ALBERT

Beschreibung: Estimating method for lift interference of wing- body combinations at supersonic speeds

Beschreibung: Conference on aerodynamics of high speed aircraft

Beschreibung: Drag measurements at low lift of four-nacelle aircraft configuration with longitudinal distribution of cross-sectional area conducive to low transonic drag rise

Beschreibung: During the flight program on the Bell X-5 airplane with 59 deg sweepback to determine the practical Mach number and normal-force coefficient limits of this configuration, a reduction in static longitudinal stability was encountered in maneuvering flight. A determination of the boundary for reduction of longitudinal stability extending to a Mach number of 0.98 is presented in this paper. A reduction of static longitudinal stability existed for all elevator and all stabilizer-executed maneuvers. The reduction of stability existed for maneuvers executed with elevator near a normal-force coefficient of 0.6 for a Mach number range of about 0.31 to 0.76. Above a Mach number of 0.76 the normal-force coefficient for reduction of stability gradually decreased to a value of 0.2 at a Mach number of 0.98. For stabilizer-executed maneuvers the stability boundary was the same as for elevator maneuvers up to a Mach number of 0.88. Above this Mach number the reduction of stability occurred at slightly higher normal-force coefficients decreasing from about 0.51 at a Mach number of 0.92 to a value of 0.311 at a Mach number of 0.97. The airplane has been flown to a Mach number of 1.04 at a normal-force coefficient of about 0.15 without encountering any reduction of stability. The pilot did not consider the reduction of stability to be dangerous at altitudes above 30,000 feet; however, precise flight was impossible. At angles of attack above that at which the reduction of longitudinal stability occurred, directional instability and aileron control overbalance were encountered.

Beschreibung: During the acceptance tests of the Bell X-5 airplane, measurements of the static stability and control characteristics and horizontal-tail loads were obtained by the NACA High-Speed Flight Research Station. The results of the stability and control measurements are presented in this paper. A change in sweep angle between 20 deg and 59 deg had a minor effect on the longitudinal trim, with a maximum change of about 2.5 deg in elevator deflection being required at a Mach number near 0.85; however, sweeping the wings produced a total stick-force change of about 40 pounds. At low Mach numbers there was a rapid increase in stability at high normal-force coefficients for both 20 0 and 1100 sweepback, whereas a condition of neutral stability existed for 58 0 sweepback at high normal-force coefficients. At Mach numbers near 0.8 there was an instability at normal-force coefficients above 0.5 for all sweep angles tested. In the low normal-force-coefficient range a high degree of stability resulted in high stick forces which limited the maximum load factors attainable in the demonstration flights to values under 5g for all sweep angles at a Mach number near 0.8 and an altitude of 12,000 feet. The aileron effectiveness at 200 sweepback was found to be low over the Mach number range tested.

Beschreibung: Flight measurements of the stability characteristics of the Bell X-5 research airplane at 59 deg sweepback were made in steady sideslips at Mach numbers from 0.62 to 0.97 at altitudes ranging between 35,000 and 40,000 feet. The results showed that the apparent directional stability was positive and increased at Mach numbers above 0.90. The apparent effective dihedral was positive and high, increasing at Mach numbers above 0.75. The cross-wind force coefficient per degree of sideslip was positive and increased rapidly at Mach numbers above 0.94.

Beschreibung: An experimental investigation was made of a preloaded spring-tab flutter model to determine the effects on flutter speed of aspect ratio, tab frequency, and preloaded spring constant. The rudder was mass-balanced, and the flutter mode studied was essentially one of three degrees of freedom (fin bending coupled with rudder and tab oscillations). Inasmuch as the spring was preloaded, the tab-spring system was a nonlinear one. Frequency of the tab was the most significant parameter in this study, and an increase in flutter speed with increasing frequency is indicated. At a given frequency, the tab of high aspect ratio is shown to have a slightly lower flutter speed than the one of low aspect ratio. Because the frequency of the preloaded spring tab was found to vary radically with amplitude, the flutter speed decreased with increase in initial displacement of the tab.

Beschreibung: Low-speed tests of a pilotless aircraft were conducted in the Langley propeller-research tunnel to provide information for the estimation of the longitudinal stability and. control, to measure the aileron effectiveness, and to calibrate the radome and the Machmeter pitot-static orifices. It was found that the model possessed a stEb.le variation of elevator angle required for trim throughout the speed range at the design angle of attack. A comparison of the airplane with and without JATO units and with an alternate rocket booster showed that a large loss in longitudinal stability and control resulting from the addition of the rocket booster to the aircraft was sufficient to make the rocket-booster assembly unsatisfactory as an alternate for the JATO units. Reversal of the aileron effectiveness was evident at positive deflections of the vertical wing flap indicating that the roll-stabilization system would produce roiling moments in a tight right turn contrary to its design purpose. Vertical-wing-flap deflections caused large errors in the static-pressure reading obtained by the original static-tube installation. A practical installation point on the fuselage was located which should yield reliable measurement of the free-stream static pressure.

Beschreibung: A wind-tunnel investigation has been made to determine the effects of unsymmetrical horizontal-tail arrangements on the power-on static longitudinal stability of a single-engine single-rotation airplane model. Although the tests and analyses showed that extreme asymmetry in the horizontal tail indicated a reduction in power effects on longitudinal stability for single-engine single-rotation airplanes, the particular "practical" arrangement tested did not show marked improvement. Differences in average downwash between the normal tail arrangement and various other tail arrangements estimated from computed values of propeller-slipstream rotation agreed with values estimated from pitching-moment test data for the flaps-up condition (low thrust and torque) and disagreed for the flaps-down condition (high thrust and torque). This disagreement indicated the necessity for continued research to determine the characteristics of the slip-stream behind various propeller-fuselage-wing combinations. Out-of-trim lateral forces and moments of the unsymmetrical tail arrangements that were best from consideration of longitudinal stability were no greater than those of the normal tail arrangement.

Beschreibung: Several groups of new airfoil sections, designated as the NACA 8-series, are derived analytically to have lift characteristics at supercritical Mach numbers which are favorable in the sense that the abrupt loss of lift, characteristic of the usual airfoil section at Mach numbers above the critical, is avoided. Aerodynamic characteristics determined, from two-dimensional windtunnel tests at Mach numbers up to approximately 0.9 are presented for each of the derived airfoils. Comparisons are made between the characteristics of these airfoils and the corresponding characteristics of representative NPiCA 6-series airfoils. The experimental results confirm the design expectations in demonstrating for the NACA S-series airfoils either no variation, or an Increase from the low-speed design value, In the lift coefficient at a constant angle of attack with increasing Mach number above the critical. It was not found possible to improve the variation with Mach number of the slope of the lift curve for these airfoils above that for the NACA 6-series airfoils. The drag characteristics of the new airfoils are somewhat inferior to those of the NACA 6- series with respect to divergence with Mach number, but the pitching-moment characteristics are more favorable for the thinner new sections In demonstrating somewhat smaller variations of moment coefficient with both angle of attack and Mach number. The effect on the aero&ynamic characteristics at high Mach numbers of removing the cusp from the trailing-edge regions of two 10-percent-chord-thick NACA 6-series airfoils is determined to be negligible.

Beschreibung: Tests of a partial-span model of a large bomber-type air1ane were conducted to determine the. aerodynamic characteristics of the wing equipped with full-span flaps and a retractable spoiler end aileron lateral control system. The arrangement consisted of (1) a double slotted flap extending over aproximate1y 86 percent of the wing semispan, (2) a 20-percent constant-percentage-chord aileron extending from the outboard end of the flap to the wing tip, and (3) a retractable spoiler, located at the 65-percent wing-chord station and extending from approximately 63 percent of the wing semispan to the wing tip. In addition, tests were made of a wing vent (of 1 and 2 percent of the wing chord located directly behind the spoiler), perforations in the spoiler, a blot or cut-out along the lower edge of the spoiler and spoilers of various spans. With full-span flaps deflected and with the 2-percent vent open or closed the initial stalling of the wing occurred at the tips, but with the vents closed there probably would be no appreciable loss in lateral control until maximum lift was reached. The l-percent vent increased the rolling effectiveness of the spoiler at small spoi1er deflections, particularly at high angles of attack with flaps deflected. With flaps deflected the 2-percent vent caused a large reduction in both the wing lift and rolling effectiveness of the spoiler at large angles of attack. However, at small angle of attack the 2-percent vent increased the rolling effectiveness of the spoiler at small spoiler deflections. The simultaneous operation of the spoiler and vent (in contrast to a vent fixed in the wing) would result in a large increase in the effectiveness of the spoiler and would avoid any loss in wing lift as in a fixed vent arrangement. The tests of the spoiler modifications revealed that (1) the spoiler perforations reduced the rolling-moment and yawing-moment coefficients but caused the spoiler hinge-moment coefficients to become more positive; (2) the spoiler slot had no notable effect on the rolling-moment and yawing-moment characteristics but produced a positive increase in the spoiler hinge-moment coefficients at large spoiler deflections; (3) the effects produced by the individual modifications were additive when the various modifications were combined. In general, progressively decreasing the spoiler span by removing the segments from the inboard end of the spoiler caused a decrease in rolling effectiveness approximately proportional to the span of the segment.

Beschreibung: The results of a theoretical analysis of the hinge-moment characteristics of various sealed-internal-balance arrangements for control surfaces are presented. The analysis considered overhands sealed to various types of wing structure by flexible seals spanning gaps of various widths or sealed to the wing structure by a flexible system of linked plates. Leakage was not considered; the seal was assumed to extend the full spanwise length of the control surface. The effect of the developed width of the flexible seal and of the geometry of the structure to which the seal was anchored was investigated, as well as the effect of the gap width that is sealed. The results of the investigation indicated that the most nearly linear control-surface hinge-moment characteristics can probably be obtained from a flexible seal over a narrow gap (about 0.1 of the overhang chord), which is so installed that the motion of the seal is restricted to a region behind the point of attachment of the seal to the wing structure. Control-surface hinge moments that tend to be high at large deflections and low or overbalanced at small deflections will result if a very narrow seal is used.

Beschreibung: Tests were made of a model representative of a single-engine tractor-type airplane for the purpose of determining the stability and control effects of a propeller used as an aerodynamic brake. The tests were made with single-and dual-rotation propellers to show the effect of type of propeller rotation, and with positive thrust to provide basic data with which to compare the effects of negative thrust. Four configurations of the model were used to give the effects of tilting the propeller thrust axis down 5 deg., raising the horizontal tail, and combining both tilt and raised tail. Results of the tests are reported herein. The effects of negative thrust were found to be significant. The longitudinal stability was increased because of the loss of wing lift and increase of the angle of attack of the tail. Directional stability and both longitudinal and directional control were decreased because of the reduced velocity at the tail. These effects are moderate for moderate braking but become pronounced with full-power braking, particularly at high values of lift coefficient. The effects of model configuration changes were small when compared with the over-all effects of negative-thrust operation; however, improved stability and control characteristics were exhibited by the model with the tilted thrust axis. Raising the horizontal tail improved the longitudinal characteristics, but was detrimental to directional characteristics. The use of dual-rotation propeller reduced the directional trim charges resulting from the braking operation. A prototype airplane was assumed and handling qualities were computed and analyzed for normal (positive thrust) and braking operation with full and partial power. The results of these analyses are presented for the longitudinal characteristics in steady and accelerated flight, and for the directional characteristics in high- and low-speed flight. It was found that by limiting the power output of the engine (assuming the constant-speed propeller will function in the range of blade angles required for negative thrust) the stability and control characteristics may be held within the limits required for safe operation. Braking with full power, particularly at low speeds, is dangerous, but braking with very small power output is satisfactory from the standpoint of control. The amount of braking produced with zero power output is equal to or better than that produced by conventional spoiler-type brakes.

Beschreibung: NACA instrumentation has been installed ii the X-J4 airplanes to obtain stability and control data during the acceptance tests conducted by the Northrop Aircraft Corporation. This report presents data obtained on the stalling characteristics of the airplane in the clean and gear- down configurations. The center of gravity was located at approximately 18 percent of the mean aerodynamic chord during the tests. The results indicated that the airplane was not completely stalled when stall was gradually approached during nominally U accelerated flight but that it was completely stalled during a more abruptly approached stall in accelerated flight. The stall in accelerated flight was relatively mild, and this was attributed to the nature of the variation of lift with angle of attack for the 001-614 airfoil section, the plan form of the wing, and to the fact that the initial sideslip at the stall produced (as shown by wind-tunnel tests of a model of the airplane) a more symmetrical stall pattern.

Beschreibung: The general characteristics of the flow field in a submerged air inlet are investigated by theoretical, wind-tunnel, and visual-flow studies. Equations are developed for calculating the laminar and turbulent boundary-layer growth along the ramp floor for parallel, divergent, and convergent ramp walls, and a general equation is derived relating the boundary-layer pressure losses to the boundary-layer thickness. It is demonstrated that the growth of the boundary layer on the floor of the divergent-ramp inlet is retarded and that a vortex pair is generated in such an inlet. Functional relationships are established between the pressure losses in the vortices and the geometry of the inlet. A general discussion of the boundary layer and vortex formations is included, in which variations of the various losses and of the incremental external drag with mass-flow ratio are considered. Effects of compressibility are also discussed.

Beschreibung: An investigation has been conducted in the NACA Cleveland icing research tunnel to determine the aerodynamic and icing characteristics of several recessed fuel-vent configurations. The vents were investigated aerodynamically to obtain vent-tube pressures and pressure distributions on the ramp surface as functions of tunnel-air velocity and angle of attack. Icing investigations were made to determine the vent-tube pressure losses for several icing conditions at tunnel-air velocities ranging from 220 to 440 feet per second. In general, under nonicing conditions, the configurations with diverging ramp walls maintained, vent-tube pressures greater than the required marginal value of 2 inches of water positive pressure differential between the fuel cell and the compartment containing the fuel cell for a range of angles of attack from 0 to 14deg at a tunnel-air velocity of approximately 240 feet per second. A configuration haying divergIng ramp sldewalls, a 7deg ramp angle; and vent tubes manifold,ed to a common plenum chamber opening through a slot In the ramp floor gave the greatest vent-tube pressures for all the configurations investigated. The use of the plenum chamber resulted in uniform pressures in all vent tubes. In a cloud-icing condition, roughness caused by ice formations on the airfoil surface ahead of the vent ramp, rather than icing of the vent configuration, caused a rapid loss in vent-tube pressures during the first few minutes of an icing period. Only the configuration having diverging ramp sidewalls, a 7 ramp angle, and a common plenum chamber maintained the required vent-tube pressures throughout a 60-minute icing period, although the ice formations on this configuration were more severe than those observed for the other configurations. No complete closure of vent-tube openings occurred for the configurations investigated. A simulated freezing-rain condition caused a greater and more rapid vent-tube pressure loss than was observed for a cloud-icing condition.

Beschreibung: The effects of primary and runback ice formations on the section drag of a 36 deg swept NACA 63A-009 airfoil section with a partial-span leading-edge slat were studied over a range of angles of attack from 2 to 8 deg and airspeeds up to 260 miles per hour for icing conditions with liquid-water contents ranging from 0.39 to 1.23 grams per cubic meter and datum air temperatures from 10 to 25 F. The results with slat retracted showed that glaze-ice formations caused large and rapid increases in section drag coefficient and that the rate of change in section drag coefficient for the swept 63A-009 airfoil was about 2-1 times that for an unswept 651-212 airfoil. Removal of the primary ice formations by cyclic de-icing caused the drag to return almost to the bare-airfoil drag value. A comprehensive study of the slat icing and de-icing characteristics was prevented by limitations of the heating system and wake interference caused by the slat tracks and hot-gas supply duct to the slat. In general, the studies showed that icing on a thin swept airfoil will result in more detrimental aerodynamic characteristics than on a thick unswept airfoil.

Beschreibung: Calculations have been made for the icing limit of a diamond airfoil at zero angle of attack in terms of the stream Mach number, stream temperature, and pressure altitude. The icing limit is defined as a wetted-surface temperature of 320 F and is related to the stream conditions by the method of Hardy. The results show that the point most likely to ice on the airfoil lies immediately behind the shoulder and is subject to possible icing at Mach numbers as high as 1.4.

Beschreibung: The effects of primary and. runback icing and frost formations on the drag of an 8-foot-chord NACA 651-212 airfoil section were investigated over a range of angles of attack from 20 to 80 and airspeeds up to 260 miles per hour for icing conditions with liquid-water contents ranging from 0.25 to 1.4 grams per cubic meter and datum air temperatures of -30 to 30 F. The results showed that glaze-ice formations, either primary or runback, on the upper surface near the leading edge of the airfoil caused large and rapid increases in drag, especially at datum air temperatures approaching 32 F and in the presence of high rates of water catch. Ice formations at lower temperatures (rime ice) did not appreciably increase the drag coefficient over the initial (standard roughness) drag coefficient. Cyclic de-icing of the primary Ice formations on the airfoil leading-edge section permitted the drag coefficient to return almost to the bare airfoil drag value. Runback icing on the lower surface did not present a serious drag problem except when heavy spanwise ridges of runback ice occurred aft of the heatable area. Frost formations caused rapid and large increases in drag with incipient stalling of the airfoil.

Beschreibung: The mechanics of laminar boundary layer transition are reviewed. Drag possibilities for boundary layer control are analyzed using assumed conditions of transition Reynolds number, inlet loss, number of slots, blower efficiency, and duct losses. Although the results of such analysis are highly favorable, those obtained by experimental investigations yield conflicting results, showing only small gains, and sometimes losses. Reduction of this data indicates that there is a lower limit to the quantity of air which must be removed at the slot in order to stabilize the laminar flow. The removal of insufficient air permits transition to occur while the removal of excessive amounts of air results in high power costs, with a net drag increases. With the estimated value of flow coefficient and duct losses equal to half the dynamic pressure, drag reductions of 50% may be obtained; with twice this flow coefficient, the drag saving is reduced to 25%.

Beschreibung: The problem of the minimum induced drag of wings having a given lift and a given span is extended to include cases in which the bending moment to be supported by the wing is also given. The theory is limited to lifting surfaces traveling at subsonic speeds. It is found that the required shape of the downwash distribution can be obtained in an elementary way which is applicable to a variety of such problems. Expressions for the minimum drag and the corresponding spanwise load distributions are also given for the case in which the lift and the bending moment about the wing root are fixed while the span is allowed to vary. The results show a 15-percent reduction of the induced drag with a 15-percent increase in span as compared with results for an elliptically loaded wing having the same total lift and bending moment.

Beschreibung: An investigation has been made in the NACA Lewis icing research tunnel to determine the aerodynamic and icing characteristics of a full-scale induction-system air-scoop assembly incorporating a flush alternate inlet. The flush inlet was located immediately downstream of the offset ram inlet and included a 180 deg reversal and a 90 deg elbow in the ducting between inlet and carburetor top deck. The model also had a preheat-air inlet. The investigation was made over a range of mass-air- flow ratios of 0 to 0.8, angles of attack of 0 and 4 deg airspeeds of 150 to 270 miles per hour, air temperatures of 0 and 25 F various liquid-water contents, and droplet sizes. The ram inlet gave good pressure recovery in both clear air and icing but rapid blockage of the top-deck screen occurred during icing. The flush alternate inlet had poor pressure recovery in both clear air and icing. The greatest decreases in the alternate-inlet pressure recovery were obtained at icing conditions of low air temperature and high liquid-water content. No serious screen icing was observed with the alternate inlet. Pressure and temperature distributions on the carburetor top deck were determined using the preheat-air supply with the preheat- and alternate-inlet doors in various positions. No screen icing occurred when the preheat-air system was operated in combination with alternate-inlet air flow.

Beschreibung: The results of wind tunnel tests at NASA Langley targeted at the performance and configurational characteristics of 0.1 and 0.13 scale model spanwise blowing (SWB) jet wing concepts are reported. The concept involves redirection of engine compressor bleed air to provide SWB at the fuselage-wing juncture near the wing leading edge. The tests covered the orientation of the outer panel nozzles, the effects of SWB operation on the performance of leading and trailing edge flaps and the effects of SWB on lateral stability. The trials were run at low speeds and angles of attack from 24-45 deg (landing). Both lift and longitudinal stability improved with the SWB, stall and leading edge vortex breakdown were delayed and performance increased with the SWB rate. Lateral stability was degraded below 20 deg angle of attack while instabilities were delayed above 20 deg due to roll damping.

Beschreibung: A 1/5-scale model of the Republic x-84 airplane (Army Project MX-578) was tested in the Langley 300 MPH 7- by 10-foot tunnel. The primary object of the tests was twofold: to determine, a practicable method of increasing the longitudinal 3tability in the landing configuration, and to investigate the effects on longitudinal and lateral Stability of various external stored (fuel tanks, bombs, and rockets). The effects of the fuselage dive brakes were also determined, and the critical Mach numbers of certain of the airplane components were estimated. The use of the revised horizontal tail (of larger aspect ratio and area than the original) seemed to be the most feasible expedient for materially increasing the longitudinal stability in the landing configuration. The neutral-point shifts produced by the various external stores were unstable, the largest shift being about 2.5 percent mean aerodynamic chord. No appreciable aerodynamic trim changes were caused by the external stores. From the standpoint of range, maximum s peed, and rate of climb, the advantages of mounting the fuel tanks at the wing tips rather than inboard beneath the wings were clearly demonstrated by the tests. The effective dihedral parameter was the only static lateral-stability derivative appreciably affected by the external stores. At high lift coefficients, the tip-mounted tanks caused a large increase in the effective dihedral parameter (about 40 increase at a lift coefficient of 1.0). This increase was held undesirable, because the tendency toward oscillatory instability that it would cause would be heightened by the increased moments of inertia resulting from the weight of the tanks when carrying fuel. The fuselage dive brakes, when deflected, caused a change in trim tending to nose the airplane up; the neutral point also moved rearward upon deflecting the dive brakes. The amount of elevator required to overcome the change in trim was well within the available range of deflection. It was estimated that a drive-brake deflection of 900 would.decrease the terminal Mach number in a vertical dive by about 0.1. The estimated critical Mach number of the V-front canopy was about 0.04 greater than that of the original canopy. Pressure-distribution tests disclosed severe pressure peaks inside the nose of the jet entrance duct. These peaks, which would lead to separation and consequently poor pressure recovery at, the engine, could be reduced by, using a smaller nose,radius and: a modified internal lip shape

Beschreibung: Low Mach number longitudinal-stability and control characteristics as predicted by use of wind tunnel data from a powered 3/16-scale model are compared with flight test measurements of a Navy BTD-1 airplane. The accuracy of the wind tunnel data and the discrepancies involved in attempting to correlate with flight data are discussed and analyzed. The comparison showed that wind tunnel predictions were, in general, in good agreement with flight test data. The predicted values were for the most part sufficiently accurate to show the satisfactory and unsatisfactory characteristics in the preliminary design stage and to indicate possible methods of improvement. The discrepancies which did occur were attributed principally to physical dissimilarities between model and airplane and the instability to determine accurately the flight power conditions. The effect of Mach number was considered negligible since the maximum flight test value was about 0.5. In order to simulate more closely the flight conditions and hence obtain more accurate data for predictions, it appears desirable to perform large-scale tests of unorthodox control surfaces such as the sealed vaned elevators with which the airplane was equipped.

Beschreibung: An investigation of a 1/7-scale powered model of the Kaiser Fleetwing all-wing airplane was made in the Langley full-scale tunnel to provide data for an estimation of the flying qualities of the airplane. The analysis of the stability and control characteristics of the airplane has been made as closely as possible in accordance with the requirements of the Bureau of Aeronautics, Navy Department's specifications, and a summary of the more significant conclusions is presented as follows. With the normal center of gravity located at 20 percent of the mean aerodynamic chord, the airplane will have adequate static longitudinal stability, elevator fixed, for all flight conditions except for low-power operation at low speeds where the stability will be about neutral. There will not be sufficient down-elevator deflection available for trim above speeds of about 130 miles per hour. It is probable that the reduction in the up-elevator deflections required for trim will be accompanied by reduced elevator hinge moments for low-power operation at low flight speeds. The static directional stability for this airplane will be low for all rudder-fixed or rudder-free flight conditions. The maximum rudder deflection of 30 deg will trim only about 15 deg yaw for most flight conditions and only 10 deg yaw for the condition with low power at low speeds. Also, at low powers and low speeds, it is estimated that the rudders will not trim the total adverse yaw resulting from an abrupt aileron roll using maximum aileron deflection. The airplane will meet the requirements for stability and control for asymmetric power operation with one outboard engine inoperative. The airplane would have no tendency for directional divergence but would probably be spirally unstable, with rudders fixed. The static lateral stability of the airplane will probably be about neutral for the high-speed flight conditions and will be only slightly increased for the low-power operation in low-speed flight. The airplane will not roll against the ailerons in a side-slip maneuver. Although the airplane would probably meet the minimum requirements of pb/2V of 0.07 at all speeds, there will be a loss in rolling ability of the airplane at high aileron deflections and at low flight speeds. It is estimated that the wing stall will be a gradual movement forward from the trailing edge and will be accompanied by no sudden pitching or rolling accelerations. Some stall warning may be indicated by reduction in the elevator and aileron force gradients and by the shaking of the controls caused by unsteady flow over the surfaces near the stall.

Beschreibung: Buffet boundaries, buffeting-load increments for the stabilizers and elevators, and buffeting bending-moment increments for the stabilizers and wings as measured in gradual maneuvers for a jet-powered bomber airplane are presented. The buffeting-load increments were determined from strain-gage measurements at the roots or hinge supports of the various surfaces considered. The Mach numbers of the tests ranged from 0.19 to 0.78 at altitudes close to 30,000 feet. The predominant buffet frequencies were close to the natural frequencies of the structural components. The buffeting-load data, when extrapolated to low-altitude conditions, indicated loads on the elevators and stabilizers near the design limit loads. When the airplane was held in buffeting, the load increments were larger than when recovery was made immediately.

Beschreibung: The hydrodynamic characteristics of a 1/10-size powered dynamic model of the XP5Y-1 flying boat were determined in Langley tank no. 1. Stable take-offs were possible at all practicable positions of the center of gravity and flap deflections. An increase in gross load from 123.5 to 150.0 pounds (21.5 percent) had only a slight effect on the stable range for take-off. A decrease in forward acceleration from 3.0 to 1.0 feet per second per second had only a very small effect on the stable range for take-off. In general, the landings were free from skipping except at trims below 6 deg where one skip was encountered at an aft position of the center of gravity. The model porpoised during the landing runout at all positions of the center of gravity when landed at trims above 10 deg. Spray in the propellers was light at the design gross load, and was not considered excessive,at a gross load of 136.0 pounds.

Beschreibung: The effects of several wing leading-edge camber and deflected-tip modifications on the force and moment characteristics of a 1/20-scale model of the Convair F-102 airplane have been determined at Mach numbers from 0.60 t o 1.14 for angles of attack up to 14 deg. in the Langley 8-foot transonic tunnel. The effects of elevator deflections from 0 deg. to -10 deg. were also obtained for a configuration incorporating favorable leading- edge and tip modifications. Leading-edge modifications which had a small amount of constant-chord camber obtained by vertically adjusting the thickness distribution over the forward (3.9 percent of the mean aerodynamic chord) portion of the wing were ineffective in reducing the drag at lifting conditions at transonic speeds. Leading edges with relatively large cambers designed to support nearly elliptical span load distributions at lift coefficients of 0.15 and 0.22 near a Mach number of 1.0 produced substantial reductions in drag at most lift coefficients.

Beschreibung: A model of the Consolidated Vultee Aircraft Corporation Skate 7 seaplane was tested in Langley tank no. 2. Presented without discussion in this paper are landing stability in smooth water, maximum normal accelerations occurring during rough-water landings, and take-off behavior in waves.

Beschreibung: An investigation has been conducted in the Langley 20-foot free-spinning tunnel to determine the effects of decreasing the rudder deflection, of decreasing the rudder span, and of differential rudder movements on the spin and recovery characteristics of a 0.057-scale model of the Chance Vought XF7U-1 airplane. The results indicated that decreasing the rudder span or the rudder deflections, individually or jointly, did not seriously alter the spin or recovery characteristics of the model; and recovery by normal use of controls (full rapid rudder reversal followed l/2 to 1 turn later by movement of the stick forward of neutral) remained satisfactory. Linking the original rudders so that the inboard rudder moves from full with the spin to neutral while the outboard rudder moves from neutral to full against the spin will also result in satisfactory spin and recovery characteristics. Calculations of rudder-pedal forces for recovery showed that the expected forces would probably be within the capabilities of a pilot but that it would be advisable to install some type of boost in the control system to insure easy and rapid movement of the rudders.

Beschreibung: The static longitudinal stability characteristics of a 0.15-scale model of the Hermes A-lE2 missile have been determined in the Langley high-speed 7- by 10-foot tunnel over a Mach number range of 0.50 to 0.98, corresponding to Reynolds numbers, based on body length, of 12.3 x 10(exp 6) to 17.1 x 10(exp 6). This paper presents results obtained with body alone and body-fins combinations at 0 degrees (one set of fins vertical and the other set horizontal) and 45 degree angle of roll. The results indicate that the addition of the fins to the body insures static longitudinal stability and provides essentially linear variations of the lift and pitching moment at small angles of attack throughout the Mach number range. The slopes of the lift and pitching-moment curves vary slightly with Mach number and show only small effects due to the angle of roll.

Beschreibung: The present report of Mr. Dupleich is the summary of a very extensive experimental study of the well-known mechanical phenomenon: the rotation in free fall (* air, for instance) of more or less elongated rectangles cut out of paper or pasteboard. This phenomenon, the conditions for existence of which depend chiefly on the elongated of the small plate and its weight per unit area, is essentially an aerodynamic phenomenon and as such, raises questions of a certain interest to our department.We believe that the modern concepts of the mechanics of fluids do not have the range attributed to them.

Beschreibung: At the request of the Bureau of Aeronautics, Department of the Navy, an investigation at transonic and low supersonic speeds of the drag and longitudinal trim characteristics of the Douglas XF4D-1 airplane is being conducted by the Langley Pilotless Aircraft Research Division. The Douglas XF4D-1 is a jet-propelled, low-aspect-ratio, swept-wing, tailless, interceptor-type airplane designed to fly at low supersonic speeds. As a part of this investigation, flight tests were made using rocket- propelled 1/10- scale models to determine the effect of the addition of 10 external stores and rocket packets on the drag at low lift coefficients. In addition to these data, some qualitative values of the directional stability parameter C(sub n beta) and duct total-pressure recovery are also presented.

Beschreibung: An investigation was made to determine the static lateral stability and control characteristics of a l/6-scale model of the Republic XF-84H airplane with the propeller operating. The model had a 40deg swept wing of aspect ratio 3.45 and had a thin 3-blade supersonic-type propeller. Many modifications to the basic configuration were investigated in attempts to alleviate lateral and directional trim problems which appeared to be associated with propeller slipstream rotation. Although significant benefits were realized with several modifications, none of those tested would be expected to afford satisfactory behavior for all normal flight conditions. A marked left-wing roll-off tendency was indicated at high angles of attack for the basic model configuration. Projection of only the left slat was the most effective remedy found for this problem with the propeller operating. The use of differential wing-flap deflection also appeared to offer a promising means for reducing the roll-off tendency with power on. The large sidewash over the vertical tail, associated with slip- stream rotation, severely restricted the conditions for which directional , trim could be maintained. A small triangular dorsal fin, oriented opposite to the slipstream rotation, was found very effective in reducing the adverse sidewash flow at the tail.

Beschreibung: An investigation of the low-speed, power-off stability and control characteristics of a 1/10-scale model of the Convair YF-102 airplane has been made in the Langley free-flight tunnel. The model was flown over a lift-coefficient range from 0.5 to the stall in its basic configuration and with several modifications involving leading-edge slats and increases in vertical-tail size. Only relatively low-altitude conditions were simulated and no attempt was made to determine the effect of freeing the controls. The longitudinal stability characteristics of the model were considered satisfactory for all conditions investigated. The lateral stability characteristics were considered satisfactory for the basic configuration over the speed range investigated except near the stall, where large values of static directional instability caused the model to be directionally divergent. The addition of leading-edge slats or an 8-percent increase in vertical-tail area increased the angle of attack at which the model became directionally divergent. The use of leading-edge slats in combination with a 40-percent increase in vertical-tail size eliminated the directional divergence and produced satisfactory stability characteristics through the stall. The longitudinal and lateral control characteristics were generally satisfactory. Although the adverse sideslip characteristics for the model were considered satisfactory over the angle-of-attack range, analysis indicates that the adverse sideslip characteristics of the airplane may be objectionable at high angles of attack.

Beschreibung: A series of flight tests have been made at the Langley Flight Research Division at the request of the Bureau of Aeronautics, Department of the Navy, to determine the flying qualities of the Grumman F8F-1 air- plane. This paper presents the test results necessary to determine the longitudinal stability and control characteristics end the stalling characteristics. These tests were made between February and June of 1947- The range of Mach numbers covered in this investigation was approximately 0.10 to 0.62, and no attempt was made to investigate compressibility effects at higher Mach numbers. The lateral and directional stability and control characteristics of the subject airplane have already been reported (reference 1). Also presented in this paper is a discussion of the normal accelerations induced by yawing velocity and sideslip which were considered ob,jectionable by the pilot for this airplane. A discussion of the undesirable accelerations has been included with a view towards formulating some flying-qualities requirements limiting them.

Beschreibung: An experimental investigation has been conducted in the Langley stability tunnel at low speed to determine the pitching stability derivatives of a 1/9-scale powered model of the Convair XFY-1 vertically rising airplane. Effects of thrust coefficient, control deflections, and propeller blade angle were investigated. The tests were made through an angle-of-attack range from about -4deg to 29deg, and the thrust coefficient range was from 0 to 0.7. In order to expedite distribution of these data, no analysis of the data has been prepared for this paper.

Beschreibung: A spin investigation has been conducted in the Langley 20-foot free-spinning tunnel on a 1/24-scale model of the North American XP-86 airplane. The effects of control settings and movements upon the erect and inverted spin and recovery characteristics of the model were determined for the design gross weight loading. The long-range loading was also investigated and the effects of extending slats and dive flaps were determined. In addition, the investigation included the determination of the size of spin-recovery parachute required for emergency recovery from demonstration spins, the rudder force required to move the rudder for recovery, and the best method for the pilot to escape if it should become necessary to do so during a spin. The results of the investigation indicated that the XP-86 airplane will probably recover satisfactorily from erect and inverted spins for all possible loadings. It was found that fully extending both slats would be beneficial but that extending the dive brakes would cause unsatisfactory recoveries. It was determined that a 10.0-foot-diameter tail parachute with a drag coefficient of 0.7 and with a towline 30.0 feet long attached below the jet exit or a 6.0-foot-diameter wingtip parachute opened on the outer wing tip with a towline 6.0 feet long would insure recoveries from any spins obtainable. The rudder-pedal force necessary to move the rudder for satisfactory recovery was found to be within the physical capabilities of the pilot.

Beschreibung: This paper presents the results of measurements of longitudinal stability of a 1/50-scale model of the XP-88 airplane by the wing-flow method. Lift, rolling-moment, hinge-moment, and pitching-moment characteristics as well as the downwash at the tail were measured over a Mach number range from approximately 0.5 to 1.05 at Reynolds numbers below 1,000,000. No measurements of drag were obtained. No abrupt changes due to Mach number were noted in any of the parameters measured. The data indicated that the wing was subject to early tip stalling; that the tail effectiveness decreased gradually with increasing Mach number up to M = 0.9, but increased again at higher Mach numbers; that the variation of downwash with angle of attack did not change appreciably with Mach number except between 0.95 and 1.0 where d(epsilon)/d(alpha), decreased from 0.46 to 0.32; that at zero lift with a stabilizer setting of -1.5 deg there was a gradually increasing nosing-up tendency with increasing Mach number; and that the control-fixed stability in maneuvers at constant speed gradually increased with increasing Mach number.

Beschreibung: An investigation of the spin and recovery characteristics of a 0.057-scale model of the Chance Vought XF7U-1 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 were determined, as were also the effects of extending the wing slats, of center-of-gravity movement, and-of variation in the mass distribution. The investigation also included wing-tip spin-recovery-parachute tests, pilot-escape tests, and rudder-control-force tests. The investigation indicated that the spin and recovery characteristics of the airplane will be satisfactory for all conditions. It was found that a single 4.24-foot (full-scale) parachute when opened alone from the outboard wing tip or two 8.77-foot (full-scale) parachutes when opened simultaneously, one from each wing tip, would effect satisfactory emergency recoveries (the drag coefficients of the parachutes, based on the surface area of the parachute, were 0.83 and 0.70 for the 4.24- and 8.77-foot parachutes, respectively). The towline length in both cases was 25 feet (full scale). Tests results showed that, if the pilot should have to leave the airplane during a spin, he should jump from the outboard side (left side in a right spin) of the cockpit. The rudder-control force necessary for recovery from a spin was found to be rather high but appeared to be within the upper limits of a pilot's capabilities.

Beschreibung: Tests of a 1/20-scale dynamically similar model of the Northrop B-35 airplane were made to study its ditching characteristics. The model was ditched in calm water at the Langley tank no. 2 monorail. Various landing attitudes, speeds,and conditions of damage were simulated during the investigation. The ditching characteristics were determined by visual observation and from motion-picture records and time-history acceleration records. Both longitudinal and lateral accelerations were measured. Results are given in tabular form and time-history acceleration curves and sequence photographs are presented. Conclusions based on the model investigation are as follows: 1. The best ditching of the B-35 airplane probably can be made by contacting the water in a near normal landing attitude of about 9 deg with the landing flaps full down so as to have a low horizontal speed. 2. The airplane usually will turn or yaw but the motion will not be violent. The maximum lateral acceleration will be about 2g. 3. If the airplane does not turn or yaw immediately after landing, it probably will trim up and then make a smooth run or porpoise slightly. The maximum longitudinal decelerations that will be encountered are about 6g or 7g. 4. Although the decelerations are not indicated to be especially large, the construction of the airplane is such that extensive damage is to be expected, and it probably will be difficult to find ditching stations where crew members can adequately brace themselves and be reasonably sure of avoiding a large inrush of water.

Beschreibung: An investigation of the low-speed; power-off stability and control characteristics of a 1/20-scale model of the Consolidated Vultee XB-53 airplane equipped with full-span leading-edge slats has been conducted in the Langley free-flight tunnel. In this investigation it was found that the-full-span leading-edge slat gave about the same maximum lift coefficient as was obtained with the outboard single slotted flap and inboard slat. The stability and control characteristics were greatly improved except near the stall where the characteristics with the full-span slat were considered unsatisfactory because of a loss of directional stability and a slight nosing-up tendency.

Beschreibung: The problem of turbulence in aerodynamics is at present being attacked both theoretically and experimentally. In view of the fact however that purely theoretical considerations have not thus far led to satisfactory results the experimental treatment of the problem is of great importance. Among the different measuring procedures the hot wire methods are so far recognized as the most suitable for investigating the turbulence structure. The several disadvantages of these methods however, in particular those arising from the temperature lag of the wire can greatly impair the measurements and may easily render questionable the entire value of the experiment. The name turbulence is applied to that flow condition in which at any point of the stream the magnitude and direction of the velocity fluctuate arbitrarily about a well definable mean value. This fluctuation imparts a certain whirling characteristic to the flow.

Beschreibung: A flight test was conducted at the Flight Test Station of the Pilotless Aircraft Research Division at Wallops Island, Va., to determine the longitudinal control and stability characteristics of a 0.5-scale model of the Fairchild Lerk Pilotless aircraft with the horizontal wing flaps deflected 15 degrees. The data were obtained by the use of a telemeter and also by radar tracking. The results show an increase of effectiveness of the longitudinal control in producing normal accelerations up to a Mach number of 0.75 where this effectiveness gradually decreased becoming negative at a Mach number of 0.89. Previous tests with wing flaps undeflected an increase in effectiveness up to Mach number of 0.93 where a sudden loss of control occurred. The model was dynamically stable throughout the speed range. The data confirmed the drag increase at the critical Mach number for large angles of attack is indicated in high-speed wind-tunnel tests.

Beschreibung: The NACA is conducting a general investigation of servo-mechanisms for use in powering aircraft control surfaces. This paper presents a theoretical analysis and the results of bench tests of a control-booster system which employs a variable displacement hydraulic pump. The booster is intended for use in a flight investigation to determine the effects of various booster parameters on the handling qualities of airplanes. Such a flight investigation would aid in formulating specific requirements concerning the design of control boosters in general. Results of the theoretical analysis and the bench tests indicate that the subject booster is representative of types which show promise of satisfactory performance. The bench tests showed that the following desirable features were inherent in this booster system: (1) No lost motion or play in any part of the system; (2) no detectable lag between motion of the contra1 stick and control surface; and (3) Good agreement between control displacements and stick-force variations with no hysteresis in the stick-force characteristics. The final design configuration of this booster system showed no tendency to oscillate, overshoot, or have other undesirable transient characteristics common to boosters.

Beschreibung: An investigation of the low-speed, power-off stability and control characteristics of a 1/20-scale model of the Consolidated Vultee XB-53 airplane has been conducted in the Langley free-flight tunnel. In the investigation it was found that with flaps neutral satisfactory flight behavior at low speeds was obtainable with an increase in height of the vertical tail and with the inboard slats opened. In the flap-down slat-open condition the longitudinal stability was satisfactory, but it was impossible to obtain satisfactory lateral-flight characteristics even with the increase in height of the vertical tail because of the negative effective dihedral, low directional stability, and large-adverse yawing moments of the ailerons.

Beschreibung: This report presents the results of wind-tunnel force tests which were conducted to determine the low-speed stability and control characteristics of a full-scale Northrop XSSM-A-3 missile. Tests were made through a range of angles of attack, sideslip, and control deflection, and at various Reynolds numbers. Characteristics of the complete missile are compared with the characteristics of the missile with the landing skids extended, with the vertical tail removed, and with the fuselage alone. No analysis of the data has been made in order to make the results available as soon as possible.

Beschreibung: An investigation of a vortex-generator configuration on the wings of a l/4-scale model of the X-1 airplane having a 10-percent-thick wing was conducted in the Langley 16-foot transonic tunnel. The effect of the vortex generators was determined by comparing the model aerodynamic characteristics, wing-pressure distributions, and wing-wake patterns for model configurations with and without vortex generators on the wings. Results are presented from tests at 0.1 increments in Mach number from about 0.69 to 0.99, at Reynolds numbers of about 4.1 x 10(exp 6) to 4.7 x 10(exp 6), and through an angle-of-attack range up to 1.5 deg at lower speeds and up to 5 deg at the highest speed. In general, little difference in the aerodynamic characteristics was observed, except at a Mach number of 0.90 where a rearward movement of the shock on the upper surface of the wing with the vortex generators installed resulted in less separation.

Beschreibung: A 1/7 scale semispan model of the XB-35 airplane was tested in the Langley 10 foot pressure tunnel, primarily for the purpose of investigating the effectiveness of a leading-edge slot for alleviation of stick-fixed longitudinal instability at high angles of attack caused by early tip stalling and a device for relief of stick-free instability caused by elevon up-floating tendencies at high angles of attack. Results indicated that the slot was not adequate to provide the desired improvement in stick-fixed stability. The tab-flipper device provided improvement in stick-free stability abd two of the linkage combinations tested gave satisfactory variations of control force with airspeed for all conditions except that in which the wing-tip "pitch-control" flap was fully deflected. However, the improvement in control force characteristics was accompanied by a detrimental effect on stick-fixed stability because of the pitching moments produced by the elevon tab deflection.

Beschreibung: The results obtained from gust and draft velocity measurements within thunderstorms for the period July 24, 1946 to August 6, 1946 at Orlando, Florida are presented herein. These data are summarized in tables I and II and are of the type presented in reference 1 for previous flights. In two thunderstorm traverses, indications of ambient-air temperature were obtained from photo-observer records. These data are summarized in table III.

Beschreibung: The spin and recovery characteristics of the Northrop XF-89 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 mainly estimated rather than determined by model tests because the XF-89 dimensional and mass characteristics were such as to make this airplane similar to several others, models of which have previously been tested. Brief tests were made on an available model of similar design to augment the estimation. The results indicate that the recovery characteristics will be satisfactory for all airplane loadings if recovery is attempted by use of rudder followed by moving the elevator down. The rudder pedal forces will be within the capabilities of the pilot but the elevator stick forces will be beyond the pilot's capabilities unless a trim tab, or a booster is used. A 9.5-foot-diameter flat-type tail parachute or a 5.0-foot-diameter flat-type wing-tip parachute with a drag coefficient of 0.7 will be a satisfactory emergency spin-recovery device for spin demonstrations and if it is necessary for the crew to abandon the spinning airplane, they should leave from the outboard side of the cockpit.

Beschreibung: At the request of the Henschel Aircraft Works. A. G. Berlin. three models of the missile "Zitterrochen" were investigated at subsonic velocities.(open jet 215-millimeter diameter) and at supersonic velocities (open jet 110 by 130 millimeters) in order to determine the effect of various wing forms on the air forces and moments. Three-component measurements were taken, and one model was also investigated with deflected control plates.

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

Beschreibung: Estimates of the static stick-fixed stability and control characteristics of the Consolidated Vultee model 240 airplane are presented in this report. The estimates are based on tests of a 0.092-scale powered model in the 10-foot wind tunnel of the Guggenheim Aeronautical Laboratory of the California Institute of Technology. Results of the analysis are evaluated in terms of the Army specifications for stability and control characteristics which are more specific and, in general, equal to or more rigid than the Civil Aeronautics Administration requirements. The stick-fixed stability and control characteristics of the Consolidated Vultee model 240 were found to be satisfactory except for the following: 1) Marginal longitudinal stability in the landing approach (flaps 30 deg, 50% minimum continuous power) with aft center of gravity (31% M.A.C.); 2) Marginal rudder control to hold zero sideslip in a climb after take-off with asymmetric power (flaps 30 deg, left engine inoperative, right engine delivering take-off power) with maximum rudder throw limited to +/- 18 deg; 3) Marginal dihedral effect with flaps 40 deg and engines delivering maximum continuous power.

Beschreibung: To determine the trim range in which a seaplane can take off without porpoising, stability tests were made of a Plexiglas model, composed of float, wing, and tailplane, which corresponded to a full-size research airplane. The model and full-size stability limits are in good agreement. After all structural parts pertaining to the air frame were removed gradually, the aerodynamic forces replaced by weight forces, and the moment of inertia and position of the center of gravity changed, no marked change of limits of the stable zone was noticeable. The latter, therefore, is for practical purposes affected only by hydrodynamic phenomena. The stability limits of the DVL family of floats were determined by a systematic investigation independent of any particular sea-plane design, thus a seaplane may be designed to give a run free from porpoising.

Beschreibung: Wind-tunnel measurements on projectiles are discussed. Tests at the Gottingen Tunnel are described. The tunnel operates on the Prandtl principle, that is, a brief stationary air stream produced in an evacuated tank by induction of atmospheric air.

Beschreibung: The compression plane is intended for operation on or close to the surface of the water, and has a hull with a concave bottom which forms the upper surface of a tunnel into which air is forced under pressure to support part of the load. The results of the tests made in Langley tank no. 1 include values of the horizontal forces, trimming moment, and static pressure in the tunnel for a wide range of loads and speeds and two power conditions, and are presented in the form of curves against speed with load as a parameter. The results are scaled up to 10 times the model size for three conditions at which the model is self-propelled at a steady speed. Lift is obtained from the static pressure of air in the tunnel. In general, the ratio of the gross load to the total resistance increases with increase in load and decrease in speed. This ratio varies between l-7 and 5.7 at high speeds and has a maximum value of 7. The total resistance is nearly the same for both power conditions except at low speeds and heavy loads. No abrupt change in forces on the hull or flow around the hull occurs in. the region of zero draft. The centers of pressure are generally far aft. At the most efficient trim (1.2'), considerable bow-up moment would be required for practicable operation. There is no abrupt transition from the air-borne to the water- borne condition.

Beschreibung: An investigation was made of a 1/10-scale dynamically similar model of the North American F-86 airplane to study its behavior when ditched. The model was landed in calm water at the Langley tank no. 2 monorail. Various landing attitudes, speeds, and conditions of damage were simulated. The behavior of the model was determined from visual observations, acceleration records, and motion-picture records of the ditchings. Data are presented in tabular form, sequence photographs, and time-history acceleration curves. From the results of the investigation it was concluded that the airplane should be ditched at the nose-high, 14 deg attitude to avoid the violent dive which occurs at the 4 deg attitude. The flaps and leading-edge slats should be fully extended to obtain the lowest possible landing speed. The wing tanks should be jettisoned to avoid the undesirable behavior which occurs with the tanks attached. In a calm-water ditching under these conditions the airplane will run smoothly for about 600 feet. Maximum longitudinal and vertical decelerations of about 3g will be encountered.

Beschreibung: An investigation is being conducted to determine the dynamic stability and control characteristics of a 0.13-scale flying model of Convair XFY-1 vertically rising airplane. This paper presents the results of flight and force tests to determine the stability and control characteristics of the model in vertical descent and landings in still air. The tests indicated that landings, including vertical descent from altitudes representing up to 400 feet for the full-scale airplane and at rates of descent up to 15 or 20 feet per second (full scale), can be performed satisfactorily. Sustained vertical descent in still air probably will be more difficult to perform because of large random trim changes that become greater as the descent velocity is increased. A slight steady head wind or cross wind might be sufficient to eliminate the random trim changes.

Beschreibung: Measurements were made, in dives to transonic speeds, of the static-pressure position error at a distance of one chord ahead of the McDonnell XF-88 airplane. The airplane incorporates a wing which is swept back 35 deg along the 0.22 chord line and utilizes a 65-series airfoil with a 9-percent-thick section perpendicular to the 0.25-chord line. The section in the stream direction is approximately 8-percent thick. Data up to a Mach number of about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.68. Data at Mach numbers above about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.68. Results of the measurements indicate that the static-pressure error, within the accuracy of measurement, is negligible from a Mach number of 0.65 to a Mach number of about 0.97. With a further increase in Mach number, the static-pressure error increases rapidly; at the highest Mach number attained in these tests (about M = 1.038), the error increases to about 8 percent of the impact pressure. Above a Mach number of about 0.975, the recorded Mach number remains substantially constant with increasing true Mach number; the installation is of no value between a Mach number of about 0.975 and at least 1.038, as the true Mach number cannot be obtained from the recorded Mach number in this range. Previously published data have shown that at 0.96 chord ahead of the wing tip of the straight-wing X-l airplanes, a rapid rise of position error started at a Mach number of about 0.8. In the case of the XF-88 airplane, this rise of position error was delayed, presumably by the sweep of the wing, to a Mach number of about 0.97.

Beschreibung: A flight investigation has been made to determine the external drag and pressure recovery of a 1/8.25 - scale flight model of the Consolidated Vultee XF-92 from Mach numbers 0.7 to 1.4 and Reynolds numbers from 8.5 x 10(exp 6) to 19.2 x 10(exp 6) at or near zero lift. Relative mass flow, average pressure recovery, total drag, internal drag, and external drag are presented as functions of Mach number. Between Mach numbers of 0.90 and 0.975, the external drag of the configuration (including base drag of the inner body and additive drag) was about equal to that of a similar model with a faired nose and no mass flow; however, at supersonic speeds the drag coefficient for the faired-nose model remained relatively constant whereas the drag coefficient for the ducted model continued to increase sharply. The internal drag coefficient of the duct was roughly constant at 0.013 up to a Mach number of 1.20; after which it decreased to 0.0075 at a Mach number of 1.4. The over-all pressure recovery of the inlet and duct varied from 94 percent at a Mach number of 0.7 to about 91 percent at a Mach number of 1.4 at a relative-mass-flow ratio of about 0.30. The losses in pressure recovery were believed to be caused by the possible occurrence of separation of flow from the inner body and by an aerodynamically unclean internal configuration which did not duplicate the form proposed for the original XF-92 airplane.

Beschreibung: An investigation of the low-speed, power-off stability and control characteristics of a 1/10-scale model of the Douglas XF4D-1 airplane has been made in the Langley free-flight tunnel. The model was flown with leading-edge slats retracted and extended over a lift-coefficient range from 0.5 to the stall. Only relatively low-altitude conditions were simulated and no attempt was made to determine the effect on the stability characteristics of freeing the controls. The longitudinal stability and control characteristics of the model were satisfactory for all conditions investigated except near the stall with slats extended, where the model had a slight nosing-up tendency. The lateral stability and control characteristics of the model were considered satisfactory for all conditions investigated except near the stall with slats retracted, where a change in sign of the static- directional-stability parameter Cn(sub beta) caused the model to be directionally divergent. The addition of an extension to the top of the vertical tail did not increase Cn(sub beta) enough to eliminate the directional divergence of the model, but a large increase in Cn(sub beta) that was obtainable by artificial means appeared to eliminate the divergence and flights near the stall could be made. Artificially increasing the stability derivative-Cn(sub r) (yawing moment due to yawing) and Cn(sub p) (yawing moment due to rolling) had little effect on the divergence for the range of these parameters investigated. Calculations indicate that the damping of the lateral oscillation of the airplane with slats retracted or extended will be satisfactory at sea level but will be only marginally satisfactory at 40,000 feet.

Beschreibung: 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 show the model to be longitudinally unstable in the angle-of-attack range around zero angle of attack and to become stable at moderate angles of attack. The results of the present investigation agree reasonably well with results obtained in other facilities at low speed. The present pitching-moment results at low Mach numbers also agree reasonably well with unpublished results of tests of the model at supersonic Mach numbers (up to Mach number 1.86). Unpublished results at moderate and high subsonic speeds, however, indicate considerably greater instability at low angles of attack than is indicated by low-speed results. The results of the present tests also showed that the pitching-moment coefficients for angles of attack up to 12deg remained fairly constant with sideslip angle up to 12deg. The elevators tested produced relatively large pitching moments at zero angle of attack but, as the angle of attack was increased, the elevator effectiveness decreased. The rate of decrease of elevator effectiveness with angle of attack was less for 8deg than for 20deg elevator deflection. Therefore although 8deg deflection caused an appreciable change in longitudinal trim angle and trim lift coefficient a deflection of 20deg caused only a small additional increase in trim angle and trim lift coefficient.

Beschreibung: A supplementary investigation was conducted in the Langley 20-foot free-spinning tunnel on a 1/24-scale model of the Grumman F9F-6 airplane. The primary purpose of the investigation was to reevaluate the spin-recovery characteristics of the airplane in view of the fact that the ailerons had been eliminated from the flaperon-aileron lateral control system of the airplane. A spin-tunnel investigation on a model of the earlier version of the F9F-6 airplane had indicated that use of ailerons with the spin (stick right in a right spin) was essential to insure recovery. The results indicate that with.ailerons eliminated, it may be difficult to obtain an erect developed spin but if a fully developed spin is obtained on the airplane, recovery therefrom may be difficult or impossible. Flaperon deflection should have little effect on spins or recoveries.

Beschreibung: An investigation has been conducted in the Langley 20-foot free-spinning tunnel on a l/23-scale model of the Lockheed XFV-1 airplane to determine the effects of control setting and movement upon the erect-spin and recovery characteristics for a range of airplane loading conditions. A windmilling propeller was simulated on the model for some of the tests. The investigation included determination of the size of tail parachute required for emergency recovery from demonstration spins. The tumbling tendencies of the model were also investigated. The results indicated that any erect or inverted spin obtained on the airplane will be satisfactorily terminated if recovery is attempted by full rudder reversal accompanied by simultaneous lateral and longitudinal movement of the stick to neutral, The model test results showed that an 11.5-foot flat-type tail parachute (drag coefficient approximately 0.73) with a 27.5-foot towline will be effective as an emergency spin-recovery device during demonstration spins of the airplane. The model results also indicate that the airplane will not tumble for any.loading condition indicated possible.

Beschreibung: Flight tests have been made to determine the longitudinal stability and control and stalling characteristics of a North American P-51H airplane. The results indicate that the airplane has satisfactory longitudinal stability in all the flight conditions tested at normal loadings up to 25,000 feet altitude. At Mach numbers above 0.7, the elevator push force required for longitudinal trim decreased somewhat because of compressibility effects. The elevator stick force per g in accelerated turns at the forward center-of-gravity position of 24 percent mean aerodynamic chord above 250 miles per hour was in excess of the required limits at both 5,000 and 25,OOO feet altitude. The longitudinal-trim-force changes due to flaps and power were small, but the rudder-trim-force change with power change was high. The stalling characteristics in all the conditions tested were satisfactory.

Beschreibung: The ditching characteristics of the Lockheed XR60-1 airplane were determined by tests of a 1/24-scale dynamic model in calm water at the Langley tank no. 2 monorail. Various landing attitudes, flap settings, speeds, and conditions of damager were investigated. The ditching behavior was evaluated from recordings of decelerations, length of runs, and motions of the model. Scale-strength bottoms and simulated crumpled bottoms were used to reproduce probable damage to the fuselage. It was concluded that the airplane should be ditched at a landing attitude of about 5 deg with flaps full down. At this attitude, the maximum longitudinal deceleration should not exceed 2g and the landing run will be bout three fuselage lengths. Damage to the fuselage will not be excessive and will be greatest near the point of initial contact with the water.

Beschreibung: An experimental investigation has been conducted to determine the stability and control characteristics of a 0.13-scale free-flight model of the Convair XFY-1 airplane during take-offs and landings in steady winds. The tests indicated that take-offs in headwinds up to at least 20 knots (full scale) will be fairly easy to perform although the airplane may be blown downstream as much as 3 spans before a trim condition can be established. The distance that the airplane will be blown down-stream can be reduced by restraining the upwind landing gear until the instant of take-off. The tests also indicated that spot landings in headwinds up to at least 30 knots (full scale) and in crosswinds up to at least 20 knots (full scale) can be accomplished with reasonable accuracy although, during the landing approach, there will probably be an undesirable nosing-up tendency caused by ground effect and by the change in angle of attack resulting from vertical descent. Some form of arresting gear will probably be required to prevent the airplane from rolling downwind or tipping over after contact. This rolling and tipping can be prevented by a snubbing line attached to the tip of the upwind' wing or tail or by an arresting gear consisting of a wire mesh on the ground and hooks on the landing gear to engage the mesh.

Beschreibung: The present report deals with the aerodynamic, constructive, and instrumental development of a spoiler control for remote control of flying missiles.

Beschreibung: The lift coefficient of!a wing of small span at first shows a linear increase for the increasing angle of attack, but to a lesser degree then was to be expected according to the theory of the lifting line; thereafter the lift coefficient increases more rapidly than linearity, as contrasted with the the theory of the lifting line. The induced drag coefficient for a given lift coefficient, on the other hand, is obviously much smaller than it would be according to the theory. A mall change in the theory of the lifting line will cover these deviations.

Beschreibung: The stability and control characteristics of an 0.08-scale model of the Chance Vought XF7U-1 airplane have been investigated over a Mach number range from 0.40 to 0.91. Results of the basic longitudinal tests of the complete model with undeflected control surfaces are given in the present report with a very limited analysis of the results.