astsyshansatlauvehsec9thru13_073007090105

8/6/2019 astsyshansatlauvehsec9thru13_073007090105

1/31

Astrionics Syste

CHAPTER 9EMERGENCY DETECTION SYSTEM

TABLE OF CONTENTS

Section Page9.1 CREW SAFETY SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.1.1

9.1.1 Guidelines Fo r Crew Safety . . . . . . . . . . . . . . . . . . . 9.1.19.1.2 Failure Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . 9.1.1

. . . . . . . . . . . . . . . . . . . ..2 EMERGENCY DETECTION SYSTEM 9.2.19.3 EDS OPERATION FOR SATURN V VEHICLES . . . . . . . . . . . . . 9.3.1. . . . . . . . . . . . . . . . . . . . . . ..3.1 General Consid eration s 9 3-1

9.3.2 Emergency Detection Pa rame ters . . . . . . . . . . . . . . . .or Automatic Abort (Saturn V) 9 3-19.3. 3 Emergency Detection Parameters. . . . . . . . . . . . . . . . . .or Manual Abort (Saturn V) 9 3-2. . . . . . . . . . . . . . . . . . . . . ..3.4 Saturn IB EDS Operation 9 3-3


2/31

Astrionics SystemSection 9.1

SECTION 9.1CREW SAFETY SYSTEM

The emergency detection syste m detects mal-functions in the launch vehicle which lead to eme rgencysituations. The EDS is one pa rt of the crew sa fetysystem. The other pa rt of the crew safety syst emis the launch escape syste m which remo ves the flightcr ew fr om the vicinity of the malfunctioning vehicle.Essentially , the re ar e two kinds of emergenc y situa-tions: explosion of the vehicle prop ellan ts and breakupof the vehi cle by aerodynamic for ces.

The crew safety system for Saturn/ApolloVehicles is semiautomatic. The system can sensefailure modes that slowly lead to catastrophic condi-tions and indicate these failures on a display panelto the flight crew to allow them to make the abortdecision.

The system can also automatically initiate anabort when it sen ses a failure mode that will lead toa rap id vehi cle breakup. Studies of many fai lur emodes have shown that despite the tremendous massand ine rt ia of the Saturn Launch Vehicles, the tim efrom oc cur ren ce of a malfunction until the vehiclerea che s a "breakup" angle of attac k can be ver yshort. Therefore, a fas t responding automatic ca-pability must be included in the crew safety system .

The Saturn V Apollo crew safety system isver y sim ila r to the Saturn IB system but has 3 stagesto monitor. Because of incr ease d loading, the sys -tem must also monitor propellant tank pre ssu res inthe S-I1 and S-IVB Stages.9.1.1 GUIDELINES FOR CREW SAFETY

The EDS must be as simple a s possible be-ca us e of s pac e and weight limi tatio ns of flight hard -ware and because any detection equipment has aposs ibili ty of malfunctioning and causing the abor tof an otherwise success ful mission. Fo r this reas onthe number of s en so rs in the detection syst em mustbe kept to a minimum. This decrees that the effectof a malfunction is sensed rather than the malfunc-

tion itself. As an example, t he EDS can be simplifiedby monitoring the angle of attac k and attitude ra te s ofthe vehicle ra ther than the position of each engineactuator. (Engine actuator failu res lead to loss ofcont rol and eventual vehicle breakup because of anexcessive angle of at ta ck ) Since other vehicle failurelead to this s am e effect (excessive angle of atta ckand/or attitude rate s), the monitoring of failu re effectyields a considera ble simplification of the crew safetysystem. This method of simplification i s possible onlywhen the time between failur e sensin g and catastrophicvehicle los s i s enough to allow a safe es cape dis tancefrom the vehicle.

An important guideline used in the developmentof the Saturn/Apollo crew saf ety system i s that, whenever enough time is available, the abort decision willbe left to the flight crew rat her than automaticallyinitiated. It is felt that no matter how reliable theautomatic abort sy stem is made, it can never repl acethe logic, judgement, and observation powers of theflight crew. However, many of the fa ilur e modes ofthe vehicle do not allow sufficient time fo r the flightcrew to make a decision and rea ct to the emergency;in these ca ses, the crew must rel y on the automaticabort system.

Since a falsely initiated abort could ruin anotherwise successful mission, the sys tem has beendesigned to reduce the probability of a fals e automati cabor t to an insignificant level. It is extremely im-portant that the crew safety system does not degradethe mission success probability.

Table 9.1-1 shows some of the mo re importantdesign guidelines.9.1.2 FAILURE ANALYSIS

The design of the emergency detection sys temis based on failur e mode and effect analysis. This isa complete analysis of each stage, syst em, subsystem


3/31

1

Table 9.1-1 Importa nt Guidelines for the Crew SafetySystem for Saturn Apollo Vehicles

Simplicityr - Use minimum number of s en so rs and se ns e effect of fai lur erathe r than cause - hen time perm its.Abort Decision - Made by flight crew whenever tim e per mi ts - manual rather

than automatic.Mission Degradation - Probabili ty of fal se automatic abor t reduced to insignificant Ilevel.

Initiation of manual abo rt will be based on at leas t two sep ara te and distinctindications.In the event of conflicting informat ion fro m the onboard crew safet y syste m andtelem etere d data relayed to the Spacecraft from the ground, the onboard in-formation shall always take precedence.

e Tri ple redundant e lect rical circu its utilizing majorit y voting logic will be usedfor all automatic abort signals. IAs a design objective, redundant ci rcu itr y will be used for manual abort indi-cations f rom th e Saturn Launch Vehicle to the Spacecraft.

0 It s hall be a design objective that no single point ele ctri cal fai lure i n the on-board crew safety system will r esult in an abort. II

and component within the vehicle to determine whichcomponent fai lur e modes can cause a fai lur e of thesubsystem, which subsystem failures can cause failureof the sy stem , and which system lo ss es can cause lossof the stage and/or vehicle. A fai lur e mode and effectanaly sis begins at the component level and investigateseach possible way in which the component can fail(i. e. , open, sho rt, rupture, leak). The effect of thepart icul ar f ailu re i s analyzed on higher levels ofass embly until the effect of that par ti cul ar componentfailure on the space vehicle is determined.

Once the fai lur e analy sis of a component iscompleted, a critical ity number is assigned to thecomponent. Fo r example, a number 10 indicates thatthis component can be expected to cause a vehicle lo ssabout 10 tim es out of 1 million flights.

After the criticality number has been derivedfor each component, the numbers ar e summarized foreach subsyste m, syste m, and sta ge of the vehicle-vehicle dynamics must be analyzed and struc tural limitsdetermined. Fo r example, a failure mode and effectanalysis shows that a par tic ula r group of controlcomponent fail ure s can ca use the engines to gimbal"hardover"; this would mean vehicle los s and acrit ical ity number could be derived to show the ex-pected frequency of th is failu re.

The criticality numbers are summarized on asummary chart which is refer red to a s a "failure treef' .Figure 9.1-1 repr esent s the lates t failure tre e sum-mar y of the IU Astr ionics s yst em of the Saturn IBVehicle. In this figure, the fail ure is trace d backonly 3 o r 4 leve ls and does not r each the componentmode in most cases.

Beyond this a naly sis of fa ilu res , the length oftim e between the occu rren ce of the fai lur e and thetim e of rea ching the cri ti ca l angle of at tac k must beknown. This is a function of the tim e of flight. Ifthe fail ure occ urs about maximum Q, we may havethe worst case condition and le ss tim e is availablebetween the occurrence of the failu re and stru ctu ralbreakup of the vehicle. To deter min e the se limit s,the s truct ural limit curves must be drawn for eachtype of fail ure (control engines hardo ver, enginesnull, engines out, etc. ). These curves show thevar iou s combinations of vehic le angle of at ta ck andengine gimbal angle at which the st ruc tur al l imi ts ofthe vehicle ar e exceeded, a s a function of ti me offlight. As an example of the se curv es, re fe r toFigur e 9. 3-1. Curv es must also be drawn of the actualexpected excursio ns of vehic le angle of a tta ck andengine gimbal angle during each of th ese vehiclefai lur e modes. The two se ts of curv es must thenbe analyzed to determ ine whether (and at what time )


4/31

Fge911SunBUAocSemFueMo

BeyDl0Fue

1

FgTmjeoyDeaoB

3

MisoRremeLmis

PmayPwFue

12

GaBnSy

26

S1MPaom

8

In

yoDemin

8

SowDeaoomOpmumTaeoy

8

InaVoy

(PoeMisoL

ScSeo

1

7

In

yos

1

In

yoS

1

~~~A/~wich

Seonea

1S

nCmma

NoeCrc~l~nmbsaauemogv

aeycaseoc

IBMB

S1MPaom

1

PaomEeocAmby

1 65

Plato~APwSy

5VPwSy

1

.

In

yoDemin

1

VceAu

LV

2

ol

ln

yoCmpea

29

LV

l21IsSenCmma

2

ln

yoDeminaAm

CooCmpe

2

CoCmpe

in

yoDemina&

5

EReGoC

CoCmpe

7-

2

CooAeomeeYw

CoAeomeePc

2

ln

yoDeming

63

CoCmpe

1

In

yos~i

1


5/31

Ast r i on i c s Sys t emSection 9. 1the str uc tur al l im its of the vehicle wil l be exceededfor each par t icu lar fa i lure condit ion.

Fr om the inform ation on the effect of the sefa i lu res and the informat ion on the vehic le dynamicso r s t ruc tura l l imi t s , the t ime available be tweenoccu rrence of the fa i lu re and ca tas trophic lo ss ofthe vehic le can be der ived. Th is t ime wi l l indica tethe resp onse t ime r equ i red of the de tec t ion mechan-i sm. The t ime wi ll de termine whether the c rew hassuff icient t ime to reco gnize the em ergency warningand make a decis ion as to when to abort. If humanrespo nse would be too s low, the abort o r escape mustbe autom atical ly ini t iated. A rang e of safe l im its isass i gned t o each pa ra me t e r s e l ec t ed for moni tor ing .Pe r formance wi t h in t he se l i m i t s a s sum es t he pa ram-et er to be functioning normally. General ly, these l ec t ed pa rame t e r coul d have a range of acceptabletole ran ce bounded by both an upper and lower leve l ofsa fe t y (e . g ., ove r p re s sur e and unde r p re s su re i n apr ess ur ize d tank). However, in so m e condit ions only

one safe ty l imi t is needed (e. g., high angle s of attack).As long as condi tions of the m easur ed param eter s taywithin the sa fety zone, it is assumed tha t a l l componentsar e opera t ing sa t i s fac tor i ly. If t he m easured va l ueapproaches the danger leve l, a ca tas t rophic fa i lure isimminent .The problem is to dec ide at what level to placethe ab ort level. If the abort l eve l is placed too c lo seto the tolerance l imit , an abort might be needlesslyt r iggered. Lf, on the other hand, it is moved too closeto the danger lev el (to hedge again st a needl e ss abor t ) ,the re may not be t ime enough for a safe escape be-cause of var ious sys tem s de lays . The tas k is oftencompl ica ted because the danger leve ls , a s wel l a s thetolerance zone, of cer ta in para me ters may change a s

a function of t ime. Othe r fac tor s that fu rth er com-pl icate the problem a re t rans ien ts which mom entar i lyexceed the danger level but offer no catastrophicthreat . The se must be taken into account during crewsafe ty sys tem s design.


6/31

Astrionics System, Section 9.2

SECTION 9.2EMERGENCY DETECTION SYSTEM

The ~aturn /Apo llo mergency detection sys-tem has triple redundant sensors and majority votinglogic for all the automatic abor t para mete rs. Dualredundancy is used for most of the manual abor tsen sor s. The redundancy is so arrange d that thepredomin ant fail ure mode of the sensi ng sys tem hasbeen pro tec ted against. The guideline ru le of havingat l ea st two separa te and distinct indications of f ailur ebefore initiating a manual abort is a protection againstany inadvertent ab orts fro m malfunctioning sensorsyste ms. The displays used in the Spacecraft ar e"fail safe" wherever practical. Where possible, theme te rs used for display of analog signals ar e "zerooffset". This means that the predominant failuremodes of the sens ing syst em (loss of power, etc. )wil l indicate off sca le conditions (either high or low)ra th er than readings within the scale of interest.The indicator lights used for discrete indications tothe flight crew a r e dual-bulb lights and so arrangedthat th e crew cannot distinguish whether one or bothligh ts a r e on. A fai lur e of one light will not benoticed by the flight crew. The other failure mode(inadvertent light) is protected against by requiringtwo se pa ra te and distinct indications of failu re.

The Saturn IB/V crew safety sys tem is shownin the functional diagrams, Figu res 9.2-1 and 9.2-2.Monitored param eter s ar e: stage thru st for both stagesguidance computer status, angular attitude rates ,attitude er ro r, and angle of attack for Spacecraftdisplay. Automatic abort is initiated for S-IB andS-IC two-engine-out or for excessive angular r at esin pitch, roll, or yaw; these automatic abort limi tsare switched out either automatically by the flightprog ramm er o r manually by the crew according tomission rules. Provision is also made for an abortrequest light to be energized from the ground controlcenter. At the ground control center , all of thevarious crew safety system parameters ar e monitoredby using telemet ry information from the vehicle. Othertelemetered data i s available at the MCC so that theflight director can scan all flight critical data andwar n the flight crew by voice communication of anyimpending danger. This facility provides an ear lywarning to the flight crew by giving informatio n ontren ds of various vehicle param eters , thereby ale rt-ing them to ce rta in types of f ailu re indications whichmay appe ar on the ir display panel. The flight crewwill not abort on the telemetry information alone butthey may use it a s backup data for an abort decision.


7/31

9. 2

Launch Pad RangeMission Control Center

SPACECRAFT

IBM B2O9

Figure 9 .2 -1 Crew Safety System ( Saturn IB )


8/31

Astrionics SystemSection 9 . 2

Launch Pad RangeMission Control Center

Figure 9 . 2 - 2 Crew Safety System ( Saturn V )


9/31

Astr ionics Sys temSection 9.

SECTION 9.3EDS SYSTEM OPERATION FOR SATURN V VEHICLES

9.3.1 GENERAL CONSIDERATIONS Provis ion is made to manually sh ut down theengines on any s tage f rom the Sp acecraf t wi thoutA s igna l f ro m the Sp acecraf t to the ground wi l l ini ti a t ing an abor t . A redundant means for accom-indicate that no automatic abort signal is ac t ive pr io r plishing EDS-commanded engine shutdown is provideto launch rele ase . A signal at li ft -off wil l automat-ica l ly ac tiva te the EDS automat ic abor t mode. Pro - Provis ion is a lso m ade in the Spacecraf t andvis ion s wi l l a l so be made to manual ly in te rrupt the launch vehicle for a man ual initiat ion of s tag in g (S-11ent i r e automat ic abo r t s igna l in the Spacecraf t . and S-IVB). Thes e provision s wil l enable abo rt intoorbit.An automa tic abor t wil l ini t iate engine shut-down on the ac t ive s tag e and the Spacecraf t abor t 9.3.2 EMERGENCY DETECTION PARAMETERS FORsequence .Initiation of a m anual a bo rt will be based on atlea s t two sep ara te and di s t inc t indica tions. Thesema y be a comb inat ion of EDS se ns or displays, physiol-ogical indicat ions, and ground information to theAstronaut . In som e case s, the two indicat ions maycome fro m the sam e param eter , but the indicat ionand se ns or sys tem s wi l l be independent .In the even t of conflicting inform ation fro m

the onbo ard EDS and te lemetered da ta re layed to theSpa cecraft from the ground, the onboard informationwi l l a lways take precedence .P r i o r t o a cer ta in f light t ime ( to be de te rminedby range sa fe ty cons t ra int s ) , abor t commanded engineshutdown capabil i ty wil l be locked out by provision sin the launch vehic le c i rcui t ry .A s ign a l wi l l be provided to the S pacecraf t toenerg ize the abor t re qu es t indica tor l ight when thera nge sa fe t y o f f ic e r c omma nds de s t ruc t sys t e m a r m-ing. Th is command a l s o te rmin a tes engine thrus t .Tr ip le redundant e lec t r ica l c i rcui t s ut il iz ingmajo r i ty voting logic wi l l be used fo r a l l automat icabort s igna ls . Redundant c i r cui t ry wi l l be used formanual ab or t indica t ion f rom the launch vehic le tothe Spacecraft . Ba t te r ie s used f or EDS wi l l be e lec-t r ica l ly i sola ted f ro m each other . No s ingle -pointe lec t r ica l fa i lu re in the onboard EDS wi l l resu l t inan abor t .

AUTOMATIC ABORT (SATURN V )Angular Overra tes . This automat ic abor t par am etercove rs a l l cont rol fa i lures which rapidly lead to anexce ssive angle of at ta ck and subsequent vehiclebreakup. Information wil l be supplied by the contro lEDS Rate Gyro package which co nsis ts of th re e ra tegyros fo r e a c h c on t ro l p la ne . The se gyros a r elocated in the Instru men t Unit . An automatic abo rtwil l be initiated w hen 2 out of 3 gyros i n a ny p l a neindicate tha t permiss ib le angular r a te s a re exceeded.

Provis ions a re made to manual ly deac t iva tethe automat ic abo r t s igna l for all 3 planes s imul ta -neously with 1 switch located in the S pacecraft .Manual deact ivat ion t ime is to be es tabl i shed bymiss ion rules . Capabil ity fo r deac t iva t ing e i therthe rol l or pitch/yaw/rol l signa ls by sequencing,p r i o r t o first stag e inbo ard engine cutoff, will beavai lable within launch vehicle circ ui try.Adj ust a bl e s e nso r l i m i t s e t t i ngs a r e p rov i de din pitch an d yaw within 2 t o 1 0 d e g r e e d s e c o n d a nd

in ro l l within 5 to 20 degrees/second. Sensing har d-war e m ust be removed fro m the launch vehic le toaccomplish these adjustments.S-IC Two-Engines-Out. Th e lo ss of th ru st on two o rmo re engines wil l ini t iate an automatic abort . Th isautomat ic abor t mode cove rs fa i lures occu rr ing nearthe pad and possible ran ge safety act ion. Deact ivat ioof the automat ic abor t capabi l ity pr io r to inboard


10/31

Astrionics SystemSection 9. 3engine cutoff arming is provided by the launch vehiclesequencer.

Thi s automatic abor t capability may bemanually deactivated fr om the Spacecraft, Manualdeactivation times will be established by missionrules.

A minimum of two thrus t sen sor s a re used oneach S-IC engine to provide inputs into the EDScircui try. Action of at le ast 2 sensors a re requiredto indicate los s of engine thrust.9.3.3 EMERGENCY DETECTION PARAMETERS FOR

MANUAL ABORT (SATURN V)S-IC Stage Thrust. The s ta tu s of each engine of theS-IC Stage is displayed in the Spacecraft (five indicatorlights). Upon lo ss of engine thrust , these engine stat uslights ar e energized by a discrete signal.

An abor t fo r one-engine-out will begoverned by mission rules.

e A minimum of two th ru st sens ors ar eused on each S-IC engine to activate theengine-out status lights.

S-II Stage Thrust . The sta tus of each engine of theS-I1 Stage is to be displayed in the Spacecraft (fiveindica tor lights). Upon lo ss of engine thrus t, the seengine status lights a r e energized by a discr ete signal.

Abort on one-engine-out will be governedby mission rules.A minimum of two thr ust s en sor s a r eused on each S-I1 engine to activate theengine-out status lights.

e Tra nsi tio n fro m S-IC monitoring to S-11monitoring at staging is accomplishedwithin the launch vehicle circuitry.

S-IVB Stage Thrust . The sta tu s of the S-IVB enginethrust is displayed in the Spacecraft (one indicatorlight). Upon lo ss of engine thrust , the engine-outstatus light is energized by a d isc ret e signal.

Engine thrust is monitored throughoutS-IVB burn.

@ Abort because of th rust loss is governedby mission rules.A minimum of two thrust se nso rs ar eused on the S-IVB engine to activate theengine-out status light.

Transi tion f rom S-I1 monitoring to S-IVBmonitoring at staging is accomplishedwith the launch vehicle circuitry.

Staging Sequence. Physical separa tio n of stages in-cluding the S-I1 second plane separat ion, a re indicatedin the Spacecraft by lights, or other suitable disc reteindications. In ca se of no separa tion, abort will begoverned by mission rules.Launch Vehicle Attitude Reference Failure. Improperoperation of the launch vehicle attitude refe ren ce(sensed by the IU guidance and control sys tem) willenergize an indicator light in the Spacecraft.

Abort o r switchover to Spacecraft guidancewill be governed by mission ru les.Angle of At tack An angle-of-attack function is dis-played by an analog indicator in the Spacecraft. Thisparameter is an indication of slow control fa ilureswhich lead to excessive angles of a tt ac k

The mea sured pitch and yaw components ar ecombined in vector form into a total angle-of-attackindication

The t y ~ ef measurement to be displayed andthe limi t settings (if necessary , a s a function offlight time) will be determined later. Limit settingswill govern abo rt action,S-11 Propellant Tank Pre ssu re s. LO2 nd LH2 tankpress ur es in the S-11 Stage are displayed in the Space-craf t by means of an analog display. Th is parame terreq uir es a redundant sensor and display system.S-IVB Propellant Tank Pressures. LO2 and LH2tank pressures in the S-IVB Stage are displayed inthe Spacec raf t by means of an analog display. Thisparame ter requir es a redundant sensor and displaysystem.Attitude Err or (Spacecraft). Attitude er r or s from theSpacecraft guidance and navigation sys tem a re dis -played in analog form on the flight director attitudeindicator. This parame ter is an indication of slowcontrol failu res leading to excessive angles of attackor excessive attitudes.

The Spacecraft guidance and navigation systemwill be preprogrammed with the launch vehicle tiltprogram for the S-IC flight period. Limit settings asa function of flight time (to be determined lat er ) willgovern abort actions.


11/31

Astrionics SystemSection 9. 3

i

Angular Rates. A single launch vehicle ove rra teindicator light in the Spacecraft is energized by theIU Control EDS Rate Gyro package when permissibleangular ra te s a r e exceeded in any plane. This indi-cation primarily covers the flight period in which theoverrate automatic abort capability is deactivated.

Spacecraft angular rat es ar e presented byanalog display on the flight dir ecto r attitude indicator.The sensing device fo r th is information is the Apollora te gyro package. Limi t settings , a s a function offlight time, determ ine abort actions.9.3.4 SATURN IB hDS OPERATION SYSTEM

The EDS operation for Saturn IB is differentonly in the following areas:

a In the S-IB Stage, eight engines ar emonitored (five in Saturn V).

a Since the Sat urn IB has no S-11 Stage,any function connected with thi s s tag edoes not exist in Saturn IB EDS.The S-IVB fuel tank pressure is notmonitored in Saturn IB Vehicles.

Figure 9. 3-1 is a plot of the c ri ti cal angle ofattack (a) versus flight time for the Saturn IB VehicleVarious engine deflection angles (0) r e included.Limit s for the cri tic al angle of a ttack (during thetime of maximum dynamic pr ess ur e) a r e shown inFigure 9. 3-2. Abort cri te ri a is summarized inTable 9.3-1 and EDS design ground rules a r e listedin Table 9. 3-2.

Table 9. 3-1 Abort Crit er ia and Ground Rules

a Angular overrat esa Los s of t hrust on two or m or e S-IB Engines

Indications for Manual Abort:a S-IB th rus t (eight indicator lights)a S-IVB thr ust (one of the eight S-IB indicator ligh ts)e Staging sequence (use of S-IB th ru st indic ator ligh ts)a Launch vehicle attitude r eferen ce fail (one indicator light)a Angle-of-attack (Q-ball on met er)a Attitude error (SC-G&N system on flight director attitude indicator)a Angular overrates (one indicator light)a Abort req uest (one indicator light fo r range safety cutoff and dest ruc t armi ng

signal path is either through Spacecraft up-data R F link or through launchvehicle hardwire prior to lift-off)

Manual Abor t Ground RulesManual abo rt will be governed by miss ion r ules within the following ground rule s:

a Manual abort will be initiated based on at least two se para te and distinct indi-

a No abort will be initiated over RF; an abort request will be given based on


12/31

Table 9.3-2 Saturn IB EDS Design Ground Rules

EDS Design Ground Rules:a Reliability goals:

Probab ility of detecting failu re - 0. 9973Probab ility of not detecting fal se failure - 0. 9997

a No single failure i n EDS circ uit s will cause a tr ue f ailur e from being detectedor a fal se fail ure being detected.

a All failu res th at jeopardize cr ew safe ty will be designed out if possible.How Design Ground Rules a r e Being Achieved:

a Use of 2 out of 3 voting circ ui ts in the automatic abo rt capabilitya Use of redundancy in the indications for manual ab ort capabilitye Complete EDS checkout durin g countdowna Interlocking EDS-ready in lift-off ci rc ui tse Interlocking 5-2 engine-ready with ignitiona Interlocking s eparat ion with S-IVB ignition

P a = A n g le o f A t t a c k- /3 = Gimbal Ang le

16

14Ye -zi 12O F-ePa

10

830 40 50 60 70 80 90 100Flight Time

(Seconds)IBM B211

Figure 9.3-1 Saturn IB Critical Angle of Attack Versus Flight Time

9.3-4


13/31

Astrionics Syste, Section 9.

18

16

huVY0 T 14z $5 3o eQ-0-J

12d

10

8-8 -6 -4 -2 0 +2 +4 +6 +8

Gimbal Angle ( P )(Degrees)

IBM B21

b- - - - - _ _ _

+ -2--% EDS sensor limi t must be i n this mnge----- 4s;-' " L;,,,;,---

5 -

4,--\+-.- ----

Figure 9.3-2 Saturn IB Cri tic al Angle of Attack Versu s Gimbal Angle ( 76 Seconds )


14/31

Astr ionics SystemI

CHAPTER 10LAUNCH SITE SUPPORT SYSTEMS

( To be supplied a t a la te r da te )


15/31

Astrionics System

CHAPTER 11OPERATIONAL PHASES OF THE

ASTRIONICS SYSTEM

(To be supplied at a later date)


16/31

Astrionics System

PART IIHARDWARE DESCRIPTION


17/31

Astrionics Systemi

CHAPTER 12INSTRUMENT UNIT

TABLE OF CONTENTS

Section Page12.1 INSTRUMENT UNIT . .. . . . . . . . .. . . . . . . . . . . . . . . 12.1-1


18/31

Astrionics SysteSection 12.

SECTION 12.1INSTRUMENT UNIT

The Instru ment Unit is a cylindrical structu re6. 6 me te rs (260 inches) in diame ter and 0. 9 meters(36 inches) in height, mounted on top of the S-IVBStage as illustrated in Figure 12.1-1.

The el ectr oni c equipment boxes of the SaturnAstrion ics System a r e mounted on cold plates whicha r e attached to the inner side of the cylindrical stru c-ture. The electronic equipment in the S-IVB Stageis mounted in a si mi lar way (Figure 12. 1-1). Thisarrangement provides clearance for the landing gearof the Lunar Excursion Module sitting on top of theIU and fo r the bulkhead of the S-IVB tank extendinginto th e IU.

The str uc tu re of the IU consi sts of thr ee120-degree segments of aluminum honeycombsandwich-joined to form a cylindrical ring. Afterassembly of the IU, a door provides acce ss to theelectronic equipment inside the struc ture. Thisacce ss door has been designed to act as a loadcar ryin g pa rt of the str uc tur e in flight. In additionthe st ruct ure contains an umbilical door which isspring-loaded and will close af te r retrac tion of theumbilical ar m at lift-off. The IU struc ture providea path fo r stati c and dynamic loads resulting fr omthe payload above the IU.

Typical arrangement of the equipment in theIU for operational and R & D Saturn Vehicles is il-lustr ate d on Figures 12.1-2 through 12.1-7.


19/31

Astrionics SystemSection 12.1

Instrument Unit

IBM B4

Figure 12. 1-1 Saturn IB and V Instrumen t Unit Physical Location

12.1-2


20/31

I

r

r

A

E

(PT

X2

UNIT6

UNIT6

A

C

NS

S

MU

M

2

B B

C

A

R

DGT

C

UNIT603 C

PEDVDC

A

SG

C

TO

"NOTS:

IPBVEE

C

C

(6

PBVS

E

2TBF

NOS

10O

m0m

Fge12.1-2SunBnumeUEpmeL

R&D


21/31

UT6

C

NS

ME

DSB

UT6

M

2

CB

C

40STME

R

DGT

UT6

\cB

A

~

L-S

%

'"ELV"

astsyshansatlauvehsec9thru13_073007090105

Documents