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Space Robotics

Thermal Materials

Coatings, insulation, heaters, thermal switches and phase change materials selected by flown and qualified planetary robotics programs, and the ground and orbital evidence on how those materials degrade.

A surface robot has to hold electronics and batteries inside a narrow band while the environment swings outside it. The lunar surface runs from 350 to 400 K by day to 50 to 100 K at night, and a night lasts about 14 days [14]. A crewed lunar terrain vehicle concept adds a requirement to survive a night of at least 85 hours and preferably 125 hours [22]. Mars ambient air runs from -90 to +10 C, and the Athena rover allowable range for electronics and science was -40 to +40 C, with the lithium-ion battery the most sensitive item: aging above 40 C and electrolyte freezing below -30 C [13]. Dust adds a second problem. It raises the absorptivity of a radiator and cannot always be brushed off. A review of the six Apollo landings sorts the recorded dust problems into nine categories, among them thermal control, and finds that simple mitigation worked only for the easy ones [5]. On Apollo 17 the two LRV batteries read 95 and 110 F at first power-up against 80 F predicted, and a battery temperature sensor later failed from electrolyte leaking through its bond [3].

Design space

Section titled “Design space”

Five families of approach appear in the held record.

Passive surfaces and radiators. White paints and second surface mirrors set the ratio of solar absorptance to emittance. They are the cheapest option and the one most exposed to dust, as the measurements below show. A lander study sizes radioisotope heater unit power and radiator area for the 14.75 day night at five candidate latitudes, using a single-node placeholder lander, and is simulation only [16].

Insulation. Pathfinder’s microrover warm electronics box used solid silica aerogel inside an E-glass/epoxy sheet-and-spar structure, so the insulation also carried the structural load; conductivity was measured on coupons, not on the assembled flight box [17]. A JPL abstract proposes a carbon dioxide filled multilayer insulation enclosure for Mars landers, with test results announced but not in the abstract [13]. Mars Science Laboratory flew gas-gap insulation: the rover avionics mounting plate sits in keep-out gaps of 7.6 cm (top), 5.0 cm (bottom) and 3.8 cm (sides) that fill with 8 Torr carbon dioxide, whose conductivity is 0.010 to 0.016 W/m-K, and about a quarter of the 150 W coldest-case heat loss passes through those gaps [31]. A bell jar test saw convection onset near 7.6 cm against about 6.4 cm predicted by correlation; it was horizontal, at 1 g, on flat plates only, so the vertical side gaps and Mars gravity rest on correlation [31]. The Ingenuity helicopter’s thermal vacuum test at Mars pressure found gas conduction across internal gaps, not radiation, to be the largest driver of survival energy, which led to enlarging the fuselage by 25 mm in one direction; the saving from removing test cabling is a model prediction, not a test result [26]. Spacerless multilayer insulation and low conductance isolators are among ten lunar night innovations at development status [14].

Active and stored heat. Thermostatic Kapton heaters have flown on Perseverance, in a redundant pair per Mastcam-Z camera head, holding a nominal -40 to +40 C [2]. Phase change storage appears at three maturities: n-hexadecane on IPEx [1], a Mars rover and lander concept pairing phase change storage with a variable conductance loop heat pipe and radioisotope heater units [13], and a stationary molten salt unit meant to heat other lunar assets through the night, tested only as a scaled prototype at ambient conditions [23]. Mars Science Laboratory flew the pumped alternative: two single-phase mechanically pumped fluid loops, one for cruise heat rejection and one that moves the 2000 W of MMRTG waste heat to hold rover electronics at -40 to +50 C against a -120 to +30 C surface [25]. The two heat exchanger panels around the MMRTG are aerogel-filled Nomex honeycomb sandwiches with a through-thickness conductance specified at 0.35 to 0.55 W/C; in the January 2009 cruise thermal vacuum test the liquid-to-liquid exchanger moved only 0.1 C across the cold plate because radiation to the backshell removed the heat, and model-to-test agreement averaged 5 C, with the MMRTG model off by about 10 C. That test was cruise only and does not verify surface performance [33]. The 16-day rover system thermal test ran 18 cases in 8 Torr nitrogen, not carbon dioxide, with shroud temperatures from +20 to -105 C, found no allowable flight temperature violation, and correlated models to within 5 C; Mars gravity, wind and the carbon dioxide atmosphere were absent, so the deliverable was a correlated model and not a validation [32]. Development testing found the conductive epoxy bonding loop tubing to the heat exchanger plates about an order of magnitude below its published conductivity, which forced a change of bonding method [25]. Early Mars surface telemetry compared with the ground solar thermal test showed the models conservative in the cold, with rover avionics mounting plate interfaces 3 to 10 C warmer in test than predicted, and accurate in the hot, within 0.2 C; the winter case the heat recovery exists for had not yet been flown when that was reported [27]. The precedent was Pathfinder’s Refrigerant 11 pumped loop, which flew only in cruise and was vented before entry; its 14000 hour ground life test led to a monthly one hour backup pump run to flush bearing particulates, and a wax valve bypass held the electronics shelf near -4 C while the radiator fell to -70 C [28]. Milliwatt radioisotope heater and thermoelectric modules for small probes are a concept study with no hardware built [18].

Switchable thermal paths. Thermally switched enclosures, reverse-operation thermal switches and miniaturized loop heat pipes are developed to reject heat by day and retain it at night without radioisotopes; the reported on/off ratios are headline numbers, and the mini switch’s 1000:1 is a target [14]. A crewed lunar terrain vehicle concept instead proposes a fully passive system of insulation, coatings, dust covers and a nested double box, with no analysis shown [22]. IPEx’s cover with an insulation blanket as thermal switch is the one flown-class example of the mechanism approach here [1].

Surface protection against dust and erosion. Brushing, coatings and cover mechanisms address the way dust degrades the first family. Brushing is the Apollo method and tests below show it recovers only part of the lost performance [7]. A nano-textured hydrophobic coating has been tried on radiator paint [15]. IPEx replaces a static radiator with an actuated cover [1]. Mars 2020 qualification also exposed white thermal paint to regolith thrown by the descent engine plume [9]. Thermal models of a 10 to 100 micrometer dust layer on an airless body (Ryugu, simulation only) show apparent thermal inertia shifting by more than 20 percent, with masking complete at about one diurnal skin depth [29]. A crewed-module counterpart is the KONTUR-2 joystick on the ISS, which has no active cooling: a finite element model predicted a 34.6 C base plate against a 40 C limit, and a nine-sensor control system holds the limit because the idle steady state of 46.4 C exceeds it [30].

What has flown: Kapton heaters on Mars [2], the MSL pumped fluid loop [25], the Apollo LRV batteries and their sensors [3][4], Ingenuity’s thermal design [26], the Pathfinder aerogel box design [17], and orbital exposure of coatings and polymers on LDEF and MISSE [11][12][19][20]. IPEx’s materials are at laboratory and qualification level [1][24]. The switch, loop heat pipe, molten salt and milliwatt radioisotope concepts have not flown [13][14][18][23].

Thermal materials selected and qualified

Section titled “Thermal materials selected and qualified”
PartUsed byStatusSource
S13GLO white thermal paintIPEx radiator plateQualification, tribo-charging measured[1]
n-Hexadecane, 99+ percentIPEx avionics heat sinkHousing built, vacuum heat transfer test pending[1]
Polyetherimide (Ultem), Cerakote coatedIPEx radiator coverBallistic dust test[1]
Etched-foil Kapton heatersPerseverance Mastcam-ZFlown since February 2021[2]
Silica aerogel in E-glass/epoxyPathfinder microrover warm electronics boxCoupon conductivity[17]
Pumped single-phase fluid loopMSL roverFlown[25]

IPEx radiator and heat sink. S13GLO is a zinc oxide pigment in an RTV602 silicone binder, the low-outgassing variant of the S13G family, applied to the 6061 aluminum cover that forms the top of the phase change housing and doubles as the radiator surface [1]. In a tribo-charging test at 1e-5 Torr against an MLI sheet on a 100 pF capacitor it produced peaks of 0.004 V and 0.006 V, or surface charge densities of 5.25e-6 and 7.87e-6 microcoulomb per square meter. Those values fed a COMSOL particle trajectory model of whether separation of the cover lofts dust onto the radiator [1]. The n-hexadecane melts at 18 C, with a calculated latent heat capacity of 300.5 kJ, and is held in a 6061 aluminum housing with internal fins bonded with thermal epoxy; the assembly weighs 2.5 kg. It was sized by transient Thermal Desktop analysis, and vacuum chamber testing under the time-varying non-uniform avionics load was still planned when the paper was written [1]. The Ultem cover is non-metallic so it can carry the wireless charger antenna, and an MLI blanket under it works as a thermal switch that blocks heat flow to the phase change material during operation. In a ballistic test with silica sand the cover was actuated 90 degrees and particles flowed off, with minor accumulation near the cover motor from particle deflection, recorded as a design change for the next iteration [1]. The IPEx actuators were characterized separately at hot and cold motor temperatures in a vacuum dynamometer, and that test did not exercise the radiator cover [24].

Thermal control surfaces measured against dust, erosion and orbit

Section titled “Thermal control surfaces measured against dust, erosion and orbit”

Every coating below has a measured optical property and a measured degradation of it. The degradation comes from dust coverage or particle impact rather than from ultraviolet or dose.

SurfaceTest environmentMeasured resultSource
S-13 white paintBasalt dust, simulated lunar conditionsAbsorptance 0.195 clean to 0.848 near full coverage[10]
Aluminized TeflonBasalt dust, simulated lunar conditionsAbsorptance 0.185 clean to 0.864 dusted[10]
AZ-93 on aluminumGlenn bell jar, JSC-1AFAbsorptivity to emissivity 0.22 pristine, 0.65 dusted[6][7]
S13GP:6N/LO-IDayton erosion rigRatio up about 37 percent after regolith exposure[9]
Aluminized Teflon FEP, 5 milMISSE 9 zenith67 percent of elongation lost in 0.54 years[12]
Sputtered coatings on KaptonLDEF tray S1003No spalling after about 33600 cycles[11]

Dust on radiators. In 1967 basalt dust measurements, absorptance of S-13 rose nonlinearly, and 11 percent areal coverage roughly doubled it [10]. Emittance barely moved because basalt dust emits near 0.88, close to the coatings themselves, and the same work tested four removal concepts but could not fabricate the electrostatic curtain for lack of funds [10]. Bell jar work found that a monolayer of activated simulant raises the absorptivity to emissivity ratio of a surface by about a factor of 3 regardless of coating or substrate, so a radiator expected to be dusted must be about three times larger. That extraction rests on a back-solved lamp intensity of 29.1 kW/m2 and AZ-93 on aluminum with pristine emissivity 0.886 and absorptivity 0.173 [6]. A thermal engineer’s synthesis reports solar absorptivity rising by as much as 50 percent at 12 percent area coverage and emissivity falling by as much as 16 percent at 54 percent coverage, so a sub-monolayer already degrades both [8].

Brushing. Bench top brushing effectiveness did not transfer to the chamber. In the bell jar AZ-93 recovered from 0.65 to 0.30 after 20 strokes of a fiberglass round brush and 0.26 after 200, with residual coverage of 3 percent for that brush and 10 percent for a nylon fan on AZ-93, against 29 to 31 percent on silvered FEP; the spread between samples is about a factor of two [7]. Apollo experience points the same way, since ground tests that predicted brushing would work were contradicted on the lunar surface [5].

Erosion by landing plumes. Mars 2020 qualification campaigns used three rigs for particles from 38 microns to 15.8 mm, and the S13GP:6N/LO-I absorptivity to emissivity ratio rose about 37 percent after exposure at 38 to 44 microns and 250 m/s as particles embedded in the soft silicone binder, plateauing near 0.55 [9]. Under 177 to 250 micron sand at 160 m/s and 30 degrees, erosion was linear until about 18.5 percent of initial weight was lost, where the paint separated from the aluminum, above the expected landing environment [9].

Passive dust coatings. The Lotus coating was applied to coupons including white radiator paint and tested with JSC-1 simulant. The layer was inconsistently present across coupons of the same formulation, and exposure times were short [15].

Low Earth orbit. LDEF returned the only measured exposure envelope in the held record: 5.8 years, atomic oxygen from 1e3 to 8.81e21 atoms/cm2 between wake and ram, and 4500 to 14500 equivalent sun hours [20]. Polymers were attacked by atomic oxygen while very thin inorganic coatings were not: ion beam deposited Al2O3 at 700 angstroms and SiO2 at 650 angstroms on Kapton held adhesion over 69 months [11]. The systems investigation attributed no system anomaly on LDEF to long exposure, and the two most significant active failures were relays [21]. MISSE 1 to 7 carried more than 540 samples; the MISSE 2 table gives erosion yields for 41 samples of 40 polymers at one fluence of 8.43e21 atoms/cm2, and ground asher testing eroded polymers faster than space in every case, so ground yields overstate erosion [19]. The MISSE 9 to 15 decks report 365 samples and silicon contamination reaching 15.1 atomic percent on a ram surface in 0.89 years [12].

What is not established

Section titled “What is not established”

The IPEx phase change housing had not completed vacuum chamber testing of heat transfer under a non-uniform avionics load when it was reported [1]. The Lotus coating’s reproducibility and long-duration stability are unresolved [15]. The dust measurements on AZ-93 and S13 are simulant-based and coupon-scale, the 1967 coverage fractions are visual estimates, and the bell jar extractions inherit an uncalibrated lamp intensity [6][10]. The Apollo 17 sensor failure cause is stated as most probable, not confirmed [3]. The switch and loop heat pipe figures [14], the milliwatt radioisotope designs [18], the molten salt unit [23] and the lunar terrain vehicle concept [22] have no flight or lunar thermal vacuum result in the held record. The CO2 insulation and phase change figures are expectations from an abstract with no test data [13]. The MSL fluid loop design paper gives no flow rate, pressure or measured temperature [25], and the surface performance paper prints no uncertainties [27]. The Ryugu dust result is a model [29], and the KONTUR-2 case is closed by control logic, not margin [30]. The Pathfinder aerogel conductivities are coupon values [17]. The MISSE 3 to 7 entries in the 2008 overview are plans, and the 2024 deck withholds its erosion yields [12][19]. The lander night survival study uses a placeholder lander [16]. The LDEF no-anomaly finding does not bound risk for actuated mechanisms [21].

References

  1. Schuler, J. M., Smith, J. D., Nick, A. J., Buckles, B. C., Dyas, J. E., Ortega, V. V., Cloud, J. M., Dokos, A. G., Zhang, E. L., Wang, J. J., Baron, M. A., Muller, T. J., Clark, C. J. and Howe, M. W. (2024). ISRU Pilot Excavator (IPEx) Technology Readiness Level 5 Design Overview . AIAA AVIATION Forum and ASCEND, 20240008162. Source
    BibTeX
    @inproceedings{schuler2024isru,
  title = {ISRU Pilot Excavator (IPEx) Technology Readiness Level 5 Design Overview},
  author = {Schuler, Jason M. and Smith, Jonathan D. and Nick, Andrew J. and Buckles, Bradley C. and Dyas, Jeffrey E. and Ortega, Victoria V. and Cloud, Joseph M. and Dokos, Adam G. and Zhang, Elizabeth L. and Wang, Jerry J. and Baron, Michael A. and Muller, Thomas J. and Clark, Casey J. and Howe, Musashi W.},
  booktitle = {AIAA AVIATION Forum and ASCEND},
  number = {20240008162},
  institution = {NASA},
  year = {2024},
  doi = {10.2514/6.2024-4890},
  abstract = {This paper details the mechanical and mechatronic design of the Technology Readiness Level (TRL)-5 In-Situ Resource Utilization (ISRU) Pilot Excavator (IPEx). IPEx is a robotic excavator designed for a technology demonstration of regolith mining in the lunar south pole region. The novel design uses pairs of counter-acting excavation tools called bucket drums, that dig at the same time in opposing directions to reduce the reaction force needed, thereby enabling mining with a small, low-mass, robotic system. IPEx builds on the prior work of the Regolith Advanced Surface Systems Operations Robot (RASSOR), which is the TRL-4 implementation of this concept. The TRL-5 IPEx subsystems that are discussed in this paper include: Regolith Delivery Subsystem (RDS), Mobility Subsystem (MS), Cameras and Dust Mitigation Subsystem (CDMS), and Thermal Control Subsystem (TCS). Each subsystem is described in detail with rationale for design selections. Dust tolerance is a key feature for IPEx, and this paper details a thermal control system with an actuated radiator cover and phase change material as well as camera modules with removable electrodynamic dust shields (EDS). Additional components such as actuators, wheels, and bucket drums are discussed in detail. Due to their complexity, the avionics and software subsystems will be discussed in a separate publication.}
}
  2. Bell, I. J., Maki, J. N., Mehall, G. L., Ravine, M. A., Caplinger, M. A., Bailey, Z., Brylow, S., Schaffner, J. A., Kinch, K. M., Madsen, M., Winhold, A. G., Hayes, A. G., Corlies, P., Tate, C., Barrington, M., Cisneros, E., Jensen, E., Paris, K., Crawford, K., Rojas, C., Mehall, L., Joseph, J., Proton, J., Cluff, N., Deen, R. G., Betts, B., Cloutis, E., Coates, A. J., Colaprete, A., Edgett, K. S., Ehlmann, B. L., Fagents, S. A., Grotzinger, J. P., Hardgrove, C., Herkenhoff, K., Horgan, B., Jaumann, R., Johnson, J. R., Lemmon, M., Paar, G., Caballo-Perucha, M., Gupta, S., Traxler, C., Preusker, F., Rice, M. S., Robinson, M. S., Schmitz, N., Sullivan, R. and Wolff, M. (2021). The Mars 2020 Perseverance Rover Mast Camera Zoom (Mastcam-Z) Multispectral, Stereoscopic Imaging Investigation . Space Science Reviews, 24. Source
    BibTeX
    @article{bell2021mars,
  title = {The Mars 2020 Perseverance Rover Mast Camera Zoom (Mastcam-Z) Multispectral, Stereoscopic Imaging Investigation},
  author = {Bell, III, J.F. and Maki, Justin N. and Mehall, Greg L. and Ravine, Michael A. and Caplinger, Michael A. and Bailey, Z.J. and Brylow, S. and Schaffner, Jacob A. and Kinch, Kjartan M. and Madsen, M.B. and Winhold, Andrew G. and Hayes, Alexander G. and Corlies, Paul and Tate, C. and Barrington, Megan and Cisneros, Ernest and Jensen, E. and Paris, Kristen and Crawford, Kelsie and Rojas, Corrine and Mehall, Laura and Joseph, Jonathan and Proton, J.B. and Cluff, Nathan and Deen, Robert G. and Betts, B. and Cloutis, E. and Coates, Andrew J. and Colaprete, Anthony and Edgett, Kenneth S. and Ehlmann, Bethany L. and Fagents, Sarah A. and Grotzinger, John P. and Hardgrove, Craig and Herkenhoff, K.E. and Horgan, B. and Jaumann, Ralf and Johnson, Jeffrey R. and Lemmon, M. and Paar, Gerhard and Caballo-Perucha, M. and Gupta, Sanjeev and Traxler, Christoph and Preusker, Frank and Rice, Melissa S. and Robinson, Mark S. and Schmitz, Nicole and Sullivan, R. and Wolff, M.J.},
  journal = {Space Science Reviews},
  volume = {217},
  number = {24},
  pages = {24--24},
  year = {2021},
  doi = {10.1007/s11214-020-00755-x},
  abstract = {Abstract Mastcam-Z is a multispectral, stereoscopic imaging investigation on the Mars 2020 mission’s Perseverance rover. Mastcam-Z consists of a pair of focusable, 4:1 zoomable cameras that provide broadband red/green/blue and narrowband 400-1000 nm color imaging with fields of view from 25.6° × 19.2° (26 mm focal length at 283 μrad/pixel) to 6.2° × 4.6° (110 mm focal length at 67.4 μrad/pixel). The cameras can resolve (≥ 5 pixels) ∼0.7 mm features at 2 m and ∼3.3 cm features at 100 m distance. Mastcam-Z shares significant heritage with the Mastcam instruments on the Mars Science Laboratory Curiosity rover. Each Mastcam-Z camera consists of zoom, focus, and filter wheel mechanisms and a 1648 × 1214 pixel charge-coupled device detector and electronics. The two Mastcam-Z cameras are mounted with a 24.4 cm stereo baseline and 2.3° total toe-in on a camera plate ∼2 m above the surface on the rover’s Remote Sensing Mast, which provides azimuth and elevation actuation. A separate digital electronics assembly inside the rover provides power, data processing and storage, and the interface to the rover computer. Primary and secondary Mastcam-Z calibration targets mounted on the rover top deck enable tactical reflectance calibration. Mastcam-Z multispectral, stereo, and panoramic images will be used to provide detailed morphology, topography, and geologic context along the rover’s traverse; constrain mineralogic, photometric, and physical properties of surface materials; monitor and characterize atmospheric and astronomical phenomena; and document the rover’s sample extraction and caching locations. Mastcam-Z images will also provide key engineering information to support sample selection and other rover driving and tool/instrument operations decisions.}
}
  3. MIssion Evaluation Team. (1973). Apollo 17 Mission Report . NASA, NASA-TM-. Source
    BibTeX
    @techreport{anon1973apollo,
  title = {Apollo 17 Mission Report},
  author = {{MIssion Evaluation Team}},
  number = {NASA-TM-},
  institution = {NASA},
  year = {1973},
  url = {https://ntrs.nasa.gov/citations/19730015117},
  abstract = {Operational and engineering aspects of the Apollo 17 mission are outlined. The vehicle configuration was similar to those of Apollo 15 and 16. There were significant differences in the science payload for Apollo 17 and spacecraft hardware differences and experiment equipment are described. The mission achieved a landing in the Taurus-Littrow region of the moon and returned samples of the pre-Imbrium highlands and young craters.}
}
  4. NASA Manned Spacecraft Center. (1971). Apollo 15 Mission Report . NASA Manned Spacecraft Center, MSC-05161. Source
    BibTeX
    @techreport{anon1971apollo,
  title = {Apollo 15 Mission Report},
  author = {{NASA Manned Spacecraft Center}},
  number = {MSC-05161},
  institution = {NASA Manned Spacecraft Center},
  year = {1971},
  url = {https://ntrs.nasa.gov/citations/19720021182},
  abstract = {A detailed discussion is presented of the Apollo 15 mission, which conducted exploration of the moon over longer periods, greater ranges, and with more instruments of scientific data acquisition than previous missions. The topics include trajectory, lunar surface science, inflight science and photography, command and service module performance, lunar module performance, lunar surface operational equipment, pilot's report, biomedical evaluation, mission support performance, assessment of mission objectives, launch phase summary, anomaly summary, and vehicle and equipment descriptions. The capability of transporting larger payloads and extending time on the moon were demonstrated. The ground-controlled TV camera allowed greater real-time participation by earth-bound personnel. The crew operated more as scientists and relied more on ground support team for systems monitoring. The modified pressure garment and portable life support system provided better mobility and extended EVA time. The lunar roving vehicle and the lunar communications relay unit were also demonstrated.}
}
  5. Gaier, J. R. (2007). The Effects of Lunar Dust on EVA Systems During the Apollo Missions . NASA, NASA/TM-2005-213610/REV1. Source
    BibTeX
    @techreport{gaier2007effects,
  title = {The Effects of Lunar Dust on EVA Systems During the Apollo Missions},
  author = {Gaier, James R.},
  number = {NASA/TM-2005-213610/REV1},
  institution = {NASA},
  year = {2007},
  url = {https://ntrs.nasa.gov/citations/20070021819},
  abstract = {Mission documents from the six Apollo missions that landed on the lunar surface have been studied in order to catalog the effects of lunar dust on Extra-Vehicular Activity (EVA) systems, primarily the Apollo surface space suit. It was found that the effects could be sorted into nine categories: vision obscuration, false instrument readings, dust coating and contamination, loss of traction, clogging of mechanisms, abrasion, thermal control problems, seal failures, and inhalation and irritation. Although simple dust mitigation measures were sufficient to mitigate some of the problems (i.e., loss of traction) it was found that these measures were ineffective to mitigate many of the more serious problems (i.e., clogging, abrasion, diminished heat rejection). The severity of the dust problems were consistently underestimated by ground tests, indicating a need to develop better simulation facilities and procedures.}
}
  6. Gaier, J. R., Siamidis, J. and Larkin, E. M. G. (2010). Extraction of Thermal Performance Values from Samples in the Lunar Dust Adhesion Bell Jar . Space Simulation Conference, NASA/TM-2010-216828. Source
    BibTeX
    @inproceedings{gaier2010extraction,
  title = {Extraction of Thermal Performance Values from Samples in the Lunar Dust Adhesion Bell Jar},
  author = {Gaier, James R. and Siamidis, John and Larkin, Elizabeth M. G.},
  booktitle = {Space Simulation Conference},
  number = {NASA/TM-2010-216828},
  institution = {NASA},
  address = {Annapolis, MD},
  year = {2010},
  url = {https://ntrs.nasa.gov/citations/20100039312},
  abstract = {A simulation chamber has been developed to test the performance of thermal control surfaces under dusty lunar conditions. The lunar dust adhesion bell jar (LDAB) is a diffusion pumped vacuum chamber (10(exp -8) Torr) built to test material samples less than about 7 cm in diameter. The LDAB has the following lunar dust simulant processing capabilities: heating and cooling while stirring in order to degas and remove adsorbed water; RF air-plasma for activating the dust and for organic contaminant removal; RF H/He-plasma to simulate solar wind; dust sieving system for controlling particle sizes; and a controlled means of introducing the activated dust to the samples under study. The LDAB is also fitted with an in situ Xe arc lamp solar simulator, and a cold box that can reach 30 K. Samples of thermal control surfaces (2.5 cm diameter) are introduced into the chamber for calorimetric evaluation using thermocouple instrumentation. The object of this paper is to present a thermal model of the samples under test conditions and to outline the procedure to extract the absorptance, emittance, and thermal efficiency from the pristine and sub-monolayer dust covered samples.}
}
  7. Gaier, J. R., Journey, K., Christopher, S. and Davis, S. (2011). Evaluation of Brushing as a Lunar Dust Mitigation Strategy for Thermal Control Surfaces . International Conference on Environmental Systems, NASA/TM-2011-217231. Source
    BibTeX
    @inproceedings{gaier2011evaluation,
  title = {Evaluation of Brushing as a Lunar Dust Mitigation Strategy for Thermal Control Surfaces},
  author = {Gaier, James R. and Journey, Khrissaundra and Christopher, Steven and Davis, Shanon},
  booktitle = {International Conference on Environmental Systems},
  number = {NASA/TM-2011-217231},
  institution = {NASA},
  address = {Portland, OR},
  year = {2011},
  url = {https://ntrs.nasa.gov/citations/20120000070},
  abstract = {Evaluation of brushing to remove lunar simulant dust from thermal control surfaces is described. First, strip brushes made with nylon, PTFE, or Thunderon (Nihon Sanmo Dyeing Company Ltd.) bristles were used to remove JSC-1AF dust from AZ93 thermal control paint or aluminized FEP (AlFEP) thermal control surface under ambient laboratory conditions. Nylon and PTFE bristles removed a promising amount of dust from AZ93, and nylon and Thunderon bristles from AlFEP. But when these were tested under simulated lunar conditions in the lunar dust adhesion bell jar (LDAB), they were not effective. In a third effort, seven brushes made up of three different materials, two different geometries, and different bristle lengths and thicknesses were tested under laboratory conditions against AZ93 and AlFEP. Two of these brushes, the Zephyr fiberglass fingerprint brush and the Escoda nylon fan brush, removed over 90 percent of the dust, and so were tested in the fourth effort in the LDAB. They also performed well under these conditions recovering 80 percent or more of the original thermal performance (solar absorptance/thermal emittance) of both AZ93 and AgFEP after 20 strokes, and 90 or more percent after 200 strokes}
}
  8. Howard, A. Z. and Stewart, S. (2024). Thermal Impact of Lunar Dust on Rovers . Thermal and Fluids Analysis Workshop, 20240009829. Source
    BibTeX
    @inproceedings{howard2024thermal,
  title = {Thermal Impact of Lunar Dust on Rovers},
  author = {Howard, Abby Zinecker and Stewart, Sarah},
  booktitle = {Thermal and Fluids Analysis Workshop},
  number = {20240009829},
  institution = {NASA},
  address = {Cleveland, Ohio},
  year = {2024},
  url = {https://ntrs.nasa.gov/citations/20240009829},
  abstract = {Experience and ground tests have shown that lunar dust coverage can severely degrade thermal system performance, and with the push to go back to the Moon for longer than a few days, lunar dust is being recognized as a significant technical challenge. The Lunar Terrain Vehicle (LTV) and Pressurized Rover (PR) will be operating on the lunar surface for long durations and roving at high speeds with interaction with astronauts and other robotics, which will cause dust to transfer to the vehicle, potentially to critical thermal surfaces. Dust coverage results in a change of overall optical properties, increased resistance to heat rejection due to the insulating effect of a dust layer, and even abrasion to thermal surfaces and soft goods. This paper provides an overview of what is currently known and unknown about what will happen to thermal surfaces exposed to dust on the Lunar South Pole, some dust mitigation options and testing guidance, and what resources can be used to help overcome this problem. }
}
  9. Bradford, E., Rabinovitch, J. and Abid, M. (2019). Regolith Particle Erosion of Material in Aerospace Environments . IEEE Aerospace Conference. Source
    BibTeX
    @inproceedings{bradford2019regolith,
  title = {Regolith Particle Erosion of Material in Aerospace Environments},
  author = {Bradford, Emma and Rabinovitch, Jason and Abid, Mohamed},
  booktitle = {IEEE Aerospace Conference},
  pages = {1-15},
  publisher = {IEEE},
  year = {2019},
  doi = {10.1109/aero.2019.8741563}
}
  10. Tatom, F. B., Srepel, V., Johnson, R. D., Contaxes, N. A., Adams, J. G., Seaman, H. and Cline, B. L. (1967). Lunar Dust Degradation Effects and Removal/Prevention Concepts, Volume 2: Detailed Report . Northrop Space Laboratories, Huntsville, Technical Report 323, Volume 2. Source
    BibTeX
    @techreport{tatom1967lunarb,
  title = {Lunar Dust Degradation Effects and Removal/Prevention Concepts, Volume 2: Detailed Report},
  author = {Tatom, F. B. and Srepel, V. and Johnson, R. D. and Contaxes, N. A. and Adams, J. G. and Seaman, H. and Cline, B. L.},
  number = {Technical Report 323, Volume 2},
  institution = {Northrop Space Laboratories, Huntsville},
  year = {1967},
  url = {https://ntrs.nasa.gov/citations/19680011610},
  abstract = {Lunar dust degradation effects and dust removal/prevention concept}
}
  11. Levine, A. S. (1992). LDEF: 69 Months in Space. First Post-Retrieval Symposium, part 3 . Langley Research Center, 19920017840. Source
    BibTeX
    @techreport{levine1992ldef,
  title = {LDEF: 69 Months in Space. First Post-Retrieval Symposium, part 3},
  author = {Levine, Arlene S},
  number = {19920017840},
  institution = {Langley Research Center},
  year = {1992},
  url = {https://ntrs.nasa.gov/citations/19920017840},
  abstract = {This document is a compilation of papers presented at the First Long Duration Exposure Facility (LDEF) Post-Retrieval Symposium. The papers represent the preliminary data analysis of the 57 experiments flown on the LDEF. The experiments include materials, coatings, thermal systems, power and propulsion, science (cosmic ray, interstellar gas, heavy ions, and micrometeoroid), electronics, optics, and life science.}
}
  12. de Groh, K. K. (2024). Overview of Results from the MISSE 9-15 Polymers and Composites Experiment 1-4 (PCE 1-4), 20240009151. ntrs.nasa.gov/citations/20240009151
    BibTeX
    @misc{degroh2024overview,
  title = {Overview of Results from the MISSE 9-15 Polymers and Composites Experiment 1-4 (PCE 1-4)},
  author = {de Groh, Kim K.},
  howpublished = {Presentation at the International Symposium on Materials in the Space Environment and International Conference on Protection of Materials and Structures in the Space Environment},
  number = {20240009151},
  address = {St Raphael},
  year = {2024},
  url = {https://ntrs.nasa.gov/citations/20240009151},
  abstract = {Spacecraft in low Earth orbit (LEO) and other planetary orbits are subjected to harsh environmental conditions that can result in erosion, embrittlement and optical property degradation of susceptible materials threatening spacecraft performance and durability. To increase our understanding of space environmental effects on spacecraft materials, NASA Glenn Research Center has developed a series of experiments flown as part of the Materials International Space Station Experiment (MISSE) missions on the exterior of the International Space Station (ISS). Four Glenn experiments with 365 flight samples were flown on ISS’s MISSE-Flight Facility (MISSE-FF). These experiments are the Polymers and Composites Experiment-1 (PCE-1) flown as part of the MISSE-9 mission, the PCE-2 flown as part of the MISSE-10 mission, the PCE-3 flown as part of the MISSE-12 and MISSE-15 missions, and the PCE-4 flown as part of the MISSE-13 mission. The PCE 1-4 flight samples were flown in either ram, wake, zenith or nadir flight orientations. The primary objectives of the PCE 1-4 experiments are to determine the LEO atomic oxygen (AO) erosion yield, Ey (volume loss per incident oxygen atom) and radiation durability of spacecraft polymers, composites and coatings. In addition, each experiment has a wide variety of materials with numerous sample objectives.  This presentation provides an overview of the PCE 1-4 experiments, the numerous sample collaborators, the space environmental exposure of the flight samples, and an overview of the post-flight observations and results. The results include AO fluence and contamination analyses for each mission flight orientation, erosion data for the PCE 1-4 AO Ey samples, optical property data, and tensile results for the MISSE-9 and MISSE-13 tensile samples. In addition, an introduction of results from other unique samples such as indium tin oxide (ITO) coated thermal control samples, shape memory polymer composites, melanin based composites, and specialty coatings, will be provided.}
}
  13. Birur, G. and Novak, K. (2024). Advanced Thermal Control Technologies for Mars Rovers and Landers. dataverse.jpl.nasa.gov/dataset.xhtml
    BibTeX
    @misc{birur2024advanced,
  title = {Advanced Thermal Control Technologies for Mars Rovers and Landers},
  author = {Birur, G. and Novak, K.},
  year = {2024},
  url = {https://dataverse.jpl.nasa.gov/dataset.xhtml?persistentId=hdl:2014/15520}
}
  14. Bugby, D. and de Jesús Rivera, J. (2023). Thermal Technology Advancements for Extended-Duration Lunar Operation . International Conference on Environmental Systems. Source
    BibTeX
    @inproceedings{david2023a,
  title = {Thermal Technology Advancements for Extended-Duration Lunar Operation},
  author = {Bugby, David and de Jesús Rivera, José},
  booktitle = {International Conference on Environmental Systems},
  publisher = {International Conference on Environmental Systems},
  year = {2023},
  doi = {10.32865/2346/94671},
  abstract = {The desire for extended-duration robotic exploration of the moon without radioisotopes has highlighted a need for improved thermal capabilities. Described in this paper are ten innovations to enable multi-day/night operability/survivability of lunar instruments/systems. Extended-duration lunar operation is highly challenging due to the 50-100 K temperatures during the (15 Earth-day) lunar night and the 350-400 K temperatures (at low-to-mid latitudes) during the (15 Earth-day) lunar day. The innovations and the reasons why they’re needed are as follows: (1) thermally-switched enclosure (TSE), which houses the instrument/system and (using innovations 2-5, 8-9 below) passively transitions from daytime thermal coupling to nighttime thermal isolation; (2) parabolic reflector radiator (PRR), which provides a low sink temperature during lunar day at low-to-mid latitude sites; (3) spacerless MLI (SMLI), which is a low e* MLI concept involving nested double-aluminized Mylar layers hung from tension cable supports without spacers; (4) low conductance thermal isolator (LCTI), which conductively isolates systems from carriers and subsystems from each other; (5) reverse-operation DTE thermal switch (ROD-TSW), which is a high ON/OFF ratio, fully passive thermal switch used within the TSE; (6) extended duration ROD-TSW (ES-ROD-TSW), which is an extended-stroke version of the ROD-TSW for non-vacuum (or vacuum) environments; (7) miniaturized ROD-TSW (mini-ROD-TSW), which is needed for highly compact thermal-switching applications; (8) miniaturized loop heat pipe (mini-LHP), which transports heat from the TSE internal housing to the radiator with high conductance and minimal cross-sectional area; (9) Vectran tension cables (VTC), which provide conductive isolation and structural support; and (10) lunar magnetometer thermal enclosure (LMTE), which is the TSE for a future JPL lunar magnetometer. These ten advancements were developed at JPL under the PALETTE, ARTEMIS-T, and ROD-TSW projects. PALETTE is funded by NASA GCD, while ARTEMIS-T/ROD-TSW were funded by JPL. Also introduced in the paper are a few related future advancements.}
}
  15. Margiotta, D. V., Peters, W. C., Straka, S. A., Rodriguez, M., McKittrick, K. R. and Jones, C. B. (2010). The Lotus coating for space exploration: a dust mitigation tool . Optical System Contamination: Effects, Measurements, and Control. Source
    BibTeX
    @inproceedings{margiotta2010lotus,
  title = {The Lotus coating for space exploration: a dust mitigation tool},
  author = {Margiotta, Danielle V. and Peters, Wanda C. and Straka, Sharon A. and Rodriguez, Marcello and McKittrick, Kristin R. and Jones, Craig B.},
  booktitle = {Optical System Contamination: Effects, Measurements, and Control},
  series = {SPIE Proceedings},
  volume = {7794},
  pages = {77940I},
  year = {2010},
  doi = {10.1117/12.864480},
  abstract = {The "Lotus" dust mitigation coating is a new technology that is currently being developed and tested, at NASA Goddard Space Flight Center (GSFC), as a countermeasure for addressing dust accumulation issues for long-duration human space exploration. This coating sheds dust particles utilizing anti-contamination and self-cleaning properties that minimize dust accumulation on spacecraft surfaces. Shedding of dust particles is accomplished by reducing the surface energy and the amount of surface available for attachment. The Lotus coating is designed to preserve optimal long-term performance of critical spacecraft surfaces and systems, while minimizing and/or eliminating dust accumulation. NASA is exceedingly interested in simplistic and innovative ways to mitigate dust accumulation while minimizing the impact to spacecraft mass and power requirements. Preliminary research and development indicates that the Lotus Coating has the potential to be a viable passive tool for mitigating dust on: radiator surfaces, solar array panels, habitation airlock walls, mechanism shields, astronaut EVA suits, and astronaut visors exterior coating.}
}
  16. Park, T.-Y., Lee, J.-J., Kim, J.-H. and Oh, H.-U. (2018). Preliminary Thermal Design and Analysis of Lunar Lander for Night Survival . International Journal of Aerospace Engineering. Source
    BibTeX
    @article{park2018preliminary,
  title = {Preliminary Thermal Design and Analysis of Lunar Lander for Night Survival},
  author = {Park, Tae-Yong and Lee, Jang-Joon and Kim, Jung-Hoon and Oh, Hyun-Ung},
  journal = {International Journal of Aerospace Engineering},
  volume = {2018},
  pages = {1--13},
  publisher = {Wiley},
  year = {2018},
  doi = {10.1155/2018/4236396},
  abstract = {A lunar lander is exposed to extreme lunar thermal environments with a nighttime of 14.75 earth days. Thus, a proper thermal design is an important task to guarantee a successful lunar mission. This paper describes a preliminary thermal design and analysis results of a lunar lander to ensure its survivability during lunar night. The effectiveness of the thermal designs of a lunar lander with various thermal hardwares was numerically investigated according to the landing candidate areas to determine which design is the most feasible for night survival. In addition, we analyzed the mechanical safety of the solder joint of electronic components in accordance with the operating temperature range, because it is an important factor for reducing the system power budget during night survival.}
}
  17. Hickey, G. S., Braun, D., Wen, L.-C. and Eisen, H. J. (1995). Integrated Lightweight Structure and Thermal Insulation for Mars Rover . IEEE International Conference on Robotics and Automation. Source
    BibTeX
    @inproceedings{hickey1995integrated,
  title = {Integrated Lightweight Structure and Thermal Insulation for Mars Rover},
  author = {Hickey, Gregory S. and Braun, David and Wen, Laing-Chi and Eisen, Howard Jay},
  booktitle = {IEEE International Conference on Robotics and Automation},
  volume = {1},
  publisher = {SAE International},
  year = {1995},
  doi = {10.4271/951523},
  abstract = {<div class="htmlview paragraph">An integrated lightweight structure and thermal insulation has been developed for the Mars Rover Warm Electronics Box (WEB). It combines a low thermal conductivity fiber reinforced composite structure with co-cured thermal control surfaces and an ultralightweight hydrophobic solid silica aerogel which eliminates convection and minimizes conduction and radiation. It provides excellent thermal insulation both at low gas pressures and in vacuum; and meets the structural requirements for spacecraft launch loads and for a 60 g impact landing at Mars without damage to the insulation or structure. The integrated design provides for variable conductance to meet the different thermal requirements for the different phases of operation. The system incorporates heat conduction paths to provide thermal control on the ground and during interplanetary cruise, which are then disconnected for Mars operation. The thermal performance of the design was evaluated at prototype level, and as a full structural engineering model and flight system. The thermal conductivity for the structure with insulation was measured at the prototype level to have a value of 0.033 W/mK and 24 °C at 1 atmosphere pressure, and in a10 torr CO<sub>2</sub> simulated Mars environment to be 0.0163 W/mK at 24 °C, 0.0126 W/mK at -27 °C and 0.0089 W/mK at -80 °C.</div>}
}
  18. Bahrami, P., Nesmith, B. and Carpenter, K. (2017). Milli-watt Radioisotope Power to Enable Small, Long-Term Robotic “Probe” Space Exploration . IEEE Aerospace Conference. Source
    BibTeX
    @inproceedings{bahrami2017milli,
  title = {Milli-watt Radioisotope Power to Enable Small, Long-Term Robotic “Probe” Space Exploration},
  author = {Bahrami, Poyan and Nesmith, Bill and Carpenter, Kalind},
  booktitle = {IEEE Aerospace Conference},
  pages = {1-8},
  year = {2017},
  doi = {10.1109/aero.2017.7943899},
  abstract = {Milli-watt Radioisotope Power Systems (RPS) based on Radioisotope Heater Units (RHUs) could be an ideal power source for certain spacecraft that cannot use solar power due to large distances from the sun, or other environmental constraints, and where they enable or significantly enhance the ability of a mission to meet its scientific or operational goals. Various modular, compact RHU-based thermoelectric (TE) generator concepts developed or derived from current NASA Small Business Innovation Research (SBIR) projects satisfying this need have been investigated. These modular, compact and low mass power systems could support small, highly-mobile robotic exploration packages, and could be incorporated into different robotic package concepts, spacecraft or satellites. Current modular RPS design concepts with 40mW, 80mW and 120mW power levels use RHUs and Bi2Te3TE converters. Skutterudites materials could be used in the future if new higher thermal energy output and higher temperature miniature heat sources were developed, for example, using technologies currently in the General Purpose Heat Source (GPHS) used in higher electric power output RTGs. Small (a.k.a., “mice-like”) robotic packages could effectively utilize these RHU-driven power levels to accommodate crawling, climbing, monitoring, taking measurements, and communicating during long-term planetary missions aimed at gathering environmental and geologic data (i.e., over multiple decades). Waste heat from the cold side of the TE converter could also be directed toward the electronics and / or energy storage (e.g. batteries) to keep them within design temperature ranges. In addition to power generation and electronics / battery heating, the RHU / TE configuration could be designed to survive an external 500°C bake out procedure for critical spacecraft sterilization, environmental certification and planetary protection. Analytical studies have been performed to optimize various design configurations for power, mass, volume and robotic mobility. Specific power (mW/kg) and volumetric specific power (mW/cm3) characteristics of various design configurations will be presented and key conceptual design tradeoffs will be discussed. Hot- and cold-side thermal interfaces required to meet power and mass goals and associated design sensitivities will also be discussed. RHU / TE systems must overcome critical design challenges to survive high-g loadings in some robotic applications and we will examine the mass impacts required to satisfy various dynamic loading environments up to 10,000 g's. Power can be generated for a minimum of 30 years or more using plutonium-238 dioxide heat sources (given that Pu-238 has an 87.7 year half-life) with some reduction in power as the heat source naturally degrades.}
}
  19. deGroh, K. K., Banks, B. A., Dever, J. A., Jaworske, D. A., Miller, S. K., Sechkar, E. A. and Panko, S. R. (2008). NASA Glenn Research Center's Materials International Space Station Experiments (MISSE 1-7) . International Symposium on SM/MPAC and SEED Experiments, NASA/TM-2008-215482. Source
    BibTeX
    @inproceedings{degroh2008nasa,
  title = {NASA Glenn Research Center's Materials International Space Station Experiments (MISSE 1-7)},
  author = {deGroh, Kim K. and Banks, Bruce a. and Dever, Joyce A. and Jaworske, Donald A. and Miller, Sharon K. and Sechkar, Edward A. and Panko, Scott R.},
  booktitle = {International Symposium on SM/MPAC and SEED Experiments},
  number = {NASA/TM-2008-215482},
  institution = {NASA},
  address = {Tsukuba},
  year = {2008},
  url = {https://ntrs.nasa.gov/citations/20090005995},
  abstract = {NASA Glenn Research Center (Glenn) has 39 individual materials flight experiments (>540 samples) flown as part of the Materials International Space Station Experiment (MISSE) to address long duration environmental durability of spacecraft materials in low Earth orbit (LEO). MISSE is a series of materials flight experiments consisting of trays, called Passive Experiment Carriers (PECs) that are exposed to the space environment on the exterior of the International Space Station (ISS). MISSE 1-5 have been successfully flown and retrieved and were exposed to the space environment from one to four years. MISSE 6A & 6B were deployed during the STS-123 shuttle mission in March 2008, and MISSE 7A & 7B are being prepared for launch in 2009. The Glenn MISSE experiments address atomic oxygen (AO) effects such as erosion and undercutting of polymers, AO scattering, stress effects on AO erosion, and in-situ AO fluence monitoring. Experiments also address solar radiation effects such as radiation induced polymer shrinkage, stress effects on radiation degradation of polymers, and radiation degradation of indium tin oxide (ITO) coatings and spacesuit fabrics. Additional experiments address combined AO and solar radiation effects on thermal control films, paints and cermet coatings. Experiments with Orion Crew Exploration Vehicle (CEV) seals and UltraFlex solar array materials are also being flown. Several experiments were designed to provide ground-facility to in-space calibration data thus enabling more accurate in-space performance predictions based on ground-laboratory testing. This paper provides an overview of Glenn s MISSE 1-7 flight experiments along with a summary of results from Glenn s MISSE 1 & 2 experiments.}
}
  20. Stein, B. A. (1992). An interim overview of LDEF materials findings . NASA, NASA-TM-107664. Source
    BibTeX
    @techreport{stein1992interim,
  title = {An interim overview of LDEF materials findings},
  author = {Stein, Brad A.},
  number = {NASA-TM-107664},
  institution = {NASA},
  year = {1992},
  url = {https://ntrs.nasa.gov/citations/19930009140},
  abstract = {The flight and retrieval of the National Aeronautics and Space Administration's Long Duration Exposure Facility (LDEF) provided an opportunity for the study of the low-Earth orbit (LEO) environment and long-duration space environmental effects (SEE) on materials that is unparalleled in the history of the U.S. Space Program. The remarkable flight attitude stability of LDEF enables specific analyses of various individual and combined effects of LEO environmental parameters on identical materials on the same space vehicle. This paper provides an overview of the interim LDEF materials findings of the Principal Investigators and the Materials Special Investigation Group. In general, the LDEF data is remarkably consistent; LDEF will provide a 'benchmark' for materials design data bases for satellites in low-Earth orbit. Some materials were identified to be encouragingly resistant to LEO SEE for 5.8 years; other 'space qualified' materials displayed significant environmental degradation. Molecular contamination was widespread; LDEF offers an unprecedented opportunity to provide a unified perspective of unmanned LEO spacecraft contamination mechanisms. New material development requirements for long-term LEO missions have been identified and current ground simulation testing methods/data for new, durable materials concepts can be validated with LDEF results. LDEF findings are already being integrated into the design of Space Station Freedom.}
}
  21. Dursch, H. W., Spear, W. S., Miller, E. A., Bohnhoff-Hlavacek, G. L. and Edelman, J. (1992). Analysis of systems hardware flown on LDEF. Results of the systems special investigation group . Legacy CDMS, 19920022433. Source
    BibTeX
    @techreport{dursch1992analysis,
  title = {Analysis of systems hardware flown on LDEF. Results of the systems special investigation group},
  author = {Dursch, Harry W. and Spear, W. Steve and Miller, Emmett A. and Bohnhoff-Hlavacek, Gail L. and Edelman, Joel},
  number = {19920022433},
  institution = {Legacy CDMS},
  year = {1992},
  url = {https://ntrs.nasa.gov/citations/19920022433},
  abstract = {The Long Duration Exposure Facility (LDEF) was retrieved after spending 69 months in low Earth orbit (LEO). LDEF carried a remarkable variety of mechanical, electrical, thermal, and optical systems, subsystems, and components. The Systems Special Investigation Group (Systems SIG) was formed to investigate the effects of the long duration exposure to LEO on systems related hardware and to coordinate and collate all systems analysis of LDEF hardware. Discussed here is the status of the LDEF Systems SIG investigation through the end of 1991.}
}
  22. Hernandez, Z. Y., Wilcox, S. and Slusser, T. (2023). Designing a Lunar Terrain Vehicle: Thermal Challenges . Thermal and Fluids Analysis Workshop. Source
    BibTeX
    @inproceedings{hernandez2023designing,
  title = {Designing a Lunar Terrain Vehicle: Thermal Challenges},
  author = {Hernandez, Zaida Y. and Wilcox, Samuel and Slusser, Thomas},
  booktitle = {Thermal and Fluids Analysis Workshop},
  organization = {NASA Johnson Space Center},
  year = {2023},
  url = {https://ntrs.nasa.gov/citations/20230002802},
  abstract = {The Lunar Terrain Vehicle (LTV) is an unpressurized rover that will be used to transport astronauts on the South Pole of the moon under the Artemis Program. The National Aeronautics and Space Administration (NASA) will be partnering with an American industry partner to develop the LTV. This paper discusses the thermal challenges faced when designing a crewed lunar vehicle concept and shares a concept that addresses these constraints. One of the most difficult aspects of the design is surviving the cold lunar night on the South Pole which includes permanently shadowed regions (PSR) that can have an extended lunar night duration of 125 hours. Other challenges are critical component temperature limits and programmatic constraints.}
}
  23. Flecht, T., Zormpas, S., Mashhood, M., Muthuvijayan, I., Stober, J. and Cyrus, J. (2025). LUX-Thermal: A Heat Storage, Management and Distribution System on the Lunar Surface . International Conference on Environmental Systems. Source
    BibTeX
    @inproceedings{flecht2025lux,
  title = {LUX-Thermal: A Heat Storage, Management and Distribution System on the Lunar Surface},
  author = {Flecht, Tobias and Zormpas, Sotirios and Mashhood, Muhammad and Muthuvijayan, Indra and Stober, Javier and Cyrus, Julian},
  booktitle = {International Conference on Environmental Systems},
  publisher = {International Conference on Environmental Systems},
  year = {2025},
  doi = {10.32865/2346/102737},
  abstract = {The next decades will see a growing presence on the Lunar surface. Not only will robotic systems continue to explore the Moon, but also assist in establishing a permanent human presence. Various stationary and mobile assets will be needed to enable such a growth in Lunar exploration. Different systems have emerged during the last years that include instruments, vehicles, and habitats [1]. Due to the extreme thermal environment on the Lunar surface [2], the long-term survival constitutes the most important challenge during the designing of these systems as depicted by NASA in their shortfall ranking list [3]. Studies have identified that the electrical power needs of various systems during the night could range between 11 W to 6 kW ranging from rovers to habitats [4]. An attempt to meet these energy demands with a conventional combination of battery and PV systems would result in a significant mass penalty, with the additional disadvantage of not producing sufficient heat to meet the system’s demand during the Lunar night [5]. Dedicated technology is required to allow human-made objects to survive multiple diurnal cycles. Different approaches and technologies have been proposed by multiple entities, which might include radiator covers, variable conductance heat pipes and thermal switches [6, 7]. Lunar Outpost EU is developing LUX-Thermal, an autonomous, self-contained energy generation technology that provides thermal and electrical energy on demand. Its operation is based on the harnessing and storage of heat from multiple inputs during the day. During the Lunar night, it would provide power to customers in proximity, thus enabling their survival without additional internal subsystems. The testing of the first prototype showcasing the basic functionalities concluded in September 2024. The proposed paper might cover the ongoing development of the system, the latest results of the testing campaign, as well as the identified user needs.}
}
  24. Clark, C. J., Smith, J. D., Nick, A. J., Ortega, V. V., Schuler, J. M., Dyas, J. E. and Lahl, J. (2025). Design and Testing of TRL5 IPEx Actuators . IEEE Aerospace Conference, 20250000003. Source
    BibTeX
    @inproceedings{clark2025design,
  title = {Design and Testing of TRL5 IPEx Actuators},
  author = {Clark, Casey J. and Smith, Jonathan Drew and Nick, Andrew J. and Ortega, Victoria V. and Schuler, Jason M. and Dyas, Jeffrey E. and Lahl, John},
  booktitle = {IEEE Aerospace Conference},
  number = {20250000003},
  institution = {NASA},
  address = {Big Sky, MT},
  year = {2025},
  url = {https://ntrs.nasa.gov/citations/20250000003},
  abstract = {NASA is advancing In-Situ Resource Utilization (ISRU) by focusing on missions aimed at establishing sustainable infrastructure on the Moon and Mars. On the Moon, regolith serves as the most abundant resource. To support ISRU objectives, a 30-kg-class robot called ISRU Pilot Excavator (IPEx) is being developed to excavate 10,000 kg of lunar regolith during a future technology demonstration mission. IPEx uses novel excavation tools, called bucket drums, which are hollow cylinders with scoops staggered around the outside. Regolith is collected with the scoops and flows into the drum where it is captured by an internal baffle system. The excavator can then transport the regolith in the drum and reverse the direction of the drum rotation to dispense the regolith out. IPEx uses two sets of bucket drums that dig simultaneously in opposing directions and results in counteracting excavation forces. This combination of bucket drum excavation tools and counteracting excavation forces enables low mass robotic excavators to effectively dig in reduced gravity environments. This is a significant departure from terrestrial excavators that rely on high mass to produce tractive forces to counteract the forces of excavation.

IPEx is made up of several custom actuators that all need to be verified for their intended application. This paper focuses on the initial design, testing and modifications of three actuators: the mobility actuator, the shoulder actuator, and the excavation actuator.

Each actuator was tested under four separate test profiles: ambient motor characterization, hot and cold motor characterization, accelerated life test (ALT), and concept of operations (ConOps) test. The motor characterization tests enabled derivation of torque equations for each actuator to estimate output torque without implementing sensors. The accelerated life tests were successful for each actuator and verified the motors’ ability to survive the mission. For the ConOps test, only the bucket drum actuator performed a complete ConOps without the need to restart the test. Overall, the first series of testing resulted in various failure modes and minor design alterations for each actuator. The test results, failure modes, and design alterations are highlighted within this paper. This paper focuses on discussing the mobility, excavation, and shoulder actuator design, testing principles, and results that qualified it as a TRL 5 system.}
}
  25. Birur, G. C., Bhandari, P., Bame, D., Karlmann, P., Mastropietro, A. J., Liu, Y., Miller, J., Pauken, M. and Lyra, J. (2012). From Concept to Flight: An Active Fluid Loop Based Thermal Control System for Mars Science Laboratory Rover . International Conference on Environmental Systems. Source
    BibTeX
    @inproceedings{birur2012concept,
  title = {From Concept to Flight: An Active Fluid Loop Based Thermal Control System for Mars Science Laboratory Rover},
  author = {Birur, Gajanana C. and Bhandari, Pradeep and Bame, David and Karlmann, Paul and Mastropietro, A. J. and Liu, Yuanming and Miller, Jennifer and Pauken, Michael and Lyra, Jacqueline},
  booktitle = {International Conference on Environmental Systems},
  publisher = {American Institute of Aeronautics and Astronautics},
  year = {2012},
  doi = {10.2514/6.2012-3514}
}
  26. Cappucci, S. and Pauken, M. (2020). Thermal system and environmental testing of the Mars Helicopter . JPL Open Repository. Source
    BibTeX
    @inproceedings{cappucci2020thermal,
  title = {Thermal system and environmental testing of the Mars Helicopter},
  author = {Cappucci, Stefano and Pauken, Michael},
  publisher = {JPL Open Repository},
  year = {2020},
  url = {https://dataverse.jpl.nasa.gov/dataset.xhtml?persistentId=hdl:2014/53049}
}
  27. Bhandari, P., Birur, G., Bame, D., Mastropietro, A. J., Miller, J., Karlmann, P., Liu, Y. and Anderson, K. (2013). Performance of the mechanically pumped fluid loop rover heat rejection system used for thermal control of the Mars Science Laboratory Curiosity Rover on the surface of Mars . International Conference on Environmental Systems. Source
    BibTeX
    @inproceedings{bhandari2013performance,
  title = {Performance of the mechanically pumped fluid loop rover heat rejection system used for thermal control of the Mars Science Laboratory Curiosity Rover on the surface of Mars},
  author = {Bhandari, Pradeep and Birur, Gajanana and Bame, David and Mastropietro, A. J. and Miller, Jennifer and Karlmann, Paul and Liu, Yuanming and Anderson, Kevin},
  booktitle = {International Conference on Environmental Systems},
  publisher = {American Institute of Aeronautics and Astronautics},
  year = {2013},
  doi = {10.2514/6.2013-3323}
}
  28. Birur, G. C. and Bhandari, P. (1998). Mars Pathfinder Active Heat Rejection System: Successful Flight Demonstration of a Mechanically Pumped Cooling Loop . SAE Technical Paper Series. Source
    BibTeX
    @inproceedings{birur1998mars,
  title = {Mars Pathfinder Active Heat Rejection System: Successful Flight Demonstration of a Mechanically Pumped Cooling Loop},
  author = {Birur, Gajanana C. and Bhandari, Pradeep},
  booktitle = {SAE Technical Paper Series},
  volume = {1},
  year = {1998},
  doi = {10.4271/981684},
  abstract = {<div class="htmlview paragraph">One of the new technologies successfully demonstrated on the recent Mars Pathfinder mission was the active Heat Rejection System (HRS). This system consisted of a mechanically pumped cooling loop, which actively controlled the temperatures of the various parts of the spacecraft. A single phase Refrigerant 11 liquid was mechanically circulated through the lander and cruise electronics box heat exchangers. This liquid transferred the excess heat to an external radiator on the cruise stage. This is the first time in unmanned spacecraft history that an active heat rejection system of this type has been used on a long duration spacecraft mission. Pathfinder was launched in December 1996 and landed on the Martian surface on July 4, 1997. The system functioned flawlessly during the entire seven months of flight from Earth to Mars.</div> <div class="htmlview paragraph">A life test set up of the cooling loop was used to verify the life of the system. The life test system was run for over 14, 000 hours before complete examination of the components used in the life test was made. Some of the components used in the system were tested in the life test set up. The results from the life test loop indicate no major issues that would hinder the pumped loop operation for many more years.</div>}
}
  29. Biele, J., Kührt, E., Senshu, H., Sakatani, N., Ogawa, K., Hamm, M., Grott, M., Okada, T. and Arai, T. (2019). Effects of dust layers on thermal emission from airless bodies . Progress in Earth and Planetary Science, 73. Source
    BibTeX
    @article{biele2019effects,
  title = {Effects of dust layers on thermal emission from airless bodies},
  author = {Biele, Jens and Kührt, Ekkehard and Senshu, Hiroki and Sakatani, Naoya and Ogawa, Kazunori and Hamm, Maximilian and Grott, Matthias and Okada, Tatsuaki and Arai, Takehiko},
  journal = {Progress in Earth and Planetary Science},
  volume = {6},
  number = {73},
  pages = {48--48},
  year = {2019},
  doi = {10.1186/s40645-019-0291-0},
  abstract = {We have investigated the influence of thin thermally opaque dust layers on the thermal emission of rocks and regolith and determined the thermal response of these dust-covered surfaces to diurnal insolation cycles. Results are computed for Hayabusa2’s target asteroid (162173) Ryugu, which was observed by thermal infrared instruments on the orbiter and in situ. We show that even a very thin (10..100 μm) fine-grained porous dust layer with thermal inertia of 25 J m−2 K−1 s−1/2 can have a significant influence on surface temperatures and alter the apparent thermal inertia of the underlying material derived under the simplified assumption of a homogenous half space by more than 20%. The masking of the underlying material is complete at about 1 diurnal skin depth, corresponding to ~ 10 mm on Ryugu. Between 0.1 and 1 diurnal skin depths, we find a thermal lag smaller than what would be predicted for a surface consisting of dust only. If a dust cover were present on Ryugu, this should be clearly visible in the data returned by the orbiter’s thermal infrared imager (TIR) and the MASCOT lander’s radiometer (MARA), which observed a single boulder at the landing site. However, this appears not to be the case, and dust seems to play a minor role in the thermal emission from the asteroid.}
}
  30. Bayer, R. (2015). The Thermal Design of the KONTUR-2 Force Feedback Joystick. elib.dlr.de/112503
    BibTeX
    @misc{bayer2015thermal,
  title = {The Thermal Design of the KONTUR-2 Force Feedback Joystick},
  author = {Bayer, Ralph},
  howpublished = {Presentation at the European Space Mechanisms and Tribology Symposium},
  year = {2015},
  url = {https://elib.dlr.de/112503/}
}
  31. Bhandari, P., Karlmann, P., Anderson, K. and Novak, K. (2011). CO2 insulation for thermal control of the Mars Science Laboratory . International Conference on Environmental Systems. Source
    BibTeX
    @inproceedings{bhandari2011insulation,
  title = {CO2 insulation for thermal control of the Mars Science Laboratory},
  author = {Bhandari, Pradeep and Karlmann, Paul and Anderson, Kevin and Novak, Keith},
  booktitle = {International Conference on Environmental Systems},
  publisher = {American Institute of Aeronautics and Astronautics},
  year = {2011},
  doi = {10.2514/6.2011-5119}
}
  32. Novak, K. S., Kempenaar, J. E., Liu, Y., Bhandari, P. and Dudik, B. A. (2012). Mars Science Laboratory Rover System Thermal Test . International Conference on Environmental Systems. Source
    BibTeX
    @inproceedings{novak2012mars,
  title = {Mars Science Laboratory Rover System Thermal Test},
  author = {Novak, Keith S. and Kempenaar, Joshua E. and Liu, Yuanming and Bhandari, Pradeep and Dudik, Brenda A.},
  booktitle = {International Conference on Environmental Systems},
  year = {2012},
  doi = {10.2514/6.2012-3516},
  abstract = {On November 26, 2011, NASA launched a large (900 kg) rover as part of the Mars Science Laboratory (MSL) mission to Mars. The MSL rover is scheduled to land on Mars on August 5, 2012. Prior to launch, the Rover was successfully operated in simulated mission extreme environments during a 16-day long Rover System Thermal Test (STT). This paper describes the MSL Rover STT, test planning, test execution, test results, thermal model correlation and flight predictions. The rover was tested in the JPL 25-Foot Diameter Space Simulator Facility at the Jet Propulsion Laboratory (JPL). The Rover operated in simulated Cruise (vacuum) and Mars Surface environments (8 Torr nitrogen gas) with mission extreme hot and cold boundary conditions. A Xenon lamp solar simulator was used to impose simulated solar loads on the rover during a bounding hot case and during a simulated Mars diurnal test case. All thermal hardware was exercised and performed nominally. The Rover Heat Rejection System, a liquid-phase fluid loop used to transport heat in and out of the electronics boxes inside the rover chassis, performed better than predicted. Steady state and transient data were collected to allow correlation of analytical thermal models. These thermal models were subsequently used to predict rover thermal performance for the MSL Gale Crater landing site. Models predict that critical hardware temperatures will be maintained within allowable flight limits over the entire 669 Sol surface mission.}
}
  33. Mastropietro, A. J., Beatty, J., Kelly, F., Birur, G., Bhandari, P., Pauken, M., Illsley, P., Liu, Y., Bame, D. and Miller, J. (2012). Design and Preliminary Thermal Performance of the Mars Science Laboratory Rover Heat Exchangers . International Conference on Environmental Systems. Source
    BibTeX
    @inproceedings{mastropietro2012design,
  title = {Design and Preliminary Thermal Performance of the Mars Science Laboratory Rover Heat Exchangers},
  author = {Mastropietro, A. J. and Beatty, John and Kelly, Frank and Birur, Gajanana and Bhandari, Pradeep and Pauken, Michael and Illsley, Peter and Liu, Yuanming and Bame, David and Miller, Jennifer},
  booktitle = {International Conference on Environmental Systems},
  address = {San Diego, California},
  year = {2012},
  url = {https://dataverse.jpl.nasa.gov/dataset.xhtml?persistentId=hdl:2014/45114}
}