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NASA NTRS1957

Investigation at High Subsonic Speeds of the Static Longitudinal and Lateral Stability Characteristics of Two Canard Airplane Configurations

The present investigation was conducted in the Langley high-speed 7-by 10-foot tunnel to determine the static longitudinal and lateral stability characteristics at high subsonic speeds of two canard airplane configurations previously tested at supersonic speeds. The Mach number range of this investigation extended from 0.60 to 0.94 and a maximum angle-of-attack range of -2dewg to 24deg was obtained at the lowest test Mach number. Two wing plan forms of equal area were studied in the present tests; one was a 60deg delta wing and the other was a trapezoid wing having an aspect ratio of 3, taper ratio of 0.143, and an unswept 80-percent-chord line. The canard control had a trapezoidal plan form and its area was approximately 11.5 percent of the wing area. The model also had a low-aspect-ratio highly swept vertical tail and twin ventral fins. The longitudinal control characteristics of the models were consistent with past experience at low speed on canard configurations in that stalling of the canard surface occurred at moderate and high control deflections for moderate values of angle of attack. This stalling could impose appreciable limitations on the maximum trim-lift coefficient attainable. The control effectiveness and maximum value of trim-lift was significantly increased by addition of a body flap having a conical shape and located slightly behind the canard surface on the bottom of the body. Addition of the canard surface at 0deg deflection had relatively little effect on overall directional stability of the delta-wing configuration; however, deflection of the canard surface from 0deg to 10deg had a large favorable effect on directional stability at high angles of attack for both the trapezoid- and delta-wing configurations.

Sleeman, William C., Jr.↗
NASA NTRS1953

Lateral Oscillatory Characteristics of the Republic F-91 Airplane Calculated by Using Low-Speed Experimental Static and Rotary Derivatives

The present investigation was conducted to determine, from low-speed tests in the Langley stability tunnel, the static and rotary derivatives of a 1/9-scale model of the Republic F-91 airplane and various of its components (including the effects of wing incidence) and to determine the accuracy with which the period and damping of the lateral oscillation of the airplane could be calculated by using these experimentally between flight and calculated period and damping of the lateral oscillation were made for Mach numbers from 0.4 to 0.9 at an altitude of 20,OOO feet for 0deg wing incidence and several other wing incidences. Some comparisons were made of the static and rotary derivatives of the model and derivatives estimated by available procedures. determined derivatives (corrected for Mach number effects). Comparisons The results of the investigation have indicated that the model did not have unusual aerodynamic characteristics except for a large (about -0.125) increment in the damping in yaw contributed by the fuselage. Changes in wing incidence, in general, had little effect on the static and rotary derivatives of the model. The static and rotary derivatives of the model could be estimated with good accuracy only in the low angle-of-attack range by using available procedures.

Jaquet, Byron M.↗
NASA NTRS1948

Investigation of the Flying Mock-Up of Consolidated Vultee XP-92 Airplane in the Ames 40- by 80-Foot Wind Tunnel: Pressure Distributions

This report contains the results of the wind tunnel investigation of the pressure distribution on the flying mock-up of the Consolidated Vultee XP-92 airplane. Data are presented for the pressure distribution over the wing, vertical tail and the fuselage, and for the pressure loss and rate of flow through the ducted fuselage. Data are also presented for the calibration of two airspeed indicators, and for the calibration of angle-of-attack and sideslip-angle indicator vanes.

Graham, David↗
NASA NTRS1947

Longitudinal Stability and Control Characteristics of a Semispan Model of the XF7U-1 Tailless Airplane at Transonic Speeds by the NACA Wing-Flow Method, TED No. NACA DE307

An investigation was made by the NACA wing-flow method to determine the longitudinal stability and control characteristics at transonic speeds of a semispan model of the XF7U-1 tailless airplane. The 25-percent chord line of the wing of the model was swept back 35 deg. The airfoil sections of the wing perpendicular to the 25-percent chord line were 12 percent thick. Measurements were made of the normal force and pitching moment through an angle-of-attack range from about -3 deg to 14 deg for several ailavator deflections at Mach numbers from 0.65 to about 1.08. The results of the tests indicated no adverse effects of compressibility up to a Mach number of at least 0.85 at low normal-force coefficients and small ailavator deflections. Up to a Mach number of 0.85, the neutral point at low normal-force coefficients was at about 25 percent of the mean aerodynamic chord and moved rearward irregularly to 41 or 42 percent with a further increase in Mach number to about 1.05. For deflections up to -8.0 percent, the ailavator was effective in changing the pitching moment except at Mach numbers from 0.93 to about 1.0 where ineffectiveness or reversal was indicated for deflections and normal-force coefficients. With -13.2 deg deflection at normal-force coefficients above about 0.3, reversal of ailavator effectiveness occurred at Mach numbers as low as 0.81. A nose-down trim change, which began at a Mach number of about 0.85, together with the loss in effectiveness of the ailavator, indicated that with increase in the Mach number from about 0.95 to 1.05 an abrupt ailavator movement of 5 deg or 6 deg first up and then down would be required to maintain level flight.

Sawyer, Richard H.↗
NASA NTRS1953

Investigation of the Effects of Propeller Operation on the Low-Speed Stability and Control Characteristics of a 1/6-Scale Model of a Revised Configuration of the Republic XF-84H Airplane

An investigation was made to determine the static longitudinal and lateral stability and control characteristics of a l/6-scale model of the revised Republic XF-84H airplane with and without the propeller operating. The model had a 40deg swept wing of aspect ratio 3.45 and was equipped with a thin, three-blade supersonic-type propeller. Modifications incorporated in the revised model included a raised horizontal tail, increased rudder size, wing fences at 65 percent semispan, and a modified wing leading edge outboard of the fences. The test results for flap-retracted and flap-deflected conditions indicated that the revised configuration should be satisfactory for most normal flight conditions provided the angle of attack does not exceed the angle for pitch-up. An abrupt pitch-up tendency of the model was evident for the zero thrust condition above approximately 15' angle of attack. Although the effects of power were destabilizing, power-on longitudinal stability was satisfactory through the angle-of-attack range for which the model was stable with zero thrust. Above the angle of attack for pitch-up, an uncontrollable left roll-off tendency would be expected with power on and slats retracted. Projection of wing slats or use of leading-edge chord-extensions with only the left extension drooped were found beneficial in controlling the roll-off tendency with power on; however the most effective means found was projection of only the left slat.

Sleeman, William C., Jr.↗
NASA NTRS1950

Effect of Various Modifications on Drag and Longitudinal Stability and Control Characteristics at Transonic Speeds of a Model of the XF7U-1 Tailless Airplane: NACA Wing-FLow Method, TED No. NACA DE 307

An investigation was made by the NACA wing-flow method to determine the drag, pitching-moment, lift, and angle-of-attack characteristics at transonic speeds of various configurations of a semispan model of an early configuration of the XF7U-1 tailless airplane. The results of the tests indicated that for the basic configuration with undeflected ailavator, the zero-lift drag rise occurred at a Mach number of about 0.85 and that about a five-fold increase in drag occurred through the transonic speed range. The results of the tests also indicated that the drag increment produced by -8.0 degrees deflection of the ailavator increased with increase in normal-force coefficient and was smaller at speeds above than at speeds below the drag rise. The drag increment produced by 35 degree deflection of the speed brakes varied from 0.040 to 0.074 depending on the normal-force coefficient and Mach number. These values correspond to drag coefficients of about 0.40 and 0.75 based on speed-brake frontal area. Removal of the fin produced a small positive drag increment at a given normal-force coefficient at speeds during the drag rise. A large forward shift of the neutral-point location occurred at Mach numbers above about 0.90 upon removal of the fin, and also a considerable forward shift throughout the Mach number range occurred upon deflection of the speed brakes. Ailavator ineffectiveness or reversal at low deflections, similar to that determined in previous tests of the basic configuration of the model in the Mach number range from about 0.93 to 1.0, was found for the fin-off configuration and for the model equipped with skewed (more highly sweptback) hinge-line ailavators. With the speed brakes deflected, little or no loss in the incremental pitching moment produced by deflection of the ailavator from O degrees to -8.00 degrees occurred in the Mach number range from 0.85 to 1.0 in contrast to a considerable loss found in previous tests with the speed brakes off.

Sawyer, Richard H.↗
NASA NTRS1954

Effects of Fuselage Modifications on the Drag Characteristics of a 1/20-Scale Model of the Convair F-102 Airplane at Transonic Speeds

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.

Kelly, Thomas C.↗
NASA NTRS1955

Investigation in the Langley Free-Flight Tunnel of the Low-Speed Stability and Control Characteristics of a 1/10-Scale Model Simulating the Convair F-102A Airplane

An investigation of the low-speed, power-off stability and control characteristics of a 1/10-scale model simulating the Convair F-102A airplane has been made in the Langley free-flight tunnel. The model in its basic configuration and with two modifications involving leading- edge slats and an increase in vertical-tail size was flown through a lift-coefficient range from 0.7 to the stall. Only relatively low-altitude conditions were simulated. The longitudinal stability characteristics of the model were considered satisfactory for all conditions investigated. The lateral stability characteristics were considered satisfactory for the basic configuration over the lift-coefficient range investigated, except near the stall, where large values of static directional instability caused the model to be directionally divergent. An 80-percent increase in vertical-tail area increased the angle of attack at which the model became directionally divergent. The longitudinal and lateral control characteristics were generally satisfactory. Although the adverse sideslip characteristics for the model were considered acceptable over the angle-of-attack range, analysis indicates that the adverse sideslip characteristics of the airplane may be objectionable at high angles of attack.

Boisseau, Peter C.↗
NASA NTRS1955

Aerodynamic Characteristics of a 0.04956-Scale Model of the Convair F-102A Airplane at Transonic Speeds

Tests have been conducted in the Langley 8-foot transonic tunnel on a 0.04956-scale model of the Convair F-102A airplane which employed an indented and extended fuselage, cambered wing leading edges, and deflected wing tips. Force and moment characteristics were obtained for Mach numbers from 0.60 to 1.135 at angles of attack up to 20 . In addition, tests were made over a limited angle-of-attack range to determine the effects of the cambered leading edges, deflected tips, and a nose section with a smooth area distribution. Fuselage modifications employed on the F-102A were responsible for a 25.percent reduction in the minimum drag-coefficient rise between the Mach numbers of 0.85 and 1.075 when compared with that for the earlier versions of the F-102. Although the wing modifications increased the F-102A subsonic minimum drag-coefficient level approximately 0.0020, they produced large decreases in drag at lifting conditions over that for the original (plane-wing) F-102. The F-102A had 15 to 25 percent higher maximum lift-drag ratios than did the original F-102. The F-102A had about 15 percent lower maximum lift-drag ratios at Mach numbers below 0.95 and slightly higher maximum lift-drag ratios at supersonic speeds when compared with those ratios for sn earlier modified-wing version of the F-102. Chordwise wing fences which provided suitable longitudinal stability for the original F-102 were not adequate for the cambered-wing F-102A The pitching-moment curves indicated a region of near neutral stability with possible pitch-up tendencies for the F-102A at high subsonic Mach numbers for lift coefficients between about 0.4 and 0.5.

Tempelmeyer, Kenneth E.↗
NASA NTRS1953

Wind-Tunnel Investigation at Low Speed of the Pitching Stability Derivatives of a 1/9-Scale Powered Model of the Convair XFY-1 Vertically Rising Airplane, TED No. NACA DE 373

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

Queijo, M. J.↗
NASA NTRS1953

Investigation of the Low-Speed Stability and Control Characteristics of a 1/10-Scale Model of the Convair YF-102 Airplane in the Langley Free-Flight Tunnel

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

Johnson, Joseph L., Jr.↗
NASA NTRS1949

Wind-Tunnel Investigation of the Low-Speed Static Stability and Control Characteristics of a Model of the Bell MX-776 (RASCAL) in Combined Angle of Attack and Sideslip

An investigation has been made in the Langley stability tunnel to determine the low-speed static stability and control characteristics of a model of the Bell MX-776. The results show the model to be longitudinally unstable in the angle-of-attack range around zero angle of attack and to become stable at moderate angles of attack. The results of the present investigation agree reasonably well with results obtained in other facilities at low speed. The present pitching-moment results at low Mach numbers also agree reasonably well with unpublished results of tests of the model at supersonic Mach numbers (up to Mach number 1.86). Unpublished results at moderate and high subsonic speeds, however, indicate considerably greater instability at low angles of attack than is indicated by low-speed results. The results of the present tests also showed that the pitching-moment coefficients for angles of attack up to 12deg remained fairly constant with sideslip angle up to 12deg. The elevators tested produced relatively large pitching moments at zero angle of attack but, as the angle of attack was increased, the elevator effectiveness decreased. The rate of decrease of elevator effectiveness with angle of attack was less for 8deg than for 20deg elevator deflection. Therefore although 8deg deflection caused an appreciable change in longitudinal trim angle and trim lift coefficient a deflection of 20deg caused only a small additional increase in trim angle and trim lift coefficient.

Letko, William↗
NASA NTRS1953

Wind-Tunnel Investigation at Low Speed of the Yawing Stability Derivatives of a 1/9-Scale Powered Model of the Convair XFY-1 Vertically Rising Airplane, TED No. NACA DE 373

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.

Queijo, M. J.↗
NASA NTRS1953

Wind-Tunnel Investigation at Low Speed of the Rolling Stability Derivatives of a 1/9-Scale Powered Model of the Convair XFY-1 Vertically Rising Airplane, TED No. NACA DE 373

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.

Queijo, M. J.↗
NASA NTRS1950

Transonic Stability and Control Investigation of a 1/80-Scale Model of the Consolidated Vultee Skate 9 Seaplane, TED No. NACA DE 345: Transonic-Bump Method

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.

Riebe, John M.↗
NASA NTRS1948

Longitudinal Stability Characteristics of a 1/40-Scale Model of a Proposed Configuration of the XF-91 Airplane Measured by the Wing-Flow Method

This report presents the results of an investigation in the transonic speed range of the longitudinal stability characteristics of a proposed configuration for the Republic XF-91 airplane. The tests covered a Mach number range of 0.55 to 1.05 and a Reynolds number range from 400,000 to 1,375,000. Lift, pitching-moment, and rolling-moment characteristics of the half model and the hinge moments on the all-moving tail were measured. The downwash factor delta x epsilon / delta x alpha at the tail was determined from the pitching-moment data. A calculation of the elevator deflection and stick force required for trim was also made. It was found that the variation of force and moment coefficients was linear through the test angle-of-attack range of -1 deg to 8 deg at any Mach number; that the stability increased markedly at Mach numbers above 0.85; that the effectiveness of the tail in producing pitching moments decreased about one-third with increasing Mach numbers and that the value of the downwash factor, delta x epsilon / delta x alpha, at the tail decreased from about 0.35 at a Mach number of 0.85 to about zero at a Mach number near 0.95 and became slightly negative at higher Mach numbers. The calculated values of stick force per g and elevator deflection per g, assuming no aerodynamic balance, increased rapidly above a Mach number of 0.85.

Crane, Harold L.↗
NASA NTRS1949

Wind-Tunnel Investigation of the Low-Speed Static Stability and Control Characteristics of a Model of Bell MX-776

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

Queijo, M. J.↗
NASA NTRS1952

An Investigation of the Effects of a Vortex-Generator Configuration on the Aerodynamic Characteristics of a 1/4-Scale Model of the X-1 Airplane (10-Percent-Thick Wing)

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

Runckel, Jack F.↗