I?.? FLIGHT AT MACH NUMBERS UP TO 1.11, AND COMPARISON WITH THEORY

Flight tests were conducted to determine the pressure distribution over the wing of a swept-wing jet-propelled airplane over the flight range of lift coefficients for Mach numbers up to l.ll. At a constant normal-force coefficient the principal effect of increasing Mach number on the chordwise distribution of pressure was the reduction of the peak pressures near the leading edge. At subsonic speeds the wing was subject to stalling, which progressed inboard from the tip. The lift coefficient at which stall first occurred decreased with increasing Mach number to 0.3 at 0.9 Mach number. At supersonic speeds the lift effectiveness of the outer portion of the wing was very low. Tuft studies showed the flow to be separated over the trailing edge of this portion of the ting. The spanwise distribution of additional loading at subsonic speeds was compared with that calculated by the Weissinger method. Up to the stall the agreement with theory was fair. The Weissinger method could be used to obtain a satisfactory approximation to the spanwise loading for structural design purposes. After the tip stalled the loading shifted inboard, departing well from the theoretical loading. At supersonic speeds the experimentalsp.anwise distribution of loading was compared with that calculated from linearized supersonic theory. The agreement was not good because of the excessively low loading on the tip portion of the wing. In this case, the span load distribution calculated-simply by the Weissinger method for Mach number of zero could be used for struc-tural design purposes throughout the entire speed Large trailing-edge loads, both up and down, particularly when the flow was separated. range. were encountere