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Discontinuous wing leading-edge droop designs have been evaluated as a means of modifying wing autorotative characteristics and thus improving airplane spin resistance. Addition of a discontinuous outboard wing leading-edge droop to three typical light airplanes having NACA 6-series wing sections produced significant improvements in stall characteristics and spin resistance. Wind tunnel tests of two wings having advanced natural laminar flow airfoil sections indicated that a discontinuous leading-edge droop can delay the onset of autorotation at high angles of attack without adversely affecting the development of laminar flow at cruise angles of attack.

A two-component LV system was used to make detailed measurements of the flowfield around both a single-rotation and a counter-rotation propeller/nacelle. The conditions measured for the single-rotation tractor configuration include two different blade angles and two propeller advance ratios, and for the counter-rotation propeller configuration include both pusher and tractor mounts. The measurements show the increasing slipstream velocities and contraction with increasing blade angle and with decreasing advance ratio. Data for the counter-rotation system show that the aft propeller turns the flow in the opposite direction from the front propeller. Additionally, the LV system was used as a diagnostic tool to provide data to explain the large side force measured on the propeller/nacelle at angle-of-attack.

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.

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.

Results of a recent low-speed wind-tunnel investigation conducted to define the forebody flow on a 16% scale model of the NASA High Angle-of-Attack Research Vehicle (HARV), an F-18 configuration, are presented with analysis. Measurements include force and moment data, oil-flow visualizations, and surface pressure data taken at angles of attack near and above maximum lift (36° to 52°) at a Reynolds number of one million based on mean aerodynamic chord. The results presented identify the key flow-field features on the forebody including the wing-body strake.

An experiment was conducted to measure the surface pressures and sectional side forces on a 5° cone with three nose tips. The nose tips included a sharp, an 8.7% blunt, and a 17.5% blunt nose tip. Rings of pressure orifices were located at 40% and 80% of the model length and the model was rolled from ±180° in 9° increments to determine roll dependence. The sectional side force data for the sharp cone showed a strong dependence on the roll orientation of the model. The blunt nose cone configurations also showed a dependence on roll orientation. The blunt nose configurations were effective in reducing the sectional side force for angles of attack up to 25°. However, at angles of attack greater than 35°, the reduction was no longer significant. Pressure distributions for three angles of attack are presented to highlight details of the flow when: vortex asymmetries are just beginning; the vortices are in a steady asymmetric state; a vortex has shed between the 40% and 80% stations.

An investigation was conducted to determine the fore and aft elastic response characteristics and footprint geometrical properties of 30 × 11.5 −14.5, Type VIII, bias-ply and radial-belted aircraft tires. Stiffness and damping characteristics of each tire were obtained from load-deflection curves generated from static tests. Tire footprints were obtained for various vertical loads, and geometrical measurements were obtained from the resulting silhouettes. Results of this investigation show considerable differences in stiffness and damping characteristics between the bias-ply and radial-belted tire designs. Footprint geometrical data indicate that footprint aspect ratio effects may interfere with improved hydroplaning potential associated with the radial-belted tire operating at higher inflation pressures. Tire-wheel slippage problems encountered when testing the radial-belted tire design required special attention.

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.

Market growth and technological advances are expected to lead to a new generation of long-range transports that cruise at supersonic or even hypersonic speeds. Current NASA/industry studies will define the market windows in terms of time frame, Mach number, and technology requirements for these aircraft. Initial results indicate that, for the years 2000 to 2020, economically attractive vehicles could have a cruise speed up to Mach 6. The resulting research challenges are unique. They must be met with new technologies that will produce commercially successful and environmentally compatible vehicles where none have existed. Several important areas of research have been identified for the high-speed civil transports. Among these are sonic boom, takeoff noise, thermal management, lightweight structures with long life, unique propulsion concepts, unconventional fuels, and supersonic laminar flow.

Although the number of possible applications of boundary-layer control is large, a discussion is given only of those that have received the most attention recently at NASA Langley Research Center to improve airfoil drag characteristics. This research concerns stabilizing the laminar boundary layer through geometric shaping (natural laminar flow, NLF) and active control involving the removal of a portion of the laminar boundary layer (laminar flow control, LFC) either through discrete slots or a perforated surface. At low Reynolds numbers, a combination of shaping and forced transition has been used to achieve the desired run of laminar flow and control of laminar separation. In the design of both natural laminar flow and laminar flow control airfoils and wings, boundary layer stability codes play an important role. A discussion of some recent stability calculations using both incompressible and compressible codes is given.