NASA Technical Reports Server (NTRS) 20110013691: Thermal Optimization and Assessment of a Long Duration Cryogenic Propellant Depot

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·Thermal Optimization and Assessment of a
Long Duration Cryogenic Propellant Depot
RyanHonour1andRobertKwas2
ANALEX, KennedySpace Center, FL32899
GaryO'Nei13
NASA, Kennedy Space Center, FL32899
and
BernardKutter4
UnitedLaunchAlliance
A Cryogenic Propellant Depot (CPD) operating in Low Earth Orbit (LEO) could provide many
near term benefits to NASA space exploration efforts. These benefits include elongation/extension
of spacecraft missions and reduction of launch vehicle up-mass requirements. Some of the
challenges include controlling cryogenic propellant evaporation and managing the high costs and
long schedules associated with new spacecraft hardware development. This paper describes a
conceptual CPD design that is thermally optimized to achieve extremely low propellant boil-off
rates. The CPD design is based on existing launch vehicle architecture, and its thermal
optimization is achieved using current passive thermal control technology. Results from an
integrated thermal model are presented showing that this conceptual CPD design can achieve
propellant boil-off rates well under 0.05% per day, even when subjected to the LEO thermal
environment.
Nomenclature
<X Absorptivity
£ Emissivity
Beta Anglebetweenthe.OrbitalPlaneand the EclipticPlane .
BTU BritishThermal Unit
CPD CryogenicPropellantDepot
ft Feet
GMM GeometryMathModel
Hc ConvectionHeatTransferCoefficient
Hv HeatofVaporization
in inches
IR Infrared Radiation
ITM IntegratedThermalModel
KSC KennedySpaceCenter
Lbs pounds
LH2 LiquidHydrogen
L02 LiquidOxygen
MLI Multi LayerInsulation
I Thermal Ailaiyst, Environments& LaunchApprovalBranch,LaunchServices Program
2SeniorThettrIal Analyst,Environments& Launch Approval Branch,LaunchServices Program
3 Launch Vehicle Thermal Analysis Team Lead, Environments & Launch Approval Branch, Launch Services
Program,VA-l'B
4ManagerAdvancedPrograms
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AmericanInstituteofAeronauticsandAstronautics
mT MetricTon
nm Nautical Mile
PSI Pound perSquareInch
Q Average HeatLeakRate
RAAN Right AscensionofAscending Node
RadK RadiationConductor
Theta AnglebetweentheCPO MinorAxis and theEclipticPlane
I. Introduction
Along duration CPO is a conceptual vehicle that can store large quantities ofcryogenic propellant in space for
extended periods. A CPO would function as an on-orbit refueling station for in situ spacecraft and/or launch
vehicles. A CPO could potentially provide many benefits to
the NASA space exploration program including the
+--r---Atlas V 5-meter Payload
extension/elongation of spacecraft missions and reduction of
____ _
t~Dg
launch vehicle up mass requirements. As shown in Fig. 1, the
CPOconceptconsidered in this analysis was adual propellant
Elongated storage configuration, which is based on the current Atlas V
LHlTank Centaur design. In this CPO concept, the forward end ofan
Modified Atlas V Centauris mated to the aftend ofa modified Atlas V
Centaur Centaur. The modified Centaur would consist of an
(LHl elongated LH2 tank connected to a small boil-off storage
Module)
tank. Both the Centaur and modified Centaur would be
encapsulated within the Atlas V 5-meter payload fairing at
launch. Onceonorbit, residual LH2 within Centaurwould be
.......JJIiiIii~--1r-BOil-Off Storage transferred to the modified Centaur; the residual H2 would be
Tank(GH2)
purged with Helium. The Centaur would then be refilled, on
orbit, with L02. Consequently, the modified Centaur
functions as the on-orbit LH2 storage module, and the
Centaur functions as the on-orbit L02 storage module. The
_+-__Centaur dual propellant CPD concept has the advantage ofbeing able
Centaur
(LH2Tank) to store both LH2 and L02. Further, the concept utilizes
(L02
existing, or slightly modified, flight hardware. To achieve its
Module)
target boil-off rate, the CPO will need to store cryogenic
propellant for extended periods with minimal evaporative
losses. This will require, in part, a Thermal Protection
-1---Centaur System (TPS) specifically designed to minimize the
(L02Tank)
cryogenic propellantboil-offassociated witheach module.
Aside from the TPS design, the orbit in which the CPO flies
will also play an important role in determining the propellant
Figure 1.Cryogenic Propellant Depot Concept boil-offrate. From alogistics standpoint, LEO represents the
(Launch Configuration) most desirable orbit in which to fly the CPO. LEO would
provide easy access for both on-orbit refueling and CPO
maintenance/resupply. Conversely, LEO represents the least
desirable orbit from a thermal management standpoint.
Having significant amounts of solar, abedo, and earth IR,
LEO constitutes the most severe on-orbit thermal
environment and, undoubtedly, would be the most
challengingorbit from which to manage cryogenic propellant boil-off. Because it is both logistically desirable and
thermallychallenging,LEO wasconsidered tobethe ideal orbitto investigatein this analysis. Itwasfelt thataCPO
design that could achieve the desired boil-offrate in LEO could also achieve, orexceed, this boil-offrate in higher,
less thermally severe, earth orbits (i.e., geosynchronous, Lagrangian, etc.). In this sense, LEO provides the ultimate
testofCPDstoragecapabilities.
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II. Developmentofthe CPDITM
TheCPDITM was developed inThennalDesktopTM Version5.3. Thennal DesktopTM is a thermal analysis tool that
facilitates the developmentofsophisticatedcomputeraideddesignbasedthennal models. Thedevelopmentprocess
involves generation of a Geometric Math Model (GMM) using the Thennal DesktopTM built-in AutoCAD tool.
Thermal and optical properties are assigned to the GMM using the Thennal DesktopTM property edit forms. The
Thermal DesktopTM pre-processor is used to compile the above data into a Systems Improved Numerical
Differencing Analyzer (SINDA) thennal network. Radiation conductors are added to the SINDA thennal network
using Thennal Desktop'sTM radiation analyzer tool, RadCADTM. The combined RadCADTM - SINDA thennal
network can be solved using any of the various Thennal DesktopTM built-in numerical solvers. The CPD ITM
developed for this analysis consists ofthe CPD structures, cryogenic propellant, and on-orbitthennal environment.
DevelopmentoftheCPDITMisdescribedindetail in thefollowingsections.
A. OnOrbitThermalEnvironment
As previouslystated, the CPD was assumed to be in LEO for this analysis. The LEO thennal environmentconsists
of the following heating parameters: 1) solar radiation, 2) albedo, 3) earth IR and, 4) deep space IR. Table 1
summarizesthespecificvalues used fortheLEOheatingparameters.I Forthis analysis, theCPDwasassumedtobe
in a circular orbit (i.e., orbital eccentricity equals zero) in which the Right Ascension of the Ascending Node
(RAAN) was always equal to 90° (i.e., orbital procession rate equals zero). These simplifying assumptions ensure
thatthe solarangle(beta) will always beequal to theorbital inclinationand, further, thatthe radiantflux incidenton
the CPD will be constant as a function oforbital cycle. Table 1summarizes the specific values used for the CPO
orbitalparameters.
Table1.• SummaryofCDPOrbit'aIParameters
Altitude BetaAngle RAAN OrbitalPeriod Orbital Theta
Tracking
(nm) (degrees) (degrees) (hr) Eccentricity (degrees)5
365 0-30 90 1.6378 0 -10-+10 +ZSolar
TheattitudeoftheCPDisdefinedinTable 1andFig. 2. Thebasiccriteriaforthedesigningthe CPDattitudewasto
try to minimize the amount oforbital flux that can reach the CPD tanks. Preliminary analysis indicated that this
couldbestbeachievedbyorientatingthe CPDsuchthatits primaryaxis was perpendiculartothe eclipticplane and,
further, to selectbeta angles that would protect the aftend ofthe LH2 module from incoming albedo and earth IR.
Orientatingthe CPD such that its primary axis is perpendicularto the ecliptic planeensures that the aftendofeach
storage module (L02 and LH2) has a constant view ofdeep space. Further, this attitude minimizes the amount of
directsolarradiation thatcan reach the propellanttanks. However, earthIRand albedocan stillbe problematicfor
this attitude. As both areemitteddiffusely, significant amountsofearth IR and albedo can traverse the open endof
each sunshield and, subsequently, impinge upon the propellanttanks. Preliminary analysis suggested that the LH2
module was particularlysensitiveto earthIRand albedo during the illuminationphaseofthe orbit. To mitigatethis
impact, beta angles were chosen such that, during illumination, the open end ofthe LH2 module sunshield faced
awayfrom the incoming earth IR and albedo. While this orientation minimizes the amount ofearth IR and albedo
incident on the LH2 module during illumination, it has the opposite effect during eclipse. Fortunately, both the
albedoandearthIRflux areverylowduringtheeclipseportionoftheorbit.
Within Thermal Desktop, the CPO orbit was simulated by using 18 discrete steady-state orbital positions. To
simulate the on-orbit radiation exchange, 5000 rays per node were shot to calculate the CPD radiation conductors
(RadK's) and orbital heating rates (solar, earth IR and al~edo). A sensitivity study determined that further
increasingthe numberofrays and/ororbital positions had littleeffectonthe fidelity ofthe solution. Itis important
to note that all predicted heating rates in this analysis represent orbital averages (average from all eighteen orbital
positions).
5ThetaistheanglebetweentheCPOminoraxesandeclipticplane(seeFig. I).
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Table2. Summar) 0fLEOThermaIEnV.lronment
SolarRadiation Earth IR(BTUlhr/ft2/oF) DeepSpace
Albedo IR
(BTUlhr/fel°F)
Dlumination Ellipse (degR)
429.2 70.2 35.1 0.35 4
Earth
OrbitalPlane
CPOPrImae
Figure2. CPD Concept- DualPropellantStorage Configuration (BroadsideVerticalAttitude)
B. Structures
As shown in Fig. 3, the dual propellantCPOconcept utilizes two separate modules to store cryogenic propellanton
orbit: the LH2 module and L02 module. TheL02module is basedon an Atlas VCentaurdesign. TheLH2 module
is based on a modified Atlas V Centaur design. To thermally isolate the two cryogenic propellants, the LH2 and
L02 modules are separated by three pairs ofcomposite struts.? The CPO is launched from inside an Atlas V 5
meter fairing. To save weight, the LH2 module is launched without propellant. Once the CPO is on orbit, the
remaining LH2 in the Centaur LH2 tank is transferred to the empty LH2 tank in the modified Centaur. After the
LH2 transfer is complete, the remainingL02in the CentaurL02 tank is transferred to the CentaurLH2 tank. Thus,
the CentaurLH2 tank functions as an on-orbitL02 storage tank. Further, the Centaur L02 tankfunctions as an on
orbit G02 storage tank, collecting boil-offfrom the CentaurL02 storage tank6. The G02 could be used to collect
heat from the CPO structure and dissipate the heat back out to space. For this report it was assumed that the
avionics boxes and batteries on the Block II avionics shelf were powered off, and the RL-lO was not operating.
Duringon-orbit steady state operation, use ofavionics and batteries would be kept to a minimumin orderto reduce
theamountoflatentheatgeneration.
6Note that inFig. 3and Fig. 4, theCentaurtanks are referred to bytheiron-orbitstoragefunction.
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Once 6norbit, the CPD sun shields are deployed. As showninFig. 4, the LH2 moduleis fully enclosed byits sun
shield;however, theL02 module isonly partiallyenclosedbyitssun shield. Thisisdue to the fact that, inorderto
mitigate plume heatingfrom the CentaurReaction Control System(RCS), the maximumlengthofthe L02 module
sunshieldislimitedtojust24.55 feet,8 Becauseitisfully enclosedbyits sunshield,theLH2moduleis primarilyan
IRand albedbdominantradiationenvironment(i.e., virtuallyno directsolar). Theopticalpropertiesforthe exterior
surfaces ofthe LH2 and the GH2 tanks were chosen as to have low alpha and very low emissivity values. These'
optical properties enable the exteriorsurfaces ofthe LH2 and GH2 tanks to effectively reflect incident albedo and
earth IR back to space. Because it is not fully enclosed by its sunshield, the L02 module is subjected to much
higheramounts ofsolarflux. Consequently, the optical properties for the L02 and G02 tanks were selected as to
have low alpha values and high emissivityvalues (Le., low ale ratio). These optical properties allow the L02 and
G02 tank to effectively reflect incident solar radiation back to space, and effectively re-emit radiation in the IR
spectrum. Inorderto protect the CPD from the LEO radiation environment, all propellant tanks were enclosed in
ten layers ofMLI. Due to its very low emissivity, aluminized Kapton® was selected as the optical coating for all
MLI internal layers.5 This selection takes advantage of the fact that MLI internal layers participate only in IR
exchange(Le., absorptivityis notimportantforMLIinternallayers).
Table2. OpiicalPropertiesforL02ModuleSunShield
StrocturelEquipmentItem Material AbsorptivitylEmissivity Ref.
CentaurSunShield7
Inner: VDAKapton® Inner: 0.14/0.05
SunShieldLayer#I [4]
Outer: SilverTeflon® Outer: 0.07/0.80
Inner: VDAKapton® Inner: 0.14/0.05
SunShieldLayer#2 [4]
•..._._.... Outer: SV5 Outer: 0.08/0.81
Inner: VDAKapton® Inner: 0.14/0.05
SunShieldLayer#3 [4]
_.....•_... Outer: SV5 Outer: 0.08/0.81
PrimaryPro~!fJntTank
Inner: VDA Kapton® Inner: 0.14/0.05
SunShieldLayer#1 [4]
Outer: Silver Teflon® Outer: 0.08/0.70
•.......
Inner: VDA Kapton® Inner: 0.14/0.05
SunShieldLayer#2 [4]
Ou~: SV5 Outer: 0.08/0.81
Inner: VDA Kapton® Inner: 0.14/0.05
SunShieldLayer#3 [4]
Outer: SV5 Outer: 0.08/0.81
Forthe LH2 and GH2 tanks, aluminized KaPton®was also selected as the optical coatingforthe MLIouter surface.
This was acceptable due to the fact that very little solarradiation impinges on the LH2 and GH2 tanks and, thus, a
low absorptivity to emissivity (ale) ratio was not required for these surfaces. Conversely, for the L02 and G02
tanks, silverTeflon®was selected as the optical coatingforthe MLIoutersurface. This was due to the fact that the
L02 and G02 tanks are subjected to high amounts ofsolarradiation and, consequently, require an optical coating
havingalowWeratio.
7Sunshieldhalfconeangleandlengthare60°and 24.55 feet, respectively.
8Thislengthisbasedonthe assumptionthatthesunshieldhasaconeangleof60°.
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-------------------------------------------- -------------~---
Payload
Adaoter
CentaurAftEnd
(RL-10,HeliumBorrles,
Propel/amFeedlines,etc..)
CompositeStruts
(ThreepairsofCarboni
CarbonsTruTs)
Figure 3. CPD Concept- DualPropellantConfiguration (SunShieldRemoved to Reveal TankDetail)
ModifiedCentaur
(LH2Module)
Centaur
(L02Module)
Figure 4. CPD Concept- DualPropellantConfiguration (Sun Shield ShowntoScale)
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C. CryogenicPropellant
It was consetvatively assumed in this analysis that the entire internal area of both the Centaur L02 tank and
modified CentaurLH2 tank are in constant contact with their cryogenic propellant. Further, itis assumed that the
heattransfetcoefficientbetweenthetanks and cryogenic propellantsis auniformvalueof0.75BTUIhr-ft2-R, which
is representative of a low gravity orbital environment. Table 3 contains the physical properties for the
cryogenic propellants that were used in this analysis. Thecryogenicpropellantsinsidethetankswere modeled
as constanttemperatureboundary nodes. Itwasconservatively assumedthatthe cryogenic propellantswere already
atsaturationtemperature when the CPD was insertedintoLEO. Consequently, all heatabsorbedbythe propellants
whileon-orbitimmediatelycontributedtothe vaporization process (i.e., noheatrequired to first raise the propellant
tothesaturationtemperature).
.
Table 3 CPD Cryoj!em.c FIUI·d Saturat"IOn Propert·les
Pressure BoundaryTemperature HeatofVaporization
CryogenicFlUid Ref.
(PSI) (R) (BTUllbm)
L02 35 180 87 [3]
....
G02 35 180 NA [3]
......._...
LH2 35 42.67 182 [3]
...
GH2 35 42.67 NA [3]
III. Results
Figure 5 through Fig. 7 summarize the predicted average heat leak rates entering the LH2 and L02. The boil-off
rates(% perday) thatare presentedinFig. 7werecalculatedusingthefollowing formula:
Q*100
Boil Off (% perday) =----'-~-- (1)
H *24*M
v
where M is the mass ofthe cryogenic propellant,Hv is the heat ofvaporization, and Q is the average heatleak
rate. The massofthe LH2andtheL02 wereassumed to be5mT(11060lbs) and55 mT (121660lbs), respectively
[8].
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LH2 Boil-offSummary
140
-Beta0
-Beta 5
120 -Beta 10
-Beta 15
-Beta 20
-100 Beta 25
...
.~......
i=
80
al
..:.:
III
~... 60
III
QI
J:
40
20
0
o
-10 -5 5 10
ThetaAngle (degrees)
Figure 5.LH2HeatLeakversus ThetaAngle
L02 Boil-offSummarY
50
45 .-
-
40
....
35
...
-
3~'
30
~
al
:;25 - -Beta0
III
QI -Beta 5
..J
'lii 20 -Beta 10
QI
J: -Beta 15
15 -Beta 20
Beta 25
10
5
o
-10 -5 o 5 10
ThetaAngle (degrees)
--:..- ---.J
Figure 6.L02HeatLeakversusThetaAngle
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-------- ------- ._-- --------------------
Total Boil-offSummar
-10 -5 0 5 10
ThetaAngle (degrees)
---
Figure7.TotalBoil-off(% PerDay) versusThetaAngle
IV. Conclusion
Figure 7contains the summaryofthe total boil off(% perday) for an on-orbitCPD. From Fig. 7, itis clearthatthe
currentCPDdesign has achieved the goal ofa total boil offrate thatisless than 0.05% perday. Figure5and Fig. 6
are also critical because it can determine ifa certain CPD attitude (i.e. Beta, Theta) is desired based on a specific
mission need to storeone propellantlonger/shorterthan theother, orstorebothpropellantsatthe same boil offrate.
Fromthe results presented in this report, the optimal combination ofCPD Beta and Theta angles for reducing L02
boil-offis 00and 00, respectively. Because ofthe reduced length ofthe Centaur sunshield, the L02 tank has direct
solar radiation impingement. At Beta and Theta equal to 0, the direct solar radiation on the L02 tank is minimal.
The optimal combination ofBeta and Theta with regards to reducing LH2 boil-offis 300 and 50, respectively. In
Fig. 6, there wereseveral observed trends. ForaconstantTheta, theheatleakenteringthe Centaurtends to increase
as Beta increases. Conversely, for a constant Beta the heat leak entering the Centaur tends to decrease as Theta
increases. These trends underscore the fact that the heatleakentering the L02 tank was more sensitive to Beta than
Theta. AtBetaand Theta equal to zero, theLH2 module's sunshield is able to fully enclose the LH2 tank from any
direct solar radiation. Therefore, the thermal environment inside the LH2 module sunshield is dominated by IR.
However, because the CPD is composed ofboth the Centaur and the LH2 module, the overall lowest heat leak was
for aBetaandThetaof200and 00,respectively. There were several trends observedin Fig. 5. ForaconstantTheta
angle, the heat leakenteringthe LH2 module tends to decrease as Beta angle increases. Conversely, for a constant
Beta, the heatleakenteringthe LH2 module tends to increaseasThetadeviatesfrom zero.
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·
.
' ,
Acknowledgments
This wotk has been funded and supported by NASA, Launch Services Program, through an advanced special
studies conttact. The authors would like to thank Paul A. Schallhom9for his invaluable advice and funding for the
CPDspecialstudy. Theauthors alsowantto acknowledgeC&RTech™ fortheirversatilesoftwareandprofessional
assistance.
References
I. UnitedLaunchAlliance(2007).AtlasFamiliarization. None
2. Kutter, Zegler, O'Neil, and Pitchford, "A Practical, Affordable Cryogenic Propellant Depot Based on ULA's Fight
Experience",September2009
3. Retrieved0729,2010,fromThermophysicalPropertiesofAuidSystems:http://webbook.nist.gov/chemistry/fluid/
4. Gilmore G. David, Spacecraft Thermal Control Handbook: Volume I: Fundamental Technologies,EL Segundo,
CalIfornia:TheAerospaceCorporation,2002
5. Section 4 In-space Propellant Transfer and Storage Demonstration: Flagship Technology Demonstration, RFI
NNHI0ZTT003L,June10,2010
6. ILCDoverFinalReport,PhaseIIIR&D-CentaurSunShieldDevelopment,December18,2009.
7. United Launch Alliance, LLC. (10 June 2010). Section 4 In-Space Propellant Transfer andStorage Demonstration
FlagshipTechnologyDemonstrationRFI-NNHlOZTTOO3L Houston,TX: NASA/JohnsonSpaceCenter.
9NASA, LaunchServicesProgram,Environments&Launch Approval BranchManager
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