Report presenting an investigation of four axial-flow single-stage compressors with a hub-tip radius ratio of 0.5 with three objectives related to the ideal symmetrical-velocity-diagram design, the comparison of blade-element turning angle and loss data with cascade data, and the validity of the simplified radial-equilibrium equation. Results regarding the overall performance, blade-element performance, and radial equilibrium considerations are provided.

Report presenting the effects of deflecting full-span, constant-chord, leading-edge flaps with round or sharp leading edges on the lift, drag, and pitching-moment characteristics of a model of an interceptor-type aircraft at supersonic and subsonic speeds.

Memorandum presenting an investigation of the effects of orientation by multi-axial stretching on properties of various plastic glazing materials. The materials studied were Lucite HC-222, Plexiglass 55, Gafite, and resin C. Some of the tests conducted included dimensional stability at elevated temperatures, surface abrasion, standard tensile tests, and stress-solvent crazing tests using ethylene dichloride.

The acceleration characteristics, in the region of maximum acceleration and compressor stall and surge, of an axial-flow turbojet engine with a fixed-area exhaust nozzle were determined by subjecting the engine to fuel flow steps, ramps, and ramps with a sine wave superimposed. From the data obtained, the effectiveness of an optimalizer type of control for this engine was evaluated. At all speeds above 40 percent of rated, a maximum acceleration was not obtained until the engine reached the point of stall or surge. A sharp drop, as high as 80 percent of maximum, in acceleration then occurred as the compressor entered surge of stall. With the maximum acceleration occurring at the point of surge or stall, the optimalizer-type control could not prevent the engine from entering surge or stall. Effective operation of the control may still be possible by sensing the sharp drop in acceleration experienced at the point of stall or surge and using this signal to limit fuel flow. The success of this type of operation would depend on the magnitude of the stall-recovery hysteresis.

Memorandum presenting an axial-flow turbojet engine operated on a sea-level static test stand to determine whether or not detectable signals were present in pressures and blade stresses that are usable as stall warnings in control applications. Surge and stall behavior were closely examined for the unique characteristics that could be used to control engine behavior in the vicinity of surge and stall. Results regarding a steady-state survey, acceleration study, investigation of an optimalizing technique, and an examination of engine damage are provided.

"An experimental investigation of the stage performance of the J71 three-stage turbine at design speed and pressure ratio was conducted as part of a program involving research on the design and operational characteristics of high-work-output multistage axial-flow turbines. The investigation indicated that the over-all work output of the turbine based on temperature measurements at design speed and pressure ratio was 95 percent of the design value, with the greatest loss in work output and efficiency occurring in the first stage. Choking in the second-stage stator over most of the pressure ratios investigated prevented the attainment of design first-stage work output and also prevented any increase in work output of the first stage with increasing over-all pressure ratio at design speed" (p. 1).

Engine tests were made on a transpiration-cooled blade that was fabricated from an internal load-carrying member with an external surface of wire cloth. After operation in the engine, some damage was noted at the tip region of the trailing edge of the blades. On other sections of the blade, the wire cloth did not appear greatly overheated, and it appeared that satisfactory chordwise temperature distribution was provided by orifices in the blade base.

Memorandum presenting an investigation of flow deflectors which extend forward of an open-nose inlet for improving positive angle-of-attack performance and auxiliary scoops for use at off-design engine air-flow conditions at low angles of attack to determine their effect on net inlet performance. The results show that a deflector inlet and a basic open-nose inlet have about the same net performance at low angles of attack although the flow steadiness characteristics of the deflector inlet are the less desirable.

Report presenting a program to investigate the aerodynamic characteristics of a family of delta wings with a blunt double-wedge section in the hypersonic tunnel at Mach number 6.9. The wings varied at semiapex angle and were tested over a range of angles of attack and Reynolds numbers. Results regarding lift and drag characteristics, center of pressure and moment coefficient, Schlieren photographs, and surface film flow studies are presented.

Report presenting shock-position controls designed to actuate the translating spike and variable bypass of a variable-geometry inlet in the 8- by 6-foot supersonic tunnel. The operation of the inlet was observed in combination with a J34 turbojet engine at a range of Mach numbers. Results regarding terminal-shock-positioning control systems and combination control are provided.

Report presenting an investigation of the performance of four divergent ejector configurations with shroud-exit to primary-nozzle-exit diameter ratios of 1.2 and 1.3 with minimum-shroud to primary-nozzle-exit diameter ratios of 1.1 and 1.2. A comparison of the performance of conical, cylindrical, and divergent ejectors showed that on the basis of similar design variables each configuration has merit. Results regarding the divergent-ejector performance and comparison of ejectors are provided.

Report presenting chordwise pressure distributions measured over the left wing of a Convair XF-92A delta-wing airplane at low lift to determine the effect of Mach number on the wing characteristics at transonic speeds. The critical Mach number of the basic wing was 0.83 for an airplane normal-force coefficient at 0.09. The spanwise load distributions were roughly triangular in shape.