Report discussing an investigation of a fingered semaphore spoiler and an equivalent solid spoiler configuration to determine the effects of spoiler projection, porosity, and chordwise position on the rolling effectiveness and drag of spoiler arrangements at a range of Mach numbers. The fingered semaphore spoilers were found to have less rolling effectiveness and drag than the solid spoilers.

"A 1/10-scale rocket model of the Lockheed XF-104 with faired inlets has been flown over a Mach number range from 0.80 to 1.45 to determine low-lift drag and a limited amount of stability data. The center-of-gravity locations were 4.0 and 1.5 percent of the mean aerodynamic chord before and after sustainer firing, respectively. Oscillations induced by pulse rockets were used to determine stability data. The external transonic drag coefficient increased from a value of 0.0160 at Mach number 0.80 to a maximum of 0.0432 near Mach number 1-13, with a drag rise Mach number of about 0.93. At Mach numbers where it could be determined, the model exhibited stable dynamic and static stability characteristics at low lift" (p. 1).

The primary purpose of this investigation was to obtain simultaneous data on the loads and stress experience in flight by the U. S. S. Los Angeles which could be used in rigid airship structure design. A secondary object of the investigation was to determine the turning and drag characteristics of the airship. The aerodynamic loading was obtained by measuring the pressure at 95 locations on the tail surfaces, 54 on the hull, and 5 on the passenger car. These measurements were made during a series of maneuvers consisting of turns and reversals in smooth air and during a cruise in rough air which was just short of squall proportions.

Memorandum presenting an investigation conducted in the Lewis 8- by 6-foot supersonic wind tunnel to determine the force and pressure characteristics of an all-external compression inlet with a conical spike and a supersonic cowl lip. Measurements of lift, drag, pitching moment, and internal and external pressures were made at free-stream Mach numbers of 1.59, 1.79, and 1.99 for a range of mass-flow ratios and angles of attack to 10 degrees.

A wind-tunnel investigation of the Boeing XB-47 full-scale empennage was conducted to provide, prior to flight tests, data required on the effectiveness of the elevator and rudder. The XB-47 airplane is a jet-propelled medium bomber having wing and tail surfaces swept back 35 degrees. The investigation included tests of the effectiveness of the elevator with normal straight sides, with a buldged trailing edge, and with a modified hinge-line gap and tests of the effectiveness of the rudder with a normal straight-sided tab and with a bulged tab.

Report presenting cold-flow free-flight tests at zero lift to investigate the possibility of flutter of simulated wing-ram-jet-nacelle configurations which were 1/9-scale models of the wing and ramjet nacelles of the Grumman XSSM-N-6a missile. Testing was performed on three different configurations. Results regarding Mach number, velocity, aerodynamic density, and flutter are provided.

From Summary: "A 1/5-scale, rocket-propelled model of the Convair F-102 configuration was tested in free flight to determine zero-lift drag at Mach numbers up to 1.34 and at Reynolds numbers comparable to those of the full-scale airplane. This large-scale model corresponded to the prototype airplane and had air flow through the duct. Additional zero-lift drag tests involved a series of small equivalent bodies of revolution which were launched by means of a helium gun."

From Summary: "An investigation of a 1/21-scale model of the Chance Vought F7U-3 airplane in the combat-loading condition has been conducted in the Langley 20-foot free-spinning tunnel. The recovery characteristics of the model were determined by use of spin-recovery rockets for the erect and inverted spinning condition. The rockets were so placed as to provide either a yawing or rolling moment about the model center of gravity. Also included in the investigation were tests to determine the effect of simulated engine thrust on the recovery characteristics of the model."

"The principle of operation of a device to control the distribution of fuel to any number of discharge nozzles of a gas-turbine engine is presented. A description of an experimental model of the device and the results of a bench investigation are presented. This device controlled the flow to four discharge nozzles within 2 percent of perfect distribution over a wide range of fuel flow and was unaffected by uneven discharge-nozzle pressures" (p. 1).

Report presenting gas-to-blade heat-transfer coefficients and turbine heat-rejection rates obtained for a wide range of one-spool turbojet engine designs believed to be representative of engines that will employ turbine cooling. The values of heat-transfer coefficient and heat-rejection rates are applicable to both liquid- and air-cooled turbines. Results regarding the heat-transfer coefficients and heat-rejection rates for turbojet engine designs, total heat-rejection rates for various turbojet engine and flight conditions, and heat-transfer coefficients and heat-rejection rates for turboprop engine design are provided.

This study of gaseous explosive reaction has brought out a number of important fundamental characteristics of the explosive reaction indicating that the basal processes of the transformation are much simpler and corresponds more closely to the general laws and principles of ordinary transformations than is usually supposed. The report calls attention to the point that the rate of molecular transformation within the zone was found in all cases to be proportional to pressure, that the transformation within the zone is the result of binary impacts. This result is of unusual interest in the case of the reaction of heavy hydrocarbon fuels and the reaction mechanism proposed by the recent kinetic theory of chain reactions.

Results are presented of an experimental and theoretical investigation of the growth of transverse velocity disturbances in the shear region caused by a flame in a duct. In the theoretical stability analysis, a flow field arising from a flame in a duct was analyzed. The flow was neutrally stable to symmetric disturbances and unstable to antisymmetric ones. In the experimental part of the program disturbances of various frequencies were imposed on a flame stabilized in a duct, and the effects were measured by shadow photography and photomultiplier-probe surveys.

From Summary: "Heating requirements for satisfactory cyclic de-icing over a wide range of icing and operating conditions have been determined for a gas-heated, 36 degree swept airfoil of 6.9-foot chord with a partial-span leading-edge slat. Comparisons of heating requirements and effectiveness were made between the slatted and unslatted portions of the airfoil. Studies were also made comparing cyclic de-icing with continuous anti-icing, and cyclic de-icing systems with and without leading-edge ice-free parting strips. De-icing heat requirements were approximately the same with either heated or unheated parting strips because of the aerodynamic effects of the 36 degrees sweep angle and the spanwise saw-tooth profile of leading-edge glaze-ice deposits. Cyclic de-icing heat-source requirements were found to be one-fourth or less of the heat requirements for complete anti-icing. The primary factors that affected the performance of the cyclic de-icing heating system were ambient air temperature, heat distribution, and thermal lag."

Report presents the results of an investigation to determine the equilibrium temperature and heat-transfer coefficients for transverse cylinders in a high-speed stream of rarefied gas measured over a range of Knudsen numbers (ratio of molecular-mean-free path to cylinder diameter) from 0.025 to 11.8 and for Mach numbers from 2.0 to 3.3. The range of free-stream Reynolds numbers was from 0.28 to 203. The models tested were 0.0010-, 0.0050-, 0.030-, 0.051-, 0.080-, and 0.126-inch -diameter cylinders held normal to the stream.

Memorandum presenting wind-tunnel tests conducted to determine the high-speed stability and control characteristics of an airplane model with a sweptback wing and tail. The aerodynamic coefficients and the longitudinal- and lateral-control characteristics of the plain wing-fuselage-tail combination are included. Results regarding lift, drag, and pitching moment, longitudinal characteristics, lateral characteristics, wing leading-edge slate, and fuselage-side dive brakes are provided.

Report discussing an investigation of two sweptback wings of different plan form in order to determine the effects of adding a triangular area to the inboard section of a conventional sweptback wing as a way to create a wing with two stages of sweepback. Lift, drag, and pitching-moment characteristics are provided for a range of Mach numbers.

"This report contains the results of tests of a 1/3-scale model of the Lockheed YP-90A "Shooting Star" airplane and a comparison of drag, maximum lift coefficient, and elevator angle required for level flight as measured in the wind tunnel and in flight. Included in the report are the general aerodynamic characteristics of the model and of two types of dive-recovery flaps, one at several positions along the chord on the lower surface of the wing and the other on the lower surface of the fuselage. The results show good agreement between the flight and wind-tunnel measurements at all Mach numbers" (p. 1).

Memorandum presenting an investigation of the hydrodynamic characteristics of a six-engine hull-type seaplane designed for supersonic flight. The configuration had four engines mounted on top of the wing in individual nacelles and two engines mounted on top of the aft section of the hull with their inlets located on top of the wing.

Report presenting an investigation of the hydrodynamic characteristics of a six-engine hull-type seaplane designed for supersonic flight. Four engines were mounted on top of the wing in individual nacelles and two engines were mounted on the top of the aft section of the hull. Results regarding the acceleration, gross-load-total-resistance ratio, stabilizer deflections, and reaction of the inlets to water are presented.

Report presents the results of a study of combustion in a spark-ignition engine given in NACA Technical Reports 704 and 727. The present investigation was made with the NACA high-speed motion-picture camera, operating at 40,000 photographs a second, and with a cathode-ray oscillograph operating on a piezoelectric pick-up in the combustion chamber. Photographs are presented showing that the origin of knock is not necessarily in the end gas. The data obtained indicates that knock takes place only in a part of the cylinder charge which has been previously ignited either by autoignition or by the passage of the flame fronts but which has not burned to completion. Mottled regions in the high-speed Schlieren photographs are demonstrated to represent combustion regions.

Memorandum presenting an analytical method developed that gives two independent means of obtaining the total-temperature ratio across a ram jet or across a turbojet tail-pipe burner without direct measurement of the final gas temperature. Experimental verification of the analysis has been obtained with a 20-inch ram jet over a wide range of operating conditions.

Report presenting an investigation in the transonic tunnel to determine the influence of a change in body shape on the effects of twist and camber. Twist and camber were found to greatly improve the maximum lift-drag ratios of the wing-modified fuselage configuration through the speed range, but reduced the ratios for a wing-curved fuselage configuration. Results regarding basic aerodynamic data are also provided.

Report presenting an investigation with the dual purpose of determining the effect of two bodies in various positions, symmetrically located from the plane of symmetry, on the aerodynamic characteristics of a wing-fuselage model and determining the aerodynamic loads on one of two bodies. Some of the most significant effects were obtained for the direct-mounted tip bodies which gave a large increase in the lift-curve slope of the basic model.

Report presenting an investigation to determine the aerodynamic characteristics in sideslip of a triangular wing of aspect ratio 2.04 in combination with a body of fineness ratio 12.5 and a vertical tail surface. The body combined with the triangular plan-form wing caused no significant changes in the lift characteristics of the wing and only a 1-percent decrease in the static margin. Results regarding the longitudinal characteristics, lateral and directional characteristics, and estimation of tail and rudder effectiveness are provided.