ALBERT

Description: No abstract available

Description: The material given in this report summarizes some of the results of recent research that will aid the designers of an airplane in selecting or modifying a configuration to provide satisfactory stability and control characteristics. The requirements of the NACA for satisfactory flying qualities, which specify the important stability and control characteristics of an airplane from the pilot's standpoint, are used as the main topics of the report. A discussion is given of the reasons for the requirements, of the factors involved in obtaining satisfactory flying qualities, and of the methods used in predicting the stability and control characteristics of an airplane. The material is based on lecture notes for a training course for research workers engaged in airplane stability and control investigations.

Description: The aerodynamic forces on an oscillating airfoil or airfoil-aileron combination of three independent degrees of freedom have been determined. The problem resolves itself into the solution of certain definite integrals, which have been identified as Bessel functions of the first and second kind and of zero and first order. The theory, being based on potential flow and the Kutta condition, is fundamentally equivalent to the conventional wing-section theory relating to the steady case. The air forces being known, the mechanism of aerodynamic instability has been analyzed in detail. An exact solution, involving potential flow and the adoption of the Kutta condition, has been analyzed in detail. An exact solution, involving potential flow and the adoption of the Kutta condition, has been arrived at. The solution is of a simple form and is expressed by means of an auxiliary parameter K.

Description: The purpose of this presentation is to give you a survey of a field of aerodynamics which has for a number of years been attracting an ever growing interest. The subject is the theory of flows with friction, and, within that field, particularly the theory of friction layers, or boundary layers. As you know, a great many considerations of aerodynamics are based on the so-called ideal fluid, that is, the frictionless incompressible fluid. By neglect of compressibility and friction the extensive mathematical theory of the ideal fluid (potential theory) has been made possible.

Description: An investigation has been made in the Langley stability tunnel to determine the low-speed static stability and control characteristics of a model of the Bell MX-776. The results of the investigation indicated that the basic model configuration was longitudinally stable in the angle-of-attack range from about -16 deg. to 16 deg. but that the stability was a minimum near O deg angle of attack. The data indicated an aerodynamic-center position about 0.64 body diameters behind the center of gravity at low angles of attack. Reduction in the size of the front horizontal fins increased the longitudinal stability. With 20 percent of the span of the normal front horizontal fins cut off the aerodynamic center was about 1.04 body diameters behind the center of gravity, and with front horizontal fins having the same area as the front vertical fins, the aerodynamic center was 2.26 body diameters behind the center of gravity (at low angles of attack).

Description: Rocket-powered models were flown at high-subsonic, transonic, and supersonic speeds to determine the zero-lift drag of fin-stabilized parabolic bodies of revolution differing in fineness ratio and in position of maximum diameter. The present paper presents the results for fineness ratio 12.5, 8.91 and 6.04 bodies having maximum diameters located at stations of 20, 40, 60, and 80 percent of body length. All configurations had cut-off sterns and all had equal base, frontal, and exposed fin areas. For most of the supersonic-speed range models having their maximum diameters at the 60-percent station gave the lowest values of drag coefficient. At supersonic speeds, increasing the fineness ratio generally reduced the drag coefficient for a given position of maximum diameter.

Description: Wind tunnel tests of the 0.16-scale Douglas MX-656 model were made at low and high subsonic Mach numbers to investigate the static longitudinal- and lateral stability characteristics. The tests shows that undesirable changes in longitudinal stability at the stall were apparently caused by an altered downwash pattern at the tail. The jettisonable nose fins were highly destabilizing. Compressibility effects for the test Mach numbers were not detrimental to the longitudinal- or lateral-stability characteristics.

Description: Measurements of wing and fuselage pressure distributions were made at low and high subsonic Much numbers on a 0.16-scale model of the projected MX-656 research airplane. The MX-656 is a supersonic design utilizing a low-aspect-ratio wing and tail. Pressure-distribution measurements indicated that, although the critical Mach number of the wing was approximately 0.81 at 0 degree angle of attack, compressibility effects were of little significance below a Mach number of at least 0.90. The principal effect of compressibility was an increase in the pressure gradient over the after 30 percent of the wing chord, causing a tendency for the flow to separate. At 0.40 Mach number, the wing stalled abruptly at approximately 12 deg, angle of attack. The wing-pressure distribution showed this stall was a result of complete separation of the flow from the upper surface of the wing, Deflecting the leading-edge flaps delayed the stall to a higher angle of attack with some increase in the maximum section normal force,

Description: A supersonic compressor design having supersonic velocity at the entrance of the stator is analyzed on the assumption of two-dimensional flow. The rotor and stator losses assumed in the analysis are based on the results of preliminary supersonic cascade tests. The results of the analysis show that compression ratios per stage of 6 to 10 can be obtained with adiabatic efficiency between 70 and 80 percent. Consideration is also given in the analysis to the starting, stability, and range of efficient performance of this type of compressor. The desirability of employing variable-geometry stators and adjustable inlet guide vanes is indicated. Although either supersonic or subsonic axial component of velocity at the stator entrance can be used, the cascade test results suggest that higher pressure recovery can be obtained if the axial component is supersonic.

Description: In the Institute for Flight Mechanics of the DVL a reactor arrangement with a maximum output of 100 kg was investigated as an expedient for the termination of dangerous spins on an airplane of the FW 56 type. reproduce the influence of a disturbance of the steady spin condition by a pitching or yawing moment. The tests were meant to reproduce the influence of a disturbance of the steady spin condition by a pitching and yawing moment.

Description: The present report, which deals with pressure-distribution measurements made on a sweptback wing with a jet engine nacelle, is similar to a report on pressure-distribution measurements on a rectangular wing with a jet engine nacelle (second partial report). Here, in investigations preliminary to high-speed measurements, as in the second partial report, useful arrangements and fillet designs have been discovered.

Description: The present report, which forms the first of six articles on experiments with airfoils of aspect ratio from 1 to 3 and various planforms, deals with the three- and six-component measurements made on the trapezoidal wing series in the 2.15 x 3-meter wind tunnel of the DVL at the request of the Henschel Aircraft Company.

Description: An investigation was conducted on an 0.08-scale semispan model of the Chance Vought XF7U-1 airplane in the Langley high-speed 7- by 10-foot tunnel in the Mach number range from 0.40 to 0.97. The results are compared with those obtained with an 0.08-scale sting-mounted complete model tested in the same tunnel and with an 0.026-scale semispan model tested by the wing-flow method. The lift-curve slopes obtained for the 0.08-scale semispan model and the 0.026-scale wing-flow model were in good agreement but both were generally lower than the values obtained for the sting model. The results of an unpublished investigation have shown that tunnel-wall boundary-layer and strut-leakage effects can came the difference noted between the lift-curve slopes of the sting and the semispan data. Fair agreement was obtained among the data of the three models as regard the variation of pitching-moment coefficients with lift coefficient. The agreement between the complete and the semispan models was more favorable with the vertical fine on, because the wall-boundary-layer and strut leakage effects were less severe. In the Mach number range between 0.94 and 0.97, ailavator-control reversal was indicated in the wing-flow data near zero lift; Whereas, these same trends were indicated in the larger scale semispan data at somewhat higher lift coefficients.

Description: The present report is concerned with a series of tests on a model airplane fitted with four types of dive flaps of various shapes, positions, and incidence located near the leading edge of the wing (from 5 to 20 percent of the wing chord). Tests were also made on a stub airfoil fitted with a ventral dive (located at 8 percent of the wing chord). The hinge moments of the dive flaps were measured.

Description: The flexural vibration of a rotating propeller blade with clamped shank is analyzed with the object of presenting, in matrix form, equations for the elastic bending moments in forced vibration resulting from aerodynamic forces applied at a fixed multiple of rotational speed. Matrix equations are also derived which define the critical speeds end mode shapes for any excitation order and the relation between critical speed and blade angle. Reference is given to standard works on the numerical solution of matrix equations of the forms derived. The use of a segmented blade as an approximation to a continuous blade provides a simple means for obtaining the matrix solution from the integral equation of equilibrium, so that, in the numerical application of the method presented, the several matrix arrays of the basic physical characteristics of the propeller blade are of simple form, end their simplicity is preserved until, with the solution in sight, numerical manipulations well-known in matrix algebra yield the desired critical speeds and mode shapes frame which the vibration at any operating condition may be synthesized. A close correspondence between the familiar Stodola method and the matrix method is pointed out, indicating that any features of novelty are characteristic not of the analytical procedure but only of the abbreviation, condensation, and efficient organization of the numerical procedure made possible by the use of classical matrix theory.

Description: When auxiliary jet engines are installed on airframes; as well as in some new designs, the jet engines are mounted in such a way that the jet stream exhausts in close proximity to the fuselage. This report deals with the behavior of the jet in close proximity to a two-dimensional surface. The experiments were made to find out whether the axially symmetric stream tends to approach the flat surface. This report is the last of a series of four partial test reports of the Goettingen program for the installation of jet engines, dated October 12, 1943. This report is the complement of the report on intake in close proximity to a wall.

Description: In an earlier report UM No.1117 by Gothert,the single-source method was applied to the compressible flow around circles, ellipses, lunes, and around an elongated body of revolution at different Mach numbers and the results compared as far as possible with the calculations by Lamla ad Busemann. Essentially, it was found that with favorable source arrangement the single-source method is in good agreement with the calculations of the same degree of approximation by.Lamla and Busemann. Near sonic velocity the number of steps must be increased considerably in order to sufficiently approximate the adiabatic curve. After exceeding a certain Mach number where local supersonic fields occur already, it was no longer possible, in spite of the substantially increased number of steps, to obtain a systematic solution because the calculation diverged. This result,was interpreted to mean that above this point of divergence the symmetrical type of flow ceases to exist and changes into the unsymmetrical type characterized by compressibility shocks.

Description: The problem of the motion of an elongated body of revolution in an incompressible fluid may, as is known, be solved approximately with the aid of the distribution of sources along the axis of the body. In determining the velocity field, the question of whether the body moves uniformly or with an acceleration is no factor in the problem. The presence of acceleration must be taken into account in determining the pressures acting on the body. The resistance of the body arising from the accelerated motion may be computed either directly on the basis of these pressures or with the aid of the so-called associated masses (inertia coefficients). A different condition holds in the case of the motion of bodies in a compressible gas. In this case the finite velocity of sound must be taken into account.

Description: Various ways were tried recently to decrease the friction drag of a body in a flow; they all employ influencing the boundary layer. One of them consists in keeping the boundary layer Laminar by suction; promising tests have been carried out. Since for large Reynolds numbers the friction drag of the laminar boundary layer is much lower than that of the turbulent boundary layer, a considerable saving in drag results from keeping the boundary layer laminar, even with the blower power required for suction taken into account. The boundary layer is kept laminar by suction in two ways: first, by reduction of the thickness of the boundary layer and second, by the fact that the suction changes the form of the velocity distribution so that it becomes more stable, in a manner similar to the change by a pressure drop. There by the critical Reynolds number of the boundary layer (USigma*/V) (sub crit) becomes considerably higher than for the case without suction. This latter circumstance takes full effect only if continuous suction is applied which one might visualize realized through a porous wall. Thus the suction quantities required for keeping the boundary layer laminar become so small that the suction must be regarded as a very promising auxiliary means for drag reduction.

Description: A model of a Consolidated Vultee Aircraft Corporation Skate 7 sea-plane:was tested in Langley tank no= 2. Resistance data, 'spray photographs, and underwater photographs,are given in this report without discussion.

Description: A vane-type angle-of-attack indicator suitable for measurements at both subsonic and supersonic speeds has been developed by the National Advisory Committee for Aeronautics. A brief history is given of the development, and a wind-tunnel calibration of the indicator is presented, together with a discussion of the corrections to be applied to the indicated readings.

Description: Four component measurements of 12 wings of symmetric profile having flaps with chord ratios t(sub R)/t(sub L) = 0.3 and t(sub R)/t(sub L) = 0.2 are treated in this report. As a result of the investigations, the effects of plan form and gap between fixed surface and control surface have been clarified. Lift, drag, pitching moment, and hinge moment were measured in the control-surface deflection range: -23 deg 〈 or = beta 〈 or = 23 deg and the range of angle of attack: -20 deg 〈 or = alpha 〈 or = 20 deg. Six wings with flaps of small chord (t(sub R)/t(sub L) 〈 0.1) were investigated at large flap settings.

Description: The purpose of the model tests is to clarify the motions in the alighting on water of a land plane. After discussion of the model laws, the test method and test procedure are described. The deceleration-time-diagrams of the landing of a model of the Bf 109 show a high deceleration peek of greater than 20g which can be lowered to 4 to 6g by radiator cowling and brake skid.

Description: Wind-tunnel tests of a full-scale model of the Republic XF-91 airplane were conducted to determine the distribution of pressure over the external wing fuel tank installation and over the vee tail and ventral fin. The data were obtained for a range of angles of attack and sideslip and elerudder deflection angles; the presentation is in tabular form.

Description: The present report describes a new method for the prediction of the flow pattern of a gas in the two-dimensional and axially symmetrical case. It is assumed that the expansion of the gas is adiabatic and the flow stationary. The several assumptions necessary of the nozzle shape effect, in general, no essential limitation on the conventional nozzles. The method is applicable throughout the entire speed range; the velocity of sound itself plays no singular part. The principal weight is placed on the treatment of the flow near the throat of a converging-diverging nozzle. For slender nozzles formulas are derived for the calculation of the velocity components as function of the location.

Description: An investigation has been conducted in the Langley 20-foot free-spinning tunnel of a 1/29-scale model of the Republic XF-91 airplane with a.conventional-tail arrangement installed. Previously, tests were made on the model with a vee tail installed. The erect spin and recovery characteristics of the model were determined for the normal loading with the wing installed at various amounts of incidence. The spin investigation also included inverted-spin tests, spin-recovery-parachute tests, tests with the center of gravity moved rearward, and tests with external fuel tanks added to the model. In addition, several tail.modifications were tested,on the model in an attempt, to improve the model's spin-recovery characteristics. The results indicate that any fully developed spin obtained on the airplane with the conventional tail installed will be satisfactorily terminated if rudder reversal is accompanied by moving the ailerons with the spin (stick right in a right spin).Decreasing the wing incidence from 6deg to -2deg should have a beneficial effect on the recovery characteristics of the airplane. Recovery characteristics by normal use of controls (full rudder reversal followed by moving the elevators down) will be satisfactory if the wing incidence,of the airplane is -2deg. Installation of external fuel tanks (with or without fuel) will have a somewhat adverse effect on the recovery characteristics of the airplane, but if the recovery technique includes movement of the ailerons to full with the spin, the spin rotation will be terminated rapidly. Varying the position of the center of gravity within the limits indicated to be possible on the airplane should not affect the recovery characteristics.

Description: In order to obtain insight into the flow conditions on tail surfaces on airplanes during spins, pressure-distribution measurements were performed on a rotating model of the design BFW-M31. For the time being, the tests were made for only one angle of attack (alpha = 60 degrees) and various angles of yaw and rudder angles. The results of these measurements are given; the construction of the model, and the test arrangement used are described. Measurements to be performed later and alterations planned in the test arrangement are pointed out.

Description: The effect of application of power, so far not clarified, is investigated in the present report in order to give the pilot, in addition to the control measures, an expedient for spin recovery of multiengine airplanes. To this end, a series of spins was performed with an airplane of the Go 150 type. It was possible to set up a uniform rule regarding the effect of power, for right end left spins as well as for any combination of the direction of rotation of the propellers.

Description: The plane problem of the vibrating airfoil in supersonic flow is dealt with and solved within the scope of a linearized theory by the method of the acceleration potential.

Description: Results of previous model ditching tests of the Lockheed Constellation airplane are reported. Further model tests have been made to determine the probable ditching characteristics and the proper ditching technique for the airplane with the Speedpak attached. This paper presents the results of these tests. Design information was furnished by the Lockheed Aircraft Corporation. A three-vies drawing of the airplane with the Speedpak attached is shown. The tests were made in calm water at the Langley tank no.2 monorail.

Description: A supplementary investigation on the stabilization of the Jettisonable nose section of the X-2 airplane has been conducted in the Langley 20-foot free-spinning tunnel. It was found that the nose section could be stabilized by the addition of curved fins which could be folded against the fuselage for normal flight.

Description: Flight tests were conducted at the NACA Pilotless Aircraft Research Station, Wallops Island, to determine the characteristics of the Allegany Ballistics Laboratory's 6.2inch-diameter Deacon and lO-inch-scale model solid-propellant rocket motors. The tests were performed to assist in the development of these rockets which were designed for, and urgently needed to propel supersonic research models and pilotless aircraft. The tests showed that the rocket motors functioned properly under various flight- acceleration loads over a range of pre-ignition grain temperatures. A maximum velocity of 4180 feet per second was obtained at an elapsed time of 2.9 seconds with the 6.25-inch Deacon rocket motor at a gross weight of l9O pounds. Free-flight data of drag coefficient for the Deacon configuration for a Mach number range of 1.1 to 3.6 have been obtained from flight tests of several pounds. Camera studies of the take-off and flights of the Deacon rocket shared no evidence of breakup of propellant grains. An analysis of the forces to which the Deacon rocket grain is subjected was made. The analysis shows that the grain loading is most severe near the beginning and near the end of the rocket action time. The 10-inch-scale model rocket motor is a scaled model of the l6-inch- diameter multi-perforated, cast-grain rocket motor. A maximum velocity of 1625 feet per second at a time of 1.075 seconds was obtained at a gross weight of 309 pounds.

Description: The characteristics of a cargo-dropping device having extensible rotating blades as load-carrying surfaces have been studied in simulated vertical descent in the Langley 20-foot free-spinning tunnel. The investigation included tests to determine the variation in vertical sinking speed with load. A study of the blade characteristics and of the test results indicated a method of dynamically balancing the blades to permit proper functioning of the device.

Description: A spin-recovery investigation has been conducted in the Langley 20-foot free-spinning tunnel on a 20-scale model modified to represent the McDonnell XF2H-1 airplane. The project included tests both with tip tanks installed and with the tanks removed. The results indicated that the recovery characteristics of the airplane would be satisfactory for all loadings by normal recovery technique (full reversal of the rudder, followed 1/2 turn later by movement of the elevator down). The rudder pedal and the elevator stick forces likely to be encountered in a spin should be within the capabilities of the pilot.

Description: Studies have been made by the NACA wing-flow method of the use of fuselage static orifices between the wing and tail of a swept-wing airplane for possible application to service airspeed installations. The tests were made at zero angle of attack. The results indicate that, although the maximum errors are large, these locations are usable from the consideration that the local Mach numbers at the locations studied are sensitive to variation of the true Mach number within the test Mach number range of 0.7 to 1.2. The maximum errors in Mach number in the subsonic range varied from zero for the most forward location to -0.05 for the most rearward location (indicated Mach number less than true). At Mach numbers above 1.0, the maximum errors were from 0.14 for the most forward location to 0.04 for the most rearward location.

Description: The take-off stability characteristics of a Consolidated Vultee Aircraft Corporation Skate 7 seaplane were determined in the Langley tank no. 2. Trim limits of stability, trim tracks, and elevator limits of stability are presented.

Description: High-speed wind-tunnel tests were conducted of two versions of a 0.17-scale model of the McDonnell XF2H-1 airplane to ascertain the high-speed stability and control characteristics and to study means for raising the high-speed buffet limit of the airplane, The results for the revised model, employing a thinner wing and tail than the original model, revealed a mild diving tendency from 0.75 to 0.80 Mach number, followed by a marked climbing tendency from 0.80 to 0.875 Mach number. The high-speed climbing tendency was caused principally by the pitching-moment characteristics of the wing. At 0.875 Mach number the results for the revised model indicated stick-fixed directional instability over a limited range of yaw angles, apparently caused by separated flow over the vertical tail. The test results indicate that the high-speed buffet limit of the airplane can probably be raised by reducing the thickness and changing the relative location of the horizontal and vertical tails, and by revising the inner portion of the wing to have a lower thickness-to-chord ratio and reduced trailing-edge angle. The addition of the wing-tip tanks to the revised model resulted in a forward shift in the neutral point below 0.82 Mach number.

Description: An investigation was conducted to determine the longitudinal- and lateral-stability characteristics of a 0.5-scale moue1 of the Fairchild Lark missile, The model was tested with 0 deg and with 22.5 deg of roll. Three horizontal wings having NACA 16-009, 16-209, and 64A-209 sections were tested. Pressures were measured on both pointed and blunt noses. The wind-tunnel-test data indicate that rolling the missile 22.5 deg. had no serious effect on the static longitudinal stability. The desired maneuvering acceleration could not be attained with any of the horizontal wings tested, even with the horizontal wing flaps deflected 50 deg. The flaps on the 64A-209 wing (with small trailing-edge angles and flat sides) were effective at all flap deflections, while the flaps on the 16-series wings (with large trailing-edge angles) lost effectiveness at small flap deflections. The data showed that rolling moment existed when the vertical wing flaps were deflected with the model at other than zero angle of attack. A similar rolling moment probably would be found . with the horizontal wing flaps deflected and the model yawed.

Description: An investigation has been conducted with the Langley helicopter tower to obtain basic performance and control characteristics of the Raman rotor system. Blade-pitch control is obtained in this configuration by utilizing an auxiliary flap to twist the blades. Rotor thrust and power required were measured for the hovering condition and over a range of wind velocities from 0 to 30 miles per hour. The control characteristics and the transient response of the rotor to various control movements were also measured. The hovering-performance data are presented as a survey of the wake velocities and the variation of torque coefficient with thrust coefficient. The power required for the test rotor to hover at a thrust of 1350 pounds and a rotor speed of 240 rpm is approximately 6.5 percent greater than that estimated for a conventional rotor of the same diameter and solidity. It is believed that most of this difference is caused by th e flap servomechanism. The reduction in total power required for sustentation of the single-rotor configuration tested at various wind velocities and at the normal operating rotor thrust was found to be similar to the theoretical and experimental results for ro tors with conventionally actuated pitch. The control effectiveness was determined as a function of rotor speed. Sufficient control was available to give a thrust range of 0 to 1500 pounds and a rotor tilt of plus or minus 7 degrees. The time lag between flap motion and blade-pitch response is approximately 0.02 to 0.03 second. The response of the rotor following the blade-pitch response is similar to that of a rotor with conventionally actuated pitch changes. The over-all characteristics of the rotor investigated indicate that satisfactory performance and control characteristics were obtained.

Description: An investigation was made to evaluate the hydrodynamic qualities of a 0.425-scale model of the Navy XP5M-1 hull, which was installed on a modified Navy J4F-2 amphibian. Longitudinal and directional stability during take-off and landing, low-speed maneuverability, spray characteristics, and take-off performance were investigated. The behavior of the airplane in moderately rough water was also observed. The opinions of three pilots have been correlated with the data.

Description: A 0.1-size powered dynamic model of a large, high-speed flying boat was landed in Langley tank no. 1 into oncoming waves 4 feet high (full size). The model was tested with two afterbodies of differing lengths (4.12 and 6.63 beams). The short afterbody had a constant angle of dead rise of 22.5deg and a keel angle of 6.5deg. The long afterbody had warped dead rise and a keel angle of 8.5deg. The vertical accelerations were slightly greater and the maximum angular accelerations and maxim= trims were slightly less for the model with the long afterbody than for the model with -the short afterbody. A wave length of 210 feet (full size) imposed the highest accelerations on the model with either the long or the short afterbody.

Description: A flight test of the Aero jet Engineering Corporation's 7KS-6000 T-27 Jato rocket motor was conducted at the Langley Pilotless Aircraft Research Station at Wallops Island, Va, to determine the flight performance characteristics of the motor. The flight test imposed an absolute longitudinal acceleration of 9.8 g upon the rocket motor at 2.8 seconds after launching. The total impulse developed by the motor was 43,400 pound-seconds, and the thrusting time was 7.58 seconds. The maximum thrust was 7200 pounds and occurred at 4.8 seconds after launching. No thrust irregularities attributable to effects of the flight longitudinal acceleration were observed. Certain small thrust irregularities occurred in the flight test which appear to correspond to irregularities observed in static tests conducted elsewhere. A hypothesis regarding the origin of these small irregularities is presented.

Description: The inlet wide vanes for the supersonic compressor of the XJ55-FF-1 engine were studied as a separate component in order to determine the performance prior to installation in the compressor test rig. Turning angles approached design values, and increased approximately to through the inlet Mach number range from 0.30 to choke. A sharp break in turning angle was experienced when the choke condition was reached. The total-pressure loss through the guide vanes was approximately 1 percent for the unchoked conditions and from 5 to 6 percent when choked.

Description: During the past several years it has been necessary for aeronautical research workers to exert a good portion of their effort in developing the means for conducting research in the high-speed range. The transonic range particularly has presented a very acute problem because of the choking phenomena in wind tunnels at speeds close to the speed of sound. At the same time, the multiplicity of design problems for aircraft introduced by the peculiar flow problems of the transonic speed range has given rise to an enormous demand for detail design data. Substantial progress has been made, however, in developing the required research techniques and in supplying the demand for aerodynamic data required for design purposes. In meeting this demand, it has been necessary to resort to new techniques possessing such novel features that the results obtained have had to be viewed with caution. Furthermore, the kinds of measurements possible with these various techniques are so varied that the correlation of results obtained by different techniques generally becomes an indirect process that can only be accomplished in conjunction with the application of estimates of the extent to which the results of measurements by any given technique are modified by differences that are inherent in the techniques. Thus, in the establishment of the validity and applicability of data obtained by any given technique, direct comparisons between data from different sources are a supplement to but not a substitute for the detailed knowledge required of the characteristics of each technique and fundamental aerodynamic flow phenomena.

Description: A preliminary analysis of the flying qualities of the Consolidated Vultee MX-813 delta-wing airplane configuration has been made based on the results obtained from the first two 1/8 scale models flown at the NACA Pilotless Aircraft Research Station, Wallop's Island, VA. The Mach number range covered in the tests was from 0.9 to 1.2. The analysis indicates adequate elevator control for trim in level flight over the speed range investigated. Through the transonic range there is a mild trim change with a slight tucking-under tendency. The elevator control effectiveness in the supersonic range is reduced to about one-half the subsonic value although sufficient control for maneuvering is available as indicated by the fact that 10 deg elevator deflection produced 5g acceleration at Mach number of 1.2 at 40,000 feet.The elevator control forces are high and indicate the power required of the boost system. The damping. of the short-period oscillation is adequate at sea-level but is reduced at 40,000 feet. The directional stability appears adequate for the speed range and angles of attack covered.

Description: Wind-tunnel tests at low Mach number of a Republic F-84C airplane were conducted to determine by pressure-distribution measurements the air loads on wing-tip tanks and the change in wing load distribution due to the presence of tip tanks. Measurements of the aeroelastic twist of the wing were also obtained. Results are presented in the form of loading coefficient, center-of- pressure location, pitching-moment coefficient, aerodynamic-center location, and aeroelastic twist. The investigation revealed that the redistributions in loading brought about by either the tip tanks or elastic deformation of the wing were relatively small when compared with the chnnges in loading normally associated with the deflection of an aileron.

Description: An additional series of high-speed wind-tunnel tests of a modified 0.17-scale model of the McDonnell XF2H-1 airplane was conducted to evaluate the effects of a reduction in the thickness-to-chord ratios of the tail planes, the displacement of the horizontal tail relative to the vertical tail, and the extension of the trailing edge of the wing. Two tail-intersection fairings designed to improve the flow at the tail were also tested. The pitching-moment characteristics of the model were improved slightly by the use of the thinner tail sections. Rearward or rearward and downward displacements of the horizontal tail increased the critical Mach number at the tail intersection from 0.725 to a maximum of 0.80, but caused an excessive change in pitching-moment coefficient at the higher Mach numbers. Extending the trailing edge of the wing did not improve the static longitudinal-stability characteristics, but increased the pitching-down tendency between 0.725 and 0.825 Mach numbers prior to the pitching-up tendency. The extended wing did, however, increase the Mach numbers at which these tendencies occurred. The increase in the Mach numbers of divergence and the tuft studies indicate a probable increase in the buffet limit of the prototype airplane. No perceptible improvement of flow at the tail intersection was observed with the two fairings tested on the forward tail configuration.

Description: Lately it has been proposed to reduce the friction drag of a body in a flow for the technically important large Reynolds numbers by the following expedient: the boundary layer, normally turbulent, is artificially kept laminar up to high Reynolds numbers by suction. The reduction in friction drag thus obtained is of the order of magnitude of 60 to 80 percent of the turbulent friction drag, since the latter, for large Reynolds numbers, is several times the laminar friction drag. In considering the idea mentioned one has first to consider whether suction is a possible means of keeping the boundary layer laminar. This question can be answered by a theoretical investigation of the stability of the laminar boundary layer with suction. A knowledge, as accurate as possible, of the velocity distribution in the laminar boundary layer with suction forms the starting point for the stability investigation. E. Schlichting recently gave a survey of the present state of calculation of the laminar boundary layer with suction.

Description: For the normal range of engine power the impeller provided marked improvement over the standard spray-bar injection system. Mixture distribution at cruising was excellent, maximum cylinder temperatures were reduced about 30 degrees F, and general temperature distribution was improved. The uniform mixture distribution restored the normal response of cylinder temperature to mixture enrichment and it reduced the possibility of carburetor icing, while no serious loss in supercharger pressure rise resulted from injection of fuel near the impeller outlet. The injection impeller also furnished a convenient means of adding water to the charge mixture for internal cooling.

Description: Behaviors of both model and full-scale airplanes were ascertained by making visual observations, by recording time histories of decelerations, and by taking motion picture records of ditchings. Results are presented in form of sequence photographs and time-history curves for attitudes, vertical and horizontal displacements, and longitudinal decelerations. Time-history curves for attitudes and horizontal and vertical displacements for model and full-scale tests were in agreement; maximum longitudinal decelerations for both ditchings did not occur at same part of run; full-scale maximum deceleration was 50 percent greater.

Description: Tests show that at inlet-air temperatures of 250 deg F and 100 deg F the knock-limited performance of the base fuel of blends, leaded with 4 ml TEL per gallon and containing 20 percent spiropentane, was reduced at fuel/air ratios below 0.085. The 20 percent methylenecyclobutane reduced the knock-limited power of the base fuel at fuel/air ratios below 0.112. Di-tert-butyl ether, methyl-tert-butyl ether, and triptane increased the knock-limited power of the base fuel at all fuel/air ratios and at both temperatures.

Description: An investigation was conducted to compare the performance of two 25-ft-diam rotors which had identical dimensions and were similar in construction but different in blade airfoil-sections. Tests were conducted at indicated blade pitch angles from 3 degrees to 11.5 degrees and rotor speeds of 200, 290, and 371 rpm. The 23012.6 rotor required 2 percent less power to hover than the 0012.6. At thrust coefficients above design, the performance of the 23012.6 became better than the 0012.6 rotor.

Description: A relatively simple equation has been found to express with fair accuracy, variation in manifold-charge temperature with charge in engine operating conditions. This equation and associated curves have been checked by multi cylinder-engine data, both test stand and flight, over a wide range of operating conditions. Average mixture temperatures, predicted by the equations of this report, agree reasonably well with results within the same range of carburetor-air temperatures from laboratories and test stands other than the NACA.

Description: An investigation has been conducted in the Langley rectangular high-speed tunnel to determine the effect of compressibility on the pressure distribution for a modified NACA 65,3-019 airfoil having a 0.20-chord flap. The investigation was made for an angle-of-attack range extending from -2 to 12 deg at .20 flap deflections from 0 to -12 deg. Test data were obtained for Mach numbers from 0.28 to approximately 0.74. The results show that the effectiveness of the trailing-edge-type control surface rapidly decreased and approached zero as the Mach number increased above the critical value.

Description: Results of flight tests of a control-feel aid presented. This device consisted of a spring and dashpot connected in series between the control stick and airplane structure. The device was tested in combination with an experimental elevator and bobweight which had given unsatisfactory dynamic stability and control-feel characteristics in previous tests. The control-feel aid effected marked improvement in both the control-feel characteristics and the control-feel dynamic longitudinal stability of the airplane.

Description: A theoretical investigation was conducted on jet-induced flow deviation. Analysis is given of flow inclination induced outside cold and hot jets and jet deflection caused by angle of attack. Applications to computation of effects of jet on longitudinal stability and trim are explained. Effect of jet temperature on flow inclination was found small when thrust coefficient is used as criterion for similitude. The average jet-induced downwash over tail plane was obtained geometrically.

Description: At the request of the Air Technical Service Command, U.S. Army Air Forces, a 0.22-scale model of a twin-fuselae pursuit airplane was built and tested at the Ames Aeronautical Laboratory. The tests of this model were made in order that the aerodynamic characteristics of the airplane, especially at high speed, might be predicted. The results shown in this report consist of force data for the model and critical Mach numbers of parts of the model as determined from pressure-distribution measurements. The results indicate that a diving tendency of the airplane can be expected at Mach numbers above 0.70 at lift co-efficients from 0 to 0.4. There is an indication that the Mach number at which the airpolane would first experience a diving tendency for lift coefficients from 0 to 0.2 can be increased if the critical speed of the radiator enclosures is increased, and the wing-fuselage-juncture fillets are improved.

Description: Wing section outboard of flap was tested by wake surveys in Mach range of 0.25 - 0.78 and lift coefficient range 0.06 - 0.69. Results indicated that minimum profile-drag coefficient of 0.0097 was attained for lift coefficients from 0.16 to 0.25 at Mach less than 0.67. Below Mach number at which compressibility shock occurred, variations in Mach of 0.2 had negligible effect on profile drag coefficient. Shock was not evident until critical Mach was exceeded by 0.025.

Description: Tests in Langley pressure tunnel of model XA-26 bomber were compared with those of A-26B (twin-engine attack bomber) and showed that static longitudinal stability, indicated by elevator-fixed neutral points, and variation of elevator deflection in straight and turning flight were good. Airplane possessed improved stability at low speeds which was attributed to pronounced stalling at root of production wing. At rudder-force reversal at speeds higher than those in flight tests, agreement in rudder-fixed and rudder-free static directional stability was good. Hinge moment obtained at zero sideslip was satisfactory for determining aileron forces in sideslip.

Description: Availability data obtained on SNB-1 trainer-class airplanes were analyzed and results presented as flight envelopes which predict occurrences of large values of air speed and acceleration. Comparison is made with SNJ-4 trainer-class airplane data analyzed by the same method. It is concluded that flight envelopes are satisfactory; that the two types show large differences in flight loads and speeds experience; and that SNB-1 will seldom, if ever, exceed design limit load factor and restricted speed, which SNJ-4 can be expected to exceed design-limit load factor and restricted speed in a very small number of flight hours.

Description: Results are reported of knock-limited tests of five aromatics, each individually blended with selected base fuels and tested with and without TEL, using 17.6, F-4, and F-3 small-scale engines. The five aromatics rated in the following order of decreasing antiknock effectiveness at fuel/air ratio 0.10: m-xylene, 1-isopropyl-4-methylbenzene, n-propylbenzene, isobutylbenzene, and n-butylbenzene.

Description: Data are presented of the flow conditions in the vicinity of an NACA D sub S -type cowling. Tests were made of a 1/2 scale-nacelle model at inlet-velocity ratios ranging from 0.23 to 1.02 and angles of attack from 6 deg to 10 deg. The velocity and direction of flow in the vertical plane of symmetry of the cowling were determined from orifices and tufts installed on a board aligned with the flow. Diagrams showing velocity ratio contours and lines of constant flow angles are given.

Description: Lift, drag, internal flow, and pressure distribution measurements were made on a low-drag airfoil incorporating various air inlet designs. Two leading-edge air inlets are developed which feature higher lift coefficients and critical Mach than the basic airfoil. Higher lift coefficients and critical speeds are obtained for leading half of these inlet sections but because of high suction pressures near exist, slightly lower critical speeds are obtained for the entire inlet section than the basic airfoil.

Description: Critical Mach number as function of lift coefficient is determined for certain moderately thick NACA low-drag airfoils. Results, given graphically, included calculations on same airfoil sections with plain flaps for small flap deflections. Curves indicate optimum critical conditions for airfoils with flaps in such form that they can be compared with corresponding results for zero flap deflections. Plain flaps increase life-coefficient range for which critical Mach number is in region of high values characteristic of low-drag airfoils.

Description: In conjunction with a program of research on the general problem of stability of airplanes in the climbing condition, tests have been made of a spring-loaded tb which. is referred to as a ?springy tab,? installed on the elevator of a low-wing scout bomber. The tab was arranged to deflect upward with decrease in speed which caused an increase in the pull force required to trim at low speeds and thereby increased the stick-free static longitudinal stability of the airplane. It was found that the springy tab would increase the stick-free stability in all flight conditions, would reduce the danger of inadvertent stalling because of the definite pull force required to stall the airplane with power on, would reduce the effect of center-of-gravity position on stick-free static stability, and would have little effect on the elevator stick forces in accelerated f11ght. Another advantage of the springy tab is that it might be used to provide almost any desired variation of elevator stick force with speed by adjusting the tab hinge-moment characteristics and the variation of spring moment with tab deflection. Unlike the bungee and the bobweight, the springy tab would provide stick-free static stability without requiring a pull force to hold the stick back while taxying. A device similar to the springy tab may be used on the rudder or ailerons to eliminate undesirable trim-force variations with speed.

Description: An analysis was made to determine the effect of rolling pull-out maneuvers on the wing and aileron loads of a typical fighter airplane, the P-47B. The results obtained indicate that higher loads are imposed upon wings and ailerons because of the rolling pull-out maneuver, than would be obtained by application of the loading requirements to which the airplane was designed. An increase of 102 lb or 15 percent of wing weight would be required if the wing were designed for rolling pull-out maneuver. It was also determined that the requirements by which the aileron was originally designed were inadequate.

Description: As part of an investigation to increase the power output of the V-1710-93 engine at altitude, the engine-stage supercharger was combined with a constant-area vaneless diffuser designed to improve the performance of the engine-stage supercharger at the rated engine operating point. The performance of the modified supercharger was investigated in a variable-component supercharger test rig and compared with that of the standard supercharger with an 8-vaned diffuser. A separate evaluation of the component efficiencies and a study of the flow characteristics of the modified supercharger was made possible by internal diffuser instrumentation. At the volume flow required by the engine for rated operating conditions, the modified supercharger increased the over-all adiabatic efficiency 0.05 and the over-all pressure coefficient 0.035. Furthermore, the capacity of the engine-stage supercharger was increased by replacing the standard 8-vaned diffuser with the vaneless diffuser. The peak over-all adiabatic efficiency for the modified supercharger, however, was 0.05 to 0.07 lower than that of the standard unit over the range of tip speeds investigated. The improved performance of the modified supercharger at rated engine operating conditions resulted from a shift of the point of peak adiabatic efficiency and pressure coefficient of the standard supercharger to a higher flow. The energy loss through the vaneless diffuser was found to be small. Because of the restricted diffuser diameter, however, diffusion was inadequate, which resulted in a relatively small static-pressure rise through the diffuser, high diffuser-exit velocities, and excessive collector-case losses.

Description: A dynamically similar model of the Army P-38 airplane was tested to determine the best way to land this airplane on the water and to determine its probable ditching performance. The tests consisted of ditching the model at various landing attitudes, flap settings, speeds, weights, and conditions of simulated damage. The model was ditched in calm water from the tank towing carriage and a few ditching were made in both calm and rough water at the outdoor catapult. The performance of the model was determined by making visual observations, by recording lengths of run and time histories of decelerations, and by taking motion pictures of the ditchings.

Description: In the present report the true weight distribution law of the wing structure along the span is investigated. It is shown that the triangular distribution and that based on the proportionality to the chords do not correspond to the actual weight distribution, On the basis of extensive data on wings of the CAHI type airplane formulas are obtained from which it is possible to determine the true diagram of the structural weight distribution along the span from a knowledge of only the geometrical dimensions of the wing. At the end of the paper data are presented showing how the structural weight is distributed between the straight center portion and the tapered portion as a function of their areas.

Description: An investigation was conducted in the Langley impact basin of the water loads on a half scale model of the XJL-1 hull whose forebody has a vee bottom with exaggerated chine flare. The impact loads, moments, and pressures were determined for a range of landing conditions. A normal full-scale landing speed of 86 miles per hour was represented with effective flight paths ranging from 0.6deg to 11.6deg. Landings were made with both fixed trim and free-to-trim mounting of the float over a trim range of -15deg to 12deg into smooth water and into waves having equivalent full-scale length. of 120 feet and heights ranging from 1 to 4 feet. All data and results presented in this report are given in terms of equivalent full-scale values. Summary tables and illustrative plots are used in presenting the material. The following maximum values of load and pressure are those which are apropos for effective flight paths less than 6.5deg which was the maximum value obtained in tests with the XJL-1 hull model representing full-scale landings with vertical velocity of 4.5 feet per second into 4-foot waves. The maximum local pressure on the flat portion of the bottom is 130 pounds per square inch which was measured on a 2-inch-diameter circular area near the step. The maximum local pressure obtained in the curved area near the chines is 200 pounds per square inch. This pressure was also measured near the step.

Description: Tables I and II of this report summarize the gust and draft velocity data for thunderstorm flights 25 and 26 of August 21, 1946 and August 22, 1946, respectively. These dta were evaluated from records of NACA instruments installed in P-61C airplanes and are of the type presented in reference 1 for previous flights. Table III summarizes the readings of a milliammeter which was used in conjunction with other equipment to indicate ambient air temperature during thunderstorm surveys. These data were read from motion-picture records of the instrument and include all cases in which variations in the instrument indications were noted during the present flights.

Description: A spin investigation has been conducted in the Langley 20-foot free-spinning tunnel on a 1/20-scale model of the Chance Vought XF6U-1 airplane, The effects of control settings and movements upon the erect and inverted spin and recovery characteristics of the model were determined for the normal-fighter condition. The investigation also included tests for the take-off fighter condition (wing-tip tanks plus fuel added) spin-recovery parachutes, and simulated pilot escape. In general, for the normal-fighter condition, the model was extremely oscillatory in roll, pitch, and yaw. The angles of the fuselage varied from extremely flat to inverted attitudes, and the model rotated with the rudder in a series of short turns and glides. Recoveries by rudder reversal were rapid but the model would immediately go into a spin in the other direction. Recoveries by merely neutralizing the rudder were satisfactory when the elevator and ailerons were set to neutral, the ensuing flight path being a steep glide. Thus, it is recommended that all controls be neutralized for safe recovery from spins obtained on the airplane. With the external wing-tip tanks installed, the spins were somewhat less oscillatory in roll but recovery could not be obtained unless full-down elevator was used in conjunction with the rudder. If a spin is entered inadvertently with the full-scale airplane with external wing-tip tanks installed and if recovery is not imminent after a recovery attempt is made, it is recommended that the tanks be jettisoned and the controls neutralized.

Description: Spin tests of a 1/20-scale model of the Northrop N-9M airplane have been performed in the Langley 20-foot free-spinning tunnel. The erect and inverted spin and recovery characteristics were determined for various loading conditions and the effect of deflecting the flaps and of extending the landing gear was investigated. The investigation also included tests to determine the size parachute required for satisfactory spin recovery by parachute action alone. The tests were performed at an equivalent spin altitude of 15,000 feet. A specialized recovery technique consisting of rapid full reversal of the rudder pedals against the spin combined with turning the wheel against the spin and movement of the stick forward is recommended for all loadings and configurations of the airplane. The results also indicated that a 7-foot-diameter spin-recovery parachute having a drag coefficient of 0.7 attached to the outboard wing tip with a towline of 10 to 30 feet or an 8.8-foot-diameter parachute attached to the fixed portion of the wing between the elevons and the pitch flaps with a 30-foot towline would provide satisfactory recovery from demonstration spins by parachute action alone. It appears possible that the first N-9M airplane may have crashed because of failure to recover from a spin.

Description: An investigation was made in the Cleveland Altitude wind tunnel to determine the performance of an Aeroproducts H20C-162-X11M2 four-blade propeller on a YP-47M airplane at high blade loadings and high engine powers. The propeller characteristics were obtained for a range of power coefficients from 0.30 to 1.00 at free-stream Mach numbers of 0.40 and 0.50. The results of the force measurements are indicative only of trends in propeller efficiency with changes in power coefficient and advance-diameter ratio because unknown interference effects existed during the investigation. At a free-stream Mach number of 0.40, the envelopes of the efficiency curves decreased about 11% between advance-diameter ratios of 2.40 and 4.40. An increase in power coefficient from 0.30 to 0.80 at an advance-diameter ratio of 2.40 had little effect on the propeller efficiency. A change in power coefficient from 0.40 to 1.00 at an advance-diameter ratio of 4.40 increased the propeller efficiency by about 40%. For conditions below the stall the thrust loading on the outboard blade sections increased more rapidly than on the inboard sections as the power coefficient was increased or as the advance-diameter ratio was decreased. For conditions beyond the stall, the thrust loading decreased on the outboard sections and increased on the inboard sections.

Description: 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. When the carburetor was used, any increase in equivalent volume flow was masked by choking in the carburetor. A small decrease in the peak adiabatic efficiency resulted from using the 12-vane diffuser. At the high volume flows where the supercharger is required to operate, the performance was improved by the 12-vane diffuser. With the vaneless discharge passage, the surge-free range of the supercharger was increased 35 percent at the lowest tip speed investigated by increasing the maximum air flow. The maximum air flow at high tip speeds was again limited by choking in the carburetor, which masked the effect of the vaneless discharge passage on the maximum air flow. At the high volume flows near the operating point of the supercharger, the performance with the vaneless discharge passage was better than that with the standard 11-vane diffuser. At the low volume flows when the standard 11-vane diffuser gave better performance. The modified scroll gave performance characteristics that were practically the same as those of the standard scroll except at high tip speeds, where the peak performance was improved.

Description: For the nvestigation of measuring instruments at higher speeds up to a Mach number 0.7 a tunnel with closed test section was built in 1942 which was as simple and cheap as possble. The blower was a radial blower with straight sheet vanes of 800-millimeter diameter the tips of which were bent backward a little. The blower sucks the air through a honeycomb of diameter 1.2 neter with wide meshes. The air is then accelerated in a short cone with smooth transition to the test section. The cylindrical test section of 200-milimeter diameter has two windows (which are displaced 180 deg from each other. The instruments may be introduced and observed through and observed through these windows. . The cross section is then enlarged by a straight diffuser 3.5 meters long and reaches the ninefold cross section. The air flows back into the room through a disk diffuser of 2-meter diameter. The maximum speed in the jet is 250 m/s for a drive power of 35 kT., if there are no installations in the jet. The velocity is determined by pressure holed along the test section.

Description: An altitude-wind-tunnel investigation has been made to determine the performance of Hamilton Standard 6507A-2 four-blade and three-blade propellers on a YP-47M airplane at high blade loadings and high engine powers. Characteristics of the four-blase propeller were obtained for a range of power coefficients from 0.10 to 1.00 at free-stream Mach numbers of 0.20, 0.30, 0.40. Characteristics of the three-blade propeller were obtained for a range of power coefficients from 0.30 to 1.00 at a free-stream Mach number of 0.40. Results of the force measurements indicate primarily the trend of propeller efficiency for changes in power coefficient or advance-diameter ratio because no corrections for the effects of tunnel-wall constriction on the installation were applied. Slipstream surveys are presented to illustrate blade thrust load distribution for certain operating conditions. Within the range of advance-diameter ratios investigated at each free-stream Mach number, the efficiency of the four-blade propeller decreased as the power coefficient was increased from 0.10 to 1.00. For the three-blade propeller, nearly constant maximum efficiencies were obtained for power coefficients from 0.32 to 0.63 at advance-diameter ratios between 1.90 and 3.00. In general, for conditions below the stall and critical tip Mach number, the maximum thrust load shifted from the inboard sections toward the tip sections as the power coefficient was increased or as the advance-diameter ratio was decreased. For conditions beyond the stall or critical tip Mach number, losses in thrust occurred on the outboard blade sections owing to flow break-down; the thrust load increased slightly on the inboard sections.

Description: 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 to decrease with increasing airfoil thickness ratio.

Description: 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 ot the production design. In addition the take-off and landing stability of the model with the alternate hull were satisfactory.

Description: Tests of a powered dynamic model of the Columbia XJL-1 amphibian were made in Langley tank no.1 to determine the hydrodynamic stability and spray characteristics of the basic hull and to investigate the effects of modifications on these characteristics. Modifications to the forebody chime flare, the step, and the afterbody, and an increase in the angle of incidence of the wing were included in the test program. The seaworthiness and spray characteristics were studied from simulated taxi runs in smooth and rough water. The trim limits of stability, the range of stable positions of the enter of gravity for take-off, and the landing stability were determined in smooth water. The aerodynamic lift, pitching moment, and thrust were determined at speeds up to take-off speed.

Description: The XF-12 airplane is a high performance, photo-reconnaissance aircraft designed by the Republic Aviation Corporation for Army Air Forces. A series of tests of a 1/8.33-scale powered model was conducted in the Langley 9-foot pressure tunnel to obtain information relative to the aerodynamic design of the airplane. This report presents the results of tests to determine the static longitudinal stability and stalling characteristics of the model. From this investigation it was indicated that the airplane will possess a positive static margin for all probable flight conditions. The stalling characteristics are considered satisfactory in that the stall initiates near the root section and progresses toward the tips. Early root section stalling occurs, with the flaps retracted and may cause undesirable tail buffeting and erratic elevator control in the normal flight range. From considerations of sinking speed landing flap deflections of 40 degrees may be preferable to 55 degrees of 65 degrees.

Description: Tests were performed on a partial span of the wing of a McDonnell XFD-1 airplane to determine a combination of sealed internal balance and spring-tab stiffness for the aileron that would give satisfactory stick-force characteristics for the airplane. Two sealed internal balances were tested in combination with spring tabs of various stiffnesses. One of the combinations was tested at several speeds to determine the variation of stick force with speed. Estimates, based on the results of the tests, indicate that for this airplane any reduction of stick force by use of the spring tab reduces the helix angle pb/2V below the required value of 0.09. The estimates show that, of the configurations tested, the most satisfactory combination for obtaining a stick force of 30 pounds at 300 miles per hour indicated airspeed is a 0.48-chord internal balance in combination with a spring-tab stiffness of 500 pounds per inch. With this combination, a wing-tip helix angle of 0.078 is estimated. Stick-force curves for all configurations show a rapid increase in stick force above approximately 20 deg. total aileron deflection.

Description: An investigation was conducted in the Cleveland altitude wind tunnel to determine the aerodynamic characteristics and the oil delivery critical altitude of the oil-cooler installation of an XTB2D-1 airplane. The investigation was made with the propeller removed end with the engine operating at 1800 brake horsepower, an altitude of 15,000 feet (except for tests of oil-delivery critical altitude), oil-cooler flap deflections from -20 degrees to 20 degrees and inclinations of the thrust axis of 0 degrees, 1.5 degrees, and 6 degrees. At an inclination of the thrust axis of 0 degrees and with the propeller operating, the total-pressure recovery coefficient at the face of the oil cooler varied from 0.84 to 1.10 depending on the flap deflection. With the propeller removed, the best pressure recovery at the face of the oil cooler was obtained at an inclination of the thrust axis of 1.5 degrees. Air-flow separation occurred on the inner surface of the upper lip of the oil-cooler duct inlet at an inclination of the thrust axis of 0 degrees and on the inner surface of the lower lip at 6 degrees. Static pressure coefficients over the duct lips were sufficiently low that no trouble from compressibility would be encountered in level flight. The oil-delivery critical altitude at cruising power (2230 rpm, 1675 bhp) was approximately 18,500 feet for the oil system tested.

Description: For two- and three-dimensional flow in a compressible medium, a simple relation is given by which, to a first approximation, the quantitative influence of compressibility upon the velocities and pressures can be understood in a clear manner. In the application of this relation the distinct behaviors of two-dimensional and axially symmetric three-dimensional flow with increasing Mach number are brought out. For slender elliptic cylinders and ellipsoids of revolution, calculations are made of the critical Mach number; that is, the Mach number at which local sonic velocity is achieved on the body. As a further example, the lifting wing of finite span is considered, and it is shown that the increase of wing lift with Mach number at a given angle of attack is greatly dependent upon the aspect ratio b(exp 2)/F.

Description: This paper presents the results of the aileron investigation and includes rolling-moment, yawing-moment, and aileron hinge-moment coefficients and pressure coefficients across the aileron-balance seal through a range of angle of attack, tab deflection, and aileron deflection with flaps neutral and deflected 20 degrees and 55 degrees. Some of the effects of wing roughness and balance seal leakage on the aileron and tab characteristics are also presented.

Description: By the method of images the horizontal steady motion of a wing at small heights above the ground was investigated in the wind tunnel, A rectangular wing with Clark Y-H profile was tested with and without flaps. The distance from the trailing edge of the wing to the ground was varied within the limits 0.75 less than or = s/c less than or = 0.25. Measurements were made of the lift, the drag, the pitching moment, and the pressure distribution at one section. For a wing without flaps and one with flaps a considereble decrease in the lift force and a,drop in the drag was obtained at angles of attack below stalling. The flow separation near the ground occurs at smaller angles of attack than is the case for a great height above the ground. At horizontal steady flight for practical values of the height above the ground the maximum lift coefficient for the wing without flaps changes little, but markedly decreases for the wing with flaps. Analysis of these phenomena involves the investigation of the pressure distribution. The pressure distribution curves showed that the changes occurring near the ground are not equivalent to a change in the angle of attack. At the lower surface of the section a very strong increase in the pressures is observed. The pressure changes on the upper surface at angles of attack below stalling are insignificant and lead mainly to an increase in the unfavorable pressure gradient, resulting in the earlier occurrence of separation. For a wing with flaps at large angles of attack for distances from the trailing edge of the flap to the ground less than 0.5 chord, the flow between the wing end the ground is retarded so greatly that the pressure coefficient at the lower surface of the section is very near its limiting value (P = 1), and any further possibility of increase in the pressure is very small. In the application an approximate computation procedure is given of the change of certain aerodynamic characteristics for horizontal steady flight near the ground.

Description: The results obtained from an evaluation for gust and draft velocities of acceleration and airspeed-altitude records taken by NACA recording instruments installed in P-61c airplanes participating in thunderstorm flights 6, 7, and 8 of July 9, 1946, July 10, 1946, and July 11, 1946, respectively, are presented herein. These data are summarized in tables I and II. In accordance with a recent discussion with a member of the U.S. Weather Bureau staff, the tabulated results for the present flight include in addition to data of the type presented in reference 1, the initial heading of the airplane for each traverse, the pressure altitude at the start of each traverse in increments of 500 feet, and the gust gradient distance when it could be evaluated. The cloud entry and exit times for the present data were taken from motion-picture records of the pilot's instrument panels whenever such records were available while the length fo the traverses in seconds and feet was taken from the airspeed-altitude records. In many cases, however, poor agreement is indicated between the duration of the cloud traverses as obtained from the motion-picture records and from the airspeed-altitude records. This result is believed to be due to camera stoppages, inaccurate spring mechanisms of the clocks, and loss of motion-picture record in exposure or development. With reference to the evaluation of gust data, the nominal threshold was about 2 feet per second. In making gust counts to this threshold, some gusts below that threshold have been included due to limitations of the procedure used. Thus, it will be noted that in some instances gust counts are given in table I although now corresponding gust velocities are listed.

Description: The results of an investigation of a 1/3-scale model of the Chance Vought XF5U-1 airplane in the Langley full-scale tunnel are presented in this report. The maximum lift and stalling characteristics of several model configurations, the longitudinal stability characteristics of the model, and the effectiveness of the control surfaces were determined with the propellers removed. The propulsive characteristics, the effect of propeller operation on the lift, and the static thrust of the model propellers were determined at several propeller-blade angles. The results with the propellers removed showed that the maximum lift coefficient of the complete model configuration was only 0.97 was compared with the value of 1.31 for the model configuration in which the engine-air ducts and canopy are removed. The model with the propellers removed (normal center-of-gravity position) has a positive static margin, stick fixed, varying from 5 to 13 percent of the mean aerodynamic chord throughout the unstalled range of lift coefficients. The unit horizontal tail is sufficiently powerful to trim the airplane with the propellers removed throughout the unstalled range of lift coefficients. The peak propulsive efficiencies for beta = 20 degrees and beta = 30 degrees were increased 7 percent at C(sub L) congruent to 0.67 and 20 percent at C(sub L) congruent to 0.74, respectively, with the propellers rotating upward in the center than with the propellers rotating downward in the center. Indications are that the minimum forward-flight speed of the airplane for full-power operation at sea level will be about 90 miles per hour. Decreasing the weight and increasing the power reduced this value of minimum speed and there were no indications from the results of a lower limit to the minimum speed.