Report presenting a wind-tunnel investigation of a model equipped with a 60 degree delta wing and a 60 degree delta horizontal all-movable canard control surface to determine the stability, control, and canard-surface hinge-moment characteristics at low speeds and at a Reynolds number of 9 x 10(exp 6). Two longitudinal positions of the canard surface were tested. Results regarding lift and pitching moment, drag, and hinge moment are provided.

Memorandum presenting an investigation to determine the effects of jet exhaust and Reynolds number on the flow over the vertical stabilizer and rudder of the Douglas D-558-II research airplane with special attention to an understanding of the rudder reversals experienced in flight on the full-scale airplane. Tests were made at Mach numbers of 1.62, 1.93, and 2.41 over a range of ratios of jet static pressure to stream static pressure from the jet-off ratio to about 40 and for a maximum range of sideslip angles of 6 degrees. Results regarding the effect of varying jet static-pressure ratio, effect of varying sideslip angle, effect of varying Reynolds number, and effect of varying angle of attack.

Memorandum presenting a study of the effects of automatic augmentation or controlling systems on the tail loads experienced in rolling maneuvers and in pitch-up. The results were calculated on an analog computer and the equations of five degrees of freedom were used for the rolling maneuvers and three degrees of freedom for pitch-up.

Memorandum presenting a discussion of oscillating flap-type and all-movable controls with particular emphasis on the aerodynamic forces and moments at transonic speeds. Hinge-moment results from recent wind-tunnel and rocket-powered-model tests are summarized for trailing-edge flap-type controls to illustrate the effects of control hinge-line position and profile shape on one-degree-of-freedom flutter of this type of control.

Report presenting an investigation in the 18- by 18-inch supersonic tunnel to determine the total-pressure recovery obtainable at Mach number 1.85 with a shock diffuser with projective cones designed to produce two oblique shocks ahead of the diffuser inlet. The variation of total-pressure recovery with tip projection was investigated for four cones with different included angles. A maximum total-pressure recovery of 94.5 percent was attained with the best configuration at an angle of 0 degrees.

Report presenting an investigation of the variables pertinent to the control of gas temperatures of gas-turbine engines during starting conducted by obtaining time records of the variables involved during actual starting conditions. Poor control of gas temperatures during starting was found to be caused by an accumulation of fuel in the engine before ignition and excessive fuel-flow rates at the time of ignition.

Report presenting the design and performance of a turbine to drive a supersonic compressor in order to determine the problems that would be encountered in turbines required to drive high-speed, high-weight-flow compressors. A model of the turbine was built and operated over a range of speed and total-pressure ratios. Results regarding performance, the effects of rotor choking, and possible modifications to the design are provided.

Report presenting an investigation of internal film cooling of the exhaust nozzle of a 1000-pound-thrust liquid ammonia liquid-oxygen rocket engine. With water as a coolant, approximately 16 percent of the total propellant and coolant flow was required to film-cool the entire nozzle and with anhydrous liquid ammonia, approximately 19 percent of the total flow was required. Results regarding the coolant results and performance results are provided.

Report presenting pressure and temperature changes in the cooling air flowing through air-cooled gas-turbine blades, which have been measured experimentally in a static cascade on a 10-tube blade for isothermal and nonisothermal flows, and on a 13-fine blade for isothermal flows. Values of pressure and temperature were obtained for a range of cooling-air flows over a range of Reynolds numbers, Mach numbers, and heat-transfer rates.

Memorandum presenting a study made to ascertain some of the operational characteristics of four types of air-induction systems for a turbojet-powered airplane intended to operate at a range of Mach numbers at an altitude of 35,332 feet. The investigation included air-induction systems with and wthout ramps to produce oblique shock waves, with fixed and variable inlet areas, and with divergent diffusers.

Report presenting the aerodynamic stability derivatives and pitch and yaw control effectiveness of two full-scale guided bombs over a range of Mach numbers in free-flight drop tests. Results regarding performance information, roll control operation, and aerodynamic characteristics are provided.

Report discussing testing on a research model to obtain experimental drag and heat-transfer data under constant and varying conditions of Reynolds number and heating effects. Average body turbulent skin-friction-drag coefficients were determined for both constant and varying heating effect.

Report presenting an investigation on a canard ramjet missile configuration to determine its lift and drag characteristics at low values of lift. Two rocket-boosted models that differed only in the size and deflection of the canard were used. Information about the trim lift characteristics, drag, and stability is provided.

Report presenting testing of a modified 1/10-power nose shape in free flight at Mach numbers up to 6.7. Measured heating rates were presented and compared with calculated values. Results regarding transition and comparison with previous investigations are provided.

"Data on atmospheric turbulence in the vicinity of thunderstorms obtained during a flight evaluation of an experimental C band (5.5 cm) airborne radar are summarized. The turbulence data were obtained with an NACA VGH recorder installed in a United Air Lines DC-3 airplane" (p. 1).

Report presenting an investigation of a five-stage rocket model at mach number 13.9 and an altitude of 81,500 feet. Temperatures were measured at 12 stations on the flat-face copper nose. Heating rates calculated from temperature time histories are compared to theoretical predictions of the rates.

Report presenting an investigation in the 10- by 10-foot supersonic wind tunnel on a 16-inch ramjet equipped with an afterburner and using gaseous hydrogen for the primary and afterburner fuels. The primary nozzle had a contraction ratio of 0.6 and the exit nozzle had a contraction ratio of 0.9. The results indicated that at a constant diffuser-exit Mach number, afterburner operation produces more than twice the thrust available without afterburning.

Report presenting an investigation in the 18- by 18-inch supersonic tunnel to determine design conditions for optimum performance of shock diffusers at Mach number 1.85 with a series of projecting single-shock cones with varying angles. Each cone was tested with a curved and a straight diffuser-inlet section. The effect of angle of attack was investigated for several configurations.

The evaluation of the altitude operational characteristics was part of the over-all investigation of the early developmental Iroquois engine. Engine steady-state windmilling characteristics were evaluated over a range of flight Mach numbers from 0.48 to 1.72 at altitudes of 35,000 and 50,000 feet. Engine altitude ignition limits were obtained over a range of flight Mach numbers from 0.5 to 1.5 with the standard engine ignition system and also with an oxygen boost system. A short investigation of high-speed altitude reignition following combustor blowout was conducted.