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 a flight test of pentaborane fuel in an air-launched ramjet engine with a design free-stream Mach number of 1.8. The ramjet used in the investigation had a revised spray bar-flameholder configuration. Results regarding propulsive thrust and propulsive-thrust specific fuel consumption are provided.

"An approximate method for the development of flow and thermal boundary layers in the laminar region on cylinders with arbitrary cross section and transpiration-cooled walls is obtained by the use of Karman's integrated momentum equation and an analogous heat-flow equation. Incompressible flow with constant property values throughout the boundary layer is assumed. The velocity and temperature profiles within the boundary layer are approximated by expressions composed of trigonometric functions" (p. 339).

"The external wave drag of bodies of revolution moving at supersonic speeds can be expressed either in terms of the geometry of the body, or in terms of the body-simulating axial source distribution. For purposes of deriving optimum bodies under various given conditions, it is found that the second of the methods mentioned is the more tractable. By use of a quasi-cylindrical theory, that is, the boundary conditions are applied on the surface of a cylinder rather than on the body itself, the variational problems of the optimum bodies having prescribed volume or caliber are solved" (p. 131).

Report presenting testing of gaseous hydrogen fuel burned in a connected-pipe combustor with a cross section equal to 35 degrees sector of a 28-inch diameter. Eleven shrouded fuel-injector configurations were used to obtain combustion data at specified high-altitude ramjet combustor conditions. Results regarding shroud air blockage, mixing tabs, higher equivalence-ratio burning, injector size, flameholder durability, burner length, combustion efficiency, and temperature profile are provided.

"Several methods of cyclic de-icing of a gas-heated airfoil were investigated to determine ice-removal characteristics and heating requirements. The cyclic de-icing system with a spanwise ice-free parting strip in the stagnation region and a constant-temperature gas-supply duct gave the quickest and most reliable ice removal. Heating requirements for the several methods of cyclic de-icing are compared, and the savings over continuous ice prevention are shown. Data are presented to show the relation of surface temperature, rate of surface heating, and heating time to the removal of ice" (p. 1).

Memorandum presenting performance calculations for hypothetical hydrogen expansion engines in which only a small part of the engine airflow passes through the heat exchanger. The flight path was representative of those for long-range, high-speed aircraft. Results regarding the off-design operation and engine performance are provided.

Report described five short multielement turbojet combustors for use with hydrogen fuel and their performance. Information about the combustion efficiency, temperature profiles, pressure loss, durability, and comparison with previous short combustors is included.

Flight investigation of pentaborane fuel in 9.75- inch-diameter ramjet engine with downstream fuel injection.

Double-ramp side inlet with combinations of fuselage, ramp, and throat boundary layer removal.

"A turbojet engine incorporating a conventional-type afterburner was operated over a range of afterburner pressure levels from 1000 to 2000 pounds per square foot absolute and a range of engine-inlet water-air ratios from 0 to 0.08. At each pressure level and water-air ratio, the afterburner fuel flow was varied from lean blowout to maximum burner-outlet temperature" (p. 1).

"In the present paper attention is directed to the aerodynamic parameters, the so-called stability derivatives, that affect the lateral behavior of airplanes and missiles. The discussion is centered on three important quantities, the effective-dihedral derivative, the directional-stability derivative, and the damping-in-roll derivative. These quantities are considered for a large angle-of-attack range at subsonic speeds" (p. 1).

The third report in a series of five that present unanalyzed pressure data obtained in tests of five full-scale propellers with NACA 16-series blade sections. Pressure distributions on the blade sections were measured under operating conditions to determine the aerodynamic characteristics of each blade section. This particular report presents information for nine radial stations of the NACA 10-(3)(090)-03 propeller.

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.

Report presenting wind-tunnel measurements of the pressure distribution on a semispan wing with a modified NACA 0012-64 airfoil section and a plain trailing-edge aileron with the wing unswept and also swept back 45 degrees. Adverse changes in loading and loss of aileron effectiveness developed with the wing unswept as the Mach number was increased, but did not occur at Mach numbers up to 0.925 with the swept wing. Results regarding section characteristics and effect of fences are provided.

From Summary: "Ogive-cylindrical fuselages of fineness ratio 11 have been flight-tested with and without stabilizing fins. Base-pressure measurements over a range of free-stream Mach numbers from 0.7 to 1.3 indicated that the fins reduced the base drag. A rear-support sting similar to those used in wind tunnels was tested with one of the fuselages and found to reduce base suction by 40 percent at subsonic speeds, but to have little effect at Mach numbers above 1.15."

Report presenting an investigation of the performance of several conical cooling-air ejectors at primary jet pressure ratios from 1 to 10, secondary shroud pressure ratios from 0.60 to 4.00, and primary to secondary temperature ratio of 1.0. The investigation centered on primary air flow, secondary air flow, primary nozzle thrust, and gross ejector thrust.

Report presenting testing of JP-3 fuel and a slurry of 60 percent powdered magnesium in JP-3 fuel in a small-scale afterburner in the presence of large quantities of water vapor. The results indicated that total temperature, combustion efficiency, and air specific impulse were improved when the magnesium slurry was used, but the improvements came at the cost of increased liquid consumption.

Report presenting an analytical and experimental evaluation of diborane as a ram-jet fuel, which seems to offer greater flight range, thrust, and combustion stability beyond what is attainable with petroleum fuels. Results regarding ideal thrust, stability, deposits, and flight range with several blends of diborane 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.

Report presenting an investigation in the free-flight tunnel to determine the dynamic longitudinal stability and control characteristics of tandem-coupled bomber-fighter airplane models. Results regarding bomber alone tests, freely coupled combination, and rigidly coupled combination are provided.

An investigation was conducted to study the mechanisms of blade failure of forged Hastelloy B and cast Stellite 21. The blades were mounted in a 16-25-6 alloy rotor and subjected to 20-minute cycles consisting of 15 minutes at rated speed and approximately 5 minutes at idle. The first failures of the Hastelloy B and Stellite 21 blades were probably the result of excessive vibratory stresses and occurred after 14.25 and 16.75 hours, respectively. After 28.75 hours of operation, all but 3 of the original 25 Hastelloy B blades had either failed or contained stress-rupture-type cracks and four of the original 27 Stellite 21 blades contained stress-rupture-type cracks.

Memorandum presenting the results of an experimental investigation to determine the effects of Reynolds number on the flow characteristics over a delta wing at a Mach number 2.41. The wing streamwise airfoil sections are based on the NACA 00-thickness series with the maximum thickness varying from 4 percent at the root section to 6.24 percent at the 90-percent spanwise station. Results regarding force data and pressure-distribution data are provided.

Report presenting the low-speed lateral control and hinge-moment characteristics of a 20-percent-chord plain aileron on a 47.7 degree sweptback wing of aspect ratio 5.1 in the pressure tunnel. The tests were performed with and without high-lift and stall-control devices. Results regarding the basic aileron data, rolling-moment characteristics, and pitching-moment characteristics are also provided.