2010-TBD

Overview of NASA's Thermal Control SystemDevelopment for Exploration Project

Ryan A. StephanNASA Johnson Space Center

ABSTRACT

NASA's Constellation Program includes the Orion, Altair,and Lunar Surface Systems project offices. The first twoelements, Orion and Altair, are manned space vehicleswhile the third element is broader and includes severalsubelements including Rovers and a Lunar Habitat. Theupcoming planned missions involving these systems andvehicles include several risks and design challenges.Due to the unique thermal environment, many of theserisks and challenges are associated with the vehicles'thermal control system. NASA's Exploration SystemsMission Directorate (ESMD) includes the ExplorationTechnology Development Program (ETDP). ETDPconsists of several technology development projects.The project chartered with mitigating the aforementionedrisks and design challenges is the Thermal ControlSystem Development for Exploration Project. The risksand design challenges are addressed through a rigoroustechnology development process that culminates with anintegrated thermal control system test. The resultinghardware typically has a Technology Readiness Level(TRL) of six. This paper summarizes the developmentefforts being performed by the technology developmentproject. The development efforts involve heat acquisitionand heat rejection hardware including radiators, heatexchangers, and evaporators. The project has also beendeveloping advanced phase change material heat sinksand performing assessments for thermal control systemfluids.

INTRODUCTION

In early 2004, President Bush announced a bold visionfor space exploration. One of the goals included in thisvision is a human return to the moon by 2020. Inresponse to this vision, NASA established theConstellation Program, which includes several projectoffices. NASA has also established a separate programoffice whose charter is to advance technologies to aTechnology Readiness Level (TRL) of six to supportfuture exploration missions. This technologydevelopment program is referred to as the ExplorationTechnology Development Program (ETDP). Theaforementioned Constellation Program serves as theprimary customer for ETDP.

ETDP currently consists of 24 separate projects rangingfrom software development to entry descent and landing.Also included in this program portfolio is the ThermalControl System Development for Exploration project.This project, herein referred to as the Advanced Thermalproject, is chartered with mitigating thermal risks anddesign challenges for various elements within theConstellation program.

The Advanced Thermal project is currently developingtechnologies for three difference Constellation elements.These elements include Orion, Altair, and Lunar SurfaceSystems (LSS). Orion is the manned capsule that will beused to transport crew to the International Space Station(ISS) and Lunar orbit. The second Advanced Thermalproject customer is Altair. Altair is the Lunar landerelement that will be used to transport crewmembers toand support them while living on the Lunar surface forshort mission durations. The third, and final, customer isthe Lunar Surface Systems project. The AdvancedThermal project is developing technologies for twoelements within the LSS project. The first element is theLunar Habitat which serves as the long-term habitat forastronauts while located on the Lunar surface. Thesecond element is the Lunar Electric Rover (LER), whichis a pressurized rover that will be used to house andtransport crewmembers_

The current paper will describe the process forgenerating the Advanced Thermal project's technicalcontent and the technology development process used toadvance the technology readiness level to six. Thecurrent document will also briefly introduce the varioustechnology developments currently underway. Themajority of these efforts are also described in otherpapers"'"'" at the current conference. These papersshould be referenced for a more exhaustive descriptionof a particular development task.

PROJECT CONTENT GENERATION

In addition to mitigating key Constellation Program Office(CxPO) thermal risks, the Advanced Thermal projectalso performs technology development based ondiscussions with members of the space thermalcommunity. These development opportunities focus onfurther investigating novel concepts such as Sublimator

https://ntrs.nasa.gov/search.jsp?R=20100021079 2018-07-13T22:46:02+00:00Z

Driven Coldplate l that result from literature surveys andconversations amongst project members and the NASAcommunity. The technology development project alsoseeks to identify unique applications using traditionalhardware (i.e. transient sublimator) within the CxPOprojects. The final source for project content is theidentification of performance improvement opportunities(i.e. advanced phase change material heat sinks,variable heat rejection radiators, etc...). Theseperformance advancements typically involve reducinghardware mass, but also include improvements inhardware maturity, reliability, and thermal performance.

To ensure that the technology development program ismeeting the customer's technology needs, CxPO hasestablished a Technology Prioritization Process (TPP).This "thermal discipline specific" part of this process isdepicted graphically in Figure 1.

• Technology Needs

Technology Needs TechnologyNeeds

AdvancedThermalProject

Figure 1. The Constellation Program's TechnologyPrioritization Process used to deliver technologydevelopment needs to the technology developmentproject.

In the TPP, the Constellation thermal discipline leads foreach of the three projects (Orion, LSS, and Altair) submittheir technology development needs to a CxPO groupthat is responsible for prioritizing these needs. Once thelist of technologies is prioritized amongst all of thevehicle disciplines, it is submitted to the technologydevelopment program, ETDP. ETDP then provides thatlist of needs to the relevant projects (i.e. thermal needsare submitted to the thermal project). The AdvancedThermal project is currently addressing a total of eleventechnology needs. Of these needs, nine were classifiedas having a ranking of "critical", which is the highestclassification. Another important step to ensuring thatthe CxPO technology needs are being satisfied iscontinual communication between the various CxPOthermal leads and the Advanced Thermal projectmanager. This communication path allows thetechnology development project to be flexible in meetingthe customer's needs. It also helps to avoid anyconfusion or loss of information that could occur with theaforementioned technology prioritization process.

TECHNOLOGY DEVELOPMENT PROCESS

The project's technology development process typicallybegins with the hardware component possessing atechnology readiness level of two or three. The first stepin the process is the completion of coupon-level bench

top tests. The objective of these tests is to betterunderstand the basic physics and the criticaldevelopment challenges associated with the technology.This phase is often followed by a design and analysiscycle focusing on addressing the previously defineddevelopment challenges. The design and analysis cycleculminates in detailed drawings for an EngineeringDevelopment Unit (EDU), which is a scaled-down modeladdressing the key technology issues. The EDU goesthrough a rigorous test program and new, correlated,thermal models are developed based on the previoustest results. After completing the initial EDU tests andthe subsequent thermal models, another design andanalysis cycle is performed. During this second designand analysis cycle, the EDU performance data isassessed, performance improvements arerecommended, and detailed requirements are defined.The result of the second design and analysis cycle is thegeneration of drawings and the fabrication of prototypehardware. The prototype hardware is then tested as astand-alone entity to verify its performance.

The technology development process finally culminateswith an integrated thermal test. The integrated test isperformed in an environment relevant to the supportingvehicle (i.e. Altair) and simultaneously includes all of thepreviously developed hardware prototypes. The primarydifference between the integrated test and the previousprototype tests is that the integrated test involves all ofthe developed hardware as an integrated thermal controlsystem. At the conclusion of the integrated test, thedeveloped technologies possess a TRL six which isdefined as system/subsystem model or prototypedemonstration in a relevant environment (ground orspace).

The project is required to complete the integrated test inadvance of the customer's Preliminary Design Review(PDR). Generally, the Advanced Thermal project strivesto complete the integrated test approximately one yearprior to the customer's PDR. After completing this test,the prototype hardware and the associated data packageare delivered to the customer. The Advanced thermalproject currently has four CxPO customers as shown inTable 1. This table also includes the planned dates forthe customer's PDR.

Table 1. Advanced Thermal Project's Customers andthe Corresponding PDR Dates.

CxPO Customer Planned PDR Date

Altair July 2013

Orion LunarBlock Upgrade

July 2013

Lunar Electric Rover April 2015

Lunar Habitat November 2014

PROJECT CONTENT AND TECHNOLOGYDEVELOPMENT TASKS

The project includes several technical tasks spanning thethree functions of an effective thermal control system.The three critical functions are heat acquisition, heattransport, and heat rejection. The following section willbe divided into five subelements in accordance with theproject's Work Breakdown Structure (WBS) as shown inFigure 2.

Project

TC5 Fluids Heat Acqu Kvaporative Radiators PCM HeatHeat Sinks Exchangers

Figure 2. Partial Work Breakdown Structure (WBS)for Advanced Thermal Project.

THERMAL CONTROL SYSTEM FLUIDS

The overwhelming majority of United States' mannedspace vehicles have used a pumped fluid loop as theprimary means of thermal control during the mission.These fluids have ranged from an ethylene glycol/watermixture on Apollo to the Orbiter with water (internal loop)and Freon® 21 (external loop). The current baselinethermal control system design for Orion, Altair, and theLunar Habitat also depend on a mechanically pumpedfluid loop. Currently, a mixture of propylene glycol andwater is the baseline internal working fluid for theaforementioned CxPO projects. This fluid formulationhas never been used for a manned vehicle so there islittle long duration data for this fluid in a flight-like thermalcontrol system.

As discussed in last year's overview paper, theAdvanced Thermal project planned and executed asuccessful fluids life test. This test was originallyplanned to demonstrate fluid compatibility over a periodof ten years. Resultantly, the test was designed to runcontinuously for the ten year demonstration. Theworking fluid for this test was a 50/50 (by mass) mixtureof DowFrostTm HD and water. The fluid loop wasdesigned to include all of the materials inherent inOrion's thermal control system design. In order to savemass, Orion had baselined the use of aluminum tubing,aluminum heat exchangers, and aluminum coldplates.Therefore, the fluids life test stakeholders wereespecially interested in replicating the expected ratio ofwetted aluminum surface area to fluid volume. The lifetest design included two Surface Area Modules (SAMs).The SAMs were fabricated using a four inch squarestainless steel housing and included several layers ofstacked aluminum fin stock. The stainless steel housingwas isolated from the aluminum fin stock by Teflonsheets in order to eliminate any potential for galvaniccoupling. As reported last year, the life test wasterminated after only two months because the systemfilters became clogged and rendered the systeminoperable. Furthermore, the coolant pH increased from10.1 at the beginning of the test to 12.2 at the testtermination. Subsequent analyses have shown that the

selected coolant was not compatible with the large ratioof wetted aluminum surface area to fluid volume.

Based on the results of this test, DowFroSt Tm HD wasremoved as the working fluid for the Orion thermalcontrol system design. Orion personnel are currently inthe process of evaluating alternative formulations ofpropylene glycol and water for use as the coolant for theinternal pumped fluid loop. In parallel with the updatedfluid evaluation, the Advanced Thermal project hasdesigned and started fabrication of an updated fluids lifetest stand.

The updated test stand design leverages the previouslife test cart, but includes several improvements basedon lessons-learned. The updated test stand designincorporating several improvements is shown in Figure 3.

Figure 3. Updated fluid schematic for the thermalcontrol system fluid life test.

There were four major design updates incorporated intothe new test stand design. The first update involvedredesigning the aluminum surface area modules toaccurately represent the current Orion wetted surfacearea to fluid volume ratio. An additional surface areamodule was also designed. The second surface areamodule included nickel rather than aluminum. Thesecond design modification incorporated gas vents atvarious locations around the test loop. These vents willbe used to sample any gas that forms within the loop.The final design improvement is the inclusion of sightglasses to aid in the detection of gas or precipitategeneration3.

HEAT ACQUISITIONAn effective thermal control system must accomplishthree basic functions (heat acquisition, heat transport,and heat rejection). One of these functions is theacquisition of excess thermal energy from the cabin airand other heat-generating devices. There are threesubelements within this WBS element and each of themseeks to advance the state of the art by reducinghardware mass and volume.

SUBLIMATOR DRIVEN COLDPLATE — As reported lastyear, project personnel have invented and fabricated an

advanced technology concept that combines all threethermal control system functions into a single piece ofhardware. This novel hardware component is called theSublimator Driven Coldplate (SDC) and has the potentialof replacing the entire thermal control system with onehardware component. Another unique feature of thisconcept is that it does not use any power and includesno moving parts. The target application for the SDC isthe Altair Ascent Module due to the low heat load, shorttransport distance, and short mission durationrequirements for this vehicle. This technology wouldeliminate the need for a pumped fluid loop as all of theheat loads would be conductively coupled to asublimator.The first generation Engineering Development Unit(EDU) is shown in Figure 4.

4

Figure 4. Sublimator driven coldplate engineeringdevelopment unit.

The EDU that was delivered to JSC was approximately6" x 6" x 2". In the preceding photograph, the SDC wasdesigned so that the heat-generating components wouldbe directly mounted to the two 6" x 6" surfaces. For thepurposes of the vacuum test program, a simpleresistance heater was used to simulate the avionics heatload. The port extending to the top left of the photographis the feedwater inlet tube and the SDC incorporatesthree '/" diameter stainless steel porous cylindrical tubesthat serve as the sublimation plates.The SDC design requirement was to maintain a surfacetemperature of less than 40°C with an evenly appliedsurface flux of 2 Win 2 . The SDC was exposed to arigorous test program designed to verify that the unit metthe design requirements and to understand its operationoutside of the design envelope. The maximum recordedsurface temperature when the SDC was subjected to auniform heat flux of 2 Win was only 9°C far exceedingthe design requirements. The steady state feedwaterutilization was near unity for the duration of the testpoints.

COMPOSITE HEAT EXCHANGERS — Heat exchangersare traditionally used to transfer energy from one fluidloop to a second fluid loop. The use of composites forthese types of hardware is an attractive alternative totraditional metallic heat exchangers. Carbon basedcomposites have very high thermal conductivities makingthem more effective heat transfer devices. In addition,composites also have a high strength to mass ratio,which has the potential of reducing the hardware mass.The project has previously explored the benefit of usingthese advanced materials to fabricate radiators.However, heat exchanger construction is significantlydifferent and requires a more detailed investigation.

Project personnel have successfully completed thedesign and fabrication of a composite air/liquid heatexchanger 4 . The performance requirements for this heatexchanger were based on the performance of anexisting, mass-optimized, metallic air/liquid heatexchanger. The metallic heat exchanger was designedto transfer approximately 3.4 kW for a prescribed set ofinlet conditions (inlet temperature and fluid flowrates).The composite air/liquid heat exchanger EDU is shownin Figure 5.

Figure 5. Composite air/liquid heat exchangerengineering development unit.

Note in the preceding photograph that the liquid inlet/exitmanifolds are not representative of flight hardware.Rather these were intended to be functional in naturebecause the project does not envision the developmentof these to be overly challenging. The composite heatexchanger was 37% lighter than the metallic baseline.Unfortunately, the thermal performance of the compositeheat exchanger was 13% less than the metallic baseline.To achieve a more representative comparison betweenthe two heat exchangers, a "specific heat transfer" wasquantified by simply dividing the heat transfer rate by theheat exchanger mass. The metallic heat exchanger hada specific heat transfer of 89 W/kg while the compositeheat exchanger's specific heat transfer was 121 W/kgfurther providing evidence that this is a promisingtechnology development pursuit.The Advanced Thermal project is currently in the processof developing a prototype composite heat exchangerdesigned to meet the original heat exchangerrequirements. The prototype heat exchanger design will

be improved based on lessons-learned from the previousdevelopment effort. The next generation heat exchangerdesign is also focused on improving the design to allowfor simpler heat exchanger fabrication.

MICROCHANNEL HEAT EXCHANGER — Thetechnology development project has continued apreviously started collaboration with Pacific NorthwestNational Laboratories (PNNL) to develop microchannelliquid/liquid heat exchangers.

The first step in the assessment of this technology wasto establish a baseline for future comparison. To thatend, a test cart was designed and used to assess theperformance of a mass-optimized flight heat exchangerdesign for use on NASA's Crew Return Vehicle (X-38).This heat exchanger was designed to transfer energyfrom a warm de-ionized water loop to a cooler loopcontaining a mixture of ethylene glycol and water. The2.7 kg heat exchanger was designed to transferapproximately 3.1 kW between the two loops. The X-38heat exchanger performance specifications weresupplied to PNNL as the requirements for amicrochannel heat exchanger. PNNL designed,fabricated, and delivered a microchannel heat exchangerintended to meet those same performancespecifications. Both the microchannel and baseline heatexchangers are shown in Figure 6.

f

- .J

Figure 6. Baseline X-38 liquid/liquid heat exchanger(left) and PNNL's microchannel heat exchanger(right).

It is apparent from Figure 6 that the microchannel heatexchanger is significantly smaller than the baseline unit.In fact, the core volume of the microchannel heatexchanger is only 303 cm 3 versus 793 cm 3 for the X-38baseline. In addition, the mass savings associated withthis unit is approximately 0.65 kg or 24 %5.

The microchannel heat exchanger test data has beenshared with PNNL. The project was especiallyconcerned about PNNL's failure to meet the pressuredrop requirements on the cold loop. To that end, PNNLhas used the test data to develop correlated heatexchanger models. These models were then used to

conceptually design a next generation microchannel heatexchanger. The conceptual design shows improvedmass and volume as compared to the first generationmicrochannel heat exchanger. This improvement wasachieved while sacrificing the thermal performance of theunit. This sacrifice was acceptable because the first unitexceeded the thermal performance specifications. Themass and core volume for the conceptual design is 1.2kg and 188 cm 3 , respectively.The project is in the process of running a life test for themicrochannel heat exchanger. This test is scheduled torun the baseline test point continuously for at least sixmonths. The ensuing life test will provide insight into theperformance of a microchannel heat exchanger over along period of time similar to that expected for futurespacecraft thermal control systems. The project isconcerned that the microchannel heat exchanger may besusceptible to performance degradation due to theextremely small flow passages.

HEAT REJECTION

The third and final, critical function for an effectivethermal control system is heat rejection. As the nameimplies, heat rejection is the process of rejecting thevehicle's waste heat to the local environment. Thisfunction is typically accomplished using radiators, butevaporators or Phase Change Material heat sinks canalso be used to reject energy.

TRANSIENT SUBLIMATOR — The Lunar orbitalenvironment presents very unique challenges for thethermal control system. Figure 7 shows the spatialvariation of the Lunar surface temperature.

4001; '501; <100K

Figure 7. Spatial distribution of the Lunar surfacetemperature.The hottest portion of the Lunar surface corresponds tothe point directly aligned with the sun (subsolar point). Inthe preceding figure, the maximum surface temperatureis approximately 400 Kelvin while the minimumtemperature is less than 100 Kelvin on the dark side.

The extreme surface variation results in a large swing inradiator sink temperatures while the vehicle is operatingin Low Lunar Orbit (LLO). The large sink temperaturevariations are problematic because it is impractical(sometimes even impossible) to use a radiator as thesole means of heat rejection during LLO if the vehicle'sradiators have a large IR incident load from the Lunarsurface. The sink temperature corresponding to thelocation immediately above the subsoIar point exceedsthe setpoint temperature of the thermal control system.As a result, the vehicle's thermal control system mustuse a Supplemental Heat Rejection Device (SHReD)while the vehicle is in LLO. Figure 8 shows an exampleof the variability of a vehicle's heat rejection capabilityusing only radiators for a beta angle of zero degrees andan orbital altitude of 100 km. For Altair, these orbitalparameters result in the worst case hot LLOenvironment_

9000 — Vehide Req-e—t

8000

7000

60003

s000t5 y 4000

x 3000

2000

1000

0 1800 3600 5400 7200

Time, Semnds

Figure 8. Vehicle heat rejection requirement (blue),radiator capability (red), and supplemental heatrejection requirement (green) as a function ofmission time in Low Lunar Orbit.

The blue curve in Figure 6 represents the heat rejectionrequirement for the thermal control system. In thisexample, the vehicle heat rejection requirement isconstant throughout the orbit. The red curve representsthe vehicle radiator capability assuming a constantaverage radiator temperature where the radiatorcapability is defined as:

4U z: EoA(T11 -T_ I Equation 1

WhereQ = Radiator heat rejection (Watts)

£ = Infrared emissivity (Dimensionless)

6 = Stefan-Boltzmann constant (W/m2-K4)A = Radiator surface area (m)TR = Average radiator temperature (Kelvin)

T. = Radiator sink temperature (Kelvin)This figure clearly shows that the radiator capabilityvaries throughout the Lunar orbit. This variability can beexplained by studying Figure 7. It is apparent that the IRbackload incident upon the vehicle will change

throughout the orbit due to the changing Lunar surfacetemperature. In addition, the incident solar load will alsovary throughout the orbit. The combination of theseeffects leads to a wide sink temperature variation in LLO.As depicted in Equation 1, the heat rejection capability ofthe proposed Altair radiator will vary as the sinktemperature changes.

The thermal control system must be capable of rejectingthe full vehicle heat load (approximately 4.8 kW in thisexample) throughout the entire orbit. Therefore, thesupplemental heat rejection device must dissipate thedifference between the heat rejection requirement andthe radiator capability, which is shown by the green curvein Figure 8.

The selection of the proper SHReD depends on theduration of the LLO mission phase. For short missiondurations, an evaporative heat sink would be used as thesupplemental heat rejection device. A phase changematerial heat exchanger would likely be selected forlonger mission durations. The baseline Altair thermalcontrol system requires a sublimator for both Lunarascent and descent. In addition, the same sublimatorhas been chosen as the SHReD during LLO. Thethermal engineers designing the Orion thermal controlsystem also identified a Low Earth Orbit (LEO)application for using a sublimator as a SHReD.

A spacecraft's thermal control system has never beendesigned to use a sublimator in this cyclical fashion. Thetechnology development project completed a trade studyand uncovered two potential problems with using asublimator as a supplemental heat rejection device.Current sublimators have a minimum heat loadrequirement which would result in a poor orbit-averagedfeedwater efficiency (or utilization). In addition, there isconcern that the hardware may burst during periodswhen the sink temperature is relatively cold and the heatload on the sublimator is quite low or possibly non-existent.

Throughout the previous year, the technologydevelopment project has executed a plan to mitigate therisk associated with rupturing the hardware whilequantifying the expected Orbital Average FeedwaterUtilization (OAFU). A sublimator coupon was fabricatedand several operational modes were executed to quantifyOAFU while assessing the hardware's susceptibility torupturing during periods of low (or no) heat loads causedby ice expansion. The sublimator coupon is shown inFigure 9.

Figure 9. Photograph of the transient sublimatorcoupon used during the GFY 2009 test program.

The test program was run over a period of several weeksand resulted in several useful data points. Thesubsequent data analysis appeared to show a trendbetween feedwater valve timing and OAFU. For a givensimulated orbit, OAFU appeared to be higher when thefeedwater was stopped prior to the end of the heat loadas compared to operational scenarios where thefeedwater supply valve was left open for the entire two-hour orbit period 6 . This likely occurs because feedwatercontinues to sublimate even during the period of zeroapplied heat load because the water is exposed to thevacuum environment. The test coupon did not appear tobe susceptible to failure caused by the expanding icelocated in the reservoir. Resultantly, the project plans torepeat the tests in GFY 2010 using a flight sublimatordesigned for X-38. Rather than using a heater to applythe heat load (as was done during the coupon tests), avarying temperature coolant loop will be used.

RADIATORS — Radiator advancement is perhaps themost critical technology development for the upcomingLunar missions (and most other future spacecrafts, forthat matter). NASA has no history of using radiators toreject the excess heat from a habitable vehicle on theLunar surface. The Lunar Excursion Module (LEM)relied on an evaporator, specifically a sublimator, duringits relatively short surface stay and the Lunar rovers onlyused a radiator to refreeze the PCM heat exchangers.However, because of the longer surface stayrequirement of seven days (LEM was only three days)

the use of an evaporator is not mass efficient for theAltair and Lunar habitat heat rejection requirements. Theproject's radiator development is further divided into twosub-elements. The first sub-element is the developmentof a variable heat rejection radiator and the second is anassessment of radiator performance with dustaccumulation.

One of the most significant design challengesencountered when developing a radiator is liquid freezingwithin the coolant lines attached to the radiator surface.Typically radiators remove energy from the coolant linesflowing through the radiator and reject that energy tospace. Radiator surface area is one of the key factorscontributing to the rate at which energy is rejected tospace. Generally speaking, radiators are sized for themaximum heat load in the warmest continuous thermalenvironment. In order to dissipate a high heat load in arelatively warm environment, it is necessary to design aradiator system with a large surface area. However,when that same large radiator is required to dissipate amuch lower heat load in a cold environment, the surfacetemperature dramatically decreases. This decreasedsurface temperature can lead to fluid freeze within theradiator coolant lines if the radiator is not correctlydesigned. The resultant frozen fluid can be problematicand both Altair and Lunar Habitat face this exact designchallenge. The radiator must be designed to dissipate ahigh heat load during Lunar surface operations, but mustalso be capable of operations at very low heat loadsduring the translunar coast. Translunar coast isextremely cold because the Altair radiators (and possiblythe Lunar Habitat radiators) will be shadowed from thesun during the entire mission phase. The requirement tooperate at both a high load and low load is referred to asthe system turndown ratio (Qmax/Qmin). Both Altair andthe Lunar Habitat have approximately the sameturndown ratio requirement. The requirement for thesevehicles is an order of magnitude greater than the Apollocondition. The previously described design requirementsare defined in Figure 10.

Ascent Asrent and

0 50 100 150 200 250 300 350 900

M ission Duration, Hours—Sink Temperature

___Q Rejecdm

Figure 10. Altair heat rejection requirements andradiator sink temperatures plotted as a function ofMission Elapsed Time (MET).

The preceding figure shows both the radiator sinktemperature and the vehicle's heat rejection requirementas a function of Mission Elapsed Time (MET). Theradiator sink temperature, which is defined as thetemperature that would be achieved for an adiabaticbody with similar optical properties, is shown along theleft, vertical axis. Due to the abscissa resolution, it isdifficult to discern the transient nature. However, theradiator sink temperature repeats during each orbitperiod. The orbit period during LEO and LLO isapproximately 1'/2 and two hours, respectively. The sinktemperature varies from approximately 65 Kelvin duringTLC and is as high as 290 Kelvin during LLO. Asmentioned above, the radiator system is sized for Lunarsurface operations, which corresponds to a sinktemperature of 210 Kelvin.

The vehicle's heat rejection requirement is shown alongthe right, vertical axis is as low as 1000 Watts duringTLC. However, there are times during the mission wherethe heat load exceeds 7000 Watts. For the design point,the heat rejection requirement is approximately 6 kW.

The Advanced Thermal project is seeking to usetechnology development to overcome this extremelydifficult design challenge. The project is pursuing thedevelopment of three separate variable heat rejectionradiator technologies. These technologies includevariable emissivity electrochromicss, a digital radiator9,and a freezable radiator 10 design. In the preceding year,the project has design and tested thermal vacuum testsamples or bench-top apparatuses to evaluate thesetechnologies. Presently, the project is designing threefull-scale radiators to meet the requirements shown inFigure 10. These designs will be compared andengineering development units will be developed tofurther evaluate the three promising technologies.

In addition to the previously described challenge, theLunar surface environment presents yet anotherchallenge to the radiator design process. From athermal perspective, the negative impact of excessiveregolith build-up on the radiator is twofold. First, thepresence of regolith adds an extra thermal resistancebetween the radiator surface and the heat rejectionenvironment. The thermal resistance of the Lunarregolith is extremely high due to the voids between theregolith particles. In other words, the regolith is a verygood thermal insulator. The second negative effectcaused by the deposition of regolith on a radiator surfaceis an increase in the solar absorptivity (or possiblereduction in IR emissivity). Radiator surfaces aretypically covered with selective thermal coatings tominimize the amount of solar energy absorbed by thesurface. The Lunar regolith has a very high solarabsorptivity, which results in the radiator surfaceabsorbing excess thermal energy.

The project has completed thermal vacuum tests and thedetailed results are presented in another paper". Thetests have shown that monolayer dust accumulationdoes not significantly impact the radiator's infraredemissivity as shown in Figure 11.

Relative Emisivity of AgFEP Dusted with Simulants

350

3M

25D

d2M

E 1sD

10D

5D

1.5

For most manned vehicles, a high turndown ratiorequirement usually results in a two-loop thermal controlsystem architecture. A two-loop architecture isadvantageous because the external loop can include afluid with a very low freezing temperature. Unfortunately,fluids with low freezing points (Freon®, ammonia, etc...)are typically toxic and cannot be located inside thepressurized volume due to crew safety concerns. Thebiggest drawback to a dual-loop system is the increasedmass associated with both loops. The addition of asecond loop requires several additional hardwarecomponents and a slight increase in the required radiatorarea due to inefficiencies associated with an interchangeheat exchanger. Orion originally baselined a single loopsystem, but quickly switched to a two-loop system. Theaddition of the second loop increased the thermal controlsystem mass by approximately 18%'.

1.3

1.1 ♦ •. • ♦ O E:

^ D.9

0.7

0.5

0.0 0.2 0.4 0.6 0.8 1.0

Fractional Dust Coverage

n Pris 0NU-LHT-1D ♦ JSC-NULHT ♦ JSC-1AF

Figure 11. Relative infrared emissivity of silverizedteflon tape as a function of fractional dust coverage.

Figure 11 shows the relative emissivity, which is a ratioof the dusted sample emissivity to the pre-dusted sampleemissivity as a function of fractional dust coverage. Asshown in this figure, the ratio remains near unity for

fractional dust coverage up to 1, which representscomplete coverage. In addition, these tests also showedthat accumulation does not adversely increase thethermal resistance.

However, the same tests have shown that monolayerdust accumulation can have a significant impact on theradiator's solar absorptivity. The relative absorptivity asa function of fractional dust coverage is shown in Figure12.

Relative Absorptivity of AgFEP Dusted with Simulants

3.4

2.9

°

>o ° o°O

0.90.0 0.2 0.4 0.6 0.8 1.0

Fractional Dust Coverage

n Pris ONUAHT-1D OJSC-NULHT ♦ JSC-1AF

Figure 12. Relative solar absorptivity of silverizedteflon tape as a function of fractional dust coverage.

Unlike Figure 11, Figure 12 shows a strong correlationbetween the fractional dust coverage and the radiator'soptical property. The relative solar absorptivity tends toincrease dramatically as the fractional dust coverage isincreased. In addition, there appears to be a strongrelationship between the absorptivity of the simulant andthe impact of the accumulation. For example, the effectof accumulated JSC-1AF is much more severe than theimpact of NU-LHT-1 D. This is likely due to the fact thatthe solar absorptivity of the latter is lower than JSC-1AF.These results show that the solar absorptivity of aradiator surface can increase by as much as 50% whenless than 20% of the radiator is covered with regolith.This increase would result in 1.5 times the amount ofsolar energy being absorbed by the radiator, which wouldseverely impact the radiator's thermal performance.

The Advanced Thermal project is collaborating withanother ETDP project to evaluate possible dustmitigation techniques. These mitigation techniquesinclude workfunction matching, surface texturing, surfacebrushing, and electrodynamic shield technologies.

PHASE CHANGE MATERIAL HEAT EXCHANGERS — Atypical PCM heat exchanger is used to store excessthermal energy during periods of high heat loads (or hotthermal environments) by melting a material andrejecting the stored energy at a later time. During therejection period, the material is frozen again preparing itfor the next heat load period.

The Advanced Thermal project has identified two PCMapplications within the planned Lunar missions. The firstknown application is the Lunar Electric Rover (LER).

The baseline LER includes the use of a PCM heatexchanger conductively attached to a radiator. Thesecond known PCM application is Orion. Orion'sconcept of operations includes a six-month LLO loiter.As mentioned in the preceding section discussing thetransient sublimator, the LLO thermal environment isquite unique and results in the need of a supplementalheat rejection device. A PCM heat exchanger is wellsuited for this application due to the relatively long andcyclic LLO mission phase (currently scheduled for sixmonths). In addition to these possibilities, a PCM heatexchanger development effort also provides riskmitigation for the transient sublimator development.

The amount of PCM mass required to provide thermalcontrol for a PCM application is inversely related to thematerial's heat of fusion. The higher the heat of fusion,the lower the required PCM mass as shown in Equation2.

M = hr

Equation 2

Where-

m = Mass of phase change material (kilograms)

E = PCM energy storage requirement (Joules)

hf = PCM heat of fusion (Joules/kilogram)

Water has a heat of fusion almost 70% higher than atypical PCM with the appropriate control (melt)temperature. Therefore, the use of water as the PCMwould significantly reduce the required heat exchangermass. Of course, there are some unique challengesassociated with the use of water as the PCM. Unlikemost fluids, water expands when it freezes which resultsin unique structural design challenges.

The Advanced Thermal project is developing two typesof ice PCM heat exchangers. The two types of heatexchangers use the same phase change material(water), but have a subtle difference in the method whichthe energy is added to and removed from the heatexchanger.

The first heat exchanger will be designed to interfacewith a traditional active thermal control system. Theenergy would be added to and removed from the PCMby a pumped fluid loop flowing through the PCMhardware. This type of heat exchanger is planned to beused on Orion. The project has worked with EnergySciences Laboratory (ESLI) to develop several ice PCMheat exchangers for evaluation. The first heatexchanger, Replicative Ice PCM (RIP), was designed toreplicate the energy storage of an existing paraffin-basedPCM heat exchanger. Due to water's higher heat offusion, RIP is much lighter and smaller than the baseline.The two units are shown in Figure 13.

2.4

^ 1.9

1.4

Figure 13. Size comparison between paraffin-basedPCM and Replicative Ice PCM (RIP).

The final RIP mass is approximately 5.4 kg while thebaseline heat exchanger mass is 8.4 kg. In addition tothis significant mass savings, the RIP heat exchangervolume is only 3500 cm 3 compared to 6600 cm 3 for theparaffin-based PCM heat exchanger'Z.

In addition to the replicative ice PCM heat exchanger, theproject collaborated with ESLI to develop two smaller ice-based PCM heat exchangers as shown in Figure 14.

F

^cT`

a.

Figure 14. PCM heat exchangers developed incollaboration with ESLI. From left to right; SHRIMP-1, SHRIMP-2, RIP, and the paraffin-based PCM heatexchanger.

In the preceding figure, RIP is shown as the third heatexchanger from the left. The two smaller ice-based heatexchangers, SHRIMP-1 and SHRIMP-2 (Small Heat sinkof Replicative Ice Material for Phase change) weredeveloped to assess the feasibility of using water as thephase change material in a very cost effective manner.SHRIMP-1 incorporated the same interstitial materialconfiguration as RIP. SHRIMP-2, on the other hand, wasdesigned to create a void space between the aluminumfins rather than randomly located void spaces throughoutthe heat exchanger. All of the water-based heatexchangers were filled with water to approximately 80%

of the void volume. Each of the preceding heatexchangers was exposed to several freeze/thaw cyclesin both favorable and adverse (if applicable) gravityorientations. RIP was exposed to a total of fivefreeze/thaw cycles while both SHRIMPS were cycled 45times during the test program. Unfortunately, each of theheat exchangers failed due to the expansion of iceduring the freeze cycles. A detailed description of thistest program is provided by Leimkuehler, et. al.

The second water-based PCM heat exchanger beingdeveloped by the Advanced Thermal project is critical forthe Lunar Electric Rover. This heat exchanger will notinterface with a coolant loop. The energy will betransferred to the phase change material through an airduct. The PCM heat exchanger is integrated with aradiator to refreeze the phase change material when theheat load is decreased or the thermal environment ismore benign. The LER is an ideal application due to thetransient nature of the vehicle's heat rejectionrequirement as shown in Figure 15.

zuuu

1800 —Equipment

—Metabolic

Total

— Average (Radiator)

1600

1400 -

1200

looa

soo

----- - -------- --

---- ------ -----

600

400

200

0

EVA Exe (m EVA Aeep

16 20 24

Time (hours)

Figure 15. Lunar Electric Rover (LER) heat rejectionprofile.

Typically, radiators are sized for the maximum heat loadin the warmest continuous thermal environment. Usingthe heat rejection profile shown in Figure 15 as anexample, the vehicle's radiator would have to be sizedfor nearly 1900 Wafts. However, because the vehicle'sthermal control system includes a phase change materialheat sink, the radiator can be sized for the average heatrejection requirement which is only 950 Watts therebyresulting in a smaller radiator.

The project designed and fabricated a total of three PCMheat sinks that were included in a thermal vacuum testprogram. One of the test articles is shown in Figure 16_

Figure 16. Photograph of one of the PCM heat sinktest articles showing the nomex insulation and the"radiator" lid.

While the preceding figure clearly shows the sides andthe lid, the heater is not shown. The test article can bethought of as a six-sided box with the "radiator" lid as thetop surface. For this test article, the heat load is appliedusing a strip heater adhered to the bottom surface. Thetest articles were installed into the thermal vacuumchamber as shown in Figure 17_

Figure 17. The three test articles installed in thethermal vacuum chamber. The infrared lamps usedto control the environmental temperature are shownon the top.

A test article transient heat load, which was scaled usingthe data shown in Figure 15, was applied to the bottomside of the test article using the aforementioned stripheaters. In addition to this simulated mission profile,several freeze/thaw cycles were also performed. Thefabrication of the test articles and the subsequent testprogram demonstrated proof-of-concept that a phasechange material heat sink can be integrated with aradiator and used to provide the heat rejection/storagefunction of a thermal control system13.

CONCLUSION

NASA's Exploration Technology Development Programincludes several projects performing technologydevelopment for the Constellation Program. One ofthese projects is the Thermal Control SystemDevelopment for Exploration project (Advanced Thermalproject). The Advanced Thermal project's objectives areto develop viable solutions for thermal design challengesand to mitigate key risks for Orion, Altair, and LunarSurface Systems through technology development.While the project is currently focused on theConstellation Program, the overwhelming majority of thedevelopment efforts are generically applicable to anyspacecraft thermal control system.

The technology development process begins withtechnologies possessing a TRL of approximately two orthree and advances them to a TRL of six. The TRL six isachieved by completing an integrated thermal testapproximately one year prior to the customer'sPreliminary Design Review.

The project's portfolio is very broad and includes thermalcontrol system fluids, heat acquisition hardware, andevaporative heat sinks. In addition to these elements,the project is also focused on developing spacecraftradiators. The final element within the project's portfoliois the advancement of Phase Change Material heatexchanger and heat sinks. These devices are currentlybeing developed for Orion and the Lunar Electric Rover.

ACKNOWLEDGMENTS

The author of this paper is the project manager for thetechnology development project. The author would liketo thank all of the individuals involved with making this asuccessful project. The project members are from avariety of NASA field centers including GRC, JPL, andJSC. In addition to this dedicated workforce, the projectalso collaborates with various industry partners includingPacific Northwest National Laboratories, Energy ScienceLaboratories, Paragon Space Development, JacobsEngineering, Oceaneering Space Systems, HamiltonSundstrand, and Ashwin-Ushas Corporation. The authorwould also like to thank his division management andAltair project office personnel for giving him theopportunity to work on the next generation Lunar lander.None of this could have been accomplished without thesupport of the program office and all of the helpful andenthusiastic individuals working in that office.

REFERENCES

1. Sheth, R. B., Stephan, R. A., Leimkuehler, T. O."Sublimator Driven Coldplate EngineeringDevelopment Unit Test Results and Development ofSecond Generation SDC." AIAA ICES PaperxxICES-xxxx. July 2010.

2. Stephan, R. A. "Overview of NASA's ThermalControl System Development for ExplorationProject." SAE ICES Paper 09ICES-0351. July 2009.

3 Lee, S. A. "Evaluation of a Thermal Control Fluid forConstellation Vehicles." AIAA ICES Paper xxICES-xxxx. July 2010.

4 Shin, E. E., Johnston, J. C. "Design andDevelopment of Lightweight Air-Liquid CompositeHeat Exchangers for Space ExplorationApplications." AIAA ICES Paper xxICES-xxx. July2010.

5 Hawkins-Reynolds, E., Le, H. V., Stephan, R. A."Development, Fabrication, and Testing of aLiquid/Liquid Microchannel Heat Exchanger forConstellation Spacecrafts." AIAA ICES PaperxxICES-xxx. July 2010.

6 Sheth, R. B., Stephan, R. A., Leimkuehler, T. O."Investigation of Transient Performance for aSublimator." AIAA ICES Paper xxICES-xxxx. July2010.

7 Ochoa, D. A., Ewert, M. "A Comparison BetweenOne- and Two-Loop ATCS Architectures Proposedfor CEV_" SAE ICES Paper 09ICES-0353. July2009.

8 Bannon, E. T., Bower, C. E., Sheth, R. B."Electrochromic Radiator Coupon Level Testing andFull Scale Thermal Math Modeling for Use on AltairLunar Lander." AIAA ICES Paper xxICES-xxxx. July2010.

9

Sunada, E., Birur, G. C., Ganapathi, G. B., Miller, J.,Berisford, D., Stephan, R. A. "Design and Testing ofan Active Heat Rejection Radiator with Digital Turn-Down Capability." AIAA ICES Paper xxICES-xxxx.July 2010.

10 Lillibridge, S. T., Navarro, M., Cognata, T., Guinn, J."Freezable Radiator Testing." AIAA ICES PaperxxICES-xxxx. July 2010.

11 Gaier, J. R. "Effect of Simulant Type on theAbsorptivity and Emissivity of Dusted ThermalControl Coatings in a Simulated Lunar Environment."AIAA ICES Paper xxICES-xxxx. July 2010.

12 Leimkuehler, T. O., Hansen, S., Stephan, R. A."Development, Testing, and Failure Mechanisms of aReplicative Ice Phase Change Material HeatExchanger." AIAA ICES Paper xxICES-xxxx. July2010.

13 Lee, S. A., Leimkuehler, T. O., Stephan, R. A., Le, H.V. "Thermal Vacuum Test of Ice as a PhaseChange Material Integrated with a Radiator." AIAAICES Paper xxICES-xxxx. July 2010.

CONTACT

Ryan A. StephanRyan.A.Stephan@nasa.gov

ACRONYMS & ABBREVIATIONS

ATCS: Active Thermal Control SystemCXPO: Constellation Program OfficeEDU: Engineering Development UnitESLI: Energy Science Laboratories, Inc.ESMD: Exploration Systems Mission DirectorateETDP: Exploration Technology Development ProgramISS: International Space StationJSC: Johnson Space CenterLEM: Lunar Excursion ModuleLER: Lunar Electric RoverLLO: Low Lunar OrbitLSS: Lunar Surface SystemsMIPR: Military Interdepartamental Purchase RequestOAFU: Orbit-Averaged Feedwater UtilizationPCM: Phase Change MaterialPDR: Preliminary Design ReviewPNNL: Pacific Northwest National LaboratoriesRIP: Replicative Ice PCMSDC: Sublimator Driven ColdplateSHReD: Supplemental Heat Rejection DeviceTCS: Thermal Control SystemTRL: Technology Readiness LevelVIO: Vehicle Integration OfficeWBS: Work Breakdown Structure

A: Radiator surface areas: Infrared emissivityE: Phase change material energy storage requirementh f9 : Phase change material heat of fusionK: Kelvinm: Mass of phase change materialm2 : Square metersQ: Radiator heat rejectionQmax: Maximum system heat rejection requirementQmin : Minimum system heat rejection requirement6: Stefan-Boltzman constantT.: Radiator sink temperatureTR : Average radiator temperatureW:Watts