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NASA NTRS1999

Thermal and Mechanical Microspacecraft Technologies for Deep Space Systems Program X2000 Future Deliveries

Thermal and mechanical technologies are an important part of the Deep Space Systems Technology (DSST) Program X2000 Future Deliveries (FD) microspacecraft. A wide range of future space missions are expected to utilize the technologies and the architecture developed by DSST FD. These technologies, besides being small in physical size, make the tiny spacecraft robust and flexible. The DSST FD architecture is designed to be highly reliable and suitable for a wide range of missions such as planetary landers/orbiters/flybys, earth orbiters, cometary flybys/landers/sample returns, etc. Two of the key ideas used in the development of thermal and mechanical technologies and architectures are: 1) to include several of the thermal and mechanical functions in any given single spacecraft element and 2) the architecture be modular so that it can easily be adapted to any of the future missions. One of the thermal architectures being explored for the DSST FD microspacecraft is the integrated thermal energy management of the complete spacecraft using a fluid loop. The robustness and the simplicity of the loop and the flexibility with which it can be integrated in the spacecraft have made it attractive for applications to DSST FD. Some of the thermal technologies to be developed as a part of this architecture are passive and active cooling loops, electrically variable emittance surfaces, miniature thermal switches, and specific high density electronic cooling technologies. In the mechanical area, multifunction architecture for the structural elements will be developed. The multifunction aspect is expected to substantially reduce the mass and volume of the spacecraft. Some of the technologies that will be developed are composite material panels incorporating electronics, cabling, and thermal elements in them. The paper describes the current state of the technologies and progress to be made in the thermal and mechanical technologies and approaches for the DSST Future Deliveries microspacecraft.

Birur, Gajanana C.↗

NASA NTRS2001

Solar-Thermal Engine Testing

A solar-thermal engine serves as a high-temperature solar-radiation absorber, heat exchanger, and rocket nozzle. collecting concentrated solar radiation into an absorber cavity and transferring this energy to a propellant as heat. Propellant gas can be heated to temperatures approaching 4,500 F and expanded in a rocket nozzle, creating low thrust with a high specific impulse (I(sub sp)). The Shooting Star Experiment (SSE) solar-thermal engine is made of 100 percent chemical vapor deposited (CVD) rhenium. The engine 'module' consists of an engine assembly, propellant feedline, engine support structure, thermal insulation, and instrumentation. Engine thermal performance tests consist of a series of high-temperature thermal cycles intended to characterize the propulsive performance of the engines and the thermal effectiveness of the engine support structure and insulation system. A silicone-carbide electrical resistance heater, placed inside the inner shell, substitutes for solar radiation and heats the engine. Although the preferred propellant is hydrogen, the propellant used in these tests is gaseous nitrogen. Because rhenium oxidizes at elevated temperatures, the tests are performed in a vacuum chamber. Test data will include transient and steady state temperatures on selected engine surfaces, propellant pressures and flow rates, and engine thrust levels. The engine propellant-feed system is designed to Supply GN2 to the engine at a constant inlet pressure of 60 psia, producing a near-constant thrust of 1.0 lb. Gaseous hydrogen will be used in subsequent tests. The propellant flow rate decreases with increasing propellant temperature, while maintaining constant thrust, increasing engine I(sub sp). In conjunction with analytical models of the heat exchanger, the temperature data will provide insight into the effectiveness of the insulation system, the structural support system, and the overall engine performance. These tests also provide experience on operational aspects of the engine and associated subsystems, and will include independent variation of both steady slate heat-exchanger temperature prior to thrust operation and nitrogen inlet pressure (flow rate) during thrust operation. Although the Shooting Star engines were designed as thermal-storage engines to accommodate mission parameters, they are fully capable of operating as scalable, direct-gain engines. Tests are conducted in both operational modes. Engine thrust and propellant flow rate will be measured and thereby I(sub sp). The objective of these tests is to investigate the effectiveness of the solar engine as a heat exchanger and a rocket. Of particular interest is the effectiveness of the support structure as a thermal insulator, the integrity of both the insulation system and the insulation containment system, the overall temperature distribution throughout the engine module, and the thermal power required to sustain steady state fluid temperatures at various flow rates.

Tucker, Stephen↗

NASA NTRS2002

Thermal Gradient Cyclic Behavior of a Thermal/Environmental Barrier Coating System on SiC/SiC Ceramic Matrix Composites

Thermal barrier and environmental barrier coatings (TBCs and EBCs) will play a crucial role in future advanced gas turbine engines because of their ability to significantly extend the temperature capability of the ceramic matrix composite (CMC) engine components in harsh combustion environments. In order to develop high performance, robust coating systems for effective thermal and environmental protection of the engine components, appropriate test approaches for evaluating the critical coating properties must be established. In this paper, a laser high-heat-flux, thermal gradient approach for testing the coatings will be described. Thermal cyclic behavior of plasma-sprayed coating systems, consisting of ZrO2-8wt%Y2O3 thermal barrier and NASA Enabling Propulsion Materials (EPM) Program developed mullite+BSAS/Si type environmental barrier coatings on SiC/SiC ceramic matrix composites, was investigated under thermal gradients using the laser heat-flux rig in conjunction with the furnace thermal cyclic tests in water-vapor environments. The coating sintering and interface damage were assessed by monitoring the real-time thermal conductivity changes during the laser heat-flux tests and by examining the microstructural changes after the tests. The coating failure mechanisms are discussed based on the cyclic test results and are correlated to the sintering, creep, and thermal stress behavior under simulated engine temperature and heat flux conditions.

Zhu, Dongming↗

NASA NTRS2002

Non-Thermal Hard X-Ray Emission in Galaxy Clusters Observed with the BeppoSAX PDS

We studied the X-ray emission in a sample of clusters using the BeppoSAX PDS instrument in the 20 -- 80 keV energy band. We estimated the non-thermal cluster emission (HXR) by modeling the thermal contribution from the cluster gas and the non-thermal contamination from the AGN in the field, and propagating the corresponding uncertainties. We also evaluated and propagated the systematic uncertainties due to the background fluctuations. The resulting non-thermal component is detected at a sigma level in approx. 50 % of the non-significantly AGN-contaminated clusters, i.e. in clusters A2142, A2256, A3376, Coma, Ophiuchus and Virgo. Furthermore, Virgo is detected at a 4 sigma level. All the clusters detected at a 2 sigma level exhibit some degree of merger signatures, i.e. deviations from the azimuthally symmetric brightness and temperature distributions, while the relaxed clusters are detected at a lower confidence. The data are consistent with a scenario whereby relaxed clusters have no non-thermal hard X-ray component, whereas merger clusters do, with a 20 -- 80 keV luminosity of approx. 10(exp 42-44)((h(sub 50))(exp -2))(erg/s). Consistent with merger boosting of cluster temperatures, the non-thermal luminosity increases by 2-3 orders of magnitude between the average cluster temperatures 2 and 10 keV, as L(sub NTE) is proportional to T(sup j) with j = 2.4+/-0.3. These results corroborate the assumption which is the essential element in most non-thermal hard X-ray emission models. The co-added spectrum of all non-significantly AGN-contaminated clusters indicates a power-law spectrum for the non-thermal component with a photon index of 1.5+/-0.25 at 1 sigma confidence level. Unless there is a high energy cut-off in the electron velocity distribution, the total spectrum implies that Inverse Compton scatter of Cosmic Microwave Background photons from electron population dominates over the non-thermal bremsstrahlung in producing hard X-rays in clusters on the merger shock acceleration of electrons in clusters.

Nevalainen, Jukka H.↗

NASA NTRS2003

Nonlinear Thermal Analyses of a Liquid Hydrogen Tank Wall

A thermal evaluation of a composite tank wall design for a liquid hydrogen tank was performed in the present study. The primary focus of the current effort was to perform one-dimensional, temperature nonlinear, transient thermal analyses to determine the through-the-thickness temperature profiles. These profiles were used to identify critical points within the flight envelope that could have detrimental effects on the adhesive bondlines used in the construction of the tank wall. Additionally, this paper presents the finite element models, analysis strategies, and thermal analysis results that were determined for several vehicle flight conditions. The basic tank wall configuration used to perform the thermal analyses consisted of carbon-epoxy facesheets and a Korex honeycomb core sandwich that was insulated with an Airex cryogenic foam and an Alumina Enhanced Thermal Barrier (AETB-12). Nonlinear, transient thermal analyses were conducted using the ABAQUS finite element code. Tank wall models at a windward side location on the fuel tank were analyzed for three basic flight conditions: cold-soak (ground-hold), ascent, and re-entry. Additionally, three ambient temperature boundary conditions were applied to the tank wall for the cold-soak condition, which simulated the launch pad cooldown process. Time-dependent heating rates were used in the analyses of the ascent and reentry segments of the flight history along with temperature dependent material properties. The steady-state through-the-thickness temperature profile from the cold-soak condition was used as the initial condition for the ascent analyses. Results from the nonlinear thermal analyses demonstrated very good correlation with results from similar models evaluated by Northrop- Grumman using a different analysis tool. Wall through-the-thickness temperature gradients as a function of flight time were obtained for future incorporation into a full-scale thermostructural analysis to evaluate the adhesive bondlines. As a result of the thermal analyses conducted, a sufficient level of confidence was demonstrated in the thermal modeling and analysis capabilities of ABAQUS to warrant future use as a thermo-structural analysis tool to evaluate cryogenic tank wall designs.

Smeltzer, Stanley S., III↗

NASA NTRS2004

Development and Fatigue Testing of Ceramic Thermal Barrier Coatings

Ceramic thermal barrier coatings will play an increasingly important role in future gas turbine engines because of their ability to effectively protect the engine components and further raise engine temperatures. Durability of the coating systems remains a critical issue with the ever-increasing temperature requirements. Thermal conductivity increase and coating degradation due to sintering and phase changes are known to be detrimental to coating performance. There is a need to characterize the coating thermal fatigue behavior and temperature limit, in order to potentially take full advantage of the current coating capability. In this study, thermal conductivity and cyclic fatigue behaviors of plasma-sprayed ZrO2-8wt%Y2O3 thermal barrier coatings were evaluated under high temperature, large thermal gradient and thermal cycling conditions. The coating degradation and failure processes were assessed by real-time monitoring of the coating thermal conductivity under the test conditions. The ceramic coating crack initiation and propagation driving forces and failure modes under the cyclic thermal loads will be discussed in light of the high temperature mechanical fatigue and fracture testing results.

Zhu, Dong-Ming↗

NASA NTRS2001

Phase change thermal control materials, method and apparatus

An apparatus and method for metabolic cooling and insulation of a user in a cold environment. In its preferred embodiment the apparatus is a highly flexible composite material having a flexible matrix containing a phase change thermal storage material. The apparatus can be made to heat or cool the body or to act as a thermal buffer to protect the wearer from changing environmental conditions. The apparatus may also include an external thermal insulation layer and/or an internal thermal control layer to regulate the rate of heat exchange between the composite and the skin of the wearer. Other embodiments of the apparatus also provide 1) a path for evaporation or direct absorption of perspiration from the skin of the wearer for improved comfort and thermal control, 2) heat conductive pathways within the material for thermal equalization, 3) surface treatments for improved absorption or rejection of heat by the material, and 4) means for quickly regenerating the thermal storage capacity for reuse of the material. Applications of the composite materials are also described which take advantage of the composite's thermal characteristics. The examples described include a diver's wet suit, ski boot liners, thermal socks, gloves and a face mask for cold weather activities, and a metabolic heating or cooling blanket useful for treating hypothermia or fever patients in a medical setting and therapeutic heating or cooling orthopedic joint supports.

Buckley, Theresa M.↗

NASA NTRS1999

Flexible composite material with phase change thermal storage

A highly flexible composite material having a flexible matrix containing a phase change thermal storage material. The composite material can be made to heat or cool the body or to act as a thermal buffer to protect the wearer from changing environmental conditions. The composite may also include an external thermal insulation layer and/or an internal thermal control layer to regulate the rate of heat exchange between the composite and the skin of the wearer. Other embodiments of the PCM composite also provide 1) a path for evaporation or direct absorption of perspiration from the skin of the wearer for improved comfort and thermal control, 2) heat conductive pathways within the material for thermal equalization, 3) surface treatments for improved absorption or rejection of heat by the material, and 4) means for quickly regenerating the thermal storage capacity for reuse of the material. Applications of the composite materials are also described which take advantage of the composite's thermal characteristics. The examples described include a diver's wet suit, ski boot liners, thermal socks, gloves and a face mask for cold weather activities, and a metabolic heating or cooling blanket useful for treating hypothermia or fever patients in a medical setting and therapeutic heating or cooling orthopedic joint supports.

Buckley, Theresa M.↗

NASA NTRS2001

Flexible composite material with phase change thermal storage

A highly flexible composite material having a flexible matrix containing a phase change thermal storage material. The composite material can be made to heat or cool the body or to act as a thermal buffer to protect the wearer from changing environmental conditions. The composite may also include an external thermal insulation layer and/or an internal thermal control layer to regulate the rate of heat exchange between the composite and the skin of the wearer. Other embodiments of the PCM composite also provide 1) a path for evaporation or direct absorption of perspiration from the skin of the wearer for improved comfort and thermal control, 2) heat conductive pathways within the material for thermal equalization, 3) surface treatments for improved absorption or rejection of heat by the material, and 4) means for quickly regenerating the thermal storage capacity for reuse of the material. Applications of the composite materials are also described which take advantage of the composite's thermal characteristics. The examples described include a diver's wet suit, ski boot liners, thermal socks, ,gloves and a face mask for cold weather activities, and a metabolic heating or cooling blanket useful for treating hypothermia or fever patients in a medical setting and therapeutic heating or cooling orthopedic joint supports.

Buckley, Theresa M.↗

NASA NTRS2005

Phase change material thermal capacitor clothing

Buckley, Theresa M.↗

NASA NTRS2008

Current Issues in Human Spacecraft Thermal Control Technology

Efficient thermal management of Earth-orbiting human spacecraft, lunar transit spacecraft and landers, as well as a lunar habitat will require advanced thermal technology. These future spacecraft will require more sophisticated thermal control systems that can dissipate or reject greater heat loads at higher input heat fluxes while using fewer of the limited spacecraft mass, volume and power resources. The thermal control designs also must accommodate the harsh environments associated with these missions including dust and high sink temperatures. The lunar environment presents several challenges to the design and operation of active thermal control systems. During the Apollo program, landings were located and timed to occur at lunar twilight, resulting in a benign thermal environment. The long duration polar lunar bases that are foreseen in 15 years will see extremely cold thermal environments. Long sojourns remote from low-Earth orbit will require lightweight, but robust and reliable systems. Innovative thermal management components and systems are needed to accomplish the rejection of heat from lunar bases. Advances are required in the general areas of radiators, thermal control loops and equipment. Radiators on the Moon's poles must operate and survive in very cold environments. Also, the dusty environment of an active lunar base may require dust mitigation and removal techniques to maintain radiator performance over the long term.

Ungar, Eugene K.↗

NASA NTRS2009

CEQATR Thermal Test Overview

A thermal test overview of the Constellation Environmental Qualification and Acceptance Test Requirement (CEQATR) is presented. The contents include: 1) CEQATR Thermal Test Overview; 2) CxP Environments; 3) CEQATR Table 1.2-1; 4) Levels of Assembly; 5) Definitions for Levels of Assembly; 6) Hardware Applicability; 7) CEQATR Thermal-Related Definitions; 8) Requirements for unit-level thermal testing; 9) Requirements for major assembly level thermal testing; 10) General thermal testing requirements; 11) General thermal cycle, thermal vacuum profiles; 12) Test tolerances; 13) Vacuum vs Ambient; 14) Thermal Gradient; 15) Sequence of Testing; 16) Alternative Strategies; 17) Protoflight; 18) Halt/Hass; 19) Humidity; and 20) Tailoring.

Balusek, Alan R.↗

NASA NTRS2010

Analysis of Solar-Heated Thermal Wadis to Support Extended-Duration Lunar Exploration

The realization of the renewed exploration of the Moon presents many technical challenges; among them is the survival of lunar surface assets during periods of darkness when the lunar environment is very cold. Thermal wadis are engineered sources of stored solar energy using modified lunar regolith as a thermal storage mass that can enable the operation of lightweight robotic rovers or other assets in cold, dark environments without incurring potential mass, cost, and risk penalties associated with various onboard sources of thermal energy. Thermal wadi-assisted lunar rovers can conduct a variety of long-duration missions including exploration site surveys; teleoperated, crew-directed, or autonomous scientific expeditions; and logistics support for crewed exploration. This paper describes a thermal analysis of thermal wadi performance based on the known solar illumination of the moon and estimates of producible thermal properties of modified lunar regolith. Analysis was performed for the lunar equatorial region and for a potential Outpost location near the lunar south pole. The results are presented in some detail in the paper and indicate that thermal wadis can provide the desired thermal energy reserve, with significant margin, for the survival of rovers or other equipment during periods of darkness.

Balasubramaniam, R.↗

NASA NTRS2010

Multidimensional Tests of Thermal Protection Materials in the Arcjet Test Facility

Many thermal protection system materials used for spacecraft heatshields have anisotropic thermal properties, causing them to display significantly different thermal characteristics in different directions, when subjected to a heating environment during flight or arcjet tests. This paper investigates the effects of sidewall heating coupled with anisotropic thermal properties of thermal protection materials in the arcjet environment. Phenolic Impregnated Carbon Ablator (PICA) and LI-2200 materials (the insulation material of Shuttle tiles) were used for this study. First, conduction-based thermal response simulations were carried out, using the Marc.Mentat finite element solver, to study the effects of sidewall heating on PICA arcjet coupons. The simulation showed that sidewall heating plays a significant role in thermal response of these models. Arcjet tests at the Aerodynamic Heating Facility (AHF) at NASA Ames Research Center were performed later on instrumented coupons to obtain temperature history at sidewall and various radial locations. The details of instrumentation and experimental technique are the prime focus of this paper. The results obtained from testing confirmed that sidewall heating plays a significant role in thermal response of these models. The test results were later used to verify the two-dimensional ablation, thermal response, and sizing program, TITAN. The test data and model predictions were found to be in excellent agreement

Agrawal, Parul↗

NASA NTRS2010

Multidimensional Testing of Thermal Protection Materials in the Arcjet Test Facility

Many thermal protection system materials used for spacecraft heatshields have anisotropic thermal properties, causing them to display significantly different thermal characteristics in different directions, when subjected to a heating environment during flight or arcjet tests. The anisotropic effects are enhanced in the presence of sidewall heating. This paper investigates the effects of anisotropic thermal properties of thermal protection materials coupled with sidewall heating in the arcjet environment. Phenolic Impregnated Carbon Ablator (PICA) and LI-2200 materials (the insulation material of Shuttle tiles) were used for this study. First, conduction-based thermal response simulations were carried out, using the Marc.Mentat finite element solver, to study the effects of sidewall heating on PICA arcjet coupons. The simulation showed that sidewall heating plays a significant role in thermal response of these models. Arcjet tests at the Aerodynamic Heating Facility (AHF) at NASA Ames Research Center were performed later on instrumented coupons to obtain temperature history at sidewall and various radial locations. The details of instrumentation and experimental technique are the prime focus of this paper. The results obtained from testing confirmed that sidewall heating plays a significant role in thermal response of these models. The test results were later used to validate the two-dimensional ablation, thermal response, and sizing program, TITAN. The test data and model predictions were found to be in excellent agreement

Agrawal, Parul↗

NASA NTRS2010

Shape memory thermal conduction switch

A thermal conduction switch includes a thermally-conductive first member having a first thermal contacting structure for securing the first member as a stationary member to a thermally regulated body or a body requiring thermal regulation. A movable thermally-conductive second member has a second thermal contacting surface. A thermally conductive coupler is interposed between the first member and the second member for thermally coupling the first member to the second member. At least one control spring is coupled between the first member and the second member. The control spring includes a NiTiFe comprising shape memory (SM) material that provides a phase change temperature <273 K, a transformation range <40 K, and a hysteresis of <10 K. A bias spring is between the first member and the second member. At the phase change the switch provides a distance change (displacement) between first and second member by at least 1 mm, such as 2 to 4 mm.

Vaidyanathan, Rajan↗

NASA NTRS2011

Design and Test of Advanced Thermal Simulators for an Alkali Metal-Cooled Reactor Simulator

The Early Flight Fission Test Facility (EFF-TF) at NASA Marshall Space Flight Center (MSFC) has as one of its primary missions the development and testing of fission reactor simulators for space applications. A key component in these simulated reactors is the thermal simulator, designed to closely mimic the form and function of a nuclear fuel pin using electric heating. Continuing effort has been made to design simple, robust, inexpensive thermal simulators that closely match the steady-state and transient performance of a nuclear fuel pin. A series of these simulators have been designed, developed, fabricated and tested individually and in a number of simulated reactor systems at the EFF-TF. The purpose of the thermal simulators developed under the Fission Surface Power (FSP) task is to ensure that non-nuclear testing can be performed at sufficiently high fidelity to allow a cost-effective qualification and acceptance strategy to be used. Prototype thermal simulator design is founded on the baseline Fission Surface Power reactor design. Recent efforts have been focused on the design, fabrication and test of a prototype thermal simulator appropriate for use in the Technology Demonstration Unit (TDU). While designing the thermal simulators described in this paper, effort were made to improve the axial power profile matching of the thermal simulators. Simultaneously, a search was conducted for graphite materials with higher resistivities than had been employed in the past. The combination of these two efforts resulted in the creation of thermal simulators with power capacities of 2300-3300 W per unit. Six of these elements were installed in a simulated core and tested in the alkali metal-cooled Fission Surface Power Primary Test Circuit (FSP-PTC) at a variety of liquid metal flow rates and temperatures. This paper documents the design of the thermal simulators, test program, and test results.

Garber, Anne E.↗

NASA NTRS2011

Autonomous Aerobraking: Thermal Analysis and Response Surface Development

A high-fidelity thermal model of the Mars Reconnaissance Orbiter was developed for use in an autonomous aerobraking simulation study. Response surface equations were derived from the high-fidelity thermal model and integrated into the autonomous aerobraking simulation software. The high-fidelity thermal model was developed using the Thermal Desktop software and used in all phases of the analysis. The use of Thermal Desktop exclusively, represented a change from previously developed aerobraking thermal analysis methodologies. Comparisons were made between the Thermal Desktop solutions and those developed for the previous aerobraking thermal analyses performed on the Mars Reconnaissance Orbiter during aerobraking operations. A variable sensitivity screening study was performed to reduce the number of variables carried in the response surface equations. Thermal analysis and response surface equation development were performed for autonomous aerobraking missions at Mars and Venus.

Dec, John A.↗