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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

An Improved Formulation for Calorimetric Emittance Testing of Spacecraft Thermal Control Coatings

Spacecraft often really heavily on passive thermal control to maintain operating temperature. An important parameter in the spacecraft heat balance equation is the emittance of thermal control coatings as a function of coating temperature. One method for determining the emittance of spacecraft thermal control from elevated temperature to cryogenic temperatures relies on a calorimetric technique. The fundamental equation governing this test method can be found in numerous places in the literature and although it generally provides reasonable results, its formulation is based on a conceptual flaw that only becomes apparent when the sample temperature approaches the wall temperature during testing. This paper investigates the cause for this error and develops the correct formulation for calorimetric emittance testing. Experimental data will also be presented that illustrates the difference between the two formulations and the resulting difference in the calculated emittance.

Kauder, Lonny R.↗
NASA NTRS1973

Handbook on passive thermal control coatings

A handbook of passive thermal control surfaces data pertaining to the heat transfer requirements of spacecraft is presented. Passive temperature control techniques and the selection of control surfaces are analyzed. The space environmental damage mechanisms in passive thermal control surfaces are examined. Data on the coatings for which technical information is available are presented in tabular form. Emphasis was placed on consulting only those references where the experimental simulation of the space environment appeared to be more appropriate.

Mookherji, T. K.↗
NASA NTRS2008

The Effect of Simulated Lunar Dust on the Absorptivity, Emissivity, and Operating Temperature on AZ-93 and Ag/FEP Thermal Control Surfaces

JSC-1AF lunar simulant has been applied to AZ-93 and AgFEP thermal control surfaces on aluminum or composite substrates in a simulated lunar environment. The temperature of these surfaces was monitored as they were heated with a solar simulator and cooled in a 30 K coldbox. Thermal modeling was used to determine the absorptivity ( ) and emissivity ( ) of the thermal control surfaces in both their clean and dusted states. Then, a known amount of power was applied to the samples while in the coldbox and the steady state temperatures measured. It was found that even a submonolayer of simulated lunar dust can significantly degrade the performance of both white paint and second-surface mirror type thermal control surfaces under these conditions. Contrary to earlier studies, dust was found to affect as well as . Dust lowered the emissivity by as much as 16 percent in the case of AZ-93, and raised it by as much as 11 percent in the case of AgFEP. The degradation of thermal control surface by dust as measured by / rose linearly regardless of the thermal control coating or substrate, and extrapolated to degradation by a factor 3 at full coverage by dust. Submonolayer coatings of dust were found to not significantly change the steady state temperature at which a shadowed thermal control surface will radiate.

Gaier, James R.↗
NASA NTRS1973

Thermal control of the solar electric propulsion stage

The thermal control requirements consist of functional requirements related to the various mission phase natural environments, operational requirements of induced power loadings by the solar electric propulsion stage subsystems, and design temperature limits for performance and reliability. The design approach utilizes passive thermal control techniques combining insulation, surface coatings, and sunshields with thermostatically controlled louvers. Heaters are used to regulate certain temperatures for extreme conditions. Details regarding the thruster array thermal control design are discussed, giving attention to the parameters used in the mathematical model, questions of conductive coupling, and thruster estimated power distribution.

Ruttner, L. E.↗
NASA NTRSDate not supplied

Thermal Control Coating Charging Tests for the Europa Clipper Mission

Thermal control coatings are an essential part of spacecraft systems. Typically, these coatings are directly exposed to the space environment, which means they must be designed and applied in a manner that can withstand exposure to radiation, ultra-violet light, thermal cycles, etc. In the case of the Europa Clipper mission, the thermal control coatings must also have relatively low resistivity in order to minimize electrostatic charging (and discharging) due to the abundant low-energy plasma populations in the Jovian environment. This paper will present the results of a test campaign led by NASA’s Jet Propulsion Laboratory (JPL) to find thermal control coating solutions that minimized the occurrence of electrostatic discharges – even at cryogenic temperature conditions. The customized laboratory test system at NASA’s Marshall Space Flight Center will be described, along with the test requirements and conditions. Finally, a summary of the coatings with the best electrostatic characteristics will be provided.

Thermal Control Coatings↗
NASA NTRS1974

Thermal control for planetary probes

Thermal control of planetary probes during the descent phase of a mission is examined. Emphasis was placed on the problems of atmospheric uncertainties to be encountered by the probes, insulation performance, and equipment temperature limits.

Mcmordie, R.↗
NASA NTRS1993

Selected results for LDEF thermal control coatings

Several different thermal control coatings were analyzed as part of the Long Duration Exposure Facility (LDEF) Materials Special Investigation Group activity and as part of the Space Environment Effects on Spacecraft Materials Experiment M0003. A brief discussion of the results obtained for these materials is presented.

Golden, Johnny L.↗
NASA NTRS2010

Long Duration Life Test of Propylene Glycol Water Based Thermal Fluid Within Thermal Control Loop

Evaluations of thermal properties and resistance to microbial growth concluded that 50% Propylene Glycol (PG)-based fluid and 50% de-ionized water mixture was desirable for use as a fluid within a vehicle s thermal control loop. However, previous testing with a commercial mixture of PG and water containing phosphate corrosion inhibitors resulted in corrosion of aluminum within the test system and instability of the test fluid. This paper describes a follow-on long duration testing and analysis of 50% Propylene Glycol (PG)-based fluid and 50% de-ionized water mixture with inorganic corrosion inhibitors used in place of phosphates. The test evaluates the long-term fluid stability and resistance to microbial and chemical changes

Le, Hung↗
NASA NTRS2014

A Morphing Radiator for High-Turndown Thermal Control of Crewed Space Exploration Vehicles

Spacecraft designed for missions beyond low earth orbit (LEO) face a difficult thermal control challenge, particularly in the case of crewed vehicles where the thermal control system (TCS) must maintain a relatively constant internal environment temperature despite a vastly varying external thermal environment and despite heat rejection needs that are contrary to the potential of the environment. A thermal control system may be required to reject a higher heat load to warm environments and a lower heat load to cold environments, necessitating a relatively high turndown ratio. A modern thermal control system is capable of a turndown ratio of on the order of 12:1, but crew safety and environment compatibility have constrained these solutions to massive multi-loop fluid systems. This paper discusses the analysis of a unique radiator design that employs the behavior of shape memory alloys (SMAs) to vary the turndown of, and thus enable, a single-loop vehicle thermal control system for space exploration vehicles. This design, a morphing radiator, varies its shape in response to facesheet temperature to control view of space and primary surface emissivity. Because temperature dependence is inherent to SMA behavior, the design requires no accommodation for control, instrumentation, or power supply in order to operate. Thermal and radiation modeling of the morphing radiator predict a turndown ranging from 11.9:1 to 35:1 independent of TCS configuration. Coupled thermal-stress analyses predict that the desired morphing behavior of the concept is attainable. A system level mass analysis shows that by enabling a single loop architecture this design could reduce the TCS mass by between 139 kg and 225 kg. The concept has been demonstrated in proof-of-concept benchtop tests.

Cognata, Thomas J.↗
NASA NTRS2015

A Morphing Radiator for High-Turndown Thermal Control of Crewed Space Exploration Vehicles

Spacecraft designed for missions beyond low earth orbit (LEO) face a difficult thermal control challenge, particularly in the case of crewed vehicles where the thermal control system (TCS) must maintain a relatively constant internal environment temperature despite a vastly varying external thermal environment and despite heat rejection needs that are contrary to the potential of the environment. A thermal control system is in other words required to reject a higher heat load to warm environments and a lower heat load to cold environments, necessitating a quite high turndown ratio. A modern thermal control system is capable of a turndown ratio of on the order of 12:1, but for crew safety and environment compatibility these are massive multi-loop fluid systems. This paper discusses the analysis of a unique radiator design which employs the behavior of shape memory alloys (SMA) to vary the turndown of, and thus enable, a single-loop vehicle thermal control system for space exploration vehicles. This design, a morphing radiator, varies its shape in response to facesheet temperature to control view of space and primary surface emissivity. Because temperature dependence is inherent to SMA behavior, the design requires no accommodation for control, instrumentation, nor power supply in order to operate. Thermal and radiation modeling of the morphing radiator predict a turndown ranging from 11.9:1 to 35:1 independent of TCS configuration. Stress and deformation analyses predict the desired morphing behavior of the concept. A system level mass analysis shows that by enabling a single loop architecture this design could reduce the TCS mass by between 139 kg and 225 kg. The concept is demonstrated in proof-of-concept benchtop tests.

Cognata, Thomas J.↗
NASA NTRS1992

Plasma effects on the passive external thermal control coating of Space Station Freedom

The current baseline chromic acid anodized thermal control coating on 6061-T6 aluminum meteoroid debris (M/D) shields for SSF has been evaluated. The degradation of the solar absorptance, alpha, and the thermal emittance, epsilon, of chromic acid anodized aluminum due to dielectric breakdown in plasma was measured to predict the on-orbit lifetime of the SSF M/D shields. The lifetime of the thermal control coating was based on the surface temperatures achieved with degradation of the thermal control properties, alpha and epsilon. The temperatures of each M/D shield from first element launch (FEL) through FEL+15 years were analyzed. It is shown that the baseline thermal control coating cannot withstand the -140 V potential between the conductive structure of the SSF and the current plasma environment.

Carruth, Ralph, Jr.↗
NASA NTRS1992

Conceptual design of a lunar base thermal control system

Space station and alternate thermal control technologies were evaluated for lunar base applications. The space station technologies consisted of single-phase, pumped water loops for sensible and latent heat removal from the cabin internal environment and two-phase ammonia loops for the transportation and rejection of these heat loads to the external environment. Alternate technologies were identified for those areas where space station technologies proved to be incompatible with the lunar environment. Areas were also identified where lunar resources could enhance the thermal control system. The internal acquisition subsystem essentially remained the same, while modifications were needed for the transport and rejection subsystems because of the extreme temperature variations on the lunar surface. The alternate technologies examined to accommodate the high daytime temperatures incorporated lunar surface insulating blankets, heat pump system, shading, and lunar soil. Other heat management techniques, such as louvers, were examined to prevent the radiators from freezing. The impact of the geographic location of the lunar base and the orientation of the radiators was also examined. A baseline design was generated that included weight, power, and volume estimates.

Simonsen, Lisa C.↗
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 NTRS1976

Long term performance of the advanced thermal control flight experiment

The Advanced Thermal Control Flight Experiment (ATFE) is providing the first zero-g flight data for the performance of a thermal diode heat pipe and an electrical Feedback Controlled Heat Pipe (FCHP). The temperature stability derived from the melting and freezing of octadecane is also being evaluated in this experiment. Flight data for the different operational modes are compared to ground test data, and the performance of the individual components is analyzed. Also, the ATFE's long term behavior from launch through September 30, 1975, is evaluated. The temperature control capability of a thermal diode, phase change material, and FCHP was demonstrated over more than 500 days of flight operation and each of these components should be considered in future applications. Both the diode and FCHP transport more than 20 W in the normal and passive modes during maximum conditions. Up to 30 W has been carried by the FCHP when the auxiliary heater is applied. The "OFF-conductance" of the diode has improved substantially in flight. A 1.4 W decrease in backflow has been realized by more complete blockage. The variable conductance behavior of the FCHP is essentially unchanged with time and relative to ground tests. Degradation of the radiator and reservoir surfaces have resulted in a partial loss of control by the FCHP during normal mode operation; however, manual control at an elevated set point has demonstrated the FCHP's ability to provide regulated temperature control to within + or - 2 C. The same pipe when operated as a passive VCHP controlled at 35 + or - 10 C with the same test conditions.

Brennan, P. J.↗