ALBERT

Description: The relation between the elevator hinge-moment parameters and the control-forces for changes in forward speed and in maneuvers is shown for several values of static stability and elevator mass balance.

Description: An investigation at a Mach number of 1.62 was made in the Langley 9-inch supersonic tunnel of a series of missile configurations having tandem lifting surfaces of low aspect ratio and of newly equal span. Some of the variables investigated were interdigitation angle, wing and tail plan form, and longitudinal location of wing with respect to tail. All configurations were tested through an angle-of-attack range from -5 deg to 15 deg at roll angles of 0 deg and 45 deg. Lift, drag, and pitching moment data are presented, together with center-of-pressure locations and tail-lift efficiency factors.

Description: Three rocket-propelled buffet-research models have been flight tested to determine the buffeting characteristics of a swept-wing- airplane configuration with the horizontal tail operating near the wing wake. The models consisted of parabolic bodies having 45deg sweptback wings of aspect ratio 3.56, at aspect ratio of 0.3, NACA 64A007 airfoil sections, and tail surfaces of geometry and section identical to the wings. Two tests were conducted with the horizontal tail located in the wing chord plane with fixed incidence angles of -1.5deg on one model and 0deg on the other model. The third test was conducted with no horizontal tail. Results of these tests are presented as incremental accelerations in the body due to buffeting, trim angles of attack, trim normal- and side-force coefficients, wing-tip helix angles, static-directional-stability derivatives , and drag coefficients plotted against Mach number. These data indicate that mild low-lift buffeting was experienced by all models over a range of Mach number from approximately 0.7 to 1.4. It is further indicated that this buffeting was probably induced by wing-body interference and was amplified at transonic speeds by the horizontal tail operating in the wing wake. A longitudinal trim change was encountered by the tail-on models at transonic speeds, but no large changes in side force and no wing dropping were indicated.

Description: 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.

Description: 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.

Description: 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.

Description: 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.

Description: Experimental measurements of the attenuation of plane shock waves moving over rough walls have been made in a shock tube. Measurements of the boundary-layer characteristics, including thickness and velocity distribution behind the shock, have also been made with the aid of new cal techniques which provide direct information on the local boundary-layer conditions at the rough walls. Measurements of shock speed and shock pressure ratio are presented for both smooth-wall and rough-wall flow over lengths of machined-smooth and rough strips which lined all four walls of the shock tube. A simplified theory based on Von Karman's expression for skin-friction coefficient for flow over rough walls, along with a wave-model concept and extensions to include time effects, is presented. In this theory, the shock-tube flow is assumed to be one-dimensional at all times and the wave-model concept is used to relate the local layer growth to decreases in shock strength. This concept assumes that local boundary-layer growths act as local mass-flow sinks, which give rise to expansion waves which, in turn, overtake the shock and lower its mass flow accordingly.

Description: 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.

Description: 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.

Description: Force tests on a proposed body shape of the Hermes A-2 missile with and without longitudinal spoilers were made at Mach number 4.04. Values of normal force coefficient, pitching-moment coefficient, and center-of-pressure position were obtained.

Description: 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.

Description: An approximate method of calculating the deformations of wings of uniform thickness having swept, M or W, Delta, and swept-tip plan forms is presented. The method employs an adjustment to the elementary beam theory to account for the effect of the triangular root portion of a swept wing on the deformation of the outboard section of the wing. To demonstrate the general applicability of the method, the modified elementary theory is applied to the more complex M or W, Delta, and swept-tip plan forms as well as to swept plan forms. For the purpose of calculating angles of attack, it is shown that the unmodified elementary beam theory applied to that part of the wing outboard of the root triangle produces satisfactory results. However, for calculating deflections it is necessary to include the effects of the root-triangle deformation.

Description: A theoretical investigation has been made to determine the effect on the lateral stability of the Douglas D-58-II airplane of an autopilot sensitive to yawing velocity. The effects of inclination of the gyro spin axis to the flight path and of tire lag in the autopilot were also determined. The flight conditions investigated included landing at sea level, approach condition at 12,000 feet, and cruising at 50,000 feet at Mach numbers of 0.80 and 1.2. The results of the investigation indicated that the lateral stability characteristics of the D-558-II airplane for the flight condition discussed should satisfy the Air Force - Navy period-damping criterion when the proposed autopilot is installed. Airplane motions in sideslip subsequent to a disturbance in sideslip are presented for several representative flight conditions in which a time lag in the autopilot of 0.10 second was assumed.

Description: An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the aerodynamic characteristics of a flying-boat hull of a length-beam ratio of 15 in the presence of a wing. The investigation was an extension of previous tests made on hulls of length-beam ratios of 6, 9, and 12; these hulls were designed to have approximately the same hydrodynamic performance with respect to spray and resistance characteristics. Comparison with the previous investigation at lower length-beam ratios indicated a reduction in minimum drag coefficients of 0.0006 (10 peroent)with fixed transition when the length-beam ratio was extended from 12 to 15. As with the hulls of lower length-beam ratio, the drag reduction with a length-beam ratio of 15 occurred throughout the range of angle of attack tested and the angle of attack for minimum drag was in the range from 2deg to 3deg. Increasing the length-beam ratio from 12 to 15 reduced the hull longitudinal instability by an mount corresponding to an aerodynamic-center shift of about 1/2 percent of the mean aerodynamic chord of the hypothetical flying boat. At an angle of attack of 2deg, the value of the variation of yawing-moment coefficient with angle of yaw for a length-beam ratio of 15 was 0.00144, which was 0.00007 larger than the value for a length-beam ratio of 12.

Description: The effect on drag of positioning symmetrically mounted Douglas Aircraft Company, Inc. stores in pairs on a parabolic fuselage of fineness ratio 10.0 has been determined by flight tests of rocket-propelled, zero-lift models through a range of Mach number from 0.9 to 1.8. The stores were mounted in half-submerged positions and on pylons and were tested in three longitudinal locations on the fuselage with the forward position being located at the maximum diameter of the fuselage. The effects on drag of removing the half-submerged stores or extending them outward on pylons also was investigated by tests of models with half-submerged-store cavities on the fuselage. Two pylons differing in airfoil section and thickness were tested at the forward position of the stores on the fuselage with cavities. The half-submerged stores gave the smallest drag increments, which were approximately equal regardless of their respective longitudinal locations. Removing the half-submerged stores to expose the cavities increased the drag increments from two to three times. For the pylon-mounted stores, the store in the midposition had less drag than in the forward or rear positions at supersonic speeds. Adding the half-submerged-store cavities to the pylon-mounted-store configurations reduced the drag at the rear position between Mach numbers 0.95 and 1.50 and increased the drag at the midposition throughout the speed range. Changing from the 6-percent-thick flat pylon to the 10-percent-thick airfoil pylon increased the total drag slightly above Mach number 1.10. Good agreement was obtain& between the experimental and theoretical interference drag coefficients for the pylon-mounted stores (without fuselage cavities} in the three longitudinal locations tested at Mach numbers 1.2 and 1.5.

Description: At the request of the Army Air Forces, Materiel Command, tests were made in the two-Dimensional low-turbulence tunnel of a mode l submitted by Bell Aircraft Corporation as the tip section of the wing of the xp-63 airplane. " The model was a NACA 66, 2x-2l6 a = 0.6 airfoil section of 50-inch chord and made of dural according to standard sheet-metal practice. The model was fitted for an internal-balance aileron with a tab and a Frise type aileron. Several alterations of the skirts and balance on the internal- balance aileron were tested to obtain the various aerodynamic characteristics of the aileron.

Description: The method for predicting wing- tail interference whereby the trailing vortex system behind lifting wings is replaced by fully rolled-up vortices has been applied to the calculation of tail efficiency parameters, lift characteristics, and center -of-pressure locations for a series of generalized missile configurations. The calculations have been carried out with assumed and experimental vortex locations, and comparisons made with experimental data. The measured spanwise locations of the vortices for the inline case were found to be in good agreement with the asymptotic values computed from the center of gravity of the vorticity using the method of Lagerstrom and Graham. For the interdigitated configurations the measured spanwise locations were in only fair agreement with the asymptotic locations computed for the inline case. The vertical displacement of the vortices with angle of attack for both inline and interdigitated configurations was small. The method utilizing the rolled -up vortex concept was shown to give good results in the prediction of tail efficiency variations with angle of attack for inline configurations. Not as good correlation with experiment was shown for the interdigitated configurations. Complete configuration lift -curve slopes and center -of-pressure locations, obtained using t ail efficiency calculations together with the characteristics of the components obtained from available theoretical methods, showed excellent correlation with experimental results.

Description: A flight investigation has been made to determine the longitudinal stability and control characteristics of a 60 0 delta-wing-canard missile configuration with an exposed wing-canard area ratio of 16:1. The results presented include the longitudinal stability derivatives, control effectiveness, and drag characteristics for a Mach number range of 0.75 to 1.80 and are compared with the results of a similar configuration having larger 6ontrols. Stability characteristics are also presented from the flights of an interdigitated canard configuration at a Mach number of 2.08 and a wing-body configuration at Mach numbers of 1.25 to 1.45. The stability derivatives varied gradually with Mach number with the exception of the damping-in-pitch derivative. Aerodynamic damping in pitch decreased to a minimum at a Mach number of 1.0 3, then increased to a peak value at a Mach number of 1.26 followed by a gradual decrease at higher Mach numbers. The aerodynamic-center location of the in-line canard configuration shifted rearward 13 percent of the mean aerodynamic chord at transonic speeds. The pitching-moment curve slope was 25 percent greater for the model having no canards than for the in-line configuration. No large effects of interdigitation were noted in the stability derivatives. Pitching effectiveness of the in-line configuration was maintained throughout the Mach number range. A comparison of the stability and control characteristics of two canard configurations having different area controls showed that decreasing the control area 44 percent decreased the pitching effectiveness proportionally, shifted the aerodynamic-center location rearward 9 to 14 percent of the mean aerodynamic chord, and reduced the total hinge moments required for 10 trimmed flight about 50 percent at transonic speeds.

Description: The lift, pitching-moment, and drag characteristics of a missile configuration having a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 were measured at a Mach number of 1.99 and a Reynolds number of 6.0 million, based on the body length. The tests were performed through an angle-of-attack range of -5 deg to 28 deg to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components (body, wing, and tail). Theoretical lift and moment characteristics of the configuration and its components were calculated by the use of existing theoretical methods which have been modified for application to high angles of attack, and these characteristics are compared with experiment. The lift and drag characteristics of all combinations of the body, wing, and tail were independent of roll angle throughout the angle-of-attack range. The pitching-moment characteristics of the body-wing and body-wing-tail combinations, however, were influenced significantly by the roll angle at large angles of attack (greater than 10 deg). A roll from 0 deg (one pair of wing panels horizontal) to 45 deg caused a forward shift in the center of pressure which was of the same magnitude for both of these combinations, indicating that this shift originated from body-wing interference effects. A favorable lift-interference effect (lift of the combination greater than the sum of the lifts of the components) and a rearward shift in the center of pressure from a position corresponding to that for the components occurred at small angles of attack when the body was combined with either the exposed wing or tail surfaces. These lift and center-of-pressure interference effects were gradually reduced to zero as the angle of attack was increased to large values. The effect of wing-tail interference, which influenced primarily the pitching-moment characteristics, is dependent on the distance between the wing trailing vortex wake and the tail surfaces and thus was a function of angle of attack, angle of roll, and wing-tail interdigitation. Although the configuration at zero roll with the wing and tail in line exhibited the least center-of-pressure travel, the configuration with the wing and tail interdigitated had the least change in wing-tail interference over the angle-of-attack range. The lift effectiveness of the variable-incidence wing was reduced by more than 70 percent as a result of an increase in the combined angle of attack and wing incidence from 0 deg to 40 deg. The wing-tail interference (effective downwash at the tail) due to wing deflection was nearly zero as a result of a region of negative vorticity shed from the inboard portion of the wing. The lift characteristics of the configuration and its components were satisfactorily predicted by the calculated results, but the pitching moments at large angles of attack were not because of the influence of factors for which no adequate theory is available, such as the variation of the crossflow drag coefficient along the body and the effect of the wing downwash field on the afterbody loading.

Description: The observed discrepancy at supersonic speeds between theoretical and apparent experimental average flat plate friction-drag coefficients calculated from boundary layer total-pressure surveys was investigated. Effects of the total-pressure probe, heat transfer through the leading edge region, change in leading-edge radius and strength of the leading-edge wave, possible early transition to turbulent flow or bursts of turbulence, and the slight stream-wise pressure gradient inherent in flat plate flow were investigated for plates with very sharp leading edges. Only one of these factors, the effect of the total-pressure probe, was found to be significant. Total-pressure probes of different tip heights, when placed in laminar boundary layers developing under identical conditions, were found to yield different values of friction drag coefficient. Extrapolation of these measurements indicates that a probe of vanishing size would yield the theoretical predicted values of average flat plate friction-drag coefficients. A correlation describing the relation between the friction-drag discrepancy and the probe tip height is presented.

Description: In many studies, especially of nonstationary flight motion, it is necessary to determine the angular velocities at which the airplane rotates about its various axes. The three-component recorder is designed to serve this purpose. If the angular velocity for one flight attitude is known, other important quantities can be derived from its time rate of change, such as the angular acceleration by differentiations, or - by integration - the angles of position of the airplane - that is, the angles formed by the airplane axes with the axis direction presented at the instant of the beginning of the motion that is to be investigated.

Description: The wall interference is obtained for a wind tunnel of elliptic section for the two cases of closed and open working sections. The approximate and exact methods used gave results in practically good agreement. Corresponding to the result given by Glauert for the case of the closed rectangular section, the interference is found to be a minimum for a ratio of minor to major axis of 1:square root of 6 This, however, is true only for the case where the span of the airfoil is small in comparison with the width of the tunnel. For a longer airfoil the favorable ellipse is flatter. In the case of the open working section the circular shape gives the minimum interference.

Description: The present report deals with a number of problems, particularly with the interaction of the fuselage with the wing and tail, on the basis of simple calculating method's derived from greatly idealized concepts. For the fuselage alone it affords, in variance with potential theory, a certain frictional lift in yawed flow, which, similar to the lift of a wing of small aspect ratio, is no longer linearly related to the angle of attack. Nevertheless there exists for this frictional lift something like a neutral stability point the position of which on oblong fuselages appears to be associated with the lift increase of the fuselage in proximity to the zero lift, according to the present experiments. The Pitching moments of the fuselage can be determined with comparatively great reliability so far as the flow conditions in the neighborhood of the axis of the fuselage can be approximated if the fuselage were absent, which, in general, is not very difficult. For the unstable contribution of the fuselage to the static longitudinal stability of the airplane it affords comparatively simple formulas, the evaluation of which offers little difficulty. On the engine nacelles there is, in addition a very substantial wing moment contribution induced by the nonuniform distribution of the transverse displacement flow of the nacelle along the wing chord; this also can be represented by a simple formula. A check on a large number of dissimilar aircraft types regarding the unstable fuselage and nacelle moments disclosed an agreement with the wind-tunnel tests, which should be sufficient for practical requirements. The errors remained throughout within the scope of instrumental accuracy.

Description: The present report deals with the determination of the impact stresses in the bulkhead floors of a seaplane bottom. The dynamic problem is solved on the assumption of a certain elastic system, the floor being assumed as a weightless elastic beam with concentrated masses at the ends (due to the mass of the float) and with a spring which replaces the elastic action of the keel in the center. The distributed load on the floor is that due to the hydrodynamic force acting over a certain portion of the bottom. The pressure distribution over the width of the float is assumed to follow the Wagner law. The formulas given for the maximum bending moment are derived on the assumption that the keel is relatively elastic, in which case it can be shown that at each instant of time the maximum bending moment is at the point of juncture of the floor with the keel. The bending moment at this point is a function of the half width of the wetted surface c and reaches its maximum value when c is approximately equal to b/2 where b is the half width of the float. In general, however, for computing the bending moment the values of the bending moment at the keel for certain values of c are determined and a curve is drawn. The illustrative sample computation gave for the stresses a result approximately equal to that obtained by the conventional factory computation.

Description: A flight investigation has been made to determine the drag and longitudinal stability of a 1/10- scale model of the Douglas XF4D-1 airplane from Mach numbers 0.7 to 1.4 at lift coefficients near zero. The drag rise occurred near M = 0.95. The external drag coefficient was a constant value of about 0.012 at subsonic speeds up to the point of drag rise where it increased abruptly to a value of 0.030 at M = 1.0 followed by a more gradual increase to a value of 0.038 at M = 1.25. The model indicated that, at 35,000 feet and a level-flight free-stream Mach number of 1.0, the drag of the full-scale airplane would exceed the thrust available from an XJ40-WE-8 engine with after-burning. The transonic trim change was small. The aerodynamic center moved gradually from the most forward location of 21.0-percent mean aerodynamic chord at M = 0.9 to the most rearward location of 40-percent mean aerodynamic chord at M = 1.25. The damping in pitch was low.

Description: An investigation was made of the trim and dynamic response characteristics of the free-floating horizontal tail of a 1/7-scale model of the complete tail of the Grumman XF10F-1 airplane in the Langley 8-foot transonic tunnel at Mach numbers up to 1.13. The complete tail was mounted in the tunnel on a 3deg conical support body. Various configurations were investigated. A loss in damping of the horizontal tail at transonic speeds was shown by both tunnel and flight tests. The loss in damping extended over a greater Mach number range and the maximum loss occurred at a higher Mach number in the tunnel tests. Large-amplitude oscillations of the horizontal tail of the basic configuration which occurred at low supersonic Mach numbers appeared to be primarily due to the vertical tail of the basic configuration and the interference effects associated with this tail. Secondary factors contributing to the development of the large-amplitude oscillations of the horizontal tail of the basic configuration were probably the loss in damping of the horizontal tail at transonic speeds and the turbulence of the airstream itself.

Description: At subsonic speeds the pressure drag arising from the thickness of the body or wings is negligible so long as the shapes are sufficiently well streamlined to avoid flow separation. In that range there exists no possibility of either favorable or adverse interference on the pressure distributions themselves. If one body is so placed as to receive a drag from the pressure field of another then the second body is sure to receive a corresponding increment of thrust from the first. At supersonic speeds this tolerance, which was permitted the designer, disappears and the drag becomes sensitive to the shape and arrangement of the bodies.To be sure, the primary factor here is the thickness ratio, but nevertheless there exist arrangements in which a large cancellation of drag occurs.

Description: Calibrations of the Friez Aerovane, Wind Measuring Set AN/GMQ-11, manufactured by the Friez Instrument Division of the Bendix Aviation Corporation, were made in the Langley 300 MPH 7- by 10-foot tunnel at the request of the Signal Corps, U, S. Army. Two propellers snd two generators were tested through a speed range of 15 to 190 knots, The results indicated that at airspeeds greater than 80 knots the instrument indicated airspeeds higher than the tunnel airspeed..

Description: The investigation of the lateral stability of an automatically controlled glide bomb led also to the attempt of clarifying the influence of a phugoid oscillation or of any general longitudinal oscillation on the lateral stability of a glide bomb. Under the assumption that its period of oscillation considerably exceeds the rolling and yawing oscillation and that c(sub a) is, at least in sections, practically constant, the result of this test is quite simple. It becomes clear that the influence of the phugoid oscillation may be replaced by suitable variation of the rolling-yawing moment on a rectilinear flight path instead of the phugoid oscillation. If the flying weight of the glide bomb of unchanged dimensions is increased, an increase of the flight velocity will be more favorable than an increase of the lift coefficient. The arrangement of the control permits lateral stability to be achieved in every case; a minimum rolling moment due to sideslip proves of great help.

Description: This paper is concerned primarily with the application of the "area rule" to the interpretation and improvement of the drag-rise characteristics of wing-body combinations at transonic and moderate supersonic speeds. Consideration of the general physical nature of the flow at transonic speeds, together with comparisons of the flow fields and drag-rise characteristics for wing-body combinations and bodies of revolution has led to the conclusion that near the speed of sound the drag rise for a thin low-aspect-ratio wing-body combination is primarily dependent on the axial distribution of cross-sectional area normal to the airstream (ref. 1). (The drag rise, sometimes referred to as pressure drag, is the difference between the drag level near the speed of sound and the drag level at subsonic speeds where the drag is due primarily to skin friction.) In order to illustrate the concept, figure 1 shows a wing-body combination and a body of revolution. A typical cross section normal to the airstream for the wing-body combination is shown at AA. The cross-sectional area of the wing is wrapped around the body of revolution so that the body has the same cross-sectional area at BB. All the other cross-sectional areas of the body of revolution are the same as those for the wing-body combination at the same axial stations. On the basis of the conclusion just stated, the drag rise for this body of revolution should be similar to that for the wing-body combination.

Description: The effects of inlet circumferential position around the fuselage on the characteristics of a half-conical scoop inlet having a 24.6deg half-angle cone have been investigated in the langley 4- by 4-foot supersonic pressure tunnel. Pressure-recovery results have been obtained at a Mach number of 2.01 for a fixed boundary-layer-bleed height which was 60 percent of the boundary-layer thickness at an angle of attack of 0deg, and for cowling position parameters of 42.4deg and 38.0deg. inlet had a capture area equal to 24.9 percent of the basic-fuselage frontal area. The angle of attack was varied from 0deg to 12deg. The most favorable pressure-recovery characteristics at angles of attack were obtained with the Inlet located on the bottom of the fuselage where the maximum recovery increased from a value of 81 percent at an angle of attack of 0deg to 87 percent at 12deg. In general, the pressure recovery decreased with increasing angle of attack for all other inlet locations. At a given angle of attack the pressure recovery decreased as the inlet location was progressively moved from the bottom to the top of the fuselage. Stable subcritical operation of the inlet with nearly constant pressure recovery was obtained for inlet mass-flow ratios from 1.0 to about 0.76 at an angle of attack of 0deg with the central body in the design position.

Description: Tests on equivalent bodies of revolution of six configurations of the Consolidated Vultee Aircraft Corporation proposed supersonic bomber (Convair MX-1964) have indicated that it is possible to reduce the drag of the configuration by designing it to have a favorable area distribution. The method of NACA RM L53I22c to predict the peak pressure drag of a configuration on the basis of its area distribution gave generally good agreement with the subject models.

Description: A flight test has been conducted to determine the longitudinal stability and control,characteristics of a 0.133-scale model of the Consolidated Vultee XFY-1 airplane without propellers for the Mach number range between 0.73 and 1.19.

Description: A flight investigation was made at high subsonic, transonic, and supersonic speeds and at high Reynolds numbers to determine the zero-lift drag of a 1/10-scale model of the Northrop MX-775A missile and a scale model of the missile fuselage. The model of the complete configuration has a 45deg swept wing of aspect ratio 5.5 and a 33deg swept vertical fin. The body model was stabilized by three 45deg swept fins. The-drag-rise Mach number for the model of the complete configuration was approximately 0.96. The drag coefficient based on total wing area was 0.0330 at Mach number 1.39. The drag coefficient of the body model less fin drag was approximately 55 percent that of the complete model at the same Mach number. Addition of the wing to the fuselage apparently resulted in a favorable drag interference near Mach number 1.0.

Description: An investigation of the longitudinal stability and of the all-movable horizontal tail and aileron control of a 1/80-scale reflection-plane model of the Consolidated Vultee Skate 9 seaplane has been made through a Mach number range of 0.6 to 1.16 on the transonic bump of the Langley high-speed 7- by 10-foot tunnel. At moderate lift coefficients (0.4 to 0.8) and below a Mach number of 1.0 the model was statically unstable longitudinally. The static longitudinal stability of the model at low lift coefficients increased with Mach number corresponding to a shift in aerodynamic center from 37 percent mean aerodynamic chord at a Mach number of 0.60 to 64 percent at a Mach number of 1.10. Estimates indicate that the tail deflection angle required for steady flight and for accelerated maneuvers of the Skate 9 airplane would probably not vary greatly with Mach number at sea level, but for accelerated maneuvers at altitude the tail deflection angle would probably vary erratically with Mach number. The variation of rolling-moment coefficient with aileron deflection angle was approximately linear, agreed well with theory, and held for the range of aileron deflections tested (-17.1 deg to 16.6 deg). At low lift coefficients the drag rise occurred at Mach numbers of 0.95 and 0.90 for the wing alone and the complete model, respectively. The effects of the canopy on the model were small. For the ranges investigated, angle-of-attack and Mach number changes caused no large pressure drops in the jet-engine duct.

Description: An investigation is being conducted to determine the dynamic stability and control characteristics of a 0.13-scale flying model of the Convair XFY-1 vertically rising airplane. This paper presents the results of flight tests to determine the stability and control characteristics of the model during constant-altitude slow transitions from hovering to normal unstalled forward flight. The tests indicated that the airplane can be flown through the transition range fairly easily although some difficulty will probably encountered in controlling the yawing motions at angles of attack between about 60 and 40. An increase in the size of the vertical tail will not materially improve the controllability of the yawing motions in this range of angle of attack but the use of a yaw damper will make the yawing motions easy to control throughout the entire transitional flight range. The tests also indicated that the airplane can probably be flown sideways satisfactorily at speeds up to approximately 33 knots (full scale) with the normal control system and up to approximately 37 knots (full scale) with both elevons and rudders rigged to move differentially for roll control. At sideways speeds above these values, the airplane will have a strong tendency to diverge uncontrollably in roll.

Description: An experimental investigation has been conducted in the Langley stability tunnel at low speed to determine the rolling stability derivatives of a 1/9-scale powered model of the Convair XFY-1 vertically rising airplane. Effects of thrust coefficient were investigated for the complete model and for certain components of the model. Effects of control deflections and of propeller blade angle were investigated for the complete model. Most of 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.

Description: An experimental investigation has been conducted in the Langley stability tunnel at low speed to deter+nine the yawing stability derivatives of a 1/9-scale powered model of the Convair XFY-1 vertically rising airplane. Effects of thrust coefficient were investigated for the complete model and for certain components of the model. Effects of control deflections and of propeller blade angle were investigated for the complete model. Most of 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.

Description: The results of free-flight drag tests of 40-millimeter shells conducted by the National Advisory Committee for Aeronautics for the Ballistic Research Laboratories, Ordnance Department, U. S. Army, are presented. A drag reduction at supersonic speeds of approximately 20 percent of the projectile's drag was obtained by combustion in the wake of the projectile in flight.

Description: An investigation of a 1/24-scale model of the Grumman F9F-6 airplane has been conducted in the Langley 20-foot free-spinning tunnel. The erect and inverted spin and recovery characteristics of the model were determined for the normal flight loading with the model in the clean condition. The effect of loading variations was investigated briefly. Spin-recovery parachute tests were also performed. The results indicate that erect spins obtained on the airplane in the clean condition will be satisfactorily terminated for all loading conditions provided full rudder reversal is accompanied by moving the ailerons and flaperons (lateral controls) to full with the spin (stick right in a right spin). Inverted spins should be satisfactorily terminated by full reversal of the rudder alone. The model tests indicate that an 11.4-foot (laid-out-flat diameter) tail parachute (drag coefficient approximately 0.73) should be effective as an emergency spin-recovery device during demonstration spins of the airplane provided the towline is attached above the horizontal stabilizer.

Description: An investigation has been conducted in the Langley 20-foot free-spinning tunnel on a l/20-scale model of the Consolidated Vultee XFY-1 airplane with a windmilling propeller simulated to determine the effects of control setting and movements upon the erect spin and recovery characteristics for a range of airplane-loading conditions. The effects on the model's spin-recovery characteristics of removing the lower vertical tail, removing the gun pods, and fixing the rudders at neutral were also investigated briefly. The investigation included determination of the size parachute required for emergency recovery from demonstration spins. The tumbling tendencies of the model were also investigated. Brief static force tests were made to determine the aerodynamic characteristics in pitch at high angles of attack. The investigation indicated that the spin and recovery characteristics of the airplane with propeller windmilling will be satisfactory for all loading conditions if recovery is attempted by full rudder reversal accompanied by simultaneous movement of the stick laterally to full with the spin (stick right in a right spin) and longitudinally to neutral. Inverted spins should be satisfactorily terminated by fully reversing the rudder followed immediately by moving the stick laterally towards the forward rudder pedal and longitudinally to neutral. Removal of the gun pods or fixing the rudders at neutral will not adversely affect the airplane's spin-recovery characteristics, but removal of the lower vertical tail will result in unsatisfactory spin-recovery characteristics. The model-test results showed that a 13.3-foot wing-tip conventional parachute (drag coefficient approximately 0.7) should be effective as an emergency spin-recovery device during demonstration spins of the airplane. It was indicated that the airplane should not tumble and that no unusual longitudinal-trim characteristics should be obtained for the center-of-gravity positions investigated.

Description: An investigation of the spin and recovery characteristics of a 1/24-scale model of the Grumman AF-2S, -2W airplane was conducted in the Langley 20-foot free-spinning tunnel. The effects of controls on the erect and inverted spin and recovery characteristics for a range of possible loadings of the.airplane were determined. The effect of a revised-tail installation (small dual fins added to the stabilizer of the original tail and the vertical-tail height of the original tail increased) and the effect of various ventral-fin and antispin-fillet installations were determined. The investigation also included spin-recovery parachute tests.

Description: A model of the Convair Y2-2 airplane was tested in Langley tank no. 2 to determine whether satisfactory stability in yawed landings was possible with a certain ventral fin. Free-body landings were made in smooth and rough water at two speeds and two rates of descent with the model yawed 15deg. The behavior of the model was determined by visual observations and from motion-picture re.cords. It was concluded that satisfactory stability was possible with the ventral fin as tested but that the characteristics of the model shock absorbers and the settings of the elevon control surfaces had an appreciable influence on behavior.

Description: An investigation has been made in the Langley 9- by 12-inch super-sonic blowdown tunnel at Mach numbers of 1.62 and 1.96 of a partial-span body with one tail surface, designed for use on the Hughes Falcon (MX-904) missile. The present paper extends the work reported in NACA-RM-SL50E10. Force and moment data including elevator hinge moment were obtained for the conditions of the tail in the presence of a small segment of the fore-shortened body, in the presence of a semi-span body and attached to a semi-span body, and for the condition of the foreshortened semi-span body alone.

Description: An investigation was conducted in the Langley high-speed 7- by 10-foot tunnel to determine effects of modifications to a bomb model (particularly with regard to drag) when mounted on a wing-fuselage model and tested at Mach numbers from 0.70 to 1.10. In addition, the static longitudinal stability characteristics of several configurations of a larger scale model of the bomb alone were obtained over a Mach number range from 0.50 to 0.95. The results obtained for the wing-fuselage-bomb model indicate that large reductions in installation drag were obtained for the wing-fuselage-bomb model when the flat nose of the basic bomb was replaced by rounded or pointed noses of various calibers. Shortening the mounting pylon gave further decreases in the installation drag. The tests of the bomb alone indicated that only the flat-nose configurations were stable over the greater part of the Mach number range. Nose-shape modifications which improved the drag also caused the bombs to become unstable at low angles of attack. The stability of the low-drag bomb configurations could be improved by lengthening the cylindrical portion of the body behind the center of gravity.

Description: An investigation has been conducted in the Langley 8-foot transonic tunnel to determine the effects of several fuselage modifications on the transonic drag-rise characteristics of a 1/20-scale model of the Convair F-102 airplane. Tests covered an angle-of-attack range from 0deg to about 10deg and a Mach number range from 0.60 to 1.14. Results indicated that the transonic drag rise .for the basic F-102 airplane could be substantially reduced by extending the fuselage after-body approximately 8 percent of the fuselage length. Tests of other bodies indicated that a shorter (4-percent) afterbody extension may have a similar effect on the drag rise. Further improvement of the axial cross-sectional-area distribution of the 8-percent extended configuration through the addition of fuselage volume resulted in additional reductions in the drag rise at a Mach number of 1.0 and caused no or only slight drag penalties at the higher Mach numbers. The results of the present tests generally substantiate the area-rule concept with respect to the prediction of the transonic drag rise through the use of an equivalent-area body of revolution for a practical delta-wing airplane configuration.

Description: An experimental investigation has been conducted to determine the dynamic stability and control characteristics in hovering and transition flight of a 0.13-scale flying model of the Convair XFY-1 vertically rising airplane with the lower vertical tail removed. The purpose of the tests was to obtain a general indication of the behavior of a vertically rising airplane of the same general type as the XFY-1 but without a lower vertical tail in order to simplify power-off belly landings in an emergency. The model was flown satisfactorily in hovering flight and in the transition from hovering to normal unstalled forward flight (angle of attack approximately 30deg). From an angle of attack of about 30 down to the lowest angle of attack covered in the flight tests (approximately 15deg) the model became progressively more difficult to control. These control difficulties were attributed partly to a lightly damped Dutch roll oscillation and partly to the fact that the control deflections required for hovering and transition flight were too great for smooth flight at high speeds. In the low-angle-of-attack range not covered in the flight tests, force tests have indicated very low static directional stability which would probably result in poor flight characteristics. It appears, therefore, that the attainment of satisfactory directional stability, at angles of attack less than 10deg, rather than in the hovering and transition ranges of flight is the critical factor in the design of the vertical tail for such a configuration.

Description: The wide use of diffusers, in various fields of technology, has resulted in several experimental projects to study the action and design of diffusers. Most of the projects dealt with steam (steam turbine nozzles). But diffusers have other applications - that is, ventilators, smoke ducts, air coolers, refrigeration, drying, and so forth. At present there is another application for diffusers in wind-tunnel design. Because of higher requirements and increased power of such installations more attention must be paid to the correctness of work and the decrease in losses due to every section of the tunnel. A diffuser, being one of the component parts of a tunnel , can in the event of faulty construction introduce considerable losses. Therefore, in the design of the new CAHI wind tunnel, it was suggested that an experimental study of diffusers be made, with a view to applying the results to wind tunnels. The experiments conducted by K. K. Baulin in the laboratories of CAHI upon models of diffusers of different cross sections, lengths, and angles of divergence, were a valuable source of experimental data. They were of no help, however, in reaching any conclusion regarding the optimum shape because of the complexity and diversity of the factors which all appeared simultaneously, thereby precluding the.study of the effects of any one factor separately. On the suggestion of the director of the CAHI,Prof. B. N. Ureff, it was decided to experiment on a two-dimensional diffuser model and determine the effect, of the angle of divergence. The author is acquainted with two experimental projects of like nature: the first was conducted with water, the other with air. The first of these works, although containing a wealth of experimental data, does not indicate the nature of flow or its relation to the angle of divergence. The second work is limited to four angles - that is, 12 deg, 24 deg, 45 deg, 90 deg. The study of this diffuser did not supply any information about the effect of smaller angles which, because of their advantages, are more commonly used, The author was able to acquaint himself with the second work only after the experiments were started. For these reasons, as well as because on the basis of those works no conclusion can be reached regarding the nature of flow distribution, of eddies, and so forth, experimental work was continued. The need for determining flow patterns follows from the fact that from them are determined methods of measurement - that is, the determination of velocities by means of the pitot tube, which, as is well known, gives correct indications only when placed with its axis parallel to the axis of flow. The data contained in this report were obtained from experiments conducted by the Aerodynamical Laboratories of the CAHI. The solutions to some. of the mathematical problems connected with the experiments are due to Prof. S. A, Chapligin.

Description: A low-speed wind-tunnel investigation has been made of some aspects of the aerodynamic problems associated with the use of air-to-air missiles when carried externally on aircraft. Measurements of the forces and moments on a missile model for a range of positions under the mid-semispan location of a 45deg sweptback wing indicated longitudinal and lateral forces with regard to both carriage and release of the missiles. Surveys of the characteristics of the flow field in the region likely to be traversed by the missiles showed abrupt gradients in both flow angularity and in local dynamic pressure. Through the use of aerodynamic data on the isolated missile and the measured flow-field characteristics, the longitudinal forces and moments acting on the missile while in the presence of the wing-fuselage combination could be estimated with fair accuracy. Although the lateral forces and moments predicted were qualitatively correct, there existed some large discrepancies in absolute magnitude.

Description: Tests were made in the Langley 8-foot high-speed tunnel to investigate the aerodynamic characteristics of the D-558-1 airplane and various wing and tail configurations on the D-558-1 fuselage. The various wing and tail configurations were tested to determine the aerodynamic effects of aspect ratio and sweep for suitable use on the second phase of the D-558 project (D-558-2). The tests were conducted through a speed range from a Mach number of 0.40 to approximately 0.94.This part of the investigation includes the lift and drag results available for the configurations tested at this rate. The D-558-1 results indicated that the lift force break would occur at a Mach number of 0.85 with some reduction in lift at speeds above this Mach number. Tests indicated that the airplane will have satisfactory lift and drag characteristics up to and including its design Mach number of 0.85. The 35deg sweptback, 35deg swept-forward, and low-aspect-ratio (2.0) wing configurations all showed pronounced improvements in maintaining lift throughout the Mach number range tested and in increasing the critical speeds above the D-558-1 value &itical to critical Mach numbers on the order of 0.9. Insofar as lift and drag characteristics are concerned level flight at speeds approaching the velocity of sound appears practical if swept or low-aspect-ratio configurations similar to those tested are used.

Description: Experimental results showing the static longitudinal-stability and control characteristics of a model of a fighter airplane employing a low-aspect-ratio unswept wing and an all-movable horizontal tail are presented. The investigation was made over a Mach number range from 0.60 to 0.90 and from 1.35 to 1.90 at a constant Reynolds number of 2.40 million, based on the wing mean aerodynamic chord. Because of the location of the horizontal tail at the tip of the vertical tail, interference was noted between the vertical tail and the horizontal tail and between the wing and the horizontal tail. This interference produced a positive pitching-moment coefficient at zero lift throughout the Mach number range of the tests, reduced the change in stability with increasing lift coefficient of the wing at moderate lift coefficients in the subsonic speed range, and reduced the stability at low lift coefficients at high supersonic speeds. The lift and pitching-moment effectiveness of the all movable tail was unaffected by the interference effects and was constant throughout the lift-coefficient range of the tests at each Mach number except 1.90.

Description: This paper presents the results of an investigation of the dynamic stability and controllability of a model which approximately represents the Lockheed XFV-1 airplane to a 1/8 scale. The investigation consisted of hovering flights in still air at a considerable height above the ground, hovering flights very close to the ground, vertical take-offs and landings, flights through the transition range from hovering to normal forward flight, and sideways translational flights. The model could be flown smoothly and easily in hovering flight despite the fact that the uncontrolled pitching and yawing motions were unstable oscillations. There was a noticeable reduction in the controllability of the model when hovered very close to the ground but take-offs could be made easily and landings on a g,ven spot could be made accurately in spite of this adverse ground effect. Flights through the transition range from hovering to normal forward flight could be performed fairly easily. The model seemed to have stability of angle of attack and angle of roll over most of the transition range. The yawing motion was divergent in the very high angle-of-attack range but could be controlled easily. At the lower angles of attack, the model seemed to become stable in yaw. In sideways flight there was an increasingly strong tendency to diverge in roll as the speed was increased and finally, at a speed of about 25 knots (full scale), the model rolled off despite efforts of the pilot to control it.

Description: A low-speed investigation was made of a 1/6-scale model of the Republic XF-84H airplane. The model had a single tractor propeller and a 40deg swept wing of aspect ratio 3.45. This investigation was undertaken to provide information on the effects of propeller operation on longitudinal stability characteristics for the XF -84H airplane and to provide an indication of slipstream effects that might be encountered on similar swept-wing configurations. Effects of propeller operation were generally destabilizing for all conditions investigated; however, the over-all stability characteristics with power on were greatly dependent on the power-off characteristics. With flaps and slats retracted, longitudinal instability was present at moderate angles of attack both with the propeller off and with power on. The longitudinal stability with flaps and slats deflected, which was satisfactory without power, was decreased by propeller operation, but no marked pitch-up tendency was indicated. Significant improvement in the power-on stability with flaps retracted was achieved by use of either a wing fence at 75 percent semispan, a leading-edge chord-extension from 65 to 94 percent semispan, or a raised horizontal tail located 65 percent semispan above the thrust line.

Description: Models of the Hermes A-3B missile were tested in the Ames supersonic free-flight wind tunnel to determine the static-longitudinal-stability characteristics at a Mach number of 5.0 and a Reynolds number based on body length of 10 million. The results indicated that the model center of pressure was 45.3 percent of the body length aft of the nose and the lift-curve slope based on body frontal area was 0.064 per degree. Estimates indicated that the effect on these characteristics of aeroelastic twisting of the model fins was small but important if a precise location of center of pressure is required. A comparison of the test results with predictions based on available theory showed that the theory was useful only for rough estimates, The drag coefficient at zero lift, based on body frontal area, was found to be 0.155.

Description: Drag and longitudinal trim at low lift of the North American YF-100A airplane at Mach numbers from 0.76 to 1.77 as determined from the flight test of a 0.11-scale rocket model are presented herein. Also included are some longitudinal stability and some qualitative pitch-damping data. The subsonic external-drag-coefficient level was about 0.012, and the supersonic level was about 0.043. The drag rise occurred at a Mach number of 0.95. The longitudinal trim change at low lift consisted basically of a mild nose-up tendency at a Mach number of 0.90. An indication of wing flutter was present at Mach numbers from 0.95 to 1.11. However, the full-scale airplane wing has approximately twice the scaled first-bending frequency as the model tested and, hence, will probably be free of this type of flutter. The aerodynamic-center location was 71 percent behind the leading edge of the mean aerodynamic chord at a Mach number of 1.03 and 62 percent at a Mach number of 1.74. Qualitative measurement of damping in pitch indicates that at low lift coefficients damping will be low at a Mach number of 1.03.

Description: An investigation has been conducted in the Langley 20-foot free-spinning tunnel on a 1/30 - scale model of the Grumman XFlOF-1 airplane to determine its spin and recovery characteristics. The investigation included erect and inverted spins for both the straight-wing and swept-wing configurations. Tests to determine the optimum size spin-recovery parachutes and the rudder forces required for recovery were also made. The results indicated that in the straight-wing configuration, satisfactory recoveries of the airplane will be obtained from erect and inverted spins by rudder reversal alone. In the swept-wing configuration recoveries will be unsatisfactory from erect spins. Unsweeping the wings during the spin and reversal of the rudder, however, will lead to eventual recovery. The test results also indicated that, if existing small ailerons are made deflectable through large angles, satisfactory recoveries will be obtained from erect spins in the swept-wing configuration by simultaneous movement of the rudder to against the spin and movement of the ailerons to with the spin. Normal-size ailerons deflected through a normal range would also be effective. Satisfactory recoveries by rudder reversal will be obtained from inverted spins in the swept-wing configuration. In the straight-wing configuration a 14.2-foot tail parachute or a 5.0-foot wing-tip parachute opened on the outer wing tip will effect satisfactory recovery of the airplane by parachute action alone; a 30.0-foot tail parachute or a 10.0-foot wing-tip parachute will be required for the swept-wing configuration. The forces required to fully reverse the rudder should be within the capabilities of the pilot.

Description: An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.41 to determine the static stability and control and drag characteristics of a l/l5-scale model of the Grunman F9F-9 airplane. The effects of alternate fuselage shapes, wing camber, wing fences, and fuselage dive brakes on the aerodynamic characteristics were also investigated. These tests were made at a Reynolds number of 1.96 x l0 (exp 6) based on the wing mean aerodynamic chord of 0.545 foot. The basic configuration had a static margin of stability of 38.4 percent of the mean aerodynamic chord and a minimum drag coefficient of 0.049. For the maximum horizontal tail deflection investigated (-l0 deg), the maximum trim lift coefficient was 0.338. The basic configuration had positive static lateral stability at zero angle of attack and positive directional control throughout the angle-of-attack range investigated up to ll deg.

Description: The longitudinal stability and control characteristics of a 1/30-scale model of the Republic XF-103 airplane were investigated in the Langley 8-foot transonic tunnel. The effect of speed brakes located at the end of the fuselage was also investigated. The main part of the investigation was made with internal flow in the model, but some data were obtained with no internal flow. The longitudinal stability and control at transonic-speeds appeared satisfactory. The transonic drag rise was small. The speed brakes had no adverse effects on longitudinal stability.

Description: Dynamic--response measurements for various conditions of displacement and rate signal input, sensitivity setting, and simulated hinge moment were made of the three control-surface servo systems of an NAES-equipped remote-controlled airplane while on the ground. The basic components of the servo systems are those of the General Electric Company type G-1 autopilot using electrical signal. sources, solenoid-operated valves, and hydraulic pistons. The test procedures and difficulties are discussed, Both frequency and transient-response data, are presented and comparisons are made. The constants describing the servo system, the undamped natural frequency, and the damping ratio, are determined by several methods. The response of the system with the addition of airframe rate signal is calculated. The transfer function of the elevator surface, linkage, and cable system is obtained. The agreement between various methods of measurement and calculation is considered very good. The data are complete enough and in such form that they may be used directly with the frequency-response data of an airplane to predict the stability of the autopilot-airplane combination.

Description: Preliminary data are presented on the variation of the maximum lift coefficient with Mach number. The data were obtained from tests in the 8-foot high-speed tunnel of three NACA 16-series airfoils of 1-foot chord. Measurements consisted primarily of pressure-distribution measurements in order to illustrate the nature of the phenomena. It was found that the maximum lift coefficient of airfoils is markedly affected by compressibility even at Mach numbers as low as 0.2. At high Mach numbers pronounced decrease of the maximum lift coefficient was found. The magnitude of the effects of compressibility on the maximum lift coefficient and the low speeds at which these effects first appear indicate clearly that consideration of the take-off thrust for propellers will give results seriously in error if these considerations are based on the usual low-speed maximum-lift-coefficient data generally used.

Description: An investigation has been made to determine the motions of and the flight paths describe by a Navy dive-bombing airplane in simulated diving attacks. The data necessary to evaluate these items, with the exception of the atmospheric wind data, were obtained from automatic recording instruments installed entirely within the airplane. The atmospheric wind data were obtained from the ground by the balloon-theodolite method. The results of typical dives at various dive angles are presented in the form of time histories of the motion of the airplane as well as flight paths calculated with respect to still air and with respect to the ground.

Description: This report covers tests of a 5-foot model of the NACA 66,2-420 low-drag airfoil at high speeds including the critical compressibility speed. Section coefficients of lift, drag, and pitching moment, and extensive pressure-distribution data are presented. The section drag coefficient at the design lift coefficient of 0.4 increased from 0.0042 at low speeds to 0.0052 at a Mach number of 0.56 (390 mph at 25,000 ft altitude). The critical Mach number was about 0.60. The results cover a Reynold number range from 4 millions to 17 millions.

Description: An investigation was conducted in the NACA 19-foot pressure wind tunnel of a tapered wing with straight railing edge having NACA 66 series low-drag airfoil sections and equipped with full-span and partial-span simple split flaps. The airfoil sections used were the NACA 66,2-116 at the root and the 66,2-216 at the tip. The primary purpose of the investigation was to determine the effect of the split flaps on the aerodynamic characteristics of the tapered wing. Complete lift, drag, and pitching-moment coefficients were determined for the plain wing and for each flap arrangement through a Reynold number range of 2,600,000 to 4,600,000. The results of this investigation indicate that values of maximum lift coefficient comparable to values obtained on tapered wings with conventional sections and similar flap installations can be obtained from wings with the NACA low-drag sections. The increment of maximum lift due to the split flap was found to vary somewhat with Reynold number over the range investigated. The C(sub L)max of the wing alone is 1.49 at a Reynolds number of 4,600,000; whereas with the partial-span simple split flap it is 2.22 and with the full-span arrangement, 2.80. Observations of wool tufts on the wing indicate that the addition of split flaps did not appreciable alter the pattern of the stall; even though the stall did occur more abruptly than with the wing alone.

Description: An investigation was conducted in the NACA 19-foot pressure wind tunnel of a rectangular wing having NACA 66, 2-216 low-drag airfoil sections and various sizes of simple split flaps. The purpose of the investigation was, primarily, to determine the influence of these flap installations on the aerodynamic characteristics of the wing. Complete lift, drag, and pitching-moment characteristics were determined for a range of test Reynolds numbers from about 2,600,000 to 4,600,000 for each of the installations and for the plain wing. The results of this investigation indicate that values of maximum lift coefficient similar to those of wings with conventional airfoil sections and split flaps can be expected of wings having the NACA 66,2-216 low-drag sections. The increment of maximum lift due to the split flap was found to be practically independent of the Reynolds number over the range investigated. The optimum split flap on the basis of maximum lift appears to have a chord about 20% of the wing chord and a deflection of 60 degrees. The C(sub L) max of the wing with the 0.20c partial-span flap deflected 60 degrees is 2.07 at a Reynolds number of 4,600,000 while with the full-span flap it is approximately 2.53; the increment of the maximum lift coefficient due to the flap is approximately proportional to the flap span. Although the addition of a split flap tends to hasten the stall and to cause it to occur more abruptly, little change in pattern is evidenced by observations of the behavior of wool tufts on the wing.

Description: An investigation was made in the NACA 19-foot pressure wind tunnel to determine the effect of various win-gun installation on the aerodynamic characteristics of a model with an NACA low-drag wing. Measurements were made of lift and drag over an angle-of-attack range and for several values of dynamic pressure on a four-tenths scale model of a high-speed airplane equipped with the low-drag wing and with various wing-gun installations. Two installations were tested: one in which the blast tube and part of the gun barrel protrude ahead of the wing and another in which the guns is mounted wholly within the wing. Two types of openings for the latter installation were tested. For each installation three simulated guns were mounted in each wing. The results are given in the form of nondimensional coefficients. The installations tested appear to have little effect on the maximum-lift coefficient of the model. However, the drag coefficient shows a definite change. The least adverse effect was obtained with the completely internal mounting and small nose entrance. The results indicate that a properly designed wing-gun installation will have very little adverse effect on the aerodynamic characteristics of the low-drag wing.

Description: At the request of the Bureau of Aeronautics, an investigation was conducted in the full-scale wind tunnel of wing installations of .50-caliber machine guns and 20-millimeter cannons. The tests were made to determine the effect of various gun installations on the maximum lift and the high-speed drag of the airplane.

Description: Results of a study to determine the effects on turning performance due to various assumed modifications to a typical Naval fighter airplane are presented. The modifications considered included flaps of various types, both part and full space, increased supercharging, and increased wing loading. The calculations indicated that near the low-speed end of the speed range, the turning performance, as defined by steady level turns at a given speed, would be improved to some extent by any of the flaps considered at altitudes up to about 25,000 feet. (If turning is not restricted to the conditions of no loss of speed or altitude, more rapid turning can, of course, be accomplished with the aid of flaps, regardless of altitude.) Fowler flaps and NACA slotted flaps appeared somewhat superior to split or perforated split flaps for maneuvering purposes, particularly if the flap position is not adjustable. Similarly, better turning performance should be realized with full-span than with part-span flaps. Turning performance over the lower half of the speed range would probably not be materially improved at any altitude by increased supercharging of the engine unless the propeller were redesigned to absorb the added power more effectively; with a suitable propeller the turning performance at high altitudes could probably be greatly improved with increased supercharging. A reduction in wing area with the aspect ratio held constant would result in impairment of turning performance over practically the entire speed range at all altitudes.

Description: The effects of direction of propeller rotation on factors affecting the longitudinal stability of the XB-28 airplane were measured on a 1/10-scale model in the 7- by 10-foot tunnel of the Ames Aeronautical Laboratory. The main effect observed was that caused by regions of high downwash behind the nacelles (power off as well as power on with flaps neutral). The optimum direction of propeller rotation, both propellers rotating up toward the fuselage, shifted this region off the horizontal tail and thus removed its destabilizing effect. Rotating both propellers downward toward the fuselage moved it inboard on the tail and accentuated the effect, while rotating both propellers right hand had an intermediate result. Comparisons are made of the tail effects as measured by force tests with those predicted from the point-by-point downwash and velocity surveys in the region of the tail. These surveys in turn are compared with the results predicted from available theory.

Description: Tests were made in the NACA 19-foot pressure tunnel of a simplified twin-engine bomber model with an NACA low-drag wing primarily to obtain an indication of the effects of engine nacelles on the characteristics of the model both with and without simple split trailing-edge flaps. Nacelles with conventional-type cowlings representative of those used on an existing high-performance airplane and with NACA high-speed type E cowlings were tested. The tests were made without propeller slipstream. The aerodynamic effects of adding the nacelles to the low-drag wing were similar to the effects commonly obtained by adding similar nacelles to conventional wings. The maximum lift coefficient without flaps was slightly increased, but the increment in maximum lift due to deflecting the flaps was somewhat decreased. The stalling characteristics were improved by the presence of the nacelles. Addition of the nacelles had a destabilizing effect on the pitching moments, as is usual for nacelles that project forward of the wing. The drag increments due to the nacelles were of the usual order of magnitude, with the increment due to the nacelles with NACA type E cowlings approximately one-third less than that of the nacelles with conventional cowlings with built-in air scoops.

Description: An investigation was made in the LWAL 7- by 10-foot tunnel of internally balanced, sealed ailerons for the Curtiss XP-60 airplane. Ailerons with tabs and. with various amounts of balance were tested. Stick forces were estimated for several aileron arrangements including an arrangement recommended for the airplane. Flight tests of the recommended arrangement are discussed briefly in an appendix, The results of the wind-tunnel and flight tests indicate that the ailerons of large or fast airplanes may be satisfactorily balanced by the method developed.

Description: As an aid to airplane designers interested in providing pursuit airplanes with decelerating devices intended to increase the firing time when overtaking another airplane, formulas are given relating the pertinent distances and speeds in horizontal flight to the drag increase required. Charts are given for a representative parasite-drag coefficient from which the drag increase, the time gained, and the closing distance may be found. The charts are made up for three values of the ratio of the final speed of the pursuing airplane to the speed of the pursued airplane and for several values of the ratio of the speed of the pursued airplane to the initial speed of the pursuing airplane. Charts are also given indicating the drag increases obtainable with double split flaps and with conventional propellers. The use of the charts is illustrated by an example in which it is indicated that either double split flaps or, under certain ideal conditions, reversible propellers should provide the speed reductions required.

Description: Tests were made in the NACA two-dimensional low turbulence pressure tunnel of a highly cambered low-drag airfoil (NACA 65,3-618) with a plain flap designed for lift control. The results indicate that such a combination offers attractive possibilities for obtaining low profile-drag coefficients over a wide range of lift coefficients without large reductions of critical speed.

Description: The development of the laminar boundary layer in a compressible fluid is considered. Formulas are given for determining the boundary-layer thickness and the ratio of the boundary-layer Reynolds number to the body Reynolds number for airfoils and bodies of revolution. It i s shown that the effect of compressibility will profoundly alter the Reynolds number corresponding to the upper limit of the range of the low-drag coefficients . The available data indicate that for low-drag and high critical compressibility speed airfoils and bodies of revolution, this effect is favorable.

Description: An investigation has been made in the NACA 7- by 10-foot wind tunnel of a large chord NACA 27-212 airfoil with a 20% chord split flap and with two arrangements of a 25.66% chord slotted flap to determine the section lift characteristics as affected by flap deflection for the split flap and as affected by flap deflection, flap position, and slot shape for the slotted flap. For the two arrangements of the slotted flap, the flap positions for maximum section lift are given. Comparable data on the NACA 23012 airfoil equipped with similar flaps are also given. On the basis of maximum section lift coefficient, the slotted flap with an easy slot entry was slightly better than either the split flap or the slotted flap with a sharp slot entry. With both types of flap the decrease in the angle of attack, for maximum section lift coefficient, with flap deflection is large for the NACA 27-212 airfoil as compared with the NACA 23012 airfoil. Also with both flaps, the maximum section lift coefficient obtained with flaps is much lower for the NACA 27-212 airfoil than for the NACA 23012 airfoil.

Description: Wind-tunnel tests of several propeller, cuff, and spinner combinations were conducted in the 20 foot propeller-research tunnel. Three propellers, which ranged in diameter from 8.4 to 11.25 feet, were tested at the front end of a streamline body incorporating spinners of two diameters. The tests covered a blade angle range from 20 deg to 65 deg. The effect of spinner diameter and propeller cuffs on the characteristics of one propeller was determined. Test were also conducted using a propeller which incorporated aerodynamically good shank sections and using one which incorporated the NACA 16 series sections for the outer 20 percent of the blades. Compressibility effects were not measured, owing to the low testing speeds. The results indicated that a conventional propeller was slightly more efficient when tested in conjunction with a 28 inch diameter spinner than with a 23 inch spinner, and that cuffs increased the efficiency as well as the power absorption characteristics. A propeller having good aerodynamic shanks was found to be definitely superior from the efficiency standpoint to a conventional round-shank propeller with or without cuffs; this propeller would probably be considered structurally impracticable, however. The propeller incorporating the NACA 16 series sections at the tims were found to have a slightly higher efficiency than a conventional propeller; the take-off characteristics appeared to be equally good. The effects noted above probably would be accentuated at helical speeds at which compressibility effects would enter.

Description: Determinations of boundary-layer transition on the NACA 0012 and 2301 airfoils were made in the 8-foot high-speed wind tunnel over a range of Reynolds Numbers from 1,600,000 to 16,800,000. The results are of particular significance as compared with flight tests and tests in wind tunnels of appreciable turbulence because of the extremely low turbulence in the high-speed tunnel. A comparison of the results obtained on NACA 0012 airfoils of 2-foot and 5-foot chord at the same Reynolds Number permitted an evaluation of the effect of compressibility on transition. The local skin friction along the surface of the NACA 0012 airfoil was measured at a Reynolds Number of 10,000,000. For all the lift coefficient at which tests were made, transition occurred in the region of estimated laminar separation at the low Reynolds Numbers and approach the point of minimum static pressure as a forward limit at the high Reynolds Numbers. The effect of compressibility on transition was slight. None of the usual parameters describing the local conditions in the boundary layer near the transition point served as an index for locating the transition point. As a consequence of the lower turbulence in the 8-foot high-speed tunnel, the transition points occurred consistently farther back along the chord than those measured in the NACA full-scale tunnel. An empirical relation for estimating the location of the transition point for conventional airfoils on the basis of static-pressure distribution and Reynolds Number is presented.

Description: The behavior of the Westinghouse electronic power regulator operating on a J34-WE-32 turbojet engine was investigated in the NACA Lewis altitude wind tunnel at the request of the Bureau of Aeronautics, Department of the Navy. The object of the program was to determine the, steady-state stability and transient characteristics of the engine under control at various altitudes and ram pressure ratios, without afterburning. Recordings of the response of the following parameters to step changes in power lever position throughout the available operating range of the engine were obtained; ram pressure ratio, compressor-discharge pressure, exhaust-nozzle area, engine speed, turbine-outlet temperature, fuel-valve position, jet thrust, air flow, turbine-discharge pressure, fuel flow, throttle position, and boost-pump pressure. Representative preliminary data showing the actual time response of these variables are presented. These data are presented in the form of reproductions of oscillographic traces.

Description: Flight tests were conducted between Mach numbers of 0.9 and 1.8 over a Reynolds number range of 9(exp 6) to 30(exp 6) to determine the zero-lift drag and some rolling-effectiveness characteristics of the Northrop MX -775B missile with small and large body. The MX-775B is a proposed long range, supersonic, ground-to-ground missile having an arrow wing with 67.5 degree leading-edge sweep, 15 deg trailing-edge sweep, and a modified NACA 0004 airfoil section. The configuration has no horizontal tail but has wing trailing-edge elevons which serve a dual purpose as elevators and ailerons. The ratio of body frontal area to wing plan-form area is 0.0127 for the small-body configuration and 0.0330 for the large-body configuration. Five 1/4-scale models were flown permitting determination of the drag coefficient for the basic small-body configuration, the incremental drag due to the large body, the incremental drag resulting from a blunt wing trailing edge, the wing-plus-interference drag, and some rolling-effectiveness data. Results indicated that the MX-775B has low supersonic zero-lift drag, the maximum zero-lift drag coefficients being respectively 0.0125 and 0.0155 at a Mach number of M = 1803 for the small- and large-body configurations. The effect of a blunt wing trailing edge, obtained by cutting off 10 percent of the wing chord, was to increase the zero-lift drag by 13 to 21 percent. Wing-plus-interference drag accounted for 78 percent of the total drag at M = 0.9 and 70 percent at M = 195 for the small-body configuration. The ailerons produced positive rolling effectiveness for the wing stiffness of the test models and the dynamic pressures of the test.

Description: The aerodynamic characteristics in pitch of an F-94C airplane, with the primary attention given to its drag characteristics, have been evaluated at low speed in the Ames 40- by 80-foot wind tunnel. The increments of drag due to various surface irregularities, ports, and component parts of the production airplane were determined. Wing-wake surveys were taken to determine the section drag coefficients at midsemispan for the smooth and the production wing. Base-pressure and internal drags of the air-induction system were measured at low inlet-velocity ratios. The characteristics of the airplane in the landing configuration are also included.

Description: The data obtained from the flight of a simplified (dummy) rocket-propelled model of the MX-656 have been analyzed to determine the booster-model characteristics and the model-alone characteristics up to a Mach number of 1.3. The data indicate that the model-booster combination is satisfactory. The model alone is longitudinally stable i n the Mach number range covered by the test (0.9 to 1.3) with the center of gravity at -15 percent of the mean aerodynamic chord. With the stabilizer setting at 0 deg. the variation of normal-force coefficient with Mach number is not large. The total-drag-coefficient variation with Mach number is not unusual. About 12 percent of the total drag at a Mach number of 1.3 can be attributed to body base drag.

Description: At the request of the Bureau of Aeronautics, Department of the Navy, National Advisory Committee for Aeronautics has conducted a preliminary investigation at high subsonic speeds of the static longitudinal and lateral stability characteristics of a 0.05-scale model of the Convair F2Y-1 water-based fighter airplane. The tests covered a Mach number range from 0.5 to 0.94 and corresponding Reynolds numbers, based on the wing mean aerodynamic chord, from 3.3 x 10(exp 6) to 4.3 x 10(exp 6). The maximum angle-of-attack range (obtained at the lower Mach numbers) was from -2 degrees to 25 degrees. Sideslip angles from -4 degrees to 12 degrees also were investigated. The investigation included effects of various arrangements of wing fences and of rocket packages.

Description: The performance of a 16-stage axial-flow compressor, in which two modifications of unloaded inlet stages were combined with loaded exit stages, has been determined. In the first modification the exit stages were loaded by decreasing the twelfth through fifteenth stage stator angles 3 deg. as compared with the blade angles in the original compressor, and the inlet stages were unloaded by increasing the blade angles the following amounts: guide vanes and first-stage stator, 6 deg; second- and third-stage stators, 4 deg.; and fourth-stage stators, 3 deg. The over-all performance of this configuration was compared with that of the compressor with the original blade angles. The peak efficiency was increased at all speeds below design and the weight flow was higher at speeds below 80 percent of design, the same at 80 percent of design, and lower at speeds abovce 80 percent of design. The maximum reduction in weight flow occurred at design speed. The surge limit line was higher at speeds between 75 and 90 percent of design when presented on a pressure ratio against weight flow basis. The second configuration was the same as the first with the exception that the second-, third-, and fourth-stage stator blade angles were the same as in the compressor with the original blade angles. A comparison of the performance of this configuration with that of the compressor with the original blade angles showed the same general trends of changes in performance as the first configuration. Comparisons were made of compressor configurations to show the effects upon the performance of decreased loading in the inlet stages. Below 75 percent of design speed, decreased loading results in increased weight flow and peak efficiency; above 80 percent of design speed, decreased loading in the inlet stages results in decreased weight flow and small changes in peak efficiencies. Between 75 and 90 percent of design the changes in surge weight flow and pressure ratio were such that the surge limit line was raised with decreased loading in the inlet stages when presented as pressure ratio against weight flow.

Description: An investigation of the aerodynamic characteristics of an 0.025-scale model of the MX-1712 configuration has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel. The tests were performed at Mach numbers of 1.41 and 2.01 at a Reynolds number of approximately 2.6 x 10(exp 6) based on the wing mean aerodynamic chord The MX-1712 is a proposed swept-wing, jet-powered supersonic bomber aircraft. The wing is of aspect ratio 3.5, taper ratio 0.2, and thickness ratio 5.5 percent (streamwise) and has 47deg sweep of the quarter-chord line. The longitudinal and lateral force characteristics of the model and various combinations of its components, including several nacelle installations, were investigated. The effects of a modified wing, two horizontal tail positions, and a shortened fuselage were also studied. The results obtained from these investigations are presented in this report. The aerodynamic investigation of this model disclosed no unusual stability characteristics or Mach number effects. The choice of nacelle installations appears to be a major decision, one greatly affecting the performance of the airplane, At M = 1.41 and C(sub L) = 0.1, the buried nacelles increased the drag of the basic model by 9 percent, while the best pod nacelles increased the drag of the basic model by 27 percent.

Description: A rocket-propelled model of the Mx-656 configuration has been flown through the Mach number range from 0.65 to 1.25. An analysis of the response of the model to rapid deflections of the horizontal tail gave information on the lift, drag, longitudinal stability and control, and longitudinal-trim change. The lift-coefficient range covered by the test was from -0.2 to 0,3 throughout most of the Mach number range, The model was statically and dynamically stable throughout the lift-coefficient and Mach number range of the test. At subsonic speeds the aerodynamic center moved f o m r d with increasing lift coefficient. The most forward position of the aerodynamic center was about 12,5 percent of the mean aerodynamic chord at a small positive lift coefficient and at a Mach number of about 0.84. A t supersonic speeds the aerodynamic center was well aft, varying from 33 to 39 percent of the mean aerodynamic chord at Mach numbers of 1.0 and 1.25, respectively. Transonic-trim change, as measured by the change in trim lift coefficient with Mach number at a constant t a i l setting, was of small magnitude (about 0.1 lift coefficient for zero tail setting). The zero-lift/drag coefficient increased about 0.042 in the region between a Mach number of 0.9 and 1.1

Description: A 1/5-scale, rocket-propelled model of the Convair F-102 configuration was tested in free flight to determine zero-lift drag at Mach numbers up to 1.34 and at Reynolds numbers comparable to those of the full-scale airplane. This large-scale model corresponded to the prototype airplane and had air flow through the duct. Additional zero-lift drag tests involved a series of small equivalent bodies of revolution which were launched by means of a helium gun. The several small-scale models tested corresponded to: the basic configuration, the 1/5-scale rocket-propelled model configuration, a 2-foot (full-scale) fuselage-extension configuration, and a 7-foot (full-scale) fuselage-extension configuration. Models designed to correspond to the area distribution at a Mach number of 1.0 were flown for each of these 'shapes and, in addition, models designed to correspond to the area distribution at a Mach number of 1.2 were flown for the 1/5-scale rocket-propelled model and the 7-foot-fuselage-extension configuration. The value of external pressure drag coefficient (including base drag) obtained from the large-scale rocket model was 0.0190 at a Mach number of 1..05 and the corresponding values from the equivalent-body tests varied from 0.0183 for the rocket-propelled model shape to 0.0137 for the 7-foot-fuselage-extension configuration. From the results of tests of equivalent bodies designed to correspond to the area distribution at a Mach number of 1.0, it is evident that the small changes in shape incorporated in the basic and 2-foot-fuselage-extension configurations from that of the rocket-propelled model configuration will provide no significant change in pressure drag. On the other hand, the data from the 7-foot-fuselage-extension model indicate a substantial reduction in pressure drag at transonic speeds.

Description: At the request of the Bureau of Aeronautics, Navy Department, tests were made in the variable-density wind tunnel of a tapered wing of 3-10-18 plan form and based on the NACA 4400R series sections. The wing was also tested with 0.2 chord spit flaps, deflected 60 deg span ratios of 0.3, 0.5, 0.7 and 1.0 respectively. In order to get data from which to calculate the characteristics of the flapped wing, the investigation was extended to include tests of the four rectangular airfoils of the NACA 4400R series (4409R, 4412R, 4415R, and 4418R) with full-span 0.2 chord, trailing edge split flaps deflected 60 deg.

Description: After conclusion of the spin investigation of the model Me 210 with elongated fuselage and central vertical tail surfaces (model condition III; reference 3), tests were performed on the same model with a vee tail (model condition IV). Here the entire tail surfaces consist of only one surface with pronounced dihedral. Since the blanketing of the vertical tail surfaces by the horizontal tail surfaces, which may occur in case of standard tail surfaces, does not occur here, one could expect for this type of tail surface favorable spin characteristics, particularly with respect to rudder effectiveness for spin recovery. However, the test results did not confirm these expectations. The steady spin was shown to be very irregular; regarding rudder effectiveness the vee tail surfaces proved to be inferior even to standard tail surfaces, thus they represent the most unfavorable of the four fuselage and tail-surface combinations investigated so far.

Description: At present there is no satisfactory theory for calculating the pressure which acts at the blunt base of an object traveling at supersonic velocity. In fact, the essential mechanism determining the base pressure is only imperfectly understood. As a result, the existing knowledge of base pressure is based almost entirely on experiments. The main object of this paper is to summarize the principal results of the many wind tunnel and free flight measurements of base pressure on both bodies of revolution and blunt trailing edge airfoils. A relatively simple method of estimating base pressure is presented, and an indication is given as to how the characteristics of base pressure play an essential role in determining the shape of an aerodynamically efficient object for supersonic flight.

Description: The three papers collected here are: 'The Effect of Nose Shape on the Drag of Bodies of Revolution at Zero Angle of Attack.', 'Base Pressure on Wings and Bodies with Turbulent Boundary Layers', and 'Flow over Inclined Bodies'. The subject of the first paper is the drag of the nose section of bodies of revolution at zero angle of attack. The main object of the second paper is to summarize the prinicpal results of the many wind tunnel and free flight measurements of base pressure on both bodies of revolution and blunt trailing edge airfoils.

Description: Hinge-moment parameters for several typical ailerons were calculated from section data with the aspect-ratio correction as usually determined from lifting-line theory. The calculations showed that the agreement between experimental and calculated results was unsatisfactory. An additional aspect-ratio correction, calculated by the method of lifting-surface theory, was applied to the slope of the curve of hinge-moment coefficient against angle of attack at small angles of attack. This so-called streamline-curvature correction brought the calculated and experimental results into satisfactory agreement.

Description: The possibility of overshooting the anticipated normal acceleration as a result of the artificial-feel characteristics of the F-89C airplane at a condition of minimum static stability was investigated analytically by means of an electronic simulator. Several methods of improving the stick-force characteristics were studied. It is shown that, due to the lag in build-up of the portion of the stick force introduced by the bobweight, it would be possible for excessive overshoots of normal acceleration to occur in abrupt maneuvers with reasonable assumed control movements. The addition of a transient stick force proportional to pitching acceleration (which leads the normal acceleration) to prevent this occurring would not be practical due to the introduction of an oscillatory mode to the stick-position response. A device to introduce a viscous damping force would Improve the stick-force characteristics so that normal acceleration overshoots would not be likely, and the variation of the maximum stick force in rapid pulse-type maneuvers with duration of the maneuver then would have a favorable trend.

Description: The transonic longitudinal aerodynamic characteristics of a 0.0858-scale model of the Lockheed XF-104 airplane have been obtained from tests at the Langley 16-foot transonic tunnel. The results of the investigation provide some general information applicable to the transonic properties of thin, low-aspect-ratio, unswept wing configurations utilizing a high horizontal tail . The model employs a horizontal tail mounted at the top of the vertical tail and a wing with an aspect ratio of 2.5, a taper ratio of 0.385, and 3.4-percent-thick airfoil sections. The lift, drag, and static longitudinal pitching moment were measured at Mach numbers from 0.80 t o 1.09 and angles of attack from -2.5 deg to 22.5 deg. Some of the dynamic longitudinal stability properties of the airplane have been predicted from the test results. In addition, some visual flow studies on the wing surfaces obtained at Mach numbers of 0.80 and 1.00 are included. Results of the investigation show that the transonic rise in drag coefficient at zero lift is about 0.030. At high angles of attack, the model becomes longitudinally unstable at Mach numbers from 0.80 t o 0.90, whereas a reduction in static stability is experienced when very high angles of attack are reached at Mach numbers above 0.90. Longitudinal dynamic stability calculations show that the longitudinal control is good at angles of attack below the unstable break in the static pitching-moment curves, but a typical corrective control applied after the occurrence of neutral stability has little effect in averting pitch-up.

Description: The effects of elevator deflections from 0deg to -20deg on the force and moment characteristics of a 1/20-scale model of the Convair F-102 airplane with chordwise fences have been determined a t Mach numbers from 0.6 to 1.1 for angles of attack up to 20deg in the Langley 8-foot transonic tunnel. The configuration exhibited static longitudinal stability throughout the range tested, although a mild pitch-up tendency was indicated a t Mach numbers from 0.85 to 0.95. Elevator pitch effectiveness decreased rapidly between the Mach numbers of 0.9 and 1.0, however, no complete loss or reversal was indicated for all conditions tested. Because of the type of longitudinal control used, trimming the configuration from the zero elevator condition resulted in substantial decreases in lift-curve slope and maximum lift-drag ratio and increases in drag due to lift. The drag at zero lift, drag due to lift, and trim drag were high for this configuration.

Description: A 1/10-scale rocket model of the Lockheed XF-104 with faired inlets has been flown over a Mach number range from 0.80 to 1.45 to determine low-lift drag and a limited amount of stability data. The center-of-gravity locations were 4.0 and 1.5 percent of the mean aerodynamic chord before and after sustainer firing, respectively. Oscillations induced by pulse rockets were used to determine stability data. The external transonic drag coefficient increased from a value of 0.0160 at Mach number 0.80 to a maximum of 0.0432 near Mach number 1-13, with a drag rise Mach number of about 0.93. At Mach numbers where it could be determined, the model exhibited stable dynamic and static stability characteristics at low lift.

Description: A flight test has been conducted to determine the longitudinal stability and control characteristics of a 0.133-scale model of the Consolidated Vultee XFY-1 airplane with windmilling propellers for the Mach number range between 0.70 and 1.13. The variation of lift-curve slope C(sub L(sub alpha) with Mach number was gradual with a maximum value of 0.074 occurring at a Mach number of 0.97. Propellers had little effect upon the values of lift-curve slope or the linearity of lift coefficient with angle of attack. At lift coefficients between approximately 0.25 and 0.45 with an elevon angle of approximately -l0 deg, there was a region of neutral longitudinal stability at Mach numbers below 0.93 introduced by the addition of windmilling propellers. Below a lift coefficient of 0.10 and above a lift coefficient of 0.45, the model was longitudinally stable throughout the Mach number range of the test. There was a forward shift in the aerodynamic center of about 3-percent mean aerodynamic chord introduced by the addition of propellers. The aerodynamic center as determined at low lift moved gradually from a value of 28.5-percent mean aerodynamic chord at a Mach number of 0.75 to a value of 47-percent mean aerodynamic chord at a Mach number of 1.10. There was an abrupt decrease in pitch damping between Mach numbers of 0.88 and 0.99 followed by a rapid increase in damping to a Mach number of 1.06. The propellers had little effect upon the pitch damping characteristics . The transonic trim change was a large pitching-down tendency with and without windmilling propellers. The elevons were effective pitch controls throughout the speed range; however, their effectiveness was reduced about 50 percent at supersonic speeds. The propellers had no appreciable effect upon the control effectiveness.

Description: The aileron characteristics of a Grumman F4F-3 airplane were determined in flight by means of NACA recording and indicating instruments. The results show that the ailerons met NACA minimum requirements for satisfactory control throughout a limited speed range. A helix angle of approximately 0.07 radian was produced with flaps down at speeds from 90 to 115 miles per hour indicated airspeed and with flaps up from 115 to 200 miles per hour. With flaps up at 90 miles per hour, the helix angle dropped to 0.055 radian; above 200 miles per hour heavy aileron stick forces seriously restricted maneuverability in roll.