UNCLASSIFIED

AD NUMBER

AD-513 619

CLASSIFICATION CHANGES

TO UNCLASSIFIED

FROM CONFIDENTIAL

AUTHORITY

OCA; JUL 31.. 1982

THIS PAGE IS UNCLASSIFIED

UNCLASSIFIED

AD NUMBER

AD-513 619

NEW LIMITATION CHANGETO DISTRIBUTION STATEMENT - A

Approved for public release;

distribution is unlimited

LIMITATION CODE: 1

FROM NO PRIOR DISTR SCTY CNTRL ST'MT ASSIGNED

AUTHORITY

AFRPL; FEB 5, 1986

THIS PAGE IS UNCLASSIFIED

I

SECURITYMARKING

The classified or limited status of this report appliesto each page, unless otherwise marked.Separate page printouts MUST be marked accordingly.

THIS DOCUMENT CONTAINS INFORMATION AFFECTING THE NATIONAL DEFENSE OFTHE UNITED STATES WITHIN THE MEANING OF THE ESPIONAGE LAWS, TITLE 18,U.S.C., SECTIONS 793 AND 794. THE TRANSMISSION OR THE REVELATION OFITS CONTENTS IN ANY MANNER TO AN UNAUTHORIZED PERSON IS PROHIBITED BYLAW.

NOTICE: When government or other drawings, specifications or otherdata are used for any purpose other than in connection with a defi-nitely related government procurement operation, the U.S. Governmentthereby incurs no responsibility, nor any obligation whatsoever; andthe fact that the Government may have formulated, furnished, or in anyway supplied the said drawings, specifications, or other data is notto be regarded by implication or otherwise as In any manner licensingthe holder or any other person or corporation, or conveying any rightsor permission to manufacture, use or sell any patented Invention thatmay in ano way be related thereto.

1i-i

CONFIDENTIALAFRtPL-TR-7046 VOLUME II

INJECTOR/CHAMBER SCALING* •FEASIBILITY PROGRAM (U)

ABLATIVE CHAMBER DESIGN AND LONG DURATION TESTING--.VOLUME II -

G.A.VOORHEES,JR s:- 1 fr I VI .

B. 0. MORTON

TRW SYSTEMS GROUPONE lPACE iPlRK 0 OOONCO 8UACH C aALIPOINIA

TECHNICAL REPORT AFRPL-TR-70-86 VOLUME II

JULY 1970

•W') SEE INSIDE FRONT COVER FOR PATENT SECRECY ORDER NOTICE

AIR FORCE ROCKET PROPULSION LABORATORY

AIR FORCE SYSTEMS COMMAND

UNITED STATES AIR FORCEEDWARDS, CALIFORNIA

CONFIDENTIAL""mM- _w Q- Y_ "..MIME ,mr wekr ,am Mololr* IllulNssll m "AM.

REPRODUCTION QUALITY NOTICE

This document is the best quality available. The copy furnishedto DTIC contained pages that may have the following qualityproblems:

"* Pages smaller or larger than normal.

"* Pages with background color or light colored printing.

e Pages with small type or poor printing; and or

"* Pages with continuous tone material or colorphotographs.

Due to various output media available these conditions may ormay not ca•,se poor legibility in the microfiche or hardcopy outputyou receive.

E If this block is checked, the copy furnished to DTICcontained pages with color printing, that when reproduced inBlack and White, may change detail of the original copy.

0600%¶NCLASS IF IED Page ii

NOTICE

(U) When U.S. Govertnt drawings, spacifications; or other data are usedfor any purpose other than a definitely related Govermmant procurmntoperation, the Goverment thereby, incurs no responsibility nor any obligationwhatsoever and the fact that the Government say have formulated, furnished,or in any way supplied the said drawings, specifications, or other data, isnot to be regarded by implication or otherwise, or in any manner licensingthe holder or any other person or cnrporation, or conveying any rights orpermission to manufacture, use, or sell any patented invention that may inany way be related thereto.

PATENT SECREC ORDER

(U) The subject matter in this document contains information which is thesubject matter of patent applications on which the United Scates PatentOffice has issued secrecy orders. 1hese secrecy orders are superimposedon the usual secrecy regulations which are in force with respect to militarycontractors' activities. Information under patent secrecy orders must notbe disclosed to unauthorized persons.

(U) By statute, violation of a Secrecy Order is punishable by a fine notto e•ceed $10,000 and/or imprisonment for not more than two years.

too UNCLASSIFIED

UNCLASSIFIED oFOREWORD

(U) This final technical report (Volume II) covers all work performed in

Task II of Contract F04bll-68-C-0085, "Injector/Chamber Scaling Feasibility

Program." The Task U1 phase of the program covered the period from

11 December 1968 to 5 February 1970. The report was prepared by G. A.

Voorhees, Jr.. Program Manager, Applied Technology Papartment of the

Technology Laboratory. The subscale Materials Evaluation Program

section (Appendix B) was prepared by B. G. Morton; Lowell Smith was

responsible for the thermal analysis and liner sizing. Richard Williams

was responsible for the data reduction. Thi3 program was carried out

under the direction oZ Dr. Harland L. Burge. Department Manager, Applied

Technology Department.

(U) Air Force technical direction was provided by M. F. Powell.

(U) This technical report has been reviewed and is approved.

Roy A. SilverProject EngineerAir Force Rocket Propulsion Laboratory

UNCLASSIFIE

F r "r ls i ~ ~I'¶ ler••••lT -iI!lI iplFII~rt r " l f

MCLASFIED' PCaLASSI FIEDABSTRACT

(U) The results of the Task II phase of an injector/cbmber scalingfeasibility prog:.•a are presented. During the fourteen month programcovering the period from 11 December 1968 to 5 February 1970 threeablative thrust chambers were designed, fabricated and test fired. Low-cost liner materials were used in three chamber designs; the materialselection was based upon subscale test data generated by both AFRPL AnaM System. Low-cost fabrication techniques were employed throughout,

ablative components were fabricated by tape-wrapping, hand lay-up, highpressure moldiz.Z and casting. Several fabrication problem, with the low-cost materials were delineated. Four ablative materials were evaluatedIn the teat program. Two of the materials evaluated (M0-2600 silica-phenolic and Dow-Corning 93-104 filled silicone rubber) had acceptableperformance for use in low-cost engines of this type and for use In multi-million pound thrust boostez engines.

# 'UNCLASSIFIED

UNCLASSIFIEU 9?43

CONTENTS

1. INTRODUCTION AND SUMMARY 1-1

1i. INTRODUCTION 1-i

1.2 S1OIARY i-1

1.2.1 Design and Fabrication-Ablative ThrustChaoer Assembly 1-1

1.2.2 Test Reaults 1-2

1.2.3 Program Evaluation 1-3

2. ENGINE DESIGN AND FABRICATION

2.1 G•NERAL 2-1

2.2 250,000 LBF THRUST DEMCNSTLATION ENGINE DESIGN 2-1

2.2.1 Engine Operation 2-32.3 THRUST CHAMBER DESIG AND FAB2ICATION 2-3

2.3.1 Pressure Shell 2-3

2.3.2 Ablative Liner No. 1 2-6

2.3.3 Ablative Liner No. 2 2-18

2.3.4 Ablative Liner Noý 3 2-26

3. TEST RESULTS 3-1

3.1 GENERAL 3-1

3.2 TEST HARDWARE 3-1

3.3 TEST VESULTS 3-1

3.3.1 Checkout Firings - S/N 003 Demonstration Injector 3-1

3.3.2 Long Duration Firing No. I - X405090-1Engine Assembly 3-4

3.3.3 Long Duration Firing No. 2 - X405090-3Engine Assembly 3-11

3.3.4 Long Duration Firing No. 3 - X405090-2Engine Asseambly 3-17

* 4. CONCLUSIONS 4-1

APPENDIX A Steam Analysis of Z5CK Long Duration Thrust, Chamber Shell Assembly (404342)

APPENDIX B Subscale Chamber Liner Materials EvaluationProgram

APPENDIX C Ablative Liner Sizing

APPENDIX D Cost Analysis

APPENDIX E Data Reduction Procedures

UNCLASSIFi Total Pages: 165

1l199-60O7?Z8-00Q1ýImM NCLASSIFIED Page vi

LIST OF FIGURES

Figure Title

2-1 X405090 Engine Ass'y, 250K Demonstration Static Test 2-22-2 X404342 Thrust Chamber Shell Ass'y Long Duration 250K 2-52-3 X404342 Pressure Shell 2-62-4 X404342 Thrust. Char.ber Shell Ass'y 2-72-5 250,000 lbf Long Duration Thrust Chamber Ablative Liuer

Thickness for Single 120-second Continuous Burn (GasRecovery Temperature-4850*F) 2-9

2-6 X404361 250K Long Durz.tion Thrust Chambez Ass'yConfiguro-ion 1 2-10

2-7 X404361-4 Tape Wrapping Operation 2-112-8 X404361-4 Throat Section Near Completion 2-112-9 X404361-4 Throat Section After Cure 2-122-10 X404361-4 Throat Section After Machining 2-122-11 X404361-4 Finished Throat Section 2-132-12 X404361-5 Exit Cone Lay-Up 2-142-13 X404361-5 Completed L-it Cone Lay-Up 2-142-14 X4043F,ji-5 Cured Exit Cone 2-142-15 X404361-5 InstaLled In Pressure Shell Set-Up for

M1achiining 2-152-16 X404361-5 Exit Cone Being Machined 2-152-17 V'j' 61-4 Throat Section Being Lowered into Chamber 2-152-18 .- 61-3 MXA-150 Chamber Section 2-172-1 '. 3,61-2 Liner Ass'y, Showing Exit Cone, Throat

dad .;Lamber Sections 2-17

2-20 X4U4361-i Long Duration Thrust Chamber Asa'y No. 1 2-172-2.1 '4043(61-1 Long Duration Thrust Chamber Assly No. 1 2-182-22 X404362 250K Long Duration Thrust Chamber Ass'y

Configuration 2 2-192-23 X404362 Chamber, DC-93-104 Casting Set-Up 2-232-24 X404362-1 DC-93-lC4 Chbmber'Throat Plaster in Place 2-242-25 X404362-1 DC-93-104 Chamber Throat Casting 2-242-26 X404362-i DC-93-104 Chamber Exit Cone Mandrels

and Casting 2-242-27 X404362-1, DC-93-I0'4 Chamber Dome Section Plaster Mold 2-242-28 X404362-1 DC-93-L04 Chamber Cast Dove Section 2-252-29 X404362-1 DC-93-I04 Chamber Mandrel 2-252-30 X404362-I DC-93-104 Chamber looking forward 2-262-31 X404362-1 DC-93-104 Chamber looking aft from chamber end 2-262-32 X404362-1 Chamber DC-93-104 Exit cone-throat section

looking forward 2-272-33 X404363 250K Long Duration Thrust Chamber Ass'y

Configuration 3 2-292-34 X404361-4 Throat Section Prior to Aasembly 2-302-35 X404361-4 Throat Section being installed in Pressure

Shell 2-302-36 X404363-7 Partial Thrust Chamber Ase'y 2-312-37 DF5-I.61 Mix at Start of Mixing Cycle 2-312-38 DP5-161 fix, Prior to Casting 2-332-39 X404363-1 Dome Section at Completion of Casting 2-33N INCLASSIFIED

I.. • F-" - I'''l"irlHtI " 1!]IlI'ILIIll•rllrlill[ll~rrR1 P1

UNCLASSIVE Pgevi

LIST OF FIGURES (Con't)

Figure Title Page

2-40 X404363 Dome Section prior to Assembly 2-332-41 X404363-1 2-inch ThiLk Cylindrical Chamber 2-342-42 X404361-1 Completion of Casting Chamber Section 2-342-43 DP5-161/MX-2600 Interface Looking Aft 2-342-44 MXA-150 Segment Punches 2-352-45 Molded HUA-150 Fwd Segment Cavity 2-352-46 Rough Molded MXA-150 Nwd Segment Set-Up for Triming 2-362-47 Molded MXA-150 Fwd Segment (rough) Set-Up for Tri-ing 2-362-48 9 Forward Panels in Place 2-372-49 Key Aft Panel Being Installed in the Exit Cone 2-372-50 X404363-1 Chamber Prior to Shipment to the AFRPL 2-382-51 X404363-1 Chamber Prior to Shipment to the AFRPL 2-383-1 Measured Thrust, Task 11 Test Program 3-33-2 Specific Impulse Efficiency, S/N 003 Ijector 3-53-3 Combustion Efficiency, S/N 003 Injector 3-53-4 Fuel Injeccor Conductance, S/N 003 Injector 3-63-5 Oxidizer Injector Conductance, SIN 003 Injector 3-63-6 Specific Impusle Efficiency, S/N 003 Injector 3-73-7 Specific Impulse Efficiency, Firing 107 3-93-8 Thrust and Stagnatiun Pressure, Test Firing 107 3-93-9 Erosion Rate of MX-2600 Throat Insert, Test Firing 107 3-103-10 Specific Impulse Efficiency, Firing Ill 3-123-11 Thrust and Stagnation Pressure, Test Firing 111 3-133-12 Post-Test Measurements, MX-2600 Throat Insert 3-133-13 Erosion Rate of MX-2600 Throat Insert, Test Firlin 111 3-143-14 Fuel Injector Conductance, S/N 001 Injector 3-163-15 Oxidizer Injector Conductance, S/N 001 Injector 3-163-16 Erosion Pattern at Throat/Exit Cone Interface Test

Firing 111 3-173-17 Specific Impulse Efficiency, Test Firing 117 3-183-18 Thrust and Stagnation Pressure, Test Firing 117 3-193-19 0ost-Test Heasurements, X404362 Throat 3-203-20 Erosion Rate of DC-93-104 Throat, Test Firing 117 3-213-21 Fuel Injector Conductance, S/N 002 Injector 3-223-22 Oxidizer Injector Conductance, S/N 002 Injector 3-22

UNCLAS?8

"UNCLASSIFIED Page Vii

LIST OF TABLES

Table Title Page

1-1 250,000 Lbf Thrust Law-Cost Ablative Engines 1-22-1 Static Test Engine Design Parameters 2-12-2 Properties of Materials Used in X404361-2

Liner Assembly 2-112-3 Properties of Materials Used in X404362-2

Liner Assembly 2-202-4 DC-93-104 Batch Mixes 2-212-5 Properties of Materials Used in 1404363-2

Liner Assembly 2-282-6 DP5-161 Batch Mixes 2-322-7 Weights of Molded MXA-150 Segments 2-363-1 Test Hardware Configuratioras 3-2

UNCLASSIFIED

1.'!9-6007-R8-OOUNCLASSIFIED,.,se ix/x

WOHENCLATURE

F thrust, ibf

P pressure, psia

£ area ratio

0 oxidizer

F fuel

I specific impulse, lbf/ibm-sec

n efficiency

Ct characteristic velocity, ft/sec

y specific heat ratio

K conductance, defined in text

KIJCF fuel injector c.,nductance

PIF fuel injection pressure

KLJO oxidizer injector conductance

AP pressure difference, psi

APr pressure difference ratio, APio/APif

r'xture ratio, Oo/Of

W flowrate, lb/sec

dr/dt-R erosion rate, (mils/sec)

Subscripts

e exit

c -- bL--

o nozzle stagnation, or oxidizer

r ratio

i injection

f fuel

UNCLASSIFIED

i• I

iii~u~eerurn 11199-6007-R8-00UNCLASSIFIED Page 1-1

SECTION IINTRODUCTION AND SUMMARY

1.1 INTROD'JCTION

(U) This report is Vlume II of a two volume final report on the Injector/Chamber Scaling Feasibility Program, Contract F04611-68-C-0085. Under therequirements of this contract TRW Systems designed and fabricated low-costinjector/chamber hardware for rest and evaluation at the AFRPL. The pro-gram consisted of two tasks: (1) Task I - 250,000 lbf thrust Injector/Chamber Development; and (2) Task I1 - 250,000 lbf thrust long DurationAblative Chamher Lvaluation.

(U) The objectives of the program were to demonstrate perfcrmance, stabilityand injector/ablative liner compatibility and to verify scaling techniquesfor rocket engines over a range of thrust levels of 100,000 lbf to 5,000,000lbf. The Task I effort Is summarized in Volume 1 of this final report.

(U) Volume II describes the Task II effort during the period from 11December 1968 to 5 February 1970 and covers the design, fabrication andtest of three low-cost ablative thrust chambers.

1.2 SUMMARY

(U) Following is a summary of the significant subtasks performe-d insupport of the Task II program objectives.

1.2.1 Design and Fabrication - Ablative Thrust ChambeX Asseoblies

(U) The material selections for the three ablative chambers were basedupon risults of subscale test programs conducted at both che AFMPL andTRW Systems. Cost-effectiveness studies %ere made of chamber liners usingthe most promising materials. Final material selections were based on thecost-effectiveness studies, fabrication processes adaptable to larger sizeablative liners, and new material technology. Cost-effectiveness ($/sq.foot installed) findings for the four materials used in the chamber linersare summrizeu as follows:

.Material Fabrication Method Liner ComoonentThroat Exit Cone

MXA-150 Layed-up 114 - -

MXA-150 Molded - - 120MX-2600 Layed-up - - 115MX-2600 Tape-wrapped - 434 -

DC-93-104 Cast 145 294 142DP-5-161 Cast 144 - -

UNCLASSI LIED

UNCLASSIFIED(U) Three thrust chamber assa.blies were a•signed and fabricated. Thepressure shells to contain ti,- iP'lative liners were designed and fabricatedof United States Steel (USS) T-I steel alloy. Table 1-1 summrrizes thethree long duration engine assemblies.

(U) Table 1-1. 250,000 Lbf Thrust Low-Cost Ablative Engines

Engine Assembly X405090-1 XAOS090-2 X405090-3Injector (S/N) 003 002 001Thrust Chamber X40436L-1 X404362-1 X404363-1Materials

Injector LCS* LCS LCSTIC Shell USS T-1 USS T-1 USS T-1Ablative Liner

Chamber Section HUA-150 DC-93-1C4 DP5-161Throa t MX-2600 DC-93-104 MX-2600Exit Cone HX-2600 DC-93-104 MIu-150

Weight (Lbs)Injector Assy. 1500 1500 1500Pressure Shell 2850 2850 2850Ablative Liner 1010 680 1130

Dimensions (Inches)Throat Diameter 26.10 26.10 26.10Contraction Ratio 1.80 2.07 1.80_,aamber Length 54.0 54.0 54,0

Chamber Length DIa. 1.54 1.44 1.54Characteristic Length,L* 89 104 89Expansion Ratio 4.0 4.0 4.0

* Low Carbon Stpel

(U) The configuration 1 ablative liner consisted of a tape-wrapped M1-2600(silica-phenolic) throat insert, an MX-2600 (silica-phenollc) exit coneliner layed up in a rosette pattern, and an MXA-150 (asbedtou-phecolic)chamber liner which was layed-up parallel-to-surface. The exit cone andchamber sections were cured in place at 100 psi while the throat insert wascured irn an autoclave at 100 psi, machined and secondarily bonded into thepressure shell using an epoxy adhesive. The chamber internal configuration

Sconsists of an 89 inch L*, a 1.54 chamber length to diasmter ratio, and a1.80 contraction ratio.

(U) The configuration 2 ablative liner consisted of a cast Dow-Corning93-104 filled-silicone rubber throughout the chamber. The liner was castin three sections; the throat-exit cone, the dome section, and the cylin-drical chamber section using internal plaster molds and sheet metal/plywoodmandrels. The cast chamber was cured at room temperature. This chamberiaer` configuration can,•ics of a 104...ch L*D a 1.44 cha~Mar lenath todiameter ratio, and a 2.07 contraction ratio.

(U) The configuration 3 ablative liner consisted of a tape--rapped KX-2600(silica-phenolic) throat insert, an MXA-150 (asbestos-phenolic) exit coneliver and a cast Ironsides Resin DP5-161 chamber liner. The exit cone wasfabricated from compression molded MXA-150 segments which were secondarilybonded into the pressure shell. The DP5-161 chamber section was cast in

UNCLASSIFIED

11199-6007-R8-00

CONFIDENTIAL Page 1-3

two sections, the dome and cylindrical chamber section, using internalplaster molds and sheet metal/plywood mandrels. The cast chamber sectionwas cured at room temperature. This chamber internal configuration consistsof an 89 inch L*, a 1.54 chamber length to diameter ratio, and a 1.80 con-traction ratio.

1.2.2 Test Results

1.2.2.1 Demonstration and Development Inlector Firings

(U) Four checkout firings of the S/N 003 demonstration injector were madeprior to start of the ablative liner firings. The data from these fourfirings were in substantial agreement with those data obtained with theS/N 001 and S/N 002 injectors. The S/N 003 demonstration injector was thenused with the configuration 1 ablative liner (X404361-1).

1.2.2.2 Ablative Liner Configuration 1

(C) The X404361-1 ablative liner assembly was fired on 4 December 1969 for66 seconds. The erosion rate of the silica-phenolic throat inserL was about6 ailo/second which was slightly lower than the predicted value. The averageerosion rate of the silica-phenolic exit cone was essentially as predicted,although there were several localized regions which indicated higher erosionrates than normal. Erosion of the MXA-150 asbestos-phenolic material in thedome-chamber section was approximately twice that predicted by the AFRPLmaterial screening program for MXA-50. In addition, the erosion rate wasnearly twice that predicted for the only aebescos-phenolic material testedin the TRW System. subscale test program.

1.2.2.3 Ablative Liner Configuration 2

(C) The X404362-1 ablative liner assembly was fired on 7 January 1970 for98 seconds. The erosion rate of the DC-93-104 material in the throat wasabout 3.0 mils/sec which was slightly lower than the predicted value. Therewas essentially no erosion in the exit cone. The char layer was estimatedat approximately 0.4 inches thick at the exit plane with 0.2 inches of virginmaterial remaining. Erosion of the silicone rubber in the chamber-domesection was essentially zero. Most of the char layer was lost during theshutdown and purging period but the char layer that remained was estimatedat 0.4 int.hes with approximately 0.4 inches of virgin material remaining.

1.2.2.4 Ablative Liner Configuration 3

(C) The X404363-1 ablative liner assembly was fired on 12 December 1969for 83 seconds. The erosion rate of the silica-phenolic throat insert wasabout 3 mils/second which was considerably lower than the predicted values.Erosion of the DP5-161 silica-phenolic material in the chamber section wasalso lower than the predicted value determined in the AFRPL materialscreening program. The chamber liner was subjec:ed to considerable crack-ing and portions of the dome and chamber liner were ejected on shutdown.

CONFIDENTIAL

11199-6007-RB-00

CONFIDENTIAL P-4

1.2.3 Program Evaluation

(C) The measured siecific impulse of the three ablative engines was 88percent + 0.5 percent of the theoretical value at the nominal design mixtureratio of 2.60. The maximum performance measured was 89 percent at a mixtureratio of 2.50. The average weight of the three long duration ablativeengines is estimated to be 5290 pounds. The injector assembly weighs 1500pounds, the thrust chamber pressure vessel weighs 2S50 pounds and theablative liners weigh from 680 to 1130 pounds with the configuration 2ablative liner being the lightest weight design.

(U) The inherent dynamic combustion stability of the three injectors wasdmonstrated during the checkout firings and three long duration firingsin this phase of the program.

(U) Acceptable ablative performance was achieved with the Fiberite NX-2600(silica-phenolic tape or broadgoods) used as a liner for two of the threethroat inserts and in one exit-cone. Acceptable ablative performance alsowas achieved with the Dow-Corning DC-93-104 silicone rubber used throughoutthe configuration 2 liner.

(U) A number of low-cost fabrication techniques were evaluated. Tapewrapping over a male mandrel, and hand lay-up of standard ablative materialsin conjunction with a low-pressure cure cycle appear applicable for fabri-cating large ablative components but would require expensive tooling.Casting of room temperature curing ablators is a feasible technique forproducing large ablative components without the use of expensive tooling.The high pressure molding of intirlocking panels appears to be a practicaltechnqiue. The joint design and secondary bonding of the molded panelsto the shell become the :ritical and expensive part of the ablative com-ponent.

CONFIDENTIAL

CONFIDENTIAL Pag-6L-

SECTION 2

ENGINE DESIGN AND FABRICATION

2.1 GENERAL

(U) The TRW low-cost, pressure-fed engine design consisted of two majorassemblies - a centrally located, coaxial injector and an ablativelycooled thrust .hamber. The engine used low-cost storable propellants(N2 04 /UM) which are compatible with conventional materials of construction.The engine was designed with a minimum of "precision" tolerances such thatindustrial fabrication techniques could be used to fabricate both theinjector and thrust chamber shell. Low-cost ablative-type thrust chamberliners, capable of being fabricated by low-cost techniques, were used toprotect the chamber shell during the long duration engine firings.

(U) The Task I1 Design and Fabrication effort consisted of the design andfabrication of three long duration thrust chamber assemblies (TCA's)comprised of an ablative liner and pressure shell. The selected ablativematerials for these TCA's were based upon subscale test results and materialstudies conducted at both TRW Systems and the AFRPL. The demonstrationinjectors to be fired with the three TCA's were fabrt.cated during Task Iof the program. These injectors were based upon the test results achievedwith the development injector configurations. The average assembly weightof the three ablative engines is estimated to be 5290 pounds with the lighestengine being the configuration 2 (5030 pounds) design. The injector assemblyweighs 1500 pounds, the thrust chamber pressure vessel weighs 2850 pounds,and the ablative liners weigh from 680 pounds to 1130 pounds.

2.2 250,000 LBF THRUST DEKONSTRATION ENGINE DESIGN

(U) The 250,000 lbf thrust (vacuum) demonstration engine design (X405090)is shown in Figure 2-1. This engine is comprised of two major assemblies,the centrally located, coaxial demonstration injector and the ablative linedthrust chamber. The design is based on achieving the parameters shown inTable 2-1. The design test duration was 120 seconds. The two primaryobjectives of the test firings were the determination of injector/chambercompatibility and the evaluation of low-cost ablative materials.

(C) Table 2-1 Static Test Engine Design Parameters (U)

F(vac), lbf 250,000

P, psia 300

co, Ae/At 4.0

C, Ac/At 2.0 (Nominal)

A/F 2.60 (Nominal)

Isp (vac), lbf-sec/lbm 283.3 (Theoretical Shifting Equil.@O/F-2.60)

nIsp 0.90C*, ft/sec 5596 (Theoretirai Shifting Equil.@O/F-2.60)

nc* 0.9375

CONFDENTIAL

lU99S-6007-S-O0

CONFIDENTIAL P-- -

C.Figur P-a14590Egie 2-2y

IH

_

2-' 1

figure 2-1. X405090 Engi~ne Asly~

250K Deontration. S~Static Test (U)

(This pagel ism unclassified. )

w;-oc

1 9-

_ _v ~ MEOW

-I1

m IVA£F r

II

1pm 0 4m

A3&V

_ MI

1 -________________

I S 41

UNCLASSIFIED Pae9-3

(U) The design criteria, including propellant flows, calculated pressuredrops, and predicted stress levels used in the development hardware designhave been detailed in Volume I. The demonstration injectors used inTask !I are maeentially identical to the development inje•tor except formethod of attachment of the oxidizer orifice ring. The pressure shellused with the ablative liners is fabricated from USS T-1 steel alloy in allareas except the long weld neck flange which is ASTH A-181 carbon steel.The stress analysis for the pressure shell used with the ablative linersis presented in Appendix k. The stress analysis wss based on a steady-stats pressure of 600 psi ia the injector and chamber.

(U) The selection of materials to be used in the ablative liners werebased upon test results obtained in the AFRPL "in-house" material screeningprogram and the TRW Systems sub-scale Materials Evaluation Program (AppendixB). The thicknesses of the various materials were determined usig theprocedure given in Appendix C.

(U) The unit costs for quantiti of 1, 3 and 10 pressure shells andablative liner assemblies are giN-.a in Appendix D.

2.2.1 Engine Operation

(U) Combustion was initiated by opening an 8-inch facility oxidizer valve(propellant valves were not supplied with the engine.) so as to proviLe adefinite oxidizer lead. Oxidizer entered Lhe injector through an 8-inch,300 lb ASA welding neck flange located on the Lnjector centerline andflowed through the oxidizer tube. The oxidizer was turned 90" by the con-tourad pintle tip and was injected tadially through multiple orifices. Fuelentered the fuel manifold through a 6-inch, 300 lb ASA welding neck flangeand then was injected as a hollow cylindrical sheet.

(U) The fuel sheet impinged with the radial oxidizer streams approximately2.5 inches from the fuel orifice. Since the propellants were hypergolic,combustion took place upon contact of the fue. with the oxidizer. After120 seconds of operation both facility propellant valves were to be se-quenced closed. Gaseous nitrogen was used to purge the residual propellantsfrom the manifold.

2.3 THRUST CHAMBER DESIGN AND FABRICATION

2.3.1 Pressure Shell

2.3.1.1 Design

(U) The pressure shell design for the ablative liners %as similar to theX403646-1 heat-sink combustion chamber design which was uqsed in the Task Iprogram. Two major changes were made in the design of the pressure shell;(1) the body was split into two sectionr with flanges to allow insertion ofthe tape-vrapped throat insert ad (2) the tý,ickness of the exit cone shellwas decreased from 0.5 inches to 0.25 inches.

UNCLASSIFIED

UNCLASSIFIED Page 2-4

(U) The X404342 pressure shell (Figure 2-2) was designed using USS T-1steel in all areas except the long weld neck flange (X404342-14) which wasASTI A-181, Grade II, carbon steel. The head (X404342-13) was a 2:1 ASMECode elliptical head of USS T-1 steel, 39.0 inch I.D. by 0.5 inch wallthickness. The plate sections of the chamber sections were 0.50 inchthickness while the exit cone plate thickness was 0.25 inches. The closureflanges (X404342-9, 10) were machined from USS T-1 forgings.

(U) The thrust aount consisted of a 1/2 in. thick support ring (X404342-11)fabricated from USS T-1 steel plate 8.7 inch wide by 130.0 inch long. Thesupport ring was welded to the cylindrical chamber section downatrera ofthe weld which connected the semi-elliptical head (X404342-13) to the foreflange (X404342-10). Shims were used to space the -11 support ring sothat a minimum 0.040 inch clearance was maintained between the support ringand semi-elliptical head. The thrust mount ring was fabricated from 1.00inch thick USS T-1 steel plate which was welded to the support s.•irt. Themounting hole pattern was 20, 1-1/8 inch diameter holes, on a 44.0 inchdiameter bolt circle. This thrust mount configuration was used on bothheat-sink combustion chambers which were fabricated during Task I.

(U) The pressure shell design incorporates twenty-two 1/4-18 NPT partsfor the purpose of thermocouple installation and pressure measurement.The two pressure taps werL located in Lhe 16-inch 300 psi ASA long weldingneck flange. This allowed for a head-end pressure measurement in theannulus between injector and flange. Twenty of the ports were located atfive axial pos-ions to measure the ablative liner backwall temperature.

(U) The major differences In the design of the three pressure shals werein the exit cone area. The X404342-2 exit cone section was designed toa 53.60 inch exit plane ID to accoodate the 0.7 inch thick ablative linerwhile both the X404342-1 and X404342-21 were fabricated to a 53.20 inchexit plane ID for the 0.5 inch thick ablative liners. Identical stiffenerrings were added to the -1 and -2 pressure shell configurations. Thestiffener ring on the X404342-21 pressure shall assembly was designed with48, 21/32 inch holes to match the end closure tooling which was used forcuring the HX-2600/HXA-150 in place.

2.3.1.2 Fabrication

(U) Three pressure shells were fabricated from the basic design. TheX404342-1 pressure shell configuration was used for the X404362-1 ablative-lined long-duration thrust chamber assembly; the X404342-2 pressure shellconfiguration was used for the X4.04363-1 thrust chamber assembly, and theX4043L2-21 pressure shell shown in Figure 2-3 was used for the X404361-1thrust chamber assembly.

(U) All pressure shells, were fabricated using rolled and welded USS T-1steel sections as shown in Figure 2-4. The two conical sections, -3 and-7, and the three cylindrical sections, -5, -8 and -11 were roiled fromsingle pieces of steel plate and had only one longitudinal weld. The 2:1elliptical head was a commercial head fabricated from a sing)e piece blank.The X40-4342-14 long weld-neck flange was a commercial 16 inch 300 AS& flange.UNCLASSIFIED

UNCLASSIFIED Pg -

f 4-4

Figure 2-3. X404342-21 Pressure Shell(U) (65542-69-10)

WU All circuunferential welds used to join the various sections were madeusing automatic, subwerged-arc welding techniques. Multiple passes wereused on all weld joints. A 100 percent X-ray requirement was imposed onall butt welds and dye penetrant inspection was used an all fillet velds.All welds were certified in accordance with applicable ASME boiler andpressure vessel codes.

(U) Two lifting eyes were added to the chamber shell, as shown in TRWdrawing X404342, to facilitate handling.

2.3.2 Ablative Liner No. 1

2.3.2.1 Dosixn

(U) The original design of the X404361-1 ablative liner was based on theuse of 11X-2600 (Fiberite) silica-phenolic material throughout. The materialthicknesses were based upon the test results obtained with the DC-93-104material during the TRW Systems sub-scale Materials Evaluation Program(Reference B). M1-2600 had not been tested in the sub-scale program butsubsequent testing at the 1500 lbf level confirmed that the design wasconserva tive.

UNCLASSI FIED

X404342-10 FL

X404342-13 HEADS~X404342-11 SUPPORTX404342-14 FLANGE X440322-12 MOUNT

I

UNCLAS!

X4043X404342-5 THROAT

X 404342-7 CONE

X4043.42-8 BODY

X404342-9 FLANGE

X404342-10 FLANGEX404342-13 HEAD

X4043,42-11 SUPPORT142-12 MOUNIT N AIUNCLum

11199-6007-R-O00UNCLASSIFIED Pageo2-7

X404342-16 RETAINER

IJ

X404342-22 RING

X404342'3l EXII CONEX404342-5 THROAT

X 404342-7 CONE

404342-8 BODY

GE

Figure 2-4. X404342-21 Thztut ChamberShell A•asely (U)

UNCLASSIFIED

____ ____ ___ ____ ____ ___j

UNCLASSIFIED(U) The TRW Systems Charring and Ablation Computer Program (AH054A),employing the erosion model of Munson and Spindler were used to size theaulative material thicknesses (Appendix C). The input data to these pro-grams Tonsisted of the erosion data, surface temperature, and effectiveheat of ablation obtained during the subscale test program of the variousmaterials. The criterion used in the liner sizing was based on a 4850"throat recovery temperature and a maximum backwall temperature of 600"Fafter 120 seconds of firing.

(U) The required thicknesses are shown in Figure 2-5 for four candidateliner materials. The AFRPL approved design, which is shown in Figure 2-6,consisted of an MX-2600 silica-phenolic throat insert, an HX-2600 silica-phenolic exit-cone liner and an KXA-150 asbestos-phenolic chamber-domeliner. The MXA-150 asbestos phenolic material was not tested in the TRWSystems sub-scale program but data furnished by the AFRPL on the erosion-char rate of the MXA-150 (0.98 inches/60 seconds) indicated that the materi-al was at least comparable to the Haveg-41 asbestos-phenolic material testedin both the TRW Systems sub-scale program (1.00 inches/60 seconds) and theAFRPL material screening program (1.04 inches/60 seconds). The materialthickness of the MXA-150 was increased to 2.0 inches from the indicated 1.60inches for the Haveg-41 at a 35.0 inch chamber I.D.

(U) Tape wrapping at 60* to centerline was chosen for fabricating thethroat insert. This method and laminate orientation provided maximumerosion resistance in an area where both heat-flux and shear-stress aremaximum. A rosette lay-up was selected for the exit cone. This methodof fabrication is used on LMDE exit cones. The chamber-dome section wasdesigned as a parallel-to-surface lay-up of material; a reduction inmaterial thickness was made in going fro-a tW4 cyliadrical section to thedome section.

2.3.2.2 Fabricatiou

(U) The materials employed in the X404361-2 liner assembly were MX-2600silica-phenolic and HXA-150 asbestos-phenolic manufactured by the FiberiteCorporation. The MX-2600 used for the throat section X404361-4 was fur-nished as bias tape while MX-.2600 broadgoods was used for the exit cone.The HXA-150 asbestos-phenolic material used in the chamber-dome section

As furnished as a broadgoods. Typical properties of the three materialsare given in Table 2-2.

(U) The X404361-4 throat section was tape-wrapped on a net, malc mandrelat 60* to centerline. Tape width war chosen to provide the proper com-ponent thickness when oriented at the desired angle to the flow, withallowances for wrapping guidance accuracy and trim, pull-down of biasmaterial in wrapping operations and loss of material in debulk during cure.

(U) The wrapping operation shown in progress in Figure 2-7 shows themandrel set-up in a lathe with a pressure roller being used to compressthe tape as it is wrapped. The as wrapped density is approximately 95percent of the cured part density. Heat-guns were used to heat the tapeand previously wrapped material during the wrapping operation. There wasno tendency for wrinkling of the inner portions of the ply during thewrapping operation even though ply widths as large as 4 inches wore employed.Figure 2-8 shows the throat section near completion of wrapping.

UNCLASSIFIED

,mn! ei'n

UNCLASSIFIED Page-2-9

I:s

-I--

in-

0 L)

IJ6 LnJ0 Z.

/ ~z

z / - z 4,____________ ~L _______ -' .

I X,-e

_L U 0 C l

-4-0 V.uu*~

PiSn

N -~ a. on___________ ___________r14

p Lei

C, 1Q I

53H:)NI SS3N)I4)IHI ý3Nfl~I1 Ml9V

UNCLASSIFIED

UNCLASSIFIEDTable 2.2

Properties of Materials Used in X404361-2 Liner Assembly (U)

MX-2600 MX-2600 HCXA-l50

Form Tape Broadgoods Broadgooda

Lot No. H-355 H-387 H-375

Resin MIL-R-9299(It) HIL-R-9299(II) MIL-R-9299(II)

Filler (Silica), % 7.8 N/A. 5.0

Fabric reinforcement, % min. 96.0 96.0 N/A

Resin solids, Z 31.0 31.4 48.5

Volatile Content, Z 7.0 7.5 12.5

Specific Gravity 1.71 N/A 1.56

SNot Available

C÷

F•gpre 2-7. X404361-4 Tape wrapping Figure 2-8. X404361-4 Throat sectiogoperation(U) near completion (U)

(65U2-69-5) (655I2-69-4)UNCLASSIFIED

I -..wf

UNCLASSIFIED 2-12

(U) Following completion of wrapping the part was prepared for curing.Nylon shrink tape was wrapped over the MX-2600 bias tape. A bleeder clothand "monkey-fur" material were then applied to the part. Finally a Mylarvacuum bag was applied to the wrapped part. The vacuum bagged part wascured in an autoclave at 100 psi and 290-310*F for five hours. The tem-perature in th. autoclave was increasod from ambient temperature to 300"Fin approximately 4 hours. The cure process is visualized as occurring fromthe inside of the mandrel outward through the layed-up part. The outersurface is protected by the insulative properties of the bleeder cloth,separation sheets and mylar vacuum bag. In this manner, volatiles arealways removed through uncured material and into the bleeder cloth, and auniformly cured part results. Figure 2-9 shows the part as removed fromthe autoclave with part of the vacuum bag and bleeder cloth removed.

(U) The machining of the cured ablative throat did nut present any parti-cular problems. Only the 0.D. sur 'ace of the throat was machined to matewith the I.D. of the pressure shell. The part was machined to allow fora glue line thickness of approximately 0.060 inches. The part was machinedon the mandrel which was used for the wrapping operation as shown in Figure2-10. The finished part prior to assembly into the X404361-2 liner assemblyin shown in Figure 2-11, with the forward part of the split mandrel in thebackground.

Figure 2-9. X404361-4 Throat Figure 2-10. X404361-4 Throat SecticmSection After Cure After Machining(U)

UNCLASSIFIED

UNCLASSIFIED(U) Prior to start of layup ofthe MX-2600 broadgoods in the exitcone section of the pressure shellthe inside surface was sandblastedand then cleaned using methyle-thyiketone (MEK). The cleanedsurfaces were than primed usingHonsanto SC 1008 phenolic resin.

(U) The X404361-5 exit cone linerwas layed-up in a rosette patternusing the steel shell as a femalemandrel. Trapezoidal plies 6 in.wide x 4 in. wide x 50 in. longwere cut from the MX-2600 broad-goods and layed in place in thepressure shell as shown in Figure2-12. The lead on the forward endwas approximately 0.110 incheswhile the lead on the aft end wasapproximately 0.190 inches. Thisresulted in a 21 ply thickness atthe aft end (exit plane) and a

S " ,55 ply thickness at the forwardend (throat extension pline). Theplies were held in place by vacuumbagging while the shell was being

Figure 2-11. X404361-4 Finished rotated. Figure 2-13 shows theThroat Section(U) completed layup prior to curing.

(65543-69) Approximately 830 plies were usedin exit cone layup.

(U) A separation sheet, bleeder cloth and Mylar vacuum bag were put inplace over the layup. The steel pressure shell served as the autoclave.The ends of the pressure shell were sealed-off; a flatplate closure wasused on the 16 in. flange end while an elliptical dome wea used to closeoff the exit end. The exit cone was cured at 100 psia and 290-310'Ffor 5 hours. The temperature was raised from ambient to 300'F in approxi-mately 4 hours. Reat applied from the metal (shell) side rauses volatilesto flow from the shell side through uncured material and into the bleedercloth and out through the vacuum system which results in a uniformly curedpart. rigure 2-14 shows the part with the end closure off after cure andremoval of separation sheet, bleeder cloth and vacuum bag.

(U) The pact was setup for machining of the recess in the exit cone lineras shown in Figures 2-15 and 2-16. The X404361-5 exit cone liner wasmachined to allow for a glue line thickness of approximately 0.060 inches.Following machining of the exit cone the throat was bonded to the shell andexit cone using Epon 901/B3 adhesive. Figure 2-17 shows the throat withadhesive being Lowered into the pressure shell. The adhesive was allowedto setup for 24 hours prior to start of the KXA-150 layup.

UNCLASSIFIED

UNCLASSIFIED Pae21

Figure 2-12. X404361-.5 Exit Figure 2-13. X404361-5 CompletedCone~ Lay-up (U) Exit Cone Lay-up (u)

(65907-69-8) (65908-69-1)

Figure 2-14. X404361-5 CuredExit Cone (U)

(67990-69-9)

UNCLASSIFIED

UNCLASSIFIED Pae21

Figure 2-15. X404361..5 Inst~alled in Pressure shall.Set-Up for Machining (U) (66777-69-1)

Figure 2-16. X404361-5 E-xit Figure 2-17. X404361-4 Throat(U)Ben acie (65543-.69-.10) Section Being LoweredUNCLASSIFIED It hme u

1.6~~~ ow-.u -

---

11199-6007-R8-00UNCLASSIFIED 2-16

"(U) The X404361-3 chamber liner was layed up using "orange-peei" sectionsof MXA-150 asbestos-phenolic from the forward openinB (16 inch flange) tothe throat liner. Plies were layed up parallel to centerline in thechamber section. The 2.0 inch thickness required approximately 55 plies.The length of the plies was increased as the thickness was tapered to 0.8inches at the forward opening. Approximately 23 plies were used to achievethe 0.8 inch thickness. The plies were layed up in the chamber section witha maximum overlap of the longitudinal joint of one-half inch. The adjacentplies did not have any coincident joints. The layup was accomplished inthree steps; the initial 16 plies were vacuum bagged and debulked at 100psi and 180*F. The layup was also subjected to the same debulking cycleafter 36 plies had been layed-up. Following layup of tha last 23 pliesrhe part was covered with a separation sheet, bleeder cloth and Mylarvacuum bag. The final cure was affected in the same manner as that usedfor the X404361-5 exit cone linar. Closures were used on both ends ofthe pressure shell. The part was cured for four hours at 100 psi and290-310"F. Approximately 4 hours were required to reach 300'F.

(U) Shrinkage of the MXA-150 asbestos-phenolic material was expected andshrinkage of about 0.030 inch on a side did occur. The dome section ofthe pressure shell was removed and Epon 934 adhesive was applied to theasbestos-phenolic part and the head was rebonded to the X404361-5 chamberliner. Shrinkage in the cylindrical chamber section was not as uniform.The void between liner and shell was back-filled with a 1:1 mixture ofVersamide 115-Epon 828 adhesive. Figure 2-18 shows the adhesive on thedome section prior to reinstallation of the dome portion of the pressureshell.

(11) Measurenents were taken of the exit plane diameter, throat diameter,chamber diameter and location of the throat plane prior to shipment. Thethroat diameter was 2b 1/8 inches t 1/16 inch while the exit plane diameterwas 52 1/8 Inches t 1/8 inch. The chamber diameter was 35 inches + 1/16inch. The thickness at the exit plane varied from 0.56 inch to 0.2 incharound the circumference. The Chickne&ws at the forward end was a ,iominal0.80 inch plus 0.060 inch bond line thickneas.

(U) Material samples were taken from each end of the HX-2600 throat,the aft end of the liX-2600 exit cone and the forward end of the MXA-150chamber liner. Specific gravity samples of the throat indicate a Sp. Gr.of 1.73 for bot'a samples. The throat was also weighed (1O + lbs); usingthe calculated volume results in an apparent bulk density of 1.74. Samplesfrom thn aft end of the exit cone indicated a Sp. Gr. of 1.66 while theSp. Gr. of the KWA-150 component was 1.50.

(U) The joint at the throat/exit interface was ground smooth an were theends at both the forward flange section and exit plane. The X404362-16retainer clips were reinstalled and tha nui"or 1 long durarion thrustchamber assembly (X404361-1) shown in Figures 2-19, 2-20, and 2-21 wasmade ready for shipment.

UNCLASSIFIED

UNCLASSIFIED Pae-1

Figure 2-18. X44361-3 MXCA-150 Figure 2-19. X404361-2 Liner Auly.(66777-69-5) Chamber Section(U) (66777-69-9) Showing Exit Cone,

Throat and ChamberSectin )

(667 77-69-11)

Figure 2-20. X404361-1 Long Duration Thrust Chamber Aau'y. fNa. 1 (U)

UNCLASSIFIED

UNCLASSIFIED Pae21

(66777-69-12)Figure 2-21. X404361-1 Loug Dur3Lion ThruSL Chamber Ass'y. No. 1 (U)

2.3.3 Ablative Liner No. 2

2.3.3.1 Der'yg

(U) ThL ee-sign of the X404362-1 ablative liner was based on the use of

Dow Corning 93-104 (filled silicone rubber) mAterial throughout. Thematerial thicknesses were based upon test: results obtained during the TRWSystems sub-scale Materials Evaluation Program (Appendix B).

(U) The erosion model of Spindler and Monson and the TRW Systemsa Charringand Ablation Computer Program (A11054A) were used to site. the ablativematerial thicknesses (Appendix C). The Input data LO there programsconsisted of the erosion data, surface temperature, and effective heat ofablation determined in the sub-scale program. The criterion used in thematerial sizing were based on a 4850*F throat recovery temperature and amaximum backwall temperature of 6C007 after 120 second* of firing.

(U) The required thicknesscii for the X404362-1 liner are shown in Figure2-5. These thicknesses were achieved by casting the material in the voidcreated by the shell and the Internal moldi and mandrels. Cold jointswere allowed in the cylindrical ch~amber section and the exit con* liner.The throat approach, throat and aft throat sections were cast as a singlethick section. Elevated temperature cures were not specified for anysections. figure 2-22 shows the AFRPL approved design.

UNCLASSIFIED

TV. M__

Ai5 aa'..,.

tr o "aN*~ ' -

m'" s4 -

j A

,.' ,,a'ap k, tj~ N kaaaa a,. .. 'a .

a a *a".lk

A AN-

- ,,

MF .. X0

&wow . .

-. - - - - ---

4 .r-IAN

F 0,

4 - . -

Aga~' .wa

§10" L.'.~fMC ^W~K~LA

a"-J

7.',J 4; ~ ~

50 9SrPLO

7 .l*A* S -

UNCLASS~iED ll99-6007-RS-O0UNCLASSIFIED page 2-20

2.3.3.2 Fabrication

(U) The X404362-2 liner assembly was cast from Dov-Corning 93-104 siliconerubber RTV using sheet metal-plywood mandrels and plaster molds to createthe internal contour. Material from three lots were used to cast theentire chamber. Typical properties of the DC-93-104 material are givenin Table 2-3.

Table 2-3

Properties of Materials Used in X404362-2 Liner Assembly (U)

Lot No. 203237 201319 301314

Amount I x 65 lbs 9 x 65 lbs 1 x 65 lbs

Viscosity, poises 2120 2160 2160

Working Time* 5 hrs 4 hrs-l0 min 4 hrs-lO min

Sp. Gravity 1.47 1.45 1.45

Durometer, Shore 75 75 75

Tensile Strength, psi 260 242 242

*Time for material to reach a preliminary elastomeric state.

(U) The X404362-2 liner assembly was cast into the X404342-1 pressureshell assembly. The interior metal surfaces were sandblasted and degreased,then rinsed with methylethylketone (MEK) to prepare a scale-free, lintfree substrate. Dow Corning 1200 primer was applied in a thin film bybrushing and allowed to air dry prior to installation of the internalmolds and mandrels.

(U) Casting of the DC-93-104 took place over a three day period. Thethroat approach, throat, throat extension and approximately 50 percentof the erit cone were cast during an 8 hour period; the balance of the exitcLuze !nk 'um section wert cast on the following day. The cylindrical chamber

t.t %'a was cast approximately one week later after receipt of additionalmaterial. The dome section was then bonded to the cha--her section usingDC-93-104 to seal the joint between the two components.

(U) DC-93-104 Is a two component silicone material which is mixed in theratio of 10 parts catalyst to 109 parts base material. The base materialwas not de-*ired prior to mixing. Catalyst was added to the base material

and mixed for approximately 5 minutes using a low speed mechanical mixer(Kol Inc., M-58). A sample from the initial batch was cured for 2 minutesat 300"F to check the mix and cure properties. Table 2-4 shows the weights

UNCLASSIFIED

M.=

UNCLASSIFIED Page 2-21

Table 2-4

DC-93-104 batch Mixtures (U)

Batch No. Baei Vetght, lbs Catalyst Weightý lbs Wt Cat/109 lbs Base

1 28.25 2.6 10.05

2 20.437 3.0 10.72

3 34.50 3.4 10.75

4 29.25 2.73 1.018

5 30.75 3.0 10.626 28.06 2.6 10.10

7 34.125 3.24 10.368 28.625 2.64 10.05

9 34.50 3.35 10.59

10 23.562 2.09 9.65

11 32.00 3.0 J1.2212 32.00 3.0 10.22

13 30.12 2.95 10.68

14 26.75 2.51 10.22

15 33.25 3.28 10.75

16 26.375 2.56 10.5917 31.437 3.14 10.89

18 30.75 2.98 10.56

19 35.50 3.46 11.25

20 33.565 3.29 10.69

21 43.00 3.96 10.05

22 2.00 O 185 10.05

TOTALS 658.806 62.965 1G.420 (Avg.)

TOTAL WEIGIIT 721.77

UNCLASSIFIED

UNCLASSIFIED(U) of base and catalyat for each of the mixtures cast into t.,e linerassembly. As each batch was mixed it was dra%,n into a vacuum pot under25 in.Hg vacuum and desired prior to injection. Samples of each sixwere taken to check the workiiig life of the material. Batches 1-4 wereused to cast the throat approach and throat section; batches 4-7 were usedto cast the throat extension section as shown in Figure 2-23. Batce ts8-11 were used to cast the forward exit cone section, while batches 12-15were used for the aft exit cone section. The dome section was cast usingmix batches 15-17, The cylindrical chamber section was cast using batches18-21 and batch 22 was used for the joint between the chamber and donesection and for repair of minor v.oids.

(U) The chamber was inverted and the plaster sold of the throat approachthroat section positioned 50.1 inches from the exit plane and centrallylocated to provide a 2.1 + .050 inch wall thickness at the throat. Allsurfaces of the plaster molds and sLeet metal mandrels were wrapped withclear, 2-inch pressure sensitive polyethylene tape to act as a releasefilm. The DC-93-104 was extruded from the deairating pot through tubingsub-surface into the space between the shell and plaster sold. Figure 2-24shows the plaster mold located in position while Figure 2-25 shows the castrubber nearly to the throat plane.

(U) The plaster sold for the throat extension section was then installedbeing keyed to the previous mold and the casting was continued using batches4-7 for the throat extension. The forward exit cone mandrel was theninstalled and batches 8-1l cast. Difficulty with the extruding pressand feed system for the catalyzed silicone rubber resulted in terminationof the casting proceus following mix 11.

(U) The casting process was resumed the following morning and the balanceof the exit cone was cast using batches 12-15. A ctntral shaft was usedto index the various sheet metal-plywood mandrels and plaster molds asshown in Figure 2-26. It Is estimated that 440 lbe of material was usedto cast the throat Approach, throat, throat extension and exit cone section.This was approximatA.y 10 percent greater than the calculated amount.

(U) A plaster mold was made for casting the dome section and is shovn inFigure 2-27. This was positioned within the dome section an shown inFigure 2-23 and thi rubber material was injected through 8 injection ports.The cost material and injection ports are shown in Figure 2-28. Batchmixes 15-17 were used for this section.

(U) The chamber-throat-exit cone was inverted following removal of thetwo plaster molds and two sheet metal-pLywood exit cone mandrels and thechamber plaster %old shown in Figure 2-29 was installed. The cylindricalportion of the chamber was cast using batches 18-21. Following cure ofthe chamber section the plaster mold was removed and the chamber domesection was fitted to the cylindrical chamber section. A 1/2-inch x 3/8inch step had previously been cast in each section. These two steps werejoined using material from batch 22 as adhesive between the sections,'

UNCLASSIFIED /

11199-6074R8-00UNCLASSIFIED Pae23

mix 12. 136 144. isSH TW L

KtVIOOD

M~Lto, I I

OMIX 4, 5, 4. 7

mix 1. 2, 3. 4

Figure 2-23. X404362 Chamber,DC-'93-104 CantingSet-up (U)

UNCLASSIFIED

UNCLASI~iED11199-6007-R8-00UNCLASIFIED Page 2-24

t.0.

Figure 2-24. X404362-1 DC-93-104 Figure 2-25. X404362-1 DC-93-104

Chamber Throat Plaster Chamber Throat Casting(65907-69-1) in Place (u) (65907-69-2) (U)

Figure 2-26. X404362-1 DC-93-104 Figure 2-27. X40..3h2-1 DC-93-104Chambe~r F.xIt Cone Chamber W~oe Sect ion

(65907-69-6) MAndrels and Casting (65907-69-4) Plaset~r .'Moid (U)(U)

UNCLASSIFIED

11199-600 7-U1-00UNCLASSIFIED Page 2-25

I41~

Figure 2-28. X404362-1 DC-93-104 Figure 2-29. X4O43b!-l DC-93-104Chamber Cast Dome Chamber Mandrel (U)

(65907-69-7) Section (U) (65907-t9-52)

(U) Measurements were taken of the exit. planau diameter, thro.&t diameterand location of the throat plane following assembly. The throat diameterwas 26.05 inches 4*.05O inch while the exit cone diametcr was 52.12 inches.The thickness at the exit plane varied from 0.5b Inch to 0.62 inch around

the circumference compared with the design value of 0.50 +-0.10 inch. Thethickness at the forward and was a nominal 0.75 inch compared to the designvalue of 0.80 +-0.10 inch. The nominal ID was 37.5 inches.

(11) A specific gravity measurement takena on .a siampie frolm bi~iteli 7 indicateda specific gravity of 1.45 by the limers iun metho~d. Examliti iion of ths,surface of the exit cone showed only very winute piln holes where ailr h-ad

been trapped between the material And poliyethy'lene rele.alik :.apf. Obviousair bubbles were cleaned out and piatched. The Joint between cured Materialaid uncured material (satart of convcrging sectiLon) w.is thoroughaly examined.

There was no nd icat ion of any poor bondinug condit ion.

-(U) F igure 2 -30 show, t he f In ishad 11inwr anssemblIy Illui qin g fo rwar d f rm' t hethreat sect ion. The mavrks left by the 1%ilyrthv1enc tape .and one of Ithe

injector ports are clearly visible. Figure 2-Jl sh1OWN theV filliNhed Assemblylooking aft from thai chamber saide. The d~ark mpot i are pl5.t14 where Void%have been repaired. Figure 22-32 ithows thL exit cont-a 4n. h1 ro'aL 84-4t lol

looking forward. The light colobratlion iti where mi-ter i-i ham.1 tivel t rimmedfrom the cast part at the joints of the sheet mctal maundre ls.

UNCLASSIFIED

UNCLASSIFIED

Filgure 2-30. X404362-1 DC-93-104 Figure 1-31. X404361-1 DC-93-104Chamber lookLng forward Chamber lookLng aft from

(65908-69-6) (U) (65908-69-7) dcmaber end. (V)

2.3.4 Ablative Liner No. 3

2.3.4.1 Design

(U) The design of the X404363-L ablative liner wa. based on the use of

a MX-2600 uLLLca-phenolic throat insert fabricated in the same manner as

the throat insert used in the number I configuration liner (set? 2.3.2.1).

The original devign of the number 3 liner was based on the u%L of either

Haveg- 41 (asbestom phenolir) or GU-223-50 (eyxiy-.iov.alak) iii the chamber

and exit cone in conjunction with the MX-260n throat insert.

(U) The AFRPL approved deigfn consisted of aii .Vi-2(u0O silica-phenclic

throat insert. , DP5-Ibl silica-plienolic chimber section and an MXA-150

asbestos-phenolic exit--one liner. Neither mat•eriai had berii tested in

the TRW Systems sub-mcale program: data furnished by the AFR.PL indicated that

the MXA-L50 -Laterial was at Ieast c(mparable, in terus of performance, Lo

the HiUveK-41l asbestos4-phs.nollc mu.tetrial used in t!"e Sizing Studien. Therefore,

the thickness for the exit cone was sized the Name A. for the Itavir-41

material. The performance of the tronsides Resin Co. DP5-lb (formerly

deLignated DPS-lhO) silica-phenolic material in the AFRPL material screening

program indicated an erosion-char rate of 0.90 inches/I0 seconds under the

most severe conditions. Therefore. the chamber linrter was des igned for 2.0

inch thickness "which would rrault in about 0.25 inches of virtin material&t the end of the 120 second firing duration.

UNCLASSIFIED

.- ..-- : . ""I'imil qlllp'l"•'P ln 'F''"l IIf• 'I ~

UNCLASSIFIED Page9 2-27

ll.. t~- w

Tx.

aK-, -.4 Y

1:..... . . .I. C 4-~ a

UNCLASSIFIED

11199-6007-U8-00UNCLASSIFIED Page

(U) The fabrication methods proposed for the number 3 chamber liner weretape-wrapping of throat insert at 60' orientation to the centerline, auto-claving, mnachining of the O.D. and secondary bonding to the steel shell.The DPS-161 silica-phenolic was cast and cured in place at room temperaturein the s manner as the configur'ation 2 liner. The MXA-150 exii coneliner was designed as compression molded (1000 psi), segmented panels,secondarily bonded into the pressure shell. Nine panels, each coveringa 40" sector, were used to an expansion ratio of 2.25 and twelve panels,each covering a 30" sector, were used to the nozzle exit. The design ofthe number 3 ablative liner Is shown in Figure 2-33.

2.3.4.2 Fabrication

(U) The 1404363-2 liner assembly consisted of an MX-2600 silica-phenolictape wrapped throat section, a cast 1DP5-161 silica-phenolic chamber and 4ome"section, and an exit cone fabricated from compression molded N4A-150 asbestos-phenolic material. The components were assembled into an X404342-2 pressureshell assembly. The MX-2600 throat section was fabricated by CompositeTechnology Inc.. Van Nuys. Calif. and assembled into the pressure shell.San Rafael Plastics Co.. San Rafael, Calif. cast the Ironsides Resin DP-5-161 directly into the pressure shell in two sectious; they also compressioniolded the exit cone gewnts and assembled then into the pressure shell.

(U) The properties of the materials used in the X404362-2 liner assemblyare given in the following Table 2-5.

Table 2-5. Properties of Materials Used in 1404363-2Liner Aaseably (U)

material DP5-16L 3,X-2600 HUA-ISO

Form Two Part Mixtsre Tape Holding Compound

Lot No. H-355 3--4,A2

Resin -IL-R-9299([I) NIL-R-9299(11)

Filler Silica-Silica flour Silica (7.8%) -

Fabric Re Ln•orcemnt .one Silica (96.0% min) Asbestos

RI.ein Soldi.. Z 31.0 49.0

Volactle conteiit. 1X 7.0 4.3

Specific Grdvity 11.40 1-71 1.b5*

**kLdctd 1_ 1000 poli

UNCLASSIFIED

1, .till

Ft o- =it 4.

s44

k_ *'". ": ..j ~ 4 . F- ý

iw ,. . q F '

'ai L. -.

',if 4j

~N( 1 ' .A~~t'c~~ ~~~-M -r

r 4 ~- Ar e~ rg~~ 4- 4f

4- 4

A"'

I,~~4' -.4~ 4

144

. 4 ,A .. . ,4i4.4 , '4 ,.,*

1`1011.A. , .-w 4

14 r. AW .* 4 2I

M MYY. -

UNCLASSIFIED Page 2-30

(U) The throat section (X404361-4) was tapa-Wrapped on a net, mile mAndfe1at 60* to centerline using techniques identical to those described inSection 2.3.2.1. Following completion of wrapping the part: van vacuumbagged and cured in an autoclave at 100 psi and 290-310*F for five hours.

(U) The part was then machinead, using the mandrel on which it was wrapped,to fit the X404342-2 pressure shell. Figure 2-34 shows the finished throatsection prior to assembly into the shell. Figure 2-35 shows the throatsection with Adhesive being lowered into the shell. The adhesive used inbonding the throat to the shell was Epon 901/B3.

Figure 2-34. X404361-4 Throat section Figure 2-35. X404361-4 Throat sectionpricr to assembly (U) being installed in

(65543-69-7) (65543-69-9) pressure shell (u)

(U) The throat assembly, X404363J-7, was cured overnight at room temperature,then cured in an oven at 240*F for approximately 2 hours. Figure 2-36shows the cured throat in place prior to shipment to San Rafael PlasticsCo. San Rafael, California.

(U) Followiug receipt of the X404363-7 partial thrust chamber assemblyby San Rafael Plastics the interior metal surfaces were gandblasted anddegreased. then rinsed with HEK to prepare a scale-free, lint-free substrate.So primer was used in the chamber-dome sections. Samples made of material

cast against both primed and unprimed surfaces showed no differences in

UNCLASSIFIED

UNCLASSIFIED(U) adhesive quality. The samples also shoved good adhesive qualitiesbetween cured and capt DP5-161. The throat invert was cleaned with HMat the interface between the DP5-161 and the throat insert.

(U) DP 5-161 is a two-component ablative system consisting of a viscousphenolic type resin (DP 5-161R) and a solid hardener (DP 5-161H). Thehardener contains silica fiber-silica flour fillers in addition to thehardening agent. The two components are mixed in the xatio of 0.7 partshardener to 2.0 parts resin. This results in a very viscous mixturewhich is difficult to mix as shown in Figure 2-37. Two different types

of mixers vere tried. A Reynolds Electric Mixer (Hodel 114) was used toachieve a smooth consistency shown in Figure 2-38. This mixer operated ata higher mixing speed than the Kol 11-58 mixer and heatcd the mix to aslightly higher temperature (about 90'F). A total of 21 batches of DPS-161were mixed and cast on a single day. A sample of batch 20 was taken andcured for two weeks at room temperature to check cure properties. Table2-6 shows the weights of resin and hardener for each of the mixtures.

Figure 2-36. X404363-7 Partial Thrust Figure 2-37. DP5-161 Mix at start(65543-69-12) Chamber Agssy (U) (66841-69-2) of mixing cycle (U)

(U) The mized-resin-filler system was cast in iuch the sme manner asthe DC-93-104 silicone rubber used in the number 2 liner assembly. Themixes were not deaerated prior to casting since laboratory tests indicated

considerable frothing of both the mix and resin under vacuum. Entrappedair em out of the mix quite readily as the mixture was cast.

UNCLASSIFIED

LUNCLASSIEiEDTable 2-6

DP5-16i Batch Mi.xtures (U)

Batch No. Resin Wt.: 3jbs Hardener Wt., lb. Wt.Resin/Wt.Hardener

1 14.50 5.063 2/0.70

2 14.88 5.186 2/0.696

3 21.06 7.312 2/0.696

4 22.06 7.75 2/0.705

5 19.13 6.75 2/0.705

6 14.63 5.125 2/0.70

7 14.06 5.00 2/0.713

8 17.88 6.25 2/0.703

9 lb.63 5.685 2/0.684

10 17.57 6.125 2/0.699

11 13.43 4.75 2/0.707

12 18.68 6.50 2/0.697

13 16.82 5.68 2/0.676

14 -"23.62 8.06 2/0.685

15 15.93 5.62 2/0.704

16 16.68 5.81 2/0-696

17 19.18 6.75 2/0.704

18 16.03 5.56 2/0.694

16.94 5.94 2/0.701

20 18.68 6.50 2/0.696

21 16.62 5.68 2/0.676

] TOTALS 365.21 127.10 2/0.696

NOTE: Estimated 14 lbs of six not used in casting chamber and domesection.

UNCLASSIFIED

UNCLASSIFIED

(U) The dome section was cast

first using batches 1-6. A plastermold was used to form the internalcontour. This mold had only fourinjection ports for introducing theDP5-161 into che void. Some dif-ficulty occurred during the casting;the resin flowed up and under the

% polyethylene tape used as a releasefilm on the plaster mold. Thissection was broken out when the moldwas broken out and later repaired.

- Figure 2-39 shows the dome sectionat the completion of the casting,

I" and Figure 2-4G shows the domesection after removal of the plastermold and prior to assembly with the

* chamber section.

Figure 2-38. DP5-161 Mix prior to(66841-69-3) caating (U)

Figure 2-39. X404363-1 Dome section Figure 2-40. X404363 Dome section

at completion of (67989-69-6) prior to assembly (U)(66841-69-1) casting (U)NCLASSIFED

11199-6007-R8-00UNCLASSIFIED Page 2-34

(u) The chamber section was cast using a sheet metal-plywood mandrelto form the internal contour. The mandrel was secured in place using theX404361-4 throat section, which had been installed previously, as an anchorfor the cylindrical section Batches 7-21 were used to cast the two-inchthick section. Figure 2-41 shows the ca,.t DP5-161 near the top of theflange; Figure 2-42 shova the completion of casting of the step-joint inthe cylindrical section. The cast material was allowed to cure for approxi-mately 10 days at room temperature prior to rwmoval of the mold and mandrel.Figure 2-43 shows the chamber (DP3-161)/throitc (MX-2600) interface afterremoval of the casting mandrel.

V . IN

Figure 2-41. X404363-1 2 inch thick figure 2-42. X404363-1 Completioncylindrical chamber of casting chamber

(66841-69-6) section. (U) (66841-69-8) section (U)

(67989-69-11)Figure 2-43. DP5 lfil/NX-2600 Interface looking aft. (U)

UNCLASSIFIED

UNCLASSIFIED Page 2-35

(U) The exit cone was fabricated from MXA-150 molded segments, &econdarfIyconded into the pressure shall using GLOM-O-N XT high te."erature adhesive.The forward exit cone segments covered a 40 degree segment while the aftexit cone segments covered a 30 degree segment. The segments were moldedIn matched cast-iron dies using a 200 ton press. Figure 2-44 shows thetwo punches (forward and aft segments) prior to clean-up. The moldingcavity was put in place on a steam-heated platen on the base of the press,while the punch was attarched to a steam heated platen on the ram. Tem-perature of the casting molds ware held at 3250F. The MXA-150 was pre-heated to 190*7 prior to charging into the cavity. Figure 2-45 shobS theforward segment after molding and prior to release from the cavity.

Figure 2-44. MXA-150 Segment Punches Figure 2-45. Molded ,XA-lO Fwd(66841-69-4) (U) (67989-69-10) segment cavity (U)

(U) Both the forward and aft exit cone segments were molded oversize andrequired trimming. Figures 2-46 and 2-47 show a molded forward exit conesegment set up for trimming. The trim liaes are clearly visible in thephotograph. The nine 40" forward cone, segments and twelve 30" aft conesegments were dry fitted in the pressure shail and final trmming -mdeon the key pieces. The weights of the various segments are given inTable 2-7.

(U) A 3/4 inch wide by 1/16 inch deep recess was routed into the backsideof each segment and 1 1/2 inch wide silica phenolic doubler strip wasattached to each longitudinal joint with adhesive.

UNCLASSIFIED

UNCLASSM'ED Pae-3

rigure 2-46. Rough molded NxA-150 Figure 2-47. ?M1ded MMf-130 rwdfwd segment met-up for segment (rough) set-up

(67989-69-7) triling MU (67989-69-8) for trimming MU

Table 2-7

Weights of Molded HIA-150 Sagments'(U

Panel No. Forvard Cone Panel No. Aft. Cone.(40-) (____ 300)

1. 22.19 1 17.502 22.19 2 16.823 22.38 3 17.004 22.19 4 17.065 22.38 5 16.636 21.82 6 16.637 21.44 7 16.828 22.13 8 16.690()21.94 9 16.82

10 16.7511 ()16.690d 16.13

Total 198.66 Total 201.54 -400.20 lb.

*rIncluden weight of silica doubler strips

Cl)Ke paelUNCLASSIFIED

11199-6007-RB-O0UNCLASSIFIED Pg 2-37

(U) Prior to the start of installation of the sagmeazs the steil shellwas sandblasted and wived clean with MEK. GLOM-ON IT adhesive mixed in

the ratio of 40 parts of A to 100 parts of B was applied to the pressureshell and backside of each panel plus the longitudinal Joint. The panelswere put in place starting with the number one panel (two doubleiw strips)and finishing with the key panel (0) which had no doubler pavel ptrips.Panels were lild forward until they butted against the Lhroat s6.;L02(Z4O4361-4) until tho measured gap betwaen throat insert ann forwa~rpanels beca eadentially zero. Figure 2-48 showe £11 forva'd pianms inplace and bottomed out on the throat iusert. A plywood ring -AhA:h fitagainst the individual panels was used to hold the panels in the properposition until the adhesive had set.

(U) The aft exit cona segments were 4netalled in a similar manner. Thepanels were put in place with an approximate 0.100 inch wedge holding thepanel from the forward segments. Figure 2-49 shows the key aft panelbeing installed in the exit cone. After the key panel was installed allwedges were pulled and the 12 segments were allowed to bottom out on theforward panels. The panels were held in place with a plywood bondingfixture and C-clamps. Approximately 35 pounds of adhesive were used inthe secondary bonding operation rasulting in an exit cone weight of ap-proximately 435 pounds.

Figure 2-48. 9 Porward panels in Figure 2-49. Kay aft panel being(70053-70-4) place (U) installed in the exit

(70053-70-6) cone (U)

(D) •1iishinU operationo consisted of cutting the aft segments to lengthand installing the retainer clips. All longitudinal J.,inLZ. and circum-ferential interfaces were grounA smooth Fi res 2-50 t,-i 2-51 show the

UNCLASSIF.I.•D

UNCLASSIFIED Pae23

st ~ . 4,

(70054-70-5)

figure' '2-51. X404363-1 CbAsber prior to shi~pmenut to ths AMlL (U)

UNCLASSIFIED

11199-6007-KS-COUNCLASSIFIED Page

(U) finished part prior to shipment to the AFRPL.

(U) Measurement& were taken of the exit plane diameter, throat diameter,chamber diameter and location of the throat plane prior to shipment. The

, throat diameter was 26 1/8 inches, t 1/16 inch, while the exit planediameter was 52 1/8 Inches, t 1/8 inch. The chamber diameter was 35 inches.

(U) Material samplea were taken from each end of the MI-2600 throat.Specific gravity samples of the throat indicate a Sp. Gr. of 1.71 for bothsamples. The throat was also weighed (183.5 lbs); using the calculatedvolume results in an apparent bulk density of 1.735. The specificationrequirement for the throat Sp. Gr. was 1.65 (min).

(U) The Sp. Cr. of the MXA-150 panels was 1.54 compared to a specificationrequirement of 1.60 (min). A small sample piece molded at 1000 psi hada Sp. Cr. of 1.65. The cause of the difference is unknown.

(U) The Sp. Gr. of the DPS-161 cast material was 1.38. The material supplierhad suggested a nominal Sp. Gr. value of 1.40.

(U) Examination of the configuration 3 liner assembly at AIUPL shoved con-siderable shrinkage of the DP5-161 material from the time the material wascast. This shrinkage maifested itself in two ways; the bond between thedome and cast material was broken and cracking of the liner was experi, aced.Prior to the test firing the repair of the chamber dome liner Vws attempted.The repair consisted of the back-filling of the void between insulation andshall with a 50:50 mixture of Epon 828-Versamd 115 adhesive using a SDIKOhand applicator to inject the adhesive mixture into the void. This repairwes only partially successful as noted in Section 3.3.3.3.

UNCLASSIFIED

UNCLASSIFIED Page 3-1SECTION 3

TEST RESULTS

3.1 GENERAL

(U) The results of the Task II test effort are summarized in this section.

Detailed discussions of each checkout test series and each long durationtest firing are given in the following sections. In addition, the summary

of the performance of the S/N 001 and SIN 002 demonstration injector@ which

erae checked out in Task I are included for completeness. The hardware

used in Task 11 is tabulated in Section 3.2. The test results presented

herein were derived from computer printout data furnished TRW System by

the AMEPL. Detailed data reduction procedures used at arriving at the

test results are given in Appendix E.

3.2 TEST HARDWARE

(U) The test hardware, including injector configurations, used in theTask II test program is tabulated as Table 3-1. The initial four testfirings (103-106) employed the X404056-1 (S/N 003) demonstration injector

and the Dev-lA heat-sink combustion chamber. The S/N 001 and S/N 002demonstration injectors had been checked out previously in Task I (seeVolum- 1). The first ablative liner firing (107) employed the SIN 003demonstration injector and the number 1 ablative thrust chamber assembly.

(U) Following test firing 107 the development injector with the 06/F2orifice configuration was checked out with the heat-sink combustion chamberto determine if acceptable performance could be achieved in a finer pattern

injector. This configuration failed to produce satisfactory performanceand the 1404056-1 (S/N 001) demonstration injector was used with thenumber 3 ablative thrust chamber assembly for firing 111.

(U) The X404056-1 (S/N 002) demonstration injector was used with the nutmer2 ablative thrust chamber assembly for the third long duration firing (117).The demonstration injectors weigh approximately 1500 pounds. The ablativethrust chamber pressure vessels weigh approximately 2850 pounds and theablative liners weigh from 680 to 1130 pounds. The number 2 ablative lineris the lightest weight design.

3.3 TEST RESULTS

3.3.1 Checkout Firings - S/N 003 Demonstration Ine ctor

3.3.1.1 Performance Data

(U) The only demnstration injector which was not test fired during theTask 1 test progrmi was the S/N 003 iuJuccor. T-&,Is IMj actor vas .. nt---allyidentical to the S/N 001 and S/N 002 injectors. This injector vwa used tocalibrate the system following facility modification and was fired fourtime (103-106) prior to Installation of the first ablative chamber.

(U) Figure 3-1 shows the measured thrust as a function of the nozzlestagnation pressure (Po). Po is computed from the average of PC-l/PC-2

(Figure I-2 of Appendix E shows the location of the inatrumentation ports)corrected to tiroat stagnation pressure for an cc - 2.25 and a y - 1.235.The data for the checkout firings of the S/N 001-ad S/N 002 demonstrationinjectors is shown for comparative purposes. Substantial agreement isshown between the S/N 001 injector checkout data and the data from firings

UNCLASSIFIED

-I

UNCLASSIFIED Page 3-2

(U) Table 3-1. Test Hardvare Configurationo (U)

_ e ing 103-106 107 108-110 111 117

Injector AscemblX

Z403829-20 I

404056-1 (S/N 1) x

X404056-1 (S/U 2) x

•404056-1 (S/N 3) x x

Pintie Tip Assembly

1404408-6 1

X404280-1 x x x x

Oxidizer Orifice Ring

X404108-1 x

1404693-1 x x I x

yuel Orifice King

X403832-4 I

Thrust Chamber Amua.

X403646-11 x

X404361-1 x

X404362-1

X404363-1 x

UNCLASSIIFEI

S...... •" .......1111I1M 11 II

11199-600? -RB-OCONFIDENTIAL Pg -

240 ti = '.wT l 77

230

210

210 *,:f pu¶-

17 M ITVi -4 4f il~~j 2 !

1600220~~~~. 24M 6 8 0 2

CONFIDETA

CONFIDENTIAL '1960-B0

(U) 103-106 while the data from firings 98-100 with the SIN 002 injectorshow a loss in the thrust coefficient of about 4 to 5 percent.

(t) The computed specif ic impulse efficiency, n1 for the S/N 003injector is shown in Figure 3-2 as a function of 'Re' operating mixtureratio. The data is wall, ordered except for the data of firing 103 whichwas of short duration and was terminated prematurely. The data from theS/N 001 injector checkout firings is shown for comparative purposes.

(U) The computed combustion efficiency, n,*, for the SIN 003 injector isshown In Figure 3-3 as a function of the operating mixture ratio.* Thethroat stagnation pressure in computed from the average of PC-1IPC-2 asdescribed in Appendix E, Section 2. Again, with the exception of testfiring 103. the data are wall ordered. The data for the SIN 001 andS/11 002 Injector checkout firings are shown for comparative purposes. Theagreement between the SIN 001 and SIN 003 injectors is very good whilethe combustion efficiency of the SIN 002 injector is about I percentgreater than either the S/H 001 or SIN 003 injectors.

3.3.1.2 Injector Characteristics

(U) The fuel Injector conductance, KIJCF, for test firings 103-106 isshown in Figure 3-4 as a function of the volumetric flow rate. The fueiorifice discharge coefficient computed from the average conductance is0.955 which In in essential agreement with the value measured with theSIN 002 injector during water-flow. The indicated conductance for theS/K 001. and S/N 002 injectors during the checkout firings in shownfor comparative purposes. both these injectors have indicated dischargecoefficients greater than 1.0. The internal manifold losses ((P17-2) -(PIF-1) I sliow good agreement with those measured durn g the S/N 001injector checkout firings..

(U) The oxidizer injector conductance, KUJCO, for test firings 103-106is shown In Figure 3-5 as a function of the volumetric flow rate. Theoxi~d~ter orifice discharge coefficient computed from the average con-ductance is 0.652 which is approximately 5 percent lower than the averageconductance measured with the S/N 001 and S/N 002 demonstration injectors.The differences In conductances between the injectors is attributed toerrors in the measurement Of injection pressure& (or chamber proos-ures)and propellant flow rates. Errors in propellant densities do not resultin significant changes in the conductances.

(C) The specific impulse efficiency for the X404056-1 (S/N 003 demon-stration injector is shown In Figure 3-6 as a function of (Aldr) (P 04

taiparin thea performance ocursulatanomna valu of 1.80 although

insufficient data was taken to define the exact peak.

3.3.2 W4nK Duration Firing No. -1 - X405090 .1 rnuime Asgembly

(U The X405090-1 250,000 lbf thrust (vacuum engine assembly was fired onDecember 1969 for 66 seconds. The engiine assembly consisted of an initialcombustion chamber length to diameter ratio of 1.54, a contraction ratio of1.8 and characteristic length of 89 inches. The thrust chamber linerconsisted of MXlA-150 (asbestos-phenolic) in the chamber-dome section, 141-2600(silica-phenolic) tape--wrapped throat insert, and HX-2600 (s ilica-phenolic)

CONFIDENTIAL

.

CONIDE TIA 11199-600748S-00CONRDLNTIA 3.5

86

88

1-T

FU?~I311199.-6007-R-O00cONFIDLNEMPage, 3-6

7.6. ... .

6.0 7.5....5.9zrVouerc lwRt. t1e

Fi5 e35.xdie7netrCodcacSN 0 netrU

5.6!qP~ai~e.

A9.

11lt-OO -K4-

890-

111844

-1;r r~ T

1.0 1.5 2.0 2.5 3.0

(AP d / 2 140.4

(C) 'Figure L-6. Specific Impulsae Efficiancy, S/1-003 In~jector (U)

(U ae-up in a rosette pattern in the exit c- -~. Zda firu (107) asche~duld forl 120 seconds but was terminated pgt '7 by a bum throughof Lhe pressure shell behind the throat insert. TL... 'bum through was be-

6lievei~ due to excessive eros ion of the HU1-150 ashes ton-phenolic mteriAlin ams quadrant of the chamber with subsequent ejection of the remaining

allaU4 as fowandresulting in a burn through. The test is discussed

X404361-1 thrust chamber assambly. Checko ut firings of the 5/1-003 Injectorwere made just pr~or to the firing (107) a&d the results are given in Section3.3.1.

CONFIDENTIAL rage 3

(U) The thru•t-throat stagnation pressure relationship calculated for theinitial 1.5 seconds of the firing is essentially in agreement with the dataobtained during the checkout firings (see Figure 3-1). This woul.d be ex-pected since erosion is not expected until a steady-state ablation processhas been established. This process has been estimated to take about 5seconds. The computed specific Impulse efficiency, n1 _, for firing 107is shown in Figure 3-7 to a function of mixture ratio. PAt comparablemixture ratios the performance is about one percent greater than meauredin the cold wall (hest-sink) combustion chamber. This is due to an increaseof 1.0 percent in combustion efficiency since the nozzle efficiency remainedessentially the sane as measured in test firings 103-106.

(U) The initial throat diameter was 26.2 inches which increased to approxi-mately 27.0 inches at the end of ':he firing. The chamber I.D. was 35.0inches at the start of the firink and was estimated to be 38.5 inches post-test. These values were used to determine the contraction ratio as afunction of time which in turn was used to correct the head-end pressureto throat stagnation pressure. The throat statnation pressure and measuredthrust are shown in Figure 3-8 as a function of time. These values weredifferentiated as discussed in Appendix E, Section 5 to %rrive at theerosion rate (dr/dt) as a function of time which Is shown in Figure 3-9.The erosion rate of the M11-2600 throat insert is somewhat lower than pre-dicted from the sub-scale test results (see Appendix B).

3.3.2.2 Injector Characteristics

(U) The fuel injector conductance, KIJC'F, for test firing 107 is shownin figure 3-4 as a function of the volumatric flow rate. The data for thecheckout firings is shown for compaative purposes.

(U) The oxidizer injector conductance, XIJCO, for test firing 107 isshown in Figure 3-5 as a function of the voluaetric flow rate. The datafor the checkout firings is shovn for comparative purposes.

3.3.2.3 Ablative Material Performance

( 6 The average erosion rate of the MX-2600 silica-phenolic throat insert(tape-vrapped at 60* to f was 6.0 mils/second. This is slightly lowerthan rho predicted value of 7.3 mils/second which was obtained in the sub-scale test program (see Appendix B). The throat erosion was nearly com-pletely symetrical with one cool streak of approximately 1.0 inches widthlocated at 270* from the fuel inlet (clockwise). Although the insert wascracked and portions were a~ected It is believed that these cracks and lossof nmaterial were the result of the failure of the material upstrem of thethroat insert.

(C) The erosion rate of the 11X-2600 sllica-phenoo -c exit cone liner(rosette lay-up) vaj slightly higher than predicted (02 mils/sec) due tolocalized heating. The 36 struak pattern created by the injector andobserved in the chamker section was obsurved in the exit cone. Theobserved erosion rate in these streak aeas was 3-4 ails/second. Some

ONE--

11199-.6007-R8-00

CONFIDENTIAL Pag 3-9

89

88

87

nlay 86

85

83

2.0 , 2.4 2.6 2.8 3.0

mixture R.atilo

(C) Pigur.e 3-7. Specific Impulse Effrcjiency, Test Firiug 107 (N)

295 220

218

290

285

Ifit

280

212 0 10 ,O 30 40 50 60 70 s0 90 100

Time, Seconds

(C) viWrs 3-8. Th-ust and Stagnation pressure, Test Firiga 107 (M)

CONRDENTIAL

CONFIDENTIAL Pal 3-i0

-4 H

ITI

I VI . OL:I

I..-!* '0

4--r

I I .:jF~iw; 1d

CONFIDENTIAL

CONFIDENTIALCONROMAL Page 3-11

(C) delamination of the exit cone plies occurred near the exit plane.This separation is typical of rosette lay-ups.

(C) The erosion rate of the MXA-150 (parallel-to-surface layup) in thechamber at a point 24 inbhes upstream of the throat averaged 17.5 mils/second (8.0 to 28.3 ails/second). The greatest erosion occurred at 150'(clockwise) from the fuel inlet which corresponds to the section where aportion of the liner was ejected during the firing. At a point 12 inchesupstream of the throat (start of convergent section) the average erosionrate was 10.8 mils/second (5.5 to 15.6 isils/second) with the greatesterosion again being at 150" from the fuel inlet. The circumferentialvariation in erosion rate is probably due to a combination of mixture ratiovariation caused by the manifold maldistribution, and material nonuniformitydue to the method of fabrication (molded insitu at 100 psi). Examinationof the test hardware post-test indicated that the portion of the chamberliner (800 to 150" clockwise from the fuel inlet) which was jected duringthe firing probably delaminated the upstream portion of the throat insert.This exposed the bond line to hot combustion gases which eventually led togas flow behind the insert and burn through of the pressure shell.

3.3.3 Loan Duration Firing No. 2 - X405090-3 Engine Assembly

(U) The X405090-3 250,000 lbf thrust (vacuum) engine assembly was fired on12 December 1969 for 83 seconds. The engine had an initial combustion chamberconfiguration identical to the first long duration firing. (L* - 89 inches,L/D - 1.54 and cc - 1.80.) The firing (111) was scheduled for 120 secondsbut was terminated prematurely by a burn through of the pressure shell in theexit cone. This burn through was believed due to excessive erosion of one ofthe forward, molded MXA-150 panels (-150" clockwise from the fuel inlet) whichresulted in ejection of the aft, molded panel just downstream. Ejection ofthe first aft segment led to the subsequent ejection of four (4) additionalsegments during ±he firing and the remaining seven (7) segments were ejectedduring the shutdown sequence. The loss of insulation in the one 40* sectorresulted in two local burn throughs of the pressure shell. The test firingis discussed in further detail in the following sections.

3.3.3.1 Performance Data

(U) The second 250,000 lbf thrust (vacuum) engine assembly tested in Task IIconsisted of the X404056-1 (S/N 001) demonstration injector and the X404363-1thrust chamber assembly. The checkout firings of this injector were madeduriug Task I and are reported in Reference 1. This injector was alsoused in the durability firing (number 78) with the low-cost insulated heat-sink combustion chamber. Prior to test firing 111 the char layer on thepintle tip was removed by grinding and the pintle tip was refurbiahed usingV-61 rubber. The thrust chamber liner consisted of DP5-151 (silica-phenolic)in the doaec.-Ae sBc~tion, L(-2^600 (sailica~=phcnolic) tape-v~rappcd throatinsert, and MXA-1ýO (asbestos-phenolic) molded segments in the exit cone.

(U) The measured thrust-throat stagnation pressure relationship shownin Figure 3-1 for firing ll indicates a thrust coefficient approximately1.5 percent greater than that measured during the checkout firings. Thisdata is for the initial 5 second period; (as noted) in Sertion 3.3.2.1no erosion of the throat was expected until a steady-state, ablation processis established. The computed spec:ific impulse efficiency, nj , for firing111 is shown in Figure 3-10 as a function of mixture ratio fA comparablemixture ratios the performance is a maximu, of two percent greater thanmeasured in the cold wall (heat-sink) combustion chamber during firings

CONRDENTIAL

I,- *N'*' =Y 1 l II1•l'• ilII•IB•I 'Im[lll •ldm

CONFID NTIAL 11199-6007-RS-00

88RD N A iag 3-129 7 :

.....86:

87

84F -. S: IM:

2.02. 24 26 .53E85ue ai

(C) igue 310. pecficImplse ffiieny, Frin il (t

(U) 2-7. Tis i du toa icrese f aproxmatly oe prcet -Ein

and fct n raid ressre ecayMaxobsrved uatilo rplat avswr

clomedsupon efiisual osranio apofimthel bune percent in h .:.tnzl

(U) Teienitiale thrat dimeteasue 2616 tsfiinghs, Pot-es-75.ctono

()The maue throat shovedzzsomstngnutioorpitysaseindicated in Figure3-.Th3-1erae diantern post-tes fiisg aproimaely 26.4pat indch.the camunerr iner-sowith time. Nobypssn of the firoatva3. insetochreds. Posttest thein di0-andtir fact estrated prossbe 36.0y inchs. Thesere aunti weroelusedto daletermiethoed cotrpton risatiobaserafntion of time bunhicuh.i unwsusdtorc

the Thead-nitpesurat throat simtewasgnat68ioncpressure Pothtestrot tinspetion opressrn thro staaue showesoenun inmigur 3-1 iniate an Fiuncinorie 3w2.Teediferaentiamted post-discuse in appendimael 26. Scionchs 5Tohrie catmter ineroso

pratsue andr/t)hicht iaus shown in Figure 3-11. The eromion rt of thme were60

throat Insert is considerably lover than predicted from the subic-ale resultsand is also less than measured on test fir.ing 107.

CONFDENIAL

CONRDENTIAL Pt31

300 226

Id

S295022

S.0 .9 4

N" 220 -

285218

0 10 20 30 40 50 60 70 80 90 100Time, Seconds

(C) Figure 3-11. Thrust and Stagnation Pressure, Test Firing 111 (U)

26.350

•J / ,-..26.575€ L-26.•57

(C) Figure 3-12. X404363-1 Pout-Test Threat Measureamnt (U)

CONRDINTIAL

CONMDENTAL Pap 3-1

I7 1L.7 1

rjifr ~ ~r'Alit!.LIif 11

tNB2LIUQL ~I.LL

r~ Qdrf 14

1 FriI ~

tlILl

K L K ~ tjj~ t .~............ ~ I

' r04'.

. .... . _

I 1129-6007-g-'00

CONFIDENTIAL Page3.3.3.2 Iniector Characteristice

(U) The fuel injector conductance, LLJCF, for test firing 1ll is shown inFigure 3-14 as a function of the volumetric flow rate. The data for thecheckout firings (72-77) is shown for comparative purposes.

(U) The oxidizer injector conductance, KIJCO, for test firing Ill isshown in Figure 3-15 as a function of the vol=Gtric flow rate. The datafor the checkout firings is shown for comparative purposes.

(U) The injector conductances measured during firing 11 are bothconsiderably lower than those measured during the checkout firings of theS/N-OOI injector. There is insufficient data to ascertain whether thedifference is due LO errors in the volumetric flow rate measurement or inthe pressure measurements, or in both measurements.

3.3.3.3 Ablative Material Performance

(C) The average erosion rate of the MX1-2600 silica-phenolic throat insert(tape-wrapped at 60' to J) was approximately 2.0 mils/second which is con-siderably lover than the measured value of 6.0 mils/second obtained on testfiring 107. The roat erosion was nearly symetrical with slightly greatererosion occurring in the 260*-300* (clockwise from fuel inlet) section ofthe insert. There is nothing in either the fabricating process or theoperating conditions to indicate a reason for this difference in erosionrate. Therefore, it is postulated that the ablated material from the com-bustlon chamber was substantially cooling the boundary layer which in turnfloved through the throat region with a greatly reduced local recoverytemperature. Both inserts used in the 107 test firing and the 111 testfiring were wrapped by the same fabricator using material from the samelot. Process variables were identical and the apparent density of the twoparts were nearly identical. The operating conditions were within theaortml design range and although the mixture ratio on firing 111 was slightlylower than on firing 107 it was not expected that the erosion rate of thesilica-phenolic throat insert would be extremely sensitive to mixture ratio.

The erosion rate of the molded HXA-150 asbestos-phenolic segments20 inches downstream of the throat averaged 10.15 mils/second (5.25 to11.15 mils/second). The 40' segment at 150* (clockwise from the fuelinlet) was completely eroded (16.15 ails/second). The high erosion inthis 40' sector caused ejection of the molded aft segment just downstreamand subsequent loss of four additional panels during the firing. Seven aft

are v ajact d upon bhutdovu. Thw erosion pattern at the throatiexitcone interface is shown in Figure 3-16. The erosion rate of the MX-2600throat insert at a point 6.5 inches downstream of the throat plane wasI ail/second; just downstream of the interface the erosion rate of theHMA-150 Is estimated at 12 mils/second. This corresponds fairly well withthe predicted erosion rate used in the preliminary design studies as dis-cussed in Appendix B.

CONFIDENTIAL

11199-6007-R8-OO

6.1 7I__3 T 1

5.6: ;=t - -Z

5.8 2

7.26 .. . ... .

7.1 r

7.05 58.0 5 .5 69 .0Oxdiel Volumeatric Flowrate, ft 3 /s.ec

(U) Figure 3-14. Fudier Injector Conductance, 51)1-001 Injector (U)

7.6 7171 (Thi . . pag; :i* 'niaei 'e.) :: . .: .- i. '

CONMIENMIAL Page3-1

(U) ligure 3-16. Erosion Pattern of X404363 Exit Cone Liner"

(C) The erosion rate of the DP5-161 silica-pheaolic material in thechamber at a point 24 inches upstreams of the throat vas 3.3 mils/secondwhich was considerably lass than would be predicted from the AMRPL material

screening program. There vere no deep gouges at the point of Impingementof the reacting streams on the chamber vwal as was experienced in test

firing 107. As has been noted previously the mixture ratio was slightlylover than the design mixture ratio; this would result In a slightly coolervail environamit than was experienced on test firing 107.

(U) One large portion of the dome liner (covsringl-.15) was ejected eitherduring the test firing or the post-test purging. The chamber lineH vasalso cracked and smaller asapnts of the liner from an area just adjacent tothe chamber liner/throat insert Interface vere ejected also. These failureswere due to the shrinkage of the DP5-161 during cure as noted in Section2.3.4.2.

3.3.4 LonW Duration Firing No. 3 - X405090-2 f-sine Asseubly

(U) The X405090-2, 2!,000 lbf thrust (vacuum). engine ausasely was fired

on 7 January 1970 for 98 seconds. The initial coubustion chmboz' configura-

tion consisted of a 104 inch L*, 1.44 chamber length to disamter ratio and

a 2.07 contraction ratio. The ablative liner, the thrust chamber, and the

total engine weights were approxiumtely 680, 3250 and 5030 pounds, respectively.

The test firing (117) was scheduled for 120 seconds but was terminated pre-

maturely by a burn through of the pressure shaell in the ronvergent sectiono th. . ....l. . '-.is bur .. thro--- was due to a high localized erosion rate of

the silicone rubber at the cylindrical/convergent section intersection whichexposed the metal shell and caused the burn through. The test firing is dis-

cussed in greater detail in the following sections.I

3.3.4.1 Y'erformsnce Data

(U) The final 250,000 lb thrust (vacuum) engine assembly tested in TaskII consisted of the X404096-1 (S/3-002) deaonstratior injector and the

__ _N___i_-_____- -

CONFIDENTIAL 3-1

(U) X404362-1 thrust chamber assembly. The checkout firings of thisinjector were made during Task I and are reported in Volume I. Thechar layer on the pintle tip from the checkout firings was reuomed bygrinding and the pintle tip was refurbished using DC-93-104 filled-siliconerubber. The thrust chamber liner consisted of DC-93-104 filled-siliacorubber throughout.

(C) The measured thrust-throat stagnation pressure relationship hbowa inFigure 3-1 wav% not verified during the initial 5.0 second period. As notedin Section 3.3.2.1 no erosion of the throat was expected until a steady-state, ablation process is established. The computed specific iinduaeefficiency, nIgp& for test firing 117 is shown in Figure 3-1 as a fmuctigoof mixture ratio. At comparable mixture ratios the performanca is ap.-roi-mately 4.0 percent greater than measured in the cold wall (heat-sink) com-bustion chamber during firings 98-100. This is due to an increase innozzle efficiency of 4.0 percent over that measured in teat firints 98-100(see Figure 3-1).

9 • •.•-• ... 1W J, . ......

' 87s -P "; 1j.... 1 - - ' ••-

84 -.... --- 4 'H--fr-,_•

S_ _4

860 2. .1762. .

(C) Figure 3-17. Specifi£c Impulse Ef•ficiency, That Fir'ng 11)7 (U)

(U) The measured thrust and throat stagnation pressure are shove in Figurer3-18 a s afunction of the firing rime. These plots show a unifoa erosion

of the throat, with tins, up to the time of burn through in the convergentsection. !No rapid pressure decay was observed until propellant valves Veraclosed upon visual obserMation of the burn through.

CONFIDENTIAL(c Fiur 3I7 spcii Imus Efiiny Tes Fiin U7 W

(U) Th mesue thus an thotsanto rsueaesoninFgr

11199-6007-R-O00CONFIDENTIAL -Pg 31

X.,.

7-7-

'45

SadJI ~'i U.... ... .. ... .

.. ... .. . .

A'.. ... .. .. ..

oC -

ofn1: F:

sm; i: 1

ris3.i 000 .... .. ...

TO:;.. . .. ..

7 ' ... ... ... .. . ..

.... ..... a~ .... 3I1 *~ Z t

CONIDNTAL'

m199-6007-18-00CONRiDETIAL Page 3-20

(C) The initial throat diameter via 26.050 inches. Post-test inspectionshowed some nonuniformity as indicated in Figure 3-19. The average diameterpost-teat Is approximately 27.330 inches, which includes the loss of charand iolten silica associated with engine shutdown and the engine purgingsequence. Post-teat inspection showed little or no change in the ID atsections where the char layer wvs still evident. These values were used todetermine the contraction ratio as a function of time which in turn was usedto correct the head-end pressure to throat stagnation pressure. The throatstarntton pressure and thrust values shown in Figure 3-18 as a function oftime were differentiated as discussed in Appendix E, Section 5.0 to arriveat the erosion rate of the DC-93-104 material at the throat is lover thanpredicted from the subscale results.

~26.624

Wnlet

27.216

S27.341

(C) Figure 3-19. Post-Test Measurements, X404362 Throat (U)

3.3.4.2 Injector Characteristics

(U) The fuel injector conductance, KIJCF, for test firing 117 Is shown inFigure 3-21 as a function of the volumetric flow rate. The data for thecheckout firings (98-100) is shown for comparative purposes.

(U) The oxidizer injector conductance, KIJCO, for test firing 117 is shownin Figure 3-22 as a function of the volumtric flow rate. The data for thecheckout firings is shown for comparative purposes.

3.3.4.3 Ablative Material Performance

(C) The average erosion rate of the DC-93-104 filled silicons rubber atthe throat plant was appronimately 3.0 mils/second which is considerably

CO-IDNT IAL

XX ~.. -L

6-4

r~~t T ~-

tj

-r-7 4I..~~

1-1 .'

;i lt I1

LWMAL

11199-6007-Rh-GOCONFIDENTIA Page 3-22

6.10 -

5.4

5. - ,

5.07. a

Fudiel Volumetric Flovra ta, ft 3 luc.

(U) Figure 3-21. Fud~el Injector Conductance, S/N-002 Injector (U)

7Ti ag sunl5ife.

1 19-60071 -R8-0C

CONFIDENTIAL Page9-23

(C) lover than the 10 mils/second value measured in the subscale testprogram. The throat erosion pattern was nonuniform as has been indicated.Some of the nonuniformity appears to be due to the persistence of streaksthrough the throat. These streaks are also apparent in the exit cone.

(C) Approximately 0.25 inch erosion had been expected at the exit plane.Preliminary evaluation of the liner indicates that the erosion of thematerial in the exit cone was almost nil. The char layer was intact andapproximately 0.20 inch of virgin material remaLned. There was considerableevidence of silica flow in the nozzle (exit cone) and one additional hotstreak (in addition to the one in line with the burn through).

(C) Very little erosion was experienced in the chamber section with theexception of the one streak in the throat approach section. It appearsthat most of the char layer from the chamber was lost during the shutdownand purging sequence. The char that remained amounted to about 0.45 inchesthick with about 0.40 inches of virgin material beneath the char. Thechar shoved very good adhesion to the virgin rubber. Prom the appearanceof the throat section the DC 93-104 material appears to be more environmntsensitive than the KX-2600 silica-phenolic material.

CONFIDENTIAL

COKFILIML Page 3-Z4

(Thi page is imclo alfied.)

CONFIDENTIAL PSECTION 4

CONCLUSIONS

(U) This section presents TRW Syitems conclusions based on the results ofdesign studies, evaluation of fabrication methods, and analysis of hardwareperformance during the Task II phase of the Injector/Chamber SealingFeasibility Program.

(C) The measured vacuum specific impulse for the three ablative enginesduring the long duration firings was 88 percent, + 0.5 percent at thenominal design mixture ratio of 2.60 (O/F). The average weight of theseengines is approximately 5290 pounds each with the all DC-93-104 ablativeliner being the lightest weight design. This engine is estimated to weigh5030 pounds and deionstrated a sea level thrust to weight ratio of 43. Thesaximum performance measured during the Task II long duration firings was89 percent which is one percent lover than that measured during Task I inthe 60-inch heat sink combustion. The one dimensional vaporization ratelimited combustion model developed in Task I for sizing low cost boosterengines predicted the one percent loss in delivered performance for the54-inch long ablative chambers. During the checkout of the S/N 002 din-stration injector in Task I a lowered performance level was observed. Sub-sequent checkout firings of the S/N 003 demonstration injector and the threelong duration firings indicate that the lowered performance was due to erron-eous thrust measurements.

(C Long duration firings of the two tape-wrapped MX-2600 silica-phenolicthroat inserts indicate an average erosion rate of 4.0 mils/aecond (2.0 to6.0 ail/second) which ia slightly better than predicted from the subscalsmaterial evaluation program. In the exit cons the performance of the MX-2600 silica-phenolic material was essentially an predicted although thepattern created by the 36 element injector was discernible. Delaxonationof the rosette plies occurred near the exit plane.

(C The perforiance of the DC93-104 filled silicone rubber was slightlybetter than predicted from the subscale material evaluation program forthe most severe condition. The erosion of the DC-93-104 in the chamberand exit cone sections was almost negligible. In the throat section thedesign thickness was adequate. The char depths ranged from 0.4 inches inthe chamber and at the exit plane to approximately 0.8 incies at the throatplane. The char layer shoved excellent adhesion to the virgin rubber whichin turn was still bonded to the pressure shell. DC-93-104 filled siliconerubber we* shown to be an effective low-cost ablator for rocket engines ofth.s size even though the current raw material cost is about 50 p,%rcantgreater than the raw material cost of the silica-phenolic material. Thisis due to the significantly lover fabrication (labor plus processing) costsfor the DC-93-104 material. The HUA-150 components were fabricated by twodifferent methods; the low-pressure, molded insitu chamber liner had amaximnum erosion rate of about 28 mils/second while the maximu erosion rateof the comp. asiaon molded exit cone segments was about 16 ails/second. BIthof these erosion rates are greater than measured in the AMPL materialscreening program at the most severe test conditions.

CONFIDEHTIAL

I' '••"* i'Ill•~ 1 Tlll '1 lll'f~~l~lUl l

CONRND IAL P.4-

(C) The fourth material evaluated in the program, Ironsides Resin Co. DP-5-161, which was used as a chamber liner in the configurstibn 3 ablativethrust chamber had acceptable performance (03.3 ails/second) as a chamberinsulator. However, the cure mechanism for the material results in a volumechange (shrinkage) which takes an indeterminate period of time and thusresults in a dimensionally unstable part. Elevated temperature cure of thecast material results in cracking of the material which is also detrimentalto performance. It is concluded that the present DP 5-161 formulation isnot satisfactory for use in ablative liners even through its erosioncharacteristics are acceptable.

(U) The casting techniques used in fabricating the ablative componentsfrom DP-5-161 and DC-93-104 were limited only by the mixing equipment.Both mixtures are viscous and there is some degree of difficulty in obtain-ins uniform mixtures. The greatest difficulty occurred in nixing the DP-5-161 which requires the additiou of a large amount of solid (0.7 lb.) to2.0 lbs of viscous resin. The DP-5-161 mixing process is exothermic andrequires jacketed mixing equipment for temperature control.

(U) The segmented, molded panel approach (often referred to as the"building-block" approach) to fabricating large ablative components appearsto be a feasible technique. Tapered joints do not provide sufficient re-tention force to hold the panels in placc and positive interlocking of eachpanel appears to be a requirement. In addition, a thick bond line technologymust be developed to hold the panels to the pressure shell.

(U) Tape-vrapping of throat Wsrts, especially for engine thrust sizesbelmy 250.000 lbf thrust, appears to be the beet technique. For booetersie engines, with burn times of approximately 120 seconds, both the tape-vrapped sllica phenolic (Oe-2600) throat insert and the cast filled-uil merubber (DC-93-104) have erosion rate ('%5 uile/second) which result in minim=throat area change. For example, the throat area of a 3 x 106 lbf thrustablative chamber would increase approximately 1 percent in 120 seconds atthose erosion rates.

(U) The program showed that acceptable hardware could be fabricatad witha minima of quality control. It is not anticipated that fabrication ofmulti-million pound thrust ablative liners would require any greater qualitycontrol then employed during this progrom.

(M) There yore no Lustances of combustion stability with any of the demon-stration injectors during any of the checkout firings or the three longduration firings.

CONFIDENTIAL

UNCLASSIFIED P

APPEIDIZX A

STRESS ANALYSIS OF 250K LONG DURATIONTHRUST CIHAMER SHELL ASSEMBLY (404342)

(U) The structural elements of the development injector and heat-sinkcombustion chamber were ana..,7-ed for a static pressure loading of 600 psiat ambient temperature with a minimum required margin of safety of oneagainst pressure loading. The analytical results are given in Reference 1.

(U) The 250,000 lbf Long DL :ion Thrust Chamber Shell Assembly, P/NX404342, is a revised design if the original heat-sink combustion chamberwhich was analyzed previously. Tha major changes in the redesign are(1) the addition of flanges in the chamber section to allow for installationof the throat inserts and (2) the reduction of material thickness in theexit cone. The stress analysis vas made for a pressure loading of 600 psiat ambient temperature.

(U) The analytical results are summarized in Table I and detailed cal-culations are given in the following pages.

Reference I. G. A. Voorhees, Jr., "Injector/Chamber ScalinS FeasibilityProgram!, AIIPL-TI-70- 86 , Volume 1; TRW Systems, Redondobeach, California; April 1970; Confideutial Import.

UNCLASSIFIED

UNCLASSIFIED. Pg -

m.

.P4

Q~U4 P4 1

I.,

V4 0

UNCLASSIFIE

uuflLwawric~u 11199-6007-RE-00IRUW PL3 A-3

p y qQ~j S'OTMO.

OMCNIERD ________150K LOP416 OWaATIOhN THQ4#ut %.~bIOOEL 44 %e

CHAu4BM Sri~j

T¶s AMOMMIr Too4.

MAT'L. T-I %T7fcL

5 3,000 ?St. 0-P t -. L.@ - I

C, 14itw. R N

Il~~i._'K

p 0, 1 fell ~

1,40 -YIC /Iim * I

UNCLASIFIEDevolve" ~ .bg

TAW Page A-4

MOWL

FMMIMGAiAfsrs~a

-W*l O- - CI *40 s+2

~'Ob 1,ht tC, I wq go-I M+

UNCLASSIFIED

UIIIL~i1.I~u~u11199-6007-RS-007WW Page A-5

Muom PA-.

t ..L.

O' - 10

CRu(.. I'&$pLACGM2riT ~Al Me-ý

gym? w, .i, ma.004

#JhuJ~k~s Mu11199-607-18-00Arm Page A-6

Vj4 1 *0

I? %Alto&

%OLT mtraiso* u

I.$l Ts A. 14 44 ,*k LtA)I

- .liO O..M )5~p

(041 M,' M4 0.4v ,Q-IF64 1O

FLAM," B.4POO(M4w1 At A0(

Overawe ý44, Nov. ~o'e ONCLIUMIrff

U11199-6007-R8-00Page A-7

rUAE v 'P1 %P&ONET NO.

CaU9calD ________________

~o'(I'400 Si~WLI Q)tO~'M,4161f~. -. A~Th1, '4%-1:16,rQ, -1loo)

110 01M4, - 16.11 Q , ~1.41 M%2 - 0115#% a-)

'L.. ibi-VWQ I M% - Li)-L

1,OS10 M,* 0.f 6~ M, 314 at- 4-1 0".f ~

-1'o' M%- 1-S061 M%1 U7 4 % ILr,' -+%IIoo

-1 s - 4-~, dz% -4ýQ Wi

- ¶Mi- O.50 Q$ 4 j.Sj M%.4,4 .%~ -. io

(t,) ~'V' 4-6 ' It X0 M p1 9~ '5*i. IL I t (oO

0.1. a%16..1AMM4 TV'l 4. XIT Poo

UNCLASSIFIEDA~VRUMS f4I NOV. b"?

7XlI 11199-6007-R8-00m ~ 4 9 P f~e-- - age A-8

ON-ne n - w-omonm

G 1 ItWL J60

UNCLASIFIEDoveTum eg *CCNv. &or

Page -

C____________1_______C________A_______L_______

A I~b

(~-X -6 1.01.. 14%*5

C '4 1.1

u U pv!5"~~,~~O

if~M4~. 1 0OO

r I

Ovl z 40Nv 4

~~~o~~~Pg A-oOSl 10W Q

t=, L41Q U-1 M40

04~i * ~aL~ -~ QA4 M

-6711 AL*1 0 CM1

+1 "1.

Q,~~ ~ ~ ~ - 1,T ., &1 1

+ LASIIIML

unto~pairmu11199-6007-RB-00

Pape A-11

Ui1 1 0 4'3.0M. 1-1 bou

qooo .6~Ie

MI.1

(1) 41.8 Ili Lab,

UNCL-ASIFIED

UIVUbLMtWNI oLU 11199-6007-18-00YRV Page A-12

00" M&at -* 0OPO SWC.. Ca&...cm.

41'4C60 t 4L~k5

61,1500 fjo. MA%. Owt&A ',vLFACC 1.00

- ~ + +V0 ~1;00 -3Io .19100

: *910?St. MAV.

mt v

-20 5, ILI,( ~

Tggft 4 ý ' 19 50 1% 1. AWAu c %b* 9000-4 'w'150 gfrs ?% o . W

UNCL ASSIFIEDove3rUwo 144o oov. 6.67

UNLwlAItlU 11199-6007-U-GO7MW Pago A-13

0-4 'OAI 5Oo 4 Q.vh4') it Sb - S*1

t C S -)~

* ~vm61 N.Ss UNCLASSIFIED

UhULbOW~IiLU 11199-6007-Rg-00TrAy Page A-14

Go or&=)C om',i - - -~o OOKckvw

MAI~~¶& L.r otC

'TAQ,4¶ AT 'oL2.LC

o lo

II e. *.*I,___

UIIUIL GmIqI ILM 11199-6007-US-CO

7Rw Page A,-15mu umSam mwAeam

IT CsAMM

We+

IKAIV AT~' 9(91U )

0.4" 1I M- a.;* -I) o

UNCLASSIFIED4VSweb 0"4d Nov. 0-09

mar Page A-16

Cog 14, cea

3 vi

IN,

-Z ICAL.

v

II

1nL& O^%" tQ&T $%%4

(1 Q0

~MSQ~sa"A

UNLASFE

UNCLSSIFIED 11960-B07NIV Page A-17

DiOf ~ tioku'.A.mtot' AT TtOROAT

to n. M O 1. is coo)

.~~~ ~~ to4 (-vn 9 .. ~

h~OlL Y~h4 IT 'no#.

-9 Iio ev . s -V~~ @

Ie C- ; 6.I 61' l

11.1 1L4 441.iYfR'L %1 00j

k1 i I* m- m

UNCLASSIFIED

11199-6007-R8-00?NIV Page A-18

0-ý W" *~ -gowcf inAc C&,*CWOA

t RNOO Ofl.

Ill 6o' 4 41a 0..-, 4 4% 4400v b

lo 4,0^

* ~~ ~j Qi'~~~j,~ ,.5. ~MINI±.'I.I*

610

9L 'I@ ra 4o

itQ~4s ~UNCL" TFQEA

UNCLASSIFIED l--

APPENDIX B

SUBSCALE CHAMBER LINER MATERIALS

EVALUAT ION PROGRAM

I. INTRODUCTION AND SUMMARY

1.1 INTRODUCTION

(U) TRW's Science and Technology Division has been investigating newtechnology related to the goal of developing low-cost space launch vehi-cles. The major technological effort has been the demonstration of anengine design concept which would be scaleable in terms of performance,stability, and durability and which could be manufactured uiing industrial!father than aerospace fabrication techniques.

(U) It support of the Air Force funded Injector/Chamber Scaling Feas.bilityProgram (Contract No. F04611-68-C-0085) TRW Systems undertook an IR&Dprogram to investigate the feasibility of low-cost ablative liners for the250,000 lbf thrust engine designs. The Subsczle Chamber Liner Materi&lsEvaluation Program had the following objectives:

e Evaluate candidate 250,000 lbf thrust ablative chamber liners at

a reduced thrust level (1500 lbf)

* Generate ablative performance data necessary for designing the120 second firing duration, 250,000 lbf thrust chamber liners.

* Recommend additional materials for future material evaluationprograms.

(U) The materials evaluated are the result of an industry search ofpotential candidate tank lining materials, reentry heat shield materials,launch pad protective coating materials and high temutrature ablativematerials. The current raw material costs ranged from $2.00/lb. to$10.00/lb. end all candidate materials showed the potential of an In-placecost of less than $10.00/lb. for production quantities.

(U) The program was conducted in two phases. Phase I (injector char-acterization phase) consisted of determining the smooth wall thermo-

dynmic environment as a function of injector operating conditions forfixed injection orifices. Heat sink hardware was used tc determine thelocai thermodynamic enviro-•-ent thhroughout the combustion chamber andconverging-diverging nozzle section. Phase It (material evaluation phase)consisted of evaluating the candidate low-cost mattrials to determine (1)char formation characteristics of the char forming materials, (2) chamber,throat, and exit cone erosion rates. (3) circumferential variations in theerosion rates, and (4) the virgin materials insulative ability from thestandpoint of minimizing chamber backwall temperature.

UNCLASSIFIED

11199-6007-RS-00UNCLASSIFIED(U) Limited laboratory investigations were also conducted to obtaiuadditional thermodynamic data on the cand-idate materipls. An oxygen/methane (0 2 /CH 4 ) torch was used (in the laboratory) to simulate oxidizer-rich and fuel-rich chemical environments. Steady-state ablation surfacetemperatures were measured with an optical pyrometer. This surface test-perature was then used with the k dn gas recovery temperature, a Bartzvalue of convective film coefficient, and the measured throat mass lossrate to calculate the effective heats of ablation of the various canditatematerials. The effective heat of ablation is a lumped parameter used todescribe the ablation process by a purely thermal model. This p%rameterincludes all fcrms of mass loss such as chemical erosion, mechanicalerosion, and thermal erosion.

1.2 SUMMARY

1.2.1 General

(U) All testing was conducted using the TRW 1500 lbf vacuum thrust coaxialinjector (see Figure B-1), operating with oxidizer (N2 04 ) flow throughthe center circuit and fuel (UDMiU) flow through the outside circuit. Theflow passages and injection orifices of this injector are similar to thoseof the 250,000 lbf injector used in the Air Force sponsored Injector/Ch•aberScaling Feasibility Program.

(U) The nominal test conditions were based on matching the 1500 lbf sub-scale thermodynamic and gas dynamic throat environments to those anticipatedin the 250.000 lbf low-cost rucket engine. Section 3 describes in detailthis scaling procedure. The resultant test conditions were:

Propellants N2 0O/UDIO

Chamber pressure 190 psia

Total flow rate 4.5 lb/sec

Mixture ratio 2.0 tO 2.7 (O/F)

Contraction ratio 2.4

Expansion ratio 2.8 (optimum at sea level)

Cavitating venturis were used to maintain constant propellant flow ratesduring each test. Redundant measurements were taken on all criticalengine performance parameters.

(U) Seven iniector characterization teast were made in a heat-sink thrust

chamber to determine the smooth wall thermodynamic environiment thi-ilou• tthe thrust chamber as a function of injector operating conditions. FigureB-2 shows the heat-sink thrust chamber assembly; the assembly was instru-mented with three steel Nanmac thermocouple probes in the chamber and sixisolated copper calorimeter plugs in the nozzle section.

(U) The results of the seven heat-mink injector characterization tests arepresented in Table 1. Test firings HAl-489-itAl-491 were used to determinethe chamber thermal environment over the mixture ratio (0/F) range of 2.0

UNCLASSIFIED

UNCLASSIFIED

STEEL~~P C PMBOPW P

NANMAO MOUS COUPLE PT

PINT PER AL A$SISEPy

l Figure B-I. 1500 lbf Thust (vacuThum) Cihl-aubet

' •' •Coaxial rnjector Assembly (U)

0XIDW~CPPE NOZZLEpql T

NANMAC THERMCOUPLE POlk

UNCLASSIFIED

lull INC*NO K

UNCLASSIFIED Page B-4

(U) to 2.7 Figure B-3 presents the combustion efficiency (n C k) based ontheoretical shifting equilibrium calculations as a function of mixtureratio. Also presented in Figure B-3 are the chamber and Lhroat heat re-jection data as measured in line with the injector fuel inlet. Both curvesmaximized at a mixture ratiu of 2.4 (oxidizer/fuel). The petasured gasrecovery temperature at this mixture ratio vail 4850*F in the chambersection. The throat copppr calorimeter plugs resulted in longitudinal heatlosses which removed the possibility of measuring the throat gas recoverytemperature. However, since a value equal to 90 percent of the theort-ticalflame temperature vas measured in the chamber section it yas assumed to beconstant throughout che thrust chamber.

Table I. Subsecale Chamber Liner Materials EvaluatiunProgram Injector Characterization Test Results (U)

Nasal. CliracterstitceTest 'Mutur. Total sttgnawaq Ealmus C orrbusuaft

R"a Ju4r aou Ratio Flow Rate 4, PlO APIV Prossu. Velocity r I(caemcyNo. (etc) (0/) C:bIetcI (p0e ) (psi&) (psia) (ftfsec) (%)

1.,A-4(L .0 L,4) 4.49 111 40 16.a 51)11 93.4

I(AI.413 1.0 2.09 4.51 10 1 51 191.$ 5447 4.I

IlAI-4-4 4. 3 1.44 4.SI 134. 39 189. 1 50 94.9

IIAI-445 6. 1 2.41 4.50 116 39 148A. $568 "4.7

IIA 1-48'. 4. 1 1. 46 4.5•4 IZZ 40 1". a $941 .

SParlorn•m It, " 'd 4an wi dol.- tr..m dch.a.heu pousmur. corrected to asisa. o taga"m e.so Percent pay~riuuu e as h6&P,4 on Uheretical a)-iltim wiluililprior caciltins.

1,0 0

C"IKI TWLL'MM RKLNTS N26 /Isa •Pmom

.. CMO•AN FMSU I" 0I[ 083 JOYA •iC• - M loal •me "s• 3MI/C-

I-I-

INRTON OFCLAfITGN..0111) -

PONXUU (33151 90 F1 I AACtjLON RO~GATIO.S 9((P

Figure 8-3. Chamber and Throat Heat Rejection Data and

Combustion Perforwanci (n A) as a Functionof Mixture Ratio for t0 ý 1500 1bf Injectorwith Fixed Injection 0 ifices (U)

UNCLASSIFIED

11199-6007-RS-0UNCLASSIFIED Page B-5

1.2.2 Ablative Haterial ,valuation Firings

(U) The purpose of the ablative material evaluation tests was to screenand evaluate various candidate low-cost ablative materials for possibleuse in the 250,000 lbf low-cost long-duration thrust chamber assemblies.All materials evaluated showed promise of less than $10/lb "in-place"costs for large quantities. The ablative performance data generated focthese candidate materials were used to design baseline thrust chamberliner assemblies capable of ablatively cooling the 250,000 lbf rocketengine for a single 120-secon' continuous burn. A mixture ratio of 2.4(oxidizer/fuel) was chosen fur evaluating the ablative materials since thiscorresponds to the most aevere thermal envirooment measured In the sub-4cale injector characterization test firings. At this mixture ratio, thethermodynamic and gas dynamic environments were:

Gas Recovery Temperature 4850OF

Smooth Wall Shear StressChamber 4pofThroat 36pafExit Plane 27psf

(U) Twelve low-cost ablative liners and a baseline silica-phenolic linerwere tested. The ablative thrust chamber assembly euployed on thesetests is shown in Figure B-4. All tests were terminated after essentiallythe same amount (percent) of throat erosion had occsrred. The calculatedchanges in the thermodynamic and gas dynamic environments because ofthis erosion are shown in Figure B-5. The throat convective film coefficientdecreased 35 percent and the throat smooth wall shear stress level de-creased 45 percent. The materials evaluated are described in Table 11.

(U) The easired head end chamber pressure deray rates for the four ooitpromising ablative materials are shown in Figure B-6. For constant pro-pellant flow rates and assuming constant combustion pecformance (character-istic exhaust velocity) the instantaneous throat erosion rate can becalculated from the measurable instantaneous chamber pressure decay rate.Figure 5-7 presents the instantaneous throat erosion rates of the fourcandidate materials. The throat eiosion rate decreases nearly linearlywith the measured head end chambtr pressure. Since the convective filmcoefficient is dependent upon chamber pressure to the 0.8 power (seeSection 3), it appears that the throat erosion can be described by a simpli-fied thermal model. In essence, tht model defines the effective heat ofablation which groups the thermal, chemical, and mechanical erosion intoa lumped parameter. Table III presents the effective heats of ablationand steady-state ablation surface temperatures for the uandidxte ablativematerials. The surface temperatures were mea1ureA with an optical pyro-meter in controlled chemical environment laboratory torch tests (seeSection 3.3.6).

(U) Figure B-8 shown the throat radius history as determined by integratingthe instantaneous throat erosion rate and the calculated value using theeffective heat of ablation presented in Table 1II. The agreement in thetwo curves verifies the validity of the thermal model for deecribing thethrý;at erowAon process.

UNCLASSIFIED

-- - -

UNCbLiASSIFIED Page B-ý

SILICA LINU OMX-2625i

I.2A

AXIA DAATO

Figre -4.150 lb Ablaivonent Thrust ofme AsseThroy

Eiroiong (Ueo) yTmeaur ~5F U

UNLASIIE

UNCLASSIFIED Pg -

Table I1. DescriptiOn of Materials Evaluated During SubscaleChamber Linar Materials Evaluation Program (U)

CU"o

Notartal Sourcoe T7 Filler Rolal Notbod (P) wig

141-3600 Frihrit Rigid silta fabric PMomelet Ta4 Wrap Jac

VI-13-104 Of Couming glastowtric C.P., Silicone Cast in place A.T. 2 "nt.3

HiAV5W-4& HAVUG Rigid Agste Fiber Pbeaolic Hold"d saas.I No- 20N-3.3-se cemiai Blect. USigd Silica Epazo Nmlded Seponts A.1. 8.

P leuO•jFib., aalakN AV1G.41P 1 MVIG Rigid Abestos Fiiber Phemlic welded Seames 3.T. 1.AVC&T 4011 AVCO ItSid Silica Fiber Iroay Nblded SRpine. no m.

NowalabT-500--i13 7bermint Sys. Rigid C.P. rbolic CastAbldD-350 Oyma ThArum lataftric RIermsil Nad. lKpy Tromelilabe R.T. Air.1-500-7 Thioras Syst. Rigid C.P. Phsealic CutAilbdAVOUT 3OO AVC0 Ilaetemric Now apea- Coa in ilace We5 A.

U,.'

"R.?7421W st-ca uis"i•d abetms Epay- Trewellable R.T. Amb.platitcs "opigm

(1) C:.p. - Ceamny FroprnteIrr(3) R.I. w P Temprrture(3) Ana. - Ambient Pressure

oc _41-

figure --6. Experimental Chamber Figure 11-7. Calculated Instantan-Preasure Decay lates eous Throat Erosionfor the Four beat Per- late for Candidateforming Ablative materials (U)Mtaterials (U)MtrasUNCLASSIFIED

-N C L I F I E DS-i-T =-"V -*--'V 'i ... .i--•• -V •-'T -~~~~IN ...NOW- - -- • • --- .. .. ... U I

UNCLASSIFIEDTable III. Subacale Chamber Liner Katerials Evaluation

Prusr•-m Allatlve Test Results for the Candi-date Katerials (U)

Steady-State Effective HeatSuzrface Temperature of Ablation

Material (OF) (Btu/lbm)

MX-2600 3340 2400

DC-93-104 3Z00 1600

HAVEG-41 Z880 1000

GE-223-50 3130 875

0.5 r -

MEASURED HEAD-END HAVEG--41-- " CHAMBER PRESSURE DATA GE-223-50

THERMAL MODEL WHICH I0.4 "'' DEFINES EFFECTIVE HEAT D 1SOF ABLATIONJ

O 0.3z

:' 0.2

0. 4

2 o.2

0.00 tO16 15 20 25 30 35 40

TIME, SECONDS

Figure 3-g. Throat Radius BHstory as Calculated From MeasuredHead End Chamber Pressure Decay (solid line). Alsodescribed by the thermal model which Includes mechan-ical, chemical and thermal erosion as a lumped paran-ater-Jefined as the effective helt of ablation(dashed Ine) (U)

UNCLASSIFIED0,

UNCLASSIFIED2. MATEZIALS EVALUATED

2.1 GE24K.AL

(U) The materials evaluated in the program are presented in Table VIand are the result of an industry search of the following:

* Chemical tank lining materials

R leentry heat shield materials

* Launch pad protective coating materials

@ Aerodynamic structure protective coating materials

* Conventional high temperature ablative materials

(U) The resin systems include phenolics, numerous epoxies, and silicone,urethane, and neoprene elsatomers. The fillers include silica fabrics,silica flour, silica fibers, asbestos fibers, and other company proprietaryfillers. The current raw material costs ranged from $2.00/lb. to $10.00/lb.and all materials evaluated showed promise of reaching $10.00/lb. or lesin-place cost for production quantities. The following paragraphs descriiethe five most promising materials.

2.1.1 MX 2600 Siliept-PhetnoLic (Fiberite Corp)

(U) The MX-2bOO maserijgl consists of a silica fabric impregnated witha silica-filled lientiuc resin. It is a conventional rocket engine ablativematerial which can be tape-wrapped, using a tape wrapping machine, or handlaid-up by building up a number of plies of previously cut fabric. Thetooling costs for tape wrapping a part and the labor costs for hand lay-utpprocesses make it a high-cost material when installed in a thrust chamber.It requires both elevated temperature (30MF) and pressure (depending uponageometry of p.art) for curing.

2.1.2 DC-93-104 Filled Silicone Rubber (Wow Corning)

(U) The DC-93-104 is a filled silicone rubber. The silicone rubber ia across-linking of phenyl-methyl polysiloxane polymers and the added fillersare company proprietary. The addition of these fillers increases thematerial's resistance to shear loads without inhibiting its elongationcharacteristics an6 reduces the thermal conductivity of the material.

(U) The material is available in both a thixotropic and nonthixotropic1vw istui r ci - teaiaratura curable an-- .h- a "10 pmr.... e-longatinn

capability. The high elongation rate allows it to expand or contract withits backup substrate. The use of DC primer solution will result in self-bonding of the material to Its back-up substrate during cure. The aviountof shrinkage during cure Os minimal because the catalyst becomes an intekralpart of the polymer.

UNCLASSIFIED

S . . . . . :" , .. . . . , ,_-

11199-6007-R8-00UNCLASSIFIED ,.,.e B--0

A -

o t

a

1.

bi d

w' aj~ A bIa

U 4

UNCL.ASSI0FIEDL.. . D. . =: .•.. . :, • •F,•q•.•• ,., , p.• i:[r• • [• •! 'l•

UNCLASSIFIED2.1.3 HAVEG-41 Asbestos-Phenolic (Have. Industries)

(U) HAVEG-41 is an asbestos fiber-filled phenolic resin system. Thuasbestos fiber content Is 40 percent by weight. The our-gassing of cata-lytic reactions during cure causes about 2 percent volumetric shrinkage ofthe material. It requires both temperature (300"F) and pressure (2CO psi)for curing.

(U) HAVEG-41 is a commercial molding compound used primarily for iLningchemical tanks and pipelines. It is a char forming material and thepresence of the asbestos fibers retain this char. Very little data havebeen generated on this material as an ablative liLner for rocket enginethrust chambers.

2.1.4 HAVEG-41F A.beetos-Phenolic ýHaveft Industries)

(U) The HAVEG-41F asbestos-phenolic is a room temperature cataiyticAllycured version of the HAVEG-41. Its primary use is as an ailativecement for bonding compression molded segments of the HAVEG-41 material.It is a two-part system consisting of a hardener solution plus the asbestosphenolic cement. The hardener solution must be thoroughly mixed with thecement to insure complete curing of the material.

2.1.5 GE-223-50 Silica Filled Epoxy Novalak(General Electric Co.)

(U) The GE-223-50 is a silica filled epoxy Novalak system. It containsa mixture of silica flour and milled silica fibers as fillers for theepoxylated NovaLak resin base. The filler content is 50 percent byveight. It is a castable material which can be cured at room temperature.The silica fibers are spread uniformly throughout the resin and increasethe resistance to shear forces for retnining the char layer.

3. EXPERIMENTAL PROGRAM

3.1 BACKGROUND

(U) To effectively evaluate, on a subscale basis, low-cost ablative linermaterials, the theraodynamic and gas gynamic environment of the subucalerocket engine must be comparable to the full-scale hardware. The AirForce sponsored lnjfctor/Chamber Scaling Feasibility Program consists ofdemonstrating the fe.asibility of the low-cost ablative liner concept ata 250,000 lbf . hr'.t level. The operatLn: conditions of this engine arm:

Thrust level 250,000 lbf

Propellants N204 /U JIH

Chameir pressure 300 psi^

Hixture ratio 2.3 to 2.9 (O/F)

Contraction ratio 2.1

Expansion ratio 4.0

UNCLASSIFIED

UNCLASSIFIED Page B-12

3.1.1 Throat Gas Dynamic Evironmnt

(U) The wwoo*C wall shear stress profile throughout any thrust chamber

may be calculated uning Equation 3-1 where this equation is applicable forgaseas with a specific heat ratio of 1.2 and a Prandtl number of 0.83.

T 2 19ý(.26a0.)1.0 + 1.64 .16i1 .1 (3-1)

(U) The Reynolds number at the threat is calculated from the followinga•quation:

Re: - 1580 Po Dt (3-2)

(U) In terms of nozzle stagnation pressure and throat diameter the smooth

wall shear stress at the throat is given by Equation 3-3

TV 0 0 (PD,- 2 + 0.0737 P0D -0.3 (33)

The units to be u ._ in Equation 3-3 are:

P , pals

DL inches

T 2wlbf/ft

3.1.2 Throat Thermodynaianc Environment

(U) Usin; Barts' equation (Reference I) foa the convective heat transferfilm coefficlant (hg) and assuming constant viscosity (U), constant specific

heat (Cp), constant Prandtl number (Pr) and constant property variationacross the boundary layer (a), the throat convective film coefficientbecomes dependent upon nozzle stagnation pressure, throat diameter andthroat radius of curvature as follove:

0O.8

hS - 2.22 x 10"5 0 (3-4)oO0.1 0.1t Cc

where:

h- in the convective heat transfer film

coefficient, Btu/inZ sec*R

r is the nozzle throat radius of curvature,inches

Reference 1. P. R. Bartz, "An Approximate Solution of CompressibleTurbulent boundary-Layer bmelopment and Convective HeatTransfer in Convergent-Divergent Nozzles," Trans.ASiM, Nov.'515

UNCLASSIFIED

11199-60,07-118-M0UNCLASSIFIEDP is the nozzle stagnation pressure, psia

D is the nozzle throat diameter, inches

(U) The nozzle -agnation pressure, throat diameter and throat radiusof curvature of the 250,000 lbf low-cost rocket engine are 300 psia,26.2 Inches, and 11.5 inches, respectively. According to Equations (3-3)and (3-4) thIs corresponds to a smooth vail shear stress at the throatof 31 lbfift atid a throat convective fila coefficient of 0.0012 Btu/in2sec*R.

3.2 1500 ENGINE OPERATIh%; CONDITIONS

(U) The 1500 lbf rocket engine was designed to operate at a nozzlestagnation pressure of 300 pals and throat diameter of 1.8 inches. Fora constant ratio of thrust level to thrust coefficient the operating levelof the 1500 lbf rocket engine is:

P oD t20 974.5 (lbf) (3-5)

Figure 3-9 shows the relationships between nozzle stagnation pressure andthroat diameter for constant smooth wall shear stress at thG throat,constant -onvective film cuefficient at the throat, and thu operatinglevel ot he 1500 lbf rocket engine. At a nozzle stagnation pressure of170 psia, and a throat diamseter of 2.30 inches, the smooth wall shear stressat the throat and the throat convective film coefficient of the 1500 lbfrocket engine are equal to the values for the 250,000 lbf rocket engine.

(U) A slightly higher nozzle stagnation pressure was selected for evaluat-ing the ablative liners. The reason for this was that as the throat erodesthis pressure will decrease and a greater initial pressure will result ina longer test duration beiore reaching a minimum head end chamber pres-sure uhutdowm criteria. A nozzle stagnation pressure of 190 pois. throatdiameter of 2.25 inches, and throat radius of curvature of 2.5 inches wereselected.

(U) Table V shows the calculated smooth wall shear stress at the throntand the throat convective film coefficient for both the 250,000 lbf andthe 1500 Ibf rocket engines.

3.3 ABLATIVE MAIERIAL EVALUATION FIRINGS

(U) Thirteen ablative liners were tested. Figure i-10 imai, zhe 1500 1bfablative rocket engine assrubly mounted in the HEFTS Al test cell.

(U) Tahle VI presents the ablative material evaluation test results. Alltosts wern conducted at a mixture ratio of 2.4 (O/F) and a total propellantflow rate of 4.5 lb/sec. This corresponds to the most seýere Qhmber andthroat thermal environments measured in the injector characterizationfiringo. All tests were terminated after essentially th- sene munt ofthroat erosion had occurred. The changes in the thermodynamic and gasdynamic environments because of the throat erosion were calculated by

UNCLASSIFIEDi i i i i i i i i i ' .... iFiED - ii i-U-

UNCLASSIFIEDU EQUATION 3-4 •

~5.0___ __ _

i 3"C•- .0204" 1.25 EUTOr© I'2-NCHESI rrc =2.5 IINCHES•

-- -13.0C

EQUATION 3-3

100 200 300

NOZZLE STAGNIATION PRESSURE, PSIA

Figure B-9. Throat Diametvr Versu- Noatle Stagnation Pressurefor ConstAnt Smooth Wall Shear Stress at the ThroatConstant Convective Film Coefficient at the Throatand the operating Level of the 1500 lbf Engine (U)

Table V. Subscale Chamber Liner Materials Evaluation ProgramTharmodynaialc and Gas Dynamic Environments (U)

EnbineParameter

250, 000 1b5 IS00 lbf

Smooth Wall Shear Stress 31 psi 36 psiEquation ( 3-3

Throat ConvectiveFilm Coefficient 0. 001z 2 Btu O. 0.00iS Btu

((Equatiun (3-4 in sec°R "n sec R

UNCLASSIFIED

A -.... --. :: ' " I1 r - k i q - - r ll ••P• i | F I '

UNCLASSIFIED PaeB1(U) assuming that the wall chemical environment did not change. FigureB-11 shows thcse changes in environments. The throat co~nvective filmcoefficient dropped 35 percent and the throat smooth wall shear stresslevel dropped 45 percent.

Figure B-10Ablative Cooled 1500 lbf

Rocket Engine MountedLn the HEPTS Al Text Cell (U)

(U) The tour best performing ablative materials were MX-2600, DC-93-104,HAVWG-41; and GE-223-50. The measured head end exanber pressura decayrates for theme materials are presented in Figure B-12. The instantaneousthroat erosion rates which were calculated from the measured head endchamber pressure decay rates are shown in Figure B-13. A nearly linearrelationship between head end chamber pressure and instantaneous throaterosion rate was observed by compering Figures B-12 and 3-13. Thisindicated that the throat erosion was highly dependent upon the heat fluxinto the material1 since the convective film "±cfcent 1z dependentUpon ch~amber pressure to the 0.8 pawer (mee Reference 1). The thermalmodel which describes this ablation rate dependence upon chamber pressure.requires a knowledge of the steady-state ablation surface temperaturefor the materials.

UNCLASSIFIED

UNCLASSIFIED ,P.e 1--6

SE -

4 4 4 . 4

4.. 0044a

lop

__ .10

II

5uz 1-- - - - - z z

U .N CLASS.IFIED: :.

S- I

UIULASSMEI -Page B-17

z --- -

U-0

S1.2

• W,•

> Z6

-, woo

z0U 0.4

S0.2[

Of 3

U,.

3: - INITIAL CHAMBER10 GEOMETRY

O FINAL CHAMBER0 CR1; .GEOMETRY

AXIAL LOCATION

Figure B-11. Ini~tial end Final Thermodynamic and GasDynamic Environments Throughout the 1500 lbfThrust Chamber Assembly (U)

(U) Controlled chemical environment torch tests vore conducted to de-tevine the steady-state ablation surface, temperature for each of the

four candidate materials. Section 3.3.6 describes theme tests aud present;

the data obtained for additional materials of interest. Table VII givesthe t'fective heat of ablation and the steady-state ablation surfacetemperature wm measured in the controlled chemical environment torch testsfor the fou indidate materials.

(U) Sections 3.3.1 through 3.3.5 discusses the results obtained during theengine testing of each material. Inspections of the photographs indicates

the -ate.T-als char formation characteristics. The baseline MX-2600 silica-phenolic liner maintained a thick char layer throughout the engine.

UNCLASSIFIED

.1r

IU~t~U~e~r~fl 11996007-RB-Q0UII~aM~)~)FIEUPage B-18

UNCLASSIFIED ,.. "1 "-99-,

LU

16 8- -- M--60

< OLC

2 6 10 14 )a 22 263 -0 34 38-42TIME AFTER FIRE SWITCH - SECONDS

Figure B-12. Measured Head-End Chamber Pressure DecayRates far Beat Performing Ablative Materials (U)

0

LU

0 GE-223-50

1 • --: ' • ...L _"V.93 -_ 1 U1Mll0- / f

2 6 10 14 18 22 26 30 34 38TIME AFTER FIRE SWITCH- SECONDS

Figure B-13. Instantaneous Throat Erosion Rate forFour Beat Performing Ablative Materials (U)

UNCLASSIFIED

UNCLASSIFIED PaeB9Table VIX. Subscale Chamber Liner Materials Evaluation Program

Ablative Test Results for the Candidate Materials (U)

Steady State Ablation Effective Heat of AblationMaterial Surface Temperature (Btuilbu)

(O F)

M-2600 3340 2400

DC-93-104 3200 1600

HAVWG-41 2880 1000

CE-223-50 ~ 3130 875

3.3.1 111-2600 Silica-Phenolic (P4.oerite Corporation)

(U Thu 111-2600 silica pheuolic ablative liner is shown in Figure B-14.The silica fabric was oriented 60 degrees to the center line of the chamber.The billet wasn cured at 320*F and a maximum pressure of 5000 psia. Theinternal chamber contour warn machined and the billet was secondarilybonded to the steel sleeve.

71 (U) Figure B-15 shown the post-test condition of the M11-2600 ablativeliner. The test duration was 37 seconds, during which time the measuredhead and chamber pressure dropped frou 205 to 132 psi&. The test was

is terminated prematurely because of possible danger to the test hardware.

-. Y

Figure B-14. MX-2600 Silica Phenolic Figure B 15. H11-2600 Silica PhenolicAblative Liner before Ablative Liner AfterTesting (U) 'resting (U)

UNCLASSI FlED

imaune aneinsr

UNLLA It~tU 11199-6007-fl-GO0UNCLASSIFIED PaeB0

(U) Figure B-'6 shows the post-test profile of the MX-2600 liner. Thisprofile is in line with the injector fuel inlet fitting and is representativeof the average ablative performance for all materials except the HAVEG-41F.Figure B-17 shows the circumferential variations in throat erosion forthis liner. The surface regression rate measured from this photograph is7.3 mils/second.

.9 ,MIi"ý• M YT, ...... ..... !

Figure B-16. Surface and Char Profiles for MX-2600 Ablative LinerTest (total on time was 37 seconds) (U)

(U) Figure B-13 shows (1) the measured head end chamber pressure decay,(2) the corrected nozzle stagnation fressure decay, (3) the calculatedfirst derivative of the throat radius, and (4) the effective throat radiusas a function of time. The effective heat of ablation for this materialis 2400 Btu/Ibm and is based on an ablation (surface) temperature of3340"F.

3.3.2 LC-93-104 Filled Silicone Rubber (Dow Corning)

(U) ThA DC-93-104 lfilled silicone rubber liner is rhova in Figure B-19.The DC-93-104 waterial wva cast into the steel sleeve which was treat•dwith WX'-1200 primer solution and cured at room temperatura and ambientprtassurz. The internal chamber contour was machined.

(U) Figure 8-20 shows the post-test condition of the DC-93-104 ablafivyllner. The test duration was 40 seconds during -i~ich ti" the •--a-urehead end chamber pressure dropped from 204 to 114 psia.

(U) FIgure d-21 shows the post-test profile of the DC-93-104 liner asmeasured in line with the injector fuel inlet. The circumferentialvariation in throat erosion is shown in Figure B-22. The surface regressionrate measured from this photograph is 10.0 +_ mils/aecond.

UNCLASSIFIED

.... .----;-. . . ..- ' - ' ' - . . -_ •

UFICLASSIFIEDPROPELLANTS N 20 4 /UDMN

TOTAL FLOW tAll 4 55 LI/5ECMIXTURE RATIO 2:4510 /F)CONTRACTION RATIO

FUL INIETIL 2

FINAL 1.97

no-60 SiiaPeolcAlt* ~ ~ ~ ~ HA IN Liner NE otSSfie Sae

goa- teiiia hot rfl(diameter ST2.G5ATIOhee). Tea

du1i40a 7seod wt .

120 bec oa lw aea

so itue ato o .( 20 4 /U1I.5

TRATHOTh avRADI e thotISina60 0d 1.40~~ nnnrn

Va. 7.1.l30cn. U

UNCLASSIFIE

UNCLASSIFIED Page B-22

44;Z4

Fibure B-19. DC-93-.104 Filled Silicone Figure B-20. DC-93-104 Filled Silicone

Rubber Ablative Liner Rubber Ablative LinerIBefore Testing (U) After Testing()

a a.~LM 7

(4J a.,. '

Figure Bi-21. Surface and Char Profiles for DC-93-104 Ablative LinerTest (total on timke was 39.4 seconds) (U)

UNCL ASSIFIED

UNCLASSIFIED

FUEL INLET Figure B-22

DC-93-104 Filled Silicone RubberAblative Liner Throat Profile.Shaded area In the initial throatprofile (diameter 2.25 inches).

"-- .. Teat duration van 40 seconds with4.5 lb/eec total flow rate at a

-'T mixture ratio of 2.4 (N204 /UDHH).SJr. ,.The average throat erosion as

measured from this photographr ,was 10 mils/see. (U)

(U) Figure B-23 shows (1) the measured head end chamber presmure decay,(2) the corrected nozzle stagnation pressure decay, (3) the calculatedfirst derivative of the throat radius, and (4) the effective throat radiusan a function of time. The effective heat of ablation for this materialis 1600 Dtu/lbm and is based on an ablation temperature of 3200'7.

3.3.3 HAVEG-41 Abeaston-Phenolic (Haveg Industries)

(U) The HAVEC-41 asbestos-phenolic ablative liner is shown in Figure B-24,The material wan compression molded at 200 psi and 300"F. The internalchamber contour was machined and the billet wan secondarily bonded to thesteel sleave.

(U) Figure B-25 shows the post-test conditiou of the HAVEG-41 ablativeliner. The test duration was 25 seconds during which time the measuredhead end chamber pressure dropped from 199 psia to 110 pasi.(U) Figure B-26 shows the post-test profile of the HlAEG-41 liner as

measured in line with the injector fuel inlet. The circumferentialvariation in throat eroslon is shown In Figure 1-27. The 6urface regressionrate measured from this photograph is 18_11 mils/second.

(U) Figure B-28 shows (1) the measured head end chamber pressure decay,(2) the corrected nozzle stagnation pressure decay, (3) the cilculatedfirst derivative of the throat radius, and (4) the effective throat radiusas a function of time. The effective heat of ablation for this materialis 1000 Btu/lbm and is based on an ablation temperature of 2880'F.

UNCLASSIFIED

UNCLASSIFIED PaeB2PROftLANII N204 IUD-ITOTAL FLOW LATEMIXTI.AE RTIO 4 5 .4/ECCONTRACTION RATIO 2:4 (0./F)

INITIAL 2Z 36FINAL 1.50

C:6

1400 Ijg

Liner ~ ~ k~t Maera Evauai m mat uga(

Fiur r24 IAVO41Aieuom Fiur .25. ut V4 .sbuto

Phenoli Abla2i. D-314Flleiiconr RubernAlaivebaiv ie

Before Tenting (U) After Tenting (U)

UNCLASSIFIED

UNCLASSIFIED Pae-2

.0. if -.P4

Figuze B-26. Surface and Char Profiles for HAVEG-4l Ablative LinerTest (total on time wasn 24.9 seconds) WU

LNL

V Figure B-27

~ .~. &VZG,-41 Asbestos Phenolic AblativeLiner Throat Profile. Shaded area1. the initial throat profile

. (dismiter 2.25 Lacharn). Test~. duration was 25 sec with 4.5 lb/mac

L totlel flow ratea t a mixture ratio

0 f 2.4 (11 0 flmm). The average~~~ v~ hroat er~intn an measuudd from

thim photograph was 18 ails/sec. (U)

UNCLAMSSIFIED

UNCLASSIFIED Page -26

PIROPELLANTS N2 0 4 /UDMHTOTAL FLOW RATEMIXTURE RATIO 4.5 LI/SECCONTRACTION RATIO 2.4 (0/F)

INITIAL 2.38FINAL 1.50

200 - ,.

lo- NN-

ISO- -_7:I -EAO END CHAM PR 'ESS URE

- NOZZLE STAGNATION P-ESSU•- 01400

10 R , AT -ADIUS 1.50

20 - 0 1.20 '

2C 2

60 - -1 .40

0 1.10

0 2 4 6 8 10 12 14 16 18 20 22 24 26Z28

TIME, SECONDS

ilgure B-28. HAVEG-41 Asbestos Phenolic Ablative LinerKaterial Evaluation Test Results (U)

3.3.4 UAVMG-4lF Asbestos-Phenolic (Havem Industries)

(U) The LAVEG-41T asbestos phenolic ablative liner is shown in Figure 1-29.The steel sleeve, was sandblasted and used as an open s•ld for curing

the HAVECG-4uF material at room teaperature and umblant pressure. Theliner bonded to the sleeve during cure; hovewor, the bond was weak andfailed during machining of the internal contour. Isolated voids were un-covered during this machining process and were patched with the parentmaterial.

(U) Figure B-30 shows the post-test condition of the HAVEG-411 ablativeliner. The teaL duration was 17 seconds during which time the mesuredhead end chamber pressure dropped froa 203 to l05 pots. i1Exrd 3-31shows the post-test profile of the HAVEG-4lI liner as measured In linewith the injector fuel Inlet. The circumferential variatton in throat

erosion si s om in Figure -.32A The surface regression rate measuredfro% this photograph is 2 .54J ,ile/second.

(U) Figure 5-33 rhows (1) the measured head nd charober pressure decay,(2) the corrected nozzle stagnation pressure decay, (3) the calculatedfirst derivative of the throat radius, and (4) the effective throat radiueas a function of tbme.

UNCLASSIFIED

------------------- =

IIMI'A('('c~rfl 11199-6OO77*u8-00UNCLASS~)iIFIED Page B-27

r4 4

....... . .: .

Figure B1-29. ILAVEG-41F Asbestos Figure B-30. HAVEC-41F AsbestosPhenolic Ablative Liner Phenolic Ablative LinerBefore Testing MU After Testing (U

ILL

Figure 5-31. Surface and :bar Profiles for HAVEC-41lF Ablative LinerTest (total on time wasn 17 seconds) (U)

UNCLASSIFIED

UNCLASSIFIED Pgl99-

FUEL INLET

Figure B-32

HAVEG-41F Asbestos Phenolic AblativeLiner Throat Profile. Shaded areais the initial throat profile(diameter 2.25 inches). Teatduration was 18 seconds with 4.5 lb/sac total flow rate at a sixtureratio of 2.4 (N204 /UDKa). Theaverage throat erosion an measuredfrom this photograph was 24.5 ails/xecond. (U)

PROPELLANT5 N2 O4 /UDMH

TOTAL FLOW RATE 4.53 LI/•£CMIXTURE RATIO 2.4 (0/F)CONTItACTION RATIO

INITIAL 2.33FINAL 1.5

2AD 4EAND CHAMBER PESSUR

60-- NOZZLE STAGNATION PRESSURE

-100 TE ROAT- -. 60

•o - " •' I l l I, i o -~4 0~ 1.50~

30 1.40

THROAT RADIUIS

I I T, 1.20

2 4 6 8 10 12 14 16 is 20TIM! SECONDS

Figure B-33. ILAVEC-41F Asbestos Phenolic Ablative Liner KaterialEvaluatiOn Test Revalta (U)

UNCL ASSIFIED

UNCLASSIFIED PaeB9

3.3.5 GE-223-50 Silica Filled Epoxy Novalak (General Electric Co.)

The GE-223-50 silica epoxy Novalak ablative liner is shown inFigure B-34. The GE-223-50 material was cast into an open mold andcured at room temperature for 16 hours. The cure process was thenaccelerated by curing at 260*F for 12 hours. The total required curetime at room temperature is approximately 48 hours. The liner wassecondarily bonded to the steel sleeve, with a high temperature epoxyadhesive. The internal contour was machined which uncovered a pc ekatof partially cured material in the chamber section. The specimen wastested in this condition.

(U) Figure B-35 shows the post-test condition of the GE-223-50 ablativeliner. The test duration was 24 seconds during which time the measuredhead end chamber pressure dropped from 200 to 112 psai. Figure B-36shows the post-test prifile of the GE-223-50 liner as measured in linewith the injector fuel inlet. The circumferential variation in throaterosion is ahown in Figure g-37. The surface regression rate measuredfrom the photograph is 18+11 mils/second.

~~, If~

#4N3 -1

Figure D3-34. GE-223-50 Silica Filled Figure B-35. GE-223-50 bilica FilledEpoxy Novalak Ablative Epoxy Novalak Ablat iveLiner Before Testing (U) Liner After Testing (U)

(U) Figure 5-38 shows (1) the measured bead end chamber pressure decay,(2) the corracted nozzle stagnation pressure decay, ('I) the calculatedfirst derivative of the throat radius, and (4) the eifective throat radiusas a function of time. The effective heat of ablation for this materialis 675 Btu/lbu and is based on an ablation temperature of 3130Fr.

UNCLASSIFIED

11199-6007-RU-00UNCLASSIFIED Pae B-30

ON=" LA ma•

a

Figure B-36. Surface and Char Profile. for GE-223-50 AblativeLiner Test (total on time was 23.9 seconds) (U)

""JR INUT

Figure 5-37

GE-223-50 Silica Filled EpoxyNovalak Ablative Liner Throat

, Profile. Shaded area is theInitial throat profile (diamter

- 2.25 :n-hea). Test durationwas 24 seconds with 4.5 lb/sec.total flow rate at a aixtureratio of 2.4 (N2 0 4 /UDWI). Theaverage thrnat erostin asm....... f!-- Chi. photograph

was 18 uILis/second. MU)

UNCLASSIFIED

UNCLASSIFIED rae31

PROMILLANTS N204 /UOMHTOTAL FLOW LATE 4.5 LB/SECMIXTiEI RATIOCONTRACTION RATIO 2.4 10/F)

FINAL 1.50

20TI4.2.

- IEC RAIU 20, 1.40THROAT

3.3.6. 1.30ucalPocdr

Figue -3. Thee-3 datic colFte ie orlated Epoxy thela Anlatples Lioneritto dtermneMaberial combinations ofx binerut a(Uile.)orsino

a.. Conticlrolpoiincudle ad Cais1EvrnetUoradotion Torc Thesutsfe rcl[ or pos3.3.6. fxtordentelPopmentofr ewpout

I - ~(U) Thet experimenta appro achiused to t "a 1-ncandiam ater byaterials

to~ia atis of enietach tochap~nd thecr torhelam degrecatedn an thfncisn hole.acb aTherse dara prepard thn the confilaedratio toe promides chrick dasan detertsinder dfferaet combnditions o idradfla.Lweoino

(9) Fiues spe0cime aer pahtorah tof th tor-ch teasting bacility.althogthe orh cylntdirme. A'se 38icamye boe ewasuatdoriopraed wiouhthe eongierztwdinare axispr af a lf offec sml and thee torhfaiest ere ran tin anoairataocalhestsecmnar .ai U CA IFigueDB3.SvrlseiesfovAc aeilwr rpae uti ofgrtont rvd hc i;

11199-6007-R8-OOUNCLASSIFIED Page B3

I lk

Figure B-39

Photograph of Typical Test Speci-* ~. mens Used for Oxygen/liethjine Torch

Teats. The smaples are machinedto cylinders and a 3/8-inch disameterhole bored through to receive thetorch flame. Percent weight losswith flame exposure time in cal-culated, and reported as torch

(U) To conduct each test, specimens were rapidly swung under the burn-ing torch (shown in a closeup photograph in figure B-41) which has beenpositioned prior to lighting the flame. and the timing started. Afterthe required exposure of 1, 3, or 5 minutes, the sample was swtung awayfrom the flame impingement zons and allowed to cool. Any remainingflames were snuffed out with a fabric mat. During the exposure, a PyroInstrument Company Model 95 optical pyrometer "as used to measure thesample surface temperatute at the drilled hole opening. This temperature,insumed to be the highest experienced by the sample, is reported inTazble VII as the equilibrium surface temperature for that samiple. Deter-mination of the heat flux was carried out by using a copper cylinderof the sa&e specimen dimensions fitted with a thermocouple.

id)A number of programs have been conducted to find a satisfactorylow-cost chamber liner composition. Initial TRW engine P~ests conductedunder an IR and D program found tour likely candidate materials. These

L ~materials, more fully describe ina Section 2 are AA followaz

F~~ j iX-2600, Fiberite Corporatio~a

e DC-93-104, Dow Corning Corporation

e HAVEG-41, Haveg Industries

*GE-223-50, General Elaectic Company

UNCLASSIFIED

11199-6007- RS-0UNCLASSIFIED PaeB3

Mi Figure B-40

Laboratory Torch Test facility (U)

Figure 1-41

Torch Test in 9Progress (U)

UNCLASSI FlED

UNCLASSIFIED PaeB3

kJ- 00 0 C) 00 U, gn

&0 kn t- tn" 4 " 4 4 " 4 "

uu

,- 1

u 0

%.4m L v nf CY00t-C" 00

0 0

00W4

N- f42d-4

"44 W6 OW &

UNCLASSIIE/ --. ~ __________

UNCLASSIFIED(U) A similar program conducted by the AFRPL has shown the followingmaterials, among others, to be attractive candidates:

a MIA-150, Fiberite Corporation

o Plastonium, Insulation Systems InLorporated

& Ironsides DP5-160, Ironsides Resins Incorporated

(U) In all, a total of 12 materials have been tested. The behavior ofthe above seven materials is described in the following section.

3.3.6.2 Test Results

(U) Temperatures at the flame impingement surface were measured byoptical pyrometer and appear to be reproducible vithin±+20"C (*I percent).This observed temperature is considered to be the maximum attained underthe test conditions and is essentially the ablation mechanism reactiontimporaturs.

(U) Additional information was obtained by weighing each specimen tothe nearest milligram before and afte.' exposure. In this way, erosion,as a result of the flame exposure, could be calculated for each specimen.Results are reported on a percent by weight basis. However, as thecandidates' specific gravity values are all between' 1.4 and 1.7 (with oneexception, Plastonium, with a specific gravity of O.B) these data are alsoreasonably comparable to percent by volume. Inherent in the erosion lossresults are weight losses due to volatile products from below the charbarrier. Although not actually shear-induced erosion, these losses arealso a significant factor in determining the ablative and Insulativepropartiae of a material and are an appropriate component of the data.

(U) Figure B-42 shows the samples of the four candidate mater;als fromthe M •TRW W materials evaluation program. Figure B-43 shows the samplesof the three candidate materials from the AFRPL "in-house" test prortam.

(U) Four of the candidate materials were each tested at three flameexposures times---, 3, and 5 minutes. Plots of erosion losses versustime Save relaLively smooth curves as in Figure B-44. The curvesdefining thim relationship era expected as the post-test specimens eachhad a significantly thick "wall" remaining. A sharp deflection in theerosion loss versus time curve could occur if the sample was inhomoge-naws or if the wall decreased its heat sink function. Apparently neitherwas the case.

(U) Both oxidizer-rich and fuel-rich torch conditions were used intesting each sample. In thn laboratory experiments, torch flame tem-perature varied from 2050C for the fuel rich condition to 2670%C f'crthe oxidizer rich condition. The expected flame condition in the angineis betweu 2000C and 2200*C. Figure B-44 shows that generally fastererosion rates were obtained under oxidizer-rich conditions than withfuel-rich conditions. This was expected due primarily to the 620%C increase

UNCLASSIFIED

UNCLASIFIED 11199-6007-R.8-00

d fi

91

.44

UNCLASIFIE

UNCLASSIFIED Page B-37

7-~ ~~~ -.. - c~u

* *~/ '.d-

Uk

L~ S'ALS.

UNLASIIE

UNCLASSIFIED I48

-E ;

SOXIDIZER-RICH-FUEL-RICH PLASTONIUM

42 23

T, EIRONSIDES

36!rl

2 24 0 /

_ _ _ _r I,. , -____orhTu I~mto _ _ _ _~t (_ _

00 DC-93-10412

012345 6

Fiue" .0 2/Cn,6 TrhTs Eoin aut.(U)

UNCLASSIFIED

............... •1 _•

UNCLASSIFIED(U) in flame temperature. One notable exception to this observation wasfound with the DC-93-104 material.

(U) The msot significant result from this text series was the excellentperformance of the filled silicone compounds. The weight lose for boththe DC-93-104 and DC-93-115 (a low flow version of DC-93-104) was lessthan the weight loss for the IX-2600 silica-phenolic material.

(U) Low weight losses found for the DC-93-104/DC-93-115 samples indicatethat the major products of high-temperature reactions are noavolatile.However. the low quantities of volatiles produced serve to thermallyinslate the samples' surface. Ablation temperatures of the surfaceexposed to the flame were similar to the other materials tested in oxidizer(16500 to 17300C) and fuel-rich (1580' to 16500C) envtirouments. Becauseof the samples' low erosion rates and the fact that polysllozanes zetainsome strength at 39000C, thie observed good results for these materialsare thought to be caused by*

(1) The absence of large amounts of volatile products fromendothermic filler and polymer internctions

(2) Inherent stability of the polymeric binder at the testtemperature

(U) Figures B-45 through i-70 shown the post-test specimens of the sevenprimary material candidates. The photographs are somewhat deceiv' g inthat the more erosion-resistant materials are moet unsightly Wt" j someof the least eroxion-resistant materials appear unmarked. The .iollovingobservations of two of the test specimess will aid in evaluation of thephotographs.

(U) The DC-93-104 samples suffered greater erosion dmage on the outersurface than adjacent to the drilled bole. These samples swelled duringthe test which caused the hole diameter to shrink and the flawe to spillaround the outside of the ample. DC-93-104, a room temperature cure(RTV) material, swelled and distorted. A white coating, assumed to be510., deposited on the spoles. The flame-impacted surface was fracturedin in irregular pattern as if this reaction relieved internal stresses.

(U) Plastouium samples wea deceptive In that very slight dimensionalchan4os were found despita the high erosion losses. Strain cracks werefew in number, but deep. Because these cracks appeared fresh and not"polished," it appears as though the cracks might have been formed dur-ing the post-test cooling period. Of the sampiles tested, riastonium badthe mest obvious hole enlILrgemat which conically tapered fro, the flownImpingement surface. Compered with the other r ,terials, Plastonium isstructurally weak. The samples emerged snow vwLte with flecks of black,indicating high volumes of filler and low volumes cf carbonaceous phenolicbinder were present, a condition leading to a weak material.

UNCLASSIFIED

11199-6007-RS00UNCLASSIFIED Page B4

ý77X7

14C

:,4..4 P

MX-60 oxdier-ic 6-Scn MX20 ue-ih,0Scn

Exoue()Epsr U

Figur B-5Fguea-

111e-260 oxidizer-Rich 60-Second 211e-260 Fuel-Rich 60-Second

Exposure (U) Expouure (U) u4

Figure B-47 Figure B-48

UNCLASSI FlED

11199-6007-R8-00UNCLASSIFIED PaeB4Pa77-4

AI

DC9314 xdie-Rc 6-ecn D-3-0 Fe-Rc 6-ecn

DC-9 3-104 Oxidizer-Rich 600-Second DC-93-104 Fuel-Rich bO0-S econdExposure (U) Exposure (Uj)

Figure 6-51 Figure 5-52

UNCLASSIFIE

----------- _

11199-6007-RB-CO

UNCLASSIFIED Pae-2

GE-223-50 Oxidizer-Rich 60-Second Exposure (U)Flgure &-53

CE-223-30 Tuel-lich 60-Second Exposure (U)Figure B-54

U NC L AlS"I FED

UNCILASSIFIIED Page B-43

44

AV.~

'.VA 7.La-~4 .~

itL t. ) .. ... .- Z-" " " I . '

I -

, A,•,.s: " • ' s.•. .. -, s ,-**1

6e 44, , 7

..... ...... .. . .. .

Ironsides Oxidizer-Rich 60-Second Ironsides Fuel-Rich 60-Second

Exposure (U) exposure (U)Figure B-55 Figure B-56

MUNC ASS-. E-.

,Y.. t .. ,, ,. . : .

it 4."-;a-'-"' ;-,-•. ''•.C--'

.-" .. .. " - .- .

• -t".-#. . ", ' '4.

' T- S' " :C'•• ,. "

'• ••,• • "• r. .. ,,T..A,• t • •,,_• ,."- .o

Ironsides Oxidiz r-Ric 120-Second ron ide Fuel-Rich 120 -Second.

Expos re.( ) Ex osur (U) .,,• , .,,. ..,,Fi ur B-5 Fi,-.pur • e ,,• "B-58 "..• -•

•,',•,.,,'-'.':b- --- U N CLA SS.•,:,.,•. : :.;"-•~ I FIED.. . .~ .... . .....

UNCLASSIFIED Paige-07-0 "4

rX.1

IO

Ironsides Oxidizer-Rich 300-Second Ironsides Fuel-Rich 300-SecondExposure (U) Exposure (U)

Figure B-59 Figure B-60 j

MXA-150 Oxidizer-Rich 60-Second KXJL-150 Fuel-Rich 60-Seco-idExposure (u) Exposure (U)

Figure B-61 Figure B-b2

UNCLASSIFIED

UNCLASSIFIED PaeB4

(a W

Irt.

.4

4.4.

MXA-150 Oxidizer-Rich 150-Second KIA-150 Fuel-Rich 180-SecondExposure (U) Fgr -3Exposure (U)FiueB6

7 Ir

A !Tr. I

jUA-150 Oxidizur-Rich 300-Second MXA-15O Fuel-Rich 300-SecondExpor a4 (U) Exposure (U)

Figure B-65 Figure 5-66

UNCLASSIFIED

UNCLASSIFIED PaeB4

P7e7-4

~1~ '747

Plasacnium Oxidizer-Rich 6O.-Second Plastonium Fuel-Rich 60--SecondExposure (U) Exposure (U)

Figure B-67 rigure B-68

i. 40.k

Plastonium Oxidizer-Rich 300-Second Plastonium Fuel-Rich 300f-SecondExposure (u) Figure B-69 Exposure (U)Fiue-7

UNCLASSIFIED

UNCLASSIFIEDAPPENDIX C

ABLATIVE LINER SIZING

1. InTRODUCTION MND SUUARY

(U) The three ablative liners designed and fabricated in Task II weremixed mccording to procedures given in the following sections. The am-terial thicknesses were determined fur a 250,000 lbf thrust (vacum)engin operating at 300 pmda with the N2 04 /UDMH propellant combination.The b ic data used in sizing the liners was developed in the TRW Systemsub-scale Materials Evaluation Program. This data was correlated usingthe TRW Systems Charring and Ablation Program (AH054A) which uses themodel of Munson and Spindler to determine the dimensional ablation.Finally, the erosion correlations were used to size the 250,000 lbf thrust(vacuum) ablative liners.

2. SUi-SCALI RESULTS

2.1 RECOVMY TDETIIU

(U) Initiol sub-scale tests, which were made with a heat mink chamber,indicated that the 1500 pound thruster operated at a recrvery terperaturethat wan representative of a streaking condition. Initial data fromNammac surface thermocouples were reduced using the TRW Heat Flux DataAmalysis Program (HYDAP). HFDAP computes the heat flux frum the productof the thernal conductivity (k) and the temperature gradient at the wall(dT/dr) from the following equation:

wall

The temperature distribution through the wall as a fu-ction of time isobtained from the numerical solution of the thermal diffusion equation.

p W ar'rr 3

This form of the diffusion equation considers radial conduction only. Theboundary conditions are the measured wall temperatures.

(U) A typical surface t'mperature response at the throat is shown InFigure C-1 and the corresponding heat flux is shown in Figure C-2. Datafrom this test indicated a throat recovery temperature of approxinmtely48501F. This tmperature was used in tho subsequent ablation correlationsto obtain the effective heats of ablation of each of the materials. Inadditioc the surface was assumed to ablate at a temperature of approximately32000F. The heat transfer coefficients were determined using the simpli'iedBartr equation.

UNCLASSIFIED

UNCLASIFIED Page C-2

I~~ _WTýj- L4:

I. ~~~7 - NLSI7 E-i

H -4

_ __ t4

-- 4.

4. 44-

i.1r Tv ..L.

1:: i

Figure~~~~~ ~ ~~~~~ C-.Ha-lxVsInrWl eiprtr(F U

UNCLASSIFIED

UNCI" VIED Page

2.2 EROSION CORREA ION

2.2.1 M

(U) Using the erosion and heat flux data frou the sub-scale MaterialsEvaluation Program, correlations were made to determine the effectiveerosion characteristics of four candidate matarials for the 250K thrustchamber. These correlations were generated using the TRW Charring sWAblation Program (AHO54A) which approximates the charring phenoussio byan Arrhaeius rate equation of the form

vhere,:

0 time

p5a- density of the decomposable resin

T - tmperature

4,.8r " Arrhenium rate constants(determined experimetntally)

The dimensional ablation is determinad by using the erosion model of YuaMo*and Spindler where the erosion rate can be specified by an analyticalrelationship, or calculated from the heat balance on the cur face nodes.When the erosion rate is specified. the rate iS given by

5 - +UVa T (4)

where:

heat of sublimation or reaction

T s surface temperature

UVY i empirical erosion constants

In general, at the betinniag of heating, the surface recession rate isgI-van by equation 2 as a function of the surface temperature T. Uponreaching mom critical temperature, such as the sublimation or melt point.the surface is then h1wld at a conustxt -t -- ratre T6 am lon as the hoatinput Is sufficient to hold it at this temperaturs.

(U) For a thrust chamber operating at the high heat fluxes associated withchamber pressures in the order of 300 psi&. the transient heat-up time forthe •iurface temperature to reach the constant ablati4.n temperature isrelatively short. For run tinm in the order of 100 seconds, this tran-sient surface temperature timI is an negligible fraction of the total burntims.

UNCLASSIFIED

UNCLASSIFIED Page C-

:U) For the rate of heat absorbed by ablation can either be approximatedhy a constant surface erosion rate from equation 2) whiere

YT

UT1 + Le Constt at T (5)

or the surface erosion rate can be approximated by solving the heat balanceequation at the surface given by

-k -L [(surface heat Input) - (H~ 9 ) (6)

where:

PC . char density

Hoff - efoective heat of ablation

The effective heat of ablation, Hoff essentially lumps chemical, thermaland mechanical erosion characteristics into one parameter that can be usedfor correlation purposes when the individual characteristics are not wellknown.

2.2.2 Resul

(U) The axperimental sub-scale tests performed during the MaterialEvaluation Program utilized the 1.5K thruster operated at 190 paie. Thegeometric size of this chamber and the 190 puia chamber pressure has beenshown to yield a shear and thermal environment approximating the environ-manst of the 250,000 lbf thrust (vacuum) angina. Therefore,, effectiveheats of ablation can be obtained from the experimental erosion rate andthe chamber pressure decay of the candidate materials. Figures C-3 and C-4show the head-end chamber pressure decay and the erosion rate for the candi-date materials.

(U) The erosion characteristics of the candidate materials tested In thesub-scale Materials Evaluation Program were correlated using equation (6).the thermal environment from Figure C-2 (4850'), and the erosion data fro*Figures C-3 and C-4. These correlations are shown in Figure C-S. Subse-quent to the sizing of the 250,000 lbf thrust (vacuum) chamber liners, thesurface temperatures of the candidate materials were experimentally deter-mined. The effective heats of ablation shown in the Table of Figure C-5are those values corresponding to the experimentally determined surfacetemeratures. The use of 3200F to obtain the calculated curves in FigureC-5 introduced little or no error in the correlations since maxiamm eurfecetemperature deviation represents negligible difference in heat up time.

3. LINKI SIZIUC

(U) The primary criteria used to size the 250,000 lbf thrust ablativeliners was a limiting value of 6007F on the chamber back-wall after 120

UNCLASSIFIED

UNCLASSIFIED ,.,.ge"C-5

120Z - .AE- 4

GE-223-50 DC-93-104

s o

II

02 6 1 '14 18 22 26 30 34 38 42TIME AFTER FIRE SWITCH SECONDS

Figure C-3. Measured H1ead-"aI Chamber Pressure DecayRates for Best Performing Ablative M4aterials (U)

1 20

•. GE-223-50

"a!

2 6 10 14 18 22 26 30 34 38

TIME AFTER FIRE SWITCH - SECONDS

ligure C-". Instantaanous Throat Erosion Rate forFour Beet Performing Ablative Katerials (U)

UNCLASSIFIED

UNCLASSIFIED PageC-6 -4.1

-CMLCULAUIO MA%1G-41EXPERLM1NTAL G1 77-50S

0.3-

MAXIRRAL * IFCCIIV =! fASLATION SuwURFA IMIAWMI ('?I

. C-92-I O i3,o141MG-41 1000 31011

Figure C-5. Thermal Modeling of Erosion Characteristics of YourCandidate Low-Cost Ablative Liner Materials (U)

(U) seconds firing duration. Using the Charring and Ablation Program, andinput data determined in Section 2, the liner thicknesses were calculatedfor three materials. The, required thicknesese are shown In Figure C-6 asa function of the intaxrnal radius. The dasbed likie shown the liner thick-iease required for erosion while the solid line shows the thickness required

to limit the backvall teu"perturn to 600*7.

(U) The test of the sub-scale MX-2600 instarial had not been ccepletod atthe tinte of the sixing of the ablative liner@. Therefore, the requiredthickness wasn assuaed to be equal to the DC93-104 material thick~ness. Sub-sequent testing of the 111-2600 sample Indicated its superiority to theDC93-104 material. I Igure C-5 shows the correlation obtained vith the

UNCLA -SSIFIED

.1J

UNCLASSIFIED Pg -/-

K z

T/ I

<i Ji o.

0 u=W6 s- Zgs

_____ _ __ ____ . I-~Y23

z V46

_ý -

S~~HDNI~ -SSN)iH coI AI~9

UNCLASI Fl Er r -- _

-~~~C C4~4~. -

UNCLASSIFIED Page D-

A.PP1IT D

COST ANALYSIS

1. ABLA7IVI LINZRS-PRELDIUN*!Y DESIGNS

(U) Three ablative thrust chamber liners were designed based on theablative performance data obtained in the TRW System IRAD sub-scalematerial evaluation program. These designs yere based on the four beetperforming materials:

1CX2600 - Silica - phenolic

DC-93-104 - Filled Silicone Rubber

Haveg 41 - Asbestos - phenolic

CK 223-50 - Silica - epoxy novalac

The tLree prellminazy designs are summrizsd a follows:

1.1 COIICURATION NO. 1 (X404361)

(U) This configuration yes designed to provide a baseline ablative lineragainst vhich the others could be compared. The throat section vwa )X-2600,tape-wrapped at 600 to centerline, while the exit cone yes HX-2600 in arosette pattern "lay-up." The chamber-dome section we a hand "lay-up" ofNK-2600 orange-peel sections parallel to the chamber axis in the cyliedri-cal section.

1.2 rrZ TUI&r.j NO. 2 (Q404362)

(U) This configuration was designed as an all DC-93-104 silicone rubberliner, cast-in-place and cured at room temperature and umbient pressure.

1.3 0ONIIGURATIOM NO. 3 (X404363)(U) The third liner vas designed as a coupsiLte structure employing a taps-

wrapped MX-2600 silica-phenolic throat insert, identical to that of configu-ration aumer 1, with either Haveg 41 or GE-223-50 in both chamber sectionand exit cone-section.

(U) Request for quotations (IIQ's) were solicited from the fabricatorslisted in Table I. The work statement in the RIQ's provided for alternatifabricatiom teehntque* or substltution of co-..parable mster-!ls .-A.lch ..- o--.1Aresult in loaer "in-place" costs but would not reduce the quality of theablative Linar. A total of 38 quotations were received on the threeconfigurations. These included quotes for alternate fabrication techniques,materials and various options. A majority of the alternate fabricationtechniques proposed reflected the fabricators facility capability and didnot result in significant coat savings.

UNCLASSIFIED

UNCLASSIFIEDTable I. Fabricators Solicited for Ablative Liner-Fabrication (U)

Composite Technology Inc. Van Nuys, California

Edler Industries Inc. Newport Beach, California

EWCO Plastics Palmdale, California

HITCO Gardena. California

'aveg Industries. Inc. Santa Fe Spring., California

General Electric Company Philadelphia, Pennsylvania

Hadbar Division of Purolator Inc. Alhambra, California

Thermal Systems Inc. St. Louis, Missouri

San Rafael Plastics Company San Rafael. California

Insulation Systems Inc. Santa Ana. alifornia

McDonnell Douglas Corporation Santa Movica, California

(U) Based on the fixed-price quotes obtained on the preliminary designs andtest results obtained in the AFRPL "in-house" maLerials evaluation program(Minimum Cost Design (MCD) Material Screening Project 3058031•RD) additionalquotations were solicited on the frlloving materials.

HcDormall-DougLas - Segmented Silica-Phenolic

J-H Thermal Mix 770 - Silica-Aimbestos-Phenolic

Plastonium 8 - Gypsum - Phenolic Hicroballoons

GE-250-50 - Silica-Epoxy Novalac

GE-227-50 - Silica-Spoxy Novalac

GE-PD-715H - Asbeutos-Epoxy Novelac

MXA-150 - Asbestos-Phenolic

DPS-161 - Silica-Phenolic

(U) A sumary of the cost data obtained for all the candidate materials ispresented in Table II. The costs are the result of a number of fabricators'bids on comon items vith the selected cost being based on the fabricator'sbid and reputation for delivery of a quality part. All costs were normal-

ized to the NXA-15OtM-2600/IXA-150 (alternate configuration three) value,since this it the most cost effective ablative chamber liner. The range ofthe quoted prices for the MX-2600 liner (configuration one) was + 50 percent

of the median quoted PrICS (fo,- qotations. The range of the quoted pricefor the DC-93-104 leer was l i percent of the median quoted price (fourquotations). The ranking of the 12 candidate ablative liner assemblies isthe same for costs with an without tooling except for items 7 and 8. Item 7requires high pressure molding tooling while item 8 required a minimum ofaccountable tooling. The three most cost effective ablative chamber linerassemblies azi item 6, 1, and 3, in that order.

(U) Item 6 Is the composite IMXA-150/IX-2600 (alternate configuration three)ablative liner. The main cost savin for this liner is in the low material

UNCL FE•" " I FE

UNCLASSIFIED(U) costa ( $2.00/1ba) of the =-150 molding compound, plus its relative easeof fabrication. The seecA most coat effective chamber liner is ite- 1.This is the all HX-2600 comfiguration one ablative liner. The materialcost (-$6.00/lbm) is higUr than that of MIA-150; however, less NX-2600material is required and It was the moat popular material evaluated. Thecontingency factor for fabricating a quality part of !X-2600 is probablyless than for the remainder of materials. The third most cost effectivechamber in Item 3 which is the all 1FC-93-104 configuration two ablativeliner. The raw material cost is approximstcly 69.00 per pound and will bereduced as the material becamue amra popular throughout industry. It showsgood potential of being reduiced to as low as $5.00 per pound. The costsavings for this liner Is In the low fabrication costs. It does not requireeither high temperature or pressure for curing, nor does it require second-ary bonding to the pressure vessel. It is an elastomaric material with asufficiently high elongation rate to withstand dimensional changes in thepressure shell.

Table IT. Subscale Chabber Liner Materials Evaluation Program CoatEffectiveauess anking of all Candidate Liners (U)

NormalizedUnit Costs

for Quantitiesof Ten

Item Chamber Ablative Material with Without-119 S2E11"rallog Chaber Thr1 W-LtCon 121nz I~121ju

1 1 N1-2600 MI-.2600 M11-2600 1.61 1.54

2 1 He=D VC=D KcDD 3.59 3.37

3 2 DO-93-lG4 DC-93-104 DC- 93-104 1.6 1.62

4 3 U&VEG-41 31-2600 HAVEG-4l 1.97* 1.91*

5 3 GIC-223-50 MX-2600 GE-223-50 2.39 2.12

6 3 IHA-150 1H-2600 HIA-150 1.10 1.00

7 3 TI-770 HX-2600 TM-770 1.92 1.7 1

a 3 PLST. 8 MX-2600 PLST. 8 1.82 1.74

9 3 GE-250-50 1M-2600 CZ-250-50 2.36 2.12

10 3 Gr-227-50 MX-2600 G1-227-50 2.12 1.88

11 3 GR-iPD71511 KI-2600 GE-PD715H 2.57 2.32

12 3 DPS-160 MX-2600 DP5-160 No data No data

* Estimate based on first unit costs

2. ABUTV•E LIMB - FAIRICATED DESIGNS

(U) Based on the results of the TRW Systems IRD Materials EvaluationProgram, the AnnL "in-house" materials program, and the Task I test re-sults the preliminary designs vere finalsied a" follows:

UNCLASSIFIED

" "1 • I " •• M N • l • U " • r• 'lr•-, I i i ! um".

UNCLASSIFIEDTable III. Task II Hardware Acquisition Costs* (U)

Number of Units(1) (3) (10)

X404361 Ablative Liner 14,273 14,053 13,832(OIA-150/MX-2600/HX-2600)

X404362 Ablative Liner 14,946 14,305 13,282(DC-93-104)

X404363 Ablative Liner 15,575 15,119 14,215(DP5-161/MX-z2600/MXA-150)

* The hardware acquisition costs do not include tooling costs.

Table IV. Task II Ablative Liner Tooling Costs (U)

X404361 Ablative Liner $9107(HYA-150/MX-2600/MX-2600)

1404362 Ablative Liner None(DC-93-1o&)

X404363 Ablative Liner $4485(DPS-161/MX-2600/IMA-150)

Table V. Cost-Effectiveness Ranking of Three Ablative Liners (U)

Ablative Material Normalized Unit Costs (10)Cofgrto Chamber Thlot Exit-Con Without Tooling With Toolinst

1 MXA-150a MI-2600 MX-2600 1.04 1.73

2 DC-93-104 DC-93-104 DC-93-104 1.00 1.00

3 DP5-161 MX-2600 MXA-150 1.07 1.41

Nole: a. Layed-up broadgoods

b. Molded sewgmnts

UNCLASSIFIED

UNCLASSIFIEDTable V1. Volume and Weights of Fabricated Liuers (U)

Liner 2 3Configurnr-ion Volumein- Surface.ft- Density.lb/in Weightlbs

X404361-i

ChgMberHIA-150 10183 36.8 .0542 551.92

MX-2600 2928 12.4 .0616 180.50

Exit ConeMX-2600 4650 41.0 .0600 279.00

Total 1011.42

X404362-I

ChamberDC-93-104 4392 37.9 .0524 230.14

Throat , .DC-93-104 2928 .124 .0524 153.41

x:xit ConeDC-93-104 4650 41.0 .0524 243.66

Total 627.21

X404363-1

CheaberDPS-161 10183 36.8 .0499 508.13

ThroatMX-2600 2928 12.4 .0627 183.50

fait ConeNXA-150 7800 41.0 .0560 436.80

Total 1128.43

3.1 MATERIAL COSTS

(U) The coat of the base material in the three ablative liners ranged from47-50 percent of the total cost to 71.5 percent of the total cost. The DC-93-104 silicone rubber chamber liner had the highest material , oat percent-age (71.5) of the three cham'ers which Is attributed to the limited pro-duction of the DC-93-104 material. The other three materials used in the

UNCLASSIFIED

UNCLASSIFIED(U) fabrication of the three chambers were (1) MX-2600 silica-phenolic (2) )XA-150 asbestos-phenolic and (3) DP5-161 ailica-phenolic. The HXA-150 asbestos-phenolic was used in both the broadgoods and molding compound forms. Thecost of the molding compound being approximately 60 perceat of the cost ofthe broadgoods. DP5-16l is an acid-catalyzed condensation product offormaldehyde (par&) and a phenol (resorcinal) with a silica fiber/silicaflour filler. As such it can be compared to the silica-phenolic broadgoodsin the same manner that the molding compound (MXA-150) is compared to theMIA-150 broadgoods. The DP5-151 with less filler than the MX-2600 costapproximately 50 percent of the coat of the MX-2600 broadgooda.

(U) Of the four materials used in the fabrication only the DC-93-104material offers the potential of any significant coat reduction when thematerial usate becomes sufficient to warrant higher production. The pro-jected price of DC-93-104 for lrge-scale production ("400,000 lbs/yr) is$5.00/lb. None cf the three oti.er materials offer cost reductions approach-ing the reduction projected for DC-93-104.

3.2 FABRICATION COSTS

(U) For similar fabrication methods, casting of the chamber sections, thelabor cost for both the DC-93-104 and DF5-161 &Le approximately the sam;the material cost per pound differs by a factor slightly greater than 3.However, this material cost difference is nearly made up by the differenceof 2.5 in the quantity of material required. Using projected prices theDC-93-104 chamber liner is more cost-effective than the DP5-161 chamberliner Lud approximately the same as the layed-up MXA-150 broadgoods.

(U) Evaluation of the sane material (MX - fabricated from two differentform for two different liner sections wc.i made. The amount of materialused and surface area covered Pare approximately the same. The exit coneusing the molding compound is more rostly than the layed-up broadgoods chan,-bar section even though the material cost per pound is only 60 percent thatof the broadgoods. The difference in cost comes from the time consumingmethod of molding segments in addition to the necessity for secondarilybondiug the segments into the chamber.

(U) A paradox exists in that the molded, segmented MXA-150 exit-cone ismore cost effective than either the DC-93-104 or MX-2600 exit cone sectionsat the 250,000 lbf thrust level size. However, using the projected pricefor the DC-93-104, the most ccst-effective exit-cone is the silicone rubberliner. The D,-93-104 throat section is twice as cost-effective as the MX-2600 tape-wrapped throat section when the projected price of DC-93-104 isused to calculate the cost of the component.

At Ac=_I1.TTO0N COSTS-PRESSURBE SH_- LS

(U) Three pressure shells were fabricated to contain the three ablativeliners. The design was similar to the heat-sink combustion chamber(X403646) design which was used in the Task I program. Two major changeswere made in the design of the pressure shell; (1) the body was split intotwo sectio-a with flanges to allow insertion of the tape-wrapped MX-2600throat and (2) the thickness of the exit cone shell us. decreased to 0.25

UNCLASSIFIED

IIIIPI-6007rR8flUNCLASSIFIED Page D-8

M() Luchas from 0.50 inches. In addition to thase changer a stiffeningring was added to the exit cone.

(U) Three pressure shells of slightly varying configuration were fabricatedfrom USS T-i steel alloy. The slight variations were made to allov use ofthe shell as an autoclave during curing and to allow use of a thicker abla-tive section in the X404363 exit cone.

(U) Request for quotationi (RFQ's) were solicited from the fabricatorslisted in Table VII. The work statement required 100 percent X-ray of allbutt wglds and dye penetrant inspection of all fillet welds. All weldingand certification of welds was performief in accordance with the applicableASME boiler and pressure vessel codes. Responses to the RPQ were receivedfrom four fabricators. The range of the quoted prices for three pressureshells was t 35 percent of the median quoted price. Based on these quotationsGrano Steel Co. was selected to fabricate the pressure shells,

Table VII. Fabricators Solicited for Pressure Shell Fabrication (U)

J. C. Fabricators Gardena, California

Capital Westward, Inc. Paramount, California

L. W. LeFort Co. Anaheim, California

Los Angeles Boiler Works, Inc. Los Angeles, California

SWCCO, Inc. Los Angeles, California

American Bridge Div., USS Los Angeles, California

Grano Steel Corporation Los Angeles, California

Aircraft Engineering Corporation Paramount, California

(U) Acquisition costs for the three pressure shells are shown in TableVIII. These costz include TRW Systems material handling and G&A charges,but do not include fee. There were no tooling charges for any of thepressure shells. As noted in Table VIII, the cost of the -21 pressurishell is slightly greater than the cost of either the -1 or -2 pressureshell. This results from the 48, 21/32 inch diameter drilled holes put inthe stiffening ring to allow closure of the exit cone for curing theX404361 liner components. This $90 cost should be charged against thetooling costs for the X404361 ablative liner. The costs given in TableVIII represent a 90-95 percent learning curve on labor indicating that thefabrication technique is well defined and does in fact represent commercial("low-cost") fabrication practices. Based on a dry weight of Z840 lbs andthe unit price for quantities of 10 the cost per lb figure is approximately$2.00/lb.

Table VIII. Pressure Shell 4cquisition Coats (U)Number of Units

(1) (3) (10)

X4043,2-1/X404342-2 $6637 $6175 $5733

X404342-21 6727 624.7 5805

UNCLASSIFIED

I1 I I l

11199-6007-RO-0OUNCLASSIFIED Page

5. PR•JECTED COSTS - 3000K ABLATIVE CItMBER LINERS

(U) Cost estimates have been made for ablative chamber liners at the3000K thrust level. These estimates were made on the basis of the250,000 lbf thrust level designs and projected material costs. All ofthe methods used in fabricating the 250,000 lbf thrust chamber linersappear applicable for the larger diameter chambers. Tape-wrapping of verylarge ccmponents has been used in fabricating nozzles for the solid motorprograms. Casaing appears to be eatirely feasible for the large sizesalthough some specialized mixing equipment will be required since mixingappears to be the lnimting factor in the use of castable materials. Thelay-up of broadgoods and cure insitu at low pressure using the pressureshell as the autoclave also appears to be practical.

(U) The cost estimates for the 3000K liners are based on ablative thi k-nesses used in the 250,000 lbf long duration liners and as such representa somewhat conservative design for the 120 second design firing duration.The following table gives weights and cost:ý for two ablative lined chawbers.One chamber liner is cast from DC-93-104 filled silicone rubber while thesecond is fabricated from NX-2600. The throat is assumed to be tape wrap-ped at a specific orientation to the centerline, while the chamber and exitcone are layed-up broadgoods material.

Table IX. Estimated Wuights and Coats for 3000K Ablative Liners

DC-93-104 MX-2600

Dome/Chamber, lbs 2045 2440

Throat, lbs 1695 2020

Exit Cone, lbs 4610 5500

Total Weight 8350 9960

Material Cost $70,017 $85,305

Delivered Cost $146,000 $171,500

The costs quoted above do not include any tooling or special handlingequipment for these very large chambers.

UNCLASSIFIED

UNCLASSIFIED ,..SO °-,APP"IDI IX

DATA cDUCTION PRC:c•ZDUES

1.. GENER,•

(U) The test results presented hbrain are derived from computer printoutdata furnished TRW Systems by the AWL. General perform&=e data list-Ings for firings 103 through 111 and 117 were reduced using standard datareduction procedures. Detalds of the computational procodires and of theapplied corrections are given in the following sections.

2. DELI-VERD SPECIFIC INFULSE AND THRUST COFFICIUNT

(U) The measured specifict impulas is defLnsd by the followinS equation

I (meas) =- Ne (1)ep W

where:

F (ms) - +l ÷ FlB/2, output of dual bridge load cell, lbf

it - total propellant flow rate, lb M/sc

(U) The thrust coefficient is a measure of the nozzle performance and isgiven by

C7 (mews) - F(was)o (2)

vhere:

P o a throat etagnation pressure, psia

At - masured geometric throat area, In 2

(U) Tae throat stagnation pressure •as obtained by two methods. The first

method corrected the average of the two injector end pressure measurementsto stagnation pressure at the throat using Figure 1-1 (Rsf. 1). This methodasumes@ that the gas velocity across the injector end preseute taps is zeroend, zh.rofoGs, w-c.. r..ding h-nd and stagnatien• pressure. The correctionassumes no combustion in the nozzle and isantropic flow in the nousale.

(U) The second method used the masured static pressure at the start ofthe nozzle convergeut section which is cnnverted to stagnstiou pressure atthe throat by assuming no combustion iu the nozzle and i-entropic flow inthe nozzle. The use of the second method results in higher nozzle stagna-tion pressures than would be computed by the first method. This result Is

UNCLASSIFIED

-- z ,e- •L -• .. •: ... '-:•: ,mF, -- I

UNCLASSIFIED Page ,-2

7- - -. -- I - - ' : "1-: C

i-Li i# 4

-I't •. .. .. ' I. . -p . • , .; . -,1 '- . En

l~ l i f i .! • , -I i-i - - h' : I - J • : -.4 .: -

f- .. ..... : - I,,

" i: : . ... . - i

diai

F T

• L ... ..4; . T I 't...$. L:.J . N.

T T An

o " 4 4 N

UNCLASSIFIED

-_.. .

UNCLASSIFIED(U) due to a measured pressure loss between injector and nassle entrance

vhich is lower than that predicted using Figure E-l.

2.1 SPECIFIC IhPIULSE AND NOZZLE EFFICIENCY

(U) The theoretical I., and C* (characteristic velocity) data for thepressure range of 200 psia to 300 psia and covering mixture ration (0/F)of 1.80 to 3.20 were curve-fit aad used in equation form to compute thetheoretical Isp and C*. The equations for the theoretical values are givenby

Iap (pa = 13"2,cn " 4.0) w 120.6491 + 379.7692/ - 124.4563p

+ 0.463773 P - 0.00061575 P 2 + 0.0017556 P -u (3)0 0 0

and C* (tneo) - 1496.4296 + 5838.8849ru - 20 7 2 .959u - U.21772 Po

- 0.0003875 P 2 + 0.379418 P (4)

where

U - c /4 /1

(U) The specific impulse efficiency is computed as follow

nn (necs)

"Iup *p On- 13.2,En - 4.0)

The hardware used in Task II had an expansion ratio (A /A ) of 4.0 and the

nominal ambient pressure at the High Thrust Facility (1-59) wan 13.2 psia.

(U) The nozzle efficiency was determined using the following equation

Fr (mess)n _ a •a (6)C " Po At CF (theo)

The theoretical Cy is computed using Equations (3) and (4) as follovs

I (pa w 13.2, c - 4.0) gC7 (thee) U C* (thea)

2.2 WIPRCTED NOZZLE KYFICIENCY

(U) The expected neC for the 4/1 expansion ratio, 15' half-angle conicalnozzle was 0.980. Theoretical calculations were made for the variouslosses (kinetic, divergence and friction) in the 4/1, 15' half-angle conicalnoeale. These computations indicata a nouala efficiency of 0.980 at a mix-ture ratio (0/F) of 2.60. Theoretical calculations for mixture ratios of0.5 to 10.0 show less than 0.3 percent variation in the theoretical thruatcoefficient. UNCLASSIFIED

, . __________I____I_

UNCLASSIFIED3. MEASURED LCACLCTERISTIC VELOCITY EIrICIDICT

(U) The characteristic velocity (C*) vua calculated as follows:

C* - A.t-- (8)W t

The characteristic velocity was computed for Po calculated from the averageof Pc-l/Pc-2 and also for Po determined from the Pc-5 (See Figure E-2)static pressitre measurement. On the ablative lined chamber firings onlythe injector end pressure measurements are available.

W C* (theo)

where the theoretical C* is determined from Equation (4).

(U) The flow rates used in both the measured I and Ce equations wereobtained from single turbine flovaeters using water flow calibrations.Propellant temperatures were measured in tas propellant tanks, .•aies Andat the injector. D.nsities for both propellants were computed from AMZPLderived equations given as Equations (10) and (11).

-f " 51.777139 - 0.0350405 (Tf) (10)

-0 " 95.8447 - 0.078033 (T 0 ) (11)

4. IMJECTION PARAMETERS

(U) The injector and system flow conductances were determined using thefollowing general equation

(12)

where

volumetric flow rate (general data ',sting). ft3 /sc

P - propellant density, lbm/ft 3

&F - measured pressure loss, psi

(U) The fuel Injection pressure lose was measured between the entrance tothe orifice (PIF-1) and the injector and chasiber pressure (PC-I/PC-2)average. The oxidizer injection pressure loss was measured between P10-1sad the injector ved chaaber pressure (PC-I/PC-2) average. The fuel andoxidizer orifica discharge coefficients are determined as follimmi

UNCLASSIFIED

I I II i I I I I I I I I I I I ... '.........

UNCLASSIFIEDPage 9-5

C 1.495 [ a p_ (13)fd A If

Cd -1.495Cd0 A 0

5. T M EROSION WATE

(U) It in desired to obtain an estimate of the rare of change of the -alSr as a function of time, due to ablation and erosion, thermalexpansion, and other thermal mechanisms. The mean throat radius is definedas r in the following equations

r (vac) P0 A t CI (vac) (15)

2

At = 2 (16)

where F (va)" F measured + Pamb A. and A. is the nozzle area.

(U) It is asdumed that F (vac) and r are both time varying functions andthat CF (vac) is a time varying function which varies only a umall amountfrom sow known reference point, so that its effects may be linearized.The approach io to compute r - drjdt as a function of the nozzle stagnationpressure, the vacuum thrust, and their derivatives.

(U) The differentiation of (15) -nd the back-substitution of At computedfrom (16), given

.ý P

e F _ r c (17)

Equation (17) may then be solved for

It is assumed that Cp (vac) is explicitly dependent on expansion ratioand aoszle stagnation pressure; then CF is given by

C1 (c' Po) " "• • +-" 1 t• (19)C C o ap o 0 c n n(9

U L In

UNCLASSI FIED ,

-- '.. "-.- -- •:• ... L I'•7-....."r • " - ' • - - ' ' 'i " • • • • i -- ' •. .. ..

UNCLASSIFIEDand i is given by

lI C i. . 1 (20)

(U) Equation (20) is the central equation of the computation. In order

to compute and integrate f, we need P,(t), k(t), t(t). F(t), as vell as the

linearized coefficients aCp/aPo and 3CFp/cn, and the initial throat radius,

r(o).

(U) eGnerally the first derivative of tha C (vac) Ls zero since neitherthe nozzle stagnation pressure, P , or the nSzzle expansion ratio, cvary significantly from theiuitialconditions.

Reference 1. Alutman, 0., et l. Liquid Propellant lockets, PrincetonUniversity Press, 1960.

UNCLASSIFIED

-- - -.- a

~ Page E-7

zoo

t.

n~o

PIF-2

021P

(so

(U i gutE2 nttetto oain UUNLAMI"U

ocuty laelfcaton DOCUMENT CONTROL DATA.- R & DrSee. ev to e... 116-01". .1 fill.. hdy .1 .60t..cI *..d Irnd.-,d -Ie~j ho.,. mr*hi .ne.,. the~ I.-ml 0. 0-1~ P.e . r0e.AM.di

t. OU'4IA:IS(*IN ACTIVIYV rcorpwatomhm aANRPRTN...)CASIIATO

* Injectur/Chamber Scaling Feasibility Program,Ablative Chamber Design and Long Duration TestingVolume II

4. bg§CA1PVKY NOTES (nyp. 1vp.en .'I.. N))

Final Revort,_CoverinR Period 11 Decem-ber 1968 to 5 February 1970S.AUTNORISII (P.... ný. nowl. 1"11861. If.i)nein)

Voorhees, G. A. , Jr.Morton, B. G.

* "GROOT OATE M3. TOTAL NOý OF~ PAGKIS 76. NO or mpEs

July 1970 165 0&G. CON TNAC T OR GRAN T NO. $4. ONSGINA ION'S RKPORT NUMMIL (S)

F04611-6S-C-0085 Vol LI - TRW 11199-6007-RB-00&.PQOJECT 04O

C.~~Sb OTirma REPORT NOISI (ARY athwr n.SdP lwrvierm OW6000efle 00'""t)

A AFRPL-TR-70-86-Vol 11

Air Foc Sytm Cmad.UA

The results of the Task UI phase of an injector/ chamber scaling feasibilityprogram are presented. During the fourteen month program covering theperiod from 11 December 1968 to 5 February 1970 three ablative thrustchambers were designed, fabricated and test fired. Low-cost liner materialswere used in three chamber designs; the material selection was based uponsubscaic test data generated by both AFRPL and TRW Systems. Low-costfabrication techniques were employed throughout, ablative components werefabricated by tape-wrapping, hand lay-up, high pressure molding and casting.Several fabrication problems with the low-cost materials were delineated.Four ablative materials were evaluated in the test program. Two of thematerials evaluated (MX-2600 silica-phenolic and Dow-Corning 93-104 filled

* silicone rubber) had acceptable performance for use in low-cost engines ofL this type and for use in multimillion pound thrust booster engines..

DD FO0R1Mws1473 - siie

Secmnty CIA. Si1I'.t-.'

PRIMP. - ---

WORS LINK a LINK a LINK

MOL. I OT MOLL WT 401 WT

Low-Cost Injector

Performance

Stability

NzO4 /UDMH

Scaling

250,000 lbf

Low-Cost Thrust Chamber

Ablative Liners

Silica-Phenolic Ablatives

Silicone Rubber AbLators

I