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At least 379 records · Page 21

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NASA NTRS2005

Linear Thermal Expansion Measurements of Lead Magnesium Niobate (PMN) Electroceramic Material for the Terrestrial Planet Finder Coronagraph

Linear thermal expansion measurements of nine samples of Lead Magnesium Niobate (PMN) electroceramic material were recently performed in support of NASA's Terrestrial Planet Finder Coronagraph (TPF-C) mission. The TPF-C mission is a visible light coronagraph designed to look at roughly 50 stars pre- selected as good candidates for possessing earth-like planets. Upon detection of an earth-like planet, TPF-C will analyze the visible-light signature of the planet's atmosphere for specific spectroscopic indicators that life may exist there. With this focus, the project's primary interest in PMN material is for use as a solid-state actuator for deformable mirrors or compensating optics. The nine test samples were machined from three distinct boules of PMN ceramic manufactured by Xinetics Inc. Thermal expansion measurements were performed in 2005 at NASA Jet Propulsion Laboratory (JPL) in their Cryogenic Dilatometer Facility. All measurements were performed in vacuum with sample temperature actively controlled over the range of 270K to 3 10K. Expansion and contraction of the test samples with temperature was measured using a JPL developed interferometric system capable of sub-nanometer accuracy. Presented in this paper is a discussion of the sample configuration, test facilities, test method, data analysis, test results, and future plans.

Terrestrial Planet Finder (TPF)↗

NASA NTRS1997

Wind Tunnel Measured Effects on a Twin-Engine Short-Haul Transport Caused by Simulated Ice Accretions: Data Report

The purpose of this report is to release the data from the NASA Langley/Lewis 14 by 22 foot wind tunnel test that examined icing effects on a 1/8 scale twin-engine short-haul jet transport model. Presented in this document are summary data from the major configurations tested. The entire test database in addition to ice shape and model measurements is available as a data supplement in CD-ROM form. Data measured and presented are: wing pressure distributions, model force and moment, and wing surface flow visualization.

Reehorst, Andrew↗

NASA NTRS2017

CFD Simulations of the IHF Arc-Jet Flow: Compression-Pad/Separation Bolt Wedge Tests

This paper reports computational analyses in support of two wedge tests in a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using two different wedge models, each placed in a free jet downstream of a corresponding different conical nozzle in the Ames 60-MW Interaction Heating Facility. Each panel test article included a metallic separation bolt imbedded in Orion compression-pad and heatshield materials, resulting in a circular protuberance over a flat plate. The protuberances produce complex model flowfields, containing shock-shock and shock-boundary layer interactions, and multiple augmented heating regions on the test plate. As part of the test calibration runs, surface pressure and heat flux measurements on water-cooled calibration plates integrated with the wedge models were also obtained. Surface heating distributions on the test articles as well as arc-jet test environment parameters for each test configuration are obtained through computational fluid dynamics simulations, consistent with the facility and calibration measurements. The present analysis comprises simulations of the nonequilibrium flowfield in the facility nozzle, test box, and flowfield over test articles, and comparisons with the measured calibration data.

CFD↗

NASA NTRSDate not supplied

Pumping Loss of Shrouded Meshed Spur Gears

High speed rotorcraft transmissions are subject to load-independent power losses consisting of drag loss and pumping loss. Tightly conforming shrouds enclosing the transmission gears are often incorporated to reduce the drag component of the total load-independent losses. However, tightly conforming axial shrouds can result in an increase in the pumping loss component. Quantifying the pumping loss of shrouded gear transmissions has been the subject of many studies. This study presents a new approach for estimating pumping loss based on the concept of swept volume and examines the applicability of the approach to various shroud configurations.The drag loss and pumping loss of a shrouded spur gear pair has been determined through testing using the NASAGlenn Research Center (GRC) Gear Windage Test Facility.The results from this testing have been compared to theoretical results using the formulations presented in this study. In addition, computational fluid dynamic (CFD) analysis has been conducted for the various shroud configurations tested at NASA GRC. The results from the CFD analysis confirm the theoretical and empirical results and provide insight into the applicability of the swept volume approach for estimating pumping power loss of shrouded gear transmissions.

Michael J. Hurrell↗

NASA NTRS1986

Multiaxis aircraft control power from thrust vectoring at high angles of attack

Extensive research programs conducted at the Langley Research Center have shown that thrust vectoring can be provided by multifunction (nonaxisymmetric) nozzles. Most of this research has been conducted on pitch vectoring at both static and forward flight conditions. Recent efforts have been aimed at evaluating yaw vectoring concepts at static (wind off) conditions. This paper summarizes results for three different twin-engine fighter configurations tested over a Mach number range of 0.15 to 2.47 at angles of attack up to 35 deg. The objective of these investigations was to determine the multiaxis control power characteristics provided by thrust vectoring. All three configurations employed two-dimensional convergent-divergent nozzles which provided pitch vectoring by differential deflection of the upper and lower nozzle divergent flaps. Three different means of yaw vectoring were tested: (1) a translating nozzle sidewall; (2) yaw flaps located in the nozzle sidewalls; and (3) canted nozzles. These investigations were conducted in the Langley 16-Foot Transonic Tunnel and the Lewis 10x10-Foot Supersonic Tunnel. Longitudinal and direction control power from thrust vectoring was greater than that provided by aerodynamic control effectors at low speed or at high angles of attack.

Capone, F. J.↗

NASA NTRS1986

Multiaxis aircraft control power from thrust vectoring at high angles of attack

Extensive research programs conducted at the Langley Research Center have shown that thrust vectoring can be provided by multifunction (nonaxisymmetric) nozzles. Most of this research has been conducted on pitch vectoring at both static and forward flight conditions. Recent efforts have been aimed at evaluating yaw vectoring concepts at static (wind off) conditions. This paper summarizes results for three different twin-engine fighter configurations tested over a Mach number range of 0.15 to 2.47 at angles of attack up to 35 deg. The objective of these investigations was to determine the multiaxis control power characteristics provided by thrust vectoring. All three configurations employed two-dimensional convergent-divergent nozzles which provided pitch vectoring by differential deflection of the upper and lower nozzle divergent flaps. Three different means of yaw vectoring were tested: (1) a translating nozzle sidewall; (2) yaw flaps located in the nozzle sidewalls; and (3) canted nozzles. These investigations were conducted in the Langley 16-Foot Transonic Tunnel and the Lewis 10 x 10-Foot Supersonic Tunnel. Longitudinal and direction control power from thrust vectoring was greater than that provided by aerodynamic control effectors at low speed or at high angles of attack.

Capone, Francis J.↗

NASA NTRS1946

Wind-Tunnel Tests to Determine Aileron Characteristics of the McDonnell XFD-1 Airplane, TED No. NACA 23102

Tests were performed on a partial span of the wing of a McDonnell XFD-1 airplane to determine a combination of sealed internal balance and spring-tab stiffness for the aileron that would give satisfactory stick-force characteristics for the airplane. Two sealed internal balances were tested in combination with spring tabs of various stiffnesses. One of the combinations was tested at several speeds to determine the variation of stick force with speed. Estimates, based on the results of the tests, indicate that for this airplane any reduction of stick force by use of the spring tab reduces the helix angle pb/2V below the required value of 0.09. The estimates show that, of the configurations tested, the most satisfactory combination for obtaining a stick force of 30 pounds at 300 miles per hour indicated airspeed is a 0.48-chord internal balance in combination with a spring-tab stiffness of 500 pounds per inch. With this combination, a wing-tip helix angle of 0.078 is estimated. Stick-force curves for all configurations show a rapid increase in stick force above approximately 20 deg. total aileron deflection.

Yates, Campbell C.↗

NASA NTRS1995

Fracture testing of 2324-T39 aluminum alloy

Fracture tests were conducted on middle crack tension (M(T)) and compact tension (C(T)) specimens of 7.6 mm-thick 2324-T39 aluminum in the L-T orientation. The configurations tested included 102 mm-wide M(T), 305 mm-wide M(T), 152 mm-wide C(T), and 102 mm-wide C(T) specimens. The measured critical crack-tip opening angle (CTOA) for the 102 mm-wide M(T) tests, the 305 mm-wide M(T) tests, and the 152 mm-wide C(T) tests, after a small amount of crack extension, were 4.4, 4.9, 5.1 degrees, respectively. The scatter involved with these measurements was about +/- 1 degree.

Dawicke, David S.↗

NASA NTRS1951

Results of Flight Tests to Determine Drag of Parabolic and Cone-cylinder Bodies of Very Large Fineness Ratios at Supersonic Speeds

Results of a free-flight investigation at supersonic speeds to determine zero-lift drag of a series of bodies of revolution are presented. Configurations tested included two parabolic bodies with fineness ratios of 17.78 and 24.5 and two 8 degree cone-cylinder bodies with fineness ratios of 17.2 and 21.2. Results of previous tests of similar parabolic bodies but with lower fineness ratios are included in this paper for comparison. All bodies for which data are presented in this paper had a base-to-maximum-diameter ratio of 0.437. Calculated drag coefficients are shown for all bodies for which data are presented. (author)

Welsh, Clement J↗

NASA NTRS1974

Static performance and noise tests on a thrust reverser for an augmentor wing aircraft

A 1/3 scale model static test program was conducted to measure the noise levels and reverse thrust performance characteristics of wing-mounted thrust reverser that could be used on an advanced augmentor wing airplane. The configuration tested represents only the most fundamental designs where installation and packaging restraints are not considered. The thrust reverser performance is presented in terms of horizontal, vertical, and resultant effectiveness ratios and the reverser noise is compared on the basis of peak perceived noise level (PNL) and one-third octave band data (OASPL). From an analysis of the model force and acoustic data, an assessment is made on the stopping distance versus noise for a 90,900 kg (200,000 lb) airplane using this type of thrust reverser.

Harkonen, D. L.↗

NASA NTRS1980

Static and transient performance of YF-102 engine with up to 14 percent core airbleed for the quiet short-haul research aircraft

An outdoor static test stand was used to measure the steady-state and transient performance of the YF-102 turbofan engine with core airbleed. The test configuration included a bellmouth inlet and a confluent-flow exhaust system similar in size to the quiet short-haul research aircraft (QSRA) exhaust system. For the steady-state tests, the engine operated satisfactorily with core bleed up to 14 percent of the core inlet flow. For the transient tests the engine accelerated and decelerated satisfactorily with no core bleed and with core bleed up to 11 percent of the core inlet flow (maximum tested). For some of the tests the core-bleed flow rate was scheduled to vary with fan discharge pressure, to simulate the QSRA bleed requirements. No stability, surge, stall, overtemperature, combustor flameout, or other operating problems were encountered in any of the tests. Steady-state and transient engine performance data are presented in graphs, and fuel-control trajectories for typical transient tests are shown.

Mcardle, J. G.↗

NASA NTRS1999

Low-Speed Aerodynamic Data for an 0.18-Scale Model of an F-16XL with Various Leading-Edge Modifications

Using the F-16XL as a test-bed, two strategies for improving the low-speed flying characteristics that had minimal impact on high-speed performance were evaluated. In addition to the basic F-16XL configuration several modifications to the baseline configuration were tested in the Langley 30- X 60-Foot Tunnel: 1) the notched area at the wing leading edge and fuselage juncture was removed resulting in a continuous 70 deg leading-edge sweep on the inboard portion of the wing; 2) an integral attached-flow leading-edge flap concept was added to the continuous leading edge; and 3) a deployable vortex flap concept was added to the continuous leading edge. The purpose of this report is simply to document the test configurations, test conditions, and data obtained in this investigation for future reference and analysis. No analysis is presented herein and the data only appear in tabulated format.

Hahne, Daniel E.↗

NASA NTRS2017

CFD Simulations of the IHF Arc-Jet Flow: Compression-Pad/Separation Bolt Wedge Tests

This paper reports computational analyses in support of two wedge tests in a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using two different wedge models, each placed in a free jet downstream of a corresponding different conical nozzle in the Ames 60-MW Interaction Heating Facility. Panel test articles included a metallic separation bolt imbedded in the compression-pad and heat shield materials, resulting in a circular protuberance over a flat plate. As part of the test calibration runs, surface pressure and heat flux measurements on water-cooled calibration plates integrated with the wedge models were also obtained. Surface heating distributions on the test articles as well as arc-jet test environment parameters for each test configuration are obtained through computational fluid dynamics simulations, consistent with the facility and calibration measurements. The present analysis comprises simulations of the non-equilibrium flow field in the facility nozzle, test box, and flow field over test articles, and comparisons with the measured calibration data.

TPS↗

NASA NTRS2017

CFD Simulations of the IHF Arc-Jet Flow: Compression-Pad Separation Bolt Wedge Tests

This paper reports computational analyses in support of two wedge tests in a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using two different wedge models, each placed in a free jet downstream of a corresponding different conical nozzle in the Ames 60-MW Interaction Heating Facility. Each panel test article included a metallic separation bolt imbedded in Orion compression-pad and heatshield materials, resulting in a circular protuberance over a flat plate. The protuberances produce complex model flowfields, containing shock-shock and shock-boundary layer interactions, and multiple augmented heating regions on the test plate. As part of the test calibration runs, surface pressure and heat flux measurements on water-cooled calibration plates integrated with the wedge models were also obtained. Surface heating distributions on the test articles as well as arc-jet test environment parameters for each test configuration are obtained through computational fluid dynamics simulations, consistent with the facility and calibration measurements. The present analysis comprises simulations of the non-equilibrium flow field in the facility nozzle, test box, and flow field over test articles, and comparisons with the measured calibration data.

CFD↗

NASA NTRS1973

Pitch paper pilot revisited

Two methods are described for predicting the handling qualities of an aircraft in pitch. Both methods are based on minimizing a pilot rating expression with respect to free pilot parameters. In one case the pilot rating expression was developed from a set of moving base simulation data. In the other case, the pilot rating expression was developed from a set of fixed base simulation data. In both cases pilot rating is primarily a function of pitch attitude and pitch rate; however, in the moving base case, the pilot rating is more sensitive to pitch rate than in the fixed base case. For each method, predicted pilot ratings and performance measures agree well with the measured data corresponding to that case. Both methods are used to predict pilot ratings for a set of flight test configurations and the results of the two methods agree to within one half a rating unit. The predicted ratings in both cases correlate well with the actual pilot ratings from the flight tests.

Arnold, J. D.↗

NASA NTRS1984

Water impact laboratory and flight test results for the space shuttle solid rocket booster aft skirt

A series of water impact tests was conducted using full-scale segment representations of the Space Shuttle Solid Rocket Booster (SRB) aft skirt structure. The baseline reinforced structural design was tested as well as various alternative design concepts. A major portion of the test program consisted of evaluating foam as a load attenuation material. Applied pressures and response strains were measured for impact velocities from 40 feet per second (ft/s) to 110 ft/s. The structural configurations, test articles, test results, and flight results are described.

Kross, D. A.↗

NASA NTRS1971

Evaluation of acoustic testing techniques for spacecraft systems

External acoustic environments, structural responses, noise reductions, and the internal acoustic environments have been predicted for a typical shroud/spacecraft system during lift-off and various critical stages of flight. Spacecraft responses caused by energy transmission from the shroud via mechanical and acoustic paths have been compared and the importance of the mechanical path has been evaluated. Theoretical predictions have been compared extensively with available laboratory and in-flight measurements. Equivalent laboratory acoustic fields for simulation of shroud response during the various phases of flight have been derived and compared in detail. Techniques for varying the time-space correlations of laboratory acoustic fields have been examined, together with methods for varying the time and spatial distribution of acoustic amplitudes. Possible acoustic testing configurations for shroud/spacecraft systems have been suggested and trade-off considerations have been reviewed. The problem of simulating the acoustic environments versus simulating the structural responses has been considered and techniques for testing without the shroud installed have been discussed.

Cockburn, J. A.↗

NASA NTRS1977

Low-speed power effects on advanced fighter configurations with two-dimensional deflected thrust

Wind-tunnel studies at the Langley Research Center have shown that significant increases in maximum lift coefficient and stability and decreases in drag due to lift are obtained when two-dimensional vectored thrust is used in conjuction with a close-coupled canard. The configuration tested was somewhat above the theoretical minimum drag due to lift because of the sharp leading edge on the biconvex airfoil used on the wing and canard. An effort to design a new configuration which will approach the minimum drag due to lift while maintaining high-lift configurations was completed. The resulting model will incorporate a realistic planform, airfoil section, and twist for a transonic maneuvering configuration.

Paulson, J. W., Jr.↗