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An inviscid incompressible model for the interaction of a wing with a single propeller slipstream is presented. The model allows the perturbation quantities to be potential even though the undisturbed flow is rotational. The governing equations for the spanwise lift distribution are derived and a simple method of solving these is indicated. Spanwise lift and induced drag distribution for two cases are computed.
A direct assessment of two-dimensional wind-tunnel wall interference using upwash component measured on two interfaces has been formulated by the Prandtl-Glauert equation of the flow field and solved by the Fourier transform technique. The analytic formulae obtained for the interference of upwash and pressure on the model are presented. The formulae have been applied successfully to the analytic models for lifting and blockage interferences induced by the general linear slotted and perforated tunnel wall boundary conditions. The formulae have been derived in terms of Fourier coefficients of the measured upwash and illustrated its merits applying to a wavy-wall model case.
Part 1 gives details of models tested and methods of testing of the Eiffel 36 wing alone and the JN2 aircraft. Characteristics and performance curves for standard JN are included. Part 2 presents a statistical analysis of the following: lift and drag contributed by body and chassis tested without wings; lift and drag contributed by tail, tested without wings; the effect on lift and drift of interference between the wings of a biplane combination; lift and drag contributed by the addition of body, chassis, and tail to a biplane combination; total parasite resistance; effect of varying size of tail, keeping angle of setting constant; effect of varying length of body and size of tail at the same time, keeping constant moment of tail surface about the center of gravity; forces on the tail and the effects of downwash; effect of size and setting of tail on statical longitudinal stability effects of length of body on stability; the effects of the various elements of an airplane on longitudinal stability and the placing of the force vectors. Part 3 presents the fundamental principals of dynamical stability; computations of resistance derivatives; solution of the stability equation; dynamical stability of the Curtiss JN2; tabulation of resistance derivatives; discussion of the resistance derivatives; formation and solution of stability equations; physical conceptions of the resistance derivatives; elements contributing to damping and an investigation of low speed conditions. Part 4 includes a summary of the results of the statistical investigation and a summary of the results for dynamic stability.
Aerodynamic equations for the longitudinal motion of an aircraft with a horizontal tail were developed. In this development emphasis was given on obtaining model structure suitable for model identification from experimental data. The resulting aerodynamic models included unsteady effects in the form of linear indicial functions. These functions represented responses in the lift on the wing and tail alone, and interference between those two lifting surfaces. The effect of the wing on the tail was formulated for two different expressions concerning the downwash angle at the tail. The first expression used the Cowley-Glauert approximation known-as "lag-in-downwash," the second took into account growth of the wing circulation and delay in the development of the lift on the tail. Both approaches were demonstrated in two examples using the geometry of a fighter aircraft and a large transport. It was shown that the differences in the two downwash formulations would increase for an aircraft with long tail arm performing low-speed, rapid maneuvers.
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 center dot 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 cross flow drag coefficient along the body and the effect of the wing downwash field on the after body loading.
A method is proposed to combine a numerical description of a jet in a crossflow with a lifting surface panel code to calculate the jet/aerodynamic-surface interference effects on a V/STOL aircraft. An iterative technique is suggested that starts with a model for the properties of a jet/flat plate configuration and modifies these properties based on the flow field calculated for the configuration of interest. The method would estimate the pressures, forces, and moments on an aircraft out of ground effect. A first-order approximation to the method suggested is developed and applied to two simple configurations. The first-order approximation is a noniterative precedure which does not allow for interactions between multiple jets in a crossflow and also does not account for the influence of lifting surfaces on the jet properties. The jet/flat plate model utilized in the examples presented is restricted to a uniform round jet injected perpendicularly into a uniform crossflow for a range of jet-to-crossflow velocity ratios from three to ten.
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.
Progress in the development of concepts for control of aircraft without moving parts or a separate source of power is described. The design and wind tunnel tests of a full scale fluidic rudder for a Cessna 172K aircraft, intended for subsequent flight tests were documented. The 172K fluidic rudder was designed to provide a control force equivalent to 3.3 degrees of deflection of the conventional rudder. In spite of an extremely thin airfoil, cascaded fluidic amplifiers were built to fit, with the capacity for generating the required level of control force. Wind tunnel tests demonstrated that the principles of lift control using ram air power are sound and reliable under all flight conditions. The tests also demonstrated that the performance of the 172K fluidic rudder is not acceptable for flight tests until the design of the scoop is modified to prevent interference with the lift control phenomenon.
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Available information on the effects of wing-fuselage-tail and wing-nacelle interference on the distribution of the air load among components of airplanes is analyzed. The effects of wing and nacelle incidence, horizontal andvertical position of wing and nacelle, fuselage shape, wing section and filleting are considered. Where sufficient data were unavailable to determine the distribution of the air load, the change in lift caused by interference between wing and fuselage was found. This increment is affected to the greatest extent by vertical wing position.
Support interference free drag, lift, and pitching moment measurements on a range of slanted base ogive cylinders were made using the NASA Langley 13 inch magnetic suspension and balance system. Typical test Mach numbers were in the range 0.04 to 0.2. Drag results are shown to be in broad agreement with previous tests with this configuration. Measurements were repeated with a dummy sting support installed in the wind tunnel. Significant support interferences were found at all test conditions and are quantified. Further comparison is made between interference free base pressures, obtained using remote telemetry, and sting cavity pressures.
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.
Results of analytical calculations and wind tunnel tests at cruise speeds of a representative four engine short haul aircraft employing upper surface blowing (USB) with a supercritical wing are discussed. Wind tunnel tests covered a range of Mach number M from 0.6 to 0.78. Tests explored the use of three USB nozzle configurations. Results are shown for the isolated wing body and for each of the three nozzle types installed. Experimental results indicate that a low angle nacelle and streamline contoured nacelle yielded the same interference drag at the design Mach number. A high angle powered lift nacelle had higher interference drag primarily because of nacelle boattail low pressures and flow separation. Results of varying the spacing between the nacelles and the use of trailing edge flap deflections, wing upper surface contouring, and a convergent-divergent nozzle to reduce potential adverse jet effects were also discussed. Analytical comparisons with experimental data, made for selected cases, indicate favorable agreement.
Wave drag, drag due to lift and aerodynamic interference effects on efficiency of supersonic cruise vehicle
An attempt was made to determine the effect of spindle interference on the lift of the airfoil by measuring moments about the axis parallel to the direction of air flow. The values obtained are of the same degree as the experimental error, and for the present this effect will be neglected. The results obtained using a U.S.A. 15 wing (plotted here) show that the correction is nearly constant from 0 degrees to 10 degrees incidence and that at greater angles its value becomes erratic. At such angles, however, the wing drag is so high that the spindle correction and its attendant errors become relatively small and unimportant.
A solution has been obtained for the complete tunnel-interference flow for a lifting vortex in a two-dimensional slotted tunnel. Curves are presented for the longitudinal distribution of tunnel-induced downwash angle for various values of the boundary openness parameter and for various heights of the vortex above the tunnel center line. Some quantitative discussion is given of the use of these results in calculating the tunnel interference for three-dimensional wings in rectangular tunnels with closed side walls and slotted top and bottom.
A systematic study has been made, experimentally and theoretically, of the effects of a vortical wake on the aerodynamic characteristics of a rectangular wing at subsonic speed. The vortex generator and wing were mounted on a reflection plane to avoid body-wing interference. Vortex position, relative to the wing, was varied both in the spanwise direction and normal to the wing. Angle of attack of the wing was varied from -40 to +60. Both chordwise and spanwise pressure distributions were obtained with the wing in uniform and vortical flow fields. Stream surveys were made to determine the flow characteristics in the vortical wake. The vortex-induced lift was calculated by several theoretical methods including strip theory, reverse-flow theory, and reverse-flow theory including a finite vortex core. In addition, the Prandtl lifting-line theory and the Weissinger theory were used to calculate the spanwise distribution of vortex-induced loads. With reverse-flow theory, predictions of the interference lift were generally good, and with Weissinger's theory the agreement between the theoretical spanwise variation of induced load and the experimental variation was good. Results of the stream survey show that the vortex generated by a lifting surface of rectangular plan form tends to trail back streamwise from the tip and does not approach the theoretical location, or centroid of circulation, given by theory. This discrepancy introduced errors in the prediction of vortex interference, especially when the vortex core passed immediately outboard of the wing tip. The wake produced by the vortex generator in these tests was not fully rolled up into a circular vortex, and so lacked symmetry in the vertical direction of the transverse plane. It was found that the direction of circulation affected the induced loads on the wing either when the wing was at angle of attack or when the vortex was some distance away from the plane of the wing.
Supersonic store interference - 1, swept-wing heavy bomber with large stores - lift & drag at mach 1.61