Exploratory Investigation at Mach Number of 2.01 of the Longitudinal Stability and Control Characteristics of a Winged Reentry Configuration

An investigation has been conducted to determine the longitudinal stability and control characteristics of a reentry configuration at a Mach number of 2.01. The configuration consisted of clipped delta wing with hinged wing-tip panels. The results indicate that deflecting the wing-tip panels from a position normal to the wing chord plane to a position coincident with the wing chord plane resulted in a stabilizing change in the pitching-moment characteristics but did not significantly affect the nonlinearity of the pitching-moment variation with angle of attack. The trailing-edge controls were effective in producing pitching moment throughout the angle-of-attack range for control deflections up to at least 600. The control deflection required for trim, however, varied nonlinearly with angle of attack. It would appear that this nonlinearity as well as the maximum deflection required for trim could be greatly decreased by utilizing a leading-edge control in conjunction with a trailing-edge control

Static Stability Characteristics of a Series of Hypersonic Boost-Glide Configurations at Mach Numbers of 1.41 and 2.01

An investigation of the static stability characteristics of several hypersonic boost-glide configurations has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at Mach numbers of 1.41 and 2.01 (with Reynolds numbers per foot of 2.90 x 10(exp 6) and 2.41 x 10(exp 6) respectively). This series of configurations consisted of a cone, with and without cruciform fins, a trihedron, two low-aspect-ratio delta wings that differed primarily in cross-sectional shape, and two wing-body configurations. All configurations indicated reasonably linear pitching-, yawing-, and rolling-moment characteristics for angles of attack to at least 12 deg. The maximum lift-drag ratio for the zero-thrust condition (base drag included) was about 3 for the delta-wing configurations and about 4 for the wing-body configurations

Investigation of Pressure Distribution over an Extended Leading-Edge Flap on a 42 Degrees Sweptback Wing

Pressure distribution over an extended leading-edge flap on a 42 degree swept-back wing was investigated. Results indicate that the flap normal-force coefficient increased almost linearly with the angle of attack to a maximum value of 3.25. The maximum section normal-force coefficient was located about 30 percent of the flap span outboard of the inboard end and had a value of 3.75. Peak negative pressures built up at the flap leading edge as the angle of attack was increased and caused the chordwise location of the flap center of pressure to be move forward