Report presenting wing loads and load distributions obtained by differential-pressure measurements between the upper and lower surfaces of the wing of the Douglas X-3 research airplane with various leading-edge-flap deflections. The load and hinge-moment characteristics of the leading-edge flap are presented for a range of Mach numbers with and without flap deflection. Results regarding chordwise load distributions, wing-section and wing-panel characteristics, span load and pitching-moment distributions, leading-edge-flap characteristics, and a comparison with wind-tunnel data are provided.

This paper contains tail and hull loads data obtained in an investigation of a l/15-scale model of the Goodyear XZP5K airship. Data are presented in the form of tabulated pressure coefficients over a pitch and yaw range of +/-20 deg and 0 deg to 30 deg respectively, with various rudder and elevator deflections. Two tail configurations of different plan forms were tested on the model. The investigation was conducted in the Langley full-scale tunnel at a Reynolds number of approximately 16.5 x 10(exp 6) based on hull length, which corresponds to a Mach number of about 0.12.

"At subsonic speeds the pressure drag arising from the thickness of the body or wings is negligible so long as the shapes are sufficiently well streamlined to avoid flow separation. In that range there exists no possibility of either favorable or adverse interference on the pressure distributions themselves. If one body is so placed as to receive a drag from the pressure field of another then the second body is sure to receive a corresponding increment of thrust from the first" (p. 1).

Report presenting an investigation of the effect of screens in a dump-type diffuser in the 2- by 2-foot supersonic wind tunnel at Mach number 3.85. Results of a slanted half screen of 0.41 solidity, positioned 0.263 inlet diameter from the cowl lip, are presented on a range basis, as this configuration allows for shortening of the subsonic diffuser.

Report presenting an experimental investigation of a two-dimensional, 6.6-percent-thick, 6-foot-chord airfoil section equipped with suction slots for laminar-boundary-layer control. The airfoil section was designed to have favorable pressure gradients between the suction slots. Results indicated that the laminar boundary layer on the airfoil had the same extreme sensitivity to minute details of the surface condition as has been observed in other similar investigations.

"Low-lift drag data are presented herein for one 1/7.5-scale rocket-boosted model and three 1/45.85-scale equivalent-body models of the Grumman F9F-9 airplane, The data were obtained over a Reynolds number range of about 5 x 10(exp 6) to 10 x 10(exp 6) based on wing mean aerodynamic chord for the rocket model and total body length for the equivalent-body models. The rocket-boosted model showed a drag rise of about 0,037 (based on included wing area) between the subsonic level and the peak supersonic drag coefficient at the maximum Mach number of this test" (p. 1).

The efficiency of a 30 degree conical diffuser was improved as much as 20 percent and separation was eliminated by the use of vortex generators. The use of splitter cones gave only small efficiency gains, but relatively uniform diffuser-outlet velocity profiles were obtained with the better designs. A configuration which incorporated both vortex generators and a splitter cone gave efficiencies higher than those obtained with any other splitter-cone configuration and also higher than those obtained when the same vortex generators were used without the splitter cone.

Characteristic exhaust velocity over a range of mixture ratios and variations in gas velocity with distance from the injector were measured for six injectors. Comparisons of injector performance showed the gains obtained from oxidant atomization, fuel atomization, and propellant mixing. The results are compared with oxygen and heptane performance and show the effect, which is qualitatively small, of spontaneous propellant ignition on the relation between injection processes and engine performance.

A spike-type nose inlet with sharp-lip cowl was investigated at a free-stream Mach number of 2.5 with water injection in its 16-inch diameter, 11-foot-long subsonic diffuser section. Inlet total temperature of exit with liquid-air ratios of about 0.04 with no apparent change in the critical pressure recovery. The observed temperature drops were less than the theoretically predicted values, and the amount of water evaporated was 35 to 50 percent less than that theoretically possible.

Report presenting power spectral densities of normal load factor obtained for two service operational training flights of a Republic F-84G airplane and three service operational training flights of a North American F-86A airplane in order to indicate the load-factor frequency content and possible uses of power spectral methods in analyzing maneuver load data. Results regarding the spectra of the two different planes, an analytical representation of maneuver load spectrum, and probability distributions are provided.

"An investigation has been made in the Langley Unitary Plan wind tunnel at Mach numbers of 1.60, 1.80, and 2.00 to determine the aerodynamic characteristics of a 0.03-scale model of the Avro CF-105 airplane. The investigation included the determination of the static longitudinal and lateral stability, the control and the hinge-moment characteristics of the elevator, rudder, and aileron, as well as the vertical-tail-load characteristics. Although the data are presented without analysis, a limited inspection of the longitudinal control results indicates a loss in maximum lift-drag ratio due to trimming of about 1.8 because of the large static margin" (p. 1).

The stator-blade angles in the first four stages of a 16-stage axial-flow compressor were increased in order to decrease the angles of attack of these stages, and thereby to improve part-speed performance. The performance of this modified compressor was compared with that of the same compressor with original blade angles.

The performance of a control system designed for variable thrust applications was determined in an altitude free-jet facility at various Mach numbers, altitudes and angles of attack for a wide range of engine operation. The results are presented as transient response characteristics for step disturbances in fuel flow and stability characteristics as a function of control constants and engine operating conditions. The results indicate that the control is capable of successful operation over the range of conditions tested, although variations in engine gains preclude optimum response characteristics at all conditions with fixed control constants.

Report presenting a linear stability analysis and flight-test investigation on a rolleron-roll-rate stabilization system for a canard missile. This type of damper provides roll damping by the action of gyro-actuated uncoupled wing-tip ailerons. A dynamic roll instability predicted by the analysis was confirmed by flight testing and was subsequently eliminated by introducing control-surface damping about the rolleron hinge line.

Report presenting measurements of aerodynamic heat transfer at several stations on the 15 degree total angle conical nose of a rocket-propelled model in free flight at Mach numbers up to 5.2. Laminar, transitional, and turbulent heat-transfer coefficients are provided. Results regarding skin temperature time histories, transition, and computed skin temperatures are provided.