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At least 1,441 records · Page 80
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Effects of nozzle interfairing modifications on longitudinal aerodynamic characteristics of a twin jet, variable wing sweep fighter model

A wind-tunnel investigation has been made to determine the effects of nozzle interfairing modifications on the longitudinal aerodynamic characteristics of a twin-jet, variable-wing-sweep fighter model. The model was tested in the Langley 16-foot transonic tunnel at Mach numbers of 0.6 to 1.3 and angles of attack from about minus 2 deg to 6 deg and in the Langley 4-foot supersonic presure tunnel at a Mach number of 2.2 and an angle of attack of 0 deg. Compressed air was used to simulate nozzle exhaust flow at jet total-pressure ratios from 1 (jet off) to about 21. The results of this investigation show that the aircraft drag can be significantly reduced by replacing the basic interfairing with a modified interfairing.

Reubush, D. E.↗
NASA NTRS1980

A status report on NASA general aviation stall/spin flight testing

The NASA Langley Research Center has undertaken a comprehensive program involving spin tunnel, static and rotary balance wind tunnel, full-scale wind tunnel, free flight radio control model, flight simulation, and full-scale testing. Work underway includes aerodynamic definition of various configurations at high angles of attack, testing of stall and spin prevention concepts, definition of spin and spin recovery characteristics, and development of test techniques and emergency spin recovery systems. This paper presents some interesting results to date for the first aircraft (low-wing, single-engine) in the program, in the areas of tail design, wing leading edge design, mass distribution, center of gravity location, and small airframe changes, with associated pilot observations. The design philosophy of the spin recovery parachute system is discussed in addition to test techniques.

Patton, J. M., Jr.↗
NASA NTRS1982

Effects of installation of F101 DFE exhaust nozzles on the afterbody-nozzle characteristics of the F-14 airplane

A wind-tunnel investigation was conducted to determine the effects of F101 DFE (derivative fighter engine) nozzle axial positioning on the afterbody-nozzle longitudinal aerodynamic characteristics of the F-14 airplane. The model was tested in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0.7 to 1.25 and angles of attack from about -2 to 6 degrees. Compressed air was used to simulate nozzle exhaust flow at jet total-pressure ratios from 1 (jet off) to about 8. The results of the investigation show that for subsonic Mach numbers the intermediate cruise nozzle position of the three positions tested resulted in the lowest drag.

Reubush, D. E.↗
NASA NTRS1981

Comparison of theory and in situ observations for electron and ion distributions in the near wake of the Explorer 31 and AE-C satellites

Measurements of electron density, plasma potential, and mean ion mass from the Explorer 31 satellite, and measurements of ion current, plasma potential, and ion composition from the Atmosphere Explorer C satellite were used in a comparative study with Parker's theory regarding the charged particle distribution in the near wake of an ionospheric satellite (1976). It is shown that theory and experiment agree fairly well in the angle-of-attack range between 90 and 135 deg. In the maximum rarefaction zone (between 145 and 180 deg), however, the theoretical model overestimates the measured ion depletion by several orders of magnitude. A comparison between theory and the Explorer 31 electron measurements shows that the theory again overestimates the electron depletion. These discrepancies are mainly due to the use of a steady-state theory and a single ion equation (using a mean ion mass). Improved agreement between theory and experiment can be obtained by the use of the time-dependent Vlasov-Poisson equations with separate equations for the various ion species.

Samir, U.↗
NASA NTRS1981

Estimation of airplane stability and control derivatives at high angles of attack

A technique for extracting approximate aircraft stability and control derivatives over a broad range of angle of attack from a single large amplitude maneuver which encompasses that range is presented. The results from the application of this technique to flight data are shown to agree well with derivatives extracted from several maneuvers which are more restricted in range of angle of attack. The technique consists of dividing the data from the large maneuver into bins that are a function of angle of attack. The binned data is then analyzed with a modified stepwise regression. The identification of nonlinear terms in the aerodynamic model is shown to discern regions of rapid change in aircraft behavior.

Batterson, J. G.↗
NASA NTRS1982

Numerical simulation of steady supersonic flow over spinning bodies of revolution

A recently reported parabolized Navier-Stokes code has been employed to compute the supersonic flowfield about a spinning cone and spinning and nonspinning ogive cylinder and boattailed bodies of revolution at moderate incidence. The computations were performed for flow conditions where extensive measurements for wall pressure, boundary-layer velocity profiles, and Magnus force had been obtained. Comparisons between the computational results and experiment indicate excellent agreement for angles of attack up to 6 deg. At angles greater than 6 deg discrepancies are noted which are tentatively attributed to turbulence modeling errors. The comparisons for Magnus effects show that the code accurately predicts the effects of body shape for the selected models.

Sturek, W. B.↗
NASA NTRS1984

An experimental investigation of nacelle-pylon installation on an unswept wing at subsonic and transonic speeds

A wind tunnel investigation was conducted to determine the aerodynamic interference associated with the installation of a long duct, flow-through nacelle on a straight unswept untapered supercritical wing. Experimental data was obtained for the verification of computational prediction techniques. The model was tested in the 16-Foot Transonic Tunnel at Mach numbers from 0.20 to 0.875 and at angles of attack from about 0 deg to 5 deg. The results of the investigation show that strong viscous and compressibility effects are present at the transonic Mach numbers. Numerical comparisons show that linear theory is adequate for subsonic Mach number flow prediction, but is inadequate for prediction of the extreme flow conditions that exist at the transonic Mach numbers.

Carlson, J. R.↗
NASA NTRS1984

Effect of airfoil mean loading on convected gust interaction noise

The present theoretical model for noise generation when a convected vortical or entropic gust encounters an airfoil at a nonzero angle-of-attack is based on a linearization of the Euler equations for the steady, subsonic flow past the airfoil. Noise generation is found to be concentrated at local regions which scale on the gust wavelength and are present at airfoil leading and trailing edges. Steady loading of the airfoil affects noise generation at the leading and trailing edges in significantly different ways. Parametric calculations are presented which illustrate that, at high frequencies, moderate levels of airfoil steady loading can drastically increase the noise level produced by airfoil-convected gust interactions.

Myers, M. R.↗
NASA NTRS1985

Boundary layer effects on the flowfield about flow-through nacelles

A numerical study has been made of boundary layer effects on the flowfield about flow-through nacelles. A viscous-inviscid interacting computational model for investigating the problem was constructed by coupling a three-dimensional explicit Euler solution procedure with a compressible, 'lag-entrainment' integral boundary layer solution technique. Solutions obtained with the interacting model for long-duct turbofan engine nacelles at a free stream Mach number of 0.8 and an angle of attack of 0 deg are compared with experimental data. The numerical experiments have revealed insight into the boundary layer effects on the interactions between the internal and external flows.

Compton, W. B.↗
NASA NTRS1986

Unsteady low-speed aerodynamic model for complete aircraft configurations

A method for converting steady state, potential panel methods into a time-dependent mode is derived and applied to several test cases. For this development, an improved vortex wake model was constructed which was also suitable for simulating the leading edge separation of slender wings at high angles of attack. Computed flow-field simulations are presented for various unsteady, and high-angle-of-attack conditions, involving geometries such as simple wings, rotors, and complete aircraft configurations.

Katz, J.↗
NASA NTRS1987

Integration effects of pylon geometry and rearward mounted nacelles for a high-wing transport

Results of a wind-tunnel study of the effect of pylon cross-sectional shape and tow angle on airplane drag and an aft-mounted nacelle are presented. The 1/24-scale wide-body high-wing transport model was tested in the Langley 16-Foot Transonic Tunnel at free-stream Mach 0.7-0.8 and angles of attack from -3 to 4 degrees. A compression-type pylon is found to have the lowest drag at both Mach 0.7 and 0.8 and to be capable of suppressing the velocities in the inboard region of the pylon-wing junction, reducing the extent of supersonic flow and the probability of flow separation. It is also shown that the D-shaped aft-mounted nacelle has a low interference drag, as do previously tested circular nacelles in the same position.

Carlson, John R.↗
NASA NTRS1988

Effects of surface perturbations on the asymmetric vortex flow over a slender body

An experimental investigation of the effects of surface perturbations on the asymmetric flow past a slender body has been conducted for laminar flow conditions. Beads with diameters ranging from 3/32 to 12/32 in. were attached near the apex of a cone/cylinder model having a base diameter of 3.5 in. and a cone semiapex angle of 9 deg at an angle of attack of 40 deg in an attempt to alter the sense of the asymmetric vortex flow pattern. Circumferential position as well as longitudinal location were varied to determine the most effective bead position. Whether or not the beads were effective in controlling the magnitude and direction of the vortex asymmetries was determined by 3 circumferential rows of pressure taps and by a helium-bubble flow visualization technique. The most effective circumferential position was found to be approximately 140 deg from the windward ray. While holding this circumferential position constant, the effect of bead size at three stations further along the body was also investigated. It was found that the size of the bead necessary to reverse the asymmetry increased more rapidly than the growth in cylinder radius. In general, these results indicate that discrete geometric imperfections on a body's surface can force asymmetry in a given direction if they are sufficiently large relative to the local radius.

Moskovitz, Cary A.↗
NASA NTRS1988

An analytical method for the ditching analysis of an airborne vehicle

A simple analytical method has been introduced for aerohydrodynamic load analysis of an airborne configuration during water ditching. The method employs an aerodynamic panel code, based on linear potential flow theory, to simulate the flow of air and water around an aircraft configuration. The free surface separating the air and water region is represented by doublet sheet singularities. Although all the theoretical load distributions are computed for air, provisions are made to correct the pressure coefficients obtained on the configuration wetted surfaces to account for the water density. As an analytical tool, the Vortex Separation Aerodynamic (VSAERO) code is chosen to carry out the present investigation. After assessing the validity of the method, its first application is to analyze the water ditching of the Space Shuttle configuration at a 12 degree attitude.

Pressure coefficient↗
NASA NTRS1988

Application of dynamical systems theory to nonlinear aircraft dynamics

Dynamical systems theory has been used to study nonlinear aircraft dynamics. A six degree of freedom model that neglects gravity has been analyzed. The aerodynamic model, supplied by NASA, is for a generic swept wing fighter and includes nonlinearities as functions of the angle of attack. A continuation method was used to calculate the steady states of the aircraft, and bifurcations of these steady states, as functions of the control deflections. Bifurcations were used to predict jump phenomena and the onset of periodic motion for roll coupling instabilities and high angle of attack maneuvers. The predictions were verified with numerical simulations.

Culick, Fred E. C.↗
NASA NTRS1990

Analytical method for the ditching analysis of an airborne vehicle

A simple analytical method has been introduced for aerohydrodynamic load analysis of an airborne configuration during water ditching. The method employs an aerodynamic panel code, based on linear potential flow theory, to simulate the flow of air and water around an aircraft configuration. The free surface separating the air and water region is represented by doublet" singularities. Although all the theoretical load distributions are computed for air, provisions are made to correct the pressure coefficients obtained on the configuration-wetted surfaces to account for the water density. As an analytical tool, a vortex aerodynamic code is chosen to carry out the present investigation. After assessing· the validity of the method, an application is presented for the water ditching of the Space Shuttle configuration at a 12-deg attitude.

Pressure coefficient↗
NASA NTRS1993

Forebody vortex control with jet and slot blowing on an F/A-18

A wind tunnel test program was conducted in October, 1992, on a 6-percent F/A-18 model, to determine the most effective methods of forebody vortex control for providing increased yaw control at high angles of attack. A series of circular jet blowing configurations were investigated with variations in blowing rate, fuselage station, circumferential position, and blowing angle. Slot configurations were also tested and included variations in blowing rate, forward fuselage station, and length. The optimum jet and slot configurations were then also tested at -10 deg of sideslip. Yawing moments from both the jet and slot blowing configurations were found to have very good trends with increasing blowing rates and the maximum yaw control available at 50-deg angle of attack (where rudder power had dropped to near zero) was found to be 40 percent greater than the level provided by the rudder at 0-deg angle of attack.

Kramer, Brian R.↗
NASA NTRS1947

Low-Speed Wind-Tunnel Tests of a Pilotless Aircraft Having Horizontal and Vertical Wings and Cruciform Tail

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

Mastrocola, N↗