Report discussing testing to determine the drag characteristics at zero lift of a wing with a circular-arc airfoil section with a maximum thickness of 9 percent chord. The results were compared to previous testing on an NACA 65-009 airfoil. It was found that the NACA airfoil had lower drag coefficients than the circular-arc airfoil tested in this experiment.

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.

"Tests of a 1/20-scale model of the Fleetwings XBTK-1 airplane have been performed in the Langley 15-foot free-spinning tunnel to determine the trim tendencies of the airplane at attitudes above the stall. The results of the tests indicated that the model would trim longitudinally only in the normal range of angles of attack and that the yaw trim tendencies for such longitudinal trim conditions were normal. Although wide oscillations in yaw were noted for some conditions, they occurred at angles of attack larger than those indicated as possible for longitudinal trim and spin equilibrium" (p. 1).

The available foreign and American data relating to Reynolds number effects on the maximum lift coefficients of swept-back wings are summarized and discussed. The data show that at low Reynolds numbers (below about 2.0 x 10(exp 6)) higher maximum lift coefficients were measured in most cases for moderately swept-back wings than for unswept wings of similar plan form; at high Reynolds numbers, however, increasing sweepback resulted in decreasing maximum lift coefficients. A smaller rate of increase of the maximum lift coefficient with Reynolds number was measured for the swept-back wings than for similar unswept wings in the critical range of Reynolds number. Increasing the Reynolds number resulted in decreases in the maximum lift coefficients of the two wings of approximately triangular plan form that were investigated.

"A method is described by which the total temperature of the gases at the combustion-chamber outlet of a ram-jet engine may be determined from the loss in total pressure measured across the combustion chamber. A working chart is presented by means of which the ratio of the total temperature of the gases at the combustion-chamber outlet to the total temperature of the gases at the combustion-chamber inlet may be determined from the measured loss of total pressure across the combustion chamber and the known values of air flow, total pressure, and total temperature at the combustion-chamber inlet. Values of total-temperature ratio across the combustion chamber of a 20-inch ram jet were obtained in the Cleveland altitude wind tunnel over a range of pressure altitudes from 6000 to 15,000 feet" (p. 1).