Comparison of NAE porous wall and NASA adaptive wall test results using the NAE CAST-10 airfoil model

Wind tunnels can now simulate flows over airfoils at high Reynolds numbers and high subsonic speeds. Methods to correct for (or reduce) test section wall interference at these test conditions must be validated. The National Aeronautics Establishment (NAE) of Canada and NASA have a cooperative agreement to study this area. The NAE designed, built, and tested a CAST-10 airfoil model in its conventional Two-Dimensional High Reynolds Number Facility. The results were corrected using classical correction techniques. NASA then tested the same model in its 0.3-meter Transonic Cryogenic Tunnel with the adaptive wall test section. The adaptive wall test section reduced the wall interference to what was expected to be an acceptable level. The corrected NAE results are compared with the uncorrected NASA results. The NAE results are also compared with NASA results after residual corrections for top and bottom wall interference. Finally, a comparison of both sets of results corrected for interference from all four walls is presented

Comparison of airfoil results from an adaptive wall test section and a porous wall test section

Two wind tunnel investigations were conducted to assess two different wall interference alleviation/correction techniques: adaptive test section walls and classical analytical corrections. The same airfoil model has been tested in the adaptive wall test section of the NASA-Langley 0.3 m Transonic Cryogenic Tunnel (TCT) and in the National Aeronautical Establishment (NAE) High Reynolds Number 2-D facility. The model has a 9 in. chord and a CAST 10-2/DOA 2 airfoil section. The 0.3 m TCT adaptive wall test section has four solid walls with flexible top and bottom walls. The NAE test section has porous top and bottom walls and solid side walls. The aerodynamic results corrected for top and bottom wall interference at Mach numbers from 0.3 to 0.8 at a Reynolds number of 10 by 1,000,000. Movement of the adaptive walls was used to alleviate the top and bottom wall interference in the test results from the NASA tunnel

Calibration of the 13- by 13-inch adaptive wall test section for the Langley 0.3-meter transonic cryogenic tunnel

A 13 by 13 inch adaptive wall test section was installed in the 0.3 Meter Transonic Cryogenic Tunnel circuit. This new test section is configured for 2-D airfoil testing. It has four solid walls. The top and bottom walls are flexible and movable whereas the sidewalls are rigid and fixed. The wall adaptation strategy employed requires the test section wall shapes associated with uniform test section Mach number distributions. Calibration tests with the test section empty were conducted with the top and bottom walls linearly diverged to approach a uniform Mach number distribution. Pressure distributions were measured in the contraction cone, the test section, and the high speed diffuser at Mach numbers from 0.20 to 0.95 and Reynolds numbers from 10 to 100 x 10 (exp 6)/per foot

Study of potential aerodynamic benefits from spanwise blowing at wingtip

Comprehensive experimental and analytical studies have been conducted to assess the potential aerodynamic benefits from spanwise blowing at the tip of a moderate-aspect-ratio swept wing. Previous studies on low-aspect-ratio wings indicated that blowing from the wingtip can diffuse the tip vortex and displace it outward. The diffused and displaced vortex will induce a smaller downwash at the wing, and consequently the wing will have increased lift and decreased induced drag at a given angle of attack. Results from the present investigation indicated that blowing from jets with a short chord had little effect on lift or drag, but blowing from jets with a longer chord increased lift near the tip and reduced drag at low Mach numbers. A Navier-Stokes solver with modified boundary conditions at the tip was used to extrapolate the results to a Mach number of 0.72. Calculations indicated that lift and drag increase with increasing jet momentum coefficient. Because the momentum of the jet is typically greater than the reduction in the wing drag and the increase in the wing lift due to spanwise blowing is small, spanwise blowing at the wingtip does not appear to be a practical means of improving the aerodynamic efficiency of moderate-aspectratio swept wings at high subsonic Mach numbers

Wall interference tests of a CAST 10-2/DOA 2 airfoil in an adaptive-wall test section

A wind-tunnel investigation of a CAST 10-2/DOA 2 airfoil model has been conducted in the adaptive-wall test section of the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT) and in the National Aeronautical Establishment High Reynolds Number Two-Dimensional Test Facility. The primary goal of the tests was to assess two different wall-interference correction techniques: adaptive test-section walls and classical analytical corrections. Tests were conducted over a Mach number range from 0.3 to 0.8 and over a chord Reynolds number range from 6 million to 70 million. The airfoil aerodynamic characteristics from the tests in the 0.3-m TCT have been corrected for wall interference by the movement of the adaptive walls. No additional corrections for any residual interference have been applied to the data, to allow comparison with the classically corrected data from the same model in the conventional National Aeronautical Establishment facility. The data are presented graphically in this report as integrated force-and-moment coefficients and chordwise pressure distributions

High Reynolds Number tests of the NASA SC(2)-0012 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

A wind-tunnel investigation of the NASA SC(2)-0012 airfoil has been conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel. This investigation supplements the two-dimensional airfoil studies of the Advanced Technology Airfoil Test Program. The Mach number was varied from 0.60 to 0.84. The stagnation temperature and pressure were varied to provide a Reynolds number range from 6 to 40 x 10 to the 6th power based on a 6.0-in. (15.24-cm) airfoil chord. No corrections for wind-tunnel wall interference have been made to the data. The aerodynamic results are presented as integrated force and moment coefficients and pressure distributions without any analysis

Hardware and operating features of the adaptive wall test section for the 0.3-meter transonic cryogenic tunnel

A 13- by 13-inch adaptive wall test section was installed in the Langley 0.3-Meter Transonic Cryogenic Tunnel circuit. This test section has four solid walls and is configured for two-dimensional airfoil testing. The top and bottom walls are flexible and movable, whereas the sidwalls are rigid and fixed. The test section has a turntable to support airfoil models, a survey mechanism to probe the model wake, and provisions for a sidewall boundary-layer-control system. Details of the adaptive wall test section, the tunnel circuit modifications, the supporting instrumentation, the monitoring and control hardware, and the wall adaptation strategy are discussed. Sample results of shakedown tests with the test section empty and with an airfoil installed are also included

Flow visualization studies of blowing from the tip of a swept wing

Flow visualization studies of blowing from the tip of a swept wing were conducted in the Langley 16- by 24-inch water tunnel. Four wing tips, each with two independent blowing slots, were tested. The two slots were located one behind the other in the chordwise direction. The wing tips were designed to vary systematically the jet length, the jet in-plane exhaust direction (sweep), and the jet out-of-plane exhaust direction (anhedral). Each blowing slot was tested separately at two angles of attack and at four ratios of jet to free stream velocity ratios. Limited tests were conducted with blowing from both slots simultaneously. Blowing from the tip inhibited inboard spanwise flow on the upper wing surface near the tip. The jet path moved farther away from the tip with increasing jet to free stream velocity ratio and moved closer to the tip with increasing angle of attack

On the Importance of Spatial Resolution for Flap Side Edge Noise Prediction

A spatial resolution study of flap tip flow and the effects on the farfield noise signature for an 18%-scale, semispan Gulfstream aircraft model are presented. The NASA FUN3D unstructured, compressible Navier-Stokes solver was used to perform the highly resolved, time-dependent, detached eddy simulations of the flow field associated with the flap for this high-fidelity aircraft model. Following our previous work on the same model, the latest computations were undertaken to determine the causes of deficiencies observed in our earlier predictions of the steady and unsteady surface pressures and off-surface flow field at the flap tip regions, in particular the outboard tip area, where the presence of a cavity at the side-edge produces very complex flow features and interactions. The present results show gradual improvement in steady loading at the outboard flap edge region with increasing spatial resolution, yielding more accurate fluctuating surface pressures, off-surface flow field, and farfield noise with improved high-frequency content when compared with wind tunnel measurements. The spatial resolution trends observed in the present study demonstrate that the deficiencies reported in our previous computations are mostly caused by inadequate spatial resolution and are not related to the turbulence model

Application of two procedures for dual-point design of transonic airfoils

Two dual-point design procedures were developed to reduce the objective function of a baseline airfoil at two design points. The first procedure to develop a redesigned airfoil used a weighted average of the shapes of two intermediate airfoils redesigned at each of the two design points. The second procedure used a weighted average of two pressure distributions obtained from an intermediate airfoil redesigned at each of the two design points. Each procedure was used to design a new airfoil with reduced wave drag at the cruise condition without increasing the wave drag or pitching moment at the climb condition. Two cycles of the airfoil shape-averaging procedure successfully designed a new airfoil that reduced the objective function and satisfied the constraints. One cycle of the target (desired) pressure-averaging procedure was used to design two new airfoils that reduced the objective function and came close to satisfying the constraints