Search Results

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.

Polymers are one of the major constituents in electrical components. A study investigating pre-combustion off-gassing and particle release by polymeric materials over a range of temperatures can provide an understanding of thermal degradation prior to failure which may result in a fire hazard. In this work, we report simultaneous measurements of pre-combustion vapor and particle release by heated polymeric materials. The polymer materials considered for the current study are silicone and Kapton. The polymer samples were heated over the range 20 to 400°C. Response to vapor releases were recorded using the JPL Electronic Nose (ENose) and Industrial Scientific's ITX gas monitor configured to detect hydrogen chloride (HCl), carbon monoxide (CO) and hydrogen cyanide (HCN). Particle release was monitored using a TSI P-TRAK particle counter.

Work is underway to deliver an instrument for analysis of the atmosphere aboard the International Space Station. The Vehicle Cabin Atmosphere Monitor (VCAM) is based on a low-mass, low-power miniature preconcentrator gas chromatograph/mass spectrometer (PCGC/MS) capable of providing sub-ppm measurements of volatile constituents in a space vehicle or outpost. VCAM is designed to operate autonomously, maintenance-free, once per day, with its own carrier and calibration gas supplies sufficient for a one-year lifetime. VCAM performance is sufficient to detect and identify 90% of the target compounds specified at their 180-day Spacecraft Maximum Allowable Concentration (SMAC) levels. The flight units will be delivered in mid-2008 and be operated in the ISS EXPRESS rack.

The Internal Active Thermal Control System (IATCS) aboard the International Space Station (ISS) contains an aqueous, alkaline fluid (pH 9.5±0.5) that aids in maintaining a habitable environment for the crew. Because microbes have significant potential to cause disease, adverse effects on astronaut health, and microbe-induced corrosion, the presence of both bacteria and viruses within IATCS fluids is of concern. This study sought to detect and identify viral populations in IATCS samples obtained from the Kennedy Space Center as a first step towards characterizing and understanding potential risks associated with them. Samples were concentrated and viral nucleic acids (NA) extracted providing solutions containing 8.87-22.67 μg NA per mL of heat transfer fluid. After further amplification viral DNA and cDNA were then pooled, fluorescently labeled, and hybridized onto a Combimatrix panvira 12K microarray containing probes for ∼1,000 known human viruses.

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.

Progress on the delivery of the Vehicle Cabin Atmosphere Monitor (VCAM) is reported. VCAM is an autonomous trace-species detector to be used aboard the International Space Station (ISS) for atmospheric analysis. The instrument is based on a low-mass, low-power miniature preconcentrator, gas chromatograph, and Paul ion trap mass spectrometer (PCGC/MS) capable of measuring volatile constituents in a space vehicle or planetary outpost at sub-ppm levels. VCAM detects and quantifies 40 target compounds at their 180-day Spacecraft Maximum Allowable Concentration (SMAC) levels. It is designed to operate autonomously, maintenance-free, with a self-contained carrier and calibration gas supplies sufficient for a one-year lifetime. Two flight units will be delivered for operation in the ISS EXPRESS rack.

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.

Simultaneous measurements were made for particle releases and off-gassing products produced by heating electrical wires. The wire samples in these experiments were heated to selected temperatures in a heating chamber and responses to vapor releases were recorded by the JPL Electronic Nose (ENose) and an Industrial Scientific ITX gas-monitor; particles released were detected by a TSI P-Trak particle counter. The temperature range considered for the experiment is room temperature (24−26°C) to 500 °C. The results were analyzed by overlapping responses from the ENose, ITX gas sensors and P-Trak, to understand the events (particle release/off-gassing) and sequence of events as a function of temperature and to determine qualitatively whether ENose may be used to detect pre-combustion event markers.

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.

A technology demonstration propylene Loop Heat Pipe (LHP) has been tested extensively in support of the implementation of this two-phase thermal control technology on NASA’s Earth Observing System (EOS) Tropospheric Emission Spectrometer (TES) instrument. This cryogenic instrument is being developed at the Jet Propulsion Laboratory (JPL) for NASA. This paper reports on the transient characterization testing results showing low frequency temperature oscillations. Steady state performance and model correlation results can be found elsewhere. Results for transient startup and shutdown are also reported elsewhere. In space applications, when LHPs are used for thermal control, the power dissipation components are typically of large mass and may operate over a wide range of power dissipations; there is a concern that the LHP evaporator may see temperature oscillations at low powers and over some temperature range.

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.

The Mars Exploration Rover (MER) flight system uses mechanical, paraffin-actuated heat switches as part of its secondary battery thermal control system. This paper describes the design, flight qualification, and performance of the heat switch. Although based on previous designs by Starsys Research Corporation1,2, the MER mission requirements have necessitated new design features and an extensive qualification program. The design utilizes the work created by the expansion of a paraffin wax by bringing into contact two aluminum surfaces, thereby forming a heat conduction path. As the paraffin freezes and contracts, compression springs separate the surfaces to remove the conduction path. The flight qualification program involved extensive thermal performance, structural, and life testing.

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.

Phoenix, NASA's first Mars Scouts mission, will be launched in 2007 and will soft-land inside the Martian Arctic Circle, between north 65° and 72° North latitude, in 2008 to study the water history and to search for habitable zones. Similar to the IDD (Instrument Deployment Device) on the Mars Exploration Rovers (MER), Phoenix has a Robotic Arm (RA) which is equipped with a scoop to dig into the icy soil and to deliver the soil samples to instruments for scientific observations and measurements. As with MER, the actuators and the bearings of the Phoenix RA in a non-operating condition can survive the cold Martian night without any electrical power or any thermal insulation. The RA actuators have a minimum operating allowable flight temperature (AFT) limit of -55°C, so, warm-up heaters are required to bring the temperatures of all the RA actuators above the operating AFT limit prior to early morning operation.

The Mars Science Laboratory (MSL1) mission to land a large rover on Mars is being planned for Launch in 2009. As currently conceived, the rover would use a Multi-mission Radioisotope Thermoelectric Generator (MMRTG) to generate about 110 W of electrical power for use in the rover and the science payload. Usage of an MMRTG allows for a large amount of nearly constant electrical power to be generated day and night for all seasons (year around) and latitudes. This offers a large advantage over solar arrays. The MMRTG by its nature dissipates about 2000 W of waste heat. The basic architecture of the thermal system utilizes this waste heat on the surface of Mars to maintain the rover's temperatures within their limits under all conditions. In addition, during cruise, this waste heat needs to be dissipated safely to protect sensitive components in the spacecraft and the rover.

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).

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.

Various candidates for nonpetroleum electric and hybrid vehicle (EHV) subsystems have been evaluated as part of the Advanced Vehicle (AV) Assessment at the Jet Propulsion Laboratory. The subsystems include battery and power-peaking energy storage, heat engine and fuel cell energy conversion devices, motor/controller subsystems, transmissions, and vehicle subsystem (structure and body) technologies. The primary objective of this effort, was to project the mature capabilities of the various components in the 1990’s for application in the systems evaluations in the next phase of this activity. This paper presents the basic characteristics of the subsystems and compares their capabilities with projected AV subsystem requirements.

Candidate Aeroshell Test models composed of a quasi-isotropic Carbon/Carbon(C/C) front face sheet (F/S), eggcrate core, C/C back F/S, Carbon Aerogel insulation, C/C radiation shield and the C/C close-out were constructed based on the analytical temperature predictions presented in Part One of this work[1]. The analytical results obtained for a simulated Mars entry of a 2.9 meter diameter cone shaped Carbon-Carbon Aeroshell demonstrated the feasibility of the design. These results showed that the maximum temperature the front F/S reached during the decent was 1752 °C with the resulting rear temperature reaching 326 °C in the thermal model. Part Two of this work documents the thermal modeling and correlation for the Mars Aeroshell test sample and fixture. A finite difference, SINDA/G, thermal math model of the test fixture and sample was generated and correlated to data from an arc jet test conducted at the NASA Ames Research Center's interactive heating facility.

This paper describes a novel thermal control system for future Mars landers and rovers designed to keep battery temperatures within the −10 °C to +25 °C temperature range. To keep the battery temperatures above the lower limit, the system uses: 1) a phase change material (PCM) thermal storage module to store and release heat and 2) a loop heat pipe (LHP) to transfer heat from a set of Radioisotope Heater Units (RHUs) to the battery. To keep the battery temperature below the upper limit, a thermal control valve in the LHP opens to redirect the working fluid to an external radiator where excess heat is dumped to the atmosphere. The PCM thermal storage module was designed and fabricated using dodecane paraffin wax (melting point, − 9.6 °C) as the phase change material. A miniature ammonia loop heat pipe with two condensers and an integrated thermal control valve was designed and fabricated for use with the PCM thermal storage unit.