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At least 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

TWINTAN: A program for transonic wall interference assessment in two-dimensional wind tunnels

A method for assessing the wall interference in transonic two dimensional wind tunnel test was developed and implemented in a computer program. The method involves three successive solutions of the transonic small disturbance potential equation to define the wind tunnel flow, the perturbation attriburable to the model, and the equivalent free air flow around the model. Input includes pressure distributions on the model and along the top and bottom tunnel walls which are used as boundary conditions for the wind tunnel flow. The wall induced perturbation fields is determined as the difference between the perturbation in the tunnel flow solution and the perturbation attributable to the model. The methodology used in the program is described and detailed descriptions of the computer program input and output are presented. Input and output for a sample case are given.

Kemp, W. B., Jr.↗
NASA NTRS1992

Progress of magnetic suspension and balance systems for wind tunnels in the USSR

Magnetic Suspension and Balance Systems (MSBS) for wind tunnels are being developed in order to solve the principal problems of aerodynamics which cannot be solved by conventional means: (1) measurements of aerodynamic loads acting on the aircraft models without the effects of mechanical supporting devices; and (2) the study of base pressure. This paper traces the progress of MSBS for wind tunnels in the Commonwealth of Independent States (CIS). The paper describes electromagnetic configuration, position sensing, and control and calibration systems of two wind tunnel MSBS existing in the CIS. The features of high-angle-of-attack control and roll control are discussed. The results of preliminary experiments on high-angle-of-attack and roll controls, digital control, and aerodynamic testing are also presented.

Kuzin, A. V.↗
NASA NTRS2016

Wind Tunnel Model Design for Sonic Boom Studies of Nozzle Jet with Shock Interactions

NASA and Industry are performing vehicle studies of configurations with low sonic boom pressure signatures. The computational analyses of modern configuration designs have matured to the point where there is confidence in the prediction of the pressure signature from the front of the vehicle, but uncertainty in the aft signatures with often greater boundary layer effects and nozzle jet pressures. Wind tunnel testing at significantly lower Reynolds numbers than in flight and without inlet and nozzle jet pressures make it difficult to accurately assess the computational solutions of flight vehicles. A wind tunnel test in the NASA Ames 9- by 7-Foot Supersonic Wind Tunnel from Mach 1.6 to 2.0 will be used to assess the effects of shocks from components passing through nozzle jet plumes on the sonic boom pressure signature and provide datasets for comparison with CFD codes. A large number of high-fidelity numerical simulations of wind tunnel test models with a variety of shock generators that simulate horizontal tails and aft decks have been studied to provide suitable models for sonic boom pressure measurements using a minimally intrusive pressure rail in the wind tunnel. The computational results are presented and the evolution of candidate wind tunnel models is summarized and discussed in this paper.

Wind↗
NASA NTRS1987

Aerodynamic characteristics of the modified 40- by 80-foot wind tunnel as measured in a 1/50th-scale model

The aerodynamic characteristics of the 40- by 80-Foot Wind Tunnel at Ames Research Center were measured by using a 1/50th-scale facility. The model was configured to closely simulate the features of the full-scale facility when it became operational in 1986. The items measured include the aerodynamic effects due to changes in the total-pressure-loss characteristics of the intake and exhaust openings of the air-exchange system, total-pressure distributions in the flow field at locations around the wind tunnel circuit, the locations of the maximum total-pressure contours, and the aerodynamic changes caused by the installation of the acoustic barrier in the southwest corner of the wind tunnel. The model tests reveal the changes in the aerodynamic performance of the 1986 version of the 40- by 80-Foot Wind Tunnel compared with the performance of the 1982 configuration.

Smith, Brian E.↗
NASA NTRS2019

Acoustic Testing of the Tiltrotor Test Rig in the National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel

The Tiltrotor Test Rig (TTR) was tested in the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel from 2017 to 2018. The rotor system can be configured in airplane mode, with the rotor plane perpendicular to the wind flow, and in helicopter mode, with the rotor plane parallel to the wind flow. Four microphones were placed around the TTR: two on the wind tunnel floor and two on struts. The primary goal of the test was to understand the operational capabilities of the TTR, while also acquiring research data as available. Limited measurements of the blade vortex interaction (BVI) noise of the TTR rotor were taken to not only understand the acoustic testing capabilities of the TTR in the NFAC 40- by 80-Foot Wind Tunnel, but to also compare to previous tests and to be used for future validation studies. In particular, data will be compared to measurements of an XV-15 rotor previously acquired in the NFAC 80- by 120-Foot Wind Tunnel.

Tiltrotor Test Rig↗
NASA NTRS1987

Development of wind tunnel techniques for the solution of problems in planetary Aeolian processes

Kutzbach reports wind profiles over a series of roughness elements on a frozen lake and how the wind profile changed as the surface roughness was varied. The approach of the current study was to duplicate Kutzbach's roughness arrays in the wind tunnel at 1/20 and 1/40 scales, and to compare the wind profiles over these scale models to those derived by Kutzbach at full scale in the field. The effects of scale differences and data reduction techniques are discussed. Although the study suggests that wind tunnel scale models can predict parameters measured in the field, the development of more definitive guidelines requires a field experiment designed specifically for comparison with wind tunnel results.

Sullivan, Robert↗
NASA NTRS1983

Cryogenic wind tunnels for high Reynolds number testing

The present investigation has the objective to provide an overview of the development of cryogenic wind tunnels and their application to high Reynolds number testing. The current need for wind tunnels reflects the fact that many complex three-dimensional flows cannot yet be adequately dealt with analytically. Imperfections of wind tunnels are partly related to the fact that for most wind tunnel tests the Reynolds number is much too low. The best solution to this problem appears to be an operation of the tunnel at cryogenic temperatures (arbitrarily defined as temperatures of 150 K or less. In addition to decreasing the temperature, there are three other approaches for increasing the Reynolds number. These approaches include the use of a heavy gas, an increase in the model size, and an increase in pressure. Attention is given to the variable density tunnel concept, problems with high dynamic pressures, low-speed cryogenic tunnel results, transonic cryogenic tunnel tests, and the injection of liquid nitrogen into the tunnel.

Kilgore, R. A.↗
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 NTRS1989

Hot-jet simulation in cryogenic wind tunnels

In order to evaluate hot jet simulation capability in cryogenic wind tunnel testing, simple theoretical calculations were performed. The similarity parameters, isentropic flow properties, and normal shock relations were calculated for a variety of jet simulation techniques. The results were compared with those estimated for a full scale flight condition. It was shown that the cryogenic wind tunnel testing provides an opportunity for the most accurate hot jet simulation technique. By using a compressed nitrogen gas at ambient or moderately elevated temperatures as a jet gas, most all of the relevant similarity parameters including the jet temperature and velocity ratios and the Reynolds numbers, can be set to the full scale flight values. The only exception is the ratio of specific heats for jet flow. In an attempt to match the ratio of specific heats for the turbojet flow, gases other than pure nitrogen were considered. It was found that a nitrogen/methane mixture at moderately elevated temperature behaves like the real combustion gas. Using this mixture as a jet gas, complete simulation of the full scale turbojet exhaust becomes possible in cryogenic wind tunnels.

Asai, Keisuke↗
NASA NTRS1998

Application of Rapid Prototyping Methods to High-Speed Wind Tunnel Testing

This study was undertaken in MSFC's 14-Inch Trisonic Wind Tunnel to determine if rapid prototyping methods could be used in the design and manufacturing of high speed wind tunnel models in direct testing applications, and if these methods would reduce model design/fabrication time and cost while providing models of high enough fidelity to provide adequate aerodynamic data, and of sufficient strength to survive the test environment. Rapid prototyping methods utilized to construct wind tunnel models in a wing-body-tail configuration were: fused deposition method using both ABS plastic and PEEK as building materials, stereolithography using the photopolymer SL-5170, selective laser sintering using glass reinforced nylon, and laminated object manufacturing using plastic reinforced with glass and 'paper'. This study revealed good agreement between the SLA model, the metal model with an FDM-ABS nose, an SLA nose, and the metal model for most operating conditions, while the FDM-ABS data diverged at higher loading conditions. Data from the initial SLS model showed poor agreement due to problems in post-processing, resulting in a different configuration. A second SLS model was tested and showed relatively good agreement. It can be concluded that rapid prototyping models show promise in preliminary aerodynamic development studies at subsonic, transonic, and supersonic speeds.

Springer, A. M.↗
NASA NTRSDate not supplied

Evaluation of CFD as a Surrogate for Wind-Tunnel Testing: Experimental Uncertainty Quantification for the UPWT Flow Survey Test

A series of wind tunnel tests are being performed at the Unitary Plan Wind Tunnel (UPWT)at NASA Langley Research Center to assess the validity of using computational fluid dynamics(CFD) as a surrogate for wind tunnel testing. In order to make proper comparisons, uncertainties in CFD results and experimental data must be well understood. The material presented highlights the methods, assumptions, and inputs used to achieve experimental uncertainty estimates. Results for a small subset of tunnel conditions and variables of interest from the first test in the series, the Flow Survey Test, are highlighted and sample CFD comparisons are shown. The bulk of the results from this work are used for comparisons in other AIAA conference papers related to this test series.

Erin P Hubbard↗
NASA NTRS1984

The 40 X 80 X 120 Wind Tunnel Project

The scope of this project includes increasing the speed of the 40 x 80 x 120 foot closed circuit wind tunnel to 300 knots to expand rotorcraft research capability, and to add an 80 x 120 foot open circuit wind tunnel with 100 knot maximum speed to expand vertical supersonic takeoff and landing research and capability. Topics discussed in viewgraph format are: wind tunnel characteristics, a cost analysis, and design of wind tunnel models.

Falarski, Michael D.↗
NASA NTRS1986

The role of wind tunnel models in helicopter noise research

A study was conducted to determine the applicability of using small-scale powered helicopter models operating in nonanechoic wind tunnels to predict the sound pressure levels of full-scale rotor harmonic noise components. The investigation included noise generation due to high-tip-speed effects, tandem-rotor blade/vortex interactions, single rotors operating on test towers, and the interaction between main rotor vortices and tail rotors. In all cases it was found that the pressure time history waveforms characteristic of different noise-generating mechanisms were properly reproduced by the models. Corrections for microphone locations, acoustical reverberation, and tunnel wind velocity were developed. Application of these corrections to the model data were found to yield satisfactory correlation with full-scale sound pressure levels except for the isolated single rotor, where highly transient data, both model and full-scale, recluded good agreement of absolute values.

Sternfeld, H., Jr.↗
NASA NTRS2001

Abe Silverstein 10- by 10-Foot Supersonic Wind Tunnel Validated for Low-Speed (Subsonic) Operation

The NASA Glenn Research Center and Lockheed Martin Corporation tested an aircraft model in two wind tunnels to compare low-speed (subsonic) flow characteristics. Objectives of the test were to determine and document the similarities and uniqueness of the tunnels and to validate that Glenn's 10- by 10-Foot Supersonic Wind Tunnel (10x10 SWT) is a viable low-speed test facility. Results from two of Glenn's wind tunnels compare very favorably and show that the 10x10 SWT is a viable low-speed wind tunnel. The Subsonic Comparison Test was a joint effort by NASA and Lockheed Martin using the Lockheed Martin's Joint Strike Fighter Concept Demonstration Aircraft model. Although Glenn's 10310 and 836 SWT's have many similarities, they also have unique characteristics. Therefore, test data were collected for multiple model configurations at various vertical locations in the test section, starting at the test section centerline and extending into the ceiling and floor boundary layers.

Hoffman, Thomas R.↗
NASA NTRS1986

Wind tunnel technology for the development of future commercial aircraft

Requirements for new technologies in the area of civil aircraft design are mainly related to the high cost involved in the purchase of modern, fuel saving aircraft. A second important factor is the long term rise in the price of fuel. The demonstration of the benefits of new technologies, as far as these are related to aerodynamics, will,for the foreseeable future, still be based on wind tunnel measurements. Theoretical computation methods are very successfully used in design work, wing optimization, and an estimation of the Reynolds number effect. However, wind tunnel tests are still needed to verify the feasibility of the considered concepts. Along with other costs, the cost for the wind tunnel tests needed for the development of an aircraft is steadily increasing. The present investigation is concerned with the effect of numerical aerodynamics and civil aircraft technology on the development of wind tunnels. Attention is given to the requirements for the wind tunnel, investigative methods, measurement technology, models, and the relation between wind tunnel experiments and theoretical methods.

Szodruch, J.↗
NASA NTRS1986

Real-time simulator for helicopter rotor wind-tunnel operations

This paper describes the elements and operation of a simulator that is being used to train operators of the Rotor Test Apparatus (RTA) in the large-scale 40- by 80-Foot Wind Tunnel at Ames Research Center. The simulator, named TUTOR (for Tunnel Utilization Trainer with Operating Rotor) duplicates the controls of the rotor and its dynamic behavior, as well as the wind-tunnel controls. The simulation software uses a preexisting blade-element model of a four-bladed rotor with flapping and lead-lag degrees of freedom. Equations were developed for all hardware and controls of the RTA and of the wind tunnel that are normally required to perform a wind-tunnel test of a helicopter rotor. The simulator hardware consists of consoles designed to have the same appearance and functions as those in the control room of the 40- by 80-Foot Wind Tunnel, allowing input from three operators who normally establish the required operating conditions during a test run. Normal operating procedures can be practiced, as well as simulated emergencies such as rotor power failure.

Talbot, P. D.↗
NASA NTRS2002

8- by 6-Foot Supersonic Wind Tunnel Compressor Inspected

The NASA Glenn Research Center's 8- by 6-Foot Supersonic Wind Tunnel (8 6 SWT) is NASA's only transonic propulsion wind tunnel. The test section speed range is between Mach 0.25 and 2.0. The 9- by 15-Foot Low-Speed Wind Tunnel (9 15 LWST), which has a speed range from 0 to 175 mph, is housed in the return leg of the 8 6 SWT and uses the same compressor. The 8 6 SWT uses a large, seven-stage axial flow compressor to drive the air through the tunnel. The compressor is 17 ft in diameter and is rated at 1600 m3 (56,600 ft3) of air/sec. It is driven by three electric motors with a combined horsepower of 87,000. A close examination of this compressor was performed in 2001, the first time since February of 1966.

Krupar, Martin J.↗
NASA NTRSDate not supplied

Wind Tunnel Experiment of Microjet-based Flow Control on a Multi-Element High-Lift Airfoil

A wind tunnel model was designed, manufactured, and tested to evaluate the potential of active aerodynamic flow control for multi-element airfoils. The active flow control system involves so-called microjets –small, low momentum, normal to the surface jets located near the trailing edge of the flap –for controlling the lift and improving the aerodynamic performance characteristics of airplane high-lift systems. The configuration selected for this test is the NLR7301 two-element airfoil and it was tested in the Oran W. Nicks Low-Speed Wind Tunnel at Texas A&M University at a chord Reynolds number of 2.0 million and Machnumber of 0.10. This paper presents wind tunnel results for the NLR7301 in the Flap 20 and Flap 30 configurations. The wind tunnel results include surface pressure distributions, lift and pitching moment, and flow visualizations.The results of this experiment validate the effectiveness of the microjet with regards to control of lift in the linear regime for multi-element airfoils.

ARMD↗