Experimental Determination of the Pressure Distribution on a Rectangular Wing Oscillating in the First Bending Mode for Mach Numbers from 0.24 to 1.30

摘要

The results of an experimental investigation in a wind tunnel to obtain the aerodynamic pressure distribution on an unswept rectangular wing oscillating in its first symmetrical bending mode are presented. The wing was of aspect ratio 3 with 5-percent-thick biconvex airfoil sections. Data were obtained at 0 deg and 5 deg angle of attack in the Mach number range from 0.24 to 1.30 at Reynolds numbers, depending on the mach number ranging from 2.2 to 4.6 million per foot. The reduced frequencies also a function of Mach number, ranged from 0.46 at M = 0.24 to 0.10 at M = 1.30. The most important results presented are the chordwise distributions of the surface pressures generated by the bending oscillations. Similar data obtained under static conditions are also presented. The results show that the phenomena causing irregularities in the static pressure such as three-dimensional tip effects, local shock waves, and separation will also produce significant changes in the oscillatory pressures. The experimental data are also compared with the oscillatory pressure distributions computed by means of the most complete linearized theories available. The comparison shows that subsonic linearized theory is adequate for predicting the pressures and associated phase angles at low subsonic speeds and low angles of attack for this wing. However, the appearance of local shock waves and flow separation as the Mach number and angle of attack are increased causes significant changes in the experimental data from that predicted by the theory. At the low supersonic speeds covered in the experimental investigation, linearized theory is completely inadequate, principally because of the detached bow wave caused by the wing thickness. Some indication of wind-tunnel resonance was noted; however, the effects on the experimental data appear to be confined to the M = 0.70 results.