Wind tunnel tests of a 20 inch diameter 1.15 pressure ratio fan engine model

Abstract

Aerodynamic and acoustic measurements at a typical STOL aircraft takeoff and landing velocity demonstrated that a 1.35 inlet lip area contraction ratio was superior to a 1.26 ratio at high nacelle incidence angles. Reverse thrust, obtained with a . variable pitch rotor, was lower at the landing velocity, and the noise level higher, than at the static"condition. High speed tests showed that, for the design cruise Mach number of 0.75, internal losses and external drag were 27 percent of the ideal fan net thrust, and propulsive efficiency was 1 estimated to be 59 percent for an 85 percent efficient fan stage. For comparison, a similar 1.55 pressure ratio fan system would have a propulsive efficiency of 62 percent.