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A new low-boom target generation approach is presented which allows the introduction of a trim requirement during the early conceptual design of supersonic aircraft. The formulation provides an approximation of the center of pressure for a presumed aircraft configuration with a reversed equivalent area matching a low-boom equivalent area target. The center of pressure is approximated from a surrogate lift distribution that is based on the lift component of the classical equivalent area. The assumptions of the formulation are verified to be sufficiently accurate for a supersonic aircraft of high fineness ratio through three case studies. The first two quantify and verify the accuracy and the sensitivity of the surrogate center of pressure corresponding to shape deformation of lifting components. The third verification case shows the capability of the approach to achieve a trim state while maintaining the low-boom characteristics of a previously untrimmed configuration. Finally, the new low-boom target generation approach is demonstrated through the early conceptual design of a demonstrator concept that is low-boom feasible, trimmed, and stable in cruise.

Subsonic aerodynamic experiments, both with and without sideslip, were performed on a strake/wing planform for both static and dynamic conditions. The strake sweep, wing sweep, and strake-to-wing fineness ratio for the model was 80 deg/60 deg/0.6. Flow visualization information was obtained by marking the strake and wing vortices with smoke in order to determine the state of the vortices (i.e., existence, preor postbreakdown). Fivecomponent force and moment data was also taken. Both the flow visualization and the force and moment information were obtained for the model at static angles of attack, as well as for the model undergoing sinusoidal pitching motions. The minimum and maximum angles of attack of the motions were varied. Also varied were the reduced frequencies of the motions. The combined information obtained from the flow visualization and force experiments was used to explain some of the dynamic effects observed for the double delta wing.

The axial singularity solution for the axisymmetric inverse problem has been extended to utilize doublet elements with linear intensity distribution. The solution converges faster than the source-based method, and is therefore quite promising. A procedure based on this solution has been used to design low-drag laminar fuselage shapes for small aircraft applications with a volumetric Reynolds number range of 10-30 million. A profile with a fineness ratio of 6, transition at 40% of body length, and volumetric drag coefficient of 0.012 at a nominal /?v of 15 million, has been developed. The present inverse procedure was shown to be a powerful alternative to optimization methods. Several transition criteria were investigated in the course of the study. The Crabtree criterion appears to be the most consistent. Experimental transition data for axisymmetric bodies at high (flight) Reynolds numbers are urgently needed. Nomenclature CDV = volumetric drag coefficient based on V 2/3 as characteristic area /,. = fineness ratio, body length/maximum diameter H = boundary-layer shape factor, 8*/9 L = body length / = length of an axial-singularity element n = number of doublet elements q = local velocity on body surface (at the edge of the boundary layer) RL = length Reynolds number based on UM and L Rs = Reynolds number based on local velocity and s Rx = Reynolds number based on local velocity and jc Re = Reynolds number based on local velocity and 9 7?v = volumetric Reynolds number based on Ux and V1/3 r. — local body radius and radial coordinate 5 = length along body surface starting at nose Ux = freestream velocity u = axial component of q v = radial component of q X = nondimensional axial coordinate, x/L

An experimental investigation to determine the effectiveness of porosity for alleviating side forces on forebodies was conducted in the NASA Langley Research Center 7 x 10 ft High-Speed Wind Tunnel. The study consisted of a comparison of experimental force, moment, and surface pressure results obtained on a fineness ratio 5.0 tangent-ogive porous forebody with 0.020 in. hole diameter and 22 percent porosity with results obtained on a solid forebody. The forebodies were tested with cylindrical afterbodies. The solid forebody was tested with transition grit to simulate fully turbulent conditions and without transition grit to simulate free transition conditions. The extent of porosity on the forebody was varied to determine the extent of porosity needed to alleviate side forces. Static longitudinal and lateral-directional stability and surface pressure data were obtained at Mach numbers of 0.2, 0.5, and 0.8. The angle of attack range was from 5 to 45 deg and roll angles from -90 to 180 deg. The solid forebody exhibited large asymmetries at moderate to high angles of attack causing large side forces and yawing moments. The porous forebody exhibited no significant side forces or yawing moments at any angle of attack tested.

A series of low-speed wind-tunnel tests on a generic airplane model with a cylindrical fuselage are made to investigate the effects of forebody shape and fineness ratio, and fuselage/wing proximity on static and dynamic lateral/directional stability. During the stability investigation ten forebodies were tested including three different cross-sectional shapes with fineness ratios of 2,3, and 4. In addition, the wing was tested at two longitudinal positions to provide a substantial variation in forebody/wing proximity. Conventional force tests were conducted to determine static stability characteristics, and single-degree-of-freedom free-to-roll tests were conducted to study the wing rock characteristics of the model with the various forebodies. Flow visualization data were obtained to aid in the analysis of the complex flow phenomena involved. The results show that the forebody cross-sectional shape and fineness ratio and forebody/wing proximity can strongly affect both static and dynamic (roll) stability at high angles of attack. These characteristics result from the impact of these factors on forebody vortex development, the behavior of the vortices in sideslip, and their interaction with the wing flowfield.

An experimental and theoretical investigation of planform effects on a low-fineness ratio multibody configuration has been conducted in NASA Langley Research Center's Unitary Plan Wind Tunnel at Mach numbers 1.6, 1.8, 2.0, and 2.16. Experimental and theoretical values of lift, drag, and pitching moment were obtained on three configurations which varied in outboard panel planform only. The three variations were at 65 deg delta, a 70/66 deg cranked arrow, and a 20 deg trapezoidal planform . The purpose of the study was to determine the effect of wing planform on the supersonic aerodynamics and to evaluate the ability of two existing linearized-theory aerodynamic methods to predict these effects. Experimental data showed that the planforms produced the lift, drag-due-to-lift, and pitching-moment characteristics typically found on single-body configurations. However, the data also showed that planform has a minimal influence on zero-lift drag, which is not typical of single-body configurations. Theoretical aerodynamic analysis indicated that codes based on linearized theory adequately predicted the effect of planform on the supersonic aerodynamics.

Experimental spanwise pressure distributions for a 60 deg delta wing/body of approximate fineness ratio 7.6 have been obtained and compared to predictions using full-potential theory. Analysis was performed at Mach 1.6 for angles of attack in the range of 0.8-10 deg, and for Mach numbers ranging from 1.4 to 1.8 at lift coefficient 0.3. The intent of the study was to examine an attached flow approach for maneuver wing design in the presence of a fuselage. For the Mach number and angle-of-attack conditions considered, the full-potential theory accurately modeled the pressure distributions, provided the flow remained attached. By combining the full-potential theory results with empirical shock-induced separation criteria, it was found that the onset of shock-induced separation can be predicted.

The optimum external configuration for an airplane to meet specific design and performance requirements is determined by use of an analytical method that considers systematic variations of wing sweep, thickness ratio, fuselage fineness ratio, cruising speed, and design weight. Relationships among the various structural, weight, and aerodynamic parameters used in the method are derived. Evaluation of the parameters is then based on limits imposed by current experience and assumptions. The specific design objectives in this study were to minimize takeoff weight and maximize the cruising speed for a specific range with a designated powerplant. The optimum configuration resulting from this study has a 36-deg wing sweep and 14.6% wing section thickness, a ramp weight of 21,575 Ib, and a range of 2500 n.mi. with 10 passengers. Besides providing an optimum design and an analytical method for its determination, the method and a number of the curves can be used to determine configurations which will meet other specific performance requirements.

The effect of shoulder radiusing and grooving (longitudinally and circumferentially) the afterbodies of bluff bodies to reduce the base drag at low speeds is investigated experimentally. Shoulder radii as large as 2.75 body diameters are examined. Reynolds number (ReD) based on body diameter varied from 20,000 to 200,000. Results indicate that increasing the shoulder radius to 2.00 body diameters can reduce the drag levels to those of a streamline body having 67 percent greater fineness ratio. For the relatively sharp shoulder case, body drag reductions as large as 50 and 33 percent are obtained using circumferential or longitudinal grooves, respectively.

The performance of hydrogen-fueled, hypersonic transports was analyzed in a computer study. The purpose of the study was to define the over-all mission capabilities of Mach 4 to 8 transports operating from conventional runways and operating within reasonable constraints set by structural, aerodynamic, and propulsive considerations. Tradeoff studies were conducted in an effort to define the vehicle characteristics which resulted in maximum payload fraction and also to indicate the areas in which additional research will be needed. The effects of variations in cruise Mach number, climb trajectory, engine sizing, inlet pressure recovery, regenerative cooling requirements, wing loading and aspect ratio, and fuselage fineness ratio were evaluated by relating them to the resulting changes in payload fractions. The results of the tradeoff studies indicated that hydrogen-fueled, hypersonic transports cruising at a Mach number of 6 can attain payload fractions (including reserve fuel) of about 20% of the vehicle's gross weight at ranges of about 5000 naut miles. However, before the technical feasibility of the hypersonic transports can be confirmed, additional analytical research and experimental investigations will be required in several areas, notably those involving the inlets, engines, and thermal protection systems.

Thus, the theoretical two-dimensional time-dependent solution, Eq. (9), has been tested as a description of the viscous cross flow for large fineness ratio bodies approximating circular cylinders at moderate angles of attack. Some agreement has been obtained, although empirical procedures are still required and the region of applicability is narrow. For slender bodies of high angles of attack, where vortices depart an appreciable distance from the body surface, Eq. (9), a boundary-layer-type solution, would be invalid.