Langley 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.

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.

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.

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.

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.

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.