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At least 163 records · Page 9

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NASA NTRS2012

Loop Heat Pipe with Thermal Control Valve as a Variable Thermal Link

Future lunar landers and rovers will require variable thermal links that allow for heat rejection during the lunar daytime and passively prevent heat rejection during the lunar night. During the lunar day, the thermal management system must reject the waste heat from the electronics and batteries to maintain them below the maximum acceptable temperature. During the lunar night, the heat rejection system must either be shut down or significant amounts of guard heat must be added to keep the electronics and batteries above the minimum acceptable temperature. Since guard heater power is unfavorable because it adds to system size and complexity, a variable thermal link is preferred to limit heat removal from the electronics and batteries during the long lunar night. Conventional loop heat pipes (LHPs) can provide the required variable thermal conductance, but they still consume electrical power to shut down the heat transfer. This innovation adds a thermal control valve (TCV) and a bypass line to a conventional LHP that proportionally allows vapor to flow back into the compensation chamber of the LHP. The addition of this valve can achieve completely passive thermal control of the LHP, eliminating the need for guard heaters and complex controls.

Hartenstine, John↗

NASA NTRSDate not supplied

Passive PCB Mounted Thermal Switch

A printed circuit board (PCB) generates heat during operation. In a vacuum this heat can only be rejected by conduction and radiation. In order to reject heat effectively the PCB needs to be thermally conductive to the electronics enclosure. However, when the enclosure temperature gets too low, too much heat is rejected from the PCB and it can fail. This thermal switch has been designed to allow for heat rejection when the PCB is sufficiently warm but then greatly reduce heat rejection when the temperature gets low enough.

Jarret Bone↗

NASA NTRS2025

ISS Radiator Face Sheet Anomaly Investigation and Return to Function

Maintaining sufficient heat rejection on the International Space Station (ISS) is critical to the function of the Low-Earth Orbit station. The External Active Thermal Control System (EATCS) rejects the excess heat generated by the US On-Orbit Segment (USOS) modules. The system uses single-phase liquid ammonia to collect the heat and reject it to six radiators (three on the Starboard side, three on the Port side). Each radiator is made up of eight panels. In September 2008, imagery of the Starboard radiators was conducted and showed one of the panels’ face sheets had peeled up from the internal honeycomb core structure. Out of an abundance of caution, the radiator was isolated from the rest of the EATCS and the ammonia was vented to space. Since the radiator was vented, periodic imagery of all radiator panels was taken to monitor any changes in the face sheet and inspect for other radiator anomalies. In January 2025, after years of trending and inspections, the radiator was reintegrated into the system to increase heat rejection capabilities. This paper will document the multi-year investigation that took place to determine root cause and mitigations implemented to reduce risk to the system. A summary of the data since the investigation will show rationale for reintegrating the radiator even with the damaged panel. The goal of the paper is to document the investigation for historical purposes and provide future programs with findings discovered during the investigation.

Aaron Rodriguez↗

NASA NTRS2026

ISS Radiator Face Sheet Anomaly Investigation and Return to Function

Aaron Rodriguez↗

NASA NTRS2008

Potential of a New Lunar Surface Radiator Concept 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 approx.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 Upright Lunar Terrain Radiator Assembly (ULTRA), has exhibited superior performance to all previously analyzed concepts in terms of heat rejection in the lunar noon thermal environment. A key benefit of the ULTRA is the absence of louvers or other moving parts and its simple geometry. Analysis of the ULTRA for a lunar extravehicular activity (EVA) portable life support system (PLSS) is shown to provide moderate heat rejection, on average, at all solar incident angles assuming an average radiator temperature of 294 K, whereas prior concepts exhibited insignificant heat rejection or heat absorption at higher incident angles. The performance of the ULTRA for a lunar lander is also discussed and compared to the performance of a vertically oriented, flat panel radiator at various lunar latitudes.

Ochoa, Dustin A.↗

NASA NTRS1972

EC/LSS thermal control system study for the space shuttle

The results of a parametric weight analysis of heat rejection systems for the space shuttle orbiter are presented. Integrating the suborbital heat rejection system with the overall heat rejection system design and the possible use of a common system for both on-orbit and suborbital operations require an overall system and parametric analyses applicable to all mission phases. The concept of equivalent weights, with weight penalties assigned for power, induced aircraft drag and radiator area is used to determine weight estimates for the following candidate systems: vapor cycle refrigeration, gas cycle refrigeration, radiators (space and atmospheric convectors), expendable heat sinks, and ram air. The orbiter power penalty, ram air penalty, and radiator weight penalty are analyzed. The vapor compression system and an expendable fluid system utilizing a multifluid spraying flash evaporator are selected as the two most promising systems. These are used for maximum on-orbit heat rejection in combination with or as a supplement to a space radiator.

Howell, H. R.↗

NASA NTRS2006

Preliminary Trade Study of Phase Change Heat Sinks

For short durations, phase change based heat rejection systems are a very effective way of removing heat from spacecraft. Future NASA vehicles, such as the Crew Exploration Vehicle (CEV), will require non-radiative heat rejection systems during at least a portion of the planned mission, just as their predecessors have. While existing technologies are available to modify, such as Apollo era sublimators, or the Space Shuttle Flash Evaporator System (FES), several new technologies are under development or investigation to progress beyond these existing heat rejection systems. Examples include the Multi-Fluid Evaporator developed by Hamilton Sundstrand, improvements upon the Contaminant Insensitive Sublimator originally developed for the X-38 program, and a Compact Flash Evaporator System (CFES). Other possibilities evaluate new ways of operating existing designs. The new developments are targeted at increasing operating life, expanding the environments in which the system can operate, improving the mass and volume characteristics, or some combination of these or other improvements. This paper captures the process and results of a preliminary trade study performed at Johnson Space Center to compare the various existing and proposed phase change based heat rejection systems for the CEV. Because the new systems are still in development, and the information on existing systems is extrapolation, this trade study is not meant to suggest a final decision for future vehicles. The results of this early trade study are targeted to aid the development efforts for the new technologies by identifying issues that could reduce the chances of selection for the CEV.

Anderson, Molly↗

NASA NTRS1977

Advanced spacecraft thermal control techniques

The problems of rejecting large amounts of heat from spacecraft were studied. Shuttle Space Laboratory heat rejection uses 1 kW for pumps and fans for every 5 kW (thermal) heat rejection. This is rather inefficient, and for future programs more efficient methods were examined. Two advanced systems were studied and compared to the present pumped-loop system. The advanced concepts are the air-cooled semipassive system, which features rejection of a large percentage of the load through the outer skin, and the heat pipe system, which incorporates heat pipes for every thermal control function.

Fritz, C. H.↗

NASA NTRS1976

Auxiliary Payload Power System thermal control

The Auxiliary Payload Power System (APPS) provides supplementary power and cooling to Space Processing Application (SPA) experiments to be mounted in the APPS and the Spacelab in the Shuttle Payload Bay. SPA experiment operations are planned for early Shuttle flights. This paper presents thermal control study results for preliminary analysis and design definition of the APPS. A 100/sq m, three-wing, pumped-fluid, deployable radiator with separate APPS equipment and SPA experiments coolant loops was selected as the baseline. The system is capable of rejecting the heat (approximately 26 kw) associated with the production and consumption of approximately 16 kw of electrical power produced by the APPS fuel cells for a worst case radiator orientation. For the most favorable orientation, the heat rejection and power capability approach 38 and 24 kw, respectively. Alternate approaches were evaluated, such as heat pipes for the radiator and alternate fluids for the coolant loops. Emphasis was placed on using Shuttle developed hardware: coolant pumps, heat exchangers, fluids, and radiator technology.

Nagel, R. G.↗

NASA NTRS2012

Experimental Investigation of Transient Sublimator Performance

Sublimators have been used as heat rejection devices for a variety of space applications including the Apollo Lunar Module and the Extravehicular Mobility Unit (EMU). Sublimators typically operate with steady-state feedwater utilization at or near 100%. However, sublimators are currently being considered for operations in a cyclical topping mode, which represents a new mode of operation for sublimators. Sublimators can be used as a supplemental heat rejection device during mission phases where the environmental temperature or heat rejection requirement changes rapidly. This scenario may occur during low lunar orbit, low earth orbit, or other planetary orbits. In these mission phases, the need for supplemental heat rejection will vary between zero and some fraction of the overall heat load. In particular, supplemental heat rejection is required for the portion of the orbit where the radiative sink temperature exceeds the system setpoint temperature. This paper will describe the effects of these transient starts and stops on the feedwater utilization during various feedwater timing scenarios. Experimental data from various scenarios is analyzed to investigate feedwater consumption efficiency under the cyclical conditions. Start up utilization tests were conducted to better understand the transient performance. This paper also provides recommendations for future sublimator design and transient operation.

Sheth, Rubik B.↗

NASA NTRS2010

Hollow Fiber Spacesuit Water Membrane Evaporator Development and Testing for Advanced Spacesuits

The spacesuit water membrane evaporator (SWME) is being developed to perform the thermal control function for advanced spacesuits to take advantage of recent advances in micropore membrane technology in providing a robust heat-rejection device that is potentially less sensitive to contamination than is the sublimator. Principles of a sheet membrane SWME design were demonstrated using a prototypic test article that was tested in a vacuum chamber at JSC in July 1999. The Membrana Celgard X50-215 microporous hollow fiber (HoFi) membrane was selected after recent contamination tests as the most suitable candidate among commercial alternatives for HoFi SWME prototype development. A design that grouped the fiber layers into stacks, which were separated by small spaces and packaged into a cylindrical shape, was developed into a full-scale prototype consisting 14,300 tube bundled into 30 stacks, each of which are formed into a chevron shape and separated by spacers and organized into three sectors of ten nested stacks. Vacuum chamber testing has been performed characterize heat rejection as a function of inlet water temperature and water vapor backpressure and to show contamination resistance to the constituents expected to be found in potable water produced by the distillation processes. Other tests showed the tolerance to freezing and suitability to reject heat in a Mars pressure environment.

Bue, Grant C.↗

NASA NTRS2011

Hollow Fiber Space Water Membrane Evaporator Flight Prototype Design and Testing

The spacesuit water membrane evaporator (SWME) is being developed to perform thermal control for advanced spacesuits and to take advantage of recent advances in micropore membrane technology. This results in a robust heat-rejection device that is potentially less sensitive to contamination than is the sublimator. The current design was based on a previous design that grouped the fiber layers into stacks, which were separated by small spaces and packaged into a cylindrical shape. This was developed into a full-scale prototype consisting of 14,300 tube bundled into 30 stacks, each of which is formed into a chevron shape and separated by spacers and organized into three sectors of 10 nested stacks. The new design replaced metal components with plastic ones, eliminated the spacers, and has a custom built flight like backpressure valve mounted on the side of the SWME housing to reduce backpressure when fully open. A number of tests were performed in order to improve the strength of the polyurethane header that holds the fibers in place while the system is pressurized. Vacuum chamber testing showed similar heat rejection as a function of inlet water temperature and water vapor backpressure was similar to the previous design. Other tests pushed the limits of tolerance to freezing and showed suitability to reject heat in a Mars pressure environment with and without a sweep gas. Tolerance to contamination by constituents expected to be found in potable water produced by distillation processes was tested in a conventional way by allowing constituents to accumulate in the coolant as evaporation occurs. For this purpose, the SWME cartridge has endured an equivalent of 30 EVAs exposure and demonstrated acceptable performance decline.

Bue, Grant C.↗

NASA NTRS2010

Hollow Fiber Flight Prototype Spacesuit Water Membrane Evaporator Design and Testing

The spacesuit water membrane evaporator (SWME) is being developed to perform thermal control for advanced spacesuits and to take advantage of recent advances in micropore membrane technology. This results in a robust heat-rejection device that is potentially less sensitive to contamination than is the sublimator. The Membrana Celgard X50-215 microporous hollow-fiber (HoFi) membrane was selected after recent extensive testing as the most suitable candidate among commercial alternatives for continued SWME prototype development. The current design was based on a previous design that grouped the fiber layers into stacks, which were separated by small spaces and packaged into a cylindrical shape. This was developed into a full-scale prototype consisting of 14,300 tube bundled into 30 stacks, each of which is formed into a chevron shape and separated by spacers and organized into three sectors of 10 nested stacks. The new design replaced metal components with plastic ones, and has a custom built flight like backpressure valve mounted on the side of the SWME housing to reduce backpressure when fully open. The spacers that provided separation of the chevron fiber stacks were eliminated. Vacuum chamber testing showed improved heat rejection as a function of inlet water temperature and water vapor backpressure compared with the previous design. Other tests pushed the limits of tolerance to freezing and showed suitability to reject heat in a Mars pressure environment with and without a sweep gas. Tolerance to contamination by constituents expected to be found in potable water produced by distillation processes was tested in a conventional way by allowing constituents to accumulate in the coolant as evaporation occurs. For this purpose, the SWME cartridge has endured an equivalent of 30 EVAs exposure and demonstrated minimal performance decline.

Bue, Grant↗

NASA NTRS2008

Spacesuit Water Membrane Evaporator Development for Lunar Missions

For future lunar extra-vehicular activities (EVA), one method under consideration for rejecting crew and electronics heat involves evaporating water through a hydrophobic, porous Teflon membrane. A Spacesuit Water Membrane Evaporator (SWME) prototype using the Teflon membrane was tested successfully by Ungar and Thomas (2001) with predicted performance matching test data well. The above referenced work laid the foundation for the design of the SWME development unit, which is being considered for service in the Constellation System Spacesuit Element (CSSE) Portable Life Support System (PLSS). Multiple PLSS SWME configurations were considered on the basis of thermal performance, mass, volume, and performance and manufacturing risk. All configurations were a variation of an alternating concentric water and vapor channel configuration or a stack of alternating rectangular water and vapor channels. Supporting thermal performance trades mapped maximum SWME heat rejection as a function of water channel thickness, vapor channel thickness, channel length, number of water channels, porosity of the membrane structural support, and backpressure valve throat area. Preliminary designs of each configuration were developed to determine total mass and volume as well as to understand manufacturing issues. Review of configurations led to the selection of a concentric annulus configuration that meets the requirements of 800 watts (W) of heat rejection. Detailed design of the SWME development unit will be followed by fabrication of a prototype test unit, with thermal testing expected to start in 2008.

Vogel, Matt R.↗

NASA NTRS1999

Thermal Radiator Pointing for International Space Station

In order to provide thermal radiation environments that result in adequate beat rejection, the single-phase, liquid ammonia (NH3) heat rejection system on the International Space Station (ISS) requires that its two thermal radiator wings be dynamically rotated as the ISS travels through its orbit. This paper discusses the closed-loop, thermal radiator pointing system that is used on ISS to ensure adequate heat rejection by the radiators, while preventing freezing of the ammonia under low heat loads and cold-environmental conditions. Although initial designs used an open-loop approach for radiator pointing, concerns about performance robustness, algorithm complexity, memory requirements, and sustaining support drove the development of a more robust, simpler, closed-loop system. Hence, the challenge of the closed-loop system was to utilize existing sensors, actuators and computers to fit into the existing hardware and software architecture of the ISS. Using a proportional-integral (PI) control architecture with limited output and an anti-windup integrator, the temperature of the ammonia coming out of the radiator is measured and controlled by adjusting the radiator wing orientation. The radiator wing orientation for the local minimum environment is fed forward to the control system, and the closed-loop controller is used to generate a bias off of that local minimum environment in order to heat up the ammonia when necessary to avoid freezing. In the earth's shadow, the controller is suspended and the radiator wing is oriented to face the earth, the local maximum thermal environment which further prevents freezing of the ammonia. This control architecture is shown to provide adequate heat rejection and avoid freezing of the ammonia, even though the physical system consists of large transport delays and time-varying dynamics which change dramatically due to orbit motion and variable heat loads.

Green, Scott↗

NASA NTRS1989

Numerical model of solar dynamic radiator for parametric analysis

Growth power requirements for Space Station Freedom will be met through addition of 25 kW solar dynamic (SD) power modules. The SD module rejects waste heat from the power conversion cycle to space through a pumped-loop, multi-panel, deployable radiator. The baseline radiator configuration was defined during the Space Station conceptual design phase and is a function of the state point and heat rejection requirements of the power conversion unit. Requirements determined by the overall station design such as mass, system redundancy, micrometeoroid and space debris impact survivability, launch packaging, costs, and thermal and structural interaction with other station components have also been design drivers for the radiator configuration. Extensive thermal and power cycle modeling capabilities have been developed which are powerful tools in Station design and analysis, but which prove cumbersome and costly for simple component preliminary design studies. In order to aid in refining the SD radiator to the mature design stage, a simple and flexible numerical model was developed. The model simulates heat transfer and fluid flow performance of the radiator and calculates area mass and impact survivability for many combinations of flow tube and panel configurations, fluid and material properties, and environmental and cycle variations. A brief description and discussion of the numerical model, it's capabilities and limitations, and results of the parametric studies performed is presented.

Rhatigan, Jennifer L.↗

NASA NTRS1977

Instrument canister thermal control

A transient thermal analysis and test of a thermal control canister is described. The 1 x 1 x 3 m canister provides a uniform thermal environment for shuttle instrument payloads requiring fine temperature control, the design goal being operation between 0 C and 20 C with a range of plus or minus 1 C at any selected set-point temperature. The canister side walls are isothermalized by a system of longitudinal and circumferential heat pipes rejecting heat through feedback controlled, variable conductance heat pipes to side mounted radiators. A breadboard model of two side walls and two radiators was tested in a thermal vacuum chamber. The breadboard was stable over a wide range of effective environments, experiment dissipations, and control point temperature levels.

Harwell, W.↗

NASA NTRS2000

A Continuous Adiabatic Demagnetization Refrigerator for Use with Mechanical Coolers

We have begun developing an adiabatic demagnetization refrigerator (ADR) which can produce continuous cooling at temperatures of 50 mK or lower, with high cooling power (goal of 10 PW). The design uses multiple stages to cascade heat from a continuously-cooled stage up to a heat sink. The serial arrangement makes it possible to add stages to extend the operating range to lower temperature, or to raise the heat rejection temperature. Compared to conventional single-shot ADRS, this system achieves higher cooling power per unit mass and is able to reject its heat at a more uniform rate. For operation with a mechanical cryocooler, this latter feature stabilizes the heat sink temperature and allows both the ADR and cryocooler to operate more efficiently. The ADR is being designed to operate with a heat sink as warm as 10-12 K to make it compatible with a wide variety of mechanical coolers as part of a versatile, cryogen-free low temperature cooling system. A two-stage system has been constructed and a proof-of-principle demonstration was conducted at 100 mK. Details of the design and test results, as well as the direction of future work, are discussed.

Shirron, P.↗