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At least 1,603 records · Page 89
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

Two- and three-dimensional model and wall data from a flexible-walled transonic test section

Both two- and three-dimensional model testing is being carried out in the transonic flexible-walled wind tunnel test section. The test section has flexible top and bottom walls with rigid sidewalls. Interference is eliminated by adjustments based on data taken at walls in two dimensional models. Cast-7 data will illustrate agreement between various flexible-walled tunnels. In three-dimensional models interference cannot be eliminated but wall adjustments can control and relieve the principal sources of wall-induced errors. Estimates of magnitudes of the control which may be exercised on flow by movement of one wall jack are presented. A wall control algorithm (still in analytic development stage) based on use of this data is described. Brief examples of control of wall-induced perturbations in region of model are given.

Goodyer, M. J.↗
NASA NTRS1945

Effect of Propeller-Axis Angle of Attack on Thrust Distribution over the Propeller Disk in Relation to Wake-Survey Measurement of Thrust

Tests were made to investigate the variation of thrust distribution over the propeller disk with angle of pitch of the propeller thrust axis and to determine the disposition and the minimum number of rakes necessary to measure the propeller thrust. The tests were made at a low Mach number for a low and a high blade angle with the propeller operating at three small angles of pitch, and some of the tests were repeated at a higher Mach number. The data obtained show that, for small angles of pitch, large changes occur in the energy distribution in the wake which prohibit the use of a single survey rake for thrust measurement in flight tests and limit the use of a single rake in wind-tunnel tests. Under certain conditions, the energy distribution in the wake took on a symmetrical form and two diametrically opposed survey rakes were shown to be satisfactory for obtaining propeller thrust. (author)

Pendley, Robert E↗
NASA NTRS2002

Fan Noise Source Diagnostic Test Computation of Rotor Wake Turbulence Noise

An important source mechanism of fan broadband noise is the interaction of rotor wake turbulence with the fan outlet guide vanes. A broadband noise model that utilizes computed rotor flow turbulence from a RANS code is used to predict fan broadband noise spectra. The noise model is employed to examine the broadband noise characteristics of the 22-inch Source Diagnostic Test fan rig for which broadband noise data were obtained in wind tunnel tests at the NASA Glenn Research Center. A 9-case matrix of three outlet guide vane configurations at three representative fan tip speeds are considered. For all cases inlet and exhaust acoustic power spectra are computed and compared with the measured spectra where possible. In general, the acoustic power levels and shape of the predicted spectra are in good agreement with the measured data. The predicted spectra show the experimentally observed trends with fan tip speed, vane count, and vane sweep. The results also demonstrate the validity of using CFD-based turbulence information for fan broadband noise calculations.

Nallasamy, M.↗
NASA NTRS2003

X-33 Computational Aeroheating/Aerodynamic Predictions and Comparisons With Experimental Data

This report details a computational fluid dynamics study conducted in support of the phase II development of the X-33 vehicle. Aerodynamic and aeroheating predictions were generated for the X-33 vehicle at both flight and wind-tunnel test conditions using two finite-volume, Navier-Stokes solvers. Aerodynamic computations were performed at Mach 6 and Mach 10 wind-tunnel conditions for angles of attack from 10 to 50 with body-flap deflections of 0 to 20. Additional aerodynamic computations were performed over a parametric range of free-stream conditions at Mach numbers of 4 to 10 and angles of attack from 10 to 50. Laminar and turbulent wind-tunnel aeroheating computations were performed at Mach 6 for angles of attack of 20 to 40 with body-flap deflections of 0 to 20. Aeroheating computations were performed at four flight conditions with Mach numbers of 6.6 to 8.9 and angles of attack of 10 to 40. Surface heating and pressure distributions, surface streamlines, flow field information, and aerodynamic coefficients from these computations are presented, and comparisons are made with wind-tunnel data.

Hollis, Brian R.↗
NASA NTRS2014

Shielding Characteristics Using an Ultrasonic Configurable Fan Artificial Noise Source to Generate Modes - Experimental Measurements and Analytical Predictions

An Ultrasonic Configurable Fan Artificial Noise Source (UCFANS) was designed, built, and tested in support of the NASA Langley Research Center's 14x22 wind tunnel test of the Hybrid Wing Body (HWB) full 3-D 5.8% scale model. The UCFANS is a 5.8% rapid prototype scale model of a high-bypass turbofan engine that can generate the tonal signature of proposed engines using artificial sources (no flow). The purpose of the program was to provide an estimate of the acoustic shielding benefits possible from mounting an engine on the upper surface of a wing; a flat plate model was used as the shielding surface. Simple analytical simulations were used to preview the radiation patterns - Fresnel knife-edge diffraction was coupled with a dense phased array of point sources to compute shielded and unshielded sound pressure distributions for potential test geometries and excitation modes. Contour plots of sound pressure levels, and integrated power levels, from nacelle alone and shielded configurations for both the experimental measurements and the analytical predictions are presented in this paper.

Artifical Noise↗
NASA NTRS2014

Hybrid Wing Body Shielding Studies Using an Ultrasonic Configurable Fan Artificial Noise Source Generating Typical Turbofan Modes

An Ultrasonic Configurable Fan Artificial Noise Source (UCFANS) was designed, built, and tested in support of the NASA Langley Research Center's 14x22 wind tunnel test of the Hybrid Wing Body (HWB) full 3-D 5.8% scale model. The UCFANS is a 5.8% rapid prototype scale model of a high-bypass turbofan engine that can generate the tonal signature of proposed engines using artificial sources (no flow). The purpose of the test was to provide an estimate of the acoustic shielding benefits possible from mounting the engine on the upper surface of an HWB aircraft using the projected signature of the engine currently proposed for the HWB. The modal structures at the rating points were generated from inlet and exhaust nacelle configurations - a flat plate model was used as the shielding surface and vertical control surfaces with correct plan form shapes were also tested to determine their additional impact on shielding. Radiated acoustic data were acquired from a traversing linear array of 13 microphones, spanning 36 inches. Two planes perpendicular, and two planes parallel, to the axis of the nacelle were acquired from the array sweep. In each plane the linear array traversed 4 sweeps, for a total span of 168 inches acquired. The resolution of the sweep is variable, so that points closer to the model are taken at a higher resolution. Contour plots of Sound Pressure Levels, and integrated Power Levels, from nacelle alone and shielded configurations are presented in this paper; as well as the in-duct mode power levels.

Sutliff, Daniel L.↗
NASA NTRS2019

Preliminary Results from an Experimental Assessment of a Natural Laminar Flow Design Method

A 5.2% scale semispan model of the new Common Research Model with Natural Laminar Flow (CRM-NLF) was tested in the National Transonic Facility (NTF) at the NASA Langley Research Center. The model was tested at transonic cruise flight conditions with Reynolds numbers based on mean aerodynamic chord ranging from 10 to 30 million. The goal of the test was to experimentally validate a new design method, referred to as Crossflow Attenuated NLF (CATNLF), which shapes airfoils to have pressure distributions that delay transition on wings with high sweep and Reynolds numbers. Additionally, the test aimed to characterize the NTF laminar flow testing capabilities, as well as establish best practices for laminar flow wind tunnel testing. Preliminary results regarding the first goal of validating the new design method are presented in this paper. Experimental data analyzed in this assessment include surface pressure data and transition images. The surface pressure data acquired during the test agree well with computational fluid dynamics (CFD) results. Transition images at a variety of Reynolds numbers and angles of attack are presented and compared to computational transition predictions. The experimental data are used to assess transition due to a turbulent attachment line, as well as crossflow and Tollmien-Schlichting modal instabilities. Preliminary results suggest the CATNLF design method is successful at delaying transition on wings with high sweep. Initial analysis of the transition front images showed transition Reynolds numbers that exceed historic experimental values at similar sweep angles. , section lift

Lynde, Michelle N.↗
NASA NTRS2022

Unsteady Pressure Acquisition on the 1.75% Scale SLS Block 1B Cargo Liftoff Configuration

Time-varying pressure on the core of the Space Launch System (SLS) Block 1B Cargo vehicle was acquired using instrumentation typically used for steady, time averaged pressure. Resono Pressure Systems, Inc. in partnership with the University of Wyoming have developed a unique hardware and software system for measuring unsteady pressure, which was implemented in a large-scale wind tunnel test through a nonreimbursable Space Act Agreement with NASA. The technique was applied to 12 pressure ports on the 1.75% scale SLS wind tunnel model. Tubing lengths of approximately 14 inches separated the pressure transducer module from the pressure taps on the model surface, and surface-mount validation sensors (Kulites) were installed in close proximity to 5 of the 12 pressure ports. Up to about 500 Hz, the frequency response of pressure fluctuations agreed to within 1 dB of amplitude and 10° of phase difference compared to the simultaneously-acquired surface-mount transducers. Sample results using the reconstructed pressure from the full azimuthal ring of taps are presented.

unsteady pressure↗
NASA NTRS1981

Jet V/STOL wind-tunnel simulation and groundplane effects

Low-speed wind-tunnel testing of V/STOL aircraft concepts to determine the aerodynamic-propulsion interaction effects during the transition between hover and wingborne flight is a necessary step in the development cycle of this type of aircraft. Powered models are normally used to determine the aerodynamic performance characteristics. This paper examines some of the pretest preparation necessary to define the objectives of an appropriate investigation. Several factors which influence the selection of the model concept and the engine simulator are discussed. In addition, some of the test techniques important for this class of aircraft model are examined. Finally, the paper reviews some of the wind-tunnel wall effects important to this type of aircraft testing with special emphasis on groundplane effects.

Margason, R. J.↗
NASA NTRS1997

Heat Transfer Measurements on Surfaces with Natural Ice Castings and Modeled Roughness

An experimental method is described to measure and compare the convective heat transfer coefficient of natural and simulated ice accretion roughness and to provide a rational means for determining accretion-related enhanced heat transfer coefficients. The natural ice accretion roughness was a sample casting made from accretions at the NASA Lewis Icing Research Tunnel (IRT). One of these castings was modeled using a Spectral Estimation Technique (SET) to produce three roughness elements patterns that simulate the actual accretion. All four samples were tested in a flat-plate boundary layer at angle of attack in a "dry" wind tunnel test. The convective heat transfer coefficient was measured using infrared thermography. It is shown that, dispite some problems in the current data set, the method does show considerable promise in determining roughness-induced heat transfer coefficients, and that, in addition to the roughness height and spacing in the flow direction, the concentration and spacing of elements in the spanwise direction are important parameters.

Breuer, Kenneth S.↗
NASA NTRS1975

High Reynolds number test of a NACA 651-213, a equals 0.5 airfoil at transonic speeds

Wind-Tunnel tests were conducted in the Lockheed-Georgia Company's compressible flow facility to determine the transonic two-dimensional aerodynamic characteristics of a NACA 65 sub 1-213 a = 0.50 airfoil. The results are correlated with data obtained in the NASA-Langley 8-foot transonic pressure tunnel and the NAE high Reynolds number 15x60-inch two-dimensional test facility. The tests were conducted over a Mach number range from 0.60 to 0.80 and an angle of attack range from -1 deg to 8 deg. Reynolds numbers, based on the airfoil chord, were varied.

Burdges, K. P.↗
NASA NTRS1988

Scale model acoustic testing of counterrotating fans

The UDF contrarotating propfan has been subjected to scale model wind tunnel testing to ascertain both general performance and acoustic characteristics data bases. Model Propulsion Simulator test rigs able to mount contrarotating fan blades of up to 24.5-inch diameter were used, and one of these was installed in a large anechoic test chamber for acoustic measurement of conditions simulating representative takeoffs, power cutbacks, and landing approaches. Attention is presently given to the data acquisition/reduction systems, the scaling criteria used to obtain engine size acoustic data, and comparisons with demonstrator aircraft in-flight acoustic test results.

Janardan, B. A.↗
NASA NTRS2008

Aerothermodynamic Testing and Boundary Layer Trip Sizing of the HIFiRE Flight 1 Vehicle

An experimental wind tunnel test was conducted in the NASA Langley Research Center s 20-Inch Mach 6 Air Tunnel in support of the Hypersonic International Flight Research Experimentation Program. The information in this report is focused on the Flight 1 configuration, the first in a series of flight experiments. This report documents experimental measurements made over a range of Reynolds numbers and angles of attack on several scaled ceramic heat transfer models of the Flight 1 payload. Global heat transfer was measured using phosphor thermography and the resulting images and heat transfer distributions were used to infer the state of the boundary layer on the vehicle windside and leeside surfaces. Boundary layer trips were used to force the boundary layer turbulent, and a brief study was conducted to determine the effectiveness of the trips with various heights. The experimental data highlighted in this test report were used to size and place the boundary layer trip for the flight test vehicle.

Berger, Karen T.↗
NASA NTRS2008

Aerothermodynamic Testing and Boundary Layer Trip Sizing of the HIFiRE Flight 1 Vehicle

An experimental wind tunnel test was conducted in the NASA Langley Research Center s 20-Inch Mach 6 Air Tunnel in support of the Hypersonic International Flight Research Experimentation Program. The information in this article is focused on the Flight 1 configuration, the first in a series of flight experiments. The article documents experimental measurements made over a Reynolds numbers range of 2.1x10(exp 6)/ft to 5.6x10(exp 6)/ft and angles of attack of -5 to +5 deg on several scaled ceramic heat transfer models of the Flight 1 configuration. Global heat transfer was measured using phosphor thermography and the resulting images and heat transfer distributions were used to infer the state of the boundary layer on the vehicle windside and leeside surfaces. Boundary layer trips were used to force the boundary layer turbulent and the experimental data highlighted in this article were used to size and place the boundary layer trip for the flight vehicle. The required height of the flight boundary layer trip was determined to be 0.079 in and the trip was moved from the design location of 7.87 in to 20.47 in to ensure that augmented heating would not impact the laminar side of the vehicle. Allowable roughness was selected to be 3.2x10(exp -3) in.

Berger, Karen T.↗
NASA NTRS1939

Preliminary Model Tests of a Wing-Duct Cooling System for Radial Engines, Special Report

Wind-tunnel tests were conducted on a model wing-nacelle combination to determine the practicability of cooling radial engines by forcing the cooling air into wing-duct entrances located in the propeller slipstream, passing the air through the engine baffles from rear to front, and ejecting the air through an annular slot near the front of the nacelle. The tests, which were of a preliminary nature, were made on a 5-foot-chord wing and a 20-inch-diameter nacelle. A 3-blade, 4-foot-diameter propeller was used. The tests indicated that this method of cooling and cowling radial engines is entirely practicable providing the wing of the prospective airplane is sufficiently thick to accommodate efficient entrance ducts , The drag of the cowlings tested was definitely less than for the conventional N.A.C.A. cowling, and the pressure available at low air speed corresponding to operation on the ground and at low flying speeds was apparently sufficient for cooling most present-day radial engines.

Biermann, David↗