Michael Pauken

Mars Science Laboratory thermal control architecture

The Mars Science Laboratory (MSL 1 ) mission to land a large rover on Mars is being planned for Launch in 2009. As currently conceived, the rover would use a Multimission 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 rovers 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. Mechanically pumped fluid loops 2 are used to both harness the MMRTG heat during surface operations as well as reject it to space during cruise. This paper will describe the basic architecture of the thermal control system, the challenges and the methods used to overcome them by the use of an innovative architecture to maximize the use of heritage from past projects while meeting the requirements for the design.

Development of a Thermal Control Architecture for the Mars Exploration Rovers

In May and June of 2003, the National Aeronautics and Space Administration (NASA) will launch two roving science vehicles on their way to Mars. They will land on Mars in January and February of 2004 and carry out 90‐Sol missions. This paper addresses the thermal design architecture of the Mars Exploration Rover (MER) developed for Mars surface operations. The surface atmosphere temperature on Mars can vary from 0°C in the heat of the day to −100°C in the early morning, prior to sunrise. Heater usage at night must be minimized in order to conserve battery energy. The desire to minimize nighttime heater energy led to a design in which all temperature sensitive electronics and the battery were placed inside a well‐insulated (carbon‐opacified aerogel lined) Warm Electronics Box (WEB). In addition, radioisotope heater units (RHU’s, non‐electric heat sources) were mounted on the battery and electronics inside the WEB. During the Martian day, the electronics inside the WEB dissipate a large amount of energy (ove...

From Concept-to-Flight: An Active Active Fluid Loop Based Thermal Control System for Mars Science Laboratory Rover

The Mars Science Laboratory (MSL) rover, Curiosity, which was launched on November 26, 2011, incorporates a novel active thermal control system to keep the sensitive electronics and science instruments at safe operating and survival temperatures. While the diurnal temperature variations on the Mars surface range from -120 C to +30 C, the sensitive equipment are kept within -40 C to +50 C. The active thermal control system is based on a single-phase mechanically pumped fluid loop (MPFL) system which removes or recovers excess waste heat and manages it to maintain the sensitive equipment inside the rover at safe temperatures. This paper will describe the entire process of developing this active thermal control system for the MSL rover from concept to flight implementation. The development of the rover thermal control system during its architecture, design, fabrication, integration, testing, and launch is described.

Thermal performance evaluation of a small loop heat pipe for space applications

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 a 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 total mass of the small Loop Heat Pipe, including the evaporator body and the radiator, is only 1.4 kg. This paper describes the steady state and transient performance of the small LHP in four different orientations: vertical, horizontal, adverse and reflux. The tests include start up and shut off results for the four orientations at hot and cold case conditions. The results of the test program indicate that the small LHP can successfully transport moderate heat loads for many space related applications.

Design and Preliminary Thermal Performance of the Mars Science Laboratory Rover Heat Exchangers

The challenging range of proposed landing sites for the Mars Science Laboratory Rover requires a rover thermal management system that is capable of keeping temperatures controlled across a wide variety of environmental conditions. On the Martian surface where temperatures can be as cold as -123 degrees Centigrade and as warm as 38 degrees Centigrade, the Rover relies upon a Mechanically Pumped Fluid Loop (MPFL) and external radiators to maintain the temperature of sensitive electronics and science instruments within a -40 degrees Centigrade to 50 degrees Centigrade range. The MPFL also manages significant waste heat generated from the Rover power source, known as the Multi Mission Radioisotope Thermoelectric Generator (MMRTG). The MMRTG produces 110 Watts of electrical power while generating waste heat equivalent to approximately 2000 Watts. Two similar Heat Exchanger (HX) assemblies were designed to both acquire the heat from the MMRTG and radiate waste heat from the onboard electronics to the surrounding Martian environment. Heat acquisition is accomplished on the interior surface of each HX while heat rejection is accomplished on the exterior surface of each HX. Since these two surfaces need to be at very different temperatures in order for the MPFL to perform efficiently, they need to be thermally isolated from one another. The HXs were therefore designed for high in-plane thermal conductivity and extremely low through-thickness thermal conductivity by using aerogel as an insulator inside composite honeycomb sandwich panels. A complex assembly of hand welded and uniquely bent aluminum tubes are bonded onto the HX panels and were specifically designed to be easily mated and demated to the rest of the Rover Heat Recovery and Rejection System (RHRS) in order to ease the integration effort. During the cruise phase to Mars, the HX assemblies serve the additional function of transferring heat from the Rover MPFL to the separate Cruise Stage MPFL so that heat generated deep inside the Rover can be dissipated via the Cruise Stage radiators. Significant fabrication challenges had to be overcome in order to make the HX design a reality. The cruise phase thermal performance of the Rover HXs was verified in the cruise phase system level thermal vacuum test that was performed at JPL in January of 2009. The Rover HXs were modeled in I-DEAS TMG and predictions are compared to actual data from the test.

Insulation Materials Development for Potential Venus Surface Missions

Any potential mission that operates on the surface of Venus requires a unique thermal protection system unlike those found on traditional spacecraft. This paper describes the results of thermal and mechanical testing of insulation materials that may be used to protect a conceptual Lander for Venus surface operations. The thermal control strategy for a conceptual Venus Lander does not rely upon developing new technology; it utilizes the efforts pioneered by the Soviet Venera missions. A new investigation of insulation materials is warranted because many of the practical design and implementation details of the Venera thermal system are unavailable and the present generation of thermal expertise lacks specific knowledge to claim heritage with validity. Thermal conductivity testing of three different classes of insulation materials was made at earth-like and Venus-like conditions. The material classes were porous silica, TiO2 filled aerogel, and silica fiber blankets. Earthlike test conditions were carried out in an oven at ambient pressure air at 470C. Venus-like test conditions were performed in a heated pressure vessel capable of providing a CO2 atmosphere up to 470C and 92 bar pressure. In all material classes, the thermal conductivity under Venus-like conditions increased significantly over the earth-like values. Data from these tests can be used in thermal design to determine the required insulation thickness to protect the Lander for the mission operating life. Mechanical testing of the insulation materials was performed because the atmospheric entry and landing of the vehicle can generate significant deceleration forces. Insulation bonding techniques have been developed for attaching insulation to the exterior surface of a Lander pressure vessel. Data on shear and tensile loading capacity have been collected for various adhesives and bonding techniques. Furthermore, insulation restraint using an exterior skin of stainless steel foil has been developed to enable the insulation material to handle up to 150g’s of deceleration forces. Data from these tests can be used in the mechanical design of the insulation system to ensure it would survive the high body forces of atmospheric entry and landing.

A 5 kWht Lab-Scale Demonstration of a Novel Thermal Energy Storage Concept With Supercritical Fluids

An alternate to the two-tank molten salt thermal energy storage system using supercritical fluids is presented. This technology can enhance the production of electrical power generation and high temperature technologies for commercial use by lowering the cost of energy storage in comparison to current state-of-the-art molten salt energy storage systems. The volumetric energy density of a single-tank supercritical fluid energy storage system is significantly higher than a two-tank molten salt energy storage system due to the high compressibilities in the supercritical state. As a result, the single-tank energy storage system design can lead to almost a factor of ten decrease in fluid costs. This paper presents results from a test performed on a 5 kWht storage tank with a naphthalene energy storage fluid as part of a small preliminary demonstration of the concept of supercritical thermal energy storage. Thermal energy is stored within naphthalene filled tubes designed to handle the temperature (500 °C) and pressure (6.9 MPa or 1000 psia) of the supercritical fluid state. The tubes are enclosed within an insulated shell heat exchanger which serves as the thermal energy storage tank. The storage tank is thermally charged by flowing air at >500 °C over the storage tube bank. Discharging the tank can provide energy to a Rankine cycle (or any other thermodynamic process) over a temperature range from 480 °C to 290 °C. Tests were performed over three stages, starting with a low temperature (200 °C) shake-out test and progressing to a high temperature single cycle test cycling between room temperature and 480 °C and concluding a two-cycle test cycling between 290 °C and 480 °C. The test results indicate a successful demonstration of high energy storage using supercritical fluids.Copyright

Mars Exploration Rover Thermal Test Program Overview

In January 2004, two Mars Exploration Rovers (MER) landed on the surface of Mars to begin their mission as robotic geologists. A year prior to these historic landings, both rovers and the spacecraft that delivered them to Mars, were completing a series of environmental tests in facilities at the Jet Propulsion Laboratory. This paper describes the test program undertaken to validate the thermal design and verify the workmanship integrity of both rovers and the spacecraft. The spacecraft, which contained the rover within the aeroshell, were tested in a 7.5 m diameter thermal vacuum chamber. Thermal balance was performed for the near earth (hot case) condition and for the near Mars (cold case) condition. A solar simulator was used to provide the solar boundary condition on the solar array. IR lamps were used to simulate the solar heat load on the aeroshell for the off-sun attitudes experienced by the spacecraft during its cruise to Mars. Each rover was tested separately in a 3.0 m diameter thermal vacuum chamber over conditions simulating the warmest and coldest expected Mars diurnal temperature cycles. The environmental tests were conducted in a quiescent nitrogen atmosphere at a pressure of 8 to 10 Torr. In addition to thermal balance testing, the science instruments on board the rovers were tested successfully in the extreme environmental conditions anticipated for the mission. A solar simulator was not used in these tests.

Development Testing of a Paraffin-Actuated Heat Switch for Mars Rover Applications

A paraffin-actuated heat switch has been developed for thermal control of the batteries used on the 2003 Mars Exploration Rovers.

Design and flight qualification of a paraffin-actuated heat switch for Mars surface applications

Missions to the surface of Mars pose unique thermal control challenges to rover and lander systems.