Theodore D. Swanson

NASA thermal control technologies for robotic spacecraft

Technology development is inevitably a dynamic process in search of an elusive goal. It is never truly clear whether the need for a particular technology drives its development, or the existence of a new capability initiates new applications. Technology development for the thermal control of spacecraft presents an excellent example of this situation. Nevertheless, it is imperative to have a basic plan to help guide and focus such an effort. Although this plan will be a living document that changes with time to reflect technological developments, perceived needs, perceived opportunities, and the ever-changing funding environment, it is still a very useful tool. This presentation summarizes the current efforts at National Aeronautics and Space Administration (NASA)/Goddard and NASA/JPL to develop new thermal control technology for future robotic NASA missions.

Large, Switchable Electrochromism in the Visible Through Far‐Infrared in Conducting Polymer Devices

Advanced materials with large and dynamic variation in thermal properties, sought for urgent defense and space applications, have heretofore been elusive. Conducting polymers (CPs) have shown some intrinsic variation of mid- to far-infrared (IR) signature in this respect, but the practical utilization of this has remained elusive. We report herein the first significant IR electrochromism in any material, to our knowledge, in the 0.4 through 45 μm region. This is seen in practical CP devices in the form of thin (<0.5 mm), flexible, entirely solid-state, variable area (1 cm2 to 1 m2) flat panels. Typical properties include: very high reflectance variation; switching times <2 s; cyclabilities of 105 cycles; emittance variation from 0.32 to 0.79; solar absorptance variation from 0.39 to 0.79; operating temperatures of –35 to +85 °C; durability against γ-radiation to 7.6 Mrad, vacuum to 10–6 torr, and simulated solar wind (e.g., 6.5 × 1016 e/cm2 @ 10 keV).

Miniature Loop Heat Pipes for Electronic Cooling

Thermal management of modern electronics has become a problem of significant interest due to the demand for power and reduction in packaging size. Requirements of next-generation microprocessors in terms of power dissipation and heat flux will certainly outgrow the capability of today’s thermal control technology. LHPs, like conventional heat pipes, are capillary pumped heat transport devices. They contain no mechanical moving part to wear out or require electrical power to operate. But unlike heat pipes, LHPs possess much higher heat transport capabilities enabling them to transport large amounts of heat over long distances in small flexible lines for heat rejection. In fact, a miniature ammonia LHP developed for a NASA space program is capable of transporting 60W over a distance of 1 meter in 1/16”O.D. stainless steel tubing. Therefore, miniature LHPs using water as the working fluid are excellent candidates to replace heat pipes as heat transports in electronic cooling systems. However, a number of operational issues regarding system performance, cost, and integration/packaging must be resolved before water LHPs can become a viable option for commercial electronics.Copyright

Development of the variable emittance thermal suite for the space technology 5 microsatellite

The advent of very small satellites, such as nano and microsatellites, logically leads to a requirement for smaller thermal control subsystems. In addition, the thermal control needs of the smaller spacecraft/instrument may well be different from more traditional situations. For example, power for traditional heaters may be very limited or unavailable, mass allocations may be severely limited, and fleets of nano/microsatellites will require a generic thermal design as the cost of unique designs will be prohibitive. Some applications may require significantly increased power levels while others may require extremely low heat loss for extended periods. Small spacecraft will have low thermal capacitance thus subjecting them to large temperature swings when either the heat generation rate changes or the thermal sink temperature changes. This situation, combined with the need for tighter temperature control, will present a challenging situation during transient operation. The use of “off-the-shelf” commercial spacecraft buses for science instruments will also present challenges. Older thermal technology, such as heaters, thermostats, and heat pipes, will almost certainly not be sufficient to meet the requirements of these new spacecraft/instruments. They are generally too heavy, not scalable to very small sizes, and may consume inordinate amounts of power. Hence there is a strong driver to develop new technology to meet these emerging needs. Variable emittance coatings offer an exciting alternative to traditional control methodologies and are one of the technologies that will be flown on Space Technology 5, a mission of three microsatellites designed to validate “enabling” technologies. Several studies have identified variable emittance coatings as applicable to a wide range of spacecraft, and to potentially offer substantial savings in mass and/or power over traditional approaches. This paper discusses the development of the variable emittance thermal suite for ST-5. More specifically, it provides a description of and the infusion and validation plans for the variable emittance coatings.

Loop heat pipes and capillary pumped loops-an applications perspective

Capillary pumped loops (CPLs) and loop heat pipes (LHPs) are versatile two-phase heat transfer devices which have recently gained increasing acceptance in space applications. Both systems work based on the same principles and have very similar designs. Nevertheless, some differences exist in the construction of the evaporator and the hydro-accumulator, and these differences lead to very distinct operating characteristics for each loop. This paper presents comparisons of the two loops from an applications perspective, and addresses their impact on spacecraft design, integration, and test. Some technical challenges and issues for both loops are also addressed.

Low-temperature thermal control for a lunar base

The generic problem of rejecting low- to moderate-temperature heat from space facilities located in a hot thermal sink environment is studied, and the example of a lunar base located near the equator is described. The effective thermal sink temperature is often above or near nominal room temperature. A three heat pump assisted thermal bus concept appears to be the most viable as they are the least sensitive to environmental conditions. Weight estimates are also developed for each of the five thermal control concepts studied: (1) 149kg/kW for a central thermal loop with unitary heat pumps; (2) 133 kg/kW for a conventional bus connected to large, central heat pumps at the radiator; (3) 134 kg/kW for a central, dual loop heat pump concept; (4) 95 kg/kW for the selective field-of-view radiator; and (5) 126 kg/kW for the regolith concept.

Design optimization of a hydrogen advanced loop heat pipe for space-based IR sensor and detector cryocooling

Next generation space infrared sensing instruments and spacecraft will require drastic improvements in cryocooling technology in terms of performance and ease of integration. Projected requirements for cryogenic thermal control systems are: high duty cycle heat loads, low parasitic heat penalty, long transport distances, highly flexible transport lines, and lower cooling temperatures. In the current state of cryocooling transport technology, cryogenic Loop Heat Pipes (CLHPs) are at the forefront of intensive research and development. CLHPs are capable of dispersing heat quickly from an IR heat source and transporting it to remotely located cryocoolers via small and flexible transport lines. Circulation of working fluid in a CLHP is accomplished entirely by capillary action developed in fine pore wicks of the system capillary pumps. Thus they contain no mechanical moving parts to wear out or to introduce unwanted vibrations to the spacecraft. A recently developed CLHP using Hydrogen as the working fluid performed extremely well in the temperature range of 20-30K under the most severe operating conditions. However, it was not optimized for spacecraft applications due to cost and schedule constraints of the initial research phase. Design optimization of the Hydrogen Advanced Loop Heat Pipe is the main objective of the follow-on research. Chief among the system improvements is the weight and volume reduction of the loop components.

Electrostatic radiator for satellite temperature control

An objective for advanced satellites and spacecraft is to continually reduce both their size and mass. This reduction can place severe constraints on the thermal control systems. In addition, mission requirements also may dictate the need to alter the spacecraft energy profile during the course of the mission. To facilitate these advances, significant research has been conducted to develop variable emittance coatings and devices to provide active spacecraft thermal control. Several of these technologies have matured to a level where space based testing is feasible and will be performed as part of the ST5 new millennium spacecraft mission. One of these technologies utilizes electrostatic hold-down of a high emissivity composite film to actively control spacecraft skin temperature. This electrostatic radiator (ESR) device functions as a thermal switch and changes the mode of heat transfer between the spacecraft skin and the radiator film from conduction to radiation. This device has demonstrated large changes in effective emissivity in laboratory cold thermal vacuum testing. In this paper, the theory of operation of the ESR and the construction of the device as configured for demonstration on the ST5 mission is presented. The role of the ST5 mission in the maturation process for VEC technologies is discussed. This paper also describes test results for the ESR through flight qualification testing for the ST5 mission. Finally, anticipated operational characteristics and mission reliability estimates are provided.

Parametric study of variable emissivity radiator surfaces

The goal of spacecraft thermal design is to accommodate a high function satellite in a low weight and real estate package. The extreme environments that the satellite is exposed during its orbit are handled using passive and active control techniques. Heritage passive heat rejection designs are sized for the hot conditions and augmented for the cold end with heaters. The active heat rejection designs to date are heavy, expensive and/or complex. Incorporating an active radiator into the design that is lighter, cheaper and more simplistic will allow designers to meet the previously stated goal of thermal spacecraft design. Varying the radiator’s surface properties without changing the radiating area (as with VCHP), or changing the radiators’ views (traditional louvers) is the objective of the variable emissivity (vary-e) radiator technologies. A parametric evaluation of the thermal performance of three such technologies is documented in this paper. Comparisons of the Micro-Electromechanical Systems (MEMS), ...

Thermal Performance of Capillary Pumped Loops Onboard Terra Spacecraft

The Terra spacecraft is the flagship of NASAs Earth Science Enterprise. It provides global data on the state of atmosphere, land and oceans, as well as their interactions with solar radiation and one another. Three Terra instruments utilize Capillary Pumped Heat Transport System (CPHTS) for temperature control: Each CPHTS, consisting of two capillary pumped loops (CPLs) and several heat pipes and electrical heaters, is designed for instrument heat loads ranging from 25W to 264W. The working fluid is ammonia. Since the launch of the Terra spacecraft, each CPHTS has been providing a stable interface temperature specified by the instrument under all modes of spacecraft and instrument operations. The ability to change the CPHTS operating temperature upon demand while in service has also extended the useful life of one instrument. This paper describes the design and on-orbit performance of the CPHTS thermal systems.