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An investigation was conducted in the Langley 14- By 22-Foot Subsonic Tunnel to determine the low-speed aerodynamic characteristics of a powered generic NASP-like configuration in ground effect. The model was a simplified configuration consisting of a triangular wedge forebody, a rectangular mid-section which housed the propulsion simulation system, and a rectangular wedge aftbody. Additional model components included a delta wing, exhaust flow deflectors, and aftbody fences. Six-component force and moment data were obtained over an angle of attack range from −4° to 18° while model height above the tunnel floor was varied from 1/4 inch to 6 feet. Variations in freestream dynamic pressure, from 10 psf to 80 psf, and engine ejector pressure yielded a range of thrust coefficients from 0 to 0.8. Flow visualization was obtained by injecting water into the engine simulator inlets and using a laser light sheet to illuminate the resulting exhaust flow.

Wind tunnel investigations were conducted as part of an effort to develop a stability and control database for an aerospace plane concept across a broad range of Mach numbers. The generic conical design used in these studies represents one of a number of concepts being studied for this class of vehicle. The baseline configuration incorporated a 5° cone forebody, a 75.96° delta wing, a 16°leading-edge sweep deployable canard and a centerline vertical tail. Tests were conducted in the following NASA-Langley facilities spanning a Mach range of 0.1 to 6:30- by 60-Foot Tunnel,14- by 22-Foot Subsonic Tunnel, Low Turbulence Pressure Tunnel, National Transonic Facility, Unitary Plan Wind Tunnel, and the 20 Inch Mach 6 Tunnel. Data were collected for a number of model geometry variations and test conditions in each facility. This paper highlights some of the key results of these investigations pertinent to stability considerations about all three axes.

Supersonic wind tunnels with much lower stream disturbance levels than in conventional tunnels are required to advance transition research. The ultimate objectives of this research are to provide reliable predictions of transition from laminar to turbulent flow on supersonic flight vehicles and to develop techniques for the control and reduction of viscous drag and heat transfer. The experimental and theoretical methods used at NASA Langley to develop a low-disturbance pilot tunnel are described. Typical transition data obtained in this tunnel are compared with flight and previous wind-tunnel data and with predictions from linear stability theory,

The results of a design study of a Hybrid Laminar Flow Control (HLFC) wing at transonic speed and correlative studies for finite, swept supersonic wings are discussed in this paper. Transonic HLFC wing was designed such as to obtain laminar laminar flow on the the wing upper surface for X/C of 0.6 and with suction applied from the leading edge to 60% of the chord and with suction applied from just aft of the leading edge to twenty-five percent of the chord. New theoretical methods have been recently developed for predicting pressure distributions, supersonic wave drag and transition location for finite swept wings at transonic and supersonic Mach number conditions and are illustrative computations are given. Results for laminar and turbulent boundary-layer parameters consisting of the displacement effects and skin friction drag are also presented.

The ability of modern airplane surfaces to achieve laminar flow over a wide range of subsonic and transonic cruise flight conditions has been well-documented in recent years. Current laminar flow flight research conducted by NASA explores the limits of practical applications of laminar flow drag reduction technology. Past laminar flow flight research focused on measurements of transition location, without exploring the dominant instability(ies) responsible for initiating the transition process. Today, it is important to understand the specific causes(s) of laminar to turbulent boundary layer transition. This paper presents results of research on advanced devices for measuring the phenomenon of viscous Tollmien-Schlichting (T-S) instability in the flight environment. In previous flight tests, T-S instability could only be inferred from theoretical calculations based on measured pressure distributions.