Table Of Content"_'_ 1
AIAA-2002-4206
Oxidizer Selection for the ISTAR Program
(Liquid Oxygen versus Hydrogen Peroxide)
Jason Eugene Quinn
NASA Marshall Space Flight Center, Huntsville, Alabama
AIAA/ASME/SAEIASEE
Joint Propulsion Conference and Exhibit
38 th
7-10 July 2002
Indianapolis, Indiana
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OXIDIZER SELECTION FOR THE ISTAR PROGRAM
(LIQUID ()XYGEN VERSUS HYDROGEN PEROXIDE)
Jason Eugene Quinn, NASA Marshall Space Flight Center, Huntsville, Alabama
flight velocity range would demonstrate all of the
Abstract RBCC engine operational modes: Air Augmented
Rocket (AAR), Ramjet and Scramjet. Although only
This paper discusses astudy of two alternate the ISTAR propulsion system development and the
oxidizers, liquid oxygen and hydrogen peroxide, for use initial ground test program is currently funded, due to
in a rocket based combined cycle (RBCC) demonstrator the tightly integrated nature of a hypersonic vehicle,
vehicle. The flight vehicle is b.Lselined as an initial demonstrator vehicle design is being performed
airlaunched self-powered Mach I).7 to 7demonstration to define the propulsion system requirements.
of an RBCC engine through all _f its air breathing Initial conceptual design of hypersonic vehicles
propulsion modes. Selection of an alternate oxidizer requires the designer to assume many system
has the potential to lower overall vehicle size, system characteristics such as propellants, thrust, and vehicle
complexity. /cost and ultimateb the total program risk. size to satisfy mission objectives. The designer uses
This trade study examined the oxidizer selection effects their experience on previous similar vehicles to decide
upon the overall vehicle performance, safety and what to select. Although these initial "guesses" are
operations. After consideration of all the technical and often quite good, they can limit the ability of the design
programmatic details available at this time, 90% if one does not go back and check their assumptions.
hydrogen peroxide was selected over liquid oxygen for For example, one of the first items to decide upon is the
use in this program. propellant combination that the system will use.
Typically the designer would select the highest ISP
Introduction combination propellant, but this might not be the best
selection for non-orbital missions. A lower performing
The Advanced Space Transportation Program at (in terms of ISP) propellant combination that is more
NASA Marshall Space Flight Center has assembled a dense or non-cryogenic could result in a smaller, lower
government / industry team to conduct the system cost, more reliable or more operable vehicle system.
development and ground test of a RBCC propulsion The ISTAR program has conducted several trade
system (and potentially flight). Government team studies on the "right" propellant combination to select
membership includes participants from several NASA for this program. Selection of a hydrocarbon (HC) as
centers: Dryden Flight Test (e_ter, Glenn Research the fuel for the ISTAR program over liquid hydrogen,
Center, Langley Research Center, Marshall Space although an interesting trade study, will not be
Flight Center and Stennis Test ('enter. The primary discussed in this paper. Once the decision to use a
industry team member is the Rocket Based Combined hydrocarbon fuel was made, there were still at least two
Cycle Consortium (RBCJ), whi,:h includes Boeing oxidizers that could potentially meet the system
Rocketdyne, Gencorp Aerojet and United Technologies requirements: liquid oxygen (LOX) and hydrogen
Pratt & Whitney _. Additional industry team members peroxide (HTP). The purpose of this paper is to detail
include Boeing for vehicle actix ities and several other the trades that the ISTAR program went through in
companies. selecting between these two oxidizers.
Current plans for the prop_dsion system
development and ground test ot the RBCC engine General Considerations
system are funded through NASA MSFC as the
Integrated System Test of an Airbreathing Rocket Most of the rocket industry moved away from the
(ISTAR) program. The primar,. ISTAR program use of hydrogen peroxide in the early 1960's for several
objective is to develop a propulsion system, which reasons: large all rocket propulsion systems went to
would be capable of powering z_flight demonstrator LOX due to the increase inperformance (ISP), long
vehicle from launch off a B-52 aircraft (approximately term storage users went to NTO/MMH and mono-
Mach 0.7) up to scramjet speed_ of about Mach 7. This propellant customers switched to Hydrazine. The
recent resurgence of interest in HTP in applications
where LOX would typically be used is in part due to the
This material is declared a wo] kof the U.S.
realization that higher ISP is not the entire story.
Government and is not subject 1,,copyright protection
Higher propellant density and storability can, in some
in the United States.
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cases, make HTP a better choicc than LOX. This is In order to determine if our vehicle would be
especially true for systems that have aerodynamic drag "better" with a particular propellant combination we
losses or are severely volume constrain•ed 2;ISTAR has can calculate the propellant volume required for our
large aerodynamic drag (relative to an all rocket) due to given mission AV and assumed mission averaged
its airbreathing trajectory and severe volume constraints effective specific impulse (I*) _.
IA/
due to its slender hypersonic shape.
Hydrogen Peroxide has physical properties very
egl - 1 Masst_,,a/
similar to water (i.e. density, cok_r, viscosity, etc.) 2.
The primary exception is that th_ molecular structure is Volume =
only meta-stable and will exotht:rmically decompose
iOmlxlltre
from H202 (basically a water m,,lecule with an extra In the case of ISTAR our initial vehicle utilized
oxygen atom attached) at some _ateto a water the rocket thrusters only for AAR Mode which is Mach
molecule, oxygen and energy. "lhis hot steam and 0.7 (air drop) to Mach 3.0-4.0 where the rocket is to be
oxygen (1300 °F) can then be expanded out a nozzle as turned off (see Figure 10) and the engine transitioned to
a monopropellant thruster or the hot oxygen can be Ramjet mode. The vehicle is assumed to perform
combusted with a fuel. In either case the hydrogen in identically for either oxidizer above Mach 3.0-4.0 as
HTP istightly bound in a water molecule and is no__!t Ramjet/Scramjet modes only use tanked HC fuel and
combusted. When the concentJ ation is quoted as 90% air. For a first cut analysis assume that the mass of the
that means 90% Hydrogen Peroxide in solution with vehicle isthe same at the end of AAR mode (Mach 3.0-
10% water and traces of stabilizers. In a compatible 4.0) for either oxidizer (which approximately
storage container 90% HTP has been observed to corresponds to AV of 2000 to 4000 ft/sec). With this
decompose at less than 1% per decade but in the
assumption we can plot, Figure 4 which shows the
presence of a catalyst HTP can he caused to decompose propellant volume required with either oxidizer for the
extremely rapidly 7. HTP has been used as a
Air-Augmented Rocket (AAR) mode of operation. The
turbomachinery drive gas in many systems: V-2, X-l,
point made by Figure 4 isthat below a mission AV of
Redstone, Jupiter, Centaur, Viking, X-1 and the X-15
approximately 4000 ft/sec HTP/HC will require less
,_.4 HTP has also been used as the primary oxidizer (bi-
propellant volume than LOX/HC for the I* value
propellant) in many propulsion _ystems: Me 163 assumed. These I* values were generated from our
Komet, Gamma 201./301, AR series of rocket engines,
three degree-of-freedom trajectory model. If you
LR-40 and others 2"4"5'6(as the m._jority of these flight assume the same total propellant volume is available
systems used 90% HTP it was tiae first concentration
with either oxidizer, then one can plot Figure 5 which
considered). HTP physical properties of interest are
shows the I* required for each propellant combination
compared to LOX in Table 1and density compared to
at several values of mission AV. For example, Figure 5
several propellants in Figure l.
shows that for a AV of 2000 ft/sec if the I* for the
Figures 2 & 3show the id,:al specific impulse and LOX/HC vehicle is 240 sec than the I*of the HTP/HC
density specific impulse of LOX and several
vehicle needs to be 196 sec or higher to have the same
concentrations of HTP with hydrocarbon fuel. Density
or less propellant volume required. From these graphs
specific impulse is the density ef the propellant
we can conclude that below _4000 ft/sec AV a 90%
combination (i.e. bulk density or Pmi×t,_) multiplied by
HTP/HC thruster will require less tank volume than a
the specific impulse. It isoften t_sed to select
LOX/HC thruster including the effect of the additional
propellants when volume considerations are taken into
mass of the more dense propellant provided the I* of
account. 90% HTP isonly 15lhJft3denser than LOX,
the HTP/HC vehicle in AAR mode exceeds the I* of
but the peak ISP occurs at a mu_h higher mixture ratio
the LOX/HC value given in Figure 5.
(typically 2.6 for LOX/HC versLis 7 for HTP/HC). This
results in amuch higher bulk density for HTP/HC over
Background - Initial System Definition
LOX/HC. The argument can b_ made that for a
sufficiently low mission velocity change (AV), a lower In order to understand the conclusions drawn from the
performing (in terms of ISP) denser propellant
trades performed some familiarity with the ISTAR
combination would yield a smaller propellant volume
vehicle (X-43B) and engine system is needed. The
required. Typically we assume that the mission cost is
ISTAR vehicle was designed around an existing
roughly proportional to vehicle _ize - thus the lower
hypersonic vehicle shape designed for liquid hydrogen
performing propellant would b_ a "better" choice.
which was then modified for our particular mission and
propellant combination. Figure 6 shows three views of
Generic Comparison
the current configuration with the propellant tank
location /volume emphasized. Generically this
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configuratioisnaliftingbody__thexternafolrebody components including the ignition system (baselined
compressiomno,vingcowlflapinletwithfixedinternal combustion wave ignition [CWI] tbr the LOX/HC
geometaryndexternaalftexpansion. system). The components highlighted with ared
background are those that are different between the
Agenerimcission(showinnFi_,.u7re)beginwshenthe LOX and the HTP system - note that the primary
ISTARvehicleisdroppefdromtheNASAB-52at difference between figures 8and 9 is the LOX top-off
Mach0.7and40,000ft,acceleratteosapproximately system and the ignition system (CWI for LOX/HC and
Mach7at90,00f0tbeforeshutlingdowntheengine
catalyst packs for the HTP/HC system)
andglidingbackforre-useA.fterbeingdroppefdrom
theB-52vehiclethevehiclefreclhllsforafewseconds
Figure 9 shows the HTP/HC engine system. This
beforestartintgherocketthrusterasndacceleratiinng system is very similar to the LOX system but replaces
Air-AugmentRedocke(tAAR)mode. In this the CWI system with a catalyst pack arrangement to
propulsion mode the rocket thrusters are firing at full provide auto-ignition in the rocket thrusters and doesn't
thrust and additional fuel is inje,:ted to bum with the require LOX top-off. These two changes reduced the
incoming air. As the vehicle accelerates through Mach number of fluids on the vehicle and the complexity of
3-4 the rocket thrusters are turucd off and the vehicle the functional schematic considerably.
continues to accelerate in Ramjet mode. Upon reaching
Mach 6-7 the vehicle transitions fully into Scramjet
mode before shutting down - completing the Detailed Oxidizer Trade
demonstration of all the airbreal hing propulsion modes
and the transitions between them for an RBCC engine. The ISTAR engine system was originally baselined to
be LOX/HC as these propellants are familiar and were
The ISTAR team baselined a simple engine system believed to provide enough performance without the
based upon the type used in the X-15 (and many other severe volume penalty of LH2. Prior to the initial
previous systems) using a high-pressure tank of HTP to formation of the RBC 3team a conceptual trade study on
drive the fuel and oxidizer pump. The functional the propellant selection indicated that replacing LOX
schematic for this type of system with either LOX or with 90% HTP would allow a smaller propellant
HTP as the primary oxidizer isshown in Figures 8& 9. volume to complete the mission (Mach 0.7 to 7). This
This type of system was selected to keep the rudimentary study results and several discussions
development costs of the engine system as low as between NASA and RBC 3provided motivation to
possible and still allow the RB( (2multi-mode complete a comprehensive trade study examining in
operation. The ISTAR program is focused on engine detail the system impacts of switching the oxidizer from
flowpath performance throughout the mission LOX to HTP.
trajectory, especially in mode transitions, and not in
engine system development. [)_e to flight experience The oxidizer trade study team brainstormed a detailed
with many previous programs, !_!)%HTP was baselined list of the important criteria that were judged to affect
to be used for the turbine drive _,.asrather than a higher the entire system design. Table 2 shows these criteria
concentration of HTP, which m_ght have increased grouped into five sections: Safety, Programmatic,
performance (and increased risl,.). Mission Success/Engine System Design, Mission
Success/Vehicle Integration and Operations. Within
Figure 8 shows the LOX/HC engine system and all of these five sections the criteria were also grouped into
the functional components required to operate in the categories (i.e. vehicle design impacts). Each of the
engines different modes. As previously stated the five sections was assigned a weighting factor, which
system uses a high pressure HTI' tank to provide hot attempted to capture program management's preference
gas to the turbine drive (yellow_ As this tank is not or importance level for that section. Each criteria was
linked to the main oxidizer tank we can run just the fuel to be assigned a 1, 3or 9 score for how beneficial an
pump (no rocket thruster oxidizer needed in Ram / oxidizer was to the system on this criteria. Scores were
Scramjet modes). Figure 8shows the vehicle systems then combined with the weighting to produce a single
in the top left portion of the figure with the main fuel ranking for each oxidizer.
tank, high-pressure HTP tank and main LOX tank along
with the gray purge gas tanks. I he top right portion of Each of the criteria was assigned a criteria owner(s)
the figure shows the systems required on the carrier who was responsible for investigating the criteria,
aircraft (B-52) primarily the LOX top offtank (note the selecting a score and providing agroup presentation to
purge gas supply was the same for both the HTP and back up that score. The criteria owner(s) then presented
LOX systems and isnot showni. Finally the bottom the score to the entire trade study team for discussion
left side of figure 8shows the engine system and the trade team selected a final consensus score.
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Theprocesosfassigninagteamuonsenssucsore allot a slight advantage to HTP in facilities /test costs.
requireadlloftheteammembetro,becomeeducated The final criteria of programmatic risk determined that
aboutthecriteriabeingconsidereTdh.iseducational selecting HTP would add more overall risk to the
procesesnhancethdeteamobjectivitayndtheentire program. This risk would be primarily in the beginning
tradestudyC.riteria owners we_e selected from the of the project during the thruster development while
RBC 3team members, NASA centers and Boeing. The selecting LOX would potentially delay the risk of the
entire team leaned heavily on the recent operational LOX top-off system to the flight phase of the program.
experience with both LOX and IITP at NASA
STENNIS and the flight operati_ms experience (both Mission Success / Engine System Design
recent and historical) at NASA 1)RC.
The single most controversial criteria in this oxidizer
Very early in this trade study th,: trade study team trade study is the technology readiness level and
explored the potential for use ot _)8% HTP rather than associated risk in developing the rocket thrusters with
90% HTP. 98% HTP was considered as there was either oxidizer including oxidizer cooling. Due to the
additional performance over 90" 0HTP. However, for need to cool the significant surface area of the duct with
this low AV mission, the additienal unknowns and the fuel being fed to the rocket both the LOX/HC and
development risk with 98% HTt' was judged to be not HTP/HC were assumed to use oxidizer cooling of the
worth this small performance boost. The remainder of rocket thrusters.
the trade study was performed cumparing 90% HTP to
LOX only. Selection of LOX allows us the comfort level as the
thruster would be similar to existing experience base
Safety (combustion chamber parameters) but the requirement
for LOX cooling introduces a number of system issues
This category was assigned the heaviest weight at 25% which increase mission success risk. NASA GRC has
(see Table 2) but was the least c_:,ntroversial in team test fired LOX cooled thrusters, which addressed the
discussions. Peroxide was judg,-d to be safer overall majority of the issues with LOX cooling, but they did
than LOX primarily due to the m:ed for a LOX top-off not address the closed loop issues (cooling LOX was
system to transfer LOX from thu B-52 to the X-43B not injected and burned in this test series). The closed
vehicle (similar to how the X-15 worked). This LOX loop issues like the two phase regenerative cooling
top-off system would have had !ooperate on the during the start transient is risky and potentially would
manned carrier aircraft during the entire flight up to X- not be discovered until late in the engine system
43B drop. HTP has a higher degree of risk with leaks development. Other potential risks with LOX cooling
and spills in the engine / vehicle system but was include freezing of the HC in the lines and the difficulty
considered safer than LOX when considering the entire of dealing with fuel temperature changes as the vehicle
propellant handling process, l"lLese two criteria accelerates through the Mach number range.
balanced one another out in the scoring leaving HTP
Selection of HTP limits the material selection of the
scoring safer than LOX for the iS;TAR project.
thruster as many metals (including copper) catylize the
Programmatic decomposition of HTP and are typically not considered
for use in HTP systems. After several iterations NASA
The next three categories were considered to be of and RBC _developed the design for two different rocket
equal importance and were all _wen the same weight of thrusters (using 347 stainless steel) each of which was
20%. The programmatic criteria required much capable of performing the mission with some film
discussion and work before the team could agree on a cooling. Other difficulties introduced by the selection
consensus score in the three areas of Schedule, Cost and of HTP is the difficulty of packaging the catalyst packs,
Risk. After significant discussi,:,n the schedule for and a lower thermal margin in the thruster design.
either oxidizer was determined tobe equivalent After several meetings on this criterion the team
provided additional money was made available and /or determined that LOX was preferred over HTP in the
more risk was accepted for HTI' The initial thruster design.
development cost with HTP wa, considered to be
significantly more (on the order of several million Pump and Ignition Systems were judged to be of equal
dollars) primarily due to the need to develop the HTP level of difficulty. The inherent risk in the system due
cooled thruster. Considering thu entire system HTP to the startup /shutdown transient and mixture ratio of
would save significant dollars (_3-5 million) due to not the engine is higher for the LOX system (more
needing a LOX top-off system bat the additional cost difficult) primarily due to the need to start the thrusters
for facility modification influen,:ed the team to only
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LOXrichandpassthrougshtokhometriacndhowthe prevent the formation of ice on the vehicle (also would
LOXdensitcyhangedsuringlh(start. reduce the amount of top-off needed from B-52).
Assuming an insulation thickness of 1to 2 inches yields
Theenginseystemdesigcnriter+faavorinLgOXover an amount of insulation equivalent to 10-30% of the
HTPare:lessknownissuewsiththeoxidizetrhatmight total LOX volume. HTP was preferred over LOX for
requiredevelopmeenfftort(knownunknownsa)b,etter the vehicle integration performance criteria due to this
currendtesigenxperienbcaeseh,igherrockeItSP higher potential volume for propellant.
(smalaldvantagaen)dlikelylongehrardwarliefe
(primarildyuetohighetrherntamlarginintheLOX The remaining vehicle integration criteria dealt with the
system). Criteria favoring HTP t,ver LOX are primarily carrier aircraft. The oxidizer selection was felt to
due to the lower system comple _ity. Selection of HTP heavily influence the impacts on the carrier aircraft on
significantly reduced the engine system complexity in terms of consumables required to be in-flight
terms of number of propellants and complexity of the transferred. HTP would require only nitrogen (or
ignition system as can be seen i1_Figures 8& 9. The helium) while LOX was assumed to require a higher
LOX/HC system shown concep_ ually in Figure 8needs amount of nitrogen as well as LOX for top-off and chill
an ignition system (CWI) and a I,OX top-off system. down. Additionally the avionics /control onboard the
While Figure 9shows that the I-tI'P/HC system uses a B-52 would be much more critical as it would be a
catalyst pack for ignition and dc,csn't require a LOX manned system with LOX venting. All of the above
top-off system. Also HTP is likely to require less purge considerations resulted in the entire section of the trade
gas and no chill preconditionin_ of the oxidizer matrix associated with vehicle integration to be heavily
hardware. weighted toward the selection of HTP.
Mission Success / Vehicle Integration Operations
Unlike the Engine System Design criteria discussed Operations had the lowest weighting factor at 15%.
above all the Vehicle Integratiol_ criteria trades favored This lowest weight does not mean that the team did not
the selection of HTP over LOX Due to the wedge consider operations important (it is) but this is an
shape of hypersonic vehicle resulting in a low available engine system for an X-vehicle only meant for _25
volume for propellant, the prowllant tanks are typically flights+ Most of the criteria considered in the operations
required to be integral (the tank isthe vehicle). Integral area favored HTP over LOX with the exception of
cryogenic LOX tanks have never been developed and operational procedures being well established. NASA
were judged to be more difficub than the material has had recent experience at STENNIS in the E3 test
compatibility issues with integr_t tanks for HTP. With stand with both LOX and 90% HTP but LOX
these considerations HTP was c_bviously preferred from procedures are simply better know and more
a structural criteria. established. HTP was considered to have slightly easier
handling than LOX (NASA STENNIS & DRC
Examining the engine system p;sckaging and propellant experience), easier ground operations with less
feed system issues introduced vta each oxidizer came equipment and less vehicle ground operational
down to the non-cryogenic nature of HTP difficulties (i.e. simpler setup/servicing).
removing/reducing the need lbr vacuum jacketed lines,
cryogenic insulation, with the p,nential for removal of Summary and Conclusions
the boost pump (considered part of the vehicle). HTP
would still require some thermal management but the A trade study considering two alternate oxidizers,
non-cryogenic nature was felt to be much easier to deal liquid oxygen or 90% hydrogen peroxide, for a rocket
with. based combined cycle demonstrator vehicle was
completed. This trade study considered the overall
As previously discussed the hig her density of HTP/HC system performance from both a technical and
in a volume limited, low AV vehicle like X-43B programmatic viewpoint, to select the lowest risk
compensates for the lower ISP ,>fthe rocket thrusters. solution. Given the limited energy requirement (AV)
This assumes that the same volume is available for the of the demonstrator vehicle (Mach 0.7 to 7), the higher
propellant and the vehicle weight won't change with the density and mass ratio of 90% hydrogen peroxide
higher propellant weight. While the assumption that yielded similar vehicle performance when compared to
vehicle weight is relatively con<ant with propellant LOX. Additionally, hydrogen peroxide provided
weight is valid for this vehicle, ihe assumption of the system simplification, increased flight safety and
same propellant volume available is not. LOX would packaging advantages. After consideration of the
likely require some amount of (t)'ogenic insulation to technical and programmatic details, 90% hydrogen
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peroxidweasselecteodverliqudoxygefnoruseinthe
Transportation Applications," AIAA 95-2474,
ISTARprogram. July 1995, 22 pgs.
4. Wiswell, R., "X-15 Propulsion System,"
AIAA 97-2682, July 1997, 18pgs.
References 5. Butler, K., "AR2-3 Engine Refurbishment and
Gas Generator Testing," AIAA 99-2738, June
1. Faulkner, Robert F., "INTEGRATED 1999, 6 pgs.
SYSTEM TEST OF AN AIRBREATHING 6. Ventura, M. and Wernimont, E., "History of
ROCKET (ISTAR)," A[AA 2001-1812. the Reaction Motors Super Performance 90%
2. Ventura, M. and Mullens, P., "The Use of H202/Kerosene LR-40 Rocket Engine,"
Hydrogen Peroxide fol Propulsion and AIAA 01-3838, July 2001, 10pgs.
Power," AIAA 99-2880, June 1999, 19pgs. 7. "Hydrogen Peroxide Handbook," AFRPL-TR-
3. Escher, William J. D., Ilyde, Eric H. and 67-144, July 1967.
Anderson, David M., "A User's Primer for 8. Kit, Boris and Evered, Douglas S., Rocket
Comparative Assessm_,nts of All-Rocket and Propellant Handbook, The Macmillian Co.,
Rocket-Based Combined-Cycle Propulsion New York, 1960.
Systems for Advanced Earth-to-Orbit Space
Table 1: Oxidizer properties tompared 7's.
Fluid Properties 90% HTP LOX
Boiling Point, °F (exirap0iated _t1atm for H202) 286.7 -297.4
Freezing Point, °F (1 atm) 11.3 -362
Bulk Decomposition Temperature, °F (red line) 275 NA
Density, g/cc (H202 @ 77 °F, 1,t.7 psia; LOX @ -297 °F, 14.7 psia) 1.387 1.14
Density, lbnv_ft3(H202 @ 77 °E 14.7 psia; LOX @ -297 °F, 14.7 psia) 86.6 71.2
Heat Sink, BTU/Ibm (77 °F to 2'_0 °F for H202; -300 °F to 140 °F for LOX) 114.7 178
Critical Pressure, psia (estimated for H202) 3556 730.4
Critical Temperature, °F (estimated for H202) 833 -181.8
Cost, $/lbm 3-4 .042-.068
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American Institute of Aeronautics and Astronautics
Table 2: Liquid oxygen versus hydrogen peroxide criteria trade matrix.
Score
Category Criteria HTP LOX
Section
Test Personnel/Facilities 1 1
Safety Er,gine/Vehicle ISTAR Encjine/Vehicle 1 3
3 1
Handling Personnel/Hardware
Weighting Factor: 25% Carrier Aircraft B-52 Crew/Aircraft 9 3
14 8
Schedule SSC Ground Test 9 9
3 9
Programmatic Cost Development
Facilities Modifications/Test 3 1
3 9
Weighting Factor: 20% -_isk Cost &Schedule (Confidence in)
4 18 28
Thrusters 1 3
Mission Success I Engine Pumps 9 9
System Design Technology Development Ignition System 3 3
Level/Risk
Transient Operations 9 3
Known Unknowns 1 9
E_perience Base Expert Knowledge/Experience 3 9
Engine System Performance 1 3
Performance Design Complexity (No. Values, Pumps, etc) 9 1
Hardware Life 3 9
Purge Requirements 3 1
Resources
9 1
Weighting Factor: 20% Chill & Conditioning
51 51
Structural 3 1
Mission Success I Vehicle Ve.hicle Design Impacts Engine System Packaging 3 1
Integration Propellant Feed System 3 1
Performance Fuel Mar_in Remaining (FMR) 3 1
Consumables (Propellant, purge transfer) 9 3
Carrier Aircraft
Weighting Factor: 20% Avionics/Control Interface 9 3
6 3O 10
Known Hazards 1 1
Operations I_andling Operational Knowledge/Procedures 1 3
Transportaion, Handling, Storage 3 1
Test Operations 3 1
_;round Operations
Vehicle Ground Operations 3 1
Weighting Factor: 15% Logistics GSE, Expendables, Soft Goods, etc 3 3
__ 14. 10
Raw Score 31 127 107
3.5 2.0
Safety
3.6 5.6
Programmatic
10.2 10.2
Mission Success/Engine System r_esign
6.0 2.0
Mission SuccessNehicle Integrati(,n
2.1 1.5
Operations
126 106
Weighted/Normalized Score
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American Institute of Aeronautics and Astronautics
IO0
89.4
86.6
90 84.9
All Propellant Densities at
80 77 Deg. F and 14.7 psia
71.3
A unless noted below
< 70 62.4
_: 6o
49.5 50
e_
_ 50 45.5
•_- 40
30.8
c 26.4
_ 30
2o
lO 4.43
o T
H2 CH4 C3H8 C3H8 JP-7 RP-1 Water LOX 85% 90% 98%
(20K) (111K) (200 (90K)
(9OK) Hydrogen Peroxide
psia) r,f: HydroQinPe_xlde Handbook.A.CRPL-TR-'_7-144.auty196Z
Figure I: Densities of various propellants compared 7's.
34O 3
cp
_320 eE_
"_ 2.5
_3oo ,!m 2
_E
Vacuum
280 /- " Relative Specific v Propellant
®
/r--_-_ Density . Impulse _ Density
_260
_*- _:)_'oH/w Specific = Reference
E " --"- 90% HTP
® 1.5 14- 90% HTP Impulse Vacuum Reference
a 24o i H'P ._>
::ea :::o PS:o _..._ HTP Spp:i_ c X DensityPr°pellant
>
22O 1 i r
2 3 4 5 6 7 8 2 3 4 5 6 7 8 9
Mixture Ratio Mixture Ratio
Figure 2: Ideal vacuum specilie impulse for Figure 3: Ideal relative density specific impulse for
hydrocarbon (HC) fuel and various oxidizers. HC and various oxidizers, (note reference ISP =
417.2 sec, reference density = 20.3 lbm/ft^3).
8
American Institute of Aeronautics and Astronautics
200 260
' i
I
240
90% HTP/HC
150
-- - LOX/HC _220
0
• 200
13..
b-
100 " -_ i --3ooo_siec
S 180 i .......... ......
0
> , _-160
50i
j. ,f", "
140
_ i/.'" ...... ................... ii
120
0 d 1
150 200 250 300
0 1000 2000 30:)0 4000 5000
ISP LOX/HC (sec)
AV(ft/sec)
Figure 4: Propellant volume required on X-43B for
Figure 5: ISP required at given AV for equal
oxidizer selected plotted against mission AV,
propellant volume with either oxidizer.
jl
f _J
./" [
/
f
i'
Ox Fuel
Figure 6: X-43B RBCC vehicle with oxidizer and fuel tank volume /locations shown.
9
American Institute of Aeronautics and Astronautics
