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Three dimensional inviscid supersonic flow fields with primary and embedded shock and expansion waves determined over and behind wings and wing-body configurations
Comparisons of hinge moments for simple delta wing and delta wing orbiter concept at Mach 6
Experimental aerodynamic investigations have been made on a .0035 scale model North American Rockwell/General Dynamics version of the space shuttle. Static stability and control data were obtained on the delta wing booster alone (B-20) and with the delta wing orbiter (134D) mounted in various positions on the booster. Six component aerodynamic force and moment data were recorded over an angle of attack range from -10 deg to 24 deg at 0 deg and 6 deg sideslip angles and from -10 deg to +10 deg sideslip at 0 deg angle of attack. Mach number ranged from 0.6 to 4.96.
An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the static aerodynamic characteristics of a two-staged space-shuttle system consisting of a delta-wing orbiter mated atop a winged booster. The tests were performed at Mach numbers from 2.30 to 4.60 at a Reynolds number of one million per foot.
Experimental aerodynamic investigations have been made on a .0035 scale model North American Rockwell/General Dynamics version of the space shuttle in the NASA/MSFC 14 x 14 Inch Trisonic Wind Tunnel. Static stability and control data were obtained on the delta wing booster alone (B-20) and with the delta wing orbiter (134D) mounted in various positions on the booster. Six component aerodynamic force and moment data were recorded over an angle of attack range from -10 to 24 deg at 0 and 6 deg sideslip angles and from -10 to +10 deg sideslip at 0 deg angle of attack. Mach number ranged from 0.6 to 4.96.
Wind tunnel tests to determine the pressure distribution on a space shuttle launch configuration consisting of a delta wing orbiter and a swept wing booster with canard and tip fins were conducted. Pressure data were obtained for the combined orbiter and booster and for the booster alone at Mach numbers from 0.6 to 1.3, angles of attack from minus 8 degrees to plus 10 degrees, and sideslip angles from minus 6 degrees to plus 6 degrees. Pressure data were also obtained for the booster alone without canard at Mach numbers of 0.9 and 1.1. The pressure taps were distributed primarily over the booster upper surface and the orbiter lower surface.
The low-speed flying qualities of an experimental STOL configuration were evaluated by using a fixed-base six-degree-of-freedom simulation. The configuration had either an externally blown flap (EBF) wing or an augmentor wing (AW). The AW configuration was investigated with two tails, one sized for the AW configuration and a larger one sized for the EBF configuration. The emphasis of the study was on the 70-knot approach task. The stability and control characteristics were compared with existing criteria. Several control systems were investigated for the normal four-engine condition and for the engine-out transient condition. Minimum control and stall speeds were determined for both the three- and four-engine operation.
The Langley 1/8-scale V/STOL model tunnel was used to qualitatively determine the flow fields associated with semi-span close coupled canard wing and strake wing models. Small helium filled bubbles were injected upstream of the models to make the flow visible. Photographs were taken over the angle-of-attack ranges of -10 deg to 40 deg.
An experimental and analytical aerodynamic program to develop predesign guides for irregular planform wings is reported. The benefits are linearization of subsonic lift curve slope to high angles of attack and avoidance of subsonic pitch instabilities at high lift by proper tailoring of the planform fillet wing combination while providing the desired hypersonic trim angle and stability. The two prime areas of concern are to optimize shuttle orbiter landing and entry characteristics. Basic longitudinal aerodynamic characteristics at high supersonic speeds are developed.
Tests have been conducted in a full scale tunnel to determine the low speed aerodynamic characteristics of a large scale advanced arrow wing supersonic transport configuration with engines mounted above the wing for upper surface blowing. Tests were made over an angle of attack range of -10 deg to 32 deg, sideslip angles of + or - 5 deg, and a Reynolds number range of 3,530,000 to 7,330,000. Configuration variables included trailing edge flap deflection, engine jet nozzle angle, engine thrust coefficient, engine out operation, and asymmetrical trailing edge boundary layer control for providing roll trim. Downwash measurements at the tail were obtained for different thrust coefficients, tail heights, and at two fuselage stations.
Wing leading-edge deflection effects on the low-speed aerodynamic characteristics of a low-aspect-ratio highly swept arrow-wing configuration were determined. Static force tests were conducted in a V/STOL tunnel at a Reynolds number of about 2.5 x 1 million for an angle-of-attack range from -10 deg to 17 deg and an angle-of-sideslip range from -5 deg to 5 deg. Limited flow visualization studies were also conducted in order to provide a qualitative assessment of leading-edge upwash characteristics.
Results are presented of the theory of wings and of wing sections which are of immediate practical value. They are proven and demonstrated by the use of the simple conceptions of kinetic energy and momentum only.
Wind tunnel and flight tests were conducted to determine the effects of several discontinuous drooped wing leading-edge configurations on the spinning characteristics of a light, single-engine, low-wing research airplane. Particular emphasis was placed on the identification of modifications which would improve the spinning characteristics. The spanwise length of a discontinuous outboard droop was varied and several additional inboard segments were added to determine the influence of such leading-edge configurations on the spin behavior. Results of the study indicated that the use of only the discontinuous outboard droop, over a specific spanwise area, was most effective towards improving spin and spin recovery characteristics, whereas the segmented configurations having both inboard and outboard droop exhibited a tendency to enter a flat spin.
An analytical approach to the study of flight dynamics of aircraft operating in a high-angle-of-attack flight regime and of helicopters operating in extreme thrust conditions is presented. Steady coordinated high-g turns are used to establish the initial equilibrium flight conditions near stall angles of attack. The kinematic properties of the aircraft in steady coordinated turns are examined: in high-g turns, pitch rate (independent of the angle of attack) is of a much larger magnitude than roll and yaw rate; a substantial roll rate is found to develop in steep turns for all angles of attack; the angle of attack also has a significant effect on the pitch attitude, with decreasing influence as the normal load factor increases. The exact small disturbance equations of motion of the aircraft in general steady turns are also developed for application to both rotary-wing and fixed-wing aircraft in extreme conditions. These equations are in a first-order, vector-matrix format, and are thus compatible with many efficient software packages developed in modern system theory.
The methods used to compute the mass, structural stiffness, and aerodynamic forces in the form of influence coefficient matrices as applied to a flutter analysis of the Drones for Aerodynamic and Structural Testing (DAST) Aeroelastic Research Wing. The DAST wing was chosen because wind tunnel flutter test data and zero speed vibration data of the modes and frequencies exist and are available for comparison. A derivation of the equations of motion that can be used to apply the modal method for flutter suppression is included. A comparison of the open loop flutter predictions with both wind tunnel data and other analytical methods is presented.
An investigation was made to assess the effect of wing-mounted nacelles on a supercritical wing configuration at near-sonic Mach numbers. The investigation was made by utilizing the Cornell Aeronautical Laboratory 8-foot transonic tunnel and covered a Mach number range from 0.90 to 0.99. Force data and pressure measurements at selected locations were obtained. The investigation with the nacelles on included the effect of spanwise location of the nacelle (semispan locations of 35 and 70 percent) and the effect of area ruling for the nacelles located at the 35 percent semispan station. Tests were also made with the outboard nacelle extended forward so that it was directly adjacent to the inboard nacelle location. These tests provided a direct assessment of the extent of the nacelle interference flow field in a lateral direction.
The NASTRAN model plan for the wing structure was expanded in detail to generate the NASTRAN model for this substructure. The grid point coordinates were coded for each element. The material properties and sizing data for each element were specified. The wing substructure model was thoroughly checked out for continuity, connectivity, and constraints. This substructure was processed for structural influence coefficients (SIC) point loadings and the deflections were compared to those computed for the aircraft detail model. Finally, a demonstration and validation processing of this substructure was accomplished using the NASTRAN finite element program. The bulk data deck, stiffness matrices, and SIC output data were delivered.
An advanced fighter configuration with a forward-swept wing of aspect ratio 3.28 is tested in the Langley 7 by 10 Foot High Speed Tunnel at a Mach number of 0.3. The wing has 29.5 degrees of forward sweep of the quarter chord line and is equipped with 15 percent chord leading edge and 30 percent chord trailing edge flaps. The canard is sweptback 45 degrees. Tests were made through a range of angle of attack from about -2 degrees to 22 degrees. Deflecting the flaps significantly improves the lift drag characteristics at the higher angles of attack. The canard is able to trim the configurations with different flap deflections over most of the range of angle of attack. The penalty in maximum lift coefficient due to trimming is about 0.10.