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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 199 records · Page 11
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NASA NTRS1926

Description and Laboratory Tests of a Roots Type Aircraft Engine Supercharger

This report describes a roots type aircraft engine supercharger and presents the results of some tests made with it at the Langley Field Laboratories of the National Advisory Committee for Aeronautics. The supercharger used in these tests was constructed largely of aluminum, weighed 88 pounds and was arranged to be operated from the rear of a standard aircraft engine at a speed of 1 1/2 engine crankshaft speed. The rotors of the supercharger were cycloidal in form and were 11 inches long and 9 1/2 inches in diameter. The displacement of the supercharger was 0.51 cubic feet of air per revolution of the rotors. The supercharger was tested in the laboratory, independently and in combination with a Liberty-12 aircraft engine, under simulated altitude pressure conditions in order to obtain information on its operation and performance. From these tests it seems evident that the Roots blower compares favorably with other compressor types used as aircraft engine superchargers and that it has several features that make it particularly attractive for such use.

Ware, Marsden↗
NASA NTRS1959

Transonic Flutter Investigation of Models of T-Tail of Blackburn NA-39 Airplane

A transonic flutter investigation has been made of models of the T-tail of the Blackburn NA-39 airplane. The models were dynamically and elastically scaled from measured airplane data in accordance with criteria which include a flutter safety margin. The investigation was made in the Langley transonic blowdown tunnel and covered a Mach number range from 0.73 to 1.09 at simulated altitudes extending to below sea level. The results of the investigation indicated that, if differences between the measured model and scaled airplane properties are disregarded, the airplane with the normal value of stabilizer pitching stiffness should have a stiffness margin of safety of at least 32 percent at all Mach numbers and altitudes within the flight boundary. However, the airplane with the emergency value of stabilizer pitching stiffness would not have the required margin of safety from symmetrical flutter at Mach numbers greater than about 0.85 at low altitudes. First-order corrections for some differences between the measured model and scaled airplane properties indicated that the airplane with the normal value of stabilizer pitching stiffness would still have an adequate margin of safety from flutter and that the flutter safety margin for the airplane with the emergency value of stabilizer pitching stiffness would be changed from inadequate to adequate. However, the validity of the corrections is questionable.

Jones, George W., Jr.↗
NASA NTRS1999

Rocket-in-a-Duct Performance Analysis

An axisymmetric, 110 N class, rocket configured with a free expansion between the rocket nozzle and a surrounding duct was tested in an altitude simulation facility. The propellants were gaseous hydrogen and gaseous oxygen and the hardware consisted of a heat sink type copper rocket firing through copper ducts of various diameters and lengths. A secondary flow of nitrogen was introduced at the blind end of the duct to mix with the primary rocket mass flow in the duct. This flow was in the range of 0 to 10% of the primary massflow and its effect on nozzle performance was measured. The random measurement errors on thrust and massflow were within +/-1%. One dimensional equilibrium calculations were used to establish the possible theoretical performance of these rocket-in-a-duct nozzles. Although the scale of these tests was small, they simulated the relevant flow expansion physics at a modest experimental cost. Test results indicated that lower performance was obtained at higher free expansion area ratios and longer ducts, while, higher performance was obtained with the addition of secondary flow. There was a discernable peak in specific impulse efficiency at 4% secondary flow. The small scale of these tests resulted in low performance efficiencies, but prior numerical modeling of larger rocket-in-a-duct engines predicted performance that was comparable to that of optimized rocket nozzles. This remains to be proven in large-scale, rocket-in-a-duct tests.

Schneider, Steven J.↗
NASA NTRS1999

Engineered Solutions to Reduce Occupational Noise Exposure at the NASA Glenn Research Center: A Five-Year Progress Summary (1994-1999)

At the NASA John H. Glenn Research Center at Lewis Field (formerly the Lewis Research Center), experimental research in aircraft and space propulsion systems is conducted in more than 100 test cells and laboratories. These facilities are supported by a central process air system that supplies high-volume, high-pressure compressed air and vacuum at various conditions that simulate altitude flight. Nearly 100,000 square feet of metalworking and specialized fabrication shops located on-site produce prototypes, models, and test hardware in support of experimental research operations. These activities, comprising numerous individual noise sources and operational scenarios, result in a varied and complex noise exposure environment, which is the responsibility of the Glenn Research Center Noise Exposure Management Program. Hearing conservation, community noise complaint response and noise control engineering services are included under the umbrella of this Program, which encompasses the Occupational Safety and Health Administration (OSHA) standard on occupational noise exposure, Sec. 29CFR 1910.95, as well as the more stringent NASA Health Standard on Hearing Conservation. Prior to 1994, in the absence of feasible engineering controls, strong emphasis had been placed on personal hearing protection as the primary mechanism for assuring compliance with Sec. 29CFR 1910.95 as well as NASA's more conservative policy, which prohibits unprotected exposure to noise levels above 85 dB(A). Center policy and prudent engineering practice required, however, that these efforts be extended to engineered noise controls in order to bring existing work areas into compliance with Sec. 29CFR 1910.95 and NASA's own policies and to ensure compliance for new installations. Coincident with the establishment in 1995 of a NASA wide multi-year commitment of funding for environmental abatement projects, the Noise Exposure Management Program was established, with its focus on engineering approaches to reducing occupational and community noise exposure. The organization and mission of this Program were documented shortly after its inception, and individual programmatic components have been the subject of subsequent papers. This paper summarizes the status and accomplishments of the engineering aspects of the Program from a five-year retrospective viewpoint and includes a review of retrofit noise control solution strategies, which have not been previously documented.

Cooper, Beth A.↗
NASA NTRS1961

Aerodynamic Characteristics of Parachutes at Mach Numbers from 1.6 to 3

A wind-tunnel investigation has been conducted to determine the parameters affecting the aerodynamic performance of drogue parachutes in the Mach number range from 1.6 to 3. Flow studies of both rigid and flexible-parachute models were made by means of high-speed schlieren motion pictures and drag coefficients of the flexible-parachute models were measured at simulated altitudes from about 50,000 to 120,000 feet. Porosity and Mach number were found to be the most important factors influencing the drag and stability of flexible porous parachutes. Such parachutes have a limited range of stable'operation at supersonic speeds, except for those with very high porosities, but the drag coefficient decreases rapidly with increasing porosity.

Maynard, Julian D.↗
NASA NTRS2002

Liquid Engine Test Facilities Assessment

The John C. Stennis Space Center (SSC) requested The Aerospace Corporation to examine the current testing capability of all existing large liquid engine test facilities located in the United States. That information along with projected liquid rocket engine development was used to examine future liquid rocket engine testing facilities needs in the coming decade. Current domestic liquid engine test facilities capabilities, when examined against engine concepts for the coming decade, indicate there are ample facilities offering altitude simulation during test. In addition, it was observed that many contractor facilities have limited ambient test capability of larger thrust engines under current consideration. Finally, it was concluded that diminished contractor participation engine development testing will drive this activity to the government sector. Only three facilities are seen as key contributors to engine testing in the coming decade, namely John C. Stennis Space Center (SSC), Marshall Space Flight Center (MSFC), and Air Force Research Laboratory (AFRL). Past rocket engine test experience was evaluated as a possible resource for projecting future engine test needs. A database comprised of various engine models and the level of testing performed to flight qualify those systems for their first flight was constructed. For comparison purposes in this study, development and qualification efforts were totaled and treated as one test program. Based on experience with past Air Force programs, the time on the test stand accounts for typically 50% or more of the total program time. Historical data show that the time to design and develop new engines has increased over the last 40 years, most likely due to scarcer resources in today's funding environment.

Adams, Michael J.↗
NASA NTRS2002

Determinants of erythropoietin release in response to short-term hypobaric hypoxia

We measured blood erythropoietin (EPO) concentration, arterial O(2) saturation (Sa(O(2))), and urine PO(2) in 48 subjects (32 men and 16 women) at sea level and after 6 and 24 h at simulated altitudes of 1,780, 2,085, 2,454, and 2,800 m. Renal blood flow (Doppler) and Hb were determined at sea level and after 6 h at each altitude (n = 24) to calculate renal O(2) delivery. EPO increased significantly after 6 h at all altitudes and continued to increase after 24 h at 2,454 and 2,800 m, although not at 1,780 or 2,085 m. The increase in EPO varied markedly among individuals, ranging from -41 to 400% after 24 h at 2,800 m. Similar to EPO, urine PO(2) decreased after 6 h at all altitudes and returned to baseline by 24 h at the two lowest altitudes but remained decreased at the two highest altitudes. Urine PO(2) was closely related to EPO via a curvilinear relationship (r(2) = 0.99), although also with prominent individual variability. Renal blood flow remained unchanged at all altitudes. Sa(O(2)) decreased slightly after 6 h at the lowest altitudes but decreased more prominently at the highest altitudes. There were only modest, albeit statistically significant, relationships between EPO and Sa(O(2)) (r = 0.41, P < 0.05) and no significant relationship with renal O(2) delivery. These data suggest that 1) the altitude-induced increase in EPO is "dose" dependent: altitudes > or =2,100-2,500 m appear to be a threshold for stimulating sustained EPO release in most subjects; 2) short-term acclimatization may restore renal tissue oxygenation and restrain the rise in EPO at the lowest altitudes; and 3) there is marked individual variability in the erythropoietic response to altitude that is only partially explained by "upstream" physiological factors such as those reflecting O(2) delivery to EPO-producing tissues.

Non-NASA Center↗
NASA NTRS1998

F100 Engine Emissions Tested in NASA Lewis' Propulsion Systems Laboratory

Recent advances in atmospheric sciences have shown that the chemical composition of the entire atmosphere of the planet (gases and airborne particles) has been changed due to human activity and that these changes have changed the heat balance of the planet. National Research Council findings indicate that anthropogenic aerosols1 reduce the amount of solar radiation reaching the Earth's surface. Atmospheric global models suggest that sulfate aerosols change the energy balance of the Northern Hemisphere as much as anthropogenic greenhouse gases have. In response to these findings, NASA initiated the Atmospheric Effects of Aviation Project (AEAP) to advance the research needed to define present and future aircraft emissions and their effects on the Earth's atmosphere. Although the importance of aerosols and their precursors is now well recognized, the characterization of current subsonic engines for these emissions is far from complete. Furthermore, since the relationship of engine operating parameters to aerosol emissions is not known, extrapolation to untested and unbuilt engines necessarily remains highly uncertain. Tests in 1997-an engine test at the NASA Lewis Research Center and the corresponding flight measurement test at the NASA Langley Research Center-attempted to address both issues by measuring emissions when fuels containing different levels of sulfur were burned. Measurement systems from four research groups were involved in the Lewis engine test: A Lewis gas analyzer suite to measure the concentration of gaseous species 1. including NO, NOx, CO, CO2, O2, THC, and SO2 as well as the smoke number; 2. A University of Missouri-Rolla Mobile Aerosol Sampling System to measure aerosol and particulate properties including the total concentration, size distribution, volatility, and hydration property; 3. An Air Force Research Laboratory Chemical Ionization Mass Spectrometer to measure the concentration of SO2 and SO3/H2SO4; and 4. An Aerodyne Research Inc. Tunable Diode Laser System to measure the concentrations of SO2, SO3, NO, NO2, CO2, and H2O. By September 1997, an F100 engine operating at several power levels at sea level and up to six simulated altitudes had been tested with commercial jet fuels with three levels of sulfur content and one military jet fuel. The data are being vigorously analyzed. A complete report is anticipated for the 1998 Atmospheric Effects of Aviation Project Annual Conference.

Wey, Chowen C.↗
NASA NTRS1961

Aerodynamic Characteristics of Parachutes at Mach Numbers from 1.6 to 3

A wind-tunnel investigation was conducted to determine the parameters affecting the aerodynamic performance of drogue parachutes in the Mach number range from 1.6 to 3. Flow studies of both rigid and flexible-parachute models were made by means of high-speed schlieren motion pictures and drag coefficients of the flexible-parachute models were measured at simulated altitudes from about 50,000 to 120,000 feet.

[Tunnel] UPWT↗
NASA NTRS2008

Feasibility of Conducting J-2X Engine Testing at the Glenn Research Center Plum Brook Station B-2 Facility

A trade study of the feasibility of conducting J-2X testing in the Glenn Research Center (GRC) Plum Brook Station (PBS) B-2 facility was initiated in May 2006 with results available in October 2006. The Propulsion Test Integration Group (PTIG) led the study with support from Marshall Space Flight Center (MSFC) and Jacobs Sverdrup Engineering. The primary focus of the trade study was on facility design concepts and their capability to satisfy the J-2X altitude simulation test requirements. The propulsion systems tested in the B-2 facility were in the 30,000-pound (30K) thrust class. The J-2X thrust is approximately 10 times larger. Therefore, concepts significantly different from the current configuration are necessary for the diffuser, spray chamber subsystems, and cooling water. Steam exhaust condensation in the spray chamber is judged to be the key risk consideration relative to acceptable spray chamber pressure. Further assessment via computational fluid dynamics (CFD) and other simulation capabilities (e.g. methodology for anchoring predictions with actual test data and subscale testing to support investigation.

Schafer, Charles F.↗
NASA NTRS2008

Future NTP Development Synergy Leveraged from Current J-2X Engine Development

This paper is a discussion of how the many long-lead development elements required for the realization of a future nuclear thermal propulsion (NTP) system can be effectively leveraged from the ongoing work being conducted on the J-2X engine program for the Constellation Program. Development studies conducted to date for NTP forward planning have identified a number of technical areas that will require advancement to acceptable technology readiness levels (TRLs) before they can be utilized in NTP system development. These include high-temperature, high-area ratio nozzle extension; long-life, low-NPSP. turbomachinery; and low-boiloff propellant management; and a qualified nuclear fuel element. The current J-2X program is working many of these areas that can be leveraged to support NTP development in a highly compatible and synergistic fashion. In addition to supporting technical development, there are other programmatic issues being worked in the J-2X program that can be leveraged by a future NTP development program. These include compliance with recently-evolved space system requirements such as human-rating, fault tolerance and fracture control. These and other similar mandatory system requirements have been adopted by NASA and can result in a significant technical impact beyond elevation of the root technologies required by NTP. Finally, the exploitation of experience, methodologies, and procedures developed by the J-2X program in the areas of verification, qualification, certification, altitude simulation testing, and facility definition will be especially applicable to a future NTP system. The similarities in system mission (in-space propulsion) and operational environment (vacuum, zero-gee) between J-2X and NTP make this highly synergistic. Thus, it can be $hown that the collective benefit of leveraging experience and technologies developed during the J-2X program can result in significant savings in development cost and schedule for NTP.

Ballard, Richard O.↗
NASA NTRS1941

Performance Characteristics of an Aircraft Engine with Exhaust Turbine Supercharger, Special Report

The Pratt and Whitney Aircraft company and the Naval Aircraft Factory of the United States Navy cooperated in a laboratory and flight program of tests on an exhaust turbine supercharger. Two series of dynamometer tests of the engine super-charger combination were completed under simulated altitude conditions. One series of hot gas-chamber tests was conducted by the manufacturer of the supercharger. Flight demonstrations of the supercharger installed in a twin-engine flying boat were terminated by failure of the turbine wheels. The analysis of the results indicated that a two-stage supercharger with the first-stage exhaust turbine driven will deliver rated power for a given indicated power to a higher altitude, will operate more efficiently, and will require simpler controls than a similar engine with the first stage of the supercharger driven from the crankshaft through multispeed gears.

Lester, E. M.↗
NASA NTRS2009

An Overview of Follow-On Testing Activities of the A-3 Subscale Diffuser Test Project

An overview of NASA Stennis Space Center's (SSC) A-3 Subscale Diffuser Test (SDT) Project is presented. The original scope of the SDT Project, conducted from April 2007 to January 2008, collected data to support mitigation of risk associated with design and procurement activities of the A-3 Test Stand Project, an effort to construct a simulated altitude test facility at SSC in support of NASA's Constellation Program. Follow-on tests were conducted from May 2008 through August 2009, utilizing the SDT test setup as a testbed for additional risk mitigation activities. Included are descriptions of the Subscale Diffuser (SD) test article, the test facility configuration, and test approaches.

Ryan, James E.↗
NASA NTRS2011

Liquid Oxygen/Liquid Methane Integrated Propulsion System Test Bed

In support of NASA?s Propulsion and Cryogenic Advanced Development (PCAD) project, a liquid oxygen (LO2)/liquid methane (LCH4) Integrated Propulsion System Test Bed (IPSTB) was designed and advanced to the Critical Design Review (CDR) stage at the Johnson Space Center. The IPSTB?s primary objectives are to study LO2/LCH4 propulsion system steady state and transient performance, operational characteristics and to validate fluid and thermal models of a LO2/LCH4 propulsion system for use in future flight design work. Two phase thermal and dynamic fluid flow models of the IPSTB were built to predict the system performance characteristics under a variety of operating modes and to aid in the overall system design work. While at ambient temperature and simulated altitude conditions at the White Sands Test Facility, the IPSTB and its approximately 600 channels of system instrumentation would be operated to perform a variety of integrated main engine and reaction control engine hot fire tests. The pressure, temperature, and flow rate data collected during this testing would then be used to validate the analytical models of the IPSTB?s thermal and dynamic fluid flow performance. An overview of the IPSTB design and analytical model development will be presented.

Flynn, Howard↗
NASA NTRS2015

ME8373 Spring 2015 ICME Proposal ICME Analysis of Fatigue Crack Growth Through a Weld in SA-516 Grade 70 Plate

This paper describes work accomplished to predict the service life of a flexure joint design which is a component of a diffuser duct in the A3 Test Stand, an altitude simulation rocket engine test facility at NASA's Stennis Space Center. The duct has two pressure shells separated by cooling water passages and connected by stiffening ribs and flexure joints. Rocket exhaust flows within the duct and heats the inner pressure shell while the outer pressure shell remains at ambient temperature. The flexure joints allow for differential thermal expansion of the inner and outer pressure shells and are subject to in-service loading by this thermal expansion along with water pressure in the cooling water passage, atmospheric pressure outside the duct, near vacuum conditions within the duct, and vibrational loads from operation of the facility and rocket engine. Figure 1 shows a schematic axisymmetric cross section of the diffuser pressure shells and flexure joints with a zoomed in view of the flexure joint. The flexure joints are expected to eventually fail by fatigue cracking leading to leaks from the cooling water passages to the outside. The zoomed in view in Figure 1 indicates where cracking is expected to occur, namely through a weld bead between two plates of SA-516 Grade 70 steel. This weld bead acts as the fulcrum of the flexure joint and it is clear from inspection of the geometry and loading represented in the zoomed in portion of Figure 1 that inherent in the design there is a severe notch formed between the flexure plate, weld bead, and stiffening ring that will be the site of crack initiation and location from which the crack grows to the outer surface of the weld bead.

Woods, Jody L.↗
NASA NTRS2017

Integrated Cryogenic Propulsion Test Article Thermal Vacuum Hotfire Testing

In support of a facility characterization test, the Integrated Cryogenic Propulsion Test Article (ICPTA) was hotfire tested at a variety of simulated altitude and thermal conditions in the NASA Glenn Research Center Plum Brook Station In-Space Propulsion Thermal Vacuum Chamber (formerly B2). The ICPTA utilizes liquid oxygen and liquid methane propellants for its main engine and four reaction control engines, and uses a cold helium system for tank pressurization. The hotfire test series included high altitude, high vacuum, ambient temperature, and deep cryogenic environments, and several hundred sensors on the vehicle collected a range of system level data useful to characterize the operation of an integrated LOX/Methane spacecraft in the space environment - a unique data set for this propellant combination.

Morehead, Robert L.↗
NASA NTRS2016

Extreme Environments Test Capabilities at NASA GRC for Parker Hannifin Visit

The presentation includes general description on the following test facilities: Fuel Cell Testing Lab, Structural Dynamics Lab, Thermal Vacuum Test Facilities - including a description of the proposed Kinetic High Altitude Simulator concept, EMI Test Lab, and the Creek Road Cryogenic Complex - specifically the Small Multi-purpose Research Facility (SMiRF) and the Cryogenics Components Lab 7 (CCL-7).

Extreme Environments Test Capabilities↗