Low Turbulence Pressure Tunnel Research Articles

Acoustic Beamforming Array Design Methods Over Irregular Shaped Areas

Abstract Acoustic beamforming array design methods are typically suited for circular and rectangular areas. A comparison of three array design methods is presented in this paper over irregular shaped areas, including L-shapes and arches. Partial-logarithmic spiral arrays that possess their geometric center either at the origin of the array area or the centroid of the irregular shaped area are compared against randomized array designs based on maximum sidelobe level (MSL) parameters and arrays generated using a recently published array design method named the adaptive array reduction method (AARM). In the AARM, a large array is reduced to a smaller array by seeking the removed microphone that possesses the minimum value of the MSL, the main lobe width (MLW), and a lobe distortion term. The AARM is also tested in two practical cases against a partial spiral array design used at the NASA Langley low-turbulence pressure tunnel and a hypothetical rectangular wall case. In both cases, the AARM showed superior performance to the logarithmic spiral arrays in all cases based on MSL and MLW criteria. Of the three methods compared, the AARM best utilizes the full potential array aperture of an irregular area and therefore produces the best MSL, MLW, and lobe distortion values.

Experimental study of surface curvature effects on aerodynamic performance of a low Reynolds number airfoil for use in small wind turbines

This paper presents the wind tunnel experimental results to investigate the effects of surface gradient-of-curvature on aerodynamic performance of a low Reynolds number airfoil Eppler 387 for use in small-scale wind turbines. The prescribed surface curvature distribution blade design method is applied to the airfoil E387 to remove the gradient-of-curvature discontinuities and the redesigned airfoil is denoted as A7. Both airfoils are manufactured with high precision to reflect the design. Low-speed wind tunnel experiments are conducted to both airfoils at chord based Reynolds numbers 100 000, 200 000, and 300 000. Surface pressure measurements are used to calculate the lift and pitching-moment data, and the wake survey method is applied to obtain the drag data. The experimental results of E387 are compared with NASA Low Turbulence Pressure Tunnel (LTPT) results for validation. The gradient-of-curvature discontinuities of E387 result in a larger laminar separation bubble which causes higher drag at lower angles of attack. As the angle of attack increases the separation bubble of the airfoil E387 moves faster towards the leading edge than that of A7, resulting in a premature bubble bursting and earlier stall on E387. The impact of the gradient-of-curvature distribution on the airfoil performance is more profound at higher angles of attack and lower Reynolds number. The aerodynamic improvements are integrated over the 3D geometry of a 3 kW small wind turbine, resulting in up to 10% increase in instantaneous power and 1.6% increase in annual energy production. It is experimentally concluded that an improved curvature distribution results in a better airfoil performance, leading to higher energy output efficiency.

Numerical Analysis of NACA64-418 Airfoil with Blunt Trailing Edge

The aerodynamic performance of blunt trailing edge airfoils was investigated. The flow fields around the modified NACA64-418, which consists of the tip blade of the wind turbine and Mexico model of IEA wind, were analyzed. To imitate the repaired airfoil, the original NACA64-418 airfoil, a cambered airfoil, is modified by the adding thickness method, which is accomplished by adding the thickness symmetrically to both sides of the camber line. The thickness ratio of the blunt trailing edge of the modified airfoil, t TE /t max , is newly defined to analyze the effects of the blunt trailing edge. The shape functions describing the upper and lower surfaces of the modified NACA64-418 with blunt trailing edge are obtained from the curve fitting of the least square method. To verify the accuracy of the present numerical analysis, the results are first compared with the experimental data of NACA64-418 with high Reynolds number, Re=6×10 6 , measured in the Langley low-turbulence pressure tunnel. Then, the aerodynamic performance of the modified NACA64-418 is analyzed. The numerical results show that the drag increases, but the lift increases insignificantly, as the trailing edge of the airfoil is thickened. Re-circulation bubbles also develop and increase gradually in size as the thickness ratio of the trailing edge is increased. These re-circulations result in an increase in the drag of the airfoil. The pressure distributions around the modified NACA64-418 are similar, regardless of the thickness ratio of the blunt trailing edge.

Wind Tunnel Aerodynamic Tests of Six Airfoils for Use on Small Wind Turbines

This paper presents detailed wind tunnel tests data taken on six airfoils having application to small wind turbines. In particular, lift, drag and moment measurements were taken at Reynolds numbers of 100,000, 200,000, 350,000 and 500,000 for both clean and rough conditions. In some cases, data were also taken at a Reynolds number of 150,000. The airfoils included the E387, FX 63-137, S822, S834, SD2030, and SH3055. Prior to carrying out the tests, wind tunnel flow quality measurements were taken to document the low Reynolds number test environment. Oil flow visualization data and performance data taken on the E387 compare favorably with measurements taken at NASA Langley in the Low Turbulence Pressure Tunnel. Highlights of the performance characteristics of the other five airfoils are presented.

Airfoil Validation Using Coupled Navier-Stokes and e Transition Prediction Methods

Navier-Stokes airfoil computations coupled to e* transition prediction are feasible, provided the limiting N factor is known beforehand. In the present study a procedure is outlined to validate free-transition airfoil experiments in wind tunnels for which the limiting N factor is not known a priori. The approach does not rely on mesh adaption procedures to obtain adequate laminar viscous layer data from Navier-Stokes computations for the stability analysis. To the contrary, the laminar viscous layer is computed by a boundary-layer method applying as input the pressure distribution from Navier-Stokes computations on initial meshes. Two measurement campaigns are validated: the NLF(1)-0416 laminar airfoil in the low-speed NASA Langley Low-Turbulence Pressure Tunnel LTPT and the NACA 64*A015 airfoil in the NASA Ames 12-Foot Pressure Tunnel. The free-transition measurements in both tunnels include pressure distributions and transition locations and, in supplement, for the NLF(1)-0416 laminar airfoil lift and drag measurements. The computational results document the validity of the present approach, the existence of a constant limiting N factor for a specific wind tunnel, and an excellent agreement with the experimental findings.

Unsteady Flow Computations of a Slat with a Blunt Trailing Edge

A detailed computational study of a high-lift configuration was conducted to understand the source mechanism behind a dominant acoustic tone observed in recent experiments on slat noise. The unsteady Reynolds-averaged Navier-Stokes computations focused on accurate simulation of the local flowfield of a slat with a blunt trailing edge. At a slat deflection angle of 30 deg relative to the main element, the simulations revealed the presence of strong vortex shedding behind the slat trailing edge. The resulting flow unsteadiness produced large-amplitude acoustic waves propagating away from the trailing-edge region. The local spatial resolution of the computed solution was sufficiently fine to capture both the near-field structure and propagation direction of the generated sound. The calculated shedding frequency is in good agreement with the measured acoustic frequencies obtained at NASA Langley Research Center's Low Turbulence Pressure Tunnel. In contrast, computational results at a slat deflection angle of 20 deg indicated that the shedding process was severely damped, and, therefore, in agreement with the corresponding acoustic measurements during the experiment. There was no evidence of a strong acoustic source near the trailing edge at this lower slat deflection.

Parametric Investigation of a High-Lift Airfoil at High Reynolds Numbers

A new two-dimensional, three-element, advanced high-lift research airfoil has been tested in the NASA Langley Research Center s Low-Turbulence Pressure Tunnel at a chord Reynolds number up to 1.6 x 107. The components of this high-lift airfoil have been designed using a incompressible computational code (INS2D). The design was to provide high maximum-lift values while maintaining attached flow on the single-segment flap at landing conditions. The performance of the new NASA research airfoil is compared to a similar reference high-lift airfoil. On the new high-lift airfoil the effects of Reynolds number on slat and flap rigging have been studied experimentally, as well as the Mach number effects. The performance trend of the high-lift design is comparable to that predicted by INS2D over much of the angle-of-attack range. However, the code did not accurately predict the airfoil performance or the configuration-based trends near maximum lift where the compressibility effect could play a major role.

Freestream Velocity Corrections for Two-Dimensional Testing with Splitter Plates

The velocity correction method for two-dimensional wind-tunnel testing is expanded for the case of testing airfoils between splitter plates. Despite the common nse of splitter plates, the peculiar effects of this testing arrangement on the velocity correction are not well documented. The use of splitter plates in the test section has several advantages, but there are unfortunate consequences (such as spillage, entrainment, and blockage to varying degrees) that require the freestream velocity to be measured directly between the splitter plates with the airfoil in the test section. In this case, the standard velocity correction for wall interference is not sufficient. An additional correction must be made to account for the change in velocity ahead of the airfoil model caused by the circulation about the airfoil. The results reveal that failing to correct for the circulation effect can yield significant errors in the aerodynamic coefficients. The proposed velocity measurement and correction method provides results that are in agreement with data obtained in the NASA Langley Research Center low-turbulence pressure tunnel facility. The method can be readily implemented in the data acquisition and reduction process.

Separation control on high-lift airfoils via micro-vortex generators

An experimental investigation has been conducted to evaluate boundary-layer separation control on a twodimensional single-flap, three-element, high-lift system at near-flight Reynolds numbers with small surfacemounted vortex generators. The wind-tunnel testing was carried out in the NASA Langley Low-Turbulence Pressure Tunnel as part of a cooperative program between McDonnell Douglas Aerospace and NASA Langley Research Center to develop code validation data bases and to improve physical understanding of multielement airfoil flows. This article describes results obtained for small (subboundary-layer) vane-type vortex generators mounted on a multielement airfoil in a landing configuration. Measurements include lift, drag, surface pressure, wake profile, and fluctuating surface heat fluxes. The results reveal that vortex generators as small as 0.18% of reference (slat and flap stowed) wing chord (micro-vortex generators) can effectively reduce boundarylayer separation on the flap for landing configurations. Reduction of flap separation can significantly improve performance of the high-lift system by reducing drag and increasing lift for a given approach angle of attack. At their optimum chordwise placement on the flap, the micro-vortex generators are hidden inside the wing when the flap is retracted, thus extracting no cruise drag penalty.

Effect of underwing frost on a transport aircraft airfoil at flight Reynolds number

The effect of underwing frost on a transport aircraft airfoil in a takeoff configuration was studied. Underwing frost can occur when the lower surface of the wing is cooled by fuel cold-soaked in the wing tanks during cruise. Frost may accrete on the wing lower surface while the aircraft is awaiting takeoff. A two-dimensional test was performed in the NASA Langley Low-Turbulence Pressure Tunnel on a representative high-lift airfoil with a leading-edge slat and trailing-edge flap. Frost was simulated on the lower surface using distributed roughness particles. The test was conducted at M - 0.2 and Re = 5 x 106 to 1.6 x 10 7. The effects of the frost on performance were generally small, with the largest effects occurring for the open-slat case with the frost starting at 12% chord. In this situation, it was found that the frost contaminated the upper surface boundary layer at high angles of attack, increasing drag and reducing maximum lift.

Wind-tunnel wall effects on delta wings

ACOMPUTATIONAL study was performed to assess the effects of wind-tunnel wall interference on the flowfield around delta wings. The free vortex sheet theory was applied to three sizes of a 65-deg delta wing surrounded by representative wind-tunnel walls to determine the extent to which the walls displace the leading-edge vortex core and alter its circulation distribution, hence, altering detailed flowfield properties. A wing size is determined such that these details can be corrected through a simple global adjustment to the angle of attack. Contents A wind-tunnel experiment is under consideration for measuring Reynolds number effects on the subcritical flow around a 65-deg delta wing in the NASA Langley Research Center Low-Turbulence Pressure Tunnel (LTPT) using a three-component laser velocimetry (LV) system. The primary factors affecting the determination of model size are the contrasting requirements for 1) a large model for adequate LV resolution of flow details, with 2) a small wing to minimize wall interference effects. This issue is of particular importance here because the LTPT, primarily an airfoil test facility, has a narrow test section with solid walls, which limits the allowable model span. This study focused on determination of a model size that best meets the test requirements by using an available computational tool to estimate the wall effects on the flowfield. The presence of the walls around a lifting wing induces variations in the magnitude and direction of the flow that are not present around the wing in free air. These effects increase substantially as model size increases. Induced curvature in the flow is manifested in the wing as artificial camber and twist. Because the proposed wing will primarily experience vortex flow, the induced camber and twist will alter the flow nonlinearly. With such effects present, details in the flow are difficult to correct because the principles of superposition are no longer valid. From a practical standpoint, it is desirable to determine the largest wing for which flow curvature effects are sufficiently small so that flow details in the confined flow can be directly related to a free-air flow through a simple adjustment to the global angle of attack. The vortex flowfield was computed for several sizes of the 65-deg delta planform confined within simulated solid testsection walls using the free vortex sheet (FVS) theory of Ref. 1. The FVS code utilizes a potential-flow panel formulation to model steady subsonic, in viscid flow about wings or wingbody configurations with separation-induced leading- and/or

Design and experimental results for a high-altitude, long-endurance airfoil

Currently, there is interest in the development of high-altitude, long-endurance vehicles for a number of missions including communications relaying, weather monitoring, and provision of targeting information for cruise missiles. The preliminary design of such aircraft is complicated, however, by the lack of suitable airfoils. This is due to the fact that such vehicles, unlike those for which the majority of airfoils have been developed in the past, operate at fairly high lift coefficients and at relatively low Reynolds numbers. Thus, to provide realistic airfoil performance information for preliminary design efforts, an airfoil has been designed for an aircraft with missions similar to those noted. The airfoil is unflapped and has a thickness of 15% chord. The design Reynolds number range is 7 x 10 to 2 x 10. Low drag is predicted for lift coefficients ranging from 0.4, which corresponds to a high-speed dash, to 1.5, the maximum endurance condition. Further, the airfoil is designed specifically such that the maximum lift coefficient is unaffected by surface contamination. Consequently, takeoff and landing in rain or with insect residue on the wings should present no special difficulties. The airfoil has been tested in the NASA Langley Low-Turbulence Pressure Tunnel and, with the exception of the maximum lift coefficient prediction, the results generally confirm the theoretical predictions.

Venusian Probes Subsonic Drag Determination from Flight Data

Subsonic drag coefficients have been obtained from flight data for the Pioneer Venus multiprobes. The technique used to extract the information from the data consisted of utilizing in situ pressure and temperature measurements. Analysis of the major model parameter error sources indicates overall error levels of 5% or less in the flight values of CD. Comparisons of the flight coefficients with preflight wind-tunnel test data showed generally good agreement for the five configurations investigated, although all flight-derived values are slightly larger than preflight wind-tunnel test values. Additional wind-tunnel tests were performed on the Sounder descent probe as a control experiment on the data extraction technique. A special attempt was made to accurately duplicate the probe geometry for tests in a high Reynolds number environment in order to achieve as realistic model and flight conditions as practical. Results from this testing in the NASA Langley low-turbulence pressure tunnel produced CD =0.68 at a~0 deg which is within the expected accuracy limits of the flight derived CD value of 0.72 ±0.04.

Experimental Verification of Turbulent Skin Friction Reduction with Compliant Walls

Preliminary compliant wall skin friction test results obtained in a low-turbulence pressure tunnel are reported. Compliant surface skins consisted of 0.0025 cm thick mylar, stretched under tension and area-bonded or longitudinally strip-bonded with silicone rubber adhesive to polyurethane foam. Mean velocity and fluctuating survey data were obtained with a single slanted hot wire. Mean velocity profiles for the area-bonded mylar surface skins indicated up to a 20% reduction in boundary-layer thickness (and lower momentum thickness) over rigid surfaces. This reduction in boundary-layer thickness indicates that a drag reduction occurred. In addition, a 16% reduction in wall shear was evident for the mylar/compressor foam compliant surface.

A low-speed experimental investigation of the effect of a sandpaper type of roughness on boundary-layer transition : Albert E. von Doenhoff and Elmer A. Horton. NACA Rept. 1349, 1958. ii, 16 pp., diagrs., photos

An investigation was made in the Langley low-turbulence pressure tunnel to determine the effect of size and location of a sandpaper type of roughness on the Reynolds number for transition. Transition was observed by means of a hot-wire anemometer located at various chordwise stations for each position of the roughness. These observations indicated that when the roughness is sufficiently submerged in the boundary layer to provide a substantially linear variation of boundary-layer velocity with distance from the surface up to the top of the roughness, turbulent spots begin to appear immediately behind the roughness when the Reynolds number based on the velocity at the top of the roughness height exceeds a value of approximately 600. At Reynolds numbers even slightly below the critical value (value for transition), the sandpaper type of roughness introduced no measurable disturbances into the laminar layer downstream of the roughness. The extent of the roughness area does not appear to have an important effect on the critical value of the roughness Reynolds number.