A procedure is described for the correction of wind-tunnel wall interference effects on the experimental measurement of aerodynamic coef cients. The correction is given by the difference between the values obtained in two different numerical simulations: In the rst one the ow over the model in free-air conditions is simulated, whereas in the second one, the measured pressure values over the wind tunnel walls are used as boundary conditions. A necessary preliminarly step is the choice of the number, location, and accuracy of the pressure measurements. A strategy is proposed to determine these parameters, based on the same correction procedure in which the experimental part is replaced by numerical simulation. This strategy is applied to the subsonic ow around a complete aircraft con guration by means of a potential ow solver. The sensitivity to the number and location of sensors, as well as to the transducer accuracy, is investigated. Given the desired correction accuracy, the proposed strategy permits identi cation of suitable con gurations with reduced time and computational costs. Nomenclature a = slot width, m C L = lift coef cient C m = pitching moment coef cient c p = pressure coef cient E = pressure transducer error, kPa h = test section height, m k = parameter depending on the geometry of the slots for the de nition of slotted wall boundary conditions L = distance between two slot centers, m l = test section length, m M = Mach number N c = number of sensors in the cross direction N l = number of sensors in the longitudinal direction U 1 = freestream velocity, m/s V n = velocity normal to the wind tunnel walls, m/s x = streamwise coordinate with origin at the model rotation point, m x 0 = streamwise coordinate of the in ow tunnel section, m z = vertical coordinate with origin at the model rotation point, m ®= angle of attack " pr = residual error due to wall pressure representation " tot = global error on aerodynamic coef cients ¾ = standard deviation of the Gaussian function used to de ne the longitudinal sensor distribution
DiscussionIn Table 1 the row C M 1/4 = -0.09 corresponds to a light airplane with a NACA-4412 airfoil, while the bottom row Q/,i/4 = -0.18, represents the same airplane with a typical Wortman airfoil. Similar calculations show that at the same airspeed the use of the Wortman airfoil would nearly double the tail download. Equation (8) predicts L t = 0 for h = 0.454 with the NACA-4412 airfoil. However, Eqs. (13) and (14) show that this is unstable since h np = 0.3264, for q t lq = 1.
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