Aerospace & Aviation

Spin and recovery characteristics of north american x-15 aircraft

Static stability and control characteristics of aircraft model with tail surfaces outboard of wing tips

Thrust modulation techniques - cold flow tests of three variable-throat-area convergent-divergent nozzles

Hub-tip radius ratio transonic turbine for secondary flow - stator configurations designed to eliminate blade wakes and secondary-flow effects

Full scale turbojet engine testing of fuels - pentaborane-jp-4 mixtures and trimethylborate azeotrope performance

Low-temperature chemical starting of combined jp-4 nitric acid propellant for low-thrust rocket engine using three-fluid propellant valve

Proportional navigation guidance systems - radar antenna positioning effects on speed of response

Buffeting tests of attack aircraft model at transonic speeds

Combined compressor coolant injection and afterburning for turbojet engine thrust augmentation

Supersonic performance of variable two-dimensional inlets designed for engine-inlet matching

Off-design performance of divergent ejectors

Effect of pressure level on afterburner wall temperature

Rotating-stall limits in supersonic turbofan engine

Limitations on turbojet engine design by turbines with downstream stators

The flow about slender flat-top wing-body configurations traveling at high supersonic speeds and small angles of attack is investigated analytically. In the case of conical configurations, approximate algebraic solutions to the flow field are obtained. In the case of configurations which are conical at the vertex but curved in the stream direction, these solutions are combined with a slender-body approximation to the generalized shock-expansion method to obtain the flow downstream of the vertex. Surface pressures were obtained experimentally at Mach numbers from 3.0 to 6.0 and angles of attack up to 6 deg for several flat-top wing-body configurations. These configurations consisted of half-bodies of revolution mounted beneath thin highly swept wings. Three different bodies were employed. The two conical bodies consisted of one-half of a fineness-ratio-5 cone and one-half of a fineness-ratio-2-1/2 cone. The body of the third configuration consisted of one-half of a fineness-ratio-5 ogive. For the ogive configuration, the leading edges of the wing were curved and designed to just maintain the theoretically determined bow shock along the leading edge at a Mach number of 5.0 and an angle of attack of 3 deg. The predictions of the conical flow theory of this paper for the surface pressures are found to be in good agreement with experiment at Mach numbers of 5.0 and 6.0 up to angles of attack of approximately 3 deg. Estimated lift, drag, and pitching-moment coefficients, as well as maximum lift-drag ratio, are also in good agreement with existing experimental data at a Mach number of 5.0 for a conical configuration having an arrow plan-form wing. It is also found that the generalized shock-expansion method yields reasonable good agreement with experiment for the surface pressures on the half-ogive configuration at a Mach number of 5.0 and an angle of attack of 3 deg.