Comparison of free-molecule-flow theory with the results of wind-tunnel tests performed to determine the drag and temperature-rise characteristics of a transverse circular cylinder.

Instead of solving the nonlinear differential equation which governs the compressible flow, an approximate method of solution by means of the variational method is used. The general problem of steady irrotational flow past an arbitrary body is formulated. Two examples were carried out, namely, the flow past a circular cylinder and the flow past a thin curved surface. The variational method yields results of velocity and pressure distributions which compare excellently with those found by existing methods. These results indicate that the variational method will yield good approximate solution for flow past both thick and thin bodies at both high and low Mach numbers.

Heat-transfer investigation conducted on cooled as well heated impulse-type turbine blades in a static cascade to determine the effect of direction of heat flow on convective heat-transfer coefficients.

An experimental investigation of a highly loaded two-stage turbine with a downstream stator revealed that (1) at equivalent design work and speed the brake internal efficiency was 0.81 and the maximum efficiency obtained was 0.85, and (2) the downstream stator left very little energy in the form of tangential velocity in the gas and, in general, performed well with 0.78 recovery being obtained at equivalent design work and speed.

Report describing an investigation conducted with a single combustor from a J47 turbojet engine using weathered aviation gasoline and several spark-plug modifications to determine altitude ignition, acceleration, and steady-state operating characteristics. Results regarding ignition, acceleration, altitude operational limits, combustion efficiency, and pressure loss are provided.

Report presenting testing of a 6.5-inch-diameter ramjet engine with a design Mach number of 2.13 and a short-flame-length combustor. Performance characteristics are provided for a range of Mach numbers and fuel-air ratios. Results regarding thrust, drag tests, static pressure, and combustion efficiencies are provided.

Report presenting a determination of control-surface flutter derivatives for a sinusoidally oscillating control surface mounted on a two-dimensional fixed airfoil for a range of reduced frequency. Results regarding subsonic results and transonic results are provided.

The durability of five different structural or cooling configurations or combination of both of air-cooled blades made of noncritical materials was investigated in a modified turbojet engine. The greater part of the investigation was conducted at an engine speed of 11,500 rpm, a turbine-inlet temperature of approximately 1670 degrees F, and a cooling-air to combustion gas flow ratio per blade of 0.05. The results of the investigation indicated that air-cooled blades made of noncritical metals can be operated for extended periods of time in engines at current inlet temperatures; however, before these blades are considered completely satisfactory for gas-turbine application, some means such as coatings is required to inhibit the oxidation of the blades.

Report discussing the spin and spin-recovery characteristics of a model of the Douglas X-3 airplane in a free-spinning tunnel. The pitching-moment characteristics of the scale model were not found to be typical of larger scale results. The results indicated that at an altitude range of 15,000 to 30,000 feet, the X-3 will not exhibit unusual trim tendencies and it will be difficult to obtain erect spins unless the ailerons are full with the spin.

Report presenting a preliminary investigation in a duct system with fully developed pipe flow to determine the effectiveness of vortex generators in improving the performance of a 21-inch-diameter straight-outer-wall annual diffuser with an overall equivalent conical expansion angle of 15 degrees and a and 1.9:1 area ratio. The vortex generators used were rectangular noncambered airfoils that varied in chord, span, angle of attack, number, and location. Results regarding a diffuser with no vortex generators, a diffuser with vortex generators on inner body, and a comparison of the diffuser performance with and without vortex generators are provided.

Report presenting a measurement of the magnitudes of pressure fluctuations in the turbulent wakes of restrained two-dimensional airfoils at transonic speeds. Measurements indicated that the total-pressure fluctuations occurred only within the wake. Results regarding the magnitude of pressure oscillation in the wake, effect stream-angle fluctuations in the wake, and frequencies of pressure oscillation on the wing and in the wake are provided.

Memorandum presenting the observed gust experience of an airplane and information on the associated meteorological conditions to obtain a simple empirical relation for estimating the intensity of turbulence in the earth's friction layer. The relation does not discriminate between differences in turbulence intensity resulting from variations of flight altitude or diurnal variations of turbulence.

Report presenting an investigation of two large-scale, semispan, wing-fuselage models with the 0.25-chord line swept back 45 degrees to determine and compare the effects of partial-span, double-slotted flaps on the characteristics of a 45 degree sweptback wing with and without camber and twist. A secondary investigation was also conducted to determine the effects of various full-span, leading-edge modifications on the characteristics of the models with and without the flaps.

Report presenting an investigation of a 1/21-scale model of the Chance Vought F7U-3 airplane in the 20-foot free-spinning tunnel to determine the erect and inverted spin and recovery characteristics for combat loading in the clean condition. Results regarding erect spins, inverted spins, spin-recovery parachutes, and recommended recovery technique are provided.

A 24-inch-diameter shock-in-rotor supersonic compressor rotor designed to obtain simple radial equilibrium behind the normal shock obtained a maximum total-pressure ratio of 2.02 at an adiabatic efficiency of 0.74 and a weight flow of 61.5 pounds per second at design speed. Although this design method considerably reduced the radial forces, there was no apparent reduction in the radial forces, there was no apparent reduction in the radial redistribution of mass flow over the radial redistribution obtained in the previous investigation. This investigation indicates that the shock boundary-layer interaction pattern and associated effects have a very large effect on the radial redistribution of mass flow through the shock-in-rotor supersonic-compressor rotor.

Report discussing the results of an investigation to determine the stability and control characteristics of a flying model of an unswept-wing vertically rising airplane. Information about stabilization, maneuverability, rolling motions, and landings is provided.

Report presenting results and analysis of the longitudinal stability and control characteristics of the Douglas D-558-I airplane at a range of altitudes and Mach numbers. It was found that large and rapid changes in elevator deflection and force were required for balance at Mach numbers above 0.84 due to a sharp decrease in elevator-stabilizer effectiveness.

Report presenting an investigation of a model of an ejectable pilot seat combination with and without stabilizing fins. The main purpose of the investigation was to determine the static aerodynamic characteristics and effectiveness of various stabilizing fins at a high subsonic Mach number. Results regarding stability and force characteristics for the combinations are provided.

Report presenting an investigation of some effects of aileron spanwise location and wing structural rigidity on the rolling power of 0.3-chord plain, flap-type ailerons on a wing with a taper ratio of 0.25, an aspect ratio of 3.5, and swept back 63 degrees at the leading edge. Results regarding the aeroelastic effects and aileron span and spanwise location are provided.

Memorandum presenting results obtained during Air Force testing of the Northrop X-4 airplane. Information is included on the stalling characteristics, the static and dynamic longitudinal and lateral-stability characteristics, and the lateral-control characteristics.

Elevated-temperature properties of titanium carbide base ceramals containing nickel or iron were determined in oxidation, modulus of rupture, tensile strength, and thermal-shock resistance. These materials followed the general growth law and exhibited two stages in oxidation. The following tensile strengths were found at 2000 degrees F: 13.3 weight percent nickel, 16, 150 pounds per square inch; 11.8 weight percent iron, 12,500 pounds per square inch; unalloyed titanium carbide, 16,450 pounds per square inch. Nickel or iron additions to titanium carbide improved the thermal-shock resistance, nickel more. The path of fracture in tensile and thermal-shock specimens was found to progress approximately 50 percent intergranularly and 50 percent transgranularly.

Fundamental flame velocities are reported for pentane air, ethylene-air, and propylene-air mixtures for the concentration range 60 to 130 percent of stoichiometric. A form of the Tanford and Pease equation, which includes a small constant velocity term independent of diffusion, will predict the observed changes in flame velocity.

Memorandum presenting the results of an investigation to establish rupture and total-deformation strengths at 1200 and 1800 degrees Fahrenheit for typical commercial treatments of low-carbon N-155 alloy. The results show that there were large differences in strength between the heats of bar stock at temperatures above 1200 degrees Fahrenheit except when a 2200 degree Fahrenheit solution treatment was used.

In an experimental investigation of local recovery factors for a blade having a pressure distribution similar to that of a typical reaction-type turbine blade, it a was found that the recovery factors were essentially independent of Mach number, Reynolds number, pressure gradient, and position on the blade surface except for regions where the boundary layer was probably in the transition range from laminar to turbulent. The recommended value of local subsonic recovery factor for use in calculating the effective gas temperature for gas turbine blades was 0.89.