Eric Sunada

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...

Exo-C: a Probe-Scale Space Mission to Directly Image and Spectroscopically Characterize Exoplanetary Systems Using an Internal Coronagraph

“Exo-C” is NASA’s first community study of a modest aperture space telescope designed for high contrast observations of exoplanetary systems. The mission will be capable of taking optical spectra of nearby exoplanets in reflected light, discover previously undetected planets, and imaging structure in a large sample of circumstellar disks. It will obtain unique science results on planets down to super-Earth sizes and serve as a technology pathfinder toward an eventual flagship-class mission to find and characterize habitable exoplanets. We present the mission/payload design and highlight steps to reduce mission cost/risk relative to previous mission concepts. At the study conclusion in 2015, NASA will evaluate it for potential development at the end of this decade.

Mars Exploration Rover Surface Mission Flight Thermal Performance

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).

Design and Testing of an Active Heat Rejection Radiator with Digital Turn-Down Capability

NASAs proposed lunar lander, Altair, will be exposed to vastly different external environment temperatures. The challenges to the active thermal control system (ATCS) are compounded by unfavorable transients in the internal waste heat dissipation profile: the lowest heat load occurs in the coldest environment while peak loads coincide with the warmest environment. The current baseline for this fluid is a 50/50 inhibited propylene glycol/water mixture with a freeze temperature around -35 C. While the overall size of the radiators heat rejection area is dictated by the worst case hot scenario, a turn-down feature is necessary to tolerate the worst case cold scenario. A radiator with digital turn-down capability is being designed as a robust means to maintain cabin environment and equipment temperatures while minimizing mass and power consumption. It utilizes active valving to isolate and render ineffective any number of parallel flow tubes which span across the ATCS radiator. Several options were assessed in a trade-study to accommodate flow tube isolation and how to deal with the stagnant fluid that would otherwise remain in the tube. Bread-board environmental tests were conducted for options to drain the fluid from a turned-down leg as well an option to allow a leg to freeze/thaw. Each drain option involved a positive displacement gear pump with different methods of providing a pressure head to feed it. Test results showed that a start-up heater used to generate vapor at the tube inlet held the most promise for tube evacuation. Based on these test results and conclusions drawn from the trade-study, a full-scale radiator design is being worked for the Altair mission profile.

Vapor Compression Hybrid Two‐Phase Loop Technology for Lunar Surface Applications

NASAs vision for Space Exploration that would return humans to the Moon by 2020 in preparation for human explorations of Mars. This requires innovative technical advances. The lunar mission requires a temperature‐lift (heat pump) technology to reject waste heat to hot lunar surface (heat sink) environments during lunar daytime. The lunar outpost and Lunar Surface Access Module (LSAM) to operate anywhere during the hot lunar daytime require a high performance and energy‐efficient, yet reliable refrigeration technology. A vapor compressor‐driven hybrid two‐phase loop was developed for such high temperature‐lift applications. The vapor compression loop used an advanced porous wick evaporator capable of gravity‐insensitive capillary phase separation and excess liquid management to achieve high temperature‐lift, large‐area, isothermal and high heat flux cooling capability and efficient compression. The high temperature lift will allow the lunar surface systems use compact radiators by increased heat rejection...

Fluid Line Evacuation and Freezing Experiments for Digital Radiator Concept

The digital radiator technology is one of three variable heat rejection technologies being investigated for future human-rated NASA missions. The digital radiator concept is based on a mechanically pumped fluid loop with parallel tubes carrying coolant to reject heat from the radiator surface. A series of valves actuate to start and stop fluid flow to di erent combinations of tubes, in order to vary the heat rejection capability of the radiator by a factor of 10 or more. When the flow in a particular leg is stopped, the fluid temperature drops and the fluid can freeze, causing damage or preventing flow from restarting. For this reason, the liquid in a stopped leg must be partially or fully evacuated upon shutdown. One of the challenges facing fluid evacuation from closed tubes arises from the vapor generated during pumping to low pressure, which can cause pump cavitation and incomplete evacuation. Here we present a series of laboratory experiments demonstrating fluid evacuation techniques to overcome these challenges by applying heat and pumping to partial vacuum. Also presented are results from qualitative testing of the freezing characteristics of several di erent candidate fluids, which demonstrate significant di erences in freezing properties, and give insight to the evacuation process.

Thermal Design and Flight Experience of the Mars Exploration Rover Spacecraft Computer-Controlled, Propulsion Line Heaters

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. The MER cruise stage thermal design employed a computer-controlled thermostatic heater system to keep the propellant lines within their allowable flight temperature limits (17°C to 50°C). The MER propellant line thermal design differed from previous propellant line heater designs in that the line heaters were placed only in areas of highest potential heat loss (not along the entire length of the lines) and that computer-controlled thermostats were used instead of mechanical thermostats. Computer-controlled thermostats enabled setpoint flexibility; adjustments to setpoints were made after solar thermal vacuum testing and during flight. This paper covers the design, thermal testing and flight experiences with the computer-controlled thermostats on the propulsion line heaters. Flight experience revealed heater control behavior with propellant loaded into the system and during thruster firings that was not observable during system level testing. Explanations of flight behavior, lessons learned and suggestions for improvement of the propellant line heater design are presented in this paper.

Wax-actuated heat switch for Mars surface applications

Missions to the surface of Mars pose unique thermal control challenges to rover and lander systems. With diurnal temperature changes greater than 100 °C, the presence of a Mars atmosphere, and limited power for night time heating the thermal control engineer is faced with a fundamental problem: how to successfully keep components above their survival or operating temperatures at night while managing higher environmental temperatures and dissipation rates during the day. Payload and avionics elements, among others, must be well insulated to survive night conditions at the risk of overheating during the day. This problem will be magnified in future missions as higher demand on electrical components will result in increased dissipations. One solution is a heat switch that changes thermal conductance to reject excess heat during the day and conserve heat during the night.

Wide Field Infrared Survey Telescope [WFIRST]: telescope design and simulated performance

The Wide Field Infrared Survey Telescope (WFIRST) mission concept was ranked first in new space astrophysics missions by the Astro2010 Decadal Survey, incorporating the Joint Dark Energy Mission payload concept and multiple science white papers. This mission is based on a space telescope at L2 studying exoplanets [via gravitational microlensing], probing dark energy, and surveying the near infrared sky. Since the release of the Astro2010 Decadal Survey, the team has been working with the WFIRST Science Definition Team to refine mission and payload concepts. We present the current interim reference mission point design of the payload, based on the use of a 1.3m unobscured aperture three mirror anastigmat form, with focal imaging and slit-less spectroscopy science channels. We also present the first results of Structural/Thermal/Optical performance modeling of the telescope point design.

Design and Modeling of a Radiator with Digital Turn-Down Capability under Variable Heat Rejection Requirements

Future NASA human-rated near Earth missions impose severe demands on a radiator to be capable of responding to variable heat loads in challenging environments. The capability requires the rejection of peak loads during the warmest environment as well as the rejection of the lowest loads during the coldest environment. Conservation of mass and power becomes a challenge while maintaining the cabin environment and equipment temperatures. A radiator with digital turn-down capability is under development at JPL to manage the transient environments and heat loads. Designed for the worst case hot scenario, the radiator can satisfy any turn-down ratio by isolating sections of the radiator using active valving and varying mass flow rate. Of particular complication is the demand during Low Lunar Orbit (LLO) when environmental temperatures vary the most. Additional concerns involve stagnation of fluid in the closed lines and the time needed to perform turn-down and turn-up operations. A closed loop thermal management system with a representative lunar mission profile and heat loads has been modeled using Thermal Desktop/SINDA Fluint. The fluid selected for the study was a 50/50 propylene glycol/water mixture by weight. Panel size, tube quantity, tube spacing, and tube diameter were traded in order to satisfy the worst case hot scenario and provide options for various other mission phases. Stepping through the mission timeline allows the performance to be predicted as the radiator “reacts” to changes in heat load and environment. Results from the trade study will be used to generate a prototype digital radiator system for testing purposes.