From Summary: "Tests have been made in Langley tank no. I of a dynamic model of the Consolidated Vultee PB2Y-3 airplane. These tests were made using an alternate hull form, the purpose of which was to reduce the bow spray and eliminate the landing instability which are objectionable features of the production design. The major differences from the PB2Y-3 hull included a deeper step to improve the landing stability , and a lengthened forebody and increased beam to reduce the sway in the propellers and on the flaps. The tests showed that the spray characteristics of the revised hull form were much better than that to the production design. In addition the take-off and landing stability of the model with the alternate hull were satisfactory."

"Three modifications of the auxiliary-stage supercharger for the V-1710-93 engine were designed and tested as part of an investigation to improve the power output and the altitude performance of the engine. A 12-vane diffuser was substituted for the standard 11-vane diffuser, and a vaneless discharge passage and a modified scroll were designed to increase the flow capacity of the supercharger and thereby to increase the performance at the high volume flows required by the engine. With the 12-vane diffuser installed and the carburetor replaced by an adapter, the equivalent volume flow at the peak efficiency point was increased 25 percent at the lowest speed investigated and 9.5 percent at the highest speed" (p. 1).

The NACA 6A-series airfoil sections were designed to eliminate the trailing-edge cusp which is characteristic of the NACA 6-series sections. Theoretical data are presented for NACA 6A-series basic thickness forms having the position of minimum pressure at 30-, 40-, and 50-percent chord and with thickness ratios varying from 6 percent to 15 percent. Also presented are data for a mean line designed to maintain straight sides on the cambered sections. The experimental results of a two dimensional wind tunnel investigation of the aerodynamic characteristics of five NACA 64A-series airfoil sections and two NACA 63A-series airfoil sections are presented. An analysis of these results, which were obtained at Reynolds numbers of 3 x 10(exp 6), 6 x 10(exp 6), and 9 x 10(exp 6), indicates that the section minimum drag and maximum lift characteristics of comparable NACA 6-series and 6A-series airfoil sections are essentially the same. The quarter-chord pitching-moment coefficients and angles of zero lift of NACA 6A-series airfoil sections are slightly more negative than those of corresponding NACA 6-series airfoil sections. The position of the aerodynamic center and the lift-curve slope of smooth NACA 6-series sections. The addition of standard leading-edge roughness causes the lift-curve slope of the newer sections …

Report presenting an investigation of the performance of a turbojet engine equipped with a tail-pipe burner at zero ram over a range of rotor speeds and tail-pipe-burner fuel flows. A thrust augmentation of 40 percent was obtained at zero ram for a tail-pipe-burner fuel-air ratio of 0.043 or a total fuel-air ratio of 0.056. Results with an engine with standard tail pipe, engine with tail pipe burner and no afterburning, and engine with tail pipe burner and afterburning are provided.

From Summary: "Additional tests of a 1/7-size model of the Grumman XJR2F-1 amphibian were made in Langley tank no. 1 to compare the behavior during take-off of the model equipped with split- and slotted-type flaps. The slotted flag had a large effect on locating the forward center-of-gravity limits for stable take-offs. Stable take-offs within the normal operating range of positions of the center of gravity could be made with the split flaps deflected 45 degrees or with the slotted flaps deflected less than 20 degrees."

Report presenting measurements made at transonic speeds by the freely-falling-body method to compare the drag of a rectangular plan-form airfoil of aspect ratio 7.6 with an NACA 65-006 airfoil section. Results regarding the velocity measurements, base-pressure measurements, and airfoil drag measurements are provided.

Tests were conducted in calm water in Langley tank no. 2 and in calm and rough water at an outdoor catapult in order to determine the best way to make a forced landing of an Army A-26 airplane and to determine its probable ditching behavior. These tests were requested by the Air Materiel Command, Army Air Forces, in their letter of March 26, 1943, WEL:AW:50.

The teat results showed that with either of the three tail arrangements, the model usually spun in flat attitudes with oscillations about the lateral and longitudinal axes. In general, full reversal of the rudder pedals did not stop the spinning rotation. To make the model satisfactorily meet-the spin-recovery requirements it was found that installation of either a very large ventral fin (l7.9 square feet, full scale) below the tail or a somewhat smaller ventral fin and rudder (12.4 square feet, total . full-scale area) with a rudder throw of at least +/-22deg was required. Either a 21.3-foot tail parachute or a 6.4-foot wing-tip parachute (drag coefficient approximately 0.70) appears necessary as an emergency spin-recovery device during demonstration spins.

From Summary: "Tests of a 1/7 size model of the Grumman XJR2F-1 amphibian were made in Langley tank no.1 to examine the landing behavior in rough water and to measure the normal and angular accelerations experienced by the model during these landings. All landings were made normal to the direction of wave advance, a condition assumed to produce the greatest accelerations. Wave heights of 4.4 and 8.0 inches (2.5 and 4.7 ft, full size) were used in the tests and the wave lengths were varied between 10 and 50 feet (70 and 350 ft, full size)."

"A Douglas C-74 airplane, during a test dive at about 0.525 Mach number, experienced uncontrollable longitudinal oscillations sufficient to cause shedding of the outer wing panels and the subsequent crash of the airplane. Tests of a section of the horizontal tail plane from a C-74 airplane were conducted in the Ames 16-foot high-speed wind tunnel to investigate the possibility of the tail as a contributing factor to the accident. The results of the investigations of fabric-covered elevators in various conditions of surface deformation are presented in this report" (p. 1).

"An analysis has been made of the lift control effectiveness of a 20-percent-chord plain trailing-edge flap on the NACA 65-210 airfoil section from section lift-coefficient data obtained at Mach numbers from 0.3 to 0.875. In addition, the effectiveness of the plain flap as a lift-control device has been compared with the corresponding effectiveness of both a spoiler and a dive-recovery flap on the NACA 65-210 airfoil section. The analysis indicates that the plain trailing-edge flap employed on the 10-percent-thick airfoil at Mach numbers as high as 0.875 retains at least 50-percent of its low-speed lift-control effectiveness, and is sufficiently effective in lateral control application, assuming a rigid wing, to provide adequate airplane rolling characteristics" (p. 1).

"An analysis has been made of the lift-control effectiveness of a 20-percent-chord plain trailing-edge flap on the NACA 65-210 airfoil section from section lift-coefficient data obtained at Mach numbers from 0.3 to 0.875. In addition, the effectiveness of the plain flap as a lift-control device has been compared with the corresponding effectiveness of both a spoiler and a dive-recovery flag on the INCA 65-210 airfoil section" (p. 1).

"It is shown that for velocities slightly in excess of sonic, the position of detached shock wave located in front of a given body at zero angle of attack may be estimated theoretically to a reasonable degree of accuracy. In case of bodies of revolution the result was simple, but for two-dimensional bodies, pressure coefficient varies with Mach number, and slight difficulty appears. Theory developed compares favorably with available experimental data" (p. 1).

"At the request of the Air Material Command, Army Air Forces, an investigation was conducted by the NACA Cleveland laboratory to determine the performance characteristics of the compressor of the XJ-41-V turbojet engine. This report is the second in a series presenting the compressor performance and analysis of flow conditions in the compressor. The static-pressure variation in the direction of flow through the compressor and the location and the cause of the maximum flow restriction at an equivalent speed of 8000 rpm are presented" (p. 1).

"A fuel-distribution control for continuous-flow manifold injection on reciprocating engines is described. A method is installation of the control on an engine is suggested. The device controlled the flow to four spring-loaded nozzles within 2 percent of perfect distribution over a wide range of fuel-flow rates and the distribution was practically unaffected by uneven discharge-nozzle pressures" (p. 1).

Report presenting an investigation of two sharp-edge annular inlets with conical central bodies at low airspeeds in the propeller research tunnel to obtain information concerning the surface-pressure, drag, and pressure-recovery characteristics of the inlets at subsonic flight. Results regarding the nose and inner surface of the diffuser, inlet lips, minimum surface pressures and critical Mach numbers, pressure surveys in inlets, and pressure surveys in diffusers are provided.

Report presenting results of chordwise pressure-distribution measurements on a 45 degree sweptback wing at transonic speeds. The two stations for measuring were located at the 18 percent chord from the root and the 87 percent chord. The changes in pressure distribution with Mach number did not indicate any appreciable net loss in section lift, but did indicate large increases in the section drag and diving moment.

"On the basis of a recently developed theory for sweptback wings at supersonic velocities, equations are derived for the wave drag of sweptback tapered wings with thin symmetrical double-wedge sections at zero lift. Calculations of section wave-drag distributions and wing wave drag are presented for families of tapered plan forms. Distributions of section wave drag along the span of tapered wings are, in general, very similar in shape to those of untapered plan forms" (p. 1).

From Introduction: "A 12-percent-chord-thick wedge section and a reversed NACA 0012 section were chosen for these tests as they are representative of sections having no boat tailing and appreciable boat tailing (i.e., blunt and rounded trailing edges, respectively), and the results of this investigation are compared with those obtained from a previous investigation of the NACA 0012 section. Conclusions are drawn regarding the relative merits of the two unconventional sections and the conventional section in transonic speed range."

Report presenting flight tests to determine the longitudinal stability and control characteristics of models of the Lark pilotless aircraft. Testing included the standard configuration with the horizontal wing flaps deflected 0 degrees, 15 degrees, and 60 degrees, and a test with the tail surfaces in line with the wings with the wing flaps not deflected. Results regarding time-history records, drag, longitudinal stability, and control effectiveness are provided.

Presents results of the U. S. Air Forces' accelerated transonic flight tests of the XS-1 No. 1 airplane for the Mach number range from 0.70 to 1.25 at altitudes from 30,000 to 49,000 feet. Data are included on horizontal-tail loads and buffeting, longitudinal trim changes, elevator effectiveness and control forces, and lateral trim characteristics.

Report presenting testing of six flutter wings, two swept 45 degrees and four unswept, attached to three freely falling bodies. Flutter frequency and phasing were recorded.

"An altitude-test-chamber investigation was conducted to determine the operational characteristics and altitude blow-out limits of a Solar afterburner in a 24C engine. At rated engine speed and maximum permissible turbine-discharge temperature, the altitude limit as determined by combustion blow-out occurred as a band of unstable operation of about 8000 feet altitude in width with maximum altitude limits from 32,000 feet at a Mach number of 0.3 to about 42,000 feet at a Mach number of 1.0. The maximum fuel-air ratio of the afterburner, as limited by maximum permissible turbine-discharge gas temperatures at rated engine speed, varied between 0.0295 and 0.0380 over a range of flight Mach numbers from 0.25 to 1.0 and at altitudes of 20,000 and 30,000 feet" (p. 1).

Contains charts for use in determining preliminary values of the spanwise-load, shear, bending-moment, and accumulated-torque distributions of swept wings. The charts are based on strip theory and include four aerodynamic-load distributions, two section-moment distributions, and two inertia-load distributions. The taper ratios considered cover the range from 1.0 to 0 and the results are applicable to any angle of sweep.