depressurize the cabin. The cabin regulator
will close_ stopping the oxygen supply to the cabin, permitting the escape of
toxic contaminants and preventing oxygen assistance to combustion in the event
of fire. The cabin repressurization valve permits repressurization of the
spacecraft cabin.
The control knob for the cabin repressurization valve is located on the lower
console and is rotated counterclockwise to open the valve. It is rotated
6-44
SEDR 300
clockwise to close the valve when cabin pressure is between 4.3 and 5.3 psia.
Cabin pressure is then automatically controlled at 5.1 +0.2 psia by cabin-0.i
pressure regulator valve.
Egress Oxygen (Figure 6-13)
This system is installed on spacecraft 5 and 6 only. Operation of the egress
oxygen system is initiated by three of the four lanyards which are pulled when
the seat leaves the spacecraft. One lanyard pulls a pin in the composite dis-
connect allowing it to separate and close the normal suit circuit. Two of the re-
maining Imlyards open the container shutoff valve and circuit relief valve acti-
vating the egress oxygen system.
Each of the egress oxygen containers is pressurized to 1800 psig with gaseous
oxygen. The oxygen flows fro_ the containers through a pressure regulator, where
the pressure is reduced to 40 psia. It then flows through a shutoff valve and a
flow restrictor, which allows a flow of 0.052 to 0.063 Ib/_in, then through a
check valve to the suit. After leaving the suit, oxygen flows through the shutoff
and relief valve, which du_s the oxygen overboard, as well as controls the suit
+0.6pressure to 3.5 -0.0 psia if ejection occurs at an altitude above 31,500 feet, and
2 to 8.23 inches of water above ambient at an altitude below 31,_00 feet.
S_ UNITS
DUAL SECONDARY OXYGEN RA_ AND SUIT SYSteM SEu_OFF VALVE (Figure 6-14)
The dual secondary oxygen rate and suit system shutoff valve provides a constant
flow rate of oxygen directly to the pilot's suit during re-entry or in the
event the suit circuit malfunctions during launch or orbit.
6-_.5
,-_ _ SEDR 300
PROJECTGEMINI
SU_T OUTLET
TANK
SUIT
ASSEMBLY
TO SUIT iNLET
_ECONOA,YOXYGEN_SUITOUTLET
i
_(_ LANYARD
_ _,_.. _J CONNECTION
•SHUTOFF AND ESSUREGAGERELIEFVALVE
SSEMELY CONNECTION
OXYGEN "_"_',
ASSEMBLY|_.- --
-CHECK /VALVE
SHUTOFF AND
RELIEF VALVE ___ LANYARDS
INSTALLED iNORMAL CONDITION NO FLOW SCHEMATIC I
[.J
ACTIVATED SCHEMATICCHECK
SHUTOFF AND PRESSURERELIEF VALVE GAGE
Figure 6-13 Egress Oxygen Flow Diagram (S/C 5 & 6)
6-46
.. -_ SEDR 300
f_
9RSHUTOFF SWITCH
0 2 SUPPLy
Figure 6-14 Dual Secondary Oxygen Rate and Shutoff Valve
6-47
SEDR300
MINI
The valve is designed for manual and automatic initiation. The recirculating
suit oxygen circuit flows through the shutoff section of the valve, which is
,_n,_ally opened and is spring loaded to the closed position. The shutoff valve
is held open by a 2_ vdc solenoi_ pin, as long as the solenoid is de-energized.
The secondary flow poppet valve, held closed by spring tension, remains closed
whenever the shutoff valve is in the open position. When the solenoid is
energized, the butterfly arm is released and rotates by spring tension, closing
the suit circuit valve and mech-nically opening the secondary oxygen flow rate
poppet valves. Opening the poppet valves allows oxygen to flow to each pilot's
suit through fixed orifices at a rate of 0.08 + 0.008 Ib/min per _n (total
flow 0.16 Ib/min.). The butterf_v arm simultaneously actuates a switch that
de-energizes the solenoid, turns off the suit compressor and cabin fan, and
illuml nares a SECO]_RY FLOW RA_ lamp on the pilots' center display panel.
A pressure sensor switch attached to each pilot's suit circuit will energize
the solenoid if the suit circuit pressure in either suit decreases below 3.0
+O.i psia, auto_-tically shutting off, the suit circuit flow azzd initiating-O.0
the secondary flow rate. A manual control is provided for resetting the valve
to the normal position. The secondary flow rate is used during re-entry.
Prior to retro-grade the pilots ,_n,z=11y disengage the solenoid initiating
the secondary flow rate.
SUIT OXYGEN DEMAND REGULATOR (Figure 6-1p)
The suit oxygen demand regulator controls the oxygen to the suit circuit from
the primary or secondary oxygen system and replenishes oxygen used by the pilots
or lost by lea_age.
6-48
SEDR 300
o M,N,
OXYGEN OUTLET
_ VENT TO CABIN
SUIT LOOP
Figure 6-15 Suit Oxygen Demand Regulator
6-49
SEDR 300 __.__[PROJECT GEMINI
Cabin pressure is sensed on one side of the diaphragm and suit pressure is
sensed on the opposite side of the diaphragm. The differential pressure across
this diaphragm opens or closes a poppet valve admitting or stopping oxygen flow
into the suit circuit. With cabin pressure of 5.0 psia the suit regulator main-
tains suit pressure at 2.5 to 3.5 inches of water below cabin pressure.
A resilient diaphragm type valve relieves pressure in the suit during ascent and
l_m_ts excess pressure to between 2.0 and 9.0 inches of water above cabin pressure.
During descent, the suit demand regulator relieves the secondary oxygen rate flow
through the relief portion of the valve, maintaining suit pressure 2 to 9 inches
of water above cabin pressure.
+O.4 psiaA constant bleed and aneroid elements maintain the suit pressure at 3.5 -0.0
if cabin pressure decreases below this pressure. The bleed flow by-passes the
tilt valve through a bleed orifice and is directed to the cabin pressure sensing
side of the pressure demaud diaphragm. A metering valve, controlled by an aneroid,
regulates the reference pressure on the demand diaphragm. The regulator returns
+0.2to normal operation when cabin pressure returns to 5.1 -0.i psia. In the event
that cabin decompression and a ruptured relief diaphragm in the regulator occur
simultaneously, an aneroid over the relief diaphragm extends to control suit
pressure at 3.9 psia maw_mum.
CABIN PRESSURE REW.VE_VALVE (Figure 6-16)
The cabin pressure relief valve automatically controls the cabin-to-amblent
differential pressure during launch, orbit and re-entry. Duplicate spring
loaded poppet valves are controlled by servo elements within the valve.
The servo elements control spring loaded metering valves which determine the
6-5o
__ SEDR300 __
PROJECT GEMINI
CABIN PRESSURERELIEF VALVE
ING
CHAMBER (TYP) (TYP)
SENSING
METERING VALVE (TYP)
SENSING CHAMBER (TYP)
CABIN PRESSURE(TYP)
CABIN AIR PORTAND FILTER (TYP)
[TYP)
BLEED ORiFiCE (
DIAPHRAGM (TYP)
SERVO ELEMENT (TYP)
_ET VALVE
POPPET VALVE
t CABIN
Sh'_LL PRESSURE J AMBIENT
BULKHEAD (REF)
" SCREEN ASSEMBLY MANUAL SHUTOFF VALVE
Figure 6-16 Cabin Pressure Relief Valve
6- 51
SEDR 300 -__ ---1PROJECT GEMINI f-
pressure within the diaphragm chamber behind the poppet, controlling the poppet
position. A small inlet bleed orifice admits cabin pressure to the diaphragm
chamber. When the poppet opens, a large orifice permits rapid change in pressure
ensuring quick closure of the poppet.
During ascent the valve will relieve cabin pressure as emblent pressure decreases
until cabin differential pressure is 5.5 to 6.0 psia. The valve closes maintain-
ing differential pressure in this range. When cabin pressure decresses below 5-5
psia the servo element closes the metering valve meintaining reference pressure
within the diaphragm chamber at cabin pressure. The poppet is held closed by
spring force and the zero differential between the diaphragm and the cabin prevents
cabin pressure release. If cabin differential pressure exceeds 5.5 psia the zero
element retracts, opening the metering valves, allowing the diaphragm chamber to
discharge to ambient. The discharge port being larger than the inlet bleed orifice
permits the diaphragm chamber to approach external pressure. The cabin pressure
reacting on the diaphragm overrides the poppet spring force, which opens per-
mlttlng cabin pressure relief to ambient. During descent, as external pressure
increases, ambient air is admitted to the cabin by the valve to reduce the differ-
ential pressure. As external pressure increases above the cabin pressure the
metering valves are held on their seats, preventing external pressure from enter-
ing the diaphragm chamber and retaining cabin pressure in the chamber. The poppet
valve senses diaphra_n chamber pressure versus ambient pressure. When the ambient
pressure is 15 inches of water greater than cabin pressure the poppet begins to
open permitting ambient air to enter the cabin. The poppet opens fully when the
differential pressure is 20 inches of water.....
To preclude water entering the cabin during postlandlng, a manual shutoff valve
6-52
___ SEDR300
PROJECT GEMINI
is provided.
SUIT CIRCUIT COMPRESSOR (Figure 6-17)
Two electric motor driven, single stage compressors are incorporated in the suit
circuit. One compressor is utilized for circulation of the gases within the suit
circuit, supplying both suits. The other compressor functions as a backup and is
activated only by manual selection by the pilots. Either c_pressor can be
manually selected by a switch on the center display panel, and both compressors
can be selected simultaneously.
When secondary oxygen flow rate is selected, the compressor is automatically de-
energized. Re-entry is made using the secondary rate. At an altitude of 25,600
_ feet or below the manual inflow valve is opened which re-energlzes the compressor.
The suit compressor provides ventilation during landing and for a twelve hour
postlanding period, or until the batteries fail.
SOLIDS TRAP (Figure 6-18)
A solids trap is located in the oxygen outlet duct of each suit. A cylindrical
40 micron filter strains the gaseous flow in the suit circuit removing the
solid matter. In the event that the trap becomes choaked with collected solids,
an integral by-pass opens when the differential pressure across the screen exceeds
0.50 inches of water.
DUAL CABIN PRESSUI_ REGULATOR (Figure 6-L9)
The cabin pressure regulator maintains cabin pressurization by providing makeup
oxygen to the cabin on demand. The regulator contains two aneroid elements _hich
individually sense cabin pressure. When cabin pressure decreases, the anerolds
6-53
s oR300PROJECT GEMINI
ELECI_ICAL
CON MOTOR
FAN
Figure 6-17 Suit Circuit Compressor
6-54
___ SEDR300 ___
PROJECT GEMINIf_
Jz_ _.
SOLIDSTRAP
/
Figure 6-18 Suit Circuit Solids Trap
6-5,5
__ SEDR300
CABIN PRESSURESENSING CHAMBER--
CABIN AIRFILTER_
Jp TO PRESSURESUIT
FROM OXYGEN SUPPLY
Figure 6-19 Dual Cabin Pressure Regulator
6-56
SEDR 300
PMINI
expand, forcing metering pins open and permitting oxygen flow into the cabin,
+0.2maintaining cabin pressure at 5.1 -O.i psia. If the cabin is punctured or develops
leakage greater than the flow capacity of the valve (4.79 + 0.48) 10-3 ib/mln,
oxygen flow to the cabin is stopped when the cabin pressure decreases to 4.0 +0.2-0.i
psia, by the aneroids expanding enough to cause the metering to close off the
oxygen.
PRIMARY SUPERCRITICAL OXYGEN CONTAINER (Figure 6-20)
The prlm-ry oxygen container is a double walled tank. A dual concentric cylinder,
quantity measuring devices I heaters and heat transfer spheres are internal to the
container. The tank contains two heaters. The first is a 12.0 + 2 watt heater
which is activated either manually by a switch located on the center panel 3 or
automatically by a pressure switch. The pressure switch controls the activation
of the heating element in the tank to automatically ,u_1ntain the cryogen in a
supercritical state. The switch de-energizeS the heater circuit when the pressure
in the tank is between 875 to 910 pslg, and closes the circuit 15 to 75 psig be-
+50low the opening pressure. The second heater is a 325 -0 watt heater manually
controlled by a switch located on the overhead switch/circuit breaker panel.
The pressure relief valve maintains the oxygen pressure within the container at
i000 -_5 psig, and prevents overpressurization of the containers.
Provisions for servicing the primary oxygen container from a ground supply source
of oxygen are provided.
SEC0_DARY OXYGEN CONTAINER (Figure 6-21)
The secondary oxygen container is a cylindrical shaped container, having a
useful oxygen capacity of 6.5 pounds at an operating pressure of 5000 psig.
6-57
-+. SEDR300
_-_'__ PROJECT GEMINICONTROL PRESSURE AND
PRESSURE ANDQUAF_ITY GAGINGSYSTEM POWERINVERTER
TRANSDUCER
OXYGEN PRESSURECO NTgOI. _VITCH
OUTLET PORT
TO HEAT EXCHANGER
HEATEI_'_
TRANSDUCER SENSE PORT
PRESSURIZATION
Figure 6-20 Supercritical Primary Oxygen Container
6-58
sEo,,ooPROJECT GEMINI
3ULATOR
OXYGEN CYLINDER
VALVE
PORT
PORT
RELIE
DXYGEN FILL VALVE
PRESSURETRAN:
/_ ELECTRICAL RE(
Figure 6-21 Secondary Oxygen Tartk
6-59/60
COOLING SYSTEM
eetionVII
TABLE OF CONTENTS
TITLE PAGE __bQOtOQ_- _ -_ ......
SYSTEM DESCRIPTION 7-3 _"":""_':_'_"'-"__............................... _--ff_:-_.:-..-:.:-::._..-.---_
SYSTEM OPERATIONS ................................ 7-5 _.:"T'"'"".:'"..'..'".:"IPRELAUNCH...............................................i!iiiiiiiiiii!iiiiiiiiiiiii
i:.:'"'":"_'""'.-"_'.:_"_LA U N CH ................................................... 7-10 iiii!i_iiiii[!![iiiiii_iii
::::::::::::::::::::::::::::::::::::::::::::::::::::::0 RBIT ....................................................... 7 -11 :::::::::::::::::::::::::::
SYSTEM UNITS ........................................... 7-11 ii[[iiii_[[[iii[i[i![_i[ii_PUMP PACKAGE ......................................... 7-11 iiiiiiiiiiiiii'_iii_iii!ii
RADIATOR ................................................ 7-15 ..........................--..-.----.---..-.----...............................°°.o°o°...°..° ...... ,°°oo
COLD PLATES 5 ...........................,, .............., .....................,..,.. 7-1 -.°.----°......,-.-°......"......."'"'""°'"°o..°.°oo ....... °°,°o.°..°...... °° ....... ° ...... °.°........ °° ....... °° ...... .°
HEAT EXCHANGERS ..........................718 .........................._ -°° ........ .. ...... ..o .....° .................. . .....• .. 0.--o, **° • **......°....*......... °.o .......................
.... ........... ° ...... ...,.°.... ......... .oo0...o.,.
TEMPERATURE CONTROL VALVE ..........................• °..... ....... .. ...... °•....... 7-18 ..........................
LAUNCH COOLING HEAT ANGER ..........................EXCH 7 22 ..........................• •.• ........ • .......... •,•- ° ........... . .............• • • °.. °o.. ..°• ....... o,. ....... o°••,
• .° ........ • .......... •..o• .............. °° ....... ••
.:-_. SEDR300
PROJECT GEMINI
"(VIEW LOOKING FROM Z0.00) (VIEW LOOKING FROM Z0.00)
ECS COOLANT
/ HEAT EXCHANGER
/ TO GROUND COOLANT
EQUIPMENT ADAPTER
RADIATOR SECTION (TYF)_
SECTION RADIATOR (TYP)
PRESSURIZEDAREA
EQUIPMENT BAYCGLDPLATES (UNPRESSURIZED)
EXCHANGER
RECORDERCOLD PLATES
ELECTRONIC MODULE
COLDpLATES _.
TEMPERATURECONTROL VALVES
FROM PRESSURIZEDAREA
SUIT HEAT EXCHANGER /
RIGHT EQUIPMENT BAY _ _ /
COLDPLATES (UNPRESSURIZED)- BAY COLDPLATES(UNPRESSURIZED) "_
SUIT AND CABINCONTROL VALVE
Figure 7-1 Spacecraft Coolant System
%2
sEo300PROJ EC-'T GEMINI
SECTION Vll COOT.TNG SYSTEM
S_STEM DESCRIPTION (Figure 7-I)
The spacecraft cooling system consists basically of two identical temperature
control circuits functioning independently of each other to provide the cooling
requirements for the spacecraft. Each cooling circuit consists of a pump
package, thermostatic and directional control valves, various type heat exchangers,
rad/ators, filters, and the necessary plmabing required to provide a closed
circuit. The coollng system may be operated in either the primary and/or
secondary circuit, and is capable of carrying maximum heat loads in either circuit.
The equipment coldplates, cabin and suit heat exchangers are located in the re-
entry module. The upper radiator panels are located in the retrograde section.
The pump package, battery coldplates, filters, electronic equipment coldplates,
ground launch cooling and regenerative heat exchangers and the lower radiator
panels are located in the adapter equipment section. System manual controls are
located on the pilots' pedestal console and the control switches, warning lights
and indicators are located on the center panel.
The cooling systems in spacecraft 8 through 12 are provided with a means of by-
passing the coolant around the fuel cells rather than throug3 them. This pro-
vision is for ground operation when fuel cells are not in use.
During orbital flight, Monsanto MCS-198 coolant is supplied throughout the cool-
ing system and thermostatic control valves regulate the coolant temperature.
%3
.__. SEDR300
PROJECT-'G EM INI
Temperature sensors, located in the system, provided the necessary telemetering
of system temperatures to ground stations.
SYS_ DISPLAYS AW9 COi2_Z,S
Cooling system displays and controls are located on panels in the cabin as shown
in section 3 and function as specified below.
SUIT And CABIN _ Controls
Dual concentric knobs are _unted between the ejection seats for suit and cabin
temperature control. These knobs control the operation of valves regulating
the flov rate of primary and secondary coolant through the suit and cabin heat
exchangers for spacecraft 5 and 6. Spacecraft 8 does not have the cabin heat
exchanger. Clockwise rotation results in increased temperatures.
CABIN And SUIT _ INdicator
A dual indicator provides for monitoring temperatures in the suit and cabin
circuits. P_e markings are calibrated in degrees Fahrenheit.
PRIMARY And SECOKDARY Pu_ Switches
These switches are connected to the coolant pumps power supplies, one switch
for each power supply. Each switch has two positions; ON and OFF. The switches
are located on the center panel. On spacecraft 8 through 12 B pump switch in each
loop changes the flow rate from 183 ib/hr to 140 ib/hr.
7-4
___ SEDR 300 ___ __.__
PROJ GEMINI
Pump lights ill-m4uate when the pumps are activated. They are located above
their respective switches near the top of the center panel. The _S LO lights
ill-,_uate when the coolant level in the reservoir is low.
EVAP PRESS Indicator
+0.0 pstgThis light illuminates when pressure in the evaporator builds up to _.0 -0.3
+0.3 psig.and is extinguished when the pressure falla to 3.1 -0.3
EVAP PRESS Heater Switch
This switch is connected to the evaporator heater and is used to heat the w_ter
in the evaporator before dumping.
SYSTEM OPERATION
The cooling circuit in which the cooling system operates is dependent upon the
temyerature loads generated by the equipment, spacecraft phase of flight and the
temperature within the spacecraft cabin. Cooling is provided throu6hout the
mission up to pre-retrograde firing. At this time the coolant pump packages are
Jettisoned with the adapter equipment section, terminating spacecraft cooling.
Spacecraft 5 and 8 through 12 require both loops to be operaC¢'_ oontlnuous],7. In
spacecraft 6 the primary circuit operates continuous_7 provldi_ the required
cooling during low temperature loads. The secona=ry elreult is used, in con-
Junction with the primly circuit, during phases of high temperature loads;
f_ namely - launch, rendezvous, and pre-retrograde. Under normal heat loads, the
7-2
SEDR300
number i pump in the primary circuit provides the required cooling. Under peak
heat loads, the number 1 pump in the secondary circuit is used with the primary
circuit number 1 pump to provide maximum cooling. In the event of a number 1
pump malfunction in either circuit, the number 2 pump in that circuit is used.
In the event of both pumps failing in one circuit, both pumps of the remaining
circuit can be used to provide the required cooling. (Spacecraft 6 does not have
the number 2 pump in either circuit. )
PRE-LAUNCH (Figure 7-2)
During pre-launch an external supply of Monsanto MCS-198 coolant is circulated
through the spacecraft ground cooling heat exchanger providing temperature control
of the cooling system coolant. The number i pumps of the primary and secondary _
cooling circuits are activated, using an external power source, to provide the
required cooling for spacecraft equipment and cabin. The spacecraft radiator
switch, located on the center panel, is placed in the BYPASS position so the
cooling system coolant by-passes the radiators and is directed through the
ground cooling heat exchanger.
Coolant is circulated through each coolant loop by a positive-displacement gear
pump. Spacecraft 5, and 8 through 12 are provided with 2 pumps in each loop.
Spacecraft 6 has only one pump in each loop. Selection of loops and number of
pumps is controlled manually.
The coolant is filtered_ as it leaves the pump, and simultaneously flows to
the inlet of the battery coldplate or fuel cell temperature control valve and
primary oxygen heat exchanger.
7-6
SEO3o0I I COLDP_TEI _ _l;_/ COLO_TE[ _ . _]=_f
_LJ o-',-o, ":_J w::::':':':':':':':-_':':':'::::':':::"'_/ll ...... COLD P E !_:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:_
TAPERECORDERlI _DI PLATE 5EI.ECTORAND
II _ I ADAPTER RELIEFVALVE _oco_oocoo
I ................._
SELECTORAND/r. CA,BIN HEAT RELIEFVALVE
EXCHANGER SUITHEATAND FAN(S/C 5 & 6 EXCHANGERONLY)
TEMPERA]lURE
CABIN ! CONTROLVALVESTEMP I +2°
40= _4oF
TEMPVALVE
L PRESSURIZED CABIN DISCONNECTS /m
H20 EVAPORATOR
TEMPBELLOWS
LEGEND
.....-.---...---.-.-.-.... PRIMARYCOOLANT
SECONDARyCOOLANT
/_ _ PRIMARYDRAIN LINE
......... SECONDARYDRAIN LINE tOVERBOARD
Figure 7-2 Cooling System Flow Schematic Pre-Launch and Launch (Sheet 1 of 2)
7-7
-_ SEDR 300
"'" FUE FUEL CELLi1 8&UP
BYPASS !• -' I POWER SUPPLy EXTERNAL EXTERNAL
__1 j_ RESERVOIR RESERVOIR
FILL PORT FILL PORT
RSSHEATEXCHANGER
WATER PUELC6.L I I
II PUMPA II IpUMP |
. BATTERy ._-.-.-.-.-.-.-.-.-.-.-.-COLD PLATE COLD PLATE
(_/C 6 ONLY) (_/C 6 ONLY)
(S/C 6 ONLY)
REGENERATIVE REGENERATIVEHEAT HEATEXCHANGER EXCHANGER
ADAPTER
_O_TOR 1 1
CIRCUIT
REGEN_TIVE REGENERATIVEHEAT HEATEXCHANGER EXCHANGER
GROUND GROUNDCOOLANT COOLANTINLET OUTLET
Figure 7-2 Cooling System Flow Schematic Pre-Launch and Launch (Sheet 2 of 2)
7-8
s=o3OcGEPRO,JE MINI
The temperature control valve maintains the cooling temperature at the fuel cell
+2 °or battery eoldplate inlet at 75° .to F. Temperature increasing above setting
will reduce by-pass flow. Coolant temperature from by-pass line varies from 80°F
to 165°F. Coolant temperature from equipment lines varies from 60°F to 125°F.
Coolant enters the primary o_gen heat exchanger and then is routed around the
steam discharge lines in the water boiler before it passes through the regenerative
heat exchanger. It then passes through the selector and pressure relief valve.
This selector valve is electrically actuated and when in the radiator by-pass
position allows the coolant to pass through the ground cooling heat exchanger
_here the exterT_al supply of coolant flo_.Tingthrough the ground cooling heat
exchanger absorbs the heat from the spacecraft's coolant system.
The ground coolant heat exchanger has an airborne flow capacity of 336 ib/hr, per
coolant loop, at 125°F. It has a ground coolant flow capacity of 425 Ib/hr at
O°F.
_le coolant is now ready to pass through the terzperature control valve. This
+2 °valve maintains the outlet temperature at 40° _40 F. If the coolant entering the
valve from the ground heat exchanger is below this range, a portion of the coolant
is directed through the regenerative heat exchanger and then mixed at the valve.
The coolant then fl_s through the water evaporator to the cabin and suit manual
temperature control valves. These valves meter the coolant flow through the
cabin and suit heat exchangers. The evaporator selector valve relief portion
allows part of the coolant to by-pass the cabin and suit heat exchangers depending
7-9
SEDR300
PROJECT GEMINI
on the setting of the manual control valves. The selector portion of this
valve allows the by-pass fluid to come from either downstream or upstream of the
evaporator. The coolant continues through the various coldplates until it reaches
the battery eoldplates for spacecraft 6 or through the fuel cells on spacecraft
5, and 8 through 12. The coolant has now returned to the reservoir where the cycle
is ready to be repeated.
Shortly before launch, the external cooling and electrical power are disconnected.
zucs 7-2)
During launchx the launch cooling heat exchanger goes into operation in the
following sequences. The heat trRnafer characteristics a_idcapabilities of the
ground cooling heat exchanger no longer exist. The Monsanto MCS 198 coolant
fluid now with no place to dissipate its internal heat, which is constantly being
generated by _nd absorbed from the loop components, circulates about the tempera-
ture control valve of the heat exchanger. When the coolant temperature exceeds
+4 °46 _2o F the temperature control valve opens to pressurize a donut shaped bellows
which unseats the poppet valve exposing the water in the heat exchanger core to
reduced pressure as altitude increases during launch.
When spacecraft altitude exceeds 1OO,O00 feet, water in the heat exchanger will
boil absorbing heat from the coolant. This absorbed heat is then expelled
overboard in the form of steam.
When the coolant reaches a temperature of 46°Fj the temperature control valve re-
positions to relieve pressure to the donut shaped bellows holding the poppet open.
As this pressure diminishes, a spring behind the poppet will reposition it to the
7-10
_. SEDR300 _'---1
PROJECT GEMINI.
closed position. The evaporator selector valve is positioned to allow all flow
to go through the evaporator.
The water boiler water reservoir is constantly replenished from the suit heat
exchanger water separator, and if the need arises, from the drinking water
supply tank.
(Figure7-S)
After orbiting for approximately 30 minutes, to allow the radiator to cool after
being subject to launch heating, the coolant flow is directed through the space
radiators by manual selection of the radiator switch located on the center panel.
This by-passes the ground cooling heat exchanger. The evaporator selector valve
is also positioned so that only the flow to the suit and cabin heat exchangers
pass through the evaporator.
Prior to retrograde firing, the coolant pump packages, radiators, batteries and
various heat exchangers are jettisoned with the adapter equipment section. Prior
to adapter jettisoning and retrograde firing the number i coolant pumps for both
the primary and secondary coolant circuits are activated. The suit, cabin, and
equipment bays are cooled to as low a temperature as possible, before the adapter
equipment section is Jettisoned.
SYSteM UNITSw
PUMP PAC_E (Figure 7-4)
The pump package for each coolant circuit incorporates two constant displacement
electrical pumps, two pump inverters, an external reservoir, filters, relief and
7-11
SEOR30o
.;- PROJECT GEMINI
COLD PLATE COLD PLATE
COOLANT L
I SYSTEM ' DISCONNECTSjHIGH POINT CONTROLVALVEFILLPOAT - o÷2°
75 .,4oF
J (VC 8a l I ONLY )
I
COLD PLATE
j SELECTORANDADAPTER RELIEFVALVE _
24V DC
SELECTORANDCABIN HEAT RELIEFVALVE " "_
EXCHANGER SUITHEAT J
AND(S/C 5 & 6FAN EXCHANGER J |ONLY)
J TEMPERATURE_1_ .Ir_ CONTROLVALVESCABIN +2°TEMP I 40_ _4oF
VALVE _ 24V DC
I
L mPRESSURIZED CABIN DISCONNECTS
H20 EVAPORATOR
TEMPBELLOWS
LEGEND
PRIMARyCOOLANT
SECONDARyCOOLANT
PRIMARyDRAIN LINE
......... SECONDARYDRAIN LINEOVERBOARD
Figure 7-3 Cooling System Flow Schematic-Orbit (Sheet 1 of 2)
7-12
sEo30o•" FUE FUELCELL --
_1 8&UP
! BYPASS
___ POWER SUPPLy EXTERNAL EXT_.NAL
" RESERVOIR
FILL PORT FILL FORT
i BSS HEAT
i EXCHANGER
WATER ( FUEL CELL "_ I | 6 ONLY)n2 . I e
I PUMP A | IpuMp j I
BATTERY _ BATTERY _ I l l |j COLDmATE[ 1COLD_TEr I(_C 6 ONLY) _[ 1 (S/C 6 ONLY) I
(s/c6ONLY)
REGENERATIVE REGENERATIVEHEAT HEAT
ADAPTER EXCHANGER EXCHANGER
RADIATOR _J
CIRCUIT
REGENI_.ATiVE REGENERATIVEHEAT HEATEXCHANGER EXCHANGER
GROUND GROUND
COOLANT I COOLANTINLET OUTLET
Figure %3 Cooling System Flow Schematic-Orbit (Sheet 2 of 2)
7-13
SEDR 300
_ LIMIT SWITCH
BELLOWS
\
\FLOW SWITCH
FLUlD RESERVOIR FILL VALVE --
FILLPORT -- CHECK -_
'[-- -1"| I ELECTRICAL
'2_Mp--/ @ pUMpview A-A
DRT
(OPERATING)
xJI
CHECK VALVE" SECTION R-B
Figure 7-4 Coolant Pump Package
7-14
SEDR300
PROJECT"GEMINI
check valves. The pump package is located in the adapter equipment section.
Pump selection is provided by switches on the pilots' center pa_-el. A pump
failure warning light is provided on the center panel. When a pump is activated
the coolant flows from the reservoir to the pump, which circulates the coolant
through the cooling circuit. The cooIAut returns to an external reservoir that
compensates for thermal expansion, contraction, and leakage of the coolant. A
i00 micron filter downstream of the pump prevents contamination of the cooling
system. Check valves in the pump package prevent the operating pump from pumping
coolant into the redundant pump. Flow sensing switches il1--_nate a pump failure
lamp on the pilots' center panel in the event of pump failure.
The spacecraft radiator consists of two circnmferential radtiator panels made of
0.25 inch diameter cooling tubes. There are four sections o£ tubing %0 each
radiator panel. The tubing is -_nufactured as part of the spacecraft structure.
Each panel incorporates two parallel cooling circuits, one for the primary
cooling circuit and the other for the secondary circuit. Duri_ orbit the cool-
ing system coolant is circulated through the radiator. The heat of the coolant
radiates into space, lowering the temperature of the coolant.
coumu s (Fibre7-6)
The coldplates, other than the battery coldplates, are plate fin constructed
units incorporating parallel coolant system passages. Coldplates are fabricated
from al-,_num. Battery, electrical, electronic and other heat generating
7-15
___ SEDR 300 __
PROJECT GEMINI
COOLANT FLOW PASSAGE
(TYP) _'--ADAPTER MOLD LINE
_ .-...
INLET
PRIMARY OUTLET-
EQUIPMENTSECTION
RETROGRADESECTION
Y OUTLET _
QUARTER PANELS (TYP 4 PLACES)
Figure 7-5 Radiator Stringer Assembly
7-16
.__ SEDR300 _---'-]
PROJECT GEMINIf---
COLDPLATE(TYPICAL)
INLET
/
PRIMARYSECTION
VIEW OF COLD PLATESEPARATEDTO CLEARLY ILLUSTRATE THE PRIMARYAND SECONDARY FLOW,
,f_x
SECONDARY SECTION
Figure 7-6 Cold Plate
7-17
___ SEDR3O0 ___ ..__
PROJECT GEMINI
components are mounted on coldplates. The coolant flowing through the coldplates
absorbs the heat generated by the components2 preventing overheating of the
operating equipment.
SEAT EXCHANGI_RS (Figure 7-7, 7-8)
Two types of heat exchangers are used in the spacecraft; namely, plate fin
constructed and shell and tube constructed heat exchangers. The suit, cabin,
water evaporator, ground cooling and regenerative heat exchangers are of plate
fin construction. The primary oxygen heat exchanger is of shell and tube
construction. The coolant absorbs heat from the cabin, suit and regenerative
heat exchangers. The ground cooling and water evaporator heat exchangers permit
heat transfer to cool the coolant. The primary oxygen heat exchanger is designed
so heat transfer will heat the primary oxygen to a desired temperature.
T_4PERATURE CONTROL VALVE (Figure 7-9)
Temperature control valves are provided in both the primary and secondary cooling
circuits. These valves are located at the radiator outlets and at the inlets
to the battery coldplates or fuel cells.
The temperature control valve located in the coolant system radiator outlet
automatically maintains the coolant outlet temperature at 40 +_F as long as
the radiator capacity has not been exceeded.
The temperature control valve located in the battery coldplate inlet automatically
maintains the coolant l-!et teaperature at 75 - F or above.
7-18
_._._. SEDR300
f_
EN OUTLET
0 U;_E_TLANT ....
(SEC
__ V_ATEROUTLET
COOLANT
(PRIMARy) SUITOXYGEN
INLET
VEIW ROTATED 180 °
TYPICALPLATEFIN CONSTRUCTION
!
Figure 7-7 Heat Exchanger-Suit
7-19
._ SEDR300 __
PROJECT GEMINI
OUTLEt PORTCOOLANT LOOP ONE
OXYGEN _INLET
INLET PORTCOOkANT LOOP ONE
OUTLEt PORTCOOlaNT LOOP TWO
INLEt PORT
COOLANT LOOP TWO
DRT
OUTLETPORT
VALVERELIEF
COOLANT TUBES
OXYGEN TEMPERATURE L
SENSOR PORT
OXYGEN TESTPORT
Figure7-8 Coolant Tube Type Heat Exchanger
7-20
___. SEDR300 __
PROJECT GEMINI
INLET PORT
(FROM ELECTRONIC EQUIPMENT)
9
4_1 BY-PASS PORT
/
FLUID MIXINGSECTION
S_NSING ELEMENT
CALIBRATION SCREW
F-_ _,OUTLET PORT
Figure 7-9 Coolant Temperature Control Valve
7-21
__. SEDR300PROGEMINI ..
The t_perature control valve contains a piston that regulates the inlet flow
to the valve. The piston is spring loaded on one side. A thermostatic actuator
on the opposite side of the piston determines piston movement, which in turn
regulates the coolant flow through the valve. The thermostatic actuator, which
is located to accuratel_ sense mixing temperature, consists of an encapsulated
wax pellet that expands or contracts as temperature varies. As temperature around
the pellet increases, the wax expands exerting pressure. on the diaphragm. The
diaphragm moves a piston, which in turn controls the inlet flow to the valve.
Temperature reduction aroun_ the wax decreases the pressure in the pellet cup
allowing the spring to reposition the piston regulating the flow of cool-nt
through the valve.
LAUNCHCOOT.r,G._ATEXCEA_GE_(Figure7-10,7-11)
The launch cooling heat exchanger is located in the adapter section. Via its
relief valve it can dump liquids overboard, or if the temperature control valve
senses temperature greater than 50°F, it can control the outlet temperature
of the pr_,_y and secondary coolants to _6 . . In addition, it serves as
a water reservoir, storing water until it is needed for cooling.
This evaporator consists of a wicking type heat exchanger and is capable of
storing seven pounds of water. A temperature control valve has been set to
+_ocontrol the outlet coo_,t temperature to _6° _20 F. A relief valve opens and
allows excess water to be dumped overboard at 2.75 + 0.25 psi differential and
reseats at 2.0 psi differential minimum. An electrical heater is provided in
the poppet to prevent ice formation. Coolant flow capacity is 366 ib/hr at 40°F. .....
7-22
, SEDR300 __
PROJECT GEMINI
COOLANT PROMPUMPS Jl
STEAMOVERBOARD
CIRCUITS REGENERATIVE
f_
COOLANTTO COOLAIXrTTORE-ENTRY _11__ -- I_' RE-ENTRYMODULE SEC PRI MODULE
COOLANT LOOP(PRIMARY) _@_:_:_::
COOLANT LOOP(SECONDARY) . - .
Figure 7-10 Launch Cooling Heat Exchanger Schematic
7-23
___ SEDR300 __r
PROJECT GEMINI
ELECTRICAL FLUID HEATERCOOLANT
LOOP (PRIMARY), OUTLETWARNING LIGHT J - FLUID HEATER COOLANT
PRESSURESWITCH--_ / LOOP (PRIMARY) iNLET\STEAM OUTLET
WATER 11_ TO AMBIENT
WATER
CORE COOLANT 3 HEATERLOOP (PRIMARY) COOLANT LOOP
OUTLET (SECONDARY) INLET
CORE COOLANT EATERLOOP (SECONDARY)OUTLE1
PRESSURECONTROLVALVE
CORE COOLANT LOOP
(SECONDARY) INLET
CORE COOLANTLOOP (PRIMARY)TO RESERVOIR
CORE COOLANTLOOP (SECONDARY)TO
Figure 7-11 Launch Cooling Heat Exchanger
7-24
__. SEDR300
PROJECT GEMINI
Water flow capacity is 3 Ib/min when coolin_ is not required from the evaporator.
Maximum operating pressure in the fluid heater coolant circuits is 230 psig, and
lO0 psig in the core circuits. Maximum operatin_ pressure in the water circuit
is 20 psig with exit port relief valve in normal operation.
The steam exit duct is continuously heated by coolant coming from the pri_-_y
oxygen heat exchanger to prevent ice form.tion.
A loss of pressure in either coolant loop will not affect the operation of the
valve.
7-2/26
GUIDANCE andCONTROL SYSTEM
VIII
r •
REFER TO THE SEDR 300CONFIDENTIALSUPPLEMENT FOR INFORMATION CON.................CERNING THE GEMINI GUIDANCE AND ::::::::::::::_:"'_-_"_CONTROL SYSTEM ...........................,°*o_o_o_°°°°°ooo_o_oo.°o,
,Jo.ooI°.°°°°°.°°°tetoo_,oo°6o_oo°oo°°°°o..Hooooo_
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°................... °............................... °......................... °°.
.................. °°°.o ....
81/2 ....................................... .° ...............
COMMUNICATIONSYSTEM
SectionIX
TITLE PAGE
SYSTEM DESCRIPTION ................................. 9-3ANTENNAS ................................................. 9-3
BEACONS ................................................... 9-5
VOICE COMMUNICATIONS ........................... 9-8
TELEMETRYTRANSMITTERS ............................ 9-8
FLASHING RECOVERY LIGHT ..... ............ 9-8
DIGITAL COMMAND SYSTEM........................ 9-8 i'.::_.;_--'.:_'_SYSTEM OPERATION .................................. 9-9 i_-_-_!_ _
VOICE TAPE RECORDER ............................... 9-9 ff_"_._'_--_--":_"::{i_-'_
PRE-LAUN CH .............................................. 9-10 i_:_i:_ff:_iii._i.__i:_:_SPACECRAFT/LAUNCHVEHICLE SEPARATION ...9-13 iiiiii!ii!_iiii_ii:":_":"'_:_
_°°*t°l._ettoo_o°o*°_o.°o°.
i:'.:'::".7:':::'::-::}F:-.:.:}{_0 RBIT ........................................................ 9-13 iiiiiiiiiiiiiii_iiiii_iiiiADAPTER SEPARATION ................................. 9-17 _--_.-_--..-.:--.:_--..:.:----._
:::::::::::::::::::::::::::
LANDING THROUGH RECOVERY .................... 9-19 iiiiii[_!iiiii_iiiiiiiii[ii
SYSTEM UNITS ............................................ 9-20 iiii_iiii_i_iiii!iliiiiiiil....°..o..°oo.°°°°°.°.°..°.
ANT ENNAS ................................................. 9- 20 iiiiiiii!ii!i_!_i!_iiiiill
8EAC0 NS................................................... 9- 39 i_i_iiii!ii_iiiiiiiiiiiiiii,o.°° .°o°.°°°....°°.°o°°o.,
VOICE COMMUNICATION ............................. 9-49 _.:_...:._.:.._z._._..:.._!_TELEMETRYTRANSMITTERS ............................ 9-63 iiiiiiiiiiiiiiiiiii_i_iiiiFLASHING RECOVERY LIGHT AND "'""'":.:_':"_:':'_'_i
POWER SUPPLY.......................................... 9-68 iHiiii!!!!ilHiiiiii!!i!iii.... °........,°°.°..°o° ............. . ........ ..°° .....
DIGITAL COMMAND SYSTEM ........................ 9-70 _i_iiiiiiiiiiiiiiiii_i_!!::9-]_ :::::::::::::::::::::::::::
SEDR 300
Figure 9-1 Communication System
9-2
___ SEDR300 ____j
PROJECT GEMINIf_
SECTION IX CGMMURICATION SYST_
SYST_ DESCRIPTION
The Cummanication System is the only c_.,,_,m4cation link between the ground and the
Gemini Spacecraft. The system has the following capabilities: radar tracking of
the spacecraft; two-way voice c_]n_cations between the ground and the spacecraft,
and between the crew; grom_ c@,_,_ndto the spacecraft; Instrmmentation System data
transmission; and postlanding and recovery aid data transmission. To make possible
these various capabilities, the C_manication System contains components that may
be divided into the following categories: antennas, including multiplexers and
coaxial switches; beacons; voice c_._._mications; telemetry transmitters; flashing
recovery light; and Digital C_nd System. The flashing recovery light and the
_A uhf recovery beacon are grouped together in a category called the Electronic
RecoveryAids (ERA).
The C_munication System components are located throughout the spacecraft with
the largest concentration being in the right equipment bay of the re-entry
module and the electronic module of the adapter equipment section as illustrated
in Figure 9-i.
ANTENNAS
Eight antennas and one antenna system provide transmission and/or reception capa-
bilities for the various Communication System components. The spacecraft Com-
munication System (Figure 9-2) contains the following antennas: uhf recovery;
uhf stub; uhf descent; two uh_ whips; two hf whips; C-band annular slot; and a
C-band antenna system consisting of a power divider, a phase shifter, a phase
shifter power supply, and three helical antennas. Antenna usage is illustrated
in Figure 9-3 and described in the individual antenna description.
9-3
300PROJECT GEMINI
4_ZN
I "
_ I I _°
0_
Z i-. i-.
u_ z
I -- 'I 't 'r
I _ Ia.
/\ /\ I_LX/\I_ix /\" I_ I! .... "
Figure 9-2 Communications System Block Diagram9-4
__ SEDR300
PROJECt GEMINIf_
To achieve the most efficient antenna usage, a diplexer and a quadriplexer are
used with the ,,hf whips and the uhf stub antenna. The multiplexers make it
possible to use more than one transmitter and/or receiver with a single antenna.
Five coaxial switches permit antenna and transmltter/receiver switching for best
ccmnunication coverage during the various phases of the mission (launch, orbit,
re-entry and recovery).
BEACONS
Four beacons in the Comunication System establish the capability of locating
and tracking the spacecraft during the mission. The four beacons are: An
acquisition aid beacon and a recovery beacon used to locate the spacecraft, and
two C-band beacons used to track the spacecraft. The acquisition aid beacon,
operating on a fixed frequency, is used to determine when the spacecraft is
within the range of a ground tracking station, and provides information for
orientating the ground station antennas during the orbital phase of the mission.
The recovery beacon is a transmitter that operates on the international distress
frequency, and is used by the recovery forces to determine the spacecraft
location. The C-band beacons are transponders which, when properXy interrogated
by a ground station, transmit signals for accurate spacecraft tracking.
During the recovery phase of the mission, emergency communlcatioms man be estab-
lished by connecting one of the uhf rescue beacon transceivers to the uhf recovery
antenna. The rescue beacon transceivers are Govei_ent Furnished Equipment (GFE),
stowed in pilot's survival kits.
9-5
__ SEDR300 ___
PROJECT GEMINI
RE-ENTRY EQUIPMENTADAPTERPREPARATION SEPARATION
® ®180"YAW
SPACECRAFTTURN/_IOUND___
SPACECRAFTSEPARATION_ _ 1_ PRE.-LAUNCHI RE,ENTRYc-BAND BEACON-ON (CONT.)
...... 7_ / ")o _ _>UHPT/RNO.R-OF_
........ ( / HF TIR -OFF$_ AUDIO NO. 1 AND NO_ 2 -ON
l REALTIMETELEMETRY(R/T-TM) -ONjf DELAYEDTIMETELEMETRY(DiT-TM) -ON
STANDBYTELEMETRY(STBY-TM)-OFF
HF WHIPANTENNA-RETRACTEDSECONDSTAGE UHFWHIPANTENNA-RETEACTEDCUT-OFF DIPLEXERWHIP ANTENNA-RETRACTED
RECOVERYANTENNA-RETRACTEDDESCENTANTENNA-JRETRACTEDFLASHINGRECOVERYLIGHT-OFF
I ANTENNA USAGE.* )
UHFVOICET/R ""-_'P-/T-TM UHFSTUBANT.D.C.S, RCVRoNO.. 2RE-ENTRYC-BAND BEACONAND ANTENNA SYSTEM
NO5_ FAIBJNG_ _ ADAPTERC-BAND BEACONAND SLOTAI_EN NAJEFTISON
S/C SEPARATIONDIPLEXERWHIPANT,-EXTENDED(AUTOMATIC)UHFWHIPANT. -EXTENDED(AUTOMATIC)ACQUISITION AID BEACON-ON (AUTOMATIC)
SECONDSTAGEIGNITION ANTENNA USAGE:
UHFVOICE T/R _J
DP_.T.C-,'rMRCVR.NO, 2_ UHFSTUBANT,D.C.S° RCVR.NO. 1 "lACQ,."AID BEACON J_DIPLEXERWHIPANT.RE-ENTRYC-BAND BEACONAND ANTENNA SYSTEM
I ('_ORBIT: [_>ADAPTERC-BAND BEACONAND SLOTANTENNAVOICE: HFT/R-ONFIRSTSTAGECUT-OFF _ UHF NO. I T/R ( _ OR UHFNO. 2 TIR)-ON
AND JET]'ISON TRACKING:'_JADAPTERC-BAND BEACON( [_ OR RE-ENTRyC-BAND BBACON)-ON _[]ACQ° AID BEACON'ON (WHEND/T-TM IS OFP)_[]RECOVERYBEACON-OFF
(_ TELEMETRY: R/T-TM( _> ORSTBYTM)-ON'_]D/T-TM ( E_ OR STBVTM)*-ON (WHENACQ, AID BEACONIS OFF) _[]
COMMAND:
D.C.S. -ON
LIFT-OFF _ _ ANTENNA USAGE:T=O UHFVOICE T/R "l UHEWHIPANT°
E/T-TM _ ORD.C.S. RCVR.NO. 2) UHFSTUBANT.HFVOICE TE-ORBITALWHIPANTENNAACQ. AID BEACON "11D/T-TM _ DIPLEXERWHIPANT. .--._D°C.S. RCVR.NO. 1_1ADAPTERC.-BANDBEACONAND SLOTANTENNA
_>RE-ENTRY C-BAND BEACONAND ANTENNA SYSTEM
Figure 9-3 Communication System Sequential Diagram (Sheet 1 of 2)
9-6
SEDR 300
RETROFIRE RETROADAPTER
(TR) SEPARATION TR+ 1310SEC.TR+ 1775SEC.
JETTISONED: _ _ '_ __
UHFWHIPANTENNA & _'_ D __L_
ORBITALHF WHIPANTENNA NOM PREDICTEDUHFCO_. BLACKOUTPERIO DROGUECHUTEDEPLOY
RE-ENTRYPREPARATION: %
(_ [_ HF T/R-OFF
ANTENNA USAGE: %.UffF VOICE T/R NO. 1 ( [_ UHFVOICE T/R NO. 2)I UHFP-/T-TM ( E_ STBYTM) ,_ STUBD.C.S. RCVR.NO. 2 ,) ANT*ACQ. AiD BEACON _1
D/T-TM _ DIPLEXERWHIPANT.D.C.S. RCVR.NO. I PILOTCHUTEDEPLOYRE-ENTRYC-BAND BEACONAND ANTENNA SYSTEM
EQUIPMENTADAPTERSEPARATION:
EQUIPMENTJETTISONED_D.C.S. JETTISONED:
D/T-TM UHFSTUBANTENNA RENDEZVOUSAND RECOVERYDIPLEXER SECTIONJETTISONDIPLEXERWHIPANT°C-BAND SLOT ANTENNAADAPTERC-BAND BEACONACQ. AID BEACON
\COAXIAL SW. NO. 2
ANTENNA USAGE:
UHFVOICE T/R NO. I ( [_ UHFT/R NO. 2)'k UHFR/T-TM ( _ STBY-TM) J STUB.ANT.
/_-_, RE-ENTRYC-BAND BEACONAND ANTENNA SYSTEM
(_MAIN CHUTETWO POINTSUSPENSION: DURING THISPERIOD
ANTENNA USAGE: COMMUNICATION ISLIMITEDTOC-BAND BEACON
DESCENTANT. EXTENDED(AUTOMATIC) WHILEANTENNASRECOVERYANT.-EXTENDED _,_UTOMATIC) AREAUTOMATICALLyUHFRECOVERYBBACON'-ON (MANUAL SWITCH) SWITCHEDAND EXTENDED
UHFVOICE T/RNO. 1 ( [_ UHFT/R NO. 2) _ DESCENTANT. MAIN CHUTEDEPLOYP./T-TM( _> STBY-TM)RECOVERYBEACON-RECOVERYANT.RE-ENTRYC-BAND BEACONAND ANTENNA SYSTEM
RE-ENTRYC-BAND BEACON-OFFRECOVERY BBACON_N
UHFT/R NO. I ( []_ ORUHFT/R NO. 2),,ONHF T/R-ONHF/DF (TONEGENERATOR-ON) ORVOICETRS-OFFP_/T-TM( _ STBY-TM)-MANUAL OFFFLASHING RECOVERYLIGHT-ON (MANUAL SWITCH)RECOVERYHF WHIPANT. -EXTENDED(MANUAL SWITCH)RECOVERYANT. -EXTENDEDDESCENTANT. -EXTENDED
ANTENNA USAGEz
UHFVOICE T/R DESCENTANT. (_ tWO POINTSUSPENSIONHF T/RHF/DF OR VOICE_ RECOVERYHI: WHIPANTENNARECOVERYBEACON-RECOVERYANT.
_> BACK-UPEQUIPMENT(USEDtN THEEVENTOF A MALFUNCTIONIN THE PRIMARYUNIT) .
C_>EQUlPMENTTURNEDON ANDOEF OVERGROUND STATIONS MAIN CHUTEJEI"rlSONBYTHEFLIGHTCREWOR A GROUND COMMANDTRANSMITTEDTO THED.C.S. IN THESPACECRAFT.
_FREFERREDANTENNA-USED EXCEPTDURING ,_*'I__ LANDING/_ UNCONTROLLEDFLIGHT.
[_>USED DURING ORBITALMANEUVERSANDAS BACK-UPEQUIPMENT.
Figure 9-3 Communication System Sequential Diagram (Sheet 2 of 2)
9-7
._@ SEDR300 ___
PROJ E INI
VOICE COMMUNICATIONS
Voice communications is maintained by one hf and two uhf transmitter/receivers
and the Voice Control Center (VCC). The V_C has ell the necessary controls and
switches required for various keying modes, transmitter/receiver selection,
squelch, volume control, and voice recording. The hf voice transmltter/receiver
may also be used for Direction Finding (DF) purposes during the postlanding
phase of the mission.
An intercom connector is available for communication between the crew and recovery
testa,prior to opening the spacecraft hatches during the recovery phase of the
mission. Lightweight headsets are supplied for use when the spacesuit helmets
are removed during orbit, or during postlanding if the helmets or entire space-
suit is removed prior to recovery.
TELEMETRY TRANSMIT'fERS
Receiving inputs from the Pulse Code Modulated (PCM) programmer and the on-board
tape recorder, three telemetry transmitters transmit vital spacecraft system
parameters to the ground stations. The three transmitters operate on different
frequencies and are identified as real-time, delayed-time, and stand-by trans-
mitters. The stand-by transmitter is only used in case of real-time, or delayed-
time transmitter failure.
FLASHING RECOVERY LIGHT
The flashing recovery light, used during the postlanding phase of the mission,
contains its own power supply and improves visual spacecraft location.
DIGITAL CO_4AND SYS___
The Digital Command System (DCS) is the command llnk between the ground and the
9-8
SEDR 300
P R 0 J-'E'G-T GEMINI
spacecraft. The DCS consists of two uhf receivers, a decoder, and two relay
packages and Is operational from pre-launch until adapter equipment section
separation. Basically, the DCS receives and decodes two types of c......_nds: a
discrete or Real-Time Cammand (RTC) for spacecraft equipment utilization, and
Stored Program C_nds (SPC) that supply digital information to various space-
craft systems. Real-time commands operate DCS relays that control power directly
or energize relays in the spacecraft Electrical System that determine equilment
usage. Stored program c_-r;_nds are received and decoded for use by the Time
Reference System (TRS), or the computer.
SYSTH4 OPERATION
The C_-..-_nication System is semi-automatic in operation. The sequence and theory
of operation of the Cum_unication System is described in the following paragraphs
and referenced in Figures 9-2 and 9-3- Individual components are described in
System Units.
VOICE TAPE RECORDER
Voice tape recordings are made during the mission by placing the RECORD switch
on the VCC to the CONT or MGM position. The TONE VOX, AUDIO & UHF T/R i and 2
circuit breakers must be in the ON position. Each tape cartridge allows approxi-
mately one hour of recording time and is easily changed. An end-of-tape light
on the voice recorder illuminates for two seconds when two minutes of recording
tlme remains on the tape. The end-of-tape light will remain on when the end of
the tape is reached. A digital timing sigv_l is applied to one channel of the
tape for time correlation of the voice recording.
9-9
SEDR300 ___PROJECT GEMINI
PRE-LAUNCH
C-Band Radar Beacons
During pre-launch the BEACONS-C circuit breaker is placed to the ON position
to arm the C-RNTY and C-ADPT BEACON CORTROL switches. The C-RNTY switch is
placed in the CONT position during pre-launch to enable the re-entry C-band
beacon to reply when properly interrogated by a ground station. The
C-ADPT switch is placed in the CMD position during pre-launch. The CMD position
enables the ground station during launch, to activate the adapter C-band beacon
via a DCS channel if the need arises. After the adapter C-band beacon is acti-
vated, it will reply when properly interrogated by a ground station.
The C-band antenna system, used with the re-entry C-band beacon, is energized
when the ANT SEL switch is placed in the R_TY position. The ART SEL switch is
armed when the COAX CBTL circuit breaker is positioned to ON. The ANT SEL switch
controls application of po,er to the phase shifter power supply in the C-band
antenna system.
UHF Transmitter/Receiver
The number 1 ,,h_ voice transmitter/recelver will be utilized during pre-launch
unless some malfunction occurs in which case the number 2 transmitter/receiver
can be selected. For operation of either uhf voice transmitter/receiver_ stand-
by power is applied through the AUDIO & UHF T/R circuit breakers i and 2, which
must be in the ON position. The selected transmltter/receiver will be powered
by placing the UHF select switch to the number 1 or number 2 position and
the MOI_ switch of number 2 AUDIO to the UHF position. The UHF select switch also
controls coaxial switch 1 which connects the uhf transmitter/receiver to thef_
quadriplexer. Coaxial switch power is obtained from the c_-,on control bus
9-10
SEDR300
through the ON position of the UHF _AY circuit breaker. The method of key-
ing the ,,bf transmitter/receiver is selected by positioning the KEYING switch
on the VCC to VOX (voice operated relay), PTT (push-to-talk), or CONT INT/PTT
(continuous intercom/push-to-talk transmitter keying).
The desired antenna usage is obtained by placing the ANT CI_ circuit breaker
to the ON position. This places coaxial switch 3 to position i; thus connecting
the quadriplexer to the IN position of the coaxial switch 5- Coaxial switch 5
was placed in position i when the ANT _. switch was placed in the _ position
during the C-band beacon operation. With coaxial switch 5 in position i, the
uh_ stub antenna is available for ,,bf voice transmission and reception. Prior
F_ to _bilical release, voice c_munication is maintained between the spacecraft
and the ground complex through a hardline using the headset and microphone ampli-
fiers of the VCC. After umbilical release, voice transmission to the ground
complex is accsmplished by means of the uhf voice transmitter/receiver.
Real-Time Telemetr_ Transmitter
The real-time telemetry transmitter will be operating during the pre-launch
phase of the mission. The real-time telemetry transmitter is powered by placing
the RT _MTR circuit breaker in the ON position and placing the _ CONTROL switch
to the R/T & AC_ position.
The real-time telemetry transmitter uses the ,,hf stub antenna via the quadz_-
plexer and coaxial switches 3 and 5, the same as the uhf transmitter/receiver.
In case of real-time telemetry transmitter failure, the stand-by telemetry
transmitter may be used for real-time transmission. The STBY XMTR CNTL and
9-11
SEDR 300 ____PROJECT GEMINI
circuit breakers must be in the ON position to operate the stand-by telemetry
transmitter. If selection is made by the crew, the STBY TM CONTROL switch is
placed to the R/T position. Selection can be made by a ground command via the
DCS when the _4 CONTROL switch is in the OFF position. When operating as the
real-tlme telemetry transmitter, the stand-by transmitter uses the stub-antenna
for transmission.
N0n-Operation Components
The following Communication System components will be non-operational during
the pre-launchphase of the mission. To assure the off condition of these
components, the following switches should be in the position specified below:
(onvcc) OFF
BEACON CONTROL - RESC OFF
HFANT OFF
To assure proper sequential actuation of the various communication components,
the following circuit breakers (in addition to those previously described) must
be placed to the position listed prior to launch:
WHIP ANTENNAS - HF
WHIP ANTENNAS - UHF
WHIP ANTENNAS - DIPLEX
HF T/R - ON
]_EACONS - ACQ
BEACONS - RESC
XMTRS - DT
TAPERCDR - CNTL
9-12
,7'. _.__ SEDR 300
PROJECT GEMINIf_
SPACECRAFT/LAUNCH _EHICLE SEPARATION
Equipment usage after spacecraft/launch vehicle separation is identical to that
described under Pre-Launch except for the following: Upon closure of any two of
the three spacecraft separation sensors the acquisition aid beacon is energized.
The uhf whip ante_n_ solenoid actuators are powered and release the latch mecha-
nism of the ,h_ whip antennas, allowing them to self extend.
The acquisition aid beacon transmits via the diplexer and ,h_ whip ante-_a on
the adapter equipment section. Placing the TAPE RCDR-CTfL circuit breaker to ON
and the _I_CORTROL switch to R/T & ACQ during pre-launch places coaxial _wltch
2 in position i which connects the acquisition aid beacon to the diple_er.
./_- ORBIT
During orbit, operation of the telemetry transmitters and beacons will normally
be controlled by ground c_nds via DCS channels. To operate from ground com-
mands the C-ADPT, C-RNTY and T/M CONTRO_ switches must be in the CHD position.
HF Voice Transmitter/Receiver
During orbit hf communications is via the hf whip antenna on the adapter retro-
grade section. At insertion the adapter hf whip is extended by placing the
LANDING switch to the SAFE position, and the HF ANT switch to the EX_ position.
This will place coaxial switch _ in position 2 and allow hf voice transmission
and reception via the adapter hf whip. After extension (approximately one minute)
the HF ANT switch is returned to the OFF position.
Stand-by power is applied to the hf transmitter/receiver by the SF T/R circuit
breaker which was positioned to ON during pre-launch. The hf tranmmltter/receiver
9-13
__ SEDR 300 _____
PROJECT GEMINI
is powered by positioning the HF select switch (on the VCC) to RNTY and
audio MODE switch i or 2 to the HF position. The method of keying the hf
transmltter/receiver is selected by positioning the KEYING switch on the VCC
to VOX, PTT, or COBT IRT/PTT. During orbit, any of the three keying modes may
be selected.
URF Voice Transmitter_eceiver
The uhf voice trsn_,It_er/receiver operation is identlcal to that described
under Pre-Launch with the following exception. Preferred antenna usage during
orbit for uhf trsn_,4_sion and reception is via the adapter retrograde uhf whip
antenna. The retrograde uhf whip antenna is selected by placing the ANT SEL
switch to the ADPT position which places coaxial switch 5 to position 2. Although
preferred uhf tran_-1_sion and reception is via the retrograde uhf whip antenna,
the uhf stub antenna may be used during orbit by placing the ANT SEL switch to
the RRTY position.
Delay Time Telemetr_ TransmiSter
The acquisition ald beacon operates continuously during the orbital phase of
the mission except when the delayed-time telemetry transmitter is operating_
When the ground station receives the acquisition aid beacon signal, it initiates
a DCS command for the delayod-time telemetry transmitter to transmit data stored
by the on-board rec_er while the spacecraft was between ground stations.
Delayed-time tran_-_ssion may also be initiated by placing the T/M CONTROL switch
to the R/T-D/T position. This will initiate real-time as well as delayed-time
telemetry tra_ salon.
9-14
_. SEDR300 ____
PROJECT GEMINI
Real-time and delayed-time transmission will nox_nallybe initiated from the
ground station via DCS channels. At the time the delayed-time telemetry
transmitter is selected, the acquisition aid beacon is turned off and coaxial
switch 2 is placed in position 2, allowing telemetry transmission via the
diplexer and uhf whip antenna on the adapter equipment section.
As the spacecraft goes out of range the delayed-time telemetry transmitter is
turned off and the acquisition aid beacon res_es transmission. This is
normally performed by the ground station but may be accomplished by placing
the T/M CONTROL switch to the C_D, or the _T & ACQ position. If the _T &
AOQ position is selected, the delayed-time transmitter is turned off and the
real-time transmitter and the acquisition aid beacon begin transmitting. If
_ the CMD position is selected, only the acquisition aid beacon will operate;
however, the ground station has the capability of energizing the real-time tele-
metry transmitter via a DCS channel.
Any of the three previously described methods of disabling the delayed-time tele-
metry transmitter will place coaxial switch 2 to position I, and allow acqui-
sition aid beacon transmission via the diplexer and uhf whip antenna.
The stand-by telemetry transmitter may be used for delayed-time transmission
should failure of the delayed-time telemetry transmitter occur. The stand-by
transmitter is switched to delayed-time transmission by a ground c_-,-_ndvia a
DCS channel (if the STBY _ CONTROL switch is in the C_F position), or by placing
the STBY _ C01_I_OLswitch to the D/T position. Delayed-time transmission via
the stand-by telemetry transmitter uses the ,h_ stub or the uhf whip antenna on
the retrograde adapter depending upon the setting of the ANT SEL switch.
9-15
___ SEDR3O0 ____ ____
PROJECT GEMINI
Real-Time Telemetr_ Transmitter
Orbital operation of the real-time telemetry transmitter is similar to that of
the delayed-time telemetry transmitter in that the real-time telemetry trans-
mitter is operated only during the period that the spacecraft is within range
of a ground station. The real-time telemetry transmitter is turned on by a
DOS c_ud from the ground station or by placing the T/M COl_l_£_switch to the
R/T & ACQ or _T-D/T position. Real-time transmission is by the ,,bfstub or the
retrograde ,b_ whip antenna, depending upon the position of the AR"_SET.
switch. In case of failure of the real-time telemetry transmitter, the stand-
by transmitter may be used for real-time transmission. The stand-by transmitter
is switched to real-time transmission by a ground c_d via a DCS channel (if
the STBY _4 COI_OL switch is in the OFF position), or by placing the STBY _4
C0_ROL switch to the R/T position. The stand-by telemetry transmitter trans-
mits via the _Jhfstub or the retrograde uhf whip antenna, depending upon
the position of the _ _cB_!.switch.
It should be noted that the stand-by telemetry transmitter may be used for
delayed-time, or real-time transmission, but mmy not be used simultaneously for
both. In the event that both the real-time and delayed-time transmitters fail,
it is up to the ground station to determine the purpose for which the stand-by
transmitter will be used.
C-Band P_dar Beacons
During orbit, the C-band beacons are used only while the spacecraft is within
range of a ground station. Normally, the adapter C-band beacon will be used
during stabilized flight and the re-entry C-band beacon used during roll maneuvers.
9-16
__ SEDR300
PROJ EC'T" GEMINI
Operation of the beacons is similar to that described under Pre-Launch. The
C-RNTY and C-ADPT BEACON CONTROL switches are normally kept in the C_ position.
When the spacecraft comes within range of a ground station, as determ_ued from
the acquisition aid beacon sisnal, power to the desired C-band beacon is applied
by ground comm-_d via a DCS channel. The desired beacon may also be selected by
placing the C-ADPT or C-RNTT _EACON CON_ROL switch to the CORT position. After
power is applied_ the selected C-band beacon will transpond when properly inter-
rogated by a ground station. When the re-entry C-band beacon is selected, the
ANT SEL switch should be placed in the RNTY position to energize the phase
shifter and provide optimum radiation coverage.
SEPARATION
_ Prior to adapter equipment section separation, the re-entry module antennas are
selected by placing the T/M CONTROL switch to R/T & ACQ, the ART SEL switch to
R_PI"/,and the C-RNTY ]_ACON CORTROL switch to the CO_T position. Tr-ngmission
and reception during re-entry is via the C-band antenna system and the uhf stub
antenna. The acquisition aid beacon will operate until it is Jettisoned with the
adapter equipment section. The hf voice co,m_un_cations is disabled by placing the
HF select switch to the OFF position. On spacecraft 5, the hf whip on the adapter
retrograde section will remain extended. On later spacecraft, the hf whip may
be retracted by holding the HIP ANT switch in the RET position for approximately
i.5 minutes for complete retraction.
The following c@._,unications components will be Jettisoned with the adapter sec-
tion¢
_ Digital Co._aud System (DCS)
Delayed-time telemetry transmitter
9"17
00
Diplexer
C-band annu!_r slot antenna
Adapter C-band radar beacon
Diplexer ,_b_ whip antenna
Acquisition aid beacon
Co_al switch 2
T_is 1_mlts telemetry data transmission to real-time, voice conmmnication to uhf,
and tracking data to the re-entry C-band beacon.
Following equipment section separation and retro firing, retrograde section sepa-
ration will occur at which time the retrograde uhf whip and the adapter hf whip
antenn_ w_ _I be Jettisoned.
RE-Eh_RY
During the re-entry phase of the mission, two short duration communication black-
out periods occur. The first period, from approximately 1310 seconds after
retrofire time (TR) to 1775 seconds after TR, is caused by an ionization shield
around the spacecraft. This ionization is due to the extremely high temperatures
created upon re-entry into the earths atmosphere. The second blackout period
occurs at Rendezvous and Recovery (R & R) section separation when the uhf stub
antenna is Jettisoned. This period is terminated at two-point suspension which
occurs shortly after main parachute deployment.
At R & R separation, energized parachute deploy time delay relays energize coaxial
switch 3, placing it to position 2. This makes the uhf descent antenna avail-
able for real-time telemetry transmission and uhf voice communications.
9 -18
___ SEDR 300 ____ ___
PROJECT GEMINI
At two-point suspension the uhf recovery and uhf descent antennas are auto-
matically extended. The ,,_ recovery beacon is turned on by placing the RESC
BFAC(_ CONTRC_ switch to the W/O LT position.
Antenna usage during re-entry is as follows: Prior to R & R separation, real-
time telemetry transmission and ,,hf voice e_,,-_cation is via the ,,h_ stub
antenna. After two-point suspension, the uhf descent antenna is used instead
of the uh_ stub. The re-entry C-band beacon and C-band antenna system is used
for tracking and the uhf recovery beacon will use the uhf recovery antenna.
LANDING THROUGH RECOVERY
Upon impact the main parachute is Jettisoned by actuating the PARA JETT switch.
This extends the flashing recovery light. The light is energized by eha-_ging
the RESC BEACON CONTROL switch from the W/O LT position to the ON position.
The re-entry C-band beacon and real-time telemetry transmitter is turned off
by placing the C-RNTY BEACON CONTROL and the T/M CGRTROL switch to the CMD
position. If the stand-by telemetry transmitter was selected for real-time
transmission, the stand-by transmitter will be turned off by placing the STBY
TM CONTROL switch to the OFF position.
The recovery hf whip antev-_ is extended by placing the HF ANT switch to the
PST LDG position for spacecraft 5, or on later spacecraft by holding the }IF
ANT switch in the EXT position for approximately one minute. Voice c_un_ca-
tion via the hf transmltter/recelver is then possible by placing the HF select
switch to the RRTY position and either MODE switch to HF. The hf transmitter/
receiver can also be used to transmit a direction finding signal by placing
either MODE switch to HF/DF.
9 -19
__ $EDR300 ___._
PROJECT GEMINI
During the recovery phase of the mission the uhf rescue beacon transceiver may
be connected to the uhf recovery antenna. The uhf recovery beacon can be turned
off by positioning the RESC BEACON CONTROL switch to OFF. Lightweight headsets
are provided to replace the spacesuit helmets if the helmets or spacesuits are
removed and the crew remains inside the spacecraft. A recovery team disconnect
is used for intercom conversation between the crew and recovery team prior to
opening the spacecraft hatches.
SYSTEM UNITS
ANTENNAS
UHFDescent and U]IF Recovery Antennas
Purpose: The 1_bfdescent antenna is used for simultaneous transmission of the
real-time and stand-by telemetry transmitters, and transmission and reception for
the uhf voice transmitter/receiver. The uhf recovery antenna provides trans-
mission capability for the uhf recovery beacon. The two antennas are used from
two-point suspension of the main parachute through final recovery of the space-
craft.
Physical Characteristics: The two antennas, being similar in physical appearance#
are shown in Figure 9-4. Both antennas are mounted in the parachute cable trough
where they are stowed until main parachute t_o point suspension during the landing
phase of the mlssion.
The antenna element consists of two one-half inch wide gold plated steel blades
bolted together at two places. The uhf descer_ antenna is approximately lg
inches long. The uhf recovery antenna is approximately 18 inches long.
9-20
__ SEDR 300 __
PROJECT GEMINI
ANIENNA •ELEMENT _ _ ANTENNA
_R%%E._(@OLE EL_E'---_
PARACHUTE CABLE 1
TROUGH
SEAL_ _ ! _/
JDESCENT ANTENNA
PARACHUTE BRIDLE
RETAINING STRAP(TYPICAL 2 PLACE
(STOWED POSITION) RECOVERY ANTENNA
COAXIAL
(STOWED POSITION)
IDLETROUGHCOV_
MAIN CONSOLE
I PARACHUTE BRIDLETROUGH
LOWER CONSOLE DISCONNECT
Figure 9-4 UHF Descent and Recovery Antennas
9-21
___ SEDR300 ___ .__
PROJECT GEMINI
Mechanical Characteristics: For rigidity, the antenna element is shsped in a
0.5 inch wide src having a radius of 1.5 inches. The two laminations of steel
blades, compounding a single antenna element, are rigidly secured at the lower
half of the antenna. To allow a slight displacement of the two laminations with
respect to each other during stowage and deployment, two nuts and bolts placed
through elongated holes secure the two IA_inations together at the upper half of
the antenna element.
The antennas are bent towards the small end of the spacecraft for stowage and
are held in place by a retaining strap. The strap is broken when the Landing
System shifts from single point to two point suspension, allowing the antennas
to extend.
Each of the two antennas have a radiation pattern which is identical to that
of a quarter wave stub. _-_
UHF Stub Antenn8
Purpose: The uhf stub antenna (Figure 9-5) allows simultaneous transmission of
the real-time and stand-by telemetry transmitters, transmission and reception for
the uhf voice transmitter/receivers, and reception for DCS receiver number 2. The
antenna may be used from pre-launch until separation of the R & R section during
re-entry, but is normally used from pre-launch to insertion and from re-entry
preparation to R & R section separation.
Physical Characteristics: The uhf stub antenna, physically constructed as
illustrated in Figure 9-5, is mounted in the nose of the R & R section. The
antenna protrudes forward from the R & R section and is covered by a nose fairing
during the boost phase of the mission. The antenna consists of a mast and base
which weighs approximately i.I pounds. The mast is constructed of 3/_ inch cobalt
steel, machined to tubular form, and covered by a Teflon ablation shield for
9-22
ABLATION SHIELD ABLATION SHIELD $OCKI PIN SPACER
,__ CONNECTOR
L-- CABLE ASSEMBLY
Figure 9-5 UHF Stub Antenna
9-23
SEDR300
GEMINi
protection during re-entry. The antenna is approximately iS. 5 inches long
including the connector, and X.25 inches in diameter over the ablation material.
The mast consists of two sections. The front section is mounted on a cobalt
steel ball Joint and retained to the rear section by a spring loaded cable.
Electrical contact between the mast sections is made through the ball Joint
and the spring loaded cable assembly. The ball Joint allows the front section
of mast to be deflected to approximately 90 degrees in any direction around the
antenna axis. The spring of the cable assembly is pre-loaded to approximately
_5 pounds to cause the front section, when deflected, to return to the erected
position.
The rf connector is press fitted into a socket and makes contact to the mast
through the socket and sleeve, which are the same material as the mast. The
shell of the rf connector is mounted to the base which is isolated from the mast
by a Teflon spacer and sleeve.
Mechanical Characteristics: The uhf stub is a quarter wav_ length antenna. The
radiating length of the antenna is approximately 11.2 inches.
UHF Whip Antennas
Purpose: Two identical uhf whip antennas (Figure 9-6) supply the required ,Jbf
transmission and reception facilities during orbit. One of the uhf antennas is
located on the adapter equipment section and serves the DCS receiver number i,
and the acquisition aid beacon or delayed-time telemetry transmitter. The
second uhf antenna, mounted on the adapter retrograde section, serves the real-
time and standby telemetry transmitters, the uhf voice transmltter/receivers,
and DCS receiver number 2.
9-2_
SEDR 300
POWERCONNECTOR
COVER_ COAXIALCONNECTOR
_. COONTERWE'OHT--_ "_LATCH
-- LATCHRETAINER
LATCH CATCH
[ ._--COVER_O_E_O,DJ
RELEASEMECHANISM(SHOWN IN LATCHED CONDITION)
EXTENSION OF ANTENNA
: ANTENNA FULLY EXTENDED
Figure 9-6 UHF Whip Antennas
9-25
__ 5EDR3O0 _.__
PROJECT GEMINI
Physical Characteristics: The ,,he whip a_ten-a is self extendable and requires
no power other than that required for initial release. The antenna element Is
a tubular device made from a 2 inch wide berylliu_ copper strip processed in
the form of a tube. _e antenna, when fully extended, forms an element that is
approximately 12 inches long and i/2 inch in diameter. During stowage, the
tube is opened flat, wound inside of a retaining dr_n, and latched in position.
Upon release of the latch by a solenoid, the extension of the antenna depends
entirely on the energy stored in the rolled strip material. This energy is
sufficient to erect the antenna at a rate of 5 feet/second into its tubular
form. In the stored condition, the antenna is flush with the outer skin of the
spacecraft.
Mechanical Characteristics: The antenna element is retained inside the housing "_
by a metal lid. A metal post Is attached to the lid and passes through the
center of the coiled antenna. The bottom of the post is grooved to accept a
forked latch which holds the catch post assembly firmly in position prior to
release. The forked latch is attached to a miniature pull-solenoid which is
spring loaded in the extended position to ensure that launch shock and vibration
loads will not cause inadvertent antenna extension. _hen a voltage from the
Sequence System is applied to the antenna solenoid, the latch will be withdrawn
allowing the antenna cap to eject and the antenna to extend. As the catch post
assembly is ejected, a microswitch in series with the solenoid coll opens the
circuit to the coil to prevent further current drain from the power source.
9 -26
__. SEDR 300 ____
PROJECT GEMINIS _
The two antennas are Jettisoned with the corresponding adapter section.
HF Whip Antennas
Purpose: The hf whip antennas provide transmission and reception for the hf voice
transmitter/receiver during the orbital and postlanding phases of the mission.
Physical Characteristics : The hf whip ante-n_s are physically constructed as
illustrated in Figure 9-7. The recovery hf whip antenna is mounted on the small
pressure bulkhead, outside the pressurized area of the spacecraft re-entry module.
The other hf whip antenna is located on the adapter retrograde section. The antenna
mechanism housing, approximately 6.25 inches wide and 22._ inches high, completely
encloses all parts of the antenna_ including storage space for the antenna elements.
_ The recovery hf whip antenna contains six elements which, when fully extended,
comprise a single antenna mast appro_mately 13 feet 3 inches long. The adapter
hf whip antenna contains three elements which, when fully extended t comprise a
single antenna mast approximately 16 feet long on spacecraft 5 and 6, and approxi-
mately 13 feet long on later spacecraft. The mast is one inch in diameter on
all spacecraft. Two connectors, supported by the anten-a body, provide a means
of applying power and connecting the antenna to the rf connector on the hf voice
transmitter/receiver. The recovery hf whip antenna weighs approximately 9.0 pounds.
The 16-foot version of the adapter hf whip antenna weighs approximately 7.5 pounds
and the 13-foot version 6.0 pounds. The main supporting structure of the antenna
mechanism housing is the antenna body consisting of a thin fiberglass shell.
9-Z7
F-_ ..oJ_cTGEM.N.
IN LIMITSWITCH
III i rENNA
| I] CONNECTOR
II I
,_ SWITCH
/ ,/III ,OWER/
OARD
\
MOTORC
COVER
2.ASSETTE ROLLERS i
Figure 9-7 HF Whip Antennas
9-28
___ SEDR 300
PROJECT GEMINIf_
The outer shell is made in two sections which mate together and form a completely
sealed envelope around all moving parts. The antenna mast elements are heat
treated stainless steel strips and are stored in adc motor driven cassette.
Mechanical Characteristics : The strip material comprising the antenna elements
is heat treated into a material circular section in such a manner that the e_es
of the material overlap approximately 180 degrees. When the antenna is retracted,
the tubular elements are continuously transformed by guide rollers into a flattened
condition, and stored in a strained manner in a cassette. Extension and retrac-
tion of the antenna is accomplished by a motor which, by means of a chain, drives
the storage cassette core. Because of the natural physical shape of the antenna
elements, the antenna has a tendency to self-extend; thus giving an extension
time of approximately 25 seconds. Retraction time is approximately 40 seconds.
The antenna is stopped within its desired ]_m_ts by two microswitehes, one for
extension and one for retraction, which automatically cut the power applied to
the motor at the time of extreme limits of the antenna are reached.
The rf connection to the antenna is obtained by a wiper arm sliding on the cas-
sette core drive shaft.
On spacecraft 53 the hf whip antennas sre operated as follows : Spacecraft control
bus voltage is supplied through the WHIP ANTENNAS -HF circuit breaker to the HF
ANT switch. The adapter hf whip antenna is extended during orbit by positioning
the HF ANT switch to EXT. The adapter hf whip antenna is not retracted during
9-29
SEDR300 _.___[
PROJ EC"T" GEMINI
orbit, but is Jettisoned in the extended position with the retrograde section.
After landing, the recovery bf whip antenna is extended by positioning the HF
ANT switch to PST LDG, and is retracted by positioning the HF _ switch to EXT.
On spacecraft 6 thro_ 12, extension of the hf whip antep_nRsis controlled through
the HF ANT switch A,_ LAWllZNGswitch. The hf antennas are operated as follows:
Spacecraft control bus voltage is supplied through the WHIP ANTENNAS - HF circuit
breaker to the HF ART switch, which has momentary type contacts. During orbit,
the LANDING switch is in the SAFE position and adapter hf whip antenna can be
extended or retracted by holding the I_ ANT switch in the EXT or RET position
respectively. During re-entry, the LANDING switch is placed in the AI_ position.
After landing, the recov_X'y hf whip antenna can be extended or retracted by
holding the HF ANT switch in the _ or SET position respectively. The HF AFf _
switch should be held in the EXT position for approximately one minute for full
extension of the antennas, and in the SET position for approxlwately 1.5 minutes
for full retraction.
C-Band Annular Slot _utenr_
Purpose: The C-band A-n_lar slot antem_ (Figure 9-8) serves the adapter C-band
radar beacon and is normally used during stabilized flight.
Physical Characteristics: The C-band annular slot antenna is mounted on the
equipment section of thp adapter. The physical construction is such that the
antenna is flush with the outer skin of the spacecraft. The antenna is approxi-
mately I._ inches in diameter, 1.35 inches long and weighs 8 ounces maximum.
9-30
.._. SEOR3OO _--__--_/) PROJECT GEMINI
!iliJiiiiiiiiiiiiiiiiiiiiiii!iiiiii
-- C-BAND HELICAL iiii;iiiiiiiii
iiiiiii• iiiiiii
iiiiiiiiiiiiiii!!iiiiiiiiiiiiiiiiiiiiiiiii
iiiii!iiiiiiiiiiiiiii
iiiii!i
_WER DIVIDER -- iiiiiii
L iiilgiiPHASE SHIFTER i!iii!iPOWER SUPPLY ii_i!!i ,f
iiiiiii_ .iiiiiiiiii;iiiig!iiii
iii[iii _ii!iiii
iiiiiii PHASESHIFTERPOWER SUPPLY
_iiiiil
_.A_E_.,E,_ iiiiiii_iiiii!iliiii!
POWER DIVIDER C-BAND HELICAL ANTENNA
Figure 9-8 C-Band An_nna Sys_m
9-31
.__ $EDR 300 _._
PROJECT GEMINI
Mechanical Characteristics: _he antenna radiation pattern is identical to that
of a quarter wave stub on a ground plane. The antenna is used for both reception
and transmission of the adapter C-hcud radar beacon during the orbital phase of
the mission. The _ute_a is Jettisoned with the equipment section of the adapter.
C-Band Antenna _stem
Purpose: The C-hand ante_ system, consisting of a power divider, a phase
shifter, and three helical antennas, provides transmission and reception capa-
bility for the re-entry C-band radar beacon. The power divider supplies equal
transmission power to the three helical antennas. A phase shifter is in series
with one of the antennas to ccmpensate for areas of low or no radiation coverage
between lobes of the three individual radiatio_ patterns. A phase shifter power
supply supplies the phase shifter with 26 vac 453 cps power. The antenna system ....
gives the circ,,1_rradiation pattern around the spacecraft longitudinal axis
required for ascent, descent and roll spacecraft attitudes.
Physical Characteristics: The power divider, phase shifter, phase shifter power
supply, and helical antennas are shown in Figure 9-8. The power divider, phase
shifter, and phase shifter power supply are mounted on the small pressure
bulkhead, outside the pressurized area of the spacecraft. The power divider
measures approximately 3.86 inches over the connectors, 4.0 inches over the tun-
ing knobs and weighs appro_tely 6.5 ounces. The phase shifter is approximately
5.8 inches long, 2.8_ inches wide at the large end, 1.4 inches high, has a
diameter at the small end of about 1.5 inches, and weighs approximately 12 ounces.
The phase shifter power supply measures appro_mately 1.5 inches wide, 1.75 inches
high, 3.5 inches long over the eo_ucctor_ and weighs approximately 8 ounces.
9-32
___ SEDR 300
PROJECT GEMINI
The three C-baud helical antennas are mounted flush with the outside skin of the
spacecraft and spaced approximately 120 degrees apart. Each antenna unit is approx-
imately 3.4 inches long, 1.8 inches wide, has a depth of 2.21 inches over the
connector and weighs approximately 3.5 ounces.
Electrical Characteristics : The power divider, phase shifter, and helical antennas
comprise an antenna system that satisfies the transmission and reception require-
ments for the re-entry C-band radar beacon during the launch a_ re-entry phases
of the mission.
The power divider is basically a cavity type power splitter. During beacon trans-
mission, power is delivered to the power divider where it is divided equally among
the C-band helical antennas. The power divider compensates for loss of power
due to the phase shifter in series with the right ante,_a. The power divider co_-
+_ins a double stub tuner to compensate for mismatch between the re-entry C-band
beacon, the C-band helical antennas, and the phase shifter. Tu,lug is accomplish-
ed by means of a self-locking tuning shell located underneath each tunlng stub cap.
The phase shifter has its own ac power supply. The input to the phase shifter, is
half wave rectified and applied across a coil wound around a ferrite material.
Due to the characteristics of the ferrite material, the rf signal from the power
divider is delayed 0 to 180 degrees + 20 degrees at the rate of 453 cycles per
second. The changing phase shift of the rf power on one of the C-band helical
antennas with respect to the other two, shifts the lobe of that antenna by approx-
imately + 45 degrees; thus giving the effect of an almost ideal circular radiation
pattern around the longitudinal axis of the spacecraft. The combination of the
three antenna elements gives a radiation pattern which extends in all directions
except forward and aft of the spacecraft.
9-33
$EDR300 __PROJECT GEMINI
The phase shifter power supply is a dc-ac inverter which supplies a nominal 26 vac,
453 cps power to operate the phase shifter. The power supply is a hermetically
sealed solid-state unit consisting of a volt_e re_lator_ single-stage oscilAator,
buffer stage, and a push-pull output stage vith transformer coupled output. The
power supply provides a minimtm output of 21 volts rms at 453 + 17 cps with an
input voltage range fr_n 20 to 30 vdc. Input voltage is applied from the space-
craft main bus via the BEACON-C circuit breaker, C-RF_Y BEACON C0_TRCL switch and
the I_TY position of the ANT 8EL switch. Ma_Imtun input current is 370 m111iamperes.
Multiplexers (UHY Diplexer and UHF Quadri_lexer)
Purpose: The uh_ diplexer provides isolation between DCS receiver number i, and
the acquisition aid beacon or the delayed-time telemetry transmitter operating
into a c_umon antenna. The uhf quadrlplexer provides isolation between the stand-
by telemetry transmitter, the real-time telemetry transmitter, a ,,h_ voice trans-
mitter/receiver_ and DCS receiver n,_er 2 operating into a c_i._._onantenna via
coaxial switches.
Physical Characteristics: The physical representation and approximate location of
the uh_ diplexer and the ,,h_ quadriplexer is shown in Figure 9-9. The diplexer is
located on the electronic module of the adapter equipment section. The quad-
riplexer is located forward of the small pressurized bulkhead outside the pres-
surized area of the cabin.
The diplexer is approximately 4. 5 inches wide, 4 inches high_ and 2.7 inches deep;
contains two input and one output connectors, and weighs approximately 1.25 pounds.
The uhf quadriplexer is approximately 5.75 inches wide, 5.5 inches deep, and _.i
inches high; weighs approximately 2.75 pounds, and has four input and one output
connectors.
9-34
3oo_)F 1
STANDBYTELEMETRY TRANSMITTER
REAL-TIMETELEMETRY TRANSMITTER
J5 O'---- _
ANTENNAVIA COAXIALSWITCH
O J3 ' " 'DIGITAL COMMAND ".SYSTEM RECEIVER NO. 2
-T--_- T I_ o_0_vo,cETRANSCEIVERS
VIA COAXIAL
_'_ __
x\
_( 0s_ACQUISITION AIDBEACON ANDDELAYED - TIME
T TM TRANSMITTERJ3 O _ VIA COAXIAL
o UHF WHIP J SWITCH0 _._ ANTENNA
° ° '" _( o_C o DIGITAL COMMANDSYSTEMRECEIVI_ NO. I
/'_" U
Figure 9-9 UHF Diplexer/UHF Quadriplexer
9-35
SEDR 3O0
Electrical Characteristics: Figure 9-9 shows the schematic of the uhf diplexer
and the uhf quadriplexer. Each channel consists of a high Q cavity, tuued to the
corresponding operating frequency. All channels are isolated from each other with-
out appreciably attenuating the rf signals passing through it. Each channel can be
re-tuned if the assi_ed operating frequency is chs_ed. The diplexer isolates DCS
receiver number l, and the acquisition aid beacon or the delayed-time telemetry
transmitter, depending upon the position of coaxial switch n_ber 2. The diplexer
operates into the ,bf whip antenna on the adapter equipment section.
The uhf quadriplexer isolates the real-time telemetry transmitter, the stand-by
telemetry transmitter, one of the two uhf voice transmitter/receivers, and DCS
receiver number 2. The quadriplexer operates into one of the following three ubf
antennas, depending on the position of the coaxial switches in series with the
antennas: ,,hfstub antenna, uhf descent antenna, or the uhf whip antenna on the
adapter retrograde section.
Coaxial Switches
Purpose: Five coaxial switches are used to perform the following functions :
(i) select the acquisition aid beacon or the delayed-time telemetry transmitter
output as the input to the diplexer; (2) select one of the two uhf voice trans-
mitter/receiver outputs as the input to the quadriplexer; (3) connect the hf voice
transmitter/receiver to the adapter hf whip antenna on the retrograde section, or
to the recovery hf whip antenna on the re-entry module; (4) connect the output of
the quad_iplexer to the uhf descent antennAj or through coaxial switch 5 to the
-_f stub or the retrograde adapter uhf whip ante___r__.
9-36
___ SEDR300 -___
PROJECT GEMINI
ITEM FUNCTION
1 UHF TRANSCEIVERS
2 TELEMETRY TRAN SMITTER/ACQUlSI-TION AID BE_ON
3 DESCENT ANTENNA
_" 4 HF TRANSCEIVER
5 UHF WHIP/UHF STUB COPTERAND INDICATORCIRCUIT CONNECTOR
_.... _ _ _ _*:o_._:._._:_:_.'_%_'_ ._i_*_:_ _.,__;_'_._:_-_*_:_ :<.-_'%_ _._ _
+28V POSITION NO. 2
1 j o,J L_ ____
+28V POSITION NO. 1 g
INDICATOR CIRCUIT NO. 1 E I
SHOWN IN ENERGIZED PiN B POSITION
POWER AND iNDICATORCIRCUIT CONNECTOR
/
Figure 9-10 RF Coaxial Switches
9-37
SEDR 300
Physical Characteristics: The physical construction and approxlmate location
of the coaxial _witches is shown in Figure 9-10. The location of the switches
is as follows:
Coaxial switch I: approximately five inches from the small end of the cabin, in
the fourth quadrant.
Coaxial switch 2: appro_w_tely I0 inches from the forward (small) end of the
adapter equipment section, in the third quadrant.
Coaxial switch 3: appro_tely i0 inches from the small end of the cabin, in
the third quadrant.
Coaxial switch I_: loca_d adjacent to coaxial switch i
Coaxial svitch 5: approxt_tely 7 inches fram the small end of the cabin, in
the third quadrant.
Each switch co_tains a power connector, an input connector, two output connectors,
and weighs appr_mately 0-5 pounds. The dimensions of each switch are approxi-
matel_ 2.65 inches lo_, 1.82 inches high, and 1 inch wide.
Electrical Characteristics: The five coaxial switches are identical and may be
used interchangeably. Basically, the coaxial switches supply single pole double
throw switching action as illustrated in Figure 9-10. The switch, having a 20
millisecond ma_ operation time, operates on 3 amperes at 28 vdc and uses a
latching solenoid break-before-make switching action. The coaxial switches are
designed to operate _ 15 mc to 500 mc, and from 5500 mc to 5900 inc. Pins D
9-38
___ SEDR 300
PROJECT GEMINI
and E of each switch are brought out to AGE test points to permit monitoring of
the switch positions prior to lift-off. Pins A and B of each switch are utilized
to accomplish the switching action.
BE_ACONS
Re-entry C-Band Radar Beacon
Purpose : The re-entry C-band radar beacon provides tracking capability of the
spacecraft from lift-off to insertion and from retrograde to landing. The
re-entry C-band beacon may be used during roll maneuvers or in the event of
adapter C-band beacon failure°
Physical Characteristics : The re-entry C-band radar beacon is a sealed unit
which measures approximately 7.64 x 6.14 x 3.02 inches and weighs about 8.3
pounds. As shown in Figure 9-11, the beacon has power, antenna, and test con-
nectors. Located on the rear of the beacon are various adjustments for trans-
mitter, preselector, and local oscillator tuning. Solid-state modular circuitry
is used throughout the beacon with the exception of the transmitter magnetron
and the local oscillator. The beacon is mounted on the right forward equipment
bay, and uses the C-band antenna system for reception and transmission.
Electrical Characteristics: The re-entry C-band radar beacon is a transponder
which upon reception of a properly coded interrogation signal from a ground radar
tracking station, transmits a pulse modulated signal back to the tracking station.
By measuring the elapsed time between transmission and reception at the tracking
stations, and compensating for the time delay of the beacon, the position of the
spacecraft can be determined. The block diagram of the beacon is shown in/_
9-39
SEO300PROJECT GEMINI
_ ADAPTER C-BAND RADAR BEACON
TUNING
PRESELECTORTUNING
RE-ENTRY C-BAND RADAR BEACON
Figure 9-11 C-Band Radar Beacons
9-40
SEDR 300
PROJECT GEMINI
Figure 9-12. The signal arriving at the antemaa is routed through the directional
coupler to one half of a dual ferrite circulator. The ferrite circulator isolates
the transmitter from the receiver, allowing a si_le antenna system to be used for
both reception and transmission. The beacon utilizes a superhetrodyne receiver
which is tunable, by means of a three stage preselector, over a range of 5600 mc
to 5800 me. The assigned receiver center frequency is 5690 me.
The output of the preselector is combined with the local oscillator frequency
in the crystal mixer to produce an output intermediate frequency of 80 inc. The
local oscillator is of the metal-ceramic triode cavity type. The mixer contains
a ferrite circulator for isolation between the local oscillator, mixer and pre-
selector. The output of the mixer is ampliflod by three tuned intermediate
P_ frequency amplifier stages, followed by a video detector and a video preamplifier.
Additional amplification is obtained by a pulse amplifier whose output is supplied
to the decoder. The purpose of the decoder is to initiate triggering of the trans-
mitter after a correctly coded signal has been received. _e system delay, in
conjunction with the delay variation correctic_ circuitry, provides for a constant
fixed delay used in determining the exact position of the spacecraft. The beacon
incorporates a cw _w_unity circuit that prevents the transmitter from being
triggered by random noise. The noise level is reduced below the triggering level
of the transmitter by controlling the gain of the pulse amplifier. The trans-
mitter uses a magnetron and provides a one kilowatt peak pul_ modulated signal
at a frequency of 5765 mc to the power divider. _he beacon is powered by a dc-dc
converter employing a magnetic amplifier and silicon controlled rectifiers. The
converter provides voltage regulation for input voltage variations between L8 and
9 -_I
_@ SEDR 300 __
PROJECT GEMINI
ANTENNA
I7 VOLTAGd _L
PEAK PO
RECEIVER__._I RECEIVER TEST POINTS
COUPLER AMPLIFIER -- TRANSMITTERPRF
PEA)( FOWER T
(_ MONITOR _ TRANSMITTER PRF
t 1 1 ' I-- A_"ODUo_OTL "i OR,VERAND="SM'_"R O:E_:XT"RO
I_ I
T=NSM,.ER O'eR _ __ FREQUENCY INTERROGATION
,, t /-II H H t£ l
CRYSTAL PULSE SHAPERAND GATE SYSTEMCIRCULATOR I.F.
MIXER DECODER DELAY COINCIDENCE DELAY(DIPLEXEe)
CIRCULATOR C.W.PRE-SELECTOR (MIXER)
®®® I
RECEIVERFREQUENCY
LOCAL VARIATION VARIATIONOSC. CORRECTION _ CORRECTION
#2
® RECE,V_R ® @ ®FREQUENCYSLOPE ll SLOPE #2
UNE POWER VARIOUS_1 FILTER SUPPLY VOLTAGES
Figure 9-12 Re-Entry C-Band Radar Beacon Block Diagram
9-42
___ SEDR 300 _____
PROJECT GEMINI
32.5 vdc. The input to the converter is filtered by a pl-type filter to minimize
any line voltage disturbances.
Adapter C-Band Radar Beacon
Purpose. The adapter C-band radar beacon provides tracking capability of the
spacecraft during the orbital phase of the mission and is jettisoned with the
adapter equipment section.
Physical Characteristics : The adapter C-band beacon is a sealed unit and measures
approximately 9.3_ x 8.03 x 3.26 inches. As shown in Figure 9-i1, the adapter
beacon has a power and test connector, an antenna connector, and a crystal
current test point connector. The beacon contains external adjustments for local
oscillator, preselector (rf filter), and transmitter tuning; switches for selecting
the desired interrogation code, and one of two preset transponder fixed delay
times. These adjustments and switches are accessible by removing pressure sealing
screws. The beacon employs solid-state circuitry, except for the transmitter mag-
netron and receiver local oscillator. The adapter beacon is located on the
electronic module of the adapter equipment section and uses the C-band annular
slot antenna for reception and transmission.
Electrical Characterlsties: The adapter C-band radar beacon is a transponder,
which employs the same basic operating principles as the re-entry C-band beacon
to provide spacecraft location data upon receipt of a properly coded interro-
gation signal. A block diagram of the adapter C-band beacon is shown in Figure
9-13. The interrogation signal is fed from the antenna to the duplexer. The
duplexer is a ferrite circulator which couples the received signal to the rf
filter preselector and also isolates the receiver from the transmitter to permit
9-43
SEDR 300
• _ mll Im _ _ " II
ANTENNA
I J2 DUPLEXER TRANSMITTER-- -- _ PULSEFORMING = MODULATOR
I (CIRCULATOR) (MAGNETRON) NETWORK (PEN)
I, , ]i_I
jADe. 1 r --
' I l=l 'I I,P I_ PU_EEEOR_MOOO_TORI M,×ER= _MPL,F,_RI RES'ORER¢ONTRO'
I
, [ ' 1 1ILOCAL _ =1 DECODER :OSCILLATOR -" DRIVER
1 I I I
J u___ -]
DC/DC PULSE ASSEMBLYCGNVERTER
I REGULATOR AND
LINE FILTERS
L_ _ _ _ - r_24.5 ro 30v cc POWER
Figure 9-13 Adapter C-Band Radar Beacon Block Diagram
9-44
J
__ SEDR300
PROJEC:r GEMINI
use of a common antenna for reception and transmission. The superhetrodyne
receiver frequency is tunable from 5395 me to 5905 inc. The assigned operating
center frequency is 5690 mc and is selected by adjustment of the rf filter.
The rf filter is a three-stage preselector, employing three separately tuned
coaxial resonator cavities to provide adequate rf selectivity and to protect
the mixer crystal from damage due to transmitter power reflected by the antenna.
The output of the preselector is combined with the local oscillator output in
the mixer stage to provide a 60 mc output to the intermediate frequency amplifier.
The mixer consists of a coaxial directional coupler and a mixer crystal. The
directional coupler isolates the local oscillator output from the antenna and
directs it to the mixer crystal. The local oscillator is a re-entrant cavity type
employing a planar triode to generate the cw signal required to operate the mixer.
The intermediate frequency amplifier is a high gain amplifier composed of an input
stage, five amplifier stages, and a video amplifier. The amplified video output
is fed to the pulse form restorer circuits which prevent a ranging error due to
variations in receiver input signal levels, and also provides a standard amplitude
pulse to the decoder for each input signal exceeding its triggering threshold.
The decoder determines when a correctly coded signal is received and supplies an
output to the modulator driver. The type code to he accepted is selected by the
CODE switch. Single pulse, two pulse or three pulse codes may be selected. The
modulator driver and control circuits initiate and control triggering of the trans-
mitter modulator. The modulator driver supplies two fixed values of overall system
9-45
$EDR 300 ______PROJECT GEMINI
delay. The desired dela_r is selected by the position of the DLY switch. An al-
ternate value of maximum dela_ is available by removing an internal Jumper lead.
The modulator control furnishes the trigger And turn-off pulse for the modulator"
and limits modulator triggers to prevent the magnetron duty cycle from being
exceeded, regardless of the interroKating sign,! frequency. The modulator cir-
cuit employs silicon controlled rectifiers which function similar to a thyratron,
but require a much shorter recovery time.
The associated modulator Pulse Formlng Network (PFN) and transformer provide the
necessary pulse to drive the transmitter magnetron. The desired pulse width is
selected by the internal co--_ctlons made to the PFN. The transmitter _netron
frequency is tunable from _00 me to _ inc. The assigned transmltter center
frequency is 5765 inc. A m_n_m_m of 500 watts peak pulse power is supplied to the ....•
antenna under all conditions of rated operation.
The transponder power stq_p_Vcellists of input line filters, a series regulator,
and a dc-dc co_verter. The _ supply furnishes the required regulated output
voltages with the --_egulated input voltage between 21 and 30 vdc. The converter
e_ploys a multivibrator aM fnll wave rectifier circuits.
Acquisition Aid Beacon
Purposes Unltke the C-band beacons that supply accurate tracking data, the
acquisition aid beacon is mere_V a tr=-_tter used to determtue when the space-
craft comes within range of a ground tracking station. When the spacecraft comes
within the range of a ground tracki_ station 3 the acquisition aid beacon is dis-
abled and remains off until the spacecraft is again out of range.
9-_
___ SEDR300
PROJECT GEMINI
f_
Physical Characteristics: The acquisition aid beacon, shown in Figure 9-14, is
cylindrical, having a diameter of approximately 2.6 inches, and a height of
approximately 3.5 inches. The acquisition aid beacon is located as shown in Figure
9-I_. The beacon contains a power connector, a coaxial antenna connector and
weighs appro_mately 17 ounces.
Electrical Characteristics: The acquisition aid beacon consists of a transmitter,
de-de voltage regulator, and a low pass output filter.
The transmitter is an all transistorized unit, containing a push-pull output stage
to obtain a minim_ output of 200 milliwatts at a frequency of 2_6.3 me. The
transmitter frequency is derived by taking the basic frequency of an oscillator
and multiplying it through a series of tripler and doubler stages.
The transmitter is powered by a de-de voltage regulator. The regulator is com-
pletely transistorized and supplies a regulated output voltage of 28 vdc. To
reduce the probability of obtaining a spurious output signal, a band pass filter
is placed in the output circuit.
UHF Recover_ Beacon
Purpose: The uhf recovery beacon, operating on the international distress fre-
quency of 2_3 me, serves as a recovery aid by providing information regarding
location of the spacecraft.
Physical Characteristics: The uhf recovery beacon and its approximate location
is shown in Figure 9-1_. The beacon is mounted on the aft right equipment bay
of the spacecraft re-entry module. The beacon is approximately 9.0 inches long,
9-47
SEDR 300
RE OUTPUT
CONNECTOR
_---FILTER
,,/_ TRANSM11T ER
S POWER SUPPLY'_'-_ POWER CONNECTOR _"
ACQUISITION AID BEACON
UHF RECOVERY BEACON
Figure 9-14 Acquisition Aid and UHF Recovery Beacons
9-48
___ SEDR300
PRO,JEC GEMINIf_
4.0 inches wide, 2.5 inches high, and weighs 3.9 pounds ma_mum. The beacon con-
tains one multlpin power connector and one coaxial connector.
Electrical Characteristics. The uhf recovery beacon consists of a spike elimi-
nator, a regulator, a dc-dc converter, a pulse coder, a modulator, and a trans-
mitter.
Spacecraft main bus voltage is fed to the switching type regulator through the
spike eliminator filter. The voltage regulator provides a dc regulated output
voltage of 12 vdc to the dc-dc converter, the transmitter tube filaments, and the
pulse coder.
The dc-dc converter is a solid-state device providing two high voltage outputs to
F--, the transmitter and modulator. The pulse coder, a solid-state device, operates
with the modulator to apply correctly coded high voltage pulses to the transmitter
for plate modulation of the power amplifiers.
The transmitter consists of an oscillator stage, a doubler stage, and a power
amplifier. The transmitter power amplifier provides a uhf pulse coded output hav-
ing a peak power of at least 50 watts to the uhf recovery anten_. An external rf
band-pass filter is installed between the transmitter output and the antenna to
reduce spurious rf radiations, especially at the uhf voice transmitter frequencies.
VOICE COMMUNICATION
Voice Control Center
Purpose: The Voice Control Center contains switches and controls for selecting
the type of voice communication and the desired operating mode. The VCC also
_" contains microphone and heatset amplifiers, an alarm tone generator, and voice
9-49
•ALARM TONE GENERATOR COMMON_)" MODULE
R/RECEIVER
CONNECTOR
HEADSET AND MIKE
AMPLIFIER LOAD TRANSMITTER/RECEIVERCONNECTOR
O J UHFINT HE _ , INT HF HP C)"
UHF_ (_J _J_, OUHP_D F'...""MODEl ORE MODE"-.. .'_
1_0.2NO.
_]UHP, HE , UHF rR_l
Ngt, E ,NpNc • q J
_] HF OFF, J ADIPT_,NTy HF[_]
o NORM FIT OFF o
VOX INT :ONT MOMNO.i 0.2 _-_
SILENCE o KEYING :, RECORD rc
AGE TEST ALARM TONE GENERATOR, TEST POINTS,POINTS PUSH-TO-TALK SWITCH, AND HEADSET AND
MIKE AMPLIFIER LOAD CONNECTOR
Figure 9-15 Voice Control Center
9-50
___ SEDR300
PROJECT GEMINI
actuated transmitter keying circuitry.
Physical Characteristics: The VCC and its approximate location is shown in
Figure 9-15. The VCC is mounted in the center instrument panel of the space-
craft cabin. The VCC is modular constructed, approximately 6.4 inches wide,
6.4 inches high, 5.5 inches deep, and weighs approximately 6.5 pounds.
Five connectors located on the rear of the unit provide connection to the other
voice communication system components and test connectors. The function of each
connector is listed on Figure 9-15.
The switches and controls of the VCC are located on the front panel. The number
1 and number 2 audio MODE switches are for selection of UHF, INT, HF, or HF/DF
.... _ transmission. Below the MODE switches are three thumb-wheel-type multidetent
volume controls, one for each of the above mentioned modes.
In the center is the KEYING switch, a HF select switch, a UHF select switch, and
thumb-wheel-type squelch controls for uhf and hf circuitry. The KEYING switch
provides for selection of PTT, VOX, or CONT INT/PTT for the voice transmitters.
The UHF and HF select switches provide capability of selecting the desired trans-
mitter/receiver. The ADPT position of the HF select switch is not used.
The record switch, lower right, permits recordings to be made in any mode of
operation. Continuous (CONT) or Momentary (MOM) recording can be selected.
9-51
SEO..OOCOMMAND f iPILOT J
/ I
.J >';;;'9 -_::__''_'_" M Ic lO PItO N E$ #/ +._
,i_._ > .' L i!it _<_( / -"
"\_'_, .... ,S-" HEADSET
HEADSET I IAMPLIFIERS
HELMEt IGHTj//,_ko,:O., <qZ voxRELEASE
KEYING
__ NORMSUIT DISCONNECT _ __ O. 2
_ATTITUDE
HAND
_'_ --0 1 _._ONTROLLER l O
I O
SUIT DISCONNECT -
oo,PUSH-TO-TALK _ _" _l
VOX
SWITCH I I _ RELEASE
HELMET _ O - HEADSET
o, .- --I '</ lLIGHTWEIGHT
// ......._,,,.'-'EADSET\ -_//#_ J
t_ :" /,L<_, \
HIES-
• ""_='_l_ MICRO pHO N ES tLL__ _J.
"_ MICROPHONEPILOT AMPLIFIERS "'-_'
Figure 9-16 Voice Control Center Functional Diagram (Sheet 1 of 2)
9-52
,-_ SEDR 300
f U.E_ ,AUD,O."N.",E," I FI ,ooRo0o0C_E_
CONT,-o-o 1";3_, o....-@.... _ _RoMGROUNOCOMPLEX
'DICE RECORDER FWR
......, o I _E__IZ]H
-- | _ECOROER AUDIO INPUT
1 _1_0 0 D_ _ jJ_ UHFNO. 2AUDIO(RECEIVER)_--_oOC _UHF UHE NO. 1 CHASS'S GROUND
UHF NO. 2 CHASSIS GROUND
UHF NO. 2 KEY LINE
UHF NO. 1
_N OFF
y._ #o.oo, O"F'_IOPOWER_O'DOFF
I C NO. 2 UHF MAIN F_OWERNO. 2
_ UHF RECEIVER POWER NO. 2
TO COAXIAL SWITCH !1
U.E .E/DE0_T 0 ,0CO_'ALSW'TC.'l
_INT O_o:tHF
e" _ _ O RE-ENTRY HF AUDIO (TRANSMIT)I
_''_'-'_HFI I "
| PADI RE-ENTRY HF KEY LINE
ADAPTER HF KEY LINE
RE-ENTRY HF AUDIO (RECEIVER)
' _E-ENTRy HF CHASSIS GROUND
J _ I I I : _'_ E RNTY HF
_J (_ _ OFFpT_ pRNIYJRE-ENTRy HF RECEIVER POWER
=_ OAD RE-ENTRY HF MAIN POWER
_ " GOI o
jj RE-EN1T_Y HF SQUELCH
if--\ -IAUDIO NO. 2 PWRFROM GROUND C OMR,.EX
J TO GROUND COMPLEX-- _ -- J_ TONE GENERATOR & VOX POWER
Figure 9-16 Voice Control Center Functional Diagram (Sheet 2 of 2)
9-53
SEDR 300
The SILENCE _wltch, lower left, is to permit uninterrupted sleep during extended
spacecraft missions. The N01_ position allows reception for both pilots. The
NO. i position removes power fran the command pilot's headset amplifiers and the
NO. 2 position removes power from the pilot's headset amplifiers; thus, making re.
ception impossible.
Electrical Characteristics: The VCC contains two headset and two microphone
amplifiers for each of the audio channels.
Figure 9-16 shows a functional block diagram of the VCC. An audio signal, fr_n
the microphone in the helmets or lightweight headsets, is amplified by two micro-
phone amplifiers and then applied to the MODE switch. With the MODE switch in the
HI_position, the output of the microphone amplifiers is applied to the hf trane-
mitrer. When the MODE switch is in the INT position, the output of the microphone
is applied to the four headset amplifiers, via the two TI_Tvol_e controls. With
the MODE switch in the I_F position, the output of the microphone amplifiers is
applied to the -b_ transmitters. The _ switch selects uhf transmitter
n_ber i or n_mber 2 and also operates coaxial switch i to connect the selected
transmitter output to the uhf quadriplexer. The desired keying mode is selected
by a e_....on_IN@ switch. Three methods may be selected to key the voice trans-
mitters. _e VOX position enables keying of the selected transmitter at the instant
the microphone has an output signal. The PTT position eanbles keying of the trans-
mitter when either push-to-talk switch, on the suit disconnect cables of the
attitude control handle, is depressed. The CONT INT/PTT position gives continuous
intere_=u_nication between the crew and push-to-talk keying for tra_ssion fr_
9-5k
___ SEDR 300
PROJECT GEMINI
the spacecraft to the ground station.
The VCC also controls the power supplies of the transmitter/receivers by means of
ground switching. With the MODE switeh i_ a position other than HF and the HF
select switch in the I_NTYposition, a ground is supplied to the hf transmitter/
receiver auxiliary power supply to power the hf receiver.
With the HF select switch in RNTY and the MODE switch in the HF position, a ground
is supplied to the hf transmitter/receiver main power supply to power the hf
receiver and transmitter. The uhf circuitry operates on the same principle as the
hf. The UHF select switch supplies power ground for the selected receiver.
The MOI_ switch (UHF position) together with the UHF select switch, supplies a
_ power return for the uhf transmitter and receiver.
The HF/DF position of the MODE switch is used for direction finding purposes.
With the MODE mrltch in HF/DF and the HF select in the RNTY position, the hf trans-
mitter is modulated by a 1,000 eps tone which is utilized to determine spacecraft
location.
UHF Voice Transmitter/Receivers
Purpose: Two ub_ voice transmitter/receivers are provided for redundant line-of-
sight voice c_uication between the spacecraft and the ground.
Physical Characteristics: The uhf voice transmitter/receivers and their approxi-
mate location is shown in Figure 9-17. Both t_tter/receivers are identical
and are mounted side by side in the forward right equipment bay of the re-entry
9-55
__ SEDR300 ____._j
PROJEC:I" GEMINI
module. Each transmitter/receiver is a modular constructed, hermetically sealed
--It approximately 7.7 inches long, 2.8 inches wide, 2._ inches deep and weighs
appro_mately 3.0 pounds. Each unit has a multipin audio and power connector, and
a coaxial connector.
Electrical Characteristics: The ,,hf voice t_tter/receiver consists of a
transmitter, receiver, and power supply.
The transmitter consists of a crystal controlled oscillator, two rf amplifiers, a
driver, and a push-pull power amplifier. All stages except the driver and power
a_plifier are transistorized. The transuitter is fixed-tuned at 296.8 mc and is
capable of producing an rf power output of 3.0 watts into a 50 ohm resistive load.
The transmitter is amplitude modulated (s_n)by a transistorized modulator stage.
The am superhetrodyne receiver is fully transistorized, is fixed-tuned at a fre-
quency of 296.8 mc, and contains a squelch circuit for noise limiting. The squelch
threshold is manually controlled. An automatic vol_e control stage is also
incorporated to provide a constant audio output with input signal variations.
The ,,b_voice transmitter/receiver is powered by two dc-dc converters co_prising
an auxiliary and a main power supply. Operating power for the two power supplies
is limited by two circuit breakers located on the left switch/circuit breaker panel.
One circuit breaker is provided for each unit. Actuation of the power supplies is
accomplished by ground return switching through the Voice Control Center. If the
UP_ select switch is in the NO. i or NO. 2 position and the MODE switch is
in a position other than UHF, a ground is supplied to the auxiliary power supply
9-56
.__ SEDR 300
PROJECT GEMINIf-
AUXILIARY AND MAIN cOVER
MAIN POWER J
SUPPLY I
AUDIO SECIIO N _-,-.,....,_1 _CHASSIS
AUDIO A NDPOWER
ILATOR
TO UHF COAXIAL
SWITCH MODULAR CONSTRUCTIONII
,h_TTER
Figure 9-17 UHF Voice Transmitter/Receiver
9-57
__ SEDR 300
PROJECT GEMINI
only, placing the tra-mnitter/receiver into a receive condition. With the MOI_
switch in the _ position, a ground is supplied to the main power supply, plac-
ing the selected uhf voice transmitter/receiver into a receive and transmit condition.
It should be noted that when the uhf transmitter is keyed, the uhf receiver is dis-
abled and ,,b_voice trs_ssions from the ground station can not be received.
HF Voice Transmitter/Receiver
Purpose: The hf voice transmitter/receiver is provided to enable beyond the line-
of-sight voice c_=-_,-_cationbetween the spacecraft and the ground.
Physical Characteristics: Figure 9-18 shows the modular construction and approxi-
mate location of the hf voice transmitter/receiver in the forward right equipment
hay of the re-entry module. The unit weighs approximately 62 ounces, is approxi-
mately 8.5 inches long, 3.3 inches wide, and 2.9 inches deep. One multipin audio
connector and one rf connector are provided.
Electrical Characteristics: Basically, the hf voice transmitter/receiver is
electrically identical to the uh_ transmitter/receiver except for the operating
frequency and power output. The hf transmitter and receiver are fixed tuned to a
frequency of 15.016 mc and the hf transmitter provides an rf power output of 5
watts.
Actuation of the hf receiver and transmitter is accomplished through the VCC. If
the HF select switch is in RNTY and the MODE switch is in a position other than HF,
the hf transmitter/receiver is in a receive condition. With the MODE switch in the
HF position, the hf transmitter/receiver is placed in a receive and transmit condition.
9-58
._./. SEDR 300 __
PROJECT GEMINI
MAIN POWERSUPPLy
TO HF COAXIAL
MODULAR CONSTRUCTION
AUDIO ANDPOWER AUDIO AMPLIFIERCONNECTOR
/_. tRANSMITTER
Figure 9-18 HF Voice Transmitter/Receiver
9-59
SEDR300
When the hf transmitter is keyed, the hf receiver is disabled.
Voice Tape Recorder
Purpose: The voice tape recorder is provided so recordings can be made during the
spacecraft mission.
Physical Characteristics: The physical construction and approximate location of
the voice tape recorder is shown in Figure 9-19. The voice tape recorder is located
inside the cabin in a vertical position between the pilot's seat and the right-hand
side wall on spacecraft 5 and 6. On spacecraft 8 through 12 the recorder is located
on the left-hand side wall aft of the abort handle. The voice tape recorder as-
sembly consists of the recorder, tape cartridge, and shock absorber mounting plate
and is supplied as GFE equipment. The recorder is approximately 6.25 inches long,
2.87 inches wide, one inch thick, and weighs 30 ounces maximum without the tape
cartridge. The shock absorber mounting plate is approximately 6.3 inches long,
three inches wide, and weighs 20 ounces maximum. The tape cartridge is approxi-
mately 2.25 inches square, 3/8 inch thickI and weighs two ounces.
The recorder contains a power connector and a signal connector located on the end.
as shown in Figure 9-19. The recorder is retained in the shock mount by guides
and two allen-head bolts for easy removal. The door contains a red plastic lens
so that light from the end-of-tape bulb is visible. A safety latch prevents
accidental opening of the door. The door is opened by pressing down on the latch
and sliding it sideways. When the latch is released, the spring loaded hinge
causes the door to open, exposing the cartridge tab. Flat pressure springs on the
door hold the inserted cartridge in place and maintains tape contact with the
9/oO
_- SEDR300
_RTRIOGE
MOUNTING PLATE
END-OF-TAPELIGHT
SAFETY LATCH J
Q
ji o(DOOR OPEN)
I
POWER CONNECTOR
/_, VOICE AND TiMESIGNAL CONNECTOR
Figure 9-19 Voice Tape Recorder
9-61
SEDR300
recorder head and end-of-tape contact.
The tape cartridge is guided into the recorder by step rails on each side of the
cartridge. When the recorder door is opened, a heavy tab on the cartridge springs
up to provide easy removal. The cartridge contains approximately 180 feet of mag-
netic tape, a supply reel, _ake-up reel, and associated gears and clutches.
Electrical Specifications: The recorder is a two-channel transistorized ,,n_tcon-
sisting of the cartridge hold-down mechanism, voltage regulator, voice amplifier,
time signal amplifier, bias oscillator, motor drive elrcuitt synchronous drive
motor, speed reducti_ unit, capstan, magnetic record head, and end-of-tape circuit.
When the tape cartridge is inserted and secured in the tape recorder, the pressure
roller in the cartridge contacts the capstan and the tape is pressed against the "-_
record head and the e_x1-of-tape contact.
The voice tape recorder is energized by spacecraft main bus power applied through
the TOHE VOX circuit b_r and the CONT or MCN position of the _CORD switch on
the VCC. The voXtage regulator supplies 15 vde to the motor drive cireuits2 bias
oscillator and szplifiers. With the ¥CC and recorder energized, voice signals
from the microphones are applied thro_Bh microphone amplifiers in the VCC to the
recorder voice slpllfier. _ voice signal is amplified and applied to the lower
record head for recording on the magnetic tape. The time channel receives a digi-
tal timing signal from a time correlation buffer in the TRS. The timinB signml
is amplified by the recorder time signal amplifier and applied to the upper record
head for recording on the magnetic tape.
9-62
[__ SEDR300 __ .___
PROJECT GEMINI
Simultaneously with the voice or timing signal, a 20 kc bias current from the
bias oscillator is applied to the recorder heads to make a linear recording.
The motor drive circuit consists of a iBB cps oscillator, a driver and push-pull
output stage used to drive the synchronous motor. Phase-shift capacitors are
connected to one motor winding for self-starting. The motor speed of 8000 rpm
is reduced through the speed reduction unit to a capstan speed of 122 rpm.
The end-of-tape circuit is energized by conductive foil on the tape contacting
the recorder head and end-of-tape contact, causing the end-of-tape light to
illuminate. The end-of-tape light will illuminate for two seconds when two
minutes of recording time remains on the tape. The light will remain on when
_-- the end-of-tape is reached. Recordings cannot be made when the light is illumi-
nated. The pilot may remove the used tape cartridge, insert another cartridge
and continue recording. Each cartridge provides approximately one hour of re-
cording. The tape speed is approximately 0.6 inches per second.
_ELE_ TRY TRANSMITTERS
Purpose: The three telemetry transmitters provide a radio frequency (rf) link
from the spacecraft to ground communication facilities for transmission of various
data obtained by the Instrumentation System.
Physical Characteristics: The three telemetry transmitters are identical except
for the operating frequency. The physical construction and approximate location
of the transmitters in the spacecraft is shown in Figure 9-20. The transmitters
are approximately 2.75 inches high, 2.25 inches wide, 6.5 inches long, and weigh
9-63
__ SEDR 300
,- PRO JE'-E-C'T'--GE M IN I
approximately 41 ounces. Each transmitter contains a dc power connector, an rf
output power connector, and a video connector. Two of the transmitters are
located in the right forward equipment bay of the re-entry module, the third is
located on the electronic module in the adapter equipment section.
Electrical Characteristics: The three telemetry transmitters are classified by
their operating frequency or by their function.
The real-time (low-frequency) telemetry transmitter operates at 230.4 inc. The
delayed-time (mid-frequency) telemetry transmitter operates at a frequency of
246.3 inc. The stand-by (hlgh-frequency) transmitter, operating at 259.7 mc,
may be used for real-time or delayed-time transmission in case one of the trans-
mitters fails.
The telemetry transmitters are solid-state fmtransmitters. After a 30 second
warm-up, the transmitters are capable of continuous uninterrupted operation for
500hours. Information is transmitted to the ground in digital format by deviating
the carrier frequency to the higher frequency deviation limit to transmit s I,
and to the lower deviation limit to transmit a 0.
The transmitters receive Non-Return to Zero (NRZ) PCMpulse trains from the PCM
programmer and voice tape recorder. The real-time transmitter provides the
ground monitoring stations with current real-time data at a rate of 51.2 kilobits
per second. The delayed time transmitter provides the ground monitoring station
with data stored on the tape recorder while the spacecraft was between ground
stations. The delayed-time data is transmitted at a rate of 112.6 kilobits
per second. The stand-by transmitter is used as backup for the real-time
or delayed-time transmitters in event of a failure in eithertransmitter. _
Transmission of the real-time and delayed-time data provide essentially full-time
9-64
____ SEDR300 ___
PROJECT GEMINI
FROM FCM PRO(
TO ANTENNA VIACOAXIAL SWITCH
POWER CONNECTOR
Figure 9-20 Telemetry Transmitters
9-65
SEDR 300 ___]
PROJECT GEMINI
coverage throughout the spacecraft mission. The transmitters can be energized
by a command from the ground station via the DCS or by controls on the instru-
ment panel.
Each transmitter consists of five subassemblies as shown in the block diagram of
Figure 9-21. The subassemblies are an oscillator-modulator, a times 12 (x12)
multiplier and power amplifier, a bandpass output filter, a line filter and a
de-de converter. The oscillator-modulator and the times 12 multiplier and power
amplifier subassemblies contain the variable resistors, inductors, transformers
and tr_-_er capacitors for tuning the transmitter frequency and power output. The
subassembly components are point-to-point wired.
The oscillator-modulator consists of a video amplifier, crystal controlled oscil-
lator, phase shift networks and buffer amplifiers. The oscillator frequency is
modulated by the video amplifier output. The phase shift networks provide impe-
dance matching of the crystal oscillator to improve signal linearity for large
deviations of frequency. The buffer amplifiers increase signal levels and isolate
the crystal circuit from the frequency multipliers.
The times 12 multiplier and power amplifier consists of a buffer amplifiers, times
4 multiplier, power amplifier and times 3 multiplier which increase the carrier
frequency and power to the desired output values. The power amplifier develops
6 to 7 watts of power at a frequency from 75 to 80 mc into the output tripler
circuit.
The bandpass filter is used to minimize spurious radiations at the output of the
transmitter. The real-time, stand-by and delayed-time transmitters each contain a _
filter with a different frequency bandpass. The filter has a minimum S db
9-66
,EO,300PROJECT GEMINI//_- \
m m
FILTER ANTENNA
MULTIPLIER _1 POWER _ BUFFER _ MULTIPUER _1" BUFFERX3 _1 AMPLIFIER I AMPLIFIER X4 AMPLIFIER
i
X12 MULTIPLIER AND POWER AMPLIFIER
BUFFER PHASE i,, BUFFER PHASE BUFFERAMPUPIER _ ,- i'--'--II IISHIFTER AMPLIFIER SHIFTER AMPLIFIER
II
VIDEO _ iNPUTOSCILLATOR AMPLIFIER
OSCILLATOR-MODULATOR
H i I_ 70 VDC
V
18 TO 30.5 VDC '% LINE DC-DCINPUT POWER f FI LTER CONVERTER
l_ 30 VDC
Figure 9-21 Telemetry Transmitter Block Diagram
9-67
SEDR300
bandwidth of 16 me, an impedance of 50 ohms and a vswr of less than 1.5 to I. The
rf output co--ector J3 is an integral part of the bandpass filter.
The line filter prevents noise on the input power bus from affecting transmitter
operation and prevents transients generated within the transmitter from feeding
back to the input power bus. The multipin power connector J2 is an integral part
of the line filter.
The de-de converter is a completely encapsulated unit employing transistors, diodes
and a transformer to provide regulated outputs of 30 vdc and 70 vdc from an u-_eg-
ulated input voltage of 18 to 30.5 vdc. The converter is a constant power input
type, thus minimizing the heat dissipation caused by high voltage inputs.
FLASHING RECOVERY LIGHT AND POWER SUPPLY
Purpose: The flashing recovery light and power supply provide visual spacecraft
location information.
Physical Characteristics: Figure 9-22 shows the physical representation and
approximate location of the flashing recovery light and its power supply. The
light is self-extended by a torsion spring. The plug applying power to the light
is kept in place by a compression spring. The recovery light will be automatical-
ly extended at the time the main parachute is Jettisoned.
The flashing recovery light power supply is mounted in the cabin, aft of the
ejection seats. The power supply is approximately 7 inches long, _ inches wide,
3 inches deep and contains one connector. The flashing recovery light is approx-
imately 1.25 inches wide, 0.75 inches thick, and 3.25 inches high, excluding tube
and erecting mechanism. The overall length of the light and erecting mechanism is
9-68
_,__ SEDR300 ____
PROJECT GEMINIS _
RECOVERY LIGHT __
ERING
.FLASHING RECOVERY
SPRING LIGHT POWER SUPPLY
(2 REQUIRED)
I
PLUG (2 REQUIRED)
COTTER PIN_ HINGE PiN
NUT
(2 REQUIRED),
Figure 9-22 Flashing Recovery Light and Power Supply
9-69
SEDR300
approximately 6.5 inches.
Electrical Characteristics: The recoveIV light is automatically extended at main
parachute jettison. The extended recovery light is energized by positioning the
RESC BEACON CONTROL switch to ON.
The power supply consists of a battery pack and converter. The battery pack
consists of several mercury cells to comprise a power source of 6.75 vdc to a
dc-dc converter whose output is fed to a voltage doubler and a capacitive network.
The 450 vdc output of the voltage doubler is used to power the flashing light
while the capacitive network in conjunction with a thyratron, provides trigger
pulses to accomplish switching or flashing action of the light. The trigger
pulses occur at a rate of 15 triggers per minute,
DIGITAL COMMAND SYSTEM
Purpose
The DCS provides a discrete command link and a digital data updating capability
for the com_uter and _RS.
The discrete command link enables the ground to control radar tracking beacons,
selection of telemetry transmitters, instrumentation data acquisition_ and abort
indications.
The capability of digital data updating enables the mission control center to
update the computer and _RS to bring about a controlled re-entry at a pre-deter-
mined point, and allows timed shutdown of equipment controlled by DCS relays.
9-70
___ SEDR300
PEMINI
Physical Characteristics
The DCS consists of a receiver/decoder package and two relay boxes as illustrated
in Figure 9-23 and 9-24, respectively. The three con_onents are located in the
electronic module of the adapter equipment section.
The receiver/decoder package is approxJn_tely 8 inches high, 8 inches wide, and
12 inches long. Both relay boxes are identical. Each relay box is approximately
2.25 inches wide, 5 inches_high, and 3 inches deep. The combined weight of the
receiver/decoder package and the two relay boxes is approximately 2B pounds. The
receiver/decoder package contains two uhf receivers and a decoder while each of the
two relay boxes contain eight relays.
General Descriptio n
The DCS receives Phase Shift Keyed (PSK) signals composed of a reference and an
information signal. The information signal is in phase with the reference for a
logical 1 and 180 degrees out of phase with the reference for a logical O; thus
establishing the necessary requirements for digital data.
Types of Commands
The DCS receives two types of digital commands: Real Time Commands (RTC) and
Stored Program Co_ands (SPC). RTC causes relays within the DCS to be actuated.
Nine of the 16 relays available for RTC are utilized to perform the following
functions:
(i) Select the stand-by telemetry transmitter for real-tlme transmission.
(2) Select the stand-by telemetry transmitter for delayed-time transmission.
(3) Select the real-time telemetry and acquisition aid beacon transmission.
(4) Select real-time and delayed-tlme telemetry transmission.
(5) Actuate the adapter C-band radar beacon.
9-71
SEDR300
Figure 9-23 DCS Receiver/Decoder
9-72
__. SEDR300 __
PROJECT GEMINI
DCS RECEIVER-
Figure 9-24 DCS Relay Box
9-73
SEDR 300
PROMINI
(6) Actuate the re-entry C-band radar beacon.
(7) llluminate the abort indicators.
(8) Actuate the playback tape recorder.
(9) Initiate calibration voltage for the PCM programmer.
The remaining seven relays are not utilized and perform no mission function. DCS
channel assignments for the nine functions listed above may be different on each
spacecraft.
When the spacecraft goes out of range of the ground station, equipment controlled
by DCS channels may be shutdown by a signal applied from the TRS to reset the DCS
relays. This condition is known as salvo. The DCS relays in one relay box may
be reset by momentarily positioning the TAPE PLY BK switch to RESET.
Message Format and Modulation
The ground station transmlts a 30-bit message for SPC and a 12-bit message for
RTC. Each bit consists of flve sub-blts. The five sub-blts are coded to repre-
sent a logical 1 or O. The first three bits of each message designate the
vehicle address. If the vehicle address is not correct, the DCS will reset itself
and will not accept the message. If the vehicle address is accepted the sub-bit
code will be automatically changed for the remainder of the message to reduce the
probability of accepting an improper message.
The second three bits of each message designate the system address and identify
the re-_Inder of the message as being a RTC or one of the following SPC :
computer update, TRS time to go (TTG) to TR, or TRS TI_ to equipment reset (Tx).
If the message is a SPC, the last 24 bits will be a data word. If the SPC is a
TRS TTG TO TX command, the last eight bits are ignored by the TRS. In case of a
9-74
SEDR300
PROJEC---'T-'GEMINI
computer message, six bits of the data word contains the internal computer
address and the remaining 18 contains information. Since a RTC consists of
12 bits, the six bits following the system address contain a 5-bit relay number
and a 1-bit relay set-reset discrete.
The PSK modulation signals are i kc reference and a 2 kc information signal. The
receiver output is the composite audio of the I kc and the 2 kc signals. The
composite audio output is filtered to recover the 1 kc and the 2 kc signals. The
phase co_arator compares the 2 kc to the 1 kc signal. The output of the phase
comparator is used to trigger a flip-flop to produce either a logical 1 or 0
sub-bit. The i kc reference signal is used to synchronize the DCS.
_ Operational Description
A block diagram of the DCS receiver/decoder is shown in Figure 9-25. Basically,
the block diagram consists of a receiver, a decoder, and a power supply common
to both sections.
The audio outputs of the two receivers are linearly summed in an emitter follower
of the sub-blt detector module. The sub-bit detector converts the audio to sub-
bits. The 5-stage shift register provides buffer storage for the output of the
sub-blt code. When a proper sub-bit code exists in the shift register, the
bit detector produces a corresponding i or 0 bit. The output of the bit detector
is applied to the 24 stage shift register. The operation for RTC and SPC is
identical up to the input to the 24 stage shift register.
The sub-bit sync counter produces a bit sync output for every five sub-bits. The
_ bit sync is used to gate the 24 stage shift register.
9-75
SEDR300
PROWl
When a message is received, the vehicle address is inserted into the first three
stages of the 24 stage shift register. If the vehicle address is correct, the
vehicle address decoder circuit will produce an output to the bit detector which
changes the acceptable sub-bit code for the remainder of the message. The next
three bits of the message, the system address, are inserted into the first three
stages of the 24 stage shift register, displacing the vehicle address to the next
three stages. The system address decoder circuit identifies the specific address
and sets up the DCS to handle the remainder of the message.
When the system address is recognized to be a RT% the message is inserted into
the first six stages of the 24 stage shift register and the system address and
vehicle address are shifted into the next six stages. The RTC selection circuit
recognizes the first stage of the 24 stage shift register to be a relay set or
reset function and will apply a positive voltage to all set or reset relay coils,
as applicable. The RTC selection gates select the proper relay from the relay
number stored in the 24-stage shift register and provides an output which applies
a power return to the coil of the selected relay.
;_en the system address is a SPC, the six address bits in the 24 stage shift
registers are cleared and the remaining 24 bits of the message are placed into
the register.
Assuming that the system address recognizes a TRS TTG to TR message, the data
flow would be as follows: The TRS TR isolation amplifier, in the inter_ace cir-
cuit, will apply a READY pulse to the TRS. The READY pulse sets up the TRS
to transfer TRS TTG to TR data from the DCS. When the TRS is ready to accept the
data, it sends 24 shift pulses, at the TRS data rate, to the TRS input of the DCS.
_ SEOR 3O0 __
PROJECT GEMINI
T ?SIGNAL STRENGTH RECEIVER RECEIVER _ SIGNAL STRENGTHTO TELEMETRY R1 R2 TO TELEMETRY
' I
I I [CHANG COO'NGISUB-BIT SUB-BITS b 5 STAGE SUB-BITS s= ADDRESSSTORAGE _ BIT DETECTOR DECODER
DETECTOR REGISTER
TIMING ERROR
INHIBIT
ADDRESS
I _ ADDRESS NO.
DATA
SUB-BIT SET RELAYSYNCHRONtZER STORAGE
FAILURE AND RESET REGISTER DATA SELECTION
I 1RELAY DRIVE
_ 2
L ENDIPOWER 20-30V DC PROGRAM .,= TRANSFER IN INTERFACESUPPLY JI INPUT POWER R 'kDy CONTROL PROGRESS
VALIDIIY
' - ' _ _ =o/f_ POWERSUPPLY NOTE _ : _; _-<
VOLTAGES TO >I O U '_.TELEMETRY 1. HEAVY LINES DENOTE DATA FLOW. _ U _ --
>-_ _Figure 9-25 DCS Block Diagram
9-77
SEDR 300
PROJEMINI
The data in the 2_ stage shift register ie then shifted out of the register
through the DCS data isolation amplifier to the _RS. The DCS operations for
computer updating and _ _ to TX messages _ s_mllar to _ TTG to TR
operations.
Salvo occurs when _S TTG to TX reaches zero. At TX : O, the TRS applies a signal
to the TRS TX input line of the DCS which causes the RTC selection circuits to
reset the DCS relays.
After a Sl_ or _ has been carried out by the DCS, a verification signal is
supplied to the telemetry system for transmission to a ground station. The DCS
indicator, on the instrument panel, illuminates when a SPC is transferred to the
appropriate system. _
Upon completion of data transfer or if the system to which the data was transferred
fails to respond within 100 milliseconds, the DCS will reset in preparation for
the next message. The DCS will also reset in the event of a timing error in trans-
mission of data, _or if the DCS power supply voltages become out of tolerance.
9-78
INSTRUMENTATIONSYSTEM
, ecfionTITLE PAGE XSYSTEM DESCRIPTION ................................. lO-5
SYSTEM OPERATION ................................... lO-5SEQUENTIAL SYSTEM PARAMETERS ................ 10-11ELECTRICAL POWER SYSTEM PARAMETERS ...... 10-16ECS PARAMETERS ........................................ 10-19
INERTIAL GUIDANCE SYSTEM PARAMETERS ..... 10-24ACME PARAMETERS ..................................... 10-26
OAMS PARAMETERS .................................... 10-28
RE-ENTRY CONTROL SYSTEM PARAMETERS .... 10-31AERODYNAMIC AND CREW
CONTROL PARAMETERS............................... 10-33
COMMUNICATION SYSTEM PARAMETERS ....... 10-35
INSTRUMENTATION SYSTEM PARAMETERS ...... 10-35
PHYSIOLOGICAL PARAMETERS ...................... 10-39RENDEZVOUS RADAR PARAMETERS ............... 10-41 i"_"_'_"_._.
SYSTEM UNITS............................................. 10-43 :'::::..--:_::."._::'m..=:;.:.".'._:.:.:.:.._::::._..':___PRESSURE TRAN SDUCERS ............................. 1O-43 iiiii_!iii!ii'_-_'-_
ff'.:'"'.:'.:."'.:.:"_-:"7--"_:TEM PERATURE SENS0 RS ............................... 1O-45 !i_iiiiii_!_![i!_[ii_i_ii
SY NCHR0 REPEA TERS .................................. 10- 47 iiiiii!!iiiiiiiiii!!!!!_i!i.oo..oo°.°°.oo.°°,o. oo,.o.I.oo..o.o°°°..o°°..o....._*,
co2 PARTIALPRESSUREDETECTOR...............10-47 i!ii!!iiiiii!i!Hiiiiiiiiii,°o.°........°°,°.o.°,.._,_,oo....o....°...,.°_...._
ACCELERO M ETERS........................................ 10- 50 ii_iiii_iiii_!iiiiiiiiiiiil,o°°°°o°°_°°.°°.°°°°°°°°°.,
INSTRUM ENTA TION PA CKA GES .................... 10- 50 i_i_!_iiH_!i!iii_i_iMULTIPLEXER/ENCODER SYSTEM .................. 10-54 iii!_iiii_i!i!i_!iiiiiiii..... ....°°.o......°°°°°°o,,.....°°°°..°°°°.°°.°°.°.°_
TRANSMITTERS ............................................ 10-61 !i!ii!iliHii!i!!iii!HiiiiPCMTAPE RECORDER .................................. 10-61 i_!i!ii_!i_!!iiiiiiiiiii
DC- DC C0 NV ERTERS..................................... 1O-6 5 iiiiiiiiii_iiiiiiiiii!iiill
BI0- M ED TAPE REC0 RDERS iiiiiiiiiiiiiiiiiiilH!iiiii..., ................... ...,
AND POWER SUPPLY ................................. 10-67 iiiiiiiiiiiii!iiiiiiiiiiiii10-1 ::::::::::::::::::::::::::::
M_ -;_. SEDR 300
(REF)
/./
/I ,,:'D
// /._-_ ! i
I ::2/ /"
// ,i1 / ,1
; //t: ;i \ ':',i ,_: ! i' " \i :7 ' -' -_>---"' -'"!i ii:! iil "...... i #l /: """ "-J ' _\! /! : I _ s i _::_'i,i !/ i : i ,_,;,_ j _
/ :/
!/ /\ RETRO
ROCKET, #4(REF)\\. /
OAMS " "
:(REF) \
i ®//
/"
ELECTRONICS i_." /%MODULE(REF) /:
/MID FREQUENCE "/
TELEMETRY TRANSMITTER
INSTRUMENTATIONASSEMBLY NO. 2
•HIGH LEVELMULTIPLEXER MULTIPLEXER
Figure 10-1 Instrumentation System Components (Sheet 1 of 3)
10-2
j_=-_. SEDR 300
I i _-__ PROJECT GEMINI_t..,.._._ _-_-................. _
.\
RECORDER NO. 2
BiGMEDPOWERSUPPLY \\
BIO MED TAPE \
RECORDER NO, I \
\
/ \J \, ,\
//
/ ........... \/
/ _.,_ [ , .......®2
%.
\\.\"x,, _i \LOW LEVEL _ 1 l
l\
, /\ /
\. \ _\\
\/ /
Zf I
INSTRUMENTATIONASSEMBLY
PACKAGE LOW FREQUENCYNO. 2 TELEMETRy
HIGH LEVEL TRANSMITTER L_fHIGH FRE(DC-DC CONVERTER TELEMETRY& REGULATOR TRANSMITTER
PCM PROGRAMMER" )
TAPE RECORDER//REPRODUCER
_..' /
Figure 10-1 Instrumentation System Components (Sheet 2 of 3)
10-3
LEGEND LEGEND
ITEM PARAMETER NOMENCLATURE ITEM PARAMETER NOMENCLATURE
1 GC04 REG He AT OX1D TANK TEMPERATURE 31 LAe5 DCS PACKAGE TEMPERATURE
2 GB02 OXIDIZER FEEDTEMPERATURE 32 LD01 ACQ AID ECN CASE TEMPERATURE
3 GB01 FUEL FEED TEMPERATURE 33 MC02 MID FREQ TM XMTR CASE TEMPERATURE
4 GC0_3 REG He AT FUEL TANK TEMPERATURE 34- CC05 OXYGEN HIGH RATE
5 GC01 SOURCE He PRESSURE 35 GC02 SOURCE He TEMPERATURE
6 GD06 TCAEI0 HEAD TEMPERATURE 36 IC_.01 Z ACCELERATION
7 CJ01 PRIMARY COOLANT PUMP INLET PRESSURE 37 KA02 X ACCELERATION _
38 KA03 Y ACCELERATION8 CJ02 SECONDARY COOLANT PUMP INLET PRESSURE
39 CB07 EWD COMPARTMENT ABSOLUTE PRESS9 GC05 REGULATED He PRESSURE
40 KB02 STATIC PRESSURE10 CH0_ SECONDARy COOLANT RADIATOR OUTLET TEMPERATURE
41 CBOI CABIN PRESSTO FWD COMP.11 CD04 SECONDARY COOLANT TEMP AT OUTLET OF RADIATOR
12 CD0¢3 PRIMARy COOLANT TEMP AT OUTLET OF RADIATOR 42 CB02 CABIN AIR TEMPERATURE
43 HC03 REG N 2 PRESSURE-SYSTA13 CA06 PRIMARY ECS 0 2 SUPPLY ROTTLE TEMPERATURE
44 DQ07 PITCH ATTITUDE-SYNCHRO REPEATER14 CA02 PRIMARy ECS 0 2 TANK PRESSURE
15 CH02 PRIMARY COOLANT RADIATOR OUTLET TEMPERATURE 45 DQ08 ROLL ATTITUDE-SYNCHRO REPEATER
16 CA09 CRYO MASS QUANTITY (RSS-ECS) 46 DQ09 YAW ATTITUDE-SYNCHRO REPEATER
47 HC04 REG N2 PRESSURE-SYST B17 CD02 SECONDARY COOLANT INLET TO F.C. SECT 2 TEMP
48 HC06 SOURCE N2 PRESS-SYST B18 CD01 PRIMARY COOLANT INLET TO F.C. SECT 1 TEMP
49 HC02 N2 SOURCE PRESS-RCS SYST B19 CL01 WATER PRESSURE
20 BC03 F.C. H2 TEMP AT HEAT EXCHANGER OUTLET 50 HA02 RCS OXIDIZER FEED TEMP-SYST A
51 HC01 N2 SOURCE PRESS-RCS SYST A21 BA04 HYDROGEN TANK PRESSURE
52 HCD5 SOURCE N 2 PRESS-SYST A22 BB05 F.C. 0 2 TEMP AT HEAT EXCHANGER OUTLET
23 BA06 RSS H2 SUPPLY BOTTLE TEMPERATURE 53 CC06 CO 2 PARTIAL PRESSURESENSOR
24 CA09 CRYO MASS QUANTITY (RSS-ECS) 54 CC03 LEFTSUIT INLET AIR TEMP
25 HH01 RETRO ROCKET CASE TEMPERATURE 55 CA03 ECS 0 2 SUPPLY PRESSNO. 1-SEC
26 CA09 CRYO MASS QUANTITY (RSS-ECS) 56 CK06 SUIT HEAT EXCHANGER INLET TEMP-PRI
57 CC01 LEFTSUIT PRESSURE27 LC09 ADAPTER C-BAND BCN PACKAGE TEMPERATURE
28 CA09 CRYO MASS QUANTITY (RSS-ECS) 58 CC02 RIGHT SUIT PRESSURE
29 BA05 _SS 02 SUPPLy BOTTLE TEMPERATURE 59 CA04 ECS 0 2 SUPPLY PRESSNO. 2-SEC __,
60 CC04 RIGHT SUIT INLET AIR TEMP30 BA02 OXYGEN TANK PRESSURE
Figure 10-1 Instrumentation System Components (Sheet 3 of 3)
10-4
_. SEDR300
PROJECT GEMINI
S_ SECTIONX INS__ION SYS_
STS_M nRSCRIPTION
The Instrumentation System provides a means of data acquisition with respect
to the performance and operation of the spacecraft throughout its mission.
Data acquisition is defined as the sensing of specific conditions or events
on board the spacecraft, displaying the derived data from these inputs to
the crew and ground operation personnel, and recording and later processing
this data for use in post flight reports and analysis. In this respect the
data acquisition function is shared by all spacecraft systems, the ground
operational support system, and the data processing facility.
Basically, the instrumentation Parameters are divided into two categories :
operational and nonoperational. Operational parameters are those which are
necessary for determing the progress of the mission, assessing spacecraft
status, and makin E decisions concerning flight safety. Nonoperational psram-
eters are those which are required for post mission ana]jsis and evaluation.
The basic components comprising the Instrumentation System are: sensors, signal
conditioners, Multiplexer-Encoder System, and transmitters. Because the system
is used to sense Parameters of every spacecraft system, its components are
located throughout the spacecraft as shown in Figure i0-i.
SYSTm_ OPERATION
The purpose of the Instrumentation System is data acquisition with respect to
the progress and condition of the spacecraft, necessitating its operation
so300PROJECT GEMINI
throughout the mission. The Instrumentation System provides the capability of
data acquisition and transmission to the ground stations. The data is supplied
by all spacecraft systems. The basic operations by which the system fulfills
its purpose are: to sense the various conditions and functions; convert them to
proportional electrical signals (if applicable); condition the resulting signal
(when necessary) to make it compatible with the encoding and multiplexing equip-
ment; display pertinent data in the cabin; record data for delayed time (data-
dump) transmission; and provide signals for real-time transmission to the ground
station. An overall block diagram of the Instrumentation System is shown in
Figure 10-2 and the power distribution is shown in Figure i0-3.
The system senses the prescribed parameters through the use of sensors which
may be contained within the Instrumentation System or which may be an integral ....
part of another system. Typical sensors include pressure transducers, accelero-
meters, and temperature sensors. Signals may also be obtained from such functions
as switch and relay actuations, and from electronic package monitor points. Sen-
sors and signal sources are shown in block diagram form on the applicable data
source system illustration.
The majority of the signals acquired are usable for the spacecraft cabin indi-
cators and/or the encoding equipment without alteration. Some of them, however_
are routed to signal conditioning packages (instrumentation assemblies) where
their characteristics and/or amplitudes are changed. The resulting signals,
as well as those from the other sensors, are of four basic types: low-level
.../__. SEDR300
[c_AOI I QC27 BA02
ADO6 REF "BA05 6.A02 I REFQC28 BA04AD08 FIGURE 10-5"( BA06 : FIGURE ---- REF
QC29AD09 BB05 c_,AL_ i 10-11 8003 -FIGUREAD10 BC03 _.B01 i QC30 fiB04 10-5AEI3 CA06 6_B0_ = ' BC01
REF AF04 CD01 &.B04 = BC02FIGURE _ AG02 REF CD02 _D02 CA0210-4 AG03 F_GURE I0_-_ CD03 _.D03 CJ01
AGO4, CD04 _ _,E02 CJ02AG13 CH02 _ _E27 !CJI6 RL_F
AG14 CH03 _ a,E28 _ CJ17 .FIGURE
AG,5 CB0, :_" F,GuREREF10-4. A.F0, = CJISCJI8 10-6AG,8 CB02 = _F02 REAL TIME n_ I_T_--
REF BG01 REF GC02 __ AFO3 TELEMETRY _._ CL01FIGURE _ BG02 FIGURE 10-94 GC03 _ AGO5 TRANSMITTER _ GC0110-5 BG03 GC04 _' AGI0 _ GC05
AG11 _ OF01BG04 GD06 _ AG12 _, GE02CA03 REF FIGURE 10-7-- DWOl _ AG16CA04 REF FIGURE 10-10.-- HH01 > _ GE03REF CB01 ./" LA02 _ AG17 GE04 REF
FIGURE- CCOI t LA03 _ AG19 _ GE_ - FIGURE
10-6 CC02 REF LA0.4 O AG21 ">' GE06 10-9
CC05 _ FIGURE 10-12 LA05 AG22 _ GE07= co, GE0DAOl _Z LD01 REF BD04 _ GE09DA02 REF MB02 FIGURE 10-5= BD06 STANDBY GEllDA03 FIGURE 10-13 MB03 BE04 TELEMETRY LA06 REF
REF DD0I MC02 BE06 TRANSMITfER LA07 FIGURE
_ FIGURE- DD02 REF FIGURE 10-6_ CB07 LA08 _" 10-1210-7 DD03 DBCG _0T REF
DE01 DB_ M80_ FIGUREDC01 MC01 _" 10-13
DED2 REF DC02 _ DW02 REFDE05 FIGURE I0-7=EB01 DC03 _ DW03 " FIGURE
REF EB02 ( BD01 DQ07
FIGURE = EB03 REF J BD02 DQ08 O_ DW04GC22 10-710-8 FIGURE 10-5_ BE01 DQ09 O FIGUREEA01 10-9
L BE02I BH0t _ EA02 DELAYED
BH02 __ EA03 TIME
r EG01 TELEMETRYHE04 REF CB02 _ REF EG02 TRANSMITTERFIGURE 10-6 " CC03 _ FIGURE 10-8 " EG03
_._ CC04 _ EGO4CK06 x
REE HE07 REF _ EGD5FIGURE_ HE08 FIGURE I0-I0" _- HA02 _ EG06
HC05 _ EG07
I0-I0 REF HC06 :3FIGURE 10-13 MA21 _ REF FA01
MA38 _ FIGURE 10-11 FA02
JA04 _ FAg3
JAI3 _ REF HC01REF JB01 FIGURE 10-10"_ HC02
FIGURE 10-15 JB02 HC03 PCM TAPEJB04 HC04 RECORDERJD01 KA01
REF j MAlT REF KA02 JB(_ -FIGURE MA22 FIGURE I0-11 =(
10-13 =_ MA37 KA03 ]C01REF MD06 KB02 JCD2
._ REF FIGURE 10-12-- LA01 JC03FIGURE NA06
10-14. NB06 if NA0| JC04 REF
REF =_ JE09 1 NA02 JE01
NAI_3 JE03 -FIGUREFIGURE AG06 REF NA04 JE04 10-1510-15 FIGURE 10-14 NB01 JE05
NB02 3E06NB0_ JE07NBC_ JE08
J
Figure 10-2 Instrumentation System Signal Flow Block Diagram
10-7
SEOR3OOPROJECT GEMINI
_N_ o_o _I _,__L_s_
0c_o o,PC-PCCO_. 0 0 ) 0
I 'T ,, I To o o o o ,_ _ o o o o o5 ,_
o o o> > o o o_ o > > >> > > > a. > >
PRLMARY DC-DC CONVERTER SECONDARY DC-DC CONVERTER
II I102sEcT i I ;ISuPPLYI ATTITUDE
E E E.c.H2S_RCERCSN2S_RCERCS",'2S'_CEIN2REGP'ESSN_,REGPRESS HANDSYST 8 PRESS TRANSDUCER TRANSDUCERPRESS SYST A PRESS PRESS
TRANSDUCER CONTROLLERTRANSDUCERTRANSDUCER TRANSDUCER SYST A SYST BSYST1 ROLL
BA04 HC01 HC02 HC03 HC04 CA03 FA02
COOLANT PUMP ATTITUDE 'RE PRI INLETRIGHT SUIT PRESS CRYO PRESS& PRI 02 PRESS HAND
TRANSDUCER QUANTITY INDCC02 CA09 * TRANSDUCER PRESS CONTROLLERCA02 PITCHTRANSDUCER
CJ01 FA01
Iw, T ,L IT!o ll .c.o!T-PITRANSDUCER PRESS PRESS F.C. 02 F.C. H2 AT H-X OUTLET PRESS
TRANSDUCER TRANSDUCER RA05 RA06 TRANSDUCERCL01 GCOI GC05 CB07
T _ T T T T T
i III IICONT. VALVE CONT. VALVE OUTLET TEMP KB02PR' F.C. COST VAIVF TEMP COOLANT
F.C. TEMp TEMP BRIDGE SEC F.C, COSTAT H-X _)UTLET PRI 0 2 - ECS VALVE TEMP TRANSDUCER
BC03 CA06 CD01 CD03 CD04 CH02
T T T T T T T TCOOLANT TEMP TEMP TEMP IFUELTANK TEMPj OX_D TANK RETRO ROCKET ACQ AND BCN
OUTLEI" TEMP GB01 GB02 GC02 GC03 TEMP CASE TEMP CASE TEMPCH03 GCO4 HHOI LD01
T T T T T T T
1 1 I I I lI I JTEMP TO SUIT H-X TEMP SYST A TEMP SYST A TEMP SYST B DCS PACKAGE ClB_.CON CASE TEMP _,CB02 CKD6 HA02 HC(_ HC06 LA(_ LC09 MCG2
Figure 10-3 Instrumentation System Power Control Circuitry Functional Diagram (Sheet 1 of 2)
10-8
7T_ SEDR300
/_" 0 NO. I I INSTCRYO. QTY. PACKAGE
• _--0 SYSTEM _ NO. 1C | O C_09
I NO.2
CONTROL _ I _O ! PACKAGE
8US _ CALIB CALI8 O NO.2 NO 2
t H"LEVEL I
-- MULTIPLEXER
TAPE PLYBK (RE-ENTRY)
RESET OCMD K-D/T STBY
TAME_CDR _ ]' K-D/TII TAPE- II RECO.OERCNTL
CONTROL _ CMDBUS
TO COAX I O Pet & D/T I
MAIN _ XMTR IXMFR [ TMcoOT_oL
•11 XMTRD/T XMTR K-D/F
MAIN K-D/TSTBY
INSTRUk_EN - P/T iTATION IlK _ STBY TM CONTROL OPACKAGE NO. I _ OFF PROGRAMMER
TR O D/T _I
cO_.BO, _ _BUS
F-, ISYNC._REPE.ER_.TC.oo0._, ' _,NBUSSTBYXMTRMR.,O-_EDRCOX.._F_oFF TMXMFR--_''O''_O ! O CONT ,IO-MED ]
BIO-MED TAPE RCDR
MAIN INST NO, 1
I BUS
SYNCHRO ; O CONT BIO-MEDREPeATER-ROLL TAPE RCOR
DQ08 b NO. 2
i ! j BIO-MED
I BIO-MED . INSTRUMEN -POWER SUPPLY TATION
I SYNCHRO 4
REPEATER-YAW
He TEMP He TEMPAT He TEMPATiNDiCATOR " FUEL TANK OXID TANK
INDICATOR INDICATOR
LEFT SEAT I LEFT & RIGHT I T PRESSURE EGONE SW SUit FEMP CC06 ,
NA02 I TRANSDUCERBRIDGE _1
T , SOURCE N2 TEMP SOURCE N2 TEMP QUANTITYI INDICATOR INDICATOR INDICATOR
I II I i 11 rAFT.UDE 'ND ÷/_--_., CONTROLLER- INSTRUMEN- NX NY NZ
YAW TATION 'FA0_ PACKAGE NO. 2 i ACCELEROMETER ACCELEROMETER ACCELEROMETERKA02 KA03 KA01
Figure 10-3 Instrumentation System Power Control Circuitry Functional Diagram (Sheet 2 of 2)
10-9
SEDR 300
PROGEMINI
(0-20 my tic),high-level (0-5 vdc), bi-level (0 or 28 vdc), and hi-level pulse
(28 or 0 vdc). Signals of selected parameters are supplied to the cabin indi-
cators, while signals of all parameters are supplied to the l_ltiplexer/Encoder
System. The Multiplexer/Eneoder System converts the various spacecraft analog
and digital signals to a serial binary-coded digital signal for presentation
to the data-dump tape recorder and the real-time telemetry transmitter. The
tape recorder records a portion of the real-time data from the progra_r at
a tape speed of 1 7/8 inches per second and, upon command, will play back the
data for transmission to a ground station, at a speed of _1.25 ips (22 times
the recording speed).
Four physiological functions are monitored for each pilot. All of the measure-
ments are supplied as real-time data, while only one is supplied as delayed-time
data. In addition, most of the measurements are recorded by two special (bio-
mad) tape recorders.
During pre-launch operations, data acquisition is accomplished by use of hard-
lines attached to the spacecraft umbilical and by telemetry. Between launch
and orbital insertion, data acquisition is via the real-time telemetry trans-
mitter. While the spacecraft is in orbit, data is acquired via the real-
time telemetry transmitter for the period while the spacecraft is within range
of a ground station. Data during the period while the spacecraft is out of
range of a ground station is recorded On the PCM recorder and played back via
the delayed-time telemetry transmitter while the spacecraft is within range of
a ground station. A more detailed description of the telemetry transmitters
is given in Section IX.
i0-I0
__@ SEDR300
PGEMINI
The paragraphs to follow, present a brief description of all instrumentation
parameters. The parameters are described in groups identified by their appli-
cable data source system. It should be noted that although most of the para-
meters are applicable to all spacecraft, the following parameters is for space-
craft 8 speeifically_
SEQUENTIAL SYS_M P_R_ME_ERS
A functional diagram showing the Sequential System parameters is presented in
Figure 10-4. The Instrumentation System monitors 41 sequential events and Se-
quential System parameters. Each parameter is described below individually,
or as part of a group of related parameters.
The Time Reference System (TRS) supplies three 24-bit digital words to the 24-bitj_
shift register of the PCM progr_er. These three signals are: time since lift-
off (_01, _02) and time to retrograde (_03). Time since lift-off is referenced
to the launch vehicle lift-off signal and provides time correlation for the data
tape recorders. Time to retrograde (MOB) indicates the time remaining before
retrofire initiation by the TRS. This signal is used to verify that the correct
retrofire time has been inserted into the TRS by ground command or by the pilots.
l_unch vehicle _econd stage cut-off (ABOI) is monitored for ground station indi-
cation of this event. This parameter is provided by a signal from the space-
craft IGS computer to a bi-level channel of the progra_er.
Launch vehicle/spacecraft separation (AB03) is indicated to the ground station
when any two of the three spacecraft/launch vehicle limit switches close, ener-
i0-ii
:--_=_ _SEDR 300 I _: __
__ PROJECT GEMINI____/OFF IEROM,'TO ABORT _ " ELECTRONIC
J U RSS H2
sw, j o_ ITMERI,INSTRUMENTATION FROM PLATFORM '
COMMON ABORT RELAY J MODE SW-ON POS
CONTROL CRYO QUANTITY SW | INPUT POWER FROM BIO MED _
BUS BOOSTER j I j RECORDERNO. 2 ABOi: ; FROM COMPUTER AF02
FROM EJECTION _ RIGHT,3
SEAT LIMIT SWS LEFT $ ABORT
AFO3 !; _ AF06
I LIGHT z AG21jSEP SWS SPACECRAFT SEP AG23
SENSOR RELAY I
AB03%=
L AB04COMMON
FROM BOOSTER IN AE28 R
SHUTDOWN SW _, !_ OORVV t/V ENGINE I J O C ; AE27 G
J SHUT DOWN SW =_ !!OG)UyE I AC03
RELAY _ DOCKING CATCH RELEASEi R.L AE02 A
1 FROM ADO2 M" DROGUE DEPLOY M
E
C_ AD03 R
FROM EQUIP l,I . -_ ! ]ADAPTER SEP. _ 2 -p II 4k YAW RATE
SWITCHES ADAPTER SEP PRIMARYSENSOR FROM DROGUE ROLL RATE RATE GYRORELAY DEPLOY SW _'_ _ DROGUE PITCH PATE REE. /
J DEPLOY J1 DISC
INST.
F : 1ROCKET FIRE I SECONDARY
CIRCUITRY ==_ ROCKET FROM PILOT J
AUTO FIRE CHUTE DEPLOY 4--- REF.RELAY SW J "-'_'
iFROM RETRO AGIO m_ I.._A__I
ROCKET FIRE [ _ RETRO ROCKET _ AGI I
CONTROL CIRCUITRY ==_ MANUAL
FIRE/2 _Y AG12
COMMON SCANNER AG17 AG17 /
FROMREI'ROT-_uT_ IcoNT_ sw SEC POS [ROCKET F RE BUS LDG & REC I LCONTROL J _ INST INSTRUMENTATION I J
c, c0, .Y, , oo oor7COMMON MANUAL O NO. 2
CONTROL ,,,L--1JFIt_N_LAYI _ FROM . PITCH-PRI _
RETRO INST " JETTISON ROLL-PR[ f AG03 AG03_ H
! A; SW AG04 AG04 GFROM RATE YAW-PRI [
IN CHUTE H)..,..,,J.FROM RETRO f _ JETTISON GYRO AG]3 AGI3 / I
._ SWITCHES PITCH-SEC g I
ROCKET FIRE " RETRO ROCKEP RELAY
MANUAL ROLL-SEE F 4 ICIRCUITRY FIRE E4 RELAy
_J YAW-SEC$ AG15 AG]5 R L I.] E EE E
_ AEI3 N L IT M I
_ _ ADO9R U
_' I • Y L IFROM MANUAL I ADO8 T IRETRO FIRE F ,J.l) REI"RO ROCKET
i
SW 2 MANUAL ADI0 P IFIRE El RELAY L
ADO6 E1 X" AF04 E
--O mm_'_ INPUT POWER AGI8II • FROM BIO MED RECORDER
FROM ABORT _ REF. SEQUENTIAL SYSTEM NO. I
_9 1SW SALVO AG06 '
RETRO RADAR PRIMARy POWER ;.
Figure 10-4 Sequential System Parameters Functional Diagram
10-12
___ SEDR 300 _____
PROJ E'-'CT GEMINI
gizing the spacecraft separation relays. Actuation of any two of the three
relays applies 28 vdc to a bi-level channel of the programmer.
Rendezvous Radar prlmarypower (AGO6) is a high-level signal applied to the re-
entry high-level multiplexer. This signal originates when the Rendezvous Radar
primary power switch is energized. Docking catch release (AC03) originates
during the separation sequence after docking has occured. Actuation of the
release catch mlcroswitch energizes the docking catch release relay in the
instrumentation relay panel, and provides a bi-level signal to the programmer.
Equipment section separation (AD02) is monitored to indicate a safe condition
for retrograde prior to manual initiation or ground command of retrofire as
,_ a backup to the automatic system. This signal is originated when any two of
the three separation sensors close, energizing the equipment section separation
relays. Actuation of two of the three relays applies 28 vdc to a bi-level
channel of the programmer.
The retrorocket ignition commands are monitored by ground stations to obtain
data for calculation of expected re-entry trajectory. Automatic (ADOB) and
manual (ADO6) ignition commands are monitored. Parameters are obtained from
the ignition command of the four retrorockets individually; ADO9, rocket 2;
ADO8; rocket 3; ADIO, rocket 4. The manual and automatic retrofire commands
indicate retrorocket I fire. The signals, 28 vdc, are applied to the re-entry
hlgh-level multiplexer.
Channel i0 of the Digital Command System is used by the ground station to relay
_ the abort command to the spacecraft. Verification of ABORT light lllumluatlon
lO-13
SEDR300
is by (_06) parameter.
Indication that the pilot actuated abort (AFOI) is supplied to the ground station.
The signal is originated when the abort handle is moved to the ABORT position
actuating a limit switch which energizes the instrumentation abort relays.
Actuation of one of the relays applies a signal to a hi-level channel of the
programmer.
In case of pilot ejection during an abort, left (AFOB) and right (AFO2) ejection
seat gone signals are relayed to the ground station. The signals are origi-
nated at the time the ejection seats leave the spacecraft closing the corres-
ponding limit switch and applying the signals to the bi-level channels of the
progr_-,ner.
Confirmation of salvo retrofire is given to the ground station in case of an
abort. A signal is applied to a hi-level channel of the re-entry high-level
multiplexer when the salvo retrograde relay is energized.
Indication of booster cut-off cap--and(ABO$) is given to the ground station
when pilots move the ABORT handle to the SHUTDOWN position, actuating a limit
switch. This energizes a relay applying 28 vdc to a bi-level channel of the
programmer.
Ground indication of pilot parachute deployment (AE_2) is provided via a bl-
level channel of the programmer. The signal is originated when a lanyard from
the parachute actuates a toggle switch, energizing the pilot parachute deployed
instrumentation relay.
Io-I
SEDR300
.... PROEMINI
The parachute Jettisoned (AEI3) signal is initiated when the pilot depresses
the CHU_ JETT switch energizing redundant m_in parachute jettison relays.
The relays apply a 28-vdc signal to a bi-level channel of the re-entry high-
level multiplexer.
Platform mode selection (_G05) is indicated to a ground station. Any position
other than OFF on the PLATFORM mode switch will apply a signal to a bi-level
channel of the programmer.
Primary (AGI6) or secondary (AGIT) horizon scanner operation can be monitored
by the ground station via bi-level channels of the prograswaer.
Primary pitch (AGO2), roll (AGO3), and yaw (AGO4) and second-v-j pitch (AGI3),
roll (AGI4), and yaw (AGIS), rate gyro operation is monitored to indicate an on or
off condition. Each signal is applied to a sign_! conditioner whose output
is applied to a hi-level channel of the high-level multiplexer.
Pitch (AGIO), roll (AGII), and yaw (AGI2) rate gyro (prJ._ry or secondaz_ depending
which is operational) outputs are applied to three signal conditioners. Each of
the signal conditioners is a transistor switch providing no output for an input
of 0-0.325 volts and a 16. 5 volt output for an input greater than 0.325 volts.
The conditioned signals are applied to bi-level input channels of the progran,,er.
Bio-medical tape recorder on-off signals (AGIS, AGI9) are used for time corre-
lation of the recorded bio-medical data with the telemetry data. An on-off
indication is provided to the playback recorder and to telemetry by a hi-level
10-15
SEDR 300
PRoJ-ae-e- -oEM,N'
signal to the programmer (AGI9) and re-entry high-level multiplexer (AG18).
Drogue parachute deployment (AE27) and drogue release (AE28) can be verified by
the ground station via hi-level channels of the programmer. The signals are
initiated when the HI-ALT DROGUE switch is depressed.
The selected cryogenic quantity switch position is indicated to the ground station
by AG21 (Reactant Supply System oxygen), AG22 (Reactant Supply System hydrogen),
and AG2B (Environmental Control System oxygen) to allow the ground station to
identify the reading of CA09 described under Environmental Control System.
ELECTRICAL POWER SYSTEM PARAMETERS
Figure 10-5 shows a functional diagram of the Electrical Power System parameters.
Approximately 24 Electrical Power System parameters are monitored by the Instru-
mentation System. The parameters are listed and described in the following sub-
paragraphs.
Fuel cell oxygen (BA02) and hydrogen (BA04) tank pressures are monitored by
dual potentiometer pressure transducers installed as part of the fuel cell
system. Each dual transducer provides one output to the adapter hlgh-level
multiplexer and the other output drives an indicator on the instrument panel
in the cabin.
To evaluate proper operation of the fuel cell# stack IA (BDO1), IB (BD02),
2A (_E01), 2B (_E02) and section I (BHOI) and 2 (BH02) currents are monitored
and transmitted to the ground station. Stack C currents are obtained mathe-
matically by subtracting section A and B currents from the corresponding section
current. The signals being monitored originate from 50 millivolt shunts. The
Io-16
.---'2=_- SEDR 300
/
F.C.HEATH2TEMP J_ F.C. O_ SUPPLY JATEXCH, OUT. PRE_'SURE
J IBC03 BA02 HIGH -LEVEL
F.C. 0 2 TEMP PLEXER
AT HEAT EXCH. OUT (ADAPTER)fiB05 LOW- F.C. H SUPPLY
LEVEL PRE_?SUREMULTI- BAD4
J I PLEXER
F.C. H SUPPLYT_P (ADAPTER)BA06
F.C. SECT I 0 2. t
H2 L_P SWITCH BC05
F.C,0 2 SUPPLY
TEMP
BA05
COMMONBUS AP I
CONTROL MAIN F*C. SECT I O -_
FUEL CELL _ H20 SW iT_:2FI BB07
CNTL I H2 DO6 I BUS FUEL CELLJ_ PANEL
V° o0, o-V-o--iSECT I •
F.C. SECT 2 O -
PRO- H2/_P _VITCH 2 BC06COMMON GRAMMER
CONTROL FUEL CELLBUS
BE
O_)_O O BE F.C. SECT 20 2-
SECT2 _22 T _ _ H20 _ SWITCH BB_
PURGE
TO O PURGE REF.K V'_ LVES ELECTRICAL
TO ,I _- POWER _ _ _ )STACK CONTROL SYSTEM Y
,f--\ _ SWITCHES ] TO S/C INDICATORSREF ELECTRICAL
COMMON POWER SYSTEM
co% oL SEOLTBO_'_O CNTL BG04- BG04--
O. o,N0LTBUS TEST BO01 BG01 HIGH-OAMS LEVEL
[_ SQUIB OAMS SQUIB MULTI-
BUS 1 BUS MON I PLEXERO_O -BG02 -- -BG02 - (RE-ENTRY)
OAMS
QUIB OAMS SQUIB
US2 tf'_ _BUSMON2 BG03 BG(_F.C. SHUNT STACK IA
INSTRU-F.C. SHUNT -BD01 MENTATION -BD01 -
PACKAGE
NO. 2(RE-ENTRY)
ELECTRICALSYSTEMPOWERREF. _ MAINSHUNT ! BD02 BD02
F.C. SHUNT NO. 1 BH01 BH01 LOW-LEVELMULTI-PLEXER(RE-ENTRY)
F.C. SHUNT BE01 BE01
MAIN J_l --BE02 BE02
SHUNT.._ -BH02
/_" NO. 2 BH02
Figure 10-5 Electrical Power System Parameters Functional Diagram
10-17
__ SEDR300
PROJECT GEMINI
shunts are installed at the main buses for the section, and in the lines from
stacks A and B to the main buses for stack A and B currents. Each of these
signals is conditioned to a 0 to 20 millivolt signal which is directly propor-
tional to the input current and then applied to the re-entry low-level multi-
plexer.
The following parameters relate to the ground station information regarding
spacecraft main, squib and control bus voltages: BGOI (main), BG02 (squib l),
BG03 (squib 2), BC_ (control bus). Each of these parameters is conditioned
and then applied to the re-entry high-level multiplexer.
The Reactant Supply System (P_S) 02 (BA05) and H2 (BA06) supply bottle tempera-
tures are monitored by means of two temperature sensors located on each supply .....
bottle. The output of the sensors is applied to the adapter low-level multiplexer.
Fuel cell section I 02 to H2 (BCO?), section 1 02 to H20 (BB07)3 section 2 02
to H2 (BOO6), and section 2 02 to H20 (BB08) differential pressures are
monitored by a pressure-sensitlve switch installed within the fuel cell to provide
for safe operation monitoring capability of the fuel cell by the ground station.
The outputs of the pressure switch is applied to hi-level channels of the
adapter high-level multiplexer.
Oxygen (BB05) and hydrogen (BC_3) temperatures at the outlet of the heat ex-
changer are monitored and relayed to the ground station via the adapter low-
level multiplexer.
To provide an aid in evaluating fuel cell operation by the ground station,
lO-18
___ SEDR300
PROJE-C GEMINIf_
section 1 02 (_D0_), section 2 02 (_0_), section 1 _ (BD0_), and section 2
H2 (BE06) purging is monitored. The signals are actuated by the pilot by
placing the corresponding section purge switch to the H2 or 02 position. The
signals are applied to the hi-level channels of the programner.
ENVIRONMBPrAL CONTROL SYS_ PARAME_
A functional diagram showing the Enviro_ntal Control System (ECS) instrumen-
tation parameters is presented in Figure 10-6. TwentT-eight parameters and
RSS/ECS qu-ntities associated with the ECS are monitored by the Instrumentation
System and relayed to the ground station for ana_sis.
The primary oxygen tank pressure (CA02) is telemetered to the ground station and
_-- displayed in the spacecraft cabin. The signals originate from a dual poten-
tiometer pressure transducer installed as part of the ECS. The signal is re-
layed to the ground station via the adapter high-level multiplexer.
A differential pressure transducer is used to sense cabin to forward compart-
ment pressure differential (CBOI). The transducer has a dual output used for
cabin indications and for transmission to the ground station via the re-entry
high-level multiplexer.
Left (CCOI) and right (CC02) suit to cabin differential pressure is displayed
in the spacecraft cabin and telemetered to the ground station. Dual potentio-
meter pressure transducers serve as the signal source. The output of each
transducer is applied to the cabin indicator and to the re-entry high-level
multiplexer.
t
z0-19
i :'_- SEDR 300
"!_ PROJECT GEMINI
TEMPERATURES PRESSURES
PRIORY LOOP I PRESSURE J
CH02 CLOI
SECONDARY OUTLET I INLET-SEC JCH0_ CJ02
I RADIATOR CONTROL I COOLANT PUMP
VALVE-PRIMARY I INLET-PRI ICDO3 LOW-LEVEL CJOI
I MULTIPLEXER J(ADAPTER)
INLET TO F C SECT 2 I PRESSURE JCD02 " '
I _°_ I
, ,II c°°' i I i
CA06 -- __ J CAoBSUPPLYPRESSSYSTEM I --I TE/_° SENSOR
AG2_
i I , [ lI VALVE -SEC CABIN PRESSCD04 I AG22 CC01 I
I I I iII CB02 CABIN PRESS --hCC02
I AIR INLET J PRESSUREDETECTOR
CC03 CC06
LOW - LEVEL
I I _L,,,LE_R I(REENTRY)
I AIR INLET j PRESSURE
CC04 CB01
I Ii i io co_I HEAT EXCHANGER I SUPPLY PRESS, SYBT §2
CK06 CA04
L J L _J
Figure 10-6 Environmental Control System Parameter Functional Diagram (Sheet I of 2)
10-20
3ooPROJECT GEMINIL_
OFF J
ECS 02 ECS 02 PRIMARY 02_ O_O iI'-. O QUANTITY TANK QUANTITY
CONTROL UNIT SENSOR
CRYQTY J dF.C. 02 TO 02MAIN IBUS l OF'C° H2 r ¢ PRESSURE
TRANSDUCER
I--_ F .C. 02 F.C. 02 TANKQUANTITY QUANTITY
I CONTROLUNIT SENSORI O
TO H2
l 4 PRESSURE
O (REF FIG 10-4)
F,C. H2 F.C. H2 TANK
SELECT SW O O_ QUANTITY QUANTITYCONTROL UNIT SENSOR
DC-AC AND QUANTITY DC-ACINVERTER INDICATOR INVERTER
J F .C. 02 I
DC-ACINVERTER
J'_ _ _ _ BI-LEVEL I
f-_" J HIGH-LEVEL J
MULTI PLEXER
(ADAPTER)
PRIA ,o i,. co ' 0-U
,_ON TOPRIPUMPAO O
ECS IND LTS vtvi PRIMARy r POWER SUPPLY
C_O PUMPA" 0 0 _ TO PRI PUMP A CHECKf VALVE LIMIT SWITCH
PRI B
O O _ ON TO PRI PUMP BI PRIMARY I POWER SUPPLYI PUMP B
TO PRI PUMP B CHECKt _ I _ VALVE LIMIT SWITCH REF
SEC A ECS
O O _ ON s TO SEC PUMP A SYSTEMI SECONDARY POWER SUPPLY] PUMP A
_ TO SEC PUMPA CHECK' "{VALVE LIMIT SWITCH
SEC B
0_...,_0ON0 O _ TO SEC PUMP B
I SECONDARY POWER SUPPLYI PUMP B
MAIN _ I TO SEC PUMP B CHECKBUS VALVE LIMIT SWITCHVOLTAGE
COMMONHIGH-LEVEL MULTIPLEXER O O CONTROL(RE-[ NTRY) CC05 I BUS
I
ECS (! T "( 02 RATE VALVE
I DISCONNECT CIRCUITRYRELAY REFECS
' m 02 RATE VALVE SYSTEM
L__ iFigure 10-6 Environmental Control System Parameter Functional Diagram (Sheet 2 of 2)
10-21
__ SEDR300
PROMINI
_e _o_ station is informed of am 02 high z_te co-dttto_ by CCO_. This signal
is orlslzmted when the s_ceere_t ¢J_ FAN mrltch is placed in the 02 HI RATE
pOSltlC_, when :m41_i 02 high rate Is selected by the pilot, or 1_en the sult
pressure drops below 3-3 psla and 02 high rate is automatically selected.
The signal is al_I/ed to a hi-level channel of the re-entz_ high-leTel multi-
plexer.
TO assure that a safe st_pplyof oxygen is available tO the pilots, CO2 partial
pressure (C006) is monitored indicating the percen_e of carbon dioxide with
respect to the total pressure of gas in the suits. C02 partial pressure is dis-
played in the spacecraft cabin _ applied to the re-entry hi6h-level multiplexer.
Primary and secondary coolant temperatures are monitore8 at various loc_tions
within the coolant loop to evaluate systmn performance. Coolant temperatures are
monitored at the primary coolant t_let to .ection 1 of the fuel cell (CDO1),
secondary coolant Inlet to section 2 of the fuel cell (CD02), the radiator control
valve in the primary loop (CD03), the secondary loop (CDO_), radiator outlet in the
primary loop (C_), and radiator outlet in the secondary loop (CH03).
To relay inforaation concerning proper operatic_ of the coolant loop and pumps,
to the ground statio_, primary (C_OI) and secondary (CJ02) coolant pump inle_
pressures are monitored. The outputs of the transducers are applied to the
adapter high-level multiplexer.
The condition of the primary and seco_lary coolant pmnps is monitored by CJI6
A), c.n7 p,p B), C.U8(seconc ryp =pA), andC.T'-9
10-22
__. SEDR300
PROJECT GEMINI
(secoedary _ B). The signal is orlgLJ_,,_ v'aen the eorrespc_Lng eonl.ant
Pump iS ae_t;ed, and Is applied to bi-level _-ls c_ the a_sl_er bl_-level
multiplexer.
To i_vze safe opersti_ of the fuel cell, _ter. pressure (C_Ol) is monitored
at the OUtl_ _ the fuel cell. _ sl_l Is appUed to the adai_er bl_=
level multiplexer •
The coo].m_ inlet _mperature to the suit ]_at exchanger ((_) is monLtored to
relay to Ilzom_ starless information concerning the environmental conditio_ of
the pilots. _e eutput e_ the temperature mensar is applied to the re-entry
low.level multlplez_r.
/-_ The poslttcn of the eryopnic quantity select switch Is mcmitored to identify
l_rame_er ¢_0_. The parameter CAO_ t_tca_es ECS 02, _ 02, or _SS H2 quantity
dependi_ upon the poeittc_ of the cryogenic quantity select swatch. The
position o_ the selector swatch is indicated to the ground station by AG21
(_.c. 02), A_ (_.C. _), ana A_Z3(_CS02). m_ s±_nals a_ appl±ea to b_-
level cha--els of the pro_ra_er. The parameter CAO_ is also applied to the
pr_r and is di_pl_yed in the spacecre_t cabin.
Secondary 02 supply pressures are monitored in the number i (CA03) and number
2 (CAO_) systems. The transducers are installed as part of ECS secondary 02
supply assemblies. The outputs of the pressure transducers are applied to the
re-entry high-level multiplexer.
As an aid in calculating ECS 02 quantity, the primary 02 supply bottle tempera-
_o -23
_@ SEDR300 _..__
PROJECT GEMINI
ture (CA06) is monitored and applied to the adapter Low-level multiplexer.
To provide the capability for the Kround station to monitor the environmental
conditic_ of the cabin and to provide an aid for evaluating suit pressure, a
cabin air teEperature tz_nsducer (CBOe), and a folw_rd compartment absolute
pressure tz_-ndueer (CBOT) is provided. Absolute pressure is applied to the
pr_r and cabin temperature is applied to the re-entry low-level multi-
plexer. Cabin temperature is also displayed in the spacecraft cabin.
To further evaluate system performance and pilot envir_ental condition,
the air entering the suit circuit Is monitored with respect to te=perature by
2 dual temperature sensors (1 for each suit circuit). The te:tperatuzes are
displayed in l_e spacecraft cabin and are applied to the re-entry low-level
_ltiple_er as _03 (left suit), and CCO_ (right suit).
IREm_AL GUIDANC__ P_
Figure 10-7 shows a block diagram of the Inertial Guidance 8ygtem (IGS) param-
eters except the digital computer ftmctt_n,_-. The Instr_entatton System
monitors 8 IGS parameters and handles appro_wately 200 conputer _ords.
The Instrumentation System monitors the computer modes of operation; pre-launch,
ascent, catch-up, rendezvous, re-entry, and touchdown. Important functions or
parameters (approximatel_ 200) are monitored during each mode of operation.
This information is used during post mission analysis and is applied to the
programmer.
In addition to the digital computer words, the Instrumentation System monitors _'_
the following IGS parameters.
lo-z_
_": _'_ SEDR 300
DE04
HIGH-COMPUTER DE05 LEVEL
MULTIPLEXER(RE"ENTRY)
DC04
YAW SYNCHROREPEATERDQ09
INERTIALMEASURING PITCH SYNCHROUNIT REPEATER PROGRAMMER
DQ07
ROLL SYNCHROREPEATERDQ_
DW01 LOW LEVELMULTIPLEXERADAPTER
HiGH-AUXILIARY LEVELTAPE MULTIPLEXERMEMORY DW02 (ADAPTER)
DW03
DW04
Figure 10-7 Inertial Guidance System Parameters Block Diagram
10-25
__ SEDR 300 __
PROJEC:lr" GEMINI
Inertial platform attitudes are monitored to provide ground stations with
attitude data during flight. Roll (DQ08), pitch (DQ07), and yaw (DQ09) signals
are taken from the Inertial Measuring Unit (IMU), conditioned by synchro re-
peaters, and applied to the progrs_er.
IGS regulated power is monitored at two points : 26 vac (DE04) and 10.2 vdc
(I_05). These voltages are conditioned and then applied to the re-entry high-
level multiplexer.
Computer START light (DC04) malfUnction signal is the only malfUnction detection
parameter monitored. This signal is used for display in the spacecraft cabin
and applied to the bi-level channel of the programmer.
Auxiliary Tape Memory environmental conditions are monitored by case temperature __
(DWOI) and internal pressure (DW02). Motor drive inhibit (DWO3) and verification-
reproduction plus 20 volts (DW04) parameters indicate the mode of operation (off,
standby, read, or write).
ATTITUDE CONTROL J_NDMANEUVERING ELECTRONICS PJ_RAME_ERS
A block diagram showing the Attitude Control and Maneuvering Electronics (ACME)
System parameters is shown in Figure 10-8. Fifteen ACME parameters are monitored
by the Instrumentation System.
Spacecraft rates in pitch (F&OI), roll (EA02), and yaw (F&O3) are monitored
to allow evaluation of the rate control portion of the ACME. Each signal
from the rate gyro package is conditioned by a phase sensitive demodulator and
then applied to the high-level channels of the programmer. Primary and secondary
rate gyro signals are parallel summed and monitored on the same channels.
10-26
J--:_- SEDR300
!"
J SEARCH EBO3
I PR ffv'_ARy TJ HORIZONSENSOR ROLL EB02 EB02
J (REF) HIGH-LEVEL
J PITCH EB01 EB01 MULTIPLEXER
J _J J (RE-ENTRY)
j ECOI ECOI
EC02 EC(Y2
SECONDARYHORIZONSENSOR
(REF)
_J
I 1j SIGNALCONDITIONER
J AC_E PACKAGEINV. NO. 2
l (REF)
IIl J
1PITCH EA01 EA01
PRI RATE ROLL EA02 EA(]@
" GYRO T(REF) J YAW _ I EA03 EA03
[I 1__I SECRA,E II GYRO
I <REEl I
I-.._ I
1I EG01
IEG02
J EG03
IACE
(REF) _" EG04
J EG05
[J. EG06
IEGO7
I_ J
Figure 10-8 Attitude Control & Maneuvering Electronics Parameters Block Diagram
10-27
o00PROJECT GEMINI
Horizon Sensor operation is monitored with respect to pitch (EB01) and roll
(EB02) outputs, and search mode of operation (EB03). Pitch and roll paremeters are
monitored to verify inertial platform alignment for the retrograde phase of the
mission. These parameters (EBO1, EB02) provide pitch and roll attitudes from the
Horizon Sensor during orbital flight when the platform has been shut down to con-
serve electrical power. The signals originate when the SCANNER switch is in the
PRI or SEC position. The pitch and roll outputs are conditioned and then applied
to the re-entry hlgh-level multiplexer. The search mode of operation is monitored
to determine whether the Horizon Sensor unit is in the search mode, or has sensed
the horizon. This signal illuminates the SCANNER light in the cabin and is also
applied to a bi-level channel of the re-entry high-level multiplexer.
ACME inverter 26 vac voltage (EC01) and frequency (EC02) is monitored for post
mission analysls. The signals are conditioned and then applied to the re-entry
high-level multiplexer.
The following attitude control modes are monitored depending upon the position of
the A_'ITITUDECONTROL switch; HOR SCAN (EG01), RATE CMD ORBIT (EG02), DIRECT
(EG03), _ (EGO4), RATE CMD RNTY (EGO5), RE-ENTRY (EG06), and PLATFORM (EG07).
The signals are applied to bi-level channels of the progra_,er.
ORBIT, ATTITUD_ AND MANEUVERING SYSTEM PARAMETERS
The Orbit, Attitude and Maneuvering System (0A_8) parameters are shown in Figure
10-9 in block diagram form. A brief description of each of the 19 parameters is
given in the paragraphs to follow.
To insure that adequate propellant pressure is available for OA_8, helium source
10-28
__ SEDR 300 _ "----[
PROJECT GEMINI
1TEMPERATURES j
[ F,,o ]lHEAD TEMP GD06 J
II REG.oA1. I I
OX/DIZERTANK GC04 j LOW- LEVEL
L REG He AT l I MULTIPLEXERFUEL TANK GC03 I (ADAPTER)
I
I , sou cEj ITEMPERATURE GC02 I
I OXIDIZER FEED l lTEMP GB02 II
i FUELFEEOI;TEMP GB01 I
I
f FIRING CMD FROM OAME (1YP 14 PLACES)
GE011.O TCA NO. I FUEL & OXIDIZER VALVES
GE02• S 1"O TEA NO. 2 FUEL & OXIDIZER VALVES
GE03• ( TO TEA NO. 3 FUEL & OXIDIZER VALVES
GE04lid _ TO TCA NO. 4 FUEL & OXIDIZER VALVES
GE05• _ TO TCA NO. 5 FUEL & OXIDIZER VALVES
GE06 • f TO TEA NO. 6 FUEL & OXIDIZER VALVES REFORBIT
GE07• .( TO TCA NO. 7 FUEL & OXIDIZER VALVES ATTITUDE
AND
GE08 • f TO TEA NO. 8 FUEL & OXIDIZER VALVES MANEUVERSYSTEM
GE09• _ TOTCANO. 9&IOFUEL&OXIDIZERVALVES
HIGH - LEVEL GEl I• _ TOTCANO. 11 &I2FUEL&OX[DIZERVALVES
MULTIPLEXER(ADAPTER) GEl 3 • 5 TO TCA NO. 13 FUEL & OXIDIZER VALVES
GEl4• _ TO TCA NO- 14 FUEL & OXIDIZER VALVES
GEl5• f TO TCA NO. 15 FUEL & OXIDIZER VALVES
GEl6• J TO TCA NO. '6 FUEL & OXIDIZER VALVES
r - pRESSURES |I I
RESERVETANK ! I"F" PACKAGE GC22
I Ii l sEE_UcMEPRESS" GCOl l J
I Ii L .EL,O__G JPRESS-PKG B GC(_ I
F- I It J
Figure 10-9 Orbit, Attitude & Maneuvering System Parameters Block Diagram
10-29
__ SEDR300
PMINI
pressures (GCO1) is monitored. The signal ortgt_tes fz_la a duel potenttrmeter
pressure transducer at the helium pressura:t _ks. One output is applied to
the adapter blab-level multiplexer and the other is used to drive an indicator
in the spacecraft cabin.
The propellant feed teaperature at the fuel (GBO1) and oxidizer (GB02) feed
lines is monitored to verify that propellant aboard is above freeztn8 tempera-
ture and is available for use. The signals ori_Lnate _m two individual temp-
erature sensors and are applied to the adapter low-level multiplexer.
To allow monitoring capability of the he].t_a source temperature (GC02)_ a
temperature sensor is installed on the helium suppl_ line at the supply tank.
The output is applied to the adapter low-level amlttplexer. A separate sensor
is installed to drive an indicator in the spacecraft cabin.
Temperature of the pressure regulated helium is monitored at the fuel (GO03)
and oxidizer (GC0_) tank inlet lines. The outputs of these temperature sensors
is applied to the adapter low-level multiplexer. Two additional sensors are
installed to drive indicators in the spacecraft cabin.
Regulated helium pressure (GC05) is monitored by a dual potenticueter pressure
transducer. One of the outputs is applied to the adapter high-level multi-
plexer, and the other is used to drive a cabin indicator. Reserve tank pressure
(GC22) monitors pressure available to the reserve fuel tank.
To provide an indication of ma_m_n Thrust Chamber Assembly (TCA) temperature,
TCA I0 (GD06) injector head temperature is monitored. This signal is applied
to the adapter low-level multiplexer.
io-3o
__. SEDR 300
PROJ EC"T--'GEM IN I
To provide ground station monitoring capability of TCA firing, the following
TCA solenoid command signals are applied to hi-level channels of the adapter
high-level multiplexer: GEOI (TCA i), GE02 (TCA 2), GEO3 (TCA S), GE04 (TCA 4),
GE05 (TCA 5), GE06 (TCA 6), GE07 (TCA 7), GE08 (TCA 8), GE09 (TCA 9, IO), GEII
(TCAii, 12),GEl3 (TCA13),GEl4 (_A 14),GEl5 (TCA15),and GEl6 (TCA16).
HE-ENTRY CONTROL SYSTEM PARAMEteRS
Figure iO-I0 shows in block diagram form the Re-entry Control System parameters.
Some 2_ parameters are monitored by the Instrumentation System to provide for
ground station observation of proper system performance.
Nitrogen source pressure, HC01 (system A) and HC02 (system B), and nitrogen
_ source temperature, HC05 (system A) and HC06 (system B) are monitored. Pressure
is sensed by two dual pressure transducers. One of the outputs of each trans-
ducer is used to drive a cabin indicator, and the other is applied to the pro-
grammer. Outputs of the temperature sensors are applied to the re-entry low-
level multiplexer and are used to drive a spacecraft cabin indicator.
Because the oxidizer has a more critical temperature range than fuel, its tem-
perature is measured to insure that both fuel and oxidizer are within the proper
temperature range for use in the Re-entry Control System. The oxidizer feed
temperature (HA02) is applied to the re-entry low-level multiplexer.
Regulated nitrogen pressure is monitored for system A (HCO3) and system B
(HC04). The outputs of the pressure transducers is applied to the programmer.
To provide for ground station monitoring capability of proper RCS TCA firing,
firing commands are applied to bi-level channels of the re-entry high-level
lO-31
__ SEDR300 __
PROJECT GEMINI
IPRESSURES j
ITO FUEL & OXIDIZER
REG PRESS J j. t VALVES TYPICALHC04 I HEOI , A'CE FIRING CMDS 16 PLACES
FROM RCSMODEI _ SWITCHII HE02
I
[ I'RING 8 N2 J HE03SOURCE PRESS I
HC02
HE04
PROGRAMMERRCS
SYSTEM A
REG PRESS IHCC_ I
II HE06
II
I I'RING A N2 I HE07SOURCE PRESS I
.col II HEO8
I _ _L l HIGH- LEVEL -_ -_
MULTI PLEXER 0 _
(RE-ENTRY) HF01 ._
TEMPERATURES ZL
OXIDIZER FEED I _"
TEMP SYST A I HF02 ,_
HA02
HF(132
N2 SOURCE TEMP I LOW- LEVEL HF04
SYST A I MULTIPLEXER "
HC05 J (RE-ENTRY) RCSl )" SYSTEM B
II HF05 _'
II
I I HFO6 ,
N2 SOURCE TEMP ISYST B IHco_ I HFO7
I EIII HF08
I' l f 'RETRO ROCKET J LOW- LEVEL
14 CASE TEMP I MULTIPLEXER
HHOI I (ADAPTER)
I _"J
Figure 10-10 Re-Entry Control System Parameters Block Diagram
10-32
SEDR300
multiplexer. RCS system A thrusters, 1A thru 8A have been assigned parameters
HEO1 thru HEO8 respectively and system B thrusters IB thru 8B are designated
by HFO1 thru HF08 respectively.
The retrorocket case temperature (HHO1) was monitored on spacecraft 5. The
signal originated from a surface mounted temperature sensor located on retro-
rocket number 4 and was applied to the adapter l_.level multiplexer.
AERODYNAMIC AND CREW CONTROL PARAMETERS
Aerodynamic and crew control parameters are monitored as shown in block diagram
form in Figure I0-ii.
t_ Spacecraft longitudinal (KAOI), lateral (KAD2), and vertical (KAOS) accelerations
are monitored to provide ground station indications during the launch and re-entry
phases of the mission. The accelerometer outputs are applied to the programmer.
Static pressure (KB02) is monitored by a potentlometer type absolute pressure
transducer. Static pressure is obtained from four static pressure ports equally
spaced around the forward part of the conical section and connected in parallel
to the transducer. The transducer output is applied to the programmer.
Pitch (FAOI), roll (FA02), and yaw (FADS) attitude control stick positions are
monitored to indicate pilot manual control usage and to evaluate thruster opera-
tion. Signals originate from the attitude hand controller potentiometers and
are applied to the programmer.
Two bi-level channels are reserved for events to be monitored as required by
the experiments of each particular spacecraft mission. Electrical provisions
for monitoring these parameters are provided at the right (FD01) and left
10-33
.._T:=_. SEDR 300
RIGHT AUX HIGH - LEVELRECEPTACLE MULTIPLEXER
FD01 (RE-ENTRY)
LEFT AUXRECEPTACLE
FEOl
STATICPRESSURE
KB02
PITCH FA01
ATTITUDEROLL FA02
HANDCONTROLLER
YAW FA03
X VIBRATION0.02GQC28
PROGRAMME R
Nx
ACCELEROMETERKA02
y VIBRATION0.02GQC30
NyACCELEROMETER
KA03 XVIBRATION0.2GQC27
Y VIBRATIONNz 0.2G
ACCELEROMETER QC29KA01
Figure 10-11 Aerodynamic and Crew Control Parameters Block Diagram
10-34
__ SEDR300
PROJ EC'T GEMINI
(FEOI) utility receptacles.
COb_GNICATION SYSq_M PARA_q_S
The Instrumentation System monitors ll Co_munlcation System par--,_ters. These
parameters are shown in block diagram form In Figure 10-12. A brief descrip-
tion of each of the parameters is presented in the _phs that follow.
To verify proper DCS performance and aid in malfunction isolation, the following
DCS parameters are monitored: diplexer (LAO_) and quadriplexer (LA03) receiver
signal strength, package temperature (LA05), 6 vdc reg-lated power (LA02), 28
vdc regulated power (LA06), -18 vdc regulated power (LAO7), 23 vdc regulated power
(LA08), and-6 vdc regulated power (LA09). Parameters LA02, LA06, LA07, LAO8, and
LA09 are conditioned and then applied to the adapter high-level multiplexer except
_ LA02 which is applied to the adapter low-level multiplexer. Parameters LAO3, LAO_,
and LA05 are applied directl_ to the adapter low-level multiplexer.
Acquisition aid beacon (LD01) and adapter C-band beacon (LCOg) ease temperatures
are also monitored to assure proper equipment performance. These temperature
signals are applied to the adapter low.level multiplexer.
INS_TION SYS_ PARA_TERS
To insure proper operation of the Instrumentation System, various reference
voltages and other pertinent data is telemetered to ground stations for --a_ysls.
The Instrumentation System parameters are shown in Figure 10-13 in block diagram
form. A brief description of each parameter follows.
10-35
_.--;--_. SEDR 300
l._k09 LA09
L_,06 LA06
DIGITAL LA07 INSTRU - LA07 HIGH - LEVELCOMMAND MENTATION MULTIPLEXER
SYSTEM PKG. NO. ] (APT)
LAO_ LAO8
LA02 LA02
LA03
ACQUISITION AIDBEACON TEMP
LD01
LA04
LAOI
LOW- LEVEL ADPT C-BANDMULTIPLEXER BEACON TEMP
(ADPT) LCO?
INSTRU-MENTATION LA01 PROGRAMMER
PKG NO, 2
DCS PACKAG ETEMPERATURE
LA05
Figure 10-12 Communication System Parameters Block Diagram
10-36
______ $EDR300
PROJECT'-GEMINI
High (MALT) and low (MA38) level zero reference voltages are monitored to
insure that proper scaling is being employed by the Multiplexer/Encoding
System. The low-level zero reference originates from the 5 vde output of the
dc-dc converter which is attenuated by a signal conditioner to 3 millivolts
(the zero reference point) and is then applied to a channel in each of the
multiplexers in the re-entry vehicle. This signal is also applied to the
progra.,.,eras MA38.
High (MA37)and low (MA21) level full scale reference voltages, as the zero
reference voltages, are required to insure that proper scaling is being employed
by the Multiplexer/Encoding System. The 5 vdc output of the de-de converter is
attenuated to 4.5 vdc and to 15 millivolts prior to application to channels of
the high and low-level multiplexers in the re-entry vehicle, respectively.
These parameters are required to provide a measurement of the reference voltage
for potentiometer type transducers and resistive element temperature sensors.
Reference voltages for the high and low-level multiplexers located in the equip-
ment adapter are provided by _BOI (full scale) _B02 (zero) and _03 (full scale).
These have the svme characteristics as the reference parameters described above.
Parameter MA22 (calibrate signal) is provided to indicate that a calibration
voltage is being applied, thus eliminating the confusion between a data and
a calibrate signal. Parameter MA22 will exist whenever the CALIB switch in
the spacecraft cabin is actuated or a calibration is eo_nded by the DCS.
IO-37
_._- _. SEDR300
MB01
INSTRUMENTATION HIGH- LEVEL LOW- LEVELPACKAGE -- -- M U'LTIPLEXER MULTIPLEXER
NO. 1 (ADAPTER) (ADAPTER)
MC01 MC02
HIGH -LEVELMULTIPLEXER MD0_ CAMERA. EVENT D/T TM
(RE-ENTRY) MECHANISM TRANSMITTERMA3;
MA21LOW - LEVEL
INSTRUMENTATION _ MULTIPLEXERPACKAGE MA38
NO. 2 j _ (RE-ENTRY)
HIGH - LEVEL MA22 MAg'3 PCM TAPEMULTIPLEXER PROGRAMMER _ECORDER
(ADAPTER)
Figure 10-13 Instrumentation System Parameters Block Diagram
10-38
___ SEDR300
PROJECT GEMINI
An indication of proper functioning of the PCM tape reco_er is provided by
monitoring tape motion (MA95). This is accomplished by providing a signal to a
hi-level channel of the programmer when the recorder tape reels are in motion.
The rf _ power output (MC01) and the case temperature (MC02) of the delayed.time
telemetry transmitter is monitored to provide an indicatic_ of tranmnitter
operation. The transmitter physically located in the adapter is chosen for
these measurements because it is subject to more extreme environmental tempera-
ture changes than the other two transmitters. Temperature signals are applied to
th_ adapter low-level multiplexer, and the rf power output is applied to the
adapter high-level multiplexer.
_ A camera event (MD06) is indicated to the Kround station when the pilot initiates
the camera event mech_ism on the onboard camera. This signal is applied to a
_i-level c_-nel of the re-entry high-level multiplexer.
PHYSXOLOGXCALPARAMETEBS
The physiological functions of the crew are monitored by sensors which are
attached at various points to their skin. A block diagram showing the physio-
logical parameters is shown in Figure I0-I_. Signal conditioners, located in
pockets of the underwear_ condition the signal_ from the sensors to make them
c_atible with the recording and multiplexin_ equipment. All parameters except
the oral temperature are recorded on bio-medical recorders. _11 signals, except
oral temperature are applied to the programmer. Oral temperature is applied to
the re-entry high-level multipleKer. The following command pilot parameters are
_ monitored: electrocardiograms i and 2 (NAOI, NA02), respiration rate and depth
lo-39
___.% PROJECT GEMINI _l •
HIG H- LEVELMULTIPLEXER
(RE-ENTRY)
(RA03
NA06 DURING EVA)
S/C 5&6 ONLY NB07
NAOl
(NOT MONITORED DUP_ING EVA') NB01
NA02
PILOTCOMMAND NB(_PILOT
r,u_o3I ]
I NB03NA04 _ /C5&6 ONLY _ 'C5&6 ONLY- NB05
I
Ij irll
BIOMEDICAL BIO'MED ICALRECORDER g2 PROGRAMMER RECORDER II
Figure 10-14 Physiological Parameters Block Diagram
10-40
__ SEDR300
PROJECrF GEMINI
(NA03), and oral temperature (NA06). The followlng pilot parameters are monitored:
electrocardiograms 1 and 2 (1_01, N]_2), respiration rate and depth (N_03), and
oral temperature (1_06). The com_nd pilot's blood pressure (NA04) and the
pilot's blood pressure (HB05) And suit inlet temperature (I¢B07) was monitored
on spacecraft 5 and 6 missions.
Electrocardiogra= 1 (NBO1) is not monitored during Extra-Vehicular Activities
(EVA). To allow the ground station to evaluate the pilot's enviro_=ental
condition while outside the spacecraft, suit pressure (RA03) is monitored in
place of oral temperature (NB06). All other physiological parameters are moni-
tored as previous_v described.
I:_lfl_ZV'OUS RADAR PAI_L'_R$
p_ To permit post-mission anaS_sis of the Radar System operation during the rendez-
vous portion of the mission, 19 radar parameters are monitored. A functional
block diagram of the radar parameters is shown in Figure 10-15.
The target verification (JC04) parameter provides telemetry information that
the spacecraft radar has located the target vehicle. The radar LOCK-ON
indicator on the pilot's pedestal will illum_uate when this event occurs. Tar-
get range rate (JA04) information is relayed to the ground station via telemetry
transmission and displayed to the pilots on the range rate indicator. During
rendezvous, the range to target (JAI3) parameter provides close-ra_e data
which is not available through the computer.
Radar inner skin (JBOI), antenna faceplate (JB02), and transmitter tube (JBO4)
temperatures are sensed by resisitive-element temperature sensors. A output
/r_ which is proportional to the temperature sensed is produced by integral bridges
i0 -41
-:-_ SEDR 300
__ PROJECT GEMINI
JA04
JA13
JB01 LOW-LEVELMULTI PLEXER
(RE-ENTRY)JB02
JB04
IDOI
JO08
JC01
JC02
JCCG
RENDEZVOUS
RADAR JC04
JEOI PROGRAMMER
JE03
JEO4
JE05
JE06
JE07
JE_
JE09
AGO6 HIGH - LEVELMULTIPLEXER
(RE-ENTRY)
Figure 10-15 Rendezvous Radar Parameters Block Diagram
10-42
___.____ SEDR300 __.__j
PROJ EGT GEMINI
within the sensor. _ pressure transducer within the radar package senses radar
pressurization (JB03). This signal is used to determine the validity of the radar
operation, which is dependent on the internal pressure being maintained.
Oscillator crystal current (JC01) and radar regulated power (JEO1 and JEO3
through JE09) parameters provide information relative to the Radar System opera-
tion and monitor for specific functions which occur. Additional parameters
monitored to evaluate the operation of the radar are: rf power (JC01), agc
voltage (JCOR), and narrow-band agc voltage (JC03).
SYSTEM UNITS
PRESSURE TRANSDUCERS
The purpose of the pressure transducer is to sense pressure, and to convert this
pressure into a proportional electrical signal. There are six physically different
configured pressure transducers as shown in Figure lO-16. Transducers have
different physical appearances and different pressure ranges to accommodate the
specific application of use. The numerical call outs below each transducer in
Figure lO-16 identifies the location and application of the transducer as shown
in Figure lO-1. The numbers correspond to those on Figure lO-1.
The sizes of the units vary from about 1 1/4 inches x i 1/_ inches x 3 inches to
2 1/2 inches x R 1/2 inches x 4 inches; the weights vary from approximately 5
ounce to 1 pound. The unit construction utilizes a bellows or Bourdon tube which
varies the wiper position of a potentiometer, proportionally, with the input pres-
sure. Two potentiometers are used in the dual-output units to separate the cabin
10-43
SEDR 300
CASE
GROUND
-.QIo R c AI
SINGLE POTENTIOMETER WATER PRESSURE TRANSDUCER ABSOLUTE AND STATICTRANSDUCER SCHEMATIC REEFIGURE10-1INDEXNO. 19 PRESSURE TRANSDUCER
(TYPICAL) FOR LOCATION REFFIGURE 10-1 INDEX NO.39 & 40 FOR LOCATION
H_GH RESISTANCE LOW RESISTANCE ._..-'-- _ELEMENT ELEMENT
GROUND _".
IA _ C o E G_ I
DUAL POTENTIOMETER ECS SECONDARY SUPPLYTRANSDUCER SCHEMATIC PRESSURETRANSDUCER
(TYPICAL) REF FIGURE I0-I iNDEX NO.55 & 59 FOR LOCATION
RSS AND PRIMARY ECS SUPPLY OAMS PROPELLANT QUANTITY CABIN AND SUITPRESSURE TRANSDUCER PRESSURETRANSDUCER
REF FIGURE 10-I iNDEX NO. REF FIGURE 10-1 INDEX NO.
14, 21 & 30 FOR LOCATION 41, 57 & 58 FOR LOCATION
Figure 10-16 Pressure Transducers
10-44
SEDR300
PROJECT--G'EMINI
indicator circuit from the multiplexer/encoder (telemetry) circuit, thus avoiding
a possible loading error in the latter. With one exception, pressure transducer
outputs range from 0 to 5 vdc. The OAMS quantity system pressure-temperature
sensor, driving a cabin indicator, has an output of 0 to 24 VDC.
TEMPERATURE SENSORS
Temperature sensors are used to convert temperatures into directly proportional
electrical signals. Basically there are two types of temperature sensors:
a probe type and a surface mounted type. Variations exist within each type
to accommodate specific mounting requirements. Nine physically different
types of temperature sensors are shown in Figure 10-17. With respect to tem-
perature range, approximately 20 different sensors are used. The numbers benesthf_
each sensor in Figure lO-17 corresponds to the sensors locations and application
as shown in Figure lO-1.
Spacecraft temperatures are monitored by platinum element temperature sensors.
The sensors vary somewhat in size but are roughly 0.4 x 0.75 x 2.0 inches. There
are two types of resistive-element sensors, a probe type and a surface-mounted
type. Probes _e used to monitor fluid temperatures, and surface-mounted sensors
are used to monitor surface temperatures. Both types utilize a fully-annealed
pure-platinum wire, encased in ceramic insulation. The sensors form one leg
of a bridge network whose unbalance will produce an output of 0 to 20 mv dc or
0 to 400 my dc. The 0-20 mv dc outputs sre used for data transmission purposes
and the 0-400 mv dc outputs for cabin displsys.
/_ In some applicstions_ mounting and space requirements necessitate thet the bridge
i0-45
_.-_--_. SEDR 300
INRDT_ [
OUTPUT O # _1__ r" - -'I
INPUTO _:EN_ 'E_?J---
OUTPUT 0
TYPICAL SCHEMATIC
SURFACEMOUNTED PROBEAIR TEMPERATURE MOUNTEDSENSORAND BRIDGESENSOR AND BRIDGE SENSOR AND BRIDGE REFERTOFIGURE10-1
REFERTO FIGURE 10-1, REFERTO FIGURE 10-1 INDEX NO. 2, 3, I|, 17, 20,INDEX NO. 6,27,31 &34 INDEX NO. 42,54, & 60 22, 35, 48, 50,52 &56FOR LOCATION FOR LOCATION FOR LOCATION
RETROROCKET CASE SURFAC_I_ SURFACE MOUNTEDMOUNTED SENSOR AND BRIDGE BRIDGE PACKAGE SENSOR ELEMENT
REFERTO FIGURE 10-t REFERTO FIGURE 10-1 OAMS QUANTITYINDEX NO. I, 4, 25, 32 & 33 INDEX NO. 54 & 60FOR LOCATION FOR LOCATION
'"DGEPACKAGE :ELP:,Vp , LBEEREFERTO FIGURE I0-I REFERTO FIGURE I0-1 REFERTO FIGURE I0-]INDEX NO. 13, 23 & 29 INDEX NO. _3, 23, 29 INDEX NO. I0, 12, 15 & ]8FOR LOCATION FOR LOCATION FOR LOCATION
Figure 10-17 Temperature Sensors
10-46
___ SEDR30O
PROJECT GEMINI
is remotely located from the sensing element. In most cases, however, the bridge
and sensing element are housed in the same case.
Regardless of how the bridge and sensing element are housed, combined, they
comprise a schematic as shown in Figure 10-17.
SYNCHR0 REPEA_RS
Three synchro repeater assemblies, mounted in the upper portion of the aft land-
ing gear will as shown in Figure 10-18, monitor the synchros on the IGS platform
gimbals. Each synchro repeater output is adc signal proportional to the space-
craft roll, yaw and pitch attitude in terms of platform coordinates. Two outputs
are available per repeater; a course output, which provides 0-5 vdc output for
0-350 degrees of synchro travel and a fine output, which gives 0-5 output forf_
every 35 degrees of synchro travel; only the coarse output is monitored as shown
in Figure 10-18. A dead band of i0 degrees maximum exists, centered around the
iSS-degree position in the synchro repeater potentiometers. Control of the synchro
repeaters is achieved by pilot actuation of the P_RM mode select switch.
co2 PARTZALPEESSUEE_ECTOR
A carbon dioxide partial pressure detector, as shown in Figure 10-19, is utilized
to insure that there is a safe level of C02 in the pilots suit circuits. The
detector is located in the EC8 module. The gaseous mixture to be sampled is
obtained as it exits from the ECS carbon dioxide and odor absorbers. The
sample stream is divided through two separate passages, both filtering water
vapor, but only one filtering carbon dioxide. The streams then pass into
10-_7
SEDR 300
v
F m u ,i , N u m I m m M e_j
AMPLIFIER [
- Box II [
-- -- -- MECHANICAL CONNECTION _ _ J COARSE
Figure 10-18 Synchro Repeaters and Schematic Diagram
10-48
j:_. SEDR300
ml_l_, wain i
F ..... /_i RADIOACTIVE jh/_TI_RIA L GAS OUTPUT
I I _)ABSORBING r_
I fr i _,oo_r_9:_I _ J III III _ CONVERTER - TELEMETRY
G k_LGk_ OUTPUT TO
[NPLIT i MEG OHM w PANEL METER
j -. RES,STOR . ,_L_p?E_ETER- IJ ,,
B N -E_]-H20VAPOR &
i_RTR'0oEI--! _r--],oNc.AMB_BI '-t, II ORFICE/ GAS OUTPUT
_-RADIOACTIVE
MATERIAL /
Figure 10-19 C02 Partial Pressure Detector and Schematic Diagram
10-49
SEDR 300
identical ion chambers which are polarized with + 50 vdc obtained from a dc-dcw
converter contained in the detector assembly; there a radioactive source
ionizes the gases. The difference of the electrical outputs is amplified and
provides a voltage which is proportional to the partial pressure of the mixture.
The gas is then returned to the inlets of the suit compressors. The system
provides two outputs: O-5 vdc into a nominal 2.5 megohm load for telemetry
use and 0-I00 microamps into a 4000 ohm cabin indicator.
ACCELEROME _RS
Three linear accelerometers are provided to measure the accelerations along each
of the spacecraft axes. The units are appro_mately 1.2 x 1.2 x B inches and are
pictured in Figure 10-20. The accelerometers are electrically-damped, force-
balance, servo-type units with outputs of 0-5 vdc. The unit which is used for
longitudinal measurements has a range of -B to +19 g and the other two have
ranges of + B g. The accelerometer is a torque-balanced, closed.loop systemm
with a pendulous mass supported by an extremely low friction Jewel bearing.
The schematic of the accelerometer is shown in Figure 10-20. An electromagnetic
position detector notes the slightest movement of the mass and supplies a
directly proportional electrical signal to a servo amplifier. The output of
the servo amplifier is applied to a torque generator which tends to restore the
mass to its equilibrium position. The output of the accelerometer is obtained
by sensing the voltage _rop across the resistor in the system loop.
INSTNUMENTATION PACKAGES
A number of the sign_Is in the various spacecraft systems are not compatible
Io-5o
"_ . PROJECT GEMINI
Figure 10-20 Servo Accelerometer and Schematic Diagram
10-51
.._,JJ--_,,, SEDR 300 _"_
, 4_L,,_.:-\-\_
='__ _ PROJECT GEMINI
SIGNAL CONDITIONER --SIGNAL CONDITIONERCARDS SANK B CARDS BANK C vC'-.
PACKAGE NO:. I
PACKAGE NO. 2
Figure 10-21 Instrumentation Package Assemblies
10-52
SEDR300 __,PROJECT GEMINI .' . ..
with the instruRentation circuitry, and therefore, must be conditioned for their
use. Two signal-conditioning packages (instrtanentation assemblies) are pro-
vided for this purpose. Instrumentation assembly number i is appro_tely
8 x i0 x 3 inches and is located in the adapter section. InstruBentation assembly
number 2 is approximately i0 x 10 x 8 inches and is located in the upper right hand
eqnil=ent bay of the re-entry section. Both units utilize sealed containers
with an operating pressure of _.5 paid and are shown in Figure 10-21. The
assemblies employ a modular construction with p_ug-in modules that may be replaced,
individually. A module consists of one or two standard printed circuit boards
with the necessary component parts and a connector for attachment to a mother
board within the package. There are 18 modules spaces in assembly number i and
51 in assembly number 2. Some modules provide for one data channel and others for
two. There are six basic types of modules, and several of these have additional
variations for different signal handling capabilities.
There are six variations of the Phase Sensitive Demodulators (PSD). Basically,
the PSD accepts two input voltages: one signal voltage and one reference. It
provides adc output of five volts for a full scale input signal that is in
phase with the reference and an output of zero volts for a full scale signal that
is out of phase with the reference. The various configurations of this unit
provide different full-scale sensitivities including special calibration curves
for rate gyros.
The twelve types of dc voltage monitors are designed to accept various positive
and negative dc voltage inputs and provide outputs of 0 to 5 vdc.
10-53
PROJECT GEMINI
The ac voltage monitor accepts a signal ranging from 23 to 29 volts r=s over
a frequency range of 380 to 420 cycles. The output is from 0-5 vdc, varying
only with the input voltage.
There are nine types of attenuator modules._ These modules have various dc
inputs which are changed to signals in the 0-20 my dc range or the 0-5 vdc
range. S_ne attenuator modules contain two data channels.
The dc milllvolt monitor accepts an input of 0 to 50 mv dc and provides a pro-
portional output of 0 to 20 my dc.
The ac frequency sensor provides a 0 to 5 vdc output proportional to an input
frequency varying fro= 380 cps to 420 cps. The voltage level of the input is
26 volts rms and does not affect the output of the module. _,
MULTIPLEXE_ENCODER SYSTEM
The multiplexer/encoder is divided into five packages to allow the data signals
to be sampled near their sources. Its purpose is to accept signals fro= the
signal pickups or signal conditioners, combine and convert these signals to a
serial binary-coded digital signal, and supply this signal to the tape recorder
and to the real-time telemetry transmitter. The units are shown in their respective
locations in Figures 10-22 and i0-23.
The multiplexer/encoder consists of a programmer, two identical high-level multi-
plexers, and two identical low-level multiplexers. These five units are dis-
cussed as a single unit because of their close function relationship.
The low-level multiplexers each weigh 2.6 pounds and measure approximately 5.75 x
5.25 x 2.75 inches. The high-level multiplexers are 5.25 x 4.75 x 2.5 inches and
weigh approximately 3.25 pounds each. The multiplexers might be considered as an
i0- 54
____ SEDR 300 __
PROJECT GEMINI
LOW - LEVELMULTIPLEXER
"ULTIPLEXER
Figure 10-22 Instrumentation System Multiplexers
10-55
..---:-_-_. SEDR 300
-_ PROJECT GEMINI
MOTHER BOARDS
(13)_
rIMEINDICATOR
Figure 10-23 Instrumentation System Programmer
10-56
___ SEDR 300 ____
PROJECT GEMINI
expansion to the data handling capsbility of the programmer and are dependent on
the programmer for encoding snd timing. The sampling rates of the multiplexer
inputs are established by the timing chain from the programmer. The low-level
multiplexers each ssmple 32 low-level (0-20 millivolts dc) signals: 24 at 0.416
samples per second and 8 at 1.25 samples per second. In the high-level multi-
plexer, 32 high-level (0-5 vdc) are samples at 1.25 samples per second, and the
24 bi-level and 16 bi-level-pulsed signals are sampled in sets of eight at a
sample rste of ten per second. For the bi-level signals, a binary one (nominally
28 vdc, but at least 15 vdc) may indicate that an event or function has or has
not taken place. For example, the indication that the bio-med tape recorders
are on is a one but the indication that the computer is on is a binary zero
..... (nominally zero vdc, but less than 5 vdc). For bi-level-pulsed signals, 15 vdc
or more represents a binary zero, while 5 vdc or less for at least ten milliseconds
is a binary one. The pulse conditioning circuitry in the multiplexer senses these
pulses and holds the voltage level until it is sampled by the progrA_er.
The programmer weighs approximately 20 pounds and measures approximately ll x ll x
_.5 inches. The programmer may operate as a seLf-contained data handling unit.
The basic functions of the progr_er are: data multiplexing; timing, to support
the multiplexing functions; and analog to digital conversion. The progr_er
output is a PCM pulsetrain to the tape recorder and the real-time telemetry trans-
mitter.
The bssic components of the programmer are a high-level analog subcommutator,
.... prime subcommutator, master commutator, analog to digital converter, output shift
register, dlgital i_ut circuitry and input selector, special timing, output fil-
ter, and t_De recorder input converter (Figure 10-24).
lo-57
. .. .
- _-_ SEDR 300
INPUTSLOW LEVEL SIGNALS
(6 AT 640 SAMPLESPER SECOND)
(6 AT 160 SAMPLESPER SECOND)
(9 AT 60 SAMPLES NON-RETURNPER SECOND) TO ZERO
DIGITAL SIGNALS OUTPUT TO(24 AT 0.416 SAMPLES TELEMETRYPER SECOND) TRANSMITTER
_,..__.._._..._.._....__
i
r'- ...... _'--'OGRAMM_R_ _ "--I
(32 AT 1.25 SAMPLES
PERSECO.D)I J i ......... !
INPUTS JHIGH LEVEL SIGNALS |
(3 AT 40 SAMPLES JPER SECOND) I
(3 AT 20 SAMPLES ANALOG TO OUTPUTPER SECOND) PRIME SUB- MASTER
(6 AT 10 SAMPLES COMMUTATOR COMMUTATOR CONVERTERDIGITAL REGISTERSHIPT lPER SECOND) I
BI-LEVE L SIGNALS(40 AT 10 SAMPLES
PER SECOND) J J
I I -TAPE jJ SPECIAL INPUT
TIMING CONVERTER/REPRODUCER
I DIGITAL INPUT J
INPUTS -- I
DIGITALBI-LEVELOUTPUTS
L_....... __I
4 SEC. RESET- RETURNTO ZEROOUTPUTTO TAPE
RECORDER
LOW-LEVEL LOW-LEVEL HIGH-LEVEL HIGH-LEVELMULTIPLEXER MULTIPLEXER MULTIPLEXER MULTIPLEXER
TINPUTS(EACH) INPUTS(EACH)
LOW LEVEL SIGNALS HIGH LEVEL SIGNALS(24 AT 0.416 SAMPLES (32 AT 1.25 SAMPLESPER SECOND) PER SECOND)
(8 AT 1.25 SAMPLES BI-LEVEL SIGNALSPER SECOND) (24 AT 10 SAMPLES
PER SECOND)8I-LEVE L PU LSE SIGNALS P'-_"
(16 AT I0 SAMPLESPER SECOND)
Figure 10-24 Multiplexer/Encoder System Block Diagram
10-58
oo0PROJECT GEMINIf_
The programmer high-level analog subcommutator segment has the capability of
sampling B2 high-level signal inputs at 1.25 samples per second. The analog
subcommutator receives its inputs directly from the signal sources, or from
the signal conditioners. The output of the analog subcommutator is applied to
the prime subcoramutator for further multiplexing.
The prime subcormmutator, in addition to accepting the sampled high-level sub-
commutator data output, has the capability of sampling40 bi-level signals at
lO samples per second, 3 high-level signals at 40 samples per second, B high-
level signals at 20 samples per second and 6 hlgh-level signals at lO sampler per
second. The prime subcommutator supplies its output to the master commutator.
_-_ The inputs to the prime subcom_utator, low-level outputs from the two low-level
multiplexers, and high-level outputs from the two high-level multiplexers are
combined into a prime subframe which is applied to the master commutator. In
addition, the master commutator has the capability of sampling 6 low-level
signals at 640 samples at 80 samples per second, and 24 digital signals at O.416
samples per second.
The output of the master commutator is applied to the analog to digital converter
where the analog output from the master commutator is converted to a digital
presentation.
The digital data from the output of the analog to digital converter, digital in-
put circuitry, and the two hlgh-level multiplexers (bi-level signals) Is combined
in the output shift register into a continuous Non-Return-to-Zero (NRZ) Pulse Code
lO- 59
sEo3ooPROJECT GEMINI
Modulated (PCM) wavetrain of 51.2 kilobits per second.
The output of the output shift register is applied to the real-time telemetry
transmitter, and to the tape recorder input converter/conditioner.
The tape recorder input converter/conditioner selects data from the 51.2 kilobits
per second NRZ output of the output shift register, converts the data to an Return-
to-Zero (RZ) output and applies the data to the tape recorder input conditioner
module at 5.12 kilobits per second.
The serial outputs, provided from the programmer, all have positive voltages
for ones and zero or negative voltages for zeros. The output for the tape recorder
is a 5.12 kilobit per second serial Return-to-Zero (RZ) signal with a +5 volt
transition for data ones and a -5 volt transition for data zeros. A clock signal _-_
at 5.12 kilohits per second is also provided for the tape recorder. This output
is a pulse train of 50 percent duty cycle at a peak amplitude of 5 volts. The
timing of the positive excursion is coincident with data one pulses. The pro-
grammer output for the real-time transmitter is a 51.2 kllobit NRZ signal with
a voltage which is adjustable between 0.I volt and 1.0 volt peak. Separate
hardline outputs are provided to allow various test equipment to be used without
degradation of the transmitter or tspe recorder outputs. The hardline
outputs are real-time PCM signal, basic PCM clock rate signal, and master reset
pulse signal. The signals are two volts peak-to-peak and are fed over twinex
coaxial or video cables.
The programmer message format includes a master frame and prime subframe; the
complete format is transmitted as real-time and only the prime subframe is
io-6o
SEDR 300
tape recorded. The master frame consists of 160 woNs, each word consisting of
eight data bits, sampled 40 times per second. Ninety-six master frames are
required to c_letely cycle throug_ all data inputs. Every tenth word in the
master frame contains prime subframe data. The prime subframe consists of 64
words sampled ten times per second. _enty-four prime subframes are required
to cycle through all data inputs of this part of the system. Information bits
are obtained from analog data, arranged with the most significant bit first,
digital data, broken into groups of eight bits with the most significant bit
first, or bi-level data grouped as eight consecutive data bits (referred to
as a bi-level set).
TRANSMI_RS
Three telemetry transmitters are used to transmit the Instrumentation System
/- data to the ground stations. Although the transmitters serve the Instrumenta-
tion System, its antennas, and associated switching is part of the Communica-
tion System; therefore, the transmitters are described in detail in Section IX,
Cc_._.,AnicationSystem.
PCM TAPE RECORI_R
The tape recorder is designed for monitoring and for producing a recording of
the signals received from the PCM programmer. The tape recorder records PCM
data at a tape speed of i 7/8 inches per second and playback, on command, of
this recorded data, will, on command, stop, reverse tape direction and playback
the recorded data at a tape speed of 41.25 inches per second. Erasure of data
will occur only during record mode. The power control circuitry is described in
detail in Section IX, Co_unication System.
IK-=_- SEDR 300
Figure 10-25 PCM Tape Recorder
10-62
___ SEDR 300
PRO,.I EC"T GEMINI
Telemetered signals recorded are RZ. The PCM tape recorder reproducer shown in
Figure 10-24 consists of one completely enclosed tape recorder which is approxi-
mately 4.3 inches high, i0.0 inches wide and i0.0 inches deep. The tape recorder
consists of the cover assembly, capstan drive assembly and tape transport assembly.
Connectors on the side of the case assembly provide signal connections, power
connections and test connections.
Record
The magnetic tape recorder is capable of providing a minimum of four hours of
recording time at a tape speed of 1 7/8 inches per second. Two tracks of
simultaneous PCM data can be recorded at 1 7/8 inches per second. Four hours
of RZ data at 5120 bits per second can be recorded at 1 7/8 inches per second.
Playback
On command, the recorder is to rewind the tape into the supply reel at 22 times
the record speed (41.25 inches per second) while reading and playing back the
informalion recorded on the tape. Final output of recorded data is in NRZ form.
Diphase System
The diphase signal processing technique permits the maximum tape utilization
efficiency, while avoiding certain serious problems encountered with use of
conventional NRZ recording at high packing density. It involves the encoding
of the digital information prior to recording and decoding of the playback and
conversion of the reproduced signal into standard NRZ form.
10-63
_. SEDR300 .__.__
PROJ EC'T GEMINI
The diphase technique is essentially a phase-modulated carrier process. The
digital data format to be recorded in RZ with an accompanying clock, and the
desired output in the reproduce mode is of the standard NRZ form.
The diphase signal to be recorded is created in the following manner. Inverted
RZ data and clock signal are OR gated into a binary flip-flop such that a transi-
tion of the flip-flop occurs on every negative going edge. A logical zero is
represented in the diphase code by a square wave at 1/2 the data rate. Each
time a logical one is received, a phase transition occurs in the center of the
bit cell so that a logic one is represented by a square wave at the data rate.
The output of the flip-flop is the diphase signal. This signal is then fed to
the record amplifier which drives the diphase signal into the record head.
f_
Record Mode
During the record mode, the input signal is sent to a preamplifier, encoder and
amplifier. A clock signal is applied to the input of the triggerable flip-flop.
The diphase code produced is recorded on magnetic tape.
The magnetic tape is dc erased prior to recording. The magnetic head utilized
is a high-quality instrumentation recording head has a gap width approximately
i/3 the recorded wave length. The gap width is not critical, but if it is much
wider than I/3 of the recorded wave length, the high frequency playback components
are attenuated and if it is much narrower, the high frequency components are
accentuated, causing a difficult equalization problem.
lO-6#
___ SEDR 300 _.____
PROJECT GEMINIf
Reproduce Mode
During the reproduce mode, the signal is picked up by the magnetic head and
applied to the playback amplifier, where it is amplified approximately 60 db,
filtered, and equalized to compensate for the effects of the head to tape system.
The equalized signal is then fed to an input coupler where approximately 40 db
of hard limiting is provided, thus providing extremely high immunity from ampli-
tude variation in the reproduced signal.
The ability of this system to operate satisfactorily through such a large varia-
tion in playback signal amplitude assures a high degree of reliability and
extremely low data drop out. The outputs of the input coupler, the recorded
diphase signal and its complement, are fed to the one shot timing extractorf_
circuitry and simultaneously to the decoder circuitry.
The function of the timing extractor and decoder is to produce timing pulses
from the amplified and limited dlphase playback signal. This circuitry detects
the data, using the timing pulses and dlphase sign_!, and produces the final
NRZ output.
The output filter is fed the decoder output and filters out some of the higher
harmonics of the NRZ output signal. The hardline output amplifiers produce a
hardline output with good square wave characteristics at high frequencies.
DC-DC CONVERTERS
The two dc-dc converters (one of which is a standby unit) supply the Instru-
zo-65
j-- _. _ SEDR 300
Figure 10-26 DC-DC Converter & Regulators
10-66
__. SEDR300
PROJ E(T GEMINI
mentation System _-ith regulated dc power. The units are approximately 5.5 x 5.5
x 7 inches weigh approximstely seven pounds each, and are located in the right-
hand equipment bay of the re-entry section as shown in Figure 10-25. The con-
verters are essentially voltage regulators which operate on 18 to 30.5 vdc and
supply output voltages of +5 vdc, +24 vdc and -24 vdc.
The power control circuitry for the dc-dc converters is shown in Figure 10-3.
Essentially, input power to the dc-dc converters is supplied through the on
position of the DC-DC CONV circuit breaker on the overhead swltch/circuit breaker
panel. This arms the DC-DC CONV switch. Placing the DC-DC CONV switch on the
overhead swltch/circuit breaker panel, to the SEC or PRI position, will apply
power to the corresponding converter. Usage of the dc-dc converter regulsted
_- output voltages is illustrated in Figure lO-B.
BIO-MED TAPE RECORDERS AND POWER SUPPLY
The two tape recorders used in the physiological Instrumentation System are
identical. Each one is approximately 9 x 6 I/2 x 1 3/4 inches (excluding con-
nector and mounting projections) and weighs about three pounds. One external
connector provides termination points for all inputs and outputs. The circuitry
is made up of 19 printed circuit boards with solid-state components. The recorder
uses recording tape with a width of 0.497 + 0.001 inches. The reel capacity is
880 feet. All physiological functions_ except oral temperature and blood pressure
of each pilot are recorded on the tape recorders. Each recorder has six data
channels and one timing channel. The timing input is a pulse-coded pulsetrain
derived from the Time Reference System through the Time Correlation Buffer
f_
10-67
PROJECT GEMINI
This signal is used for time correlation during post mission analysis.
The recorders willoperate for a total of i00 hours at a normal tape speed of
0.029S inches per second. Recorder operation is controlled by the crew during
the mission without playing back the data. Upon completion of the mission, the
recorders are removed from the spacecraft so thai the tape can be removed and
the data extracted. The total power requirement of each recorder is 1.2watts
at 24 vdc.
The electrical control circuitry for the bio-med instrumentation is shown in
Figure 10-2 and the location of the components is shown in Figure i0-i. The
recorders are Government Furnished Equipment and are actuated from the space-
craft main bus through the BIO-MED INST circuit breaker and the CONTposition
of the BIO-MED RCDR switch (I and 2). _
The bio-med power supply, similar in construction to the dc-dc converters,
supplies dc regulated voltage to the bio-med instrumentation. Input power for
the converter is obtained from the main bus through the BIO-MED INST circuit
breaker.
io-68
PYROTECHNICS andRETRO ROCKET
SYSTEM
,SectionXI
TABLE OF CONTENTS
TITLE PAGE
GENERAL INFORMATION ............................. 11-3RECURRENT COMPONENTS .......................... 11-3 _-_a_._iiiii_SEPARATION ASSEMBLIES AND _4:"..._-__.'=_._
DEVICES ....................................................... 11- 8 Ii!ii._.T:_.y:.:_i:_:_N:--EGRESS SYSTEMS AND DEVICES .................. 11-34 !iiiii_]iiiiiiijijiiiiilPA RA CHUTE LA N DIN G SY STEM ii!ii_iiiii!iiiiff:_i._iil..°..**°..°°°°°°°.°.°°.....
:::::::::::::::::::::::::::
PYROTE CHN ICS ........................................... 11-56 _!iiiiiiiiiiiiiiiiiii!ilH'_.°°°, • ,,,,,,°,,,°,, ,°_°, ,4.
PYROTECHNIC VALVES ................................. 11-67 ::ii_i_i_!_ili_iii!i_i!i,°°°.°..°°°°° ...... °°°°°°°,
RETROGRADEROCKETSYSTEM...................11-7oiiiiiiiiiiiiiiHiiiiiiiiiiiDOCKING SYSTEM PYROTECHNIC :::::::::::::::::::::::::::::::::::::::::::::::::::::::
:::::::::::::::::::::::::::............ • .., ...........
DEVICES ...................................................... 11-77 iiiiiiiiiiiiiiiiiiiiiiiiiiil
............... • ..o.....,..
..°...,.°.o, ............ ,.,
........ ,, ........ ,,°°..,,,......... ,....,....,,° .....
............... ,....,...,..
..., ............................. ° ....................
:::::::::::::::::::::::::::
........... ° ...............
.°°° ........ , ............ °
............... °°°°.°.°°.................... ° ........
................ ° ..........
.... ° .......... °...°°..°°°
::::::::::::::::::::::::::................... ,....,°
11-]- ::::::::::::::::::::::::::
l_L_r_,_ SEDR3O0-, PROJECT GEMINI
SINGLE BRIDGEWlRE DETONATOR(ELECTRICAL RECEPTACLE)
ELECTKICAL
INSULATION -- FIRINGClRCUI1PIN
TIME DELAY COLUMN
(WHEN APPLICABLE) _xCASE
DUAL BRIDGEWIRE DETONATOR(ELECTRICAL RECEPTACLE)
Figure 11-1 Detonator (Typical)
11-2
SEDR 300
PROJEC'T GEMINI
SECTION XI PI'RO,TEC_ICS AND,RETROGRADEROC_-TS
Dr o A, ,IO
The pyrotechnic devices and retrograde rockets, installed in the Gemini Spacecraft,
provide the escape system propulsion modes, enable and disable systems, and sep-
arate various sections and assemblies. Pyrotechnics are installed in each of the
major sections and in numerous locations throughout the spacecraft. The retrograde
rockets retard the spacecrafts orbital velocity to initiate re-entry into the earths
atmosphere. The retrograde rockets are located in the retrograde section of the
adapter.
RECURRENT COMPONENTS
"_ Some pyrotechnic items are used extensively throughout the spacecraft. To avoid
repetition in subsequent paragraphs, their description and operation will be
presented at this time. When describing the various systems, these components
shall be mentioned by name only.
DETONATOR
Description
The typical detonator (Figure ii-i) is a machined steel or aluminum cylinder
containing an ignition mix, booster charge and an output charge. In some instances
a pyrotechnic time delay column is used to provide a time delay between ignition
and detonation. The case is threaded at one end for installation purposes. An
electrical receptacle is provided at the other end. Electrically, the detonators
are provided in two different configurations. One incorporates two independent,
_" identical Siring circuits. The other incorporates only one firing circuit. The
Ii "3
SEDR 300
circuits of both detonators are electrically insulated f_om and independent of the
detonator body. Each firip4_ circuit consists of two electrical connector pins,
across which a bridge wire is incorporated. The detonator is used to initiate
high explosive components.
oeration
Upon receipt of a 28 vdc electrical signal, the firing circuit or circuits will
cause the detonator to fire. Either circuit (detonators with dual circuits) will
initiate the charge with the same performance characteristics as exist when both
circuits are operative. The bridge wire ignites the ignition mix which in turn
ignites the booster charge. The booster charge then propagates detonation to
the output charge. If a delay column is installed, the ignition mix will ignite
the delay column which ignites the booster charge. The output charge detonates
and transmits the detonation wave to the assembly to which it is attached.
CARTRIDGE
Description
The typical cartridge (Figure 11-2) is a machined steel sylinder containing an
ignition mix and an output charge. In some instances a pyrotechnic time delay
column is used to provide a specific time delay between ignition and output.
The cartridge is threaded at one end for installation purposes. An electrical
receptacle is provided at the opposite end. Electrically, the cartridges are
provided in two different configurations. One incorporates two independent,
identical firing circuits. The other incorporates only firing circuit. The
circuits of both cartridges are electrically insulated from and independent of
the cartridge body. Each firing circuit consists of two electrical connector _,
pins mounted in a high strength ceramic dielectric base and with a bridge wire
11-4
-,_'_c:_% SEOR3oo._ PROJECT GEMINI
FIRING
J ,_gg'_gN
SINGLE BRIDGEWlRECARTRIDGE(ELECTRICALRECEPTACLE) FIRING
CIRCUITPINS
\f_
FIRING
ELECTRICAL
INSERT
Ill lJlTIME DELAy COLUMN c. o 3E WIRE
(WHEN APPLICABLE) O O o °o O
OOo°o • 0 o O Oo
o_o o ° o o oa
ELECTRICAL
INSULATION
DUAL BRIDGEWIRECARTRIDGE
(ELECTRICAL RECEPTACLE)CLOSURE
Figure 11-2 Cartridge (Gas Pressure)
11-5
, SEDR300 ____
PROJECT GEMINI
connected between the two pins. The cartridge output is a hot gas pressure.
0_eration
When initiated by a 28 vdc electrical sign_!, the firing circuits will cause
the cartridge to function. Either circuit (cartridges with dual circuits) will
fire the charge with the same performance characteristics as exist when both
circuits are operative. The bridge wire ignites the ignition mix which propa-
gates burning to the delay co],,mnjif applicable, and to the output charge. The
output charge produces gas pressure that is used to operate the specific device
in which the cartridge is installed.
FLEXIN T.vw RSHAPED
Desc tion
Flexible Linear Shaped Charge (FLSC) is a V-shaped, flexible lead sheathing con-
tainlng a high explosive core. FLSC is used in separation assemblies to sever
various types, thicknesses, and shapes of ,_terials. The specific type, shape
and thickness of the material to be separatea, dictates the amount of explosive
contained in the FLSC. In the Gemini Spacecraft and Agena Adapter, the FLSC is
provided in four different core leadings: 7, 10, 20, and 25 grains per foot.
O_eration
When installed, the open portion of the V-shaped FLSC is placed towards the item
to be severed. The FI_C is detonated by a booster charge that has been initiated
by a detonator. The explosive core of the FI_C detonates, resulting in collapse
of the sheathing in the V groove, which produces a cutting jet composed of
explosive products and minute metal particles. This Jet produces extremely high _
localized pressures resulting in stress far above the yield strength of the target
L-6
SEDR300
PROd E-C'T GEMINI
material.
MILD DETONATING FUSE
Description
Mild Detonating Fuse (MDF) is a strand of high explosive encased in a lead sheath-
ing with s circular cross section. MDF is used as a separation device and as an
explosive interconnect. As a separation device, the strand contains 5 grains of
explosive per foot. As an explosive interconnect, the strand contains 2 or 3.3
grains of explosive per foot. The interconnect type MDF is installed in either
flexible woven steel mesh or nylon hose and rigid stainless steel tubing. Both
rigid and flexible MDF have a small booster charge incorporated at each end. The
booster charges are referred to as acceptor and donor. The acceptor being on the
end that receives a detonation wave from an initiator. The donor being on the end
that transmits a detonation wave to a component or other acceptor. The inter-
connects are attached to various devices by AN type or Bendix type electrical
COnnectors.
Operation
The MUF used as a separation device is placed in a groove milled in a magnesium
ring. The ring is formed to the shape of the items to be separated and is placed
between the mating surfaces. The assembly to be Jettisoned is attached to the
main structure by frangible bolts. The bolts have been axially drilled to reduce
tensile strength to a specified breaking point. When detonated, the MDF exerts
a force against the mating surfaces greater than the tensile strength of the
frangible bolts. The MDF, used as an explosive interconnect, is initiated when a
_ detonator or booster charge propagates a detonation wave to the MDF booster. The
if-?
SEDR300 ,
PROJECT GEMINI
booster strengthens the wave and transmits it linearly through the length of
the _F strand. The booster, at the opposite end, propagates the detonation wave
to the device to which it is attached.
E ARA IOASSEMBLIES DEVICES
There are several different types of separation assemblies and devices used in
the Gemini Spacecraft (Figure ii-3). These assemblies and devices are presented
indlvidually in the followlng paragraphs.
SPACECRAgVf/LAUNCHVEHICLE SEPARATION ASSEMBLY
Description
The spacecraft/launch vehicle separation assembly (Figure ii-_) separates the
spacecraft from the launch vehicle by severing the mating ring. The separation _
assembly primarily consists of two flexible linear shaped charges (FLSC) installed
around the periphery of the mating ring, three detonators, three detonator blocks,
three dual boosters, a molded backup retainer and a back blast shield. The dual
boosters are inserted in the detonator blocks. The dual booster protrude into the
molded backup retainer, indexed directly above the FLSC, when the detonator blocks
are installed. The detonators are inserted in the detonator blocks with the out-
put charge adjacent to the dual boosters. The back blast shield attaches the
molded backup retainer and FLSC to the mating ring.
O eratlon
Upon receipt of the 28 vdc electrical signal, the detonators transmit a detonation
wave that is propagated to the dual boosters. The dual boosters strengthen the
detonation wave to achieve proper detonation of the FLSC The FLSC detonates and
severs the mating r_ng redundantly. The backup retainer absorbs the shock in the
11-8
- _ SEDR300 __]
PROJECT GEMINI
RENDEZVOUS AND RECOVERY --DOCKING BAR ASSEMBLY
SECTION SEPARATION (S/C 6-8 & UP)
WIRE BIGUILLOTINE
MAIN PARACHUTEREEFING CUITERS
MAIN PARAC HU
MORTAR CARTRIDGE
CARTRIDGE
PARACHUTE REEFING CUTTERS
LiNE GUILLOTINE
MORTAR
WIRE BUNDLE
DROGUE PARACHUIEREEFING CUTTERS _ _
DROGUE MORTAR EMERGENCY DOCKINGRELEASE SYSTEM
NOSE FAIRING (S,/C6-B& UP)
DOCKING DOOR
RELEASE CABLE BCUTTERS (3 TYPICAL) A(S/C 6-8 & UP) -DROGUE PARACHUTE
BRIDLE RELEASE
(3 TYPICAL)
Figure 11-3 Spacecraft Pyrotechnic Devices (R & R Section) (Sheet 1 of 3)
11-9
--_. SEDR 300
PROJECT GEMINI
HORIZON SCANNERHEAD EJECTOR
ACTUATOR __
(2 BEQ)
BACKBOARD MDE _,
ROCKET CATAPULT
(2
-.%DROGUE MORTAi_ ".
_"_ ROTECHNIC
: 2::::} SWITCHES _._ _.
HORIZON SCANNER FAIRINGEJECTOR
i.._
h
AIR
DOOR ACTUATOR
PACKAGE CA SYSTEM
._,._"COMPONENT
PACKAGE DB SYST EM
":_ D,SCONNECT
i_ PACKAGE AB SYSTEM
J
-PYROTECHNICSWITCHES
MDF INTERCONNEC1 -SINGLE POINTPYROTECHNIC DISCONNECTWITCH
MDF CROSS
MDF MANUAL PACKAGE C DETONATORFIRING INITIATOR B SYSTEM(2 REQ) COMP ONENT
PACKAGE A (A)A
-COMPONENT PACKAGE DDETC A SYSTEM
Figure 11-3 Spacecraft Pyrotechnic Devices (Landing Module) (Sheet 2 of 3)
11-10
SEO.3ooPROJECT GEMINI
ASSY.
R.E.P. COVER GUILLOTINE VALVEZ13.44
D-10 HAND HOLO GUILLOTINE (S/C 8 ONLY) )k VALVED-IO ATTITUDE
(S/C I0 & 12) ISOLATION VALVEUHF-VHF POLARIZATION
MEASUREMENT EXPE
D-14 (S/C 8 & 9)
TUBE CUT TER/SE ALER_
R.E.P EJECTOR (S/C 5 ONLY)-
FOOTREST GUILLOTINE
-Z]3 DETONATOR (S/C 8 & 9)
._, > ASSY
i/_._ CUTTER ASSY
L-:J$
GUILLOTINE
D-4,
(3 REQ) (S/C 5, 7 & 9)
UHF, "/HE POLARIZATION MEASUREMENT UAD IIIGUILLOTINE (S,/C 8 ONLY)D-4, D-7 EQUIPMENT RELEASE
"_ GUILLOTINE (3 REQ) (S/C ,-/D-I0 HAND HOLD GU JJ(S/C 8 ONLY)
D-10 MINIMUM REACTION ROCKET #! QUAD II
POWER TOOL _._GUILLOTINE_
DETONATOR(TYP3 "
_ ROCKET #4
H2 •ACTUATOR
ROCK ET13
ROCKET MOTOR
TYPICAL 4 PLACES D-7GUILLOTINES INITIATOR ASSYS/C-L/V (2 REQD/MOTOR DOOR RELEASE
GUILLOTINE (REF)RADIONATAR
ADAPTER EQUIP GUILLOTINES (S/C S&7)
EI.EASE GUILLOTINE (REF)
-NUCLEAR EMULSIONGUILLOTINE (S/C 8 & 9)• : FAIRING JETTISON
D-4, D-7 DOOR RELEASE ( GU ILLOTINE (S,/C SON LY)GUILLOTINE (REF) ELECTRO STATIC CHARGE EXP.
FBY INTERCONNECT
SPECTROMETER/INTERFEROMETER GU(2 REQ)
D-,4& D-7 EQUIP RELEASED-15 LOW LIGHT LEVEL T.V. GUILLOTINE (S/C 8 AND]I )
(TYP 3 PLACES) DOOR RELEASEGUILLOTINE (S/C S, & 9)
Figure 11-3 Spacecraft Pyrotechnic Devices (Adapter) (Sheet 3 of 3)
11-11
_._..S_ SEDR 300
r_='%_: ,_ PROJECT GEMINI
BACK BLAST
BACK-UP I
SHAPED CHARGE(10 GRAINS PER FOOT
HOUSINGFIBERGLAS
bNATOR
DUAL BOOSTER
LAUNCHVEHICL
SECTION A-A
Figure 11-4 Spacecraft/Launch Vehicle Separation Assembly
11-12
SEDR300
PRI
back blast. The back blast shield protects the structure and equipment from
shrapnel. Proper detonation of only one strand of FLSC is sufficient to sever
the mating ring.
EQUIPMENT SECTION/RETROGRADE SECTION SEPARATION ASSEMBLY
Description
The equipment section/retrograde section separation assembly (Figure 11-5) sep-
arates the equipment section of the adapter from the retrograde section of the
adapter. The assembly basically consists of two main units: the shaped charge
assembly and the tubing cutter assembly. The shaped charge assembly primarily
consists of two flexible linear shaped charges (FLSC), three detonator blocks,
containing three crossovers and six boosters, three detonators, ten segmented back-f_
up strips and a molded backup retainer. The detonator blocks provide for installa-
tion of the detonators. One detonator block provides for the installation of the
tubing cutter explosive interconnect. The tubing cutter assembly primarily con-
sists of an explosive interconnect (MDF), two formed aluminum parallel housings,
molded backup retainer, two flexible linear shaped charges with boosters attached,
a detonator block and a detonator. The explosive interconnect (MDF) is a flexible
nylon hose containing a strand of high explosive and end mounted booster charges.
The interconnect has Bendix type connectors incorporated at each end for attaching
the interconnect to the cutter and shaped charge detonator blocks. The inter-
conuect is attached to the cutter detonator block with its booster charge adjacent
to one of the boosters on the FLSC. The detonator is installed in the cutter
detonator block with its output end adjacent to the other booster on the FLSC.
The cutter assembly is bracket mounted to the inside of the retrograde sectiont_,
of the adapter, forward of the parting line. The shaped charge assembly is
11-13
. __ SEDR 300
L '_ PROJECT GEMINI
SEE DETAIL
"A"\--_ A ;RADESECTION (REP)RX_,
A
By /LX
J INTERCONNECT MDF---_ SECTION (REF)
DETONATOR
C
TUBE CUTTER ASS
MDP SHAPED CHARGE ASSEMBLYINTERCONNECT
SEVERED (REF)
MOUNTING BOLTS--_
DETONATOR
V_EW A-A
(20 GRAINS PER FOOT)
SHAPED CHARGE(10 GRAINS PERPOOT)-'_ SECTION C-C
EXPEOS,VE __
,NTERCONNECT
DETAIL A
Figure 11-5 Separation Assembly-Equipment Section/Retrograde Section
11-14
._. SEDR300
PROJ EC'T" GEMINIif-- _
installed around the outer periphery of the adapter at the equipment section and
retrograde section parting llne.
Operation
When initiated by a 28 vdc electrical signal, the detonators Mill fire. The deto-
nators of the shaped charge assembly transmit a shock or detonation wave to the
crossovers which in turn initiates the boosters. The boosters propsgate the wave
to the FLSC. The FLSC detonates and functions to sever the adapter at the parting
llne redundantly. The detonator of the tubing cutter assembly propagates detona-
tion to the booster on one strand of FLSC in the cutter assembly. The explosive
interconnect transmits detonation from the shaped charge assembly to the booster
on the other strand of FLSC in the cutter assembly. The two boosters propagate
the shock wave to the FLSC. The two strands of FI_C in the cutter assembly
detonate and sever the twelve aluminum tubes and one nylon tube. Proper detona-
tion of only one strand of FLSC_ in both the shaped charge assembly and tubing
cutter assembly, is sufficient to achieve separation.
RETROGRADE SECTION/RE-ENTRY MODULE SEPARATION ASSEMBLY
Description
The retrograde sectlon/re-entry module separation assembly (Figure ll-6) functions
to separate the retrograde section of the adapter from the re-entry module. Sepa-
ration is accomplished by severing the three titanium straps and various tubes
and wire bundles. The separation assembly primarily consists of three cutter
assemblies, three detonator housings, three detonators, three parallel booster
col,-,ns, six explosive interconnects, and three unions. The detonator housings
comtain a booster column and a parallel booster column. The cutter assemblies
11-15
--__ SEDR 300
'_ PROJECT GEMINI
SHAPED CHARGE
(25 GRAINS PER FOOT) _1_I_,
-_2-_--_I_L_ _DETONATOR--"°--U
HOUSING
SECTION A-A
(REF)
RE-ENTRY MODULE (RE_
C
_.__ (TYPICAL 2 pLACES)
SECTION B-BSTRAP
HOUSING
ADAPTER_,_ SHAPED
TIE FAIRING _- CHARGE
IDETONATOR
EXPLOSIVE BOOSTER_
,_ERcc V//////////////J_CUTTER ASSEMBLY.
(3 REQ)
SECTION C-C
Figure 11-6 Retrograde Section/Re-Entry Module Separation Assembly
11-16
_@ SEDR300
PROJECT GEMINI
consist of two parallel machined aluminum bars that contain four strips of FLSC.
The bars are Joined by the detonator housings with the parallel boosters. A
detonator is installed in each of the three detonator housings. The cutter assem-
blies are located in three places around the parting llne and are linked by the
explosive interconnects.
Operation
When initiated by a 28 vdc electrical signal, the detonators propagate a detona-
tion or shock wave to the boosters which relay propagation to cutter FLSC and
simultaneously the shock wave is propagated to the explosive interconnects. The
interconnects transmit the wave to all three cutter assemblies. This is to ensure
detonation of all three cutters FLSC, in the event one or even two detonators do
not function. Detonation of the cutter FLSC completely severs the titanium
straps, wire bundles and tubing redundantly. Proper detonation of only two
opposing strips of FLSC in each cutter is sufficient to achieve separation.
RENDEZVOUS AND RECOVERY SECTION SEPARATION ASSEMBLY
Description
The rendezvous and recovery section separation assembly (Figure ll-7) separates
the Rendezvous and Recovery (R & R) section from the Re-entry Control System (RCS)
section. The assembly primarily consists of Mild Detonating Fuse (MDF), MDF
housing ring, two detonators, two detonator housings and two booster charges. Two
strands of MDF are installed in parallel grooves milled in the housing ring face.
The grooves intersect at the booster charges which are installed approximately 180 °
apart. The R & R section is attached to the RCS section by frangible bolts, with
_ the MDF ring fastened to the R & R section at the mating surface. The detonator
ii-17
_I_ SEDR300 ___i__ ,:,,_oJ_,::-rG_,,,,,
FRANGIBLE BOLT (24 REQ)
RENDEZVOUS AND RECOVERY SECTION {RBF)
=;(Tu._.7R,No_--RE-ENTRY CONTROL SYSTEM SECTION
RENDEZVOUS AND RECOVERY SECTION (RBF)
SECT'ON A-A MDF HOUSING RING--_ __ __ _,MDFBOOSTER CHARGE _ FRANGIBLE BOLT
_i _ _--_WASHER
RE-_.TR¥CO._OL_YSTEM
SECTION (REF)-_ .f" _"
BOC _ B"" "_ 'Ji I SCREW
Y
'_- SCREW
SECTION B-B
Figure 11-7 Rendezvous and Recovery Section Separation Assembly
11-18
[_______@ SEDR300
PROJECT GEMINI
housings are installed in the RCS section, with the detonators indexed directly
above the booster charges, when the sections are mated.
0peration
When initiated by a 28 vdc electrical signal, the detonators propagate a detonation
wave to the two booster charges. The booster charges strengthen the detonation
wave and transmit it to the dual strands of MDF. The MDF detonates, exerting a
force against the RCS and R & R section mating surfaces. The force breaks all
the frangible bolts and allows the pilot chute to pull the R & R section free of
the spacecraft. Satisfactory propagation of either strand of MDF will successfully
separate the R & R section.
WIRE BUNDLE GUILLOTINE
_ Description
The wire bundle guillotine (Figure i1-8) is used throughout the spacecraft to
sever various sized bundles of electrical wires. The guillotines are used in two
sizes. One size can sever a wire bundle up to one and one quarter inches in dia-
meter and the other can sever a wire bundle up to two and one half inches in dia-
meter. Both sizes are similar in design, appearance and operation. The guillo-
tines primarily consist of a body, end cap or anvil, plston/cutter blade, shear
pln(s) and an electrically fired gas pressure cartridge. The body houses the
plston/cutter blade, provides for installation of the cartridge, and attachment
of the anvil. The anvil is removable to facilitate removal and installation of
either the guillotine or wire bundle. Two guillotines are used on a wire bundle,
one on each side of the separation plane. Lugs, for attaching the guillotine
to the spacecraft structure, are an integral part of the guillotine body.
l1-19
f- _-_-_ SEDR 300
__ .RoJ_._..,,.
WIRE BUNDLE _, _/
PISTON/CUTTER BLADE --MOUNTING LUGS::: _¢_
N _,xx,._asvJL "--_:ANVIL"
Figure 11-8 Wire Bundle Guillotine
11-20
___ SEDR300 __
PROJECT GEMINI
Operation
When initiated by the proper electrical signal, the cratridge produces gas pressure.
This gas pressure, exerts force on the piston cutter blade. When sufficient force
is applied, the piston/cutter blade will sever the shear pin(s). As the pin(s)
shear, the piston/cutter blade strokes, and completely severs the wire bundle. The
wire bundle is then free to pull out of the guillotine body.
WIRE BUNDLE GUILLOTINE (CABLE CUTTING)
Description
The wire bundle guillotine (cable cutting) (Figure 11-9) is used to sever _oven
stainless steel cables. The guillotine primarily consists of the body, piston/
cutter blade, shear pin, anvil and end cap, and two electrically fired gas pressure
cartridges. The body provides a piston actuation area and provides for cartridge
installation. The anvil is retained in the barrel section of the body by the end
cap. The anvil and end cap is removable to permit guillotine and cable installa-
tion and removal. Lugs, for attaching the guillotine to the spacecraft structure,
are an integral part of the body. The shear pin is provided to retain the piston/
cutter blade in a retracted position.
0_oeration
When Inltlatedbya 28 vde electrical signal, the two cartridges fire and produce
gas pressure. The gas pressure exerts force on the plston/cutter blade. When
sufficient force is applied, the plston/cutter blade severs the shear pin. The
plston/cutter blade travels the length of the barrel section and severs the cable
installed in the guillotine. The cable is then free to pull out of the guillotine.
11-21
/ _ SEDR 300
I I _,_ PROJECT GEMINI
CABLE (REF)
ANv'L 'sT° cuTTE HEARB DEP,N"I/CUTTE R BLADE
.=¢.-.----- BODY
1 [I II
L CARTRIDGE
Figure 11-9 Wire Bundle Guillotine (Cable Cutting)
11-22
______ SEDR300 _j__
PROJECT GEMINI
TUBING C%_R/_A_
Description
The tubing cutter/sealer (Figure ll-10) is used to cut and seal two stainless
steel, Teflon lined tubes. The tubes contain hypergolic propellants used in the
Orbit Attitude and Maneuvering System (OAMS). Two tubing cutter/sealer assemblies
are located in the adapter, one on each side of the retrograde/equlpment section
separation llne. The tubing cutter/sealer assembly primarily consists of the body,
anvil, one electrically fired gas pressure cartridge, four shear pins and cutter
assembly. The cutter assembly consists of the piston, crimper and blade. The
crimper and blade are attached to the piston by two of the shear pins, (sequencing
pins). The piston is secured in the body by the other two shear pins, (initial
_- lock pins). The body provides for the installation of the cartridge, attachment
of the anvil, and housing for the cutter assembly. Lugs, for attaching the tubing
cutter/sealer to the spacecraft structure, are an integral part of the body.
Operation
When initlatedby a 28 vdc electrical signal, the cartridge fires and generates
gas pressure. The gas pressure exerts a force on the piston of the cutter assem-
bly. When sufficient force is applied to the piston, the initial lock pins are
severed and the cutter assembly strokes to seal and cut the two tubes. The blade
and crimper, extending past the end of the piston, contact the tubing first. As
the cutter assembly moves down, the crimper flattens the tubing against the
raised portion of the anvil. As the cutter assembly continues its travel, the
sequencing pins are severed between the crimper and blade, stopping the travel of
the crimper. The base of the piston and blade further crimp and seal the tubing
with the blade severing the tubing. The sealed portion of the tubing remains in
11-23
_ _ SEDR 300
__ _3_)j PROJECT GEMINI
TUBES
A
CARTRIDGE'_
S.,ELD i_?i!ii:'!i!ilii _'"
; PIN _ _/_-PROPE LkANT
iNITiAL TUBELOCK PiN
g ;
VIEW A-A (BEFORE FIRING)
SEVEREDTUBE"
FREE--_
ANVIL/'_'_ VIEW A-A (AFTER FIRING)
Figure 11-10 Tubing Cutter/Sealer
11-24
__ SEDR300 _ _.__
PROJECT GEMINI
the tubing cutter/sealers at adapter separation. The severed portion of the tubing
between the tubing cutter/sealers is free to pull out at adapter separation.
PYROTECHNIC SWITCH
Description
The pyrotechnic switch (Figure ll-ll) functions to positively open electrical
circuits and prevent current flow in various wire bundles prior to their being
severed. The switches are located in various places throughout the re-entry
module. The switches primarily consist of the body, actuator (piston), shear
pin, spring lock, and electrically fired gas pressure cartridge. The shear pin
secures the actuator in the switch closed position prior to switch actuation.
Incorporated in opposite ends of the switch body are two electrical receptacles.
_'-- The end mounted receptacles contain hollow spring leaf contacts. The contacts are
axially connected by pins mounted in the actuator. All switches are identical in
design and operation with the exception of the number of contacts in the recep-
tacles. One model contains 41 contacts, and the other model contains 55 contacts.
Lugs, for attaching the switch to the spacecraft structure, are an integral part
of the body.
Operation
When initiated by a 28 vdc electrical signal, the cartridge fires and generates
gas pressure that is ported through the switch body to the actuator. The pressure
exerts a force against a flange of the actuator. The force causes the actuator to
sever the shear pin and move axially in the body. As the actuator moves, the
connecting pins mounted in the actuator are disengaged from the hollow contacts
,_ at one end and are driven further into the hollow contacts at the other end. The
spring lock _ops into place behind the actuator and prevents it from returning to
L_-25
SEDR 300
__,__"_," PROJECT GEMINI +
/ BODY "--_
SWITCH
LOCKED_
_ ELECTRICALRECEPTACLE
SHEAR PIN (ROTAIED
90° FOR CLARITY) CONNECTINGPINS
SWITCH CLOSED SWITCH OPEN
Figure 11-11 Pyrotechnic Switch
11-26
._ SEDR300
its original position. The actuator is thus held _ the electrically open pos_ion.
HORIZON SCANNER FAEEG RE_A_ASSEMBLY
Description
The horizon scanner fairing release assemb_ (Fibre ll-12) secures the horizon
sca_er falri_ to the spacecraft, and when initiated, Jettisons th_ fa_i_. The
assemb_ pri_rily consists of the actuator housing, actuator, actuator e_ension,
_in piston, release piston, eight locking pins and Go electrically fired _s
pressure ca_ridges. The actuator e_enslon forms a positive tie between the
actuator and the scanner fair_g. The actuator is lo_ed to the ma_ piston by
four lock_g pins. The main piston is locked _ the _se of the actuator h_s_g
by four locking p_s, that are held in p_ce by the release piston. The release
..... piston is sprig energized in the locked _sitlon. The actuator h_slng provides
for installation of the cartridges and mounting for the assemb_.
Operation
When init_ted by a 28vdc electrical signal, the caPriCes fire and produce gas
press_e. The pressure is po_ed throu_ a mi_ed passage in the actuator h_s_g,
to the base of the piston. The gas pressure moves the release piston forwa_,
which enables the four locking pins to c_ _board, releas_g the main piston.
The _s pressure causes the main piston, w_h at_ched actuator, to move through
the len_h of the actuator h_s_g. As the piston reaches the end of the housing,
a sh_lder stops the piston travel. The f_r lock_g pins, searing the actuator
e_ension to the piston, cam outboard into a recess and release the actuator ex-
tension. The act_tor e_ension bei_ thus freed is Jettisoned w_h the scanner
..... fairi_ attached.
ii -27
.._,_--__ SEDR300
m
I
CARIRI ACTUATOR) LOCK PIN
(4 TYP)
©1 EXTENSION
,'T---I m J
• • • • p
BEFOREFIRING(ACTUATORLOCKED)
ACTUATOR
AFTERFIRING(ACTUATORRELEASED)
Figure 11-12 Horizon Scanner Fairing Release Assembly
11-28
___ SEOR 300 __
PROJECT GEMINI
HORIZON SCANNER RELEASE ASSEMBLY
Description
The horizon scanner release assembly (Figure ll-13) secures the horizon scanners
to the spacecraft and Jettisons the scanners when initiated. The horizon scanner
release assembly primarily consists of the actuator housing, actuator, locking
mechanism, cartridge housing, and two electrically fired gas pressure cartridges.
The actuator is secured in the actuator housing by the locking mechanism. The
locking mechanism consists of a tang lock, tang lock retainer and a shear pin.
The tang lock is secured to and is located in the base of the actuator housing.
The actuator housing is attached to and becomes a part of the spacecraft structure.
The scanner base support and mounting platform are attached to the actuator prior
to installing the cartridge housing on the actuator. The two cartridges are in-
stalled in the cartridge housing.
Operation
When initiated by a 28 vdc electrical signal, the two cartridges fire and produce
gas pressure which is ported through the hollow actuator to the base of the actua-
tor housing. Slots in the tang lock allow the gas pressure to flow to the base of
the tang lock retainer. The gas pressure exerts a force against the base of the
retainer. The retainer moves axially in the actuator housing, severing the
shear pin and exposing the tines of the tang lock. The tines cam open, releasing
the actuator and allowing the gas pressure to Jettison the actuator and horizon
scanners.
ii -29
.il_..l_._._\,,l_ -. SEDR 300 ____,, PROJECT GEMINI ,
CARTRIDGE , /
(2 REO) _ / HOUSING
J_; ,7.m_7 "
_.: I _ """ ..M ......... i ..,
i "7"_'-' _"'._*P"_7............ ' LOCKNUT
.OR_ZONSCANNERi I _:_ i i.................................._................._i _!:[_;->ii
ASSEMBLY(_EF)_ ! S i {_. i ......::..:..::!_.i.:-,,,_, j ,, _ ......._ i f'_':7_._"....
i \:
i ........... ' ........................ ".'.L'.'.'.L ........ / _"'_;_-.--._-,<SPACECRAFT 1 s ....... L ......... "'.'.................... _-._,.-;_ ( 7 L.
:.... I ..................... z:::=.7........... ".............. ./.: ........... L.. i ..v'Lz.-':............... 'x'zti.. ..........................
ii'" _"......................i'__!<(}i! i(
.............../' _ i....22.Z::::::2:::222........................IANG LOCKRETAINER
SHEAR PI
DR OCK
HORIZON SCANNER (LOCKED) -COTIER PIN
7i
•, ,........................ ..] , fT....._j, _N_LOC_r"r---.-,-.-d_x _ j ..............-_--i . L _" '? ! F....... • __ i _7' i:i_::i:i::
i t i !
i :\ /
! i
] HORIZON SCANNER; ASSEMBLY (REF)
, i.... _............. j
LOCK•_., RETAINf:R (RRF)
7_ i f
............J ! _............................ _[..._._.2..TT.L_.-7ACTUATOR
HORIZON SCANNER (RELEASED)TANG
Figure 11-13 Horizon Scanner Release Assembly
11-30
___ SEDR300 _-_
PROJECT GEMINIf---.
FHESH AIR DOOR AC_JATOR
Description
The fresh air door actuator (Figure 11-14) is provided to retain the fresh air
door to the spacecraft and to eject the door when initiated. The fresh air door
actuator is located forward of the egress hatches, to the left of the spacecraft
centerline and below the outer mold llne. The actuator prln_arily consists of the
breech, plunger, screw and two electrically fired gas pressure cartridges. The
plunger forms a positive tie between the fresh air door and the breech. The
plunger is retained in the breech by the screw which acts as a shear pin. The
breech provides for installation of the two cartridges. Lugs, for attaching the
actuator to the spacecraft structure, are an integral part of the breech.
Operation
When initiated by a 28 v_c electrical signal, the cartridges are caused to fire
and generate gas pressure that exerts a force on the plunger. When sufficient
force is applied, the plunger severs the screw and is ejected out of the breech.
The plunger and fresh air door are then Jettisoned free of the spacecraft.
NOSE FAIRING EJECTOR
Description
The nose fairing ejector (Figure 11-15) is used to secure the rendezvous and
recovery nose fairing to the spacecraft until initiated by a 28 vdc signal.
When initiated the pyrotechnic ejector will positively Jettison the nose fairing.
The nose fairing ejector assembly consists of a breech, ballistic hose, actuator
assembly, crank assembly, and an electrically fired gas pressure cartridge. The
f_nose fairing is attached to the crank assembly. An actuator shaft forms a
ii-31
_.-;_._ SEDR 300
1
SCREW
CARTRIDG E
PLUNGER
EJECTED_ __
SCREWSCRfW (SHEAR PIN)
S[VERED
ACTUATOR BEFOREFIRING ACTUATOR AFTERFIRING
Figure 11-14 Fresh Air Door Actuator
11-32
_....,_f_.--:_. SEDR 300
ACTUAI"OR ASSEMBLY
J
.__ BALLISTIC HOSE CRANK MECHANISM
GAS IMPULSE /ACTUATOR _ S
INLET BODY END CAP CRANK MECHANISM
r_ /--LOCK,NGP_N
F_'--'_"_l"........................."_................."'"_'_,.',_ / _=---:--...... I I
NOSE FAIRING INSTALLED
SHAFT
SHEAR PIN --_ LCRANK MECHANISM
NOSE FAIRING EJECTED
Figure 1]-15 Nose Fairing Ejector Assembly
11-33
___ SEDR300 _-_
PROJECT GEMINI
positive tie between the actuator body and the crank assembly. The actuator
shaft is locked to a piston in the actuator by two locking pins and held in
place by a shear pin in the end cap of the actuator. The actuator assembly is
connected to a breech by a ballistic hose. The breech provides for installation
of the cartridges and is positioned approximately nine inches from the actuator.
The actuator is installed on the antenna support and fairing actuator fitting
of the R and R section and is located on the X axis, five inches from Y zero.
Operation
When initiated by a 28 vdc signal, the cartridge generates gas pressure which
is transferred through a ballistic hose to the actuator housing and exerts a
force on the actuator piston. The gas pressure causes the piston, with attached
shaft, to move, severing the shear pin and continuing through the length of
the actuator housing. As the piston reaches the end of the housing, the two
locking pins, securing the shaft to the piston, cam outboard into a recess
and release the actuator shaft: The actuator shaft, now free, is Jettisoned,
with the nose fairing attached, by the crank mechanism. The crank mechanism
provides an angled Jettisoning of the fairing from the axial movement of the
ejector shaft, without recontact with the spacecraft. A hinge on the nose
fairing, located on the outer mold line, releases and directs the path of the
fairing away from the spacecraft.
EGHESS SYSTEMS AND_EVICES
The egress systems and devices (Figure 11-16) provide the pilots with a rgpid
and positive method of escaping the spacecraft, should an emergency arise. The
system is manually initiated and is used below an altitude of 15,000 feet only.
Each system and device is presented in the sequence of their operation
ii-3_
__ SEDR 3OO
SEAT EJECTORROCKET
• RELEASESYSTEM
,,/ \,+HATCH • , +/
ACTUATOR _ j_, j_"
, !__ • ;'_
II
.,\ +X% •, ..
-.?.-._'.. . _j +:+;\-__-
m
• \
\ /-Z+++p j-_+.. _zz
DROGUE MORTARBACKBOARD JETTISONASSEMBLY
, /i
ACTUATOR ASSEMBLY
EJECTLON
CONTROL +HATCH ACTUATOR t' _+INITIATION SYSTEM(D-RING) MDF MANUAL _"
FIRING MECHANISM
Figure 11-16 Egress System and Devices
11-35
sEo3o0PROJECT GEMINI
HATCH AC_'JATOR INITIATION SYSTEM (MDF)
Description
The hatch actuator initiation system (Figure ii-12) is used to initiate the
firing mechanisms of both hatch actuators. The system is manually activated by
either pilot. The system primarily consists of 8 MDF interconnects, two MDF
crossovers and two manual firing mechanisms. The interconnects consists of four
rigid and four flexible MDF assemblies that connect the firing mechanisms to the
hatch actuators. The two crossovers are rigid MDF assemblies that cross connect
the two initiation system firing mechanisms. The firing mechanisms each contain
dual firing pins, dual percussion primers, and a booster charge. The firing
mechanism is drilled and tapped for installing two MDF interconnects and two cross-
overs. The MDF interconnects and crossovers are installed so that the small _-_
booster on the end of each MDF is adjacent to the booster charge of the firing
mechanism. The firing mechanism is attached to the spacecraft structure, located
below the pilots feet.
Operation
The hatch actuator initiation system is activated when either pilot pulls the
ejection control handle (D-ring) located between the pilots knees and connected
to the firing mechanism. Approximately one-half inch travel and approximately a
_0 pound pull of the lanyard connecting the ejection control (D-ring) to the
firing mechanism will cock and release the dual firing pins. The firing pins
strike the dual percussion primers, causing the booster charge to detonate. The
firing mechanism booster charge propagates detonation to the four MDF ends. The
interconnecting MDF propagates the detonation wave to the firing pins of the hatch _-_
actuator breech assembly. The crossover MDF propagates the detonation wave to
the other pilots firing mechanism. This insures initiation of both hatch actuators.
ii-36
._. SEDR 300 _-_1"_1 __
CTUATOR
\ (REF)./
/'° \
// \\ INTERCONNECT
/ ,'" \ MDF
" j/
f /"
HATCH ACTUATOR / /,"/
iNTERCONNECTMDF
HATCH ACTUATOR /BREECHi
, i i '_/ _ EJECTION CONTROL
\ HANDLE (O-RING)
FLEXIBLE J_'"...........INTERCONNECT "//_ --_
R_GiD \INIERCONNECT /
/ !MDF "\-- \\\ / i
\ , /\\\ /
EJECTIONCONTROLS\ \\i
HANDLR (D_RING) --/ \-_. Z2
/,"--- LANYARD CONNECTION /
TO EJECTION CONTROL /_
D-RiNG CRO OVER f/
iMDF :::
S :::MANUALFIRING
;:i
: :
LINTERCONNECT CROSSOVER i
(REF)(REF)
FIRING MECHANISM
/-_ BEFOREFIRING ::ii_(AFTER FIRING)!i!i!ii!
Figure ] 1-17 Hatch Actuator Initiation System
11-37
, SEDR300 --__3
PROJECT GEMINI
HATCH AC_JATOR ASSEMBLY
Description
The hatch actuator assembly (Figure 11-18) unlocks, opens and mechanically
restrains the hatch in the open position. The assembly also furnishes
sufficient pressure to initiate the firing mechanism of the seat ejector rocket/
catapult. The assembly primarily consists of the breech end cap, breech, cylinder,
stretcher assembly, end cap (base) and rod end assembly. The breech end cap
assembly contains the locking mechanism for mechanically restraining the hatch
in the open position; provides for installation of the seat ejector rocket/cata-
pult ballistic hose; provides for installation of the breech assembly, and is
thread mounted to the top of the cylinder. The breech contains two firing pins,
two percussion fired cartridges, and a gas producing propellant charge. Two _
interconnects, from the hatch actuator initiation system, are attached to the
breech adjacent to the firing pins. The stretcher assembly primarily consists of
the piston and stretch link, and is located inside the cylinder. One end of the
stretch link is attached to a web inside the piston. The other end is attached
to the rod end assembly. The rod end assembly connects the stretcher assembly to
the hatch. The end cap is attached to the lower end of the cylinder, and provides
for attaching the hatch actuator assembly to the spacecraft structure. The end
cap contains a latch piston that actuates the hatch unlock mechanism.
Operation
The hatch actuator functions when initiated by the initiation system MDF inter-
connects. The shock wave, propagated by the MDF interconnects, causes the two
zz-38
.--__--_ SEDR 300
ii::
_.... i II AssE_.y
i I EXTENDED(HATCH OPEN)
I
GAS PRESSURETOHATCH STRUCTURE _:_:_ SEAT EJECT
(REF) !:: ORZASSEMBLY ::::::
i:: END CAP
ACTUATOR LOCK ::: ACUTATOR
LOCKED
END CAP
" ASSEMBLY_
(EXPENDED) il"_
PISTON _
ACTUATOR ASSEMBLY ACTUAlfOR ASSEMBLY i,_'_ \ _,."BEFOREFIRING t AFTER FIRING
_>; i:: i /
SPACECRAFT
SPACECRAEI" i:" STRUCTURE(REF) _2,,_,
STRUCTURE (REF)_] TRIPPERASSEMBLY _ l_._) ) TRIPPERASSEMBLY
' ._ LOCKED POSITION --
i 7_: .... (UNLOCKED POSITION)--./ =.J
\
Figure 11-18 Hatch Actuator Assembly
11-39
.____ SEDR 300 ___
PROJECT GEMINI
firing pins of the breech assembly to sever shear pins and strike the primers of
the two percussion fired cartridges. The cartridges ignite and generate hot gas
which ignites the main propellant charge of the breech. The propellant charge
produces a large volume of high pressure gas. The gas pressure is exhausted into
the area between the piston of the stretcher assembly and the cylinder. Orifices
in the lower end of the piston wall admit the gas pressure to the base of the
stretcher assembly. The gas pressure is ported through a drilled passage to the
latch piston. The gas pressure extends the latch piston, which unlocks the hatch
through a bellcrank/pushrod mechanism. The gas pressure then acts on the base of
the stretcher assembly, moving it through the length of the cylinder. Tmmedlately
prior to the stretcher assembly reaching full extension, gas pressure is exhausted
through a port to the ballistic hose. The ballistic hose delivers the pressure
to the firing mechanism of the seat ejector-rocket/catapult. As the stretcher
assembly reaches the f_lly extended position, the lock pin of the locking mechanism
engages the piston of the stretcher assembly and holds the hatch open. The locking
mechanism is also operative when the hatch is fully opened by hand. A lanyard,
attached to the locking mechanism, permits the hatch to be unlocked, when manually
actuated.
SEAT EJECTOR-ROCKET/CATAPULT
Description
The seat ejector-rocket/catapult (Figure ll-19) is used to eject the man-seat mass
from the spacecraft. The seat ejector-rocket/catapult basically consists of the
catapult assembly and the rocket motor assembly. The catapult assembly primarily
consists of the catapult housing, firing mechanism, main charge (gas producer),
and locking assembly. The catapult housing contains all of the listed components
ll-40
f:---. SEDR 300
f--
LOCKING RING
...._SEAT AITACH
I. _EF LIN_)
_:_..::_:__ ._ _ .-
_SPACECRAFTATTACH, (REF)
•RELAY CHARGE
BALLISTICHOSE
Figur6 11-19 Seat Ejector - Rocket/Catapult
11-41
j___ SEDR300 -__-_
PROJECT GEMINI
in its base. The firing mechanism consists of dual firing pins, dual percussion
fired primers, and _ relay charge. The firing plns are secured in place by retain-
ing pins. The locking assembly consists of the lock rlng and a spring to hold the
ring in place. The base of the catapult assembly is attached to the spacecraft
structure. The rocket motor assembly primarily consists of the motor case, nozzle,
motor lock housing, lock ring, shear pins, upper and lower auxiliary igniters, and
the main propellant charge. The nozzle is threaded to the motor case and is secur-
ed by four set screws. The nozzle is secured to the motor lock housing by locking
tangs. The locking tangs are held in place by a lock rlng that is retained by four
shear pins. The motor lock housing is secured in the base of the catapult by tang
locks. The tangs are held in place by the lock rlng of the catapult. The maln
propellant charge is located in the motor case with an auxiliary igniter at each
end of the charge. The top end of the rocket motor assembly Is attached to the
upper aft portion of the seat.
O_eration
The seat ejection cycle is initiated when gas pressure is received via the ballis-
tic hose from the hatch actuator. Sufficient gas pressure will cause the dual
firing plns to shear their retaining plns and strike the dual percussion primers.
The primers ignite the relay and main charges. Hot gas pressure, produced by the
maln charge, releases the motor lock housing by displacing the lock ring against
the spring through piston action. Wlth the motor lock housing released, the gas
press1_re propels the rocket motor through the length of the catapult housing.
Prior to complete ejection from the catapult housing, the lock rlng of the motor
lock housing makes contact with a stop whlch severs its four shear pins. The tang
locks of the motor lock housing cam open and release the rocket motor. Separation
of the rocket motor from the motor lock housing allows the hot gas from the
11-42
.__ $EDR 300 __j
PROJECT GEMINI
catapult main charge to ba1!_stic hose to initiate the thruster assembly.
THRUSTER ASSEMBLY-SEAT/MAN SEPARATOR
Description
The thruster assembly - seat man separator (Figure 11-21) is the active portion
of the seat/man separation assembly. The thruster supplies a stroke of adequate
length and power to a webbed strap that accomplishes seat/man separation. The
thruster assembly primarily consists of the thruster body, thruster piston,
firing mechanism and percussion fired gas pressure cartridge. The cartridge and
firing mechanism is installed in the upper end of the thruster body. The firing
mechanism contains a firing pin, retained by a shear pin. The ba114stic hose
from the harness release actuator is attached to the firing mechanism. The
_-.... thruster piston is located in the thruster body and is retained in the retracted
position by a shear pin. The thruster body is mounted on the front of the seat
structure, between the pilots feet.
Operation
High pressure gas from the harness release actuator is transmitted through the
ba_llstic hose to the thruster firing mechanism. The gas pressure causes the
firing pin to sever its shear pin and strike the primer of the cartridge. The
cartridge is ignited and generates gas pressure. The gas pressure exerts force on
the thruster piston, causing the piston to sever its shear pin. As the piston
extends out of the thruster body, the strap is pulled taut effecting seat/man
separation.
B_TJ_TE DEPLOY AND EEL_ASE SYSTEM
Descriptionf_
The ballute deploy and release system (Figure I1-22) primarily consists of the
11-_3
.j_ . SEOR300 __
_ORTTOBALL,ST,CHOSE li!' il _ii
j_ L_ _EIRING MECHANISM
BALLISTIC HOSE TO THRUSTER (REF)
RELEASE ACIUATOR BEFORE FIRING
TIME DELay CC
_CARTRIDGE FIRED FIRING
LANYARD CONNECT
TO THRUSTER
FIRING MECHANISM AND CARTRIDGE
RELEASE ACTUATOR AFTER FIRING
Figure 11-20 Harness Release Actuator Assembly
11-44
sEoR3004;. PROJECT GEMINI
IHRUSTER ASSEMBLY THRUSTER EXTENDED
(BEFORE FIRING) (FIRED)
HOSE (REF) (REF)
PERCUSSION
THRUSTER
j• THRUSTER ASSEMBLY
SEAT/MAN SEPARATOR (TYP)
HARNESSRELEASEACTUATOR BALLISTIC
(BEE) HOSE (REF)
II (REF)
STRAP ASSEMBLY (REF)--
Figure 11-21 Thruster Assembly-Seat/Man Separator
11-45
j- -_. SEDR 300
PROJECT GEMINI
PIN
SEQUENCING
i::ii ANEROID-FIRING
CUTTER
,_ ASSEMBLY
• _ LAN_'ABDLOCK'_'_\ DEPLOY CABLE BALL (RELEASED)-
SYSTEM BEFORE ACTIVATION
TODEPLOY CUTTER
II SEQUENCING
LEGEND
S EPLOYCUTTER
WORKING GAS PRESSURE
(ACTIVATED)i':ii OEeLOV
C ABLE
BLOCKED GAS PRESSURE BALLUTE DEPLOY ACTIVATED
:,ii!! ABOVE 7500 FEETONLY
Figure 11-22 Ballute Deploy and Release System (Sheet 1 of 2)
11-46
.i-=_. SEDR 300
_-_i_ ] PROJECT GEMINI
_-'\ _BALLUTE RISER {SEVERED)
PISTON/CUTTER (ACTIVATED)
41'
(ACTIVATED)
pEPLOYGASBLOCKED
PISTON
I _ (ACTIVATED)
f--" _ CARTRIDGE FIRED
_iii _! HOSE
iiill PISTON/CUTTER
iliiii!ii_iiii_iiiii ASSEMBLY ACTIVATEDiiiii
iiHi BELOW 7500 FEETii_,iEiiili
iiiiiiiiiiiiiii
iiiliiiiii
iiiiiiiiii!iliiiiili
iiiiiiiiiiii_
" iiiii
!!i[iBALLUTE RELEASEACTIVATED
ABOVE 7S00"FEET _ii
Figure 11-22 Ballute Deploy and Release System (Sheet 2 of 2)
11-47
__ SEDR 300 __
PROJECT GEMINI
firing assembly, deploy cutter and hose, and release guillotine and hose. Con-
ta_ed within the firing assembly, is the release aneroid firing mechanism and
cartridge, the deploy firing mechanism and cartridge, and the sequencing housing
and piston. The basic function of the system is to deploy and release the ballute
between specified altitudes and prevent ballute deployment below specified altl-
tudes. The system is located on the upper left side of each pilots backboard.
The deploy firing mechanism and the release aneroid firing mechanism is linked
to the pilots seat by individual lanyards.
.Operation
The system is _n_tiated by the lanyard pull as seat/man separation is effected.
When initiated above 7500 feet, the release aneroid is armed and the deploy firing
mechanism Is activated. The firing pin of the deploy firing mechanism strikes the
primer of the cartridge and causes ignition. The cartridge generates gas pressure
after burning through the time delay column. The pressure is ported through the
deploy hose to the deploy cutter assembly. The cutter severs a nylon strap that
allc_s the ballute to deploy. The armed aneroid functions when an altitude pres-
sure level of 7500 feet is reached. The aneroid sear releases the cocked firing
pin of the ballute release firing mechanism. The firing pin strikes the primer,:
which ignites the cartridge and causes it to generate gas pressure. The pressure
is ported through the release hose to the release guillotines The guillotine
severs the ballute riser strap and allows the ballute to be carried away. When the _
system is activated by the lanyard pull below 7500 feet, both cartridges are im- ;
mediately initiated. The time delay incorporated in the deploy cartridge permits
the release cartridge to generate gas pressure first. The pressure is ported ;
through the release hose to the release guillotine, _hlch severs the ballute riser ."
ll-_8
sEoR3ooPROJECT GEMINI
strap. Simultaneously gas pressure is ported to the sequencing housing and
sequencing piston. The piston is actuated, causing it to block the gas exit
of the deploy cartridge. The gas pressure, generated by the deploy cartridge,
does not reach the deploy cutter, preventing deployment of the ballute.
DROGUE MORTAR-BACKBOARD JETTISON ASSEMBLY
The drogue mortar-backboard jettison assembly is provided to deploy the per-
sonnel drogue parachute and to separate the backboard and seat from the pilot.
Description
Drogue Mortar
The drogue mortar (Figure 11-23) functions to fire a weighted slug with sufficient
velocity to forcibly deploy the personnel parachute and to initiate the backboard
Jettison assembly firing mechanism. The drogue mortar primarily consists of the
mortar body, mortar barrel, drogue slug_ main cartridge (gas pressure), initiator
cartridge (detonator), aneroid assembly, main lanyard, manual lanyard, and the
main and manual firing mechanisms. The mortar barrel is threaded into the mortar
body and contains the drogue slug. The drogue slug is retained in the barrel by
a shear pin. The aneroid assembly is attached to the mortar body and contains
the main firing mechanism. The main lanyard is enclosed in a rigid housing to
prevent Inadvertant pulling of the lanyard. The housing is attached to the main
firing mechanism housing at one end and to a take-up reel at the other. The main
lanyard, a fixed length of cable, is attached to the main firing mechanism at one
end and to the take-up reel at the other. The take-up reel incorporates an ex-
tendable cable that is attached to the ejection seat. The main cartridge is
threaded into the mortar bodyt with the primer end, adjacent to the main firing
mechanlsm_ and the output end in the mortar body pressure cavity. The manual
ll -_9
k__ PROJECT GEMINI
MORTAR
lUTE CONTAINER
:HUTE
lING MECHANISM
_SLUG MAIN (
EJECTED
BACKBOARD FIRINOME(
//
ANEROID RELEASESARMED /
/'(REF) MANUAL LANYARD_
RELEASEDAND FIRED ARMED AND COCKED _ .//_ATTACHED TOSEAT
©/_FIRING PIN RELEASED
(ANEROID ACTION)
Figure 11-23 Drogue Mortar
11-50
s°300PROJECT GEMINI
lanyard is enclosed in a flexible conduit to prevent inadvertant pulling of the
lonyard. The lanyard Is attached to the manual firing mechanism at one end and
to a manual pull handle at the other. The manual firing mechanism Is threaded
Into the mortar body. The primer end of the detonator is threaded into the
manual firing mechanism, and its output end 90 degrees and adjacent to the maln
cartridge output area. The drogue mortar is attached to the upper right slde
of each pilots backboard.
Backboard Jettison Assembly
The backboard jettison assembly (Figure 11-24), functions to separate the back-
board and seat from the pilot, when initiated by the pressure from the drogue
mortar. The backboard Jettison assembly primarily consists of the MDF firlng
i_ .
mechanism, MDF time delay cartridge (detonator), interconnect (time delay MDF),
MDF manifold assembly, Jetelox release pin, interconnect (Jetelox pin MDF), lap
belt disconnect, interconnect (belt disconnect MDF), restraint strap cutter (FLSC),
and interconnect (strap cutter MDF). The MDF firing mechanism Is attached to the
drogue mortar body and contains a shear pln retained firing pin. The MDF time
delay cartridge is a percussion fired cartridge and Is installed in the MDF firing
mechanism. The interconnect (time delay MDF) is connected to the MDF firing
mechanism and the MDF manifold. The interconnect (jetelox pin MDF) is connected
to the MDF manifold and the Jetelox release pin. The interconnect (belt dis-
connect MDF) is connected to the MDF manifold and the lap belt disconnect. The
interconnect (strap cutter MDF) Is connected to the MDF manifold and the restraint
strap cutter (FLSC). The three component interconnects terminate in the MDF mani-
fold wlth their acceptor end adjacent to the interconnect (time delay MDF) donor
end. The Jetelox release pin retains the Jetelox Joint to the seat until
i1-51
,.-_ SEDR 300
MECHAN SM MECHANISM "_IHOUSING HOUSING
GAS PRESSUREFROM DROGUE
SHEARPIN MORTAR
NG PIN PIN (SEVERED)
CARTRIDGE CARTRIDGE
INTERCONNECT _ i_ RE DELAy MDF) CONNECT
SECTION B-B BEFORE FIRING SECTION B-B AFTER FIRING
BEFORE FIRING _ AFTER FIRING RESTRAINT
ST_PS--_ ¢FB" CUTTER
MDFFIRINGMECHANISM iliil " FLSC (7GRAINS _HOUSING
+7 i!il _+ 'VIEWA-A ' B .
(REF) RESTRAINT STRAP CUTTJ:R (FLSC)
MDF FIRING - ,'
MECHANISM_
cuT_FER (FLSC)
ERCONNECT
(STRAP CUTTER MDF)
(BELTDISCONNECT MDF)
/--- LAP BEL'r DISCONNEC_
_hc__" _%x,%xINTERCONNECT _lJC
gETELOX PIN MDI VIEWC+C ...._%_i::_::_....."-:'::_!!_...
_ LAP BELTADJUSTER--_ //=- PIN (2 REQ)
"- _ D SCONNECf "_ / CAM (2 REQ'
% +, \//- ,::_ INTERCONNECT
_ !ii_ (BELTD'SC_)I_E_ _"_ _o_ ' ' ' ' 'I I +_INTERCONNECT i::iii MDF)_. x_._'_)gll I I I I
BOO_ _,STON _1 /--FISTON(ACTUATED)iiiiBEFORE_:/":_--J[II II_1_ /---SHEAR PIN NN$/ /--SHEARPIN i!_i LAPRELTADJUSTER _ --
LOCK BALL _ / _F,,,=-_ (SEVERED) ii!i! (DISCONNECTED)_ ,,'--CAM (ROTATED)(4 REQ)_ /--JETELOX JOINT _,_,_x_._ _ _ /
!!PIN L / _
. _1_"_1 _ l_x_x_ (SEPAP-:ATED)-7 ::ii! U__ :+'--_LOC_BAL.ii++• AFTERF,R,NG+_,t__ _ +--.BEFORE FIRING AFTER FIRING RELEASED) _ ........ J_iii_ _A" IF///A }
_ETELOXRELEASEI'_N H+:i ,APBE_.TOISCONN,CT
Figure 11-24 Backboard Jettison Assembly (Sheet 1 of 2)
11-52
f_., SEDR 300
INTERCONNECT LOCK BALL
(STRAPCUTTER
(BELTDISCONNECT
TO EGRESS KIT
U_¢Co°x"Z_0_,-/lI ,.TE_CON.E_ _,.G_O,?RELEASEBEFORE FIRING (JETELOX PIN MDF) -J AFTER FIRING
OX RELEASE PIN
Figure 11-24 Backboard Jettison Assembly (Sheet 2 of 2)
11-53
___ SEDR 300 ___
PROJECT GEMINI
initiated. The Jetelox release pin primarily consists of the body, piston, four
lock balls, and a shear pin. The lap belt disconnect is provided to unfasten the
lap belt when properly initiated. The lap belt disconnect primarily consists of
the housing, two lock pins, two cams, piston and a shear pin. The restraint strap
cutter is provided to sever the pilots shoulder harness. The cutter primarily
consists of the housing, two strips of FLSC and a booster.
Operation
Drogue Mortar
The drogue mortar is initiated by the pull of the main lanyard, at seat/man sep-
aration. The extendable cable, attached to the seat, uncoils from the take-up
reel. Upon reaching the end of its travel, the cable pulls the take-up reel free
of the rigid housing. The fixed length main lanyard attached to the reel is
pulled, and if in excess of 5,700 feet, cocks the main firing mechanism and arms
the aneroid. At an altitude pressure level of 5,700 feet, the aneroid releases
the cocked main firing pin. The firing pin strikes the primer and ignites the
main cartridge, which produces gas pressure. The gas pressure causes the drogue
slug to sever its shear pin and travel out of the mortar barrel. Simultaneously,
the gas pressure initiates the backboard firing mechanism. When initiated by the
main lanyard below 5,700 feet, the main firing mechanism is cocked and immediately
released to fire the main cartridge. The aneroid is in the release position
because of the altitude pressure level, therefore is not armed and does not
delay the cartridge firing. The drogue mortar may be initiated manually by
pulling the manual lanyard handle at any altitude. The lanyard cocks and releases
the manual firing pin, which strikes the primer of the initiator cartridge
(detonator). The initiator cartridge detonates and ignites the output charge _-".
_i-54
sEo300PROJECT GEMINI
of the main cartridge, which produces the gas pressure for drogue slug ejection
and backboard firing mechanism initiation.
Backboard Jettison Assembly
The backboard jettison assembly is caused to function when the main cartridge of
the drogue mortar is fired. Gas pressure from the drogue mortar main cartridge,
causes the firing pin of the backboard firing mechanism, to sever its shear pin
and strike the primer of the time delay cartridge. After the proper time delay,
the cartridge propagates a detonation wave to the MDF interconnect, which trans-
mits the wave simultaneously, to the three MDF interconnects attached to the MDF
manifold assembly. Simultaneously, the detonation wave is propagated by the three
MDF interconnects to the restraint strap cutter (FLSC), lap belt disconnect, and
the Jetelox release pin. The detonation wave propagated by the interconnect
(jetelox pin MDF) acts upon the piston of the Jetelox pin, causing it to sever the
shear pin. As the piston moves, a recess in the piston is aligned with the lock
balls. The pressure exerted by the Jetelox Joint, forces the lock balls into the
piston recess, and releases the Jetelox joint and egress kit. The detonation wave,
propagated by the interconnect (belt disconnect MDF), is directed against the pis-
ton of the lap belt disconnect. The detonation wave moves the piston causing it
to sever the shear pin. As the piston moves, the cams rotate and retract the pins
from the lap belt adjuster. The lap belt separates and permits the pilot to be
partially free of the backboard. The detonation wave, propagated by the inter-
connect (strap cutter MDF), is transmitted to the booster of the restraint strap
cutter (FLSC). The booster strengthens and increases the reliability of the deto-
nation wave for proper detonation of the two strips of FLSC. The FLSC detonates
ii-55
r --IPROJECT GEMINI
and severs the two restraint straps allowing the pilot to be completely free of
the backboard. The seat may be released m_nually by the pilot actuating the
seat single point release b-halle. Effective spacecraft 5 and 6 a cable from the
single point release, pulls a ball retaining pin from the MDF manifold. The
pressure of the interconnect (jetelox pin MDF), moves the ball aside and pulls
out of the MDF manifold.
PARACHUTE LANDING SYS_ PYR(Y_NICS
The Parachute Landing System (Figure 11-25) is provided to safely recover and land
the re-entry module, after its entry into the earths atmosphera_. The pyrotechnic
portion of the system consists of the drogue, pilot, and main parachute reefing
cutters; the drogue and pilot parachute mortars; the drogue parachute bridle re-
lease guillotines; the pilot parachute apex line guillotine; and the main para-
chute disconnects. Each of these pyrotechnic devices are presented in the follow°
lng paragraphs.
DROGUE PARACHU_ MORTAR ASSEMBLY
The drogue parachute mortar assembly is provided to positively deploy the drogue
chute. The assembly is similiar to the pilot parachute mortar assembly (Figure
11-25) in design and operation.
DROG_E PARACHUTE RF_._ING CUTIERS
The drogue parachute reefing cutters are provided to disreef the drogue chute.
The cutters are simillar to the pilot parachute reefing cutters (Figure 11-26)
in design, operation, and number.
DROGUEPARACHUTEBRIDLERELEASEGUILLOTINES
The drogue parachute bridle release guillotines are provided to sever the bridle
ii-56
__;i._,,_,<_.,. SEC>R300 I_!_ --j.,. PROJECT GEMINI
BRIDLE DISCONNECT
\.-\ \
/._J ./ .... \. PARACHUTE"I j . \ \ BRIDLE (REF)/
// .. \
(. \ ,,
/ y _,,',.,-... "\
_;_i ::i ",i 7-. _. S_NGLEPOINT". /? ii 1/ ' DISCONNECT
_ _ /, f : h: l..ll / \
"_ "_'" ............... / \ PILOT PARACHUTE"_\ _ _"_ _'_ "" "'" "" / '_ MORTAR PRESSURE
"'" / // REEFING LINE CUTTERSPARACHUTE
//
/
/ •
X\ ] _ PARACHUTE" MORTAR
FWD BRIDLE DISCONNECT
_AIN PARACHUTEREEFING LINE CUTTERS --
DROGUEMORTARPRESSURE
r PARACHU I_EAPEX LI NE
DROGUE PARACHUTE GUILLOTINEREEFING LINE C
DROGUE PARACHUTEDROGUE BRIDLE RELEASE
PARACHUTE GUILLOTINE (3 REQ)MORTAR-
Figure 11-25 Parachute Landing System Pyrotechnics
11-57
___ SEDR 300 __
PROJECT GEMINI
at its three attach points. The release guillotines are similar in design and
operation to the cable and wire bundle guillotine (Figure I1-9).
PILOT PARACHUTE MORTAR ASSEMBLY
De script ion
The pilot parachute mortar assembly (Figure ].I-26)functions to deploy the pilot
parachute in the event of a malfunctlonlng drogue chute. The mortar assembly is
located in the forward end of the rendezvous and recovery section. The mortar
assembly primarily consists of the mortar tube, sabot, breech, orifice, frangible
bolt, washer and two electrically fired gas pressure cartridges. The base of the
mortar tube is attached to the breech and the breech is attached to the rendezvous
and recovery section structure. The flanged orifice passes through the base of
the mortar tube and breech. The orifice is secured beneath the breech by a
locknut. One end of the breech is drilled and tapped to provide for installation
of the two cartridges. The sabot is located in the lower section of the mortar
tube and is secured by a washer and a frangible bolt. The frangible bolt passes
through the washers, the center of the sabot mortar tube base, and is threaded
into the base of the orifice. The pilot chute is installed in the sabot.
Operation
The mortar functions when the cartridges are initiated by a 28 vdc electrical
signal. The cartridges generate gas pressure that is ported through the breech
and orifice to the base of the sabot. When sufficient pressure is exerted on the
sabot, the frangible bolt will part and release the sabot. The gas pressure
propels the sabot and pilot chute out of the mortar tube, thus effecting positive
chute deployment. _-_
11-58
SEDR 300
DROGUE PARACI lUTE
MORTAR
CARTRIDGE
RTAR TUB
BREECHASSEMBLY _
MORTAR TUBE
SABOT
CARTRIDGE_
FRANGIBLE BOLl
ORIFICE
SECTIONA-A
Figure 11-26 Pilot Parachute Mortar Assembly
11-59
__. SEDR300 __
PROJECT GEMINI
PILOT PARACHUTE REEFING CUTTERS
pescriptlon
The pilot parachute reefing cutters (Figure 11-27) are provided to disreef the
pilo t chute by severing the reefing line. The reefing cutters primarily consist
of the cutter body, cutter blade, firing mechanism and percussion fired time delay
cartridge. Two cutter assemblies are sewn to the inside of the parachute skirt
band 180 degrees apart. The reefing cutter is a tubular device, with all its
components contained within the cutter body. The firing mechanism is contained
in one end of the cutter body and consists of a firing pin, lock ball, spring, and
sear pin. The firing pin is retained in the cocked position by the lock ball. The
lock ball is held in place by the sear pin. A lanyard is attached to the sear
pin and to the parachute canopy. The spring is precocked and energizes the firing
pin when the sear pin is pulled. The cartridge is installed in the center portion
of the cutter body and is roll crimped in place. The cartridge consists of a
percussion primer, time delay column and output charge. The cutter blade is stake
locked in the cutter bod_, below and adjacent to the output end of the cartridge.
A washer is crimp locked in the end of the cutter body, and serves as the anvil
and stop for the cutter blade. A hole in each side of the cutter body, between
the cutter blade and washer, permits installation of the reefing cable.
Operation
Deployment of the pilot chute causes the reefing cutters to be initiated. As the
canopy extends, the lanyard is pulled taut and pulls the sear pin from the firing
mechanism. The lock ball moves inboard and unlocks the firing pin. The spring
energized firing pin is driven into the primer of the cartridge and ignites the
ll-60
f _-z>. SEDR300
/ i \ ,. i \'%.,,.i>.,,. __
i\ ;".i
\.i \ 7 ,__,," .{
,tt \ •REE.,NO _ I i iCUTTERS _
CUTTERSPARACHUTEREELED PARACHUTEDISREEFED
_,............---,EEF,NOL,NEI_EEI
_CUTTERBLADE--FIRING PIN
__............/ /-_.:_/_.oo°_/r _--_
_--ANVI L WASHER _ SAFETY PINPIN HOLE
REEFINGCUTTERBEFOREFIRING
7U£EE" O_J_UT .,R,.O.,N
COLUMN PRIMER
I'_-_ REEFINGCUTTERFIRED
Figure 11-27 Pilot Parachute Reefing Cutters
11-61
sEo30oPROJECT GEMINI
time delay column. After the specified t_me delay, the cartridge produces gas
pressure that exerts force on the cutter blade. When sufficient force is exerted_
the cutter blade shears the stake lock and strokes to sever the reefing cable.
Proper functioning of only one of the two cutters is sufficient to perform pilot
chute disreeflng.
PILOT PARACHUTE APEX LINE GUILLOTINE
Description
The pilot parachute apex llne guillotine (Figure 11-28) is provided to sever the
pilot chute apex llne, in the event of a drogue chute malfunction. The guillotine
primarily consists of the body, cutter blade, and two electrically fired gas pres-
sure cartridges. The cutter blade is retained by a shear pin. The body provides
for the installation of the two cartridges, and incorporates drilled passages from "--_
the cartridges to the cutter blade. Design of the guillotine allows the apex line
to pull free, when the pilot chute is deployed by the drogue chute. The guillotine
is located in the forward section of the rendezvous and recovery section.
Operat. !on
When initiated by a 28 vdc electrical signal, the cartridges produce gas pressure.
The pressure is ported through a drilled passage to the head of the cutter blade.
Sufficient pressure causes the cutter blade to sever the shear pin and stroke
to cut the apex llne. The apex line is thus free of the malfunctioned drogue
chute, permitting mortar deployment of the pilot chute.
MAIN PARACHUTE REEFING CU_'2ER
The main parachute reef£ug cutter (Figure 11-29) is provided to disreef the main
11-62
SEDR 300
_ MOUNIlNG LUGS
z
BEFOREFIRING AFTERFIRING
Figure 11-28 Pilot Parachute Apex Line Guillotine
11-63
__ PROJECT GEMINI
i i i: "+" "....//, ...... , ',, 'b
I_,\t ;i\JJ ,//!';'";'k;x,----_._'':_','_,
r, _1 _,. _t {i f t :: _1 A';2"
2)_ '__REEE NO
'C'5. \\ ...... i_"-''-'_7''+'._ // i?' CUTTERS
PARACHUTE REEFED PARACHUTE DISREEFED
REEFING LINE (REF)
_CUITER BLADE CUTTER BODYGAS FIRING FIN
IM10_ NT_-, IJ-,_.m-_flt_,,,,-,,,_._'" " _OUTPUT _TIME DELAy "--PERCUSSION \ _SEAR
_ANVIL --STAKE CHARGE COLUMN PRIMER _- SAFETY PINWASHER LOCK PIN HOLE
REEFING CUTTERBEFOREFIRING
,//_ CUTTER
BLADE FIRING PIN
...... /-o.I_E /- ,- Loc__ALL
_--TIME DELAy '_ PERCUSSIONCOLUMN PRIMER
REEFING CUTTER FIRED .-_,
Figure 11-29 Main Parachute Reefing Cutter
11-64
___ SEDR300 ___j_ _
PROJECT GEMINI
parachute. Three reefing cutters are located on the inside of the canopy skirt
band 120 degrees apart. The reefing cutters are slm_lar in design and identical in
operation to the pilot chute reefing cutters. Proper operation of only one of
the three cutters is sufficient to dlsreef the parachute.
MAIN PARACHUTE DISCONNECT
De script ion
The main parachute disconnects (Figure ii-30) include the s_ngle point disconnect
assembly and the forward and aft bridle disconnect assemblies. The disconnect
assemblies are identical, in design and function. The disconnect assemblies pri-
marily consist of the breech assembly, arm, and two electrically fired gas pres-
sure cartridges. The breech assembly consists of the adapter, piston, lead slug,
/-_ snubber disc, and plunger. The piston is retained in the adapter by a shear pin.
The lead slug is located on the end of the piston. The snubber disc is located
under the head of the plunger. The adapter is threaded into the spacecraft
structure with the piston extending into the arm. The breech is threaded onto
the adapter and the cartridges are installed in the breech. The single point
disconnect is mounted on the hub of the main parachute adapter assembly. The
adapter assembly is located on the forward ring of the Re-entry Control System
section. The forward bridle disconnect is mounted at the top of the forward ring
of the Re-entry Control System section. The aft bridle disconnect is located
fox-ward of the heat shield between the crew hatches.
Operation
When initiated by a 28 vdc electrical signal, the disconnect cartridges are ig-
F_ nited. The cartridges produce gas pressure that is ported through drilled passages
ii-65
..- _ _ SEDR300
•, PROJECT GEMINI
'i'_ ADAPTER DISC
SRACECRAF] BREECHSTRUCTURE (REF)
LEAD SLUC
SHEAR PIN"
SNUBBER
_ I l_///'//'/_/, _J CAETR,OOESstructure reE_ l\\\\\_//,_\\\x/ / _.'/._\'_×\\'-.\\\\\1 \
DIscONNECTASSEMBLYBEFOREFIRING
PIN SHEARED
PISTON
SLUG MUSHROOME 7ARM RELEASED
SPACECRAFT
STRUCTURE (REF) Ij
_____._L J , _-_DISC
DISCONNECTASSEMBLYFIRED
Figure 11-30 Main Parachute Disconnects
11-66
_,-. SEDR3O0
in the breech, to a common chamber at the head of the plunger. The gas pressure
exerts a force on the head of the plunger, which in turn propels the piston by phy-
sical contact. The piston severs the shear pin and is driven into the arm. The
plunger is prevented from following the piston, since the head of the plunger
strikes a shoulder in the adapter. The snubber disc provides a cushioning effect,
to pre_nt shearing the plunger head. As the piston strikes the back of the arm,
the lead slug at the end of the piston mushrooms. Mushrooming of the slug, retains
the piston in the arm, preventing the piston from hindering arm operation. The
pull of the parachute causes the arm to cam open, thus releasing the riser or
bridle.
PYROTECHNIC VALVES
DES CRIPTION
Pyrotechnic valves(Figure ll-31)are installed in the Orbit Attitude and Maneuver-
ing System(OAMS)and in the Re-entry Control System(RCS). The valves are one time
actuating devices, used to control the flow of fluids. The spacecraft contains
pyrotechnic valves that consist of the electrically fired high explosive cartridge,
valve body, nipple, ram, seal, and screw. The nipple, either open or closed depend-
Ing on the particular valve, is installed in the valve body and welded into place.
The ram, incorporating the seal and screw at its head is located in the valve body,
indexed directly above the center of the nipple. The cartridge is installed in the
valve body at the top of the ram head. Two types of valves are used; normally open
and normally closed. The "A" packages of the RCS and OAMS, contain a normally
closed non-replaceable valve. Tne "E" package of the OAMS contain a normally open,
and a normally closed, non-replaceable valve. If the valves in the "A" and "E"
11-67
___ SEDR 300 __
-- :_i_PROJECT GEMINI
OAMS PACKAGE"E"
BEFORE FIRING
P,
FLOW /
RCS PACKAGES AFTER FIRING"A" "C" "D'
A AND B RINGS (REF)
NIPP
NORMALLY OPEN VALVE (NON-REPLACEABLE)
Figure 11-31 Pyrotechnic Valves (Sheet 1of2)
11-68
k__,_ _ PROJECT GEMINI
OAMS PACKAGE "A" AND "E" OAMS PACKAGE "C" AND "D"RCS PACKAGE"A" RCS PACKAGE "C" AND "D"
BEFOREFIRING BEFOREFIRING
i'#,iSECTIONS IPPL[ SECTIONS
FLOW REMOVED FLOW REMOVED
NORMALLY CLOSED VALVE (NON-REPLACEABLE) NORMALLY CLOSED VALVE (REPLACEABLE)
/
Figure 11-31 Pyrotechnic Valves (Sheet 2 of 2)
11-69
f.._ SEDR 300
packages are defective, or the cartridge has been fired, the packages must be
changed. The "C" and "D" packages of the RCS and the 0AMS contain normally
closed replaceable valves. These valves are attached to the exterior of the
package, and if defective or the cartridge fired, may be changed individually.
OPERATION
Normally open valve; the pyrotechnic valve is caused to function when the car-
tridge is initiated by a 28 vdc electrical signal. Ignition of the cartridge
produces gas pressure that acts on the head of the ram. The ram is driven down
on the nipple, severing and removing a section of the nipple. The ram, having a
tapered cross section, is wedged in the nippled opening, completely sealing the
nipple, thus stopping the flow of fluid. Normally closed pyrotechnic valves are
all basically identical except for nipple design. The non-replaceable valve has
two closed end nipples butted together. The ram severs and removes the end of
each nipple and wedges itself between the ends. A hole is incorporated in the
ram, allowing fluid flow after ram actuation. The replaceable pyrotechnic valve
has a nipple installed with a bulkhead in the cross section that stops fluid flow.
The ram removes the section of the nipple containing the bulkhead and wedges it-
self in place. A hole incorporated in the ram allows fluid to flow.
RETR0_RADE ROCEET SYSTEM
The retrograde rocket system (Figure 11-32) primarily consists of four solid pro-
pellant rocket motors and eight igniter assemblies. The retrograde rockets are
provided to retard spacecraft orbital velocity for re-entry and to provide
distance and velocity to clear the launch vehicle in the event of an abort during
ascent. The rocket motors are s_,_etrically located about the longitudinal axis _-_
SEO30oADAPTER,
GRADE
ROCKET
//
/
RETROGRADE ROCKETIGNITER ASSEMBLY
RETROGRADE ROCKETMOTOR
MOUNTING LUG
NozzLE ASSEMBLY
TEST ADAPTER/
Figure 11-32 Retrograde Rocket System
11-71
___ SEDR300 __
PROJECT GEMINI
of the spacecraft and are mounted in the retrograde section of the adapter. The
rocket motors are individually, optically aligned prior to mating the adapter to
the re-entry module.
RETROGRADE ROCKET MOTOR ASSEMBLY
Description
The spacecraft contains four retrograde rocket motor assemblies (Figure 11-33)
that are identical in design and performance, spherical in shape, and are approx-
imately 13 inches in diameter.
Rocket Motor Case
The motor case is formed from titaniumalloy in two hemispherical sections. The
halves are forged, machined, and welded together at the equator. Each hemisphere _
is insulated to reduce heat transfer during motor operation. The aft hemisphere
is drilled and tapped to provide a mating flange for the nozzle assembly. The
nozzle assembly, a partially submerged type, consists of the expansion cone, throat
insert and nozzle bulkhead. The nozzle bulkhead is a machined titanium alloy,
bolted to the flange at the aft end of the motor case. The bulkhead is threaded
to provide for expansion cone installation. The expansion cone is compression
molded of vitreous silica phenolic resin and is threaded into the nozzle bulkhead.
The throat insert is machined from high density graphite snd is pressed into the
nozzle bulkhead. The throat insert is insulated from the bulkhead by a plastic
material to reduce heat transfer during motor operation. The throat insert is re-
cessed into the motor case to reduce nozzle assembly length. A rubber nozzle
closure is sandwiched between the throat insert and the nozzle bulkhead. The
closure incorporates a shear groove that permits ejection at a predetermined
11-72
---_. SEDR 300
iNSERT
<..ANOZZLE BULKHEAD "_,
LANT
TES_ GRAIN
%.
INSULATION
ASSEMBLY
(2 RE(_)
SHEAR GROOVE
EXPANSION CONENOZZLE CLOSURE
NOZZLE CLOSURE
SHEAR GROOVEEXPANSION _::,.CONE :_ii_'::";..:....
ROATINSERT
INSULATION
SECTION A-AIGNIT
GRAIN CONFIGURATION
NOZZLE BULKHEAD " J
/_" MOTORCASE
Figure 11-33 Retrograde Rocket Motor Assembly
11-73
sEoR30oPROJECT GEMINI
internal pressure level, or basically at motor ignition. A test adapter fitting
is incorporated in the closure to permit pressure checking of the rocket motor.
Rocket Motor Propellant
The motor case is lined with a rubber material that provides propellant grain to
case adhesion. The rocket motor propellant is cast and cured in the motor case.
The propellant grain is cast in an internal burning eight pointed star configura-
tion. The propellant grain is ignited by the two igniter assemblies, mounted 180
degrees apart on the aft end of the motor case, adjacent to the nozzle assembly.
Operation
The retrograde rocket motors function in two modes: normal and abort. In the
normal mode of operation, the rocket motors are used to initiate spacecraft re- _-_
entry. The rocket motors are fired at 5.5 second intervals in 1-2-3-4 order.
The propellant grain of the rocket motor is ignited by the hot gases from the
igniter assemblies. The propellant grain burns over the entire surface of its
eight pointed star configuration until exhausted. The thrust produced by the
motors is transmitted to the spacecraft structure and retards spacecraft velocity.
In the abort mode of operation, the rocket motors are fired in salve or as mission
requirements may direct.
RETROGRADE ROCKET IGNITER ASSEMBLY
Description
The retrograde rocket igniter assemblies (Figure 11-34) are used to ignite the
propellant grain of the retrograde rocket motor. The spacecraft contains eight
igniter assemblies that primarily consist of the case, head cap, grain, booster _
ii-7_
PROJECTOEM,N,
IGNITER ASSEMBLY2 PLACE!
f--
j;U TONBOOSTER ELLETSBOO E PELLE, SK..........__ _ _ S _N_cT_
Figure 11-34 Retrograde Rocket Igniter Assembly
11-75
___ SEDR 300 __
PROJECT GEMINI
pellets, pellet basket, and initiator. The igniter assembly is essentially a
small solid propellant rocket motor. The case and head cap are individually mach-
ined from a stainless steel alloy and have a threaded interface. On the inter-
nal surface of the case, at the gas exit, a silica-phenolic insulating material
is bonded to reduce heat transfer during igniter firing. The grain is cast and
cured in a phenolic paper tube. The grain is inserted into the igniter case
prior to case and head cap assembly. The booster pellets, consisting of boron
potassium nitrate, are contained in the pellet basket, located in the head cap.
The pellet basket is held in place by the head cap and is installed prior to
case and head cap mating. The initiator cartridge consists of the body, one
firing circuit (bridge wire), ignition mix, and output charge. The basic func-
tion of the initiator is to fire the igniter. The initiator is threaded into
the head cap of the igniter at the time of igniter assembly build up at the
vendors.
Operation
When initiated by a 28 vdc electrical signal, the initiator of the igniter
assembly is activated. The initiator ignites the boron pellets, which boosts
the burning to the igniter grain. The igniter grain generates hot gas which is
exhausted into the retrograde rocket motor cavity. The hot gases provide the
temperature and pressure for retrograde rocket motor propellant grain ignition.
Either igniter is sufficient to initiate burning of the rocket motor.
11-76
__ SEDR300 -__
PROJECT GEMINI
DOCKING SYSTEM PYROTECq_NIC _VICES
The pyrotechnic devices utilized for docking are located in the R and E section
of the spacecraft (Figure 11-3). The pyrotechnic devices utilized for docking
consist of:
(a) A pyrotechnically actuated indexing bar to aid the astronauts in matlng
the spacecraft with the TDA.
(b) Three pyrotechnically ejected latch receptacles.
(e) Three pyrotechnically actuated cable cutters to release the three latch
covers.
DOCKING BAR ASSEMBLY
Descriptlonl-
The docking bar assembly (Figure 11-35) primarily consists of the cylinder, inner
piston, inner piston extension, cutter piston, indexing bar, housing/manlfold,
two breeches_ extension cartridge, two jettison cartridges, and a locking mecha-
nism. The basic function of the assembly is to extend and lock the indexing bar,
prior to the docking maneuver and to jettison the indexing bar after the docking
operations are completed. The purpose of the indexing bar is to aid both visually
and mechanically in the docking maneuver. The docking bar assembly is located
in the rendezvous and recovery section and attached to the section structure.
The housing/manifold, mounted at the top of the cylinder, contains the locking
mechanism and provides for installation of two breeches. The breeches provide
for the _nstallation of cartridges. One extension cartridge is installed in the
extension breech on the left hand side of the manifold and two jettison cartridges
are in_talled in the jettison breech on the right hand side of the manifold.
11-77
1-- -_.. SEDR 300
RETAINING
LOCKING INDEXING BAREXTENDED
LOCK
PIN INDEXINGBAR
CARTRIDGE
RETAINING PiN
LOCKING
SHEAR
PIN
.CYLINDER
PISTON
PISTON EXTENSION
ORIFICE
NNER PISTONEXTI_NSION
BEFORE DOCKING BAR INDEXING BARACTIVATION ASSEMBLY EXTENDED
EXTENSION CYCLE
ACTIVATED
_-..,
Figure 11-35 Docking Bar Assembly (Sheet i of 2)
11-78
I-T_ _. SEDR 300
INDEXING EARDOCKING BARASSEMBLY
JEITISI_ED
PLUG _"
,so. ,CARTRIDGES
SHEARED
EXTI_ISION
PINENGAGED
SHEAR
_-- sHEAR PIN_
LOCKING
MECHANISM LOCKINGMECHANISM
%_ OUTFRPISiON
PISTON
' INDEXING BAR EXTENDED INDEXING BAR
JETTISON CYCLE JETTISONED (RE-ENTRY)ACTIVATED
Figure 11-35 Docking" Bar Assembly (Sheet 2 of 2)
11-79
__ SEDR 300 __
PROJECT GEMINI
The hollow inner piston extension protrudes through the cutter piston base
and is attached to the inner piston. The indexing bar is secured to the
outer piston by a shear pin and retained in the cylinder by a retaining pin.
Operation
The indexing bar is extended by the gas pressure generated when the extension
cartridge is initiated by a 28 vdc signal. The gas pressure is ported into the
cylinder and enters an orifice in the base of the indexing bar. The gas pressure
is then ported through the hollow inner piston extension and exerts force on the
bottom of the outer piston, causing it to sever the retaining pin and extend the
outer piston with indexing bar attached. When fully extended, the indexing bar
is secured in place by the pin of the locking mechanism engaging a groove in
the outer piston. The extension cartridge has dual bridge wires with a separate
electrical circuit to each bridge wire for redundancy. The indexing bar is
jettisoned when the dual jettison cartridges are initiated by a 28 vdc electrical
signal. The gas pressure generated by the jettison cartridges is ported into the
cylinder and through an orifice in the outer piston to a cavity between the inner
and outer pistons. The thrusting action of the inner piston causes the indexi1_
bar to sever the shear pin and be jettisoned. Both the inner piston and the
indexing bar are Jettisoned. Initiation of only one jettison cartridge is suffic-
ient to jettison the indexing bar. The jettison cs_-tridgeshave a two second
pyrotechnic time delay to assure that during an abort mode the extend cartridge
will be firing first to extend the bar before the Jettison cycle iz initiated.
EMERGENCY DOCKING RELEASE SYSTEM
Description
The emergency docking release system (Figure 11-36) contains three release assem-
i1-8o
SEDR 300
EMERGENCY DOCKINGRELEASESYSTEMINSTALLATION
LATCH RECEPTACLE STUD
LATC PA ' GASPRESSOREPORT
_--_. _ ELECTRICALCATTACHED_---.....__
• CONNECTOR
j--
CARTRIDGE
SHEAR PI COLLAR
EMERGENCY DOCKING RELEASEASSEMBLY (UNACTUATED)
C LOSURE CAPSPACECRAFT GAS PRESSURE
STRUCTURE,_EE_--_/
GAS PRESSUREPORT
ELECTRICAL
PAWL _---- " _ CONNECTOR
(SHEARED)
SHEAR PIN (SHEARED). COLLAR
EMERGENCY DOCKING RELEASEASSEMBLY (ACTUATED)
/f-_.
Figure 11-36 Emergency Docking Release System
11-81
__ SEDR300 _____
PROJECT GEMINI
blies located s2mmetrically about the longitudinal axis and attached to th,
rendezvous and recovery structure at the outer mold lineo The basiefunction
of the release system is to positively release the spacecraft from the docking
vehicle in the event of normal release system failure. The release assemblies
primarily consist of a latch pen, body, piston closure cap stud assembly and
two gas pressure cartridges. The release body is attached at the opposite side
of a common structural member of the rendezvous and recovery section. A stud
assembly, torqued into the dockir_ release piston protrudes out of the unit body.
The latch pan is secured to the stud by a collar and shear pin. Installation
of the closure cap seals the area above the release piston. Each unit is acti-
vated by two pyrotechnic cartridges, installed in the base of the release body.
Operation
Malfunction of the normal docking release system requires the use of the emergency
docking release system. Upon receipt of a 28vdc electrical signal, the cart-
ridges are ignited in all three release assemblies. The cartridges generate gas
pressure that is ported through drilled passages in the body to the underside of
the piston. The pressure displaces the piston, separating the shear pin fro_
the latch pawl and moving the piston stud up into the body to clear the latch pawl.
The latch pawl being held by the docking vehicle latch, is thus parted from the
emergency release body. The spacecraft is then free to move out of the docked
condition.
TRANSPONDER COVER
Description
The Target Docking Adapter has a single Flexible Linear Shape Charge (FLSC)
ring assembly to open and jettison the transponder cover. The jettison assembly
ii-82
,____ SEDR300 "[_ ---_
PROJECT GEMINI
consists of one FLSC, installed around the inside edge of the transponder cover,
two detonator blocks, two detonators and an explosive ring. The explosive ring
is bolted to the TDA under the transponder door. Detonators are inserted in
the two detonator blocks, which are installed on the TDA under the explosive ring.
Operati on
Upon receipt of a 28 vdc electrical signal, the detonators transmit a detonation
wave that is propagated to the FLSC. The FI_C detonates, severs and jettisons
the transponder door skin. The explosive ring absorbs the shock in back of the
blast and protects the transponder and structure from shrapnel. One strand of
FLSC is adequate to sever and Jettison the transponder door.
_- FUEL CELL HYDROGEN TANK VENT ACTUATOR
Description
The fuel cell hydrogen tank vent actuator (Figure 11-37) is utilized to puncture
the hydrogen tank plnch-off tube to allow any gas in the void between the inner
and outer tank walls to escape after the spacecraft is in orbit, thus increasing
the thermal efficiency of system. The hydrogen tank vent actuator assembly con-
sists of a body, piston/cutter blade, shear pin, guard, breech, ballistic hose,
and an electrically fired gas pressure cartridge. The body houses the piston/
cutter blade which is retained by a shear pin and provides for attachment of
the blade guard. The actuator body assembly is bonded and strapped in place on
the fuel cell hydrogen tank and connected to the breech by a ballistic hose. The
breech contains the gas generating cartridge and is positioned approximately nine
inches from the actuator.
_i-83
BODY SHEAR PIN
PISTON BLADE
VENT TUBE
BLADE GUARD
NK (REF)
BODY SHEAR PiN
PISTON BLADE GUARD
i
! ENT TUBE (SPLIT)
/ _lt II _ _) I .- BLADE
/ i\ .......... /
H2 TANK(REF)
Figure 11-37 H 2 Tank Vent Actuator Assembly
11-84
___ SEDR 300 _ ___._
PROJECT GEMINI
Operation
When initiated by a 28 vdc electrical signal, the cartridge generates gas
pressure that is transferred through a ballistic hose to the actuator body
which exerts a force on the piston/cutter blade. The gas pressure causes
the piston/cutter blade to mov% severing the shear pin and spliting the
vent tube. The blade guard is attached to the actuator body and forms a
barrier which will stop the blade when the piston reaches the end of its
stroke.
11-85/86
LANDING SYSTEM
TABLE OF CONTENTS
TITLE PAGE
SYSTEM DESCRIPTION 12 3 ==:=:====::"=:=_°°°o°.°c:;=;=,o.o_° *..._.......=..= .................
SYSTEMOPERATION.................................. _2=3 !N_;__:-?_d_i$i_._EMERGENCYOPERAnON............................12-Z i!i_-iHii'_'_'.--'-'_:=-_
SYSTEM UNITS ............................................ 12-7 !-_HilHi!iiiiHiii-_lii-_DROGUE PARACHUTE ASSEMBLY ................. 12-7 iiiiiiiiiiii_ii_i_iii_i***.°o..o.°°oooo...°°o.o**.
PILOTPARACHUTEASSEMBLY...................... 12-12 iiiiii_ii_i!i!i!!iiii_!i!MAtHPARACHUTEANDR_SERASSEMBLY..... ]2-_4 ii_iiii_ii!iiiiiii!!iiiiiil
:::::::::::::::::::::::::::
ig!_iiiiiiiiiiiiiiii_ii_i:::::::::::::::::::::::::::
iiii_ii_!iiiiiiiiiiiiiiiiilo°.°°°.°°o.°°oo..°°°°°°°°°.
:::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::
:::::::::::::::::::::::::::
iiiiiii_i_i;_iiii!ii!iii_il:::::::::::::::::::::::::::_-" iiiiiiiiiiiiiiiiii_i!_i!!ii
°°**°° .....................::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::
i_iiiiii!i_i!ii_ii_iiiiii!...................... o ....
12-1 :::::::::::::::::::::::::::
_ SEDR 300
PILOT MORTAR ASSEMBLY
DROOl
UE MORTAR ASSEMBLYBRIDLE
DISCONNECT
",,,/ LEG
/
// _PILOT CABLE/
// .."_ _ SECTION A-A
RISER MAiNGUARD RING ASSEMBLY
/.\ , ADAPTERASSEM,LY'_" _"_'_'_. ...... _/ RECOVERY SECTION
(R &R)
// CARTRIDGES
I/
MORTAR
CABli FORWARD BRIDLE(REF) DISCONNECT
RC S SECTION
(REF)J
LEG
SINGLE POI!DISCONNECT
MAIN pARACHUTE NOSE FAIRING
CONTAINER AFT BRIDLE
MAIN PARACHUTE DEPLOyIVENT _LE_I I
BAG AND CONTAINER ASSEMBLY CABLE TROUGH
FORWARD BRIDLE LEG
MAIN DISCONNECT BRIDLE LEGTRAy
BRIDLEASSEMBL
TO MAIN pARACHUTE
SlCllOljB-B SECTIONC-C
Figure 12-1 Parachute Landing System
12-2
SEDR300 ___
PROJECT GEMINI
f'\
SECTION XII LANDING SYSTEM
SYSTEM DESCRIPTION
The parachute Landing System (Figure 12-1) provides a safe rate of descent to
return the re-entry module to the earth's surface and furnishes the proper atti-
tude to the landing module for a water impact. A system of three parachutes in
series is utilized for stabilizing and retarding the velocity of the spacecraft
from 50,000 feet. During the final stage of descent, the main parachute sus-
pension is changed from a single point to a two point system in order to achieve
a more favorable attitude for a water landing. The Landing System consists of
three parachute assemblies (a drogue parachute assembly, a pilot parachute
assembly, and a main parachute assembly), two mortar assemblies, reefing cutters,
disconnect assemblies, riser assemblies, and attaching hardware. The entire
Landing System, with the exception of the aft bridle leg and disconnect assembly,
is located in the rendezvous and recovery section of the spacecraft. Figure 12-2
illustrates the sequence of events from re-entry to impact in block diagram form.
Figure 12-3 illustrates the electrical sequence of the landing system.
SYSTEM OPERATION
Prior to re-entry, the landing and postlandlng common control electrical buses
are armed by positioning the LANDING switch to ARM. This also applies power to
the two barometric pressure switches for illumination of the lO.6K and 40K warning
Indicaters.
When the altimeter indicates an altitude of 50,000 feet, the HI-ALT DROGUE switch is
manually activated. The drogue switch energizes two single pyrotechnic cartridges
_ in the drogue mortar. The drogue mortar deploys the reefed drogue parachute. The
12-3
___ SEDR300 __
PROJECT GEMINI
BAROMETRIC UGHT _ _PRESSURESWITCH ILLUMINATESACTIVATES
DROGUE DROGUE DROGUE CHUTESWITCH MORTAR CHUTE DISREEPED
DEPLOYED (16 $EC. T.D,)
BAROMETRIC 10.6K WARNINGSWITCH LIGHTACTIVATES ILLUMI SATES
UTE DISREEFED
DEPLOYED (6. SEC. PYRO T. D
DROGUECHUTE
DISCONNECTGUILLOTINES(80MS PYRO T.D.)
SWITCHI
2.5 SEC. JELEC.
ITIME DELAy
I MDF RING REEFED
FIRED MAI N CHUT E
(50-70 MS T.D.) DEPLOYED
9DISREEFED "-_
(10 SEC. PYRO T.D
LDG ATT SINGLE POINT SUSPENSION ANTENNA
SWITCH DISCONNECT J_l/I EXTENDED
UHF DESCENTANTENNAEXTENDED
BEACON
ACTIVATED Q
I IMPACT +
"_! HOIST LOOP J
AND FLASHINGRECOVERY LIGHTRELEASE LEGEND
_IIP PILOr ACTUATED @
PAPA JET/ PARACHUTESWITCH JETTISONED MECHANICAL CONNECTION
RECOVERYLIGHT
ACTIVATED -_._._,_____; -,. ._
Figure 12-2 Landing System Sequential Block Diagram
12-4
___ SEDR 300 .__]
PROJECT GEMINIf_" r- .................... DROGUE .................... "-1
' r"l 'DROGUE CHUTE _ _ O O
O O i C_ _ T -- MORTAR --LDG. SEQ. LDG. SEQ.
CONT. Ri K5-93 "-_ _ "CONT. /2
I CABINAIR t " K5-94t'__ INLET IGN. --,,._
I-- ................... PARA DEPLOY ................... 7I I
K5-87 t DROGUE CHUTE K5-88"ANDING SQUIB I _ = - DISCONNECT -- ( LANDING SQUIB
I
Ks._ APEXCAR_ _5__ ....... GU,LLOT,NE.........
m PILOT CHUTE
MORTAR KSB-86K5-85 K5-89 K5-90
__,B:_-"I MA,NCHUTEoEPLOyI--__,_L- .................. EMERG DROGUE ...................
o._ - S,NGLEPO,NT-- _ 'I RELEASE II I
L ................... LDG. ATT. -- .................. J
I- .................... pAPA JETT .................... 1
FLASHING J jI RECOVERYI LIGHT RELEASE K5-22 I
- I_ ---- MAIN CHUTE -- _ "JETTISON FWD
f_ MAIN CHUTEJETTISON AFT
_" _J I-_- _5__._SEC I ;•, _ TIMEDELAY --
-- UMBILICAL --
PYROSWITCH
R & R SECTION 1
K5-91 COAXIAL / K5-92
A K5-7 GUI LLOTIN E K5"g A
' I I ' "_"_- _,RSECt'ON--"_--_ II" WIRE
GUILLOTINE
K5-91 j _ K5-91
IB K5-97 DESCENT B; I I ' /"_--_-- - ANTENNA -- II I
._ SELECT
J R &r SECTION i K5-6
K5-7 K5-5 JETTISON PRIMER 3B K5-8
I I ,-'q_-__ co_D(_) -__n'-Cq_ I I- / R'RSECT'ON/ ---pYRO SWITCH
10.6K
INDICATOR LANDING AND
PG_TLANDINGMAIN BUS 10.6K CONTROL BUS
BAROMETRIC
oo h 60pAPA CONT.SEQ. LIGHT 8AROSTATPWR.
BAROMETRIC 4OKsw_rcn _nD,CAtOR
Figure 12-3 Landing System Schematic
12-5
SEDR 300
parachute is reefed to 31m_t the opening shock load. Sixteen seconds after de-
ployment, two pyrotech-lc reefing cutters disreef the drogue parachute. The drogue
parachute s+__b_iLtzes the re-entry module.
At appro×lwately 10,600 feet, the PARA switch is activated. The PARA switch fires
the three dro_e cable guillotines and sets a 2.5 second time delay to the MDF
ring detonators. After the drogue riser legs have been cut, the drogue parachute
pulls away from the re-entry module extracting the pilot parachute fr_n the pilot
mortar tube with the apex line. When deployed, the pilot parachute is reefed to
limit the initial shock load. Two lanyard initiated pyrotechnic reefing cutters
disreef the pilot chute six seconds after deployment. 2.5 seconds after the pilot
chute has been deployed, the MDF ring fires separating the rendezvous and recovery
section frcB the landing vehicle. The pilot parachute functions to decelerate the
re-entry module, remove the rendezvous and recovery section, and deploy the main
parachute.
As the landing module falls away from the rendezvous and recovery section, the
main parachute is deployed in a reefed condition. The main parachute is disreefed
by three lanyard initiated pyrotechnic reefing cutters ten seconds after deployment.
The two decelerations provided by the main parachute divide the retarding shock
load. After the main parachute has been disreefed, the manually operated L_G ATT
switch is actuated to chan_e the single point suspension system to a two point
suspension system. The two point suspension system provides a more favorable
attitude for impact. As soon as the landing module contacts the ocean surface,
the PARA _ switch is activated. The PARA JETT switch energizes the forward and
aft bridle disconnects releasing the main parachute. Upon c_pletion of the
_.__ SEDR 300
PROJECT GEMINI
landing, the landing module is prepared for transmitting data and recovery
information through the erected recovery antenna.
_ENOY OPERATION
In the event the drogue parachute does not deploy or deploys improperly, the
PHE-MAIN IO.6K switch is actuated. The closure of this switch fires the three
drogue cable guillotines, the apex line guillotine, and the pilot parachute
mortar and starts the 2.5 second time delay to the MDF rings. The pilot mortar
deploys the pilot parachute in a reefed condition. From this point, the emer-
gency sequence of events is exactly the same as used during a normal landing.
Figure 12-4 illustrates the emergency sequence of events in block diagram form,
and Figure 12-5 illustrates the emergency deployment.
DROGUE PARAC_ ASSEMBLY
The drogue parachute assembly (Figure 12-6) stabilizes the re-entry module and
deploys the pilot parachute. This assembly consists of an 8.3 ft 441ameterconi-
cal ribbon parachute with twelve 750-pound tensile strength suspension lines.
A three legged riser assembly attaches the parachute assembly to the rendezvous
and recovery section.
When initially deployed, the drogue chute is reefed to 43 percent of the parachute
diameter in order to reduce the opening shock load. Sixteen seconds after deploy-
ment, two pyrotechnic reefing cutters disreef the drogue chute. Initiation of the
PARA switch fires three cable guillotines located at the base of the three riser
legs. As the drogue chute pulls away from the rendezvous and recovery section, an
12-7
_;,_:. SEDR300
BARO. SW 10.6K WARNINGACTIVATES LIGHT ILLUMINATES
i oouEMERG 10.6K PARACHUTE DISCONNECTSWITCH APEX LINE GUILLOTINES
GUI LLOTI NE (80 MS PYRO T .D. )
QPARACHUTE PILOT PARACHUTEMORTAR PARACHUTE DISREEEED(.5 SEC PYRO T .D.) DEPLOYED (6 SEC PYRO T .D.
ELECT. FIRED SECTIONTIME DELAy (50-70MS T.D.) SEPARATES
MAIN CHUTE
LEGEND DEPLOYED
(_II_P_ PILOT ACTUATED
ELECTRICAL CONNECTION (IOSEC T.D.)
q H H
LOG ATT MAIN PARACHUTE BRIDLE UHF DESCENTSWITCH SINGLE POINT SUSPENSION ANTENNA
DISCONNECT ACTIVATED EXTENDED
U.ERESCUE¢ ®ANTENNAEXTENDED
QUHF DESCENT AND FLASHINGANTENNA RECOVERY LIGHT
EXTENDED RELEASE ! IMPACT +
pARA JETT PARACHUTESWITCH JETTISONED
FLASHING l
RECOVERY .... ,7_ _ _._.L,G.T ®'' "iiit/':_ACTIVATED
Figure 12-4 Emergency Landing Sequential Block Diagram -- ->_
12-8
-_-.. SEDR 300
69" _EF)-_
(TYPICAL 12 pLACES)
(TYPICAL 12 PLACES)
96" (REF)
/PENSION LINE-RISER JOINT /
._/_----'RISER SEPARATION JOINT
110" (REF)
._RISER FITTING JOINT
(TYPICAL 3 PLACES)
/,
PARACHUTE IN REEFED CONDITION PARACHUTE IN DISREEFED CONDITION
Figure 12-6 Drogue Parachute Assembly
12-10
__ SEDR 300 ___
PROJECT GEMINI
apex line, which is attached to one of the riser legs, extracts the pilot
parachute from the pilot mortar tube. The drogue parachute remains attached to
the pilot parachute during the entire descent of the rendezvous and recovery
section of the re-entry module.
Drogue Parachute Morts_,Assembly
The drogue parachute mortar assembly stores and protects the drogue parachute
during flight and deploys the drogue parachute when activated by the HI-ALT
DROGUE switch. An insulated metal pan retains the parachute in the mortar tube
which has a diameter of 7.15 inches and is 9.12 inches long. The breech assembly,
located at the base of the mortar tube, contains two electrically actuated pyro-
technic cartridges and an orifice. The cartridges generate gases which enter the
.... mortar tube through the orifice and eject the drogue parachute and sabot.
Drogue Mortar Sabot
The drogue mortar sabot is an aluminum cup located in the base of the mortar and
functions to eject the drogue parachute with a piston llke action. In order to
insure the most effective ejection, the sabot is fastened to the base of the ori-
fice by a frangible bolt, and an O-ring, located near the base of the sabot which
contacts the inner wall of the mortar tube to prevent any escape of gases gen-
erated by the two pyrotechnic cartridges. When enough pressure to break the
frangible bolt has built up, the sabot and parachute are expelled from the mortar
tube. After ejection, the sabot remains attached to the parachute bag and aids in
stripping the bag from the parachute.
Drogue Parachute Deployment Bag
The drogue parachute deployment bag protects the drogue parachute during deployment
12-11
.@ SEDR 300 _---1
PROJECT GEMINI
and allows for an orderly deployment of the parachute. The bag is fabricated
from cotton sateen and nylon. A 0.35 pound aluminum plate, sewn into the top
of the bag, aids in stripping the bag from the canopy during the deployment.
PILOT PARACHUTE ASSEMBLY
The pilot parachute assembly (Figure 12-7) decelerates the re-entry module and
removes the rendezvous and recovery section from the lending module which results
in the deployment of the main parachute. During flight, the pilot parachute
assembly is stowed in the pilot mortar tube. The 18.3 foot dieaeter canopy is of
the ringsall type having 16 gores and fabricated from I.I and 2.25 ounce per
square yard nylon. Sixteen nylon cord suspension lines, which are 17 foot long
and have a tensile strength of 550 pounds each, attach the canopy to the riser
assembly. A iO.7 foot long split riser, constructed of four layers of 2600 pound
tensile strength dacron webbing, holds the pilot parachute assembly to the rendez-
vous and recovery section of the spacecraft. When initially deployed, the pilot
parachute is reefed to 11.5 percent in order to limit the opening shock load to
SO00 pounds. Two pyrotechnic reefing cutters disreef the parachute 6 seconds
after deployment. The pilot parachute remains attached to the rendezvous and
recovery section throughout the entire descent.
Pilot Parachute Mortar Assembly
The pilot parachute mortar assembly is similar in design and operation to the
drogue parachute mortar assembly. During normal operation of the Landing System,
this assembly serves only to store and protect the pilot parachute. In the event
of a failure in the deployment of the drogue parachute, the pilot parachute mortar
can be activated to deploy the pilot parachute by initiation of the PI_E-_IN IO.6_T_ _"
12-12
t_ SEDR300
JOINT (16)
LINES (16)
APEX LINE
13.0 IFT(REF)
RISER
JOINT (2)
RECOVERY SECTION
16.5 FT(REF)
PARACHUTE INREEFED CONDITION
PARACHUTE INDISREEFED CONDITION
Figure 12-7 Pilot Parachute Assembly
12-]3
.__ $EDR300 _______
PROJECT GEMINI
switch. Actuation of the PEE-MAIN IO.6K switch fires the three drogue cable
guillotines, the apex line guillotine, and the pilot parachute mortar. After
the pilot parachute has been deployed, the landing is completed through the normal
sequence of events. Figure 12-5 illustrates the pilot parachute deployment.
Pilot Mortar Sabot
The pilot mortar sabot functions are the same as those of the drogue mortar sabot.
Refer to the description of the Drogue Mortar Sabot.
Pilot Parachute Depl0_ment Ba 6
The pilot parachute deployment bag is similar to the drogue parachute deployment
bag in design and use, except for the bag handles attached to the apex line for
extraction by the drogue parachute.7--
MAIN PARACHUTE AND RISER ASSemBLY
The main parachute (Figure 12-8 and 12-9) is of the ringsail type with a diameter
of 84.2 feet. The nylon canopy has seventy-two gores alternating in colors of
international orange and white. Seventy-two suspension lines are attached to
eight legs of a single inte6_l riser. Each suspension line has a tensile strength
of 550 pounds. The 3.25 foot integral riser consists of eight layers of 5,500
pound tensile strength nylon webbing. The canopy is fabricated from i.i and 2.25
ounce per square yard nylon and is designed to operate at a dynamic pressure of
120 pounds per sqn-_e foot. However, by reefing the main parachute, a maximum
load of 16,0OO pounds is experienced at deployment. When initially deployed, the
parachute is reefed to i0.5 percent. The disreefed main parachute allows a maxi-
mum rate of descent of 31.6 feet per second for a module weight of 4,400 pounds.
12-14
.-_ SEDR 300
LINE
949.6 IN. (REF)
MAIN PARACHUTE
26.4 IN. (REF)I
/_11_ PARACHUTE IN REEFEDCONDITION
Figure 12-8 Main Parachute and Single Point Suspension System
12-15
SEOR300_ PROJECT GEMINI
•58 FT (REF)
22 FT
(REF)
SUSPENSION LINEJOINTS
LINES (72)
MAIN pARACHUTE
r _'-MAIN PARACHUTE
BRIDLE LEGS
061N.
PARACHUTEcoNDITIoNINDISREEFEDLINE OF VERTICAL DESCENT
Figure 12-9 Main Parachute and Two Point Suspension System
12-16
__ SEDR 300
PROJEC GEMINI
Main Parachute De_lo_ment Ba_ and Container Assembly
The main parachute deployment bag and container assembly (Figure 12-1) stows the
maln parachute. This assembly is located in the aft end of the rendezvous and
recovery section of the spacecraft. The deployment bag is fabricated from a
cotton sateen material reinforced with nylon webbing. In order to insure a full
and orderl_vdeployment of the maln parachute, the suspension lines must be
stretched out prior to the release of the canopy. Therefore, transverse locking
flaps are incorporated in the bag to separate the canopy from the suspension
lines. Four restraining straps hold the deployment bag in the container until
deployment.
The main parachute container is 22.25 inches in diameter and 21.32 inches long.
The container is closed on the forward end and is secured to the rendezvous and
recovery section by four vertical reinforcing brackets. At deployment, the re-
straining straps of the deployment bag are unlocked, the risers and suspension
lines are extended, and the canopy is pulled from the deployment bag. The deploy-
ment bag remains attached to the container by four bag handles.
Main Parachute Bridle Assembl_m
The main parachute bridle assembly (Figure 12-9).provides a two point suspension
system in order to achieve the optimum attitude for a water landing. Two separate
bridle straps constitute the main parachute bridle assembly. The foi_ard bridle
strap is an 85 inch long nylon strap with a looped end connected to the forward
bridle disconnect. Prior to single point release, the forward bridle is stowed
in the bridle tray (Figure 12-10). The aft bridle is 106 inches long and connects
f-_ to the aft disconnect which is located _-..ediatelyforward of the single point
12-1T
SEDR 300
LEG)DISCONNECT ASSEMBLY
SINGLE POINT HOISTLOOP COVER
TROUGH
DISCONNECT ASSEMBLY
TRAy
MAIN GUARD R_NGASSEMBLY
SUPPORT
DISCONNECT ASSEMBLY
Figure 12-10 Main Parachute Support Assembly
12-18
SEDR 300
hoist loop (Figure 12-10). Constructed of heat resistant nylon, the aft bridle
is stowed in a trough that extends from the front of the Re-entry Control System
section to the aft disconnect during flight. An insulating cover shields the eft
strap in the cable trough until the single point suspension is released, at which
time the bridle leg tears through the insulation.
Main Parachute Release
Upon landing in the water, the main parachute is released from the landing module
by activation of the P_RAJETT switch. This initiates the forward and aft dis-
connect pyrotechnics and allows the chute to pull away from the landing module.
/-
DOCKING SYSTEM
TABLE OF CONTENTS
_ TITLE PAGE
SYSTEM DESCRIPTION ............................... 13-3RENDEZVOUS AND RECOVERY
SECTION .................................................. 13-3
APPROACH AND MOORING SEQUENCE ......... 13-6
RIGIDIZING AND UNRIGIDIZING
SEQ U EN CE............................................... 13- 7 iiiii:':'-'_-i'_;"_SYSTEM OPERATION ................................. 13-9 !ii_iiiiii!iiiii:_?'_ii
DO CK IN G SEQ UEN CE.................................. 13- 9 iiiiiiiiiiiiii_ii:lii!l._i'i..°O''***'H°*'°°'°°H,O*O..°
MANEUVERING SEQUENCE ......................... 13-14 iiiiiiiiiiiiiiiii!i_iiH!i!SEPARATION SEQUENCE .............................. 13-15 !i!iiii!!!!!!!ii_iii_iiiii
H*...o..°.°...°o..** .o.o,.°o...*o.....o...oo. Ho° oo°*
SYSTEM UNITS ........................................... 13-18 _i!i_!!H_i!H_Hiiiiiiiii_°.o.. ....... ....°o°.,o.o°..o...o ...... °.....,,.o..,o,.
MOORING LATCH RECEPTACLES .................... 13-18 iiiiiiiiiiii_iiiiiiiiiii_il:::::::::::::::::::::::::::
MOORING LATCH COVER ASSEMBLIES .......... 13-18 iiiiiiiiiiiiiiiiiH!iiii!!!UMBILICAL CONNECTORS ............................ 13-20 iiiiiiiiiiiiiiiiiiiiHiii!i
............... ,...**......
INDEX BAR ASSEMBLY ................................ 13-20 iiiii!iiiiiiiii!ii!i!iiiii!RADAR SY STEM AN D PYR OTECHNIC iiiiiiiiiiiiiiiiiiiiiiiiiii
r_ _ _ _ _ _
DEVICES .................................................. 13-22 :::::::::::::::::::::::::::...... o ....................:::::::::::::::::::::::::::
TARGET VEHICLE STATUS DISPLAY PANEL ...... 13-22 iiiiiiiiiiiiiiiiiiiiiiiiiii............ .,.o,°o.oo..,°.
_3-_ :::::::::::::::::::::::::::
s oR30o" PROJECT GEMINI
VEHICLE
DOCKING ADAPTER
CONE
UMBILICALPAD ; BAR
US AND
RECOVERY (R & R)SECTION
GEMINI
SPACECRAFT 1 DOCKING
RECEPTACLE
LATCH HOOK
Figure 13-1 Docking System
13-2
___ SEDR 300 __ .__]
PROJECT GEMINI
SECTION XIII DOCKING SYSTem4
SYST_ DESCRIPTION
The Docking System (Figure 13-1) is used on spacecraft 8 through 12 and utilizes
two major assemblies: the spacecraft Rendezvous and Recovery (R & R) section and
the Target Docking Adapter (TDA) assembly. These assemblies permit the spacecraft
to mate with the target vehicle during orbital flight. The TDA assembly may be
mounted on either the Agena Target Vehicle (ATV) or the Augmented Target Docking
Adapter (ATDA). The docking cone receives the spacecraft R & R section during
docking. Electrc_echanical devices in the spacecraft and the docking adapter
control the docking operations automatically or as directed by the pilots or the
_-_ ground control stations. For convenience, these major assemblies are considered
separately. In this section, details of the R & R section of the spacecraft,
approach and mooring, rigidizing and unrigidizing to the target vehicle are pre-
sented. Details of the TDA are presented in the Target Docking Adapter section.
RENDEZVOUS AND RECOVERY SECTION
The R & R section of the spacecraft (Figure 13-2) contains the following Docking
System units: the support structure, a nose fairing, an index bar, 3 mooring
latch receptacles, 3 latch-receptacle covers, an umbilical connector (receptacle),
two rendezvous umbilical connectors (pads), pyrotech-Ic cable cutters and ejection
devices. Other components of the R & R section are not concerned with docking.
The main structure is secured to the spacecraft and serves as a mounting structure
for other units.
".... The nose fairing is rigidly but temporarily fastened to the forward end of the
13-3
.--- SEDR 300
N PARACHUTESTORAGE (REP)
(_(_) R & R SECTIONCHUTE (REF)
1 / \_. CHUTE (REF)
//
///
/ /t ,_ _ RECE,TACLE
I "Q (REF)%
\\
RADAR (RRF.)
--SLOT AND /LATCH
ROD .//
-r -_"_ _---'-'----- "-------;Z _ .....t LATCH RECEPTACLE
-TL .......: xNOSE F,_
(REF) ."*JETTISONED AT S/C SEPARATION
CABLE _'_ _ FROM LAUNCH VEHICLEASSEMBLY
LATCH RECEPTACLE
I___ COVER• _-_ CABLE CUTTER
(pYROTECHNIC)
LATCH
COVER MOORING
ROD ASSEMBLY _ CLOSE LATCH RECEPTACLE
I
,,,J.._.............._-tJ_ ,,, _ _ - R & R SECTION
SECTION A-A
Figure 13-2 R & R Section, S/C
13-4
__.__ SEDR 300 _____
PROJECT GEMINI//
R & R section. The fairing is Jettisoned at spacecraft separation from the launch
vehicle. A pyrotechnic igniter severs a shear pin and swings the fairing clear of
the spacecraft. The docking latch receptacles are exposed as the fairing is
Jettisoned. The index bar is so located that it coincides with the slot in the
docking cone when properly aligned. The bar is extended pyrotechn_cally by the
pilots and remains extended until Jettisoned at retrograde adapter separation.
The pilots use the bar to align the spacecraft with the docking cone. The pilots
use another set of igniters to Jettison the bar. The bar can neither be extended
or Jettisoned prior to nose fairing Jettison.
The three latch receptacles are equally spaced around the forward perimeter of
the R & R section. The lower latch serves as a pivot for the nose fairing when
it is Jettisoned. The exposed receptacles latch with the mooring latch hooks on
the docking cone as the two units come in contact during the docking maneuver.
The latch receptacles will remain latched until unlatched automatically by the
separation sequence. The latch receptacles are Jettisoned Just before re-entry or
during an emergency. The cavities left by the Jettisoned receptacles are covered
by pyrotechnlcally released covers before re-entry. Relays prevent these operations
until after the nose fairing has been Jettisoned.
The Rendezvous Radar is located in the front part of the R & R section and consists
of the transmitter-receiver unit plus 3 receiving antennas and one transmitting
antenna. The radar (and encoder) are used to control the target vehicle before
docking. The pilots can initiate a number of encoded c_ands via the Command Link
System to accomplish this.
13-5
__ SEDR300 ______
PROJECT GEMINI
The rendezvous umbilical pads mate with the rendezvous umbilical lever assemblies
on the docking cone, at the same time the latch hooks engage the latch receptacles.
The main ,n_iLical receptacle mates with its umbilical plug during the rigidizing
sequence of the docking maneuver. _he umbilical connection provides direct hard-
line control of the target vehicle. This control includes simple switching and
encoded commands.
Pyrotechnic igniters are used to extend the index bar, Jettison the index bar and
the mooring latches, Jettison the nose fairing and release the latch covers. The
pyrotechnic igniters are fired electrically from the crew station.
APPRO_m AND MOORING S_UENCZ
The approach and mooring sequence is initiated by a ground station co_and to the
target vehicle. Using the Digital Cr._._d System, the ground station will turn
on electrical power and radar transponder circuits and extend the dipole antenna
on the TDA. This allows *-_ pilots of the spacecraft to make contact with the
target vehicle using the Rendezvous Radar System.
The pilots will begin tracking and control the target vehicle at approximately
IOO n_u_ieal miles separation. When the separation is about 20 nautical miles,
the pilots can use the C_mand Link System to unrigidize the docking cone and
turn on the acquisition Lights. At about 500 feet separation, the pilots use
the C_ Link to turn on the approach lights and status display panel lights.
The pilots will extend the index bar pyrotechnically and decrease the spacecraft
closing rate to 0.7 feet per second. Later, the pilots turn on the suacecraft
docking light. As the target vehicle is approached, the pilots use spacecraft
13-6
_l_- SEDR 300
PROO-----O-ECTGEM IN If
thrusters and the index bar to align the spacecraft with the docking cone.
When proper aliment is achieved the pilots begin the mooring sequence. The
mooring sequence is completed when the spacecraft is firmly connected to the dock-
ing cone in the un_igidized position. First contact between the R & R section
and the docking cone is with three electrical discharge fingers protruding fraa
the docking cone. These discharge fingers are so wired that any static charge
between the spacecraft and the target vehicle will be dissipated at a controlled
rate. As the pilots thrust the R & R section further into the docking coos the
spring loaded mooring latch hooks are punhed aside, then drop behind the latch
receptacles to create the latched condition. At the same time the two rendezvous
_ilical connectors are mated and provide direct hardllne control to the TDA
mooring drive system.
RIGIDIZING ABD UNRIGIDIZ_INGS_UENCZ
The rigidize sequence and the unrigidize sequence are autaaatie ones they are
initiated. The rigidize sequence is initiated when the spacecraft R & R section
contacts the latch hooks, pushes them aside and then allows the hooks to lock on
to the latch receptacles. Latch engagement is sensed by three limit switches,
c_e at each latch position. The sensing switches route target vehicle electrical
power through the rigidize limit switches to the rigidized power control relay.
The relay energizes, connecting target vehicle electrical power to the moorlng
drive system motor. When the system is rigidized, the rigidized l_mlt switches
open, de-energizing the rigidized power control relay and stopping the motor. The
main uubilical plug is extended during the rigidize sequence.
13-7
, BAR
f _ (EXTENDED)
TARGET VEHICLE IONE
R & R SECTION
DISCHARGE DEVICE
,PACECRAFT
LATCH HOOK MOORINGLATCH
APPROACH RECEPTACLE
I
I iPACECRAFT
MOORING
TARGET VEHICLE
; CONE
I
[_> NOT OPERATIONAL ON ATDA.SPACECRAFT
.,°,o,z,+ iW++ IIIISTATUS DISPLAY PANEL +_'_
DETAIL A
Figure 13-3 Docking System Operation
13-8
___ SEDR 300
PROJECT GEMINI
The unrigidize sequence is automatic once it is initiated by either ground command
or by the pilots. When the mooring drive system is unrigidized, the unrigidized
limit switches route target vehicle electrical power to the latch actuator. When
the mechanism is unlatched, the unlatch sensing switch routes target vehicle
electrical power to the separation timer. The timer, after a 30-second delay,
applies power to the reset side of the latch actuator through the latch reset
control relay and the latch reset relay. The latch hooks are reset so that the
docking maneuver can be repeated. Normally the cone does not rigidize after
reset of the latch hooks but if the pilot does not back away within the 30-second
delay before reset, the cone will rigidize. The unrigidize seq_nce would then
have to be repeated.
The Docking System operation (Figure 13-3) illustrates the sequence in _aich dock-
ing is accomplished. Control of the Docking System is primarily electrical but
pyrotechnic devices and mechanical sensing switches are integrated into the
system. The Docking System electrical block diagram (Figure 13-_) ll-_ together
various assemblies used in the docking operation. The electrical systma provides
the capability for the pilots or ground station to control the docking operation.
The operation of the Docking System consists of three basic functions:
(i) Docking sequence, (2) Maneuvering sequence, (3) Separatic_ sequence.
_e docking sequence consists of approach, mooring, rigidizing and main _ilieal
connection.
13-9
I- :_- SEDR300
TARGETS/C TDA VEHICLE
DOCKING _ DOCKING
LIGHT LIGHTSWITCH
INDEX BAR
MBAR EXTEND EXTEND
SWITCH RELAY IGNITER
_T VEHICLE
TRANSPONDER ORPROGRA_ER
I
CONTROL ENCODER NDDE_R l LATCH RESET SEPARATIONCONTROL TIMER
I (30 SEC DELAY)
I
t _ UNRIGIDIZED
MAIN MAIN LIMIT LATCHUMBILICAL UMBILICAL SWITCHES 1 ACTUATOR
(3) _,
s,o_ _ L___I
tCONTROL RENDEZVOUS __ RENDEZVOUS DRIVE POWERSWITCHES UMBILICAL UMBILICAL SYSTEM CONTROL
RELAY
TARGETVEHICLEPOWER
hqHLATCH RELEASE LATCH LATCH SENSING LIMITRELEASE IGNII"ER RECEPTACLE HOOKS SWITCHES SWITCHESSWITCH (3) (3)
TLEGEND
PILOT FUNCTION
MECHANICAL CONNECTION
ELECTRICAL CONNECTION
F/gure 13-4 Docking System Electrical Block Diagram
13-10
• . : •
PROJECT GEMINI
A_roach
The pilots in coordination with the ground station control the Docking System
(Figure 13-3 and Figure 13-4).
During the approach phase, the Digital Co_ud System is used to switch on the
target vehicle covmaandcircuits, the docking adapter radar transponder and to
extend the docking adaoter dipole antenna boom. With this accomplished, the
pilots and the ground stations can control other docking adapter/target vehicle
functions via the Co.,.;.andLink System. Typical commands sent during the approach
phase are UNRIGIDT_R and ACQUISITION LIGHTS ON at about 20 nautical miles
separation; APPROACH LIGHTS 0N, STATUS DISPLAY ON and STATUS DISPLAY _RIGHT at
near approach. When the DOCK light on the status display panel is illuminated,
the system is un_igidized and the latch hooks in docking position. If the light
is not on, the UNRIGIDIZE command is given again, either by the pilots or the
ground station. In addition, the pilots operate the following switches during
the approach phase: DOCK LT (or EXT IE on spacecraft 8 through 12), _JS ARM and
INDEX EXTEND/POD EJECT. In the ON position, the DOCK LT (or EXT LT) switch
applies main bus power through the CABIN LIGHTS circuit breaker to the dock light.
In the INDEX EXTEND position, the _ EXTEND/POD EJECT switch applies 0AMS
squib bus 1 and 2 power to the index bar extend relay coils through the BUS ARM
switch. The energized index bar extend relays apply docking squib bus power to
the extended igniters on the index bar assembly. The pilots use the extended
index bar to align the spacecraft with the docklng cone.
Mooring
_- As proper alignment is achieved_ the pilots maneuver the spacecraft to slide the
indexing bar into the slot (at the tip of the V-notch) in the docking cone (Figure
13 -11
- • _ •.... ,i. ¸ •
SEDR 300
?CONE
TARGET VEHICLE ,f_
FINGERS
SPACECEAFT
T"II
ilI[ARGET VEHICLE SPACECRAFT
I
!
' ///..
Figure 13-5 Rigidizing Operation
13-12
____ SEDR 300
PROJECT GEMINI
13-3). The pilots increase the forward thrust as the spacecraft enters the docking
cone and makes contact with the electrical discharge fingers. This temporary
thrust increase provides enough momentum for the spacecraft R & R section to
actuate the spring loaded mooring latch hooks on the interior of the docking cone.
Actuation of the latch hooks captivates the spacecraft R & R section in the docking
cone and automatically initiates the rigidize sequence. At the same time the two
rendezvous umbilical connectors are mated and provide direct hardline control to
the TDA mooring drive system. The completed mooring operation connects the
spacecraft firmly to the docking cone.
_dizing
Rigidizing (Figure 13-5) is the process of pulling the docking cone in toward the
F-_ adapter main structure until it reaches a firm seating against the pads (part of
the adapter main et_cture). This operation is required because the cone dsmpers
form only a flexible connection between the cone and the adapter structure.
Maneuvering the spacecraft/target vehicle ccmbtnetion is impraetieal under
unrlgidized conditions.
The docking cone is rigidized as soon as the mooring operation is completed. As
mooring is coRpleted_ three limit switches mounted on the TDA latch assemblies
aut_tically activate the mooring drive (rigldizing) mechanism motor, which in
turn, drives the rigldizing linkages through the fl_ble drive cables, gear boxes
and drive arms. The drive arms apply tension to the rigtdizi_ Ltuka_s connected
to the target docking cone. As the Linkages are retracted, the docking cone is
pulled toward the docking adapter main structure and the main _ilieal pl_ is
extended. _en the docking cone bottoms an the adapter pads, the moor_ drive
mecha_ima stops and the RIGID light on the ,_tus cltaplay _li:tllmLt_tes. _he
13-13
SEDR300
PROJECT GEMINI
fizl e_nn__et$on between the 8paeecr_t and the target docking vehicle enables
Joint manewqers vlth the rye 'vehicles.
The rlgidizlng sequence can also be initiated by the pilots or the ground stations
as bask-up to the automatic system. In tl_Ls ease, the pilots have two ways to
rigidize the doekZng eoRe. (1) Position the RTGTD-OI_-STOP switch, on the main
instx_m_nt _n__nel, to I_(]ID. _his by-I_eses the dock lstch. ],trait _rltehes and
applies power to the moori_ dr_ve mo_Or. (2) Send the I_GZDZZE c_m_and to the
target vehicle c_l p_r via the _--_,d Li_.System, _ch routes power
to the moOr_nK d_Te mo_or..The remainder of the seq_nee is the same as when
If the ri_ldize litLt sw£teh .should fail to remove p_r from the moorlng driwe
meter once the docking eerie is ri[idized, :thepilot Can r_oye power by placing
the RIGID-OFF'STOP mrlteh to the _OP peS_tion.
Main r_i:llcsl C_tlon
The main .umbilical p_u8 in the doek_ng adap_r ,as.s_!_:.is sC_u_a b_ the moorin_
drive mee_a._m, In the ri_dized condition, the moori_-driP, seeh_n_sz e_ends
the plu_ out :of tts castn_ :so that it _es ._th the reeep_e_e on the sl_eere_
R & R section. The:p:u_ _Ins extended and m_ted :until the cone is unriEidized.
The maneuvering sequence is sub_ect to considerable va_f_tion except for a few
ec_mon Ope_tlons. Host Of %he maneuvertn8 sequence, is sub_eet, to individual
mission requZz_aents; hoverer, the _ollowing operatlons:are _p_l on most
d_ng m_ss_ons Y_th _ :A_ena Target Yeb_ele (ATV). _.
___ SEOR300 _.__._
PROJECT GEMINI
When the docking cone is rigidized, the pilots can use the Command Link System
to turn off the acquisition lights, the approach lights and the target vehicle
radar transponder. In addition, they will dim the status display lights and send
maneuvering co_ands to the ATV. When the pilots have the ENGIRE switch in the
A_4 position, the target vehicle propulsion systems can be used for attitude
changes, orbit adjustments, etc. At the end of a maneuvering sequence, the
ENGIBE switch is set to STOP and the Co_and Link System is used to return the
target vehicle to the pre-docking status if desired. A co._.._ndvia the Digital
Command System may be substituted for any pilot co_nand if necessary. However,
ground c_nd cannot ove_-_idepilot-swltched functions such as engine arm.
No maneuvering sequence has been considered when docked with the ATDA.
f--.,
SEPARATIONSEQUENCE
As the maneuvering phase ends, the pilots will perform the separation sequence.
Existing conditions at the time will determine whether the mode is normal or
emergency (Figure 13-6).
Normal
The normal separation sequence is initiated by a coamand to the target vehicle
cuwmand progremmer to unrigidize the docking cone. The pilots or ground control
may initiate this cu_.and. Receipt of this co_,nandby the target vehicle c_nd
progr-mmer starts the mooring drive mechanism motor and lights the MSG ACPT light
in the crew station. When the cone is unrigidized, the mooring drive motor stops
and the latch mechanism actuator is energized to retract the latch hooks and
release the spacecraft. The unlatch limit switch, within the latch mechanism,
stops the actuator and starts a 30-second time delay relay. The 30-second time
delay normally allows the pilots sufficient time to back away from the docking cone
13-15
+. -_ SEDR 300
i ":__"'= PROJECT GEMINI
_/ (MOORED POSITION)
LATCH HOOK
LINKAGE
_ LATCH
MOORED AND RIGIDIZED ,'" +,111 -- --
_---CONE
LATCH RECEPTACLE
NORMAL SEPARAIlON _ --_]
EMERGEN &TION _--_/" _ I I_LATCH HOOK
'L _! f"
Figure 1g-6 Separation Sequence
13-16
___ SEDR 300 ___._
PROJECT GEMINIf--,
before the latch hooks are reset to the mooring position. At retrograde adapter
Jettison, the pilots will Jettison the index bar and the docking latches and
release the docking latch covers (Figure 13-2).
An exception to the normal sequence as presented can arise if the pilots should
fail to thrust clear of the docking cone during the 30-second time delay before
the latch hooks are automatically reset. In this case, the rigidize sequence Is
initiated automatically after the latch-reset function. _ pilots would then
have to send the UNMIGIDIZE command a second time. The pilots will then have
another 30-second period in which to thrust free of the docking cone. These
operations may be repeated several times if necessary.
A backup operation to the normal sequence may be initiated by the pilots if the
C_,._andLink System should fail. In this case, the pilots would place the UNDOCK-
OFF switch on the main instrument panel to the URDOCK position. This applies
power to the mooring drive mechanism, which starts the unrigidized sequence. The
remainder of the operation is the same as the normal sequence.
_sergency
If the pilots cannot thrust free of the docking cone by using the normal separation
sequence or if emergency conditions exist, they may employ the emergency
separation sequence (Figure 13-6) to rapidly separate the spacecraft from the target
vehicle. The pilots initiate the emergency separation sequence by pressing an
emergency release switch in the crew station. This switch fires pyrotech-4c
igniters at each of the three docking latch locations, separating the latch
_-_ receptacle from the spacecraft. The pilots then fire thrusters to back out of
the docking cone. The latch covers are released and the index bar Jettisoned at
13-17
___ SEDR 300 ___._j
PROJECT GEMINI
retrograde adapter Jettison. The emergency separation sequence permanently removes
the docki_ capabilities of the spacecraft.
SYST_ _TS
The Docking System ,.n_t_ are those units _t_Lch are concerned solely with docking
and in addition, are not part of another system° Those units which have docking
functions but are _ of other systems are not discussed in detail in this
section. Some -n!ts vhtch are mounted on the R & R section of the spacecraft
are not concerned vith docking at all and therefore are not discussed.
MOORINGLATCEI_Am._
The mooring latch receptacles (Figure 13-2) are fixed ,m4ts, shaped to mate vith
the latch hooks on the docking cone. They are securely but tenporarily fastened
to the main structure of the R & R section. Pyrotechnic igniters are electrically
fired to separate the la_h receptacles loose fram the R & R section. When docking,
the receptacles easily push the latch hooks aside as the spacecraft enters the
docking cone and then let the hooks drop behind the catches to create the latched
condition. The la_h receptacles are protected by the nose fairing during the
boost phase.
MOORINGLATCH CO_ER
Each mooring latch cover asmmbly (Figure 1_-2) consists of a cover, a spring
assembly, a chat% sad latch assembly, a cable and a pyroteeh,4c cable cutter.
shaft, latch and cable keep the spring assembly campressed and the eover
retracted until the cable Is cut by the pyrotechnic cable cutter (guillotine).
The guillotines are electrically fired. As each guillotine cuts it associated
sable, the latch drops dagn, alloying the spring assembly to slide the cover over
the cavity left by the Jettisoned receptacles. The covers prevent v_=rheatl_ of
13-18
PLUG
EPTACLE " FITTING
_BLY
©
? ViEW A-A
SECTION
/i/
/
LATCH
LATCH
SPACECRAFT FREE
INSULATION _ ELECTRICAL EPTACLE ASSEMBLYWIRING
RENDEZVOUS UMBILICAL MATED
CONTACT (TYP)8 PLACES
UMBILICAL MATED
Figure 13-7 Umbilical Connectors
13-19
_@ 5EDR 300 _--_
PROJECT GEMINI
the latch cavities during re-entry.
_BILICAL CONNECTORS
The main ,_ilical connector (Figure 13-7) is a 9 conductor receptacle-plug assembly.
The receptacle is designed to mate with the plug as the Docking System is rigid-
ized. The plug is extended and retracted by the rigidizing mechanism. As the plug
enters the receptacle, it forces the insulator block into the receptacle. This
action allows the receptacle and plug contacts to touch. When the plug is re-
trected, the spring returns the insulator block to the closed position. The in-
sulator block covers the end of the receptacle.
The two rendezvous umbilicals are mated automatically when the R & R section
latches to the TDA. The rendezvous umbilical consists of an electrical slide
contact on the R & R section and a single pin lever assembly on the docking
cone. During the docking sequence, the electrical slide contact depresses the
spring loaded lever assembly slightly providing an electrical connection between
the spacecraft and the _@..
INE_X BAR ASSEMBLY
The index bar assembly (Figure 13-8) includes the bar, a housing, a pyrotechnic
extension mechanism and a pyrotechnic jettison mechanism. The pyrotechnic
igniters are fired electrically from the crew station. The extension igniters
generate a gas pressure which forces the bar up until the socket in which the bar
fits reaches a stop. This is the extended position of the bar. The bar is
jettisoned in the same way, except that the gas pressure forces the bar out of the
socket and ejects it into space. The Jettison igniters also seal off the opening
left by the Jettisoned bar. The jettison igniters have a l-second pyrotechnic
delay to assure that the extend igniters will be fired first.
13-20
.--_. SEDR 300
£XTEND IGNITERS
Figure 13- 8 Index Bar Assembly
13-21
SEOR3OOPROJECT GEMINI
RADAR SYSTEM AND PYROTECHNIC DEVICES
The details of the Rendezvous Radar System units and the Pyrotechnic Devices are
presented in Section VIII and XI respectively. The necessary information to ex-
plalm how these units fit into the Docking System operation is presented elsewhere
in this section.
TARGET V_ICLE STATUS DISPLAY PANEL
The status display panel (Figure 13-3) is mounted on the TDA of the target vehicle.
It is located on the docking adapter Just back of the V-notch in the docking cone.
This gives the pilots a full view of the panel during the approach and docking
sequence. The displays on the panel and their functions when docking with either
the ATV or the ATDA are explained in the following paragraphs.
ATV Displays
(l) DOCK (green light) when illuminated, indicates that the docking cone is
--_igldlzed and that the latch hooks are reset.
(2) RIGID (green light) when ill,_4nated, indicates that the docking cone is
rlgidized.
(3) _ (green llght) when ill,m_nated, indicates that +28 volts dc unregulated,
+28 volts de regulated, -28 volts dc regulated, 115 volts _0 cps single
phase and I15 volts 400 eps 3 phase power are operating.
(4) MAIN (red light) when illmnlnated, indicates the following:
a. With the main engine firing, the turbine has exceeded 27,000 rpm (1046
sensor level), the hydraulic pressure is below 1500 + 20 psi, or the
differential pressure between the fuel and oxidizer tanks is below
2 + 2 psi .....
b. With the main engine not firing, the differential pressure between the
fuel and oxidizer tsars is below 2 + 2 psi (fuel above oxidizer).
13-22
___ SEDR300
PROJECT'-GEMINI
(5) MAIN (green light) when illmminated, indicates that the main fuel ta,_ is
above 15 + 2 psia, the oxidizer tank is abo_ 15 + 2 psia and hydraulic
pressure is above 50 _ 5 psia.
(6) _ (amber light) when ill_mminated,_leates that the engine control cir-
cuits are closed and either the main or secondary engines may be fired by
command •
(7) SEC HI (green light) when illt_Linated,indicates that more than lllO + 20 psi
expulsion gas pressure exists in both nitrogen spheres for a 50 second
thruster firing and that more than ITO + 5 psi regulated pressure exists in
both propellant tank gas manifolds.
(8) SEC LO (green light) when illmzinated, indicates that more than 360 + 20 psi
expulsion gas pressure exists in both nitrogen spheres for a 150 seconds
thruster firing and more than 170 + 5 psi regulated pressure exists in both
propellant _nk gas manifolds.
(9) ATT (green light) when ill,_Inated, indicates that the Agena attitude control
sys_ is active.
(lO) MAIN TIME (clock display) indicates by minute and second hands the time
remaining for gin engine burn. The regulated 28-volt dc power is applied
to the display unlt when the main engine is rn-n_ug. This causes the display
unit to decrease the time re_ining indieation at a rate of one second per
second of burning time.
(ll) SEC TIME (clock display) indicates by minute and second hands the n_sber
of seconds of 200-pound secondary propulslon system burn time remaining.
The regulated 28-volt DC power is applied on separate wires for high and
r- low thrusters of the secondary propulsion system. This causes the display
unit to decrease the time-remainin8 indication at a rate of one second per
13-23
__. SEDR 300
PROJECT GEMINI
second of burn time for the high eons_ption and a rate of one twelfth
(1/12) seeomd per second of burn time for the low eonmmption rates.
(12) ATT GAS (synehro display) indicates the percentage of total pressure re-
maining in the A_ena attitude control system gas spheres.
ATDA Displays
When docking with the ATDA only four displays are used, the others are inoperative.
(1) DOCK(green light) when illuminated, indicates that the docking cone is
unrigidlzed and that the latch hooks are reset.
(2) RIGID (green light) when ill-m_uated, indicates that the docking cone is
rigidized.
(3) A_ (amber light) when ill_insted, indicates that ring A of the reaction
control system has been activated.
(_) ATT (green light) when illmainated, indicates that 0 degree per second rate
control has been selected in all three axes.
13-2_
TARGET DOCKINGADAPTER (TDA)
XIVTABLE OF CONTENTS
TITLE PAGE
SYSTEM DESCRIPTION ................................. 14-3SYSTEM OPERATION ................................... 14-3
DOCKING CONE ........................................... 14-3
DOCKING ADAPTER ..................................... 14-7
SYSTEM UNITS ............................................. 14-9DOCKING CONE .......................................... 14-9
CONE DAMPER ASSEMBLIES ........................... 14-9
APPROACH LIGHTS ....................................... 14-10
MOORING LATCH ASSEMBLIES ....................... 14-10 i:i:i:i:i:!:i:i:!:i:i:i:!:i:i:i:!:i:i:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:TARGET VEHICLE STATUS DISPLAY PANEL ........ 14-12 :::::::::::::::::::::::::::::::::::::::.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.:.
°%%%%%%%°°%°°%%%-.%%..%.
MOORING DRIVE SYSTEM ......... 14-12 ::::::::::::::::::::::::::::::::::::::..................... :::::::::::::::::::::::::::::::::::::- .o° °°.°.°-°°.°°°.-.-°°.-.-.-.-.-.°,
ACQUISITION LIGHTS .................................... 14-13 ::::::::::::::::::::::::::::::::::::::!:_:!:_:_:_:_:_:_:_:i:_:!:!:i:_:!:_:i:RADAR TRANSPONDER AND ANTENNAS ......... 14-14 ..-.-.-...-.-.....-.-.....-...........':':':':':':':':':':':':':':':':':':':
:.:-:.:.:-:-:-:-:.:-:°:.:.:.:-:-:.:.:..-,*,%%°.*.o.,...,°-.-.*.*.-.,.-,**.
STATIC DISCHARGE DEVICE ............................ 14-14 ii!_!?ii!_i_i!i!_!_i!i!i_i_i_?_i.-°,.-.*.-.%%-.%-...-..o..%.°..-...:-:.:.:.:.:-:.:.:.:-:-:-:-:.:.:.:.:.:..-...*.-.-.-.-.%-.-.....-...*...-.,...-,-.-...,.-...*...-,*.-.*.-o-.-.-.,.
. •..., -.-...,...o..-.-...-o-.-.,.• -., *.,.-.,.,.-.-,-.-.-.,.*.*.*.-.*,..-.-.-.-.-.-.-...-.-.*...**...,....-,:-:-:.:.:-:.:°:.:.:.:-:.:.:-:-:-:-:.:..:.:.:.:.:-:.:.:.:-:-:-:.:.:-:.:-:-:.:::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::*.*,***-.-...-.%-,°....,.......::::::::::::::::::::::::::::::::::::::• • *.......,***-...-.*...*...°,*.-.*.• **..-***-.-.-.-***-...*.*.-.-**.,.-.
14-1 :-:-:-:-:-:':':':':':'1":-:-:':-:':':
...--.;_._ SEDR300
,=.. o..,,=c-rCLE .
ANIENNA
IISITIONLIGHT (2 2EQD)
RENDEZVOUS ANTENNA
DISPLAy
LATCH CA_'LE i NDICATORS
ASSEMBLY _
LINKAGEDIPOLE ANT.EXTENSIOMOTOR DRIVE
POWER UNIT
MOORING
LATCH - _.
E_\
DOCKINGADAPTER
TRAN!COVER
APPROACH
(2 REQD)
DAMPERS(7 REOD)
SPIRAL ANIENNA DISCHARGEFINGERS
STATIC DISCHARGE DEVICE
LATCH HOOK -CONE MOORING(3REQD) SURFACE
Figure 14-1 Target Docking Adapter Assembly
14-2
___ SEDR 300
PROJ EC-"T GEMINI
SECTION XIV TARGET DOCKING ADAP_R
SYS_ DESCRIPTION
The Target Docking Adapter (Figure 14-1) may be mounted on either the Agena
Target Vehicle or the Augmented Target Docking Adapter. The Target Docking
Adapter is mounted on the forward equipment section of the target vehicle and
becomes the mounting structure for all other components of the Target Docking
Adapter.
The docking cone receives the spacecraft during docking maneuvers. Electro-
mechanical devices in the spacecraft and the docking adapter rigidlze the
docking cone as soon as the mooring operation is completed. Rigidizing
operations may be automatic or as directed by the pilots or the ground control
stations.
SYSTEM OPERATION
The Target Docking Adapter (TDA) consists of two major subassemblies (Figure
14-1): the docking cone and the docking adapter.
DOCKING CO_E
The docking cone (Figure 14-1) is connected to the docking adapter by seven
cone damper assemblies, three lateral and four longitudinal, and is the
mounting structure for the mooring latch assemblies_ the umbilical plug
assembly, two rendezvous umbilical lever assemblies, the latch release
actuator and a static discharge device. The latch assemblies are released
and reset by the latch release actuator as a part of the unrigidize sequence.
f_ With the target vehicle control circuits operating, the cone unrigidized and
I_-3
S,DR30O_-_,i_ _ _ PROJECT GEMINI
RIGIDIZING
_LINK ASSEMBLY(TYP 3 PLACES)
UMalLICALPLUG ASSEMBLY (TYP 3 PLACES )
DOCKINGLATCHASSEMBLY
ASSEMBLY
ELECTRIC
_CONESTRUCTURE
f
DISCHARGEDEVICE
\
STBUCTURE
LATCH HOOK RETRACTED {DISENGAGED) POSITION LATCH HOOK EXTENDED (SET) POSITION
Figure 14-2 Cone Assembly
14-4
___ SEDR300
PROJECT GEMINI
ready for docking, the latch assemblies are reset. In the reset condition,
the latch assemblies are spring loaded so that when the Rendezvous and Recovery
section enters the docking cone, the latch hooks engage the latch receptacles.
The cone has a V-notch, with a slot at the bottom of the V, cut into the
cone's upper surface. This notch and slot, together with the index bar,
indexes the spacecraft to obtain proper alignment between the spacecraft and
the docking target.
The seven cone damper assemblies are attached in three sets and are equally
spaced at 120-degree intervals around the rear perimeter of the docking cone.
One set, has a single lateral and two longitudinal damper assemblies and is
connected to the cone behind the V-notch. Each of the two lower sets consists
of one lateral and one longitudinal damper assembly. The _-mpers connect
the cone to the adapter structure and absorb the shock of the actual docking
thrusts. They are arranged to absorb thrust from any direction and to compensate
for any misalig_aent of the two vehicles.
Three umbilical connectors provide the connection for electrical circuits
between the spacecraft and the TDA so that harallne c_s may be tran_Itted
from the spacecraft to the TDA.
The main umbilical plug assembly (Figure 14-2) is mounte<1on one of the cone
structural members which supports one of the latch assemblies. In the
rigidlzing operation, the mooring drive mechanism extends the plug out of its
casing so that it mates with the receptacle on the spacecraft. The plug remains
i
14-5
SEDR 300
MAIN
ARMED
[_ NOT OPERATIONAL ON ATDA.
Figure 14-3 Status Display Panel
14-6
SEDR300
extended and mated until the cone is unrigidlzed.
Two rendezvous umbilical lever assemblies (Figure 14-1) are mounted on the
cone assembly and are mated with an electrical contact pad on the spacecraft
when the latch hooks engage the latch receptacles.
DOCKING ADAPTER
The following docking units are mounted on the docking adapter (Figure 14-1):
the mooring drive system, the target vehicle status display panel, the
acquisition lights, the radar transponder, three radar antennas and two
approach lights.
The mooring drive (rigidizlng) system draws the docking cone tightly against
the adapter structure and extends the umbilical plug. The cone is drawn into
the adapter structure by the rigidizing linkages. The linkages are driven by
three gear boxes. The gear boxes are driven by a single electric motor,
through flexible drive cables and an H-drive gear box. The motor is started
automatically and the cone rlgidlzed when the spacecraft is latched into the
cone. The docking adapter Is unrlgidlzed (motor reversed) and the spacecraft
released by a single pilot command through the umbilical connection. Prior
to docking, the unrigidizing process is initiated by rf con_nand (radar-
transponder Command Link or uhf ground radio link).
The target vehicle status display panel (Figure 14-3) is mounted on the docking
adapter structure snd on one cone dsmper, just above the V-notch in the cone. It
visually displays information to the pilot on twelve functions when docking
with the Agena Target Vehicle (ATV). The display consists of nine lights,
14-7
SEDR 300
two clocks and one synchro indicator. _11 power required by the panel is
supplied from the Agena power system. The ATV displays are as follows:
DOCK (green light), RIGID (green light), _ (green light), MAIN (red light),
MAIN (green light), ARJ_D (amber light), SEC HI (green light), SEC LO (green
light), ATT (green light), MAIN T_ (clock display), SEC TIME (clock display)
and ATT GAS (synchro display).
The status display p_nel via-ally displays information to the pilot on four
functions when docking with the Augmented Target Docking Adapter (ATDA). The
display consists of four lights: DOCK (green light), RIGID (green light),
ARMED (ember light) and ATT (green light). Power required by the panel is
supplied from the equil_nentsection of the ATDA.
The two acquisition lights aid the pilots in visual tracking of the ATV.
They produce a flashing light that can be seen for approximately 20 nautical
miles. They are mounted at the outer edges of the adapter structure so that
they are visible to the pilots around the outer edge of the cone. The lights
are mounted so that they rotate outward as the cone is unrigidized. Once the
cone is unrigidized the lights are spring held in the outward position.
The radm_ transponder and the radar antenn-_ (2 spiral and i dipole) are used
for pilot cc_ and tracking of the target vehicle. The ra_-_ transponder
is energized and the dipole antenna (boom) is extended by uhf ground command.
The two approach lights are mounted on the lower inside structure of the adapter.
They are arranged to shine through the rear of the cone and on the upper inside
I_-8
__ SEDR300 __
PROJEC- GEMINIf_
surface near the V-notch. The lights are turned off and on by pilot or ground
command. _en on, they aid the pilot in attaining proper all_,-,,entof the
spacecraft as they approach the docking cone.
sYs
The TDA system units are those units which are concerned with mooring,
rigidizin_ and unrigidizing the dockinz cone to the docking adapter. Units
_Thichare part of the other systems (Radar, Command Link and Docking) are not
discussed in detail in this section.
DOCKING CONE
The docking cone (Figure l_-l and 1_-2) is designed to mate wlth the spacecraft
R & R section. In the moored condition, the spacecraft R & R section flts tight
against the cone bottom and is held there by the mooring latch hooks. The cone
shape tends to guide the spacecraft to the proper mooring position. This action
is aided by a V-notch and its terminating slot. When the spacecraft index bar
is aligned with the slot, the spacecraft and target vehicle are properly
oriented in roll for mooring. The cone is connected to the docking adapter by
three cone d_mper assemblies. A rigidizing linkage links the cone to the
mooring drive (rigidizing) system mounted on the docking adapter main structure.
CONE DAMPER ASSEMBLIES
_o sets of damper assemblies are composed of one lateral and one longitudinal
assembly and one set is composed of one lateral and two longitudinal assemblies
(7 dampers in all). When subjected to an impact, the dampers (Figure 14-1)
f_ compress slowly, absorbing energy as they move. In this _ay, most of the
14-9
SEDR 300
PROJE-'E'-G EMI N I
energy of impact is absorbed by the fluid in the dampers and very little
transmitted to the target vehicle. Springs and gas pressure return the dampers
to their extended positions when the impact is dissipated.
Two approach lights are mounted on the docking adapter (Figure 14-1). They
are positioned so that they shine through the rear opening in the cone to
illuminate the notch during final approach of the spacecraft to the target
docking vehicle. Some light, however, reaches the entire inner surface of
the cone. The pilots turn the lights off and on by using the C<-_.,_,_udLink.
Electric power is supplied by the target vehicle power system.
MOORINGLATCH ASSEMBLIES
There are three mooring latch assemblies (Figure 14-2). Each assembly is mounted
to the cone as shown in the illustration. Each assembly contains a bellcrank
which actuates the mechanism. The bellcranks of the three assemblies are
connected together by a cable assembly. One bellcrank is connected to the
latching actuator by a linkage. When the actuator is extended, the bellcrRnks
are rotated so that the latch assemblies are in the reset condition. This
is the condition which must exist Just prior to mooring. As the spacecraft
slides into the docking cone and contacts the latch hooks, the hooks are pushed
aside to allow the latch catches to get in front of the hooks. As this happens,
the hooks drop in behind the catches (elevated part of the receptacle) and hold
the spacecraft in place. _%e umbilical plug is mounted on one of the cone
structure members which supports a latch assembly. The latch actuator is
electrically driven and either retracts or resets the latch hooks. Power _
14-10
FLEXIgLE D
//
ARM
\
DRIVE SHAFT
ONE
OVERCENTER LINK ASSEMBLY
SPRING
RETRACTED) -_=
OVERCENTE,UNK_ / _'4_ // FI I1"11'._ ""V_.
DRIVEARM- _ I
! INOCONE
c__. _\J_k_________o.,,o,o,z_o,,_E_.OEO,,O_,,,o."---LIMIT SWITCH _MOORING
(CONEEX_ENOEO) SUPROU(RE_I
Figure 14-4 Mooring Drive System
14-11
.__ SEDR300 _l__ "_
PROJECT GEMINI
to the latch actuator is switched by the latched sensing switch, the unlatched
limit switch and the unrigidized limit switch. During the unrigidize sequence,
the actuator is energized automatically to rotate the bellcranks and retract
the latch hooks to the disengaged condition. When the latches are disengaged,
there is a 30-second delay before the latch hooks are automatically reset.
The latched sensing switch is operated only when the latch receptacle is present.
If the latch receptacle is present, the latched sensing switch will initiate
the rigidize sequence when the latch hooks are reset.
TARGET VEHICLE STATUS DISPLAY PAREL
The status display panel (Figure i_-3) is located on the docking adapter Just
back of the V-notch in the docking cone. The notch gives the pilots a full view
of the docking panel during near approach to the docking cone. The displays
on the panel and their functions are presented in detail in the Docking System
Section.
MOORING DRIVE SYS_.I
The mooring drive system (Figure 14-_) is composed of a power unit (DC motor
and gears), four flexible drive shafts, an H-drive, three gear boxes, six limit
switches and a mooring drive linkage. A drive arm is attached to the output
shaft of each gear box.
The power unit supplies power to the gear boxes through the flexible drive
shafts and the H-drive unit. The gear boxes rotate the drive arms which in
turn, retract or extend the rigldizing linkage. When the linkage is retracted,
the docking cone is rigidized to the docking adapter and the spacecraft/target
vehicle combination can be operated as a single unit. The linkage also moves an
ik-12
___ SEDR 300 __
PROJECMINI
actuator which extends the umbilical plug out of its casing. When the linkage
is extended, the umbilical plug is retracted into its casing and the docking
cone is moved aw_y from the adapter so that it is completely supported by the
dampers. The limits switches sense when the linkage is either fully extended
or fully retracted and in each case remove power from the drive motor when the
limit is reached.
The power unit is operated to retract (rigidlze) the system by the actuation of
three sensing switches on the mooring latch assemblies. The sensing switches
energize the mooring drive power unit when the spacecraft is latched to the
docking cone. The linkage can also be retracted by the RIGIDIZE co_nd to
the target vehicle command programmer. Limit switches stop the power unit when
the linkage is fully extended. The same limit switches also apply power to the
unlatch side of the latch release actuator of the mooring latch assemblies.
ACQUISITION LIGHTS
The acquisition lights (Figure lh-l) are used for visual guidance and tracking
of the target vehicle when the vehicles are 20 nautical miles or less apart.
Two lights are provided. They are mounted on the docking adapter and are held
in the retracted position during the boost and insertion phases of a mission.
The docking cone holds the lights in the retracted position until the cone is
unrigidizedby the pilots via the Co_mandLink.
Each light consists of a capacitor discharge flashing light system. The lamp
flashes at a rate of 65 flashes per minute and has a minimum of I00 candles
14-!3
PROJECT GEMINI
effective intensity through an included angle of _90 degrees from the lamp
longitudinal axis. A reflector increases the intensity so that the lamp is
visible from 20 nautical miles with the intensity of a third magnitude star.
The pilots turn the lights off and on via the Command Link,
RADAR TRANSPONDERANDANTENNAS
The radar transponder and antennas are considered part of the Rendezvous Radar
System, the details are presented in Section VIII.
STATIC DISCHARGE DEVICE
A discharge device (Figure l_-l) is mounted on the docking cone to neutralize
the electrostatic potential between the spacecraft and the target vehicle.
This device consists of three flexible metal fingers that protrude from the
docking cone. They make the first contact between the spacecraft and target
vehicle. The device is so wired that any static charge between the two orbiting
vehicles will be dissipated at a controlled rate.
14-14
AUGMENTED TARGET
DOCKING ADAPTER
SeofionXV
TABLE OF CONTENTS
TITLE PAGE
AUGMENTED TARGET DOCKING
ADAPTER-GENERAL ...................................... 15-3SEQUENTIAL SYSTEM ....................................... 15-13ELECTRICAL POWER SYSTEM ........................... 15-27COMMUNICATION SYSTEM ............................ 15-35INSTRUMENTATION SYSTEM ........................... 15-63
PYR OTECHNIC DEVICES :'_iii_i_i:."_i!i_i!i:iiiiiiiiiii_i:iiiA N D SEPA RA TIO N A SS EM BLIES ................... 15- 81 :!:i:i'i:!:i:i._!:i:i:i:i:i:i:i:i:i:i:i
TA RG ET STA BILIZ A TIO N SY STEM .................... 15-95 :::::::::::::::::::::::::::::::::::::
REACTION CONTR OL SYSTEM .......................... 15-1o7 i::iiiiiiiii!i::ili::ii!!iii!i!i::i::ili::ii
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::::::::::::::::::::::::::::::::::::::::::::::::::::::::,_ .%°°%o°%%°°%°°°°°°°°°°%%°°°°%°
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AUGMENTED TARGET DOCKINGADAPTER-GENERAL
TABLE OF CONTENTS
TITLE PAGE
ATDA DESCRIPTION _' 5 5ASCENT SHROUD........... 15-6TARGET DOCKING ADAPTER ....... 15-6EQUIPMENTSECTION. • . . 15-6
__ REACTION CONTROL SY_;TEM'MODI;LE . . . 15-6BATTERY MODULE ........... 15-7
ATDA OPERATION ............. 15-7ATDA SYSTEMS ........ . .... 15-9
DOCKING SYSTEM ....... • • • • 15-9SEQUENTIAL SYSTEM ......... . 15-9ELECTRICAL POWER SYSTEM ....... 15-9COMMUNICATION SYSTEM......... 15-10INSTRUMENTATION SYSTEM ..... 15-10TARGET STABILIZATION SYSTE:M: .... 15-IOREACTION CONTROL SYSTEM ....... 15-10
5LIGHTING SYSTEM ........... I -10
SEO3ooPROJECT GEMINI
AS
139.17
117.00
DOCKING
TARGETDOCKING ADAPTERI I/_241.52 ............ U
1146.05
AS
(TYPICAL6 PLACES)
PLACES) _*EQUIPMENTSECTION
RCSMODULE
-72.00 J
BATTERYMODULE
-102.35
LAUNCH CONFIGURATION
Figure 15-1 Augmented Target Docking Adapter
15-4
___ SEDR300 ___ .___
PROJECT GEMINI
AUGMENTED TARGET DOCKING ADAPTER - GENERAL
ATDA DESCRIPTIONJ ,
The An_,ented Target Docking Adapter (ATDA) is an alternate target vehicle for
the Gemini Spacecraft. The ATIIA may be substituted for the Agena Target Vehicle
provided sufficient mission objectives can be accomplished. The ATDA is capable
of performing all but one of the major functions of the Agena. The ATDA is not
capable of a translational maneuver. It does, however, have small thrusters to
control attitude rates.
The ATDA (Figure 15-1) consists of a Target Docking Adapter (TDA) augmented with
power, cn_mm4cation, instrumentation and stabilization systems. It is made up
/-_ of five major sections or modules. They are: the ascent shroud, the TDA, the
equil_nent section, the Reaction Control System (RCS) module, and the battery
module. Only the equipment section, part of the TDA, and the ascent shroud are
visible when the ATDA is mated with the launch vehicle. A mating ring attached
to the equipment section, mates the ATDA to the launch vehicle.
The ATDA is cylindrical in shape; approximately ii feet long, 5 feet in diameter
at the greatest cross section and 3 feet at the _llest. Although constructed
for the most part of al_num alloy, the ATDA weighs a Little over 2000 pounds
with all equipment aboard. To aid in locating the ATDA during daylight hours,
the equipment section is painted black with white longitudinal stripes. The
other sections and modules are aluminuR color.
15-5
SEDR300 __ ____
PROJECT GEMINI
ASCENT SHROUD
The ascent shroud is a weather tight Jettlsonable fairing, constructed of phenolic
fiberglass. It is 9 feet 9 inches high and 5 feet 5 inches in diameter. The
shroud is used to protect the TDA from aerodynamic pressures and thermal damage
during launch. At the end of the boost phase, the shroud is Jettisoned by a
ground command to expose the docking cone for mission operation.
TARGET DOCKING ADA_
The TDA is an electrmmechanical Imlt constructed to receive the Gemini Spacecraft
and to rigidly dock with it. The TDA is 5 feet in diameter and approximately
4 feet long. The TDA consists of two major assemblies; a docking cone and an
adapter. The docking cone has a V-shaped notch to provide a roll alig_nent path
for the spacecraft during the docking maneuver. The adapter assembly mates the
TDA to the equipment section and is the mounting structure for all TDA components.
A complete description of the TDA is contained in Section XIV of this manual.
EQUII_ENT SECTION
The equipment section is 60 inches in diameter and 54 inches long. This section
contains the largest concentration of the ATDA electronic components. The c_n-
ponents are mounted on internal cross beams. Entry into the equipment section
is provided by four access doors.
REACTION CC_ROL SYSTEM MODULE
The RCS module is approximately 40 inches in diameter and 18 inches long. The
module contains two side-by-slde rings of eight thrust chamber assemblies and
duplicate fuel, oxidizer, and pressurant supplies. The RCS supply tanks are
mounted on the RCS module but extend into the equipment section.
15-6
___ $EDR 300
PROJECT GEMINI
BATTERY MODULE
The battery module contains the three main batteriem and two squib batteries.
The module is 36 inches in diameter and approximately 18 inches long.
ATDA OPERATION
Before the ascent shroud is installed on the ATDA, the docking cone is rigidized.
During pre-launch operations, ground personnel check out the ATDA using ore-board
equilmnentin conjunction with Aerospace Ground Equipment (AGE). AGE power cables
are used to supply power to the ATDA systems during pre-lannch to prevent
depletion of the batteries. The cables are removed and internal power selected
Just prior to lift-off.
The ATDA is boosted by an Atlas Launch Vehicle into a circular orbit. (See Figure
15-2). At the end of the boost phase, the ATDA is separated from the launch
vehicle; the ascent shroud that covers the TDA is Jettisoned; sad all systems
that are used in orbit are turned on or put in readiness to operate.
As the ATDA passes over the launch facility, a Gemini Spacecraft is launched
into an elliptical coplanar orbit. A catch-up maneuver is performed by the
spacecraft. When the spacecraft gets within radar range (appro_mately 250
miles), rendezvous computations are performed and the spacecraft is maneuvered
within 100 feet of the ATDA. At this time relative velocities and attitudes
are adjusted in preparation for docking.
While the docking maneuver is being performed, ATDA systems stabilize the rotation
of the ATDA in all three axes. When the spacecraft enters the doc_ng cone,
f_ the docking latches engage and the rigldize motors are actuated. The spacecraft
is docked and rigidly connected to the ATDA. C_.,._ndsto un_igidize and ,mlatch
15-T
s o.3oo .oJ,:c'r#
_1_ =-_='_r REPARATION ORBIT RENDEZVOUS
_F VERNIER ENGINE CUTOFF
(WCO) +10 SEC
UNDOCKCMD®OO@ "TSS OFF CMD (IF REQD)® C_
SUSTAINER ENGINE UNRIGIDIZE AND UNLATCH(_
(SECO) • ACQ LTS-OFF _)
I2°/SECRATE-DRACnVAT_®@I=/BEC RATE-SELECT(_ (_)
RENDEZVOUS RADAR TRANSPONDER CMD_RIG[DIZE STOP CMD®
/ RIGIDIZECMD@®® ®ARMED LT--ON I(1(1_(IF REQD)
ALL SEPARATION EVENTS RELAyS-DRENERGIZED DOCK LT-ON@DCS CMD INTERLOCK RELAY-_ESET OK TO DOCK
POWER DROP RElAY-ENERGIZED RCS POWER OFF CMD(_D2-RECOND TIME DELAy RELAY-ENERGIZED RENDEZVOUS AND DOCKDOCKING CONE-UNRIGIDIZED_ACQUl Sn'lo N UGHTS-ON
STATUS, RUN AND APPROACH LIGHTS-ONSPiN RATE MODE-SELECTEDRCS POWER-ON
L-RAND BOOM ANTENNA-IEXTENDEO(_TDA ASCENT ANTENNA*OFFSTUBAblTENNA-SELECTED
BOOSTER ENGINE 10 SEC. TIME DELAY RElAy-ENERGIZED C-RAND BRACON-GN
TOFF (RECO} SEPARATION LIMIT SWITCH-ACTIVATED(_ TM-ONLV-ATDA SEPARATION AV-3FPS BY BUNGEE CORDS L-RAND RADAR-ONRCS RING B SQUIB-IGNITED RENDEZVOUS RADAR TRANSPONDER._ENABLED
FLSC SQUIB-IGNITED ACQUISITION LIGHTS-ON
2 SECOND TIME DELAy RELAy-ENERGIZED SQUIB BUS-ARMED
SHROUD EXPLOSIVE BOLT SQUIB-IGNITED RCS RING A SELECTED (IF REQD)(_ _)
I SHROUD JET[ SQUIB FIRE RElAYS-ENERGIZED SQUIB BUS-DISARMED
SQUIB BUSES-ARMED (_) TSS 0°/SECOND RATES (3 AXES)-SELECTEDSQUIB ARM/DISARM RELAyS-ENERGIZED(_) STATUS, BUN AND APPROACH LTS-O NINTERLOCK RELAYS-LATCHED(_ _,DCS SEPARATION INITIATE COMMANDS SENT(_
VERNIER ENGINE CUTOFF (VECO) +10 SEC. PRE RENDEZVOUS
SEPARATION EVENTS I
_t TDA LTS-ON/OFFC-BAND BEACON El OR #2-ON/OFFL-BAND TRANSPONDER-OFF TM P1 OR P2-OI_OFF
BOOST TDA CONE-RIGIDIZED TSS li OR #2-ON/OFFPRIMARY C-RAND BEACON-ON TSS RATES (3 AXES)-MONITORED (VIA TM)TDA BLADE ASCENT ANTENNA-_LECTED RCS POWER-OI_/OFFDCS--ON (USING LAUNCH ANTENNAS) ACQUISITION LIGHTS-ON/OFFRCS HEATER SWITCH-ON TM ANTENNA-SELECT WHIP OR STUBPRI TM-ON (USING IAUNCHAf',ITENNA) RCS RING-SELECT A OR BRCS POWER-OFF RATE CNTL 2°/SEC-ON/OFFRCS RINGS-INACTIVE SQUIB BUS-ARM/DISARM
j SQU[8 BUS-DISARMED L-BAND RADAR-ON/OFF
ALL TDA EQUIPMENT-OFF (UNLESS NOTED) L-RAND CMD-ENABLE/D[SABLEPRI TSS-ON; 2°/SECOND MODERCS RING B -SELECTED ORBIT CMDS (AS REQD)UHF WHIP ANTENNA.EXTENDEDSQUIB BATTERIES-ONMAIN BATTERIES-ON
L-BAND DIPOLE ANTENNA-RETRACTEDACQsAPPROACH , RUN & STATUS DISPLAY LTS-OFF
UFT-OFF
IRCS HEATER SWlTCIt-OFF LEGEND
DCS-ON (USING lAUNCH ANTENNAS) TELEMETERED-REAL TIME! LIFT-OFF AGE CHECKOUT (USING TM, DCS AND LOCAL AGE) DIGITAL COMMAND SYSTEM• j
RF COMMAND LINKHARDLINE UMBILICAL CMD SYSTEM
PIlE-lAUNCH TSS - TARGET STABILIZATION SYSTEMFLSC - FLEXIBLE LINEAR SHAPED CHARGE
Figure 15-2 Sequence Of Events
15-8
___ SEDR300 _____
PROJECT GEMINI
can be given via the hardllne mnbilieal which mates during the docking sequence
or by ground e<-,_,_a_nds.
ATDA SYST_
Eight systems establish the rendezvous and docking capability of the ATDA. These
systems are: Docking, Sequential, Electrical Power, Co,_un_cation, Instrumenta-
tion, Target Stabilization, Reaction Control and Lighting. Sequential functions,
such as ATDA/launeh vehicle separation, ascent shroud separation, and RCS acti-
vation, are performed by pyrotechnic devices and mechanical separation assemblies.
These are described in a separate subsection.
DOCEING SYST_
._ The Docking System is utilized to mate, secure, and rigidize the spacecraft with
the ATDA. A detailed description of the Docking System is contained in Section
XIII •
SEQUENTIAL SYSTEM
The Sequential System prepares the orbiting ATDA for rendezvous and docking with
the Gemini Spacecraft. The system controls the sequence and t_mlng of events
which occur at ATDA/launch vehicle separation.
ELECTRICAL POWER SYST_24
The Electrical Power System consists of the batteries and buses which supply and
distribute dc power to all the systems of the ATDA. The system includes a means
of substituting ground power for the on-board batteries in order to checkout the
ATDA systems before l_unch.
15-9
___ SEDR 300 _._
PROJ GEMINI
C_UNICATION SYS_
The C_._unication System provides the means of tracking the ATDA_ transmitting
information as to the condition and progress of the ATDA. The C_unications
System also accepts switching c,_-_nds from the ground station.
INS_ATION SYS_
The Instrumentation System monitors and collects information on the enviro_ental
conditiona of the ATDA and on the operationsl conditions of its electronic and
electromechanical equipment. This informati_ is assembled in the proper format
and signal conditioned (if necessary) by the Instrumentation System, for trans-
mission to the ground station by telemetry.
TARGETSTABILIZATI0_SYST_ ....
The Target Stabilization System (TBS) monitors the attitude rates of the ATDA
and generates thrusting cuawands to control it at preselected rates.
REACTI_ CC_I_OL SYSTD_
The RCS is the dual rings of eight small thrusters each. These thrusters are
fired by TSS c_-...,.ands.They produce the yaw, pitch, and roll thrust to control
the ATDA at the selected attitude rate.
LIGHTING SYST_4
The Lighting System consists of the acquisition, approach and status display
panel lights on the TDA, and the running lights on the equipment section. A
description of the TDA lights is contained in Sections _III and XlV
Description of Runnin_ Lights _.
The six running lights are mounted on the surface of the A_I)A equipment section.
15-10
..__ SEDR300 _._
PROJECT GEMINI
Two are green, two red, and two amber. Viewing the space vehicle from the TDA
end, the green lights are on the top left, the red on the top right, and the
amber at the bottom. Lights of the same color are arranged on a longitudinal
line with each other. The unfiltered intensity of each lamp is approximately
21 candlepower. The dome-shaped color filters which cover each lamp reduce the
light transmission less than 15 percent. Each lamp and color filter is covered
with a clear protective dome.
Operation of Running Lights
The operational readiness of the running lights is checked during pre-launch.
At lift-off, all lights on the ATDA are off. They remain off during the boost
phase. At the end of the separation events, all TDA lights and the x_mnlngs
lights are turned on together. The lights may be left on until the C_ni
Spacecraft and ATDA rendezvous. If the rendezvous exercise is not carried out
as scheduled or if there is some reason for conserving electrical power, all
ATDA lights can be turned off by a ground c_,--and. Ground station digital
c_.,_._andcan turn these lights on again. After docking, the acquisition lamps
can be turned off by hardline umbilical c_....ands.
15-11/le
SEQUENTIAL SYSTEM
TABLE OF CONTENTS
T i TLE PAGE
SYSTEM DESCRIPTION ........... 15-15SYSTEM OPERATION ...........• 15-15
_ SQUIB BUS ARMING . ..... 15-15' SEPARATION INITIATE COMMAND: .... 15-16
ASCENT SHROUD JETTISON ...... 15-16FIRST 2-SECOND TIME DELAY_ • • • 15-16ATDA-LAUNCH VEHICLE SEPARATiOI_ . . . 15-18SEPARATION SENSING ........ 15-18RCS RING B ACTIVATION_ ....... 15-18TEN-SECOND TIME DELAY........ 15-19RCS POWERAPPLICATION ........ 15-19UHF ANTENNA SELECTION ........ 15-19TRANSPONDER ANTENNA EXTENSION .... 15-19DOCKING CONE UNRIGIDIZING ...... 15-20ACQUISITION LIGHTS ON........ 15_20RUNNING LIGHTS ON......... 15-21STATUS DISPLAY LIGHTS ON : ..... 15-21APPROACHLIGHTS ON ......... 15-22LAST 2-SECOND TIME DELAY . ..... 15-22SEQUENTIAL POWERDROP OUT...... 15-22
SYSTEMUNITS......... 15"23POWER AND SEQUENTIAL I_EI'AyI_ANEI_. . 15-24SYSTEMS CONTROL RELAY PANEL ..... 15-24SEPARATION SENSORS ......... 15-24
15-18
SEDR 300
latch coils of the squib bus arm relays and the separation command initiate relays
to the DCS pulse bus. An open set of contacts on both separation command initiate
relays is also ready to connect the latch coils of the squib bus arm relays to
the common control bus.
When the primary DCS execute on command (channel 18) is sent at VECO plus lO
seconds, the DCS pulse bus is armed. The DCS pulse bus ant the separate command
initiate relays redundantly latch the squib bus arm relays. (See Figure 15-4. )
SEPARATION INITIATE COMWAND
The primary DCS execute on command latches the 1 and 2 DCS command separation
interlock relays, K31-2 and K31-1. K31-2 and K31-1 interlock the Sequential
System in the on condition until a cycle of operation has been completed. K31-2
and K_I-I energize shroud Jettison relays K31-3 and K31-4, separation 2-second
timer relays K31-15 and K31-14, and separation events 10-second time delay relays
K31-5 and K31-6. The ATDA separation events begin to happen with the energizing
of the shroud jettison relays.
ASCENT SHROUD JETTISON
The shroud jettison squib fire relays, K31-3 and K31-4, energize immediately.
Their C, D, E and F contacts connect squib bus power to redundant igniters of
four shroud explosive bolts. Explosive charges cause the bolts to break as if
sheared off. This allows the shroud separation assembly, which is described in
the Pyrotechnics and Separation Assemblies subsection_ to function. The shroud
separates and falls away from the target vehicle.
FIRST 2-SECOND TIME DELAY
Three time delays are built into the ATDA Sequential System. The ti_e delays
15 -16
sEo,3ooPROJECT GEMINI-_-_"
DCS DCS COMMAND RELAY DCS COMMAND RELAY DCS PULSEPULSE BUS CHANNEL 9 BUS
SQUIB SQUIB
BUS I K31-:_ /
41 +iiisK31-3
i"++u°+'°+'K31-3 BOLT IGNITER I-I BOLT IGNITER 2-2
__'L_] BOLTIGNmR2-2 _K31-a K31 K3']_
' ''3SHROUOEXP'OS'VEI1'3SHROUOEXELO'IVEK31-3_rl T I BOLT IGNITER 1-1BOLT IGNITER 2-2
K31-3
_R+S+OODEXPLOSIVEI I"SHROUO°XPLOSIVE+'BOLTIGN,TER2_2BOLT IGNITER I-I _,,K31 _ _,=,-- K31-4
K_I} -9 _ IGNITER 2-1 l [_ATDA SHAPED CHARGE
I ATDASHAPEDCHARGE
1- IGNITER 1-2 --
--q_/'q_ K31_9_D _ ATDA SHAPED CHARGE I [ ATDA SHAPED CHARGE
-- IGNITER 3-I IGNITER 3-2 --
COMMON K31-5 _ INSTR'UMENTATIO N
g TROL B
:<+, I - L__,m,,_K3!:13I® T ®,--t J ATOA/L_V_..._PARAnONSENSORSmTCHE_
_-_.IA}__+-_+T:_ IK:++-:+I--'SQUIB SQUIB
BUS ! KB31-9 :431-10
K31-9 FUELISOL S_ I-I FUEL ISOL SQB 1-2
11-8J© E _ K31 -I0
__ RCS RING B PKG C RCS RING B PKG C
K31-9 OXID ISOL SQB 1-1 OXID ISOL SQB 1-2
RCSRINGBI_GA J I RCSRINGBPKGAPRESS ISOL SQB 1-1 PRESSISOL SQB 1-2
RELAy REDUNDANTRELAY NOMENCLATURE RELAY PANEL
RELAYK31--2 #l DCS COMMAND SEPARATION EVENTS INTERLOCK ATDA POWER & SEQUENTIAL
LATCH COIL OF K31-1 #2 DCS COMMAND SEPARATION EVENTS INTERLOCK ATDA PWR & SQLLATCHING RELAY K31-3 #1 SHROUD JETTISON SQUIB FIRE ATDA PvVR& SQL
K_I-4 #2 SHROUD JETTISON SQUIB FIRE ATDA PWR & SQL(_ K31-5 lI SEPARATION EVENTS 10-SECOND TIME DELAY ATDA PWR & SQL
RESET COIL OF
LATCHING RELAY K31-6 #2 SEPARATION EVENTS 10-SECOND TIME DELAY ATDA PWR & SQL
K31-7 SEPARATION SENSOR POWER DROP ATDA PWE & SQLK31-9 #! ATDA SHAPED CHARge & RCS RING B SQUIB FIRE ATDA PWR & SQL
K31-10 #2ATDASI'IAPEDCHARGE&RCSRING B SQUIBFIRE ATDAPWR&SQLK31-I I R1 SEPARATION POWER DROP 2-SECOND TIME DELAy ATDA PWE & SQL
K31-12 I/2 SEPARATION POWER DROP 2-SECOND TIME DELAY ATDA PWR & SQLK31-13 #1 SEPARATION EVENTS POWER DROP SLOW RELEASE ATDA _ & SQL
K31-8 #2 SEPARATION EVENTS POWER DROP SLOW RELEASE ATDA PWR & SQLK31-15 El SEPARATION 2-SECOND TIMER ATDA PWR & SQL
K31-14 R2 SEPARATION 2-SECOND TIMER ATDA PWR & SQLK32-13 ANTENNA CONTROL ATDA SYSTEMS CONTROLK34-4 RCS RING "B" CONTROL ATDA SYSTEMS CONTROL
K16-2 K16-17 UNRIGIDIZE COM.MAND MOORING SQL CONTROLK35-2 ACQUISITION LIGHTS POWER ATDA SYSTEMS CONTROL
/'_ K35-3 BOOM ANTENNA EXTEND ATDA SYSTEMS CONTROLK36-1 STATUS DISPLAY & RUN LIGHTS POWER ATDA SYSTEMS CONTROLK36-2 RUN LIGHTS POWER ATDA SYSTEMS CONTROL
K36-3 J., APPROACHLIGHTSPOWER ATDASYSTEMSCONTROL
Figure 15-4 Sequential System Schematic
15-17
___ SEDR300 __
PROJECT GEMINI
prevent the violence or rapid occurrence of one event from disruptin_ any other
event. The first time delay is inserted between ascent shroud Jettison and
mating ring separation. This delay is controlled by the separation 2-second
timer relays, K31-14 and K31-15.
ATDA/LAUNCH VEHICLE SEPARATION
Two seconds after ascent shroud Jettison, K31-9 and K31-10 are energized. These
are the ATDA shaped charge and the RCS ring B squib fire relays. Their C and D
contacts connect the squib buses to the shaped charge igniters (or squibs). The
shaped charges explode, severing the mating ring. The mechanical separation
device now performs !_ke a catapult. The bungee cords contract and shove the
ATDA out of the launch vehicle with a separation velocity of 3 feet per second.
SEPARATION SENSING
As the ATDA starts its movement away from the Atlas Launch Vehicle, the blast
shield surface that held the plungers of the separation sensors in the preloaded
position is left behind. The compressed springs in the sensors thrust the plun-
gers downward and the linkages pull the toggles down to the closed positions.
Onl_ two of the three sensor switches need close to complete the sensing circuit.
The closed switches connect common control bus power to the Instrumentation
System Programmer. This bi-level signal is telemetered to the ground monitoring
station as the ATDA separation parameter.
RCS RING B ACTIVATION
At lift-off, the RCS fuel and oxidizer tanks are isolated from the thrusters,
and the nitrogen pressurant is isolated from the fuel and oxidizer ta_ by --
15-18
.__ SEDR300
PROJECT GEMINI
pyrotechnic valves. The B, E and F contacts of relays K31-9 and K31-10 connect
squib bus power to six RCS ring B pyrotechnic valves. The pyrotechnic valves
fire, activating ring B of the RCS.
_N-SECOND TI_ DELAY
The second time delay begins when the execute command to initiate separation is
given. It is timed redundantly by the separation events iO-second time delay
relays, K31-5 and K31-6. When these relays energize, they connect the common
control bus to the separation events relays. The separation events relays
apply RCS power, select the uhf antennas, extend the L-band b_ antenna, turn
on the ATDA lights, and tu_rigidize the docking cone. The B contacts of K31-5
and K31-6 control the last 2-second time delay.
RCS POWER APPLICATION
The RCS latching control relay, K_-4, connects common control bus power to all
RCS solenoid valves. RCS ring A/ring B select relays were set at lift-off, by
DCS command, to select ring B. The Target Stabilization System has been operating
since lift-off, generating rate stabilization commands. These commands now switch
the various RCS ring B thrusters on and off to stabilize the ATDA.
UKF ANTENNA SELECTION
The antenna control relay, K32-13, connects the common control bus to the position
1 input on coax switch 3. This switches the primary telemetry transmitter and
one of the DCS receivers from the uhf ascent blade antenn- to the uhf stub antenna.
TRANSPONDER ANTENNA EXTENSION
The L-band transponder antenna extend relsy, K35-3, connects the common control
bus to the extend terminals of the sntenns drive motor. Within 30 seconds, the
15-19
___ SEDR300 ___ ---1
PROJECT GEMINI
transponder antenna rises to e height of approximately 85 inches and closes the
extend limit switches. One of the switches removes the extension voltage from
the drive motor. The other switch connects the common control bus to the
Instrumentation System programmer. This parameter is telemetered to the ground
monitoring station, confirming that the L-band transponder antenna has been fully
extended.
DOCKING CO_E UNRIGIDIZING
The unrigidize comm_nd relays, K16-2 and K16-17, are latched, energizing the
unrigidize winding of the mooring drive motor and the unlatch winding of the
latch release motor. The mooring drive motor runs,activating three sets of
unrigidized limit switches (one open and two closed). The latch release motor
runs, closing the latch actuator limit switch. The opened set of limit switches -_
causes power to be removed from the drive motor unrigidize winding. One closed
set of limit switches resets the undock circuit. The latch actuator limit
switch and the other set of closed limit switches completes a circuit between
contacts of K36-I and the DOCK light on the status display panel. This circuit
also causes the unrigidize parameter to be applied, through contacts of K16-20,
to the Instrumentation System programmer. This parameter is telemetered to the
ground station for telemetry confirmation of cone position. One set of rigidized
limit switches opens to prevent the RIGID light on the status display panel from
illuml nating.
ACQUISITION LIGHTS ON
The acquisition lights are located on the top left and bottom right of the docking
cone. The acquisition lights, like the docking cone behind which they are mounted,
are held in a rigid retracted position during ascent. When the mooring drive motor
15-20
__ SEDR300 ____
PROJECT GEMINI
unrigidizes the docking cone, the movement causes the lights to move outward
from behind the cone into their operating positions. The acquisition lights
power relay, K35-2, applies common control bus power to the lights.
RUNNING LIGHTS ON
The running lights power relay, K36-2, and status display and running lights
power relay, K36-I,are latched. K36-2 connects main bus power to the red,
green and amber running lights at station -5.00 which is near the Target Dock-
ing Adapter. K36-I applies the same power to the three running lights at
station -49.00 near the RCS module. The running lights can be turned off by
DCS command.
_ STATUS DISPLAY LIGHTS ON
The status display panel on the ATDA is the same panel as the one used on the
Agena Target Vehicle. However, the ATDA uses only four of the status display
lights: DOCK, RIGID, ARMED and ATT (attitude).
The status display and running lights power relay, K36-1, which powered up three
of the running lights, also connects power to the circuits which control the
status display lights. At this time, only the DOCK light illuminates. The
DOCK light illuminates green to indicate that the TDA cone is u_rigidized and
the docking latches are ready to receive an entering spacecraft. The green
RIGID light which indicates a rigidized docking cone, the amber ARMED light
which indicates RCS ring A has been activated, and the green ATT light which
indicates that zero degree per second rates ha_ been selected cannot illuminate
f_ now. These indicators will illuminate when the appropriate function has been
selected by DCS command.
15-21
___ SEDR300 ___
PROJECT GEMINI
APPROACH LIGHTS ON
The approach light power relay, K36-3, connects main bus power to the two
series-wired floodlights inside the TDA. These floodlights, called approach
lights, are located near the bottom of the TDA and illuminate the space behind
the docking cone for a visual approach. The two lights are turned on and off
by the commands which control the running lights.
LAST 2-SECOND TIME DELAY
The third time delay is a 2-second interval between the start of the separa-
tion events and _the removal of Sequential System relay power. This delay is
provided by the separation power drop time delay relays, E31-11 and K31-12.
The B contacts of K31-11 and K31-12 connect squib bus power, switched by K31-1
and K31-2, to the power drop slow release relays.
SEQUENTIAL POWER DROP OUT
Many relays, energized by the ten second time delay relays, are latching relays.
Such relays are latched by permanent magnets in the positions to which they have
been switched. Hence, they no longer need electrical power to maintain their
selected positions. Relays K32-13, K34-4, K35-I, K35-2, K35-3, K36-I, K36-2 and
K36-3 are of this kind. Power which was applied to their coils can be removed
without changing their contact positions.
Besides the latching relays, there are twelve relays, the contacts of which
change when the relays are deenergized. These relays are used to fire the
squibs and to produce time delays. After their functions are completed, power
to their coils also may be removed.
15-22
___ SEDR300 __
PROJECT GEMINI
The power drop sequence is accomplished in seven steps: First, a 2-second time
delay is placed between the application of the separation events signal and the
start of the power drop sequence. The delay permits the separation events relays
to reach a steady state. Second, power drop 2-second time delay relays KSl-ll
and K31-12 close, connecting squib bus power to power drop slow release relays
K31-13 and K31-8. Third, the power drop relays connect the common control bus
to the reset coils of K31-1 and K31-2. Fourth, K31-1 and K31-2 reset, opening
the circuit from the squib buses to KBl-B, K31-4, KBi-5, K31-6, E31-9, K31-lO, K31-
14 and K31-15. Fifth, K31-5 8nd KS1-6 deenergize, disconnecting squib bus power
from eight separation events latching relays and from KSl-ll and I(31-12. Sixth,
KS1-11 and KSl-12 deenergize, disconnecting squib bus power from K31-1S and
f- K31-8, the slow release relays. Seventh, in approximately 40 to 70 milliseconds,
K31-13 and K31-8 deenergize, disconnecting the common control bus from the
reset coils of KBl-1 and KSI-2. This step completes the power drop-out sequence.
SYSTEM UNITS
The units of the Sequential System are the relays, the relay panels and sensor
switches which control or monitor the separation events. At] of the Sequential
System relays are located on tilepower and sequential relay panel and the systems
control panel.
Although some unrigidizing relays are operated by the separation events signal,
the relays and their relay panels are properly part of the Docking System. The
sensor switches which monitor separation are discussed here, the pyrotechnic
devices which cause separation and some of the separation events are discussed
in the Pyrotechnics and Separation Assemblies subsection.
15-23
_._.._@ SEDR 300 ___ ___
PROJECT GEMINI
POWER AND SEQUENTIAL RELAY PANEL
The power and sequential relay panel (Figure 15-3) is approximately 8.7 inches
hlgh by 9.8 inches wide by 3.5 inches thick. This panel is located in the
equipment section of the ATDA and provides a mounting structure for fifteen
sequential relays. These relays and their functions are identified in Figure
15-4. Seven connectors provide electrical access to the relay solenoids and
contacts. Relays mounted on this panel are identified by a K31 series reference
designation.
SYSTE_ CONTROL RELAY PANEL
The systems control relay panel (Figure 15-3) is approximately 7.5 inches high
by ll.5 inches wide by 3.5 inches thick. This panel is located in the equipment
section of the ATDA and provides a mounting structure for eight sequential relays.
Seven of these relays and their functions are identified in Figure I_-_. The
remaining relay, KSS-1 is used to complete a telemetry path, confirming when the
docking cone is unrigidized. In addition to the sequential relays, eleven system
control relays are located on this panel. These relays are used to select secon-
dary systems or different modes for the various ATDA systems. Six connectors
provide electrical access to relay solenoids and contacts. Relays mounted on
this panel are identified by a KS2 through K36 series reference designation.
SEPARATION SENSORS
The separation sensors are electromechanical devices used to detect the separation
of the ATDA from the launch vehicle. The electrical part of the sensor is a
double-pole double-throw toggle switch. 0nly switch 2 uses both poles. (See
Figure 15-4. ) ....
15 -24
___ SEDR300
PROJECT GEMINIf \
The mechanical part of the sensor is a bracket, on which the switch is mounted,
and a linkage designed to operate the toggle. (See Figure 15-3. ) The linkage
is basically a plunger or piston confined in a housing. Shafts attached to both
faces of the plunger extend out the top and bottom of the housing. The bottom
shaft, which is narrow, controls the position of the plunger. The top shaft, which
is thicker, has an oblong hole, somewhat larger than a toggle switch handle, machin-
ed into it. A compression spring fits over the top shaft inside the housing and
holds the plunger against the bottom of the housing. The toggle of the switch
extends into the top shaft hole which pulls the toggle into the down or closed
position.
Three separation sensors are mounted in the end of the equipment adapter where it
interfaces with the launch vehicle mating ring. Each sensor fits through a hole
cut in the end of the equipment section, so that the control shaft extends into
the mating ring. A heat and shrapnel protective shield surrounds the inside of
the mating ring and forms a trough. The sensor shafts rest against the bottom
of this trough. When the vehicles are mated, the trough pushes the shafts and
plungers upward in the housing, compressing the springs and moving the toggles
to the up position. These are the launch positions of the separation sensors.
When the shaped charge severs the mating ring, the trough stays with the launch
vehicle and the sensors leave with the A_DA. The spring loaded sensor switches
now complete the path from the co_on control bus to the Instrumentation System.
f_
5-25/26
ELECTRICAL POWER SYSTEM
TABLE OF CONTENTS
TITLE PAGE
SYSTEM DESCRIPTION ........... 15-29SYSTEM OPERATION ............ 15-31
,_ PRE-LAUNCH ............. 15-31ORBIT. ......... - . - - - - 15-31svs, om, s..............SILVER-ZINC BATTERIE . i;i_i.:s • • • 15-32CONTROL AND MONITORING A E • • • 15-33D,OOEPANEL.. • 15-33POWER AND SEQUEI_TIAI"I_ELAY I_AICIEI_: . 15-33
15 -27
_ SEOR3OO I__--'_PROJECT GEMINI
) SEQUENTIALRELAY PANEL
MONITORING
PANEL
EQU|PMENI"
SECTION(REF.) SQUIBSILVER-ZINCBATTER_ BATTERIES
MODULE (2 REQ*D)
MAIN SILVER-ZINCBA'r'fERIES(3 REQ'D)
MAIN BATTERY SQUIB BA11|KY I IL_S h_i_j
_ _ _ _ _ lPOWER CONTROl. SWITCHES
Figure 15-5 Electrical Power System Installation
15-28
$EDR 300
P ROJ--JE"C"T GEMINI
ELECTRICAL POWER SYSTEM
SYS_.M _SCRIPTION
The Electrical Power System for the Augmented Target Docking Adapter (ATDA)
basically consists of three silver-zinc main batteries, two silver-zinc squib
batteries, a control and monitoring panel, a diode panel and relays for con-
trolling squib bus power. Refer to Figures 15-5 and 15-6.
The three main batteries provide dc power to the ATDA main power bus. The two
squib batteries provide dc power to the common control bus and the two squib
power buses which are isolated from the _n and co.on control bus. The con-
trol and monitoring panel provides switching for the power system and the
capability for utilizing Aerospace Ground Equipment (AGE) external dc power and
remote monitoring of the power system parameters.
The diode panel provides electrical isolation between the two squib batteries
for individual battery fault protection. All of the power system circuits are
of a redundant nature. No primary ac electrical power system is provided for
the ATDA. Devices requiring ac power will obtain this power from self-contained
inverters within the individual systems.
Both squib batteries supply power to the common control bus. Squib battery i
and squib battery 2 are separately connected to squib bus i and 2 respectively
via the squib bus arm relays. These relays are controlled by the Digital Co-_and
System (DCS).
15-29
_-_. SEDR300
,-RoJcTMAINBATTERIES MAIN BATTERY
SWITCHES
ON MAIN BUS
o "-_
MI OFF OVOLTAGE OMONITOR ORD PWR
T/M
ON
---'-----------0
CURRENT{ PWR iMONITOR 1
M2 {VOLTAGE
MONITOR j
{ _M3 OFF O_VOLTAGEMONITOR PWR " "
MAIN I
GRD PWR
DCS PULSE BUS IS CONNECTEDZ
TO COMMON CONTROL BUS
VIA DCS.
DCS PULSE BUS --_.
o_. _ T/M
_ --"-1Z_GOL_ _C SQUIB BUS
CONTROL
MOD'_ 1 '_ SQUIBCHANNEL 1
BUS 1
RESET
GRD PWR #I SQUIBARM RELAy
{ sou,Sl PWR r_ BUS 2
VOLTAGE AG-ZN liJMONITOR DFF"---O I_1
$I ---OON
K31-I
#2 DCS CMD SEP
INTERLOCK RELAySQUIB
SQUIB BATTERY T/M S •
- "1BATTERIES_ SWITCHES I mR2SQUIB"_ T/M
GRD PWR ARM RELAy
AG-ZN C _ _
VOLIAGE POWER & SEQUENTIALMONITOR RELAY PANEL
Y
Figure 15-6ElectricalPower System Schematic
15-30
___ SEDR300
PROJEC-T GEMINI
The three main batteries and two squib batteries are installed in the battery
module, which is located in the opposite end of the ATDA from the docking
adapter. The control and monitoring panel, diode panel and power and sequential
relay panel (containing the squib bus arm relays) are located in the equipment
section of the AT_A.
SYSteM OPERATION
PHE-LAUNCH
In order to conserve the ATDA batteries, AGE external dc electrical power is
utilized during pre-launch checkout of the ATDA systems. Exte _r__! power is
supplied to the ATDA through AGE cables connected to the control ..R monitoring
panel receptacles. Remote monitoring of main bus current, individual main battery
voltage and individual squib battery voltage is also accomplished through the AGE
cables.
External power is applied to the main power bus and c_u control bus by setting
the MAIN BATTERY and SQUIB _ switches to the GRD _ position. Just prior
to launch, all battery switches are set to ON position and the AGE cables are
removed from the ATDA. The squib power buses are not armed prior to launch.
ORBIT
The squib power buses are armed via the two squib arm rel_Vs Just prior to
orbital insertion. The relays are energized to latch position by a con_Rud from
the DCS. Ccmnon control bus voltage, squib bus 1 and 2 voltage and nmln bus
voltage and current are monitored by the Instrumentation System.
After the required squib functions for orbital insertion are accomplished, the
_5-31
__ SEDR 300 _._._
PROJECT GEMINI
squib buses are disarmed by a command from the DCS and may be rearmed if
required. The squib buses are disarmed prior to rendezvous of the spacecraft
with the A_DA.
SYSTEM _SITS| ,a , ,,
SILVER-EriC BATSERIES
The three ma!n batteries are 400 ampere/hour, 16 cell, silver-zinc batteries and
are identical to the adapter power supply batteries used on spacecraft 6. The
two squib batteries are 15 ampere/hour, 16 cell, silver-zinc batteries and are
identical to the squib batteries used in the spacecraft.
The main battery cases are constructed of magnesium. The approximate activated
(wet) weight of each main battery is 118 Ibs. The squib battery cases are
constructed of titanium. The approximate wet weight of each squib battery is 8
ibs. The squib batteries are special high-discharge-rate batteries which will
maintain a terminal voltage of 18 volts for one second under a 75 Ampere load.
All of the silver-zinc batteries have an open circuit terminal voltage of 28.8 to
29.9 volts.
The battery electrolyte consists of a 70 percent solution of reagent grade pot-
assium hydroxide and distilled water. The squib batteries have a vent valve in
each cell designed to prevent electrolyte loss. The valve will vent the cell to
atmospheric pressure in the event that a pressure in excess of 40 psig builds up
within the cell.
All of the silver-zinc batteries are equipped with relief valves which maintain a
15-32
3ooPROJECT GEMINI
tolerable interior to exterior differential pressure in the battery cases. The
batteries are capable of operating in any attitude in a weightless state. Prior
to installation into the battery module, the batteries are activated and sealed
at sea level pressure.
CONTROL AND MONITORING PANEL
The control and monitoring panel contains the main and squib battery switches
and the Reaction Control System (RCS) heater switch. This panel also provides
receptacles for connecting AGE external power and test cables to the ATDA.
DIODE PANEL
The diode panel contains diodes required to provide individual fault protection
for the squib batteries. The squib batteries are connected to the common control
bus via these blocking diodes to provide battery isolation.
POWER AND SEQUENTIAL RELAY PANEL
The power and sequential relay panel, which is essentially part of the Sequential
System, contains the squib bus arm relays in addition to the other sequencing
and control relays. The squib arm relays are controlled by the DCS and redundantly
by the Sequential System.
15-3313_
f--\
COMMUNICATION SYSTEM
TABLE OF CONTENTS
T I TLE PAGE
SYSTEM DESCRIPTION ........... 15-37ANTENNAS 15-37
f_" BEACONS. : : : : : : : ! ! ! i i ! ! 15-3'TELEMETRY TRANSMITTERS 1 5"39L DIGITAL COHHAND SYSTEM ....... 15"39
SYSTEM OPERATION ....... . . . . 15"39RADAR TRACKING ........... 15-39DIGITAL COMHAND........... 15-4_TELEHETRY .............. 15 -4_
SYSTEM UNITS .............. 15-51ANTENNAS .............. 15-51BEACONS.............. 1 5-57TELEMETRY TRANSMITTERS _ ...... 15-57DIGITAL COMMANDSYSTEM ....... 15-59
15 -35
____ PROJECT GEMINI
ASCENTANTENNA
RADARBEACONS
-DIPLEXERS
C-BANDANTENNA
" " TELEMETRYTRANSMITTERS (2)
/" SWITCH (3)
ANTENNA
DCS RELAY
/ MODULES (3)
.f-,/
• .///"
] // ,
/ _;y .."/ /[ ./
[/ ..."
f -'/ r""C-_'N0 1 il 'POWER DIVIDE_S \ :';12) , , _ /
C-_NOANTENNA
DCSRECEIVER-DECODER
UHF UHFWHIP STUBANTENNA ANTENNA
Figure 15-7 Communication System
15-36
_@ SEDR 300 _.___j
PROJECT GEMINI
f_
CC_@gJ_CATI ON SYSTEMHJ ,,,,
SY_ DESCRIPTION
The C_-....,n_cation System is the only c_unication l_nk between the ground and
the Augmented Target Docking Adapter (ATDA). The C_--Ication System provides
the following capabilities: radar tracking of the ATDA, ground C,:--..AUdto the
ATDA, and telemetry transmission. To make possible these various capabilities,
the C.-,..._mlcation System components may be divided into the following categories:
antennas, including diplexers and coaxial switches; beacons; telemetry trans-
mitters; and a Digital C.._aud System (DCS).
The C_-.Im_cation System components are located throughout the ATDA with the
largest concentration being in the equipment section. The location of thef_
C_m._mieation System components is illustrated in Figure 15-7.
Three uhf antennas and two sets of C-band helical antennas provide transmission
_nd/or reception capabilities for the various C_unication System components.
The Co_.unication System (Figure 15-8) contains the following antennas: uhf
stub, uhf whip, uhf ascent, and two sets of C-band helical ante_n-s. Each set
of C-band antennas consist of a power divider and three helical antennas.
Three coaxial switches permit antenna and transmitter/receiver switching for best
DCS and _elemetry e.....unication with the ground stations during the various phases
of the mission (pre-launeh, launch, orbit, and docking).
15-37
- _-_.. SEDR 300
__i__ PROJECT GEMINI
> i rmnc__o.> i ,o_Bi i c-_.oIB,ACON
l A PWR
_- I C _HANNEL -ANTENNA PRIMARy
i ,,,..._1
-C-BANDBEACON ! _ r DCS_I • PWR
C-BAND _ SECONDARY I ' C_
AN_NNA_ _'-- ..I
DIPLEXER COAXIA IN T/M J
,2 Bw,,cH BE,_>N_I,_J
I il _ilI
ISEP. +8 SEC.) I ,A_
U o_t K32-13 E WHIP/STUB J C PWRAN'_NNA I'=.__..ICONTROL ANTENNAELAY ELAy
i I
(iAL IN T/M I
UHF WHIP r_ I DIPLEXER iAN_NNA _ _ II
/i i I I n 1 1
DCS RECEIVER-DECODER "lll
Figure 15-8Communication System Block Diagram
15-38
___ SEDR 300
PROJECT GEMINI
BEACONS
Two C-band radar beacons establish the capability of tracking the ATDA during
the mission. The beacons are transponders which, when properly interrogated by
the ground station, transmit signals for accurate ATDA tracking. The radar
beacons are controlled by DCS cow,ands frc_ the ground stations.
TRT._ETRY TRAN_I T_RS
Two identical transmitters supply the radio frequency link from the ATDA to the
ground for transmission of instrumentation data. During pre-launch the telemetry
transmitters are used for pad checkout of the ATDA. During orbit, telemetry
transmission is made while the ATDA is in range of a ground station.
_ DIGITAL CQMMAND SYST_f
The DCS is the c_auication link for ground c@.,_-audsto the ATDA. The DCS
receives and decodes c.--,,-_dtransmissions from the ground stations. These
c.!N_.*._ndsare used to operate relays which control the operation of various ATDA
equipment.
The IES consists of a receiver-decoder and three relay units located in the equip-
ment section. The DCS operates from pre-launch throu@hout the mission.
SYST_ OPERATION
The AT_A C_mm,m_ cation System operates from pre-launch throughout the _ission.
The sequence and theory of operation of the C_u_xication System is as described
in the following paragraphs and as illustrated in Figures 15-7 and 15-8.
RADAR TRACKINGS _
_adar tracking of the ATDA is accomplished by the use of two C-hand radar
beacons. Complete redundancy of the tracking capability is available by using
15-39
__ SEDR 300 _j___
PROJECT GEMINI
two beacons_ each controlled by a DCS channel and operating with its own antennas.
The radar beacons are transponders which, upon reception of a properly coded
interrogation signal from the ground station_ transmit a pulse modulated return
signal. The location of the ATDA is determined by measuring the elapsed time
between transmission and reception at the tracking station, conpensating for the
known time delay of the beacon.
At lift-off, the primary C-band beacon is activated via channel 14 of the DCS.
The pri_y beacon is used for tracking during the launch phase and orbital
insertion. The DCS is used to control the operation of the radar beacons during
the reminder of the mission. The secondary beacon can be selected via DCS
channel 17 should it be needed as a result of pr_ beacon malfunction.
The ground interrogation signal is coupled from the antenna to the receiver via
a ferrite circulator (Figure 15-9). The circulator isolates the receiver from
the tr_-m4tter to permit the use of a c_on antenna for reception and trans-
mission. The rf filter is a three stage preselector, employing three separately
tuned resonator cavities which gives adequate rf selectivity and protects the
mi_er crystal fr_ damage by the reflected tranm4 tter power from the antenna.
The output of the preselector is c rmbined with the local oscillator output in
the m1_er. The mix_-.r consists of a coaxial directional coupler and a crystal.
The directio-a1_cou_ler isolates the local oscillator output from the antenna and
directs it to the mi_er crystal. The local oscillator is a re-entrant cavity type
which uses a planar triode to generate the signal required to operate the mixer.
15-_0
_-=,_ SEDR300
_ I _ _I_° IJ2 DUPLEXER TRANSMITTER TRANSFORMER &
_ _ _ PULSE FORMING MODULATOR J
(CIRCULATOR) (MAG NETRON) I NETWORK (PEN)
i 1 Jt lRF FILTRR(PRESELECTOR)
TUNING
ADJ. $
/..._ MIXER AMPLIFIER RESTORER CONTROL
t , I 1LOCAL _ I DECODER MODULATOR
I OSCI LLATOR _ DRIVERII
I l I I ' ICONVERTER
I I
I REGULATOR AND ILINE FILTERS
i I
!24.5 TO 30V OC POWER
Figure 15-9 C-Band Radar Transponder Block Diagram
15-41
.@ SEDR300 _._
PROJECT GEMINI
The intermediate frequency amplifier is a high gain amplifier composed of an input
stage, five amplifier stagesj and a video amplifier. The amplified video output
is fed to the pulse form restorer circuits which prevent a ranging error due to
variations in receiver input signal levels, and also provides a standard amplitude
pulse to the decoder for each input signal exceeding its triggering threshold.
The decoder determines when a correctly ceded signal is received and supplies an
output to the modulator driver. The type code to be accepted is selected by the
CODE switch. Single pulse, two pulse, or three pulse codes may be selected. The
modulator driver and control circuits initiate and control triggering of the trans-
mitter modulator. The modulator driver supplies two fixed values of overall system
delay. The desired delay is selected by the position of the DLY switch. The
modulator control furnishes the trigger and turn-off pulse for the modulator
and limits modulator triggers to prevent the magnetron duty cycle from being
exceeded, regardless of the interrogating signal frequency. The modulator circuit
employs silicon controlled rectifiers which function s_m_lar to a thyratron, but
require a much shorter recovery time.
The associated modulator pulse forming network and transformer provide the
necessary pulse to drive the transmitter magnetron. The desired pulse width is
selected by the internal connections made to the pulse forming network.
The transponder power supply consists of input line filters, a series regula-
tor, and a de-de converter. The power supply furnishes the required regulated
output voltages with the unregulated input voltage between 21 and 30 vdc. The
converter employs a multivibrator and full wave rectifier circuits.
15-_e
__ SEDR300
PROJE--C"T'--GEMINI
The output of the transmitter is applied to the power divider via the circulator.
The power divider gives equsl transmission power to the three helical antenna
radiating elements.
DIGITAL CC_¢_ND
The DCS provides a discrete link between ground cc_nd and the ATDA. _e
discrete link enables the ground command to control the operatio_ of the various
ATDA systems. The c_nd transmissions from the ground are received and decoded
by the DCS and are used to operate DCS relays that control power directly or
energize relays ih the ATDA that determine equipment usage. During pre-launch
the DCS is used for pad checkout of the ATDA systems. Discrete c.,_._de are
sent to the ATDA via the DCS and their execution is verified from instrmnentation
data.
The uhf ascent and the uhf whip antennas are used for the reception of DCS
ground cu_nds during pre-launch and lift-off through ATDA separation. At
separation coaxial switch 3 changes DCS reception from the uhf ascent antenna
to the uhf stub antenna. The two diplexers permit the use of a c_--,_onantenna
for reception of DCS cu_.._andsand telemetry transmissions.
The DCS consists of a receiver-decoder and three relay ,mlts. The receiver-
decoder contains two uhf receivers and the decoder. The two receivers are
redundant, and the system will operate properly if only one receiver is function-
ing. Each of the three DCS relay units contains eight magnetically latched
relays. Each relay is set or reset by a DCS cuK_and.
The DCS receives phase shift keyed frequency modulated signals composed of a
reference and an infol_nation signal. The information signal is in phase with
15-43
SEDR 300
the reference for a logical I and 180 degrees out of phase with the reference
for a logical O, thus establishing the necessary requirements for digital data.
The ground e,-_v_nd to the DCS is a 12-bit message. Each bit consists of five
sub-bits. The five sub-bits are coded to represent a logical i or O. The
first three bits of each message is the vehicle address. The second three bits
of the message is the system address specifying real time c,_!_and. The last six
bits in the message contain a five-bit relay number and a one-bit relay set-
reset discrete c_=,,,_. Table 15-i lists the DCS channels and contains a brief
description of their function.
The DCS uses two types of c_-ds. One type cn-_and is executed the mument it
is received. _-_s type c_-,-_nd operates DCS ch_-nel relays which directly control
power to the selected ATDA system. Power is applied in the set condition and
removed in the reset condition.
The other type c_ is received but not immediately executed. These co_a_ds
operate DCS channel relays that connect the control circuits of selected ATDA
systems to a pulse bus. The c.-_,_udsare executed when the pulse bus is energized
by an execution c_..-_nd. The channels marked by an asterisk in Table 15-1 indicate
the DCS c....._nds which require _- execution c,.,_i_-d.
A block diagram of the DOS recelver-decoder is shown in Figure 15-10. Basically,
the block diagrsm consists of a receiver, a decoder, and a power supply c--,mon
to both sections.
15-_h
.-_ SEDR 300
DCS COMMANDS
CHANNEL FUNCTION
'1 ARM AND DISARM BUS FOR SQUIBPOWER
2 SELECTSPIRAL ANTENNA FOR L-BAND TRANSPONDER
*3 COMMAND UNRIGIDIZE AND UNLATCH
*4 COMMAND R|GIDIZATION OF TDA CONE WHEN LIMIT SWITCHES CLOSE
*5 TURN ACQUISITION LIGHTS ON OR OFF
*6 TURN PRIMARY CONVERTER AND TRANSMITTER ON OR OFF
*7 TURN PRIMARY STABILIZATION SYSTEM ON OR OFF
*8 SELECTRCS RING A OR B
*9 COMMAND SEPARATION SEQUENCE
"10 TURN SECONDARY CONVERTER AND TRANSMITTER ON OR OFF
"11 TURN SECONDARY STABILIZATION SYSTEM ON OR OFF
"12 TURN RCS RING POWER ON OR OFF
"13 SELECT BIASED RATE DAMPING OR NORMAL RATE DAMPING TSS MODE
14 TURN PRIMARY C-BAND BEACON ON OR OFF
*'15 EXECUTE PRESELECTEDDCS COMMANDS
16 PRIMARY OR SECONDARY DCS EXECUTE SELECT
17 TURN SECONDARY C-BAND BEACON ON OR OFF
*'18 EXECUTE PRESELECTEDDCS COMMANDS
"19 TURN L-BAND TRANSPONDER ON OR OFF
*20 CONTROL SELECTIONOF WHIP OR STUB ANTENNA
"21 COMMAND SEPARATION SEQUENCE
22 ENABLE OR DISABLE L-BAND TRANSPONDER
23 TURN OFF TDA LIGHTS WITH THE EXCEPTION OF ACQUISITION LIGHTS
24 NOT ASSIGNED
* INDICATES DCS COMMANDS THAT REQUIRE AN EXECUTION COMMAND
** INDICATES DCS EXECUTION COMMAND
Table 15-1
15-45
;_> SEDR300 ......_'1"_ _
PROJECT GEMINI
Y YSIGNAL STRENGTH RECEIVER RECEIVER := SIGNAL STRENGTHTO TELEMETRY _ El /?2 TO TELEMETRY
' I
I I K'EcISUB-BIT SUB-BITS ." 5-STAGE SUB-BITS _. ADDRESSSTORAGE BIT DETECTOR DECOOERDETECTOR REGISTER
TIMINGERR___.__ j |INHIBIT j _<
ADDRESS
I DATA i
i sit Ii 24-STAGE RELAY NO. RELAY
SUB-BIT SYNC
SUB-BIT SETSYNCHRONIZER : _ _ STORAGE ) SELECTION
FAILURE - AND RESET J REGISTER
I
J RELAy DRIVE l 1TO RELAYS
:I
°1 IPOWER PROGRAM VALIDITY _ INTERFACECONTROL
SUPPLY READY
20-30V DC TELEMETRYINPUT pOWER VERIFICATt ON
NOTE1. HEAVY LINES DENOTE DATA FLOW.
Figure 15-10 Digital Command System Block Diagram
15-46
___ SEDR 300 __
PROJECT GEMINIf--
The audio outputs of the two receivers are linearly s_.,,-,_din an emitter follower
of the sub-bit detector module. The sub-bit detector converts the audlo to sub-
bits. The 5-stage shift register provides buffer storage for the output of the
sub-bit detector. The states of the five stages of the shift register repre-
sent the sub-bit code. When a proper sub-bit code exists in the shift register,
the bit detector produces a corresponding I or 0 bit. The output of the bit
detector is applied to the 24-stage shift register.
The sub-bit synchronizer counter produces a synchronizing bit output for every
five sub-bits. The synchronizing bit is used to gate the 2_-stage shift register.
When a message is received, the vehicle address is inserted into the first three
stages of the 24-stage shift register. If the vehicle address is correct, the
vehicle address decoder circuit will produce an output to the bit detector which
changes the acceptable sub-bit code for the r_._inder of the message. The next
three bits of the message, the system address, are inserted into the first three
stages of the 2_-stage shift register, displacing the vehicle address to the next
three stages. The system address decoder circuit identifies the specific address
and sets up the I_S to handle the remainder of the message.
When the system address is a real time e,:-,,_,_a.d,the message is inserted into the
first six stages of the 2_-stage shift register Rr_ the system address and v_h_cle
address are shifted into the next six stages. The real time c_-_ selection
circuit recognizes the first stage of the 2_-stage shift register to be
a relay set or reset function and will apply a positive voltage to set or reset
all relay coils, as applicable. The real time c.-,,_Indselection gates selectf_
the proper relay from the relay n_m_oerstored in the 2_-stage shift register and
15-h7
__ SEDR300 __
PROJECT GEMINI
provides an output which applies power to the coil of the selected relay.
Upon completion of data transfer or if the system to which the data was transferred
fails to respond within I00 milllseconds, the DCS will reset in preparation for
the next message. The DCS will also reset in the event of a timing error in
transmission of data, or if the DCS power supply voltages become out of tolerance.
The DCS power supply operates from the ATDA main power bus and supplies the
receivers and decoder with regulated dc voltages.
A verification signal is supplied by the decoder for telemetry transmission when
a ground command has been received.
T_._RY
Transmission of instrumentation data is accomplished by the telemetry transmitters.
Operation of the transmltters is controlled by the ground station via DCS channels
6 and i0. An interlock in the ATDA prevents both transmitters from operating at
the same time.
Coaxial switches select the antenna used for telemetry transmission. The uhf
ascent and uhf whip antennas are used during pre-launch checkout and lift-off
through ATDA separation. After separation the output connection of coaxial
switch 3 is switched from the uhf ascent antenna to the ,,bf stub antenna.
Data from the Instr=_entation System progr_er is supplied to the telemetry
transmitters in non-return-to-zero pulse code modulated pulse trains. The
transmitters relay the information to the ground stations in a digital format
composed of l's and O's at a rate of 51.2 kiloblts per second. The carrier
15-48
___ SEDR 300
PROJ EC--'T GEMINI
frequency is deviated to the higher frequency deviation limit in order to
transmit a i and to the lower deviation limit to transmit a O. Figure
15-II is a block diagram of the telemetry transmitters. The transmitters
operate on the same frequency and are identical in operation.
:, The input signal is amplified by the video amplifier and is used to modulate the
Output of the oscillator. The oscillator is crystal controlled for good frequency
stability. The modulated signal is passed through a series of buffer amplifiers
and tw0"phase shifters. The buffer amplifiers increase the signal level and iso-
late the crystal circuit from the frequency multipliers. The phase shifters
provide impedance matching of the crystal oscillator to improve signal linearity
for large deviations in frequency.
A times-four multiplier, a power amplifier, and a times-three multiplier increase
the carrier frequency and power to the desired output values. The bandpass fil-
ter minimizes spurious radiations at the output of the transmitter.
The line filter prevents noise on the input power bus frc_ affecting transmitter
operation and prevents transients generated within the transmitter from feeding
back to the input power bus.
The transmitter dc-dc converter is a c_npletely encapsulated unit employing tran-
sistors, diodes and a transformer to provide regulated outputs of 30 vdc and 70
vdc from ,,nregulatedinput voltage of 18 to 30.5 vdc. The converter is a constant
power input type, thus minimizing the heat dissipation caused by high voltage
inputs •
15-49
.._.-_. SEDR300
i___,_1_ PROJECT GEMINI
I
I
, i II
I iNPUT
I/ ,.,o_o,,,,o,_oou,,,o,,o,,o/I I ....
MULTIPLIER _ POWER _ BUFFER MULTIPLIER _ BUFFER
J AMPLIFIER AMPLIFIER _ X4 AMPLIFIER J
I _ -- IA-2X12MULTIPLIERANDPOWERAMPEIFIERBOARD|m m m m i _ m _ m i _ m m m m m m m
I18TO30,5VDC A-5 ! A-4 _ 70VOC @ 162MAINPUT POWER _ LINE -_ DC-DC
FILTER I CONVERTER _ 30 VDC @ 42MA
Figure 15-11 Telemetry Transmitter Block Diagram
15 - 5 0
E@ sE°R3ooPROJ EC"T GEMINIF _
Antennas
The three uhf antennas (ascent, stub and whip) provide s_ultaneous t_-Az_ssion
for the telemetry transmitters and reception for the DCS receivers. The uhf
ascent antenna is used during pre-launch for pad checkout and from lift-off
until ATDA separation. The uhf stub antenna gives _directianal roll coverage
after ATDA separation. Switching from the ascent antenna to the stub antenna is
accomplished by coaxial switch 3 at ATDA separation. The uhf whip antenna pro-
rides omnidirectional yaw coverage and is used from pre-launch throughout the
mission. The antennas have a quarter-wavelength radiation pattern.
DHF Ascent Antenna
The ,,_fascent antenna (Figure 15-12) is mounted on the Target Docking Adapter
cone where it is protected from wind blast and launch temperatures during
lift-off by the ascent shroud.
The uhf ascent antenna is appro_mately 16 inches long. The antenna element con-
sists of two I/2-inch wide gold,plated steel blades which are bolted together.
For rigidity, the antenna element is shaped in a O.5 inch wide arc having a
radius of 1.5 inches. The two laminations of steel blades, compounding a single
antenna element, are rigidly secured together with nuts and bolts at the top
and bottom.
D_F Stub Antenna
The uhf stub antenna, p_ysically constructed as illustrated in Figure lS-l_, is
mounted on the battery module. The antenna protrudes aft _ the battery module
15-51
-_ SEDR 300
POWERCONi_
COVER _-_
JUHF WHIP ANTENNA
// ,CONNECTOR
BASE
/" -_ \\ _- -_\\\\\\
Figure 15-12 UHF Ascent, Stub, and Whip Antennas
15-52
__ SEDR300 ____
PROJECT GEMINI
into the forward section of the launch vehicle. _,I s protects it during the
boost phase of the mission.
The antenna consists of a mast and base which weighs approximately i.i pounds.
The mast is constructed of 3/4-inch cobalt steel_ machined to tubular form, and
covered by a Teflon ablation shield. The antenna is approximately 13.5 inches long
including the connector, and 1.25 inches in diameter over the ablation material.
The radiating length of the ante_-_ is approwlmately 11.2 inches long. The
mast consists of two sections. The front section is mounted on a cobalt steel
ball Joint and retained to the rear section by a spring loaded cable. Electrical
contact between the mast sections is made through the ball Joint and the spring
loaded cable assembly. The ball Joint allows the front section of mast to bef_
deflected to appro_wlmately 90 degrees in any direction around the antenna axis.
The spring of the cable assembly is preloaded to approximately 45 pounds to
cause the front section, when deflected, to return to the erected position.
The rf connector is press fitted into a socket and makes contact to the mast
through the socket and sleeve, which are the same material as the mast. The
shell of the rf connector is mounted to the base which is isolated from the mast
by a Teflon spacer and sleeve.
UHF Whip Antenna
The uhf whip antenna is mounted on the equipment section of the ATDA. The whip
antenna is extended during pre-launch and is used throughout the mission. A
fiberglass fairing protects the antenna from damage during launch.
The antenna element (Figure 15-12) is a tubular device made from a 2-inch wide
beryllimm copper strip processed in the form of a tube. The antenna forms an
15-53
PROJECT GEMINI
element that is approximately 9.2 inches long and i/2 inch in diameter.
C-Band Antennas and Power Dividers
Two sets of C-band helical antenna supply the transmission and reception
capability for the C-band radar beacons. Each set of helical antennas consists
of three radiating elements (antennas) and a power divider. The three antennas
have a radiation pattern with three s>_etrically located lobes. Two of the
lobes are oriented toward the earth when the ATDA is roll-stabilized in flight.
The C-band helical antennas (Figure 15-13) are mounted on the equipment section
flush with the outside skin of the ATDA and spaced 120 degrees apart. Each
antenna ,,-_t is approximately 3.4 inches long, 1.8 inches wide, has a depth
of 2.21 inches over the connector, and weighs approximately 3.5 ounces.
The power divider (Figure 15-13) measures approximately 3.86 inches over the
connectors, 4.0 inches over the tuning knobs, and weighs about 6.5 ounces.
The power divider is basically a cavity type power splitter. During beacon
transmission, power is delivered to the power divider where it is divided
equally among the C-band radiating elements. The power divider contains a
double stub tuner to compensate for mismatch between the C-band beacon
and the C-band radiating elements. Tuning is accomplished by means of a self-
locking t,m!ng shell located underneath each tuning stub cap.
Diplexers
The two uhf diplexers provide isolation between the DCS receivers and the tele-
metry transmitters to permit the use of a c_on antenna for both transmission
and reception. The diplexers are located in the ATDA equipment section. The
15-5_
...dr:__. SEDR 300
UHF "DIPLEXERSpowER iC- BAND HELICAL ANTENNA
SWITCHES
•C-SAND ANTENNAS i
(TYPICAL 3 PLACES)
_il ,OWE,DIV,DE,
_ ! ( O J1
±J3ANTENNA O
_j.__ ( O_11EUEMETRYTRANSMITTER
UHF DIPLEXER
SOLENOID COM/_ON C RF POSITIONS
-__ - _ j .....+28V POSITION NO, I
INDICATOR CIRCUIT NO. I
SHOWN IN ENERGIZEDPOWER AND INDICATOR PIN B POSITIONCIRCUIT CONNECTOR
_" RF COAXIAL SWITCHES
Figure 15-13C-Band Antennas, Diplexers,And Coaxial Switches
15-55
SEDR 300
GEMINI
physical representation and a schematic of the diplexers is illustrated in
Figure 15-13.
The diplexer is approx_m-tely 4.5 inches wide, 4 inches high, and 2.7 inches
deep, contains two input and one output connectors, and weighs approximately
1.25 pounds. Each channel consists of a high Q cavity, tuned to the corres-
ponding operating frequency. All channels are isolated from each other without
appreciably attenuating the rf signals passing through it. Each channel can
be re-tuned if the assigned operating frequency is changed.
Coaxial Switches
The three coaxial switches permit antenna and transmitter/receiver selection.
The coaxial switches are located in the ATDA equipment section. The physical .....
construction and approximate location of the coaxial switches is illustrated in
Figure 15-13.
Coaxial switch 1 is used for antenna selection for the primary telemetry trans-
mitter.
Coaxial switch 2 is for antenna selection for the secondary telemetry transmitter.
Coaxial switch B provides antenna selection for one of the DCS receivers and
either of the telemetry transmitters.
Each switch contains a power connector, an input connector, two output connectors,
and weighs approximately 0.5 pounds. The dimensions of each switch are approxi-
mately 2.65 inches long, 1.82 inches high, and i inch wide. The three coaxial
switches are identical and may be used interchangesbly. Basically, the coaxial
15-56
__. SEDR 300 ______j
PROJECT GEMINI
switches supply single-pole double-throw switching action as illustrated in
Figure 15-13. The switch, having a 20 millisecond maximum operation time,
operates on 3 amperes at 28 vdc and uses a latching solenoid break-before-make
switching action. The coaxial switches are designed to operate from 15 mc to
500 mc, and from 5500 mc to 5900 mc. Pins A and B of each switch are utilized to
accomplished the switching action.
BEACONS
The two C-band radar beacons supply the tracking capability for the ATDA. The
two radar beacons are transponders which transmit a reply when properly inter-
rogated. The beacons are located in the equipment section of the ATDA. Each
beacon consists of a superheterodyne receiver, a decoder section, and a trans-
mitter. The receivers are tuned to a center frequency of 5690 megacycles and
the transmitters to 5765 megacycles. The transmitter peak power output is 500
watts mi_m_n to the antennas. Code spacing for the beacons is 3.0 microseconds.
The C-band beacon is a sealed unit and measures approximately 9.34 inches by 8.03
inches by 3.26 inches. As illustrated in Figure 15-14, the beacon has a power
and test connector, an anten_ connector, and a crystal current test point
connector. The beacon contains external adjustments for local oscillator, pre-
selector (rf filter), and trans_nittertunlng; switches for selecting the desired
interrogation code; and one of two preset transponder fixed delay times. These
adjustments and switches are accessible by removing pressure sealing screws.
The beacon _nploys solid-state circuitry, except for the transmitter magnetron
and receiver local oscillator.
TELEMETRYTRANSMITTERS
The telemetry trau_dtters provide the instrunentation data tran_mlssion capabil-
15-57
_,:_;_, SEOR30o __
//
/
@
COAXIAL SWITCH
POWERCONNECTOR
TELEMETRYTRANSMITTER
C-BAND RADAR BEACON
Figure 15-14 C-Band Beacon and Telemetry Transmitters
15-58
SEDR 300
ity. The telemetry transmitters are located in the ATDA equipment section. The
physical construction and approximate location of the telemetry transmitters is
illustrated in Figure 15-14.
The two telemetry transmitters are identical. They weigh about _i ounces and
are 2.75 inches high, 2.25 inches wide and are 6.5 inches long. Each transmitter
contains adc power connector, an rf output power connector, and a video input
connector.
The telemetry transmitters are solid-state frequency modulated transmitters.
Each transmitter consists of an oscillator-modulator, a times-12 (x12) multiplier
and power amplifier, a bandpass output filter, a line filter and a dc-dc converter.
After a 30-second warm-up, the transmitters are capable of continuous _m_nterrupt-
ed operation for 500 hours. The transmitters operate at 2_6.3 megacycles with 2.0
watts minimum power output. Peak carrier frequency deviation is + 150 kilocycles.
DIGITAL CC_ND SYSTH_
The DCS consists of a receiver-decoder and three relays and supplies the link for
ground cc_mands to the ATDA. The DCS components (Figure 15-15) are located in
the equipment section of the ATDA.
The two DCS receivers operate on a fixed frequency in the 406-450 megacycle
range. Each receiver consists of a preselector, local oscillator and multiplier,
two intermediate frequency strips, a discriminator, and an output audio amplifier.
The decoder contains a sub-blt detector, a bit detector, an address decoder, a 5-
stage and a 24-stage storage register, relay selection circuits, and a syachronizing
and reset circuit.
15-59
s oR300L-t_ F'RoJEcTG_,N,
(TYPICAL 3 PLACES)
'DCS RECEIVER/bECODER
Figure 15-15 DCS Receiver/Decoder and Relay Boxes
15- 60
__ $EDR300 _____
PROJECT GEMINI
The receiver-decoder package is approximately 8 inches high, 8 inches wide, and
12 inches long. The relay boxes ax_ identical. Each relay box is approximately
2.25 inches wide, 5 inches high, and 3 inches deep. The combined weight of the
receiver-decoder package and the relay boxes is approximately 23 pounds. The
receiver-decoder pacMage contains two uhf receivers and a decoder while each of
the relay boxes contain eight relays.
zS-6z/6
INSTRUMENTATION SYSTEM
TABLE OF CONTENTS
T I TLE PAGE
SYSTEM DESCRIPTION ........ . . . 15-65SYSTEM OPERATION ....... 15-66
PCMPROGRAMMER6P R T[O ..... 15-66SEQUENTIAL SYSTEM PARAMETERS : . . . 15-70ELECTRICAL SYSTEM PARAMETERS .... 15-71TARGET STABILIZATION SYSTEMPARAMETERS ....... 1 5"72DIGITAL COMI_ND" SC(S÷EMPARAMETERS ...... 1 5-72REACTION CONTI_OI"SYSTEM"PARAMETERS . . . . 15-73RENDEZOUS RADAR" TI_AI_SPONDER"PARAMETERS ..... . . • 15"73I NSTRUMENTAT I(_N "SYSI_EI_PARAMETERS 15"74
PAI_AMETEI_S'"" " ..... 5 74STRUCTURAL ....... 1 "SYSTEM UNITS ........ . .... . 15-75
TEMPERATURE SENSORS... ° .... , 15-75SIGNAL CONDITIONER . . . . ..... 15-75DC-DC CONVERTERS .......... 15-77PCM PROGRAMMER.......... ° 1 5-79TELEMETRY TRANSMITTERS ....... 1 5-79
15-63
-_.. SEDR300
LIMIT _/VtTC HES
(TO09)
ANTENNA(TC0|)
JSRADARTRANSPONDER
('rBI3 e TBi4, TB15)
(TA07, TAGS, TB01, TB02)
SENSORS (3 PLACES)(TC03)
/SENSOR (TA02)
\
---- RIGIDIZED LATCH
(3 PLACES)(TC04)
DCS RECEiVER-DECOD ER
PC_'ER RELAY AND {TA03, TA0,(-a TA06, TC20)SEQUENTIAL PANEL(TC08)
ION SENSOR
(TB08, TB09_ TB10)
EQUIPMENT BEY TEMPERATURESENSOR (TA01)
Figure 15-16 Instrumentation System
15-64
__ SEDR 300 -___ "---1
PROJ EC"]" GEMINIS _
INSTRUMENTATION SYSTEM
SYSTEM _ESCRIPTION
The Instrumentation System provides a means of data acquisition with respect
to the progress and condition of the ATDA. Data acquisition in the ATDA is the
sensing of specific conditions or events on board the ATDA and displaying the
data derived from these inputs to ground operation personnel. In this respect
the data acquisition is shared by all ATDA systems and the ground operation
support system.
The basic components of the Instrumentation System are: temperature sensors,
signal conditioners, dc-dc converters, and a programmer. The location of the
Instrumentation System components and various signal monitoring points are
illustrated in Figure 15-16.
The Instrumentation System supplies 43 parameters (measurements) to the ground
station. The system primarily monitors ATDA system parameters. These parameters
are used for determining the progress of the ATDA and the performance of the
various systems, and for making decisions concerning mission safety and success.
During pre-launch checkout of the ATDA, ground commands are given and their exe-
cution is verified from instrumentation data. Equipment status and sequential
switch positions are monitored to insure all ATDA systems are in lift-off con-
figuration. During the orbital phase of the mission, the Instrumentation System
supplies real-tlme telemetry information while the ATDA is within range of the
ground station.
15 -65
___ SEDR3OO _'-_
PROJECT GEMINI
SYSTEM OPEI_TI ON
•The Instrumentation System is controlled by ground cow,handsvia the Digital Com-
mand System (DCS). DCS channel 6 is used to apply main bus power to the primary
dc-dc converter and telemetry transmitter; while channel i0 controls power to the
secondary converter and telemetry transmitter. An interlock in the wiring arrange-
ment of the power control relays (K33-I and EB3-2) prevent application of power
to the primary and secondary units simultaneously. A block diagram of the ATDA
Instrumentation System is illustrated in Figure 15-17.
The Instrumentation System obtains data from temperature sensors, pressure trans-
ducers, switch and relay actuations, and various monitoring points in the ATDA
systems. The majority of the signals obtained are compatible with the multiplex-
ing and encoding equipment without alteration. Some of the signals, however, are --r
routed to a signal conditioner where their characteristics and/or amplitudes are
changed. The resulting signals, as well as those from other sensors, are of three
basic types: low-level (0 to 20 millivolts de), high-level (O to 5 vdc), and
bi-level (0 or 28 vdc). The programmer combines and converts these inputs into
a serial binary-coded digital signal and supplies it to the telemetry transmitters.
PC_ PROGP_J_ER OPERATION
The basic functions of the Pulse Code Modulated (PC_) progr_u_er are: data multi-
plexing; timing, to support the multiplexing functions; and analog to digital
conversion. The output of the prog_er to the transmitter is a pulse code
modulated pulse train. A block diagram of the PC_ programmer is illustrated in
_gure 15-18.
The basic components of the prog_er are a high-level analog subc_utator,
prime subc_tator, master c_-._-utator,analog to digital converter, output
15-66
SEO3OOPROJECT GEMINI
SYSTEM RADAR BAY CONVERTERSYSTEM TEMP SENSOR PRIMARy
IDCS PULSE K33-1
PRIMARYT/M PWR
CONTROL
. RELAy
HIGH LEVEL ZERO REF TBO1 .EQUIPMENT BAY TEMPERATURE TA01
TRANSPONDER RF POWER TB13m
TRANSPONDER CASE TEMPERATURE 18t4 I
ANTENNA SELECTION (DIPOLE OR SPIRAL) TBI5
BOOM EXTENDED(DIPOLE ANTENNA) TC01
RCS N2 REGULATED PRESS TB12
RCS RING A N2 SOURCE PRESS TB16
RCS RING B N2 SOURCE PRESS TB17
RCS RING A/B ON TC|0
TCA NO. 1 THROUGH NO. 8FIRING TCI1TOTC}8
ATDA SEPARATION TC(]_ 1 PRIGIDIZED TC04 |
SEQUENTIAL SHROUD JETTISON TC05 I TELEMETRy
SYSTEM LATCH RELAY MONITOR TC08 I TRANSMITTER
SPACECRAFT FREE TC09 PRIMARy
UNRI GIDIZE/LATCHES RESEE TC21
DCS PULSE BUS ON-OFF TC2O
t_" I DCS RECEIVER NO. 2 STRENGTH TA04.
I DCS RECEIVER NO. 1 STRENGTH TA03
l I °DCS O
SYSTEM
BATTERY MODULE TEMPERATURE TA02
SQUIB BUS SAFE TE06
SQUIB BUS ARMED TC07
SQUIB. BUS NO. 2 VOLTAGE TB06 / TELEMEtry I
SQUIB BUS NO, I VOLTAGE TB05 L ERANSMITTER ICONTROL BUS VOLTAGE S TB04 SECONDARYMAIN BUS VOLTAGE / TB(_
MAIN BUS CURRENT G TA05N
YAW RATE A TB1O
II"l LPITCH RATE C TBO8AC VOLTAGE O TBO7
D
I
T
l I 'TARGET O TA07ELECTRICAL STABILIZATION (HIGH-LEVEL FULL SCALE) N
SYSTEM SYSTEM (LOW-LEVEL FULL SCALE) E TA08(LOW-LEVEL ZERO REF) R TIB02
DES VALIDITY TA06
MAINBusBUS"}_ O'_lIA T "-_PULSE
f_K33-2 _ DC-DCCONVERTER
SECONDARy
TIM PWRCONTROL RELAY
m
Figure 15-17 Instrumentation System Block Diagram
15-67
_f_-_.. SEDR 300
BI-LIEVEL INPUTS19 AT IO SAMPLES LOW-LEVE L INPUTSPER SECOND 3 AT 80 SAMPLES
HIGH-LEVEL INPUTS HIGH-LEVEL INPUTS PER SECOND13 AT 1.25 SAMPLES 3 AT 40 SAMFtES 5 AT 640 SAMPLES
PER SECOND t _ PER SECOND l _ PERIbECOND- t
HIGH-LEVEL PRIME SUB _ MASTER _ DIGITAL -_ SHIFTANALOG SUB _ COMMUTATOR COMMUTATORCOMMUTATOR CONVERTER REGJSTER
OUTPUTFILTER
TO TRANSMITTER
-_ [ TAPE
RECORDER
SPECIAL _ iNPUTTIMING CONVERTER/
REPRCOUCER
DIGITAL INPUT
I_,_UTS _)*-_ S;ELFCTOR CIRCUITRY
Figure 15-18 PCM Programmer Block Diagram
15-68
SEDR 3O0
shift register, digital input selector and input circuitry, special timing, output
filter, and tape recorder input converter/reproducer.
The programmer high-level analog subcommutator samples 13 high-level inputs at
1.25 samples per second. The analog subcommutator receives its inputs directly
from the signal sources, or from the signal conditioner. The output of the
analog subcommutator is applied to the prime subcom_utator for further multi-
plexing.
The prime subcommutator, in addition to accepting the sampled high-level analog
subcom_utator data output, samples 19 bi-level signals at i0 samples per second,
and 3 high-level signals at 40 samples per second. The prime subcommutator sup-
f_ plies its output to the master commutator.
In addition to accepting the inputs from the prime subcommutator, the master
commutator samples 5 low-level signals at 640 samples per second and 3 low-level
signals at 80 samples per second. The output of the master commutator is applied
to the analog to digital converter where the analog output from the master com-
mutator is converted to digital data.
The digital data from the output of the analog to digital converter is applied to
the output shift register which provides a continuous non-return-to-zero PCM
pulse train of 51.2 kilobits per second. The output of the output shift register
is applied to the telemetry transmitter through the output filter.
Timing of all operations in the PCM programmer is provided by the special timing
circuits. Two crystal controlled oscillators and a series of counters are used
to supply the gate pulses which determine the sampling rates of the programmer.
15-69
___ SEDR300 ____
PROJECT GEMINI
The digital input selector and input circuitry and the tape recorder input con-
verter/reproducer components are not utilized during ATDA operation. The Instru-
mentation System does not sample digital data or record data for data-dump
transmission.
The output signal from the PCM programmer is a 51.2 kilobit non-return-to-zero
signal with a voltage that is adjustable between 0.i volt and 1.0 volt peak. The
serial output has positive voltage for a i and zero or negative voltage for a 0.
A brief description of all AT_A instrumentation parameters is contained in the
following paragraphs. The parameters are described in groups identified by their
applicable data source system. Data flow from the signal source to the programmer
is illustrated in Figure 15-17.
SEQUENTIAL SYSEM PARAME_
The Instrumentation System monitors six Sequential System parameters. The sequen-
tial parameters are hi-level signals originating from switch or relay actuations
and are supplied directly to the PCM progrRmmer.
During pre-launch checkout of the ATDA, latch relay monitor (TC08) indicates that
all control latch relays are in the sequential mode required for launch condition.
When all control relays are latched or reset in the proper position a 28 vdc signal
is applied to the PCM programmer. One or more relays in an abnormal mode reduces
this signal to 0 vdc.
With the ascent shroud installed, a 24 vdc signal is applied to the PCM programmer.
At shroud Jettison (TC05), the 24 vdc circuit is broken thus reducing the voltage
applied to the programmer to 0 vdc. A 24 vdc signal is supplied to the PCM pro-
g_r to indicate ATDA separation (TCOS). This signal results when any two of
the three separation sensor switches close.
15-Yo
SEDR 300
PROJECT GEMINI
Unrigidize/latches reset (TC21) indicates the Target Docking Adapter cone is
unrigidized and all latches are reset. When this condition exists a 24 vdc signal
is supplied to the PCM programmer. Rigidized (TC04) is obtained from three rigi-
dized limit switches which close when the rigidized limit is reached. When all
three limit switches are closed a 28 vdc signal is applied to the PCM programmer.
Spacecraft free (TC09) indicates whether or not physical attachment exists between
the A_DA and the spacecraft. A 24 vdc signal is fed through one or both of two
limit switches to the PCM programmer when complete attac_ent does mot exist.
ELECTRICAL SYS_M P_ERS
The Electrical System supplies the ground station with 8 parameters which indicate
the condition of the ATDA batteries. Five of the signals are routed through the
_- signal conditioner to make them compatible with the PCM programmer circuits.
Battery temperature (T__02)indicates the battery case temperature. A resistive
element temperature sensor mounted on the battery supplies the low-level signal
to the PCM programmer.
Squib bus safe (TC06) and squib bus armed (TC07) are bi-level signals applied
to the PCM programmer which indicate the position of the squib control relays.
When all relays that control power to squib igniters are deenergized (TC06) a
24 vdc signal is obtained. When any of these relays are energized it removes
the 24 vdc signal from the PCM programmer. When one or more of the squib control
relays are energized (TC07) a 24 vdc signal is applied to another channel of
the PCM progr_,.mer.
f_ Main bus current (TA05) is used to monitor the total current being drawn from
15 -71
__ SEDR300 _.___
PROJECT GEMINI
the batteries. The signal originates from a 50 millivolt shunt installed at the
main bus. The signal is routed to the signal conditioner which provides a low-
level signal output to the PCM programmer.
Main bus voltage (TB03), control bus voltage (TB04), squib bus i voltage (TB05),
and squib bus 2 voltage (T_06) are measured to suppS_v the ground station with
information relative to battery condition. The high-level signals are routed
through the signal conditioner before being applied to the PCM programmer.
TARGET STA_VT,T_ATION SYS_ PARAMEteRS
The Instrumentation System monitors four parameters from the Target Stabilization
System. These signals are high-level signals conditioned by the signal conditioner
prior to being applied to the PCM programmer ......
AC voltage (TS07) gives an indication of the output of the inverter that is in
operation. Pitch, roll, and yaw rates (T_08, TB09, TSIO) are monitored to allow
evaluation of the rate control portion of the Target Stabilization System. Prlma_y
and secondary rate package signals are monitored on the same telemetry channel,
depending on which system has been selected by the DCS.
DIGITAL CO_%ND SYSTEM PARAME_2RS
The DCS supplies four parameters for ground monitoring. These signals provide
information relative to the o_eration of the DCS. Three of the four signals
are low-level signals.
Receiver i signal strength (TAO3) and receiver 2 signal strength (TA04) give an
indication of the strength of the signals received by the receivers in the DCS.
15-72
__. SEDR 300 _____
PROJECT GEMINI
DCS validity (TA06) is monitored to indicate that a command signal was received
by the DCS receiver-decoder.
DCS pulse bus (ON-OFF) (TC20) indicates that power is applied to the DCS execute
circuitry. A 28 vdc hi-level signal from the DCS pulse bus is applied to the
PCM programmer each time the pulse bus is energized by a ground command.
REACTION CONTROL SYSTEM PARAMETERS
Five parameters from the Reaction Control System are monitored to indicate the
condition of the system. Reaction Control System ring A/B (TCIO) is a hi-level
signal that indicates which thruster ring is energized. A 28 vdc signal to PCM
programmer indicates ring A operation; 0 vdc indicates ring B operation. Thruster
..... firing, TCA i through 8 (TCII through TCI8), are bi-level signals are are moni-
tored to indicate proper operation of the Reaction Control System, Bi-level
pulse signals of 0 or 28 vdc are derived from the solenoid command signals. A
0 vdc signal indicates thruster firing. Regulated nitrogen pressure (TBI2)
is a high-level signal obtained from a potentiometer-type transducer to indicate
the pressure (0-500 psi) of the Reaction Control System regulated nitrogen.
Ring A and B nitrogen source pressure (TBI6, TBIT) parameters indicate the
source pressure in each nitrogen supply. The pressure is measured by a potentio-
meter-type pressure transducer which is a part of the Reaction Control System.
EENIEZVOUS RADAR TRANSPONDER P___S
To monitor the operation and environmental condition of the rendezvous radar
transponder, four parameters are relayed to the ground station. A temperature
transducer installed in the transponder supplies a high-level signal output for
traneponder case temperature (TBI/_).
15-73
__ SEDR300 _.__
PROJECT GEMINI
Beam (dipole) antenna extended (TCOI) is monitored for operational reliability.
An output of 28 vdc from the boom actuator indicates the antenna is extended. An
output of 0 vdc indicates the antenna is not fully extended. Antenna selection
(TBIS) indicates spiral or dipole antenna usage. Selection of the antenna is
governed by the strength of the interrogating signal. An output of 1.O to 5.0
vdc indicates dipole antenna usage and O to Ou 5 vdc indicates spiral antenna
usage •
Transponder rf power (TBI3) is a high-level signal that indicates when the high
voltage power supply in the radar transponder is energized. As long as the trans-\
ponder receives interrogation signals the high-voltage power supply is energized.
Loss of the interrogating signals deenergizes the power supply. An output of
2 to 5 vdc indicates an on condition, and an output of 0 to 0.5 vdc indicates an .....
off condition.
INSTRUMENTATION SYStem4 PABAMET_RS
To insure proper scaling is being employed by the Instrumentation System, four
parameters are monitored. These reference signals, low-level full-scale (TA07),
low-level zero (TA08), high-level full-scale (TB02), and high-level zero (TBOI)
are outputs of the de-de converter that have been conditioned by the signal con-
ditioner. The data received from these parameters provide the ground station
verification that the Instrumentation System is functioning properly.
STRUCTURAL PARAMET_mS
Equipment bay temperature (TAOI) is the only structural parameter monitored and
indicates the temperature that exists within the equipment bay. The temperature
is measured by a resistive-element temperature sensor which supplies a low-level
signal to the PCM programmer_
15-7h
___ SEDR 300 ___
PROJ E-'CT" GEMINI
SYS__.M UNITS, J , ,,
TEMPERATURESENSORS
The two temperature sensors are resistive-element, surface-mounted, sensors that
convert the te_uerature into directly proportional electrical signals. The
sensing elements are fully-annealed pure-platinum wire encased in ceramic insula-
tion in a strain-free manner to provide maximum stability.
The physical construction of the sensors and a typical schematic is illustrated
in Figure 15-19. The sensors are approximately 0.4 inches by 0.75 inches by
2.0 inches and weighs a maximum of 0.075 pounds. They have a temperature range
of O to 400 degrees Fahrenheit and provide a 0 to 20 millivolt dc output.
"- One of the temperature sensors is mounted on the battery case and the other in
the equipment section at the intersection of the equipment support assemblies.
SIGNAL CONDITIONER
The signal conditioner is approximately I0 inches by i0 inches by 8 inches and
weighs about eight pounds. The signal conditioner is located in the equipment
section of the ATDA. Figure 15-19 illustrates the physi_l construction of the
signal conditioner.
The signal conditioner contains 12 signal conditioning modules: one ac voltage
monitor; four de voltage monitors; one dc m_11_volt monitor; three atten,,ators;
and three phase sensitive demodulators. The modules are constructed on circuit
boards using printed wiring techniques.
_ The ac voltage monitor accepts a signal ranging from 23 to 29 volts rms over a
15-75
. SEDR 300
F..O.,.CT
__11 MODULES
TEMPERATURE SENSOR AND BRIDGE SIGNAL CONDITIONER
INPUT C _7 t
OUTPUT C
J !SENSING |,.p_To ' I _.E_',',_TIL,.___ J
OUTPUT0
TYPICAL SCHEMATIC TEMPERATURE SENSOR
Figure 15o19 Signal Conditioner & Temperature Sensor
15-76
__ SEDR300
PROJECT GEMINIf_L
frequency range of 380 to 420 cycles. The output is from O to 5 vdc, varying
only with the input voltage.
The four types of dc voltage monitors are designed to accept various positive
and negative dc voltage inputs and provide outputs of O to 5 vdc.
The de millivolt monitor accepts an input of 0 to 50 millivolts and provides a
proportional output of O to 20 millivolts.
There are three types of attenuator modules. These modules have various dc
inputs which are attenuated to the 0 to 20 millivolt range or the O to 5 vdc
range. Some attenuator modules contain two data channels.
f_ There are three variations of the phase sensitive demodulators. Basically,
the phase sensitive demodulator accepts two input voltages: one signal voltage
and one reference. It provides a dc output of five volts for a full scale
input signal that is in phase with the reference and an output of 0 volts for
a full scale signal that is 180 degrees out of phase with the reference.
DC-DC CONVER_ERS
The two dc-dc converters supply the Instrumentation System with regulated dc
voltages of +5 and + 24 vdc. 'l_neconverters are essentially voltage regulators
which operate on 18 to 30.5 vdc and supply regulated outputs.
The dc-dc converters (Figure 15-20) are approximately 5.5 inches by 5.5 inches by
7 inches3 and weigh about seven pounds each. The converters are located in the
ATDA equipment section.
15-77
DC-DC CONVERTER
,MI3OTHERBOARDS_
T1 (TEST CONNECTOR)
PCMPROGRAMMER
Figure 15-20 DC-DC Converter and PCM Programmer
15-78
___ SEDR 300 __
PROJECT GEMINI
PCM PROG_
The PCM progra_ner is located in the equipment section of the ATDA and encodes
all instrumentation data for transmission. The progrsn_er weighs approximately
20 pounds and is ii inches by ii inches by 4.5 inches. (Figure 15-20).
The programmer is composed of 16 subassemblies. Thirteen of the subassemblies
are multilayer printed-circuit boards (mother-boards) constructed of a glass-
epoxy bass material laminated at high temperatures. Most of the programmer
subassemblies are embedded in polyurethane foam to provide rigidity and damping.
An elapsed time meter indicates total operating time.
_ELEMETRY TRANSMITIERS
Two telemetry transmitters transmit the Instrumentation System data to the
ground stations. Although the transmitters serve the Instrumentation System,
the antennas and associated switching are part of the Cu_..unicationSystem;
therefore, the transmitters are described in detail in Section XV, Cu_._m4-
cation System.
f_
PYROTECHNIC DEVICESAND SEPARATION ASSEMBLIES
TABLE OF CONTENTS
TITLE PAGE
INTRODUCTION............. 15-83EXPLOSIVE BOLTS ............ 15-83
F_, DESCRIPTION ............. 15-85OPERATION ........... 15-85
PYROTECHNIC VALVES: . o . o . . . . . . 15-85DESCRIPTION ............. 15-86OPERATION ...... 15"87
FLEXIBLE LINER'SHAPED CHARGE: ..... 15-87DESCRIPTION ............. 15"87OPERATION ....... 15"89
SHROUD SEPARATION ASSEMBLY: ...... 15"89DESCRIPTION ............. 15"90OPERATION .... 15-92
ATOA/LAUNCH VEHICLE SEPARATION'ASSEMBLY 15-92DESCRIPTION... , ......... 15-92OPERATION .............. 15"93
15-81
=-_. SEDR 300 __PROJECT GEMINI
SPRING LOADEDN
SPRING
STEEL BANDS
PRING ACTUATOR
_OLTS
(4 REQ'D.)
; "C" PACKAGE
BY PYROTECHNIC VALVE
@ PACKAGEPY_ROTECHNIC VALVE
rE SEPAKATI ON ASSEMBLY
RCS _D" PACKAGE (FLEXIBLE LINEAR SHAPED CHARGE,HEAT PROTECTION ANO SHRAPNELSHIELD AND FAIRING)
\
RCS "A" PACKAGE
NCH VEHICLE"ION ASS'Y
BUNGEE CORD
(8 REQ'D .)
SEPARATION BARS(4 REQ'D.)
RCS "C" PACKAGEPYROTECHNIC VALVE
Figure 15-21 Pyrotechnic Devices and Separation Assemblies
15-82
___ SEDR300
PROJECT GEMINI
PYROTECHN,IC DEVICE S AND SEPARATION ASSSMRT,TR.S
ODUC ION
The pyrotechnic devices and separation assemblies (Figure 15-21), installed in
the Augmented Target Docking Adapter (ATDA), are used to execute events initiated
by the Sequential System. The events executed by these devices, in sequential
order, are: release the shroud retaining bands; Jettison the ascent shroud,
sever the ATDA/launch vehicle mating ring, separate the ATDA from the launch
vehicle and activate the Reaction Control System (RCS).
The pyrotechnic devices and separation assemblies which perform these functions
and the events executed by them are as follows:
_ Explosive bolts - release shroud retaining bands.
Shroud separation assembly - separate and Jettison the shroud.
Flexible Linear Shaped Charge - sever the ATDA/launch vehicle mating ring.
ATDA/launch vehicle separation assembly - provide a separation velocity
between the launch vehicle and the ATDA.
Pyrotechnic valves - open lines and activate the RCS.
Expnosn , BOLTS
Explosive bolts (Figure 15-22) are located at four points on the ATDA. Two bolts
secure the steel retaining bands around the shroud. The other two bolts secure
the base of the shroud_ at the separation line. When _-4tiated by the Sequen-
tial System, the bolts explode, releasing their respective clamps and allowlng
the separation assembly to Jettison the shroud.
The bolts are electrically connected to the Sequential System via two cables. The
cables are located 90 degrees from the shroud separation line, near the base of
each segment.
15-83
,.-d_"_-_ SEDR 300 L_r_ __]O M,N,TYPICAL EXPLOSIVE BOLT CROSS SECTION
EXPLOSIVE CHARGE 16REQ'D)
_i_ DETONATOR_i_ CONNECTOR
(2 REQJD)
)DETONATOR (2 REQ_D)
WASHER (2 REQJD)
iiii RCS PACKAGE "C" AND "O"RCS PACKAGE "A" ii::i
ii!i BEFORE FIRING
BEFOREFIRING i_!_iI N IPPLE CARTEIDGE _:_:
_ ..... )ii ii o• i!i!i _ --
_:_:i
_oFLOW ii!ii_-- SCREW
(4 TYPICAL)
_:_:
:_:_
_:_:AFTER FIRING :::: AFTERFIRING
!!_!i
REMOVED iiili _EMOVED .....:_:_
NORMALLY CLOSED VALVE (NON.REPLACEABLE) :::! NORMALLY CLOSED VALVE (REPLACEABLE)
Figure 15-22 Pyrotechnic Valves and Explosive Bolts
15-84
SEDR300 __
PROJECT GEMINI
DESCRIPTION
The four identical explosive bolts are five inches long, 5/8 of an inch in
diameter and are made of stainless steel. A deep cylindrical cavity is machined
into each end of the bolts. The cavities extend in approximately two inches
from each end of the bolt. A solid part in the center is not drilled out. The
end of each cavity is tapped to receive the threaded electrical detonator. The
bottum of each cavity is so designed that the explosive blast will cause the bolt
to break or fail at the central solid part as though sheared off there.
The detonators used are of the bridge wire type. Redundancy is maintained by
using two detonators per bolt. A sleeve is slipped over the outside of the bolt
and held in a center position with set screws. The sleeve maintains the parting
planes of each bolt outside the clamps which it holds in place, and prevents
any bolt fra_entation from piercing the shroud.
OPERATION
When the shroud Jettison squib fire relays connect squib bus power to the eight
shroud explosive bolt detonators_ a current sufficient to set off the l_-!tion
mix, flows in the bridge wire. The energy frcm the ignition mix detonates the
Penta Erythritol Tetra Nitrate (PETN) explosive charge next to it. The stress
which the explosion puts upon the bolt structure causes the hollow parts on each
side to break away frcm the central solid part.
PYROTECHNICVALVES
Six pyrotechnic valves (Figure 15-22) are installed in the RCS. Three valves
are used in each ringj one valve for each packs4_e. Each ring has a fuel package
(called the "D" package)j an oxidizer or "C" package, and a pressurant or "A"
pac_e° _e v_ives isolate t_e fuel_ oxidizer and pressurant from the rest of
15-85
__ SEDR300
PROJ EC'" GEMINI
the RCS by keeping them in their respective supply tanks. When the valves of
either A or B ring are actuated_ pressurant is released into the _,-eland
oxidizer ta-_, and fuel and oxidizer are forced into lines to the eight
activated thrusters. The pyrotechnic valves are one-tlme actuating devices
which control the flow of these fluids.
Two types of valves are installed in the RCS. Both are normally closed. One
type however is replaceable, the other is not° The RCS package "A" uses the
nonreplaceable type. The RCS packages "C" and "D" use the replaceable type.
If a nonreplaceable valve is defective, or if the squib has been fired, the
package must be changed. Replaceable valves can be changed individually
without chapping the whole package.
DESCRIPTION
A pyrotechnic valve consists of a cartridge, a valve body, one or two nipples,
a ra_, a seal and a screw. The replaceable valve uses one closed-end nipple
which is installed within the valve body and welded into place. The nipple
is in the line from the supply. As long as the nipple is intact, it blocks
the fluid from the supply. The nonreplaceable valve has two closed-end nipples
butted together; one is the inlet, the other is the outlet port. The ram is
in a slot or channel in the valve body, indexed directly above the centers of
the nipples. The seal and screw are in the head of the ram. The cartridge is
in the valve body at the top of the ram head. The function of the seal is to
prevent blow-by of the actuation gas. The function of the cartridge is to
generate the energy to drive the ram. The ram has two functions. One is to
shear off the nipples and open the lines. The other is to c_plete the line
by aligning the hole in its body with both sides of the llne.
15-86
__. SEDR300
PROJECT GEMINIf_
OPERATION
At Vernier Engine Cut-Off (VEC0) plus 12 seconds, the squib buses are connected
to six squibs on the three pyrotechnic valves of RCS ring B. Each valve operates
in an identical manner. The squib sets off the cartridge which generates suffi-
cient gas pressure to drive the ram. The ram is propelled along the channel, and
the ram head shears off the nipples in its path. As the ram comes to rest at the
opposite end of the channel, a hole through its body is aligned with the lines
once blocked by the nipples. Fluid flows freely from the supply tank through the
llne s.
This same process is repeated on RCS ring A the first time this ring is selected
by a Digital Command System (DCS) command.
FLEXIBLE LINEAR SHAPED CHARGE
The Flexible Linear Shaped Charge (FLSC) is the pyrotechnic device which severs
the ATDA/launch vehicle mating ring. The FIASCis shown in cross section in
Figure 15-2S. The FLSC is a V-shaped flexible lead sheathing containing a core
of cyclonite (RDX) high explosive. The specific type, shape and thickness of the
material to be severed by a FLSC dictates the amount of explosive to be used. In
this explosive separation assembly, dual strands of FLSC are used.
DESCRIPTION
The FLSC assembly consists of three detonator blocks, three crossovers, three
dual boosters, a molded backup retainer and a back blast shield. Single bridge
wire detonators are used in two detonator blocks, a double bridge wire in the
f-_ third. In each detonator block, the crossovers are next to the detonators, and
the boosters are next to the crossovers. The crossovers apply the shock wave
15-87
RiNG
_UNGEE CORDS (TYP.)
SINGLE BRIDGE WIRE
TWO BRIDGE WiREDETONATORS
_ DETONATOR
_ND SHRAPNELPROTECTION SHIELD
_--- SHAPED CHARGE
_--- BACK BLASTSHIELD
DETONATOR
DETONATOR BLOCK_
D CHARGE
EQUIPMENT
MOLDED BACKUPRETAINER
BACK BLAST SHIELD-FIBERGLASSFAIRING
Figure15-23 FLSC Mechanical Separation Assembly
15-88
___ SEDR 300 ___
PRO,J GEMINI
to the booster. The booster propagates the shock wave to the FLSC to insure
detonation of both strands.
When the explosive separation assembly is installed, the dual FIaC is attached
to the periphery of the ATDA/launch vehicle mating ring. The open portion of
the V-shaped FLSC is place_ toward the surface to be cut.
The detonator blocks are installed on the inside of the mating ring and extend
through the skin into the FLSC molded backup retainer. The detonators are
wired to the ATDA sheped charge and RCS ring B squib fire relays. The back
blast shield surrounds the molded backup retainer and is attache_ to it. A
fiberglass fairing covers the retainer, shield and FLSC to protect the explosive
_ separation assembly from aerodynamic heating.
OPERATION
At VECO plus 12 seconds, the Sequential System connects the squib buses to four
ATDA shaped charge igniters, setting them off. The igniters detonate the cross-
over, the crossover detonates the boosters, and the boosters detonate the redun-
dant strands of FLSC.
As the FLSC detonates, the concussion collapses the sheathing in each V-groove.
This directed action produces a cutting Jet of explosive products and minute
metal particles. The Jet in turn produces extremely high localized pressures.
The resulting stress is far above the yield strength of the ring material. The
skin is severed and the ATDA is detached, permitting the mechanical separation
assembly to operate.
SHROUD SEPARATION ASS'_'MBLY
The shroud separation assembly is designed to Jettison the ascent shroud at
15-89
___ SEDR300 _____
PROJECT GEMINI
the completion of the launch phase. The shroud is held together by steel bands
until after VEC0, then explosive bolts are fired, releasing the steel bands and
allowing the separation assembly to function. The shroud is constructed in two
segments, one long segment (including the nose cone) and one short segment. When
Jettisoned, the shroud separates and falls away to the rear, one segment on each
side.
DESCRIPTION
The shroud separation assembly (Figure 15-24) consists of two preloaded spring
actuators; two preloaded spring thrusters; two pivots; two spring loaded nose
latches; two elastic latch cords with ball detents; two pair of split anchor
blocks; and two pair of steel bands and clamps (held together by explosive bolts).
The spring actuators are mounted at the base of the shroud at the separation
line. The spring thrusters are mounted just below the shroud center of gravity.
The nose latches hold the upper end of the shroud together before Jettison,
and unlatch to permit the segments to part at the nose. The elastic cords apply
tension to the latch handles to keep the latches engaged. The ball detents on
the cords are clamped in the anchor blocks to hold tension on the latches. The
anchor blocks, which are compressed and trapped in the parting line by the tension
of the steel bands, maintain tension on the latch cords. The base clamps, when
pulled together by explosive bolts, compress the spring actuators and hold the
shroud base in firm contact with the Target Docking Adapter (TDA). The entire
assembly is held together by the steel bands, until the explosive bolts are fired
by the Sequential System.
15 "90
s,o.30o_-. PROJECT GEMINI/
_%R_ NOB_ED_C.
STEEL BANDS
EXPLOSIVE BOLTS
PIVOT BRACKE, f S_NONUE_Elr_RCTRICAL ....._
)
LANYARD NOSE LATCH
ASCENT SHROUDtARGET DOCKING i
i SEPARATION PLANE L
ADAPTER =_ !_
SlurINO ACTUATOR
(EXPANOED)
Figure 15-24 Shroud Separation Assembly
15-91
_@ SEOR300
PROJ INI
OPERATION
Operation of the shroud separation assembly (Figure 15-24) is initiated by the
Sequential System signal which fires the explosive bolts. When the bolts fire,
they release clamps in the steel retaining bands and in the base of the shroud.
Four lanyards are used to disconnect the explosive bolt cables as the retaining
bands spread. Spring actuators spread the shroud base apart, driving pivots
against brackets on the TDA. When the base has spread far enough, the anchor
blocks separate and release the elastic latch cords. When tension is removed
from the latch cords, the spring loaded nose latches open and allow the shroud
to separate at the nose. The spring loaded thrusters now drive the two segments
apart, nose first. As the shroud segments separate, lanyards are pulled tight
on two electrical connectors near the base of each segment. The lanyards discon-
nect the electrical connectors when the shroud segments are separated by approxi-
mately 50 degrees. The shroud segments continue to swivel on the pivots until
separated by approximately 180 degrees. At this point the pivots allow the shroud
to separate from the ATDA. The two shroud segments then fall behind and away from
the ATDA.
ATDA/IAUNCH VEHICLE SEPARATION ASSemBLY
The ATDA/launch vehicle separation assembly (Figure 15-23) is used to release
the ATDA from the launch vehicle and provide it with a separation velocity.
DESCRIPTION
The separation assembly consists of the mating ring; a catapult, made up of
eight bungee cords and attaching ring; and four separation rails. The mating
ring structurally bonds the ATDA to the launch vehicle during ascent and provides
a base from which the bumgee cords are suspended. The mating ring also provides
15 -92
___ SEDR 300 __
PROJECT GEMINI
a separation line where the two vehicles are severed pyrotechnically when orbit
is achieved.
The catapult is elastically stressed to approximately 500 pounds and provides
a separation velocity when the mating ring is severed. An attaching ring sup-
plies a co, non surface for attaching the eight bungee cords which provide the
thrust for the catapult. The attaching ring also provides a surface on which
the battery module rests during launch. The bungee cord catapult is suspended
from hangers on the underside of the mating ring.
Four separation rails span the RCS and battery module to guide the ATDA during
the separation sequence. The rails protect the RCS and battery module from
damaging contact with the mating rlng or launch vehicle during separation.
OPERATION
At VECO plus 12 seconds, the Sequential System initiates the separation sequence.
The pyrotechnics sever the mating ring and allow the separation assembly to
catapult the ATDA away from the launch vehicle. The ATDA# which weighs approxi-
mately 2000 pounds, separates from the launch vehicle at a velocity of approxi-
mately three fps (_wo mph).
15-93/94
TARGET STABILIZATION SYSTEM
TABLE OF CONTENTS
TITLE PAGE
SYSTEM DESCRI PT I ON........... 15-97SYSTEM OPERATION ......... 15-1OO
-_- RATE COMMANDS........... 15-1OOSECONDARY OPERA'FION......... 15-101ELECTRICAL POWER .......... 15-101
SYSTEM UNITS ........... 15-101RATE GYRO PACI_A(_ES..... 15-101ATTITUDE CONTROL ELECTR6NiC_; .... 15-102ORBIT ATTITUDE AND MANEUVERELECTRONICS ............. 15-104POWER INVERTERS ........... 15-105
15 -95
. sEo.3oo"_ PROJECT GEMINI
SECTION (REF)
MANEUVER ELECTRONICS
/I
1
RATEGYROPACKAGES.
__. $EDR 300 -__-_
PROJECT GEMINI
f--,
TARGET STABILIZATION SYSTEM
SYS_EMDESCRIPTION
The Target Stabilization System (TSS) is used in conjunction with the Reaction
Control System (RCS) to contx_l the rotational rates of the ATDA in the pitch,
roll and yaw axes. It provides two modes of control which are selectable from
a ground station via the Digital Command System (DCS) or from the spacecraft
via the Command Link.
Prior to the docking phase of'the mission, the TSS operates in the Biased Rate
Damping (BRD) mode. In this mode, the TSS establishes nominal ATDA rotational
rates of 1.8 degrees per second in the pitch axis, 2.2 degrees per second in the
roll axis, and 0.25 degrees per second in the yaw axis. These rates are estab-r_
lished to provide uniform solar heating, minimum radar ellipticity errors, and
maximum DCS and Rendezvous Radar reception. The BRD mode is selected prior
to launch and is initiated by the Digital Cu,_and System after the ATDA is
inserted in its orbit.
For the docking phase of the mission, the TSS is placed in the Normal Rate
Damping (NRD) mode. In this mode, the ATDA rotational rates are stabilized
at less than 0.5 degrees per second in all three axes to faciliate spacecraft
docking.
The TSS incorporates basically the same equipment as the Attitude Control and
Maneuver Elect_incs in the Gemini Spacecraft. It consists of an Attitude Control
Electronics (ACE) package, an Orbit Attitude and Maneuver Electronics (OAME)
,_ package, primary and secondary Rate Gyro Packages (RGP), and primary and secondary
15 -97
_ s oR3ooPROJECT GEMINI
r ATTITUDE CONTROL ELECTRONICS (ACE)
J
-1
,SECONOAR_" I I
I_'TEG_°PAC_AGE,I I r _ _',C.C.ANNE__ --'_ I-- RATE AMP. AMP
,_)__ = - = I
, ,,_ Y I SECONDARY _
SWITCH I
IYAW _ I
400 CPS Jk 1 J
PACKAGE'_,_RY7 [-- DCBUS_R -._, _RATE GYRO -- _j ROLLCHANNEL
DC BIAS POWER _J (SAME AS PITCH CHANNEL) _"
-_ ....... ____FROLLI-......... --1 I_J YAW CHANNEL + YAW
(SAME AS PITCH CHANNEL, EXCEPT J -YAW
VOLTAGE NO MIN. PULSECKTS) ..]
DIVIDER ; -L..,. .... jDC BIAS POWERO
DAMPING J
LN BIAS I
BIASED RATE POWER SUPPLy]NO
'_- --I -- J I
j SEC CIRCUIT SELECT(PRICIRCUIT SELECT
-- DC BUS POWER(TO INSTRUMENTA-TION SYSTEM)
PRI . PULSE BUS (DC MOMENTARy)AC POWER J
SEC
PRIMA,RyPOWERINVERTER PRI
-- ' CONTROL BUS (DC)
OFF _. BUS SEC SELECT & [:_OFF
SECONDARY O_POWERINVERTER A
SECII
Figure 15-26 TSS Functional Block Diagram (Sheet 1 of 2)
15-98
_ SEDR 300
_-'_' ,_,_o..,E_C',"_,,_, _-Im
r -- -- ORBIT ATTITUDE & MANEUVER ELECTRONICS (OAME)
I PRII
J sE¢[
I S_'KESUP':_ESSO_I
J PRI
J SPIKE SUPPRESSOR
I _J VALVE DRIVER _ _ TORCST-d A
l SEC
l SPIKE SUPPRESSOR
,, II SPIKE SUPPRESSORI I,_1 ;.t--.-_ sEc/ ,, T
II, -_--I ILL], [_P,K_soPPREss_]
L__
I NOTE
J_ DCS CONTROLLED
Figure 15-26 TSS Functional Block Diagram (Sheet 2 of 2)
15-99
_@ SEDR300 __
PROJECT GEMINI
Power Inverters (PI). All equipment is located in the equipment section of
the ATX)A (Figure 15-25).
oFm A caThe T_S controls the firing of the RCS thrusters to establish the rotational
rates c_nded by a selected control mode. The existing rates in the three
ATDA axes are sensed by three rate gyros in the operating _P. Signals repre-
senting these rates are supplied to the ACE package where they are c_pared with
signals representing the commanded rates for each axis (Figure 15-26). When a
difference exists between an actual rate and a c_..._ndedone, the appropriate
rotation1 c_ud (i.e., pitch up, pitch down, etc. ) is developed to effect a
rate change. This c_ is routed from the ACE to the QAME package where it
is used to develop individual thruster firing c_._;_nds. The firing cu_,ands,
in turn, are trsmsa_tted to the RCS to fire the thrusters. The thrusters con-
tinue to fire until the rate 8yro and rate c_nd signals are of equal magni-
tude. At this time, the thruster firing c_,ands are terminated.
RATE COMMAWnS
The rate c_s, in the BRD control mode, are generated by a voltage divider
within the TSS. The voltage divider generates c_..._ndsonly for the pitch and
roll axes since the yaw axis is rate damped in both control modes.
When the BRD mode is selected, the rate control relays in the system control
relay panel are latched. This causes a 26 vac, 400 cps supply voltage to be
applied to the voltage divider. The outputs of the voltage divider are used as
cc_a_s to establish the pitch and roll rotational rates for the BRD mode.
These c_,_ds are always in phase with the basic 400 cps supply voltage.
15-I00
_ SEDR300 __.__
PROJECT GEMINIS _
Selection of th_ NRD control mode causes the rate control relays In the system
control relay panel to reset. Thus, the excitation voltage is removed from
the voltage divider, and its outputs drop to zero. This results in the ATDA
being rate damped about all three axes.
SECONDARY OPERATION
The TSS has provision for either primary or secondary system operation. Selec-
tion can be made from the ground, via the DCS. For each type of operation,
different RGP and power inverters are utilized. In addition, certain selections
of the ACE and OAME packages are redundant to provide for primary and secondary
operation.
_- ELECTRICAL POWER
The basic operating power for the TSS is 28 vde from the ATDA main bus. This
power is used by the power inserters to produce the 26 vac, 400 cps power for
the system components.
SYSTEM UNITS
RATE GYRO PACKAGE
The two rate gyro packages in the TSS are identical. Each is approximately 3.5
inches by 4 inches by 5.6 inches and weighs about 4 pounds. Each unit con-
tains three individual rate gyros which sense ATDA rotational rates in the
pitch_ roll and yaw axes. The gyros are hermetically-sealed, single-degree-of-
freedom sensors. They receive 26 vac, 400 cps power from the power inverter
and provide outputs to the ACE package and the Instrumentation System. Each
one also has a torquing capability for use during ground checkout. The output
of each gyro is _._.00cps si_;nalwhose amplitude is proportional to the angular
!5_I01
SEDR 300
rate about the sensitive axis. The signal can be either in phase or out of
phase with the supply voltage, depending on the direction of rotation. Upon
leaving the RGP, the rate signal is routed thro_sh a scale factor network which
attenuates it to the proper level for use by the ACE.
A'i'21TUDE COI_i_ROL EI_CTRON!CS
_l_leACE package is approximately 16.3 inches by 8.4 inches by 5.1 inches and
weighs about 17 pounds. It contains mode logic circuits, rate amplifiers (presmps),
switching smplifiers, demodulators, positive and negative switches, minimum pulse
circuits, and a bias power supply. All sections, except the mode logic and rate
smplifiers, are redundant to provide for prima_ 7 or secondary operation.
The ACE package generates the rotational commands to effect necessal_ changes
in ATDA rotational rates. The logic for producing specific commands is developed
by summing the rate gyro and rate command input signals together. _nis is aecon_-
]_ished by routing the two signals to a common summir_ point. If the signals are
of equal magnitude and are directly out of phase, it is indicative that the
sctusl rate is the same as the commanded one. In this instance, the signals
null each other out, amd a zero-level resultant signal is obtained. In cases
_,.herethe rate gyro si_snal has a different magnitude or phasing with respect to
the rate cor_and, either an additive or difference signal will result. If the
resultant signal is greater than a certain level (deadband threshold), it causes
a specific rotational contend to be developed. A resultant signal which is in
phase with the 400 cps supply voltage will cause a negative rotational co_m_'mnd
(pitch down, roll left or yaw left) to be produced. An out-of-phase resultant
sl6nal will cause a positive command (pitch up, roll right or yaw right) to be
produced.
15-102
SEDR300
PROJEC:I" GEMINI
I_ode Logic
_ne mode logic section perfo_s several functions within the ACE. First, it
establishes and maintains the conditions for effecting the rate command oper-
ational mode in the TSS. This basically consists of setting up the proper servo
loop gain and turning on the transistor switches which route the rate signals
from the rate preami01ifiers to the switching amplifiers. In addition, the mode
logic controls the operation of the primary and secondary circuits within the
ACE (except for the bias power supply). This is accomplished by m_plyiD_
operating, voltages only to the selected circuits. Primary and secondary operation
of the bias power supply is controlled directly by a relay in the system control
relay panel.
Amplifier Sections
Each of the three signals resulting from the summing process is routed to a
rate preamplifier. The rate preamplifier amplifies and reverses the phase of
the input signal and supplies it to either the primary or secondary switching
amplifier. The switching amplifier further 8mplifies the signal to a usable
level for the demodulator.
Demodulators
The demodulators discriminate between in-phase and out-of-phase ac signals.
If the input signal is in phase, the demodulator transmits a positive dc signal
to the positive switch. If the input signal is out of phase, the demodulator
transmits a negative dc signal to the negative switch.
/
15-io3
SEDR 300
Posltiveand _gati_ Sw'Atches
The positive switches, when turned on by the demodulator output, provide a
path to ground through transistors in the positive switch circuits. This
connection to ground acts as a logic cc-,,_.mndto the 0AME where it causes the
positive-torquing thrusters to fire. In the same manner, the negative switches
control the firing of the negative-torquing thrusters.
Minimum Pulse Circuits
The outputs of the positive and negative switches for the pitch and roll channels
are also used to trigger minimum pulse circuits. Each of these circuits generates
a pulse which is fed back to the inout of its respective switch circuit. The
pulse causes the switch circuit to be clamped on for the duration of the pulse
(nominally, 18 milliseconds). Thus, when a switch circuit is turned on, it is
held on for at least this duration.
ORBIT ATTITUDE AND MANEUVER _.RCTRONICS
The OAME package is approximately 8 inches by 10.2 inches by 3 inches and weighs
about 8 pounds. It has a removable cover and contains three replaceable module
boards (two relay boards and one cemponent board) as well as fixed components.
Electrically, the OAME consists of primary and secondary valve driver circuits
and a spike suppression circuit for each of the eight thruster solenoid valves.
The OAME converts the logic c_ands from the ACE into individual firing c_nds
for the RCS thrusters. When turned on, the valve driver provides a circuit
ground to actuate a thruster solenoid valve in the RCS. A back-to-back diode/
zener diode spike suppressor for each thruster channel is provided in the OAME
15-Io_
$EDR 300
___ PROJ E--C-T GEMINI
to minimize voltage transients when a solenoid valve is actuated or de-actuated.
POWER INVERTER
The power inverters convert main bus dc power to 26 vac, 400 cps, slngle-phase
power for use by the ACE, RGP and system control relay panel. The two units are
designated primary and secondary and operate as such in the system. Each package
is approximately 5 inches by 5.3 inches by 4 inches and weighs about 7 pounds.
15-105/106
REACTION CONTROL SYSTEM
TABLE OF CONTENTS
T ITLE PAGE
SYSTEM DESCRIPTION ........... 15-109PRESSURANT GROUP 15-11 2
f'_ FUEL/OXIDIZER GROIJP: : : : : : : : : 15-112THRUST CHAMBER ASSEMBLY GROUP .... 15-113
SYSTEM OPERATION • • • 5-11315-113p_s_o,,N,o,oov..! i i i i ! i ! 'FUEL/OXIDIZER GROUP. 15-1 1/-I.
THRUSTCHAMBERASSEMBLYGROUP.... l S-llSSYSTEM UNITS. . . 15-115
PRESSURANT ST(_RAGE I:AI_K: : : : ." . . 15-115"A" PACKAGE ..... . . ..... 15-115PRESSURE REGULATOR ......... 15"1 16"B" PACKA(;E......... . . . . 15-116FUEL TANK...... 15-1 17
15-117ox.o.z._,-r.N.,..... iiiiiiii"C" AND "D" PACKAGES 15-117PROPELLANT SUPPLY SHUTOFF VALVES . . 15-118THRUST CHAMBERASSEMBLY GROUP.... 15-1 18
15 -107
SEDR 300
__D-_-" PROJECT GEMINI ,_,,
NOTETHIS SCHEMATIC TYPICALFOR RINGS A AND B
( L-i• % " "% pc'_ccxxxxxx_ FUEL
PURGE _ OXIDIZER
FUELTANK POR' I..........ETEEcS:'c_JMANUAL TESTVALVE
VALVE
N2 FUEL TANK "B" PACKAGE I PACKAGE
VENTVALVE,..........-.-................................................J...............)J _ ............ _ _ER, ........I "D" PACKAGEBURST RELIEF
DIAPHRAGM FILTER VALVE
OVER,OARO
VENT NORMALLY CLOSEDPYROTECHNIC VALVE
t CHECK ALVERELIEF V
VALVE TEST PORT
PRESSUREREG[ L6.TOR HIGH PRESSURE
l':] | TEST _ _ NORMALLY CLOSED MANUAL
FILTER VALVE el:: PYROTECHNIC VALVE CHARGING VALVE "_X
INLET _ HIGH PRESSURE
I__ !FILTER _ B -_I _q CHARGiNG_ PORT
___%_.::__::_i ,,_.,.c.oJ...........................1_: __"_ ...........),NTERCHEC_I/ _E: P,LTERI Li:Ii_ .I-- I L,_='_--._VALVE TEST RELIEF VALVE I I_MI_
SOURCE SENSOR' I IP,_f _ _;E;_:, _ PRESSUREMANUAL TEST TRANSDUCER
........N20XIDIZERI ,_1 £'4TANK VENT PYROTECHNIC VALVE
_1 I _1" E,LTERII- _ _ ............................................................, .........__.:.:.:.,.n_ ....................................
HIGH PRESSURE MANUAL TEST /"-_
MANUAL CHARGING VALVEVALVE
Figure 15-28 Reaction Control System Schematic (Sheet I of 2)
15-110
/ "_ SEDR 300
PROJECT GEMINI
[] THESE VALVES ARE SET TO THE OPENPOSITION BEFORE LAUNCH AND REMAINOPEN THROUGHOUT THE MISSION
25_ 25l
OFF/ON VALVE []
:: ERSUPPLV•J:_l OFF/ON VALVE
:.;._,_.:.:.:.:......:._'i_._ 25# 25lVALVE
(16 TYP.)
'OXIDIZER VALVE HEATER
(16 TYP)
Figure15-28Reaction Control System Schematic (Sheet2 of2)
15-111
__ SEDR300 ____
PROJEC'T GEMINI
signal from the Sequential System. Ring B is thus activated and it begins
immediately to execute TSS co_m-nds. Ring A is not activated until it is se-
lected. For explanation purposes, the RCS is divided into three groups: the
Pressurant, Fuel/0xidizer (or propellant) and Thrust Chamber Assembly (TCA).
PRESSURANT GROUP
The pressurant group (Figure 15-28) consists of a pressurant tank, "A" package,
pressure regulator and "B" package. Valves and test ports are provided at
accessible points to permit servicing, venting, purging and testing. Filters
are provided throughout the system to prevent contamination. The pressurant is
stored and isolated from the remainder of the system by a normAlly-closed pyro-
technic actuated valve, located in the "A" package.
FUEL/OXIDIZER GROUP
The fuel/oxidizer (propellant) group (Figure 15-28) consists of expulsion bladder
storage tanks, "C" (oxidizer) and "D" (fuel) packages. Valves, ports and test
ports are provided at accessible areas to permit servicing, venting, purging
and testing. Filters are provided throughout the system to prevent contami-
nation. The propellants are isolated in the storage tanks from the remainder
of the system by normally-closed pyrotechnic actuated valves in the "C" and "D"
packages. Heaters are installed on the packages to maintain an operating tempera-
ture. The propellants used are:
Oxidizer - Nitrogen Tetroxide (N294) conforming to
Specification MIL - P - 26539A
15-112
___ $EDR300 _.___
PROJECT GEMINIf_
_EL - Monomethyl Hydrazine (N2H3CH3) confroming
to Specification MIL - P - 27403
THRUST _ ASSEMBLY GROUP
The TCA group (Figure 15-28) consists of eight 25-pound TCA's used for attitude
(rot!, pitch and yaw) control of the ATDA. Each TCA is equipped with two solenoid
valves which open simultaneously. Heaters, installed on the oxidizer solenoid
valves, maintain the oxidizer at an operating temperature.
SYSTmOPERATION
RCS thruster operation is controlled by signals from the TSS. The RCS (Figure
15-28) is activated by a signal from the Sequential System. The RCS, in con-
_-_ Junction with the TSS, provides attitude control of the target vehicle in the
roll, pitch and yaw axes.
Ring B of the RCS module is selected during pre-launch by a DCS commaud to pro-
vide attitude control and is activated when power is applied at ATDA/Launeh
Vehicle separation plus eight seconds. Selection of ring A, in the event of a
malfunction in ring B, may be accomplished by a DCS command.
The operation of the RCS is described in three groups, Pressurant, Fuel/Oxidizer
and Thrust Chamber Assembly.
PRESSURANT GROUP
(Figure 15-28) High pressure nitrogen (N2) pressurant is stored at 3000 psi in the
pressurant tank. The tank i_ serviced through the "A" package high pres-
F_ sure gas charging port. Pressure from the pressurant tank is isolated from the
15-113
__ SEDR300 _-_r
PROJECT GEMINI
remainder of the system by a normally-closed pyrotechnic actuated valve located
in the "A" package. Stored nitrogen pressure is monitored by the source pres-
sure transducer located in the "A" package and relayed to the ground station as
instrumentation data. Upon command, the "A" package pyrotechnic actuated valve
is opened and nitrogen flows to the pressure regulator and "B" package. The
"B" package routes the pressure to both propellant tanks. The regulated pressure
is sensed by the pressure transducer ("B" package) and supplies an output to
the Instrumentation System indicating pressure downstream of the regulator. The
check valves prevent backflow of propellant vapors into the pressurant system.
The "B" package has a safety feature to prevent over pressure of the fuel and
oxidizer tank bladders. Should the system be over pressurized downstream of
the regulator, the pressure would first rupture the burst diaphragms, then be
vented overboard through the relief valves.
FUEL/OXXDIZERGROUP
Fuel and oxidizer propellants are stored in their respective tanks, and are
serviced through the high pressure charging ports in the "C" and "D" packages.
The propellants are isolated from the remainder of the system by a normally
closed pyrotechnic valves in the "C" and "D" packages. Upon command, the "C"
(oxidizer) and "D" (fuel) package pyrotechnic actuated valves are opened and
propellants are distributed through their separate tubing manifold system to
the thrust chamber inlet solenoid valves.
Two motor operated valves are located in the propellant feed lines, upstream of
the TCA's. These valves are opened prior to launch and will remain open through-
out the mission. The valves are not connected electrically. Heaters are _
15-Ii_
$EDR 300
connected to the output lines of the "C" and "D" packages to prevent freeze-up
and are turned on prior to launch by the RCS HRTS switch.
THRUST CHAMBER AS_MBLY GROUP
The TCA fuel and oxidizer solenoid valves are controlled by signals from the TSS.
In response to the signals, the valves open and propellants are directed through
small injector Jets into the thrust chamber. The fuel and oxidizer impinge on
one another, where they ignite hypergolically to burn and create thrust. Heaters
are connected to each TCA oxidizer solenoid valve to prevent freeze-up and are
turned on at launch by the RCS HRTS switch.
SYSTEM UNITS
PRESSURANT STORAGE TANK
The nitrogen (N2) pressurant is stored in a welded, titanium spherical tank. The
tank is 7.25 inches in diameter and has an internal volume of 185.0 cubic inches.
Nitrogen gas is stored at 3000 psi and held therein by the "A" package pyrotech-
nic valve until released by a sequential signal. This nitrogen under pressure
is used to expel the fuel and oxidizer from their respective tanks.
"A" PACKAGE
The "A" package consists of a pressure transducer, isolation valve, filters and
two high pressure gas charging valves. The pressure transducer monitors the
source pressure and supplies an output to the Instrumentation System, indicating
the pressure of the stored gas. The normally closed isolation valve is used to
isolate the pressure from the remainder of the system.
15-_5
sEoR300PROJECT GEMINI
The valve is pyrotechnically opened to activate the system for operation. Two
dual seal, high pressure gas charging valves and ports are provided, on on each
side of the isolation valve. The upstream valve is used for servicing, venting
and purging the pressurant tank, while the downstream valve is used to test
downstream components. Filters are installed to prevent contaminants from
entering the system.
PRES_JRE REGUIATOR
The pressure regulator is conventional, mechanical-pneumatic type. The regu-
lator reduces the source pressure to system operating pressure. An inlet filter
reduces any contamination in the gas to an acceptable level.
"B" PACKAGE
The "B" package consists of filters, pressure transducer, three check valves,
two burst diaphragms, two relief valves, regulator output test port, fuel tank
vent valve, oxidizer vent valve, check valve test port and two relief valve test
ports. The inlet filter reduces contam_uation in the gas to an acceptable level.
The relief valve filters prevent particals from a ruptured burst diaphragm from
entering the relief valves. Test port filters prevent contaminants from entering
the system through the test ports. The pressure transducer monitors the regu-
lated pressure and supplies an output to the Instrumentation System. Check
valves prevent backflow of fuel or oxidizer into the pressurant system. The
burst diaphragms are safety devices that rapture when the pressure reaches the
design failure pressure, thus, prevents imposing excessive pressure on the pro-
pellant bladders.
The two relief valves are conventioan mechanical-pneumatic type with the pre-
15-n6
___ SEDR300 __
PROJECT GEMINIf-,.
set opening pressure. In the event of burst diaphragm rupture, the relief valve
opens to vent excess pressure overboard. The valve reseats when a safe level
is reached. This prevents venting the entire gas source. Manual valves and
ports are provided to vent, purge and test the regulated system.
FFEL TANK
The fuel tank is a welded, titanium cylindrical tank which contains an internal
bladder and purge port. The tank is 5.10 inches outside diameter, 30.7 inches
in length and has a fluid volume capacity of 546.0 cubic inches. The nitrogen
pressurant is imposed on the exterior of the bladder to expel fuel through the
"D" package to the TCA solenoid valves. The purge port is proveded to purge and
vent the fuel tank bladder.
OXIDIZER TANK
The oxidizer tank is a welded, titanium cylindrical tank which contains a bladder
and purge port. The tank is 5.10 inches outside diameter, 25.2 inches in length
and has a fluid volume capacity of 439.0 cubic inches. The bladder is a double
layered Teflon, positive expulsion type. The nitrogen pressurant is imposed on
the exterior of the bladder to expel the oxidizer through the "C" package to the
TCA solenoid valve. The purge port is provided for purging and venting the
oxidizer tank bladder.
"C" AND _'D"PACKAGES
The "C" and "D" packages are identical in operation and are located downstream of
the tanks of their respective system. Each package consists of filters, an isola-
F" tion valve, propellant charging valve and test valve. The in llne filter reduces
15-i17
___ SEDR 300 __
PROJECT GEMINI
partical contamination to an acceptable level. The valve filters prevent con-
taminants from entering the system. The normally closed isolation valve is
used to isolate propellants from the remainder of the system until ready for
use. The isolation valve is opened pyrotechnically to activate the system.
The propellant charging valve is located upstream of the isolation valve and
is used for sel_iclng and venting the system. A test valve Is used to test
the system, downstream of the isolation valve.
PROPELLANT SUPPLY SHUTOFF VALVES
Motor operated propellant supply shutoff valves are located downstream of the
"C" and "D" packages. The valves are opened during pre-launch and are not con-
nected electrically.
THRUST CHAMBER ASSEMBLY GROUP
Each TCA consists of two propellant valves, injection system, calibrated orifices,
combustion chamber and expansion nozzle. The fuel and oxidizer solenoid valves
are the normally-closed quick acting type_ which open simultaneously upon
command. The action permits fuel and oxidizer flow into the injector system.
The injectors use precise Jets to impinge fuel and oxidizer streams on one
another for controlled mixing and combustion. The calibrated orifices are
fixed devices used to control propellant flow. Hypergollc ignition occurs in
the combustion chamber. The combustion chamber and expansion nozzle are lined
_ith ablative materials and insulation to absorb and dissipate heat and control
external wall temperature. TCA's are installed within the RCS module mold line,
with the nozzles terminating flush with the outer mold llne. TCA's are located
at fixed points on the RCS module in a location suitable for attitude control. -_
15-i18
__ SEDR 300
PROJECT GEMINI
Thermostatically controlled electric heaters, located at various points in the
system, are used to prevent propellant from freezing.
f_
EXTRA-VEHICULARACTIVITY
XVITABLE OF CONTENTS
TITLE PAGE
SYSTEM DESCRIPTION ............................... 16-3
SYSTEM OPERATION ................................. 16-6
ELSSCHESTPACK OPERATION ....................... 16-6
HAND HELD MANEUVERING
UNiT OPERATION ..................................... 16-15
SYSTEM UNITS ........................................... 16-17
ELSSCHESTPACK ........................................ 16-17
EILSS25 FOOT UMBILICAL AND "i:!:i:i:i:i:i:i:i:i:!:i:i:!:i:i:i:i:EILECTRICAL JUMPER 16-18 "-"-'-'-'-"-'-'-'-'-'-""""................................. !_!_i!i!ii_i_i_!_!i!!i!_iii_!iii_i
ELSS50 FOOT UMBILICAL ............................ 16-20 iii_i!iiiii!iiiii!_ii_iiiiii_i!!ii!iii_ii_iiHAND HELD MANEUVERING UNIT ................ 16-20 :::::::::::::::::::::::::::::::::::::
:.:.:.:.:.:-:.:-:-:-:-:-:.'-:-:-:-:.',
GASEOUS NITROGEN MODULE .................... 16-22 :::::::::::::::::::::::::::::::::::::i:i:i:i:i:!:i:i:_:_:i:_:i:i:i:i:i:_:i
EVA HANDRAILS ........................................ 16-23 ":':':':':':':':':':':':':':':':':'::::::::::::::::::::::::::::::::::::::::::::::::::::::::HANDHOLDS AND FOOTREST ........ 16-23 ""'""'"'""'"""""""• • • • • * • • • • * • .'.-.-.°.',*.%%%%'.%°.%'.'.'*'.• • .:.:.:.:°:.:.:.:.:.:.:.:.:.:.:.:.:.:
-.- .%-.....-...-.o.*.-.-,..-.-...o.
/_ ::::::::::::::::::::::::::::::::::::-.-.-o-..,-.o°o.°.,°-°-.-.-.-°o.*...:.:-:.:-:-:-:-:.:-:-:.:.:.:-:-:-:-:-.:.:.:.:.:.:.:.:-:,:-:-:.:-:.:.:.:.:....°.*........°.°......o.°...*.°.:.:°::.:.:.:.::.:::.:.:.:.:.:.:.:_:i:i:i:i:i:i:!:!:!:i:!:i:i:_:_:_:!:::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::- °.°.*o-o.°-,.°-.O.-oO..°-°-..°O..o:.:.:.:.:-:-:-:-:°:.:.:-:.:-:.:-:,:-
]L@- ] ":':':':':':':':':':':':':':':':':
._.. SEDR 300
Figure 16-1 Typical Use of EVA Equipment
16-2
SEDR 300
SECTION XVI EX_HICU_R ACTIV!_£
SYS_H DESCRIPTION
_xtrs_ehiculsr Activity (EW!) is that portion of the mission during which the
pilot egresses from the cabin to conduct experiments and perform maneuvers.
The purpose of EVA is to: determine man's performance capabilities wearing a
pressurized suit in a free space environment, demonstrate the ability to per-
form controlled maneuvers in space independent of the spacecraft, and conduct
experiments outside the spacecraft. EVA is planned for spa cecrs_ 8 through
12 missions.
EVA is accomplished usi_ the E_ra-vehicu_r Life Support System in conjunction
with: the Extra-vehicu_r Li_ S_port Package on spacecraft 8; the Hand Held
Maneuvering Unit on spacecraft _ and ll; and the Modular Maneuvering U_t on
spacecraft 9 and 12. The ModularNaneuveri_ Unit is an expe_ment; therefore
it is not included in the EVA System description. No EVA was accomp_shed on
spacecraft 8 which utilized the Extra-vehicular Life Support Package; therefore,
information pertaining to that configuration is not included.
The Extra-vehicular _fe Support System (EI.SS) is the prima_ component of the
EVA System on each spacecraft mission. The E_S pro_des environmental control
and s_plies the necessary electrical and tether connections to support the
pilot during the entire EVA. The E_S consists basical_ of a chest pack,
umbilical_ and electrical jumper. The E_S umbilical is modified for spacecraft
l0 and ll to include a nitrogen hose to s_p_ propel_nt to the Hand He_ _neu-
vering Unit (HHMU). The NE_ is used to maneuver through space. It provides
_ forward or reverse thrust as re_ired. Typical use of EVA e_ipment is illus-
trated in Figure 16-1.
16-3
_.._. SEDR 300
STOWAGE BOX
STOWAGE
, \//
//
j// UMBILICAL
\ (s/c loa it)
\
N '(S/C 9 A 12) ,'_';_
l,_.].....OVERHEAD SWITCH AND ,I,.0.
CIRCUIT BREAKER PANEL I ,v._°'.A,s,,_
(S/C 9
COMMAND PILOT MAIN
INSTRUMENT PANEL /_'EVA 02 DISCONNECT
(S/C 8, 10 & 11) __
•_'=_'_ RIGHTSWITCHAND CIRCUIT
: _': BREAKER PANEL (S/C 8, 10 & 11)
LSUIT TEMPERATURE PANEL
Figure 16-2 Typical Locations of EVA Equipment and Support Provisions (Sheet 1 of 2)
16-4
.._.%_. SEDR 300
L_,_;--- PROJECT GEMINI
DIS
RELEASEASSEMBLY
NITROGEN DISCONNECT
PANEL (S/C 10 & 11)
HAND RAILS
HOLD(2 PLACES)
_--THERMAL COVER
HHMU NITROGEN MODULE(S/C 10 & 111
Figure 16-2 Typical Locations of EVA Equipment and Support Provisions (Sheet 2 of 2)
16-5
SEDR 300
EVA support provisions consist of handrails, handholds, footrest, lights and
tether attach points. These aid the pilot in maneuvering to the rear of the
adapter and in donning the back pack or performing experiments in that area.
Figure 16-2 illustrates the typical location of EVA equipment and support
provisions.
SYSTEM OPERATION
The EL$S provides suit pressurization and environmental control during EVA.
Oxygen, from the spacecraft primary oxygen supply, is supplied to the ELSS
chest pack via the umbilical. The umbilical oxygen attach point is located
on the center pedestal in the spacecraft cabin. Oxygen flow to the umbilical
is controlled by the REPRESS valve on the suit temperature panel. On spacecraft
missions utilizing a back pack, a limited oxygen supply in the back pack replaces
the spacecraft oxygen supply to permit farther excursions from the spacecraft.
In addition to supplying oxygen to the chest pack, the umbilical provided voice
communications and telemetry information, as well as a tether restraint. The
modified ELSS umbilical used on spacecraft i0 and 11 also includes a propellant
hose for the HHMU. The electrical jumper provides a single connecting point
for the chest pack and pressure suit to the umbilical.
ELSS CHEST PACK OPERATION
O_gen Loo_ - Normal Operation
Oxygen enters the chest pack at the spacecraft UMBILICAL 02 CONNECTOR or the
02 CONNECTOR. The connectors are self-sealing quick disconnects. In
either case the oxygen is supplied through check valves to a flow control .....
16-6
___ SEDR300
PROJEC'-'C'T-'GEM IN I,f.--,
valve. Figure 16-3 is a functional diagram of the ELSS chest pack and Figure
16-4 illustrates its physical construction and external controls and indicators.
Oxygen flow to the suit loop is manually controlled by the 02 FLOW SELECTOR
valve. The selector valve has three positions: OFF, _DIUM (restricted flow),
and HIGH (full open). A push-to-turn motion is required when changing the
valve position. The valve handle and valve position indications are illuminated
for use during darkness. The selector valve is adjusted to give maximum comfort.
Oxygen enters the suit loop at the ejector. The ejector supplies the means of
circulating the flow through the suit loop. The energy of the expanding high-
pressure oxygen (92 psia) from the selector valve provides the required energy
.... for circulation of the low-pressure (3.7 psia) gas from the evaporator-condenser.
The oxygen entering the ejector is heated by two i0 watt heaters connected in
series. One of the heaters is wrapped around the ejector throat and the other
is wrapped around the inlet line to the ejector. The heaters are energized when
power is applied to the chest pack.
Suit pressure is maintained at 3.7 psia by the suit outflow valve which vents a
portion of the exhausted suit gas overboard. In addition to maintaining Suit
pressure, the suit outflow valve accomplishes two additional functions: suf-
ficient carbon dioxide is vented to maintain a safe level within the suit loop
and a portion of the heat load is removed with the vented exhausted suit gas.
The remaining exhausted suit gas enters the evaporator-condenser to be cooled.
The evaporator-condenser is a temperature-controlled heat exchanger. Cooling
is accomplished by transferring heat to water stored in the integral metal
16-7
T_ SEDR 300
"__ PROJECT GEMINI
VENT VENT
_ Ev_P__U_T'o°_w_ SO,TOUTFLOWVALVE
TRANSDUCER
CONTROLVALVE ' _"" '' !;_
OXYGEN SUPPLY J . L_'I
CHECK _
VALVE =r_
' " " " " " ":" " " " " "" HOSE SUIT DUAL
J FLOW - - "_ • i' ' "_ CONNECTORSCONNECTORS
PRESSURE - -
• , • " , REGULATOR
CHECK MANUAL ii
LINE VALVEHEATER BYPASS _:
VALVE _;'_
TEMPERATURE i_ TO SPACECRAFT
SENSOR ;:il SU,TCIRCUITREDUCER -_
RELIEFVALVE - '_-_
IHECK "_
EJECTOR
VALVE VALVE "
FILLVALVE
ELOWCONTRO,_-- -- - _'VALVE L
FILLPORT MMU 02 UMBILICAL 02 J[_"".::i:.JJ%L_-c_
CONNECTOR CONNECTOR _ j
Figure 16-3 ELSS (]best Pack Functional Diagram
16-8
SEDR 300
ELECTRICALCONNECTOR
MANUAL SHUTOFFVALVE
Figure 16-4 ELSS Chest Pack
16-9
___ SEDR3OO __
PROJECT GEMINI
wicks of the heat exchanger. As the gas is cooled the water vapor in the
e_musted suit gas condenses and is separated and removed with evaporant _,_cks
by capillary action. The condensed water vapor is passed through the evaporant
wicks to the evaporative heat sink. %lue evaporative heat sink uses the condensed
water vapor as a coolant thereby reducing the initial amount of water which must
be stored in the heat sink.
The outlet temperature of the gas leav%ng the evaporator-condeL_ser is coi_rolled
by the EVAPORA_ FLOW CONTROL valve. A temperature sensor at the evaporator-
condenser outlet controls the operation of the EVAPO[@,}_' FLOW COntrOL valve to
maintain the temperature at a nominal 45 degrees Fahrenheit. An increase in
outlet gas temperature opens the flow control valve, lowering the boili1_ point
of the stored water, thus allowing greater heat dissipation in the heat exchanger.
The EVAPORANT FLO_T CONTROL valve is shut-off until the ELSS is in operation to
prevent loss of the water in the heat sink. If suit temperature falls to an un-
comfortable level during operatio_ the valve ma_ be closed.
Oxygen Loop - Emergency Operation
There are three conditions that are considered as emergencies. These conditions
are: loss of external oxygen supp]$ pressure, low suit pressure, and heat
exchanger failure. Each of these three conditions is cause for aborting the EVA°
The first two conditions are automagical]_ compensated for; while the third_
which is not as critical 7 requires manual action by the pilot.
In the event the external pressure falls below 67 + i0 psig (reference to
ambient conditions), the emergency pressure reducer automatically actuates
16-io
___ SEDR 300 _._.__
PROJECT--'GEMINI
to supply emergency oxygen to the ejector. The emergency oxygen supply has
the capacity of delivering oxygen to the ejector for a period of 16 to 27
minutes, depending upon the setting of the 02 FLOW SELECTOR valve. The pres-
sure reducer has a relief feature which will vent the emergency oxygen supply
overboard should the pressure reducer fail in the open position. Sufficient
oxygen will he vented to avoid pressurizing the suit loop above 5.0 pslg.
A pressure gage on the instrument panel allows monitoring of the emergency
oxygen supply pressure. The emergency oxygen, upon leaving the storage tank,
passes through a manual shutoff valve, a pressure reducer, and a line heater.
A temperature sensor located downstream of the heater is used in conjunction
with a temperature control circuit to maintain the emergency oxygen at a temp-
erature of 45 degrees Fahrenheit. A flow sensor in the emergency oxygen line
initiates an audio and visual alarm when the emergency oxygen supply is being
used.
A demand type suit-pressure regulator in parallel with the BYPASS-NORMAL and
02 FLOW SELECTOR valves opens in the event suit pressure decays below B.B psia.
Oxygen flow through the suit-pressure regulator is from the same supply which
is supplying the suit loop via the ejector. The suit-pressure regulator will
maintain suit pressure at B.B psia. A flow sensor in series with the regulator
initiates an audio and visual alarm when flow through the regulator is sensed.
A failure of the evaporator-condenser will cause a considerable rise in suit
inlet temperature. If, after increasing the 02 FLOW SELECTOR valve to HIGH,
sufficient cooling in the suit loop is no_ obtained, the BYPASS-NORMAL valve
is placed in the BYPASS position. This increases the flow of oxygen to the
16-11
so30oPROJECT GEMINI
suit loop. The BYPASS-NORMAL valve is the push-to-open and push-to-close
type. Valve position is illumlnated for use during darkness.
Electrical Operation
The electrical portion of chest pack consists of a temperature control and
oscillator module, oscillator and light controller module, battery, relay,
heaters, and associated switches -nd indicators. Figure 16-5 is a schematic
of the chest pack electrical comgonents.
The chest pack operates from battery power or external power from the spacecraft
supplied throug_h the umbilical. During normal operation external power is
utilized. A control relay in the chest pack is used to select internal or
external power. The relay is energized through the umbilical and supplies ....
external power to the chest pack when the spacecraft ELSS PWR circuit breaker
is closed. When the relay is de-energized battery power is utilized and the
EIZS PWR indicator on the chest pack instrument panel illuminates.
The battery is the silver-zinc alkaline type using potassium hydroxide as the
electrolyte. The battery consists of a group of series-parallel cells con-
nected to provide a nominal load voltage of 2_ v_c. Regardless of the power
used, the chest pack circuits are not powered until the battery power switch,
located underneath the access door on the chest pack, is placed in the ON
position.
The oscillator and light controller module consists of a voltage regulator and
a logic control circuit. The voltage regulator supplies the necessary 28 vdc
and i_ vdc re_11_te_ voltages. The logic control circuit consists of a
16-12
PROJECT GEMINIi
flip-flop and an _ gate. _e OR gate obtains trigger inputs fr_ the _it pressure
regulator and emergen_ o_gen f!_ sensors. Upon receipt of one of the trigger
_ts, the _ gate causes the flip-flop to change states and p_vide an out_t
to the audio oscillator _ the temperature control and oscil_tor module. _e
ala_ c_ be reset by _press_g the AUDIO pushb_tton on the control panel which
causes the flip-flop to reset.
_e tem_ra_re control and oscillator mo_leprovides control for the _ergency!
o_gen line heater and the audible and visible warning devices. _e emergency
o_gen l_e heater is contro_ed by the temperat_e control circuitry obtaining
in_ts from the emergency ox_en temperature sensor and emergency oxygen fl_
sensor. _en the tem_rat_e sensor indicates a t_pera_re bel_ 45 degrees
Fahre_eit, and a triter sisal fr_ the eme_ency o_gen fl_ sensor exists, the
temperature control circuit a_s power to be supplied to the line heater. _en
the temperature at the sensor exceeds _5 degrees Fahre_eit, the lye heater is
deact_ated. A sisal represent_g a tempera_re in excess of 45 degrees Fa_enheit
and an emergency o_gen fl_ sisal have to be p_sent before the Ene heater is
activated. _e EMERG 02 li_t is also i_uminated by the emergency o_gen fl_
sisal. _e oscillator is activated by the outer from the flip-fl_ of the
osci_ator and light contro_er and wi_ supp_ an audible ala_ to _e pilots
headset and to the s_cecraft via the umbilical.
The suit-pressure regulator flow sensor signal, in addition to triggering the OR
gate of the oscillator and light controller, Is used to illuminate the SUIT PRESS
indicator light.
The pressure transducer in the suit loop monitors suit pressure. The output is
].6-].4
$EDR 300
PROJECT GEMINI
returned to the spacecraft for transmission as telemetry information. The
pressure transducer is a variable reluctance transducer with an output of
0 to 5 vdc with a sensed pressure of 2.5 to 5 psig.
The wrap-on heaters at the ejector provides 20 watts of power to maintain the
suit loop inlet temperature at 45 + lO degrees Fahrenheit. When the battery
switch is placed in the ON position, power is applied to the nap-on heaters.
The panel illumination lights are turned on and off by the battery switch.
A three position switch (TEST, DIM, BRIGHT) controls the intensity of the panel
lights and the emergency indicator lights. In the TEST position, the panel
lights, emergency indicator lights, and the warning alarm audio oscillator are
_ tested. The voltage regulator provides the reduced voltage for the low-
intensity position of the light switch.
The following light indicators on the instrument panel are not operational
unless the experimental back pack is connected: H202, RCS, 02 PRESS, and FUEL
PRESS. The H202 pressure indicator is operational when either of the back
packs are connected.
HAND HELD MANEUVERING UNIT OPERATION
The HHMU (Figure 16-6) provides the necessary propulsion required to maneuver
in space. The HHMU has two tractor nozzles and a larger pusher nozzle. The
HHMU provides two pounds of thrust in the positive or braking direction.
The pusher nozzle is attached to the major assembly and is used to supply
the braking thrust. The two tractor nozzles are on arms which extend out on
each side. This positions the nozzles for thrusting in the positive or
forward direction.
16-15
SEDR 300
PROJECT GEMINI
BRAKING THRUST
DETAIL A
TRACTOR NOZZEL-
_JRROFtJLSION TRIGGER i i
(SEE DETAIL A) _: ]
TRACTOR NOZZLE
N 2
N 2 SHUTOFF VALVE
Figure16-6 Hand Held Maneuvering Unit
16-16
SEDR 300
Be trigger is pivoted in the ce_er so the tractor nozzles or the pusher nozzle
can be co_rolled with one hand. Squeezi_ the lower end of the trigger s_plies
nitrogen to the pusher nozzle and squeezi_ the upper end of the trigger s_plies
nitrogen to the tractor nozzles.
_e propulsion gas for the _,_ is provided by two bottles of gaseous nitrogen
that are mounted in the adapter equ_ment section. _e gas is routed to the
adapter skin through metal tubing. A quick disconnect connector and a lever
shutoff valve are located on the outside of the ada_er and can be reached by
opening a small door on the spacecraft skin.
SYS_M _'II_
E_S _ST PACK
_e ELSS chest pack is 18 by lO by 6 inches and weighs 42 pounds when _i_
charged. _erati_ de_ces on the chest pack used by the pilot include four
valves, three switches, two meters, seven indicators, and five eo_ectors. _e
chest pack housing is mo_ed from fiberglass with a volar finish.
_en in use, the chest pack is positioned on the pilots chest (occ_ying an
area from above the thins to be_w the chin), and the control panels are readi_
visible and accessible. _e chest pack is held in p_ce by a self-adheri_ we_
be_ attached to the pints personal harness. The be_ ends are lald across
strips of the same self-adhering, web-belt material affixed to the chest pack.
Two dual o_gen connectors are prodded with the chest pack to pe_it continuous
o_gen flow to the pilots pressure suit duri_ the egress and ingress preparation
procedures. _e dual connectors are stowed in the left hand a_ stow_e container.
16-17
SEDR300 _.___]PROJECT GEMINI
EY_SS 25 FOOT UMBIIICA_ AND ELECTRICAL JUMPER
The 25 foot umbilical (Figure 16-7) is an aluminized mylar and nylon covered
assembly that contains the pilots oxygen supply hose, electrical leads and
restraint tether.
The restraint tether consists of a flat nylon ribbon with a minimum breaking
strength of 900 pounds. The tether is connected by snap rings to the right
egress bar, located at the forward end of the right hatch opening, and to the
D-ring on the pilots harness.
Electrical leads through the ,_,_ilical are wrapped around the oxygen hose and
ca_i_¥ the bio-med instrumentation and c¢--,mmication signals, and spacecraftr_
power to the chest pack. The umbilical is protected in passing over the hatch
opening by an umbilical cor_ guard.
The oxygen supply hose routes spacecraft primary oxygen to the chest pack. It
connects to the d_sconnect on the suit temperature panel in the spacecraft
cabin. The oxygen supply hose is disconnected when the back pack is connected.
Oxygen is supplied from the supply tank contained in the back pack.
The electrical jumper provides a single connecting point between the pressure
suit and chest pack to the umbilical. The electrical Jumper interconnects the
signals passing between the chest pack, pressure suit, and spacecraft.
The 25 foot umbilical, electrical Jumper, and umbilical cord guard are stowed
in the left hand aft stowage container. The 25 foot umbilical, electrical
Jumper, and umbilical cord guarcl are used on spacecraft 8, 9 and 12 missions.
16-18
__ SEDR300 __
PROJECT GEMINIpf-_-\
P1 MATES WITH SUIT ELECTRICAL CONNECTOR SPACECRAFT
34 33 17 14 4 36 35 20 37 18 15 5 16 10 II 25 24 23
ELECTRICAL JUMPER
JI
34 33 17 14 4 36 35 20 37 18 29 28 30 15 5 16 11 25 24 23 22 21 10 31 32 27 13 26 9 12 3 7 2 19 6 8
P1i
ONLY
25 FOOT UMBILICAL
ELECTRICAL SPARES
NOT USED , • •
34 33 17 14 4 36 35 20 37 18 29 28 30 15 5 16 IO 24 23 I1 25 7 8 32 31 27 26 19 9 I
I:'2 MATES WIT H SPACECRAFT ELECTRICAL CONNEC[O_
Figure 16-7 ELSS 25 Foot Umbilical and Electrical Jumper
16-19
SEDR300 _._PROJECT GEMINI
ELSS 50 FOOT UMBILICAL
The 50 foot umbilical (Figure 16-8) is an aluminized mylar and nylon covered
assembly that contains the pilots oxygen supply hose, EH_ nitrogen hose, electri-
cal leads and restraint tether.
The restraint tether consists of a flat nylon ribbon with a minimum breaking
strength of 900 pounds. The tether is connected by a snap ring to the egress
bar in front of the right seat and by a harness clamp to the pilots harness.
Electrical leads through the umbilical are wrapped around the hoses and carry
the bio-med instrumentation and communication signals. _ne cabling monitors
an electrocardiogram and impedance pneumograph. Communications are established
through a left and right microphone and a left and right earphone. Electrical
power is supplied in + lO vdc and -lO vdc along with power for chest pack opera-
tion. The umbilical is protected in passing over the hatch opening by a nylon
chaffing guard. The oxygen supply hose connects to the oxygen disconnect on the
suit temperature panel in the spacecraft and to the bZMBILICAL 02 CON_CTOR on
the chest pack. The nitrogen hose connections are the quick disconnect aft of
the pilots hatch and the HHMU.
The 50 foot umbilical is stored in the left hand footwell stowage pouch. The
50 foot umbilicsl is used on spacecraft i0 and ii missions.
HAND I_LD MANEUVERING UNIT
The }IHMU provides the necessary propulsion required to maneuver in space. The
HHMU is stowed in the left hand aft s+owage container. The HI_dU is used with ....
16 -20
I___ PROJECT GEMINI " _.=-_
N 2 CONNECTOR TO HHMU'_
To__ECRAE, E_C,RICA' MATE_W,TH_.E_,_ACK E'ECTR.C_CONNECTOR ELECTRICAL CONNECTOR CONNECTOR
TO SPACECRAFT _. TO SUIT
18 15 16 t2 4 7 9 3 14 /
ELECTP,_CAL SPAEES
29 --EXTERNAL POWER
13
28 -- EXTERNAL POWER ...........
12
23 P,'_.IKE i 23
24 RMIKE 24
11 LMIKE _" 11
10 -- 25
10
35 --R _A_ H_ADS_T _ 35
37 _EAE.EAOSET--,k--_, '×,:_:--L 3,20 -- U EAR HEADSET--_--_ ,_ _ 20
36 --R EARHEADSET_ _"_ _ l 3618 I 18
33 ------ EKG NO, 2 _ _" ------ /(_ 33L J i J
34 --- --EKG NO, 2 L _ 34
17 17
--IMPEDANCE PNELIMOGRAPH L"_ i _ 414
.... r i
5 _IOVDC I I 515 IOVDC ........ b I 15
--10VDC RETURN _-- 16
"1" "I2_ 3o
_- MATES WiTH S/C MATES WiTH SUIT t_' ELECTRICAL CONNECTORELECTELCA_.CONNECTOR
Figurel6-8 ELSS 50 Foot Umbilical
16-21
SEDR300
the 50 foot umbilical on spacecra_ lO and Ii.
The H_ is co_tructed of fiberglass reinforced with po_ester resin. The HHMU
is _pro_mate_ _ in_es lo_, 6 in_es wide, and 1 in_ thick. A pivoted
trigger controls the operation of the two valves. The pusher nozzle is moused
on the main body of We HH_. The two tractor nozzles are on 14 inch a_ assem-
blies which swivel aw_ from the forward end of the HHMU.
GASEO_NITROGEN MODULE
The gaseous nitrogen module (Figure _-2) is bracket mounted in the _acecraft
adapter section. The modu_ consists of two bottles, a pressure regulator,
and fi_ va_e. Nitrogen from _e bottles is routed to a _ick disconnect and
shutoff valve located aft of the right hatch by a _arter inch metal tubing.
An access door (Figure _-2) protects the disconnect and valve duri_ the _unch
phase. The door is _ened _ compressi_ the _ri_ latch assemb_. Movement
of the door past the ve_ical position dise_a_emthe hinges for removal]'
Each nitrogen bottle will hold up to 438 cubic inches of gaseous nitroge_
providi_ a system c_aci_ of 8_ cubic inches, whi_ represent ll.5 pou_s
of nitrogen (_.75 pounds of the _trogen is usab_ for the pilot). The nitrogen
+i0
is stored at a pressure 5000 to 5500 pslg and is regulated to _0 -15 pslg by
a built-in reg_ator on the b_tle. The press_e at the HHMU will be appro_-
mate_ l_ psig. The system has sufficient capacity for a little over five
minutes of co_inuous usage.
16- 22
SEDR 300
EVA HANDRAILS
The two EVA handrails are located along the top of the adapter and aid the pilot
in maneuvering from the spacecraft cabin to the rear of the adapter. The hand-
rails are illustrated in Figure 16-9.
Launch vehlcle-spacecraft sel_Lration releases the handrail on the adapter equip-
ment section. The handrail is held in the stowed position by a release pin.
The release pin is physically attached to the launch vehicle. At separation,
the release pin is pulled away with the launch vehicle allo_ringthe handrail
to move slightly aft. This a]_lowsthe handrail to clear the spacecraft retain-
Ing hooks and self-extend by spring force. Upon full extension the handrail is
-_'t"_ locked in the extended positicm.
The forward handrail is extended manually by the pilot when he depresses the
release button. Depressing the release button releases the stop and all_s the
handrail to move slightly forward. This forward movement allows the handrail
to clear the retaining hooks emd self-extend by spring force.
An umbilical guide (Figure 16--2)is located at the rear of the handrail on the
adapter equipment section. The umbilical guide assembly prevents damage to the
umbilical in passing over the edge of the adapter. The umbilical guide assembly
is extended when the handholds and footrest in the adapter are extended.
HANDHOLDS AND FOOTREST
The handholds and a footrest are located in the adapter on spacecraft 8, 9, ll,f
and 12 mlssions_ These provide a means for the pilot to hold on and remain in
16-23
__ SEDR300 __
PROJECT GEMINI
HANDRAIL RELEASEBUTTON
(TYPICAL 2 PLACES)
STOWE D POSITION
////_ t
:2EXTENDED POSITION
N DRAI L RELEASE
. BUTTON
__'::::_ IRETAIN]NG BOLT
MATING LINESTOWED POSITION
SPACECRAFT ----_'1_r/
- _ RETAINER kAUNC H VEHICLE /_- RETAINING HOOKS
(TYPICAL 2 PLACES)
EXTENDED POSITION
Figure 16-9 EVA Handrails
16-24
SEOR300PROJE(T GEMINI
the adapter to don the back pack (on spacecraft mission 8, 9 and 12) or perform
experiments. Figure 16-2 illustrates the handholds and foot rest.
=_nehandholds and footrest are retracted prior to use and are self-extended by
spring action. On missions utilizing a back pack, the handholds are held in
the retracted position by the thermal covering which protects the back pack. On
missions without a back pack, a retaining bar is used.
The handholds and footrest are extended when the INDEX EXTEND-POD EJECT switch
in the cabin is positioned to POD EJECT. The switch is located on the right
switch and circuit breaker panel (Figure 16-2). The switch controls two guillo-
tines which sever the cables holding the handholds and footrest in the retracted
position. A mechanical latch locks them in position. On spacecraft 9 and 12 the
control switch is labeled INIEX EXTEND-EVA BARS EXT.
A light is installed on each handhold to illuminate the adapter area during the
dark side of the revolution. The lights are controlled by the EXT LTS switch
on the right switch and circuit breaker panel. The lights are illustrated in
Figure 16-2.
A mirror is attached to each handhold to allow the pilot to observe the back pack
during the donning procedure. The mirrors may be positioned as required by the
pilot. Figure 16-2 illustrates the physical location of the mirrors.
The back packs are attached to the spacecraft by means of a retaining bolt which
passes through a guillotine. After the back pack is donned by the pilot, the
command pilot positions the BACK PACK-DEPLOY switch (located on the maln instru-f
ment panel) to the DEPLOY position. This actuates the guillotine and severs the
retaining bolt releasing the back pack.
16-25