SPACE STORABLE ROCKET TECHNOLOGY PROGRAM FINAL … · nasa contractor report 189131 space storable rocket technology program ssrt final report - basic program may 12, 12 -accesion

AD-A251 949

NASA CONTRACTOR REPORT 189131

SPACE STORABLE

ROCKET TECHNOLOGY PROGRAM

SSRT FINAL EPORTpTIC_FINAL REPORT - BASIC PROGRAM ELFCTE

JUN05 1992

MAY 12, 1992 A1

Prepared by:

M.L. Chazen, T. Mueller, A.R. Casillas, D. HuangTRW Applied Technology DivisionRedondo Beach, California 90278

Prepared for:NASA-LeRCCleveland, Ohio 44135Contract NAS 3-26246

t-- do7u0em ha been oPprovedfcx publi rele e and sale; its

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NASANational Aeronautics and 92-14861Space Administration 92I 148l11! ! l 11111 Jlll! l

92 6i ' ' "? I,


SPACE STORABLE


SSRT

FINAL REPORT - BASIC PROGRAMMAY 12, 12 -Accesion For

NTIS CRA&I

Prepared by: DTIC TAU 1]Unanfou ccd0M.L. Chazen, T. Mueller, A.R. Casillas, D. Huang Justification

TRW Applied Technology Division --- :1Redondo Beach, California 90278 By ............

Distribution IAvailabiltty Codes

Prepared for: I ",%va(aVd ' or

NASA-LeRCCleveland, Ohio 44135 AContract NAS 3-26246

92-14778

National Aeronautics andSpac Administration


SPACE STORABLE


SSRT

FINAL REPORT - BASIC PROGRAM

MAY 1992

Prepared for:NASA-LeRC

Cleveland, Ohio 44135

Contract NAS 3-26246

Prepared by:M.L. Chazen, T. Mueller, A.R. Casiflas, D. HuangTRW Applied Technology DivisionRedondo Beach, California 90278

Approval:

Melvin Chazen/Program Manager

* Albert Solbes, ManagerCombustion and Energy Technology Department

i

TABLE OF CONTENTS

Pacre

1.0 Summaryr............................................. 1

2.0 Introduction........................................ 2

3.0 Applications Evaluation..............................5

3.1 M'issions ............................. 53 .2 System Analyses.................................123.3 Fuels Evaluation................................203.4 System Requirements.............................273.5 Applications Evaluation Conclusions............27

4.0 Analyses........................................... 33

4.1 Performance Analyses............................334.2 Thermal Analyses.............................. 37

5.0 Exploratory Tests.................................. 56

5 .1 Design Approach............................... 565.2 Engine Design Point.............................565.3 Design Description and Fabrication.............565.4 Test Summary.................................. 575.5 Test Conclusions.............................. 85

6.0 Test Plans......................................... 87

6.1 Option 1...................................... 876.2 Option 2...................................... 916.3 Option 3...................................... 91

7.0 Conclusions........................................ 90

8.0 Recommendations.................................... 91

REPORT DOCUMENTATION PAGE.................................92

LIST OF TABLES

Table No. Paae

3-1 Mission Planning ............................ 83-2 OMV Key Mission Requirements .............. 103-3 CRAF Mission AV Requirements .............. 113-4 Fuels Selection for System Studies ........ 153-5 Engine Performance .......................... 163-6 Weight into GEO .......................... . 193-7 Summary of OMV Type System Capabilities 213-8 CRAF Mission System Capabilities .......... 223-9 Mission/System Capability ................... 233-10 Exhaust Product Constituents .............. 253-11 Fuels Evaluation ............................ 263-12 Propulsion System Requirements ............ 283-13 Engine Requirements ......................... 314-1 Summary of Injector Performance Analyses .. 364-2 Dome Cooling Concepts ....................... 545-1 SSRT Instrumentation List ................... 635-2 -8 Fuel Element Test Summary .............. 645-3 -7 Fuel Element Test Summary .............. 675-4 -9, -10, -11 Fuel Element Test Summary .... 745-5 -11 Hybrid Element Test Summary ........... 81

iii

LIST OF FIGURES

Figure No. Page

2-1 Apogee Engine Applications ................ 33-1 DRM-6/STS Altitude Profile ................ 63-2 CRAF Mission Benefits Using Advanced

Propulsion .................................... 73-3 Representative Integral Dual Mode

Propulsion System for GEO Satellites ...... 133-4 Hydrazine RCS System ........................ 143-5 Weight of Payload into GEO ................ 184-1 Effect of Combustion Efficiency on

Specific Impulse ............................. 354-2 SSRT Engine Injector Housing Thermocouple

Locations .................................. 384-3 SSRT Engine Dome/Neck SINDA Model ......... 394-4 Measured vs Calculated Dome Temperatures -

Test HA2A-4000 ............................ 404-5 Measured vs Calculated Neck Temperatures -

Test HA2A-4000 ............................ 414-6 Run 4000 Boundary Temperatures ............ 424-7 Heat Flows - Run HA2A-4000 ................ 434-8 Measured vs Calculated Dome Temperatures -


Test HA2A-3999 ............................ 454-10 Run 3999 Boundary Temperatures ............ 464-11 Heat Flows - Run HA2A-3999 ................ 474-12 Measured vs Calculated Dome Temperatures -


Test HA2A-4061 ............................ 494-14 Run 4061 Boundary Temperatures ............ 504-15 Heat Flows - Run HA2A-4061 ................ 514-16 Test Data Compared to Thermal Model Result. 524-17 Predicted Steady-State Columbium Chamber

Wall Temperature .......................... 535-1 L02/Hydrazine Engine with Copper Chamber .. 585-2 Engine Hardware ........................... 595-3 Test Facility Schematic ................... 615-4 -7 Elememt C* verses Total Flow Rate ...... 685-5 -7 Element C* verses Mixture Ratio ........ 695-6 -7 Element Fuel Gap Performance Trend ..... 705-7 -7 C* verses Oxidizer Gap ................. 715-8 Wall Zone Gas Temperature verses Mixture

Ratio ............ .... ................... 725-9 C* verses Fuel Gap for 200 lbf Elements ... 755-10 C* verses Oxidizer Gap for 200 lbf

Elements .. .................................. 775-11 -11 Element C* verses Mixture Ratio ....... 785-12 C* Performance verses Momemtum Ratio ...... 79

iv

LIST OF FIGURES (continued)

Figure No. Paae

5-13 Hybrid Injector Compared to Basic -11Injector ..................................... 82

5-14 Wall Zone Gas Temperature versus MomentumRatio ........................................ 83

5-15 Wall Zone Gas Temperature versesMomentum Ratio ............................ 84

6-1 SSRT Program Logic (Option 1) ............. 886-2 SSRT Program Logic (Option 2) ............. 89

V

ABSTRACT

The Space Storable Rocket Technology Program (SSRT) wasconducted for NASA-LeRC by TRW to establish a technology basefor a new class of high performance and long-lifebipropellant engines using space storable propellants. Theresults of the initial phase of this systematic multi-year program are described. Task 1 evaluated severalcharacteristics for a number of fuels to determine the bestspace storable fuel for use with L02 . The results of thistask indicated that LO2-N2H4 is the best propellantcombination and provides the maximum mission/systemcapability-maximum payload into GEO of satellites. Task 2,Preliminary Design, developed two models-performance andthermal. The performance model indicated the performancegoal of specific impulse > 340 seconds (E = 204) could beachieved. The thermal model was developed and anchored tohot fire test data. Task 3, Exploratory Test, consisted ofdesign, fabrication and testing of a 200 lbf thrust testengine operating at a chamber pressure of 200 psia usingL02-N2H4. A total of 76 hot fire tests were conducteddemonstrating performance > 340 seconds (E = 204) which is a25 second specific impulse improvement over the existinghighest performance flight apogee type engines.

vi

1.0 SUMMARY

The Space Storable Rocket Technology (SSRT) Basic Program wasinitiated in mid February 1991 and completed on schedule inmid October 1991. The program was very successful inachieving its overall objectives.

The Applications Evaluation task (Task 1) evaluated severalcharacteristics for a number of fuels to determine the bestspace storable fuel for use with L02 oxidizer. Theseevaluation factors included mission usage, propulsion systemconfiguration and space storable fuel properties to achievepayload maximization. The evaluation task also establishedpreliminary system and engine requirements. The maximummission potential usage for the Space Storable engine isplacement into GEO of NASA, military and commercialcommunication, surveillance, tracking, earth observation andmeteorological satellites. The system analyses and fuelsevaluation indicated that LO2-N2H4 is the best propellantcombination and provides the maximum mission/systemcapability-maximum payload into GEO. The nominal enginedesign based on preliminary system/engine requirements ispresented as follows:

Propellants L02-N2H4Thrust (FO) 200 lbfChamber Pressure (Pc) 200 psiaSpecific Impulse (Isp ,) 340 lbf-sec/lbm

The Preliminary Design task (Task 2) developed a performancemodel which indicated the performance goal could be achieved.A thermal model was developed and anchored to the test dataobtained in the Exploratory Test task so it would be a usefultool. The thermal model indicated that additional injectordome cooling is required to operate for long duration at highengine performance. Therefore, overall engine dome conceptshave been identified which will be evaluated in Option 1.

The Exploratory Test task (Task 3) consisted of design,manufacturing, testing and analysis of the test data. Twoseries of tests were conducted evaluating six configurationsindicating high performance could be attained. A total of 76tests was conducted. Performance of 95% C* which projects to> 340 lbf-sec/lbm vacuum specific impulse (E = 204) wasachieved with thermal characteristics indicating thatoperation with a columbium thrust chamber is feasible. Theuse of a rhenium thrust chamber is another alternative whichwould allow performance approaching 350 lbf-sec/lbm.

2.0 INTRODUCTION

The increasingly demanding spacecraft missions and theirassociated requirements for increased payloads over the last30 years have been successfully achieved by the steadilyimproving capabilities of spacecraft propulsion systems.These systems have used earth storable propellants,principally either hydrazine as a monopropellant or nitrogentetroxide/amine fuels as bipropellant. The technology levelof these propellants and their systems have been repeatedlyimproved as mission demands have grown.

Space storable propellant usage offers the advantage of usinghigher performance propellants to achieve increased payloadweight into orbit. The results of TRW studies are in concertwith NASA-LeRC's conclusion that liquid oxygen (LO2 ) is thebest space storable oxidizer. The space storable fuels aredefined as those fuels that can be passively stored, withinmission constraints, without active cooling or refrigeration.Figure 2-1 shows the overall propulsion scheme of propellantdevelopment (Isp levels with respect to time) and where spacestorable fuels fit into this overall scheme which indicatesthe need for space storable rocket development. Spacestorable propellants provide the link between upgraded earthstorable and an integrated H/O system. Among the categoriesevaluated were alcohols, amines, cryogens and hydrocarbons.In order to adequately evaluate the propellants, selectioncriteria were established and system analyses conducted basedon representative missions and engine performance. Theresults of this Task 1 study provided the following:

* Evaluation of mission usage

e Propulsion systems and fuels evaluation to achievepayload maximization

* Evaluation and selection o. fuels

e Preliminary system and engine requirements

The space storable rocket technology (SSRT) program consistsof four phases (Basic program + three options). The firstphase (Basic Program) consisted of three tasks:

o Applications Evaluation as discussed above

e Preliminary Design

- Performance analyses- Thermal analyses- Overall engine concepts

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- Initial tests with L02- Modify hardware based on initial test results- Retest with modified hardware

This report will discuss the results of the three tasks ofthe Basic program phase.

4

3.0 APPLICATIONS EVALUATION

Space storable propellants offer the advantage of providinghigher performance to achieve greater payload weight intoorbit. The applications evaluation studied the followingareas:

* Mission evaluation usage of advanced propulsiontechnology

* Propulsion systems and fuels evaluation to achieve

payload maximization

* Evaluation and selection of fuels

e Preliminary system and engine requirements

* Conclusions

3.1 Missions

Three representative missions were investigated to utilizeadvanced propulsion technology. These three types ofmissions are defined as follows:

o Perigee/apogee integral propulsion systems are used toplace satellites into geosynchronous earth orbit (GEO)utilizing expendable launch vehicles (i.e., Atlas,Delta, etc.). These missions include NASA, militaryInaA,and commercial applications for communication,surveillance, tracking, earth observation andmeteorology. These missions constitute the greatestquantity and frequency of mission applications and areshown in Table 3-1 which average 30-40 yearly notincluding classified military missions.

e Low earth orbit is another mission application. Oneapplication uses the Orbit Maneuvering Vehicle (OMV)or another similar vehicle to go from shuttle cargobay to Space Station or from Space Station to therequired mission. Table 3-2 shows the typical OMVmissions and their requirements. Figure 3-1 shows therepresentative mission selected for the system studysince it utilizes the greatest .V requirement.

e The planetary application is another representativemission. The Comet Rendezvous Asteroid Flyby (CRAF)was selected as the typical planetary application.The mission is shown in Figure 3-2 and the 6Vrequirements are shown in Table 3-3.

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TABLE 3-1MISSION PLANNING

LAUNCH VEHICLE 1992 .1993 .1994 1951996 1997Arianespace Insat 2A Intelsat VII F4 Astra 1D2_______ ______________

Telecon 2 Fl Galaxy 7 Brazilsat B2 _______ _____________

Intelsat VI F4 Telecom 2 F2 ______

Intelsat VII F1 Astra 1 C _______

Satcom C3 Brazilsat B61TopexiPoseldon ESA Intraved

____________Space OBS _____________

Hlspasat I ________________________

Hispasat 2 ________________________ _____

___________Galazy 4_____ __ ____ _____

Long March Aussat 81 Aussat 82 ______________

Shuttle USML-01 SL-D2 IML-02 USML-02 SRL-03LAGEOS 11 SLS-02 SPTN-04 SL-D3 GEOSTAR-03EURECA-1 R TDRS-F ISEM -02 SPACEHAB-06 SPACEHAB-07ASP DEE CXM-02 SPTN-05 OAET-03DOD CTMV SFU-FtETR CXH-10 ATLAS-05TSS-01 SPACEHAB-02 CXM-03 USMP-04 SSBUV-A-04 ______

EURECA-11L OAET-01 CMSE-02 WSF-03 CONEIMAX-06 CAPL XTE SSFIMB-02 SSF/MB-04

ATLAS-01 SRADfrP1TS GEOSTAR-01 GEOSTAR-02 SSFIMB-05 ____

SSBU-04TM WS03 FAN-0 1 ERT1 SPACEHAB-03OSEA2 SSM-03

ICANEX-02 ISEM-01 HPE ATLAS-04 WSF-04JDXS SPACEHAB-01 MICROWAVE-01 WiSP DCWS______JINTELSAT VI-R OREFUS-SPAS USMP-03 SSBUV-A-03ICVTE-01 GAS BRIDGE OAET-FLYER AAFE ______ _____

JASEM SHOOT SRL-02 OAET-02 ______ _____

SL-J CRISTA-SPAS SSF/MB-Ol _______

G3AS BRIDGE ______ATLAS-03 ____________ _____

____________ _________SSBUV-A-02 _____________

____________ _________WSF-02_____________________

____________ _________SPACEHAB-04 _______ _____________

____________ _________EURECA-2L ______ ______

__________ __________ _________CXM-04 ______

MEDIUM CLASS GEOTAIL (12) POLAR (D) RADARSAT (D) GPS III (D) LIFESAT 1 (D) ACE (D2)Atlas (A) WIND (12) NOAA-J (A) LAGEOS III (D) GPS; III (D) LIFESAT 2 (D) LIFESAT-3 (D2)Delta (D) GPS 11 (D) GPS; 11 (D) NOAA-K (17-i11) GPS; III (D) NOAA-L (T-l 1) NOAA-M (T-1 1)Titan 11 (T-1 1) GPS 11 (D) GPS 11 (D) GPS 11 (D) DMSP 5D2 (T-i 11) GPS III (D) GPS IlI (D)IPS 11(D___11______1_D _GS II() SII D

GPS 11 (D) GPS 11 (D) GPSII1 (12) GPS III1(D) GPS III (D)GPS 1(D) GPSI1(D) GPS (D) 11 IID) GPS III (D)

DMSP 5122 (r- 11) ______ DMSP 5D2 (T-i1)T GPS III (D) JGPS III (D)________DMSP5D2(r-11) jGPS III (D)

I 8

TABLE 3-iMISSION PLANNING CONTINUED

LAUNCH VEHICLE 1992 199W 1994 1995 1996 1997INTERMEDIATE GOES-I (A-i) Intelsat VII MSAT (10C) SOHO (11 AS) 080$ 5D33 (11) ATDRS-i (10)CLASS F2 (11 AS)_______

Titan III MO (T-i i) Intelsat VII Telstar 4 TDRS-G (10) UHF FO (A-i) GOES-L (A-1)Atlas 1111 AS F3 (11AS) F2 (11AS) ______

GOES-J (A-i) UHF-2 (A-i) SAX (A-i) GOES-K (A-i) UHF FO (A-i) SSFGALAXY-i R(A-i) Telstar 4 ORION-2 (iiA) MARS OBS UHF FO (A-i) DSCS 5133(11)

_______ ______F1 (11AS) ______

UHF-i (A-i) ORION-i (11A) 0805D3 (11) SSF UHF FO (A-i)DSCS 503 (A-il) DSCS 503 (1i) UHF FO (A-i) UHF FO (A-i)

DSCS 5D3 (11) UHF FO (A-i) UHF FO (A-1) ______

_____________ __________UHF FO (A-i) ______

LARGE CLASS MILITARY SURV MILITARY SURV MILITARY SURV MILITARY SURV MILITARY SURV MILITARY SURVTitan IV ________MILITARY SURV MILITARY COMM MILITARY COMM MILITARY SURV MILITARY COMM

__________ ________CASSINI CRAF CASSINI CRAF MILITARY COMM_________________ ________________ SF SURV

SECONDARY P/LS; PMG SEDS-2 HETE NRGL______Delta EUV SAC-B ______ COC______

___________SEDS-i _____ ______ ______ ______

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3.2 System Analyses

System analyses were conducted for each of the three missionsof 3.1. The fuels considered and selected typical for thesemissions are shown in Table 3-4.

3.2.1 Engine Performance

The engine configuration utilized for this study consisted of100-400 lbf thrust radiation cooled engines operating at 100-400 psia chamber pressure. This selection was based on theuse of multiple liquid bipropellant engines to avoid singlepoint failures. Rhenium thrust chambers were used to achievethe high performance (4000°F wall temperatures). A highthrust version was investigated using 1000 lbf thrustregeneratively cooled engines operating at 400 psia chamberpressure. The engine performance used is presented in Table3-5 and was anchored to the TRW dual mode engine which isqualified and successfully flying.

3.2.2 System Configuration

The baseline system to be used for evaluation which ispresently flying on communication satellite applications toGEO is shown in Figure 3-3. This configuration was based onevaluation of the various system options and includes thefollowing:

e Pressurization - regulated pressure-fed system usingGHe at 7500 psia in spherical graphite/epoxyoverwrapped tank with aluminum liner.

* Propellant storage - two oxidizer and two fuel tankswhich are cylindrical with elliptical heads andare graphite/epoxy overwrapped with aluminumliner. MLI is used for the cryogenic tanks.

e Perigee/apogee engines - radiation cooled as discussedin 3.2.1. The low thrust version used a totalthrust of 400 lbf (1, 2 or 4 engines) while thehigh thrust version used a total thrust of2000 lbf (2 engines).

* Reaction control system - decomposed hydrazine systemper Figure 3-4. For L02-N2H4 system, thehydrazine is fed from main hydrazine tanks.

The ground rules used to analyze the various applicationsand fuels are:

o Residual - 1% of total propellant was consideredunusable for all applications including OMV andCRAF.

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" Boil-off - 0.4% of the cryogenic propellant wasconsidered boil-off except for high thrustcondition which used 0.25%. These losses wereused for all applications including OMV and CRAF.LH2 losses were established at 6.1% (based onAFRPL-TR-86-045).

" Startup/Shutdown - 0.5% of total propellant used forall applications. For LH2 , 2% was used.

* Thermodynamic vent system cooling - 2% total of LH2.

3.2.3 Results of System Analyses

System analyses were conducted to evaluate their weights toorbit for the three types of missions (GEO, OMV, CRAF)defined in 3.1.

3.2.3.1 Perigee/Apogee Applications

The investigation of mission applications indicates perigee/apogee applications have the greatest usage potential in theforeseeable future. Therefore, the system analyses for theseapplications are most important and the results have thegreatest impact. The analyses used the Delta 7925 (6000 ibminto LEO) and Atlas IIA (14,750 ibm into LEO) as typicallaunch vehicles. The results of the analyses are presentedin Table 3-6 which indicates LO2 -N2H4 is the best propellantcombination to achieve the greatest weight into GEO. Theseweights consider the major subsystems except for the RCSsystem and system components (regulators, valves, lines,heaters).

The hydrogen tank volumes (- 450% of amine tank volume) areso great that LH2 could only be integrated into the vehicleusing toroidal tanks necessitating aluminum toroidal tanksdue to unavailable overwrapping toroidal tank technology.This results in non-competitive weights into GEO.

The impact of the RCS subsystem was investigated. Figure 3-4shows the system used to assess RCS impact based on a&V = 1465 ft/sec for RCS and stationkeeping. Figure 3-5shows the RCS weight impact on payload into GEO.

3.2.3.2 OMV Type Applications

System analyses were conducted to determine the bestpropellant combination for a typical low earth orbitapplication and the OMV viewing mission (DRM-6) was selected.The four best fuels (amines and hydrocarbons) were selectedfor the analyses including N2H4 , MMH, CH4 and C3H8 . Thepropulsion system was similar to Figure 3-3 except four

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0 J ..jn eCD

4)~~~~~ 4) ' 0 ~ A 0

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41.

4m) -c %0! LA a%! O . .4 C) .-

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(A L&.0 M CD CO CD C) C) CD (fL)I

mr C 4 C~4 w C4 to

___ __ __ __ L- to

09

19

engines of 100 lbf thrust (operating at 100 psia chamberpressure) were utilized for this application. Two enginesoperate simultaneously for each maneuver and the other twoengines provide redundancy. The results of the systemanalyses are presented in Table 3-7 and indicate LO?-N2H4 isthe system of choice as it provides the lightest initialvehicle/system weight and highest bulk density and massfraction over the other candidates.

3.2.3.3 CRAF Application

System analyses were conducted to determine the bestpropellant combination for a typical planetary mission andthe CRAF mission was selected - current plans show the use ofa typical bipropellant system similar to our system studies.The four best fuels were evaluated using the same regulatedpressure-fed configuration of Figure 3-3 but using only one100 lbf thrust radiation cooled engine operating at 100 psiachamber pressure. The initial spacecraft weight is 11,305lbm. The results of the system analyses are presented inTable 3-8 and indicate L02-N2H4 is the best system as itprovides the maximum payload weight with the lightest systemand maximum bulk density and mass fraction over the othercandidates.

3.2.3.4 Summary/Conclusion

Based on the system analyses, the overall mission/systemcapability is summarized in Table 3-9. Using the Figure ofmerit defined and presented in Table 3-9, L02-N2H4 is thebest propellant combination.

3.3 Fuels Evaluation

The seven selected fuels were evaluated to select the bestengine and system to achieve the mission applications. Eightevaluation factors were considered in the evaluation.

e Mission/System Capability

The evaluation factor considered weight and volumeconsiderations of engine and system. The figure ofmerit and evaluation of Table 3-9 were the basis ofevaluation of this factor.

e Safety Considerations

The safety considerations included three factors--flammability, explosive potential and system safetywere the major considerations.

20

o

-J

= ca - C -CD -4

LI m~ Lfl CD ~.0 4J J Lfn eJ co C) %0

a% 00 mr. C. *n - . c%

4) 4-J f J N .

L W> d

> ~ ~ ~ ~ ~ 4 CM ~ N .

in44

I 4J 3

o 4-I.- = 44-)

ou CC% Lu0

.'-S.- (A CD

4J 4J A C +

X- 4- 1-CL 0 C) u 4-r= to 4Jh0

4A A, Cu (

O 4- t o 04 I b

4.. 01 0 - W

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co CA4-I- S. LIco

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f-) cn1 %0 m. Ch

e'kJ CD C4J 1-O r% 0 U

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" System Integration

System integration, considered two primaryconsiderations--RCS and fuel integration. The RCSmust be integrated to achieve minimum weight andcomplexity while achieving high reliability with thenecessary control. In all cases, hydrazine RCS wasused as the system due to its high reliability anddemonstrated flight data base. Fuel integrationconsidered insulation of cryogenic tanks and linesand design of regulators and valves.

" Plume Contamination

The exhaust products at the nozzle exit wereevaluated based on the engine operating conditions.The major toxic constituent was determined to be COalthough the amine fuels had traces of NO. Theexhaust products are summarized in Table 3-10.

* Logistics

The logistics considerations were based on the useof the fuel for flight at the launch facilities.This factor considered shipping, storage,availability of fuel, ground support at the launchfacilities and validated operating procedures at thelaunch facilities.

" Materials Compatibility

This factor considered seals and materials that arecompatible with the fuels.

• Cost of System

The cost assessment included development, recurringand life cycle cost of system/engine.

• System Risk

The risk assessment included development andrecurring cost and schedule risk.

The evaluation of the eight factors is presented in Table3-11. The results indicated that L02-N2H 4 is the bestpropellant combination of the eight evaluated. Therefore,TRW recommends the use of L02-N2H4 in the development of theSpace Storable Test Bed Rocket Engine.

24

tmK x

*\a 0\ ~

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4ACJ C.J

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0

-% csj i *9 45,421 04. cc

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4-J-

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4-1 4-1 to

o . *~ en .-

0.- c.. C-) C~j CNJ m~ -4-1 4~)-JI

to C .) ~ U .0 ~"a 'a 0 *; * 5

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to ' C* -4 - M; C- C4

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3.4 System Requirements

A preliminary set of system and engine requirements have beenestablished based on the system studies and TRW experiencewith systems and engines. The system analyses indicated thatthe payload to orbit optimized at 200 lbf thrust and 200 psiachamber pressure. Therefore, this was the recommended designpoint pending further testing. The preliminary systemrequirements are shown in Table 3-12. The engine preliminaryrequirements are shown in Table 3-13. These requirementswill be updated as test results necessitating change becomeavailable.

3.5 Applications Evaluation Conclusions

The maximum mission potential usage for the Space Storableengine is placement of satellites into GEO for NASA, militaryand commercial applications for communication, surveillance,tracking, earth observation and meteorology.

To achieve this mission potential, an evaluation of thevarious candidate fuels indicated that L02 -N2H4 is the bestpropellant combination and provides the maximum mission/system capability. The preliminary system and enginerequirements provided the basis for the preliminary designand indicated the nominal engine design as follows:

Propellants L02-N2HThrust (FO) 200 if

Chamber Pressure (Pc) 200 psiaSpecific Impulse (IspO) 340 lbf-sec/lbm

27

4) IA

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28

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Table 3-13. Engine Requirements

Thrust (F.)-lbf 200 ±10- (890 + 45 N)

Mixture ratio (O/F) 0.75 +0.03

Specific impulse (Isp.)-lbf-s/lbm k340 nominal (a 3334 N-sec/kg)3+0 +0 N/cm2)

Inlet pressure - psia 35010 (241 -14

Fuel inlet temperature - OF 70 +10 (excluding heat soakback) (21 + 6*C)

Oxidizer inlet temperature -F -285 (excluding heat soakback) (- 176*C)

System mixture ratio 0.75 +0.08 (includes pressure andtemperature variations)

Life - sec 10,000 (qual - 15,000)

Maximum continuous firing - sec 3000

Operation Steady state (performance at >30 s)

Operating voltage - Vdc 24-34

Engine length - inches 29 maximum (74 cm)

Engine diameter - inches 12 maximum (30.5 cm)

Heaters Required to prevent fuel fromfreezing

Valve seat leakage (scc/hr GHe) 5 per valve seat

Random vibration

Qual - 14.1 g-rms A/T - 10.0 g-rms20 Hz - 0.026 g2/Hz 20 Hz - 0.01 g2/Hz

20-50 Hz - +6 dB/oct 20-160 Hz - +3 dB/oct

50-800 Hz - 0.16 g2 /Hz 160-250 Hz - 0.08 g2 /Hz

800-2000 Hz - -6 dB/oct 250-2000 Hz - -3 dB/oct

2000 Hz - 0.026 g2/Hz 2000 Hz - 0.01 g2/Hz

Oxidizer-fuel inlet pressure Fuel = +5 psia of oxidizer pressure

variation

Alignment +0.5 degree

Engine weight - Ibm 11 maximum (5 kg)

Contamination control PR2-2-12Valve must have 25 micron inletfilters

31

Table 3-13. Engine Requirements (Continued)

Valve characteristics

Pull-in voltage (Vdc) 19 maximumDropout voltage (Vdc) k2Open response (ms) 30Close response (ms) 30Maximum pressure (psia) 400

Engine starts (cold) 25

Engine roughness +12%

Gas ingestion 2 in3 (33 cm )

Oxidizer depletion Must have capability

Heat shield Minimum impact on enginetemperatures

32

4.0 ANALYSES

The two major categories of analyses emphasized during theBasic program were performance and thermal. The performanceanalysis objectives were to establish a model to predictsensitivity to design variables and assess ability to meetperformance goals. The thermal analyses objectives were toestablish a model to assess thermal operating characteristicsof the injector aid thrust chamber.

4.1 Performance Analyses

4.1.1 Analysis of Injector

A model of the coaxial pintle injector was developed byDr. Richard Priem to calculate the performance based oncombustion characteristics using L02-N H4 . The primeconsideration was the model should predict sensitivity ofvarious combustion parameters to design variables. The modelfor the fuel centered injector incorporates the followingelements:

* Injection velocity - treat fluids as columns thatintersect each other. First spray is caused by slotsof fuel impinging with oxidizer. Second spray iscaused by fuel gap flow between slots impinging onoxidizer.

o Jet size and drop size - jet size of each stream iscalculated on the basis of a round jet having the samearea as the impinging streams. Drop size iscalculated using impinging jet correlation curve ofTR 67.

o Vaporization

- Prior to impingement of first spray

Assume a gas velocity of flow out of the domethrough the spray.Using assumed velocity calculate momentum balanceto determine radial gas velocity of this flow thatwould balance a decrease in liquid velocity of thefirst spray to the point where the radial gasvelocity equals the resultant spray velocity.ThLr calculate drag and deceleration of the sprayaloihg with the amount vaporized of the fuel andoxidizer as a function of radial position.

- Vaporization of second spray - determine amcuntvaporized before spray impinges on wall

33

- Vaporization in chamber

Assumes spray bounces off chamber wall withaverage angleBreak spray into five sections having varying massand bounce angleCalculate the amount vaporized in ten annularsections of the chamberWith the angle, calculate the length prior tomovement out of the annular sectionUse this length to determine effective lengthMass average all the different parts of the sprayI and sum each for the various annuli

• Mixing in the chamber - simulate mixing bytransferring 10% of each flow from adjacent annuliinto each other. This is done on a flux differencebasis and area of smaller annuli.

* Final performance - based on O/F in each annuli andmass flow, sum the mass averaged C* to obtain engineC* and resultant combustion efficiency.

The results of this model were used to predict the trends forcombustion efficiencies (C*) of the various elements. Themodel was established based on the results obtained on thefirst element tested (-3) with LO2-N 2H4. These results wereused to anchor the model. The model was then used onsubsequent elements to predict the performance. Table 4-1shows the results of the analyses and test results.Increasing the number of slots is the most effective way ofincreasing combustion efficiency (C*).

4.1.2 Nozzle Performance

A two dimensional kinetic analysis was conducted to assessthe thrust coefficient and potential vacuum specific impulseachievable for the L02 -N2H4 engine. The analysis was basedon a two zone model operating at mixture ratios (O/F) of0.875 in the core and 0.5 at the wall to produce an overallengine mixture ratio (O/F) of 0.8. The overall enginecharacteristics are as summarized follows:

I Thrust (F.) 200 lbfChamber Pressure (Pc) 200 psiaNozzle Expansion (E) 204Mixture Ratio (O/F) 0.8

The results indicated a vacuum thrust coefficient (CfO)including boundary layer losses of 1.89. Based on 94.6%combustion efficiency, the vacuum specific impulse (Isp,)would be 340 seconds. The effect of combustion efficiency onspecific impulse for the two zone TDK analysis is shown inI Figure 4-1.

34

W111-1 I-HT, I dill!All IIIHJ-111,1111 1 11111.44

1-R1+++-1 4 +-1,, 1 i 1 444 1 Figure 4-1.4+444+ m++ +tH,+ + fl t-*#4 ## tf# I t W t +1

Effect of Combustion Efficiency444 f 4 , 1 $10 1, TTon

Specific Impulse.+

44 +,-

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36

4.2 Thermal Analyses

Thermal analyses of the injector and thrust chamber which areshown in Figures 5-1 and 5-2 were conducted to assess areasrequiring modifications to the initial design. Thermalmodels were developed and anchored to test data prior toassessing design capabilities.

4.2.1 Injector Thermal Analyses

A SINDA model of the injector dome/neck region was developedto assess combustion gas heating loads from test data.Figure 4-2 shows the locations of the injector thermocouplesutilized in test and Figure 4-3 shows a sketch of the modelwith the thermocouple locations indicated.

The general approach used is presented as follows;

" Film coefficients for the liquid oxygen in the annulusand cone passages were calculated for the forcedconvection, nucleate boiling, transitional, and filmboiling regions using published empirical relations(e.g., Sieder & Tate, Rohsenow, Gambill, andRocketdyne cryogenic data).

" A heating load was applied from the combustion gasessuch that the resulting temperatures agreed withmeasured values.

Results for three cases - low, moderate, and highperformance - are presented in the following paragraphs.Higher performance was accompanied by higher heating loads asexpected.

o Low Performance. The correlation between measured andpredicted dome temperatures for test number, HA2A-4000(80% C*) is shown in Figure 4-4. The calculatedcurves were applied to the noted mode numbers of theSINDA model of Figure 4-3. Injector neck temperaturesare shown in Figure 4-5. The imposed gas temperaturefor all zones are shown in Figure 4-6. The higherinitial gas temperature (1650 F) resulted due to theN204-N2H4 ignition; it then decreased to 450°F at 6.5seconds when chamber pressure stabilized. The localfilm coefficients required for the outer zone (Zone 1in Figure 4-3) was significantly higher than for theother zones. However, Figure 4-7 which shows thetransient heat flows, indicated that the L02 heatabsorption requirement at steady-state was only 0.5Btu/sec.

o Moderate Performance. The correlation of dome andneck temperatures for test number HA4-3999 (87% C*) isshown in Figures 4-8 and 4-9, respectively. The

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54

(inside) and 2410°F (outside). Therefore, the columbiumthrust chamber is the primary approach. Rhenium (iridiumcoated internally) is the backup approach to the thrustchamber design which may allow even higher performance.

55

5.0 EXPLORATORY TESTS

5.1 Design Approach

The engine design approach was to maximize design andoperational flexibility to allow cost effective evaluation ofthe range of engine parameters. The injector was designed tooffer flexibility in test to evaluate the changes necessaryto achieve high performance. The goal was to maximize testinformation for minimum cost. The TRW coaxial injector wasideal for these evaluations as it allowed variations invelocity and geometry of the basic design to be readilytested and assessed.

The exploratory test engine utilized an injector whichallowed shimming of the oxidizer and fuel gaps to changevelocities and replaceable extensions to change fuelgeometry. The thrust chamber for this engine was a robustcopper heatsink thrust chamber using thermocoupleinstrumentation. The injector and thrust chamber were boltedtogether for ease of testing. Test stand valves were used atthis point in the program to eliminate the valve developmentprior to understanding the specific requirements andinterfaces. Pre and post test GN2 purges were used on allpropellants. Since the propellants were non-hypergolic, anigniter was required. The igniter used was N204 injectedthrough a port in the injector to ignite with the fuel priorto introduction of L02 . This concept was selected based onease of design and test.

5.2 Engine Design Point

The applications evaluation as discussed in 3.0 evaluated thevarious fuels and system requirements to maximize payloadinto orbit. The results indicated the system should bedesigned to the preliminary requirements of Table 3-12.Based on these preliminary requirements, the enginepreliminary requirements of Table 3-13 were developed andprovided the design point for the exploratory tests. Theserequirements also provided the design for the test bed engineas modified based on the exploratory test results.

5.3 Design Description and Fabrication

The TRW coaxial injector for the SSRT program was based onthe DM-LAE qualified and flying successfully on ANIKsatellites (E-1 and E-2). These engines produce an averagespecific impulse of 314.5 lbf-sec/lbm (E = 204) and havedemonstrated almost 25,000 seconds operating life duringqualification with N204-N2H4.

56

The SSRT injector consisted of the following elements:

e Body of columbium with aluminide coated face(oxidation protection)

* Sleeve of 15-5 pH incorporating thermal isolationof L02 and N2H4

o Pintle of 15-5 PHe Extensions of 15-5 PH incorporating various slotgeometries

e Igniter to inject N204 to react with N2H4 prior toL02 injection.

The injector in the copper heatsink thrust chamber is shownin Figure 5-1 and photographs of hardware are shown in Figure5-2. Injector configurations are changed by replacing sleeveextensions to assess variations in fuel slot geometry.Additionally velocity changes can be varied by independentlyshimming the oxidizer and fuel gaps. Six fuel geometrieswere evaluated using five different configuration sleeveextensions with the standard pintle. The highest performancesleeve with a pintle incorporating three doublets (designatedhybrid) which bleeds fuel into the center of the engine wastested to enhance performance. The slot configurationsvaried from 36-60 slots with slot widths of 8-16 thousands ofan inch and aspect ratios (slot depth/slot width) of0.67-4.8. These wide variations in fuel geometry along withvariations in fuel gaps and oxidizer gaps provided theability to test over a range of large variations to assessperformance characteristics. This flexibility provided amethod to obtain affordable test costs with major geometrychanges in the injector.

The thrust chamber used during this basic program was arobust heatsink copper chamber with type K thermocouplesbrazed into the wall at three axial locations and fourthermocouples at each station (900 apart). Thisinstrumentation allowed an assessment of the thermalconditions of the thrust chamber.

5.4 Test Summary

5.4.1 Test Plan

As part of the SSRT basic program, exploratory hot fire testswere defined to provide input to the engine design. Thesetests were performed using the TRW IR&D hardware that wastested in 1990.

The exploratory tests performed in the basic program werestructured to provide basic engineering information relatingto the performance and thermal aspects of the design. Someof the issues addressed were:

57

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" Engine combustion performance characteristics" Stable operation" Engine thermal characteristics* Injector characteristics" Comparison of L02/N2H4 to hypergolic earth

storable propellants* Ignition characteristics with N204e Hardware and system chill

Two test series were performed during the basic program. Thefirst series addressed the differences between the L02/N2H 4propellant combination and the hypergolic propellant enginesusing the -8 fuel element. Also included in the first serieswas testing of the -7 fuel element, which is the baseline 200lbf thrust fuel element (see table 4-1).

A second test series was performed, incorporating hardwaremodifications based on the initial test series results.Three new 200 lbf equivalent fuel elements were evaluated inthis series, as was a modification to the injector pintle.

5.4.1.1 Test Facility

All hot fire testing of the SSRT engine in the basic programwas performed at TRW's Capistrano Test Site (CTS) Facility inthe HEPTS HA2A vacuum capsule. A facility schematic is shownin Figure 5-3. A mechanical pumping system maintained thetest cell at less than 50 torr absolute pressure for all hotfire testing.

The fuel propellant tank was an 80 gallon hydrazine tank withan outer glycol jacket that allowed thermal conditioning ofthe propellant. Liquid oxygen propellant tankage included a150 gallon run tank, fed from a 300 gallon L02 storage tank.Both L02 tanks were vacuum insulated. The L02 in the runtank was kept at its normal boiling point (-298F) by ventingthe tank to atmospheric pressure between tests. L02propellant lines to the test capsule were insulated, and werechilled prior to a test by bleeding LO from the run tank tothe fire valve. The line downstream of the L02 fire valveand the injector were pre-chilled by liquid nitrogen prior toeach test.

The igniter fluid was supplied by a small N204 tank andcontrolled by a cavitating venturi. Propellant line heaterswere used on the fuel and igniter lines to prevent freezingof the propellants during engine start-up. All propellantlines were purged with GN2 during the start up and shutdowntransients. All valve timing was controlled by an IBM PCbased timer that allowed millisecond timing resolution of thevalve command signals.

60

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5.4.1.2 Test Instrumentation and Data Recording

Performance evaluation of the SSRT engine was based on C*performance measurements. Redundant instrumentation was usedon all performance related parameters, including propellantflow rates, chamber pressure transducers, and venturi inletpressures. Cavitating venturis were used to control the flowrates to the engine. These venturis have been water flowcalibrated. Three calibrated flowmeters in series were usedto measure the fuel flow rate. The oxidizer flow rate wasdetermined by use of a cav4.tating venturi.

Thermocouple instrumentation included 12 type K thermocouplesbrazed into the copper chamber. Also, 12 thermocouples werelocated at key locations on the injector to allow anassessment of the thermal characteristics of the injectorhead end. Other thermocouple instrumentation includedpropellant temperatures at the flowmeters, venturi inlets andengine inlets. An instrumentation list is presented in Table5-1.

Critical temperature measurements such as chamber andinjector dome temperatures were displayed on strip charts forreal time monitoring during testing. Early shutdown of atest was determined by strip chart trends. Oscillographrecording of critical parameters was available for quick lookand transient analysis of each test. All instrumentation wasrecorded on digital tape and printed in numeric format fordata reduction analysis.

5.4.2 Test Summary of -8 Fuel Element

Initial hot fire testing of the SSRT engine was performedwith the -8 fuel element. This extension was designed basedon the TRW Dual Mode Liquid Apogee Engine (DM-LAE) fuelgeometry. The -8 element was designed to match the fuelinjection geometry and flow characteristics of the DM-LAEengine as closely as possible. This allowed a directcomparison of the operating trends of the non-hypergolicL02/N2H4 propellant combination verses the well characterizedN204/N2H4 propellant combination utilizd by the DM-LAEengine. The nominal flow rate for this element wasestablished at an equivalent thrust of 125 lbf to match thefuel injection characteristics of the DM-LAE engine.

Twenty-five tests were performed with the -8 element, ac-cummulating 306.5 seconds of hot fire duration. The testresults for the -8 element are summarized in Table 5-2.Performance of the element was approximately 83% C*efficiency, compared to approximately 95% C* efficiency forthe DM-LAE Engine. Many of the DM-LAE performance trendswere non existent or not as clearly defined during testing ofthe SSRT engine with the -8 element.

62

TABLE 5-1SSRT INSTRUMENTATION LIST

RECORD/DISPLAYMETHOD

ID RANGE S/C OSC DVM PARAMETER

PC-I 0-300 PSIA X X X CHAMBER PRESSUREPC-2 0-300 PSIA CHAMBER PRESSUREPIO-1 0-1000 PSIA X X OXID INLET PRESSUREPIO-2 0-750 PSIA OXID INLET PRESSUREPID 0-500 PSIA OXID DISTRIBUTION PRESSUREPIF-1 0-1000 PSIA X X FUEL INLET PRESSUREPIF-2 0-750 PSIA FUEL INLET PRESSUREPOVI-1 0-1000 PSIA X OX VENTURI INLET PRESSUREPOVI-2 0-1000 PSIA OX VENTURI INLET PRESSUREPFVI-1 0-1000 PSIA X FU VENTURI INLET PRESSUREPFVI-2 0-1000 PSIA FU VENTURI INLET PRESSUREWO-1 0.15-0.30 LBM/S X X OXID FLOWRATEWO-2 0.15-0.30 LBM/S OXID FLOWRATEWO-3 0.15-0.30 LBM/S OXID FLOWRATEWF-1 0.20-0.40 LBM/S X X FUEL FLOWRATEWF-2 0.20-0.40 LBM/S FUEL FLOWRATEWF-3 0.20-0.40 LBM/S FUEL FLOWRATETOF -350 to -200°F OXID FEEDLINE TEMPTFF 40-100°F FUEL FEEDLINE TEMPTFI 40-100°F X FUEL INLET TEMPTOI -350 to -200"F X OXID INLET TEMPTOVI -350 to 60oF OXID VENTURI TEMPERATUREPIGT 0-1000 PSIA X IGNITION TANK PRESSUREPIGFV 0-1000 PSIA IGNITION FIRE VALVE PRESSPIGI-1 0-500 PSIA X X IGNITION INLET PRESSUREPIGI-2 0-500 PSIA IGNITION INLET PRESSURETIGN 40-100°F INGITION INLET TEMPPA-1 0-50 TORR X CELL PRESSUREPA-2 0-50 TORR CELL PRESSUREPOT 0-1000 PSIA X OXID TANK PRESSUREPFT 0-1000 PSIA X FUEL TANK PRESSURETR-1 0-2000°F X CHAMBER/NOZZLE TEMPSTHRU ITR-12 0-2000°F X CHAMBER/NOZZLE TEMPSTI-1 -300-1000°F X INJECTOR TEMPSTHRU 1 iTI-12 -300-1000"F X INJECTOR TEMPSTC-1 0-2500°F X TC PROBE TEMPSTHRUTC-16 0-2500"F X TC PROBE TEMPSACCEL 0-100 Gs X HEA ACCELEROMETER

*ALL PARAMETERS TO BE RECORDED ON DIGITAL TAPE.

63

Table 5-2-8 Fuel Element Test Summary

Test # Duration Wt PC C* Fuel OXHA2A- sec O/F lb/sec psia ftlsec Gap (Mf) Gap (do)

3992 5.0 0.717 0.3592 114.4 5443 0.0031 0.00833993 10.0 0.575 0.3126 84.8 4636 0.0031 0.00833994 10.0 0.550 0.3080 87.5 4856 0.0051 0.00833996 10.0 0.839 0.3601 111.4 5288 0.0031 0.00833998 10.0 0.786 0.3634 108.5 5124 0.0031 0.00633999 15.0 0.796 0.3641 112.8 5345 0.0031 0.00834000 15.0 0.776 0.3607 103.4 4926 0.0031 0.01034001 15.0 0.806 0.3674 106 4960 0.0042 0.00834002 15.0 0.832 0.4473 141.9 5490 0.0042 0.00834003 15.0 0.808 0.3674 109.1 5120 0.0024 0.00834004 14.8 0.830 0.4476 127.5 4912 0.0024 0.00934005 13.0 0.827 0.4498 129.2 4958 0.0042 0.00934006 15.0 0.782 0.3650 101.9 4798 0.0007 0.00934007 15.0 0.791 0.3634 105.7 5011 0.0007 0.00934008 15.0 0.789 0.3630 101.5 4805 0.0007 0.00934009 15.0 0.832 0.4475 125.6 4843 0.0007 0.00934010 15.0 0.799 0.3656 108.4 5112 0.0053 0.00624011 15.0 0.786 0.3628 109.5 5207 0.0082 0.00434012 8.9 0.787 0.3635 110.3 5215 0.0007 0.00434013 14.8 0.754 0.3544 105.6 5142 0.0031 0.00834014 15.0 0.780 0.3247 94.9 5029 0.0019 0.00834015 15.0 0.616 0.3653 106.9 5033 0.0019 0.00834016 15.0 0.841 0.4869 145.5 5185 0.0019 0.0083

64

Difficulties in obtaining single phase liquid oxygen flow tothe injector caused poor repeatability of the test data, andresulted in no clear-cut performance trends with varyinginjector parameters. The most significant factor affectingperformance was the amount of pre-chill to the injector andL02 run line bleed. Injector pressure drops and dischargecoefficients on the oxidizer circuit varied by ±35% duringtesting and averaged 20% lower than the oxidizer Cd measuredduring water flow of the injector, indicating vaporgeneration and two phase flow conditions.

Incomplete fuel vaporization was evidenced by the chamberwall thermocouple data. Row 1 measurements showed a tendencyto operate near the fuel saturation temperature, indicatingliquid fuel impingement at the wall. Throat thermocoupledata also corresponded to a low wall zone mixture ratio.Test durations for all -8 testing was limited by injectordome redline temperatures (500F) rather than chamberthermocouple redline (1000F).

The igniter for these tests was the same configuration testedin the 1990 IR&D program; a single N204 stream directedthrough the fuel spray pattern. This configuration caused ahigh heat load to one side of the dome during the igniterstage, resulting in a thermal maldistribution in the injectorat the start of the test.

On test HA2A-4002, a reaction of fuel and N204 in the igniterline (located at 6 o'clock) caused the line to rupture. Theengine was removed from the stand and a new igniterconfiguration was employed. The old igniter port was weldedshut and two new ports, located 180 degrees apart (at 9 and 3o'clock), were machined into the injector dome (see Figure 5-1 for both configurations). These igniter ports created afine spray fan directed axially down the chamber, through thefuel spray pattern. The ignition sequence with this igniterconfiguration was improved, resulting in less thermalmaldistribution to the injector during ignition. Theoriginal igniter would cause a thermal maldistribution ofapproximately 10OF during the ignition stage, while the newconfiguration had a maximum maldistribution of approximately30F. The igniter stage heat load to the injector was alsoreduced for the new configuration.

5.4.3 Test Summary of 200 lbf Elements

The remainder of the hot fire testing of the SSRT engine wasconducted with fuel elements designed for 200 lbf equivalentflow rates. These elements( -7, -9, -10 and -11) all haveequal slot flow areas, with the number of slots varying from36 to 60. Performance was improved dramatically over the -8fuel element, and test reproducibility and performance trenddefinition was also better. The higher oxidizer flow rate

65

allowed a colder oxidizer inlet temperature, resulting infewer problems with vapor generation and two phase flowconditions. This was subtantiated by the 20% higher averageoxidizer Cd measured during the -7 testing as compared to the-8 element testing, and by the lower oxidizer Cd variation of+10% for the -7 compared to ±35% for the -8 element.

5.4.3.1 -7 Element Results

Fourteen tests were performed with the -7 element,accumulating 147.3 seconds of hot fire duration. The testdata for the -7 element is summarized in Table 5-3.Performance of this element was in the 90% to 92% C*efficiency range. Although the performance was much morerepeatable than with the -8 element, there was still somescatter that probably related to the injector and L02 linepre-chill conditions. Performance was relatively insensitiveto flow rates and injector parameters, usually within thescatter of the data points. The -7 element performanceverses total flow at fixed gap conditions is shown in Figure5-4. The performance was essentially unchanged over theentire flow range tested. Performance verses mixture ratiofor the same injector gaps is shown in Figure 5-5. A slightincrease in performance with increasing mixture ratio isindicated, although the trend is within the data scatter.

The performance trends verses fuel gap and oxidizer gap ispresented in Figure 5-6 and 5-7. Again, the trend wasslight, indicating maximum performance at a fuel gap ofapproximately 0.0020 inch and for an oxidizer gap of 0.0185inch.

The injector gaps presented here are the gaps set prior tothe test based on shim changes. However, differentialthermal expansion between the fuel pintle and the oxidizersleeve caused a post-chill growth of about 0.0040 inch in thefuel gap during the burn. Thus, a set fuel gap of 0.0020inch resulted in an actual gap of approximately 0.0060 inchduring the test. The magnitude of the change was determinedby comparing hot fire fuel pressure drops to the water flowdata on the fuel injector. A thermal analysis of theinjector predicted a gap change of .0035 to .0040 inch basedon the sleeve outer diameter being chilled to -280F (from60F) prior to the run. The expansion of the fuel gap will beminimized in later hardware designs where the shim locationis moved from its present location (shown in Figure 5-1) tothe end of the sleeve near the fuel extension. This willgreatly reduce the free length for thermal expansion.

The oxidizer gap also experienced a gap increase, althoughthe magnitude of the change was more difficult to assessbecause of differences in injector body chill and LOZ densityfrom test to test. It was also likely that the oxidizer gap

66

Table 5-3-7 Fuel Element Test Summary

Test # Duration Wt PC C* Fuel OXHA2A- Sec O/F lb/sec psia ft/sec Gap (df) Gap (do)

4017 11.4 0.809 0.5814 188.0 5617 0.0019 0.01634018 11.4 0.799 0.5802 189.1 5662 0.0019 0.01634019 11.2 0.803 0.5776 186.0 5594 0.0019 0.01434020 11.2 0.798 0.5814 190.2 5687 0.0019 0.01834022 12.0 0.822 0.5861 184.4 5471 0.0000 0.01834025 10.0 0.785 0.5733 184.8 5584 0.0034 0.01834026 10.0 0.788 0.5789 184.2 5513 0.0019 0.02134027 10.0 0.789 0.5773 187.0 5618 0.0019 0.01834028 10.0 0.808 0.6400 204.8 5556 0.0019 0.01834029 10.0 0.810 0.5824 190.2 5667 0.0019 0.01834030 10.0 0.828 0.5394 174.2 5593 0.0019 0.01834031 10.0 0.679 0.5673 181.2 5531 0.0019 0.01834032 10.0 0.942 0.5729 186.2 5644 0.0019 0.01834033 10.0 0.817 0.6009 194.0 5607 0.0019 0.0183

67

SSRT Hot Fire Tests-7 Fuel Element

6000- - - _ _ _ _ _ _ _ _ _ _ _ _ - _ _ _ _

5900 . ..........

5.00..... ....... .......

5700----

n 5 400 - . .... . ........... .............

5100-. . ..

0.53 0.54 05 5 05 6 05 7 0.58 0.59 0.6 0.61 0.62 0.63 0.64Wt (Ib/sec)

Figure 5-4. -7 element C* verses total flow rate

68

SSRT Hot Fir e Tests:3-7 Fuel Element

6000. .

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5 7 0 0 ................

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-5400..........

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5 2 0 0 ....... .... ............ ........................ ......................... .........................

51001T I

0.65 0.'7 0.75 0.'8 0.85 0.,9 0.95Mixture Ratio (0/F)

Figure 5-5. -7 element C* verses mixture ratio

69


6000-

5900 .............. ..

5800 ............ ...... .-

~'5500-

5400 ........

Fdox = 0.0 183inch~

5100 -- r0 0.0005 0).001 0.0015 0.002 0.0025 0.003 0.0035

Fuel Gap (inch)Figure 5-6. -7 element fuel gap performance trend

70


5700-

5 8 0 ............................. ....................................................... ............ .. ..........

5400 - ~f _ .919..........................

5300-0. 014 0.0 15 0.01'l6 0.01'l7 0.01'l8 0.019 0.02 0.0'21 0. 02 2

Ox Gap (inch)Figure 5-7. -7 C* verses oxidizer gap

71

SSRT Hot Fire Tests-7 Fuel Element, Wt =0.59

2500.

2300-

.. ...................... ........... -------.. ......................... ...... ........

S2200-0

N:NS 2 1 0 0 -............... ....................... ... ............

1900-0.5 0.55 0.7 0.75 0.'8 0.85 0.,9 0.95

Mixture Ratio (0/F7)Figure 5-8. Wall Zone Gas Temperature verses mixture ratio

72

was changing slightly during the test due to the injectorbody warming up while the injector sleeve remained at the L02temperature. Comparing test data to water flow data of theoxidizer delta P indicates an average change in the oxidizergap of about 0.0020 inch during injector chill, which agreeswell with thermal analysis of the injector. With the domecooling concepts incorporated the oxidizer gap change shouldbe significantly reduced.

Chamber thermocouple data indicated a considerably higherheat load to the chamber for the -7 element compared to the -8. Throat thermocouples were reaching redline temperaturesof 100OF in 11 to 12 seconds for the -7 element, while theywere below 50OF after 15 seconds duration for the -8 element.A trend of increasing chamber heating rate for higher mixtureratios is presented in Figure 5-8. A wall zone gastemperature of about 2500F was calculated by comparing thethroat thermocouple data with predictions using the thermalmodel of the copper chamber. This corresponded to a wallzone mixture ratio of about 0.2, which indicated that steadystate operation with a columbium chamber with a comfortablethermal margin is feasible.

5.4.3.2 Test Series Number 2: -9, -10 and -11 Results

After testing and data analysis of the -7 element wascompleted, three new elements, -9 through -11, were designedand built. These elements were based on the -7 geometry,with modifications to some of the geometrical parameters (seesections 5.3 and 4.1), with emphasis on increasing the fuelvaporization rate.

All previous testing of the SSRT engine demonstrated that thefuel element geometry was the primary factor in engineperformance. For all fuel element geometries tested,the engine attained a certain level of performance, and onlysecondary changes in performance were observed by changinginjector and flow parameters (aside from the detrimentaleffects of two phase flow in the oxidizer circuit). Based onthis assessment, the test plan for the last three elementsincluded a limited number of tests for each element. If theelement tested didn't indicate increased performance in therange of parameters covered by these tests, the assessmentwas that no further gap or flow changes would make asignificant improvement in performance for that injectorgeometry.

The hot fire testing performed with the -9, -10 and -11elements are summarized in Table 5-4. The performance trendverses fuel gap for the three elements (-9, -10 and -11) ispresented in Figure 5-9. This trend was similar to thatdemonstrated by the -7 element, although the -11 element didperform better at larger fuel gaps than the other elements.The -9 and -11 elements matched the -7 performance element

73

Table 5-4

-9 Fuel Element Test Summary


4034 10.0 0.808 0.5737 184.6 5511 0.0020 0.01854035 10.0 0.759 0.5738 187.6 5654 0.0020 0.01854037 10.0 0.819 0.5913 189.3 5538 0.0033 0.01854038 10.0 0.812 0.5895 192.3 5641 0.0007 0.01854039 10.0 0.802 0.5897 190.8 5595 0.0007 0.01404041 10.0 0.801 0.5894 185.7 5445 0.0007 0.0210



4042 10.0 0.783 0.5848 187.9 5555 0.0020 0.01854043 10. 0.801 0.5883 187.6 5517 0.0034 0.01854044 10.0 0.795 0.5848 184.9 5470 0.0007 0.01854045 5.0 0.807 0.5913 193.2 5609 0.0020 0.0140



4046 10.0 0.791 0.5848 1,,1.9 5647 0.0018 0.01854047 10.0 0.809 0.5917 194.9 5710 0.0033 0.01854048 10.0 0.798 0.5878 189.1 5572 0.0045 0.01854049 5.0 0.800 0.5932 198.9 5769 0.0033 0.01404050 6.2 0.808 0.5935 197.0 5708 0.0033 0.01604052 9.8 0.802 0.5881 187.9 5534 0.0033 0.01204053 8.6 0.651 0.5894 193.6 5672 0.0033 0.01404054 5.0 0.923 0.6062 202.1 5741 0.00A3 0.01404055 6.2 0.821 0.5944 197.4 5723 0.0033 0.0140

74

SSRT Hot Fire TestsMR=O.8, Wt = 0.58

5800-

5750 ..... ... .... ...

57-0

'4-9

5450 ...... ... ..

-10

5405

closely, but the -10 element operated at a lower performancelevel. This demonstrated that high slot aspect ratio was notthe key to higher performance.

The performance of the elements verses oxidizer gap at thenominal flow conditions is shown in Figure 5-10. Once againthe -9 element showed the same performance trend as the -7,preferring the 0.0185 oxidizer gap. The -10 and -11 elementsdemonstrated higher performance at smaller oxidizer gaps.The -11 element was tested at the 0.0120 inch oxidizer gap,but the performance decreased dramatically at this condition.Uneven injector dome heating and unsteady chamber pressureduring this test indicated that there may have been a problemwith the oxidizer injection distribution.

A mixture ratio survey was conducted with the -11 element atthe 0.0033 inch fuel gap and the 0.0140 oxidizer gap. Theresults of this survey, Figure 5-11, indicated a trend ofincreasing performance with mixture ratio, similar to the -7element. Performance verses injector momentum ratio is shownin Figure 5-12. A general trend of increasing performancewith higher momentum ratio was observed for all the elements,with the exception of the test with highest oxidizer velocity(do = 0.0120) as discussed above.

Chamber heating rates for the -9 through -11 elements werevery similar to the -7 element, but the injector dome heatingrate was very rapid with the -11 element, especially atsmaller oxidizer gaps. At a 0.0140 inch oxidizer gap andhigh mixture ratio, the injector dome reached the 50OF redline temperature in about five seconds. The dome thermaldistribution was uneven, with a higher heating rate on oneside of the injector.

5.4.3.3 -11 Hybrid Results

The combination of relatively low wall zone mixture ratio andincreasing performance trends at higher over all mixtureratios led to the conclusion that the core combustion zone ofthe engine was operating at a high mixture ratio. Thisassessment was also confirmed by the Priem model of theinjector. A simple two-zone performance analysis indicatedthat if the wall zone was at a mixture ratio of 0.2, the coremixture ratio was approximately 1.3. Since the theoreticaloptimum mixture ratio for this propellant combination isapproximately 0.73, the assessment was made that asignificant increase in performance could be gained if marefuel could be directed into the core combustion zone,decreasing the core mixture ratio toward optimum. This wouldalso result in increased performance at a lower over allmixture ratio.

Increasing the fuel flow to the combustion core zone was thereasoning behind the hybrid injector. A sp-re pintle was

76

SSRT Hot Fire TestsMR=0.8, Wt= 0.58

5800 - _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _

5 7 5 0 .U....................

U

5 5 . . . . ................................

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

o 5550.....-----.--

5 5 0 0 ........... .... ................. ...................

5 4 5 0 ..................................

5400-0.011 0.013 0.015 0.017 0.019 0.021 0.023

Ox Gop (inch)0-7 (df=.oo 19) + -9 (df=.0007) w -10 (df=.002o) 0 -11 (df=.0o33)

Figure 5-10. C* verses oxidizer gop for 200 lbf elements

77

SSRT Hot Fire Tests-1 1 Element, Wt 0.59

5800-

5700....

U

.4-f50.0....

5600r 1T

0.6 0.65 0.7 0.75 0'8 08 5 '0.9 0.95Mixture Ratio (0/F)

Figure 5- 1. -11 element C* verses mixture ratio

78

SSRT Hot Fire Tests200 lbf Elements, 0/F =0. 8, Wt =0.59

5800 - - - __

~5650 +

,5550 +

5500 ...........

54 0 . ........ . ............................... .... .... ,... . . . . . . .

5400-0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 1.3

Momentum Ratio (FO/FF)

*-7 element + -9 element X - 10 element 0 -1 1 element

Figure 5-1 2. C* performonce verses Momentum Rotio

79

modified by drilling three like-on-like doublets into thetip, yielding a hollow cone spray pattern oriented axiallydown the centerline of the chamber. Figure 5-13 shows thepintle tip with the doublets installed. The doublets weredesigned to direct about 10% of the total fuel flow into thecore. This pintle was tested with the -11 element, sincethis element gave the highest performance.

The performance summary for the -11 hybrid injector ispresented in Table 5-5. C* efficiencies over 95% oftheroretical (ODK) were obtained, corresponding to aprojected vacuum Isp of >340 lbf-sec/lbm. However, asdiscussed below, problems with oxidizer injector delta Pvariations prevented a complete characterization of theinjector.

The performance trend of the hybrid injector compared to thebasic -11 element tests is shown in Figure 5-14. At amixture ratio of 0.8, performance of the two injectors wasapproximately equal. As the mixture ratio was decreased byincreasing the fuel flow rate, the performance of the hybridinjector increased, producing the highest performance at amixture ratio of 0.70. Decreasing the mixture ratio evenfurther, however, caused variations in the oxidizerdelta P that resulted in operation at a lower performancelevel. The onset of this condition appeared primarily at lowmixture ratios, even though the oxidizer flow conditions wereessentially unchanged from other higher mixture ratio testswhere the condition was not observed.

Post test examination of the oxidizer metering geometryrevealed a contour downstream of the minimum area that couldallow the oxidizer to diffuse to a lower velocity withattendant pressure recovery. Apparently the attachment ofthe oxidizer to this surface was not complete, resulting invariations in the injection delta P. All future SSRTinjector hardware will incorporate modifications to theoxidizer metering geometry to eliminate this condition.

The throat wall zone gas temperature verses momentum ratiofor all of the 200 lbf elements is presented in Figure 5-15.This was derived from the chamber thermal model. The gastemperature increased with momentum ratio for all of theelements except for the hybrid injector, where the data wasdistorted by the oxidizer delta P variations discussed above.The -7 element had the lowest gas temperatures at a givenmomentum ratio, even though it had equivalent or even higherperformance than the -9 and -10 elements.

Dome and chamber thermocouple heating rates during testingwith the hybrid injector were about 20% higher than with the-11 element testing. This was expected, since the hybridpintle diverted some of the fuel flow from the wall zone tothe core, resulting in a higher wall zone mixture ratio. Italso created a higher effective momentum ratio (oxidizer to

80

Table 5-5-11 Hybrid Element Test Summary


4056 5.4 0.794 0.5864 193.2 5672 0.0033 0.01404057 5.0 0.724 0.6201 207.9 5769 0.0033 0.01404058 7.8 0.796 0.5872 194.7 5732 0.0010 0.01404059 5.0 0.733 0.6214 207.0 5739 0.0010 0.01404060 6.6 0.712 0.6146 204.4 5743 0.0021 0.01404061 7.2 0.702 0.6388 216.2 5858 0.0021 0.01404062 7.4 0.674 0.6474 215.5 5764 0.0021 0.01404063 5.8 0.728 0.6375 214.5 5807 0.0021 0.01404064 8.8 0.685 0.6588 210.1 5531 0.0021 0.01404065 8.2 0.717 0.6424 204.9 5510 0.0033 0.01404066 6.2 0.720 0.5902 196.5 5737 0.0021 0.01404067 9.9 0.675 0.5931 193.4 5613 0.0021 0.0140

81

ru4

C5'

"D -oIL 41-4

L

82

SSPT Hot Fire TestsPerformance Verses Mixture Ratio

5900-

5850.........................'. . .... . .

5-0 --------- ------- - .... .. ........ ............ ............. ........... ....................

5 7 0 . ..........7.......0.................... ......... .............. ................. .......i .....

U . +

5 . ..... ......................5...........6.......................................................... ... ... .... ......... .

++

5 5 0 0.... ....................... .......... .. .................... ........... ................560.6 065..0.7.075.....0.5..09 0.95MitreRto 0F

5600 ~ ~ - E---ement---- - -1- Hybrid---- Element............. ................ ............... ...... .........

Figre -1 4. lmnt+ -Hybrid inetolomeedtmb c 1 njto

83

SSRT Hot Fire TestsChamber Thermal Environment

3000-

-2 900 .- ... . ..... .......... .. ...... .........

LL X

~2600o~2500

2 6 0 .......-........ C

o2400-N

....... .u .. .. ........ ........

2000-1900 r 1

0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 1.3Momentum Ratio (FO/FF)

w -7 element + -9 element *-1 0 element0 -1 1 element X -1 1 hybrid

Figure 5-1 5. Wall Zone Gas Temperature verses momentum ratio

84

fuel) at the primary impingement point, which caused a hotterdome heating condition. As discussed in section 4, analysisof thermocouple data from the highest performing tests (HA2A-4061 and 4063) indicated that operation with a columbiumthrust chamber at a maximum steady state temperature of 2500Fis feasible.

The heat load to the injector during testing with the -11hybrid element was too high to allow steady state operationof the engine in the current configuration (see section4.2.1). Analysis indicated that the oxidizer will experiencefilm boiling in the injector, leading to an unacceptablethermal condition. Additional work is needed with injectorcooling concepts to reduce the heat load into the oxidizermain flow.

Post test hardware condition was excellent, with no signs ofexcessive heating or distortion. The copper chamber was inexcellent condition, with no signs of damage or erosion.

5.5 Test Conclusions

Overall, the exploratory test series of the SSRT was verysuccessful. A total of 76 tests was performed, accumulatingover 700 seconds of hot fire duration. The performance goalof a vacuum specific impulse (Isp) of >340 lbf-second/lbm(E = 204) was demonstrated, while maintaining a wallenvironment compatible with long duration operation of aradiation cooled columbium thrust chamber. The thermal loadto the injector was defined, yielding information for thedesign of a thermally adequate injector in the next programoption. Reliable ignition and stable operation of the enginewas demonstrated.

Testing of an injector fuel element geometry that yieldedhigh performance with hypergolic propellants resulted in alower level of performance with the L02/N2 H4 combination.Significant differences in operating characteristics betweenthe hypergolics and L02/N2H4 were observed. Also, it wasfound that operation at lower flow rates (F = 125 lbf)resulted in difficulties in attaining single phase liquidoxygen flow to the engine.

Testing at the 200 lbf thrust equivalent flow rate allowedmore control over the LO? inlet temperature, allowing singlephase flow and improved injector cooling for short durationtesting. The results of the testing with five 200 lbfinjector element configurations indicated that the fuelelement geometry is the primary performance driver for thisengine. The engine was relatively insensitive to otherinjection parameters such as total flow and injectionvelocities, especially compared to the hypergolic engines.

85

The results of the testing indicated that an increased numberof fuel slots resulted in highest performance. The smallerslots produced a finer drop size, resulting in bettervaporization of the fuel.

Although significant accomplishments were achieved in thistest series, additional development is required. A method ofcooling the injector to allow steady state thermal operationof the engine is required. The impacts of the injectorcooling approach on performance must be assessed. Moreextensive performance mapping of the injector is required, aswell as the demonstration of long duration operation.

86

6.0 TEST PLANS

A preliminary logic plan was developed for Option 1 of theSSRT Program. This logic plan is shown in Figure 6-1.

6.1 Option 1

The major emphasis on the Option 1 program is evaluation ofof the most promising methods of dome cooling fordetermination of the effectiveness of these concepts andtheir impact on performance and thrust chamber walltemperatures. A preliminary design of the injectorintegrating the cooling concept for generation of maximumperformance will be accomplished at the conclusion of Option1.

New injector hardware will be designed using a thermallyisolated dome. This dome is designed to have the capabilityof using replaceable auxiliary sections incorporating thermalblockage and film cooling dome cooling concepts.

Pintle, sleeve, oxidizer gap and fuel gap changes areincorporated into the injector design similarly to the Basicprogram design. This injector is designed and manufacturedto allow for wide flexibility of testing in a cost effectivemanner. In addition a study will be conducted to assessignition methods so it can be incorporated into thepreliminary design.

Testing will be accomplished in the same test cell as theBasic program tests (HA2A). The preliminary logic matrix ispresented in Figure 6-2. The detailed test plan will beprepared upon completion of design. The test program isconfigured to obtain the data required to determine the domecooling concept for integration into the Option 2 injector.The test program will utilize two injector elements and testvarying oxidizer (60) and fuel (6 f) gaps and mixture ratio(0/F) and total flow (WT). Analysis of the test data withcorrelations of performance and thermal considerations willbe generated to allow an understanding for further design andtest.

A preliminary design of an injector integrating the bestcooling concept and the high performance mechanisms will beaccomplished by the completion of the Option 1 program. Thebasis of the Option 2 detailed design will be provided bythis preliminary design developed based on the test resultsof Option 1.

87

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Q) 0 LiiZ E-Lii LL.

0 &4 -I-4 4 0 a-

2m AL.(S L-) Lii : 2:z Of Li- Lii Lii

L ~ X:Lii I--0(A -i >

m(A L

:t L, UL

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a)4-*4r -.- a)1 CD ( C.) .=3 - 4J +. 4- W 04 m -

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10 ~ ~ ~ CJ. UJ a) -4U *4- )c4 .Q 0 +1>- C. D I= --LL 0C

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7.0 CONCLUSIONS

Several conclusions were reached as a result of completion ofthe basic SSRT program.

" The greatest potential usage for the Space Storable engineis utilization as an advanced dual mode apogee/perigeeengine in a dual mode propulsion system for placement andmaintenance of satellites in GEO. The best propellantcombination to achieve this usage is LO2-N 2H4 as itprovides the best system/engine capability includingmaximization of payload into orbit and achieved the bestoverall rating of the characteristics of the fuelsevaluated.

" Thermal and performance analyses indicated that highperformance of 340 lbf-sec/lbm (E = 204) could be achievedwith operation in a columbium thrust chamber.

" Testing confirmed the analyses. Six injector geometriesindicated the need to redesign the injector dome to preventtwo-phase LO . The testing demonstrated that highperformance 195% C*)(Isp 2 340 lbf-sec/lbm with E = 204)could be achieved and that operation in a columbium thrustchamber is feasible. The use of a rhenium thrust chamberis another alternative which would allow even higherperformance (approaching 97% C* to yield Isp -- 350lbf-sec/lbm).

90

8.0 RECOMMENDATIONS

The major recommendation based upon the basic program resultsis to continue the development of the L02-N2H4 Space Storableengine with Option 1. The emphasis on the Option 1 programis resolving the dome heating issues by incorporating domecooling concepts and evaluation of these concepts by test todetermine cooling capability and its impact on performance.Ignition concepts should also be studied to determine theconcept to be incorporated into the integrated injector ofOption 2. The output of the Option 1 program should be thepreliminary design of the best cooling concept in theinjector providing maximum performance.

The recommendations for the Option 2 and Option 3 programsare to complete development of the Space Storable engine toallow verification and qualification beyond Option 3. Therecommendation for the Option 2 program is to develop theintegral injector incorporating high performance and domecooling and demonstrate its characteristics by test. Theresults of the Option 2 program are factored into Option 3.The recommendation for the Option 3 program is to develop aflight-type engine and demonstrate its characteristics bytest prior to shipment to NASA-LeRC.

91

Form AppomveREPORT DOCUMENTATION PAGE OMB No. 0704.0188

Public reporing burden f" ha Coiledaon 06 infornabon is estimated to average 1 hour per response. including he tot review insrutions. searching existng data sources.gaheing and maintaining ft data noided, and completing and reviews ie collecton of inlormaton. Send conments regarding hls burden estimate or any other aspect of htcolection of infonna don. including suggestions for reduang this burden, to Washington Headquarters Servi es. Direc orate for inlorink i Operations ind IRepons. 1215 JefersonDavs Highway. Sut 1204. Artoton. VA 22202-4302. and to the Office of Management and Budget. Paperwork Reduction Project (0704-0188). Washington. DC 20503.

1. AGENCY USE ONLY (Leave blank) j2. REPORT DATE 3. REPORT TYPE AND DATEP COVERED

12 May 1 992 Final Contractor Report4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Space Storable Rocket TechnologyFinal Report - Basic Program

WU-6. AUTHOR(S) Contract NAS 3-26246

Melvin L. Chazen A. Ramon CasillasThomas Mueller David Huang

7. PERFORMING ORGANIZATION NAME(S) AND ADORESSES) 8. PERFORMING ORGANIZATIONTRW Space & Technology Group REPORT NUMBER

Applied Technology DivisionOne Space Park E- NoneRedondo Beach, CA

9. SPONSORING/MONPE T RING AGENCY NAMES(S) AND ADDRESSES ) 10. SPONSORING/MONITORINGAGENCY REPORT NUMBER

National Aeronautics and Space AdministrationLewis Research Center NASA CR- 1 89131Cleveland, Ohio 44135 - 3191

11. SUPPLEMENTARY NOTES

Project Manager - Mr. James A. BiaglowSpace Propulsion Technology DivisionNASA-Lewis Research Center

12a. OISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Unclassified -UnlimitedSubject Category

13. ABSTRACT (Maximum 200 words)The Space Storable Rocket Technology Program (SSRT) was conducted for NASA-LeRC byTRW to establish a technology base for a new class of high performance and long-lifebipropellant engines using space storable propellants. The results of the initialphase of this systematic multi-year program are described. Task I evaluated severalcharacteristics for a number of fuels to determine the best space storable fuel foruse with L02. The results of this task indicated that L02-N2H4 is the bestpropellant combination and provides the maximum mission/system capability-maximumpayload into GEO of satellites. Task 2, Preliminary Design, developed two models-performance and thermal. The performance model indicated the performance goal ofspecific impulse > 340 seconds (e = 204) could be achieved. The thermal model wasdeveloped and anchored to hot fire test data. Task 3, Exploratory Test, consistedof design, fabrication and testing of a 200 lbf thrust test engine operating at achamber pressure of 200 psia using L02-N2H4 . A total of 76 hot fire tests wereconducted demonstrating performance > 340 seconds (c = 204) which is a 25 secondspecific impulse improvement over the existing highest performance flight apogee typeengines.

14. SUBJECT TERMS 15. NUMBER OF PAGESRocket engines, Satellite propulsion, Bipropellant engines,Space storable, High performance engines, Long life egiiies. 16. PRICECODE

A0517. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT

OF REPORT OF THIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified SARNSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)

92 Prescribed by ANSI Std. Z39-18298-102

TRW Applied Technology One Space PrkDivision Pedondo Beach, C"A 902'Space & Tecnnolog, Group 3108 12 4321

SN 5637225 June 1992SSRT-022

National Aeronautics and Space AdministrationLewis Research Center21000 Brookpark RoadCleveland, Ohio 44135

Attention: Mr. James Biaglow, MS SPTD-lProject Manager

Subject: Contract NAS 3-26246Space Storable Rocket Technology (SSRT)Basic Program Final Report Errata

Dear Mr. Biaglow:

Enclosed is page 45a of the Final Repcrt for the SSRT BasicProgram which was omitted from the report. This page isbeing distributed to the receivers of the Final Report.

Sincerely,

Melvin L. Cha4,Program Manager - SSRT

MC:scCC: Saundra R. Hill, 01/1040

TRW Inc

initial effective gas temperature was the same as oftest -4000 (16500 F) but the steady-state valuedecreased to only 1300OF (Figure 4-10). Local filmcoefficients on the gas side were the same as for test-4000. Resulting heat flows are shown in Figure 4-11.For a 400°F dome (- 10 secs. into the run, Figure 4-8), the LO would have to remove 1.7 Btu/sec in thepresent injector design. The L02 was able to absorbonly - 0.5 Btu/sec due in part to film boiling over aconsiderable cooling area.

o Hiah Performance. Test number, HA2A-4061, with adifferent injector element and at a higher pressurethan -4000 and -3999, had high performance (95.1% C*).Measured vs. calculated dome and neck temperatures forthis case are shown in Figure 4-12 and 4-13respectively. The imposed gas temperature is shown inFigure 4-14. Reflecting an improved ignition design,the initial temperature was 1300OF for -4000 and-3999). Steady-state gas temperature, however,doubled to 26000 F. Gas-side film coefficients forzones 1 and 2 (see Figure 4-3) were an order ofmagnitude higher then over the rest of the area, butapproximately the same as those used for zone 1, tests-4000 and -3999. Corresponding heat flows are shownin Figure 4-15. For a dome temperature of 400°F(about 4 secs. into test), the L02 would have toabsorb 4 Btu/sec to stabilize the temperatures. Theonset of film boiling was clearly seen in the sharpdecrease in the heat absorbed by the L02 at 7 seconds.

The above results indicated boiling of L02 in the injectorpassage will be difficult to prevent in the presentconfiguration. Therefore, preliminary investigations wereconducted to identify various methods of avoiding filmboiling. Table 4-2 presents these concepts includingadvantages and disadvantages/concerns. Further evaluation ofthe best of these concepts and critical experiments will beconducted in Option 1 to assess their capabilities prior toincorporation into the design.

4.2.2 Thrust Chamber Thermal Analyses

Thermal analyses were conducted to assess the walltemperatures of the thrust chamber using the high performancetests with the -11 hybrid element. The results are shown inFigures 4-16 and 4-17 which indicate a wall zone combustiongas temperature of 29000 F. Using a columbium thrust chambercoated with R512E silicide coating, the maximum temperatureis 2444 0F (outside) and 2573 0 F (inside) which is slightlyupstream of the throat. The throat temperatures are 25530 F

II 45a

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