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NASA NTRS2000

Managing Risk for Thermal Vacuum Testing of the International Space Station Radiators

The International Space Station (ISS) is designed with large deployable radiator panels that are used to reject waste heat from the habitation modules. Qualification testing of the Heat Rejection System (HRS) radiators was performed using qualification hardware only. As a result of those tests, over 30 design changes were made to the actual flight hardware. Consequently, a system level test of the flight hardware was needed to validate its performance in the final configuration. A full thermal vacuum test was performed on the flight hardware in order to demonstrate its ability to deploy on-orbit. Since there is an increased level of risk associated with testing flight hardware, because of cost and schedule limitations, special risk mitigation procedures were developed and implemented for the test program, This paper introduces the Continuous Risk Management process that was utilized for the ISS HRS test program. Testing was performed in the Space Power Facility at the NASA Glenn Research Center, Plum Brook Station located in Sandusky, Ohio. The radiator system was installed in the 100-foot diameter by 122-foot tall vacuum chamber on a special deployment track. Radiator deployments were performed at several thermal conditions similar to those expected on-orbit using both the primary deployment mechanism and the back-up deployment mechanism. The tests were highly successful and were completed without incident.

Carek, Jerry A.↗

NASA NTRS2004

Mars Radiator Characterization Experimental Program

Radiators are an enabling technology for the human exploration and development of the moon and Mars. As standard components of the heat rejection subsystem of space vehicles, radiators are used to reject waste heat to space and/or a planetary environment. They are typically large components of the thermal control system for a space vehicle or human habitation facility, and in some cases safety factors are used to oversize them when the operating environment cannot be fully characterized. Over-sizing can impose significant weight and size penalties that might be prohibitive for future missions. Radiator performance depends on the size of the radiator surface, its emittance and absorptance, the radiator temperature, the effective sky temperature surrounding the radiator, solar radiation and atmospheric irradiation levels, convection to or from the atmosphere (on Mars), and other conditions that could affect the nature of the radiator surface, such as dust accumulation. Most particularly, dust is expected to be a major contributor to the local environmental conditions on either the lunar or Martian surface. This conclusion regarding Mars is supported by measurements of dust accumulation on the Mars Sojourner Rover solar array during the Pathfinder mission. This Final Report describes a study of the effect of Martian dust accumulation on radiator performance. It is comprised of quantitative measurements of effective emittance for a range of dust accumulation levels on surfaces of known emittance under clean conditions. The test radiator coatings were Z-93P, NS-43G, and Silver Teflon (10 mil) film. The Martian dust simulant was Carbondale Red Clay. Results were obtained under vacuum conditions sufficient to reduce convection effects virtually to zero. The experiments required the development of a calorimetric apparatus that allows simultaneous measurements of the effective emittance for all the coatings at each set of experimental conditions. A method of adding dust to multiple radiator coupons was developed and shown to be capable of depositing dust on the surfaces with acceptable uniformity. In these experiments, the dust layer accumulates under earth gravity and in the presence of an earth atmosphere. An invention disclosure for the dust deposition apparatus is being filed through NASA and University of Houston.

Witte, Larry C.↗

NASA NTRS2000

Sensor Needs for Advanced Life Support

Sensors and feedback systems are critical to life support flight systems and life support systems research. New sensor capabilities can allow for new system architectures to be considered, and can facilitate dramatic improvements in system performance. This paper will describe three opportunities for biosensor researchers to develop sensors that will enable life support system improvements. The first opportunity relates to measuring physical, chemical, and biological parameters in the Space Station Water Processing System. Measuring pH, iodine, total organic carbon, microbiological activity, total dissolved solids, or conductivity with a safe, effective, stable, reliable microsensor could benefit the water processing system considerably. Of special interest is a sensor which can monitor biological contamination rapidly. The second opportunity relates to sensing microbiological contamination and water condensation on the surface of large inflatable structures. It is the goal of large inflatable structures used for habitation to take advantage of the large surface area of the structure and reject waste heat passively through the walls of the structure. Too much heat rejection leads to a cold spot with water condensation, and eventually microbiological contamination. A distributed sensor system that can measure temperature, humidity, and microbiological contamination across a large surface would benefit designers of large inflatable habitable structures. The third opportunity relates to sensing microbial bioreactors used for waste water processing and reuse. Microbiological bioreactors offer considerable advantages in weight and power compared to adsorption bed based systems when used for long periods of time. Managing and controlling bioreactors is greatly helped if distributed microsensors measured the biological populations continuously in many locations within the bioreactor. Nitrifying bacteria are of special interest to bioreactor designers, and any sensors that could measure the populations of these types of bacteria would help the control and operation of bioreactors. J

Graf, John C.↗

NASA NTRS2017

PIPER Continuous Adiabatic Demagnetization Refrigerator

We report upon the development and testing of a 4-stage adiabatic demagnetization refrigerator (ADR) capable of continuous cooling at 0.100 Kelvin. This cooler is being built to cool the detector array aboard NASA's Primordial Inflation Polarization Explorer (PIPER) observatory. The goal of this balloon mission is to measure the primordial gravitational waves that should exist if the theory of cosmological inflation is correct. At altitude, the ADR will hold the array of transition-edge sensors at 100 mK continuously while periodically rejecting heat to a 1.2 K pumped helium bath. During testing on ground, the array is held at the same temperature but heat is rejected to a 4.2 K helium bath indicating the flexibility in this coolers design.

Magnetic Cooling Sub-Kelvin ADR↗

NASA NTRS1981

Study of Thermal Control Systems for orbiting power systems

Thermal control system designs were evaluated for the 25 kW power system. Factors considered include long operating life, high reliability, and meteoroid hazards to the space radiator. Based on a cost advantage, the bumpered pumped fluid radiator is recommended for the initial 25 kW power system and intermediate versions up to 50 kW. For advanced power systems with heat rejection rates above 50 kW the lower weight of the advanced heat pipe radiator offsets the higher cost and this design is recommended. The power system payloads heat rejection allocations studies show that a centralized heat rejection system is the most weight and cost effective approach. The thermal interface between the power system and the payloads was addressed and a concept for a contact heat exchanger that eliminates fluid transfer between the power system and the payloads was developed. Finally, a preliminary design of the thermal control system, with emphasis on the radiator and radiator deployment mechanism, is presented.

Howell, H. R.↗

NASA NTRS2016

Subscale Water Based Phase Change Material Heat Exchanger Development

Supplemental heat rejection devices are required in many spacecraft as the radiators are not sized to meet the full heat rejection demand. One means of obtaining additional heat rejection is through the use of phase change material heat exchangers (PCM HX's). PCM HX's utilize phase change to store energy in unfavorable thermal environments (melting) and reject the energy in favorable environments (freezing). Traditionally, wax has been used as a PCM on spacecraft. However, water is an attractive alternative because it is capable of storing about 40% more energy per unit mass due to its higher latent heat of fusion. The significant problem in using water as a PCM is its expansion while freezing, leading to structural integrity concerns when housed in an enclosed heat exchanger volume. Significant investigation and development has taken place over the past five years to understand and overcome the problems associated with water PCM HX's. This paper reports on the final efforts by Johnson Space Center's Thermal Systems Branch to develop a water based PCM HX. The test article developed and reported on is a subscale version of the full-scale water-based PCM HX's constructed by Mezzo Technologies. The subscale unit was designed by applying prior research on freeze front propagation and previous full-scale water PCM HX development. Design modifications to the subscale unit included use of urethane bladder, decreased aspect ratio, perforated protection sheet, and use of additional mid-plates. Testing of the subscale unit was successful and 150 cycles were completed without fail.

Sheth, Rubik↗

NASA NTRS2018

Characterization Testing of Lockheed Martin Micro1-2 Cryocoolers Optimized for 220 K Environment

The Mapping Imaging Spectrometer for Europa (MISE) Instrument on the Europa Clipper mission has baselined a Lockheed Martin high-power Micro1-2 pulse tube cryocooler operating at 135 Hz with a 220 K heat rejection temperature. This paper describes the testing and results of two Lockheed Martin Micro1-2 coolers optimized for these conditions. The thermal performance of the microcoolers was measured in vacuum for heat reject temperatures between 220 and 230 K. The coolers were operated with input powers ranging from 5 W to 40 W and drive frequency between 125 Hz and 145 Hz. The optimal drive frequency was strongly dependent on heat reject temperature. In addition, the exported forces and torques of the coolers were measured at 300 K heat rejection for input powers ranging from 5 W to 60 W and drive frequency between 120 Hz and 160 Hz. The exported forces were dependent on both piston amplitude and drive frequency. Moreover, the following were measured on one of the coolers: DC and AC magnetic fields at various locations, the off-state conductance of the pulse tube, and the effect of inclination angle of the pulse tube relative to gravity on the performance of the cooler. Finally, one of the coolers was subjected to random vibration testing.

Rodriguez, J.I.↗

NASA NTRS1990

Modular, thermal bus-to-radiator integral heat exchanger design for Space Station Freedom

The baseline concept is introduced for the 'integral heat exchanger' (IHX) which is the interface of the two-phase thermal bus with the heat-rejecting radiator panels. A direct bus-to-radiator heat-pipe integral connection replaces the present interface hardware to reduce the weight and complexity of the heat-exchange mechanism. The IHX is presented in detail and compared to the baseline system assuming certain values for heat rejection, mass per unit width, condenser capacity, contact conductance, and assembly mass. The spreadsheet comparison can be used to examine a variety of parameters such as radiator length and configuration. The IHX is shown to permit the reduction of panel size and system mass in response to better conductance and packaging efficiency. The IHX is found to be a suitable heat-rejection system for the Space Station Freedom because it uses present technology and eliminates the interface mechanisms.

Chambliss, Joe↗

NASA NTRS1972

Study of fuel cell thermal control systems for advanced missions.

This study evaluated many heat rejection and thermal control concepts which could be applied to fuel cells for long term (600 hours) orbital and lunar surface missions. The concepts considered several types of radiators which utilized pumped gas, liquid and two phase working fluids and incorporated solid conduction fins as well as heat pipe (vapor chamber) fins. The comparison of the concepts was based on weight, area and other factors such as standby power, ability to accommodate heat load variation, control complexity, and meteoroid survival capability. A design selection matrix was established and an optimum (primary) and an alternate (secondary) heat rejection concept was chosen. Heat rejection techniques utilizing self-controlled heat pipe radiators dominate the results.

Caputo, R. S.↗

NASA NTRS1991

Active thermal control system evolution

The 'restructured' baseline of the Space Station Freedom (SSF) has eliminated many of the growth options for the Active Thermal Control System (ATCS). Modular addition of baseline technology to increase heat rejection will be extremely difficult. The system design and the available real estate no longer accommodate this type of growth. As the station matures during its thirty years of operation, a demand of up to 165 kW of heat rejection can be expected. The baseline configuration will be able to provide 82.5 kW at Eight Manned Crew Capability (EMCC). The growth paths necessary to reach 165 kW have been identified. Doubling the heat rejection capability of SSF will require either the modification of existing radiator wings or the attachment of growth structure to the baseline truss for growth radiator wing placement. Radiator performance can be improved by enlarging the surface area or by boosting the operating temperature with a heat pump. The optimal solution will require both modifications. The addition of growth structure would permit the addition of a parallel ATCS using baseline technology. This growth system would simplify integration. The feasibility of incorporating these growth options to improve the heat rejection capacity of SSF is under evaluation.

Petete, Patricia A.↗

NASA NTRS1972

Development of an engineering model traveling wave tube amplifier for space communication systems

A design has been made of a 100 watt traveling-wave tube amplifier for use in space communication applications. The features of very high overall efficiency and heat rejection of waste heat at low thermal densities were predominant in the design concept. The design concept was proven by building a series of tubes, operating at efficiencies up to 50%. These tubes utilized heat pipe cooling and heat distribution such that 150 watts of waste heat was rejected at a density of less than 1.5 watts per square inch. A power supply to convert a 28 volt primary line of the needs of the TWT was built and operated at 85% efficiency.

Eallonardo, C. M.↗

NASA NTRS1990

Space Station thermal control system evolution

The thermal control system (TCS) for the space station assembly complete configuration includes a two-phase central thermal bus with a supplemental body mounted radiator system. Evolution of the space station from a heat rejection capacity of 75 kW to 300 kW will require scars to expand the thermal fluid distribution network, equipment replacement to enable greater thermal transport capacity, and enlargement of the heat rejection subsystem for increased heat rejection. The TCS requirements for assembly completion and growth are presented along with a review of the basic structure of the active and passive thermal control systems which include provisions for growth.

Bullock, Richard L.↗

NASA NTRS2011

Long Duration Testing of a Spacesuit Water Membrane Evaporator Prototype

The Spacesuit Water Membrane Evaporator (SWME) is a heat-rejection device that is being developed to perform thermal control for advanced spacesuits. Cooling is achieved by circulating water from the liquid cooling garment (LCG) through hollow fibers (HoFi?s), which are small hydrophobic tubes. Liquid water remains within the hydrophobic tubes, but water vapor is exhausted to space, thereby removing heat. A SWME test article was tested over the course of a year, for a total of 1200 cumulative hours. In order to evaluate SWME tolerance to contamination due to constituents caused by distillation processes, these constituents were allowed to accumulate in the water as evaporation occurred. A test article was tested over the course of a year for a total of 1200 cumulative hours. The heat rejection performance of the SWME degraded significantly--below 700 W, attributable to the accumulation of rust in the circulating loop and biofilm growth. Bubble elimination capability, a feature that was previously proven with SWME, was compromised during the test, most likely due to loss of hydrophobic properties of the hollow fibers. The utilization of water for heat rejection was shown not to be dependent on test article, life cycle, heat rejection rate, or freezing of the membranes.

Bue, Grant C.↗

NASA NTRS2012

Long Duration Testing of a Spacesuit Water Membrane Evaporator Prototype

The Spacesuit Water Membrane Evaporator (SWME) is a heat-rejection device that is being developed to perform thermal control for advanced spacesuits. Cooling is achieved by circulating water from the liquid cooling garment (LCG) through hollow fibers (HoFi s), which are small hydrophobic tubes. Liquid water remains within the hydrophobic tubes, but water vapor is exhausted to space, thereby removing heat. A SWME test article was tested over the course of a year, for a total of 600 cumulative hours. In order to evaluate SWME tolerance to contamination due to constituents caused by distillation processes, these constituents were allowed to accumulate in the water as evaporation occurred. A test article was tested over the course of a year for a total of 600 cumulative hours. The heat rejection performance of the SWME degraded significantly--below 700 W, attributable to the accumulation of rust in the circulating loop and biofilm growth. Bubble elimination capability, a feature that was previously proven with SWME, was compromised during the test, most likely due to loss of hydrophobic properties of the hollow fibers. The utilization of water for heat rejection was shown not to be dependent on test article, life cycle, heat rejection rate, or freezing of the membranes.

Bue, Grant C.↗

NASA NTRS2014

Controllable Multi-Functional Heat Exchanger for Space Applications

Team Spacering will design SpaceRing, a cooling and waste heat recycling system that addresses NASA’s interest for a cooling and power-generating unit in its prototype deep space habitat[1]. The merits of this system are as follows: - Controllable cooling is critical in maintaining a habitable environment and in the functionality of life support systems, powering electronics and other devices in a spacecraft. - Power is expensive to produce in a space environment. Converting wasted thermal energy into electricity will save energy and money for the space station. Space habitation is a highly complex endeavor. Consequently, the SpaceRing must adhere to additional constraints, which include: - Lack of natural fluid or gas convection due to micro-gravity environment - Limited range for heat source and cold sink temperature - Subject to high fluctuations of incoming thermal energy - Minimal weight due to the high cost of transporting items into space After consulting with NASA and experts in thermal studies, several design objectives were identified and are listed below in order of importance: - High cooling capacity - Reliable safety control and alert system - Maximization of energy generation from wasted heat - Optimization of cooling capability and power generation - Long operational life and minimization of maintenance In space, conventional cooling systems utilize passive heat pipe systems for small-scale heat rejection and pumped pipes with heat exchangers for large-scale heat rejection. Thermal power generation in space is mainly achieved by using thermoelectric material that produces electric energy from temperature differentials. We propose a design that combines cooling and thermoelectric power generation. Using a hot pipe connected to a heat source, a cold pipe connected to a radiator, a heat exchanger in contact with both pipes, and a pump in each pipe to move the fluid, we make an active cooling system. The thermoelectric material is inserted in the heat exchanger between the hot and cold pipes to extract energy from the maximum temperature difference coming from the two pipes. It also provides controllable cooling capacity through the use of the thermoelectric material as cooling modules. We have tested our system under target heat source and sink conditions and measured the temperatures and flow rates. Our preliminary results are promising, demonstrating the output of 25 Watts of power on average. We are in the process of running control tests to compare the heat transfer performance without the thermoelectric generators, which is vital in determining whether or not the generators introduce an unacceptable amount of thermal resistance. While some control tests have been run already, the results have been noisy. We have modified our testing environment and data acquisition methods to allow for simple comparison between the control test results and thermoelectric performance results. Lastly, we have modified the heat exchanger design slightly to include acrylic baffles. This will allow the heat to flow through the heat exchanger more efficiently. We have presented our preliminary results to NASA and will deliver the final product by the end of May 2014.

Lee Xiong↗

NASA NTRS1980

Spacecraft radiators for advanced mission requirements

Design requirements for spacecraft heat rejection systems are identified, and their impact on the construction of conventional pumped fluid and hybrid heat pipe/pumped fluid radiators is evaluated. Heat rejection systems to improve the performance or reduce the cost of the spacecraft are proposed. Heat rejection requirements which are large compared to those of existing systems and mission durations which are relatively long, are discussed.

Leach, J. W.↗

NASA NTRS1990

Megawatt solar power systems for lunar surface operations

Lunar surface operations require habitation, transportation, life support, scientific, and manufacturing systems, all of which require some form of power. As an alternative to nuclear power, the development of a modular one megawatt solar power system is studied, examining both photovoltaic and dynamic cycle conversion methods, along with energy storage, heat rejection, and power backup subsystems. For photovoltaic power conversion, two systems are examined. First, a substantial increase in photovoltaic conversion efficiency is realized with the use of new GaAs/GaSb tandem photovoltaic cells, offering an impressive overall array efficiency of 23.5 percent. Since these new cells are still in the experimental phase of development, a currently available GaAs cell providing 18 percent efficiency is examined as an alternate to the experimental cells. Both Brayton and Stirling cycles, powered by linear parabolic solar concentrators, are examined for dynamic cycle power conversion. The Brayton cycle is studied in depth since it is already well developed and can provide high power levels fairly efficiently in a compact, low mass system. The dynamic conversion system requires large scale waste heat rejection capability. To provide this heat rejection, a comparison is made between a heat pipe/radiative fin system using advanced composites, and a potentially less massive liquid droplet radiator system. To supply power through the lunar night, both a low temperature alkaline fuel cell system and an experimental high temperature monolithic solid-oxide fuel cell system are considered. The reactants for the fuel cells are stored cryogenically in order to avoid the high tankage mass required by conventional gaseous storage. In addition, it is proposed that the propellant tanks from a spent, prototype lunar excursion vehicle be used for this purpose, therefore resulting in a significant overall reduction in effective storage system mass.

Adams, Brian↗

NASA NTRS2013

Optimized Radiator Geometries for Hot Lunar Thermal Environments

The optimum radiator configuration in hot lunar thermal environments is one in which the radiator is parallel to the ground and has no view to the hot lunar surface. However, typical spacecraft configurations have limited real estate available for top-mounted radiators, resulting in a desire to use the spacecraft's vertically oriented sides. Vertically oriented, flat panel radiators will have a large view factor to the lunar surface, and thus will be subjected to significant incident lunar infrared heat. Consequently, radiator fluid temperatures will need to exceed approximately 325 K (assuming standard spacecraft radiator optical properties) in order to provide positive heat rejection at lunar noon. Such temperatures are too high for crewed spacecraft applications in which a heat pump is to be avoided. A recent study of vertically oriented radiator configurations subjected to lunar noon thermal environments led to the discovery of a novel radiator concept that yielded positive heat rejection at lower fluid temperatures. This radiator configuration, called the Intense Thermal Infrared Reflector (ITIR), has exhibited superior performance to all previously analyzed concepts in terms of heat rejection in the lunar noon thermal environment. A key benefit of ITIR is the absence of louvers or other moving parts and its simple geometry (no parabolic shapes). ITIR consists of a specularly reflective shielding surface and a diffuse radiating surface joined to form a horizontally oriented V-shape (shielding surface on top). The point of intersection of these surfaces is defined by two angles, those which define the tilt of each surface with respect to the local horizontal. The optimum set of these angles is determined on a case-by-case basis. The idea assumes minimal conductive heat transfer between shielding and radiating surfaces, and a practical design would likely stack sets of these surfaces on top of one another to reduce radiator thickness.

Ochoa, Dustin↗