Report presenting an investigation to determine the effects of inlet flow distortion on the compressor stall and engine acceleration characteristics of a J65-B-3 turbojet engine. A sinusoidal circumferential distortion and two abrupt radial tip flow distortions were introduced at the engine inlet at several altitudes and Mach number 0.8. Results regarding the definition of inlet flow distortions, acceleration limits, compressor pressure ratio at stall, fuel flow for stall, and maximum acceleration rate are provided.

From Summary: "The performance of the XRJ43-MA-3 flight-weight ram-jet engine has been investigated in a free-jet facility as part of the development program for the "Bomarc," ram-jet powered, interceptor-type missile. The engine performance was obtained at Mach numbers of 2.35, 2.50, and 2.70, altitudes from 50,000 to 75,000 feet, and Miami hot day and cold day inlet temperatures at an angle of attack of +4 deg. At each flight condition investigated, the engine control set an inner-ring fuel flow which was held constant as outer-ring fuel flow was varied."

An experimental investigation was conducted to determine the performance characteristics an underslung nose-scoop air-induction system for a supersonic airplane. Five different nose shapes, three lip shapes, and two internal diffusers were investigated. Tests were made at Mach numbers from 0 to 1.9, angles of attack from 0 deg to approximately l5 deg, and mass-flow ratios from 0 to maximum obtainable. It was found that the underslung nose-scoop inlet was able to operate at Mach numbers from 0.6 to 1.9 over a large positive angle-of-attack range without adverse effects on the pressure recovery. Although there was no one inlet configuration that was markedly superior over the entire range of operating variables, the arrangement having a nose designed to give increased supersonic compression at low angles of attack, and a sharp lip (configuration designated N3L3) showed the most favorable performance characteristics over the supersonic Mach number range. Inlets with sizable lip radii gave satisfactory performance up to a Mach number of 1.5; however, as a result of an increase in drag, the performance of such inlets was markedly inferior to the sharp-lip configuration above Mach numbers of 1.5. Throughout the range of test Mach numbers all inlet configurations evidenced stable air-flow characteristics …

Memorandum presenting a procedure based on the transonic area rule, which has been used to design a four-nacelle delta-wing airplane configuration. A flight test of a model of the configuration showed a zero-lift transonic drag rise of 0.010 which, when compared with estimates, indicated the absence of adverse interference effects.

Report presenting an investigation of pressure distributions and flow characteristics at low speed through a range of yaw and angle of attack through the stall for three small-scale low-aspect-ratio pointed wings with varying degrees of trailing-edge sweep. An effort was made to correlate pressure distributions with the strong conical vortex flow observed. Results regarding chordwise pressure distributions and flow characteristics, section lift characteristics at zero yaw, spanwise-load distributions, centers of pressure at zero yaw, and wing force and moment characteristics are provided.

Report presenting an investigation of a series of wings of various plan forms in the high-velocity field of the side-wall reflection plate of the 7- by 10-foot tunnel at a range of angles of attack and Mach numbers. The results indicated that the maximum lift coefficients obtainable increased with increase in sweep angle and decreased with Mach number at the lower subsonic Mach numbers.

Report discussing a flight test of a four-nacelle delta-wing airplane configuration based on the transonic area rule in order to counteract the effects of high pressure drag at transonic and supersonic speeds. The testing verified that the transonic area rule can make designing complex aircraft with low transonic and supersonic pressure drag much easier.

Report presenting wind-tunnel testing to determine the effect of leading-edge sweepback on the lift, drag, and pitching-moment characteristics of 3-percent-thick wings of aspect ratio 3 and taper ratio 0.4. Increasing the leading-edge sweepback of wings decreased both the lift-curve slope and the variation of static longitudinal stability at zero lift with Mach number. Results regarding lift, pitching moment, and drag are provided.

Report presenting a paper to consider the feasibility of attaining, through the use of non-afterburning engines, essentially the same range with an all-supersonic mission as the range currently obtained with a mission incorporating subsonic cruise and supersonic dash.

The action of a surge control that reduced fuel flow after receiving an indication of surge initiation was investigated. The control system could successfully limit surge to only 1 cycle but could not completely eliminate surge. Inability to interrupt a surge cycle before its completion was attributed to the conclusion that a surge cycle is irreversible.

From Introduction: "This paper describes the theoretical treatment that has been given the combustion process as it occurs in turbojet combustors. Various parts of this work have been previously published (refs. 1 to 3); this report presents a brief summary of this previous work together with new data which amplify the conclusions of references 1 to 3. Similar studies have been made of the ram-jet combustion process (ref. 4 and 5); however, the analysis for ram-jet combustion differs in some details and is therefore not included herein."

"The drag of bodies has now assumed greater importance because, as shown in NACA RM L53I15a, 1953 and NACA RM A53H18a, 1953, the transonic drag rise of an airplane can be the same as its equivalent body. Obviously, the airplane designer would like his airplane to have a low-drag equivalent body. This paper shows some of the factors which minimize the drag of bodies at transonic and supersonic speeds and shows some of the penalties caused by deviating from low-drag body shapes" (p. 1).

"An investigation was conducted to determine the performance characteristics of the axial-flow supersonic compressor of the XJ55-FF-1 turbojet engine. An analysis of the performance of the rotor was made based on detailed flow measurements behind the rotor. The compressor apparently did not obtain the design normal-shock configuration in this investigation. A large redistribution of mass occurred toward the root of the rotor over the entire speed range; this condition was so acute at design speed that the tip sections were completely inoperative" (p. 1).

Report presenting a theoretical and experimental study to determine the effectiveness of camber in reducing the drag due to lift resulting from pressure forces acting on low-aspect-ratio triangular and sweptback wings. Design charts are included which aid in the selection of camber for various sweepback angles and Mach numbers. Results regarding drag characteristics and lift and moment characteristics are provided.

From Summary: "An investigation was conducted in the Langley 19-foot pressure tunnel to determine the lift, drag, pitching-moment and stalling characteristics of a 1/4 -scale partial-span model of the left wing of the Republic XF-12 airplane. The effects of a duct inlet, located between the nacelles at the leading edge of the wing, on those characteristics were also investigated. The Reynolds numbers for the investigation covered a range from 4,500,000 to 8,600,000. The results of the investigation indicated that maximum lift coefficients of 1.36, 1.71, and 2.11 were measured on the model with flaps neutral and deflected 20 deg and 55 deg, respectively at a Reynolds number of 8,600,000."

From Summary: "An investigation of the low-speed, power-off stability and control characteristics of a 1/10-scale model of the Convair YF-102 airplane has been made in the Langley free-flight tunnel. The model was flown over a lift-coefficient range from 0.5 to the stall in its basic configuration and with several modifications involving leading-edge slats and increases in vertical-tail size. Only relatively low-altitude conditions were simulated and no attempt was made to determine the effect of freeing the controls."