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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
DOE OSTI2025

The Design of a Multiphase CO2 Turbine Test Facility

As the energy market continues to evolve with increased renewables, such as solar and wind energy, the need to balance energy demand with its availability has become increasingly more challenging. The country’s aggressive decarbonization goals in the upcoming decades will increase renewable energy market penetration, thus exacerbating the disparity between energy demand and renewable resource availability. Long-duration energy storage technologies, such as pumped thermal energy storage (PTES), are potential solutions to this challenge with high potential round-trip efficiency (RTE), no geological or geographical constraints, and low safety risks. The charge mode of a PTES system operates a heat pump cycle, converting input power to thermal energy in hot and cold reservoirs. Fundamentally, heat pumps are commonly used in daily life, but the operating conditions required by high RTE PTES systems are outside the design experience of these systems. The low-pressure side of the cycle for a supercritical CO2 (sCO2) PTES operates just above the vapor-dome and potentially drives the turbine to expand into the dome, resulting in multiphase operation in the rear stages of the turbine. This paper presents the design of the test facility for an advanced multiphase-tolerant sCO2 turbine for the heat pump of a PTES system. Heat rejection and pressure loss estimations for the multiphase sections of the test loop presented unforeseen challenges.

25 ENERGY STORAGE↗
NASA NTRS1991

Preliminary design of a radiator shading device for a lunar outpost

The National Aeronautics and Space Administration is designing a thermal control system for an outpost to be placed permanently on the Moon. One of the functions of the thermal control system is to reject waste heat, which can be accomplished through a radiator. At the lunar equator and during the lunar midday, an unshaded radiator absorbs more heat than it rejects. This problem can be solved by using a shading device to reduce radiation incident on the radiator. The design team was asked to develop concepts for reducing the radiation incident on the radiator and for deploying the radiator and shade system for a 10 kW and a 25 kW heat rejection system. The design team was also asked to develop the best concepts into preliminary design. From the several alternatives developed by the design team, the best one was selected using a decision matrix. Preliminary design of the best concept include support structure, stress analyses, and thermal performance. In addition, the team developed ideas for removing lunar dust from the shading device. The final design solution consisted of a winged radiator shading system with a rail support structure and a scissors mechanism for deployment. The total radiator area required was calculated to be 389 sq m for the 10 kW heat rejection system and 973 sq m for the 25 kW heat rejection system.

Barron, Carlos↗
NASA NTRS1992

Development of advanced high temperature in-cylinder components and tribological systems for low heat rejection diesel engines, phase 1

Analysis and concept design work completed in Phase 1 have identified a low heat rejection engine configuration with the potential to meet the Heavy Duty Transport Technology program specific fuel consumption goal of 152 g/kW-hr. The proposed engine configuration incorporates low heat rejection, in-cylinder components designed for operation at 24 MPa peak cylinder pressure. Water cooling is eliminated by selective oil cooling of the components. A high temperature lubricant will be required due to increased in-cylinder operating temperatures. A two-stage turbocharger air system with intercooling and aftercooling was selected to meet engine boost and BMEP requirements. A turbocompound turbine stage is incorporated for exhaust energy recovery. The concept engine cost was estimated to be 43 percent higher compared to a Caterpillar 3176 engine. The higher initial engine cost is predicted to be offset by reduced operating costs due the lower fuel consumption.

Kroeger, C. A.↗
NASA NTRS1989

Advanced radiator concepts

Two radiator systems to reject heat from future space nuclear power systems were conceptually designed. One design would dissipate 1.7 MWt of heat at 600 K, and the other would reject 2.4 MWt at 875 K. The low-temperature radiator utilized a pumped loop system constructed of titanium, and achieved a specific mass of 5.8 kg/sq m, including pumps and structure. The high-temperature radiator system utilized potassium heat pipes constructed of SiC-reinforced titanium, and achieved a specific mass of 5.5 kg/sq m. Both radiators took advantage of light, high-thermal-conductivity carbon/graphite composite fins to distribute and radiate the rejected heat.

Begg, L. L.↗
NASA NTRS1979

Heat rejection development requirements for orbital power systems

Long-term orbital applications in which large amounts of electrical power are generated and utilized will require waste heat rejection beyond the capabilities of existing radiator systems. Therefore, it will be necessary to develop a new concept for large orbital energy systems. This paper presents a discussion of the primary factors to be considered in the design and development of a large orbital power module heat rejection system. Specific design requirements are defined based on these factors, and a review of supporting technology in the advanced space radiator area is presented. A proposed development approach and a candidate baseline heat rejection system are then identified and discussed.

Rankin, J. G.↗
NASA NTRS1992

Thermal control systems for low-temperature heat rejection on a lunar base

One of the important issues in the lunar base architecture is the design of a Thermal Control System (TCS) to reject the low temperature heat from the base. The TCS ensures that the base and all components inside are maintained within the operating temperature range. A significant portion of the total mass of the TCS is due to the radiator. Shading the radiation from the sun and the hot lunar soil could decrease the radiator operating temperature significantly. Heat pumps have been in use for terrestrial applications. To optimize the mass of the heat pump augmented TCS, all promising options have to be evaluated and compared. Careful attention is given to optimizing system operating parameters, working fluids, and component masses. The systems are modeled for full load operation.

Sridhar, K. R.↗
NASA NTRS1993

Thermal control systems for low-temperature heat rejection on a lunar base

One of the important issues in the design of a lunar base is the thermal control system (TCS) used to reject low-temperature heat from the base. The TCS ensures that the base and the components inside are maintained within an acceptable temperature range. The temperature of the lunar surface peaks at 400 K during the 336-hour lunar day. Under these circumstances, direct dissipation of waste heat from the lunar base using passive radiators would be impractical. Thermal control systems based on thermal storage, shaded radiators, and heat pumps have been proposed. Based on proven technology, innovation, realistic complexity, reliability, and near-term applicability, a heat pump-based TCS was selected as a candidate for early missions. In this report, Rankine-cycle heat pumps and absorption heat pumps (ammonia water and lithium bromide-water) have been analyzed and optimized for a lunar base cooling load of 100 kW.

Sridhar, K. R.↗
NASA NTRS2018

Investigation into Cryogenic Tank Insulation Systems for the Mars Surface Environment

In order to use oxygen that is produced on the surface of Mars from In-Situ production processes in a chemical propulsion system, the oxygen must first be converted from vapor phase to liquid phase and then stored within the propellant tanks of the propulsion system. The oxygen must then be stored in the liquid phase for several years between when the liquefaction operations are initiated and when the ascent stage lifts off the Martian surface. Since the Space Exploration Initiative, NASA has been investing small sums of money into soft vacuum systems for Mars Applications. A study was done into these various insulation systems for soft vacuum insulation, to determine what types of systems might be best to further pursue. Five different architectures or cycles were considered: Aerogel-based multilayer Insulation (MLAI), Space Evacuated Mars Vacuum Jacket (SEMOV) (also known as lightweight vacuum jacket), Load Responsive-Multilayer Insulation, Spray on Foam with MLI, and MLAI in SEMOV. Models of each architecture were developed to give insight into the performance and losses of each of the options. The results were then compared across six categories: Insulation System Mass, Active System Power (both input and heat rejection), Insulation System Cost, Manufacturability, Reliability, and Operational Flexibility. The result was that a trade between reliability and mass was clearly identified. Systems with high mass, also had high perceived reliability; whereas, systems with lower mass and power had a much lower perceived reliability. In the end, the numerical trades of these systems showed nominally identical rankings. As a result it is recommended that NASA focus its Martian insulation development activities on demonstrating and improving the reliability of the lightweight identified systems.

Soft vacuum↗
NASA NTRS1985

Space heat rejection radiatiors: Meteoroid/debris consideration

The effect of the meteoroid/space debris environment on the design and evolution of spacecraft waste heat rejection radiator systems is discussed. Active radiator systems; i.e., systems in which waste heat is collected from the various heat sources within a spacecraft and delivered to the radiator system by a heat transport loop are emphasized. The heat is distributed over the radiator area and thus rejected to space. Present and future systems are discussed.

Rankin, J. G.↗
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 NTRS1972

Advanced spacecraft fuel cell systems

The development and characteristics of advanced spacecraft fuel cell systems are discussed. The system is designed to operate on low pressure, propulsion grade hydrogen and oxygen. The specific goals are 10,000 hours of operation with refurbishment, 20 pounds per kilowatt at a sustained power of 7 KW, and 21 KW peaking capability for durations of two hours. The system rejects waste heat to the spacecraft cooling system at power levels up to 7 KW. At higher powers, the system automatically transfers to open cycle operation with overboard steam venting.

Thaller, L. H.↗
NASA NTRS1991

From orbital debris capture systems through internal combustion engines on Mars

The investigation and conceptualization of an orbital debris collector was the primary area of design. In addition, an alternate structural design for Space Station Freedom and systems supporting resource utilization at Mars and the moon were studied. Hardware for production of oxygen from simulate Mars atmosphere was modified to permit more reliable operation at low pressures (down to 10 mb). An internal combustion engine was altered to study how Mars atmosphere could be used as a diluent to control combustion temperatures and avoid excess Mars propellant production requirements that would result from either methane-rich or oxygen-rich, methane-oxygen combustion. An elastic loop traction system that could be used for lunar construction vehicles was refined to permit testing. A parabolic heat rejection radiator system was designed and built to determine whether it was capable of increasing heat rejection rates during lunar daytime operation. In addition, an alternate space station truss design, utilizing a pre-integrated concept, was studied and found to reduce estimate extravehicular activity (EVA) time and increase the structural integrity when compared to the original Warren truss concept. An orbital-debris-capturing spacecraft design which could be mated with the Orbital Maneuvering Vehicle was studied. The design identified Soviet C-1B boosters as the best targets of opportunity in Earth orbits between an altitude of 900 km and 1100 km and at an inclination of 82.9 deg. A dual robot pallet, which could be spun to match the tumbling rate of the C-1B booster, was developed as the conceptual design.

Source record↗
NASA NTRS1987

Development of an emulation-simulation thermal control model for space station application

An improved capability for comparing various techniques for thermal management in the Space Station was developed. Current planning for the orbiting space station calls for a dual keel configuration. The thermal control system (TCS) for the space is composed of a central TCS and internal thermal control systems for the modules, as well as service facilities and attached payloads. The internal TCS may be attacted to the central TCS through a thermal bus. The central TCS is composed of a main transport system which collects waste thermal energy from each of the modules and transports it through coolant lines to the main rejection system. The waste heat loads in the modules arise from electrical and electronic equipment as well as metabolic loads in the manned modules. Several candidate technologies are being considered for acquiring the waste heat loads, for transporting the thermal energy between the acquisition and rejection systems, and for rejecting the waste heat to space. The analysis techniques described were developed for use in evaluating reliability, weights, costs, volumes, and power requirements for configurations using different candidates and different mission parameters.

Hartley, James G.↗
DOE OSTI2026

Hybrid Fuel Cell Systems for Heavy-Duty Trucks: Configuration, Heat Rejection, and Performance

Low-temperature polymer electrolyte membrane fuel cell systems can achieve higher efficiency than diesel engines, but heat rejection remains a major challenge in class-8 heavy-duty fuel cell trucks. For the same rated power, the radiator heat load is greater than that in a diesel engine, while the allowable operating temperatures are lower. This work proposes and evaluates 400 kWe fuel cell–battery hybrid (FCH) platforms and operating strategies that manage heat rejection without enlarging the radiator frontal area. Three FCH platforms are identified, each varying in fuel cell system (FCS) rated power, battery energy storage system (ESS) capacity, and maximum stack coolant exit temperature (T h1 ). All three satisfy key system and vehicle requirements, including 175 kWe FCS power at top sustained speed, 400 kWe FCH power on a 6% grade climb, a target stack power density (PD) of 750 mW e /cm 2 , and heat rejection constraints. The first FCH has the smallest FCS, the largest ESS, and a T h1 of 90 °C. The second achieves the highest PD of 840 mW e /cm 2 at a T h1 of 95 °C. The third has the largest FCS, the smallest ESS, and a T h1 of 102 °C. At a Th1 of 115 °C, the platform can be configured as a stand-alone 400 kWe(net) FCS without hybridization, but the achievable PD drops to 460 mW e /cm 2 .

25 ENERGY STORAGE↗
NASA NTRS2004

Weight Optimization of Active Thermal Management Using a Novel Heat Pump

Efficient lightweight power generation and thermal management are two important aspects for space applications. Weight is added to the space platforms due to the inherent weight of the onboard power generation equipment and the additional weight of the required thermal management systems. Thermal management of spacecraft relies on rejection of heat via radiation, a process that can result in large radiator mass, depending upon the heat rejection temperature. For some missions, it is advantageous to incorporate an active thermal management system, allowing the heat rejection temperature to be greater than the load temperature. This allows a reduction of radiator mass at the expense of additional system complexity. A particular type of active thermal management system is based on a thermodynamic cycle, developed by the authors, called the Solar Integrated Thermal Management and Power (SITMAP) cycle. This system has been a focus of the authors research program in the recent past (see Fig. 1). One implementation of the system requires no moving parts, which decreases the vibration level and enhances reliability. Compression of the refrigerant working fluid is accomplished in this scheme via an ejector.

Lear, William E.↗
NASA NTRS1992

Thermal control systems for low-temperature heat rejection on a lunar base

In this report, Rankine-cycle heat pumps and absorption heat pumps (ammonia-water and lithium bromide-water) have been analyzed and optimized for a lunar base cooling load of 100 kW. For the Rankine cycle, a search of several commonly used commercial refrigerants provided R11 and R717 as possible working fluids. Hence, the Rankine-cycle analysis has been performed for both R11 and R717. Two different configurations were considered for the system--one in which the heat pump is directly connected to the rejection loop and another in which a heat exchanger connects the heat pump to the rejection loop. For a marginal increase in mass, the decoupling of the rejection loop and the radiator from the heat pump provides greater reliability of the system and better control. Hence, the decoupled system is the configuration of choice. The optimal TCS mass for a 100 kW cooling load at 270 K was 5940 kg at a radiator temperature of 362 K. R11 was the working fluid in the heat pump, and R717 was the transport fluid in the rejection loop. Two TCS's based on an absorption-cycle heat pump were considered, one with an ammonia-water mixture and the other with a lithium bromide-water mixture as the working fluid. A complete cycle analysis was performed for these systems. The system components were approximated as heat exchangers with no internal pressure drop for the mass estimate. This simple approach underpredicts the mass of the systems, but is a good 'optimistic' first approximation to the TCS mass in the absence of reliable component mass data. The mass estimates of the two systems reveal that, in spite of this optimistic estimate, the absorption heat pumps are not competitive with the Rankine-cycle heat pumps. Future work at the systems level will involve similar analyses for the Brayton- and Stirling-cycle heat pumps. The analyses will also consider the operation of the pump under partial-load conditions. On the component level, a capillary evaporator will be designed, built, and tested in order to investigate its suitability in lunar base TCS and microgravity two-phase applications.

Sridhar, K. R.↗
NASA NTRS2006

Heat Rejection Concepts for Lunar Fission Surface Power Applications

This paper describes potential heat rejection design concepts for lunar surface Brayton power conversion systems. Brayton conversion systems are currently under study by NASA for surface power applications. Surface reactors may be used for the moon to power human outposts enabling extended stays and closed loop life support. The Brayton Heat Rejection System (HRS) must dissipate waste heat generated by the power conversion system due to inefficiencies in the thermal-to-electric conversion process. Space Brayton conversion system designs tend to optimize at efficiencies of about 20 to 25 percent with radiator temperatures in the 400 K to 600 K range. A notional HRS was developed for a 100 kWe-class Brayton power system that uses a pumped water heat transport loop coupled to a water heat pipe radiator. The radiator panels employ a tube and fin construction consisting of regularly-spaced circular heat pipes contained within two composite facesheets. The water heat pipes interface to the coolant through curved sections partially contained within the cooling loop. The paper evaluates various design parameters including radiator panel orientation, coolant flow path, and facesheet thickness. Parameters were varied to compare design options on the basis of H2O pump pressure rise and required power, heat pipe unit power and radial flux, radiator area, radiator panel areal mass, and overall HRS mass.

Siamidis, John↗
NASA NTRS2017

Mars Propellant Liquefaction and Storage Performance Modeling using Thermal Desktop with an Integrated Cryocooler Model

NASAs current Mars architectures are assuming the production and storage of 23 tons of liquid oxygen on the surface of Mars over a duration of 500+ days. In order to do this in a mass efficient manner, an energy efficient refrigeration system will be required. Based on previous analysis NASA has decided to do all liquefaction in the propulsion vehicle storage tanks. In order to allow for transient Martian environmental effects, a propellant liquefaction and storage system for a Mars Ascent Vehicle (MAV) was modeled using Thermal Desktop. The model consisted of a propellant tank containing a broad area cooling loop heat exchanger integrated with a reverse turbo Brayton cryocooler. Cryocooler sizing and performance modeling was conducted using MAV diurnal heat loads and radiator rejection temperatures predicted from a previous thermal model of the MAV. A system was also sized and modeled using an alternative heat rejection system that relies on a forced convection heat exchanger. Cryocooler mass, input power, and heat rejection for both systems were estimated and compared against sizing based on non-transient sizing estimates.

fluid modeling↗