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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

NASA systems autonomy demonstration project: Advanced automation demonstration of Space Station Freedom thermal control system

The NASA Systems Autonomy Demonstration Project (SADP) was initiated in response to Congressional interest in Space station automation technology demonstration. The SADP is a joint cooperative effort between Ames Research Center (ARC) and Johnson Space Center (JSC) to demonstrate advanced automation technology feasibility using the Space Station Freedom Thermal Control System (TCS) test bed. A model-based expert system and its operator interface were developed by knowledge engineers, AI researchers, and human factors researchers at ARC working with the domain experts and system integration engineers at JSC. Its target application is a prototype heat acquisition and transport subsystem of a space station TCS. The demonstration is scheduled to be conducted at JSC in August, 1989. The demonstration will consist of a detailed test of the ability of the Thermal Expert System to conduct real time normal operations (start-up, set point changes, shut-down) and to conduct fault detection, isolation, and recovery (FDIR) on the test article. The FDIR will be conducted by injecting ten component level failures that will manifest themselves as seven different system level faults. Here, the SADP goals, are described as well as the Thermal Control Expert System that has been developed for demonstration.

Dominick, Jeffrey↗
NASA NTRS1989

Space vehicle thermal rejection system

A space vehicle thermal heat rejection system 10 utilizing separate optimized heat pipe components for the functions of heat acquisition, heat transport, and heat rejection. A honeycomb panel heat pipe evaporator section 20 performs the function of heat acquisition, and forms a closed thermodynamic system with a dual channel heat pipe transport section 30, which performs the function of heat transport. A plurality of truss or channel core heat pipe rejection fins 41 form the condenser section 40, which performs the function of heat rejection. A common wall 32 separates the condenser section 40 from the transport section 30. Using the above heat pipe components and having efficient interfacing between them results in high performance factors for the overall system.

Tanzer, Herbert J.↗
NASA NTRS2017

A Comparison of System Architectures for a Mechanically Pumped Two-Phase Thermal Control System

The NASA Jet Propulsion Laboratory is developing a mechanically pumped two-phase fluid loop thermal control system to enable novel mission designs and greater science return. Pumped two-phase fluid loops have the potential to provide robust and effective thermal control that combine the best aspects of passive two-phase systems (heat pipes) and mechanically pumped single-phase fluid loops. The current program requirements include the development of a system with multiple 1 sq.m evaporators, each of which is capable of remaining spatially and temporally isothermal while accommodating heat loads of up to 500 W and local fluxes of up to 5 W/sq.cm. The goal is to attain this using less than 5 W of power. Such a system would be able to accommodate the next generation of payload and bus electronics while using minimal resources. This paper compares two different mechanically pumped two-phase fluid loop architectures in the context of these requirements. A mixed flow and separated flow architecture are compared on a theoretical and experimental basis. Test data from sub-scale, single evaporator/single condenser, mixed flow and separated flow testbeds are presented. In addition, a model is introduced to better understand separated flow systems and some expressions for the theoretical performance limits of such systems are developed. To date, the investigation suggests that a separated flow architecture is better suited to the program requirements. Separated flow systems have the potential to accommodate an isothermalizing two-phase evaporator while using lower levels of power than would be required for a mixed flow system. In addition, it is argued that separated flow systems are more robust and amenable to analysis than mixed flow systems, since they significantly reduce the occurrence of two-phase flow by separating phases in the evaporator. Future work will include developing a full-scale testbed that includes multiple evaporators and condensers in a representative flight configuration.

Furst, Benjamin↗
NASA NTRS1984

Shell tile thermal protection system

A reusable, externally applied thermal protection system for use on aerospace vehicles subject to high thermal and mechanical stresses utilizes a shell tile structure which effectively separates its primary functions as an insulator and load absorber. The tile consists of structurally strong upper and lower metallic shells manufactured from materials meeting the thermal and structural requirements incident to tile placement on the spacecraft. A lightweight, high temperature package of insulation is utilized in the upper shell while a lightweight, low temperature insulation is utilized in the lower shell. Assembly of the tile which is facilitated by a self-locking mechanism, may occur subsequent to installation of the lower shell on the spacecraft structural skin.

Macconochie, I. O.↗
NASA NTRS2007

Thermal Analysis on Cryogenic Liquid Hydrogen Tank on an Unmanned Aerial Vehicle System

Thermal analyses are performed on the liquid hydrogen (LH2) tank designed for an unmanned aerial vehicle (UAV) powered by solar arrays and a regenerative proton-exchange membrane (PEM) fuel cell. A 14-day cruise mission at a 65,000 ft altitude is considered. Thermal analysis provides the thermal loads on the tank system and the boiling-off rates of LH2. Different approaches are being considered to minimize the boiling-off rates of the LH2. It includes an evacuated multilayer insulation (MLI) versus aerogel insulation on the LH2 tank and aluminum versus stainless steel spacer rings between the inner and outer tank. The resulting boil-off rates of LH2 provided by the one-dimensional model and three-dimensional finite element analysis (FEA) on the tank system are presented and compared to validate the results of the three-dimensional FEA. It concludes that heat flux through penetrations by conduction is as significant as that through insulation around the tank. The tank system with MLI insulation and stainless steel spacer rings result in the lowest boiling-off rate of LH2.

Wang, Xiao-Yen↗
NASA NTRS1971

The relative effects of entry parameters on thermal protection system weight

Shielding a spacecraft from the severe thermal environment of an atmospheric entry requires a sophisticated thermal protection system (TPS). Thermal computer program models were developed for two such TPS designs proposed for the space shuttle orbiter. The multilayer systems, a reusable surface insulation TPS, and a re-radiative metallic skin TPS, were sized for a cross-section of trajectories in the entry corridor. This analysis indicates the relative influence of the entry parameters on the weight of each TPS concept. The results are summarized graphically. The trajectory variables considered were down-range, cross-range, orbit inclination, entry interface velocity and flight path angle, maximum heating rate level, angle of attack, and ballistic coefficient. Variations in cross-range and flight path angle over the ranges considered had virtually no effect on the required entry TPS weight. The TPS weight was significantly more sensitive to variations in angle of attack than to dispersions in the other trajectory considered.

Hirasaki, P. N.↗
NASA NTRS2012

Compact Water Vapor Exchanger for Regenerative Life Support Systems

Thermal and environmental control systems for future exploration spacecraft must meet challenging requirements for efficient operation and conservation of resources. Regenerative CO2 removal systems are attractive for these missions because they do not use consumable CO2 absorbers. However, these systems also absorb and vent water to space along with carbon dioxide. This paper describes an innovative device designed to minimize water lost from regenerative CO2 control systems. Design studies and proof-of-concept testing have shown the feasibility of a compact, efficient membrane water vapor exchanger (WVX) that will conserve water while meeting challenging requirements for operation on future spacecraft. Compared to conventional WVX designs, the innovative membrane WVX described here has the potential for high water recovery efficiency, compact size, and very low pressure losses. The key innovation is a method for maintaining highly uniform flow channels in a WVX core built from water-permeable membranes. The proof-of-concept WVX incorporates all the key design features of a prototypical unit, except that it is relatively small scale (1/23 relative to a unit sized for a crew of six) and some components were fabricated using non-prototypical methods. The proof-of-concept WVX achieved over 90% water recovery efficiency in a compact core in good agreement with analysis models. Furthermore the overall pressure drop is very small (less than 0.5 in. H2O, total for both flow streams) and meets requirements for service in environmental control and life support systems on future spacecraft. These results show that the WVX provides very uniform flow through flow channels for both the humid and dry streams. Measurements also show that CO2 diffusion through the water-permeable membranes will have negligible effect on the CO2 partial pressure in the spacecraft atmosphere.

Izenson, Michael G.↗
NASA NTRS2016

Design, Development, and Testing of a Water Vapor Exchanger for Spacecraft Life Support Systems

Thermal and environmental control systems for future exploration spacecraft must meet challenging requirements for efficient operation and conservation of resources. Maximizing the use of regenerative systems and conserving water are critical considerations. This paper describes the design, development, and testing of an innovative water vapor exchanger (WVX) that can minimize the amount of water absorbed in, and vented from, regenerative CO2 removal systems. Key design requirements for the WVX are high air flow capacity (suitable for a crew of six), very high water recovery, and very low pressure losses. We developed fabrication and assembly methods that enable high-efficiency mass transfer in a uniform and stable array of Nafion tubes. We also developed analysis and design methods to compute mass transfer and pressure losses. We built and tested subscale units sized for flow rates of 2 and 5 cu ft/min (3.4–8.5 cu m/hr). Durability testing demonstrated that a stable core geometry was sustained over many humid/dry cycles. Pressure losses were very low (less than 0.5 in. H2O (125 Pa) total) and met requirements at prototypical flow rates. We measured water recovery efficiency across a range of flow rates and humidity levels that simulate the range of possible cabin conditions. We measured water recovery efficiencies in the range of 80 to 90%, with the best efficiency at lower flow rates and higher cabin humidity levels. We compared performance of the WVX with similar units built using an unstructured Nafion tube bundle. The WVX achieves higher water recovery efficiency with nearly an order of magnitude lower pressure drop than unstructured tube bundles. These results show that the WVX provides uniform flow through flow channels for both the humid and dry streams and can meet requirements for service on future exploration spacecraft. The WVX technology will be best suited for long-duration exploration vehicles that require regenerative CO2 removal systems while needing to conserve water.

Izenson, Michael G.↗
NASA NTRS2000

Lunar Dust Contamination Effects on Lunar Base Thermal Control Systems

Many studies have been conducted to develop a thermal control system that can operate under the extreme thermal environments found on the lunar surface. While these proposed heat rejection systems use different methods to reject heat, each system contains a similar component, a thermal radiator system. These studies have always considered pristine thermal control system components and have overlooked the possible deleterious effects of lunar dust contamination. Since lunar dust has a high emissivity and absorptivity (greater than 0.9) and is opaque, dust accumulation on a surface should radically alter its optical properties and therefore alter its thermal response compared to ideal conditions. In addition, the non-specular nature of the dust particles will alter the performance of systems that employ specular surfaces to enhance heat rejection. To date, few studies have examined the effect of dust deposition on the normal control system components. These studies only focused on a single heat rejection or photovoltaic system. These studies did show that lunar dust accumulations alter the optical properties of any lunar base hardware, which in turn affects component temperatures, and heat rejection. Therefore, a new study was conducted to determine the effect of lunar dust contamination on heat rejection systems. For this study, a previously developed dust deposition model was incorporated into the Thermal Synthesizer System (TSS) model. This modeling scheme incorporates the original method of predicting dust accumulation due to vehicle landings by assuming that the thin dust layer can be treated as a semitransparent surface slightly above and in thermal contact with the pristine surface. The results of this study showed that even small amounts of dust deposits can radically alter the performance of the heat rejection systems. Furthermore. this study indicates that heat rejection systems be either located far from any landing sites or be protected from dust producing mechanisms.

Keller, John R.↗
NASA NTRS1990

Thermal control system technology discipline

Viewgraphs on thermal control systems technology discipline for Space Station Freedom are presented. Topics covered include: heat rejection; heat acquisition and transport; monitoring and control; passive thermal control; and analysis and test verification.

Ellis, Wilbert E.↗
NASA NTRS1998

Integrated Thermal Insulation System for Spacecraft

An integrated thermal protection system (TPS) for a spacecraft includes a grid that is bonded to skin of the spacecraft, e.g., to support the structural loads of the spacecraft. A plurality of thermally insulative, relatively large panels are positioned on the grid to cover the skin of the spacecraft to which the grid has been bonded. Each panel includes a rounded front edge and a front flange depending downwardly from the front edge. Also, each panel includes a rear edge formed with a rounded socket for receiving the rounded front edge of another panel therein, and a respective rear flange depends downwardly from each rear edge. Pins are formed on the front flanges, and pin receptacles are formed on the rear flanges, such that the pins of a panel mechanically interlock with the receptacles of the immediately forward panel. To reduce the transfer to the skin of heat which happens to leak through the panels to the grid, the grid includes stringers that are chair-shaped in cross-section.

Kolodziej, Paul↗
NASA NTRS1995

Preliminary control system design and analysis for the Space Station Furnace Facility thermal control system

This report presents the Space Station Furnace Facility (SSFF) thermal control system (TCS) preliminary control system design and analysis. The SSFF provides the necessary core systems to operate various materials processing furnaces. The TCS is defined as one of the core systems, and its function is to collect excess heat from furnaces and to provide precise cold temperature control of components and of certain furnace zones. Physical interconnection of parallel thermal control subsystems through a common pump implies the description of the TCS by coupled nonlinear differential equations in pressure and flow. This report formulates the system equations and develops the controllers that cause the interconnected subsystems to satisfy flow rate tracking requirements. Extensive digital simulation results are presented to show the flow rate tracking performance.

Jackson, M. E.↗
NASA NTRS2004

Heat Shielding Characteristics and Thermostructural Performance of a Superalloy Honeycomb Sandwich Thermal Protection System (TPS)

Heat-transfer, thermal bending, and mechanical buckling analyses have been performed on a superalloy "honeycomb" thermal protection system (TPS) for future hypersonic flight vehicles. The studies focus on the effect of honeycomb cell geometry on the TPS heat-shielding performance, honeycomb cell wall buckling characteristics, and the effect of boundary conditions on the TPS thermal bending behavior. The results of the study show that the heat-shielding performance of a TPS panel is very sensitive to change in honeycomb core depth, but insensitive to change in honeycomb cell cross-sectional shape. The thermal deformations and thermal stresses in the TPS panel are found to be very sensitive to the edge support conditions. Slight corrugation of the honeycomb cell walls can greatly increase their buckling strength.

Ko, William L.↗
NASA NTRS1993

Thermal control system for SSF sensor/electronics

A thermal control system was designed for the Space Station Freedom (SSF) sensor/electronics box (SSTACK). Multi-layer insulation and heaters are used to maintain the temperatures of the critical components within their operating and survival temperature limits. Detailed and simplified SSTACK thermal models were developed and temperatures were calculated for worst-case orbital conditions. A comparison between the two models showed very good agreement. Temperature predictions were also compared to measured temperatures from a thermal-vacuum test.

Akau, R. L.↗
NASA NTRS2012

Flexible Thin Metal Film Thermal Sensing System

A flexible thin metal film thermal sensing system is provided. A thermally-conductive film made from a thermally-insulating material is doped with thermally-conductive material. At least one layer of electrically-conductive metal is deposited directly onto a surface of the thermally-conductive film. One or more devices are coupled to the layer(s) to measure an electrical characteristic associated therewith as an indication of temperature.

Thomsen, Donald Laurence↗
NASA NTRS2012

An Overview of the NASA Ames Millimeter-Wave Thermal Launch System

The Millimeter-Wave Thermal Launch System (MTLS) is a beamed-energy propulsion concept being designed at NASA Ames Research Center. This effort is in response to the NASA Office of the Chief Technologist s announcement of the Ride the Light program. Our objective is to produce a design that goes beyond the feasibility analysis level of previous studies and provides a solid foundation for low cost access to space. The MTLS is designed to place a 500 lb payload into Low Earth Orbit (LEO) two times a day. This frequent launch, small payload niche is well suited for the particular advantages and constraints of beamed-energy propulsion, and has the potential to drastically increase access to space by reducing the cost per kilogram of placing payloads into LEO. This paper summarizes the findings of the MTLS study. The chemical rocket engine is in principle a simple device. It acts by releasing the chemical energy stored in propellants such as hydrogen and oxygen through combustion, then converting that thermal energy into kinetic energy by expansion through a nozzle. As such, it is fundamentally limited by the energy released in combustion reactions and the molecular weight of the products of those reactions. The highest performing conventional propellant combination, liquid oxygen and liquid hydrogen, can produce vacuum specific impulses of around 450 seconds. The design space of current launch vehicles (which tend to be large, multi-stage, and expendable) are defined by these limitations. An entirely new approach may be necessary in order to enable future launch vehicles of radically improved capabilities. Beamed-energy propulsion (BEP) is an alternative approach that bypasses the energy limitations of chemical propulsion. Instead of relying on a chemical reaction as the energy source, it is supplied externally via a beam of electromagnetic energy produced on the ground. In the concept examined in the MTLS, this energy is absorbed by a heat exchanger which then transfers the energy to the propellant. This decouples the energy source from the working fluid, vastly expanding the design space. For example, a launch vehicle could use only water as propellant, making it very safe and easy to handle. However, the most commonly proposed way to take advantage of this decoupling is to use pure hydrogen as the working fluid, which enables specific impulses of around 800 seconds1 if the heat exchanger material can operate at 2200 K. With such a significantly increased Isp, it may be possible to build single-stage to orbit vehicles, with enough mass margins left over to permit lower cost fabrication techniques.

Murakami, David↗