ALBERT

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: 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: 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: 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: 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: 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: 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: 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: 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 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: 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: A flight investigation was made to determine the effect of distance flown in the icing region, antenna length, and antenna angle on the tension occurring in aircraft antennae while in regions of aircraft icing. The experimental antennas were of lengths ranging from 15 to 43 feet and were placed at angles of 0 deg to 64 deg with the airplane thrust axis. Distances up to 256 miles were flown in diverse icing conditions at true airspeeds from 157 to 214 miles per hour and pressure altitudes at which icing conditions were encountered. The results indicate that: The effect of ice formation on antenna tension increased with the angle of the antennas with the longitudinal axis of the airplane. The maximum tension for antennae having angles from 0 deg to 15 deg was 68 pounds, whereas the maximum tension for antennas having angles of 44 deg and 64 deg was 274 and 438 pounds, respectively.

Description: The sound field of a rotating propeller is teated theoretically on the basis of aerodynamic principles. For the lower harmonics, the directional characteristics and the radiated sound energy are determined and are in conformity with existing experimental results.

Description: Tests were made of a 1/18-scale dynamically similar model of the Lockheed Constellation airplane to investigate its ditching characteristics and proper ditching technique. Scale-strength bottoms were used to reproduce probable damage to the fuselage. The model was landed in calm water at the Langley tank no. 2 monorail. Various landing attitudes, speeds, and fuselage configuration were simulated. The behavior of the model was determined from visual observations, by recording the longitudinal decelerations, and by taking motion pictures of the ditchings. Data are presented in tabular form, sequence photographs, and time-history deceleration curves. It was concluded that the airplane should be ditched at a medium nose-high landing attitude with the landing flaps full down. The airplane will probably make a deep run with heavy spray and may even dive slightly. The fuselage will be damaged and leak substantially but in calm water probably will not flood rapidly. Maximum longitudinal decelerations in a calm-water ditching will be about 4g.

Description: Flight tests of a P-51H airplane with two different vertical-tail assemblies were made to determine lateral and directional stability and control characteristics. The airplane had satisfactory directional stability in the landing, approach, and wave-off conditions with either tail. In the power-on clean and glide conditions, however, the airplane had weak directional stability with the original tail. The production tail, which had a 7-inch fin extension and a shorter span rudder, improved the directional stability in the power-on clean and glide conditions, but the stability was still weak in the power-on clean condition. Increased altitude in either case caused a slight decrease in the stability. The rudder-trim-force change with speed with either vertical-tail assembly was high. The general aileron control characteristics were satisfactory but the aileron effectiveness failed to meet the Army handling-qualities requirements.

Description: An analysis of the estimated high-speed flying qualities of the Chance Vought XF7U-1 airplane in the Mach number range from 0.40 to 0.91 has been made, based on tests of an 0.08-scale model of this airplane in the Langley high-speed 7- by 10-foot wind tunnel. The analysis indicates longitudinal control-position instability at transonic speeds, but the accompanying trim changes are not large. Control-position maneuvering stability, however, is present for all speeds. Longitudinal lateral control appear adequate, but the damping of the short-period longitudinal and lateral oscillations at high altitudes is poor and may require artificial damping.

Description: A spin investigation has been conducted in the Langley 20 -foot free-spinning tunnel on a 1/29 - scale model of the Republic XP-91 airplane with vee tail installed. The effects cf control settings and movements upon the effect spin and recovery characteristics of the model were determined for the clean condition (wing tanks removed, landing gear and flaps retracted). The tests were made at a loading simulating that following cruise at altitude and at a time when nearly all fuel was expended. The results indicated that the airplane might not spin at normal spinning-control configuration, but if a spin were obtained, recovery therefrom by full rudder reversal would be satisfactory. It was also indicated that aileron-against settings would lead to violent oscillatory motions and should be avoided.

Description: The longitudinal stability and control characteristics of a B-29 airplane have been measured with a booster incorporated in the elevator control system. Tests were made to determine the effects on the handling qualities of the test airplane of variations in pilots control-force gradients as well as the effects of variations in the maximum rate of control motion supplied by the booster system.

Description: An investigation is being conducted to determine the altitude performance characteristics of the Nene II engine and its components. The present paper presents the preliminary results obtained using jet nozzle 18.00 inches in diameter, with an area equal to 92.2 percent of the area of the standard jet nozzle for this engine. The experimental results presented are for conditions simulating altitudes from 20,000 to 60,000 feet and ram-pressure ratios from 1.1 to 3.5. These ram-pressure ratios correspond to flight Mach numbers between 0.374 and 1.466. Data obtained with the 18.00 inch-diameter jet nozzle and corrected to standard sea-level conditions showed substantially the same trends with altitude as the data previously obtained with an 18.75-inch-diameter nozzle and with an 18.41-inch-diameter nozzle. Jet thrust, air consumption, and fuel consumption, corrected to standard sea-level conditions, increased rapidly with increasing ram-pressure ratio. In general, corrected net thrust specific fuel consumption increased with increase in ram-pressure ratio. Corrected net thrust decreased with an increase in ram-pressure ratio at an engine speed of 8000 rpm. At corrected engine speeds between 8000 and 10,800 rpm, net thrust first decreased with an increase in ram-pressure ratio and then increased with further increase in ram pressure ratio; at corrected engine speeds above 10,800 rpm, net thrust increased continuously with increase in ram-pressure ratio. Tail-pipe temperature decreased with an increase in ram-pressure ratio.

Description: An investigation is being conducted to determine the altitude performance characteristics of the British Nene II engine and its components. The present paper presents the preliminary results obtained using a standard jet nozzle. The test results presented are for conditions simulating altitudes from sea level to 60,000 feet and ram pressure ratios from 1.0 to 2.3. These ram pressure ratios correspond to flight Mach numbers between zero and 1.16 assuming a 100 percent ram recovery.

Description: Design data are presented for the graphical construction of two-dimensional sharp-edge-throat supersonic nozzles of minimum length for test-section Mach numbers from 1.20 to 10.00. The method of characteristics used in the design is briefly reviewed.

Description: An investigation of the pressure distribution on the fuselage nose and the pilot canopy of a supersonic airplane model has been conducted at a Mach number of 1.90 over a wide range of angles of attack and yaw. Boundary layer separation apparently occurred from the upper surface at angles of attack above 24 degrees and from the lower surface at minus 15 degrees. No separation from the sides of the fuselage was evident at yaw angles up to 12 degrees.

Description: An investigation has been conducted in the Langley 20-foot free-spinning tunnel to evaluate the spin, longitudinal-trim, and tumbling characteristics of a 1/20-scale model of the Consolidated Vultee MX-813 airplane. The effects of control position were determined for the model ballasted to represent the airplane in its design gross weight loading. The model, in general, would not spin but demonstrated a tendency to trim at very high stalled angles of attack. Static tests substantiated the dynamic tests as regards the trim characteristics. Movement of the elevator, however, from up to slightly down was effective in pitching the model from stalled to normal trim attitudes. The model would not tumble.

Description: A summary has been made of available data on the characteristics of airfoil sections with trailing-edge high-lift devices. Data for plain, split, and slotted flaps are collected and analyzed. The effects of each of the variables involved in the design of the various types of flap are examined and, in cases where sufficient data are given, optimum configurations are deduced. Wherever possible, the effects of airfoil section, Reynolds number, and leading-edge roughness are shown. For single and double slotted flaps, where a great mass of unrelated date are available, maximum lift coefficients of a large number of configurations are presented in tables.

Description: The drag coefficients of bombs at high velocities velocity of fall was 97 percent of the speed of sound) (the highest are determined by drop tests and compared with measurements taken in the DVL high-speed closed wind tunnel and the open jet at AVA - Gottingen.

Description: An investigation has been conducted in the Langley 20-foot free- spinning tunnel scale model of the Cornelius XFG-1 glider, a tailless design having its wings swept forward 15 degrees. It was previously found to possess erratic spin and recovery characteristics, and tests were made to determine modifications which would lead to normal steady spins with consistently good recoveries. The results of the investigation indicated that modifications that aid not appreciably alter the basic design aid not appreciably improve the spin and recovery characteristics. In this instance it appears that the sweptforward wing is the cause of unsatisfactory spin and recovery characteristics.

Description: An investigation of the XJ-41-V turbojet-engine compressor with a revised vaned collector was conducted to determine the performance of the compressor and to obtain fundamental information on the aerodynamic problems associated with large centrifugal compressors of this type. The original vaned collector was revised by increasing the flow area at the vaned collector entrance. A maximum adiabatic efficiency of 0.81 was obtained et a corrected weight flow of 36.5 pounds per second and a pressure ratio of 1.90. The peak pressure ratio was 3.93 and occurred at an impeller speed of 11,500 rpm at a corrected weight flow of 65.5 pounds per second. Revision of the vaned collector resulted in an increased airflow capacity over the speed range. The design air-flow capacity of 78 pounds per second was very nearly reached at the engine design speed of 11,500 rpm. The compressor air-flow choking point occurred in the vaned collector passage; however, at speeds above 8300 rpm, the air-flow capacity of the impeller was being approached as indicated by large pressure losses in the impeller at maximum air-flow conditions. An increase in compressor air-flow capacity at the higher speeds can possibly be obtained 5y removal of the flow restriction in the impeller, which would result in an increased air density at the vaned collector entrance.

Description: The losses in the inlet air ducts, the diffusers, and the de-icing equipment associated with turbojet engine installations cause a reduction in the total pressure at the inlet of the engine and result in reduced thrust and increased specific fuel consumption. An analytical evaluation of the effects of inlet losses on the net thrust and the fuel economy of a 3000-pound-thrust axial flow turbojet engine with a two-stage turbine is presented. The analysis is based on engine performance characteristics that were determined from experiments in the NACA Cleveland altitude wind tunnel. The experimental investigation did not include tests in which inlet losses were systematically varied, but the effects of these losses can be accurately estimated from the experimentally determined performance characteristics of the engine.

Description: Two rocket-powered models representative of a fighter-type airplane were investigated in flight at Mach numbers up to 1.01 and 1.07 by the Langley Pilotless Aircraft Research Division at its testing station at Wallops Island, Va. These models incorporated an inverse-taper wing and a vee tail and were flown with controls undeflected and wing and stabilizer set at 0 deg incidence. Values of lateral acceleration, normal acceleration velocity, and drag were obtained by use of telemeters and a Doppler velocimeter radar unit. The results of this investigation indicated no unusual variation in the lateral acceleration characteristics. After the cessation of powered flight, the lateral oscillation quickly damped to zero. The data indicated that the airplane, at low lift coefficients, should not experience any abrupt trim changes until it attains a Mach number of 0.97. The change in normal-force coefficient associated with this trim change will amount to about 0.03 with the center of gravity located at 4.48% of the mean aerodynamic chord. At higher lift coefficients, on the basis of other data, the Mach number at which this trim change occurs would be expected to be decreased. The neutral point of the model at Mach numbers near 1.05 was estimated to fall at 45% of the mean aerodynamic chord, assuming a lift-curve slope of 0.05. A value of the static-directional-stability parameter dCn/d(psi) of approximately -0.002 was estimated for a Mach number of 0.93. The values of drag coefficient obtained from both model flights were in a good comparative agreement. The highest drag coefficient occurred at a Mach number of 1.01 and was equal to 0.044.

Description: An investigation was conducted on a large centrifugal compressor from an experimental turbojet engine to determine the performance of the compressor and to obtain fundamental information on the aerodynamic problems associated with large centrifugal-type compressors. The results of the research conducted on the compressor indicated that the compressor would not meet the desired engine-design air-flow requirements (78 lb/sec) because of an air-flow restriction in the vaned collector (diffuser). Revision of the vaned collector resulted in an increased air-flow capacity over the speed range and showed improved matching of the impeller and diffuser components. At maximum flow, the original compressor utilized approximately 90 percent of the available geometric throat area at the vaned-collector inlet and the revised compressor utilized approximately 94 percent, regardless of impeller speed. The ratio of the maximum weight flows of the revised and original compressors were less than the ratio of effective critical throat areas of the two compressors because of the large pressure losses in the impeller near the impeller inelt and the difference increased with an increase in impeller speed. In order to further increase the pressure ratio and maximum weight flow of the compressor, the impeller must be modified to eliminate the pressure losses therein.

Description: Wind-tunnel tests of the McDonnell XP-85 airplane were conducted to determine its longitudinal, lateral, and directional stability and the characteristics of the aileron, the ruddervator, the leading-edge droop nose flap, and the stall control vanes. The directional stability of the airplane with numerous skyhook modifications and with a ventral fin was also investigated. The results of the tests showed that the effectiveness of the droop nose flaps and the stall control vanes was negligible with regard to either the maximum lift or longitudinal stability of the airplane. Contrary to any previous small-scale results, extension of the skyhook caused a 75-percent reduction in the directional stability of the airplane for both low and high values of lift coefficient. The simplest solution to the problem short of a major redesign of the skyhook appears to be the adoption of a ventral fin.

Description: The NACA is investigating a series of J-33 turbojet-engine compressors to determine the over-all and component performances and to improve theories of flow through large centrifugal compressors, The production model J-33-A-21 was operated over a range of inlet temperatures from 80 to -40 F and inlet pressures from 14 to 5 inches mercury absolute for equivalent impeller speeds from 6000 to 13,400 rpm. At the equivalent design speed of 11,500 rpm, the compressor had a peak pressure ratio of 3.98 at an equivalent weight flow of 73.4 pounds per second and an adiabatic temperature-rise , efficiency of 0.701. When the compressor speed was reduced from the design speed to 6000 rpm, the adiabatic temperature-rise efficiency increased to 0.747. At the maximum equivalent speed investigated (13,400 rpm), a peak pressure ratio of 5.09 was obtained at an adiabatic temperature-rise efficiency of 0.617 and an equivalent weight flow of 66.O pounds per second. An increase in inlet pressure from 5.5 to 14 inches mercury absolute, with a consequent increase in Reynolds number index, improved the pressure ratio but had no apparent effect on the ratio of temperature rise through the compressor to inlet temperature. The variation of the peak adiabatic temperature-rise efficiency with inlet pressure is in the direction that would be expected from a Reynolds number effect. Decrease in the inlet temperature from 80 to -40 F, with a consequent increase in Reynolds number index, resulted in scatter of the pressure-ratio data and increased values of temperature ratio. The variation of the adiabatic temperature-rise efficiency with inlet temperature is probably the result of heat-transfer effects and scatter in the pressure ratio.

Description: As part of an investigation of the performance and operational characteristics of the axial-flow gas turbine-propeller engine, conducted in the Cleveland altitude wind tunnel, the performance characteristics of the compressor and the turbine were obtained. The data presented were obtained at a compressor-inlet ram-pressure ratio of 1.00 for altitudes from 5000 to 35,000 feet, engine speeds from 8000 to 13,000 rpm, and turbine-inlet temperatures from 1400 to 2100 R. The highest compressor pressure ratio obtained was 6.15 at a corrected air flow of 23.7 pounds per second and a corrected turbine-inlet temperature of 2475 R. Peak adiabatic compressor efficiencies of about 77 percent were obtained near the value of corrected air flow corresponding to a corrected engine speed of 13,000 rpm. This maximum efficiency may be somewhat low, however, because of dirt accumulations on the compressor blades. A maximum adiabatic turbine efficiency of 81.5 percent was obtained at rated engine speed for all altitudes and turbine-inlet temperatures investigated.

Description: The investigations of the reports to 4 on wings of small aspect ratio are continued. The present report deals with the results of the three- and six-component measurements and the flow pictures of the triangular wing series with the aspect ratio Lambda = 3 to Lambda = 1.

Description: An equation is presented for calculating the heat flow required from the surface of an internally heated windshield in order to prevent the formation of ice accretions during flight in specified icing conditions. To ascertain the validity of the equation, comparison is made between calculated values of the heat required and measured values obtained for test windshields in actual flights in icing conditions. The test windshields were internally heated and provided data applicable to two common types of windshield configurations; namely the V-type and the type installed flush with the fuselage contours. These windshields were installed on a twin-engine cargo airplane and the icing flights were conducted over a large area of the United States during the winters of 1945-46 and 1946-47. In addition to the internally heated windshield investigation, some test data were obtained for a windshield ice-prevention system in which heated air was discharged into the windshield boundary layer. The general conclusions resulting from this investigation are as follows: 1) The amount of heat required for the prevention of ice accretions on both flush- and V-type windshields during flight in specified icing conditions can be calculated with a degree of accuracy suitable for design purposes. 2) A heat flow of 2000 to 2500 Btu per hour per square foot is required for complete and continuous protection of a V-type windshield in fight at speeds up to 300 miles per hour in a moderate cumulus icing condition. For the same degree of protection and the same speed range, a value of 1000 Btu per hour per square foot suffices in a moderate stratus icing condition. 3) A heat supply of 1000 Btu per hour per square foot is adequate for a flush windshield located well aft of the fuselage stagnation region, at speeds up to 300 miles per hour, for flight in both stratus and moderate cumulus icing conditions. 4) The external air discharge system of windshield thermal ice prevention is thermally inefficient and requires a heat supply approximately 20 times that required for an internal system having the same performance.

Description: The icing characteristics, the de-icing rate with hot air, and the effect of impact ice on fuel metering and mixture distribution have been determined in a laboratory investigation of that part of the engine induction system consisting of a three-barrel injection-type carburetor and a supercharger housing with spinner-type fuel injection from an 18-cylinder radial engine used on a large twin-engine cargo airplane. The induction system remained ice-free at carburetor-air temperatures above 36 F regardless of the moisture content of the air. Between carburetor-air temperatures of 32 F and 36 F with humidity ratios in excess of saturation, serious throttling ice formed in the carburetor because of expansion cooling of the air; at carburetor-air temperatures below 32 F with humidity ratios in excess of saturation, serious impact-ice formations occurred, Spinner-type fuel injection at the entrance to the supercharger and heating of the supercharger-inlet elbow and the guide vanes by the warn oil in the rear engine housing are design features that proved effective in eliminating fuel-evaporation icing and minimized the formation of throttling ice below the carburetor. Air-flow recovery time with fixed throttle was rapidly reduced as the inlet -air wet -bulb temperature was increased to 55 F; further temperature increase produced negligible improvement in recovery time. Larger ice formations and lower icing temperatures increased the time required to restore proper air flow at a given wet-bulb temperature. Impact-ice formations on the entrance screen and the top of the carburetor reduced the over-all fuel-air ratio and increased the spread between the over-all ratio and the fuel-air ratio of the individual cylinders. The normal spread of fuel-air ratio was increased from 0.020 to 0.028 when the left quarter of the entrance screen was blocked in a manner simulating the blocking resulting from ice formations released from upstream duct walls during hot-air de-icing.

Description: Flight tests were made in natural icing conditions with two 8-ft-chord heated airfoils of different sections. Measurements of meteorological variables conducive to ice formation were made simultaneously with the procurement of airfoil thermal data. The extent of knowledge on the meteorology of icing, the impingement of water drops on airfoil surfaces, and the processes of heat transfer and evaporation from a wetted airfoil surface have been increased to a point where the design of heated wings on a fundamental, wet-air basis now can be undertaken with reasonable certainty.

Description: An investigation has been made in the Langley two-dimensional low-turbulence pressure tunnel to develop the optimum configuration of a 0.35-chord slotted flap on an NACA 65(sub 1120)-111 airfoil section modified by removing the trailing-edge cusp. The section pitching-moment characteristics and the effects of standard roughness on the section characteristics were determined for the flap retracted at Reynolds numbers ranging from 3.0 x 10(exp 6) to 9.0 x 10(exp 6).

Description: Powered models of three different flying boats were landed in oncoming wave of various heights and lengths. The resulting motions and acceleration were recorded to survey the effects of varying the trim at landing, the deceleration after landing, and the size of the waves. One of the models had an unusually long afterbody. The data for landing with normal rates of deceleration indicated that the most severe motions and accelerations were likely to occur at some period of the landing run subsequent to the initial impact. Landings made at abnormally low trims led to unusually severe bounces during the runout. The least severe landing occurred after a small lending when the model was rapidly decelerated at about 0.4 g in a simulation of the proposed use of braking devices. The severity of the landings increased with wave height and was at a maximum when the wave length was of the order of from one and one-half to twice the over-all length of the model. The models with afterbodies of moderate length frequently bounced clear of the water into a stalled attitude at speeds below flying speed. The model with the long afterbody had less tendency to bounce from the waves and consequently showed less severe accelerations during the landing run than the models with moderate lengths of afterbody.

Description: Theoretical pressure distributions and measured lift, drag, and pitching moment characteristics at three values of Reynolds number are presented for a group of NACA four-digit-series airfoil sections modified for high-speed applications. The effectiveness of flaps applied to these airfoils and the effect of standard leading-edge roughness were also investigated at one value of Reynolds number. Results are also presented of tests of three conventional NACA four-digit-series airfoil sections.

Description: An investigation of the spin and recovery characteristics of a 1/24-scale model of the McDonnell XP-88 airplane has been conducted in the Langley 20-foot free-spinning tunnel. The effects of control settings and movements on the erect and inverted spin and recovery characteristics of the model in the normal loading were determined. Tests of the model in the long-range loading also were made. The investigation included tail-modification, spin-recovery parachute, pilot-escape, and rudder-pedal-force tests. Recoveries were generally satisfactory for spins in the normal loading provided the ailerons were not held against the spin. Satisfactory recoveries were obtained regardless of the aileron setting when the leading-edge flaps were deflected and normal recovery technique was used or when the horizontal tail was raised 70 inches, full scale. Recoveries were rapid from all inverted spins obtained. In the long-range loading with tanks on, it may be necessary to jettison the tanks in order to obtain recovery. A 12.0-foot spin-recovery parachute at the tail or a 4.0-foot parachute opened on the outer wing tip (drag coefficient of 0.66) was found to be effective for recoveries from demonstration spins. Test results showed that in an emergency the pilot should attempt to escape from the outboard side of the spinning airplane. The rudder-pedal forces in a spin were indicated to be within the capabilities of the pilot.

Description: From flight tests of 0.5-scale models of the Fairchild Lark pilotless aircraft conducted at the flight test station of the Pilotless Aircraft Research Division at Wallops Island, Va., some evaluations of the static longitudinal stability were obtained by analysis of the short-period oscillations induced by the abrupt movement of the rudder elevators. The analysis shows that for the Lark configuration with wing flap deflections of 0 degrees and 15 degrees the static longitudinal stability decreases slightly up to the critical Mach number and than as the Mach number increases further the stability increases greatly.

Description: At the request of the Air Materiel Command, Army Air Forces an investigation of the low-speed, power-off stability and control characteristics of the McDonnell XP-85 airplane is being conducted in the Langley free-flight tunnel. The XP-85 airplane is a parasite fighter carried in a bomb bay of the B-36 airplane. As a part of the investigation a few force tests were made of a 1/5 scale model of the XP-85 with a conventional tail assembly installed in place of the original design five-unit tail assembly. The total area of the conventional assembly was approximately 80 percent of the area of the five-unit assembly. The results of this investigation showed that the conventional tail assembly gave about the same longitudinal stability characteristics as the original configuration and improved the directional and lateral stability.

Description: Flight tests to determine propeller performance have been made of a Curtiss No. 838-102-18 three-blade propeller having trailing-edge extensions on a Republic P-47D-28 airplane in climb and high speed. These tests are a part of a general propeller flight-test program at the Langley Laboratory of the National Advisory Committee for Aeronautics. Results of climb tests indicate that when power is changed from approximately 1475 horsepower at 2550 rpm (roughly normal power) to 2400 horsepower at 2700 rpm (approximately military power) there is a loss in propeller efficiency of 3 percent at an altitude of 7000 feet, and 4 percent at 21,000 feet. At an airplane Mach number of 0.7 there is a gain of 9 percent in propeller efficiency when the power coefficient per blade is increased from 0.06 to 0.09. Optimum power coefficient per blade at this Mach number is estimated to be approximately 0.12. An analysis to determine the effect of the addition of extensions on the performance of the basic propeller blades indicates that climb performance was increased but high-speed performance was reduced. Both effects, however, were small.

Description: An investigation of two 1/14 scale model configurations of an outboard nacelle for the XB-36 airplane was made in the Langley two-dimensional low-turbulence tunnels over a range of airplane lift coefficients (C (sub L) = 0.409 to C(sub L) = 0.943) for three representative flow conditions. The purpose of the investigation was to develop a low-drag wing-nacelle pusher combination which incorporated an internal air-flow system. The present investigation has led to the development of a nacelle which had external drag coefficients of similar order of magnitude to those obtained previously from tests of an inboard nacelle configuration at the corresponding operating lift coefficients and from approximately one-third to one-half of those of conventional tractor designs having the same ratio of wing thickness to nacelle diameter.

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

Description: This report contains the flight-test results of the stalling characteristics measured during the flying-qualities investigation of the Lockheed P-8OA airplane (Army No. 44-85099). The tests were conducted in straight and turning flight with and without wing-tip tanks. These tests showed satisfactory stalling characteristics and adequate stall warning for all configurations and conditions tested.

Description: Qualitative investigations have shown that use of the NACA injection impeller with the R-3350 engine increases the inertia of the fuel-injection system and, when the standard fuel-metering system is used, this increase in inertia results in poor engine acceleration characteristics. This investigation was therefore undertaken to determine whether satisfactory acceleration characteristics of the engine equipped with the injection impeller could be obtained by simple modifications to the fuel-monitoring system. The engine was operated with two types of carburetor; namely, a hydraulic-metering carburetor incorporating a vacuum-operated accelerating pump and a direct-metering carburetor having a throttle-actuated accelerating pump. The vacuum-operated accelerating pump of the hydraulic-metering carburetor was modified to produce satisfactory accelerations by supplementing the standard air chamber with an additional 75-cubic spring. The throttle-actuated accelerating pump of the direct-metering carburetor was modified to produce satisfactory accelerations by replacing the standard 0.028-inch-diameter bleed in the load-compensator balance line with a smaller bleed of 0.0225-inch diameter. The results of this investigation indicated that both carburetors can be easily modified to produce satisfactory acceleration characteristics of the engine and no definite choice between the types of carburetor and accelerating pump can be made. Use of the direct-metering carburetor, however, probably resulted in better fuel distribution to the cylinders during the acceleration period and reduced the backfire hazard because all the fuel is introduced through the injection impeller.

Description: 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.

Description: Attempts were made to alleviate the buffeting of external fuel tanks mounted under the wings of a twin-engine Navy fighter plane. The Mach number at which the buffeting began was increased from 0.529 to 0.640 by streamlining the sway braces and increasing the lateral rigidity of the sway brace system. Further increases of the Mach number, at which buffeting began to 0.725, was obtained by moving the external fuel tank to a position under the fuselage.

Description: This report contains the flight-test results of the lateral and directional-stability and control phase (including tests with wing-tip tanks) of a general flying-qualities investigation of the Lockheed P-80A airplane (Army No. 44-85099). These tests were conducted at indicated airspeeds up to 494 miles per hour (0.691 Mach number) at low altitude and up to 378 miles per hour (0.816 Mach number) at high altitude. These tests showed that the flying qualities of the airplane were for the most part in accordance with the requirements of the Army Air Forces Stability and Control Specifications. The only major deficiency noted was the negative lateral stability with the wing-tip tanks installed.

Description: The NACA 23012-4 airfoil was investigated for the purpose of increasing lift by means of blowing out air from the wing, in conjunction with the effect of plain flap of variable contour and slotted flap of 25 percent chord length. The wing also was provided with a hinged nose, to be deflected at will. Air was blown out frcm the wing immediately in front of the flap; also at the opening between wing and hinged nose,tangentially to the surface of the wing. Another device employed to increase maximum lift was a movable slat, to be opened to form a clot. Lift was measured in relation to the volume of blown-out air and considerable increases were observed with increasing volume.

Description: Flight and ground investigations have been made to compare an exhaust-ejector installation with a standard exhaust-collector-ring installation on air-cooled aircraft engines in a twin-engine airplane. The ground investigation allowed that, whereas the standard engine would have overheated above 600 horsepower, the engine with exhaust ejectors cooled at take-off operating conditions at zero ram. The exhaust ejectors provided as much cooling with cowl flaps closed as the conventional cowl flaps induced when full open at low airspeeds. The propulsive thrust of the exhaust-ejector installation was calculated to be slightly less than the thrust of the collector-ring-installation.

Description: The basic principles of the control of TL ongincs are developed on .the basis of a quantitative investigation of the behavior of these behavior under various operating conditions with particular consideration of the simplifications pormissible in each case. Various possible means of control of jet engines are suggested and are illustrated by schematic designs.

Description: 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. It appears, therefore, that the oscillatory motions reported for the airplane may have been the direct result of control movements rather than the result of inherent oscillatory tendencies.

Description: The spin and recovery characteristics of the Curtiss-Wright XP-87 airplane, as well as the spin-recovery parachute requirements, the control forces that would be encountered in the spin, and the best method for the crew to attempt an emergency escape, are presented in this report. The characteristics were estimated rather than determined by model tests because the XP-87 dimensional and mass characteristics were considered to be noncritical and because data were available from model tests of several similar airplanes. The study indicated that the recovery characteristics of the airplane will be satisfactory for all loadings if the controls are reversed fully and rapidly. The control forces, however, will probably be beyond the capabilities of the pilot unless some additional balance or a booster is used. A 6-foot tail parachute or a 3.5-foot wing-tip parachute with a drag coefficient of 0.7 will be a satisfactory, emergency spin-recovery device for spin demonstrations. If it is necessary for the crew to abandon the spinning airplane, they should leave from the outboard side of the cockpit.

Description: The results obtained from measurements of gust and draft velocities within thunderstorms for the period July 22, 1946 to July 23, 1946 at Orlando, Florida, are presented herein. These data are summarized in tables I and II, respectively, and are of the type presented in reference 1 for previous flights. Inspection of photo-observer records for the flights indicated that no data on ambient air temperature variations within thunderstorms were obtained.

Description: Tests with a dynamically similar model of the Navy XP4M-1 airplane were made to determine the best way to land the airplane in calm and rough water, to determine its probable ditching performance, and to determine practicable modifications which could be incorporated in the design of the airplane that would improve its ditching characteristics. The results were obtained by making visual observations, by recording longitudinal decelerations,a nd by taking motion pictures of the landings. A list of conclusions from the test results is included.

Description: An investigation has been made in the Langley two-dimensional low-turbulence tunnel to develop the optimum configuration of a 0.35-chord slotted flap on an NACA 65 (112)-111 airfoil section modified by removing the trailing-edge cusp. The results of the investigation indicate that for the optimum configuration at a Reynolds number of 2.4 x 10(exp 6), the flap deflection was 45 degrees and the flap leading-edge radius center was 0.73 percent-chord behind and 4.46 percent-chord below the slot lip. The maximum section lift coefficient for the optimum configuration at a Reynolds number of 2.4 x 10(exp 6) was 2.46 or 0.12 higher than that obtained for an NACA 65-210 airfoil section with a 0.250-chord slotted flap.

Description: This report presents the result of tests of a 0.35-scale model of the Bell P-39N-l airplane. Included are the longitudinal-stability and - control characteristics of the airplane as indicated by tests of the model equipped with each of two different sets of elevators. The results indicate good longitudinal stability and control throughout the speed range encounterable in flight. The variation of estimated stick force with speed was less when the model was equipped with elevators constructed to the theoretical design dimensions than when equipped with elevators as built to scale from measurements of the corresponding-parts of the actual airplane. The predicted stick forces required to produce the normal accelerations attainable in flight are within the limits specified by the Army Air Forces.

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

Description: A preliminary evaluation of the spin and recovery characteristics of the XF3D-1 airplane has been made, based primarily on the results of the free-spinning tunnel tests of a model which closely simulated the XF3D-1 in tail design, tail length, and mass loading. Estimates have been made of the rudder-pedal force that may be encountered in effecting recovery from a spin and of the spin recovery parachute requirements of the airplane for demonstration spins. The method of bail-out which should be used if it becomes necessary for the crew to abandon the airplane during a spin is indicated. It was indicated that the recovery characteristics of the XF3D-1 airplane in the clean condition for erect and inverted spins would be satisfactory for all loadings specified by the contractor as possible on the airplane. However, if a spin is inadvertently entered while the landing flaps are down, recovery may be slow. The slow-down brakes and the landing flaps should be retracted immediately upon the inception of a spinning condition, after which recovery from the spin should be attempted. The pedal force necessary to reverse the rudder during a spin will be within the physical capabilities of the pilot. Opening a 10-foot diameter parachute attached to the tail (laid-out-flat diameter, drag coefficient 0.7) or a 4.5-foot diameter parachute attached to the outboard wing tip will insure satisfactory spin recovery from demonstration spins. If it becomes necessary for the crew to abandon the airplane during a spin, they should leave from the outboard side of the cockpit.

Description: 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). Maximum normal accelerations of about 6.5g were obtained in 4.4 inch waves and 8.5g were obtained in 8.0 inch waves. A maximum angular acceleration corresponding to 16 radians per second per second, full size, was obtained in the higher waves. The data indicate that the airplane will experience its greatest accelerations when landing in waves of about 20 feet (140 ft, full size) in length.

Description: In the following, high-speed measurements on a swept-back wing are reported. The curves of lift, moment, and drag have been determined up to Mach numbers of M = 0.87, and they are compared to a rectangular wing. Through measurements of the total-head loss behind the wing and through schlieren pictures, an insight into the formation of the compression shock at high Mach numbers has been obtained.

Description: 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.

Description: A series of investigations of several 1/14-scale models of an inboard nacelle for the XB-36 airplane was made in the Langley two-dimensional low-turbulence tunnels. The purpose of these investigations was to develop a low-drag wing-nacelle pusher combination which incorporated an internal air-flow system. As a result of these investigations, a nacelle was developed which had external drag coefficients considerably lower than the original basic form with the external nacelle drag approximately one-half to two-thirds of those of conventional tractor designs. The largest reductions in drag resulted from sealing the gaps between the wing flaps and nacelle, reducing the thickness of the nacelle training-edge lip, and bringing the under-wing air inlet to the wing leading edge. It was found that without the engine cooling fan adequate cooling air would be available for all conditions of flight except for cruise and climb at 40,000 feet. Sufficient oil cooling at an altitude of 40,000 feet may be obtained by the use of flap-type exit doors.

Description: A 1/5.5-size powered dynamic model of the Columbia XJL-1 amphibian was landed in Langley tank no. 1 in smooth water and in oncoming waves of heights from 2.1 feet to 6.4 feet (full-size) and lengths from 50 feet to 264 feet (full-size). The motions and the vertical accelerations of the model were continuously recorded. The greatest vertical acceleration measured during the smooth-water landings was 3.1g. During landings in rough water the greatest vertical acceleration measured was 15.4g, for a landing in 6.4-foot by 165-foot waves. The impact accelerations increased with increase in wave height and, in general, decreased with increase in wave length. During the landings in waves the model bounced into the air at stalled attitudes at speeds below flying speed. The model trimmed up to the mechanical trim stop (20 deg) during landings in waves of heights greater than 2.0 feet. Solid water came over the bow and damaged the propeller during one landing in 6.4-foot waves. The vertical acceleration coefficients at first impact from the tank tests of a 1/5.5-size model were in fair agreement with data obtained at the Langley impact basin during tests of a 1/2-size model of the hull.

Description: Tests were made with a 1/16 size dynamically similar model of the Navy XP2V-1 airplane to study its performance when ditched. The model was ditched in calm water at the Langley tank no. 2 monorail. Various landing attitudes, speeds, and conditions of damage were simulated. The performance of the node1 was determined and recorded from visual observations, by recording time histories of the longitudinal decelerations, and by taking motion pictures of the ditchings From the results of the tests with the model the following conclusions were drawn: 1. The airplane should be ditched at the normal landing attitude. The flaps should be fully extended to obtain the lowest possible landing speed; 2. Extensive damage will occur in a ditching and the airplane probably will dive violently after a run of about 2 fuselage lengths. Maximum longitudinal decelerations up to about 4g will be encountered; and 3. If a trapezoidal hydroflap 4 feet by 2 feet by 1 foot is attached to the airplane at station 192.4, diving will be prevented and the airplane will probably porpoise in a run of about 4 fuselage lengths with a maximum longitudinal deceleration of less than 3.5g.

Description: An investigation of the air-stream fluctuations at the tail of the D-558-1 airplane has been made at high speed for the purpose of determining the vertical region in which the horizontal tail may be placed without becoming subject to tail buffeting. The investigation was made for a range of Mach numbers from 0.775 to 0.907, and a range of vertical positions at the tall to include two proposed horizontal-tail positions. The tests were made at two angles of attack, 0,2 deg. and 4.2 deg., representative, of the angles of attack for high-speed level flight and a pull-out condition.

Description: This report contains the results of tests of a 1/3-scale model of the Lockheed YP-90A "Shooting Star" airplane and a comparison of drag, maximum lift coefficient, and elevator angle required for level flight as measured in the wind tunnel and in flight. Included in the report are the general aerodynamic characteristics of the model and of two types of dive-recovery flaps, one at several positions along the chord on the lower surface of the wing and the other on the lower surface of the fuselage. The results show good agreement between the flight and wind-tunnel measurements at all Mach numbers. The results indicate that the YP-80A is controllable in pitch by the elevators to a Mach number of at least 0.85. The fuselage dive-recovery flaps are effective for producing a climbing moment and increasing the drag at Mach numbers up to at least 0.8. The wing dive-recovery flaps are most effective for producing a climbing moment at 0.75 Mach number. At 0.85 Mach number, their effectiveness is approximately 50 percent of the maximum. The optimum position for the wing dive-recovery flaps to produce a climbing moment is at approximately 35 percent of the chord.

Description: Spin tests of a 1/16-scale model of the Chance Vought XF5U-1 airplane have been performed in the Langley 20-foot free-spinning tunnel. The effect of control position and movement upon the erect and inverted spin and recovery characteristics ae well as the effects of propellers, of stability flaps, and of various revisions to the design configuration have been determined for the normal fighter loading. The investigation also included spin recovery parachute, tumbling, and pilot-escape tests. For the original design configuration, with or without windmilling propellers, the recovery characteristics of the model were considered unsatisfactory. Increasing the maximum upward deflection of the ailavators from 45 deg to 65 deg resulted in greatly improved recovery characteristics. Dimensional revisions to the original airplane configuration, which satisfactorily improved the general spin and recovery characteristics of the model, consisted of: (1) a supplementary vertical tail 34 inches by 59 inches (full-scale) attached to a boom 80 inches aft of the trailing edge of the airplane in the plane of symmetry, (2) a large semispan undersurface spoiler placed along the airplane quarter-chord line and opened on the outboard side in a spin, or (3) two additional vertical tails 64 inches by 52 inches (full-scale) located at the tips of the ailavators. A satisfactory parachute arrangement for emergency spin recovery from demonstration spins was found to be an arrangement consisting of a 13.3-foot parachute attached by a 30-foot towline to the arresting gear mast on the airplane and opened simultaneously with an 8-foot parachute on the outboard end of the wing attached by a 3-foot towline. Tests indicated that pilot escape from a spin would be extremely hazardous unless the pilot is mechanically ejected from the cockpit. Model tumbling tests indicated that the airplane would not tumble.

Description: Spin tests have been conducted in the Langley free-spinning tunnel on a 1/16-scale model of the McDonnell XP-85 airplane with the normal X-tail replaced with a short-coupled conventional tail arrangement. The effect of the conventional tail arrangement and the effects of various modifications upon the spin and recovery characteristics of the model were determined. The results of the tests indicated that installation of the conventional tail arrangement wil not provide satisfactory recoveries from spins of the airplane. Satisfactory recoveries will be obtainable, however, either by installing in addition a very large ventral fin (17.94 sq ft, full-scale) below the tail or by decreasing the width of the fuselage and making it flat sided rearward of the wing trailing edge.

Description: This report presents the results of wind-tunnel tests of a 0.22-scale model of the North American XP-82 airplane with several modifications designed to reduce the buffeting of the airplane. The effects of various modifications on the air flow over the model are shown by means of photographs of tufts. The drag, lift, and pitching-moment coefficients of the model with several of the modifications are shown. The result indicate that, by reflexing the trailing edge of the center section of the wing and modifying the radiator air-scoop gutter and the inboard lower-surface wing fillets, the start of buffeting can be delayed from a Mach number of 0.70 to 0.775, and that the diving tendency of the airplane would be eliminated up to a Mach number of 0.80.

Description: Measurements are described which were taken in the large wind tunnel of the AVA on a rectangular wing "Mustang 2" with nose flap of a chord of 10 percent. Besides force measurements the results of pressure-distribution measurements are given and compared with those on the same profile "without" nose flap.

Description: The results obtained from measurements of gust velocities, draft velocities, and ambient-air temperature within thunderstorms for the period from September 11, 1946 to September 16, 1946 at Orlando, Florida are presented herein. These data are summarized in.and presented.

Description: This report presents comprehensive pressure-distribution measurements on four (4) swept-back wings (phi = 0 deg, 15 deg, 30 deg, and 45 deg) of constant chord and over a large range of angles of attack with symmetrical air flow. The distributions, experimentally obtained, were compared with theoretical ones calculated by the methods of Weissinger and Multhopp.

Description: A low-speed investigation in the Langley propeller-research tunnel of annular air inlets designed to avoid compression shocks and attendant boundary-layer separation on the fuselage ahead of the inlets at transonic flight speeds by maintaining substream flow velocities on the fuselage nose was reported in NACA RM No. L6J04. In the present investigation, one of the original annular inlets was converted by the installation of a canopy and a nose-wheel fairing into a twin side inlet in order to study problems involved in applying such an inlet to a fighter-type airplane. Extensive measurements of pressures on the surface of the model and surveys of the internal flow were conducted at angles of attack of 0 degrees, 3 degrees, and 6 degrees over a wide range of inlet-velocity ratio.

Description: Tests of two-blade propellers having the NACA 4-(3)(06.3)-06 and NACA 4-(3)(06.4)-09 blade designs (blade activity factors of 179 and 263, respectively) have been made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 70 degrees for free-stream Mach numbers from 0.165 to 0.725 to determine the effects of high solidity and compressibility on propeller characteristics. The tests are part of a general investigation of propellers at high forward speeds. Results previously reported for similar tests of two-blade propellers having the NACA 4-308-03 and NACA 4-308-045 blade designs (blade activity factors of 87 and 133, respectively) are included for comparison. The results showed that the 0.06- and 0.09-solidity blades, although producing efficiencies of the order of 90 percent, were less efficient than blades of conventional solidity. The variation in average blade lift coefficient with solidity at a constant blade angle and advance-diameter ratio through the speed range of these tests was found to be analogous to the variation of wing lift coefficient with aspect ratio, indicating that high-solidity blades may be desirable at very high speeds. Because of power limitations of the test equipment, conclusive evidence of the possible favorable effects of increased blade solidity at high speeds was not obtained. Further tests are desirable.

Description: In accordance with a request of the Bureau of Aeronautics, Navy Department, tests were performed in the Langley 20-foot free-spinning tunnel to determine the spin and recovery characteristics of a 1/24 scale model of the Grumman XTB3F-1 airplane. The airplane is a two-place, midwing torpedo bomber equipped with a tractor propeller and an auxiliary jet engine. The effect of control setting and movement on the erect and inverted spin and recovery characteristics of the model were determined for the normal loading. Brief tests with mass extended slightly along the fuselage were also made, however, in order to determine the effect of such a mass variation on elevator effectiveness. Tests were performed to determine the size of emergency spin-recovery tail and wing-tip parachutes required for satisfactory recovery by parachute action alone. The investigation also included emergency pilot-escape tests and tests to determine the rudder pedal and elevator stick forces necessary to move the rudder and elevator for recovery.

Description: The probable spin and recovery characteristics of the XSN2J-1 air-plane have been estimated on the basis of the results of brief test a performed on a model of an airplane of somewhat similar design. The spin-recovery tail-parachute requirements for the airplane were also determined end, in addition, an analysis was made to determine the best method of emergency pilot escape during a spin. The results of the investigation indicate that the recovery characteristics of the airplane will be satisfactory for all probable loading conditions of the airplane. A 6-foot-diameter tall parachute attached to a 30-foot tow-line will be satisfactory as a spin-recovery device for emergency recovery from demonstration spins. If the occupants of the airplane decide to abandon the airplane in a spin, they should leave the airplane from the outboard side of the cockpit and as far rearward as possible.

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

Description: As one phase of a comprehensive canopy load investigation, conventional front and rear sliding canopies which are typified by installation on the SB2C-4E airplane, were tested in the Langley full-scale tunnel to determine the pressure distributions and the aerodynamic loads on the canopies. A preliminary analysis of the results of these tests is presented in this report. Plots are presented that show the distribution of pressure at four longitudinal stations through each canopy for a range of conditions selected to determine the effects of varying canopy position, yaw, lift coefficient, and power. The results indicate that the maximum loads, based on the external-internal pressure differential, for the front and rear canopies were obtained with the airplane simulating the high speed flight condition. The highest loading on the front canopy was in the exploding direction for the configuration with the front and rear canopies closed. The highest loads on the rear canopy were in the crushing direction with the front canopy open and the rear canopy closed. For most of the simulated flight conditions, the highest loads on the front canopy, per unit area, were over twice as great as the highest loads on the rear canopy when the comparison was made for the most critical canopy configuration in each case. The external pressure distribution over the front and rear canopies, which were fairly symmetrical to 0 degree angle of yaw, were greatly distorted at other yaw attitudes, particularly for the propeller operating conditions. These distorted pressure distributions resulted in local exploding and crushing loads on both canopies which were often considerably higher than the average canopy loads.

Description: A total of 197 hours of time histories of airspeed and altitude have been obtained on Lockheed Constellation airplanes flying between New York, N.Y. and San Francisco, Calif. during May 1946 and June 1946. Data for 130 hours were previously analyzed to determine the probability of attaining excessive airspeeds and Mach numbers and the results have been published. After the publication, data for additional 67 hours became available. All the data were obtained at altitudes less than 20,000 feet and under approximately the same conditions. The combined data have been analyzed to obtain the results given in the tables and figures contained herein. Based on the combined data, the probability of exceeding a given airspeed or Mach number is of the same order of magnitude as that given. The conclusions are, therefore, unchanged.

Description: An analysis has been made of airfoil data taken on several NACA 16-series propeller airfoils from tests of 5-inch-chord models in the Langley 24 inch high-speed tunnel and l2-inch-chord models in the Langley 8 foot high-speed tunnel, This analysis has shown that the combined effects of Reynolds number changes and variations in airfoil characteristics resulting from differences in models and tunnels are such that when 5 inch-chord and l2-inch-chord data are applied to full-scale propeller design at or near the design condition, differences of less than 1 percent in efficiency will be involved.