"This paper presents a mathematical discussion of the laminar boundary layer, which was developed with a view of facilitating the investigation of those boundary layers in particular for which the phenomenon of separation occurs. The treatment starts with a slight modification of the form of the boundary layer equation first published by Von Mises. Two approximate solutions of this equation are found, one of which is exact at the outer edge of the boundary layer while the other is exact at the wall" (p. 541).

Report presents the results of tests made to determine the drag of closed-cockpit and transport-type windshields. The tests were made at speeds corresponding to a Mach number range of approximately 0.25 to 0.58 in the NACA 8-foot high-speed wind tunnel. This speed range corresponds to a test Reynolds number range of 2,510,000 to 4,830,000 based on the mean aerodynamic chord of the full-span model (17.29 in.). The shapes of the windshield proper, the hood, and the tail fairing were systematically varied to include common types and refined design.

Wind tunnel tests haved been made at low speed to various small-scale models of swept-back, swept-forward, and yawed wings. Tests covered changes in aspect ratio, taper ratio, and tip shape.

"A general integral form of the boundary-layer equation, valid for either laminar or turbulent incompressible boundary-layer flow, is derived. By using the experimental finding that all velocity profiles of the turbulent boundary layer form essentially a single-parameter family, the general equation is changed to an equation for the space rate of change of the velocity-profile shape parameter. The lack of precise knowledge concerning the surface shear and the distribution of the shearing stress across turbulent boundary layers prevented the attainment of a reliable method for calculating the behavior of turbulent boundary layers" (p. 1067).

From Summary: "An investigation is being conducted to determine the performance of the 12-stage axial-flow compressor of the XT-46 turbine-propeller engine. This compressor was designed to produce a pressure ratio of 9 at an adiabatic efficiency of 0.86. The design pressure ratios per stage were considerably greater than any employed in current aircraft gas-turbine engines using this type of compressor. The compressor performance was evaluated at two stations."

The stage performance of the X24C-2 axial-flow compressor, as determined from radial distributions of total pressure and temperature measured in each stator-blade row, was investigated at design speed at weight flows corresponding to those of the maximum weight-flow, peak-efficiency, and surge points. The stage performance is presented as the measured stage pressure ratios and the calculated velocity diagrams. The average stage total-pressure ratio at the surge operating condition in the first stage was approximately 1.10; in the second stage, approximately 1.13; and in all remaining stages, approximately constant at a value of 1.15. Regions of inefficient flow are discussed so as to point out the reasons for the poor flow conditions.

Part I of this report covers critical-mass determinations for pseudospheres of oralloy hydride composition (approximating UH3) in 8"-thick Tu and Ni tampers and in the Tu tamper with Ni liner. The critical mass of a hydride cube in the thick Tu also is given. Data on weight and dimensional changes of hydride pieces during the period of use are included. In Part II are presented the results of Rossi time-scale measurements on the hydride assemblies. Values of alpha at delayed critical and its variation with mass in the neighborhood of delayed critical are given. Measurements on the activation of various detectors within the hydride assemblies are described in Part III. Results as a function of radial position are given for Au, for Au shielded by Au and by Cd, for S and for fission catchers with U235 and U238. Reactivity changes resulting from the introduction of foreign materials into the hydride assemblies are discussed in Part IV. Apparent regularities with respect to Z and qualitative interpretations of variations with radius are pointed out. From data for various radial positions, changes in critical mass corresponding to small changes in composition and density are computed.

The effects of fineness ratio (14.2 and 12.2) and boattailing on aerodynamic characteristics of four bodies of revolution at Mach numbers from 1.50 to 1.99 within a range of angles of attack from 0 degrees 10 degrees at an approximate Reynolds number of 35x10(superscript)6 based on body length were investigated. A comparison of experimental data with available theory is included. At zero angle of attack, fineness ratio has no appreciable effect on model characteristics while boattailing and boattail convergence significantly affect fore drag and base drag.

Introduction: "A limited test progress for the evaluation of type 347 stainless steel pipe which failed the Huey (?) test has been complicated. The tests reported herein were similar in scope to those reported in document number EW-20765, Corrosion Tests on Stainless Steel Plate, Waste Metal Recovery Process. Material tested was obtained on order number EWC-8553, National Tube Company Heat #3X9741. Lots 9789 and 9790."

This report provides fundamental equations and maps to calculate neutron behavior in a spherical cavity.

Report presenting the effects of variation of trailing-edge angle and trailing-edge thickness on the lift characteristics of a 10-percent-chord thick symmetrical NACA airfoil section with a 25-percent-chord plain flap as determined from wind tunnel tests over a range of Mach numbers. Results regarding the section lift and pitching-moment characteristics, variation of lift coefficient with flap deflection, schileren photographs, and effect of angle of attack are provided.

The factors affecting the performance of ejector suitable for aircraft cooling are investigated theoretically and experimentally. The investigation covers a range of shroud-to-nozzle diameter ratios from 1.1 to 1.6, of shroud lengths from 0.2 to 2.28 nozzle diameters, of secondary-to-primary weight-flow ratios from 0 to 0.12, and of ambient-to-nozzle pressure ratios from 0.7 to 0.06. The results of a simplified theoretical analysis based on each type of flow are in good agreement with those experimentally obtained.

Memorandum presenting an investigation conducted at Mach number 1.90 to determine if Helmholtz resonators can be used in a typical ramjet configuration to damp pressure pulsations associated with shock oscillations. Only one of the two resonators was found to effectively damp pressure pulsations that occurred when the ram jet was operating in the vicinity of its peak pressure recovery. Neither was found to effectively damp pressure pulsations over the entire range of ramjet operating conditions.

This report presents "Results of a critical literature review on pressure drop due to isothermal turbulent flow in noncircular geometries" (introduction).

Memorandum presenting an investigation of a 2:1 area ratio conical diffuser of length equal to the inlet diameter with separate variation of several basic parameters for simple nontwisted counterrotating rectangular vortex generators over a considerable range of inlet-boundary-layer thickness. The maximum values of static-pressure rise were determined for angle of attack, spacing, aspect ratio, and span-to-inlet boundary-layer thickness. Results for the diffuser with no vortex generators, effects of selectable vortex-generator variables, effect of inlet-boundary-layer displacement thickness, and use of the maximum indicated vortex-generator design values are provided.

In solving the two-group, two medium equations for neutron diffusion in a finite one-dimensional multiplying system, the equations in each medium can be solved, and the boundary conditions used to determine the coefficients. This direct attack involves eight simultaneous linear equations. The opportunities for error are numerous, and, frequently, a great deal of accuracy is lost by subtractions; hence, it is felt that an alternative method of solution, in which numerical errors would be less likely to occur, would be desirable.

Memorandum presenting flutter tests of untapered, low-aspect-ratio, sweptback-wing models and of triangular-wing models of solid magnesium, aluminum, and steel construction in an intermittent supersonic wind tunnel at Mach number 1.3. Flutter data are presented for 31 sweptback wing models and 24 triangular-wing models. The flutter boundaries are indicated on plots of streamwise thickness ratio and aspect ratio.

From abstract: General expressions relating product stream compositions to the operating conditions of countercurrent stagewise extraction columns are derived by an inductive method, and the conversion of the general equations to ones involving the particular case of constant extraction factor is demonstrated. For comparison, the derivation of simple column equations by the use of finite difference equations is included.

The decontamination of portable instruments by the Calibration Unit has presented a problem for some time. Most of the contamination is in the 100 to 500 c/m or 500 to 2000 d/m range. Higher levels of contamination can usually be reduced to these levels, but further cleaning proves quite difficult

Report presenting combustion-efficiency data for a 5-inch ramjet-type combustor using a simple V-gutter flame holder. Data for homogeneous fuel-air mixtures of 14 pure fuels and 3 fuel blends cover a range of inlet static pressures and velocities.

Motion pictures at camera speeds up to 3000 frames per second were taken of the combustion of liquid oxygen and gasoline in a 100-pound thrust rocket engine. The effect of seven methods of propellant injection on the uniformity of combustion was investigated. The flame front was generally found to extend to the injector faces and all the injection systems showed considerable nonuniformity of combustion. Pressure vibration records indicated combustion vibrations that corresponded to resonant-chamber frequencies.

Report presenting an investigation to determine the static longitudinal stability and control characteristics of a tailless wing-body combination with a pointed wing of aspect ratio 2 and trailing-edge elevons. The effectiveness of inset tabs in reducing the elevon hinge moment was also determined. Results regarding the lift, drag, pitching moment, elevon hinge moment, tab hinge moment, elevon load, and center of pressure of elevon load are provided.

Report presenting a blade-vibration and rotating-stall survey on a modified version of a production turbojet engine with a 13-stage axial-flow compressor with a design pressure ratio of 7 and an air flow of 120 pounds per second. This testing was carried out due to the problems rotating stall can cause with both experimental and production axial-flow compressors. Results regarding the rotating-stall patterns, rotor-blade vibrations, radial and axial strength of rotating stall, and maximum rotating-stall speed are provided.

Report presenting the effect of adjusting compressor inlet-guide-vane angle on overall compressor performance, stage performance, engine thrust, and specific fuel consumption in a 7000-pound-thrust turbojet engine. Data was obtained at several angles and demonstrated that increasing the guide-vane angle decreased the weight flow at the higher compressor speeds. It had no effect on stage performance other than at the first stage, where it reduced maximum flow, pressure, and temperature coefficients.