lox/hydrocarbon auxiliary propulsion system … (oms), an aft reaction control subsystem (arcs), and...
TRANSCRIPT
REPORT MDC E2548
COPY NO. _"";
(NASA-CR- I71656) iCX/HYDRCEA_ EC_ _.UXIII_F¥
_BOPULSICN S¥S_ S_UEY _i_al _e[c[t
(McDonnell-Douglas Astrona_ t[¢-c £c,) 166 p
_C AOS/M_ A01 C$C[ 21Hgncla_
G3/20 C___LT_C__
LOX/HYDROCARBONAuxiliary Propulsion System Study
Prepared By: G. F. Orton
T. D. MarkD. D. Weber
FINAL REPORTJ U L.Y 1982
SUBMITTED TO: NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
LYNDON B. JOHNSON SPACE CENTER, HOUSTON, TEXAS 7/058
UNDER CONTRACT NO. NAS9--16305
MCDONNELL DOUGLAS ASTRON,._.ITICS COMPANY-ST. LOUIS DIVISION
t Box 516, Saint Louis, Missouri63166 (314) 232.0232
t MCDONNELL DOUGL_j_,.--_-
CORPORATION
https://ntrs.nasa.gov/search.jsp?R=19830016280 2018-06-26T01:04:29+00:00Z
• .-4 _w -_
LOX/HYDROCARBON
Auxiliary Propulsion SysTem StudyFINAL REPORT
ABSTRACT
REPORT MDC E2.548
JULY 1982
This report describes a study to evaluate liquid oxygen (LOX)/hydrocarbon (HC)
propulsion concepts for a "second generation" Shuttle Orbiter auxiliary propulsion
system. The auxiliary propulsion system consists of an Orbital Maneuvering
Subsystem (OMS), an Aft Reaction Control Subsystem (ARCS), and a Forward Reaction
Control Subsystem (FRCS). The primary goals of this effort were to identify the most
attractive fuel and system design approach and to determine technology advancements
that are needed to provide high confidence for a subsequent system development. The
work was performed by the McDonnell Douglas Astronautics Company in St. Louis,
Missouri (MDAC-STL) for the NASA-Lyndon B. Johnson Space Center under contract
NAS9-16305. Aerojet liquid Rocket Company provided engine system data under asubcontract to MDAC-STL.
The study consisted of e Phase Z--Preliminary System Evaluation and a
Phase ll--In-Depth System Evaluation. The fuel candidates were ethanol, methane,
propane, and ammonia. Even though ammonia is not a hydrocarbon, it was included for
evaluation because it is clean burning and has a good technology base as a result of
its use with LOX in the X-15 rocket engine system. The major system design options
were pump versus pressure feed, cryogenic versus ambient temperature RCS propellant
feed, and the degree of OMS-RCS integration.
On the basis of the Phase I and Phase II evaluations, ethanol was determined to
be the best fuel candidate. It is an earth-storable fuel with a vapor pressure
slightly higher than monomethyl hydrazine. The LOX/ethanol propellant combination
does not produce free carbon contaminant in the engine exhaust gases and, because of
its high bulk density--specific impulse product, provides the most efficient
packaging and highest total impulse capability of all the propellants considered.
A pump-fed OMS was recommended because of its high specific impulse, enabling
greater velocity change (AV) and greater payload capability than a pressure-fed
system. Oxygen is fed to th_ OMS engine in a liquid state at cryogenic temperature,
and the OMS oxygen feedline is vented between burns. Common OMS/ARCS propellant
tanks were recommended to conserve weight, provide higher total impulse capability,
and provide increased mission flexibility.
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_4CDONNELL DOUGLA8 4STRONAUTICii COMPAN¥.BT. LOLII_ DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORTREPORT MDC E2548
JULY 1982
For the RCS, a hybrid feed system (liquid ethanol--and gaseous oxygen) was
recommended to preclude the requirement for RCS feed system insulation, ---The
recommended RCS feed system employs ambient temperature, blowdown accumulators for
supplying propellants to the thrusters. Propellants are fed to the accumulators
using small electric pumps which operate at low flowrates and low discharge
pressures. The energy to thermally condition the RCS oxygen flow to a gaseous state
is derived from a passive ethanol tank heat exchanger. The heat exchanger is a
tubular coil attached to the outsi.de surface of the ethanol tank. The electric pump
supplies liquid oxygen to the heat exchanger where the oxygen absorbs heat from the
tank wall, the liquid ethanol inside the tank, and the environment. The oxygen exits
the heat exchanger in a gaseous state and is then routed to the RCS accumulator.
This passive thermal conditioning approach is attractive because of its simplicity
(no active gas generator--heat exchanger assemblies) and high specific impulse (no
gas generator vent loss).
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MCDONNELL DOUGLAS ASTRONAUTICS COMPAftlY . BT LOUIS DIVIBIOItl
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Auxiliary Propulsion System Study
FINAL REPO RI:-
TABLE OF CONTENTS
REPORT MDC E2548
JULY 1982 B¥
1.0 Introduction ........................... -............................
2.0
3.0
Task
2.1
2.2
2.3
1.1--Phase I Groundrules ......................................
Propellant Candidates .........................................
System Design Options .........................................
Design Requirements ...........................................
Task
3.1
3.2
3.3
1.2--Phase I Component Characterization .......................
APSDS Code ....................................................
TKHEAT Code ...................................................
FDLINE Code ...................................................
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6
15
23
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32
41
4.0
5.0
Task
4.1
4.2
4.3
4.4
4.5
4.6
Task
1.3--Phase I System Evaluations ...............................
Pump Versus Pressure Feed .....................................
NBP Versus Subcooled Fuel Storage .............................
Cryogenic Versus Ambient Temperature RCS Propellant Feed ......
Common Versus Separate OMS/RCS Tanks ..........................
Tank Insulation Options .......................................
Feedline Insulation Options ...................................
ll.1--Phase II Groundrules ....................................
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60
65
68
79
6.0 Task
6.1
6.2
6.3
ll.2--Phase II Component Characterization .....................
APSDS Code ....................................................
TKHEAT Code ...................................................
FDLINE Code ...................................................
87
87
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97
7.0 Task
7.1
7.2
ll.3--Phase II System Evaluations .............................
Tank Insulation Evaluations ...................................
RCS Feedline Insulation Evaluations ...........................
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
REPORT MDC E2548
JULY 1982
TABLE OF CONTENTS (Continued)
Page
7.3 GOX/Ethanol RCS Feasibility Evaluations ....................... 109
7.4 LOX/Ethanol and LOX/Methane OMS-ARCS Sensitivity Analyses ..... 125
7.5 Separate Versus Common FRCS/Aft Propulsion Tanks ..... ; ........ 142
7.6 Conventional Versus Conical Propellant Tank Shapes ............ 142
7.7 Pump Versus Pressure Fed FRCS................................. 147
7.8 Side-by-Side OMS/ARCS Comparisons ............................. 147
8.0 Conclusions and Recommendations .................................... 149
References .............................................................. 153
LIST OF PAGES
Title Page
ii thru xiii
i thru 153
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MCDONNELL DOUGLAS AsTRONAUTICs COMPANY.IT'. LOUIS DIVI.'qlON
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
LIST OF FIGURES
REPORT MDCF.254_
JULY1982.
NUMBER
4
5
6
7
8
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10
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14
15
16
17
18
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24
25
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28
TITLE
Current Orbiter Aft Pod OMS/RCS Packaging Arrangement .........
Pump and Pressure Fed System Concepts .........................
Cryogenic and Ambient Temperature Propellant Feed System
Concepts ......................................................
Interconnected Forward and Aft Propulsion Systems .............
APSDS Flow Diagram for Fixed Volume Analysis ..................
Cryogenic Propellant Tank Model ...............................
RCS Accumulator Operation .....................................
Cryogenic Propellant Feedline Model ...........................
Thermodynamic Vent System Concept .............................
Tank Thermal Model ............................................
TKHEAT Flow Diagram ...........................................
Comparison of TKHEAT Code with Test Data ......................
Example TKHEAT Computations for OMS LOX Tank ..................
Feedline Thermal Model ........................................
FDLINE Flow Diagram ...........................................
Example FDLINE Computations for RCS LOX Feedline ..............
Pressure Fed OMS Schematic ....................................
Pump Fed OMS Schematic, .......................................
Pump FED OMS PC Sensitivity ...................................
Pump and Pressure Fed OMS Comparisons .........................
Pressure Fed RCS Schematic ................... , ................
Pump Fed RCS Schematic ........................................
Pump and Pressure Fed RCS Comparisons .........................
RCS PC Sensitivity (Pump and Pressure Fed Systems) ............
Pod Layout Sketch for Subcooled Propane Storage ...............
OMS AV Comparison for NBP and Subcooled Propane Storage .......
Ambient Propellant Temperature Feed RCS Schematic (Gaseous
02/Liquid Fuel) ...............................................
Ambient Propellant Temperature Feed RCS Schematic (Gaseous
02/Gaseous Fuel) ..............................................
PAGE
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12
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MCDONNELL DOUGLAS ASTRONAUTICS COMPANY -_IT. LOUIS DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
LIST OF FIGURES (Continued)
REPORT MDC E2548
JULY 1982
NUMBER
29
30
31
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4O
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TITLE
RCS PC Sensitivity (Ambient Propellant Temperature Feed) ......
Comparison of Cryogenic and Ambient Temperature Propellant
Feed RCS......................................................
Schematic for OMS-Aft RCS with Common Propellant Tanks ........
Comparison of OMS-Aft RCS with Common and Separate
Propellant Tanks ..............................................
Comparison of Separate and Interconnected Forward and Aft
Propulsion Systems ............................................
OMS LOX Tank Insulation Comparisons ...........................
RCS LOX Tank Insulation Comparisons ...........................
Common OMS-RCS LOX Tank Insulation Comparisons ................
RCS Feedline Insulation Performance with Current Manifold
Arrangement ...................................................
TG-15000 Insulation Performance with Re-Configured Manifold
Arrangement ...................................................
Effect of LOX Usage Rate on Feedline Temperature Response
(TG-15000 Insulation) .........................................
Effect of LOX Usage Rate on Feedline Temperature Response
(MLI) .........................................................
Passive Ethanol Feedline 02 Heat Exchanger Concept ............
Passive Ethanol Tank 02 Heat Exchanger Concept ................
LOX/Ethanol OMS Engine Specific Impulse .......................
LOX/Ethanol RCS Engine Specific Impulse .......................
LOX/Methane RCS Engine Specific Impulse .......................
LOX Electric Motor and Pump Weight ............................
Methane Electric Motor and Pump Weight ........................
Ethanol Electric Motor and Pump Weight ........................
Battery Weights for Electric RCS Pumps ........................
Ethanol Tank 02 Heat Exchanger Model ..........................
RCS Thruster Heat Soakback Model ..............................
30-Day Pressure Profile for Common OMS-ARCS Tank (Methane) ....
30-Day Temperature History for Common OMS-ARCS Tank (Methane).
PAGE
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64
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LOX/HYDROCARBON
Auxiliary Propulsion SystEm Study
FINAL REPORT
L.IST OF FIGURES (Continued)
REPORT MDC E2548
JULY 1982 Bv'
NUMBER
54
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TITLE
Methane Vent Losses for Common OMS-ARCS Tank ..................
Methane Vent Losses for FRCS Tank .............................
LOX Vent Losses for Common OMS-ARCS Tank ......................
LOX Vent Losses for FRCS Tank .................................
Methane Temperature Response in RCS Feedline (TG-15000
Insulation) ...................................................
Effect of Methane Usage Rate on Feedline Temperature Response
(TG-15000 Insulation) .........................................
Effect of Methane Usage Rate on Feedline Temperature
Response (MLI) ................................................
LOX Temperature Response in RCS Feedline (TG-15000 Insulation)
Effect of LOX Usage Rate on Feedline Temperature Response
(TG-15000 Insulation) .........................................
Effect of LO× Usage Rate on Feedline Temperature Response
(MLI) .........................................................
LOX/Ethanol OMS-ARCS Using OMS Turbopumps for RCS Supply ......
RCS Ethanol Accumulator Mission Response (OMS Turbopump
Resupply) .....................................................
RCS GOX Accumulator Mission Response (OMS Turbopump Resupply),
LOX/Ethanol OMS-ARCS Using Electric Motor Pumps for RCS Supply
02 Temperature Histories for Passive Ethanol Tank Heat
Exchanger .....................................................
Ethanol Temperature And Quantity Histories for Passive
Ethanol Tank Heat Exchanger ...................................
RCS Ethanol Accumulator Mission Response (Electric Pump
Resupply) .....................................................
RCS GOX Accumulator Mission Response (Electric Pump Resupply),
Baseline LOX/Ethanol OMS-ARCS.................................
Baseline LOX/Methane OMS-ARCS.................................
Sensitivity of LOX/Ethanol System to OMS Chamber Pressure .....
Sensitivity of LOX/Ethanol and LOX/Methane Systems to
RCS Chamber Pressure ..........................................
PAGE
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MCDOItilblELL DOLIGLAB ASTRO&I4LJTIC8 COIt4PAN_-BT. LOdJIS DIVISIOI_I
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LOX/HYDROCARBON
Auxiliary PRopulsion System study
FINAL REPORTREPORT MDC E254_
JULY 1982
LIST OF FIGURES (Continued)
NUMBER
76
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8O
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9O
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TITLE
RCS GOX Accumulator Blowdown Pressure Ratio Sensitivity .......
RCS Liquid Ethanol Accumulator Blowdown Pressure Ratio
Sensitivity ...................................................
RCS LOX/Methane Accumulator Blowdown Pressure Ratio
Sensitivity ...................................................
LOX/Ethanol OMS Isp Sensitivity (Pod AV = 500 ft/sec) .........
LOX/Etharol OMS Isp Sensitivity (Fixed Pod Volume) ............
LOX/Methane OMS Isp Sensitivity (Pod AV : 500 ft/sec) .........
LOX/Methane OMS Isp Sensitivity (Fixed Pod Volume) ............
LOX/Ethanol RCS Isp Sensitivity (Pod _V = 500 ft/scc) .........
LOX/Ethanol RCS Isp Sensitivity (Fixed Pod Volume) ............
LOX/Methane RCS Isp Sensitivity (Pod aV = 500 ft/sec) .........
LOX/Methane RCS Isp Sensitivity (Fixed Pod Volume) ............
LOX/Ethanol Tank Minimum Gage Sensitivity .....................
LOX/Methane Tank Minimum Gage Sensitivity .....................
Separate Versus Common FRCS/Aft Propulsion Tanks (LOX/Ethanol)
Conventional Versus Conical Propellant Tank Shapes
(LOX/Ethanol) .................................................
Pump Versus Pressure Fed FRCS (LOX/Methane) ...................
Comparison of LOX/HC, LOX/H2, and N204/MMH OMS-ARCS ...........
PAGE
131 --
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Auxiliary Propulsion System Study
FINAL REPORT
LIST OF TABLES
REPORT MDC E2548
JULY 1982
Table No.
I Candid.a.te Fuel
II LOX/HC OMS-RCS
III Options
IV Options
V Phase I
Vl
Vll
Vlll
IX
X
Xl
Xll
Xlll
XIV
XV
XVl
XVll
XVIII
Matrix. ..................... ;...-;.. ........
Design Options.-. ...........................
Selected for Phase I Evaluation .............. ,.,..
Deleted from Phase I Evaluation ...................
System Evaluation Matrix ..........................
Design Requirements/Constraints for Component Weight andSizing ....................................................
System Components in APSDS Computer Code ..................
Example System Weight Summary from APSDS Computer Code ....
Example APSDS Output--Propellant Tanks ....................
Example APSDS Output--RCS Accumulators ....................
Example APSDS Output--RCS LOX Feedline ....................
Properties of Candidate Insulation Materials ..............
Phase II Groundrules Summary ..............................
Phase II Design Requirements/Constraints for ComponentWeight and Sizing .........................................
Phase iI System Evaluation Tasks ..........................
Overall Study Conclusions ................................
New Technology Requirements ...............................
Recommendations for Further Feed System Effort ............
Pa__
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69
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100
150
151
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
NOMENCLATURE
AE/AT
A1
Ag
ALRC
APS
APSDS
ARCS
BTU
CH4
Cu
C2H50H
C3H8
D
F
FDLINE
FIB
FRCS
FSYS
ft
ft 3
GG
HC
He
HR
H2
in.
ISp
It
JSC
L
Ib
Ibf
Ibm
nozzle area ratio
aluminum
silver
Aerojet Liquid Rocket Company
auxiliary propulsion system or aft
advanced propulsion system design
Aft Reaction Control Subsystem
British Thermal Unit
methane
copper
ethanol (ethyl alcohol)
propane
diameter
thrust
feedline heat transfer computer code
fiberous insulation (TG-15000)
Forward Reaction Control Subsystem
system thrust, Ibf
feet
cubic feet
gas generator
hydrocarbon
helium
hours
hydrogen
inches
speclfic impulse,
total impulse,
Johnson Space
length
pounds
pounds-force
pounds-mass
Ibf-sec/Ibm
Ibf-sec
Center
propulsion system
and sizing computer code
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REPORT MDC F_2548.
JULY 1982
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Auxiliary Propulsion System Study
FINAL REPORT
REPORT MDC E2548
JULY 1982
NOMENCLATURE (Continued)
LH2
Li
liq
LOX
max
MDAC-STL
min
MLI
MMH
MR
N
NASA
NBP
NH3
Ni
NPSP
N2
N2H4
N204
O/F
OME
OMS
OX
O2
P
PC
psi
psia
RCE
RCS
RP-1
liquid hydrogen
lithium
liquid
liquid oxygen
mass flow rate
maximum
McDonnell Douglas Astronautics Company - St. Louis
minimum
multi-layer insulation
monomethyl hydrazine
mixture ratio by mass (oxidizer-to-fuel)
number of line segments
National Aeronautics and Space Administration
normal boiling point
ammonia
nickel
net positive suction pressure
nitrogen
hydrazine
nitrogen tetroxide
mixture ratio by mass (oxidizer-to-fuel)
Orbital Maneuvering Engine
Oribtal Maneuvering Subsystem
oxygen
oxygen
pressure
chamber pressure
pounds per square inch
pounds per square inch - absolute
heat transfer rate
Reaction Control Engine
Reaction Control Subsystem
rocket propellant-I
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MCDONNELL DOUGLAS ASTRONAUTICS COMPANY - ST. LOiLII8 DIVISION
LOX/HYDROCARBON
Auxiliary ProPulsion System Study
FINAL REPORT
NOMENCLATURE (continued)
R/V
OR
sec
SS
T
TG- 15000
Ti
TKHEAT
TPA
TVS
V
XFEED
Zr
Zn
&P
AV
%
relief valve
degrees Rankine
seconds
stainless steel
temperature
silica fiber insulation
titanium
tank heat transfer
turbopump assembly
thermodynamic vent system
volume
crossfeed
zirconium
zinc
incremental length
pressure drop
velocity change
nozzle area ratio
percent
computer code
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MCDONNELL DOUGLAS ASTRONAUTICS COMPANV.ST. LOUIS DIVISION
REPORT MDC F..2548
JULY 1982
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Auxiliary Propulsion System StudyFINAL REPORT
1.0 INTRODUCTION
REPORT MDC F.2548
JULY 1982
BY
During the last two decades, spacecraft propulsion systems have employed
simple pressure fed systems using earth-storable propellants such as nitrogen
tetroxide (N204) and monomethyl hydrazine (MMH). These systems have been reliable
and have afforded low development risk. However, their disadvantages are that the
propellants are highly toxic and corrosive and impose high operational costs for
reusable applications such as the Space Shuttle Orbiter. Furthermore, MMH is a
possible carcinogen and is expensive to produce.
Over the years numerous studies have considered the use of LOX/H 2 for
spacecraft auxiliary propulsion systems. However, two inherent characteristics of
liquid H2--a low density and a very low storage temperature--impose severe penalties
on a reusable system such as the Shuttle Orbiter in the form of additional spacecraft
volume and weight.
Liquid oxygen/hydrocarbon (LOX/HC) propellants possess many of the desirable
characteristics of the LOX/H 2 combination while avoiding its disadvantages. They
are low in toxicity, non-corrosive, low in cost and can be vented or purged from the
system to facilitate system maintenance. The hydrocarbon fuels also have a high
density compared to liquid H2 which allows much lower fuel tank volumes. During
evolution of the Shuttle design in the early 1970's LOX/HC propellants were
considered for the Orbiter OMS/RCS. Even though they offered operational advantages
over N204/MMH , they were not selected because they lacked the necessary technology
base to support the development schedule and development cost criteria for the
Orbiter. However, to achieve the ultimate Shuttle goal of economic, aircraft-like
operations, it will be necessary to replace the toxic and corrosive N204/MMH
propellants with a more passive LOX/HC propellant combination.
To begin building a technology base for LOX/HC engines NASA-JSC sponsored two
previous research and development efforts: Photographic Combustion Characteriza-
tion of LOX/HC Type Propellants (NAS9-15724) and Combustion Performance and Heat
Transfer Characterization of LOX/HC Type Propellants (NAS9-15958). These efforts
were a first step in addressing engine technology deficiencies.
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LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT REPORT MDC E2548
JULY 1982
The purpose-of this study was to provide a corresponding technology evaluation
for the overall system. The general study approach was to compare LOX/HC propulsion
systems applicable to a second generation Orbiter OMS/RCS and to evaluate majorsystem/component options.
The technical effort for this study was conducted in two phases. Phase I was
a preliminary evaluation to screen a large number of propellant combinations and
system concepts. Phase II was an in-depth evaluation of the most promising
propellants and system concepts resulting from Phase I. Both study phases were
divided into three major tasks. Task I defined the groundrules in terms of candidate
propellants, system/component design options, and design requirements. In Task II,
system and engine component math medels were incorporated into existing computer
codes for system evaluations. Aerojet Liquid Rocket Company (ALRC), under a
subcontract to MDAC-STL, provided characterization data for both the OMS and RCS
engines. Finally, in Task III, the detailed system evaluations and comparisons were
performed to identify the recommended propellant combination and system approach.
The detailed data dump reports for Phase I and Phase II were provided in
References (I) and (2), respectively. This report provides a summary description of
all technical _ffort conducted during the study.
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Auxiliary Propulsion System StudyFINAL REPORT
2.0 TASK I.I--PHASE I GROUNDRULES
REPORT MDC F..2548
JULY 1982
The overall study approach was to use the Space Shuttle Orbiter OMS and RCS
requirements as a framework for comparing alternate LOX/HC propulsion system
concepts. The current Orbiter aft propulsion subsystem pod is shown i_ Figure I.
Each pod contains OMS/ARCS propellant and pressurant tankage, propellant distri-
bution -networks, a 6000 Ib-thrust OMS engine, twelve 870 Ib-thrust primary RCS
thrusters, and two 25 Ib-thrust vernier RCS thrusters. The propellants are N204and MMH.
The OMS and ARCS are designed to operate independently, but are equipped with
interconnecting plumbing to allow OMS propellant tanks in either pod to supply
propellants to the OMS engines or ARCS thrusters in both pods. ARCS propellant
tanks in either pod can also supply propellants to ARCS thrusters in both pods. A
FRCS module, which is similar in design to the ARCS, is installed in the nose of theOrbiter.
Because of the large number of possible LOX/HC propulsion system alternatives
for the OMS and RCS the major challenge of the Task 1.1 groundrules effort was to
limit the number of system/propellant concepts to a manageable level. To
accomplish this effort Task I.i was divided into three primary areas:
• definition of propellant candidates
• definition of system/component design options
• definition of system design requirements and constraints.
These are discussed in the following paragraphs.
2.1 Propellant Candidates
The candidate propellant combinations selected for the study were:
• oxygen/ethanol (02/C2H50H)
• oxygen/propane (02/C3H8)
• oxygen/ammonia (02/NH3)
• oxygen/methane (02/CH4).
As shown in Table I, the candidate fuels r_present each of the major propellant
classes. Ethanol (ethyl alcohol) represents the earth storable propellant class
3
MCDONNELL DOUGLAS ASTRONAUTIC8 COMPAN'r .|T. LOUIB DIVIBION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
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MCDONNELL DOUGLAS ASTRONAUTICS COMPANY ST. LOUIS DIVISION
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JULY 1982
OR}GIN _,L P,"_'...... :_-"OF PO'.':;;-'; ""' ;" ...._ ...,,.,.,_! :"y
TABLE ICANDIDATE FUEL MATRIX
EARTH STORABLE (BOILING POINTS MUCH GREATER THAN AMBIENT)
EXAMPLES: RP-1 FUELS SELECTED FOR PHASE IETHANOL
HEPTANE ETHANOL (C2H5OH)BENZENE
METHANOL
n-OCTANE
.SPACE STORABLE (BOILING POINTS SLIGHTLY LESS THAN AMBIENT)
EXAMPLES: PROPANE
B UTANE PROPANE (C3H8)
ISOBUTANE AMMONIA (NH3}PROPYLENEAMMONIA
.CRYOGENIC (BOILING POINTS LESS THAN - 100OF)
EXAMPLES: ETHANE METHANE (CH4)METHANE
ETHYLENE
CYCLOPENTANE
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LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT ' REPORT MDC E2548
JULY 1982
because it is non-coking, has a good technology base for engin_ development (was -
used in the original X-15 engine system), and has an acceptably high vapor
pressure. (The vapor pressure-of ethanol is slightly greater than MMH.) RP-I was
not a candidate because it produces excessive free carbon in the combustion process
and does not possess good restart characteristics for a regeneratively cooled OMS
engine due to its low vapor pressure. Propane and ammonia represented the space
storable propellant class because they were being tested under engine technology
efforts sponsored by NASA-JSC (NAS9-15724 and NAS9-15958). Even though ammonia is
not a hydrocarbon, it was included because it is clean burning (no contaminating
carbon compounds in the exhaust products) and was used with LOX in the uprated X-15
rocket engine system. The final fuel candidate, methane, represents the cryogenic
storage class because it is non-coking and was also being tested unde_ NASA-JSC
engine technology contracts (NAS9-15724 and NAS9-15958).
2.2 System Design Options
A list of major system and component design options applicable to LOX/HC
propulsion systems is presented in Table II. In order to limit the number of
options to be evaluated, only the key elements (system, tankage, and feedline)
listed in Table III were selected for evaluation in Phase I. (The rationale for
deleting design options from the Phase I evaluations is presented in Table IV.) The
following paragraphs describe 2ach Phase I design option of Table III and provide
the rationale for the Phase I system evaluation matrix.
2.2.1 Pump Versus Pressur_ Feed - The primary issue associated with this option is
the development complexity associated with a turbopump feed system versus the
heavier system weight associated with a helium pressure fed system. Simplified
schematics illustrating pump and pressure fed systems concepts are shown in
Figure 2. The pressure fed concept is similar to that e_loyed in the current
OMS-RCS with the exception that the helium bottle is stored inside the LOX tank to
conserve bottle volume and weight and minimize LOX heating during propellant tank
pressurization. Net positive suction pressure (NPSP) for the pump fed concept is
provided by a small helium pressurization system. Since helium pressurization
weight is a function of propellant temperature, this option was evaluated for all
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TABLE II
LOX/HC OMS-RCS DESIGN OPTIONS
REPORT MDC E2548
JULY 1982
• Overall system options
- pump versus pressure feed
- cryogenic versus ambient-temperature propellant feed
- common versus separate OMS/RCS tanks
- helium versus boost pump NPSP
- NBP versus subcooled propellant storage
- propulsive versus non-propulsive gas generator vents
- subcritical versus supercritical propellant storage
Pressurization assembly options
- ambient versus LOX stored helium tank
- separate versus common helium supply for fuel and oxidizer tanks
hydraulic versus electric boost pumps
Propellant tankage options
- insulation options
- conventional versus non-conventional tank snape
- conventional versus thermodynamic tank vent (cryogenic tanks)
- propellant acquisition options
- propellant gaging options
- internal versus external entry propellant sumps (common OMS/aft RCS tanks)
Propellant feedline options
insulation options
- separated versus thermally shorted fuel and oxidizer lines
• Accumulator options
- blowdown versus helium pressure regulated liquid accumulators
Engine conditioner assembly options
- electric motor versus turbine pump drive
gas generator versus engine expander cycle turbine drive
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TABLE III ....
OPTIONS SELECTED FOR PHASE I EVALUATION
• pump versus pressure feed
• NBP versus subcooled fuel storage
• c._yogenic versus ambient temperature propellant feed
• common versus separate OMS/RCS tanks
• propellant tank insulation Jptions
• feedline insulation options
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TABLE _V
OPTIONS DELETED FROM PHASE I E_ALUATION
REPORT MDC E2548
JULY 198;2
Helium versus boost pump NPSP -
This option was initially deferred for Phase II evaluation. HeliumNPSP was
baselined for Phase I. (Subsequent evaluations prior to the start of
Phase II showed that helium NPSP was desirable for implementing an overboard
propellant dump in the event of an abort and for providing propellant
crossfeed between pods in the event of a turbopump failure. As such, helium
NPSP was baselined for Phase II, and this option was ultimately deleted.)
Propulsive versus non-propulsive gas generator vents -
Propulsive vents were baselined for the OMS to maximize overall system
specific impulse. Non-propulsive vents were baselined for the RCS to
preclude translational thrust during attitude control.
Subcritical versus supercritical propellant storage -
Previous Space Shuttle Auxiliary Propulsion System studies (NAS8-26248)
have shown that tank weight penalties are excessive for supercritical
propellant storage. Since the critical pressures for all the propellants of
this study are greater than 600 psia, subcritical storage assemblies were
baselined.
Ambient versus LOX stored helium tank -
LOX stored helium tanks were baselined for this study to minimize helium
tank volume and minimize the heat input to the cryogenic tanks during
pressurization.
Separate versus common helium supply for fuel and oxidizer tanks -
Since the propellant candidates considered in this study are not hyper-
golic, propellant vapor mixing upstream of the check valves in the helium
pressurization sytem will not form reaction products. Therefore, a common
helium supply was baselined to ensure accurate mixture ratio control and
minimize helium pressurization system weight. Anti-migration screens would
be provided in the propellant tank helium inlet diffusers to prevent liquid
migration into the helium pressurization lines.
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MCDONNELL DOUGLAS ASTRONAUTICB COMPANY.S_ LOLIIB DIVIJlON
LOX/HYDROCARBON
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FINAL REPORT
TABLE IV (Continued)
OPTIONS DELETED FROM PHASE I EVALUATION
REPORT MDC F.2.548
JULY 1982
• Hydraulic versus electric boost pumps -
This option was tentatively deferred for Phase II evaluation. (Subsequent
evaluations prior to the start of Phase II showed that helium NPSP was
desirable for implementing an overboard propellant dump in the event of an
abort and for providing propellant crossfeed between pods in the event of a
turbopump failure. As such, helium NPSP was baselined for Phase II, and
this option was deleted.)
Conventional versus non-conventional tank shape -
Evaluation of this option was deferred for Phase II evaluation.
tional tank shapes were baselined for Phase I.
Conven-
Conventional versus thermodynamic tank vent (cryogenic tanks)
Conventional vents in which vapor is vented from the propellant tank were
considered impractical due to the difficulty associated with positioning
the vapor bubble inside the tank in low-g. As such, a thermodynamic vent
system was baselined in which liquid is withdrawn from the tank through the
propellant acquisition system to relieve tank pressure.
Propellant acquisition system options -
Because of the current technology status of the OMS and RCS tanks and the
requirement for system reuse, surface tension screen propellant acquisition
systems were baselined for this study. The detailed design of a surface
tension screen system for cryogenic propellants is a key issue that should
be addressed in future technology efforts.
Propellant gaging options -
Because of their current technology status, a capacitance gaging system was
baselined for the OMS and a pressure-temperature-volume measurement system
was baselined for the RCS. The propellant gaging system is also a key issue
that should be addressed in future technology efforts.
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MCDOItlNEI.I,. OOUGLA8 ASTROItlAUTIC8 COMPANY-ST. LOUIII DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
TABLE IV (Continued)
OPTIONS DELETED FROM PHASE I EVALUATION
REPORT MDC E2548
JULY 1982
Internal versus external entry propellant sumps (common OMS/aft RCS tanks) -
Propellant acquisition systems for common OMS/aft RCS tanks were evaluated
previously by MDAC-STL under a contract with Rockwell International during
the Orbiter design phase, it was concluded in that evaluation that
external, in-line entry propellant sumps were superior to internal sumps.
The external sump is easier to service and check-out and remains full of
propellant during the launch and orbital mission phases. Because of these
advantages the external sump was baselined for common OMS/aft RCS tanks.
Separate versus thermally shorted fuel and oxidizer lines
Evaluation of this option was deferred for Phase II evaluation.
feedlines were baselined for Phase I.Separate
Blowdown versus helium pressure regulated liquid accumulators -
Previous auxiliary propulsion system studies conducted under Contract
NAS9-12013 have shown that blowdown liquid accumulators are superior to
helium pressure regulated accumulators. They are lower in weight, afford
lower design, development, and operational complexity and are less costly.
Because of these advantages, blowdown liquid accumulators were baselined
for this study.
Electric motor versus turbine pump drive -
Evaluation of this option was deferred for Phase II evaluation.
pump drives were baselined for Phase I.
Turbine
Gas generator versus engine expander cycle turbine drive
Evaluation of this option was also deferred for Phase II evaluation.
generator turbine cycles were baselined for Phase I.
Gas
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MCDONNELL DOUGLAS ASTRONAUTICS COMPANY.ST. LOUIS DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINALRePORt REPORT MDC E2548
JULY 1982
PUMP FED
TANK
- BURST DISC
__RELIEF VALVE
CHECK VALVE
HELIUM•BOTTLEC_- I)PRESSURE
___ _O_ATO_TURBOPUMP I_-_ ,,_ GAS
_'- L] GENERATOR
(,.) i :'LACCUMULATOR
ENGINE
PRESSURE FED
'_______
FIGURE 2 PUMP AND PRESSURE FED SYSTEM CONCEPTS
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four fuel candidates. Furthermore, since helium pressurization system weight is
also a function of the total impulse requirement, this option was evaluated for
both the OMS and RCS.
2.2.2 Normal Boilin 9 Point (NBP) Versus Subcooled Fuel Storage- The primary issue
associated with this option is the increased system total impulse capability with
subcooled fuel storage (due to increased fuel dansity) versus the lower thermal
control complexity associated with normal boiling point storage. Because of the
large density variation between its normal boiling point and freezing tempera-
tures, this option was evaluated for propane only. Since the freezing temperature
of propane is less than the normal boiling temperature of oxygen propane can be
stored at LOX temperatures with an attendant density increase of 25 percent. The
corresponding total impulse benefit is significant in large volume systems such as
the OMS, and, as such, this option was evaluated for the OMS only.
2.2.3 Cryogenic Versus Ambient Temperature RCS Propellant Feed - The primary issue
associated with this option is the feedline insulation complexity with cryogenic
propellant delivery systems versus the thermal conditioning energy penalty
(specific impulse loss) associated with ambient temperature, gaseous propellant
feed. Simplified schematics illustrating these system concepts are shown in
Figure 3. In the cryogenic feed system, the propellant is delivered to the
thrusters as a liquid at temperatures near tSenormal boiling point. In the ambient
temperature feeG system, the cryogenic propellants are thermally conditioned to a
gaseous state in a heat exchanger supplied with gas generator exhaust products.
The advantage of the ambient temperature feed system is the elimination of
insulation on the accumulators and propellant feedlines.
This design option is only applicable to the RCS feed system. OMS feedline
lengths are very short compared to the RCS and only a limited number of engine
firings are required per mission. As such, the cryogenic propellants were assumed
to be vented from the OMS feedlines between engine burns. The RCS feedlines are
very long, however, and there are a large number of thruster firings required for
attitude control. Because of this it is not practical to vent the feedlines between
firings. Therefore, insulated feedlines are required to prevent propellant
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MC_ONNEI.L _OUGI,.4S ASTROIti,4LITICS COMPANY.ST. LOUIS _IV#S#Oltl
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Auxiliary Propulsion System Study
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• SEE SCHEMATIC
REPORT MDC E2548
JULY 1982 i
OF POOR Q'jk, Li',"';
tNOMENCLATURE OF FIGURE 2
AMBIENT FEED I
',CRYGGEtlIC FEED
])
HEATEXCHANGER
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FIGURE 3 CRYOGEN!C AND AMBIENT TEMPERATURE PROPELLANTFEED SYSTEM CONCEPTS
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vaporization during idle periods, otherwise the cryogenic propellants must be
thermally conditioned to a gaseous state for thruster feed. Since thermal
condit,oning energy requirements depend on-the propellant type and thruster
mixture ratio, this option was evaluated for all four fuel candidates. In the
LOX/methane RCS both propellants are thermally conditioned-to a gaseous state.
However, in the LOX/ethanol, LOX/propane, and LOX/ammonia systems only the LOX is
thermally conditioned to a gaseous state. This is because the fuels are stored as
liquids at near ambient temperatures.
2.2.4 Common Versus Separate OMS/RCS Tanks - As mentioned previously the current
aft propulsion pods are interconnected to allow OMS tanks in either pod to supply
OMS engines and RCS thrusters in both pods and RCS tanks in either pod to supply RCS
thrusters in both pods. To further enhance propellant utilization flexibility
common tanks can be employed for OMS and ARCS propellant_, as well as for OMS, ARCS,
and FRCSpropellants. In the latter approach the FRCS would be interconnected with
the OMS and ARCS by feedlines routed along the length of the Orbiter as illustrated
in Figure 4. Because of varying performance capabilities, integrated OMS-RCS
tankage options were evaluated for all four fuel candidates.
2.2.5 Tank and Feedline Insulation Ootions - Alternate insulation materials for
the OMS and RCS tanks and RCS feedlines were selected for investigation to complete
the Phase I evaluation matrix. The candidate insulation materials were aluminized
mylar multi-layer insulation (MLI), TG-15000 silica fiber insulation, which is
currently employed on the aft pod internal moldline, and polystyrene foam
insulation. Because of the low boiling point and low heat of vaporization for LOX,
insulation options were evaluated for the LOX tanks and feedlines only.
2.2.6 Phase I System Evaluation Matrix - Based on the preceeding discussion the
Phase I system evaluation matrix of Table V was established. To complete the
Task I.i groundrules definition task system design requirements were established
as described below.
2.3 Design Requirements
Requirements employed for the Phase I system evaluations were divided into
mission, envelope, reliability, and component weight and sizing categories.
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LOX/HYDROCARBON
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" ' .,to ' . . .
..":..L..,.....•...................i............:........•/.,....................A I
FORWARD RCS
PROPELLANT CONNECTORS
!FIGURE 4 INTERCONNECTED FORWARD AND AFT PROPULSION SYSTEMS
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OF POOR QUALITY
TABLE V
PHASE I SYSTEM EVALUATION MATRIX
CANDIDATE FUELS
DESIGN OPTIONS ETHANOL PROPANE AMMONIA METHANE
I'
PUMP VERSUS PRESSURE FEED (OMS AND RCS)
COMMON VERSUS SEPARATE OMS/RCS TANKAGE
CRYOGENIC VERSUS AMBIENT TEMPERATURE RCS PROPELLANT FEED
NBP VERSUS SUBCOOLED FUEL STORAGE (OMS)
TANK INSULATION OPTIONS
FEEDLINE INSULATION OPTIONS
PLOX
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LOX/HYDROCARBON
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JULY 19_
Generic OMS and RCS mission duty cycles consisting of engine and thruster on/off
timeswereprovided by the MDAC-STL APS Project. These duty cycles were originally
developed by NASA-JSC and were employed for-the APS static firing t_sts at
NASA-White Sands Test Facility. In this study they were used to perform tank and_
feedline thermal analyses. For comparing &V and total impulse capab.ilities of the
candidate propellants and system concepts the forward RCS module and aft pod
envelopes were constrained to the current dimensions. In addition OMS engine and
RCS thruster lengths and diameters were constrained to the current values. Feed
system schematics were prepared for each system concept to reflect the same "fail
operational/fail safe" component redundancy as the current OMS and RCS. The
detailed requirements and constraints employed for Phase I component weight and
sizing are summarized in Table VI.
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TABLE VI
DESIGN REQLIIREMENTS/CONSTRAINTS FOR COMPONENTWEIGHT AND SIZING
Helium pressurization system
Common helium supply for fuel and oxidizer tanks
- Current OMS/RCS line lengths
- Line Mach number = 0.1 (maximum)
- Real gas effects
- Solubility effects
- Vapor pressure effects
- Line Materials: 2219-T87 A1 or 304L SS
- Polytropic exponent : 1.0 (helium bottle inside LOX tank)
- Regulator pressure ratio = 0.7 (outlet/minimum inlet)
- Tank shape: spherical
- Tank Materials: LOX storage--2219-T87 A1
- Storage pressure: 3000 psia
- Ultimate factor of safety = 1.5
Propellant tanks
- Propellant dump through OMS and RCS engines
- Tank volume determination:
• impulsive propellant volume
• 2% liquid residuals by _olume
• 98% vapor residuals by volume
• tank boil-off loss
• OMS line chilldown/vent loss
• 5% ullage volume at storage temperature
- Shape:
• OMS: cylindrical with oblate spheriod end domes
• RCS and entry sump: spherical
• Common OMS/RCS: cylindrical with oblate spheriod end domes
• OMS and Common OMS/RCS fuel and oxidizer tanks are constrainted to
equal lengths to permit attachment to common aft pod bulkhead.
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LOX/HYDROCARBON
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FINAL REPORT
REPORT MDC E254,8
JULY 1982
TABLE Vl (Continued)
DESIGN REQUIREMENTS/CONSTRAINTS FOR COMPONENTWEIGHT AND SIZING
Materials:
• LOX:
• Fuel:
Minimum gage:
• Aluminum: 0.03 in. I
• Titanium: 0.02 in.
Ultimate factor of safety : 1.5
2219-T87 AL
2219-T87 A1 or 6AI-4V Ti (whichever is lighter)
Per NASA Direction
Thermal control: silica fiber, foam, or multilayer insulation with
thermodynamic vent
- Propellant acquisition:
- OMS propellant gaging:
- RCS propellant gaging:
surface tension screens
capacitance probes
P-V-T
Accumulators
- Pump startup response = 0.5 sec
- Number of RCS accumulator recharge cycles : 50/mission
- Blowdown accumulator operation (isentropic blowdown process)
- Shape: spherical or cylindrical with hemispherical end domes
- Materials:
• LOX: 2219-T87 A1
• Fuel: 2219-T87 A1 or 6 AI-4V Ti (whichever is lighter)
- Minimum gage
• Aluminum: 0.03 in.
• Titanium: 0.02 in.
- Ultimate factor of safety = 1.5
- Thermal control: silica fiber or multilayer insulation (no vent)
Propellant acquisition: surface tension screens for liquid feed
Propellant feedlines
- Current OMS/RCS line lengths
- Pressure drop:
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TABLE Vl (Continued)
DESIGN REQUIREMENTS/CONSTRAINTS FOR COMPONENTWEIGHT AND SlZlNG
• 0.5 psi/ft for pressure fed system
• 1.0 psi/ft for pump fed system
- Darcy friction factor
- Isenthalpic expansion process
- Materials:
• LOX: 2219-T87 A1
• Fuel: 2219-T87 A1 or 304L SS
- Minimum gage = 0.028 in.
- Ultimate factor of safety:
• 4.0 for D < 1.5 in.
• 1.5 for D > 1.5 in.
- Thermal control: silica fiber or multi layer insulation
- Linear and angular deflection compensation joints
Gas generator exhaust vent line
- Line Mach number = 0.3 (maximum)
- Fanno line analyses
- Line length: 20 ft
- Exhaust nozzle area ratio = 2.0
- Minimum gage and ultimate factor of safety:
Propulsive vent for OMS, non-propulsive vent for RCS
Line material: 304L stainless steel
same as feedlines
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MCDONNELL DOUGLAS ASTRONAUTICS COIt4PANY.B_ LOUIB DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
3.0 TASK 1.2 --PHASE I COMPONENTCHARACTERIZATION
REPORT MDC E2548
JULY 1982
Three proprietary MDAC-STL computer codes were used to evaluate the candidate
system concepts:
• Advanced Propulsion System Design and Sizing Code (APSDS)
• Tank Heat.T-ransfer Code (TKHEAT)
• Feedline Heat Transfer Code (FDLINE)
A description of these computer codes is provided in the following paragraphs.
3.1 APSDS Code
This code sizes the propulsion system to a fixed volume or fixed total impulse
constraint. Sizing to a fixed pod volume constraint is accomplished through use of
an iteration loop within the code. To start, propellant tank volumes are
calculated based on an assumed total impulse requirement. The calculated total
tank volumes (fuel and oxidizer) are then compared to the total available tank
volume within the pod. If the calculated volume is out of tolerance a revised total
impulse estimate is made, and tank volumes are recalculated. To ensure rapid
convergence a secant numerical analysis technique is used to estimate a new total
impulse requirement. A simplified flow diagram for the fixed volume analysis is
shown in Figure 5. After the propellant tank volume is determined, the program
calculates the system pressure budget and total system weight. Table VII
identifies the system components that are modeled in the APSDS code and Table VIII
shows example output for a LOX/ethanol system with common OMS-aft RCS prope]lant
tanks. A brief description of major system component models--tanks, accumulators,
feedlines, and engines--is provided below.
3.1.1 Propellant Tanks - The cryogenic propellant tank model is shown in Figure 6.
Options for tank shape, material, and insulation are available depending on the
propellant and system type. The tank shapes employed for Phase I were:
• OMS tanks--cylindrical with oblate spheroid end domes
• RCS tanks--spherical
• Common OMS/RCS tanks--cylindrical with oblate spheriod end domes.
Aluminum, 2219-T87, was assumed for the LOX tank, while either 2219-T87 aluminum or
6AL-4V titanium was assumed for the fuel tank (whichever was lighter). The
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MCDONNELL DOUGLAS ASTRONALITICS COMPANY-BT. LOUIB DIVISION
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Auxiliary PRopulsion system studyFINAL REPORT
ORIGINAL P_:',GZ _S
OF POOR Q!JALITY
REPORT MDC E254_JULY 1982
L
INPUT DATA
PROPELLANT STORAGE CONDITIONS:
TANK TEMPERATURES & PRESSURESENGINE CHARACTERISTICS:
MR, PC, _. ETC.GAS GENERATOR CHARACTERISTICS:
MR, PC, ETC.
ACCUMULATOR CHARACTERISTICS:BLOWDOWN & SWITCHING RATIOS
TURBINE CHARACTERISTICS:INLET TEMPERATURES,PRESSURE RATIOS
INITIAL TOTAL IMPULSE
VALUES FOR OMS & RCS, ETC.
OMS SYSTEM PERFORMANCE
PROPELLANT PROPERTIES
PD82 44()
OUTPUT DATA
OMS & RCS PER FORMANCE:TOTAL IMPULSE, ISPMR, _V, IT, E]:C.
OMS & RCS SYSTEM PROPERTIESOMS, ARCS, & FRCS COMPONENT WTS.
PROPELLANT & HELIUM TANKS.ACCUMULATORS, PROPELLANT &
HELIUM DISTRIBUTION SYSTEM,ENGINES, ETC.
OMS, ARCS, & FRCS TOTAL SYSTEMWET WEIGHTS
TURBOPUMP POWER BALANCE
ENGINE & THRUSTER PERFORMANCETOTAL LINE &PPROPULSIVE VENT
USABLE & NON-USABLE PROPELLANTTANK VOLUMES, ETC.
TOTALTANK\_-_OLUME WITHIN _--I_NO_--_-
TOLERANCE /
.i i,i
RCS SYSTEM PERFORMANCACCUMULATOR OPTIM IZAT{ONSYSTEM PROPERTIESTOTAL LINE &PENGINE FLOW RATESUSABLE & NON-USABLE
PROPELLANTSTOTAL TANK VOLUME
& TOTAL IMPULSE
]TOTAL SYSTEM WET WEIGHT
PRESSURANT TANKSPROPELLANT TANKSPROPELLANT FEEDLINESVALVINGTURBOPUMPS
HEAT EXCHANGERSENGINES, ETC.
s_
FIGURE 5 APSDS FLOW DIAGRAM FOR FIXED VOLUME ANALYSIS
24
MCDoIVIVELL DOUGLAS ASTROlVAuTICS COMPANY iT LOUIS DIVISlOIIble .
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Auxiliary Propulsion System Study
FINAL REPORT
TABLE Vll
SYSTEM COMPONENTS IN APSDS COMPUTER CODE
REPORT MDC E2548
JULY 1982 Bv'
• Pressurant tanks
- metallic (monolithic)
- composite
° Propellant tanks (insulated and non-insulated)
- spherical 1
- cylindrical aluminum and titanium
- conical
• Accumulators (insulated and non-insulated)
- spherical 1- cylindrical aluminum and titanium
• Feed system components
- pressure regulators
- check valves
bust disk/relief valves
- manual valves
solenoid valves
- cryogenic valves
° Propellant feedlines
- insulated
- non-insulated aluminum and stainless steel
• OMS engine (regen-cooled)
- pump fed
- pressure fed
• RCS thruster (film-cooled)
25
MCDONNELL DOLIGLA8 4STROItlAUTICS COMPAIVY-BT. LOUOli OlVIBION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
EXAMPLE SYSTEM WEIGHTTABLE VIIISUMMARYFROM APSDS COMPUTER CODE
REPORT MDC E2,548
JULY 1982
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IAP$ SUmml_YIP00
:O_0h 0qSIRC$ L]XI[THANOL(NBP)
INT_GRAhD FRCSIAPS
PERFO_mlNCE
* OqS SYSTEm e ON) ENGINE
FSTS, LBF - bOgOoOO Fe LBF - lOOOO.3IT LBF-SEC - ,)34DEtO? PC, PS|A - B0O.0XSPe L_F-SECILBN 344.10 |SP* LBF-$E¢ILBM - 34q.0OIF - 1.?_ OIF - |tBO
AEIAT 240,0
e RC$ SYSTEm • RCS PR|_AR¥ ENGINE
FSY$_ L3F - bogo.oo F, LBF - 870.0IT LBF-SEC - tLZgO;*O? PC* PSIA - ZSO,O|$Pp LBF-$ECIL_R Z34,52 ISP* LBF-SECILBN - 301.00IF - l,|Z OIF 1.41
AEIAT - 4T,O
MEIGHTp LBN
OX FUEL TOTAL
• STSTE_ VET NE|GHT - |lqBb.4B
o USFA_LE PROPELLANT 930Z,66 $90S.30 2$_07.9&- TANKED PROP BBgO.qB SbO?,3? 1449B,34- $U_P TAN_ PROP 411.69 Zg?.q) 709.61
• NON-USEABLE PROPELtANT S_4,?3 |_5,3q qZO.!l- TANK B_IL-OFF S_.34 O,O0 S_.34- LXNE BOIL-OFF 0.00 O.O0 0.00
PROP TAN¢ LX_OVAP RESZD Z_B.39 1_1.54 349,q4AP$ ACCUqULATDR PROP _6_.33 S_,4b 2Z4.?9L[NE PR|PELLANT _,_Q _.08 _9,)?HALF CROSS-FEED PROP 7.85 1.64 15.49ARCS |NTERC3_ LZNE PROP • 5_U_P TANK L|glVAP RESID 9._7 1,73 Z.OB
6.13 lb._OFRCS |q_ERCOq L|NE P_DP 9.0L 50.99 6O,O0FRC$ A¢¢UXULATOR PROP BB,q) 59,4b |4B.39FrO RC$ LINE PROP 2,2b 1b,34 1B.61
• NELIU_ - lO.OZ
• RC$ VERNIER ENGINE
F_ LBFP _ PSIA|_P_ LBF-SECILBRDIFAEIAT
* StSTE_ DRY dHGqT- PROPELL4NT T4NK$ )Z?_BR |gq_l ZBZ$.39_ZT.Oq- A_S ACCUqULkTOR$ |_q,_8 _4,q4 Z04.$3
P_OPELL_T O|$TR[_UTION _0._3 |00,4_ |51.07HFLZUN TANKS 43.07 43.07MELIUK OIST_IBUT|ON - 48.|4ENGtN[__ vent zz_Bs zs_ _)e.zo47,99
- TPAtGG 117,74 l_,51 _Ob,?SHEAl Fx_tqGERS _1,14 O,OO .14
0NF-qlLF CROSS FEED |b,_4 _$,4_ )qeggA_C$ INTERCO_N[CT LI_ l_.qT |B,S1 31.49[_TRy SU_P T_NK 46.74 3].75 80*50FeCS lnTE_CON L|_E )q.t? q_.4O _38,6_FeC$ ACCUqULtTOe S |Z_*Q_ 44,q4 16?,B?F_CS LI_E 43,?_ ??,O? |20.BO
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MCDONNELL DOUGLAS ASTRONAUTICS CO.14PANY.liT. LOUlli IDIVllilOIV
Z$.0Z_O.O
1,8043.0
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Auxiliary Propulsion System StudyFINAL REPORT
REPORT MDC E2548
JULY 1982
OF P;35i_ g' "' "
OUTER
PROTECTIVE
INS
THERMODYNAMIC
VENT COOLING
SHROU
PRESSURE
VESSEL
PROPELLANT
ACQUISITION
SYSTEM
THERMAL STANDOFF
NG PROBES(2)
HELIUM TANK
FIGURE 6 CRYOGENICPROPELLANTTANK MODEL
27
MCOOItlNELL DOUGL4$ 4STRONALITICB CO_P,4Nvr-J'r. LOiLII8 IDIVISlOIV
LOX/HYDROCARBON
Auxiliary PRopulsion System StudyFINAL REPORT REPORT MDC E2548
JULY 1982
pressure vessel was sized for the total required propellant volume, including
allowances for vent losses, tank residuals (liquid and vapor), and a 5% ullage
volume. Pressure vessel wall thickness was calculated-applying the material
ultimate stress and the maximum operating pressure (tank relief pressure) with a
factm ' of safety of 1.5. Minimum gage wall thicknesses were specified by NASA to
be O.03 in. for aluminum and 0.02 in. for titanium. The calculated pressure vessel
weight includes a non-optimum factor to account for bosses, gage variations,
support attachments, and weldments. The cooling shroud (thermodynamic vent
system), insulation, and outer shell weights were calculated based on the taqk
surface area, while the acquisition system weight was based on the current OMS and
RCS designs. OMS gaging probe weights were calculated as a function of tank length.
A tank support structure weight was also calculated based on the total tank weight
and loaded propellant weight. An example computer output weight summary for common
OMS-aft RCS propellant tanks (LOX and ethanol) is shown in Table IX.
3.1.2 Accumulators - The RCS accumulator has three basic functions in a pump fed
system: to provide propellant to the gas generator during the pump start
transient, to provide impulsive propellant to the RCS thrusters during the mission,
and to provide an ullage volume for propellant thermal expansion due to line
heating. The OMS accumulator has one primary function--to supply propellant to the
OMS _as generator during engine startup. As such its volume is much less than theRCS accumulator.
The accumulator operates in a blowdown mode as shown in Figure 7. Initially,
th( accumulator is charged to a maximum pressure. It is then allowed to blowdown
to e switching pressure, at which time the turbopump assembly is activated. During
th( pump start transient, the accumulator supplies propellant to the gas generators
t,_d the pressure decays further to a minimum pressure. The accumulator is then
,upplied with propellant from the turbopump and recharged to the maximum pressure.
In Phase I the RCS accumulator was sized to provide 50 recharge cycles per mission.
Accumulator weights are calculated in a manner similar to the propellant tanks
but do not include a cooling shroud or gaging system. An example weight summary for
the RCS accumulators is presented in Table X. This example is for a hybrid RCS in
which the thrusters are supplied with gaseous 02 and liquid ethanol at ambient
temperature.
28
MCDONNELL DOUGLAS ASTRONAUTIC8 COMPANY.lIT. LOUIS DIVISION
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Auxiliary Propulsion System Study
FINAL REPORTREPORT MDC E2548
JULY 1982
!(
TABLE IXEXAMPLE APSDS OUTPUT -- PROPELLANT TANKS
APS PROPELLANT TAN<AGE qOOEL
COMPONENT
PRESSURE VESSEL
CODLING SHR']UD
INSULATION
OUTER SHELL
AQUISITION SYSTE_
HELIUM TANK
HELIUM TANKSUPPORT STRUCTURE
PR_PFLLANT TANKSUPPORT STRUCTURE
TOTAL
(aLl IAL)OXIDIZER FUEL
SIDE SIOE
- 80,75 77,03
- 10.83 O.OO
- 52.01 O.OO
- 36,00 O,OO
- 69,48 b4,45
- 40,_4
- 2,6_ -
- 78,81 57.73
370,94 199,_1
TOTAL PROPEL! ANT TANKAGE DRY WEIGHT o 570.15 LBS PER POD
PROPELLANT TANK VOLURE t
PROPELLANT TAN_ DIAMETER t
PROPELLANT TANK PRESSURE t
PROP TANK RELIEF PRE$SUREI
OX SIDE - 13T,_4 CUB FTFUEL SIDE - 1Z_,b2 CUB FT
OX Sift Q 4.1E FTFUEL SIDE - 3,89 FT
OX SIDE - 35,00 PSIAFUEL SIDE - 39.00 PSIA
OX SIDE - 48.94 PSlkFUEL SlOE 45.90 PSIA
HELIUM TANK VOLUME - 1,BO CUB FT
HELIUM TANK DIAMETER - 1._1 FT
29
MCDONNELL DOUGLAS ASTRONAUTICS COMPANY*IIToLOUIII DIVI|ilON
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORTREPORT MDC E2548
JULY 1982
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PUMP FLOWI DEMAND
LOW RATE THRUSTERUSAGE
o / _FORAZ_T_UDE.CON!ROL
i_ PMAXPSWITCH_/'-'-'--BLOWDOWNRATIO = PMAX/PSWITCH
""------_._i_ SWITCHINGRATIO : PSWITCHIPMINV
,.,.,
PMINi
TItlE
FIGURE 7 RCS ACCUMULATOR OPERATION
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L _UGLAS tONJ IC5 iNY SFo .
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Auxiliary Propulsion System StudyFINAL REPORT REPORT MDC E2548
JULY 1982
ORIGIN,gL F,:.C_.:.:..:OF Poor .... .-_-" Q'J'.L:_ ;_
TABLE X
EXAMPLE APSDS OUTPUT -- RCS ACCUMULATORS
APS ACCUMULATOR WEIGHT PER PO0
IGAS) (LID)COMPONENT OXIOIZER FUEL
• SIDE SIDE
PRESSURE VESSEL - 88,31 35,q8
INSULATION " 0,00 0,00
FIBERGLAS SHROUD " 0.00 2.47
AQUISITION SYSTEM - 0,00 2,39
REGS ANO VALVES " 28.08 0,00
SUPPORT STRUCTURE - 6,53 ¢,10
TOTAL 122.gE 6_o9¢
WEIGHT OF PROPELLANTIN ACCUMULATOR -
t
88,93 5_.t6
TOTAL ACCUMULATOR WET WEIGHT 31b,Z6 LBS PER PO0
ACCUMULATOR VOLUME l
ACCUMULATOR DIAMETER !
OX SIDE - I_,69 CUB FTFUEL SIDE 2._? CUB FT
OX SIDE - 3,0_ FTFUEL SIDE - 1,b8 FT
31
MCDONNELL DOUGLAS IISTROltlAILITIC8 CO,'HPANY .|T. LOUIS DIVIIION
LOX/HYDROCARBON
Auxiliary Propu6ion System StudyFINAL REPORT REPORT MOO E2,54,8
JULY 19_
3.1.3 Feedlines - In Phase I, vacuum jacketed feedlines were modeled for cryogenic
propellant feed while non-insulated lines were modeled for ambient temperature
propellant feed. _he line lengths were based on the current OMS-RCS line routings,
and each line segment was sized by iteratively solving the Darcy pressure drop and
Colebrook friction factor equations. The feedline weights were determined based on
the use of 2219-T87 aluminum with a minimum gage thickness of 0.028 in. The
cryogenic feedline model is illustrated in Figure 8 and includes weights for
multi-layer insulation, vacuum jacketing, and linear/angular compensation joints.
Feedline support weights were also calculated based on the feedline weight and
weight of propellant contained in the feedline. A typical computer output summary
is shown in Table XI for a cryogenic RCS LOX feedline. This summary includes the
weights for feedline isolation valves.
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3.1.4 Engines - Engine system component weight and performance data were developed
by ALRC and were provided in Volume ll--Parts A and B of Reference (I). Included
in the engine system are the turbopumps, gas generators, thrust chamber assemblies,
and valving. The OMS engine was assumed to be fuel regen-cooled, while the RCS
thruster was assumed to be film-cooled. OMS engine and RCS thruster lengths and
diameters were constrained to their current values.
3.2 TKHEAT Code
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Ibis code was developed to determine the thermodynamic response of cryogenic
propellant tanks and accumulators during representative orbital mission duty
,ycl s. The basic cryogenic tank model (Figure 6) consists of a pressure vessel,
insulation system, an optional thermodynamic vent system, and optional outer
cove.-. The insulation system can be a single component or a composite type system,
i._., foam and multi-layer insulation.
The operation of the thermodynamic vent system (TVS) is shown in Figure 9. In
this system propellant is vented to maintain acceptable tank pressure and
temperature levels. Vent propellant is withdrawn from the tank in a liquid phase
through the propellant acquisition assembly. After throttling to a lower pressure
and temperature it is passed through a tank heat exchanger where it is vaporized by
heat entering the tank. This TVS concept eliminates the propellant positioning
32
MCDONNELL DOUGLAS ASTRONAUTIC8 COMPANY 81" LOLII8 OIVI,_ilON• o
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Auxiliary PropulsbN System Study
FINAL REPORT ORIGINAL PACE |:_
OF POOR QUAI..ITY
REPORT MDC E254,BJULY 1982
MULTILAYER_--INSULATION
/---VACUUMJACKETPROPELLANT iNNERLINEDIAMETER,IN. WEIGHT,LB/FT
LINE 0.5 0361.5 0.54
2.5 1.00
_. • ---_j'?fr _J r .
..... b -_- -- ' ........ ' ..... i .............. \
I tJ._._ --_-" _ I, ._)
LINEARCOMPENSATOR ANGULATIONJOINTMULTILAYEINSULATION
INNER LINE COMPENSATORDIA., IN. TOTAL WEIGHT, LB
.5 3.7 LBS1.0 3.8 LBS1.5 4.0 LBS
FIGURE 8 CRYOGENIC PROPELLANT FEEDLINE MODEL
33
MCDONNELL DOUGLAS ASTRONAUTICS COMPANY. IT. LOUli DIVISION
LOX/HYDROCARBOI_
Auxiliary Propulsion System StudyFINAL REPORT
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MCDONItI£LL. DOUGLAS ASTRONAUTICS COMPANy IT LOUIS DIVIIION- ,
REPORT MOC E2548
.JULY 1982
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FINAL REPORT
REPORT MDC E2548
JULY 1982
OR._GINAL P,_.CC ._-a
OF POOR Q..,AL!_ _
THROTTLINGORIFICE
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r-,,".-...,
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FIGURE 9 THERMODYNAMIC VENT SYSTEM CONCEPT
35
MCDONNELL DOUGLAS ASTI@ON4UTIC3 COMPANY.lIT. LOUIB DIVISION
,/
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
REPORT MDC E2548
JULY 1982
problems associated with vapor phase tank venting and reduces the heat flux into
the tank. T_e TKHEAT code can evaluate the relative effectiveness of TVS heat
exchanger lines mounted directly on the pressure vessel or on a cooling shroud that
surrounds but is displaced from the pressure vessel.
The tank thermal model is illustrated in Figure 10 for an installation within
the Orbiter aft pod. The steady state one-dimensional heat transfer equations are
solved using an implicit finite difference technique. The surface boundary
condition is obtained by calculating the net solar flux entering the thermal tiles
at surface I or by specifying vehicle skin internal temperature at surface 2. A
uniform state approximation is applied to the fluid within the propellant tank.
The fluid inside the tank can be single or multi-phase, homogeneous (propellant
only), or heterogeneous (propellant and pressurant gas).
Engine, vent and gas pressurization valve operations are modeled to calculate
the thermodynamic response of fluid within the tank. Fluid pressure and
temperature response are determined by solving the unsteady flow forms of the
conservation of mass and the first law of thermodynamics using an implicit finite
difference technique. A simplified flow diagram for the TKHEATcode is provided in
Figure 11.
To validate the code computed results were compared with prior experimental
data obtained on an MDAC-STL prototype cryogenic tank using liquid nitrogen as the
test fluid. The tank was spherical in shape, had a diameter of 39 in., and was
constructed of Ipconel 718. The insulation system consisted of 2 in. of foam and
2 in. of aluminized mylar multi-layer insulation (MLI). The tank also employed a
thermodynamic vent system. As shown in Figure 12, test data obtained from this tank
compare very well with analytic predictions using the TKHEAT code.
The TKHEAT code was used to evaluate propellant tank heating and venting
during typical OMS/RCSmissions. Example LOX tank results for an OMS 30 day mission
are presented in Figure 13. LOX tank venting initiates when the tank total pressure
reaches 60 psia and terminates when the total pressure decays to 57 psia. The total
LOX vent loss is just over 200 lb.
36
MCDONNELL DOUGLAS ASTRONAUTICS COMPANV* liT. LOUIS DIVISION
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Auxiliary Propulsion System StudyFINAL-REPORT
ORDINAL .......Of POOR QUALITY
REPORT MDC E2548
JULY 1982 B,/
QSOLAR
QSPACE
1
ERMAL TILE
STRUCTURAL SUPPORT
(HEAT LEAK)
VACUUM
PROTECTIVE COVER
INSULATION
3LANT SHROUD (TVS)
ROPELLANT TANK
RUCTURAL SUPPORT
(HEAT LEAK)
_PROPELLANT
FIGURE I0 TANK THERMALMODEL
37
MCDONNELL DOUGLAS ASTRONAUTICS COMPANY.liT. LOUIJ DIVIBION
LOX/HYDROCARBON , +.;:: + I
Auxiliary Propulsion S.+_T__mStudy 0 H_ , + REPORT MDC F_2_B +
FINAL REPORT OF POoR Q:J/_L_I-" IULY 1982 I |
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
ORIGINAL PA_ i5
OF POOR QUALITY
REPORT MDC E25,48
JULY 1982
£
I
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PD82 451
EXPERIMENTAL HEAT FLUX - 8.84 BTU/HR I I_
ICALCULATED HEAT FLUX - 8.48 BTU/HR (TKHEAT) TEST
• TEST DATA _ TKHEAT PREDICTION TANK I_._IlL \\\_v^,, /_';v
J J. _ 1
QeO_II$11 .... |_t._._ loe lee•coo• eeo •ee --
I l , J U
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as
TIME - HR
116
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FIGURE 12 COMPARISON OF TKHEAT CODE WITH TEST DATA
39
MCDONNELL DOUGI. AS AS TROtIIAU TICS COMPANY- liT. LOUIS DIVISION
LOX/HYDROCARBOIN
Auxi|iARy PROpu_iON SySTEm STudy
FINAL REPORT
• INITIAL LOX LOAD _ 9350 LBM• RELIEF PRESSURE = 60 PSIA
_g
F_
J
Q:
fJ
8
a-
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FIGURE 13 EXAMPLE TKHEAT COMPUTATIONS FOR OMS LOX TANK
40
MCL)OI_IblEI.I. dr_OLIGI,. AS A 5 TRONALITICS COMP.4_tl Y . ST. L, OUI_ OIVI_IOI_
REPORT MDC E2548
JULY 1982
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Auxiliary Propulsion System Study
FINAL REPORT
3.3 FDLINE Code
REPORT MDC E25,48
JULY 1982
This code was developed to determine the.-thermal response of propellant
feedlines during orbital mission duty cycles.
The propellant thermal response is computed by first dividing the feedline
into segments as illustrated in Figure 14. The conservation of mass and energy
equations for each segment are solved simultaneously by an implicit finite
difference method for each time step. A two-dimensional transient heat transfer
analysis of the teedline segments is performed. Heat transfer is by conduction
through the insulation, along the feedline, and through structural members (such as
line supports). The outer surface insulation temperature _nd heat leak source
temperatures are fixed boundary conditions. Heat soakback from the engines into
the feedlines is computed by specifying a thrust chamber injector head temperature
and valve thermal isolation resistance. The propellant in the feedline is modeled
as a single or multiphase, homogeneous fluid. A simplified flow diagram for the
FDLINE code is presented in Figure 15.
The FDLINE code was used to compute LOX feedline temperature response during
representative RCSmission duty cycles. The mission duty cycle is broken down into
two distinct modes of operation--primary thruster firings and vernier thruster
firings. Each prima_y thruster firing (~150 per mission) is modeled as a discrete
pulse, whereas the vernier firings (~15,000 per mission) are modeled as a
continuous burn at low flowrate.
Example LOX temperature profiles for a 30 day RCS mission are shown in
Figure 16. In this example, all the RCS propellant is expended through a single
primary manifold feeding all the primary and vernier thrusters necessary fo."
Orbiter three-axis attitude control. The feedline insulation for thi_ exam#le
consists of one inch of aluminized mylar multi-layer insulation (MLI). Despite a
wide range of accumulator LOX temperatures (160 to 200OR), acceptable feedline
temperatures are achieved throughout the mission.
41
MCDONNELL DOLIGLA8 4s'rRD_ALITIC8 _._DMPANt,'-ST, IL.Otll8 DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
IREPORT MDC E2548
JULY 1982 I __ ,
ORIGINAL F.r_(_,,-:_; I
t
THRUSTER J
, VALVE J
,_;_:;.,,• . -...
___ j._:::-::.:':._'
,_,_;_'.",
OF. POOR QUALITY
L,NENODALMODEl _ ®
ACCUMULA O _--STRUCTURAL SUPPORTS MANIFOLD /
(HEAT LEAKS) VALVE / THRUSTER
(_HEAT &AMBIENT
_1111II1Ill|[][l[ll_l!llfli!111
o. MIN@' • ---}--"_out-_z/////////_//////l_/.,'//////_
.oo,,@I @ t ®
FIGURE 14 FEEDLINE THERMAL MODEL
I
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42
MCDONNELL I/JOUGLAS 4STI_tONALITIC,."; COMPANY-ST. I..OUI_ OIt/IBION
!
I
I
I
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORTREPORT MDC E2548
JULY 1982
43
MCDONNELL DOUGLAS ASTRONAUTICS COMPANY.lIT, LOfJIS _DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
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Or|Gll_/_L _,_,0": _:,-, jULY 1982 i | "
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MCDONNELL DOUGLA#i ASTRONAUTICS COMPANY-ST. LOUO_ DIVOBION I
LOX/HYDROCARBON
AuxiliaRy ProPulsioN System Study
FINAL REPORT
4.0 TASK 1.3--PHASE I SYSTEM EVALUATIONS
REPORT MDC E2548
JULY 1982
the system design options established in Task 1.1 (Table V) were evaluated in
this task. The required system weight and performance data were generated using
the APSDS computer code with engine system models supplied by ALRC. Computer
output summaries for each system concept were presented in Volume ll--Part C of
Reference (i). The following paragraphs present the results and conclusions from
these evaluations.
4.1 Pump Versus Pressure Feed
As discussed in Section 2.2.1, this option was evaluated for both the OMS and
RCS, as well as, all four fuel candidates.
4.1.1 OMS - As shown in Figure 17, the pressure fed LOX/HC OMS is similar to the
current OMS with exception that the helium bottle is stored inside the LOX tank.
The pump fed LOX/HC OMS schematic is shown in Figure 18 and incorporates the
component redundancy necessary to meet the fail operational-fail safe reliability
requirement of the current OMS. The pumps are powered Dy gas generator driven
turbines, and pump NPSP is provided by a small helium pressurization system.
During startup the gas generators are supplied with propellants from small liquid
accumulators that operate in a blowdown mode. As in the current OMS the engine is
fuel regeneratively cooled, and a separate nitrogen supply is used for engine valve
actuation. LOX and methane propellants are fed to the engine at cryogenic
temperatures, and the feedlines are vented following each burn.
Even though the bulk density-specific impulse product for the LOX/HC
propellants is less than for the current storable propellant combination (N204/
MMH) the LOX/HC OMS provided an opportunity for improved propulsion system
packaging. The reason for this can be seen by referring back to Figure I which
shows propulsion system packaging for the current system. By storing the helium
bottle inside the LOX tank the required helium volume is reduced as a result of the
low storage temperature (165OR), and the propellant tanks can be extended
11.5 inches aft. The corresponding increase in available propellant tank volume
45
MCDONNELL DOUGLAS AS TROI_IAUTICB COMPANY. BT. LOUIB DIVISION
LOX/HYDROCARBON I
Auxiliary Propulsion System Study RepoRt MDc e2_m
FINAL REPORT _ r- JULY 1982
I• SEE SCHEMATIC NOMENCLATURE OF FIGURES 2 & 3
I
'',,I
XFEED _ --, I XFEED I
!
!
I
FIGURE 17 PRESSURE FED OMS SCHEMATIC I
H
I46
MCDONNELL DOUGLAS ASTRONALITICS COMPANY- ST. LOLIIB DIVISION I
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT ORIGINAL _ "_:"OF pOOR QUAL.F;'Y
REPORT MDC E2548
JULY 1982.
• SEE SCHEMATIC NOMENCLATURE OF FIGURES 2 & 3
LOX
]XFEED _'_ L.t_ _
FUEL
1ACCUMULATOR
.C)
XFEED
FIGURE 18 PUMP FED OMS SCHEMATIC
47
MCDOftlIYEI,.L DOUGLAS ASTRONAUTICS COMPAIYV-IIT. LOUt| DIVIIIION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
REPORT MDC F..2548
JULY 1982
compensates for the lower bulk density of the LOX/HC propellants. The benefit is
most pronounced for the pump fed systems which have substantially lower helium mass
requirements for tank pressurization.
Based on the sensitivity data of Figur, 19 for LOX/propane, an OMS chamber
pressure of 800 psia was baselined for the pump fed systems in Phase I to maximize
performance (&V capability) and minimize weight. A chamber pressure of I00 psiawas
selected for the pressure fed system based on prior experience.
The pump and pressure fed systems are compared in Figure 20 for all four fuel
candidates. Three criteria are used in the comparison; OMS AV capability, OMS wet
weight, and OMS dry weight. To compare OMS 6V capability the aft pod volume was
fixed at the current value. To compare wet and dry weights the 6V capability was
set equal to the current OMS value of 500 ft/sec per pod. (The dashed line in each
comparison represents the capability of the current OMS.) From the comparisons of
Figure 20 it is seen that the pump-fed OMS offers overriding advantages in terms of
weight and performance. This is the result of the higher engine specific impulse
that can be achieved with the pump fed systems. For example, the LOX/propane pump
fed engine IsP is 363 Ibf-sec/Ibm (with a nozzle area ratio of 240), while the
pressure fed engine Isp is only 324 Ibf-sec/Ibm (with a nozzle area ratio of 44).
As discussed in Section 2.3 the overall engine envelope is constrained to the same
dimensions as the current OMS engine. Also, from Figure 20, it is seen that ethanol
offers the highest OMSAV capability and lowest system dry weight. This is because
Lhe LOX/ethanol combination offers the highest bulk density-specific impulse
.,rodvct of the candidate propellants. Although LOX/methane provides the lowest
system weight (highest payload capability) for a fixed AV requirement, the
_OX/ethanol system would be less costly since cryogenic tankage is required for the
O_ side of the system, only. On the basis of these comparisons the pump fed OMS
basedlined for Phase II. For both aft pods the pump fed OMS offers a
i)00-4000 Ib weight advantage over the pressure fed system.
4.1.2 RCS - Schematics for the pressure and pump fed aft-RCS are shown in
Figures 21 and 22, respectively. The forward-RCS Ss similar but incorporates two
additional primary thrusters. Both schematics incorporate the same component
48
MCDONNELL DOLIGLA8 ASTRDNALITIC8 COMPANt'oST. LDLII5 DIVISION
I
I
I
I
I
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I
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
O,-_,G,N,L ; ......OF PO `,;l_" Q_'""_ _
REPORT MDC F...2548
JULY 1982
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MCDONNELL DOUGLAS ASTI_rONAUTICS COMPANY* ST. LOLII5 DIVISION
LOX/HYDROCARBON IAuxiliary Propulsion System Study -_ ,,,_ RePoRtMDCE2548FINALREPORT oR_IN,aL pl_,_.t, lt_1 .'ULY 1982 _;_.'
OF POOR QUALITY
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50
MCDONNELL DOUGLAS ASTRONAUTICS COMPANY*ST. LOUIS DIVISION I
LOX/HYDROC,_RBON
Auxiliary Propulsion System Study
FINAL REPORT
OF _'C_'" _" ;':
REPORT MDC E2548
JULY 1982
• SEE SCHEMATIC NOMENCLATURE OF FIGURES 2 & 3
-F
#,
LOX FUEL
THRUSTERS
XFEED XFEED
FIGURE 21 PRESSURE FED RCS SCHEMATIC
51
MCDONNELL DOUGLAS ASTRONAUTICS CO[t4PANt" oST. LOUIS DIVIBIOFti
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
_J
• SEE SCHEMATIC NOMENCLATURE OF FIGURES 2 & 3
REPORT MDC F.2548
JULY 1982
I
L
I
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XFEED XFEED
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!
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FIGURE 22 PUMP FED RCS SCHEMATIC
52
.? . °MCDOItI_ELL DOUGLAS ASTFI'JNAUTICS COMP4NY ST LOUIS DIVISION
I
I
I
I
I
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
REPORT MDC F.,2.5.,_
JULY 1982
redundancy as the current RCS and assume liquid propellant delivery to the
thrusters. As such, the LOX and methane feed systems employ insulated feedlines
and accumulators. (Vacuum jacketed feedlines incorporating multi-layer insulation
were baselined for these evaluations.) The pump fed RCS (Figure 22) employs
OMS-type turbopumps to preclude the necessity for a second turbopump development
program. System specific impulse calculations in the APSDS computer code account
for Isp penalties associated with the use of the OMS turbopumps and for penalties
associated with pulse mode operation of the RCS thrusters. As discussed in Section
3.1.2 the accumulators of Figure 22 operate in blowdown mode while supplying
impulsive propellant to the RCS thrusters. They were sized to limit the number of
turbopump cycles to 50 per mission.
A comparison of pump and pressure fed RCS is presented in Figure 23 for all
four" fuel candidates. Similar to the OMS three criteria are used in the comparison;
percent total impulse capability, wet weight, and dry weight. (100% total impulse
capability corresponds to the capability of the current earth storable RCS, and the
dashed lines show the wet and dry weights of the current RCS.) As shown in
Figure 23, neither the pump nor pressure fed systems are able to meet the total
impulse capability of the current system when constrained to the same RCS tank
volume as the current aft pod. This is because of the lower bulk density-specific
impulse product of the LOX/HC propellants. When sized to a fixed total impulse
requirement the pump fed systems are heavier as a result of their lower system
specific impulse. The lower specific impulse is due to turbine vent loss
penalties. When sized to a fixed volume requirement the pump fed systems generally
provide higher total impulse capability. This is due to lower system mixture
ratios which permit more efficient propellant packaging within the pod. As in the
case of the OMS, the LOX/ethanol propellant combination provides the highest _V
capability and lowest dry weight.
The comparisons of Figure 23 were performed for a RCS thruster chamber
pressure of 150 psia. The performance and weight sensitivities of the pump and
pressure fed systems to chamber pressure are presented in Figure 24 for LOX/
propane. These data show that the near optimum chamber pressures are 300 and
I00 psia for the pump and pressure fed systems, respectively. When compared at
their near optimum chamber pressures it is seen that the pump and pressure fed
53
MCDONNELL DOUGLAS ASTRONAUTICS COMPANY. ST,. I.OUI$ DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
P ;, ,C:.: '_,L'_O_K_INAL ':'"'';"OF pOOR QuALtYY
REPORT MOC F..2548
JULY 1982
I--.
<o,,.
w
,=Io..
I00-
90
80 - I
i
lO -i
°°-I5O
,m
l
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FIXED VOLUME
PC = 150 PSIA
• SEPARATE FRCS
• CRYOGENIC FEED
CURRENT FRCS TOTAL IMPULSE - 100%
I
=|
PUMP
3800 "
1200
1100z_rt
- I000i,,,-
900>.
p-,
800
70O
FIXED TOTAL IMPULSE (100%)
CURRENTFRCSWET NT. - 3280LBm
m
3700 -
==_=3600
3500
3300
PRESSURE PUMP
FIXED TOTAL IMPULSE (i00%)_
CURRENT F'RCS DRY WT. - 750 LB
__ O0 _r"P :I:
R'_ U -PU O
.PUMP PRESSURE
-t-O
:_ hg
PRESSURE
CURRENT SYSTEM
FIGURE 23 PUMP AND PRESSURE FED RCS COMPARISONS
54
MCDONNELL DOUGLAS ASTRONAUTIC* .S COMPANY. ST. LOUIS DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORTREPORT MDC F_.254_
JULY 1982
lOOI.--
..I
N 90O,.<
w 80V')
70_--I<t-,
_ 60
50
FIXED VOLUME
II
I100 200
PC, PSIA
TURBOPUMP
PRESSURE FED
(LOX/PROPANE) FIXED TOTAL IMPULSE(IO0%)CRYOGENIC FEED
3700
3600 -M
3500 ..... l
3400 -- ---- ..._ I_
3300100 200 300
PC, PSIA
300
1200 F IXE_D_TOT_AL.IMPULSE (100%)
........... #II00 ........
1000 ...... "-'--"F-',r-
_" 900 ........f
s
" 800 ....
11
.......f
Jp ....
__JI
I I
200 300
PC, PSIA
lO0IO0
FIGURE 24 RCS PC SENSITIVITY (PUMP AND PRESSURE FED SYSTEMS)
55
MCDONNELL DOUGLAS ASTRONAUTICS COIffPANY.BT. L_UIB DIVISION
LOX/HYDROCARBON
AuxiliaRy Propulsion SYSTEM STudyFINAL REPORT REPORT MDC E2548
JULY 1982
systems are more competitive in terms of _might and performance than at 150 psia.
Because of the competitiveness of the two systems the pump versus pressure fed
option was selected for further evaluation in Phase II. Also, it was decided to
consider electric RCS pumps in order to increase system performance and decrease
feed system complexity.
4.2 NBP Versus Subcooled Fuel Storage
As discussed in Section 2.2.2, propane has a low freezing point and can be
stored at L0X temperatures with a 25% increase in density. This increased density
offers the potential for increased&V capability, Initial evaluations showed that
only a limited 0MS &V advantage could be achieved with subcooled propane storage
because of the constraint of equal fuel and oxidizer tank lengths (Table Vl).
Because of the high LOX/propane mixture ratio requirement the 0MS LOX tank was near
its maximum allowable diameter. Therefore, even though the propane volume could be
reduced with subcooling, the oxidizer tank diameter could not be increased
sufficiently to provide enhancedAV capability. As a result an alternate design ap-
proach was considered in which the pod aft bulkhead was divided to allow
lengthening the L0X tank. To accommodate the longer L0X tank the 02 turbomachinery
and accumulators were relocated to the fuel side of the pod as shown in Figure 25.
Figure 26 compares the OMSAV capabilities for NBP and subcooled propane storage and
shows a substantial AV improvement with the extended L0X tank. ThisAV enhancement
r_ust _e balanced against the increased thermal control complexity associated with
_ubr,_oled propane storage (cryogenic fuel tank as well as oxidizer tank) and the
es _cted access to the accumulators and turbopumps fnr maintenance as a result of
t_lecompact packaging on the fuel side of the pod.
: _ CryogeniG Versus Ambient Temperature RCS Prop_' ant Feed
As discussed in Section 2.2.3 ambient temperature propellant feed eliminates
the need for insulation on the RCS accumulators and feedlines but reduces overall
system performance as a result of ISp penalties associated with propellant thermal
conditioning. As shown in Figure 27 the ambient temperature feed system for the
LOX/ethanol, LOX/propane, and LOX/ammonia RCS is a hybrid concept in which gaseous
56
MCDONNELL DOUGLAB ASTRONAUTIC,S COMPANY.BT. LOUIS DIVISION
I
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
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REPORT MDC F..2548
JULY 1982
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57
MCDONNELL DOUGLAS ASTROIYALITIC3 COMPANY-ST, LOUIS DIVISION
LOX/HYDROCARBON I
Auxiliary Propulsion System StudyFINAL REPORT EPORT MDC F.2548
JULY 1982 i
' <=i",¢_"_'_ORIGINAL _ • • -
OF POOR 0U/aLI'i'Y I
I...... CURRENT OMS
I
FIXED VOLUME. |
'°°F _ I
,oo r ,
1,1,.i ,_j
.. 500I- ....._©.x.'_',..'l.............................................. II'-- .. cl_ .....
I _._ _ _ I_- / \_.,,] <,-> % x \\\° 400 I- I :o",CJ'_ ,_> o \\\ _ _=
/ I "" I _' J _ _ _ ,.,,'l I/ =" I _ _ ---- = _ _300 / I/ I
PUMP-FED PRESSURE-FED I
FIGURE 26 OMS _V COMPARISON FOR NBP AND SUBCOOLED PROPANE STORAGE
I
58
lifCDONN£LI. DOUGLAS AI_ITRONAUT'ICI! COlttPAN1Y . iT. LOUlil IDIVI_ilOIV
I
!
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
oF,, PO0_ Q;JA__.I_"_,"
REPORT MDC E2548
JULY 1982
• SEE SCHEMATIC NOMENCLATURE OF FIGURES 2 & 3
LOX_F :"-" C2H50HUEL C3H8
NH3
FIGURE 27 AMBIENT PROPELLANT TEMPERATURE FEED RCS
SCHEMATIC (GASEOUS 02/LIQUID FUEL)
59
MCDONNELP. DOLI_.LAS 4$TRONAUTICS COMPA_Iy.ST. LOUlli DIVIBION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
REPORT MDC E2548
JULY 1982
oxygen and liquid fuel are delivered to the thrusters. Oxygen thermal conditioning
is achieved in a tube-in-shell heat exchanger in which the hot side is supplied with
combustion products from a separate gas generator assembly. The use-of a separate
gas generator allows independent development of the heat exchanger and turbopump
assemblies. In the LOX/methane ambient temperature RCS feed system both oxygen and
fuel are thermally conditioned to a gaseous state as shown in Figure 28. The
performance and weight sensitivities of the gas/liquid and gas/gas ambient
temperature feed systems to RCS chamber pressure are presented in Figure 29. As
shown, the near optimum RCS chamber pressure for both systems is 250 psia.
The cryogenic and ambient propellant tempercture RCS feed systems are
compared in Figure 30 in terms of weight and performar,ce. As shown, the ambient
temperature feed systems are heavier and provide lower total impulse capability as
a result of Isp penalties associated with propellant thermal conditioning. The Isp
penalties are greatest for the LOX/methane system because of the requirement to
thermally condition both the oxidizer and fuel. The LOX/ethanol RCS is the most
attractive ambient temperature feed system because of its higher total impulse
capability. The thermal conditioning energy requirement is lower for the
LOX/ethanol RCS as a result of its lower thruster mixture ratio (1.4 for
LOX/ethanol as compared to 2.2 for LOX/propane and 2.4 for LOX/methane). Because
of this LOX/ethanol ambient temperature RCS feed systems were selected for further
evaluation in Phase II. In addition it was decided to consider passive 02 thermal
conditioning approaches to improve system performance and reduce weight.
4.4 Common Versus Separate OMS/RCS Tanks
I
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I
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Common propellant tanks were evaluated for the OMS and aft RCS, as well as the
OMS, aft RCS, and forward RCS, in order to increase propellant utilization
ilexibility and reduce the number of system components.
4.4.1 Common OMS-Aft RCS Propellant Tanks - A schematic for an OMS-aft RCS with
common propellant tanks is presented in Figure 31. The entry sumps, which are
located just downstream of the main propellant tanks, are sized to provide
propellants to the aft RCS thrusters during the entry phase of the mission. They
are filled to capacity prior to launch and remain full during the orbital phase.
6O
MCDONIVELL DOUGLAS ASTRONAUTICS CO.tt4PANY-STo LOUIS DIVISION
I
!
I
I
I
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
0_. POL_F! """ ' " "
REPORT MDC F..2548
JULY 1_.
• SEE SCHEMATIC NOMENCLATURE OF FIGURES 2 & 3
CH4
FIGURE 28 AMBIENT PROPELLANT TEMPERATURE FEED RCSSCHEMATIC (GASEOUS 02/GASEOUS FUEL)
61
MCDONNELL DOUGLA_ ASTRONAUTICS COMPANY-ST. LOUIS DIVISION
LOX/HYDROCARBOIN
Auxiliary Propulsion System Study
FINAL REPORT
Ol_poor QUALITY
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JULY 1982
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MCDONNELL DOUGLAS aISTR'OF#ALITIC8 CO_fPANt'oST. LOUIS DIVISION!
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT REPORT MDC E25A8
JULY 1982
ORIGINAL PA,2_ _
OF POOR QUALITY,
100
9O
D..
80-
..t
_- 7o
60-
50
460( FIXED TOTAL IMPULSE (100%)
FIX(D VOLUME
._ "T" q
-t- ._ i "r"
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4000
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3400
320(
w.
_1 _'_" oo
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m
mm _
AMBIENT
FIXED TOTAL IMPULSE (100%)
1400 F
1200 • • o ,
-r" .:r "-t-
O0
_ - - r,_ ::z 0,_ i -r.
8001 •-" _ _ ,._ =,-600 ....
CRYOGENIC AMBIENT
CURRENT RCS
PC = 250 PSIA
(PUMP FED)
FIGURE 30 COMPARISON OF CRYOGENIC AND AMBIENT TEMPERATUREPROPELLANT FEED RCS
63
MCDONiV£LL DOUGLAS 4$TROItI41.i'rICB COMP_INV.BT. LOUIS OIVISI()N
LOX/HYDROCARBON IAuxiliary Propulsion System Study
OF PooP, _J_" :' "
!
• SEE SCHEMATIC NOMENCLATURE OF FIGURES 2 & 3 I
!
ENTRY SUMPS LIQUID !
ACCUMULATORS
TURB( p i
• GAS ( ) IGENERATOR
- , |
I
VERNIER tTHRUSTERS IL
PRIMARY THRUSTERS i
FIGURE 31 SCHEMATIC FOR OMS-AFT RCS WITH COMMONPROPELLANT TANKS
I
I
I
64
MCDONNELL DOU_iLA,S ASTRONAUTICB COMPANY- ST. LOUIS DIVIBION
I
I
LOX/HYDROCARBON
AUxiliary Propulsion System Study
FINAL REPORTRI=pORT MDC E2548
JULY 1982
Weight and performance comparisons for the OMS/aft RCS with common and separate
propellant tanks are shown in Figure 32. The most significant advantage of common
tanks is the ability to provide-t00% of the current aft RCS total impulse
requirement while still exceeding the OMS AV capability of the-current storable
system. The low volumetric mixture ratio for the common LOX/ethanol system permits
efficient propellant tank packaging within the pod and affords the highest _V-total
impulse capability of the candidate propellants. The common systems also provide
a substantial reduction in wet and dry weights. Because of these advantages common
OMS-aft RCS tanks were baselined for the Phase II evaluations.
4.4.2 Common OMS, Aft RCS, and Forward RCS Propellant Tanks - To evaluate common
propellant tanks for the OMS, aft RCS, and forward RCS, a feedline model was
developed for interconnecting the forward and aft pods (Figure 4). Propellant is
transferred through this feedline from common tanks in the aft pod to RCS
accumulators in the nose pod. The advantage of this approach is increased
propellant usage flexibility. The disadvantage is the loss in propellant storage
volume in the nose pod and the complexity of installing the interconnecting
feedline.
Comparisons of separate and interconnected forward and aft propulsion systems
are presented in Figure 33 for all four fuel candidates. (A turbopump, cryogenic
propellant delivery system was assumed for the comparisons.) Because of the loss
of propellant volume in the nose module the OMS AV capability is substantially
reduced for the integrated systems. Only the LOX/ethanol system provides an OMS&V
in excess of the current capability. As such, it was decided to consider conical
propellant tank shapes in Phase II as a means of increasing propellant volume
capability for the interconnected forward and aft propulsion systems.
4.5 Tank Insulation Options
Foam, fiberous, and multi-layer (MLI) insulation concepts were evaluated for
separate OMS, separate RCS, and common OMS-aft RCS tanks. The intent of these
evaluations was not to develop a detailed insulation system design but to determine
the feasibility of candidate insulatiun materials. The LOX tank was selected for
these evaluations since LOX has the lowest storage temperature and lowest heat of
65
MCDONNELL. _DUGLA8 ASTRONAUTICB COMPANY-ST. LOUIS DIVIBION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORTREPORT MDC E2548
JULY 1982
I
I
I
600 "
5O0V3
=" 400
300
18000
=IC
. 1700'
,z211
1600('-de
15000
÷
_ 14000
. CRYOGENIC FEED SYSTEM!
SEPARATI
FIXED VOLUME
Z i
Ii
TANKS
, N_N
\-_.\ -r\ _,
E = fN
i
COMMON TANKS
70
_- 60
5C
FIXED VOLUME I
IO0- ,,,,,\,, \ . .. .
SEPARAT I _NKS COMMON TANKS
3000
FIXED A VII I
: 2500
g _ooo
£UUUSEPARATE TANKS COM_ON TANKS
I
FIXED AVII I
SEPARATE TANKS
_i_I
Cf_.iMON TANKS
I
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I
.... CURRENT 01MSIRCS I
FIGURE 32 COMPARISON OF OMS-AFT RCS WITH COMMONAND SEPARATE PROPELLANT TANKS
I
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I66
MCDOIW/tiELL DOUGLAS A_TRONAUTICS COMPA_Y.If. LOLII_ DIVIIION I
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORTREPORT MDC E2,548
JULY 1382
ORIGINAL p,.,...- [,?,_.
OF POOR QUALF_'£
1200 -
I
W
._I000 .....
d
800
600J
SEPARATE FRCS
s- 600O
L,÷
w 4500
÷
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FIXED VOLUME
COMMON OMSIARCS TANKS- TOTAL A"5 SYSTEM
CRYOGENIC tEED SYSTEM38000
,..I
36000
L_m
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34000
s-
32ooeSEPARATE FRCS
"_ _ _i _, _ ¢_
INTEGRATED FRCS
m
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%
INTEGRATED FRCS
FIXED AVIIT55_
5000L =
SEPARATE FRCS
.,,mc,Lt._0 m _
INTEGRATED FRCS
--- -- CURRENT SYSTEM
FIGURE 33 COMPARISON OF SEPARATE AND INTERCONNECTEDFORWARD AND AFT PROPULSION SYSTEMS
67
MCDOIti/VELI. DOUGLAS 4STROfI#,4UTICS COMPANIK.BToLOUll DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORTREPORT MDC F...2548
JULY 1982
vaporization compared to the other propellants. The evaluations were performed
using the TKHEAT code described in Section 3.2 applying representative OMS-RCS
engine firing cycles for a 30-day mission.
The properties for the candidate insulation materials are shown in Table XII.
The MLl-exhibits the lowest vacuum thermal conductivity but requires a vacuum cover
(dewar-type tank) to prevent moisture degradation. The TG-15000 silica fiber
insulation is an attractive material because it is easier to handle and install
than MLI and is not susceptable to moisture degradation (does not require a vacuum
cover). TG-15000 insulation is currently employed on the aft pod internal
moldline. Its disadvantage is a higher vacuum thermal conductivity compared to
MLI.
The measure of tank insulation effectiveness is the propellant boil-off (vent)
loss that occurs as a result of environmental heating during the mission. Thirty-
day LOX vent losses for the three candidate insulation materials are compared in
Figures 34 through 36 for _ separate OMS tank, separate RCS tank, and common OMS-
aft RCS tank. As shown, verst losses with foam insulation are excessive, whereas,
vent losses with both MLI and TG-15000 are considered acceptable. The MLI provides
the lowest vent loss as a result of its low vacuum thermal conductivity. It is
noteworthy that the common OMS-aft RCS tank provides a lower vent loss than the
combined totals for separate OMS and RCS tanks. On the basis of these evaluations
both '_LI and TG-15000 were selected for further evaluation in Phase II.
a.6 Feedline Insulation Options
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Because of the poor vacuum performance of foam insulation only MLI and TG-15000
ir_ulation were evaluated for cryogenic RCS feedlines. The evaluations were
,¢formed using the FDLINE code described in Section 3.3 applying representative
_CS thruster firing cycles for 7 and 30-day missions.
Initially evaluations were performed for the current RCS manifold arrangement
in which the propellant usage is divided among four primary thruster manifolds and
one vernier thruster manifold. The results are shown in Figure 37 where LOX
temperature at the thruster valve inlet (for one primary thruster manifold) is
68
MCDONNELL DOUGLAS ASTRONALITIC$ COMPANY.ST. LOUIS DIVISION
I
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORTREPORT MDC F..2548
JUuY 1982
w'
OF POOR QUALITY
TABLE Xll
PROPERTIES OF CANDIDATE INSULATION MATERIALS
INSULATION MATERIAL
TG-150OO FIBROUS INSULATION (4)
NRC-2 SINGLY ALUMINIZEDMYLAR MLI (50 LAYERS/IN)WITH 5% PERFORATION
AMBIENT (1)THERMAL (3)CONDUCTIVITY
BTIJ/(HR-FT-°R_
0.0123
0.05
EVACUATED (2)THERMAL (3)CONDUCTIVITY
BTU/(HR-FT-°R)
0.00075
0.000038
HEAT CAPACITY
BTU/(LBM-°R)
0.2
0.27
(3)
(I) GROUND HOLD CONDITIONS, PRESSURE = 14.7 PSIA.(2) ORBIT CONDITIONS, PRESSURE = VACUUM(3) PROPERTIES EVALUATED AT A MEAN TEMPERATURE OF 180°R.(4) TG-ISO00 INSULATION IS EMPLOYED ON THE ORBITER APS POD INTERNAL SURFACE.
DENSITYLBM/FT 3
2.0
1.14
69
MCDONNELL DOUGLAS ASTRONAUTICS COMPANY-ST. LOUIS DIVISION
LOX/HYDROCARBON
Auxil.iARy PRoputsiON SYSTEM STudy
FINAL REPORT
REPORT MDC E25,t',8
JULY 1982
Y
I
t
• INITIAL L0X LOAD _' 9350 LBM
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MISSIQN TIME, HR ,I0'
+ FBAM
A FIB
O NLI
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FIGURE 34 OMS LOX TANK INSULATION COMPARISONS
7O
MCDONNELL DOUGLAS ASTRONAUTICS COIt4PANt" -ST. LOUIB DIVIBION
!
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Auxiliary Propulsion System Study
FINAL REPORT
i '!
REPORT MDC E25_B
JULY 1982 B
• INITIAL LOX LOAD "_ 1650 LBM
_,
b°=:)
rn_J
°_
ILll---
ZUJ
..J
0t_
L
I 0 213 30 40 50 60 ?0 80
MISSION TIME, HR , I0'
+ FeAM
" FIB
O MLI
FIGURE 35 RCS LOX TANK INSULATION COMPARISONS
71
MCDONNELL DOUGLAS ASTRONALITIC_ COMPAI'_tY_ . ST. LOUIS DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT REPORT MOC E2548
JULY 1982
I
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e--
133
.-,I o
E_I..,UI---
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• INITIAL LOX LOAD ll,O00 LBM
lJ +
jJ .-I"-"
40 50 60 70
MISSION TI!_, HR ,10'60
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I+ FeAM
A FIB I
(D 141.]
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IFIGURE 36 COMMONOMS-RCS LOX TANK INSULATION COMPARISONS |
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MCDONNELL DOUGLAS AI'rI_oNAI.JTICS COMPI4N Y o IT. LOtLIIS DIVISION
I
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
I_._.,... ,.,_..'.
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REPORT MDC F.2548
JULY 1982
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73
MCDONNELL DOUGLAS ASTRONAUTICN COMPANV.IIT. LOUI_i DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT REPORT M)C F..2.548
JULY 1982
plotted as a function of mission time for both MLI and TG-]5000 insulation. The
plot shows that the LOX vaporization limit is exceeded for both insulation
materials. (The vaporization limit corresponds to the temperature at which
vaporization will occur in t'he injector for a 250 psia chamber pressure thruster.)
Since lower LOX feedline temperatures can be achieved with h_gher LOX usage
rates, this analysis was repeated for a re-configured manifeld arrangement in which
all the RCS propellant is consumed through a single thruster manifold having the
required number of primary and vernier thrusters for Orbiter three-axis attitude
control. (Should a thruster failure occur with this manifold arrangement--i.e.,
failed open thruster valve--the primary manifold would be isolated and a back-up
manifold activated.) A LOX temperature history for this reconfigured manifold
arrangement is presented in Figure 38 for a 7-day RCS mission. This plot is for the
LOX temperature adjacent to the thruster valve, which is the warmest point in the
feedline. As shown by this example one-inch of TG-15000 insulation provides
acceptable LOX feedline temperatures throughout the mission. As a result the
reconfigured manifo!d arrangement was baselined for all subsequent feedlinethermal analyses.
A summary plot of maximum LOX feedline temoeratures as a function of usage rate
and accumulator temperature is presented in Figure 39 for TG-15000 insulation. At
the usage rate associated with a 7-day mission the LOX vaporization limit is
exceeded for the 200OR accumulator temperature. For the low usage rate associated
with a 30-day mission the LOX vaporization limit is exceeded for all three
accumulator temperatures. On the basis of these results it was concluded that the
TG-15000 insulation provides inadequate thermal protection for the RCS LOX
feedlines. Similar data is presented in Figure 40 for one-inch of MLI. Even for
the low usage rate associated with a 30 day mission LOX temperatures are maintained
below the vaporization limit for a wide range in accumulator temperature (160 to
200°R). Based on these analyses it was concluded that a vacuum-jacketedMLl system
is required to maintain acceptable temperatures in the RCS LOX feedlines.
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MCDONNELL DOLIGLaIli ASTRONAUTICS COMPANY -SY. LOLIOB DIVOBION
I
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LOX/HYDROCARBON
Auxiliary Propulsion System St'Jdy
FINAL REPORT
ORI(3-1r.'_,L ;.,'-._.::: ' j
OF POOR ':'" ' :'£:.,'#
REPORT MDC E2548
JULY 1982
°iW=O
Or,,,.,-,-,- _I.-- I--
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75
MCDONNELL DOUGLAS ASTRONAUTICS COMPANV.ET. LOUIS DIVISION
LOX/HYDROCARBON IAuxiliary Propulsion System Study
OFPoorQ_Au_ t
!• PRESSURE : 500 PSIA
• I INCH TG-.15000 INSULATION
,_-- ,_ LOX TEMP, AD,Jt_A__N.I_.I_Q THRUSTER VALVE !
I=o. , |
30-D Y VERNI'R ._ I USAGE RATE
_u __ 7-DAY V__RNIER J i
_\- l\ \j / I::E -- LOX VAPORIZATION LIMIT FOR 250 PSI CHAMBER PRESSURE THRUSTER
c)
hr"
L'UO0-I.--OO_
II---
0CD
2 ) 4 5LOX USAGE RATE, LBM/HR
FIGURE 39 EFFECT OF LOX USAGE RATE ON FEEDLINE
TEMPERATURE RESPONSE (TG-15000 INSULATION)
ACCUMULATOR I
L0X TEMP.I
+ 200 OR
" 18o OR I
(I) 160 OR
,
|
I
I
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MCDONNELL DOUGLAS ASTRONAUTICS COMPANY 111" LOUIS L_IVlSlON
I
I
LOX/HYDROcARBON
Auxiliary PropulsiON System Study
FINAL REPORT ORIGINAL "_ " _
OF POOP qUP:LITI
REPORT MDC 52548
JULY 1982
.... m.
PRESSURE = 500 PSIA
INCH MLI
01 TEMP_ ADJAKENT. TO THRUSTER VALVE
,-,, 1
LLIo
I 30-1_AYVERNER
_ \_ 7 DAY_ERNIER
.LIMI_ -. 250 PSI CHAMBER
PRESSURE THRUSTER
ACCUMULATORLOX TEMP.
+ 20o °R
,',i8o OR
(D 160 OR
"13 _ 2 3 4 5L_X USAGE RATE, LBM/HR
FIGURE 40 EFFECT OF LOX IJSAGE RATE ON FEEDLINE
TEMPERATURE RESPONSE (MLI)
77/78
MCBONNELL DOUGLAS ASTRONAUTICS COMPANY. BT. LOUIS DIVISION
LOX/HYDROCARBON
AuxiliaRy Propulsion System Study
FINAL REPORT
5.0 TASK II.1--PHASE II GROUNDRULES
REPORT MDC F_.2548
JULY 1982
The results from the Phase I effort, described above, formed the groundrules
for the Phase II effort. These results are summarized below.
On the basis of the Phase I evaluations,.two fuel candidates were selected for
Phase ll--ethanol and methane. Ethanol is a storable fuel with a vapor pressure
slightly higher than MMH, while methane is a cryogenic fuel. Both are non-coking
and offer high performance capability. LOX/ethanol affords the highest total
impulse capability for systems constrained to the current pod volume because of its
higher bulk density-specific impulse product. LOX/methane affords the lowest
system wet weight (highest payload capability) for systems sized to a fixed total
impulse requirement because of its high engine specific impulse (Isp).
A pump fed OMS was baselined for Phase II because it offers overriding weight
and performance advantages compared to a pressure fed system. A gas generator
cycle engine was selected for LOX/ethanol, whereas an expander cycle engine was
selected for LOX/methane. (The expander cycle allows higher Isp for the
LOX/methane OMSengine.) In addition, a single turbine drive for both the fuel and
oxidizer pumps was selected to reduce the number of components and provide low
engine system weight,
Pump and pressure fed feed system options were more competitive in terms of
weight and performance for the RCS than for the OMS because of the lower RCS total
impulse requirement. As a result, both pump and pressure fed RCS were selected for
further evaluation in Phase II. However, battery powered electric pumps were
baselined for the pump fed RCS to eliminate the Isp penalty associated w_th
turbopumps and to reduce the number of feed system components.
Common OMS/ARCS propellant tanks were baselined for Phase II because they
offer improved propellant packaging in the aft pods (higher OMS _V and RCS tota]
impulse), reduce feed system weight, and provide greater flexibility in the
utilization of OMS/ARCS propellants compared to separate propellant tanks.
Because of these advantages fully integrated tankage systems for the OMS, ARCS, and
79
MCDONNELL OOLIGLAS 4STRONALITICB COP_PANY.mT. LOUlJ DIVIIIION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINALREPORT RePoRt MOc e2s_
JULY 1982
FRCS were selected for further evaluation in Phase II considering the use of
conical propellant tanks for improved aft pod packaging.
A cryogenic RCS feed system was baselined for LOX/methane, however, an ambient
propellant temperature RCS feed system was baselined for LOX/ethanol because of the
lower energy requirement for thermal conditioning. (Only one propellant--LOX--
must be thermally conditioned-_n the LOX/ethanol RCS.) The advantage of the
ambient temperature RCS feed system is the elimination of insulation and thermal
control requirements except for the storage tanks and pumps. Furthermore, because
of the low mixture ratio (1.4) for the LOX/ethanol RCS thrusters, it was felt that
passive heat exchanger approaches (without the use of fuel rich hot gas 02 heat
exchangers) would be feasible for thermally conditioning the RCS oxygen flow. As
such, two passive thermal conditioning approaches for the LOX/ethanol RCS were
selected for Phase II evaluation. The first employs an ethanol feedline heat
exchanger for gasifying the oxygen flow while the second employs an ethanol tank
heat exchanger. In the first concept (Figure 41), the fuel flow is pre-heated using
a hot gas heat exchanger, and then the heated fuel flow is used to vaporize the 02
in apassive feedline heat exchanger. In the second concept (Figure 42), the oxygen
flow is circulated through a heat exchanger coiled around the outside of the
ethanol tank where it absorbs heat from the tank wall, the environment, and theethanol within the tank.
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The final concept variable selected for Phase II evaluation was the choice of
insulation materials for the cryogenic LOX/methane tanks and RCS feedlines. On the
asl of the Phase I results for LOX, two types of insulation materials--aluminized
myl_:" multi-layer insulation and TG-15000 silica fiber insulation--were selected
or further evaluation in Phase II.
I
IIn accordance with the above discussion the Phase II groundrules are
immarized in Table XlII. The system design requirements and constraints for
component weight and sizing, which are similar to those established in Phase I, are
summarized in Table XIV. One notable change to the Phase I component weight and
sizing groundrules was that the minimum gage for tank and accumulator wall
thickness was increased to 0.060 inch because of the concern that the Phase I values
of 0.030 inch for aluminum vessels and 0.020 inch for titanium vess,_Is were very
susceptible to handling damage.
8O
MCDONNELL DOUGLAS ASTRONAu_'ICS COMPANY._ LOUIB DIVISION
!
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
REPORT MDC F.2.548
JULY 1982
ETHANOL FROMPROPELLANTTANK
-FUEL-RICH BIPROPELLANT
GAS GENERATOR
ETHANOL PASSIVE/---HEAT EXCHANGER /--ETHANOL FEEDLINE
t / /HEAT EXcHANGEr
I VENT t-_W_A/_./kl , GASEOUS 02 TO
I' .... I' RcSAcCUMuLAtOrI
LOx FROM I LIQUID ETHANOL TOPROPELLANT TANK _ RcS ACCUMULATOR
FIGURE 41 PASSIVE ETHANOL FEEDLINE 02 HEAT EXCHANGER CONCEPT
LOX FROMPROPELLANT TANK
fGASEOUS 02 TORCS ACCUMULATOR
.,,--PASSIVE ETHANOLTANK HEAT EXCHANGER
LIQUID ETHANOL TORCS ACCUMULATOR
FIGURE 42 PASSIVE ETHANOL TANK 02 HEAT EXCHANGER CONCEPT
81
It_eCDONNELL DOUGL AS A_YRONAUTICS COII_IfPANY- II11". LOUIS DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
TABLE Xlll
PHASE II GROUNDRULES SUMMARY
I. Fuels:
• ethanol
• methane
REPORT MDC E2548
JULY 1982
II.
III.
Baseline Feed System Constraints
• pump-fed OMS
- single turbine drive for both fuel and oxidizer pumps
- gas generator cycle for LOX/ethanol
- expander cycle for LOX/methane
• common propellant tanks for OMS/ARCS
• cryogenic propellant tanks for OMS (LOX and methane)
• cyrogenic propellant feed for LOX/methane RCS
• ambient temperature propellant feed for LOX/ethanol RCS
- ethanol: liquid phase
- oxygen: gas phase
Feed System Options to be Evaluated
• propellant tank insulation options for LOX and methane
- aluminized mylar multi-layer insulation (MLI)
- TG-15000 silica fiber insulation
• RCS feedline insulation options for LOX and methane
- aluminized mylar MLI
- TG-15000 silica fiber insulation
• turbopump versus electric pump-fed RCS (LOX/ethanol)
• passive 02 thermal conditioning options for LOX/ethanol RCS
- ethanol feedline heat ex_hanger
ethanol tartk heat exchanger
pump versus pressure fed FRCS (LOX/methane)
separate versus common FRCS/aft propulsion tanks (L(_X/ethanol)
conventional versus conical aft propulsion tanks (LOX/ethanol)
82
MCDONNELL DOUGLAB ASTRGNALITICli COMPANY-lIT. LOt,lIB DIVIBION
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LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
REPORT MDC E2548
JULY 1982
TABL_ XIV
PHASE II DESIGN REQUIREMENTS/CONSTRAINTS FOR COMPONENTWEIGHT AND SIZING
I • Helium Pressurization System
• common helium supply for fuel and oxidizer tanks
• current OMS/RCS line lengths
• ling Mach number = 0.I (Maximum)
• real gas effects
• solubility effects
• propellant vapor pressure effects
• line material: 304L stainless steel (SS)
• polytropic exponent = 1.0 (helium bottle inside LOX tank)
• regulator pressure ratio = 0.7 (outlet/minimum inlet)
• tank shape: spherical
• tank material: 2219-T87 aluminum (AI)
• storage pressure: 3000 psia
• ultimate factor of safety for helium tank : 1.5
II. Propellant Tanks
• propellant dumped overboard during an abort
• tank volume determination
- impulsive propellant volume
- 2% liquid residuals by volume
- 98% vapor residuals by volume
- tank boil-off loss (LOX and methane)
- OMS feedline chilldown/vent los (LOX and methane)
- 5% ullage volume at storage temperature
• Common OMS/ARCS tank shape
cylindrical with oblate spheriod end domes, or
conical with oblate spheriod e_L,i domes
Common OMS/ARCS tanks are constrained to equal lengths to permit
attachment to common aft pod bulkhead
FRCS tank and entry sump shape: spherical
83
MCDONNELL DOUGLAS ,45TRONALITICS COMPANY-BT. LOfJI8 DIVISION
LOX/HYDROCARBON
Auxiliary Propulsio_ System StudyFINAL REPORT
REPORT MDC E2548
JULY 1982
!ABLE XIV (Continued)
PHASE II DESIGN REQUIREMENTS/CONSTRAINTS FOR COMPONENTWEIGHT AND SIZING
materials
- LOX: 2219-T87 A1
- fuel: 2219-T87AI or 6A1-4V titanium (Ti) (whichever is lighter)minimum gage thickness: 0.06 in.
ultimate factor of safety : 1.5
insulation options (LOX and methane)
- aluminized mylar MLI with thermodynamic vent system
- TG-15000 silica fiber insulation with thermodynamic vent system
propellant acquisition: surface tension screens
OMS propellant gaging: capacitance probes
RCS propellant gaging: P-V-T
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III.
IV.
RCS Accumulators
• sized to provide Shuttle External Tank separation impulse withoutresupply
• shape: spherical
• blowdown accumulator operation (isentropic blowdown process)• materials
- LOX: 2219-T87 A1
- fuel: 2219-T87 A1 or 6A1-4V Ti (whichever is lighter)• minimum gage thickness: 0.06 in.
• ulitmate factor of safety : 1.5
• insulation options (LOX and methane)
aluminized mylar MLI without vent
- TG-15000 silica fiber insulation without vent
• propellant acquisition for liquid accumulators: surfac_ tension screens
Propellant Feedlines
• current OMS/RCS line lengths
• pressure drop criteria:
- 0.5 psi/ft for pressure-fed system
- 1.0 psi/ft for pump-fed system
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MCDONNELL DOUGLAs ASTRONAUTICS COMPANy.sT, LOUIS DIVISION
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Auxiliary Propulsion System Study
FINAL REPORT
REPORT MDC E2548
JULY 1982
TABLE XlV (Continued)
PHASE II DESIGN REQUIREMENTS/CONSTRAINTS FOR COMPONENTWEIGHT AND SIZING
B
, Darcy friction factor
• isenthalpic expansion process
• material: 2219-T87 A1
• minimum gage = 0.028 in.
• ultimate factor of safety:
- 4.0 for diameters < 1.5 in.
- 1.5 for diameters > 1.5 in.
• linear and angular compensation joints
• insulation options for RCS feedlines (LOX and methane)
- aluminized mylar MLI
- TG-15000 silica fiber insulation
Vo Gas Generatcr Exhaust Vent Line
• line Mach number = 0.3 (maximum)
• Fanno line analysis
• line length: 20 ft
• exhaust nozzle area ratio = 2.0
• propulsive vent for OMS; nonpropulsive vent for RCS
, line material: 304L SS
• minimum gage and ultimate factor of safety: same as feedlines
85/86
It_C£)ONNELL DOUGLAS ASTRONAUTICS COMPANY-ST. LOeJIB OIVISlON
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
6.0 TASK II.2--PHASE II COMPONENT CHARACTERIZATION
REPORT MDC E2548
JULY 1982
The three computer codes described in Section 3.0 were used to perform the
system evaluations of Task 11.3. Changes that were made to support the Phase II
effort are described below.
6.1 APSDS Code
This code sizes the overall propulsion system (OMS and RCS) to a fixed pod
volume or fixed total impulse constraint. It was changed in Phase II to include
revised performance and weight models for the OMS and RCS engines and new models for
battery powered electric motor pumps for RCS propellant feed.
6.1.1 OMS and RCS Engine Models - The data for the OMS and RCS engine models were
developed by ALRC and are provided in th_ appendicies of Reference (2). Engine
lengths and diameters were constrained to the current values. The OMS engine was
based on fuel regenerative cooling and a single turbine for driving both the fuel
and oxidizer pumps. A gas generator cycle was selected for LO×/ethanol while an
expander cycle was selected for LOX/methane to provide maximum performance. In the
expander cycle, gaseous methane leaving the engine cooling jacket is used to drive
thetu_-bine. The methane exiting the turbine is then routed directly to the engine
injector to avoid the vent loss associated with the gas generator cycle.
LOX/ethanol OMS engine specific impulse (_SP) is shown in Figure 43 as a
function of chamber pressure (Pc) for both zirconium-copper (Zr-Cu) and nickel (Ni)
chambers. The decrease in Isp with chamber pressure for the Ni chamber is the
result of high supplementary film cooling losses. Supplementary film cooling is
not required in the Zr-Cu chambers up to a chamber pressure of 600 psia, and only
0.7% film cooling is required at 800 psia. Because of its high performance
capability a Zr-Cu chamber was baselined for the OMS engine.
Similar parametric data for the LOX/methane OMS engine was not developed since
an energy balance could not be achieved with the expander cycle at chamber
pressures greater than 400 psia. (At the higher chamber pressures there is
insufficient thermal energy transferred to the methane in the cooling jacket to
87
MCDONNELL DOUGLAS ASTRONAUTICS COMPANY.lIT. LOlJllI DIVllIION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
ORIGINAL p,c.C_.%
OF POOR QIJALFF_
THRUST = 6000 I.BF
REPORT MDC E2548
JULY 1982 !
t
I
m,,,,,,,,mm
360- ,
350-
340
330 _
320-
310
o ZR-CU_RNI CHAIR
]
I..... [......
• I
[
I
t--HJ
..I II l
r
L ' I , (___._
! I .....
I
i
300,I t
400
i ! ,
r , i _ i.....4- J ', i I
j ii j! j|j500 600 700 800 900
C_ PC'PSTA
FIGURE 43 LOX/ETHANOL OMS ENGINE SPECIFIC IMPULSE
88
MCOOItlItI£LL DOUGLAS ABTRONAU rlC$ CGMPAIVY.IT, LOUIB DIVtSlOItl
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Auxiliary Propulsion System StudyFINAL REPORT REPORT MDC E2.548
JULY 1982
drive the turbine.) As a result a chamber pressure of 400 psia was baselined for
the LOX/methane OMS engine, providing a specific impulse of 364 seconds for a Zr-Cu
chamber. Despite the lower chamber pressure capability with the expander cycle the
LOX/methane OMS engine specific impulse is substantially higher than that achieved
with the LOX/ethanol gas generator cycle engine.
RCS engine specific impulse as a function of chamber pressure is presented in
Figures 44 and 45 for LOY/ethanol and LOX/methane, respectively. These data are
based on film-cooled Zr-Cu chambers. The 550 Ib-thrust chambers have a higher Isp
than the 870 Ib-thrust chambers because of the fixed envelope constraint which
permits a larger nozzle (higher expansion ratio) for the lower thrust level.
6.1.2 Electric Motor and Pump Weights - Electric motor operated RCS pump weights
were also generated by ALRC (see Reference (2) appendicies) for incorporation into
the APSDS code. The electric motor weights were based on an alternating current
design to provide minimum weight. The total weights for the motor and pump are
presented in Figures 46 through 48 for LOX, methane, and ethanol, respectively. As
shown, weights were developed as a function of flow rate and discharge pressure.
The power demand for these pumps was based on the shaft horsepower requirements and
a pump efficiency of 90%. The corresponding battery weights for" meeting the RCS
total impulse requirement are shown in Figure 49 for both silver-zinc (Ag-Zn) and
lithium (Li) batteries. These weights were developed by NASA-JSC. The lithium
batteries require new technology development but were baselined for the study
because of their low weight.
6.2 TKHEAT Code
The TKHEAT code determines the thermodynamic response of propellant contained
in storage tanks or accumulators during representative orbital mission duty
cycles. The code was changed in Phase II to include the capability to analyze
methane and to provide a subroutine for evaluating a passive ethanol tank 02 heatexchanger.
The ethanol tank 02 heat exchanger model is shown in Figure 50. The 02 heat
exchanger line is coiled around the outside wall of the ethanol tank where it
89
MCDONNELL DOIJGL/i$ ASTROttlAUTIC8 COMPANY-lIT. LOILtlB DlVlSlOItl
RCEPC'PSIA
FIGURE 44 LOXIETHANOL RCS ENGINE SPECIFIC IMPULSE
90
MCI_OIIINELL DOUGL AI 4ITltOIIO4U TICIi COMPANV . IIT. LOUll I_lVl_IlOItl
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LOX/HYDROCARBON
Auxiliary P.pulsion System StudyFINAL REPORT REPORT MDC F.2.,548
JULY 1982
°.;
Ctl_t " , : "_"_F_
OF pOO'._ ,,_--"-
RCE THRUST= 870 LBF,
,3l,J.J
I
_31o: /iii :_J i i
,li, 1r ......... ( "; ..... ; ; .......
300_ i_
330 ! _ _ I
i = i i " i i 'I ' ! _ i
• .._ ...... z ...... 1....... i.. "
i , !320" ! : _ ii , , , . t- , .......... :
t .. :.... l'." "f ..... t .... i ......... ; ,
, i
.. i I_ .... l ...l .
, I
: jT
0 100
; : ; i
l " !t t290 _ I, I i I :
200 300
• tI
RCE PC' PSIA
I' [
I
.• ,
I
I
i
i
i •..
i
I
400 500
FIGURE 45 LOX/METHANE RCS ENGINE SPECIFIC IMPULSE
91
MCDONNELL DOtlGLAB 41T_O_AuflcI COMPANY.|T. LOUIS DIVISION
LOX/I_¥DROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
ORtGINAL F#,C_ ;'*
OF POOR QUALI'T'Y
REPORT MDC E2548JULY 1982
I
t
1
I
60.0
50.0
PU_DISCHARGE |PRESSURE,PSIA f
-1 , ! , . . . _ -- - " _ . , I ; _ _ ,L[.._,l"=__±. : _=_-.===-=_-_:-.......................... - : _
40.o _____:_ __=_.__::..... , ___:= _ I30.0 _ L-.::_--4 4 , 600 "
.... _ _ _ , : - , .........J" - : : : ' I - -- . ........
_o.o i---::::_T:-i::::-:-_-::---:_'_¢C=:-_ ::_:1, _ l ; [ !_ : , I
-, , '.,:, I-,_ ,_-_..;,- 300
10.0 ,," '. l'/ .l ., _-'- :'. 1'_I_ _.',1'.--_
I
7.0 _"-----__---T _:.... _ -=---='7_---: : { i: i : : ,-=--=_-.-'--._
-_--._,-==--==:-_-_ _ ill* _:: - !uH,,, 5.0 --=-:-_.,---::- .... ---L:__,.......... '" -----L:_:Y----T I
----:--:-:'_"'_=_-=-- ............. --_--_:-- } : : ; '-_ "-_-'_-1 2"--_T L-
4.0 -.2...... --:-= :-;-':---:=:-_.:E .:-:":-; -:-_:::.-=/;---_:_:_._--_ -:: ;;'t II........ ,; " " & , ; - T _ ........... i.... : ._ _. _2 _-- -r ............... _ ._ ' :I . 2_:
• _ :-:-_: _--- ...... _ .... _-..-_ .:.. I....... _ .................. _.--:_-: -_-r ..... ::----.. •3.0 __" -: ; -::_:;-: ;_ ................. L • ............ Z_L" • - • * ...... 4 ........_-_:_ L ,,, f_ " _!-.:--i:--
I2.0
0 2 4 6 8 10FL(_VRATE,LBSISEC |
FIGURE 46 L.OXELECTRIC MOTOR AND PUMP WEIGHT I
I
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Mcr-'JI_NNELt. DOUGLAB ASTRO,_AUTICB COtt4PANIY .IITo LOUIE DlVliilOItt
I
I
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
¢. - ,--..- °.'?
OF PO_f< QUALITY
REPORT MDC F...2.548
JULY 1982
200-._ " £_llllJ,.l __1.11: t"
_J--'_J : i i ! I ! J . . I : : : i ] : : ! ; ! ] I I
1_:1111]i l_[{l!,J/ 11 _ i_ _ i ;_C.Ji]!ll:: I 'I'''
,,_i_IJ.z_L._ ,_J__L__±I_ !. L :.1 .:_i .... ! ...... LJ._L_L! :I 'i_, i_ I I l. ii
___----/.J_.._L.._ !rl j! :_!liiJ_J!il!
J:,:i[i ]: :
|[ [ II I --;
ji_i. _i] :riurr i_LJ__ ii_LL_: ,
T il;" i_--,"
!li] i"l !_]
FLOW RATE,-LBS,/SEC,
ISCHARGEPSIA
1 000
60O
300
FIGURE 47 METHANE ELECTRIC.MOTOR AND PUMP WEIGHT
93
MCBOItlNEL L LIOUGLAli ,48TROItI_IUTOCli COMPAItlt' . Jrr. LOUlll DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
oRIG1_'.tL 'fi',:.,OF pOOR (_'_.j/.kL|'("/
REPORT MDC F..254_
JULY 1982
I
!
I
100.090.080.070.0 ...............
60.0
:_ : =:..
i)I_PRESSURE,PSIA
l(JO0
I
i
I
50.0 I
40.0 600
: --:: =_:: _ - - :---_T-i: : ..._ I30.0 -:---_-÷- --:--::- _ i- -_-::::
..... ' 30020.0
3.o I
2.0
o 2 4 6 8 lbFLOWRATE,I_BS/SEC
FIGURE 48 ETHANOL ELECTRIC MOTOR AND PUMP WEIGHT
I
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I94
MCDONNELL DOUGLAS ABTROI'tlAUTICII COMPANY-IT. L(I_II DIVtIIION I
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAl.REPORT REPORT MDC E2.54,8
JULY 1982
o OXYGEII
o ETI_IOL
360
340
320'
300 -
280
260
240
220
200
18o
160
140
120
lO0
80
60
4O0
(REDUI_DAI_TBATTERIES)
, . ' I
• : ' ; • : , • J 1: i : , : : IAG_ !1982.)! : • . : • , . ; _ • i.. ,...,: ,//f, 9 t,i
......... iJ_ '. _ I I
• I . _ " . ' ;•. .'. :.i . ' , _ , . ! i
• t ' /A/" i _ t i ' _, ://Y'* .... *.i. !.... ,-.; ,! //vL ' , _ i ,
.... ,'//_.i " ; .......... i . I. i .-.;
• " r .... • " ' " -. I • ., - - *
; , • . LI ( )!
500 600 700 800 gO0 l000 IlO0
PLI,P_aP,PSIA
FIGURE 49 BATTERY WEIGHTS FOR ELECTRIC RCS PUMPS
95
MCDONNELL IIOUGLA8 41"rROIV4LJTICIi Y.MT.
LOX/HYDROCARBON t
Auxiliary Propubion System Study RePort moc E2_8F_nauREPORT ' JULY 1982
C,F POC!:¢ QUALi'FY' j
i
• 02 HEAT EXCHANGER LINE ABSORBS HEAT FROM THE TANK,
HELIUM SU PPLY-.--:=:_ . I_1@._ G OX.
-._ _-_ RADIATION .......J_J.....:-"_ L j_HEAT LOX J_
EXCHANGER __ LIN E ELEMENT d
LINE __ I .... _'
TANK THERMAL MODEL LINE HEAT EXCHANGER THERMAL MODEL
IFIGURE 50 ETHANOL TANK 02 HEAT EXCHANGER MODEL
96
MCDOIII,VELL DOUGLAII ASTRONAUTICS COMPAftlY . 8"r. LOUIS DIVIIilOIV
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
REPORT MDC E2548
JULY 1982
absorbs heat from the tank wall, the environment, and liquid ethanol within the
tank. The 02 flow enters the heat exchanger as a cryogenic liquid and exits as a
superheated vapor. The heat exchanger line is divided into segments, and the
energy and mass conservation equations are solved for each segment. The subroutine
calculates the 02 exit temperature and liquid ethanol temperature inside the tank
during specified OMS-RCS mission duty cycles. A mass inventory is made to accourlt
for the decrease in ethanol quantity during the mission.
6.3 FDLINE Code
The FDLINE code computes the thermal response of propellant contained in the
feedlines during specified mission duty cycles. For the Phase II effort the code
was changed to include the capability to analyze methane and to provide an improved
RCS thruster heat soakback model.
The RCS feedline model was shown previously in Figure 14. The feedline from
the accumulator to the thruster valves is divided into segments, and the energy and
mass conservation equations are solved for each segment. The analysis accounts for
heat conduction through the insulation, thruster heat soakback, and major heat
leaks associated with structural members such as line supports. The thruster heat
soakback model is shown in Figure 51. The heat soakback is calculated based on the
thermal resistance between the injector and valve and the temperature difference
provided by Figure 51. (The thermal resistance was provided by ARLC--see
Reference (2) appendicies.) At the start of each thruster pulse, the temperature
of the thruster injector and valve are assumed to be at ambient temperature. Upon
ignition, the valve temperature decays rapidly as a result of propellant flow. At
shutdown the injector temperature increases due to heat transfer from the chamber
wall, and then the valve temperature begins to rise as a result of heat transfer
along the thermal standoff tube. At_l x_LO 5 seconds after shutdown valve and
injmctor temperatures are nearly equal. The FDLINE code maintains an inventory of
thruster pulses and the time-between firings and then applies the injector-valve
temperature difference given by Figure 51 to compute thruster heat soakback.
97
MCDONNELL DOUGLAS ASTRONAUTICS COMPANY-B'I'. LOUIB DIVISION
FIGURE 51 RCS THRUSTER HEAT SOAKBACK MODEL
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MCDONN£'LL DOUGL AS ASTRONAUTICS COMPANy . lIT, LOUlll DIVISION
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
7.0 TASK II.3--PHASE II SYSTEM EVALUATIONS
REPORT MDC E2.548
JULY 1982 Bv'
The evaluations performed in this task are summarized in Table XV. They
include design definition of the Phase II options identified previously in
Section 5.0 (Table XIII), design point sensitivity analyses, and side-by-side
comparisons of the most attractive LOX/HC systems with a LOX/H 2 system and the
current storable propellant OMS/RCS. System weight and performance data generated
with the APSDS computer code to support these evaluations were provided in the
Appendicies of Reference (2). The following paragraphs present the results and
conclusions derived from these evaluations.
7.1 Tank Insulation Evaluations
Methane and LOX insulation concepts were evaluated for common OMS/ARCS and
separate FRCS propellant tanks. Based on the Phase I results two insulation-
candidates were selected for evaluation--TG-15000 silica fiber insulation and
aluminized mylar multi-layer insulation (MLI). The properties of the candidate
insulation materials were shown in Table XII. The evaluations were performed using
the TKHEAT code described in Section 3.2 applying representative OMS-RCS engine
firing cycles for a 30-day mission.
Example TKHEAT code analysis results for a common methane OMS-ARCS tank are
presented in Figures 52 and 53. For this example a blanket consisting of 1.5 inches
of TG-15000 insulation was assumed. As shown in Figure 52 the tank pressure
increases slowly from an initial value of 35 psia to a tank relief pressure of
60 psia. At the relief pressure liquid methane is withdrawn through the propellant
acquisition system and routed through a thermodynamic vent system. The thermo-
dynamic vent system consists of a coiled tank heat exchanger inside the tank
insulation which absorbs heat from the environment. The environmental heat input
vaporizes the methane, and the methane vapor is then vented overboard. Methane is
vented until the tank pressure decays to 57 psia, and then the vent system is
shutdown automatically. This vent process is then repeated for the remainder of
the mission. The corresponding methane temperature history for this 30-daymission
simulation is shown in Figure 53. The temperature increases slowly as a result of
environmental heating during the first 200 hours of the mission. Then, during
venting, the temperature cycles between 223 and 220OR.
99
MCDONNELL DOUGLAS ASTRONAUTICB COMPANY-5_ LOUIE OlVlSlO_
LOX/HYOROCARBON
Auxiliary Propulsion SystEm StudyFINAL REPORT
ORIGINAL F'.._<',.; ',:.;
OF POOR (d'::-., ;V'_
TABLE XV
PHASE II SYSTEM EVALUATION TASKS
REPORT MDC E2.548
JULY 1982
I
I. Tank Insulation Evaluations (Methane and LOX)
2. RCS Feedline Insulation Evaluations (Methane and LOX)
3. GOX/Ethanol RCS Feasibility Evaluations
• Turbopump RCS proRellant feed with ethanol feedline 02 heat exchanger
• Electric pump RCS prooellant feed with ethanol tank 02 heat exchanger
4. LOX/Ethanol and LOX/Methane OMS-RCS Sensitivity Analyses
• OMS and RCS chamber pressure
• RCS accumulator blowdown pressure ratio
• OMS and RCS specific impulse
• propellant tank minimum gage thickness
5. Separate versus Common FRCS/Aft Propulsion Tanks.
6. Convertional versus Conical Prop_llant Tank Shapes
7. Pump versus Pressure Fed FRCS
8. Side-by-Side OMS/ARCS Comparisons
• LOX/ethanol
• LOX/methane
• LOX/H 2
• Current N204/MMH
i00
MCDONNELL DOUGLA8 As"rRoNAuTICIJ COMPANY.ST. LOUIS DIVISION
LOX/_ VDROCARBON
Auxiliary Propubion SystEm StudyFINAL REI_ORT
REPORT MDC F..2548
JULY 1982
0,-. _, • .i
IQ')O.
LIJ...-,.
"_ j,,_.1
I--
f
. 1.5 INCHES TG-15000 _NSULATION
• 30 DAY MISSION
• INITIAL METHANE LOAD = 3005 LBM• RELIEF PRESSURE = 60 PSIA
• ENVIRONMENTAL TEMPERATURE = 500OR
!
20 20 40 50
MI_-SION TIME. HR ,I0'6O 7O 6O
FIGURE 52 30..DAY PRESSURE PROFILE FOR COMMON OMS..ARCSTANK (METHANE)
101
MCDONIV.-'LL DOUGLAS As'ro_oI_.4UTIC8 COMPANY- ST. LOtLII8 DIVI.glON
LOX/HYDROCARBON IAuxiliary Propulsion SystEm StudyFINAL RePoRt REPORT MDC E2548
• .- JULY 1982 I0_._,-,.", - T. ; ", ,OF ? "Y.,J!:_(" .- .,",'f
I
i 30 DAY MISSION I:_INITIAL METHANE LOAD = 3005 LBM
• RELIEF PRESSURE = 60 PSIA
= • ENVIRONMENTAL TEMPERATURE = 5000R I
Q:
,¢I
LU
0_°
8.
20 30 40 50 60 70NISSION TIME, HR ,10'
I
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Io.
FIGURE 5330..DAY TEMPERATURE HISTORY FOR COMMON OMS-ARCSTANK (METHANE)
I
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LOX/HYDROCARBON
Auxiliary Propulsion System STudyREPOrt MDCE2548
FINAL REPORT JULY 1982
The measure of tank insulation effectiveness is the propellant boil-off (vent
loss) that occurs during the mission. Thirty-day vent losses for the two candidate
insulation materials are compared in Figures 54 through 57 for a common OF,S-ARCS
tank and a separate FRCS tank. Results for methane are presented in Figures 54 and
55, whereas results for oxygen are presented in Figures 56 and 57. As shown in
these figures the vent loss with 1.0 inch of MLI is approximately one-half that of
1.5 inches of TG-15000 insulation. As a result of these evaluations 1.0 inch of MLI
was baselined for the LOX and methane storage tanks.
7.2 RCS Feedline Insulation Evaluations
MLI and TG-15000 insulation materials were also evaluated for LOX and methane
RCS feedlines. These evaluations were performed using the FDLINE code described in
Section 3.3 and employed the thruster heat soakback model described in Section 6.3.
The evaluations were performed for a manifold arrangement in which all the RCS
propellant is consumed through a single thruster manifold feeding the required
number of primary and vernier thrusters for Orbiter three axis attitude control.
(Should a thruster failure occur with this manifold arrangement--i.e., failed open
thruster valve--the primary manifold would be isolated and a back-up manifold
activated.)
The
for speci
a 7-day
feedline
system.
mission
FDLINE code computes propellant temperature response along the feedline
fiedmission duty cycles. As an example, methane temperature response for
RCS mission is shown in Figure 58. The temperature response is for a
node adjacent to a thruster valve which is the warmest point in the feed
The feedline was modeled assuming one-inch of TG-15000 insulation. The
duty cycle was modeled assuming two distinct operating modes--primary
thruster firings and vernier thruster firings. Each primary thruster firing (~150
per mission) was modeled as a discrete pulse, whereas the vernier firings (~15,000
per mission) were modeled as a continuous burn at very low flowrate. As shown in
Figure 58 the methane temperature at the thruster valve cools to 220OR during
primary thruster firings as warm propellant in the feedline is replaced by cold
propellant from the accumulator. During periods of low flowrate vernier thruster
limit cycle operation the methane temperature increases and stablizes at 268OR,
which is below the vaporization limit of 282°R.
103
MCDONNEI.L _OUGLAS 4BTROIVAUTICg COMPA_Y.BT. LOUIm DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
ORIGINAL " ...... "'
OF PoOR Q;-::,..,-......
• 30 DAY MISSION
. INITIAL METHANE LOAD = 3005 LBM
• RELIEF PRESSURE = 60 PSIA
• ENVIRONMENTAL TEMPERATURE = 500OR
REPORT MDC E2548
JULY 1982
I
1
t
t&
W_
rnJ
a
zI,I
T
O"
°0 l0
FIGURE 54
1.5 INCHES TG-15000 INSULATION
1.0 INCH MLI
JJ
J ,___-,JJ fJ
30 4O 50 GO 10
MISSION TIME, HR ,I0'eo
METHANE VENT LOSSES FOR COMMON OMS..ARCS TAF"
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MCDONNELL DOUGLAS A.fB'rROI_AUTICE COMPANY" .ST. LO4,IIS DIVIBION
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AuxiliaRy Propulsion System StudyFINAL REPORT '... ;?,-.-.
REPORT MDC 52.548
JULY 1982
• 30 DAY MISSION
• INITIAL METHANE LOAD = 683 LBM
• RELIEF PRESSURE = 60 PSIA
• ENVIRONMENTAL TEMPERATURE = 500OR
A1,5 INCHES TG-15000 INSULATION
ol.0 INCH MLI
0
ldJ
Zc_
Lu
i
JJ
J
I
J
II
I0 20 30 40 S(3 60 70 eO
MISSION TIME, MR ,iO'
FIGURE 55 METHANE VENT LOSSES FOR FRCS TANK
105
MCCOItlNF_LL., DOILIGLAB 4STNON4UTICS COMPANTV" . roT, LOtLIIS DIVIBION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
OF PoOr QuALrrY
REPORI MDC E25,48
JULY 1982
I
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I
I• 30 DAY MISSION
• INITIAL LOX LOAD = 9606 LBM• RELIEF PRESSURE = 60 PSIA
• ENVIRONMENTALTEMPERATURE = 5000R
1.5 INCHES TG-15000 INSULATIONo 1.0 INCH MLI
.
I
I
_0 Ix
-Jo
C_i,il-,-Zul
-,4
0
IO
I
I
fJ" |
20 30 40 50 60 70 O0
MISSION TIME, HR ,I0' !|
FIGURE 56 LOX VENT LOSSES FOR COMMON OMS ARCS TANK
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MCDONItlEI.L DOUGLAS ASTRONAUTICI COMPANY.I
LOX/HYDROCARBON
AuxiLiary PRopuLsion SysTem STudyFINAL REPORT
..': (
REPORT MDC E25,48
JULY 1982|t'
• 30 DAY MISSION
• INITIAL LOX LOAD = 2148 LBM• RELIEF PRESSURE = 60 PSIA
• ENVIRONMENTAL TEMPERATURE = 5000R
m
==
Z
-J
°0 10
1.5 INCHES TG-15000 INSULATION
o 1,0 INCH MLI
,,, J
Jf
..I f" _...__ _..-.-- .4D
20 30 40 50 60 70
MI$SI6N TIME, HR , I0'
FIGURE 57 LOX VENT LOSSES FOR FRCS TANK
107
MCDONNA"LL DOUGLAB ABTRO,_IAUTICS COMPAItlt" .BT. LOtLIIB DIVIWlON
LOX/HYDROCARBON I
Auxiliary Propulsion System Study report MOc e2_8
FINAL REPORT _- JULY 1982 I
ORIn_I,'\ I/%,:.. .- ,__:_ -,j
_F _C'qF: .v , '-:'Y I
Io 7-DAY RCSDUTYCYCLE
o 1 IN, TG-15000 INSULATION
g o PRESSURE: 500 PSIA I
_ ' -/ i
t _ Il_..c__nr-h _fh,',,,In_,l,r _1,,_, !
/i/I,/V[l/g'IIi ll ']['2'ltl'l..ll[IY l,lr/iIf I ,
MISSION TIME, HR
FIGURE 58 METHANE TEMPERATURE RESPONSE IN RCS FEEDLINE
(TG-.15000 INSULATION)
108
MCDOIWItimI.L DOIIJGLAI AITROIVAUTICS COMPAI_It'-OT. L43_LJll OIVISlOItl
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Auxiliary Propulsion System studyFINAL REPORT REPORT MDC E2548
JULY 1982
For 30-day RCS missions the vernier propellant usage rate is reduced
substantially and higher feedline temperatures are attained. This is shown by
Figure 59 which is a summary plot of maximum methane feedline temperature as a
function of vernier usage rate and accumulator temperature for TG-15000 insula-
tion. Whereas temperatures are maintained below the vaporization limit for" the
7-day vernier thruster usage rate, they exceed the vaporization limit at the lower
30-day usage rate. A similar summary plot, presented in Figure 60 for one inch of
MLI, shows that laethane feedline temperatures are maintained below the vaporiza-
tion limit for both the 7 and 30-day vernier thruster usage rates.
To complete the RCS feedline insulation evaluations similar analyses were
performed for LOX. The results are presented in Figures 61 through 63 and show that
MLI is also required for the LOX feedlines to preclude excessive temperatures for
the 30-daymission. As a result of these evaluations one inch of MLI was baselined
for the cryogenic LOX and methane RCS feedlines.
7.3 GOX/Ethanol RCS Feasibility Evaluation_
Two feed system approaches were evaluated to determine the feasibility of
gaseous 02 (GOX) feed in the oxygen/ethanol ARCS. The first uses theOMS turbopumps
to resupply the RCS accumulators and an ethanol feedline heat exchanger to gasify
the RCS oxygen flow. The second uses small electric pumps to resupply the RCS
accumulators and an ethanol tank heat exchanger to gasify the RCS oxygen flow. The
advantage of these approaches is the elimination of insulation on the RCS oxygenaccumulator and feedlines.
The first feed system approach, using the OMS turbopumps to resupply the RCS
accumulators, is shown in Figure 64. This approach uses two heat exchangers
upstream of the RCS accumulators to thermally condition the RCS 02 supply and avoid
the u_e of fuel-rich gas generator products in an 02 heat exchanger. During RCS
accumulator resupply, fuel leaving the OMS turbopump is first preheated to 660OR in
a heat exchanger by reaction products from a separate fuel-rich gas generator. The
hot fuel flow is then used to thermally condition the 02 re-supply flow from 165 to
370OR in a passive feedline heat exchanger. The passive feedline heat exchanger
operates at a low oxidizer-to-fuel flowrate ratio (i.0) to enhance its 02 heating
109
MCDONNEI.L DOUGLAS ASTRONAU'rlCB COHP4Nt,'.ST. L(_LtlB
L
LOX/HYDROCARBON I
Auxiliary Propulsion System Study
FmNALREPORT c :'_--i" ] REPORTmJuoLcYe_l_ I
t
o 1 IN,TG-15000Ir]SULATION |
._ o PRESSURE= 500PSIA_ 30-DAY V_RNIER ) '" - l
r_SAGE R kTE [ i! B_!_.,_ , • , • ,.
7_.DAY VERNLER
_. ,l li_ 'USAGE RA?E " IuJ _\ CH,,VAF DRIZATION LIMIT FO 250 PSI'-"(_ • - _-'_ r-..----_......... PRES SURE
_'ii l l l -- --, I\_ I
\\X,, I
_-, ____.__
ll-
N
ACCUMULATOR METHANE I__-._ TEMPERATURE
+ 240 2n
A220 R
im
2 3 4 5METHANE USAGE RATE, LBM/HR
FIGURE 59 EFFECT OF METHANE USAGE RATE ON FEEDLINE
TEMPERATURE RESPONSE (TG-15000 INSULATION)
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MCDONA'EL m- DOUGL,_S JhBTR,OfttAUTICB COMPANY .BT. L¢_iLIIS
LOX/HYDROCARBON
Auxil.iary Propulsion System StudyFINAL REPORT
OF PoOR :_'-_"'-'" "*'
REPORT MDC E2548
JULY 1982
o I I_l.MLIo PRESSURE= 500PSIA
30-DAY VE _NIERUSAGE R &TE
.C!'14,.VAP_ RIZATION
_., --7 ,DAY VERN
_. _ I usm_RA
,--_.
LIMIT FOF_4
I
250 PSIA CHAMBER RESSURE
ACCUMULATOR METHANETEIIPERATURE
+ 240 R
"220 R
0200 R
FIGURE 60
2 3 4 5METHANE USAGE RATE, LBM/HR
EFFECT OF METHANE USAGE RATE ON FEEDLINE
TEMPERATURE RESPONSE (MLI)
111
MCDONNELL DOUOA.A8 4B'rRoNAUTICB COMPAItiV.JT. LOti,fle DIVIBION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT ORIGt:,,t .....
OF POOR Q-_,',_-,, .REPORT MDC E2548
JULY 1982
I
lit"
1
o 7-DAYRCS DUTY CYCLEo i Irl,TG-15000INSULATIONo PRESSURE- 500 PSIA
1_.
AC ;UMULATORLOX TEMPIRATURE = 180"R
"13 20 40 60 eO I00 120 140 I_
MISSION TIME, HR' IOO
1
i
t
t
I
!
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t
I
FIGURE 61 LOX TEMPERATURE RESPONSE IN RCS FEEDLINE(TG-.15000 INSULATION)
112
MCDONNELL L DOUGL AB A8 TRONIAU TICS COMPANy . SY'. LOIJOS DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
• °
REPORT MDC E2548JULY 1982
o 1 IN,TG-1500INSULATIO!,Io PRESSURE= 500PSIA
_4
g
I--rwl :--i
.J
30-DAY VERNIER
USAGE RATE
I
_P
,__ 7-DAY VEINIERi USAGE _ATE1
' "1"- "---"_ _ ACCUMULATOR
I _ LOX TEMP.
- --" A leo R
PORIZATIO_! LIMIT FIR 250 PSI CHAMBER PRESSURE
0 160 R
.. . .
""b i 2 3 4 ,5LOX USAGE RATE, LBN/HR
FIGURE 62 EFFECT OF LOX USAGE RATE ON FEEDLINE
TEMPERATURE RESPONSE (TG15000 INSULATION)
113
It/IICOOftlNELL OOLIGLA8 4STROItlAL/TICS, COI_4PANY- liT. LOUIS DIVISION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPOR1
ORtG(h",;" {;£,r_.". '".'_
OF POOR QUALITYREPORT MDC E2548
JULY 1982
I
I
o 1 II],MLI
o PRESSURE= 500 PSIA
I
I
tl" 30-DAYVERNIER/| USAGRATE
_,_ ., ,, .,111 [7-DAY VE_N_ERI_ "USAGE tATE
_" LOX VA PO _ Z'ATTON.._"IT-'F'_
I
i
I
!
I250 PSI
ii _ PRESSURE I_.._ -__-- ---- Ix ACCUMULATOR
LOX TEMP. I_. +_:X]R
AleOR
O 160 R I
2 3 4 5 ILOX USAGE RATE, LBM/HR
FIGURE 63 EFFECT OF LOX USAGE RATE ON FEEDLINETEMPERATURE RESPONSE (MLI)
!
!
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MCDONNELL DOUGLAS ASTRONAUTICS COMPANIY*BToL(]NLIll DIVISION
'_ " T/
LOX/HYDROCARBON
Auxil.iAey PeopuIsioN SYSTEM STudyFINAL REPORT
REPORT MDC E2548
JULY 1982
CROSSFEED
@]
I:}I
J
BYPASS -----------
OMS ACCUMULATOR--
ETHANOL
_'_'_"__ CROSSFEED
SITURBOPUF'iPS
LIQUID ETHANOL
Q RCS ACCUI'IULATOR
/ ,j .. ,_.. _ _ , F-- "_.-.-.--;lORCS
' - (_!l THRUSIERS
"____..___ ELECTRONIC
GASEOUS 02 PRESSURE REGULATOR
RCS ACCUMULATOR
FIGURE 64 LOX/ETHANOL OMS.-ARCS USING OMS TURBOPUMPSFUR RCS SUPPLY
LOX/HYDROCARBON
Auxiliary Propulsion SysTem STudy
FINAL REPORT REPORT MDC E2548
JULY 1982
capability. The RCS thrusters also operate at a mixture ratio of 1.0 so that the
accumulator outflow is at the same mixture ratio a_ the resupply flow. Since the
single shaft turbopumps deliver propellants at a fixed oxidizer-to-fuel flowrate
ratio of 1.72:1, the excess 02 flow is routed back to the LOX tank, by-passing theheat exchanger.
RCS accumulator operation with this OMS turbopump resupply approach was
evaluated using the TKHEAT code. Temperatureand pressure response in the ethanol
accumulator are shown in Figure 65, while temperature and pressure response in the
GOX accumulator are shown in Figure 66. Both accumulators operate in a blowaown
mode. The ethanol accumulator contains a helium charge which expands with outflow
and compresses with resupplyflow. Resupplyof both accumulators is controlled by
pressure switches in the ethanol accumulator. As shown in Figure 65 resupply flow
is initiated when the ethanol accumulator pressure decays to 500 psia and is
terminated when the pressure rebuilds to 775 psia. The temperature of the ethanol
in the accumulator decays during resupply as a result of the cold resupply flow
(370°F) and then increases as a result of environmental heating after resupply is
cut-off. A minimum ethanol accumulator temperature of 370OR is reached 24 hours
into the mission during the period of maxlmum RCS usage. The 02 accumulator
(Figure 66) cycles with the same frequency as the ethanol accumulator but operates
over a wider pressure range. (660 to 1260 psia). The minimum 02 accumulator
temperature (350OR) also occurs 24 hours into the mission.
Despite the wide variations in accumulator pressures and temperatures
adequate control over RCS thruster mixture ratio is achieved through use of an
eIcctronic pressure regulator and thermally shorted feedlines downstream of the
accumulators (Figure 64). lhe 02 accumulator outlet pressure is controlled in
response to ethanol accumulator pressure with the electronic pressure regulator,
:vhile 02 and ethanol fluid temperatures are equalized with the thermally shortedfeedlines.
The results of Figures 65 and 66 demonstrate the feasibility of a hybrid RCS
feed system (gaseous 02 and liquid ethanol) in which OMS turbopumps are used for
accumulator resupply. However, this feed system approach has the followingdisadvantages:
116
McrJoI_INELL DOUGL A8 ,41 TROI_4u TIC8 COMP_sI_ V . B1F. LOiLIII/J _lVISllDItl
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Auxiliaay Propulsion System StudyFINAL REPORT
m,*
i--
m,.
:E,_-IJJ)-.
.... . lu. 1 _J REPORT MDC E2548
JULY 1982
• OMS TURBOPUMPS USED TO RE-SUPPLY RCS ACCUMU ATORS• LIQUID ETHANOL ACCUMULATOR VOLUME• RE-SUPPLY CONDITIONS = 2.47 FT
FLOWRATE : 10.8 LBM/SECTEMPERATURE = 370ORPRESSURE = 1250 PSIA
• ACCUMULATOR RE-SUPPLY SWITCHING PRESSURESINITIATION : 500 PSIACUT-OFF = 775 PSIA
L
I ....F
Z_Z x A A
• I Ii
• ./ I
/ ,-
#I -, - ,I/Ivv/L/I/,,v
iII
/½, f.V
140i
150 180
l tt-fl-IIII
6 .a_ i " ":"
-_,,/,,l " ".;_;; _-_._,,_ '_.....,^ III _l_.g, _1.. _ ..
¢/')
/,,_ O •
P_,_
I II It
I,! !l I
i$
' --------4,1 '_ [
HISSI6N TII,£, ,_
FIGURE 65
IJI I
100 I_O 140
/
I -II,l
RCS ETHANOL ACCUMULATOR MISSION RESPONSE(OMS TURBOPUMP RESUPPLY)
117
_DONNELL DOUGL_B ABTRO_AUTICB COIt4PANV.B_ L("JNLIOB OlkTSOON
ILOX/HYDROCARBON
Auxiliary Propulsion System Study nRtGINAL PT_,_ :-_ moc e2_
FINAL REPORT OF pOOR QUALI';" REPORT JULY 1982 I I
o oM_,URl_OpUMP_u_oTORF-,U,_PL',_,_SAC_UL_TOR_• GASEOUSO_AC,:U,'.,U._ORVO,_U,V,_: 1_._,:,_ I• RE-SUPPLY CONDTTTONS
FLOWRATE = 10.8 LBM/SECTEi,_PERATURE= 370OR iPRESSURE : 1250 PSIA
_ il i I I i I i 1
o-/I I t I I J 1 ,5_)" 20 40 50 O0 _00 IZO i,lO 150 tOO i
NISSION TIME, HR IIII
I
i'll ./\I1! l\rx/, lx '_-- '_1 _ _I"\IU'_ _I _ _I i L
P_ 20 40 50 OO IOO i2o t_ ti,o IOO
HISSI6N TIME, HR
FIGURE 66 RCS GOX ACCUMULATOR MISSION RESPONSE
(OMS TURBOPUMP RESUPPLY)
118
MCDONNIEI.L DOUGLAS ASTRONAUTIC8 COMPANY-IFr. LOIJIi DIVOE#ON
I!I
III
IOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
REPORT MDC E2,548
JULY 1982
• laFge number of turbopump cycles (~50 per mission)
• complexity associated with the use of an 02 heat exchanger by-pass
circuit and a separate gas generator for fuel pre-heating
• low RCS performance (gas generator vent losses coupled with low RCS
mixture ratio).
The second feed system approach, using dedicated electric motor pumps to
resupply the RCS accumulators, is illustrated in Figure 67. In this approach 02
thermal conditioning is achieved using a passive ethanol tank heat exchanger. The
02 enters the heat exchanger as a liquid at cryogenic temperature, cbsorbs heat
from the environment, tank wall, and ethanol inside the tank and exits the heat
exchanger as a superheated vapor. The effectiveness of the ethanol tank heat
exchanger is shown in Figures 68 and 69 for a representative 7-day OMS-RCS mission
duty cycle, lhe heat exchanger was sized for two primary RCS thrusters firing
simultaneously in order to meet the back-up RCS deorbit burn requirement.
Figure 68 shows the 02 inlet and outlet temperature histories over the 7-day priod.
The coldest 02 outlet temperature is 425OR and occurs 24 hours into the mission
during the period of maximum RCS usage. Figure 69 shows the corresponding
temperature and quaptity of ethanol remaining as a function of mission time. The
coldest ethanol temperature (430OR) also occurs at the 24 hour point.
Examples of RCS accumulator temperature-pressure response with the electric
pumpresupplyapproach are shown in Figures 70 and 71. Figure 70 shows the response
of the liquid ethanol accuml_lator, while Figure 71 shows the response of the
gaseous 02 accumulator. For these examples the temperatures of the fuel and
oxidizer resupply flows were set equal to their minimum values (430 and 425OR,
respectively). In order to minimi_e electric motor weight and power requirements
pump discharge pressures were set at 500 psia. Unlike the preceeding OMS-RCS
concept which used the OMS turbopumps for resupply, the ethanol and 02 accumulators
do not have to be resupplied at the same time, and the RCS thrusters can be operated
at optimum mixture ratio (1.3 to 1.4). Similar to the preceeding concept an
electronic pressure regulator and thermally shorted feedlines are employed
downstream of the accumulators to control RCS thruster mixture ratio (Figure 67).
The results of Figures 68 through 71 demonstrate the feasibility of a hybrid
RCS feed system (gaseous 02 and liquid ethanol) In which electric pumps are used for
119
MCDONNELL DOUGLAS ABTRON4UTICS COMPANV.B_ LO4,11S DIVIBION
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT REPORT MDC E2548
JULY 1982
ORIGINAL PkG_ _{,_
OF POOR (_UA!Lri"Y
LOX
CROoSFr..Eb- l(TYP.)
ELECTRIC RCS/ i "
43
PUMP(tYe.) _--1_ _OMS TURBOPUMPS
ETHANOL
GASEOUS 02
RCS ACCUMULA FOR
ELECTROIIIC
_t_"l PRESSURE REGULATOR
"_'" _ IHRUSTERSLIQUID ETHANOL
RCS ACCUMULATOR
OHS ACCUMULATOR
I
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l
a
O_IE
LJ
FIGURE 67 LOX/ETHANOL OMS..ARCS USING ELECTRIC MOTOR
PUMPS FOR RCS SUPPLY
!20
MCLDOIbltWEI..I,. _OLIGLAS 41TI_ORIAUYICJ COMPANV-BT. LOLIRS OlVlSlOttl
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
• -., .
• . sl
REPORT MDC F...2548
JULY 1982
• 7-DAY OMS-RCS MISSION DUTY CYCLE
• HEAT EXCHANGER 02 FLOWRATE : 3.3 LBM/SEC I• HEAT EXCHANGER INLET TEMP. : 162°R• HEAT EXCHANGER !NLET PRESS. = 500 PSIA
DURINGRE-SUPPLY
_ Q.r..ju_
X
I-=
o"o ;to
FIGURE 68
02 )UTLEI EMP
02 INLET TEMP_
_ _ I_ 120 I_ l_ 180
MISSIeN TIME. R
02 TEMPERATURE HISTORIES FOR PASSIVE ETHANOLTANK HEAT EXCHANGER
121
MCDONNELL DOUGLAH ASTRONAUTIC8 COMPANY-lIT. _t.(_LIIS DIVIBION
LOX/HYDROC&RBON
Auxiliary Propulsion System Study
FINAL REPORT oRiGIn,.:-- :.......
OF _',_- _""
, 7-DAY OMS-RCS MISSION DUTY CYCLE
HEAT EXCHANGER 02 FLOWRATE : 3.3 LBM/SEC /
i HEAT EXCHANGER 02 INLET TEMP. = 1620RHEAT EXCHANGER 02 INLET PRESS. = 500 PSIA
REPORT MDC E2548
JULY 1982
DURINGRE-SUPPLY
m.
,,o
m,,
rwJ _ i_WISSIgNTIME. h_
I1
I,=-
FIGURE 69 ETHANb; TEMPERATURE AND QUANTITY HISTORIES FOR
PASSIVE ETHANOL TANK HEAT EXCHANGER
122
MCDONNELL DOLIGLAB AITRON4UTIC8 COMPANY.roT. LOUIB DIVIBION
I
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Auxiliary Propulsion SysTeM Study
FINAL REPORT
oRIGIN_-;L PAGE i_i
OF POOR QUALITY.
• ELECTRIC PUMPS USED TO RE-SUPPLY RCS ACCUMULATORS• LIQUID ETHANOl. ACCUMULATORVOLUME=2.47 FT. 3• RE-SUPPLY CONDITIONS
FLOWRATE = 2.5 LBM/SECTEMPERATURE= 430ORPRESSURE = 500 PSIA
• ACCUMULATORRE-SUPPLY SWITCHING PRESSURESINITIATION = 350 PSIACUT-OFF = 500 PSIA
REPORT MDC E2548
JULY 1982
!f
==
/k,/I/
_" ' IJ
/I/
40
I
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i
!
i
10 l0 I00 ImNIS$1gNTIN[, MR
!
II
IJ
140 I_ xm
8
i
i
. !
i
_lil:'
\!I
, I10
I •
• - ;- =
I X ,l
. , !
i• is
11 I\ I\lil_ I-: - - : ".
\I _,I .5.1 •
'| ;, -
\,,_:_, x.! \ .. Ji.
I
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iO IIIII IZO I_ IIO li_WIS$1_INTIME.
FIGURE 70 RCS ETHANOL ACCUMULATOR MISSION RESPONSE
(ELECTRIC PUMP RESUPPLY)
123
MCDONNELL DOUGLAII 48"rRON4UTICS COMPANY.roT. LOt_OE_ D#VIBION
LOX/HYDROCARBON OR_tNAL PAG_ i_ IAuxiliary PropubioN
System Study OF PooR QUALITY REP T I
• ELECTRIC PUMPS USED TO RE-SUPPLY RCS ACCUMULATORS II"
• GASEOUS 02 ACCUMULATORVOLUME = 13.8 FT3° RE-SUPPLY CONDITIONS
FLOWRATE : 3.3 LBM/SEC |TEMPERATURE = 425OR !
PRESSURE = 500 PSIA T• ACCUNULATOR RE-SUPPLY SWITCH,NG PRESSURES I
INITIATION : 350 PSIA . ICUT-OFF = 500 PSIA
., i",ljp i i ,r_'y'_'-'_....":_'_-_'_ I
,I(C[
2O 4O
IinO oG(_ [20
WISSli_q TIME, HRJ40 ofW) llo
I
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FIGURC 71 RCS GOX ACCUMULATOR MISSION RESPONSE
(ELECTRIC PUMP RESUPPLY)
124
I
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II. I)01,1111,..411 All _JTICB IT. Li
LOx/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT REPORT MDC _548
JULY 1982
accumulator re-supply and a passive ethanol tank heat exchanger is used for 02
thermal conditioning. The advantages of this concept are its simplicity (no active
gas generator--heat exchangr assembly or by-pass circuit), high RCS specific
impulse (no vent losses), and the low number of OMS turbopump cycles. Its
disadvantages are the lower RCS flow (thrust) capability due to the passive tank
exchanger and weight/power penalties associated with electric pumps. Because of
its attractiveness, the electric pump resupply approach with passive ethanol tank
02 heat exchanger was baselined for the LOX/ethanol OMS-ARCS. An attractive
back-up thermal conditioning approach is the dual fuel heat exchanger concept of
Figure 41 which eliminates the use of hot, fuel-rich gas generator products to
thermally condition the 02 .
7.4 LOX/Ethanol and LOX/Methane OMS-ARCS Sensitivity Analyses
The selected baseline feed systems for LOX/ethanol and LOX/methane are shown
in Figures 72 and 73, respectively. Both concepts employ common OMS-ARCS
propellant tanks, dedicated electric pumps for RCS supply, and component redun-
dancy for satisfying the fail-operational/fail-safe reliability requirement.
Redundant lithium batteries were baselined for powering the electric RCS pumps.
In-line entry sumps are provided just downstream of the propellant tanks. These
sumps remain full during the mission and provide a dedicated propellant supply for
ARCS operation during entry. Overboard abort dump systems are provided just
downstream of the entry sumps. The OMS engine system employs a single turbine for
driving both the fuel and oxidizer pumps. A gas generator cycle is used for
LOX/ethanol, whereas a methane expander cycle is used for LOX/methane. The
LOX/ethanol ARCS is a hybrid feed system delivering gaseous 02 and liquid ethanol
to the thrusters through uninsulated accumulators and feedlines. The 02 thermal
conditioning is provided bya passive ethanol tank heat exchanger. The LOX/m_thane
ARCS is a liquid feed system delivering cryogenic propellants tn the thrusters
through insulated accumulators and feedlines.
Sensitivity analyses were performed for both system concepts to define
optimum chamber pressures and accumulator blowdown ratios and to determine the
impact of variations in engine specific impulse and propellant tank minimum gage
thickness. The results of these analyses are presented in the followingparagraphs.
125
MCDOItlf_ELL DOUGLAS 4aTROitIAuTICoM COMPANY .ST. LO4JIB DIVIEION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORTREPORT MDC E2548
JULY 1982
: RELIEF
LOX
ABORTDUMP
CROSSFEED
FILL
OMS ENGINESTARTACCUMULATOR
GAS GENERATOR
TURBOPUMfASSEMBLY
ENTRY SUMPS
ETHANOL
ELECTRICPUMP
ETHANOL
RCSCCUMULATOR$ FILL
-, ELECTRONIC! PRESSURE
.: REGULATOR
VERNIER
THRUSTERS
PRIMARY THRUSTERS
FUEL REGEN.COOLEDOMS ENGINE
FIGURE 72 BASELINE LOX/ETHANOL OMS..ARCS
126
MCDONItlELL DOUGLAS ABTROftl.4UTIC8 COMPAIVY.BT. A.OLIIJ DIVIBION
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AuxiliaRy Propulsion System StudyFINAL REPORT REPORT MDC E254_
JULY 1982
OR_GiT,:: L , .\.. ]
RELIEF
LOX METHANE
ABORTDUMP
CROSSFEED
FILl.
ENTRY SUMPS)_) CROSSFEED
METHANE
ELECTRICPUMP
V_RNIERTfiRUSTERS
GASEOUS METHANE
P ACCUMULATOR
RCS ACCUMULATORS
LOX
FILL
PRIMARY THRUSTERS
FIGURE 73 BASELINE LOX/METHANE OMS.-ARCS
127
MCDONNELL DOUGLA.S 4='rRoNAUTIC8 COMPAI_y.ET. L4i_iJIB I_IV#BIOIV
LOX/HYDROCARBON
AvxiliAey PeopulsioN S),sTe_ STudyFINAL REPORT
REPORT MDC E2548
JULY 1982
7.4.1 Chamber Pressure Sensitivity Analyses - The weight sensitivity of the
LOX/ethanol system to OMS engine chamber pressure is shown in Figure 74. An OMS
chamber pressure of 600 psia was selected as near optimum for the LOX/ethanol
system. Lower chamber pressures provide lower performance and higher system
weights, while higher chamber pressures require supplementary film cooling and
I more complex chamber designs. OMS chamber pressure sensitivities were not
:, developed for the LOX/methane system since chamber pressures greater than 400 psia IP
were not practical with the expander cycle due to insufficient energy for powering
the turbine. As a result, an OMS chamber pressure of 400 psia was selected for the
LOX/methane system to provide the highest practical performanc_ and minimizesystem weight.
The weight sensitivities of the LOX/ethanol and LOX/methane systems to RCS
engine chamber pressure are presented in Figure 75. An RCS chamber pressure of
100 psia was selected as near optimum for both systems because it provides low
weight and minimizes the size and power requirements for the electric motor pumps.
7.4.2 RCS Accumulator Blowdown Pressure Ratio SensitivityAna________lyses - Systemweioht
sensitivities to RCS accumulator blowdown ratio are presented in Figures 76 thro_,gh
78. These data were based on propellant quantities constrained by the current pod
envelope and the use of aluminun accumulators. The higher system weights
associated with higher blowdown ratios are primarily the result of increased
electric motor and battery weights (higher pump discharge pressures). As shown in
Figure 76 a blowdown ratio of 2.0 was selerted for the gaseous 02 accumulator in the
LOX, ethanol system since it provides minimum weight. A b_owdown ratio of 1.67 was
sel,:cted for the liquid ethanol accumulator (Figure 77) because it provides low
._ei,lht and a reasonable thruster inlet pressure range. (The gaseous 02 accumulator
_' let pressure is regulated in response to the ethanol accumulator outlet pressure
:rlg an electronic pressure regulator--Figure 72.) Similarly, blowdown pressure
atios of 1.67 were selected for the liquid RCS accumulators in the LOX/methanesystem (Figure 78).
7.4.3 Sensitivity to OMS and RCS Enqine____IIsp_ Weight and performance sensitivities
to OMS and RCS engine specific impulse are presented in Figures 79 through 86 for
the LOX/ethanol and LOX/methane systems. The weight sensitivities were computed by
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MCDONNELL DOUGLAB AS"rRONAUTICs COMPANy.ST. LOIL/IS I[_IVISION
I
LOX/HYDROCARBON
Auxiliary Propulsion SystEm Study
FINAL REPORT REPORT MDC E2548
JULY" 1982
J
OF p -,.... .,
m..I
18,000
I 17,0C0
L,2
t_0
L-<+
0
16,000
15.000
FIGURE 74
• SINGLE AFT POD
• FIXED TOTAL IMPULSE
L_._._
------..==___.
400 500 600 700 800 900
OMS PC -- PSIA
SENSITIVITY OF LOX/ETHANOL SYSTEM TO OMS CHAMBER PRESSdRE
129
MCDONNELL DOUGLAS ASTRONAILITICE COMPANY-BT. LOLII8 DIVIBION
LOX/HYDROCARBON . . ,_Auxiliary Propulsion System Study ORIGINs" P,_CiL_ t ._
LOX/ETHANOL !
* SINGLE AFT POD i
• FIXED TOTAL IMPULSE t
/
!.._! ! r _J I 1ri I ?--F--]
R S PC - PSIA l
LOX/METHANE
: ,%_%g_,%o,.,_ I
i''°" t
t
FIGURE 75 SENSITIVITY OF LOX/ETHANOL AND LOX/METHANE SYSTEMS I_: TO RCS CHAMBER PRESSURE
MCDONNELL DOUGLAH ASTRONAUTICi COMPANY.IT. L._IB
_ .... :i___m_
LOX/HYDROCARBON
AuxiliaRy PRopulsion SySTeM STudyFINAL REPORT
ORIGINAL p:%GL :'_
OF pOOR "'":' "_'g
REPORT MDC E2548
JULY 1982
_':'r
18,500
18,300
18,200
18,100
18,0001.0
• PC = 100 PSIA
• MIN, ACCUMULATOR PRESSURE = 177 PSIA
:LT'-:IZE: r_,
_ -_.--. ,..
77
,.,.,
:"-b .. |..i _"
;/_._----,..._ :---
Z::]:._ntZ:: :r.•_;.:_ :::
."
:= :=:.._ _7_
1.5 2.0 2.5 ),0 -_.5 4.0
ACCUMULATORBLOWDOWN RATIO ..........
FIGURE 76 RCS GOX ACCUMULATOR BLOWDOWN PRESSURERATIO SENSIT'IVITY
LOX/HYDROCARBON
AuxitiaRy Propulsion SysTem STudy
FINAL REPORT ORIGI: -r'_'.... -'_' 'OF lliO C'_ Q''':'-;'fi_
REPORT MDC F_2548JULY 1982
..,J
-r.
LiJ2=
I--
:3=
cr_
,,:=:
,,,t-
t,J3
18,500-
18,400
18,300
18,200
• ETHANOL ACCUMULATOR
• PC = 100 PSIA
. MIN ACCUMULATOR PRESSURE = 183 PSIA
• CURRENT OMS+ARCS WET WEIGHT = 17,125 LBS.
(PER POD)' I'l-4 ...._ -L-I--_-.•---+-..... -A--.-_
I
' t, 2;- '- • ' : I i j .! i ,
i I i _: i i 'T i
....."...............i ' J i i = '!II
I i : , i !
! i t i I ! [ I
-- i_ __ i i I I I I ,
--÷ ........-r- • _ i .-_"-F t • I_ u_......_....._ _._J
l i ! i 1 _ :
""°°_- ' 'i+4''_ .... '_ _-i ............. +--4-I : I I ii ! / !
----_ .... ,_...... 4-----+ ...... ;.... _........ i...... i ....... _ t " ; ...... "f18,_i00 ! i _ , , _1.0 1.5 2.0 2.5 3.0 3.5
...., ' I I ...........,-- t'-i .......i ......7• 1_ . I i ; "
.....!--_-_.......' I !1111...... ; II i i : - -- _ i" ' ; "+......_ _.......14-_-4-_---,_......] ! : " / - " t:i I i , ] , I i _ ! ; i
-,-r- ............ i .... -_ ..... ÷........ _ ........ • " t :I ..... i • : i : !
i :" _ :' i _:' ; 'I i£7', i ! .,:_-
7 i i ! I
' g----f- ..... I......_......+-ti i I '. i !I i ';--_ --t----f---_ ......-;....
4.0 4.5 5.0 5.5
ACCUMULATOR BLOWDOWN RATIO
FIGURE 77 RCS LIQUID ETHANOL ACCUMULATOR BLOWDOWN PRESSURERATIO SENSITIVITY
132
MCDONNELL DOK,IGLAB ABTRONAUTICS C'OMPAItlV.Slr'.I.OtlJI_B DlVOlilOIbl
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LOX/HYDROCARBON
Auxil.iARy PeopuLsio_ SYSTEM STudyFINAL REPORT REPORT MDC F..2548
JULY 1962
OF POOR _,..,.-_.:-... ,
15,800"
• OXYGEN & METHANE ACCUMULATORS
• PC = 100 PSIA
• MIN ACCUMULATOR PRESSURE: OX = 198 PSIA
FUEL-195 PSIA
• CURRENT ORS + ARCS WET WEIGHT = 17,124 LBM.
15,300'1.O 1.5 2.0 2.5 3.0 3.5 4.0 4.5
ACCUMULATOR BLOWDOWI_IRATIO
5.0 5.5
FIGURE 78 RCS LOX/METHANE ACCUMULATOR BLOWDOWN PRESSURERATIO SENSITIVITY
133
MCDONItlELL DOUGL4S ASTRONAUTICB COMPANtt BT LI_LII8 DIVIBII_I_I• °
LOX/HYDROCAREION
Auxiliary Propulsion SysTem STudyFINAL REPORT ORtl._h'J_"'- __'"
OF pOOR Q1jAL[f{REPORT MDC E254_
JULY 1982
16,600
=E:16,400
-J
'-r-
"' 16,200
I.--
c,O16,000,
÷
¢J3
E 15,800O
• FIXED AV = 500 FT/SEC
. OME DESIGN CHARACTERISTICS
= 17,125 LBM
F = 6000 LBF
PC = 600 PSIAI_ = 298MR = 1,76
• CURRENTOMS-ARCSWET WT.
+ _n', II_.i.] i _
: !'_ ..... ::. I ' ':I ! "
- : '.: J i i,' , l
" "_t_
-._. :.-_-:._ I : .: _: ,,
.......................+......r-,i/i ............15,600 - - I
330 335 340 345 350' =355 360
ORS Isp, SEC,
FIGURE 79 LOX/ETHANOL OMS IsP SENSITIVITY (POD &V - 500 FT/SEC)
134
MCDONNELL DO|JOL AB A,S TRONAU TICB COMPANy. liT. LOtiJIS DIVIBION
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Auxi.l-iAry ProPulsion System Study
FINAL REPORT
. _J*
• ¥
Of P¢:'_,"_ ' REPORT MDC E2548
JULY 1982
C,f.,;
O
620
610
580
1
--T-T; d: i
i .L,i
600 ' , i
_ i :
590 i I 1
-_ -i-_- _ '; I I
!
it r
I
• FIXED POD VOLUME
• OME DESIG_ CHARACTERISTICS
F = 6000 LBF
PC = 600 PSIA
i I i ' i i _ :
570 I.I I
f
55oL II330
; .i ! : ! : f: _ II : : i :
t I / : I : .! , _ , : : :
i' i*"_d, _ !../---i :_ .,
i' ___t ' i, i:l _.! _ l I ' '
i t f I_
Y-..] .! . _ .t ' ' l " t t• . t i. I ' l
' i, I i i I ,
'J i ' I' ' _ - ! -
, , : l f I !
335 340 345 350
l i' It !
• ' / '
I: iL'iJ . I
I:i I '
355 360
OMS Isp, SEC.
FIGURE 80 LOX/ETHANOL OMS Isp SENSITIVITY (FIXED POD VOLUME)
135
MCDONNELL DOUGLAm A_TRONAUT'ICR COMPANIY.BI".LllTtfJIIB DIVIBION
LOX/HYDROCARBON IAuxiliary Propulsion System Stu
FINAL REPORT" O_,_1_j_ pl'kG_.9. REPORT MOC E2548
OF pOOR QUALITY JULY'" I
I
! FIXED v=soo T/SEC I.. OMEDESIGN CHARACTERISTICS
r F = 6000 LBFPC = 400 PSIA I
E =195
16,000
15,800-._J
-r
"' 1.5,600
i==i,i
15,400r,,,-
+
¢..')
o 15,200
15,000350
MR = 3.50
_-!!i:iil.i!i:tiii;_:.-- :i_;tii.i'_;_,,;,.._,_i_:i:------.-,,-,,.,,-,::: -" " '-::z:_:,_.,-r_::::: _=: ,-:_.-'.,'.'-." .-:,-,','• ;_rrri-_-,.:::!: ,::-. :iiH:...! .: .ii:.T.:i:T_ : ::.... :::: "': .......................... ,. " ::::I " .' . .............
::::I:::i :!:!!!:::!_ :i :: :...j..'...:: :7.:., ..;.; : ....... _:.: ..... , .. _ ; • :
"' ' . • ...... : :-: "::; :-'-T-T._ ..... ---7-" _ .....
' _ ' ' ' ! -; ..... I';'; "':T ',; , ' ..... _" .._ .. -t ..... ,I .... i- '
_Z'LIiZI:.Z .::t.::: ::':.':;:::::':::' ": ::::T'::'L:::1'i!i!':Eii "iq'"iFt !!: i:!it:i::l.':::l::..:.::" ::',: "..'"_,. _.'-:_ :---_-._#F.--_.,--_-:m-_-_q_
::t::'" " i .'_.':::t::::,_[_ " " : .........I t :: :::" ::" ":: .... ': .........."_' ................. i ......... : """ :: _ '_ I _ ::_ ' "-- ! • . . :::; :, '.;:..: : ::: , '::: .. . j • : • .., ............. : :';: .:
iili_Cilii;l:i;ili;i!lii_ii_;il;;_i':F:i!i_q;;,;(:!:i:;:iiii;I_:iii!
:::: :::: :::: : :: :_: .:: :!:ih:!i :: _!. _" i ........ : ..... "': ' i ....... [:Z:::[i::;:
......--.. : [._ ............,I..-:t_--:: ::. ::_-I- t.--1.:-:4-:]..t....I....::::::: :::'f':::::b'::: ;.:'.':::: .':.: . :! .i! _ ..... :.;: :. ' ...... :::..:_i_:_iK
"_' ""'" " ; .... ;" ";;: : l • .i ' • ' ;.- . ' ...... : . * "
:: __'_;'Fi :! :F: ::: :::: :: :'::: ::: I: :: !:F_ :: "' :'I: !;:: :::: "':: :::: :::: ::" ":'_'_i ::_! i-:.!:+_ !;_ ; : ::_. :::: _ii,_!i: :::._:_: :;._:_i_.-.t .... t.... .... ....
....... ' ...... t. ......... | .... ;.•":tf..::._,;.:l .; . ::::::H:::::::::::::
_" : ::::: '_ : :f :;:_'--'" ..... .-...-m.. _ ...... :!it: ::::'::': -':: ::: !":: :::: ::::i::.
355 360 365 370 375 380
OMS Isp, SEC.
FIGURE 81 LOXIMETHANE OMS Isp SENSITIVITY (POD AV = 500 FTISEC)
136
MCDONNELL DOUGLAI AITRDI_AUTICEI COMPAItly . I_I. LiDeJI8 DI_.,IHIC) _
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
OE} %:?L_",-':"--'' '"
REPORT MDC E254,BJULY 1982
XED POD VOLUME: _E DESIGN CHARACTERISTICS
F = 6000 LBF
PC = 400 PSIA
= 195
MR = 3.50
• CURRENT OMS AV = 500 FT/SEC
s2o i... t.l..i _PF+P,i _POr_ I _ '; - : .
_:_ :_i:. i.._ I : I. , i. i I _ i I l jr _._ I
•i _ _ii_ -i---:. -i - i-'-.,......-_'7 _.....i............1,.... .......,I.slo_ 'i.. .,.., .I' : I : ', i .....
' _ _: :' ' :i ....... "i ..... '. -'-'I ;"-7" .... -'7 ...... :- -- ' .... ;"-_" : !:I.l.-I::_it-._t,_.l.,..t.:._.:..l_.._.--I._;.!__=;_._-_7::t...::I...IK_I1_:I:;:!:_r ._]_ _i.,_ _ "i ! ' '
:iI....l-:-I+'I--:_......I....F":i....I........_............l.-..i....i:-!.-.-I....I....i..i
490 ...;t:.,_-i= i i ,_ _ " I i. •i I I i? .... :" "_ ................. " ....... t ...... i ...... _........ ',................ i-.. _ • _
_! ":' ! i ! ! _'i ! !480 ,,,,,,,I ...... ,, , i . i .
350 355 360 365 370 375 380
OMS Isp, SEC,
FIGURE 82 LOX/METHANE OMS Isp SENSITIVITY (FIXED POD VOLUME)
137
MCIDOItlI_IL L I)OUGL A I 4111_RO_I4 L,Ilr.IC I C OIt4PA_ V . IT. L OtLIII DlVJ'_lOItl
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
ORIGINAL PAC;E "_
OE. POOR QUALI'I31'
REPORT MDC E2548
JULY 1982
16,200
,.,.,J
16,1oo.-r-
F--- 16,000
CO
+
(,f)
15,900.
15,800
260
• FIXED _V = 500 FT/SEC
• RCE DESIGN CHARACTERISTICS
F = 870 LBF
PC = 100 PSIA
e = 18MR = 1.30
• CURRENT OMS+ARCS WET WT.= 17,125 LBM
,. , ! •
.....L_. I
'I I"
i"'I i
I
,I
265
( PER POD)
-_..........,__,____1 l.... ': ..... i i
__ = , ! ""I °" .i....--I--- "-_--_ --T'-t i'" I I: :': '!.
-.!.t.....!I .........................ii [ '"
270 275 280
PRIMARY RCE Isp, SEC.
i.......I-_....!q I _,,i__
, .._.I .I
i- ii.... l'_
! '
i .-t..
' j ' . ,
285 290
FIGURE 83 LOX/ETHANOL RCS Isp SENSITIVITY (POD AV = 500 FT/SEC)
138
MCDOItlIblELL DOILIGLAH 4811"ROIYAU1rlC8 COMPAItlI_.Blr. Lg_LilIB DIVIdllil_ltl
ININlu,_lmn mm Ip_ilmnaliai "
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LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT r-_l l •
OF F_%<,_-c.Q::, ..:;'.f'
REPORT MDC F..2548JULY 1982 g
600
598
596c_
I---LI-
594
cr_
0
592
590260
FIGURE 84
• FIXED POD VOLUME
• RCE DESIGN CHARACTERISTICS
F = 870 LBF
PC = 100 PSIA
E = 18
MR = 1,20
• CURRENT OMS 6V = 500 FT/SEC
265 270 275 280 285 290
PRIMARY RCE Isp, SEC,
LOX/ETHANOL RCS Isp SENSITIVITY (FIXED POD VOLUME)
139
MCDONNELL DDLIGLAE 4J'rRoNAUTICB COMPANV.IBlr'. Lira.liB DIVImlON
LOX/HYDROCARBOR
Auxiliary Propulsion SystEm STudy
FINAL REPORT
ORiGII';A'i..PAC_ i_
POOR OUALI"I"Y
REPORT MDC E2548
JULY 1982
I
I
I
15,600
...I
15,500"-i-
F-- 15,400
co1,.,)
< 15,300+
com--0
15,200
FIGURE 85
FIXED TOTAL IMPULSE
• RCE DESIGN CHARACTERISTICS
F = 870 LBF
PC = 100 PSIA
e =18
MR = 2.49
. CURRENT OMS-ARCS WET WT. = 17,125 LBM
(PER POD)
:i....... , .......... ; -- . _ ..:1 . ..; .: _' :"-:_. .....
.:_ • - ................. , ..... .,._. ; ..............
•._:._ _ii ;i_!l_i_.._.;._': :.:.::::::_:!::_:::.ii!it:_;:::i!_l!!: :._::::_- ::_:;:_:_::'-'"'::::ti:::'.........":-:::".:-_ .... I .... "','::: " ,.': :::: :::: ::::t. : ' ! .: _ !: -! ;!: -::_:_
............................. I.. I ::::I::;: ::: : : ::'I: .........
E_Eii'i L;_ i!!! E!_i_i :::! :::: :,,; " :;, • l_tl:_l't/_l_lrl: : _ ;]'- ....
: " :: : : : I ':: ........ ' ............... t ...... : " " ] ::: :::: :_ : : : "
.: .... ._ ....... _...: :. ::I:::::::::::::::::::::::: ;:':I:::" :::lI:::: :::: _ I ..
_]_li:_il_]]]l::::iil[!il]]i]]}!!:ij]]i]:]_i;l::i]ii_i]_]]],_]]_![!iti ]]i]]];,il}_!i]:t ::!i.],_
:i."..-]i::iit!!iitiiiil!iiltii!!i_lli::iti!!iiiii!tii!iti::it_iii_i]i]i_i]::i_i!!_i!ii!__ii!_!!i-!I_ ":l'i!: !i': ..... i? -:'t'.,;t "_i_L_: :::: :.: ..i.t,._.,_•_. _;:.' ;:: :i: I: '_-_......... t::.:t...,I....t.. ::::::::::::::::::::::::::::::::::::::-:,l:::.t:::.:::t::
290 295 300 305 310 315
PRIMARY RCE Isp, SEC,
LOX/METHANE RCS Isp SENSITIVITY (POD AV = 500 FT/SEC)
140
MCDONNELl. DOUGLAB AITROItlALITfCI COMPAItlY.IT. LC_iL_IB DIVIBIOftl
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AuxiliaRy PIopulsioN System Study
FINAL REPORT °, ° - .?REPORT MDC E25,48
JULY 1982
• FIXED POD VOLUME
• RCE DESIGN CHARACTERISTICS
F = 870 LBF
PC = IO0 PSIA= 18
MR = 2.49
• CURRENT OMS A V = 500 FT/SEC
"ii:: L: ::::I:::::::! i::!
:::::::::::::::::::::::::::::::::::::::::::::::::::: l:!!:_:::q:Y.-j. : '........ ::_t!::: :Li:t:i _i:':i:2]'_'q"': ::':_:::........... :::t .............::!]i_.: ":'t": .....'::::_:]: :::4:::::::::!i::_:_:,:;_: "::!_]i:: "._-._:.t.:_:_-_ _"_:h:.::..:-::
504
l,t
503'h--LL
"':! ...... :.................. _ ................... : ::'.:::t::: :':'= ....:.,,:,:-:::!::::'.']:: :::: ::-:ft.:.: :::t:::::.._: ::::t_:::: :.:l:i:;i-i::........... , ....... :::..-::| .... ";':::::" :: ::: :::::: .- ........... _ ....* I ......... t ................ _ ...... ................
...............1, .......,j..-_. .. I ............ " • • ". .... =............ - . :_. :...... ; ............... _ ........... ;. _ :' :. • .._ ..... , ..
.......... _ ........ ; ........ _.;..._' . . ; • .... .. , . :_ , . _ -I : _--:: ;:: , : f ::: :.., ........ 1........ t ..... : . :; :.: :tt : .:.:
:!i :i!:i::::::h:::.::d':-:::::::ii:_:! ::::.. ::_:: ji _-::_: .......
-:_: ::!:!{:!!:ii:!.::.:_ !i!i !!i- :!i:-::__:::' ::-: - :!:i:!l:_:i_i: _" _ ._.:_..:::_:..: :_::;:::: ::::1:::: :.'tL!.: ::_:_:?_.1_ _fl:,:t_ J._tll! . :: | ..:. ::
:::.{:::" :::4 ::: ::::L::: :::{::_: :': ............... _................ _': _"FZ"-.......... :::::::::::::::::::::::::::::::::::::::::::::::::::::::502 ::t....:...._:,:::.::::,:::.t,.......' .........::-.:,_:i::: ::_:i_:_:li_:':-:_!l_....t!_i!t!i_t_!i_:::_i_.:-
_:: ':::: ':-'::_).............. _.:._ i:!:t:;.'.: ::::'.::_:,::_:t,:: • .......... :, .... :::.::i:
o .:::i:;:,::!ii::i;,;:;I:::::::_::i::!::::::::::::::::::::::::::::::::::::::::::::::::::::::.: :o ;.': !:. : : " "-- ',_..;,- .... ::: _'-4 .: . : _.' - ' -
:::: :.l :: :_,; ::: ::" ' " ;;:';' :: 1;;1 ;'i _ "_._ .*;*I .-., : ;- ..:;; .._,: .:_,_
::!:.:.:. : .:.:::::::..::. ::,.: ,: : i. ..:::-:::... _..:..].:_:...::::::::,_:;::-:.........................................,.... ................I.:::.::-:"-" " " ! ....... :::l'." ......... I:" :1_.:: .... --_,• , .," . _......
2.95 300 305 310 315
501
50O
290
PRIMARY RCE Isp, SEC.
FIGURE 86 LOX/METHANE RCS Isp SEHSITIVITY (FIXED POD VOLUME)
141
MCDOItlNEL L DOUGL AE A _ TRONAU TIC S COR4P4 N lY - _T. L(:_I,IIB DIVI_IOItl
LOX/HYDROCARBON
Auxiliary PRopulsion System Study
FINAL REPORT REPORT moc E2548
JULY 1982
limiting the ONS_V and aft-RCS total impulse to their current values, while the OMS
performance (&V) sensitivity was developed by constraining the system volume to thecurrent aft pod envelope.
I
!
7.4.4 Sensitivity to Tank Wall Minimum Gage Thickness - System weight sensi-
tivities to propellant tank wall minimum gage thickness are shown in Figures 87 and
88. These data were generated assuming 2219-T87 aluminum tanks for both fuel and
oxidizer and sizing the tanks to make maximum use of the current pod envelope. A
design minimum gage thickness of 0.06 inches was selected to provide resistanceagainst handling loads.
7.5 Separate versus Common FRCS/Aft Propulsion Tanks
Comparisons of separate versus common propellant tanks for the FRCS and aft
propulsion pods are shown in Figure 89. These comparisons were performed for
LOX/ethanol with the_pod volume constrained to the current dimensions. For the
common system, feedlines are routed along the length of the Orbiter to interconnect
the forward and aft pods (Figure 4) As shown in Figure 89 the common system
provides lower OMS AV capability due _he loss of available propellant volume in
the nose. (For these comparisons, 100% of the current RCS total impulse
requirement was provided.) This loss in propellant volume can be compensated for
by employing conical shaped tanks in aft pods as discussed below.
7.6 Conventional versus Conical Propellant Tank Sh_
Comparisons of OMS AV capability for conventional and conical shaped
propellant tanks are presented in Figure 90. The conical shaped tank employs a
conical barrel section with an ellipsoidal end dome ar;d hemispherical forward dome.
This geometry enables the propellant tank to conform more closely to the pod
moldline and provides increased propellant volume within the pod. As shown by
Figure 90 an OMS_V of 630 ft/sec per pod can be achieved using conical tanks in-the
integrated forward and aft propulsion system concept. This is well in excess of the
500 ft/sec provided by the current storable system. Furthermore, if conical tanks
are employed for a separate aft propulsion system (OMS and ARCS), an OMS AV of
690 ft/sec could be achieved. On the basis of this evaluation it was concluded that
142
MCDONNELL Ir'#OUGLAB ABTROIt/AUTICa COMPANy.B'r. 1_18 DIVImll_l
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
'32tC'?;" .... ": ' "
REPORT MDC E2548
JULY 1982
• FIXED POD VOLUME
18,600
18,200.04
FIGURE 87
TANK MINIMUM. GAGE THICKNESS, IN.
•i0 .11
LOX/ETHANOL TANK MINIMUM GAGE SENSITIVITY
143
MCDONNELL DOUGLAi 41TROItl4Ulr'ICB COMPANY.lIT. Lfr.._LII$ DIIVIIBI_
LOX/HYDROCARBON
Auxiliary Propulsion SysTem STudyFINAL REPORT
ORtGINAL P; '.Go 15
OF POOR QU:_LJ'Y
REPORT MDC E2548JULY 1982
!
I. FIXED POD VOLUME
16,0 O0
15,900
=E:
15,800,..I
--r"C_'O
i,i:Ic
I--i,I
o9
-I,-
CD
0
15,700
15,600
15,500
15,400
• {:'t'"'i:':t"':":_:::'_" , .''--t 4. "1"'!.:._'"' .............. ;
---_-_-:-_-4-:.'-4.._.1 ":'_ :!t'::i .... tL_"t': i I" :"iii't ::i''i_':i :: I.I i =........... " • . "-::-::: ........... _ ..-T-:: .-'_..---j .... :--
_-_._i:i:.' ...... _i_.;.L_'. ...... :T. :: ": -:-i-:-_,--'::il l! :::. .. "•:::::::: .::h:.. ':::t:::: " ::!'::_ :: ::.: ::: i_.:T:::: :::: :::" :;:: :::: ::: :::::::" t • "
IIIII.I.II.II::I::.i:.:li.:.:..II ....,;I..........:".... • ;I--,!!!f!i_iIt!i!i!iifi!!ii_ii}gitl_5i!_.-:_!t!i_tiili_i_Ii_iilt!_t:.---!_ii_f_iitii_}__:t .-_!ii} iit:i;','iiii:--;: ::!_ !::- :::: ::::-_u.,., :._ ........ _.... _:;" :_'" "'" ;'_ " "4 ,.'::_"*"-' _4_ = I ' j "i T:
::::,::::Ix:: :E; :::: :::: ."F.:-'7 ii_L".:: =:: :::: "::4:_ .:_:!'" . _:i-:""!:'_"-:I I_"-"":': .::i:::._ :_;,-..,, ..: ........ . ., - :_:_ :-.t.:..f--_-.. :_--.t ............... :: :t:
:;:': .......... :;' "H _,:_'; _:" ;;_i _V...... , " ::'t": ""'I":: ::::i :ii! ::" :::" ": "
i'!2i!!ii i!!i:!i!i ii_..::
:'i;t.:: _ !!:: ::" :-_; ::;{ :::._i:-! !iiq!..'. ,..... _-.-q,-.... i_i._.Zi _ "LI_:; "L:-i:2! : "
i_i_::,i ;:v _:, ;; .... -. ': :.... " • :i __:_ -_i ::7. !'." "'i: ._;L_!!_i ::.::.÷_-_ :::: "::: -::_:T!
::::t:!. _-:. i: I: _;" :::: :: ::::" :::: _:._:._: :.':: :::_:-:_ ._:_" --- :Z-.:::z :_ :=: ::.:_::::
-.*. £;;_ 4_.,;4 "4' *4 _. ;,44 L._4 _H_ h444 "'/_- .ad.' " h' a-" ' ' " " :"........ "1 ...... _"'_t I .... "-,-, , _-_-., _ *, ....... _._'_ _........... _'_-,-' ----:
-:.-' ..... ::.:. ::. .._:: :::i :'::: ";: ::: ::"I ......... ' .... {..,:_.-..-L.._!.?'..;:-':'"'".: : :"
•04 .05 .06 .07 .08 .09 .I0
TANK MINIMUM GAGE THCKNESS, IN,
FIGURE 88 LOX/METHANE TANK MINIMUM GAGE .¢(NSITIVITY
144
R4CDONItlIELL DOUGLAB ASTRONAUTICS COMPANY.BT. LGHLll_ DOVOBIOIV
LOX/HYDk_)CARBON
AUXI|iAR), PxopuL_io_ SySTeM STudyFINAL REPORT
f
REPORT MDC F..2548
JULY 1982
i---EL
O
600"
500"
40OSEPARATE
F-'-"
COMMON
(FIXED AFT POD VOLUME)
:EQ:
"_ 21,000-
P-":m
_- 20,000-c,J
_';19,000
÷
18,000c_
',F
17,000CD SEPARATE COMMON
mm
3500"
= 3000"
÷
÷
2500Z(D
ISEPARATE COMMON
FIGURE 89 SEPARATE VS. COMMON FRCS/AFT PROPULSION TANKS(LOY/ETHANOL)
145
MCDONItIELL DOUGLAB 4BTRO/tiAUTICB COMPANY.BIt'. LOUIS I.;)IVIBION
LOX/HYDROCARBON
Auxiliary Propulsion System Study
FINAL REPORT
ORIC-!:L::'..,,.............
OF POC"-_ :",
REPORT MDC E2548
JULY 1982
I
|
I
I
700.
400
(FIXED POD VOLUMES)
FULLY INTEGRATED
FIGURE 90 CONVENTIONAL VS. CONICAL
(LOXIETHANOL)
700.
_I <
¢.J I400 . , ,
SEPARATE FRCS/AFT PROPELLANT TANKS
PROPELLANT TANK SHAPES
LOX/HYDROCARBON
AuxiliAry Propubion SySTeM STudy
FINAL REPORT REPORT MDC E2548
JULY 1982
a conical propellant tank can provide a substantial increase inAV and total impulse
capability for a pump fed system.
gV
7.7 Pump versus Pressure Fed FRCS
Comparisons of pressure and electric pump fed FRCS are presented in Figure 91.
These comparisons are for a separate LOX/methane FRCS having a thruster chamber
pressure of 100 psia, which was found to be near optimum for both the pressure and
electric pump fed FRCS. As shown in Figure 91 the pressure fed FRCS has lower wet
and dry weights. As such, a pressure fed _ystem (Figure 21) was baselined for a
separate FRCS. It is not only lower in weight but has fewer components (no pumps,
liquid accumulators, or batteries) and provides the same performance (Isp) as the
electric pump fed system.
7.8 Side-By-Side OMS/ARCS Comparisons
The final effort in the Phase II System Evaluation task was tn perform a
side-by-side comparison of the LOX/ethanol and LOX/methane OMS-ARCS with a
similarly configured LOX/H 2 system, as well as the current storable OMS-ARCS. The
resulting weight and performance comparisons are presented in Figure 92. The
LOX/H 2 system was configured to the same groundrules as the LOX/methane OMS-ARCS
(Figure 73) and employed a cryogenic liquid feed system for the ARCS. However,
because of its low _V capability (150 ft/sec) it is not a practical contender for
a "second generation" OMS-RCS. The LOX/ethanol is the best system concept because
of its highAV and total impulse capability. Ethanol is a storable propellant which
does not require a tank insulation system. Insulation is also avoided in the RCS
feed system (accumulators and lines) by thermally conditioning the RCS 02 supply to
a superheated vapor (Figure 72).
147
MCDONNELL DOUGLAe 4BlrROI_AUTICR COMPANY.BIt.
LOX/HYDROCARBON
AuxiliAry Propulsion SYSTEM STudy
FINAL REPORT
ORIGIT_P,L F;.'.;:':= :._OF POOR _'JAL:'i'Y
(FIXED TOTAL INPULSE)
CHAMBERPRESSURE= 1_0 PSIA
REPORT MDC E2548
JULY 1982
_=-
,,.I
I--
4000 -
3500'
3000
1000,
=E"' 900'
,,, "" 800'
_ 700"L,
60(
st- t., I
FIGURE 91 PUMP VS. PRESSURE FED FRCS (LOXIMETHANE)
600"
500"
"-. 400-I,-,,
< 300-x
200"
100
z
I,-
x0
L=J
¢=:
_--_r
X
.-I _-
r--i_
18,000"
17,000-
16,000-3
15,000.
÷ 14,000.
13,000
oz,c¢.=r.p-.LI.J
x
....I
i
OMS AV = 500 FT/SE_C
""I
|I
Im
Iz
,-- _=IE
0
xo
F-q
• : 3000,
2500,e_
b,.
÷
2000
OMS AV = 500 FT/SEC
z
_:: F-I-- i_
x
o
r"-I
" DOES NOT FIT CURRENTPOD VOLUME
FIGURE 92 COMPARISON OF LOX/HC, LOX/H2, AND N204/MMHOMS..ARCS
148
MCDONNELL DOUDLAB ABTRONAUTICm COMPANV-B_ LOtWIE DIVlBlOft_
mmmmm =mm m .,.,,m,.m.
LOX/HYDROCARBON
&uxil.iARy PRopuL,_ioN SYSTEM STudy
FINAL REPORT
8.0 CONCLUSIONS AND RECOMMENDATIONS
REPORT MDC F._2548
JULY 1982
The overall study conclusions are summarized in Table XVI. An integrated
LOX/ethanol OMS-ARCS (Figure 72) was selected as the best system approach because
of its superiority in terms of OMS AV and RCS total impulse capability. The
LOX/ethanol system allows use of a simple, non-insulated RCS feed system, and
recent tests--Reference (3)--have showr that the LOX/ethanol propellant combina-
tion is clean burning (non-coking) Because the propellants are low in cost,
non-toxic, and non-corrosive, the operational costs for a LOX/ethanol OMS-RCS
would be substantially less than the current N204/MMH system.
A pump fed OMS was selected over a pressure fed system because of overriding
weight and performance advantages. For two pods the pump fed OMS is approximately
3000 Ibs lighter than a pressure fed system. In addition, a single turbine drive
for both ti,_ fuel and oxidizer pumps was recommended to reduce feed system weight
and complexity.
Common propellant tanks were recommendeo over separate tanks for the OMS and
ARCS propellants because they provide improved propellant packaging (higher 6V and
total impulse capability) and greater mission flexibility. Furthermore, to
provide maximum performance and avoid using the OMS turbopumps for ARCS propellant
feed, small, dedicated electric RCS pumps were recommended for resupplying the ARCS
accumulators.
A hybrid, ambient temperature RCS propellant feed system was recommended to
eliminate the need for insulating the RCS accumulators and feedlines. The RCS
oxygen supply is thermally conditioned to a superheated vapor using a passive
ethanol tank heat exchanger which avoids the complexity and vent penalties
associated with active hot gas generator--heat exchanger assemblies.
The new technology requirements associated with this feed system approach are
identified in Table XVlI, while recommendations for future feed system studies are
summarized in Table XVlII.
149
t'J4CDO,'tlitIELL DOUGLAS 4BTRONAUT'IC,fS COMPANt* . B'F. LgN, II,_ DIVIBION
LOX/HYDROCARBON
Auxiliary Propulsion SysTem StudyFINAL REPORT
TABLE XVi
OVERALL STUDY CONCLUSIONS
REPORT MDC F..2548
JULY 1982
Best Fuel--Ethanol
• highest AV and total impulse capability (OMS AV~600 ft/sec per pod)• non-coking
• earth storable (vapor pressure slightly higher than MMH)
• good technology base for engine development
Most Attractive System Concept
• pump fed OMS with single turbine driving both fuel and oxidizer pumps
- overriding weight and prformance advantages (pump fed OMS provides 3000 Ib
weight advantage over pressure fed OMS--2 pods)
- single turbine reduces system complexity
• common OMS/AFT RCS propellant tanks (common tanks provide 18 ft 3 more
propellant volume than separate tanks)
- high AV and total impulse capability
- greater mission flexibility
• electric pumps for AFT-RCS feed
- turbopumps cycled only during OMS burns (cycle life reduced by factor of 6)
- high RCS performance (electric pump RCS Isp is 21 seconds higher thanturbopump RCS Isp )
• hybrid ambient temperature RCS propelant feed (GOX/liquid ethanol__(no
accumulator or feedline insulation required)
• passive ethanol tank heat exchanger for 02 thermal conditioning
- low feed system complexity (no gas generators for thermal conditioning)
- no ISp penalty (gas generator vent loss)
150
MCJ'#ONI_m"LL DOUGLAI ASYRUN4UTIC8 fr_.UMPANyomy. L_la DIVISIC_ftl
i
1"
I
I
I
i
I
I
I
I
I
I
I
I
i
I
I
I
LOX/HYDROCARBON
Auxiliary Propulsion SysTEm StudyREPORT MDC F..2548
FINAL REPORT JULY 1982
TABLE XVII
NEW TECHNOLOGY REQUIREMENTS
Feed System
• thermal management system for cryogenic LOX tank
- insulation
- thermodynamic vent
- auxiliary cooling
• passive ethanol tank 02 heat exchanger
• surface tension screen propellant acquisition for common OMS-AFT RCS tank
(cryogenic)
• improved propellant gaging approach
• electronic pressure regulator" for controlling RCS GOX accumulator outlet
pressure
• lithium batteries or alternate power source for electric RCS pumps
Engines
• LOX/ethanol OME
- small high speed turbopumps
- improved heat transfer characterizations, burn-out data, and performance
correlations
• LOX/ethanol RCE
- improved heat transfer characterizations
- pulse mode performance and cycle life capability
151
MCDONNELl. OOUOL.AE A'TRONAUT'ICB COMPANY . BT. Lii_JIIJ DIVIBION
LOX/HYDROCARBON
Auxiliary Propulsion system StudyFINAL REPORT
TABLE XVlll
RECOMMENDATIONS FOR FURTHER FEED SYSTEM EFFORT
REPORT MDC F..2-348
JULY 1982
I
I'i
definition of LOX tank thermal management system (considering ground hold,
transient launch, and on-orbit heating effects)
- tank insulation materials and thicknesses
- thermodynamic vent system sizing
- auxiliary cooling capability (pumps, tank supports, etc.)
detailed evaluation of integrated forward RCS/AFT propulsion system (impact ororbiter interfaces)
evaluation of system performance over broad mission spectrum
- OMS-RCS mission duty cycle extremes
- limitations of ethanol tank 02 heat exchanger
realistic RCS thruster pressure/temperature boxes to begin thrusterdevelopment
definition of system controls and failure detection/isolation requirements
• Definition of component ROM costs and schedules
propellant tanks
- pressure regulators
- valves
- accumulators
- OME
- RCE
I
I
I
c_.'l
152
IWCDONNELL DOUGL AB AB'rRGNAU'FICB COMPANY. ST. L_le DIVIBION
I
I
LOX/HYDROCARBON
Auxiliary Propulsion System StudyFINAL REPORT
REFERENCES
REPORT MDC F..2548
JULY 1982
o
=
G.F. Orton, T.D. Mark, and D.D. Weber, "LOX Hydrocarbon Auxiliary Propulsion
System Study Phase I Report, Volume I--Technical and Volume If--Appendices,"
McDonnell Douglas Corporation Report No. MDC E2479, February 1982.
G.F. Orton, T.D. Mark, and D.D. Weber, "LOX/Hydrocarbon Auxiliary Propulsion
System Study Phase II Report," McDonnell Douglas Corporation Report No. MDC
E2547, July 1982.
o
"Combustion Performance and Heat Transfer Characterization of LOX/qydrocarbon
Type Propellants", NASA JSC Contract Number NAS 9-15958.
153
MCDONNELL DOUGL AB 48 TRONAU TIC8 COMPAItl y. BT. LOUIB DIVIBIONI