SEARCH · Engineering Papers
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Final report describes domestic water preheat installed in 148 room motel. Equipment meets 50 percent of needs when motel is 100 percent occupied; equivalently, it supplies 100 percent of hot water when occupancy is 50 percent. System consists of 1,400 square feet of flat plate liquid solar collectors, storage tanks, pump, controller, and hardware.
The Solar System was designed to provide 50 percent of the total Domestic Hot Water (DHW) demand. Liquid Flat Plate Collectors (900 square feet) are used for the collector subsystem. The collector subsystem is closed loop, using 50 percent Ethylene Glycol solution antifreeze for freeze protection. The 1,000 gallon fiber glass storage tank contains two heat exchangers. One of the heat exchangers heats the storage tank with the collector solar energy. The other heat exchanger preheats the cold supply water as it passes through on the way to the Domestic Hot Water (DHW) tank heaters. Electrical energy supplements the solar energy for the DHW. The Collector Mounting System utilizes guy wires to structurally tie the collector array to the building.
Integrated systems and missions studies are presented for an evolutionary lunar-to-Mars space transportation system (STS) based on nuclear thermal rocket (NTR) technology. A 'standardized' set of engine and stage components are identified and used in a 'building block' fashion to configure a variety of piloted and cargo, lunar and Mars vehicles. The reference NTR characteristics include a thrust of 50 thousand pounds force (klbf), specific impulse (I(sub sp)) of 900 seconds, and an engine thrust-to-weight ratio of 4. 3. For the National Aeronautics and Space Administrations (NASA) First Lunar Outpost (FLO) mission, and expendable NTR stage powered by two such engines can deliver approximately 96 metric tonnes (t) to trans-lunar injection (TLI) conditions for an initial mass in low Earth orbit (IMLEO) of approximately 198 t compared to 250 t for a cryogenic chemical system. The stage liquid hydrogen (LH2) tank has a diameter, length, and capacity of 10 m, 14.5 m and 66 t, respectively. By extending the stage length and LH2 capacity to approximately 20 m and 96 t, a single launch Mars cargo vehicle could deliver to an elliptical Mars parking orbit a 63 t Mars excursion vehicle (MEV) with a 45 t surface payload. Three 50 klbf engines and the two standardized LH2 tanks developed for the lunar and Mars cargo vehicles are used to configure the vehicles supporting piloted Mars missions as early as 2010. The 'modular' NTR vehicle approach forms the basis for an efficient STS able to handle the needs of a wide spectrum of lunar and Mars missions.
The Soft X-ray Spectrometer (SXS) is one of four instruments on the Japanese Astro-H mission, which is currently planned for launch in late 2015. The SXS will perform imaging spectroscopy in the soft X-ray band (0.3-12 keV) using a 6 6 pixel array of microcalorimeters cooled to 50 mK. The detectors are cooled by a 3-stage adiabatic demagnetization refrigerator (ADR) that rejects heat to either a superfluid helium tank (at 1.2 K) or to a 4.5 K Joule-Thomson (JT) cryocooler. Four gas-gap heat switches are used in the assembly to manage heat flow between the ADR stages and the heat sinks. The engineering model (EM) ADR was assembled and performance tested at NASA/GSFC in November 2011, and subsequently installed in the EM dewar at Sumitomo Heavy Industries, Japan. During the first cooldown in July 2012, a failure of the heat switch that linked the two colder stages of the ADR to the helium tank was observed. Operation of the ADR requires some mechanism for thermally linking the salt pills to the heat sink, and then thermally isolating them. With the failed heat switch unable to perform this function, an alternate plan was devised which used carefully controlled amounts of exchange gas in the dewar's guard vacuum to facilitate heat exchange. The process was successfully demonstrated in November 2012, allowing the ADR to cool the detectors to 50 mK for hold times in excess of 10 h. This paper describes the exchange-gas-assisted recycling process, and the strategies used to avoid helium contamination of the detectors at low temperature.
Demonstration of high-structural-performance ARDEFORM cryoformed 301 stainless-steel glass-fiber-reinforced (GFR) vessels by room temperature tests of 13 1/2-in. diam spheres. Tests verified that the structural performance of ARDEFORM spherical GFR vessels not only exceeded that of all metal construction, but also bettered previous GFR experimental results by 50%. Achievement of essentially the full strength of fiberglass in a spherical wrap pattern was again verified. Significant weight advantages for this construction are projected for O2/N2 high-pressure gas tanks for Space Shuttle environmental control/life support system missions.
An experimental investigation was conducted to determine the effects of several basic variables upon the damping of the fundamental antisymmetric mode of liquids in right-circular cylindrical tanks without baffles. The variables examined include liquid depth, efflux rate, liquid amplitude, kinematic viscosity, and tank size. The data are presented in dimensionless form and compared with available theory. For the range of variables examined, variations of efflux rate and liquid amplitude were found to have no significant effects on the liquid damping. The following theoretical relationship was found to be adequate for the prediction of the variation of damping with liquid depth, kinematic viscosity, and tank size: [for equation see full text] where v is the kinematic viscosity, R is the cylinder radius, g is the acceleration due to gravity, and h is the liquid depth. However, the constant K was experimentally found in this investigation to have the value 5.23, which is 50 percent higher than the theoretically predicted value.
The paper describes the unique pneumatic test requirements for the acoustic and shock separation testing of the Second Stage for the new Delta III Rocket at the Goddard Space Flight Center in Greenbelt, Maryland. The testing was conducted in the 45,000 cu ft (25-feet wide by 30-feet deep by 50-foot high) Acoustic Facility. The acoustic testing required that the liquid oxygen (LOX) and liquid hydrogen (LH2) tanks be filled with enough liquid nitrogen (LN2) to simulate launch fuel masses during testing. The challenge for this test dealt with designing, procuring, and fabricating the pneumatic supply systems for quick assembly while maintaining the purity requirements and minimizing costs. The pneumatic systems were designed to fill and drain the both LOX and LH2 tanks as well as to operate the fill/drain and vent valves for each of the tanks. The test criteria for the pneumatic sub-systems consisted of function, cleanliness, availability, and cost. The first criteria, function, required the tanks to be filled and drained in an efficient manner while preventing them from seeing pressures greater than 9 psig which would add a pressure cycle to the tank. An LN2 tanker, borrowed from another NASA facility, served as the pre-cool and drain tanker. Pre-cooling the tanks allowed for more efficient and cost effective transfer from the LN2 delivery tankers. Helium gas, supplied from a high purity tube trailer, was used to pressurize the vapor space above the LN2 pushing it into the drain tanker. The tube trailer also supplied high pressure helium to the vehicle for valve control and component purges. Cleanliness was maintained by proper component selection, end-use particle filtration, and any on-site cleaning determined necessary by testing. In order to meet the availability/cost juggling act, products designed for LOX delivery systems were procured to ensure system compatibility while off the shelf valves and tubing designed for the semiconductor industry were procured for the gas systems.
Results of wind tunnel heat transfer tests of 0.0175-scale Rockwell International Space Shuttle Vehicle configurations for orbiter alone, tank alone, and orbiter plus external tank are presented. Body flap shielding of SSME's during simulated entry was investigated. The tests were conducted at Mach 8 for thirteen Reynolds number.
Results of wind tunnel heat transfer tests of 0.0175-scale Rockwell International Space Shuttle Vehicle configurations for orbiter alone, tank alone, and orbiter plus external tank are presented. Body flap shielding of SSME's during simulated entry was also investigated. The tests were conducted at Mach 8 for thirteen Reynolds number per foot values ranging from 0.5 million to 3.72 million.
An experimental solar heating and cooling system model has been built and operated, combining elements that are programmable - heating and/or cooling load of a building, collected solar energy - with experimental equipment. The system model was based on the loads and components used in the Solar Building Test Facility (SBTF) which includes a 1394 square meter solar collector field, at NASA-Langley. Operations covered 5 continuous days under summer conditions. For the system model, up to 55% of the simulated collected solar energy was used for the building load. This amount constituted 35% of the building cooling load. Heat loss was significant. If tank heat loss were eliminated, 75% of the collected solar energy would be used. This amount would supply approximately 50% of the building cooling load.
Since 2001, the In-Space Propulsion Technology (ISPT) program has been developing and delivering in-space propulsion technologies for NASA's Science Mission Directorate (SMD). These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, Flagship and sample return missions currently under consideration. The ISPT program is currently developing technology in three areas that include Propulsion System Technologies, Entry Vehicle Technologies, and Systems/Mission Analysis. ISPT's propulsion technologies include: 1) the 0.6-7 kW NASA's Evolutionary Xenon Thruster (NEXT) gridded ion propulsion system; 2) a 0.3-3.9kW Halleffect electric propulsion (HEP) system for low cost and sample return missions; 3) the Xenon Flow Control Module (XFCM); 4) ultra-lightweight propellant tank technologies (ULTT); and 5) propulsion technologies for a Mars Ascent Vehicle (MAV). The NEXT Long Duration Test (LDT) recently exceeded 50,000 hours of operation and 900 kg throughput, corresponding to 34.8 MN-s of total impulse delivered. The HEP system is composed of the High Voltage Hall Accelerator (HIVHAC) thruster, a power processing unit (PPU), and the XFCM. NEXT and the HIVHAC are throttle-able electric propulsion systems for planetary science missions. The XFCM and ULTT are two component technologies which being developed with nearer-term flight infusion in mind. Several of the ISPT technologies are related to sample return missions needs: MAV propulsion and electric propulsion. And finally, one focus of the Systems/Mission Analysis area is developing tools that aid the application or operation of these technologies on wide variety of mission concepts. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness.
A 1/5.5-size powered dynamic model of the Columbia XJL-1 amphibian was landed in Langley tank no. 1 in smooth water and in oncoming waves of heights from 2.1 feet to 6.4 feet (full-size) and lengths from 50 feet to 264 feet (full-size). The motions and the vertical accelerations of the model were continuously recorded. The greatest vertical acceleration measured during the smooth-water landings was 3.1g. During landings in rough water the greatest vertical acceleration measured was 15.4g, for a landing in 6.4-foot by 165-foot waves. The impact accelerations increased with increase in wave height and, in general, decreased with increase in wave length. During the landings in waves the model bounced into the air at stalled attitudes at speeds below flying speed. The model trimmed up to the mechanical trim stop (20 deg) during landings in waves of heights greater than 2.0 feet. Solid water came over the bow and damaged the propeller during one landing in 6.4-foot waves. The vertical acceleration coefficients at first impact from the tank tests of a 1/5.5-size model were in fair agreement with data obtained at the Langley impact basin during tests of a 1/2-size model of the hull.
A study was conducted to evaluate ultrasonic cleaning systems for precision cleaning effectiveness for oxygen service hardware. This evaluation was specific for Rocketdyne Division of Rockwell Aerospace alloys and machining soils. Machining lubricants and hydraulic fluid were applied as soils to standardized complex test specimens designed to simulate typical hardware. The study consisted of tests which included 20, 25, 30, 40, 50, and 65 kHz ultrasonic cleaning systems. Two size categories of cleaning systems were evaluated, 3- to 10-gal laboratory size tanks and 35- to 320-gal industrial size tanks. The system properties of cavitation, frequency vs. cleaning effectiveness, the two types of transducers, and the power level of the system vs. size of the cleaning tank were investigated. The data obtained from this study was used to select the ultrasonic tanks for the aqueous fine clean facility installed at Rocketdyne.
A study was conducted to evaluate ultrasonic cleaning systems for precision cleaning effectiveness for oxygen service hardware. This evaluation was specific for Rocketdyne Div. of Rockwell Aerospace alloys and machining soils. Machining lubricants and hydraulic fluid were applied as soils to standardized complex test specimens designed to simulate typical hardware. The study consisted of tests which included 20, 25, 30, 40, 50, and 65 kHz ultrasonic cleaning systems. Two size categories of cleaning systems were evaluated, 3- to 10-gal laboratory size tanks and 35- to 320-gal industrial size tanks. The system properties of cavitation; frequency vs. cleaning effectiveness; the two types of transducers; and the power level of the system vs. size of the cleaning tank were investigated. The data obtained from this study was used to select the ultrasonic tanks for the aqueous fine clean facility installed at Rocketdyne.
Due to the increasing international competition and growing pressure to lower the cost of launch vehicles, enhanced technologies and competences for efficient, low cost manufacturing have to be developed. In this paper, an innovative manufacturing technology for launch vehicle applications is presented that was developed over the last few years at MT Aerospace AG (MTA) in collaboration with NASA Langley Research Center, the European Space Agency (ESA), Lockheed Martin Corporation (LMCO), Leifeld Metal Spinning and Scot Forge. The Integrally Stiffened Cylinder (ISC) manufacturing method was first investigated in 2010. Early results indicated a high mass saving potential for cylindrical structures for launch vehicle systems, such as cryogenic tank barrels and intertanks, produced using the ISC method. Cost-benefit studies conducted at NASA and MTA estimated up to 50% reduction in manufacturing costs is possible compared to conventional multi-piece, welded fabrication methods. Associated reductions in manufacturing lead times have shown that a 40% faster production rate is possible by the ISC process. Following proof of concept and process development efforts, a cylindrical structure manufactured using the ISC method had its maiden flight in 2015 on a Black Brant sounding rocket at NASA Wallops Flight Facility. Further collaboration between ESA, NASA, MTA, LMCO, and Scot Forge led to manufacturing test campaigns at MT Aerospace in 2017 and 2018 that successfully demonstrated ISC fabrication at a scale representative of commercial launch vehicles. This paper summarizes the status and relevant results of the development sequence of the ISC cylinder manufacturing method. Results of cost / benefit and structural performance analyses are also summarized. Moreover, an outlook will be given on further planned activities and development steps.
Tests were conducted in the NASA Langley Research Center 31-inch continuous Flow Hypersonic Wind Tunnel to determine RCS jet interaction effect on the hypersonic aerodynamic and stability and control characteristics prior to return to launch site (RTLS) abort separation. The model used was an 0.010-scale replica of the Space Shuttle Vehicle Configuration 3. Hypersonic stability data were obtained from tests at Mach 10.3 and dynamic pressure of 150 psf for the integrated Orbiter and external tank and the Orbiter alone. RCS modes of pitch, yaw, and roll at free flight dynamic pressure simulation of 7, 20, and 50 psf were investigated. The effects of speedbrake, bodyflap, elevon, and aileron deflections were also investigated.
Tests were conducted in the NASA Langley Research Center 31-inch continuous flow hypersonic wind tunnel from 14 February to 22 February 1974, to determine RCS jet interaction effect on the hypersonic aerodynamic and stability and control characteristics prior to RTLS abort separation. The model used was an 0.010-scale replica of the space shuttle vehicle configuration 3. Hypersonic stability data were obtained from tests at Mach 10.3 and dynamic pressure of 150 psf for the intergrated orbiter and external tank and the orbiter alone. RCS modes of pitch, yaw, and roll at free flight dynamic pressure simulation of 7, 20, and 50 psf were investigated. The effects of speedbrake, bodyflap, elevon, and aileron deflections were also investigated.
Tests were conducted in the 31-inch continuous Flow Hypersonic Wind Tunnel to determine RCS jet interaction effect on the hypersonic aerodynamic and stability and control characteristics prior to RTLS abort separation. The model used was an 0.010-scale replica of the Space Shuttle Vehicle Configuration 3. Hypersonic stability data were obtained from tests at Mach 10.3 and dynamic pressure of 150 psf for the integrated Orbiter and external tank and the Orbiter alone. RCS modes of pitch, yaw, and roll at free flight dynamic pressure simulation of 7, 20, and 50 psf were investigated. The effects of speedbrake, bodyflap, elevon, and airleron deflections were also investigated.