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At least 397 records · Page 22
All sourcesNASA NTRSDOE OSTI
NASA NTRS2019

Reduction of Wind Tunnel Contamination During Flow Visualization Experiments Using Polystyrene Microspheres

Evaluation of novel methods and materials for seeding tracer particles for particle image velocimetry (PIV) was carried out in the Basic Aerodynamic Research Tunnel (BART) at NASA’s Langley Research Center (LaRC). Seeding of polystyrene latex microspheres (PSLs) from ethanol/water suspensions and from the dry state was carried out using custom built seeders. PIV data generated using the novel methods were found to be in general agreement with data collected using the current seeding methods. Techniques for assessing PSL fouling of wind tunnel surfaces were identified and refined. Initial results suggest that dry seeding PSLs may allow comparable data quality to wet seeding while reducing wind tunnel screen fouling. Results also indicate that further developments to the dry seeding system should focus on increasing single particle flux into the wind tunnel. Modifications to PSLs and seeding equipment to achieve this have been identified and are discussed.

Schmitz, Russell C.↗
NASA NTRS2011

Comparison of Resource Requirements for a Wind Tunnel Test Designed with Conventional vs. Modern Design of Experiments Methods

The factors that determine data volume requirements in a typical wind tunnel test are identified. It is suggested that productivity in wind tunnel testing can be enhanced by managing the inference error risk associated with evaluating residuals in a response surface modeling experiment. The relationship between minimum data volume requirements and the factors upon which they depend is described and certain simplifications to this relationship are realized when specific model adequacy criteria are adopted. The question of response model residual evaluation is treated and certain practical aspects of response surface modeling are considered, including inference subspace truncation. A wind tunnel test plan developed by using the Modern Design of Experiments illustrates the advantages of an early estimate of data volume requirements. Comparisons are made with a representative One Factor At a Time (OFAT) wind tunnel test matrix developed to evaluate a surface to air missile.

DeLoach, Richard↗
NASA NTRSDate not supplied

An Efficient Approach to Collect Inlet Characterization Data

This report considers a novel process to acquire performance data for supersonic mixed compression inlets that is faster, and therefore less costly, than the previous one. In the previous process, measurements were recorded in discrete increments. The backpressuring actuator was advanced, there was a pause to allow for airflow fluctuations to diminish, and then the measurements were recorded. While this process was methodical, reliable and accurate, it consumed significant wind-on time. In general, high-speed wind tunnels are expensive to operate considering maintenance, energy and human resource requirements. Furthermore, many facilities, such as blowdown wind tunnels, are limited in the duration that they can maintain the desired test condition. Consequently, a quick method to collect the measurements is desired. A favored approach has the actuator moving in a continuous fashion with measurements recorded as the actuator continuously progresses through the increment points. This approach, identified as the dynamic inlet characteristic data acquisition procedure, uses in situ dynamic high-speed pressure sensors to measure pressure signals. While faster, continuous movement adds dynamic effects to the data. To better understand these dynamic effects and other potential influences, a study was undertaken to compare wind tunnel measurements taken at discrete actuator position points with data gathered during continuous actuator movement. Specially, plots of inlet pressure recovery versus mass capture ratio, or total pressure characteristic curves, were examined. The data were measured during experiments in the NASA Glenn Research Center Abe Silverstein 10- by 10-Foot Supersonic Wind Tunnel (SWT) with an inlet propulsion test article. In this report, the wind tunnel experiment and the study objectives are described. The processes to reduce the data and for analysis are explained. A discussion of the results along with features noted in the data is given. Finally, conclusions from this study that may be used to guide future wind tunnel experiment planning are presented.

Data Acquisition↗
NASA NTRSDate not supplied

An Efficient Approach to Collect Inlet Characterization Data

This report considers a novel process to acquire performance data for supersonic mixed compression inlets that is faster, and therefore less costly, than the previous one. In the previous process, measurements were recorded in discrete increments. The backpressuring actuator was advanced, there was a pause to allow for airflow fluctuations to diminish, and then the measurements were recorded. While this process was methodical, reliable and accurate, it consumed significant wind-on time. In general, high-speed wind tunnels are expensive to operate considering maintenance, energy and human resource requirements. Furthermore, many facilities, such as blowdown wind tunnels, are limited in the duration that they can maintain the desired test condition. Consequently, a quick method to collect the measurements is desired. A favored approach has the actuator moving in a continuous fashion with measurements recorded as the actuator continuously progresses through the increment points. This approach, identified as the dynamic inlet characteristic data acquisition procedure, uses in situ dynamic high-speed pressure sensors to measure pressure signals. While faster, continuous movement adds dynamic effects to the data. To better understand these dynamic effects and other potential influences, a study was undertaken to compare wind tunnel measurements taken at discrete actuator position points with data gathered during continuous actuator movement. Specially, plots of inlet pressure recovery versus mass capture ratio, or total pressure characteristic curves, were examined. The data were measured during experiments in the NASA Glenn Research Center Abe Silverstein 10- by 10-Foot Supersonic Wind Tunnel (SWT) with an inlet propulsion test article. In this report, the wind tunnel experiment and the study objectives are described. The processes to reduce the data and for analysis are explained. A discussion of the results along with features noted in the data is given. Finally, conclusions from this study that may be used to guide future wind tunnel experiment planning are presented.

Inlet↗
NASA NTRS1972

The high Reynolds number cryogenic wind tunnel.

Theoretical considerations indicate that cooling the wind-tunnel test gas to cryogenic temperatures will provide a large increase in test Reynolds number with no increase in dynamic pressure while reducing the tunnel drive-power requirements. Studies have been made to determine the expected variations of Reynolds number and other parameters over wide ranges of Mach number, pressure, and temperature with due regard to avoiding liquefaction and adverse real-gas effects. Practical operational procedures have been developed in a low-speed prototype cryogenic wind tunnel. Aerodynamic experiments in the facility have demonstrated the theoretically predicted variations in Reynolds number and drive power. Force and moment measurements on a wing model mounted on a water-jacketed strain-gage sting balance have demonstrated the feasibility of operation of such balances in a cryogenic environment.

Goodyer, M. J.↗
NASA NTRS1978

Transonic flow past an airfoil with condensation

In connection with investigations conducted to determine the influence of water vapor on experiments in wind tunnels, the question arose as to what changes due to vapor condensation might be expected in airfoil measurements. Density measurements on circular-arc airfoils aided by an interferometer in choked tunnels with parallel walls show that increasing humidity produces increasing changes in the flow field. The flow becomes nonstationary at high humidity. At the airfoil, however, the influence of the condensation is only felt, inasmuch as the shock bounding the local supersonic region moves upstream with increasing humidity while its intensity decreases. The density distribution upstream of the shock remains unchanged. Even if the flow becomes nonstationary in the vicinity of the airfoil, no changes occur at the airfoil.

Schmidt, B.↗
NASA NTRS1978

Sonic boom research

It is demonstrated that a supersonic airplane configuration weighing over half a million pounds while creating a maximum sonic boom of less than 1 p.s.f. can be designed. New experimental techniques are developed in the wind tunnel and experiments for the sonic boom measurements were carried out. Theoretical analyses were performed for the effects of sonic boom on structures and pollution problems associated with supersonic flights were investigated. Numerical programs were generated for the sonic boom propagations from the near field of an airplane in supersonic flight at high altitude to the ground, taking into account the nonlinear effects and the asymmetric effects due to lift and the spacewise distributions of lift and volume.

Ting, L.↗
NASA NTRS1987

Evaluation of influence coefficient matrices for an adaptive-wall wind tunnel and applications to one-step convergence scheme

Within the framework of ventilated adaptive-wall wind tunnel testing, experiments were conducted to evaluate the influence coefficients from the side-wall pressure distribution, the streamwise velocity distribution, and the vertical velocity distribution on two control surfaces. Consideration is given to the effects of Mach number, superposition, suction, blowing, and the presence of a nonlifting model. It is shown that a single-step convergence scheme reduces the testing time by eliminating the intermediate experimental iterations; however, its success depends on the accuracy of the influence coefficients.

Celik, Z. Z.↗
NASA NTRS1988

Analysis of wind-tunnel boundary-layer transition experiments on axisymmetric bodies at transonic speeds using compressible boundary-layer stability theory

The design of laminar flow fuselages and advanced swept wings at high-subsonic compressible speeds can benefit from a correlation of compressible transition experiments with the e exp n transition-prediction method. A computational analysis has been conducted to investigate the detailed transition measurements obtained by Boltz et al. (1956, 1960) for two bodies-of-revolution. Nonadiabatic wall conditions were included in the analysis when wall-temperature measurements were available.

Vijgen, P. M. H. W.↗
NASA NTRS1992

Planar Rayleigh scattering results in helium-air mixing experiments in a Mach-6 wind tunnel

Planar Rayleigh scattering measurements with an argon-fluoride excimer laser are performed to investigate helium mixing into air at supersonic speeds. The capability of the Rayleigh scattering technique for flow visualization of a turbulent environment is demonstrated in a large-scale, Mach-6 facility. The detection limit obtained with the present setup indicates that planar, quantitative measurements of density can be made over a large cross-sectional area (5 cm x 10 cm) of the flow field in the absence of clusters.

Shirinzadeh, B.↗
NASA NTRS1993

Transitional flight characteristics of a geometrically simplified STOVL model

The transitional flight characteristics of a geometrically simplified Short Take-Off Vertical Landing (STOVL) aircraft configuration were measured in the NASA Ames 7- by 10-Foot Wind Tunnel. The experiment was designed to provide detailed data for evaluating the capability of computational fluid dynamics methods to predict the important powered lift flow parameters. The model consists of a 60 deg cropped delta wing planform; a blended fuselage; and tandem, circular, high pressure air jets that exit perpendicular to the flat lower surface. Freestream Mach number is limited to a maximum 0.2. Model angle of attack ranges from -10 deg to +20 deg. The nozzle pressure ratios of both jets are varied between 1 and 3, and the jet exit temperatures are maintained at near ambient conditions. Detailed surface pressure measurements show that suction pressure peaks located on the upper surface of the wing during conventional wingborne flight for angles of attack greater than 5 deg move to the wing lower surface at angles of attack less than 0 deg. A reduction in these suction pressure peaks is observed when the lift jets are operating. With sonic jet exit conditions, a freestream Mach number of 0.14, and 0 deg angle of attack, the jet-induced suckdown is equivalent to a 3.7 deg reduction in angle of attack. Schlieren, laser light sheet flow visualization and total pressure measurements in the jet plumes provide a description of the three-dimensional jet efflux flowfield.

Roth, Karlin R.↗
NASA NTRS1992

Desirable features of an infrared imaging system for aerodynamic research

Advantage of non-intrusiveness, capability for field measurement, and increased availability of IR imaging systems have resulted in their wider use for aerodynamic research. However, certain difficulties persist while using currently available systems for such applications. A critical evaluation of the IR imaging systems is presented on the basis of the state-of-the-art of IR imaging technology and experiences in wind tunnel and flight testing at NASA's Langley Research Center. The requirements for using IR thermography as a measurement tool in aerodynamic research are examined in terms of range, sensitivity, and accuracy of temperature measurement, temporal and spatial resolution, and features of target. Deficiencies of present IR imaging systems are identified, and user precautions to avoid such problems by proper selection and operation of these units are suggested. Different aspects of imager performance such as imager optics, video capabilities, and environmental tolerance are discussed. Electronic data recording and image processing hardware and software requirements are evaluated. Slit response tests and spatial resolution are discussed with the objective of obtaining reliable, accurate, and meaningful information from IR thermography measurements for aerodynamic studies.

Wright, Robert E., Jr.↗
NASA NTRS1994

STOVL CFD model test case

The transitional flight characteristics of a geometrically simplified Short Take-Off Vertical Landing (STOVL) aircraft configuration have been measured in the NASA Ames 7- by 10-Foot Wind Tunnel. The experiment is the first in a sequence of tests designed to provide detailed data for evaluating the capability of computational fluid dynamics methods to predict the important flow parameters for powered lift. The model consists of a 60 deg cropped delta wing platform, blended fuselage and two circular in-line jets that exit perpendicularly from the flat lower surface. The measured flows have a maximum freestream Mach number of 0.2. Model angle of attack is varied between -10 and +20 deg. The flow is ambient temperature in both jet exits and the nozzle pressure ratios are varied between 1 and 3. The data presented includes forces and moments, pressures measured at 281 surface pressure ports and the pressures of the jets. Measurements of the flow are also made in the tunnel test section upstream and downstream of the model and at the jet exits to guide boundary condition selection for the planned computations. Flow visualization and total pressure measurements in the jet plumes provide a description of the three-dimensional jet efflux flowfield.

Roth, Karlin R.↗
NASA NTRS1995

Interactive flutter analysis and parametric study for conceptual wing design

An interactive computer program was developed for wing flutter analysis in the conceptual design stage. The objective was to estimate the flutter instability boundary of a flexible cantilever wing, when well defined structural and aerodynamic data are not available, and then study the effect of change in Mach number, dynamic pressure, torsional frequency, sweep, mass ratio, aspect ratio, taper ratio, center of gravity, and pitch inertia, to guide the development of the concept. The software was developed on MathCad (trademark) platform for Macintosh, with integrated documentation, graphics, database and symbolic mathematics. The analysis method was based on nondimensional parametric plots of two primary flutter parameters, namely Regier number and Flutter number, with normalization factors based on torsional stiffness, sweep, mass ratio, aspect ratio, center of gravity location and pitch inertia radius of gyration. The plots were compiled in a Vaught Corporation report from a vast database of past experiments and wind tunnel tests. The computer program was utilized for flutter analysis of the outer wing of a Blended Wing Body concept, proposed by McDonnell Douglas Corporation. Using a set of assumed data, preliminary flutter boundary and flutter dynamic pressure variation with altitude, Mach number and torsional stiffness were determined.

Mukhopadhyay, Vivek↗
NASA NTRS1995

Implementation of Combined Feather and Surface-Normal Ice Growth Models in LEWICE/X

Experimental observations have shown that discrete rime ice growths called feathers, which grow in approximately the direction of water droplet impingement, play an important role in the growth of ice on accreting surfaces for some thermodynamic conditions. An improved physical model of ice accretion has been implemented in the LEWICE 2D panel-based ice accretion code maintained by the NASA Lewis Research Center. The LEWICE/X model of ice accretion explicitly simulates regions of feather growth within the framework of the LEWICE model. Water droplets impinging on an accreting surface are withheld from the normal LEWICE mass/energy balance and handled in a separate routine; ice growth resulting from these droplets is performed with enhanced convective heat transfer approximately along droplet impingement directions. An independent underlying ice shape is grown along surface normals using the unmodified LEWICE method. The resulting dual-surface ice shape models roughness-induced feather growth observed in icing wind tunnel tests. Experiments indicate that the exact direction of feather growth is dependent on external conditions. Data is presented to support a linear variation of growth direction with temperature and cloud water content. Test runs of LEWICE/X indicate that the sizes of surface regions containing feathers are influenced by initial roughness element height. This suggests that a previous argument that feather region size is determined by boundary layer transition may be incorrect. Simulation results for two typical test cases give improved shape agreement over unmodified LEWICE.

Matthew T Velazquez↗
NASA NTRS1994

Microphone Corrections for Accurate In-Flow Acoustic Measurements at High Frequency

Acoustic analysis of aircraft components through small-scale experiments in wind tunnels requires in-flow acoustic measurements at high frequencies. This high frequency information is shifted to lower frequencies during the geometric scaling process. As a result, models of scale one tenth and smaller require frequency information of up to at least 100 kHz to resolve the midpoint of the full-scale equivalent audible range. Farfield acoustic measurements of frequencies higher than 100 kHz are prohibitively difficult to make due to the physical limitations of atmospheric propagation over large distances and the delicacy of the required instruments. Even measurements up to 100 kHz require special corrections and calibration procedures to maintain sufficient accuracy. In-flow microphone measurements at high frequencies require additional consideration due to the presence of the aerodynamic microphone forebody which protects the microphone diaphragm from flow impingement. The screen located at the forebody surface attenuates the high frequency acoustic waves and must be accounted for. The purpose of this paper is to present procedures for calibrating and correcting high frequency acoustic measurements to retain high accuracy. Specifically, this paper gives calibration procedures and examples of typical data for calibrating or correcting for five physical effects. These effects include the frequency response of the entire acoustic data acquisition system, intrusion of the microphone into the acoustic field, forebody screen attenuation, directionality of microphone forebody response and atmospheric absorption.

Allen, Christopher S.↗