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The Mars Science Laboratory (MSL) mission to land a large rover on Mars is being prepared for Launch in 2011. A Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) on the rover provides an electrical power of 110 W for use in the rover and the science payload. Unlike the solar arrays, MMRTG provides a constant electrical power during both day and night for all seasons (year around) and latitudes. The MMRTG dissipates about 2000 W of waste heat to produce the desired electrical power. One of the challenges for MSL Rover is the thermal management of the large amount of MMRTG waste heat. During operations on the surface of Mars this heat can be harnessed to maintain the rover and the science payload within their allowable limits during nights and winters without the use of electrical survival heaters. A mechanically pumped fluid loop heat rejection and recovery system (HRS) is used to pick up some of this waste heat and supply it to the rover and payload.

NASA's proposed lunar lander, Altair, will be exposed to vastly different external temperatures following launch till its final destination on the moon. In addition, the heat rejection is lowest at the lowest environmental temperatures (0.5 kW @ 4K) and highest at the highest environmental temperature (4.5 kW @ 215K). This places a severe demand on the radiator design to handle these extreme turn-down requirements. A radiator with digital turn-down capability is currently under study at JPL as a robust means to meet the heat rejection demands and provide freeze protection while minimizing mass and power consumption. Turndown is achieved by independent control of flow branches with isolating latch valves and a gear pump to evacuate the isolated branches. A bench-top test was conducted to characterize the digital radiator concept. Testing focused on the demonstration of proper valve sequencing to achieve turn-down and recharge of flow legs.

Thermal characterization was performed on a vapor compression heat pump using a novel, hybrid two phase loop design. Previous work on this technology has demonstrated its ability to provide passive phase separation and flow control based on capillary action. This provides high quality vapor to the compressor without relying on gravity-based phase separation or other active devices. This paper describes the subsequent work done to characterize evaporator performance under various startup scenarios, tilt angles, and heat loads. The use of a thermal expansion valve as a method to regulate operation was investigated. The effect of past history of use on startup behavior was also studied. Testing under various tilt angles showed evaporator performance to be affected by both adverse and favorable tilts for the given compressor. And depending on the distribution of liquid in the system upon startup, markedly different performance can result for the same system settings and heat loads.

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. A life test set up of the cooling loop was used to verify the life of the system.

Launched on NASA's Aura spacecraft on July 15, 2004, JPL's Tropospheric Emission Spectrometer (TES) has been operating successfully for over three years in space. TES is an infrared high resolution, imaging fourier transform spectrometer with spectral coverage of 3.3 to 15.4 μm to measure and profile essentially all infrared-active molecules present in the Earth's lower atmosphere. It measures the three-dimensional distribution of ozone and its precursors in the lower atmosphere on a global scale. The Aura spacecraft was successfully placed in a sun-synchronous near-circular polar orbit with a mean altitude of 705 km and 98.9 minute orbit period. The observatory is designed for a nominal 5 year mission lifetime. The instrument thermal design features include four temperature zones needed for efficient cryogenic staging to provide cooling at 65 K, 180 K, 230 K and 300 K.

The Moon Mineralogy Mapper (M3) instrument is scheduled for launch in 2008 onboard the Indian Chandrayaan-1 spacecraft. The mission is managed by the Indian Space Research Organization (ISRO) in Bangalore, India and is India's first flight to the Moon. M3 is being developed for NASA by the Jet Propulsion Laboratory under the Discovery Program Office managed by Marshall Space Flight Center. M3 is a state-of-the-art instrument designed to fulfill science and exploratory objectives. Its primary science objective is to characterize and map the lunar surface composition to better understand its geologic evolution. M3's primary exploration goal is to assess and map the Moon mineral resources at high spatial resolution to support future targeted missions. M3 is a cryogenic near infrared imaging spectrometer with spectral coverage of 0.4 to 3.0 μm at 10 nm resolution with high signal to noise ratio, spatial and spectral uniformity.

The Orbiting Carbon Observatory (OCO) instrument is scheduled for launch onboard an Orbital Sciences Corporation LEOStar-2 architecture spacecraft in December 2008. The instrument will collect data to identify CO2 sources and sinks and quantify their seasonal variability. OCO observations will permit the collection of spatially resolved, high resolution spectroscopic observations of CO2 and O2 absorption in reflected sunlight over both continents and oceans. OCO has three bore-sighted, high resolution, grating spectrometers which share a common telescope with similar optics and electronics. A 0.765 μm channel will be used for O2 observations, while the weak and strong CO2 bands will be observed with 1.61 μm and 2.06 μm channels, respectively. The OCO spacecraft circular polar orbit will be sun-synchronous with an inclination of 98.2 degrees, mean altitude of 705 km and 98.9 minute orbit period.

A Small Loop Heat Pipe (SLHP) featuring a wick of only 1.27 cm (0.5 inches) in diameter has been designed for use in spacecraft thermal control. It has several features to accommodate a wide range of environmental conditions in both operating and non-operating states. These include flexible transport lines to facilitate hardware integration, a radiator capable of sustaining over 100 freeze-thaw cycles using ammonia as a working fluid and a structural integrity to sustain acceleration loads up to 30 g. The small LHP has a maximum heat transport capacity of 120 Watts with thermal conductance ranging from 17 to 21 W/°C. The design incorporates heaters on the compensation chamber to modulate the heat transport from full-on to full-stop conditions. A set of start up heaters are attached to the evaporator body using a specially designed fin to assist the LHP in starting up when it is connected to a large thermal mass.

The Trace Gas Analyzer (TGA, Figure 1) is a self-contained, battery-powered mass spectrometer that is designed for use by astronauts during extravehicular activities (EVA) on the International Space Station (ISS). The TGA contains a miniature quadrupole mass spectrometer array (QMSA) that determines the partial pressures of ammonia, hydrazines, nitrogen, and oxygen. The QMSA ionizes the ambient gas mixture and analyzes the component species according to their charge-to-mass ratio. The QMSA and its electronics were designed, developed, and tested by the Jet Propulsion Laboratory (1,2). Oceaneering Space Systems supported JPL in QMSA detector development by performing 3D computer for optimal volumetric integration, and by performing stress and thermal analyses to parameterize environmental performance.

A paraffin-actuated heat switch has been developed for thermal control of the batteries used on the 2003 Mars Exploration Rovers. The heat switch is used to reject heat from the rover battery to a radiator. This paper describes the development test program designed, in part, to measure the thermal conductance of the heat switch in an 8 Torr CO2 environment over the expected operating temperature range of the battery. The switch has a closed conductance of about 0.6 W/°C and an open conductance of 0.019 W/°C. The test program also included measuring the battery temperature profile over a hot case and a cold case Mars diurnal cycle. The test results confirm that the battery will remain well within the upper and lower allowable flight temperatures in both cases.

Future planetary science missions planned for Mars are expected to be more complex and thermally challenging than any of the previous missions. For future rovers, the operational parameters such as landing site latitudes, mission life, distance traversed, and rover thermal energy to be managed will be significantly higher (two to five times) than the previous missions. It is a very challenging problem to provide an effective thermal control for the future rovers using traditional passive thermal control technologies. Recent investigations at the Jet Propulsion Laboratory (JPL) have shown that mechanical pump based fluid loops provide a robust and effective thermal control system needed for these future rovers. Mechanical pump based fluid loop (MPFL) technologies are currently being developed at JPL for use on such rovers. These fluid loops are planned for use during spacecraft cruise from earth to Mars and also on the Martian surface operations.

NASA launched two rovers in June and July of 2003 as a part of the Mars Exploration Rover (MER) project. MER-A (Spirit) landed on Mars in Gusev Crater at 15 degrees South latitude and 175 degrees East longitude on January 4, 2004 (Squyres, et al., Dec. 2004). MER-B (Opportunity) landed on Mars in Terra Meridiani at 2 degrees South latitude and 354 degrees East longitude on January 25, 2004 (Squyres, et al., Aug. 2004). Both rovers have well exceeded their design lifetime (90 Sols) by more than a factor of 5. Spirit and Opportunity are still healthy and continue to execute their roving science missions at the time of this writing. This paper discusses rover flight thermal performance during the surface missions of both vehicles, covering roughly the time from the MER-A landing in late Southern Summer (aereocentric longitude, Ls = 328, Sol 1A) through the Southern Winter solstice (Ls = 90, Sol 255A) to nearly Southern Vernal equinox (Ls = 160, Sol 398A).

Dry heat microbial reduction is an approved method to reduce the microbial bioburden on space-flight hardware prior to launch to meet flight project planetary protection requirements. Microbial bioburden reduction also occurs if a spacecraft enters a planetary atmosphere (e.g., Mars) and is heated by frictional forces. However, without further studies, administrative credit for this reduction cannot be applied. The killing of Bacillus subtilis var. niger spores has been examined and lethality data has been collected by placing spores in a vacuum oven or thermal spore exposure vessels (TSEV) in a constant temperature bath. Using this lethality data, a preliminary mathematical model is being developed that can be used to predict spore killing at different temperatures. This paper will present the lethality data that has been collected at this time and the planned future studies.

Previous studies indicated evidence of opportunistic pathogens in samples obtained during missions to the International Space Station (ISS). This study utilized TaqMan quantitative PCR to determine specific gene abundance in potable and non-potable ISS waters. Probe and primer sets specific to the small subunit rRNA genes were designed and used to elucidate overall bacterial rRNA gene numbers. In addition, primer-probe sets specific for Burkholderia cepacia and Stenotrophomonas maltophilia were optimized and genes of these two opportunistic pathogens quantified in the pre- and post-flight drinking water as well as coolant waters. This Q-PCR approach supports findings of previous culture-based studies however; the culture based studies may have underestimated the microbial burden of ISS drinking water.

Mechanically pumped single-phase fluid loops are well suited for transporting and rejecting large amounts of waste heat from spacecraft electronics and power supplies. While past implementations of such loops on spacecraft have used moderate operating temperatures (less than 60ºC), higher operating temperatures would allow equivalent heat loads to be rejected by smaller and less massive radiators. A high temperature (100 to 150ºC) mechanically pumped fluid loop is currently being investigated at the Jet Propulsion Laboratory (JPL) for use on future Mars missions. This paper details the trade study used to select the high temperature working fluid for the system and the initial development testing of loop components.

As part of the Mars Exploration Rover (MER) project, the National Aeronautics and Space Administration (NASA) launched two rovers in June and July of 2003 and successfully landed both of them on Mars in January of 2004. The cruise stage of each spacecraft (S/C) housed most of the hardware needed to complete the cruise from Earth to Mars, including the propulsion system. Propulsion lines brought hydrazine propellant from tanks under the cruise stage to attitude-control thrusters located on the periphery of the cruise stage. Hydrazine will freeze in the propellant lines if it reaches temperatures below 1.7°C. Thermal control of the propulsion lines was a mission critical function of the thermal subsystem; a frozen propellant line could have resulted in loss of attitude control and complete loss of the S/C.

Mars Exploration Rover (MER) mission launched two spacecraft to Mars in June and July of 2003 and landed two rovers on Mars in January 2004. A Heat Rejection System (HRS) based on a mechanically pumped single-phase liquid cooling system was used to reject heat from electronics to space during the seven months cruise from Earth to Mars. Even though most of this HRS design was similar to the system used on Mars Pathfinder in 1996, several key modifications were made in the MER HRS design. These included the heat exchanger used in removing the heat from electronics, design of venting system used to vent the liquid prior to Mars entry, inclusion of pressure transducer in the HRS, and the spacecraft radiator design. Extensive thermal/fluids modeling and analysis were performed on the MER HRS design to verify the performance and reliability of the system. The HRS design and performance was verified during the spacecraft system thermal vacuum tests.

Hydrogen, gasoline, and mixtures thereof were compared as fuels for lean-burn engines. Hydrogen for the mixed fuels tests was generated by partial oxidation of gasoline. Hydrogen combustion yielded the highest thermal efficiency at any NOx level. Gasoline yielded the second highest thermal efficiency for NOx levels greater than or approximately equal to two gm/mi. For lower NOx levels and high vehicle inertia weights, progressively more hydrogen supplementation was the second most efficient system. For vehicle inertia weights below 5000 lbm (2300 kg), the statutory NOx standard (0.4 gm/mi) could be met with one lb/hr (0.13 g/s) hydrogen supplementation.

Radiators will be used on Mars to reject excess heat from various processes and surfaces and will help temper the climate of any future manned habitats. Radiator performance is a function of the radiator size (area), the emissivity, ε, of the radiator surface, the radiator temperature, local environmental conditions, and the effective sky temperature to which it radiates. The effective sky temperature of Mars is not known. Previous estimates have ranged between 80 K to 170 K. Also, it is not known how dust accumulation and other environmental effects act to change the performance of a radiator as a function of time. The MTERC Experiment is designed to gather data to address these unknowns. This paper will describe the operational theory and the configuration of the MTERC experiment hardware and will discuss results of MTERC performance testing.

The Multi-angle Imaging SpectroRadiometer (MISR) instrument was launched aboard NASA’s Earth Observing System (EOS) Terra spacecraft on December 18, 1999. The overall mission design lifetime for the instrument is 6 years. The EOS Terra spacecraft was placed in a sun-synchronous near-circular polar orbit with an inclination of 98.3 degrees and a mean altitude of 705 km. The overall objective of MISR is to provide a means to study the ecology and climate of Earth through the acquisition of global multiangle imagery on the daylit side of Earth. MISR views the sunlit Earth simultaneously at nine widely spaced angles, collects global images with high spatial detail in four colors at every angle. The images acquired, once calibrated, provide accurate measurements of brightness, contrast and color of reflected sunlight.