Supersonic Wave Interference Affecting Stability

Some of the significant interference fields that may affect stability of aircraft at supersonic speeds are briefly summarized. Illustrations and calculations are presented to indicate the importance of interference fields created by wings, bodies, wing-body combinations, jets, and nacelles

Investigations at supersonic speeds of 22 triangular wings representing two airfoil sections for each of 11 apex angles

Data obtained from wind tunnel investigations of two series of 11 triangular wings conducted at Mach numbers of 1.62, 1.92, and 1.40 to determine the effect of leading-edge shape and to compare actual test values with the nonviscous linear theory are presented. The two series of wings had identical plan forms, a constant thickness ratio of 8 percent, a constant location of maximum-thickness point of 18 percent, and a range of apex half-angles from 10 degrees to forty-five degrees. The first series has an elliptical leading edge and the second series a wedge leading edge. Measurements were made of lift, drag, pitching moment, and pressure distribution, the latter being confined to three wings at one Mach number

A reexamination of the use of simple concepts for predicting the shape and location of detached shock waves

A reexamination has been made of the use of simple concepts for predicting the shape and location of detached shock waves. The results show that simple concepts and modifications of existing methods can yield good predictions for many nose shapes and for a wide range of Mach numbers

Investigations at Supersonic Speeds of 22 Triangular Wings Representing Two Airfoil Sections for Each of 11 Apex Angles

The results of tests of 22 triangular wings, representing two leading-edge shapes for each of 11 apex angles, at Mach numbers 1.62, 1.92, and 1.40 are presented and compared with theory. All wings have a common thickness ratio of 8 percent and a common maximum-thickness point at 18 percent chord. Lift, drag, and pitching moment are given for all wings at each Mach number. The relation of transition in the boundary layer, shocks on the wing surfaces, and characteristics of the pressure distributions is discussed for several wings

A summary of information on support interference at transonic and supersonic speeds

An experimental investigation was performed to determine the effect on base and forebody pressures of using a sting modified with varying length splitter plates and fins instead of a conventional sting to support a cone-cylinder body of revolution. The investigation was conducted at a Mach number of 3.12 for a Reynolds number range of 2 x 10 to the 6th power to 14 x 10 to the 6th power and for an angle of attack range of 0 degrees to 9 degrees. For Reynolds numbers of 8 x 10 to the 6th power and 14 x 10 to the 6th power there was a negligible effect of the splitter plate modification on the base pressure, and at Reynolds number of 2 x 10 to the 6th power there was a small effect. Positioning the leading edge of the splitter plate at or ahead of the base made no appreciable change in the influence of the modifications on base pressure at a Reynolds number of 14 x 10 to the 6th power. With the fin-type modification there was a small increase in base pressure

Aerodynamic Investigation of a Parabolic Body of Revolution at Mach Number of 1.92 and Some Effects of an Annular Supersonic Jet Exhausting from the Base

An aerodynamic investigation of a slender pointed parabolic body of revolution was conducted at Mach number of 1.92 with and without the effects of an annular supersonic jet exhausting from the base. Measurements with the jet inoperative were made of lift, drag, pitching moment, base pressures, and radial and axial pressures. With the jet in operation, pressure measurements were made over the rear of the body with the primary variables being angle of attack, ratio of jet velocity to stream velocity, and ratio of pressure at jet exit to stream pressure