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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Automated microbial metabolism laboratory

The design and rationale of an advanced labeled release experiment based on single addition of soil and multiple sequential additions of media into each of four test chambers are outlined. The feasibility for multiple addition tests was established and various details of the methodology were studied. The four chamber battery of tests include: (1) determination of the effect of various atmospheric gases and selection of that gas which produces an optimum response; (2) determination of the effect of incubation temperature and selection of the optimum temperature for performing Martian biochemical tests; (3) sterile soil is dosed with a battery of C-14 labeled substrates and subjected to experimental temperature range; and (4) determination of the possible inhibitory effects of water on Martian organisms is performed initially by dosing with 0.01 ml and 0.5 ml of medium, respectively. A series of specifically labeled substrates are then added to obtain patterns in metabolic 14CO2 (C-14)O2 evolution.

Source record↗
NASA NTRS2005

Space Suit Electrocardiographic Electrode Selection: Are commercial electrodes better than the old Apollo technology?

The NASA Manned Space Program uses an electrocardiograph (ECG) system to monitor astronauts during extravehicular activity (EVA). This ECG system, called the Operational Bioinstrumentation System (OBS), was developed during the Apollo era. Throughout the Shuttle program these electrodes experienced failures during several EVAs performed from the Space Shuttle and International Space Station (ISS) airlocks. An attempt during Shuttle Flight STS-109 to replace the old electrodes with new commercial off-the-shelf (COTS) disposable electrodes proved unsuccessful. One assumption for failure of the STS-109 COTS electrodes was the expansion of trapped gases under the foam electrode pad, causing the electrode to be displaced from the skin. Given that our current electrodes provide insufficient reliability, a number of COTS ECG electrodes were tested at the NASA Altitude Manned Chamber Test Facility. Methods: OBS disposable electrodes were tested on human test subjects in an altitude chamber simulating an Extravehicular Mobility Unit (EMU) operating pressure of 4.3 psia with the following goals: (1) to confirm the root cause of the flight certified, disposable electrode failure during flight STS-109. (2) to identify an adequate COTS replacement electrode and determine if further modifications to the electrodes are required. (3) to evaluate the adhesion of each disposable electrode without preparation of the skin with isopropyl alcohol. Results: There were several electrodes that failed the pressure testing at 4.3psia, including the electrodes used during flight STS-109. Two electrodes functioned well throughout all testing and were selected for further testing in an EMU at altitude. A vent hole placed in all electrodes was also tested as a possible solution to prevent gas expansion from causing electrode failures. Conclusions: Two failure modes were identified: (1) foam-based porous electrodes entrapped air bubbles under the pad (2) poor adhesion caused some electrodes to fail

Redmond, M.↗
NASA NTRS1974

Raman scattering applied to hypersonic air flow

The Raman scattering technique was applied to measure the local static temperature and gas number density over a sharp-edge flat-plate model in a Mach 5 nozzle of the Langley nozzle test chamber with air as the test gas. The angle of attack varied from -5 to 15 deg, and the stagnation temperature varied from 317 to 442 K, with stagnation pressures ranging from 170 kN to 2.8 MN/sq m. The measured values of static temperature and density ranged from 60 to 100 K and from 0.03 to 0.8 kg/cu m, respectively. A comparison with calculated values based on static pressure measurements along the model shows that the Raman scattering technique is a viable measurement method in applications to high-speed three-dimensional flows.

Hillard, M. E., Jr.↗
NASA NTRS2001

Testing and Optimization of Electrically Conductive Spacecraft Coatings

This is the final report discussing the work done for the Space Environments and Effects (SEE) Program. It discusses test chamber design, coating research, and test results on electrically thermal control coatings. These thermal control coatings are being developed to have several orders of magnitude higher electrical conductivity than most available thermal control coatings. Most current coatings tend to have a range in surface resistivity from 1,011 to 1,013 ohms/sq. Historically, spacecraft have had thermal control surfaces composed of dielectric materials of either polymers (paints and metalized films) or glasses (ceramic paints and optical solar reflectors). Very seldom has the thermal control surface of a spacecraft been a metal where the surface would be intrinsically electrically conductive. The poor thermal optical properties of most metals have, in most cases, stopped them from being used as a thermal control surface. Metals low infrared emittance (generally considered poor for thermal control surfaces) and/or solar absorptance, have resulted in the use of various dielectric coatings or films being applied over the substrate materials in order to obtain the required optical properties.

Mell, R. J.↗
NASA NTRS2017

Thermally-Conductive Metallic Coatings and Applications for Heat Removal on In-Space Cryogenic Vehicles

For large in-space cryogenic upper stages, substantial axial heat removal from a forward skirt by vapor-based heat interception may not be achieved by simple attachment methods unless sufficient thermal conductance from the skirt to the cooling fluid can be achieved. Preferable methods would allow for the addition of the cooling system to existing structure with minimal impact on the structure. Otherwise, significant modification to the basic structural design andor novel and complex attachment mechanisms with high effective thermal conductance are likely to be required. The approach being pursued by evolvable Cryogenics (eCryo) is to increase the thermal performance of a relatively simple attachment system by applying metallic or other thermally conductive material coatings to the mating surface area of the fluid channel where it is attached the skirt wall. The expectation of candidate materials is that the dramatic increase in conductivity of pure metals at temperatures close to liquid hydrogen vapor temperature will compensate for the reduced actual contact area typical of mechanical joints. Basic contact conductance data at low temperatures for candidate interface materials is required to enable the test approach. A test rig was designed at NASA Glenn Research Center to provide thermal contact resistance testing between small sample coupons coated with conductive material via electron beam evaporation, a low-temperature option that will not affect physical properties of base materials. Average coating thicknesses were 10 k. The test fixture was designed to mount directly to a cryocooler cold head within a vacuum test chamber. The purpose of this test was to determine qualitative contact conductance between various test samples. Results from this effort will be implemented in a sub-scale vapor-based heat interception test, where the applicability for increased heat removal on large structural skirts will be considered.

small scale testing↗
NASA NTRS2023

Low-Cost Testing in Representative Lunar Regolith Environment

Rapid prototyping and testing are key enablers of iterative designs common in early technology research and development. Testing in a dusty environment is critical to prepare the Cryogenic Magnetic Coupler for lunar operations. To enable early and iterative testing of dust mitigation concepts, a low-cost, low-fidelity representative lunar regolith environment was developed at the National Aeronautics and Space Administration (NASA) Armstrong Flight Research Center (Edwards, California). Based on preliminary testing with this test setup, a similar setup may be of interest to universities and other entities looking to develop the capability to safely test relatively small-scale components with lunar regolith simulant. The development of this lunar regolith test chamber and results from preliminary testing are presented in this paper. Also discussed are further development strategies for the potential improvement of this setup.

Scott Stebbins↗
NASA NTRS2006

Conducting Closed Habitation Experiments: Experience from the Lunar Mars Life Support Test Project

The Lunar-Mars Life Support Test Project (LMLSTP) was conducted from 1995 through 1997 at the National Aeronautics and Space Administration s (NASA) Johnson Space Center (JSC) to demonstrate increasingly longer duration operation of integrated, closed-loop life support systems that employed biological and physicochemical techniques for water recycling, waste processing, air revitalization, thermal control, and food production. An analog environment for long-duration human space travel, the conditions of isolation and confinement also enabled studies of human factors, medical sciences (both physiology and psychology) and crew training. Four tests were conducted, Phases I, II, IIa and III, with durations of 15, 30, 60 and 91 days, respectively. The first phase focused on biological air regeneration, using wheat to generate enough oxygen for one experimental subject. The systems demonstrated in the later phases were increasingly complex and interdependent, and provided life support for four crew members. The tests were conducted using two human-rated, atmospherically-closed test chambers, the Variable Pressure Growth Chamber (VPGC) and the Integrated Life Support Systems Test Facility (ILSSTF). Systems included test articles (the life support hardware under evaluation), human accommodations (living quarters, kitchen, exercise equipment, etc.) and facility systems (emergency matrix system, power, cooling, etc.). The test team was managed by a lead engineer and a test director, and included test article engineers responsible for specific systems, subsystems or test articles, test conductors, facility engineers, chamber operators and engineering technicians, medical and safety officers, and science experimenters. A crew selection committee, comprised of psychologists, engineers and managers involved in the test, evaluated male and female volunteers who applied to be test subjects. Selection was based on the skills mix anticipated for each particular test, and utilized information from psychological and medical testing, data on the knowledge, experience and skills of the applicants, and team building exercises. The design, development, buildup and operation of test hardware and documentation followed the established NASA processes and requirements for test buildup and operation.

Barta, Daniel J.↗
NASA NTRS2004

Conducting Closed Habitation Experiments: Experience from the Lunar Mars Life Support Test Project

The Lunar-Mars Life Support Test Project (LMLSTP) was conducted from 1995 through 1997 at the National Aeronautics and Space Administration s (NASA) Johnson Space Center (JSC) to demonstrate increasingly longer duration operation of integrated, closed-loop life support systems that employed biological and physicochemical techniques for water recycling, waste processing, air revitalization, thermal control, and food production. An analog environment for long-duration human space travel, the conditions of isolation and confinement also enabled studies of human factors, medical sciences (both physiology and psychology) and crew training. Four tests were conducted, Phases I, II, IIa and III, with durations of 15, 30,60 and 91 days, respectively. The first phase focused on biological air regeneration, using wheat to generate enough oxygen for one experimental subject. The systems demonstrated in the later phases were increasingly complex and interdependent, and provided life support for four crew members. The tests were conducted using two human-rated, atmospherically-closed test chambers, the Variable Pressure Growth Chamber (VPGC) and the Integrated Life Support Systems Test Facility (ILSSTF). Systems included test articles (the life support hardware under evaluation), human accommodations (living quarters, kitchen, exercise equipment, etc.) and facility systems (emergency matrix system, power, cooling, etc.). The test team was managed by a lead engineer and a test director, and included test article engineers responsible for specific systems, subsystems or test articles, test conductors, facility engineers, chamber operators and engineering technicians, medical and safety officers, and science experimenters. A crew selection committee, comprised of psychologists, engineers and managers involved in the test, evaluated male and female volunteers who applied to be test subjects. Selection was based on the skills mix anticipated for each particular test, and utilized information from psychological and medical testing, data on the knowledge, experience and skills of the applicants, and team building exercises. The design, development, buildup and operation of test hardware and documentation followed the established NASA processes and requirements for test buildup and operation.

Barta, Daniel J.↗
NASA NTRS1993

Advanced Small Rocket Chambers. Basic Program and Option 2: Fundamental Processes and Material Evaluation

Propellants, chamber materials, and processes for fabrication of small high performance radiation cooled liquid rocket engines were evaluated to determine candidates for eventual demonstration in flight-type thrusters. Both storable and cryogenic propellant systems were considered. The storable propellant systems chosen for further study were nitrogen tetroxide oxidizer with either hydrazine or monomethylhydrazine as fuel. The cryogenic propellants chosen were oxygen with either hydrogen or methane as fuel. Chamber material candidates were chemical vapor deposition (CVD) rhenium protected from oxidation by CVD iridium for the chamber hot section, and film cooled wrought platinum-rhodium or regeneratively cooled stainless steel for the front end section exposed to partially reacted propellants. Laser diagnostics of the combustion products near the hot chamber surface and measurements at the surface layer were performed in a collaborative program at Sandia National Laboratories, Livermore, CA. The Material Sample Test Apparatus, a laboratory system to simulate the combustion environment in terms of gas and material temperature, composition, and pressure up to 6 Atm, was developed for these studies. Rocket engine simulator studies were conducted to evaluate the materials under simulated combustor flow conditions, in the diagnostic test chamber. These tests used the exhaust species measurement system, a device developed to monitor optically species composition and concentration in the chamber and exhaust by emission and absorption measurements.

Jassowski, Donald M.↗
NASA NTRS1976

Flight experience with a digital integrated propulsion control system on an F-111E airplane

A digital integrated propulsion control system (IPCS) installed in the left side of an F-111 E aircraft was tested in flight. The F-111 aircraft was selected for the IPCS program because it incorporated a variable geometry inlet and an afterburning turbofan engine and had two engines, one of which could remain in the normal configuration to ensure flight safety. Flight data were compared with results of tests run in an altitude test chamber. The digital system was found to be capable of duplicating the standard engine and inlet control systems. Instabilities such as inlet buzz and afterburner rumble were detected and controlled. The usefulness of an altitude chamber for developing a software and testing hardware was proven. The flexibility of IPCS was demonstrated when an autothrottle, an in-flight thrust calculation, and a coannular noise study capability were added at the end of the flight tests.

Burcham, F. W., Jr.↗
NASA NTRS1992

TOPEX Microwave Radiometer - Thermal design verification test and analytical model validation

The testing of the TOPEX Microwave Radiometer (TMR) is described in terms of hardware development based on the modeling and thermal vacuum testing conducted. The TMR and the vacuum-test facility are described, and the thermal verification test includes a hot steady-state segment, a cold steady-state segment, and a cold survival mode segment totalling 65 hours. A graphic description is given of the test history which is related temperature tracking, and two multinode TMR test-chamber models are compared to the test results. Large discrepancies between the test data and the model predictions are attributed to contact conductance, effective emittance from the multilayer insulation, and heat leaks related to deviations from the flight configuration. The TMR thermal testing/modeling effort is shown to provide technical corrections for the procedure outlined, and the need for validating predictive models is underscored.

Lin, Edward I.↗
NASA NTRS1993

Mechanism Guides Hatch Through Hatchway

Elliptical hatch designed to move through hatchway to make pressure-assisted seal with either side of bulkhead. Compact three-degree-of-freedom mechanism guides hatch through hatchway or holds hatch off to one side to facilitate passage of crew and/or equipment. Hatches and mechanisms used in submarines, pressure chambers (including hyperbaric treatment chambers), vacuum chambers, and vacuum-or-pressure test chambers.

Barron, Daniel R.↗
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 NTRS2016

Post-Test Inspection of NASA's Evolutionary Xenon Thruster Long-Duration Test Hardware: Discharge Chamber

The NEXT Long-Duration Test is part of a comprehensive thruster service life assessment intended to demonstrate overall throughput capability, validate service life models, quantify wear rates as a function of time and operating condition, and identify any unknown life-limiting mechanisms. The test was voluntarily terminated in February 2014 after demonstrating 51,184 hours of high-voltage operation, 918 kg of propellant throughput, and 35.5 MN-s of total impulse. The post-test inspection of the thruster hardware began shortly afterwards with a combination of non-destructive and destructive analysis techniques, and is presently nearing completion. This paper presents relevant results of the post-test inspection for the discharge chamber as well as other miscellaneous components such as the high-voltage propellant isolators and electrical cabling. Comparison of magnetic field measurements taken during pretest and post-test inspections indicate that the field strength did not degrade, consistent with performance data obtained during the test. Inspection of discharge chamber mesh samples show a deposition coating primarily composed of grid material that is approximately 15 micrometers in thickness. This thickness is well within the retention capability of the mesh and is therefore not expected to present any issues. Approximately 3.1 grams of deposition flakes were found at the bottom of the discharge chamber, composed primarily of grid material and carbon. Calculated size histograms of these flakes indicate that 99% have a maximum dimension of 200 micrometers or smaller, which is significantly less than the ion optics grid gap. Larger flakes that are capable of causing a grid-to-grid short will be analyzed to determine if their formation will occur in flight or is a facility effect. The high-voltage propellant isolators as well as numerous other electrical insulators were inspected and no evidence of arcing or any other issues were found.

Electric Propulsion↗