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  1. The Lunar Thermal Environment; Issues and Options Ted Swanson, Code 540 LESWG General Meeting June 28, 2007

  2. General Perspective • All heat rejection systems are highly dependent on their environment • Delta T4 relationship, source to sink • Basic equation for radiated heat flux: (very simplified) Qradiated ~ (Aradiator) (e*rad) (Fview factor) (SSteffan-Boltzmann )(Tradiator4 – Tsink4) • Lunar radiative sink highly dependent on location and lunar time of the day • Both latitude and the specific local topography • Possible Tmax ~400K, and Tmin ~ 100K • Thermally, “The Moon is a Harsh Mistress” * * Book by Robert A. Heinlein, Tom Doherty Associates, NY, 1966

  3. Environmental Issues for Lunar/Planetary Landers • Planetary temperature extremes will present real design challenges • Requirement to design for extremes; CTE effects, need for modulation, etc. • Heat rejection during the day, and maintenance of warmth at night • Uneven thermal loading from simultaneous exposure to radically different thermal sinks (e.g., direct sun exposure and shadows in craters or from equipment; sunlight hills and cold regolith at polar locations, etc) • Partial gravity environment • Should yield results scalable to earth based testing, but in-situ demos will be needed for critical, fluid-based thermal subsystems • Unknowns of Lunar/planetary environment • Dust is a major issue for thermal and PV surfaces, movable joints, etc. • Micrometeoroids, actual sink temperatures, combined radiation/thermal effects, etc • Is a Lunar LDEF needed?

  4. Illumination on the Lunar Surface • At sub-solar point, flux is 1414 ± 7 W/m2 normal to the surface. • Because the moon is a diffuse high-emissivity body, this results in surface IR load of ~1314 W/m2 at perihelion (Gilmore, 2002) • Average albedo of ~ 8.4% (visual geometric albedo at 5% phase angle), but significant variations depending on location (Enslow, 1986) • Maria; 7-10% • Highlands; 11-18% • Photometric function; strongly directional and non-Lambertian. This may have thermal implications at polar locations • Color; strongly direction dependent • Shadows; dark but back scatter illuminated

  5. Lunar Surface Thermal Environment The lunar thermal environment is more severe than LEO, GEO, and Mars. Its low albedo results in very high planetary IR. LUNAR DAY 354 Hours with Sun Max. Surface Temp. is ~ 395K LUNAR NIGHT 354 Hours without Sun Min. Surface Temp. is ~ 100K Average temperature of the lunar surface over one Earth month Surface Temperatures Seen at Different Lunar Latitudes through a 28-Day Lunar Rotation* * Eckart, P., Lunar Base Handbook, McGraw-Hill Publishing, New York, New York, 1999, Ch. 5

  6. Mars is Very Different Thermally • The Moon not a good thermal analogue for Mars • Thin CO2 atmosphere, convection/conduction a significant issue; dust storms; lower direct solar radiation (overall about 42% of Earth’s); 3/8th G • Modeling and test simulation issues due to atmospheric convection on Mars

  7. System Issues - Power Sources • Power is critical; thermal is the flip side of power • Alternative Power Sources – each with its own thermal considerations • All PV if at a location with 100% sunlight availability • PV with batteries if modest energy storage needed • PV with regenerative fuel cells • Stirling cycle coupled to regolith • Fission based systems; RTG or reactor • Lunar night issue; significant sustained power over lunar night implies nuclear fission as best long term option. • Fission reactors operate at high temperatures • Materials and working fluid issues for higher temperatures • Degradation from radiation, high temperatures, and dust are major issues • Above H2O limits (~ 550 K), available technologies not nearly as mature, but basic concepts demonstrated • Waste heat potentially useful for process/life support warming during lunar night

  8. Apollo Missions were limited to solar heated periods during the “Lunar Morning”

  9. Apollo 17 Lunar Rover Vehicle Electronics Temperature During Lunar Cycle * Private Communication from Ron Creel, Lead Thermal Engineer for the LRV; “dust degradation of radiators, and the inability to clean them, resulted in exceeding upper temperature limits and having to resort to backup plans.” * Reference: Silk and Creel; Proc. IMechE, Vol. 221, Part G; pgs. 305-309

  10. Previous Safe Operation During Full Lunar Cycles Radioisotope Thermoelectric Generators (RTG’s) Used For Several Years On Apollo Lunar Surface Experiment Packages (ALSEP’s) Radioisotope Heating Units (RHU’s) Used On Russian Lunokhod Robotic Rovers To Survive During Several Solar Eclipse Cycles Isotope Heater

  11. Views toUneven Heat Sinks • At polar locations the sun is nearly horizontal, placing a large thermal flux on the sun-facing side of a facility or rover, but all other sides view cold sinks (regolith or space) • Presence of nearby sunlight hills, which would be “hot”, could have a significant impact on a facility or rover • Adaptive heat management techniques needed for isothermalization • Even at non-polar locations, shadows can be an issue • Example: Apollo 14 Modular Equipment Transporter had rubber tires with a lower temperature limit of – 57 0C. They could not be left in a shadow (even a self induced shadow) for fear of under-temperature damage

  12. Some Thermal Control Options for a Lunar Facility or Rover • For high latitude locations thermal control will be easier; stable, mostly cold environments • Various options for dealing with heat rejection during mid day, at non-high latitude locations • Sunshields to block IR load from surface (Garrison and Nguyen, 2007) • Coupling to regolith (Silk and Creel, 2007) • Keep radiative surfaces very clean (dust shields, operational constraints, cleaning mechanisms) • Outgassing from equipment, waste water from sublimators, and dust are significant contamination sources • Heat pump to kick up heat rejection temperature, or tracking radiators (Swanson et al, 1990) • Sunshield concept example (Garrison and Nguyen, 2007) • A thin-film sunshield was modeled to cover a representative cube as well as some of the surrounding regolith. • The upward-facing surface was modeled with a white paint finish (α/ε = 0.2/0.92) to allow heat rejection, and the underside was VDA (α/ε = 0.08/0.04) to minimize the heat radiated down to the surface. • Results; sunshield heats to approximately 350 K. By blocking the view to space, this warms the radiative sink for the facility/rover and the surface. • The 1414 W/m2 of solar flux is thus converted to a smaller flux of IR energy, which is at the same wavelength that the radiators are trying to reject heat. This would make it very difficult to cool the facility/rover. • Conclusion; sunshields are awkward and unlikely to be a viable solution.

  13. Heat Sinks Buried into the Regolith • The sunlight lunar surface undergoes a monthly cycle from hot to cold. • Lunar dust has a very low thermal conductivity (~ 0.003 W/mK ), and would act as an insulator. • The regolith below the dust layer is nearly isothermal, and could be used as a source/sink for a Stirling cycle. • Lunar regolith has a thermal conductivity of approximately 0.8 W/m/K (Silk, 2007, Heiken, 1991) • Modeling shows that a single heat sink held at 290 K would only be able to sink 15 W. Attached figure shows that the sink only heats the regolith in an area approximately 50 cm wide (Garrison, 2007) Temperature Map of a 3m Diameter by 3m Height Cylinder of Regolith with a 15 W Heat Sink.

  14. Dust Observations by Apollo Astronauts • Lunar dust is a major operational issue: behaves as if it is electrostatically charged • It is difficult to remove • Brushing just spreads it around • Behaves like wet sand • It gets everywhere • No noticeable irritation (long term exposure health issue? • Not insignificant (Vanzani, 1997) Assuming the average particle size and a 20 m2 horizontal radiator, a sortie mission could expect 1.2 impacts on its radiator per 180 days. (Garrison, 2007) Micrometeoroid Issues

  15. Highly accelerated (100s of m/sec) dust has also been detected in this complicated terminator region by Apollo 17’s Lunar Ejecta and Micrometer surface package [Berg et al., 1976] 1 Event per 2 minutes detected by LEAM But not a lot of mass moved: ~0.1 gm cm-2 every billion years [Berg et al, 1976] “Dusty sleet” – An EVA hazard? “Dusty sleet” may be especially relevant at Shackleton The lunar surface is charged positive on the dayside via photoemission and negative on the nightside due to plasma currents On the nightside the low density and elevated electron temperature of the lunar wake makes this region charge even more strongly negative Some fraction of charged surface dust will be levitated and lofted The “dust fountain” model explains the lunar streamers observed by orbiting astronauts Accelerated dust impacts detected by Apollo 17 surface package (from Berg et al., 1976). LEED is specifically designed to measure this dust component with greater sensitivity. Dust Issue: Lunar Surface Charging* * Farrell et al, NESC Dust Workshop, Jan 2007

  16. Some Potential Technology Solutions • Learn more about the environment • Traditional thermal control techniques (MLI, heaters, special coatings, etc) to the extent feasible • Isothermalize and/or isolate with two-phase thermal loops (e.g., Loop Heat Pipe) • Heat pumps to kick up heat rejection temperature • Variable emittance coatings to control effective capability of radiators to reject heat • Specialized techniques and/or coatings to minimize effect of dust

  17. l Q IN State-of-the-Art; Two-Phase Loops Ambient Temperature, Single Evaporator Technology is Operational Transition Active Zone Vapor Line Liquid Line Condenser/Subcooler Flexible designs, unmatched performance

  18. HST/SM-3B; CPL for NICMOS, launch Feb, 2002 GSFC Mission TERRA, 3 CPLs, launched 12/99 GSFC Mission Two-Phase NASA Flight Applications AURA/TES instrument Launched 7/04 JPL Mission GOES N-Q weather Spacecraft; GOES N Launched 05/06 GSFC Missions SWIFT/BAT, 2 LHPs, Launched 11/04 GSFC Mission GLAS, 2 LHPs, 01/2003 launch GSFC Mission

  19. SOA: Two-Phase Design • Proven Benefits; • Design flexibility; packaging, I&T, etc. • Tight temperature control; +/- tenths of 0C • Isothermalization over large area; 10’s of m2 • Diode function; substantial heater power savings • Considerations: • Use simplest design suitable for the full range of requirements; loops may or may not be the best approach

  20. ST-8Thermal Loop Flight Experiment • Concept: Heat transport technology with miniature dual evaporator/dual condenser LHP • PI; Jentung Ku, Code 545 • Applications include; • Isothermalization for robots, rovers, EVA suits* • Reconfigurable S/C • Common design for fleets of S/C • Flight in 2009 • Key Features • Multiple evaporators • Multiple condensers • Thermoelectric coolers (TECs) for tight temperature regulation • Heat load sharing • Analytical model with scaling laws * Option to augment capillary forces with small mechanical pump

  21. Breadboard Vapor Compression, Heat Pump Test Facility – Bldg 4

  22. Low Heat Flux Liquid Flow High Heat Flux Vapor Compression With Spray Cooling • Enhances Feasibility of Space Based, High Power Missions • Radiator size reduction by increasing allowable heat rejection temperature • Compressor provides temperature lift and mechanical pumping • Spray Cooling enables efficient localized heat removal from highly concentrated sources • Technology Challenges • Rotary compressor development • Maximize temperature lift • Zero-G fluid management

  23. ST- 5; Variable Emittance Radiator Coating Flight Data for Electrostatically Switched Concept- SBIR 2 with Sensortex, Inc - • Excellent ground performance • Up to Delta e >0.6 @ 350 VDC • Life testing/environmental testing successful • Flight demonstration on ST-5 in February, 2006; • Flight data (Biter, 2007); • S/C #224; Ground testing De ~ 0.43, Beginning of Mission De ~ 0.37, End of Mission De ~ 0.33 • Note: lower driving voltage on flight unit Flight test article: 90 cm2

  24. “Lotus Effect” Anti-Contamination Coating • Based on observed properties of the Lotus plant • Microstructure effects and surface chemistry prevent accumulation of surface contaminants • GSFC has demonstrated this effect on standard AZ93P radiator paint; • Surface contaminated with lunar stimulant demonstrated less change in absorptivity than non-treated control sample, and was able to fully restore properties after simple cleaning • Testing with EVA glove proposed • Adaptable to bearing surfaces and rotating fittings

  25. Dusty Environmental Effects Particle (DEEP) Chamber • Basic Concept: • Test facility to address both the scientific study of dust and engineering mitigation techniques • Large 4 ft. dia. X 6 ft. high chamber • Designed to be applicable for both Lunar and Mars simulations • Facility under construction in Bldg 4

  26. Some R& D Topics • Two-phase thermophysics at 1/6th G issue: extrapolation from earth data/KC-135 flights reasonable, but proof required prior to commitment. • Basic longer-term thermophysics experiments on moon highly desirable • Heat pump technologies; high lift, matched to load, etc • Techniques for dust mitigation and control • Coating regolith (a film or polymer?) for low a and low dust generation • Shielding from spray and ground operations • Specialty coatings to resist dust accumulation (e.g., “lotus coatings”) • Dust cleaning techniques (electrostatic, :air guns”, etc.) • Life testing on the moon is the best approach • Variable emittance coatings for the radiators to mitigate day/night swing in heat rejection capability

  27. R&D Topics - Continued • High temperature radiators for fission power • High e/low a, long life coating at high temperature not demonstrated; potentially big issue • Demo on moon best to include all environmental effects • Life test needed • Significant construction effort • New technology needed, and will require demo priorto full scale deployment • Long life pump to circulate distribution fluid • Materials compatibility life tests • Heat pipes/containment metal • Connection of heat pipes to APG fins • Fabrication of high k materials (Ca, Al) from regolith • Level of threat from micro-meteoroids; not totally clear at present

  28. Conclusions • Thermal control on the moon is possible, but not easy • Stark thermal contrasts between night and day; must design for both extremes • Design highly dependent on Lunar latitude/local topography • Some R&D is likely to be needed • Rovers will be particularly challenging due to continuously variable sinks • Extrapolations of existing technology will be required; but no major showstoppers foreseen • Demonstrations of such technology will be required prior to operational deployment • Subscale demonstrations on the moon will very likely be required • Dust is potentially a big issue which may very well impact thermal design and operations • Hence, dust avoidance, mitigation, and control/cleaning technologies are highly desireable

  29. References • Gilmore, D., Spacecraft Thermal Control Handbook, Volume I: Fundamental Technologies, Aerospace Press, El Segundo, CA, • 2002, pp. 53-56. • Enslow, Guide to Observing the Moon, British Astronomical Association, 1986 • Eckart, P., Lunar Base Handbook, McGraw-Hill Publishing, New York, NY, 1999, pp. 107-140 • Silk, E., and Creel, R., “Technology Development for Lunar Thermal Applications and the Next Generation of Space Exploration,” Proc. IMechE Vol. 221 Part G: J. Aerospace Engineering, pp. 305-309 • Heiken, G.H., Vaniman, D.T., French, B.M., Lunar Source Book: A User’s Guide to the Moon, Cambridge University Press, New York, NY, 1991, pp. 132-134. • Garrison, M. and Nguyen, D., “Thermal Considerations for Designing the Next Lunar Lander”, Proceedings of the 2007 Space Technology Applications International Forum, University of New Mexico, Albuquerque, MN, February, 2007 • Vanzani, V., et al., “Micrometeoroid Impacts on the Lunar Surface,” in proceedings of 28th Annual Lunar and Planetary Science Conference, 1997, pp. 481-486. • Farrell, W., Stubbs, T.J. , Vondrak, R.R., “The Lunar Dusty Plasma Environment,” NESC Dust Workshop, Jan 2007 • Swanson, T et al., “Low Temperature Thermal Control for a Lunar Base,” SAE Paper 901242, 1990 Transactions of the SAE international, Washington, D.C • Westheimer, D. T., Tuan, G. C., “Active Thermal Control System Considerations for the Next Generation of Human Rated Space Vehicles,” in the proceedings of 43rd AIAA Aerospace Sciences Meeting and Exhibit, AIAA, Reston, VA, 2005, pp.1-5. • Biter, W. and Oh, S., “Performance Results of the ESR from the Space Technology 5 Satellite”, Proceedings of the 2007 Space Technology Applications International Forum, University of New Mexico, Albuquerque, MN, February, 2007

  30. Backup

  31. Beta 0° Beta 90° Example: Lunar Reconnaissance Mission • Late 2008 launch on Delta-II to Earth “Parking Orbit” • Lunar Injection burn to 50+/-20 km circular polar orbit • Thus, strong thermal link to lunar surface • Minimum 1 year in science configuration (single ground re-trace / year) • Potential extended mission in 30x216 km elliptical “relay” orbit (to end-of-life ~5yrs) BOTTOM LINE THERMALLY: Challenging! Minimum earth viewing during transport; worst case hot at low Betas with low altitude; worst case cold is high Betas (weak function on altitude). Earth eclipse issue.

  32. where: q”IR = IR flux from Lunar surface C1 = Peak flux at subsolar point C2 = Minimum flux emitted from shaded Lunar surface  = Beta angle  = Angle from subsolar point q”IR = [(C1-C2)*cos()*cos()] + C2 Lunar Thermal Environment Lunar Orbit Environment Parameters Lunar IR Emission as a function of Beta Angle

  33. Thermal Design Overview Isothermal Panel with embedded heat pipes • Orbiter features Zenith Radiators which have minimum exposure to the lunar environments Zenith Avionics + Battery Radiator SAS HGAS Prop Module (within LRO) Instrument Anti-Sun side

  34. Embedded CCHP HP Layout in Thermal Model- Preliminary - Dual Bore CCHP U-Shaped CCHP G for 0.5” ATS type HP to FS = 1.15 W/inC G for Dual Bore HP to FS = 4.0 W/inC

  35. LROC 50 46 42 CRaTER CRaTER 38 34 30 LOLA LEND LOLA LEND CRaTER 26 DIVINER LOLA 22 18 14 10 LAMP Temperature (oC) 10 6.5 3.0 -0.5 -4.0 -7.5 -11 -14.5 --18 -21.5 -25 Time (Hours) LRO Thermal Design Challenges Extreme temperature swing during operational cases as spacecraft travels from sun side to dark side of the moon Temperature Gradients on the Optical Bench Nadir Facesheet Cold Case Hot Case

  36. Thermal Load from a Hot Surface • Radiator View to Surface • Any angle that allows the radiator to view the surface will overwhelm its heat rejection capability, since the surface is emitting ~1314 W/m2 at the same wavelengths that the radiator is emitting. • Views toUneven Heat Sinks • At polar locations the sun is nearly horizontal, placing a large thermal flux on the sun-facing side but all other sides view to a cold surfaces (regolith/space) • Presence of nearby sunlight hills could have significant impact • Adaptive heat management techniques needed • Even at non-polar locations, shadows can be an issue Ambient Load Change as the Radiator is Tilted Off the Horizontal Orientation, for a subsolar location. • Example: Apollo 14 Modular Equipment Transporter had rubber tires with a lower temperature limit of – 57 0C. They could not be left in a shadow (even a self induced shadow) for fear of under-temperature damage