Report presenting an investigation to determine the aerodynamic characteristics in pitch at a Mach number of 6.8 of hypersonic missile configurations with cruciform trailing-edge flaps and all-movable control surfaces. Testing indicated that all-movable controls on the flared-afterbody model should be capable of producing much larger values of trim lift and of normal acceleration than the trailing-edge-flap configuration. Some of the configurations tested include body alone, body with 5 degree fins and trailing-edge flaps, and body with 10 degree flare and all-movable controls.

"The static aeroelastic divergence characteristics of a delta-plan-form model of the canard control surface of a proposed air-to-ground missile have been studied both analytically and experimentally in the Mach number range from 0.6 to 3.0. The experiments indicated that divergence occurred at a nearly constant value of dynamic pressure at Mach numbers up to 1.2. At higher Mach numbers somewhat higher values of dynamic pressure were required to produce divergence" (p. 1).

From Introduction: "The present paper presents the scale effect on the longitudinal aerodynamic characteristics, the aerodynamic characteristics in yaw, and the tuft studies for 0^o and 3.7^o yaw. The results of the effect of leading-edge and trailing-edge flaps on the aerodynamic characteristics of the wing will be presented in later reports."

Report presenting a high-speed wind-tunnel investigation of a pitot-static tube to provide information on the effects of compressibility upon the pressure indications of a representative airspeed head at high subsonic speeds. The calibration factor for the instrument has been evaluated for several small angles of pitch and yaw throughout a Mach number range from 0.30 to approximately 0.925. The results indicate that the calibration factor for each combination of pitch and yaw angles tested is almost constant with Mach number up to a Mach number of approximately 0.8.

Report presenting measurements of the chordwise pressure distributions over the 8-percent-thick wing of the XS-1 research airplane at a section near the midspan of the left wing. Data are presented for a Mach number range and a normal-force coefficient of about 0.33. Results regarding the upper-surface pressure distribution, lower-surface pressure distributions, and total section loads are provided.

"Abrupt rudder-fixed aileron rolls have been made with the Bell X-1 airplane in powered flight at Mach numbers between 0.90 and 1.06 at about 40,000 feet pressure altitude. These aileron rolls indicate that the aileron effectiveness for large deflections at Mach number 0.94 is only one-fourth the value at Mach number 0.82, and the effectiveness at Mach number 1.06 is slightly greater than at Mach number 0.94" (p. 1).

"NACA instrumentation has been installed in the Northrop X-4 airplane to obtain stability and control data during the Northrop conducted acceptance tests. The results of the fourth and fifth flights of the Northrop X-4 number 1 airplane are presented in this paper. These data were obtained for a center-of-gravity position of approximately 19.5 percent of the mean aerodynamic chord. The results of this flight showed that the directional stability as measured in steadily increasing sideslips was positive and high and that the effective dihedral was positive. The results also show the airplane to be longitudinally stable, stick fixed, with the center of gravity at 19.5 percent of the mean aerodynamic chord" (p. 1).

Report presenting aerodynamic and icing investigations in the icing research tunnel on a model of a turbojet-engine nacelle with a short straight air inlet. The hot-gas blowback system consisted of several orifices peripherally located around the inlet and was investigated in both dry-air and icing conditions. Results regarding the aerodynamic investigation without bleedback, aerodynamic investigation with cold-gas bleedback, aerodynamic investigation with hot-gas bleedback, and icing with bleedback are provided.

Report presenting an investigation of the proper means of coordinating the lift-acceleration commands and the roll-attitude commands to the autopilot in order to minimize transient racking errors encountered during bomber evasion. Results regarding the pull-up maneuver, horizontal turning maneuvers, the split S vertical plane maneuver, diving turn maneuver, other considerations involved in rolling maneuvers, and maximum rates of roll, control motion, and tracking-line motion are provided.

Report presenting free-falling recoverable-model tests at transonic speeds on a model with a low-aspect-ratio thin unswept wing and a 45 degree swept tail in the chord plane of the wing. Static- and dynamic-stability data and load-distribution data were obtained at a range of angles of attack. Results regarding lift, drag, static longitudinal stability, dynamic longitudinal stability, horizontal-tail effectiveness, loading distribution over the fuselage, and buffet boundary are provided.

Report presenting a flight investigation of the 35 degree swept-wing D-558-II rocket-powered research airplane in the transonic and supersonic speed ranges. Lift and drag values obtained in exploratory flights of the basic configuration are presented.

Report presenting an investigation of the effect of body indentation on the aerodynamic characteristics of an M-plan-form-wing-body combination for a variety of angles of attack and Mach numbers. The wing was tested on a plain body and a body with an indentation that was 65 percent of that required for a Mach number of 1.0. Results regarding drag characteristics, lift characteristics, pitching-moment characteristics, and estimated results are provided.

Ignition studies of flowing gases were made to obtain information applicable to ignition problems in gas-turbine and ram-jet aircraft propulsion systems operating at altitude conditions.Spark energies required for ignition of a flowing propane-air mixture were determined for pressure of 2 to 4 inches mercury absolute, gas velocities of 5.0 to 54.2 feet per second, fuel-air ratios of 0.0607 to 0.1245, and spark durations of 1.5 to 24,400 microseconds. The results showed that at a pressure of 3 inches mercury absolute the minimum energy required for ignition occurred at fuel-air ratios of 0.08 to 0.095. The energy required for ignition increased almost linearly with increasing gas velocity. Shortening the spark duration from approximately 25,000 to 125 microseconds decreased the amount of energy required for ignition. A spark produced by the discharge of a condenser directly into the spark gap and having a duration of 1.5 microseconds required ignition energies larger than most of the long-duration sparks.

"An analysis of the performance of the types A, B, and C combustion chambers of the 4000-pound-thrust axial-flow turbojet engine is presented. The data were obtained from investigations of the complete engine over a range of pressure altitudes from 5000 to 40,000 feet and ram pressure ratios from 1.00 to 1.86. The combustion-chamber pressure losses, the effect of the losses on cycle efficiency, and the combustion efficiency are discussed" (p. 1).

Report discussing an investigation to determine the aerodynamic characteristics of hypersonic missile configurations with cruciform trailing-edge flaps with all-movable control surfaces. The all-movable controls were found to produce much larger values of trim lift and normal acceleration than the trailing-edge-flap configuration.

"Performance characteristics of the turbine of a 4000-pound-thrust axial-flow turbojet engine was determined in investigations of the complete engine in the NACA Cleveland altitude wind tunnel. Characteristics are presented as functions of the total-pressure ratio across the turbine and of turbine speed and gas flow corrected to sea-level conditions. Three turbine nozzles of different areas were used to determine the area that gave optimum performance" (p. 1).

An investigation was conducted to determine the operational and performance characteristics of the TG-100A gas turbine-propeller engine II. Windmilling characteristics were deterined for a range of altitudes from 5000 to 35,000 feet, true airspeeds from 100 to 273 miles per hour, and propeller blade angles from 4 degrees to 46 degrees.

Report presenting a flight investigation to determine the low-speed flight characteristics of a 45 degree swept-wing fighter-type airplane with boundary-layer control on the trailing-edge flaps. The effectiveness of the flap with and without boundary-layer control was determined in conjunction with several slat modifications. The study also included low-speed flying qualities and a pilot evaluation of the operational use of the boundary-layer control system in landing approaches.

From Introduction: "The results obtained in the wind-tunnel tests at Mach numbers 2.36 and 2.87 for several configurations utilizing this wing, including results on the wing alone are presented."

"The effect of rotor-blade length, inlet angle, and shrouding was investigated with four different nozzles in a single-stage modification of the Mark 25 aerial-torpedo power plant. The results obtained with the five special rotor configurations are compared with those of the standard first-stage rotor with each nozzle. Each nozzle-rotor combination was operated at nominal pressure ratios of 8, 15 (design), and 20 over a range of speeds from 6000 rpm to the design speed of 18,000 rpm" (p. 1).

"An investigation of the EX-3 pine-cone-head pellet was made in the Langley high-speed 7-by 10-foot wind tunnel to determine the static force and moment characteristics at high Mach numbers with the reference center of gravity located at 37.5 percent of the over-all length aft of the nose. For this center-of-gravity location there were no secondary trim positions, and the center-of-pressure position was not appreciably affected by Mach number" (p. 1).

From Summary: "Tests in the Ames 40- by 80-foot wind tunnel of the static longitudinal characteristics of the Republic RF-84F were made to determine both the origin and a suitable remedy for a pitch up tendency of the airplane encountered at moderate lift coefficients. The results indicated that the pitch-up at moderate lift coefficients was caused by an abrupt change in downwash at the tail which in turn was traceable presumably to flow conditions associated with the inlet-to-wing leading-edge discontinuity."