Memorandum presenting an investigation of the change in spanwise load distribution at high Mach numbers, which is of considerable interest because lateral movement of the center of lift affects the trim, stability, and structural factors of safety of an airplane. The data is not sufficient to permit isolation of the effects of changes in wing configuration, but in a majority of cases, the tests reveal a tendency for the center of lift to shift outboard with increasing Mach number.

Report presenting wind-tunnel testing of five large-scale tapered wings with a variety of angles of sweep to determine the effects of scale and sweep on the damping-in-roll parameter. Rolling moment and pressure distribution were measured for each plain wing while in steady roll for a range of angles of attack. Results regarding the effects of sweep at zero lift, effects of lift, and aurorotational characteristics are provided.

Report presenting an investigation of a German Jumo 004 combustor under conditions simulating zero-ram operation of a 24C jet-propulsion engine over ranges of altitude and engine speed to obtain the altitude operating limits and characteristics. Combustion efficiency, outlet-temperature distribution, and total-pressure drop were determined. The performance of the combustor with 62-octane and JP-1 fuels were also compared.

"Following a law of stress adopted in the Navier-Stokes equations, the configuration of the viscous flow in planes at right angles to the axis of an infinite cylinder is found to be independent of the axial motion of the cylinder. In the limiting case of a yawed or swept wing of very high aspect ratio, certain boundary-layer and separation phenomena are thus determined independently by the crosswise component of velocity. It follows that the effect of sweepback is to increase the area of stable laminar flow and to decrease the lift coefficient at which flow separation occurs" (p. 487).

As part of a study of the effects of fuel composition on the combustor performance of a turbojet engine, an investigation was made in a single I-16 combustor with the standard I-16 injection nozzle, supplied by the engine manufacturer, at simulated altitude conditions. The 10 fuels investigated included hydrocarbons of the paraffin olefin, naphthene, and aromatic classes having a boiling range from 113 degrees to 655 degrees F. They were hot-acid octane, diisobutylene, methylcyclohexane, benzene, xylene, 62-octane gasoline, kerosene, solvent 2, and Diesel fuel oil. The fuels were tested at combustor conditions simulating I-16 turbojet operation at an altitude of 45,000 feet and at a rotor speed of 12,200 rpm. At these conditions the combustor-inlet air temperature, static pressure, and velocity were 60 degrees F., 12.3 inches of mercury absolute, and 112 feet per second respectively, and were held approximately constant for the investigation. The reproducibility of the data is shown by check runs taken each day during the investigation. The combustion in the exhaust elbow was visually observed for each fuel investigated.

Report discussing about 400 references pertaining to the hydrodynamic design of seaplanes have been compiled, and the information is presented in the form of abstracts classified under six main headings.

Tests have been conducted in the Langley high speed 7- by 10-foot tunnel over a Mach number range from 0.40 to 0.91 to determine the stability and control characteristics of an 0,08-scale model of the Chance Vought XF7U-1 airplane. The longitudinal-control characteristics of the complete model are presented in the present report with a limited analysis of the results.

"Flight tests to determine propeller performance have been made of a Curtiss No. 838-102-18 three-blade propeller having trailing-edge extensions on a Republic P-47D-28 airplane in climb and high speed. These tests are a part of a general propeller flight-test program at the Langley Laboratory of the National Advisory Committee for Aeronautics. Results of climb tests indicate that when power is changed from approximately 1475 horsepower at 2550 rpm (roughly normal power) to 2400 horsepower at 2700 rpm (approximately military power) there is a loss in propeller efficiency of 3 percent at an altitude of 7000 feet, and 4 percent at 21,000 feet" (p. 1).

"Tests were made with a C-54 airplane in which airline pilots made several blind approaches to determine whether any special flying techniques were used in blind landings and whether any special handling-qualities requirements would have to be formulated because of such special techniques. It was found that the airplane was flown at all times in the normal manner; that is, all turns were banked turns that were nearly coordinated by use of the rudder so that the sideslip was held close to zero. The pilot expended considerable physical work in continually moving the controls but this wake was due in part to the large friction in the three control systems" (p. 1).

Tests have been conducted in the Langley high-speed 7- by 10-foot tunnel over a Mach number range from 0.40 to 0.91 to determine the stability and control characteristics of an 0.08-scale model of the Chance Vought XF7U-1 airplane. The basic lateral stability characteristics of the complete model with undeflected control surfaces are presented in the present report with a very limited analysis of the results.

This report contains the flight-test results of the longitudinal-stability and -control phase of a general flying qualities investigation of the Lockheed P-80A airplane (Army No. 44-85099). The tests were conducted at indicated airspeeds up to 530 miles per hour (0.76 Mach number) at low altitude and up to 350 miles per hour (0.82) Mach number) at high altitude. These tests showed that the flying qualities of the airplane were in accordance with the requirements of the Army Air Forces Stability and Control Specification except for excessive elevator control forces in maneuvering flight and the inadequacy of the longitudinal trimming control at low airspeeds.

The performance at inlet pressure of 21 inches mercury absolute and inlet temperature of 538 R for the 10-stage axial-flow X24C-2 compressor from the X24C-2 turbojet engine was investigated. the peak adiabatic temperature-rise efficiency for a given speed generally occurred at values of pressure coefficient fairly close to 0.35.For this compressor, the efficiency data at various speeds could be correlated on two converging curves by the use of a polytropic loss factor derived.

It is known that the compressibility shocks accompanying local or total supersonic flows lead to pronounced flow separations which result in unusually high energy losses on airplane wings, vanes, and in diffusers. These phenomena were investigated experimentally and theoretically.

In the vicinity of a body in a wind tunnel the displacement effect of the wake, due to the finite dimensions of the stream, produces a pressure gradient which evokes a change of drag. In incompressible flow this change of drag is so small, in general, that one does not have to take it into account in wind-tunnel measurements; however, in compressible flow it beoomes considerably larger, so that a correction factor is necessary for measured values. Correction factors for a closed tunnel and an open jet with circular cross sections are calculated and compared with the drag - corrections already bown for high-speed tunnnels.

"On the basis of the investigations so far completed on the behavior of PTL power plants under various operating conditions, in which the influence of the propeller characteristics is of considerable importance, the most important aspects of a control system for turbine-propeller jet power plants are deduced. A simple possible means for its concrete realization, which is also applicable to TL [NACA comment: TL, jet] power plants, is presented by means of examples. A control device of this kind is now being developed" (p. 1).

"With an approach of the velocity of flight of a ship to the velocity of sound, there occurs a considerable increase of the drag. The reason for this must be found in the boundary layer separation caused by formation of shock waves. It will be endeavored to reduce the drag increase by suction of the boundary layer. Experimental results showed that drag increase may be considerably reduced by this method" (p. 1).

"This report presents a treatment of the effects of the tunnel walls on the flow velocity and direction in a compressible medium at subsonic speed by an approximate method. Calculations are given for the rotationally symmetric and two-dimensional problems of the flow past bodies, as well for the downwash effect in the tunnel with circular cross section" (p. 1).

"As one phase of a comprehensive canopy load investigation, conventional front and rear sliding canopies which are typified by installation on the SB2C-4E airplane, were tested in the Langley full-scale tunnel to determine the pressure distributions and the aerodynamic loads on the canopies. A preliminary analysis of the results of these tests is presented in this report. Plots are presented that show the distribution of pressure at four longitudinal stations through each canopy for a range of conditions selected to determine the effects of varying canopy position, yaw, lift coefficient, and power" (p. 1).

Studies of the stalling characteristics show that the stall begins at the tip and moves inboard with increasing angle of attack at positive sweep; the sta11 begins at root and moves outboard at negative sweep (sweepforward). At +/-45 deg sweep the stall was less sharply defined than at the lower angles of sweep. No effect of Mach number on the.flow patterns as indicated by tufts was found in the speed range of these tests which extended to a Mach number of 0.55.

"A preliminary evaluation of the spin and recovery characteristics of the XF3D-1 airplane has been made, based primarily on the results of the free-spinning tunnel tests of a model which closely simulated the XF3D-1 in tail design, tail length, and mass loading. Estimates have been made of the rudder-pedal force that may be encountered in effecting recovery from a spin and of the spin recovery parachute requirements of the airplane for demonstration spins. The method of bail-out which should be used if it becomes necessary for the crew to abandon the airplane during a spin is indicated" (p. 1).

Spin tests have been performed in the Langley 20-foot free-spinning tunnel on a 1/35-scale model of the Douglas XB-43 airplane. The spin and recovery characteristics were determined for several loading conditions of the airplane. The effects of installing a dorsal fin and of installing a ventral fin were investigated. Emergency escape of the crew was simulated and the stick and rudder pedal forces necessary to effect recoveries on the airplane were determined.

"At the request of the Air Materiel Command, Army Air Forces, an investigation is being conducted at the NACA Cleveland laboratory to determine the performance characteristics of the XJ-41-V turbojet-engine compressor. The static-pressure variation in the direction of flow through the compressor was presented in reference 1 for an equivalent speed of 8000 rpm. An analysis of these pressure indicated that the maximum-flow limitation of the compressor was caused by separation, which reduced the effective flow area at the vaned-collector entrance" (p. 1).

"This is the second of a series of six reports dealing with three- and six-component measurements on the tapering series at small aspect ratio. The present report concerns the trapezoidal wing with fuselage" (p. 1).

"In the search for an energy source independent of air for the propulsion of underwater craft, attention was early concentrated on T-substance. It was possible to convince the OKM [NACA comment: Navy High Command] very quickly of the importance of this material. In 1934, the first experiments were undertaken" (p. 1).