I86-20493 · I86-20493 CSCE 21H 63/20 Uncias 08625 ... Several recent studies of dedicated OTVs fos LSS missions have shown the advantages of multfple perigee burns to minimize gravity

I86-20493

CSCE 21H Uncias 63/20 08625

https://ntrs.nasa.gov/search.jsp?R=19860011022 2020-06-04T20:04:10+00:00Z

GDC-SP-84-050

c

L T

Prepared under Contract L-814740

GENE IVISIBN

GDC-SP -84-050

FOREWORD

documents the results of contract L-81 740, ?*Orbital Trmsfer ne Integmtion Study." This ~ t u d y waa conducted by General

Dynamics Convdr Division (GDC) fmm March - November 1984 under contract to Aemjet Techsystems Company for NASA-LeRC.

The CDC Study manager is Bill Ketchum. Other GDC personnel contributed to this Study and the key individuals and their particular contributions are as folio w a.

Ray GorsM Mark Henley John Maloney Bill Nagy Mike Simon Dennis Stachowitz Cris Torre Kenton Whitehead

Missions, Requirements, Operations Analysis Space Station Accomodations Reliability Economics Mass Properties, Trades Design Aerobraking

For further information contact :

Bill #&chum OTV Prog-rarn Mmages 06 aesol Dynamics /Conv& Bft-ision

S m Diego, California, 92138 Telephone : ( 619) 576- 3176

hdl ZOXE PI-9530

iii

Section

1

2

3

4

5

6

App@ndiX

I

I1

GDC-SP- 84- 050

T A B L E O F C O N T E N T S

INTRODUCTION MISSIONS AND R E ~ U I ~ E ~ ~ ~ T $

O T V CONCEPTS AND ENGINES

MISSIONS, SYSTEMS, PERFORMANCE INTERACTION AND TRADES

4.1 THRUST LEVEL 4.2 SINGLE AND MULTIPLE ENGINES 4.3 TRADE-OFFS

ENGINE DESIGN

5.1 ENGINE LIFE 5.2 ENGINE MAINTENANCE

CONCLUSIONS

ENGINE PARAMETRIC DATA

ENGINB COST DATA

!%E

4-1

4-2 4-2 4-9

5-1

5 -1 5-4

6-1

1-1

11-1

UUC-SP- 84-050

L

LIST OF FIGURES

Figure

I- I 1- 2 2- 1

2- 2

3- 1

3- 2 3- 3

3- 4

3- 5

3- 6

4- I.

4- 2

4- 3

OTV E n m e Support Study Elements

Acivmced L02-LH2 QTV Bnglne DeWtion Approach OTV Missions

OTV Mission Requirements

Space Based OTV LO2-LH2 Space Based OTV

Alternate OTV Confiprationa

Advanced QTV Propulsion System Concepts

Cryogenic, Reusable Space-Based OTV Technology Payoff (Lightwdght Tank, Structure, Advanced

Economic Benefit of Advanced Engine &llfssbn, System and Performance Aiialysio Interaction

Large Space Strtacteires

Ptiyh& S ~ ~ ~ ~ i t i ~ i t y to T ~ t d ~ h ~ ~ ~ ~ - ~ ~ ~ M i ~ ~ i ~ r t s

Engine AerQaSEht)

4- 4 LSS Mission Qgerations

4- 5

4- 6

Mission Trmsfer Coast Period Propellant Losses

Payload Sensitivity to Total Thrust-Manned Mission

Page

1-1 1-2 2-1

2-2

3-1

3-2

3-4

3-6

3-7

3-7

4-1

4-3

4-4

4-4

4-5

4-5

4- 7 Radiation E ~ ~ ~ ~ u r e - ~ ~ ~ @ ~ GEO Sortie Udng Multiple

4- 8 Manned GEQ Sortie Mission Options 4-6

Perigee Burn8 4-6

4- 9 Manned GEO Sortie Ulission Operations 4-7

4- 10 Required Number of Engine Tests Needed to Demonstrate Propuledon Syetem Reliability A s A Function of Engine Configuration (Zero Percent Correlation) 4-7

i 8-11 AdcTiaonal Propellant Penalty Bequired far ACS 4-a

and Oparationcil Qualification 4-8

4-13 Manned OTV Prapulsicsn System ReliabWty 4-9

I

4- 12 OTV En@ine R-diability wEl Be Assured Through Testing

v i

i

GDC-SP- 84-050

LIST OF FIGURES, Contd

Figure

4-14

4- 15

4-16

4- 17

4- 18

4- 19

5- 1

5- 2

5- 3

5- 4

5- 5

5- 6

5- 7

5- 8

Engine Trade Study Impact on VeMcle Design

Thrust Level, Perigee Burns Trade

Area Ratio Trade

Area Ratio, Chamber Pressure Trade Fixed Vs Extendible N022les Trade

Mixture Ratio Trade

Economic Benefit of Long-Life Engine

Sample Case

Life-Cycle Benefit of Long-Life OTV Engine Sensitivity To Enghe Unit Cost

Life-Cycle Benefit of Long-Life OTV Engine Sensitivity To Engine Overhaul Frequency

OTV Engine Meintenace Baseline Assumptions Task Coots For Major CITY Engine Overhauls Task Coats For Routine QPTV Err

Benefit of OTV Engine Space-Maintainability (Routine Maintenance) As A Function of Payload Weight Penalty Incurred).

Page

4- 10

4- 11

4- 11.

4- 12

4-12

4- I3

5- 1

5- 2

5- 3

5- 3

5- 4

5- 5

5- 7

5- 7

GDC-SP- 84-050

SUMMARY

NASA-LeRC is sponsorhg industry studes to establish the technology base for an advanced engine for orbital transfer vehicles for mid-1990s IOC . Engine contractors are being assisted by vehfclo contractors Po 'define the requirements, interface conditions, and operational design cdteria for new LO 2- LH 2 propulsion eystems applicable to future orbit transfer vehicles and to assess the impacts of apace badng, man rating, and low-G transfer missions on propulsion system design requirements.

This report presents the results of a study conducted by GDC under contract to Aerojet for NASA-LeRC. The primary study emphasis was to determine. what the OTS engine thrust level should be, how many engines are required on the OTV, and how the OTV engine should be designed. This was accomplhhed by evaluating planned OTV missions and concepts to determine the requirements for the OTV propulsion system, conducting tradwffs and comparisons to optimize OTV capability, and evaluating reliability and maintenance to determine the recommended OTV en ne design for future development.

i

Mission analysis resulted in three major mission catagories. GEO Satellite missions accounted for the majority. Low thrust ESS and manned CEO misu~ons are fewer and later, approximately same time as apace based OTV EOC m d awBZ- ability of new engine, but rnme demmdhg and are, therefore, the &s.eriminatops for the QYV propulsion -ystem.

Considering the 7 to 10 year development time for a new engine and the mid- 1990s IOC of the LSS and manned mi xions, the availability of a new space based OTV is expected with advanced engines, composite structure, lightweight tanks, and aerobraking. Although several OTV concepts were considered, an orbiter c a r p bay launched, space assembled, symmetrical lifting aerobrake , single stage L02-LH2 OTV was eolected for analysis. Substantial performance and economic benefits of advanced engines, lightweight structures, and aeraarasist are ahown. The characteristics of the advanced engines being considered by Aerojet, Rocket- dyne, and Pratt & Whitney were used. Additicnal parametric data were supplied by the engine contractors for other thrust levels for use in the trade studies. The objective of establishing one engine desigii required consideration of both the manned and the LSS missions.

The most difficult mission is the manned GEO soptie mission which establishes the maximurn vehicle size and the highest thrust requirements, whi' e Large Space Structure (LSS) missions with LEO deployment and checkout dcierrnine the minimum thrust requirements. Since these are conflicting requirements for one engine, effort concentrated on resolving this by attempting to determine a desigrn thrust level that would satisfy the manned mission, and w i t h throttling, also I

GDC-SP-84-050

satisfy the LSS missions. This manned mission has previously been assumed to be limited to a high thrust level to reduce gravity losses with a single perigee burn to minimize crew radiation exposure/passes thru the Van-Allen radiation belts. Several recent studies of dedicated OTVs fos LSS missions have shown the advantages of multfple perigee burns to minimize gravity losses at the low thrust level8 needed to limit acceleration bade on large space structure missions.

Recent work by NASA-LeRC indicates that multiple passes thru the Van-Allen radiation belts would not necessarily h c u r excessive r a a t i o n doaa lower thrust could satisfy both the manned mission and the LSS mission. Send- tivity to total thrust for LSS missions waB determined showing the advantage of lower thrust levels and multiple perigee burns to obtain the largest LSS diameter; a large symmet phased array system was used for analysis. The results indicate that alth payload weight capability decreases, the diameter of the payload reachea ~ u 1 optimum at 1000 to 2000 lbf thrust 88 a result of reduced structural loads. The effects of gravity losses, Isp reduction, and mission transfer :asses were included. Sensitivity to total thrust for the manned mission was determined which also shows advantages of lower thrust levels, lighter engines and vehicle systems, and multiple perigee burns to obtain the best payload weight. Optimum total thrust for the manned mission, however, is considerably higher than for LSS missions (6000 to 12000 lbf vs. 1000 to 2000 Ibf)

Using mdiation data fmm NASA J56 /LeRC, crew exposure was determhed €OF 01218, two, and four perigee burn5 and ne week at GEO showing that up ta four burne could be tolerated without increasing the current manned ~ o d ~ l e mdhtion shield

were not included in this study.

The manned mission requires a very high probability of safe crew return. An overall propulsion system reliability of 0.9997 was oelwted (based on USA traffic statistics) which would require a single engine of exceptionally high reliability or the need for redundane*r. Multiple engines provide for single failure tolerance, eliminating the need for rescue operations, and reduces the number of tests required to demonstrate the needed reliability. A single engine design would have to demonstrate 7600 failure free tests, while a two engine configuration Fequires only 140 tests. While the ACS (if H2-02) could provide a backup to a single main engine, it is expected that its lower Isp would require additional propellant to be carried. Some OTV missions will be flown prior to the first manned mission, giving the opportunity to help demonstrate the needed reliability. For comparison, the RL-10 e n m e (based on 69 Centaur flights to date) has a predicted start probability of 0.999797 and failure sate of 509 failures per million hours of operation. U s i n g these numbers, analysis shows that two main engines will attain the desired reliability (0.3997) even with correlation factors, non- independent failure modes, as high as 5 to 10 percent.

7

odified trajectories for further reduced radiiation are possible but 1

1,

X

GDC -SP-84- 050

Payload optimization for the manned mission was evaluated for the following engine parameters :

0 Aerojet, Pratt a Whitney, Rocketdyne

o T h n S t , 3000-25000 LBF

e Number of perigee burns, 1-4

e Number of engines, 2-4

BD Nozzle area ratio, 600-3000

o Chamber pressure, 1500-2500 PSIA

e Mixture ratio, 5-7

e Fixed, extendible/ret?actuble nozzles

While several vehicle concepts were considered including modular and aft c a r p carrier concepts with various aerobrake options, the modular tanks /symmetrical lifting brake concept was selected for the trade studfes. Evduation of the aerobrake/en@ine interaction determined that doors wouki be neeesssry to cover the engines during the aeropass. The ktermUon of the OTV/En@ne/Aerobr&e was evaluated. As the engine length incceaees ( f ~ i n c t i ~ ~ ~ of thruet, QX'BR ratio, chamber pressure, fixed vs . extendible mazzfes) , the aembmke diameter (weight) must increase to prevent flow stream impingement on the payload. The number of engines and nozzle exit diameter impacts the engine support structure and aerobrake door size. Altogether, these allow trades to determine optimum engine design and sensitivity.

I

1

The advantage of lower thrust engines and multiple perigee burns i s shown. Additional trades showed the advantages of extendible nozzles, high chamber pressuses, and high mixture ratios. A nozzle area ratio of .. 1000 appeared optimum.

While there is a benefit for designing a long life engine, there appears to be little advantage €or reducing the frequency of major overhauls beyond 20 to 30 missions. Major overhaul of the engine for a space based OTV should be done on the ground to reduce cost, while routhe maintenance is shown to be advantageous in space for anticipated task manhours.

This study has shown that future OTV engine requirements will be determined by LSS and manned missions. To satisfy the manned reliability requirement, twin engines appear Po be needed. The optimum engine thrust level is in the range of

xi

GDC-SP - 84-050

3500 to 6500 Ibf each, depending on the number of perigee burns for the-manned mission. Although the lower thrust level is preferable for less vehicle and payload design impact, this is contingent on the acceptance of multiple perigee bums for the manned mission and on the ability of the engine manufacturers to produce a high performance, relhble, maintainable engine at lower thrust with additional starts and longer burn time.

GDC -SP- 84- 050

SECTION 1

INTRODUCTION

NASA LeBC is sponso&g industry studh3 to establish the technology b&se for an advanced engine for orbital transfer vel~Mes for mid-1990a ICE. Engine contractors are being assieted by vehicle contractors to define the requirements, Interface conditions, and operational deaign criteria for new L02-LN2 propulsion systems applicable to future orbit transfer vehicles and to assess the impacts of space basing, man rating, and low-G transfer missions on propulsion Bystem design requirements.

This report presents the results of a study conducted by General Dynamics/ Convair under contract to Aerojet for NASA-LeRC. The primary study emphasis was to determine what the OTV engine thrust level should be, how many engines are required on the OTV, and how the OTV engine should be designed. This was accomplished by evaluating planned OTV missions and concepts to determine the req&ensnts for the QTV propulsion system, conducting tradeoffs and comparisons to optimize OTV capability, and evaluating reliability and maintenance to determine the recommended OTV engine design for future deqelopment (Figures 1-1, 1-2).

m driving mission

* Payload dafinition * Number of flightJ

* Requirements for manned

* FWormance & operations concepts

* Engine types. numbers

* Modular & aft cargo carrier requifemenW mission. satellite placement & & arrangements

large space structures mission

Aeroassist concepts

* Vehlclelengine integrati?n * Thrust levellnumber * Life

Number af engines

of engines Maintenance

mixture ratio, etc 269.022-1 Nozzle area ratio, 2811473&1A

Figure 1-1. OTV Engine Support Study Elements

1- 1

GDC -SP - 84- 050

e e e .

ThNSI leVel Number of bums Number of engines Design concept

r, mromed hain engine

Space Station man oedica!ed low thrust eogine

Separate development

r,

e Assured checkwt mth Shuttle OT

assistarrse

zaltl(M.13

269.022-2,

a LEO deployment fl

e Man M i s t at OEO Expendable not initially available

Gmund- based

Figure 1-2. Advanced LOz-LPIz OTV En@ine Definition Approach

I

1 - . .",

1- 2

0 GEO s ~ ~ ~ ~ i i ~ ~ missions - 70% commercial & NASA mark t share - 5 to 7

missions per year (3 to 4 satellites manifested on each mission = 10,000 ib)

- Servicing - 2 missions per year - DoD - 6 missions

1 I -...

GDC-BP-84-050

SECTION 2

MISSIONS AND REQUIREMENTS

Mission anaiysis resulted in three major mission categorize (Flgures 2-1, 2-21, GEO Satellite missions accounted for the majority. Low thrrlst LSS And manned GEO missions are fewer and later, approximately same f3rn.3 as space based 0 . r V IOC , and availability of new engine, but mcrc. demanding and am, therefore, the discriminators for the 3TV propulsion system. The most zurrcnt NASA mission model and other sources were used to categorize requirements.

0 Low thrust LSS missions - 10,000 to 16,000 Ib payload - 2 to 4 missions per year

anned GEO sortie missions - 1 per year - 13,QQQ Ib payload round-trip

Figure 2-1. OTV Missions

i

I

269.022-3

IC__---

Lla)’ ...

MISSIOW GROUP EIJTAL GEO PLATFORM

S E ~ W ~ C I ~ G ~ A ~ ~ ~ E ~ GEO SQ

NED GEO STATION LOGISTICS PLANETARY

MULTIPLE GEO PAYLOAD DELIVERY SATE LLlTE DE LIVERY GEO SATELLITE

SDI AWNED PLklltETARY

I *LOdLHs SINGLE STAGE (AEROBRAKI

Figure 2-2. OTV Mission Requirements

c

GDC-SP- 84-050

PROPELLANT* REaUIREMENT

,485 SEC ISP)

- LO

1 11

0 8 2

19

12 3 3 3 2

19 27

8 131

-

a

-

18 1 '

88 200211989 137 1993/1983 -

269.0224

a

GIJC -SP- 84-050

SECTION 3

OTV CONCEPTS AND ENOINES

Considering the 7 to 10 year development ,time for a new en&e and the mid- 1990s IDG of the LSS and manned missions, the availabuty of a now space based %.,TV (Figures 3-1, 3-2, and Table 3-1) is expected with advanced enghes, compou.: 3 structure, lightweight tanks, and aerobraking. Although several OTV concepts were conoidered (Figure 3-3), an orbiter cargo bay launched, space assembled, symmetrical lifting aerobrake, single stage OTV was selected for analysis.

-- Figure 3-1. Space Based OTV I

1 f Besides higher tSp engines, several othelr technologies have been iCzntified that

will make OTV reuse econarnicelly beneficial. These include redudng the inert OTV weight and utilizing aemrnssbt.

Reduced weight can be achieved with advanced structures (eo.nposites) by decreasing the loads bposed during launch and powered opersffon and by reduced tank pressures, Decreased loads Lare possible by initially launching the OTV f r o m earth without propellant, and by low thrust powered operation.

3- 1

&gur@ 3-2. LQ2-LEI2 Space Based OTV

Space basing allows the OTV to bo launched hitimy without pmpeuants and refueled on-orbit. Since the loaded tanks will be exposed only to a vacuum, internal pressures need not exceed those resulting from propellant vapor pressures which can be just above the triple point, possibly not exceeding 3 psia, as opposed to sea-level saturation conditions (> 14.7 psia) for a ground based O T y .

Once fueled, loads can be minimized by use of low thrust during powered operation which will be needed for certain payloads, e.g., Large Space Structures.

Besides Lnert weight reduction, the technology of aeroassist can reduce the pro- pulsive AV requirement for return from GEO by 50 percent (7000 fps versus 14,000). For manned, round trip missions, this results in a 50 percent reduction in OTV propellant required.

I 5

To achieve these improvements, technology development is needed in each area.

3- 2

GDC -SP- 84-050

Table 3-1. Modular Tank Space-Based QTV, Weights Summary

Tank Sets Quad Twin

Core assembly

o Main engine & TVC

0 Docking system

o Astrionics Forward s aft service bulkheads

0 Structure

o Electrical power

e ACS 8 tank pressurization

I, 647

322 322

24 24

276 276

130 130

216 216

-- 1,647 _I_

291

218

291

218

o Tank module disconnects/attaches 20 20

0 Main propellant €eed

o Contingency

70

80

70

80

1,196 - Outrigger tank sets 2,390

Propellant tankage & fittings

Insulation

o Propellant acquisition & feed

Q Structure

o 1nstrumen:aMon

o ACS 8 tank pressurhation

@ Tank module disconnects/attaches

Q Contingency (5%)

Auxiliary fluids

314

348

648

789

49

87

35

120

180 -

151

174

324

395

25

43

18

60

o ACS usable propellant 90 60

o Fuel cell reactants 90 40

160 220

1950 Aerobrake I associated structure 1950

BURNOUT \JEIGMT 6382 5053 USABLE MPS PROPELLANT 53460 28730 USABLE IMPS PROPELLANT MASS FRACTION .893 .841

- - Residuals, boiloff 8 other losses

-

3- 3

“-L-- - - - - . - - -7- - A’ I-

\-% 3 7 L.

GDC-8P-84-050

AFT CARGO CARRIER

MOQULWA

SIDE BAL..lJPE

2139.022-7

Figure 3-3. Alternate OTV Configurations

3- 4

'Y. .

GDC- SP- 84- 050

The characteristics of the advanced engines being conaidered by Aerojet , Rocket- dyne, and Pratt & Whitney are shown in Figure 3-4. Additional parametric data were supplied by the engine contractors for other thrust level8 €or use in the trade studies. The payoff of advanced techmlag4es shows the advantages of advanced engines, lightweight Panks/atructure, and 8em assi& mpabflity (Figure 3- 5). Advanced engines and Idghtweight FwksIstmcture give high pay-

round trip mi8dons, but payload delivery missions are very sensitive to aem- brake weight.

off for payload delivery EIkdQn8. A@roafdist ve8 high payoff for payload J

Figure 3-6 shows the substantial economic benefiF of a new engine.

I

3- 5

. , - -- - I

PARAMETER

THRUST, LBF CYCLE CHAMBER PRESSURE,

NO?ZLE AREA RATIO SPECIFIC IFAPULSE,

TURCOMACHINERY SPEEDS,

P51A

LBF-SECILBIM

R P I

e "2 e 0 2

COMTROL THROWLEABILITY

4~ RAIYGE 61 MODE

KEY TECHNOLOGIES

AEROJET

GDC-SP-84-050

AEfiOJET

3,000 EXPAMDER H2-02

2,000 1,200: 1

> 480

2oopoo 75,000 CLOSED LOOP 30: 1 2 STEPS (15:l CONTINUOUS) GASEOUS OXYGEN DRIVE TURBIrUE

ANNULAR THRUST

MULTiPLE ENGINE CONTROL

ROCKETOYNE

15,000 EXPANDER H2

2,000 1,300:l

> 480

17a.000 56,200 CLOSED LOOP 30:l 3 STEPS DISCRETE HYDROGEM PUMP CRITICAL

MULTIVARIABLE CLOSED

HIGH AREA RATIO NOZZLE

SPEED

LOOP CONTROLS

ROCK€TDYNE

PRAIT & WHITNEY

15.000 EXPANDER H2

1.500 6M:l

>480

150,000 67,390 OPEN LOOP 30:l 3 STEPS DISCRETE HIGH SPEED HYDROGEN

ADVANCED THRUST

HIGH AREA RATIO NOZZLE

COOLED GEARS

PRATT & WkTNEY

ADOITIOWAL ~ A ~ A ~ E T R ~ C DATA W ~ S S ~ ~ L l ~ O FOR OTHER THRUST LEVELS

269.022-8 I I-

Figure 3-4, Advmnced OTV Propulsion System Concepts

3- 6

I

4

GDC -SP-84- 050

0 ADVANCE0 ENGINE & LIGHTWEIGHT TAMKS AN0 STRE'CTURE GIVE H I G H PAYOFF FOR PAYLOAD DELIVERY MiSSlONS

0 AEROASSIST GIVES HIGH PAYOFF FOR PAYLOAO

1 I

ROUMD-TRIP MISSIONS

MISSIONS VERY SENSITIVE TO AEROBRAKE WEIGHT

0 PAYLOAO DELIVERY

2BS.022-9A

Figure 3- 5. Cryogenic, Reusable Space-Based OTV Teehncslogy Payoff (Lightweight Tank, Structure, Crdvarteed Engineer, AemrsskQ)

-PAYLOAD OELIVEREO TO GEO (WO RETURN) -OTV ROUND TRIP (LEO-GEO-LEO) -TWIN TANK SET MOOULAP SBOTV (LO2/LH2) -485 SEC ISB MEW ENGINE - 1950 LE AEROBRWKE

S121MIYEAR BENEFIT"

S781NEAR BENEFIT"

- MEW EWGIME (+ 25 SEC IS) SAVES $80s/LBp~.

- AERBBRAKIfUG SAVES $517/LOp~:

-MEW ENGINE & AEROBRAKING SAVES $1058/LBp~ $159M/YEAR BENEFIT. -MEW ENGINE OFFERS 76% OF TOTAL BENEFIT - AEROBAAKIMG OFFERS 49% OF TOTAL BENEFIT

COST A ~ U ~ ~ I ~ ~ ~ ~ 1 ALL PAOBULSIVE 2 AEROBRAKED t I _-

- 7 Q OTV T U R N A ~ O U ~ ~ , Brd e PROPELLANT DELIVERY To LEO,

PAYLOAO DELIVERY TO LEO, $M -lL0028 X PAYLOAO WEIGHT

266.022-39 *15OpaO LBM CUMJL~TIVE P A Y L O A ~ PER YEAR TO GEO (18 OTV

Figure 3-6. Economic Benefit of Advanced Erame _- 3- 7

/

GDC-SP- 84-050

SECTION 4

MYSSIONS , SYSTEMS, PERFORMANCE INTERACTION AND TRADES

The objective of establfshirmg one engine design required mnsideratlon of both the manned and the LSS missions (Figure 4-1). The most difficult mission is the manned GEO sortie mission which establishes the maximum vehlcle Size and the highest thrust requirements, while LSS missions with LEO deployment and checkout, determine the minimum thrust requirements. Since these are conflicting requirements for one engine, effort concentrated on resolving this by attempting to determine a design thryst level tf,at would satlafy the manned mission, and with throttling, also satisfy the LSS missions.

! T I

e---

* Reduce number

Van Allen Belts of pasties through 21911470945A

269.022-10

Figure 4- 1. Mission, Systems and Performance Antilysis Interaction

The manned mission has pre4ously been assumed to be k i t e & to a high thrust level to reduce gravity losses with a single perigee burn to minimize crew raciia- tion exposure/passes thru the Van-Allen radiation belts. Several recent studies of dedicated OTVs for LSS missions have shown the advantages of multiple perigee burns to minimize gPavity losses at the low thrust levels needed to limit acceleration loads on large space structure missions,

I

4- 1

GDC-SP- 84- 050

Recent work by NASA-LeRC indicates that multiple passes thru the Van-Allen radiation belts would not necessarily incur excessive radiation dosage. Thus, lower thrust could satisfy both the manned mission and the LSS mission.

4.1 THRUST LEVEL

SensBtfvity to total thrust, for LSS missions was determined showing the advantage of lower thrust levels and multiple perigee burns to obtain the largest LSS diameter. A large symmetrical phased array system shown in Figure 4-2 was used for analysis. The results indicate that although payload weight capability decreases, the payload diameter reaches an optimum at 1000 to 2000 Ibf thrust as a result of reduced structural loads (Figures 4-3, 4-4). Effects of gravity losses, Isp reduction, and mission transfer losses have been included (Figure 4-5).

Sensitivity to total thrust for the manned mission was determined (Figure 4-6) which also shows advantages of lower thrust levels, lighter engines and vehicle systems, and multiple perigee burns to obtain the best payload weight. Optimum total thrust for the manned mission, however, is considerably higher than for LSS missions, 6000 to 12000 Ibf vs. 1000 to 2000 Ibf.

Using radiation data from NASA JSC/LeRC , crew exposure was determined for one, two, and four p e ~ g e e burns and one week at GEQ , showing that up to four burns could be tolerated without increasing the current manned module radiation shield thickness (Figuree 4- 7, 4- 8 , 4- 9) . Modified trejectories €or further reduced radiation are possible but were not included in this study.

4.2 S+NGLE AND MULTIPLE ENGINES

The manned mission requires a very high probability of safe crew return. An overall propulsion syetem reliability of 0.9997 was selected, based on USA traffic statistics, which would require a single engine of exceptionally high reliability or the need for redundancy. Multiple engines provide for single failure tolerance, eliminating the need --9r rescue operations, and reduces the number of tests required to demonb d e the needed reliabMty. Figure 4-10 shows that a single eng-ine design would have to demonstrate 7600 failure free tests, while a two engine configuration requires only 140 tests. While the ACS (if H2-02) could provide a backup to a single main engine, it is expected that its lower Isp would require additional propellant to be carried (Figure 4-11). Some OTV missions w i l l be flown prior to the first manned mission (Fig'ure 4-12), giving the opportunity to help demonstrate the needed reliability. For compnrimn, the RL-10 engine, based on 69 Centaur flights to date, has a predicted start probabdilty of 0.999797 and failure rate of 509 failures per million hours of operation. Using theae numbem, analysis shows that two main engines will attain the desired reliabsty (0.9997) even with correlation factors, non-independent failure modes, a8 high as 5 to 10 percent (Figure 4-13).

4- 2

,

I

- - 1

_?" y. . * -, , ..

GDC-SP- 84-050 O ~ ~ ~ ~ ~ L PAGE IS OF POOR QU

- -. " GEO -PLATFORM

c

PHASED ARRAY

281 14801 -9 269.022.1 1

Figure 4-2. Large Space Structures

4- 3

uQI.muuu-. T ! -. - - 4

GDC-SP- 84-050

QUAD TANK SETS TWIN TANK SETS 4

REFERENCE THRUST FOR 9 BURNS (XlOOO Ibf) REFERENCE THRUST FOR 9 BURNS(X1000 Ibf) 0.5 1 2 3 4 5 10

6O0 - - 8 24

23 c1 g 22

2 21 0

0.5 1 2 3 4 5 10 300 "

2 = 200

s 100

Bc

E

3 SE

0

e CONSTANT THRUST MISSION e THROTTLED MAIN ENGINE

7000- 0.02 0.1 0.2 0.5 1.0

3 20- 2 0.1 0.05 0.1 0.5

MAXIMUM ACCELERATION OM PAYLOAD (9) ' ~ A X I ~ U ~ ACCELERATION ON PAYLOAD (9) 289.022- 12

Figure 4- 3. Payload Sensitivity to Total Thrust-LLSS WliafGions

269.022-13

Figure 4-4. LSS Mission Operations

4- 4

i

1 L

* ASTITUOE CONTROL

m POWER

* BOILOFF, LEAKAGE

. 2 - 3 L B i m (SYMMETRICAL orv AND PAYLOAD)

0.5 - 1.0 LWHR (1 - 2 KW FUEL CELL)

- (ISOLATEO/INSULATEO TANKS; ULLAGE EXPANSION COOLING WITH MULTIPLE BURN)

269.022-14

Figure 4-5. Mission Transfer Coast Period Propellant Losses

GDC -SP-84.- 050

TOTAL THRUST (Ibf) 269.022-15

Figure 4- 6 . Payload Sensitivity to Total Thrust-Manned Missmn

4- 5

-- --I

i

a

4

Total

NASA skin dose limit

1\ NASA eye dose limit (not critical if goggles are worn)

One psriges burn Two perigee burns Three perigee burns

Thrust = 9,000 !bf Total burn duration = 00:48:00 Transfer time to GEO = 8.2 hr

Thrust = 6,000 Ibf Total bur!? duration = 01 12:OO Transfer tame :o GEO = 11 hr

Thrust = 12,000 Ibf Total burn duration = 00:36:30 Transfer time to GEO = 5.5 hr

269.022-17 28114700.464

Figure 4-8. Manned GEO Sortie Mission Options

4- 6

.. 281147f%Wh 269.022.18 Alumitwm radiation shielding ~- -

Figure 4- 7. Radiation Exposure-Manned CEO Sortie Using Multiple Perigee Burns

/' / i I I \

GEO

\ \ \

I

/

/ I I

\ / \ f

~ Burn no.

I 1 1 2,368

2 2,656

3 3,032

4 5,870

5 (deorbit) 6,095

250,000 ft aeropass perigee 6 (phasing) 500

Totel I 20,521

GDC-SP-84-050

Alt et Init Propellant

256

343

430

19,348

GZO

00: 13:48 10,248 01 :56:24

00:13:11 9,788 03:25:04

9,315 0504.1 1 00: 1 2 3 3

00:18:21 13,617 - 00: 1259 9,641 05:09:36

250 00:0:51 636 - 1:11:43 53,245

Total thrust = 6,000 Ibf

Payload weight = 13,000 Ibm Isp = 485s

269.022-18 2822470949A

Figure 4- 9. Mmned GEO Sartia Mission Operations

I / TYPl CA L MAN RATE0 7600 FOR

GLE ENGINE U A

0.99 0.999 0.9997 0.9999 10 20 10 60 100 2013 4 0 6 0 0 IOOO 2500

1 REQUIREhyENT [ J \

NUMBER OF FAILURE FREE TESTS R E Q U I R E 0 TO DEMONSTRATE 169.022.,9 ENGlNE RELIABILITY

RE5UIAED PROPULSION SYSTEM RE LlABf LlTY

!

F i p r e 4-10. Required Number of Y7ngine Tests Needed to Demonstrate Propulsion System Reliability A s A Function of Engine Configuration (Zero Percent Correlation)

4- 7

i I

i

GDC-SP-84-050

2 Main Engines

lsp, se:: 4851485 Return propellant, lbm 10,970 t Ascent propellant, lbm 44,070

I Total, lbm I 55,040

13,000 Ibm payload round trip to GEO 6,400 Ibm OTV burnout weight Main engine failure at GEO

485f425

46,955 i 2,885 12,960 4- 1,990

2lll4830(.10 26D.022-20

Figure 4-11. Additional Propellant Penalty Required for ACS

48-65 O W ~ i ~ s ~ ~ ~ ~ will

Engine confidence due to operational qualification

Engine confidence due to testing

Manned mission reliability requirornetnt

.---------

281 14709-91A 259.022.21 Engine tests Missions flcllvn

Figure 4-12. OTV Engine Reliability will B e Assured Through Testing and Operational Qualification

4- 8

(*

E

"

GDC -SP- 84- 050

R, = RZ eAt: 1 engine R2 = R: + 2R, (l-fl,)(l-C): 2 engines

Where R, = .999797 (start, stop) 1 RLIO engine X = 509 x 10" failures per hour C = 0, .OS, 0.1 (correlation) 269.022.22

% Figure 4-13. Manned OTV Propulsion System Reliability

4.3 TRADE-OFFS

Payload optimization for the manned mission was evaluated for the following engine parameters (Appendix I) :

a. Aerojet, Pratt B Whitney, Racketdyne b. Thrust, 3000 - 25000 LBF c . Number of perigee burns, 1 - 4

d. Number of engines, 2 - 4

e, Nozzle area ratio, 600 - 3000

f. Chamber pressure, 1500 - 2500 PSIA

g. Mixture ratio, 5 - 7

h . Fixed, extendible /retractable nozzles

While several vehicle concept R were considered, inclucking modular and aft cargo carder concepts with various aerobrake options, the modular tanks/symmetrical Ufting brake concept was selected for the trade studies. Evaluation of the aerobrake /engine interaction determined that doors would be necessary to cover the enfines during the aeropass, because of ce:rcerns of vehicle stability and control, flow field interactions, engine cooling, and leakage of base gasses to the OTV. 1

I

4- 9

GDC -SP-84-050

The interaction of the OTV/Engine/AerobrRke is shown (Figure 4-14). A s the engine length increases (function of thrust, area ratio, chamber pressure, fixed vs. extendible nozzles), the aerobrake dianeser (weight) must increase to prevent flow stream impingement on the payload. The number of engines and nozzle exit diameter impacts the engine support structure and aembrako door size. Altogether, these allow trade-offs to determine opthum engine design and sensitivity.

J

Figure 4-14. Engine Trade Study Impact on Vehicle Design

The advantage of lower thrust engines and multiple perigee burns is shown. (Figure 4-15), A nozzle area ratio of - 1000 appeared optimum (Figure 4-16). Additional trades (Figure 4-17 to 4-19) showed the advantages of extendible nozzles, high chamber pressures, and high mixture ratios.

i 4- 10

13

12 GEO round trip payload weight, 1,000 Ib

1’

GDC -SP- 84- 050

i

Two engines

Number of ascent perigee burns

4 3

2

Aerojet sngifle 0 1000*1 area ratio 0 Fixed nozzle

I ‘ 2.500 psia chamber pressure

I 1 I

3,000 4,000 5,000

Thrust per engine (Ibf)

Figure 4-15. Thlruat Level, Perigee Burns T~rnde

r Rocketdvne PC 1,;163 psih

l3/ L 40 inches I c

GEO round trip payload weight (1,000 Ib)

10 PC 31 1,500 psi - Area ratio

269.Q22-24 201 l4627.lJA

1

PI , . c ’. “4

6

2 engines (5,000 Ibf each) 4 perigee burns Extendable nozzles MR = 6

281 I4730-3A 269.022.25

Figure 4-16. Area Ratio Trade

4-21

I

\

GDC-SP-84-050

131

12 ~

GEO round trip payload weight 1,000 Ib

2.500 psia 2.000 psia

1 500 psia

3,000 Ibf Aerojet engine Twoengines

0 Fixed nozzle M R = 6

* Four perigee burns

I

1,000 2.000 3.000 10 L Area ‘ Ratio

Figure 4- 17. Area Ratio, Chsmber Pressure Trade

13’5 17 GEO :ound trip payload weight 11.5 (1,000 Ib)

Extendable

Fixed

0 1,000 2,000 3,000 9.5

Expansion ratio

__-_I-

$

3,000 Ibf Aerojet engine * 2engines e 2,000 psia charrber pressure * 4 perigee burns * M R = 6

Figure 4-18. Fixed Vs Extendible Nozzles Trade

4- 12

e.-

.--&

I

- .A’ 2

269.022-27 281147384A

\

\

GDC -SP - 84-050

GEO

12 - round trip

payload weight (1,OOO Ib)

11.

4 3 2

- Number of

c / perigee burns

I I I I I 5.0 5.5 6.0 6.5 7.0

Mixture ratio

10 1

3,000 Ifb Asrojet engine 0 2engines * 1,500 psia chamber pressure 0 1,OOO:l area ratio 2 U W M S h e Fixed nozzle 268.022-28

4-13

4. r"*

GDC-SP-84- 050

SECTION 5

ENGINE DESIGN

5.1 ENGINE LIFE

To define how the engine should be designed, an economic analysis was conducted for engine life and maintenance. While there is a benefit for designing a long Xife e n m e , there appears to be little advantage for reducing the frequency of major overhauls beyond 20 to 30 missions.

Figure 5-1 indicates the formula used to determine the benefit of a long-life engine and a description of the parameters used. Figure 5-2 is a sample case generated with these assumptions.

0 [D, + (Ns x Us) + N, x (T+R)] - [QL c (NL x U,) + Nr x (C+T+ RI! m = Development cost of aiternative (shorter-life) engine = Number of units of alternative engine required (over OTV life) = Unit cost of alternative engine = Number of short-life engine replacements required over OTV life = Cost of transporting one O W engine to LEO = Cost of replacing an OTV engine = Development cost of long-life engine = Number of units of tong-life engine required (over OTV life) = Unit cost of long-life engine = Number of engine refurbishments required (for rnaintenmce of IC rg-life

= Cost of refurbishing a long-life engine

0, N, Us N, T R DL NL UL Nf

C engine over OTV life)

281 14801.5 269.022-29

Figure 5-1. Economic Benefit of Long-Life En@n ;

-

5- 1

I i

e Number units required = 22 (10 mission life,

* Unit cost = $10M 200 missions total, 2 spares)

Lo e

* Number units required = 4 (100 mission life,

0 Unit cost = $10.42M 0 Number of refurbishments required = 6 (every 25 missions) 0 Cost of refurbishment = $1M (per refurbishment)

0 Cost of transporting one engine to LEO = $4M 0 Cost of replacing an engine = $0.8M

200 missions, 2 spares)

Figure 5-2. Sample Case

281 14801 4

269.022.30

I

The life-cycle economic benefit of a long-life (100 mission) OTV engine is closely

mission OTV engine. The long-life engine yields a positive undiscounted benefit

the short-life engine exeeds about $3 million (Figure 5-3). This calculation is based on data provided by Aerojet (Appendix 11, Tables 1-1 and 1-2) which indicate that the long-life engine has a $105 million. greater non-recurring cost and a $0.42 million greater recurring cost than the short-life engine.

related to the unit recurring production cost of the alternative short-life PO

over the OTV mission span, two engines in use at all times, if the unit cost of

I

t 5 i

These data also include the assumption that the long-life engine must be returned to Earth for each major refurbishment every 25 missions. Establish- ing the capability to do these refurbishing tasks in space could save $4 million in transportation costs per overhaul and hence increase the benefit of the long- life engine by as much as $24 million over the values indicated or this graph. It is expected, however, that the added costs of utilizing the Space Station for engine refurbishment wouA exceed these transportation cost savings,

The nominal refurbishment rate assumed for the long-life OTV engine, one OWF- haul per 25 missions I is sbwn to be in the optimal range (Figure 5-4). Economic benefit of the long life engine drops sharply at higher overhaul rates to $lOOM at one overhaul pep 16 missions, $53M at one overhaul per 9 missions, and to

5- 2

---r--- --- I

J - Y

I

GDC -SP-84- 050

A

t

\,

d00.

300

200 Life-cycle benefit ($M)

100

0

--I_ - Fi@-ure 5-4. Life-Cycle BeKefit of Long-Life OTV Engine Sensitidcy

5- 3

To Engine Overhaul Frequency

_--- .-_______ -.-- I -_I-----

/

.r& .<*d*?d-*!,A%***-A 3 .&

n a l l r s o l - 1 ~ Short-life engine unit cost ($!Vi) 269.022.31

Figure 5-3. Life-Cycle Benefit of Long-Life OTV En@ne Sensitivity To Engine Unit Cost

20 40 60 80 too I 20114709-18A 269.022-32 Number of OTV missions beheen overhauls

.

lj,

I

GDC-SP- 84- 050

zero at one overhaul per six missions. Benefits of reducing the refurbishment rate below the nominal rate are relatively modest Doubling the number of missions between overhauls results in a $25 million increase in life cycle benefit; the benefit of further reductions in the refurbishment rate ia barely noticeable.

5.2 ENGINE MAINTENANCE

Major overhaul of the engine for a space based OTV should be done on the ground to reduce cost (500 manhours at $100/hr on ground vs. 1000 manhours at $20000/hr in space) , while routine maintenance is shown to be advantageous in space for anticipated task manhours (20 on ground , 40 in space ),

The benefits and costa of space maintainability versus returning O T 7 engines to Earth for servichg were evaluated for two cases (Figure 5-51, "Routine maintenance" represents the most frequent and least complex type of servicing, assumed to nominally require 20 man-hours if performed on the ground and 40 hours if performed in space, with a frequency of one event every five missions. Establishment of Space Station facilities to support reutine maintenance in space is assumed to cost $10 million mone than establishment of similar facilities on Earth. Transportation costs involved in returning engines to Earth for routine maintenance are assumed to be $2.5 million per event.

Man-hours to perform task

Number of times performed (over 200 missions)

Tiansportation cost

Non-recurring cost A' for Space Station servicing equipment

Manpower cost

' Over cost of establishing same facilities on Earth

Routirie ~ ~ i n t ~ ~ ~ n c ~

Ground-20, Space-40

38 (every 5 missions)

$2SM

$1 OM (space only)

G round-$1 OO/h r Space-$20,000/hr

Major Overhaul

GrounddOO, Space-1 ,000

6 (every 25 missions)

$.EM

$1 OOM (space only)

Ground-$1 OOlhr Space-$20,000/hr

28114801. 269.022.:

!

t r

I

--

Figure 5- 5 . OTV Engine Maintenance Baseline Assumptions

5- 4

6r

,

"Maj -r overhaul" rep esent

GDC-SP- 84- 050

the opposite extrem in engin

\ I

servicing with 500 hours required if performed on Earth and 1,000 hours if done in space. Fre- quency of major overhauls is assumed to be once every 25 missions, and development of Space Station facilities to support major overhauls is assumed to cost $100 million more than providing the same capabilities on the ground. Transpor- tation costs are assumed to be $4 million per overhaul. Manpower costs for eng-he servicing are assumed to be $100/hour on the ground and $20,000fhour in space.

Plotting the costs of major engine overhauls as a function of the man-hours required to perform each overhaul clearly shows that performing major overhauls in space is very unlikely to be economical (Figure 5-6). Although amortization of nonrecurring costs of Space Station support facilities is a mejor cost factor, performing overhauls in space is over $15 million more expensive per overhaul even when Space Station facility costs are excluded. Performing overhauls in space is only economical if facility costs are excluded and manpower required is less than 200 man-hours to perform the overhaul in space.

-

CQ (with NR. cost Of c0 Station facilities

ortimodover 6 (every 25 missions)

Spak~ (with no NR costs of Space Station facilities included)

$1 5M

Task cost Cost of buying &

delivering new OlV $1OM engine to LEO

Baseline manpower requirement

I I I

I I ground - 0 250 500 750

500 1,000 1,500 space ZBllaEall.19

269.022.34

Man-hours to perform task 'NA - non-recurring

Figure 5-6. Task Costs For Major OTV Engine Overhauls

5- 5

P

GDC-SP- 84-050

For routine OTV engine maintenance tasks, servicing in space is shown to be the most economical option (Figure 5-7). Space manpower requirements and costs are much lower than for performance of major overhauls, and Space Station facilities are less expensive and amortized over a greater number of servicing events. Returning the engine to E a r t h for routine maintenance is cost-effective only if the time required to service the engine in space exceeds 56 rnanhours. Amortization of Space Station facility costs over fewer events reduces the sttrac- tiveness of space servicing, but even at low maintenance frequencies (every 20 missions) the space servicing option is more economical as long as manhours required do not exceed the baseline value (20 hours) by more than 50 percent In the baseline case, servicing in space is about $1.4 million less expensive per servicing event than ground servicing, or about $280,000 less expensive per OTV mission. The bcnefir of performing routine engine maintenance in space, calculated at $280,000 per OTV mission, could be considered partially or fuUy offset by payload weight penalties if any such penalties are incurred by designing the engine for space maintainability and if resultant pay1 capacity reductions are eonsidered to have an economic cost. With payload t/lb to GEO (assumed to be $4,000 - $10,000) used as a measure, the weight penalty costs of the space-maintainable engine beph to exceed the benefits of space maintenance when the wdght penalty reaches 28 to 70 pounds per mission, depending on the costllb to CEO used as R basis for calculation (Figure 5 - 8 ) . Calculation of the costs associated with payload weight penalties are somewhat subjective, since ev@n with very high manifesting efficiencies the QTV will probably have 500 or more pounds of excess (unused) capebility on 8 typical geosynchronous mission. If, for example, only ten percent of all OTV missions were affected. by a weight penalty in the range of Consideration, then costilb to GEO might be multiplied by 0.10 .before being used as a measure of weight pendty costs. With this methodology, weight penalties would need to be in the hundreds of pounds before their costs would approach the benefits of space maintainability.

i

5- 6

-- t

v) I- GROUND

GDC -SP- 84- 050

SPACE (WITH NR COST OF SPACE STATION FACILITIES AMORTIZED OVER 8 SERVICING TASKS=EVERY 20 MISSIONS) I

I

$ 2 I-

SPACE WITH NR COST CF SPACE STATION FACILITIES AMORTIZED OVER 38 SERVICING

1 T A S U EVERY FfVE MISSIONS) 'BASELINE CASE: SCRViClNG IN SPACE IS CHEAPER BY OVER $1.4M/TASK

I I I I I I 50 100 - 40 30 20

40 60 80 - - - 10

20 -

MAN-HOURS TO PERFORM TASKS 281147C447A

f69.GZ2.35

Figure 5-7. Task Costs For Routine OTV Engine Mlintenance

PAYLOAD CAPACITY VALGEO AT $4000/LB

VALUED AT $~O,OOO/LB

PAYLOAD WEIGHT PENALTY (Ib per mission) 2011eMl.16

269.02236

Figure 5- 8. Benefit of OTV Engine Space-Maintainability (Rclutine Maintenance) As A Function of Payload Weight Panalty Incurred).

5- 7

GDC-SP-84-050

SECTION 6

CONCLUSIONS

This study has shown (Table 6-1) that future OTV entfins requirements wi l l be determimed by LSS and manned missions. To satisfy tho mtlnnad reliability requirement, twin engines are needed. The optimum design thrust level is in the range of 3500 to 6500 lbf each, depending on the number of perigee burns for the manned mission. Although the lower thrust level is referable for less vehicle and payload design impact, this is contingent on the acceptance of multiple perigse burns for the manned mission and on the ability of the e n m e manufacturers to produce a high performance, reliable, and maintainable engine at lower thrust with additional starts and longer burn time. It is recommended that further research be planned.

Table 6-1. OTV Engine Study Findings

Manned GEO mission & LSS mission impose conflicting requirements for one engine design

Multiple perigee ascent burn trajectories offer optimal performance at low thrust levels needed for LSS missions

Multiple perigee ascent burns ( 3- 4) cm be performed without increasing manned module shielding weight above that required for stay at GEO

Optimum total thrust level for 13,000 l b m payload manned GEO mission is 6.000-7,000 Ibf (3 -4 perigee burns) vs. 13000 lbf (1 burn)

Optinium total thrust level for 10,000-20,000 lbm payfoad CSS mission is 1.000- 2 , 000 Ibf ( 8 perigee burns)

Redundant engines are required for propulsion system reliability needed for mission success & crew safety, and to reduce tests

High reliability engines will be demonstrated by testing 81 operational missions before manned requirement occurs

Backup (02-H2) APS (to a single main engine) results in ndditional propellant required due to decreased ISP. Q2-H2 APS may have logistics advantages

Recommended engine configuration is 2 main engines

6- 1 I

!

GDC-SP- 84-050

Table 6-1. OTV Engine Study Findings, Contd

e Optimum engine design thrust level is 3.000-3,500 lbf each (3-4 perigee burns) vs 6,500 lbf each (1 perigee burn)

e Long-life (100 missions) engine rwxnmended, but little economic benefit indicated for reducing frequency of major overhauls beyond one overhaul per 20-30 missions

I

e Major overhaul on ground 8 routine overhaul in space are recommended for Space-based OTV engine

o Further studies are recommended to evaluate multiple perigee burns for manned missions & to define high performance, relfable, maintainable engines at bwer thrust levels (3,000-3,500 lbf)

6- 2

GDC -S P - 84- 050

APPENDIX I

ENGINE PARAMETRIC DATA

Aerojet

Pratt & Whitney

Rocket dyne

NOTE: The OTV endne parametric data produced by Rockaell I nternntional are considered proprietary infarmation and are hence not included.

I

SDC-SP- 84-050

/

(I

Table A-1. Aerojet OTV Dual Propellant Expander Cycle Engine

NOTES 1, ISP IS VACUUM OELIVEAED AT 1POK BELL (SECONOS). 2. M19 = 6.0:l LOZILH~. 3. 1 = ENGINE LEMGTH (IM.1. 4. N - ENGIWEWEIGHT (LEM.). 5. 10 EWTHALPY PUMPIMG.

445 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0

MIXTURE RATIO (oM)

2CQ.022-37

r. I

3

269.022-38

Figure A-1. Aerojet OTV Dual Expander Cycle Engine Delivered ISP V s . Mixture Ratio

I- 2

I

GDC-SP- 84-050

Table A-2. Pratt & Whitney 3K Thrust Engine

DELHVEREB

1200 1200 1200 1200 1200 1200

I200 1500 1200 1500 1200 1500 1200 1'500 1200 1500 1200 1500

600 800

I500 2000 3000

iaoo

473,l 472. ? 492.0 468.4 463.1 450.9

Table A-3. Dratt 15 Whitney fiK Thrust Zngine

600 600 800 800

to00 1000 1500 1500 2000 2000 3000

3000

I- 3

73.6 475.8 413.2 476.0 472.5 475.6 468.9

473.3 463.7 469.4 451.5 460.1

201 208 216 236 255 294

E52 243 264 253 2 76 263

306 288 337 31 2 399 362

i

I

i.

GDC-SP- 84- 056

Table A-4. Pratt Br Whitney 10K Thrust Engine

(PS I )

1200

1506

1200

1500

I200

15QO

1200

1500

1208

1 so0

1200

1506

REA RATIO ~~ ~

600

600

800

800

1000

1000

1500

1500

2000

2000

3OOQ

3000

DELI VEREO

(SIX.) - 474.2

476.3

473.7

476.4

472.9

474.2

469.3

474.0

464.4

470 0

452 2

460.7

(LB.)

345

331

368

351

394

370

4 54

41 9

59 5

466

636

564

. . . .i'

I- 4

i

A T I 0

Table A-5. Pratt 8c Whitney 15K Thrust Engine

I200 1500

2000

1200

3500

2Q00

3 200

600

600

600

800

890

800

:ooo

474 0 5

476.5

478.6

474.2

476.7

479.2

413.5

1500 1000 476.5

1200

3500

ZOO0

1200

1 SOQ 2000

1200

1500

1000

1500

1580

1500

2000

2000

2000

3000

3000

479 5

469.8

474.0

478.2

464.8

470.3

476.0

452.6

461 . O

419

398

376

454

428

399

$91

4s7

42 1

$82

530

476

672

603

531

853

?43

2GOO 3000 469.4 642

I- 5 i

GDC -SP- 84- 050

APPENDIX 11

ENGINE COST DATA

( Aerofet)

I

I

I

1

I

1'' !

GD C - S P - 84- 0 50

Table 1-1. Cost Trade: Short vs Long Mission Life

Engine

Cost DeltP, Short Life (10 Missions) Long Life (100 Missions)

Recurring Recurring* Recurring Recurring Non- Non-

U s e oi' Hydrostatic Bearings Development and Qual Additional Fab Complexity, QC, Test

Health Monitoring - Failure Prevention Development and Qual Fab , Instrumentation, Computer

Health Monitoring - Life

Development and Qual Fab , Instrumentation,

Projection

Computer

Space-Replaceable Engine Development and Qual

Mission

Space- Replaceable Engine Design of Service Center Assembly of Service Center Space Operations (Engine Replacement)

Transfer of Engine f r o m Earth to LEO & Return*

5M

20 K

200 K

40 M

20 M

200 K i i.

I

100 M 250 M

800 K

400 K

*Per Engine

NOTE: Long life is baseline AEimjet design

i I4

GDC-SP-84-050

c

Table 1-2. Cost Trade: Low vs High Specific Impulse (Isp)

Cost Delta Low Isp

Non- Becutodng R e c u d Recurring Recurring”

Dual Pmpellant Expander Cycle

Development and Qual Fab, QC

Extendible Nozzle Development and Qual Fab, QC

or

Large Area Ratio Nozzle Development and Qual Fab, QC

Mission

Delivery of Addit io~i~

(12/480 - 2.5%) PlWp@llEUlt 600 M

- -~

15 M 200 K

10 M 50 K

2 M 25 x

1 M

*Per Bngiine

NOTE: Aerojet 3000 IbF engine design with 1200: 1 nozzle has projected l s p of 484 IbF sec/IbM and near maximum Isp. AdditionaI gains of 12 IbF sec/lbbM are doubtful. Therefore high Isp case is baseline.

11-2 %

c

I I

I I * I

Y.,. 1

I

i I

5

Orbital T r a n s f e r Vehicle dngine 30 November 1984 I n t e g r a t i o n Study

c

W. J. Ketchurn

General Dynamics - Convair D iv i s ion San Diego, C a l i f o r n i a

Nat iona l Aeroqaut ics h Space Adminis t ra t ion Lewis Research C e n t e r

,Cleveland, Ohio 44135 A_-

-- NASA-LeRC is ~~~~~~~~ s t u d i s s eo estwbafata t he ~~~~~~~~~~y base for an advanced 8 b i t a l transfer &Pehicle)s for ~ ~ ~ - ~ ~ ~ # ~ IOC. Engine c o n t r a c t o r ~~~~~t~~ by v e h i c l e c o n t r a c t o r s t o def ine t h e require cond i t ions , and operational design criteria for new L propuls ion e m s a p p l i c a b l e t o f u t u r e o rb i t t r a n s fer veh i c 1 es c t s of space bas ing , ma and low-G t r a n s f e hOn5 On propu n system dgsfzn r e q u i r e

This report presents t h e r e s u l t s of a s tudy conduct by GDC under contract t o Aerojet for SA-LeRC. The primary s tudy hasis was to determine what the OTV engine t h r u s t l s v s l should be, how many eng ines are requ i r ed on t h e OTV, and how t h e OTV e n g i n e should be designed. This was accomplished by e v a l u a t i n g planned OTV miss ions and concepts t l

de termine t h e requirements for t h e -0TQ propuls ion system, conduct ing tradeoffs and comparisons t o optimize OTV c a p a b i l i t y , and e v a l u a t i n g r e l i a b i l i t y and maintenance t o determine t h e recommended OTV eng ine des ign f o r f u t u r e development.

8 assess the

Advanced engine

O r b i t a l T r a n s f e r Vehicle Space Based

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I86-20493 · I86-20493 CSCE 21H 63/20 Uncias 08625 ... Several recent studies of dedicated OTVs fos LSS missions have shown the advantages of multfple perigee burns to minimize gravity