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At least 1,531 records · Page 85
All sourcesNASA NTRSDOE OSTI
NASA NTRS1978

Flight-determined stability and control derivatives for the F-111 Tact research aircraft

A flight investigation was conducted to provide a stability and control derivative data base for the F-111 transonic aircraft technology research aircraft. Longitudinal and lateral-directional data were obtained as functions of Mach number, angle of attack, and wing sweep. For selected derivatives, the flight results were correlated with derivatives calculated based on vehicle geometry. The validity of the angle of attack measurement was independently verified at a Mach number of 0.70 for angles of attack between 3 and 10 degrees.

Sim, A. G.↗
NASA NTRS1994

F/A-18 forebody vortex control. Volume 1: Static tests

A wind tunnel test was conducted on a six percent model of the F/A-18 at the NASA Ames 7 X 10-Foot Low Speed Wind Tunnel. The primary objective of the test was to evaluate several forebody vortex control configurations at high angles of attack in order to determine the most effective method of obtaining well behaved yawing moments, in preparation for the rotary balance test. Both mechanical and pneumatic systems were tested. Single and dual rotating nose tip strakes and a vertical nose strake were tested at different sizes and deflections. A series of jet blowing configurations were located at various fuselage stations, azimuth angles, and pointing angles ranging from straight aft to 60 deg canted inboard. Slot blowing was investigated for several slot lengths and fuselage stations. The effect of blowing rate was tested for both of these pneumatic systems. The most effective configurations were then further tested with a variation of both sideslip angle and Reynolds number over a range of angles of attack from 0 to 60 deg. It was found that a very robust system can be developed that provides yawing moments at angles of attack up to 60 deg that significantly exceeds that available from 30 deg of rudder deflection (F/A-18 maximum) at 0 deg angle of attack.

Kramer, Brian R.↗
NASA NTRS1984

Heat transfer and pressure drop in blade cooling channels with turbulence promoters

Repeated rib roughness elements have been used in advanced turbine cooling designs to enhance the internal heat transfer. Often the ribs are perpendicular to the main flow direction so that they have an angle-of-attack of 90 deg. The objective of the project was to investigate the effect of rib angle-of-attack on the pressure drop and the average heat transfer coefficients in a square duct with two opposite rib-roughned walls for Reynolds number varied from 8000 to 80,000. The rib height-to-equivalent diameter ratio (e/D) was kept at a constant value of 0.063, the rib pitch-to-height ratio (P/e) was varied from 10 to 20, and the rib angle-of-attack (alpha) was varied from 90 deg to 60 deg to 45 deg to 30 deg respectively. Two types of entrance conditions were examined, namely, long duct and sudden contraction. The heat transfer coefficient distribution on the smooth side wall and the rough side wall at the entrance and the fully developed regions were measured. Thermal performance comparison indicated that the pumping power requirement for the rib with an oblique angle to the flow (alpha = 45 deg to 30 deg) was about 20 to 50 percent lower than the rib with a 90 deg angle to the flow for a given heat transfer duty.

Han, J. C.↗
NASA NTRS1990

Conical Euler simulation and active suppression of delta wing rocking motion

A conical Euler code was developed to study unsteady vortex-dominated flows about rolling highly-swept delta wings, undergoing either forced or free-to-roll motions including active roll suppression. The flow solver of the code involves a multistage Runge-Kutta time-stepping scheme which uses a finite volume spatial discretization of the Euler equations on an unstructured grid of triangles. The code allows for the additional analysis of the free-to-roll case, by including the rigid-body equation of motion for its simultaneous time integration with the governing flow equations. Results are presented for a 75 deg swept sharp leading edge delta wing at a freestream Mach number of 1.2 and at alpha equal to 10 and 30 deg angle of attack. A forced harmonic analysis indicates that the rolling moment coefficient provides: (1) a positive damping at the lower angle of attack equal to 10 deg, which is verified in a free-to-roll calculation; (2) a negative damping at the higher angle of attack equal to 30 deg at the small roll amplitudes. A free-to-roll calculation for the latter case produces an initially divergent response, but as the amplitude of motion grows with time, the response transitions to a wing-rock type of limit cycle oscillation. The wing rocking motion may be actively suppressed, however, through the use of a rate-feedback control law and antisymmetrically deflected leading edge flaps. The descriptions of the conical Euler flow solver and the free-to-roll analysis are presented. Results are also presented which give insight into the flow physics associated with unsteady vortical flows about forced and free-to-roll delta wings, including the active roll suppression of this wing-rock phenomenon.

Lee, Elizabeth M.↗
NASA NTRS1978

Subsonic roll damping of a model with swept-back and swept-forward wings

The aerodynamic roll damping and the yawing moment due to roll rate characteristics were investigated at subsonic speeds for a model with either sweptback or swept forward wings. The tests were made in the Langley high speed 7 by 10 foot tunnel for Mach numbers between 0.3 and 0.7. The configuration with a 60 deg sweptback wing had positive damping in roll up to the maximum test angle of attack of almost 20 deg. The 32 deg swept forward wing configuration had positive damping in roll at the lower angles of attack, but there was a decrease in damping and negative damping in roll was measured at the highest angles of attack.

Boyden, R. P.↗
NASA NTRS1986

Experimental results of the control of a vortical flow by tangential blowing

The results of a wind tunnel test to investigate the controlling effects of tangential, leading edge blowing on the vortical flow over a delta wing are given. Blowing is used to directly control the crossflow separation points at the rounded leading edge and hence, the trajectory of the feeding sheet and the location of the vortex. Experiments were conducted for both co-flowing and counter-flowing configurations over a range of angles of attack from 0 to 90 degrees. Results in the form of pressure distributions, overall force coefficients and flow mappings were obtained. The emphasis is on data presentation rather than detailed analysis. The initial results indicate that the co-flowing configuration was capable of extending the regime of stable, controlled vortical flow over the upper surface by approximately 30 degrees angle of attack for modest blowing requirements. Increases in maximum normal force coefficient of approximately 30% were achieved and significant rolling moments produced at angles of attack from 30 to 60 degrees. The counter-flowing configuration indicated only minor lift augmentation with the exception of an isolated occurrence at 20 degrees angle of attack. At that condition, with very weak blowing, a lift augmentation of approximately 20 was measured.

Wood, N. J.↗
NASA NTRS1958

Elliptic Cones Alone and with Wings at Supersonic Speed

To help fill the gap in the knowledge of aerodynamics of shapes intermediate between bodies of revolution and flat triangular wings, force and moment characteristics for elliptic cones have been experimentally determined for Mach numbers of 1.97 and 2.94. Elliptic cones having cross-sectional axis ratios from 1 through 6 and with lengths and base areas equal to circular cones of fineness ratios 3.67 and 5 have been studied for angles of bank of 0 degree and 90 degrees. Elliptic and circular cones in combination with triangular wings of aspect ratios 1 and 1.5 also have been considered. The angle-of-attack range was from 0 degree to about 16 degrees, and the Reynolds number was 8 x 10(6), based on model length. In addition to the forces and moments at angle of attack, pressure distributions for elliptic cones at zero angle of attack have been determined. The results of this investigation indicate that there are distinct aerodynamic advantages to the use of elliptic cones.

Jorgensen, Leland H↗
NASA NTRS1979

Technology and operational considerations for low-heat-rate trajectories

A broad parametric study which examines several critical aspects of low-heat-rate entry trajectories is performed. Low planform loadings associated with future winged earth-entry vehicles coupled with the potential application of metallic thermal protection systems (TPS) suggest that such trajectories are of particular interest. Studied are three heating conditions - reference, stagnation, and windward centerline, for both laminar and turbulent flow; configuration-related factors including planform loading and hypersonic angle of attack; and mission-related factors such as cross-range and orbit inclination. Results indicate benefits in the design of TPS to be gained by utilizing moderate angles of attack as opposed to high-lift coefficient, high angles of attack, during entry. An assessment of design and technology implications is made.

Wurster, K. E.↗
NASA NTRS1993

Forebody vortex control on an F/A-18 using small, rotatable 'tip-strakes'

A mechanical scheme for manipulating the forebody vortices of an F/A-18, therefore creating controlled yawing moments at moderate and high angles of attack, was investigated in a wind tunnel. The technique consists of rotating miniature strakes (single or dual) about the radome centerline very close to the tip of the model. Forces, moments, and pressures were measured for angles of attack up to 60 deg and sideslip angles up to -10 deg. Results indicate that single and dual strakes can produce changes in side force and yawing moment, with magnitudes comparable to, or in some cases higher than, the directional changes produced by a 30-deg rudder deflection at 0-deg angle of attack. According to its circumferential position, the strake alters the separation location, inducing different degrees of asymmetry in the forebody vortex flow field that are translated into changes in side force and yawing moment. In comparison, the dual strakes appear to provide more gradual and better behaved changes than the single strake. Excellent correlation was found between this test and water tunnel tests performed on a similar configuration.

Suarez, Carlos J.↗
NASA NTRS1984

Low-speed investigation of effects of wing leading- and trailing-edge flap deflections and canard incidence on a fighter configuration equipped with a forward-swept wing

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.

Gainer, T. G.↗
NASA NTRS1981

Flight test experience with high-alpha control system techniques on the F-14 airplane

Improved handling qualities of fighter aircraft at high angles of attack can be provided by various stability and control augmentation techniques. NASA and the U.S. Navy are conducting a joint flight demonstration of these techniques on an F-14 airplane. This paper reports on the flight test experience with a newly designed lateral-directional control system which suppresses such high angle of attack handling qualities problems as roll reversal, wing rock, and directional divergence while simultaneously improving departure/spin resistance. The technique of integrating a piloted simulation into the flight program was used extensively in this program. This technique had not been applied previously to high angle of attack testing and required the development of a valid model to simulate the test airplane at extremely high angles of attack.

Gera, J.↗
NASA NTRS1999

Unstructured Grid Euler Method Assessment for Longitudinal and Lateral/Directional Stability Analysis of the HSR Reference H Configuration at Transonic Speeds

Transonic Euler computations, based on unstructured grid methodology, are performed for a proposed High Speed Civil Transport (HSCT) configuration, designated as the Reference H configuration within the High Speed Research (HSR) Program. The predicted results are correlated with appropriate experimental wind-tunnel data for the baseline configuration with and without control surface deflections for a range of angle of attack at M(sub infinity) = 0.95. Good correlations between the predictions and measured data have been obtained for the longitudinal aerodynamic characteristics of the baseline configuration. The incremental effects in the longitudinal aerodynamic characteristics due to horizontal rail deflections as well as wing leading-edge and trailing-edge flap deflections have also been predicted reasonably well. Computational results and correlations with data are also presented for the lateral and directional stability characteristics for a range of angle of attack at a constant sideslip angle as well as a range of sideslip angles at a constant angle of attack. In addition, the results are presented to assess the computational method performance and convergence characteristics.

Ghaffari, Farhad↗
NASA NTRSDate not supplied

A Wall-Modeled LES Perspective for the High Lift Common Research Model Using LAVA

A new immersed boundary Wall-Modelled Large Eddy Simulation (WMLES) formulation is developed to study high-lift aerodynamics on the NASA High-Lift Common Research Model (HL-CRM). A sequence of Cartesian Octree grids with sizes ranging from 100 Million through 2.02 Billion grid points is utilized to systematically assess grid-sensitivity and convergence for the in-tunnel (QinetiQ) configuration of the model, and remarkable agreement between the immersed boundary and the curvilinear body-aligned WMLES formulations is reported on grids with comparable resolutions. In the free-air configuration of the model, consistent predictions between the Curvilinear Overset and the Cartesian Octree formulations are reported for angles of attack up to C(L,max) at a=19.57. However, some differences in the onset of stall are seen between the two methods for a>20°: while the curvilinear WMLES experiences wing-root separation with increasing angle of attack (Topology A), the Cartesian Octree formulation shows a different flow topology characterized by boundary layer weakness on the main element, emanating from the pylon-wing attachment (Topology B). In order to obtain further insight into the two-distinct topologies, carefully designed numerical experiments to isolate effects of the model standoff and the tunnel wall-boundary layers are conducted using the immersed boundary WMLES formulation. The increased incidence angle-of-attack on the inboard portion of the wing due to the standoff is shown to be sufficient for triggering a switch from Topology B to Topology A in Cartesian WMLES. The role of the floor boundary layer is further examined in detail by identification of additional corner-flow vorticity generated by the viscous juncture flow interactions between the floor boundary layer and the standoff leading to formation of a strong coherent and persistent vortex on the belly-side of the fuselage. The intensity of this vortex is shown to increase with the thickness of the floor boundary layer. A further increase in the incidence angle of attack near the leading-edge strake caused by the presence of this belly-side vortex is quantified for two-distinct floor boundary layers. Both of the floor boundary layers considered result in the onset of large scale wing-root separation at a=21.47in non-confined (free-air) configurations.

TTT↗
NASA NTRS1961

Wind-Tunnel Tests of Seven Static-Pressure Probes at Transonic Speeds

Wind-tunnel tests have been conducted to determine the errors of 3 seven static-pressure probes mounted very close to the nose of a body of revolution simulating a missile forebody. The tests were conducted at Mach numbers from 0.80 to 1.08 and at angles of attack from -1.7 deg to 8.4 deg. The test Reynolds number per foot varied from 3.35 x 10(exp 6) to 4.05 x 10(exp 6). For three 4-vane, gimbaled probes, the static-pressure errors remained constant throughout the test angle-of-attack range for all Mach numbers except 1.02. For two single-vane, self-rotating probes having two orifices at +/-37.5 deg. from the plane of symmetry on the lower surface of the probe body, the static-pressure error varied as much as 1.5 percent of free-stream static pressure through the test angle-of- attack range for all Mach numbers. For two fixed, cone-cylinder probes of short length and large diameter, the static-pressure error varied over the test angle-of-attack range at constant Mach numbers as much as 8 to 10 percent of free-stream static pressure.

Capone, Francis J.↗
NASA NTRS2003

Control of Interacting Vortex Flows at Subsonic and Transonic Speeds Using Passive Porosity

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) 8-Foot Transonic Pressure Tunnel (TPT) to determine the effects of passive surface porosity on vortex flow interactions about a general research fighter configuration at subsonic and transonic speeds. Flow-through porosity was applied to a wing leading-edge extension (LEX) mounted to a 65 deg cropped delta wing model to promote large nose-down pitching moment increments at high angles of attack. Porosity decreased the vorticity shed from the LEX, which weakened the LEX vortex and altered the global interactions of the LEX and wing vortices at high angles of attack. Six-component forces and moments and wing upper surface static pressure distributions were obtained at free-stream Mach numbers of 0.50, 0.85, and 1.20, Reynolds number of 2.5(10(exp 6)) per foot, angles of attack up to 30 deg, and angles of sideslip to +/- 8 deg. The off-surface flow field was visualized in selected cross-planes using a laser vapor screen flow visualization technique. Test data were obtained with a centerline vertical tail and with alternate twin, wing-mounted vertical fins having 0 deg and 30 deg cant angles. In addition, the porosity of the LEX was compartmentalized to determine the sensitivity of the vortex-dominated aerodynamics to the location and level of porosity applied to the LEX.

Erickson, Gary E.↗
NASA NTRS1990

Application of variable-gain output feedback for high-alpha control

A variable-gain, optimal, discrete, output feedback design approach that is applied to a nonlinear flight regime is described. The flight regime covers a wide angle-of-attack range that includes stall and post stall. The paper includes brief descriptions of the variable-gain formulation, the discrete-control structure and flight equations used to apply the design approach, and the high performance airplane model used in the application. Both linear and nonlinear analysis are shown for a longitudinal four-model design case with angles of attack of 5, 15, 35, and 60 deg. Linear and nonlinear simulations are compared for a single-point longitudinal design at 60 deg angle of attack. Nonlinear simulations for the four-model, multi-mode, variable-gain design include a longitudinal pitch-up and pitch-down maneuver and high angle-of-attack regulation during a lateral maneuver.

Ostroff, Aaron J.↗
NASA NTRS1954

A Transonic Wind-Tunnel Investigation of the Longitudinal Aerodynamic Characteristics of a Model of the Lockheed XF-104 Airplane

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.

Hieser, Gerald↗
NASA NTRS1961

Flight Tests of a 1/6-Scale Model of the Hawker P 1127 Jet VTOL Airplane

An experimental investigation has been made to determine the dynamic stability and control characteristics of a 1/6-scale flying model of the Hawker P lIP7 jet vertical-take-off-and-landing (VTOL) airplane in hovering and transition flight. The model was powered by a counter-rotating ducted fan driven by compressed-air jets at the tips of the fan blades. In hovering flight the model was controlled by jet-reaction controls which consisted of yaw and pitch jets at the extremities of the fuselage and a roll jet on each wing tip. In forward flight the model was controlled by conventional ailerons and rudder and an all-movable horizontal tail. In hovering flight the model could be flown smoothly and easily, but the roll control was considered too weak for rapid maneuvering or hovering in gusty air. Transitions from hovering to normal forward flight and back to hovering could be made smoothly and consistently and with only moderate changes in longitudinal trim. The model had a static longitudinal instability or pitch-up tendency throughout the transition range, but the rate of divergence in the pitch-up was moderate and the model could be controlled easily provided the angle of attack was not allowed to become too high. In both the transition and normal forward flight conditions the lateral motions of the model were difficult to control at high angles of attack, apparently because of low directional stability at small angles of sideslip. The longitudinal stability of the model in normal forward flight was generally satisfactory, but there was a decided pitch-up tendency for the flap-down condition at high angles of attack. In the VTOL landing approach condition, with the jets directed straight down or slightly forward, the nose-down pitch trim required was greater than in the transitions from hovering to forward flight, but the longitudinal instability was about the same. Take-offs and landings in still air could be made smoothly although there was a slight unfavorable ground effect on lift and a nose-down change in pitch trim near the ground. Short take-offs and landings could be made smoothly and consistently although the model experienced a decided nose-up change in pitching moment as it climbed out of ground effect.

Smith, Charles C., Jr.↗