"A ram-jet engine with an experimental 48-inch-diameter combustor was investigated in a free-jet facility. The combustor design comprised a large-volume annular pilot region and an array of sloping baffle- or gutter-type flameholders. The combustor was intended to operate at a fuel-air ratio of about 0.037" (p. 1).

Free jet tests of 48 inch diameter ramjet combustor with annular can-type flame holder.

Memorandum presenting an investigation to determine the combustion performance characteristics of gaseous hydrogen fuel in a single tubular turbojet combustor. The combustor was operated over a range of inlet-air pressures from 3.3 to 14.3 inches of mercury absolute. Results regarding stability limits and combustion efficiency are provided.

From Summary: "An experimental investigation has been made in the Langley stability tunnel to determine the low-speed yawing, pitching, and static stability characteristics of a 1/10-scale model of the Grumman F9F-9 airplane. Tests were made to determine the effects of duct-entrance-fairing plugs on the static lateral and longitudinal stability characteristics of the complete model in the clean condition. The remaining tests were concerned with determining tail contributions as well as the effect of duct-entrance-fairing plugs, slats, flaps, and landing gear on the yawing and pitching stability derivatives. These data are presented without analysis in order to expedite distribution."

"During an investigation of the J57-P-1 turbojet engine in the Lewis altitude wind tunnel, effects of inlet-flow distortion on engine stall characteristics and operating limits were determined. In addition to a uniform inlet-flow profile, the inlet-pressure distortions imposed included two radial, two circumferential, and one combined radial-circumferential profile. Data were obtained over a range of compressor speeds at an altitude of 50,000 and a flight Mach number of 0.8; in addition, the high- and low-speed engine operating limits were investigated up to the maximum operable altitude" (p. 1).

"Some aileron load characteristics for three thin wings varying in sweep have been presented for Mach numbers from 0.80 to 1.05. For the transonic Mach number range, shock effects exert a large influence on the loading, but the exact location of each shock for a specific wing design cannot be cataloged at the present time. It is shown, however, that the aileron loading, although greater in magnitude than at subsonic speeds, nevertheless varies in as uniform a fashion as at subsonic speeds" (p. 1).

Report presenting an investigation to determine the possibility of using simplified procedures for the estimation of control loads at supersonic speeds. The results indicated that relatively simple procedures are possible for the estimation of loadings on flap-type controls at supersonic speeds for cases when no flow separation occurs ahead of the hinge line.

Report presenting a wind-tunnel investigation at subsonic Mach numbers up to 0.9 to measure the root-mean-square variation of the normal forces on 27 NACA airfoil sections. The effects of thickness-chord ratio, camber, location of minimum pressure, and leading-edge radius were investigated.

Report presenting a consideration of the problem of shaping an adjoining fuselage for a given wing and supersonic Mach number so that the combination will have a low wave drag. Only fuselages that can be simulated by singularities distributed along the body axis are studied.

Memorandum presenting the Bell X-5 research airplane tested at 58.7 degrees sweepback during the program to determine the characteristics of a variable-sweep fighter airplane at transonic speeds. This paper includes the stability and control characteristics in the stable lift range up to Mach numbers near 1.0 at an altitude of 40,000 feet and to slightly lower Mach numbers at altitudes of 25,000 feet and 15,000 feet. Results regarding general comments, static stability and control characteristics, and longitudinal dynamic stability are provided.

Report presenting an investigation in a free-jet facility of a ramjet engine with an experimental 48-inch-diameter combustor. Three combustor lengths, three lengths of the shroud which separated the bypass air from the burning stream, and four fuel-distribution systems were investigated over a range of fuel-air ratios and a range of engine air flows. Results regarding the engine performance and ignition data, effect of fuel profile on combustion efficiency, total-pressure ratio, and distribution of static pressure in the main air stream are provided.

Report presenting a wind-tunnel investigation to determine the effect of lowering the wing from the top of the fuselage to the bottom of the fuselage on the longitudinal characteristics of a wing-fuselage and a wing-fuselage-tail combination with the horizontal tail at various heights above the plane of the wing. The effects of wing height on the longitudinal characteristics of the model were small.

Report presenting an investigation to determine the effect of lowering the wing from the top of the fuselage to the bottom of the fuselage on the longitudinal characteristics of a wing-fuselage and wing-fuselage-tail combination with the horizontal tail at various heights above the plane of the wing. Testing occurred at a range of Mach and Reynolds numbers.

Report presenting an investigation in the 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.41 to determine the static stability and control and drag characteristics of a model of a 35 degree swept-wing airplane. The effects of alternate fuselage shapes, wing camber, wing fences, and fuselage dive brakes on aerodynamic characteristics were also investigated.

Report presenting an investigation of the overall performance of a modified eight-stage axial-flow compressor as part of the investigation of the problems encountered in a high-pressure-ratio axial-flow compressor with transonic inlet stages. The maximum total pressure ratio obtained at design speed was 11.0 at an equivalent weight flow of 70.9 pounds per second with an adiabatic efficiency of 0.81. Results regarding modified compressor performance and comparison with predicted performance are provided.

Memorandum presenting rolling stability derivatives for a complete model with a low-aspect-ratio wing and tail surfaces for a Mach number range of 0.70 to 0.94 and for an angle-of-attack range from 0 to 13 degrees for the lower Mach numbers. The model test results indicated regions of neutral or unstable damping in roll at Mach numbers of 0.85 and 0.90 in the higher angle-of-attack range for the basic model. Results regarding damping in roll, yawing moment and lateral force due to rolling, and aileron characteristics are provided.

Report presenting some experimental and theoretical studies of axisymmetric free jets exhausting from sonic and supersonic nozzles into still air and into supersonic streams as a way to investigate problems associated with propulsive jets. A variety of variables for each case are considered, including jet Mach number, nozzle divergence angle, jet static-pressure ratio, free-stream Mach number, and specific heat ratios.

Report presenting an investigation to determine the effect on performance of two methods of throat-area modulation of two convergent plug nozzles. Data were obtained over a range of pressure ratios. Results regarding the performance of the translatable outer-shell-type plug nozzle, performance of the iris-outer-shell-type plug nozzles, sensitivity to throat-area variation, and air-flow parameter are provided.

Report presenting a compressor-rotor-blade vibration survey on a production turbojet engine incorporating a 13-stage axial-flow compressor with a pressure ratio of approximately 7 and an air flow of 120 pounds per second. The investigation was conducted due to the major problem that fatigue failures of compressor rotor blades have caused in the development of the axial-flow compressor. Results regarding rotating stall characteristics, rotor-blade vibration, and fourth-stage rotor-blade failure are provided.

Report of an investigation of the effect of mode of engine operation, engine configuration, and airplane flight condition on the performance and design requirements of a turboprop engine with a high-pressure-ratio single-spool compressor. Both single-spool engines and gas-generating engines are considered. The free-turbine engine was found to be less versatile than the single-spool engine because the range of turbine-inlet temperature over which it can operate is restricted.

Report presenting an investigation of the estimated minimum cooling requirements and related performance of turbojet engines equipped with high-performance single-stage turbines with air-cooled corrugated-insert blades over a range of turbine-inlet temperature, tip speed, and hub-tip radius ratio for Mach number 2 at 50,000 feet. The effects of stress-ratio factor, flight Mach number, altitude, turbine rotor impeller efficiency, and outside heat-transfer coefficient on cooling requirements were also investigated.

Report presenting the lateral stability and control characteristics were investigated on the Convair XF-92A delta-wing airplane during the flights of the NACA research program. The investigation included sideslips, aileron rolls, and rudder pulses at a range of altitudes and indicated speeds. The lateral handling characteristics appear satisfactory when viewed in terms of gradually increasing sideslips, lateral control effectiveness, and period, and damping.

Report presenting an investigation of the low-speed static lateral stability characteristics of a complete model with a series of M- and W-wings. They were modified from a basic 45 degree swept wing and were designed to cover a range of spanwise locations of the sweep discontinuity.

Report presenting an investigation at low speed in the 300 mph tunnel to determine the static aerodynamic forces and moments on a missile model during simulated launching from 25-percent-semispan and wing-tip locations of a wing-fuselage combination. Results regarding the differences in location and effects induced by the fuselage are provided.