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At least 1,153 records · Page 64

All sources NASA NTRS DOE OSTI

NASA NTRS1995

Aerodynamic Performance of a Supersonic Throughflow Fan Stage With a Convergent Flow Path

The results from an experimental evaluation of a supersonic throughflow (SSTF) fan stage with a convergent flow path are presented. The stage was designed for a pressure ratio of 2.46 at a tip speed of 1350 ft/s and an inlet Mach number 2.00. The stage was tested with flows at subsonic and supersonic inlet axial Mach numbers and with rotor speeds from 40 to 110 percent of design speed. The fan stage overall performance characteristics, as well as details of its performanc characteristics downstream of the stator vanes, are presented. These measurements were obtained with radially and circumferentially traversing probes. At the design condition, the measured overall pressure ratio was 2.366, and the corresponding adiabatic efficiency was 0.610. Transition from subsonic to supersonic flow was demonstrated at 80, 100 and 110 percent of design speed. At the design condition, the radial distribution of energy addition (temperature ratio) from the computational fluid dynamics (CFD) calculations is in excellent agreement with the measured values. The predicted pressure distributions agree reasonably well with those measured.

Royce D. Moore↗

NASA NTRS2002

Advanced Control Surface Seal Development at NASA GRC for 3rd Generation RLV

NASA's Glenn Research Center (GRC) is developing advanced control surface seal technologies for the 3rd Generation RLV program. GRC has designed several new test rigs to simulate the temperatures, pressures, and scrubbing conditions that seals would have to endure during service. A hot compression test rig and hot scrub test rig are being developed to perform tests at temperatures up to 3000 F. Another new test rig allows simultaneous seal flow and scrub tests at room temperature to evaluate changes in seal performance with scrubbing. High temperature ceramic preloading devices are being developed and evaluated as ways to improve seal resiliency.

Dunlap, Patrick H., Jr.↗

NASA NTRS2003

Advanced Control Surface Seal Development at NASA GRC for Future Space Launch Vehicles

NASA s Glenn Research Center (GRC) is developing advanced control surface seal technologies for future space launch vehicles as part of the Next Generation Launch Technology project (NGLT). New resilient seal designs are currently being fabricated and high temperature seal preloading devices are being developed as a means of improving seal resiliency. GRC has designed several new test rigs to simulate the temperatures, pressures, and scrubbing conditions that seals would have to endure during service. A hot compression test rig and hot scrub test rig have been developed to perform tests at temperatures up to 3000 F. Another new test rig allows simultaneous seal flow and scrub tests at room temperature to evaluate changes in seal performance with scrubbing. These test rigs will be used to evaluate the new seal designs. The group is also performing tests on advanced TPS seal concepts for Boeing using these new test facilities.

Dunlap, Patrick H., Jr.↗

NASA NTRS2013

Flow Disturbance Measurements in the National Transonic Facility

Recent flow measurements have been acquired in the National Transonic Facility to assess the test-section unsteady flow environment. The primary purpose of the test is to determine the feasibility of the facility to conduct laminar-flow-control testing and boundary-layer transition-sensitive testing at flight-relevant operating conditions throughout the transonic Mach number range. The facility can operate in two modes, warm and cryogenic test conditions for testing full and semispan-scaled models. Data were acquired for Mach and unit Reynolds numbers ranging from 0.2 less than or equal to M less than or equal to 0.95 and 3.3 × 10(exp 6) less than Re/m less than 220×10(exp 6) collectively at air and cryogenic conditions. Measurements were made in the test section using a survey rake that was populated with 19 probes. Roll polar data at selected conditions were obtained to look at the uniformity of the flow disturbance field in the test section. Data acquired included mean total temperatures, mean and fluctuating static/total pressures, and mean and fluctuating hot-wire measurements. This paper focuses primarily on the unsteady pressure and hot-wire results. Based on the current measurements and previous data, an assessment was made that the facility may be a suitable facility for ground-based demonstrations of laminar-flow technologies at flight-relevant conditions in the cryogenic mode.

King, Rudolph A.↗

NASA NTRS1988

Measured and predicted shock shapes for AFE configuration at Mach 6 in air and in CF4

Shock shapes and stand-off distances were obtained for the Aeroassist Flight Experiment configuration from Mach 6 tests in air and in CF4. Results were plotted for an angle-of attack range from -10 to 10 degrees and comparisons were made at selected angles with inviscid-flow predictions. Tests were performed in the Langley Research Center (LaRC) 20 inch Mach 6 Tunnel (air) at unit free-stream Reynolds numbers (N sub Re, infinity) of 2 million/ft and 0.6 million/ft and in the LaRC Hypersonic CF4 Tunnel at N sub Re, infinity = 0.5 million/ft and 0.3 million/ft. Within the range of these tests, N sub Re, infinity did not affect the shock shape or stand off distance, and the predictions were in good agreement with the measurements. The shock stand-off distance in CF4 was approximately half of that in air. This effect resulted from the differences in density ratio across the normal shock, which was approximately 12 in CF4 and 5 in air. In both test gases, the shock lay progressively closer to the body as angle of attack decreased.

Wells, William L.↗

NASA NTRS1992

Analysis of convection in immiscible liquid layers with novel particle tracking velocimetry

The problem under study is convective flow in immiscible liquid layers with one or two horizontal interfaces. In one-g the flow results primarily from the buoyancy force acting perpendicular to the interfaces. This creates a fluid mechanical system in which the coupling of the fluid layers across an interface plays a fundamental role. The contribution of two horizontal interface tension forces is marginal. Interface tension driven flow requires testing in microgravity. A flight experiment on the Bubble, Drop, and Particle Unit (BDPU) is planned for the second International Microgravity Laboratory (IML-2) mission onboard the Shuttle in 1994. The flow velocity fields will be analyzed by a whole-field Particle Displacement Tracking (PDT) velocimetry technique. The capabilities of this technique to address fundamental issues, such as those regarding the flow stucture, will be discussed with a few sample experiments. Experimental and numerical flow patterns are compared.

Koster, J. N.↗

NASA NTRS2019

Solid Fuel Ignition and Extinction (SoFIE) Project on ISS

The Solid Fuel Ignition and Extinction (SoFIE) project studies ignition and flammability of solid spacecraft materials (fuels) in practical geometries and realistic atmospheric conditions. It is an experiment insert designed for use within the existing Combustion Integrated Rack (CIR). The CIR chamber provides a level of containment and permits testing at variable oxygen concentrations and pressures representative of current and planned NASA Space Exploration Atmospheres. The applications of SoFIE include: (1) Determining safer selection of cabin materials and validating NASA materials flammability selection using 1-g test protocols for low-gravity fires, (2) Improving understanding of early fire growth behavior, (3) Validating material flammability numerical models, (4) Determining optimal suppression techniques for burning materials by diluents, flow reduction, and venting, (5) Obtaining burning behavior of actual engineering materials planned for spacecraft, (6) Developing corresponding models of microgravity flame spread, flammability, and extinction, and use the results to improve normal gravity combustion models for terrestrial applications. The hardware permits a wide range of solid-material combustion and fire suppression studies. It supports multiple investigations using common infrastructure including sample holders, flow control, test sections, external radiant heaters, igniters, and diagnostics. SoFIE has been developed to meet the requirements of five unique investigations. It is currently being built and slated to begin operations on the ISS in July 2021. Given the general capabilities of the hardware insert, it is intended to be used as a facility for future researchers who can propose to NASA for related solid combustion studies.

Ferkul, Paul↗

DOE OSTI2012

Customer interface document for the Molten Salt Test Loop (MSTL) system

The National Solar Thermal Test Facility at Sandia National Laboratories has a unique test capability called the Molten Salt Test Loop (MSTL) system. MSTL is a test capability that allows customers and researchers to test components in flowing, molten nitrate salt. The components tested can range from materials samples, to individual components such as flex hoses, ball joints, and valves, up to full solar collecting systems such as central receiver panels, parabolic troughs, or linear Fresnel systems. MSTL provides realistic conditions similar to a portion of a concentrating solar power facility. The facility currently uses 60/40 nitrate 'solar salt' and can circulate the salt at pressure up to 600psi, temperature to 585 C, and flow rate of 400-600GPM depending on temperature. The purpose of this document is to provide a basis for customers to evaluate the applicability to their testing needs, and to provide an outline of expectations for conducting testing on MSTL. The document can serve as the basis for testing agreements including Work for Others (WFO) and Cooperative Research and Development Agreements (CRADA). While this document provides the basis for these agreements and describes some of the requirements for testing using MSTL and on the site at Sandia, the document is not sufficient by itself as a test agreement. The document, however, does provide customers with a uniform set of information to begin the test planning process.

14 SOLAR ENERGY↗

NASA NTRS1980

Atomizing characteristics of swirl can combustor modules with swirl blast fuel injectors

Cold flow atomization tests of several different designs of swirl can combustor modules were conducted in a 7.6 cm diameter duct at airflow rates (per unit area) of 7.3 to 25.7 g/sq cm sec and water flow rates of 6.3 to 18.9 g/sec. The effect of air and water flow rates on the mean drop size of water sprays produced with the swirl blast fuel injectors were determined. Also, from these data it was possible to determine the effect of design modifications on the atomizing performance of various fuel injector and air swirler configurations. The trend in atomizing performance, as based on the mean drop size, was then compared with the trends in the production of nitrogen oxides obtained in combustion studues with the same swirl can combustors. It was found that the fuel injector design that gave the best combustor performance in terms of a low NOx emission index also gave the best atomizing performance as characterized by a spray of relatively small mean drop diameter. It was also demonstrated that at constant inlet air stream momentum the nitrogen oxides emission index was found to vary inversely with the square of the mean drop diameter of the spray produced by the different swirl blast fuel injectors. Test conditions were inlet air static pressures of 100,000 to 200,000 N/sq m at an inlet air temperature of 293 K.

Ingebo, R. D.↗

NASA NTRS1985

LDA seeding system for the Langley Low Turbulence Pressure Tunnel

A Laser Velocimetry (LV) seeding system was specifically developed for the Langley Low Turbulence Wind Tunnel (LTPT), and it has been successfully used for LV measurements in two major tests (Juncture Flow Experiment and Gortler Experiment). The LTPT is capable of operating at Mach numbers from 0.05 to 0.50 and unit Reynolds numbers from 100,000 to 15,000,000 per foot. The test section is 3 feet wide and 7.5 feet high. The turbulence level in the test section is relatively low because of the high contraction ratio and because of the nine turbulence reduction screens in the settling chamber. A primary requirement of the seeding system was that the seeding material not contaminate or damage in any way these screens. Both solid and liquid seeding systems were evaluated, and the results are presented. They can provide some guidelines for setting up seeding systems in other similar tunnels.

Scheiman, J.↗

NASA NTRS1994

Operating characteristics of the Russian D-55 thruster with anode layer

Performance measurements of a Russian engineering-model Thruster with Anode Layer (TAL) were obtained as part of a program to evaluate the operating characteristics of Russian Hall-thruster technology. The TAL model D-55 was designed to operate in the 1-2 kW power range on xenon. When received, the thruster had undergone only a few hours of acceptance testing by the manufacturer. Direct thrust measurements were obtained at a background pressure of 0.0003 Pa (2 x 10(exp -6) torr) at power levels ranging from 0.3 kW to 2.1 kW. At the nominal power level of 1.3 kW, a specific impulse level of 1600 s with a corresponding efficiency of 0.48 was attained. At all flow rates tested, the efficiency increased linearly with specific impulse until a maximum was reached, and then the efficiency leveled off. Increasing the anode flow rate shifted the efficiency upward, reaching 0.50 at 1850 s specific impulse. The thruster was equipped with inner and outer electromagnets which were isolated from the discharge and from each other. Variation of the magnetic field, obtained by changing the currents through the magnets, had little effect on performance, except at current levels below 70 percent of nominal. For a given operating condition, the performance was slightly affected by facility pressure. As the pressure was increased by a factor of thirty to 0.008 Pa (6 x 10(exp -5) torr), the current steadily increased by 4 percent, and the thrust increased by 2 percent. Performance comparisons were made with the Stationary Plasma Thruster, and the efficiency and specific impulse values were similar at power levels ranging from 0.9 kW to 1.5 kW. Endurance testing was not performed, and comparisons of lifetime were not made.

Sankovic, John M.↗

NASA NTRS1993

Flow visualization of mast-mounted-sight/main rotor aerodynamic interactions

Flow visualization tests were conducted on a 27 percent-scale AH-64 attack helicopter model fitted with various mast-mounted-sight configurations in an attempt to identify the cause of adverse vibration encountered during full-scale flight tests of an Apache/Longbow configuration. The tests were conducted at the NASA Langley Research Center in the 14- by 22-Foot Subsonic Tunnel. A symmetric and an asymmetric mast-mounted-sight oriented at several skew angles were tested at forward and rearward flight speeds of 30 and 45 knots. A laser light sheet seeded with vaporized propylene glycol was used to visualize the wake of the sight in planes parallel and perpendicular to the freestream flow. Analysis of the flow visualization data identified the frequency of the wake shed from the sight, the angle-of-attack at the sight, and the location where the sight wake crossed the rotor plane. Differences in wake structure were observed between the various sight configurations and slew angles. Postulations into the cause of the adverse vibration found in flight test are given along with considerations for future tests.

Ghee, Terence A.↗

NASA NTRS1986

Effect of Reynolds Number and Mach Number on flow angularity probe sensitivity

Preliminary calibrations were performed on nine flow angularity probes in the Langley 7- by 10-Foot High-Speed Tunnel (7 x 10 HST) and the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). These probes will be used in surveying the test section flows of the National Transonic Facility (NTF). The probes used in this study have a pyramid head with five pressure orifices. The calibrations consisted of both isolated probe measurements and rake-mounted multiprobe measurements that covered a range of subsonic Mach numbers up to 0.90 and Reynolds numbers per foot up to 40 X 10 to the 6th power. The preliminary calibration in the 7 x 10 HST included testing the probes both individually and in a rake. The 0.3-m TCT calibration tested two probes singly at varying Reynolds numbers. The results from these tests include Mach number, Reynolds number, and rake-mounting effects. The results of these tests showed probe sensitivity to be slightly affected by Mach number. At Reynolds numbers per foot above 10 x 10 to the 6th power, the probe did not exhibit a Reynolds number sensitivity.

Smith, L. A.↗

DOE OSTI2011

Note: Ultrasonic gas flowmeter based on optimized time-of-flight algorithms

A new digital signal processor based single path ultrasonic gas flowmeter is designed, constructed, and experimentally tested. To achieve high accuracy measurements, an optimized ultrasound driven method of incorporation of the amplitude modulation and the phase modulation of the transmit-receive technique is used to stimulate the transmitter. Based on the regularities among the received envelope zero-crossings, different received signal's signal-to-noise ratio situations are discriminated and optional time-of-flight algorithms are applied to take flow rate calculations. Experimental results from the dry calibration indicate that the designed flowmeter prototype can meet the zero-flow verification test requirements of the American Gas Association Report No. 9. Furthermore, the results derived from the flow calibration prove that the proposed flowmeter prototype can measure flow rate accurately in the practical experiments, and the nominal accuracies after FWME adjustment are lower than 0.8% throughout the calibration range.

46 INSTRUMENTATION RELATED TO NUCLEAR SCIENCE AND ↗

NASA NTRS1980

Atomizing characteristics of swirl can combustor modules with swirl blast fuel injectors

Ingebo, R. D.↗

NASA NTRS2007

Solar Airplanes and Regenerative Fuel Cells

A solar electric aircraft with the potential to "fly forever" has captured NASA's interest, and the concept for such an aircraft was pursued under Aeronautics Environmental Research Aircraft and Sensor Technology (ERAST) project. Feasibility of this aircraft happens to depend on the successful development of solar power technologies critical to NASA's Exploration Initiatives; hence, there was widespread interest throughout NASA to bring these technologies to a flight demonstration. The most critical is an energy storage system to sustain mission power during night periods. For the solar airplane, whose flight capability is already limited by the diffuse nature of solar flux and subject to latitude and time of year constraints, the feasibility of long endurance flight depends on a storage density figure of merit better than 400-600 watt-hr per kilogram. This figure of merit is beyond the capability of present day storage technologies (other than nuclear) but may be achievable in the hydrogen-oxygen regenerative fuel cell (RFC). This potential has led NASA to undertake the practical development of a hydrogen-oxygen regenerative fuel cell, initially as solar energy storage for a high altitude UAV science platform but eventually to serve as the primary power source for NASAs lunar base and other planet surface installations. Potentially the highest storage capacity and lowest weight of any non-nuclear device, a flight-weight RFC aboard a solar-electric aircraft that is flown continuously through several successive day-night cycles will provide the most convincing demonstration that this technology's widespread potential has been realized. In 1998 NASA began development of a closed cycle hydrogen oxygen PEM RFC under the Aeronautics Environmental Research Aircraft and Sensor Technology (ERAST) project and continued its development, originally for a solar electric airplane flight, through FY2005 under the Low Emissions Alternative Power (LEAP) project. Construction of the closed loop system began in 2002 at the NASA Glenn Research Center in Cleveland, Ohio. System checkout was completed, and testing began, in July of 2003. The initial test sequences were done with only a fuel cell or electrolyzer in the test rig. Those tests were used to verify the test apparatus, procedures, and software. The first complete cycles of the fully closed loop, regenerative fuel cell system were successfully completed in the following September. Following some hardware upgrades to increase reactant recirculation flow, the test rig was operated at full power in December 2003 and again in January 2004. In March 2004 a newer generation of fuel cell and electrolyzer stacks was substituted for the original hardware and these stacks were successfully tested at full power under cyclic operation in June of 2004.

Bents, David J.↗

DOE OSTI2011

Effects of additives on the stability of electrolytes for all-vanadium redox flow batteries

The stability of the electrolytes for all-vanadium redox flow battery was investigated with ex-situ heating/cooling treatment and in-situ flow-battery testing methods. The effects of inorganic and organic additives have been studied. The additives containing the ions of potassium, phosphate, and polyphosphate are not suitable stabilizing agents because of their reactions with V(V) ions, forming precipitates of KVSO6 or VOPO4. Of the chemicals studied, polyacrylic acid and its mixture with CH3SO3H are the most promising stabilizing candidates which can stabilize all the four vanadium ions (V2+, V3+, VO2+, and VO2+) in electrolyte solutions up to 1.8 M. However, further effort is needed to obtain a stable electrolyte solution with >1.8 M V5+ at temperatures higher than 40 °C.

74 ATOMIC AND MOLECULAR PHYSICS↗

NASA NTRS1999

New Icing Cloud Simulation System at the NASA Glenn Research Center Icing Research Tunnel

A new spray bar system was designed, fabricated, and installed in the NASA Glenn Research Center's Icing Research Tunnel (IRT). This system is key to the IRT's ability to do aircraft in-flight icing cloud simulation. The performance goals and requirements levied on the design of the new spray bar system included increased size of the uniform icing cloud in the IRT test section, faster system response time, and increased coverage of icing conditions as defined in Appendix C of the Federal Aviation Regulation (FAR), Part 25 and Part 29. Through significant changes to the mechanical and electrical designs of the previous-generation spray bar system, the performance goals and requirements were realized. Postinstallation aerodynamic and icing cloud calibrations were performed to quantify the changes and improvements made to the IRT test section flow quality and icing cloud characteristics. The new and improved capability to simulate aircraft encounters with in-flight icing clouds ensures that the 1RT will continue to provide a satisfactory icing ground-test simulation method to the aeronautics community.

Irvine, Thomas B.↗