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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Jet pump-drive system for heat removal

The invention does away with the necessity of moving parts such as a check valve in a nuclear reactor cooling system. Instead, a jet pump, in combination with a TEMP, is employed to assure safe cooling of a nuclear reactor after shutdown. A main flow exists for a reactor coolant. A point of withdrawal is provided for a secondary flow. A TEMP, responsive to the heat from said coolant in the secondary flow path, automatically pumps said withdrawn coolant to a higher pressure and thus higher velocity compared to the main flow. The high velocity coolant is applied as a driver flow for the jet pump which has a main flow chamber located in the main flow circulation pump. Upon nuclear shutdown and loss of power for the main reactor pumping system, the TEMP/jet pump combination continues to boost the coolant flow in the direction it is already circulating. During the decay time for the nuclear reactor, the jet pump keeps running until the coolant temperature drops to a lower and safe temperature where the heat is no longer a problem. At this lower temperature, the TEMP/jet pump combination ceases its circulation boosting operation. When the nuclear reactor is restarted and the coolant again exceeds the lower temperature setting, the TEMP/jet pump automatically resumes operation. The TEMP/jet pump combination is thus automatic, self-regulating and provides an emergency pumping system free of moving parts.

French, James R.↗
NASA NTRS2012

Design of Accumulators and Liquid/Gas Charging of Single Phase Mechanically Pumped Fluid Loop Heat Rejection Systems

For single phase mechanically pumped fluid loops used for thermal control of spacecraft, a gas charged accumulator is typically used to modulate pressures within the loop. This is needed to accommodate changes in the working fluid volume due to changes in the operating temperatures as the spacecraft encounters varying thermal environments during its mission. Overall, the three key requirements on the accumulator to maintain an appropriate pressure range throughout the mission are: accommodation of the volume change of the fluid due to temperature changes, avoidance of pump cavitation and prevention of boiling in the liquid. The sizing and design of such an accumulator requires very careful and accurate accounting of temperature distribution within each element of the working fluid for the entire range of conditions expected, accurate knowledge of volume of each fluid element, assessment of corresponding pressures needed to avoid boiling in the liquid, as well as the pressures needed to avoid cavitation in the pump. The appropriate liquid and accumulator strokes required to accommodate the liquid volume change, as well as the appropriate gas volumes, require proper sizing to ensure that the correct pressure range is maintained during the mission. Additionally, a very careful assessment of the process for charging both the gas side and the liquid side of the accumulator is required to properly position the bellows and pressurize the system to a level commensurate with requirements. To achieve the accurate sizing of the accumulator and the charging of the system, sophisticated EXCEL based spreadsheets were developed to rapidly come up with an accumulator design and the corresponding charging parameters. These spreadsheets have proven to be computationally fast and accurate tools for this purpose. This paper will describe the entire process of designing and charging the system, using a case study of the Mars Science Laboratory (MSL) fluid loops, which is en route to Mars for an August 2012 landing.

HRS↗
NASA NTRS1981

A pumped, two-phase flow heat transport system for orbiting instrument payloads

A pumped two-phase (heat absorption/heat rejection) thermal transport system for orbiting instrument payloads is investigated. The thermofluid characteristics necessary for the system design are discussed. A preliminary design with a series arrangement of four instrument heat stations and six radiators in a single loop is described in detail, and the total mass is estimated to be 134 kg, with the radiators, instrument heat stations, and fluid reservoir accounting for approximately 86, 24, and 12 kg, respectively. The evaluation of preliminary test results shows that the system has potential advantages; however, further research is necessary in the areas of one-g and zero-g heat transfer coefficients/fluid regimes, fluid by-pass temperature control, and reliability of small pumps.

Fowle, A. A.↗
NASA NTRS1971

Study of a heat rejection system using capillary pumping

Results of an analytical study investigating the application of capillary pumping to the heat rejection loop of an advanced Rankine cycle power conversion system are presented. The feasibility of the concept of capillary pumping as an alternate to electromagnetic pumping is analytically demonstrated. Capillary pumping is shown to provide a potential for weight and electrical power saving and reliability through the use of redundant systems. A screen wick pump design with arterial feed lines was analytically developed. Advantages of this design are high thermodynamic and hydrodynamic efficiency, which provide a lightweight easily packaged system. Operational problems were identified which must be solved for successful application of capillary pumping. The most important are the development of start up and shutdown procedures, and development of a means of keeping noncondensibles from the system and of earth-bound testing procedures.

Neal, L. G.↗
NASA NTRS1991

Study and Development of an Air Conditioning System Operating on a Magnetic Heat Pump Cycle

This report describes the design of a laboratory scale demonstration prototype of an air conditioning system operating on a magnetic heat pump cycle. Design parameters were selected through studies performed by a Kennedy Space Center (KSC) System Simulation Computer Model. The heat pump consists of a rotor turning through four magnetic fields that are created by permanent magnets. Gadolinium was selected as the working material for this demonstration prototype. The rotor was designed to be constructed of flat parallel disks of gadolinium with very little space in between. The rotor rotates in an aluminum housing. The laboratory scale demonstration prototype is designed to provide a theoretical Carnot Cycle efficiency of 62 percent and a Coefficient of Performance of 16.55.

Wang, Pao-Lien↗
NASA NTRS2013

Performance of the Mechanically Pumped Fluid Loop Rover Heat Rejection System Used for Thermal Control of the Mars Science Laboratory Curiosity Rover on the Surface of Mars

The challenging range of landing sites for which the Mars Science Laboratory Rover was designed, required 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 C and as warm as 38 C, the Rover relies upon a Mechanically Pumped Fluid Loop (MPFL) Rover Heat Rejection System (RHRS) and external radiators to maintain the temperature of sensitive electronics and science instruments within a -40 C to +50 C range. The RHRS harnesses some of the waste heat generated from the Rover power source, known as the Multi Mission Radioisotope Thermoelectric Generator (MMRTG), for use as survival heat for the rover during cold conditions. The MMRTG produces 110 Watts of electrical power while generating waste heat equivalent to approximately 2000 Watts. Heat exchanger plates (hot plates) positioned close to the MMRTG pick up this survival heat from it by radiative heat transfer and supply it to the rover. This design is the first instance of use of a RHRS for thermal control of a rover or lander on the surface of a planet. After an extremely successful landing on Mars (August 5), the rover and the RHRS have performed flawlessly for close to an earth year (half the nominal mission life). This paper will share the performance of the RHRS on the Martian surface as well as compare it to its predictions.

Mars↗
NASA NTRS2016

A Liquid-Liquid Thermoelectric Heat Exchanger as a Heat Pump for Testing Phase Change Material Heat Exchangers

The primary objective of the Phase Change HX payload on the International Space Station (ISS) is to test and demonstrate the viability and performance of Phase Change Material Heat Exchangers (PCM HX). The system was required to pump a working fluid through a PCM HX to promote the phase change material to freeze and thaw as expected on Orion's Multipurpose Crew Vehicle. Due to limitations on ISS's Internal Thermal Control System, a heat pump was needed on the Phase Change HX payload to help with reducing the working fluid's temperature to below 0degC (32degF). This paper will review the design and development of a TEC based liquid-liquid heat exchanger as a way to vary to fluid temperature for the freeze and thaw phase of the PCM HX. Specifically, the paper will review the design of custom coldplates and sizing for the required heat removal of the HX.

Sheth, Rubik B.↗
NASA NTRS2019

Mechanical design and stress analysis challenges overcome to ensure the structural integrity of Europa Clipper's mechanical pumped fluid loop Heat Redistribution System (HRS)

Launching in 2023, NASA’s Europa Clipper mission will place a spacecraft into a long, looping orbit around Jupiter to perform a series of close flyby investigations of its moon Europa. Strong evidence suggests that Europa hosts an ocean of liquid water beneath its icy crust that could harbor conditions favorable for life. The spacecraft is currently under development by the Jet Propulsion Laboratory (JPL) and the Applied Physics Laboratory (APL) and nearing completion of its Detailed Design Phase. A Mechanical Pumped Fluid Loop (MPFL) architecture known as the Heat Redistribution System (HRS) has been baselined to provide power efficient thermal control of the sensitive components within the Avionics, Radio Frequency (RF), and Propulsion Modules. Principally, the HRS harvests waste heat from the onboard dissipating equipment and distributes it to a 1.4 m diameter by 3 m tall cylindrical structure that cocoons and keeps the propulsion subsystem warm. Numerous technical issues had to be resolved in order to securely attach nearly 100 m of HRS tubing to the spacecraft, which is predicted to experience high launch loads. This paper will review the relevant environments and mission constraints as well as outline the stress and fatigue analysis approach taken to verify the detailed mechanical design of the HRS.

Kolenz, Daniel↗
NASA NTRS2004

Thermal Performance of Capillary Pumped Loops Onboard Terra Spacecraft

The Terra spacecraft is the flagship of NASA's 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.

Ku, Jentung↗
NASA NTRS1994

Lunar base heat pump, phase 1

This report describes the Phase 1 process and analysis used to select a refrigerant and thermodynamic cycle as the basis of a vapor compression heat pump requiring a high temperature lift, then to perform a preliminary design to implement the selected concept, including major component selection. Use of a vapor compression heat pump versus other types was based on prior work performed for the Electric Power Research Institute. A high lift heat pump is needed to enable a thermal control system to remove heat down to 275 K from a habitable volume when the external thermal environment is severe. For example, a long-term lunar base habitat will reject heat from a space radiator to a 325 K environment. The first step in the selection process was to perform an optimization trade study, quantifying the effect of radiator operating temperature and heat pump efficiency on total system mass; then, select the radiator operating temperature corresponding to the lowest system mass. Total system mass included radiators, all heat pump components, and the power supply system. The study showed that lunar night operation, with no temperature lift, dictated the radiator size. To operate otherwise would require a high mass penalty to store power. With the defined radiation surface, and heat pump performances assumed to be from 40 percent to 60 percent of the Carnot ideal, the optimum heat rejection temperature ranged from 387 K to 377 K, as a function of heat pump performance. Refrigerant and thermodynamic cycles were then selected to best meet the previously determined design conditions. The system was then adapted as a ground-based prototype lifting temperature to 360 K (versus 385 K for flight unit) and using readily available commercial-grade components. Over 40 refrigerants, separated into wet and dry compression behavioral types, were considered in the selection process. Refrigerants were initially screened for acceptable critical temperature. The acceptable refrigerants were analyzed in ideal single and two-stage thermodynamic cycles. Top candidates were analyzed assuming realistic component limits and system pressure drops, and were evaluated for other considerations such as safety, environmental impact, and commercial availability. A maximum coefficient of performance (COP) of 56 percent of the Carnot ideal was achievable for a three-stage CFC-11 cycle operating under the flight conditions above. The program was completed by defining a control scheme and by researching and selecting the major components, compressor and heat exchangers, that could be used to implement the thermodynamic cycle selected. Special attention was paid to using similar technologies for the SIRF and flight heat pumps resulting in the commercially available equivalent of the flight unit. A package concept was generated for the components selected and mass and volume estimated.

Goldman, Jeffrey H.↗
NASA NTRS1992

Thermal management in space

Thermal management for space-based and lunar- or asteroid-based power generating systems is discussed. Heat pipes, pump loop systems, and advanced radiator concepts are addressed.

Hertzberg, Abraham↗
NASA NTRS1992

Cryogenic storage technology readiness for First Lunar Outpost

The topics are presented in viewgraph form and include the following: an assessment of cryogenic storage technology; cryogenic boiloff predictions; Space Shuttle/Centaur thermodynamic vent system; zero-g thermodynamic vent system; heat exchanger/mixer pump module; the thick multilayer insulation (MLI) development program; blanket geometry concept evaluations; four-inch thick MLI system on 1/4-scale test tank; combined environments of vibration, acceleration, and temperature testing (CEVAT); Centaur fixed foam insulation; insulation system design; and fixed foam on operational Atlas 2.

John R. Schuster↗
NASA NTRS1978

Installation package for integrated programmable electronic controller and hydronic subsystem - solar heating and cooling

A description is given of the Installation, Operation, and Maintenance Manual and information on the power panel and programmable microprocessor, a hydronic solar pump system and a hydronic heating hot water pumping system. These systems are integrated into various configurations for usages in solar energy management, control and monitoring, lighting control, data logging and other solar related applications.

Source record↗
NASA NTRS2013

Vapor Compression and Thermoelectric Heat Pump Heat Exchangers for a Condensate Distillation System: Design and Experiment

Maximizing the reuse of wastewater while minimizing the use of consumables is critical in long duration space exploration. One of the more promising methods of reclaiming urine is the distillation/condensation process used in the cascade distillation system (CDS). This system accepts a mixture of urine and toxic stabilizing agents, heats it to vaporize the water and condenses and cools the resulting water vapor. The CDS wastewater flow requires heating and its condensate flow requires cooling. Performing the heating and cooling processes separately requires two separate units, each of which would require large amounts of electrical power. By heating the wastewater and cooling the condensate in a single heat pump unit, mass, volume, and power efficiencies can be obtained. The present work describes and compares two competing heat pump methodologies that meet the needs of the CDS: 1) a series of mini compressor vapor compression cycles and 2) a thermoelectric heat exchanger. In the paper, the system level requirements are outlined, the designs of the two heat pumps are described in detail, and the results of heat pump performance tests are provided. A summary is provided of the heat pump mass, volume and power trades and a selection recommendation is made.

Erickson, Lisa R.↗
NASA NTRS2014

High-Capacity Spacesuit Evaporator Absorber Radiator (SEAR)

Future human space exploration missions will require advanced life support technology that can operate across a wide range of applications and environments. Thermal control systems for space suits and spacecraft will need to meet critical requirements for water conservation and adaptability to highly variable thermal environments. This paper describes a Space Evaporator Absorber Radiator (SEAR) that has been designed to meet performance requirements for future life support systems. A SEAR system comprises a lithium chloride absorber radiator (LCAR) for heat rejection coupled with a space water membrane evaporator (SWME) for heat acquisition. SEAR systems provide heat pumping to minimize radiator size, thermal storage to accommodate variable environmental conditions, and water absorption to minimize use of expendables. We have built and tested a flightlike, high-capacity LCAR, demonstrated its performance in thermal vacuum tests, and explored the feasibility of an ISS demonstration test of a SEAR system. The new LCAR design provides the same cooling capability as prior LCAR prototypes while enabling over 30% more heat absorbing capacity. Studies show that it should be feasible to demonstrate SEAR operation in flight by coupling with an existing EMU on the space station.

Izenson, Michael G.↗
NASA NTRS2015

High-Capacity Spacesuit Evaporator Absorber Radiator (SEAR)

Future human space exploration missions will require advanced life support technology that can operate across a wide range of applications and environments. Thermal control systems for space suits and spacecraft will need to meet critical requirements for water conservation and multifunctional operation. This paper describes a Space Evaporator Absorber Radiator (SEAR) that has been designed to meet performance requirements for future life support systems. A SEAR system comprises a lithium chloride absorber radiator (LCAR) for heat rejection coupled with a space water membrane evaporator (SWME) for heat acquisition. SEAR systems provide heat pumping to minimize radiator size, thermal storage to accommodate variable environmental conditions, and water absorption to minimize use of expendables. We have built and tested a flight-like, high-capacity LCAR, demonstrated its performance in thermal vacuum tests, and explored the feasibility of an ISS demonstration test of a SEAR system. The new LCAR design provides the same cooling capability as prior LCAR prototypes while enabling over 30% more heat absorbing capacity. Studies show that it should be feasible to demonstrate SEAR operation in flight by coupling with an existing EMU on the space station.

Izenson, Michael G.↗