passive zbo storage of liquid hydrogen and liquid oxygen applied to space science mission concepts

Passive ZBO storage of liquid hydrogen and liquid oxygen appliedto space science mission conceptsD.W.Plachtaa,*,R.J.Christie b ,J.M.Jurns c ,P.KitteldaNASA Glenn Research Center,Propellant Systems Branch,21000Brookpark Road,Mail Stop 86-2,Cleveland,OH 44135,USAbZin Technologies,Cleveland,OH,USAcQSS,Cleveland,OH,USA dPalo Alto,CA,94303,USAReceived 17November 2005;accepted 18November 2005AbstractLiquid hydrogen and oxygen cryogenic propulsion and storage were recently considered for application to Titan Explorer and Comet Nuclear Sample Return space science mission investigations.These missions would require up to 11years of cryogenic storage.We mod-eled and designed cryogenic propellant storage concepts for these missions.By isolating the propellant tank’s view to deep space,we were able to achieve zero boil-offfor both liquid hydrogen and oxygen propellant storage without cryocoolers.Several shades were incorpo-rated to protect the tanks from the sun and spacecraft bus,and to protect the hydrogen tank from the warmer oxygen tank.This had a dramatic effect on the surface temperatures of the propellant tank insulation.These passive storage concepts for deep space missions substantially improved this application of cryogenic propulsion.It is projected that for missions requiring larger propellant tank sizes,the results would be even more dramatic.Published by Elsevier Ltd.Keywords:Hydrogen;Oxygen;Space cryogenics;Heat transfer1.IntroductionNASA Glenn Research Center (GRC),in partnership with NASA Jet Propulsion Laboratory (JPL)and NASA Ames Research Center (ARC),was tasked to define and analyze cryogenic hydrogen and oxygen storage systems for use in the propulsion design for several space science missions,defined by JPL and our customer,NASA Mar-shall Space Flight Center’s (MSFC)In-Space Propulsion Technologies (ISPT)Project Office (TD05),managed at MSFC.A detailed report of the complete study has been published by JPL [1].This paper,describes the propellantstorage thermal analysis and design for two of the missions considered:Titan Explorer (TEx)and comet nucleus sam-ple return (CNSR).It excludes the Mars Sample Return mission,which is described in a related paper [2].TEx is a mission with science instrumentation for an orbital exploration of Titan,a moon of Saturn.It is powered by radio-isotope thermal generators (RTGs).The CNSR mission rendezvous with Comet Wirtanen,stays a month,collects a 1kg sample and returns the sam-ple to Earth.It is powered by solar arrays.A summary of the top level requirements of these missions are shown in Table 1.For each mission considered,a thermal environment was defined and zero boil-off(ZBO)cryogenic storage designs and analyses were performed.Non-vented cryo-genic storage was required by the customer and was baselined.These analyses included configuration investi-gations,radiation modeling and the use and extensive0011-2275/$-see front matter Published by Elsevier Ltd.doi:10.1016/j.cryogenics.2005.11.012*Corresponding author.Tel.:+12169777126;fax:+12169777545.E-mail address:david.w.plachta@ (D.W.Plachta)./locate/cryogenicsCryogenics 46(2006)89–97augmentation of GRC’s cryogenic analysis tool (CAT).This,along with JPL’s design work on the related cryo-genic propulsion systems and other spacecraft systems,fed into a real-time mission design effort.This included engineers from each spacecraft discipline working together in a computer integrated mission design environment.2.Modeling/methodologyZero boil-offcryogenic propellant systems analysis for this study used several different analysis tools.Results were cross-checked with simple hand calculations to verify results.Thermal loads from environmental conditions were calculated using thermal analysis system (TAS),a finite ele-ment tool from Harvard Thermal,Inc.The cryogenic anal-ysis tool (CAT),an internally developed spreadsheet (used in previous studies),[3,4],determined the time dependent fluid conditions of a particular design.CAT was used to iterate on the designs,while ensuring that the specific ZBO design did not overpressure the propellant tanks during the mission.2.1.CAT methodologyGiven that a well mixed cryogen is initially stored at sat-urated conditions under its own vapor pressure in a storage tank that has structural supports,cryocoolers,a heater and MLI insulation,the pressure in the tank will change according to the following relationship [5]:d P d t hsp ¼/ðQ ÞV where P tank pressure t time/internal energy derivative Q net heat into system V tank volumehsphomogenous self pressurizationInitial tank pressure and tank volume are given,and energy derivative is calculated from the fluid properties as per the above reference.Given a closed cryogenic storage system without boil-off,net heat into the system is determined by the energy balance:Q net ¼Q MLI þQ struts þQ penetrations þQ mixerÀQ cryocooler þQ parasiticwhere Q net net heat into the systemQ cryocooler heat rejected from propellant by operating cryo-coolerQ MLI heat leak through insulation Q struts heat leak through supportsQ mixer heat introduced to system by mixerQ penetrations heat leak through piping penetrations intotankQ parasitic parasitic heat leak through redundant,non-oper-ating cryocooler Formulae for calculating heat loads are given in Appendix A .The CAT predicts the change in fluid temperature and pressure.When the net heat leak is greater than the heat removed,propellant temperature and pressure increases.This slightly decreases liquid mass,offset by an identical increase in vapor mass.When the net heat leak is less than the heat removed,propellant temperature and pressure decreases and liquid mass increases (from condensate).Fluid conditions are calculated on a daily basis based on initial fluid conditions from the previous day using tabu-lated fluid properties.From the thermal and fluid calculations,component designs such as the mixer,tank,tank supports,penetra-tions,and insulation are performed in CAT.These designs determine their mass,volume,areas,and power require-ments,which are then fed into the Team X mission environment.3.Cryogenic propellant systems 3.1.Model descriptionThe thermal model of a zero boil-off(ZBO)cryogenic propellant system is shown in Fig.1.MLI insulation was used for the LO 2and LH 2tanks.Shades shield the propel-lant tanks from solar or planetary albedo,planetary infra-red radiation,and from the spacecraft bus.Three shading configurations were modeled and iterated upon to mini-mize the heat going into the propellant tanks.Radiation modeling included iterations on inter-tank shield concepts to minimize the heat transfer between the liquid oxygen (LO 2)tank to the liquid hydrogen (LH 2)tanks.A conduc-tion model of the support structure (struts)was also incor-porated.Obstructions were included to represent structures that block the view to space,such as the struts,thruster,other tanks,and miscellaneous objects.The avionics sec-tion of the S/C Bus was fixed at 250K (based on JPL experience).Shades and shields were modeled as 20-layer MLI blan-kets,over an undefined structural frame and are in a cone shape to behave like a ‘‘V’’groove radiator.The blanketsTable 1A summary of the significant requirements of the missions studied TExCNSRHigh delta-V,large payload Moderate to high delta-V,small payload Relatively ‘cold’thermal environment Relatively ‘cold’thermal environment RTG powered Solar poweredFlagship classNew Frontiers class90 D.W.Plachta et al./Cryogenics 46(2006)89–97were modeled with an effective emissivity of 0.004,which is conservative in comparison to the James Webb Telescope shield,which has a shield effective emissivity of 0.00038.Surface properties for absorptivity and emissivity were researched and selected as shown in Fig.1.The values used are consistent with those used by JPL in the Team X envi-ronment,which is based on years of experience in the developing and flying spacecraft.In some cases,these values may be conservative.3.2.Tank insulationFor this analysis the baseline for insulation design assumed a foam thickness of 1.8cm and 45layers of MLI.The MLI was reduced significantly in the cases where the tanks were exposed to deep space.The MLI cover sheet is a flexible optical solar reflector (FOSR).The FOSR pro-vides a low solar absorptance along with a high infrared emittance and is used extensively on many spacecraft such as the Hubble space telescope.The FOSR mass includes a Nomex scrim which adheres to the Teflon to give it strength and arrest tearing.We based insulation modeling on the Lockheed [6]equation,modified for cryogenic test data per Martin and Hastings [7].We included foam insu-lation to provide insulation during pre-launch ground operations where the MLI is ineffective.We assumed that during this phase of operations,the MLI was not in a vacuum,but was gas purged.3.3.Loads and temperaturesThe thermal loads on the vehicle consist of the absorbed Sun’s radiation,planetary albedo and infrared radiation (IR).Fixed temperature boundary conditions include thespacecraft bus (S/C Bus)temperature,and deep space tem-perature.The MLI and radiator environment temperatures are determined by applying the above loads and allowing the model to reach thermal equilibrium.These surface tem-perature equilibrium temperatures are also known as the ‘‘environment temperatures.’’By averaging the tempera-ture of each surface element,we obtain an average environ-ment temperature for groups of surfaces such as the tank MLI or the radiator.We determine the average surface temperature by taking the temperature of each element raised to the fourth power and weighting it by the element’s area,then summing the products and dividing by the total area of the elements.The fourth root of the result is the average temperature:T avg ¼P A e ÃT 4e ÀÁP A e 0:253.4.StrutsThe support structure for the tanks is a major source of heat leak.After evaluating strut designs,we decided to incorporate passive orbital disconnect struts (PODS)[8].A schematic of a PODS is sketched in Fig.2.As shown by the arrows,the heat path in space is through a longer path with a much smaller cross-section area.During launch the load path is through a shorter,stronger path with a much larger cross-sectional area.By using the PODS,the heat leak in microgravity can easily be reduced to 0.1of that of the state-of-the-art (SOTA)struts [9].PODS were used on Gravity Probe-B and may be used on many other missions.Note that PODS greatly reduce the orbital natural frequency of the structure.Therefore before choosing PODS as a support method,structuralFig.1.Titan explorer thermal model.Surface properties used were based on JPL experience.D.W.Plachta et al./Cryogenics 46(2006)89–9791and flight control analyses need to be performed to verify that they can be safely used with the launch vehicle selected.3.5.Penetrations and other heat leak sourcesWe used the multipurpose hydrogen test bed test results (Ref.[7])as the baseline for other sources of heat leak.These included vent,fill and pressurizations lines.This heat leak amounted to 0.29W.As shown in Appendix A ,heat load formulae,we scaled penetration heat leaks from MHTB test heat leak data.Small amounts of heat were added to account for mixer heat and electrical cables (some required for propulsion purposes).Typical values are shown in Fig.5and the cor-relations are given in Appendix A ,heat load formulae.3.6.MarginBecause of the uncertainty in the analysis of cryogenic systems and in order to comply with JPL design standards,we incorporated conservative design margins per Table 2.Note that JPL margins were developed for different appli-cations—with our larger application and better understood heating rates,these margins may be too conservative.In addition to thermal margin,mass margin was also incorpo-rated.We added 10%margin to the tank insulation and 25%margin for all other components.4.Results4.1.Titan ExplorerThe overall configuration of the Titan Explorer (TEx)is shown in Fig.3.The TEx concept incorporated a Sun Shade mounted on the spacecraft bus,to shield the propel-lant tanks from the spacecraft bus and from solar radia-tion.Also,an inter-tank shield was effective at reducing the radiative heat transfer from the oxygen tank to the hydrogen tank.These proved to be very effective in reduc-ing the heat leak from incident solar radiation (insolation).The boundary temperatures for TEx are shown in Fig.4.The loads are the insolation on the sun facing surfaces,a 250K fixed temperature on the spacecraft busTable 2Margin summaryConservatismMargin added RemarkMLI Uncertainty with SOTA,added 30%to scale factor 50%Q MLI overestimated by 70%,not consistent with substantial database for tank heating POD’s Heat scaled from c -Al 2O 3;T-300CFPR would reduce heat by 20%50%Flight proven technologya 0.2used;references use 0.09Absorbed heat would be reduced by halfeEffective e =0.004;SOTA used 0.00038(JWT)Order of magnitude difference.Unknown effect T environmentMissions did not consider cryogenic storage in their trajectory designNot done for consistent comparison tonon-cryogenic propulsion systems.Unknown effectMargins used comply with JPL design standards.Note that conservative properties were consistently used in themodel.Fig.2.Passive orbital disconnect struts.The use of these struts reduced our tank strut heat by a factor of ten.92 D.W.Plachta et al./Cryogenics 46(2006)89–97(S/C Bus)and deep space.The inter-tank shield,also shown in the Fig.4,helped to significantly reduce the LH 2tank MLI surface temperature.To reduce the thermal resistance of the insulation,its MLI was removed.Fur-thermore,we determined that if the temperature difference between the LO 2and LH 2was reduced,that the heat leak from struts connecting the LO 2and LH 2tanks could be reduced.Consequently,the LO 2temperature was allowed to float downward and the LH 2temperature upward.By removing half of the MLI from the LO 2tank,a heat bal-ance between the heat leaks and passive cooling was obtained at a LO 2tank temperature of 68.6K.This reduced the temperature difference across the struts from 70K to 45K and cut the heat leak by36%.Fig.3.Titan explorer vehicle configuration.Shown are the final component sizes used in the vehicledesign.Fig.4.Titan explorer thermal model,with heat loads and boundary temperatures determined.D.W.Plachta et al./Cryogenics 46(2006)89–9793The heat into and out of the LO 2and LH 2tanks is shown in Fig.5.The hydrogen tank temperature is depen-dent on the emissivity properties at deep space tempera-tures,which are not known.In this application,it is important to keep the LH 2temperature below 23.6K in order to prevent tank venting (at 35psi).If the emissivity properties are close to 0.9at roughly 10K,this will be eas-ily achievable,with substantial margin.This eliminates the need for cryocoolers,their power source,and their heat rejection system.et Nucleus Sample Return vehicleThe CNSR vehicle,shown in Fig.6,must generate power at 5.3AU’rge photovoltaic arrays are required at this distance from the sun.The tanks required protec-tion from the IR heat generated by these large solar arrays.Also while the heat load on the MLI decreases with increasing sun distance,the heat leak from the struts and other sources does not.At 5.3AU the MLI heat leak has decreased by 78%(of the 1AU value),but because of the other constant heat leak sources the total heat leak only deceased 70%.Passive shading and insulation concepts were explored in detail.After several iterations,it was found that by plac-ing two well-spaced shields between the tank and the arrays,the tanks could be cooled passively.The first shield blocked both cryogenic tank’s view of the arrays;the second blocked the LH 2tank’s view of the first shield.The shields needed to be well separated to actsimilarlyFig.5.Heat balance of hydrogen and oxygen tanks.The net heat is negative,suggesting passive ZBO ispossible.SR vehicle configuration.Shown are the component sizes of the final design selected.94 D.W.Plachta et al./Cryogenics 46(2006)89–97to V-Groove shields.The thermal model of final configura-tion for CNSR is shown in Fig.7.4.3.Historical perspectiveThese results are directly supported by previous space mission experience.However,related historical activities back up these conclusions.In 1977,analysis and a sup-porting test [10]was conducted at NASA GRC (then LeRC)that generally supports these conclusions.In that effort,a five-cubic meter tank insulated with 34layers of MLI,supported by 12fiber-glass struts,and pro-tected from a simulated spacecraft bus by two double aluminized mylar sheeted shadow shields,was enclosed in a 22K cold wall and tested in a vacuum chamber.The tank was filled with LH 2.Data extrapolations for a 1200day coast to Saturn predict that a non-vented tank would have risen in pressure by only 50psi.It is very likely that advances in tank supports,using 20MLI layer shadow shields instead of two layer,and the use of today’s insulating materials,with their better surface prop-erties,would result in a design with little or no pressure rise.5.ConclusionsIt was found,when not in low planetary orbit,zero boil-offcan be achieved through passive cooling of cryo-genic propellants.This requires the use of shading to pro-tect the cryogenic tanks from the sun,the spacecraft,and its solar arrays.Also,the thermal design requires passive orbital disconnect struts and restrictions on spacecraft attitude.Each vehicle configuration and mission scenario pre-sented unique problems that required unique solutions.These conclusions are:1.To achieve passive cooling,the LH 2tank needs protec-tion from the LO 2tank.2.LO 2tanks can be passively cooled with a single shade,protecting it from the spacecraft bus.3.Passive orbital disconnect struts are necessary to reduce strut heating.4.Reduce or eliminate MLI on the tanks to radiate heat to deep space.However,LO 2tanks may need a few layers of MLI to prevent excessive cooling.The complete study [1]found additional conclusions which are duplicated here.1.LO 2/LH 2with passive ZBO is the winner for high delta-V,large propellant fraction missions.2.Conventional propulsion systems win at lower Delta-V,lower propellant fractions.3.Even with ZBO costs added in,the cryo system should be able to meet the new frontiers cost cap of $$750M.AcknowledgementsThis effort was performed for the NASA Glenn Research center on task G380for NASA Contract No.NAS3-00145and through University Affiliated Research Center (UARC)Subcontract P0228861.The UARC is managed by the University of California,Santa Cruz under NASA Ames Research Center ContractNAS2-03144.SR thermal model.This thermal model shows the temperatures determined for one-half of a symmetrical vehicle configuration.D.W.Plachta et al./Cryogenics 46(2006)89–9795Appendix A.Heat load formulaeReferences[1]Guernsey C,Baker R,Plachta D,Kittel P,Christie R,Jurns J.Cryogenic propulsion with zero-boil-offstorage applied to outer planetary exploration.JPL D-31783,April,2005.[2]Plachta D,Christie R,Jurns J,Kittel P.ZBO cryogenic propellantstorage applied to a Mars sample return mission concept.Presented atthe Cryogenic Engineering Conference,Keystone,Colorado,August 31,2005.[3]Plachta D,Kittel P.An updated zero boil-offcryogenic propellantstorage analysis applied to upper stages or depots in an LEO environment.NASA TM-2003-211691.Presented at the2002 AIAA Joint Propulsion Conference,Indianapolis,Indiana,July8, 2002.Parameter Basis CommentsQ cryocooler Qcryocooler ¼gÃT coldheadÃP inputðT hotÀT coldheadÞg=efficiency as fraction of CarnotP input=cryocooler input power(W)T hot=rejection temperatureT cold head=cryocooler cold head temperature For discussion of estimate of efficiency see Appendix A of Ref.[1]Q MLI Q MLI¼ð8:95Â10À8ÞÂ152:6ÂT hotþT liq2ÂðT hotÀT liqÞþð5:39Â10À10ÞÂ0:031ÂT4:67hot ÀT4:67liq!Â1:8#layersAs shown in Ref.[6]Q struts(LO2)Q struts¼C1ÃC2ÃM tan kþM propellant2800ÃT hotÀT cold250À68:6C1=design margin,C=0.44Heat leak based on assuming PODS with heat leak0.10*c-Alumina SOTA heat leak.SOTA heat leak determined from strut conductivity data[1](LH2)Q struts¼C1ÃC2ÃM tan kþM propellant560ÃT hotÀT cold68:6À23:6C1=design margin,C2=0.021Q mixer Qmixer ¼C1ÃV tan k1:4Ã3:5ÃC2Ã1C3C1=design marginC2=duty cycleC3=buoyancy factorð42for LH2;64for LO2ÞMixer power calculations are based on SMIRF ZBO tests,[11]and are adjusted here for tank size andfluid buoyancyQ penetrations Qpenetrations ¼C1Ãð0:0025ÃðT hotÀT coldÞÃffiffiffiffiffiffiffiffiffiffiffiV tan kpÞPenetration heat leak scaled from testdata in Ref.[7]C1=design marginQ parasitic Qparasitic ¼0:75ÃC1ÃC2ÃT2hotÀT2coldÀÁP13inputÃ15:5Ã10C1=design margin C2=eclipse factor Offduty cryocooler heat leak defined by Kittel based on info in‘‘Spacecraft Thermal Control Handbook’’for AIRS cooler and makingfirst order adjustment for effect of coolersize.Ref.[1]96 D.W.Plachta et al./Cryogenics46(2006)89–97[4]Kittel P,Plachta DW.Propellant preservation for Mars missions.Adv cryo engin,vol.45.New York:Kluwer;2000.443.[5]Lin C,Van Dresar N,Hasan M.Pressure control analysis ofcryogenic storage systems.J Propul Power2004;20(3).[6]Thermal performance of multilayer insulations.NASA CR-134477,April1974.[7]Martin J,Hastings rge-scale liquid hydrogen testing of avariable density multilayer insulation with a foam substrate.NASA TM-2001-211089.[8]Spradley IE,Parmley RT.Design and test of a modified passiveorbital disconnect strut.Adv cryo engin,vol.33.New York:Plenum;1988.935.[9]Ross RG.Estimation of thermal conduction loads for struc-tural supports of cryogenic spacecraft assemblies.Cryogenics2004;44:421.[10]DeWitt R.Boyle R.Thermal performance of an integrated thermalprotection system for long-term storage of cryogenic propellants in space.NASA TN D-8320,NASA LeRC,May1977.[11]Plachta D.Results of an advanced development zero boil-offcryogenic propellant storage test.NASA TM-2004-213390.D.W.Plachta et al./Cryogenics46(2006)89–9797。