Report presenting a discussion of the aerodynamic forces and moments on inclined bodies of revolution. An approximate theory to allow for the effects of viscosity is developed and applied. Results regarding the variation of lift and pitching moment, and center of gravity are provided.

Report presenting an evaluation of a piston-type gas-generator engine by comparing the performance of a transport airplane powered with than engine with the same airplane powered by different engines. The engines compared were a turbojet engine, a turbine-propeller engine, a compound engine, and a turbosupercharged reciprocating engine utilizing a variable-area exhaust jet nozzle. Results regarding optimum altitude and transport performance are provided.

Report presenting the performance of a ramjet unit consisting of an intake diffuser, an exhaust nozzle, and a cluster of thin-plate burners contained in a semicircular combustion chamber as investigated in the induction aerodynamics laboratory.

Report presenting the results of an experimental investigation to determine the pressure-recovery and mass-flow characteristics of single and twin NACA submerged intakes on the sides of a fuselage at various angles of sideslip. Tests were conducted with single and twin submerged intakes on a model of a fighter-type airplane.

Report presenting the pressure-recovery characteristics of a model with two scoops situated on the aft position of a long forebody and connected through diffusors to a common settling chamber at a range of Mach numbers. The effects of interaction between the flow in the two air-induction systems and of varying the mass flow through the boundary-layer scoops were studied. Results at several different angles of attack are provided.

Report presenting tests to determine the recovery of total pressure attainable at a range of Mach numbers, which were performed with models with twin-scoop inlets situated on the sides of a long forebody. Results regarding ducts without slots and ducts with slots are provided.

Report presenting aerodynamic and icing investigations conducted in the icing research tunnel on a model of a turbojet-engine nacelle with a long straight air inlet in order to provide basic design criteria for hot-gas blowback systems. The most uniform temperature distribution was obtained with a bleedback of 4.4 percent at a gas temperature of 1000 degrees Fahrenheit and resulted in an average dry-air-temperature rise of 46 degrees Fahrenheit.

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 the specific impulse of liquid diborane and liquid oxygen over a range of mixture ratios in a 100-pound-thrust rocket engine operating at a combustion-chamber pressure of 300 pounds per square inch absolute. A faired curve through the experimental data had approximately the same shape as the theoretical curve.

Report presenting an experimental study of the pressure fluctuations at the combustion-chamber inlets of a ramjet operating a Mach number of 1.92 in the supersonic tunnel. Results are presented for two burner configurations of different operating characteristics to indicate trends in pressure magnitude, frequency, and wave form. Results regarding the perforated conical flame holder, regenerative-type burner, and diffuser total-pressure recovery are provided.

Report presenting the results of an experimental investigation of rim cracking in a welded-blade composite gas-turbine wheel, in two carbon-steel disks, and five tool-steel disks. Various characteristics were investigated, including the effectiveness of holes in preventing crack propagation and the influence of hardness and various types of notch on rim cracking.

Report presenting data from 95 subsonic flutter tests conducted in the flutter research tunnel on untapered cantilever wings with sweepback angles of 0, 45, and 60 degrees and carrying a single concentrated weight. The primary purpose of the investigation was to present experimental information to be used to evaluate analytical procedures for determining the flutter speed of weighted sweptback wings. The dynamic pressure, flutter velocity, Mach number, natural and flutter frequencies, and phase-angle relationships of the stresses for the natural and flutter frequencies are presented.

Report presenting an investigation of the experimental performance of chlorine triflouride and hydrazine measured over a range of propellant mixtures in a 100-pound-thrust rocket engine. Experimental values of specific impulse, volume specific impulse, heat rejection per propellant weight, nozzle thrust coefficient, and characteristic velocity were obtained as functions of percent fuel by weight in the propellant mixture.

Report presenting an investigation of the pressure distribution on an asymmetrical nonconical body at a Mach number of 1.90 over a wide range of angles of attack and yaw. Pressure distributions occurred at anticipated trends. Results regarding the pressure coefficients, Schileren photographs, boundary layers, and pressure distributions are provided.

An investigation was conducted at a Mach number of 1.91 to determine spanwise pressure distribution over a wing tip in a region influenced by a sharp subsonic leading edge swept back at 70 degrees. Except for pressure distribution on the top surface in the immediate vicinity of the subsonic leading edge, the maximum difference between linearized theory and experimental data was 2 1/2 percent (of free-stream dynamic pressure) for angles of attack up to 4 degrees and 7 percent for angles of attack up to 8 degrees. Pressures on the top surface nearest the subsonic edge indicated local expansions beyond values predicted by linearized theory.

Report presenting an investigation of the performance and general operating characteristics of a small single-cylinder, two-stroke-cycle, loop-scavenged engine using compression ignition at low compression ratios, high inlet-manifold temperatures and high inlet-manifold and exhaust-gas pressures. Results regarding the scavenging characteristics, power output, cylinder pressures, thermal efficiency, and exhaust measurements are provided.

Report presenting a flight investigation of rocket-powered models at high-subsonic, transonic, and supersonic speeds to determine the zero-lift drag of fin-stabilized bodies of revolution differing in maximum diameter. All bodies had 6.04 fineness ratio and cut-off sterns with equal base area. The most favorable location out of the 20-percent, 40-percent, and 60-percent positions were evaluated for different speeds.

Report presenting rocket-powered flight tests from high subsonic to supersonic speeds and at high Reynolds numbers to determine the zero-lift drag of a transonic wing-body and body-alone configuration. The configuration tested featured a wing of 45 degree sweepback, aspect ratio 4, taper ratio 0.6, and an NACA 65A006 airfoil section, and a body of fineness ratio of 10 and a frontal area equal to 6.06 percent of the wing-plan-form area.

"Flight tests were conducted at the NACA Pilotless Aircraft Research Station, Wallops Island, to determine the characteristics of the Allegany Ballistics Laboratory's 6.2inch-diameter Deacon and 10-inch-scale model solid-propellant rocket motors. The tests were performed to assist in the development of these rockets which were designed for, and urgently needed to propel supersonic research models and pilotless aircraft. The tests showed that the rocket motors functioned properly under various flight- acceleration loads over a range of pre-ignition grain temperatures" (p. 1).

Report presenting testing of three rocket-powered models with Douglas D-558-2 wings as horizontal fins with scaled structural parameters to determine whether the possibility of wing flutter due to torsion and bending exists in full-scale airplanes at transonic speeds. No wing flutter was present in the models tested, but pitching oscillation developed as it passed into the transonic range.

Memorandum presenting flight tests with a fighter airplane to study the effect of boundary-layer suction aft of the shock wave on airfoil drag at supercritical Mach numbers and high Reynolds numbers. Airfoil chord force was determined from pressure-distribution measurements obtained at a range of Mach numbers at steady dives. Results of the tests showed no measurable effect of suction for the suction coefficient available.

Report discussing testing using two rocket-propelled models to test the jettisonable-nose method of pilot escape. The nose was jettisoned successfully on the second attempt and the accelerations produced were found to be within human tolerance. The first test ended with a collision of the nose and wing, indicating that this is a possible risk during power-on flight.

Rocket-powered models were flown at high-subsonic, transonic, and supersonic speeds to determine the zero-lift drag of fin-stabilized parabolic bodies of revolution differing in fineness ratio and in position of maximum diameter. The present paper presents the results for fineness ratio 12.5, 8.91 and 6.04 bodies having maximum diameters located at stations of 20, 40, 60, and 80 percent of body length. All configurations had cut-off sterns and all had equal base, frontal, and exposed fin areas. For most of the supersonic-speed range models having their maximum diameters at the 60-percent station gave the lowest values of drag coefficient. At supersonic speeds, increasing the fineness ratio generally reduced the drag coefficient for a given position of maximum diameter.

"A flight test of the Aero jet Engineering Corporation's 7KS-6000 T-27 Jato rocket motor was conducted at the Langley Pilotless Aircraft Research Station at Wallops Island, Va, to determine the flight performance characteristics of the motor. The flight test imposed an absolute longitudinal acceleration of 9.8 g upon the rocket motor at 2.8 seconds after launching. The total impulse developed by the motor was 43,400 pound-seconds, and the thrusting time was 7.58 seconds" (p. 1).