A new facility and technique for two-dimensional aerodynamic testing

Recent experiences with pressure-sensitive paint have shown that very thin paint layers on wind-tunnel models tested at high Reynolds numbers can signi cantly alter the pressure distributions, and thus the forces and moments, on the models. This was observed during two tests of transport-like wings: a "clean" supercritical wing at transonic cruise conditions in the NASA Ames High Reynolds Number Channel 2 and a high-lift wing, complete with slats and aps, at landing conditions in the NASA Ames 12-Foot Pressure Wind Tunnel. The effect of paint on the cruise wing was to displace the shock wave slightly upstream from its no-paint position. Smoothing the paint, and even removing it entirely from the leading edge, decreased this displacement slightly. Applying paint to only the slats of the high-lift wing caused the wing to stall prematurely at the highest Reynolds number. This effect could be eliminated by smoothing the paint. Adding paint to other parts of the model had little effect. Paint intrusiveness was much smaller and more ambiguous at lower Reynolds numbers. In both tests, paint intrusiveness occurred because the model surfaces were rougher when they were painted than when they were not, and this additional roughness altered the development of the turbulent boundary layers. Nomenclaturechord length c mac = mean aerodynamic chord h = paint thickness k = roughness height k s = equivalent sand grain roughness height k C s = k s u ¿ =º equal to k s R u p .C f =2/ M = Mach number P t = total pressure R = Reynolds number, U ¢ c mac =º R u = unit Reynolds number, U=º R x = U ¢ x=º Ra = average roughness t = airfoil thickness U = freestream velocity u = streamwise velocity at height k in boundary layer u e = streamwise velocity at the edge of the boundary layer u ¿ = friction velocity, (¿ w =½/ 1=2 x = streamwise distance from leading edge y = spanwise distance or distance perpendicular to surface ® = angle of attack, deg º = kinematic viscosity ½ = density ¿ w = skin friction

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“…38). This is a transonic, blowdown wind tunnel capable of operating at total pressures of up to about 6.9 bar.…”

Section: Apparatus Cruise Wingmentioning

confidence: 99%

Effects of Pressure-Sensitive Paint on Experimentally Measured Wing Forces and Pressures

2002

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“…2 and described in detail in Ref. 6. This blowdown facility has a solid wall test section, with a 0.406x0.609m (1.6c r x2.4c r ) nominal test section area.…”

Section: Experimental Facilitymentioning

confidence: 99%

Low Aspect Ratio Wing Code Validation Experiment

Olsen

Seegmiller

1993

AIAA Journal

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A code validation experiment for transonic flow is described. The experimental geometry is a low aspect ratio wing, tested in a solid wall wind tunnel. Inflow conditions were measured far enough upstream of the model to allow simple specification. Experimental uncertainties are given for pressure, temperature, velocity, and position measurements, as well as bounds on model deflections and transition location. Cursory comparison with simple computational models provides a baseline of flow conditions that should be readily modeled as well as denotes those cases that will be more challenging. The data include a wide range of a, Moo, and two Re Cr . The Reynolds numbers reached in this experiment reproduce those experienced by an aircraft with a 5-m root chord, flying at 14-16 km, from Mach 0.6 to 0.8. The angle-of-attack range of the experiment extends into the regime of leading-edge separation (0 deg<«<8 deg). Nomenclaturea = local speed of sound a* = sound speed at sonic conditions b/2 = semispan of wing C p = coefficient of pressure, (p -p^)/qĉ = local wing chord (a function of r/) c r = wing root chord c t = wing tip chord MO, = freestream Mach number M£ m = lowest Mo, at which sonic flow is attained at some point in the flowfield p t-total (pitot) pressure measured at £ 0 = -2.25, 77 = 1.6, and f=0 Poo = static pressure measured at £ 0 = -2.525, 77 = 1.6, and f=0 #00 = freestream dynamic pressure ( l /2pu£) Re Cr = Reynolds number based on root chord (U^Cr/v) T t -freestream total temperature, measured in stagnation chamber u = axial velocity v = spanwise velocity w = vertical velocity x = axial position, origin at wing root leading edge (see Fig. 2) y = spanwise position, origin at wing root leading edge (see Fig. 2) z = vertical position (normal to xy plane), origin at wing root leading edge (see Fig. 2) f = vertical distance normalized by c r , z/c r 71 = spanwise distance, normalized by b/2, 2y/b v = kinematic viscosity £ = axial distance from leading edge, normalized by local wing chord, [x-x fe (i;)]/c(i/) £o = axial distance from root leading edge, normalized by root chord, [x -Xi e (Q)]/c r p = mass density Presented as Paper 92-0402 at * Research Scientist, Modelling and Validation Branch. Senior Member AIAA. tResearch Scientist, Modelling and Validation Branch. Subscripts and Superscripts c/4 = wing quarter chord = wing leading edge = total (stagnation) conditions upstream of wing = wing trailing edge = static conditions at £o--2.525 = sonic conditions le t te oo

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