Memorandum presenting tests at two supersonic speeds to obtain experimentally the aerodynamic damping characteristics of a control surface on a two-dimensional wing. The control surface had a chord of 1.67 inches and a span of 7.25 inches and was supplied in three materials with different mass, inertia, and stiffness properties. Results regarding the presentation of data and comparison with theory and comparison with control-surface data for a triangular wing are provided.

Report presenting the unsteady-lift functions for a wing undergoing a sudden change in sinking speed for delta wings with three aspect ratios and rectangular and elliptical wings with three aspect ratios. Results regarding the delta-plan-form K1 function, indical spanwise loading distributions, effect of plan-form shape on the K1 function, and effect of aspect ratio on the K1 and K2 functions are provided.

From Introduction: "The fundamental, practically the most important branch of the modern mechanics of a viscous fluid or a gas, is that branch which concerns itself with the study of the boundary layer. The presence of a boundary layer accounts for the origin of the resistance and lift force, the breakdown of the smooth flow about bodies, and other phenomena that are associated with the motion of a body in a real fluid. The concept of boundary layer was clearly formulated by the founder of aerodynamics, N. E. Joukowsky, in his well-known work "On the Form of Ships" published as early as 1890."

Note presenting a determination of the boundary layer behind a shock or thin expansion wave advancing into a stationary fluid. The assumption of a thin expansion wave was found to be valid for weak expansions but becomes progressively less accurate for strong expansion waves. The turbulent-boundary-layer solutions in this report represent an extension of empirical, semi-infinite flat-plate, boundary-layer data to the case where the wall is moving.

A study to evaluate a method for predicting the downwash in a transverse plane behind a wing-body combination throughout a range of angles of attack. The wing-body combination had a ratio of maximum body diameter to wing span of 0.259, a triangular wing with an aspect ratio of 2.0, and a body of revolution with a fineness ratio of 12.5. Results regarding span loading, downwash, and vortex cores are provided.

Effects of Reynolds number and angle of attack on the pressure distribution and normal-force characteristics of a body of revolution consisting of a fineness ratio 3 ogival nose tangent to a cylindrical afterbody 7 diameters long have been determined. The test Mach number was 1.98 and the angle-of-attack range from 0 degree to 20 degrees. The Reynolds numbers, based on body diameter, were 0.15 x 10(6) and 0.45 x 10(6). The experimental results are compared with theory.

Report presenting pressure distributions and force characteristics for a body of revolution consisting of a fineness ratio of 5.75, circular-arc, ogival nose tangent to a cylindrical afterbody for a range of angles of attack. Comparison of the theoretical and experimental pressure distributions shows that for zero lift, both slender-body theory and higher-order theories yield results that are in good agreement with theory.

Report presenting an investigation of the Reynolds number for transition on the outside of a hollow cylinder with heat transfer from the boundary layer to the wall at Mach number 6.9. At a given Mach number, it appears that the Reynolds number based on leading-edge thickness is an important parameter in comparisons of flat-plate transition data from various installations.

A brief review of exact two-dimensional supersonic flow theory and Ackeret's linearized theory are first presented. The lift and drag coefficients of a cascade of flat plates are calculated exactly and compared to those obtained using the linearized theory. The forces on the cascade are determined for unsteady inlet flow. The flat plate cascade theory is extended to compute the efficiency of a supersonic propeller with friction and finite blade thickness.

This report is concerned with fluid mechanics of two-dimensional cascades, particularly turbine cascades. Methods of solving the incompressible ideal flow in cascades are presented. The causes and the order of magnitude of the two-dimensional losses at subsonic velocities are discussed. Methods are presented for estimating the flow and losses at high subsonic velocities. Transonic and supersonic flows in lattices are then analyzed. Some three-dimensional features of the flow in turbines are noted.

Note presenting a nonlinear plate theory of motion, valid for large deflections, in the light of the three-dimensional theory and other nonlinear plate theories. The nonlinear equations are solved for the case of propagation of straight-crested waves and the wave velocities are computed for various values of the parameters involved.

Report presenting an investigation in the transonic blowdown tunnel to determine the effects of variations in inlet lip swagger from 0 to 60 degrees on the internal-flow characteristics of an unswept semielliptical scoop-type air-inlet model without boundary-layer control. Tests were made for a range of Mach numbers and mass-flow ratios. Results regarding the flow over the fuselage nose, total-pressure recovery at inlet, flow distortions at inlet, and inlet-design considerations are presented.

Report presenting an investigation of a leading-edge slat as a possible longitudinal control device for vertically rising airplanes that utilize the redirected-slipstream principle at zero forward speed in a static-thrust facility. A semispan wing model equipped with large-chord slotted flaps and two large-diameter overlapping propellers was used. Results regarding the effect of slat position, effect of propeller position, characteristics of the modified model, and a comparison of characteristics with one or two propellers are provided.

Report presenting an investigation to determine the transition Reynolds numbers on a 10 degree cone in the 4- by 4-foot supersonic pressure tunnel at three Mach numbers and a range of Reynolds numbers. Results regarding a smooth cone and a cone with roughness are provided.

Report presenting the experimental results of an investigation to determine compressive strength and creep lifetime of 2024-T aluminum-alloy skin-stringer panels at room temperature and 400 degrees Fahrenheit. A method which makes use of time-dependent compressive stress-strain curves for predicting creep lifetime of panels is presented.

Report presenting an investigation of the low-speed cornering characteristics of two 26 x 6.6, type VII, 12-ply-rating tires under straight-yawed rolling over a range of inflation pressures and yaw angles for two vertical loads. Static testing and vibration testing were also carried out. The quantities measured included lateral or cornering force, drag force, torsional moment or self-alining torque, pneumatic caster, vertical tire deflection, lateral tire deflection, wheel torsion or yaw angle, rolling radius, and relaxation length.

Report presenting measurements of the oscillating air forces on a two-dimensional wing oscillating in pitch about the midchord that have been measured at various mean angles of attack and at two different Mach numbers. The magnitudes of normal-force and pitching-moment coefficients were found to be much higher at large angles of attack than at low angles of attack for some conditions. Results regarding the effects at low angles of attack, high angles of attack, and a description of finite-span wings oscillating in first bending mode are provided.

"Some insight into the range of applicability of the transonic area rule has been gained by comparison with the appropriate similarity rule of transonic flow theory and with available experimental data for a large family of rectangular wings having NACA 63AXXX profieles. In spite of the small number of geometric variables available for such a family, the range is sufficient that cases both compatible and incompatible with the area rule are included" (p. 1).

An investigation of the effectiveness of a wing equipped with a sliding flap and a leading-edge slat in deflecting a propeller slipstream downward for vertical take-off was conducted in a static-thrust facility. Results regarding the characteristics with leading-edge slat and with extended flap chord are provided.

Report presenting a description of a temperature-measuring system using two sonic-flow orifices in series, which is used to measure the exhaust gas temperature of an afterburning jet engine mounted in a swinging pitot-static probe. It was found to improve the reliability and reduce the computational time required of the previously used system.

Report presenting an investigation to determine the effects of the various components and combinations of components on the static longitudinal and lateral stability characteristics of unswept-midwing models having wings with a variety of aspect ratios for a range of angles of attack. Results regarding static longitudinal stability characteristics and static lateral stability characteristics are provided.

Report presenting an investigation of the static longitudinal-stability characteristics of a 45 degree sweptback wing-fuselage configuration with and without a sweptback horizontal tail for a range of angles of attack and Mach numbers. Special focus is given to the pitching-moment characteristics, location of the aerodynamic center, and slope of the lift-coefficient curve.

Note presenting the effect of inclination of the airstream on the measured pressures of 54 total-pressure tubes for angles of attack up to 60 degrees and over a Mach number range from 0.26 to 1.62. The investigation was conducted in five wind tunnels. Results regarding effect of impact-opening size, effect of varying the shape of the internal chamber, effect of external shape, effect of slant profile, effect of venting of shielded tubes, effect of varying the probe position and the throat diameter of shielded tubes, effect of slant profile on shielded tubes, effect of the shape of the entry of shielded tubes, effect of Mach number on nonshielded tubes, and effect of Mach number on shielded tubes are provided.

This volume continues the NACA study of combustion principles for aircraft propulsion. The various aspects of combustion pertinent to jet engines are organized and interpreted with quite extensive information, particularly for basic or fundamental. subject matter. The report concerns only air-breathing engines and hydrocarbon fuels, and not rocket engines and high-energy fuels.