Showing papers on "Pitching moment published in 1988"

Experimental results for the Eppler 387 airfoil at low Reynolds numbers in the Langley low-turbulence pressure tunnel

Abstract: Experimental results were obtained for an Eppler 387 airfoil in the Langley Low Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.03 to 0.13 and a chord Reynolds number range for 60,000 to 460,000. Lift and pitching moment data were obtained from airfoil surface pressure measurements and drag data for wake surveys. Oil flow visualization was used to determine laminar separation and turbulent reattachment locations. Comparisons of these results with data on the Eppler 387 airfoil from two other facilities as well as the Eppler airfoil code are included.

Actively controlled automotive suspension system with adjustable response characteristics in pitching-suppressive control

Abstract: A suspension control system for an automotive vehicle is designed for effectively suppressing pitching motion and whereby regulating vehicular attitude. The suspension control system monitors pitching moment to be exerted on the vehicle for adjusting response characteristics in pitching-suppressive control depending upon the magnitude of the pitching moment to be exerted. The suspension control system adjusts the response characteristics of the front and rear suspensions independently of each other according to preset pitching-suppressive control coefficients respectively for the front and rear suspensions, which are set depending upon the suspension characteristics and/or suspension geometory of the vehicle, to which the suspension control system is to be applied.

Airfoil pressure measurements during a blade vortex interaction and a comparison with theory

Abstract: A two-dimensional, airfoil-vortex interaction experiment was conducted to obtain the aerodynamic behavior of the airfoil during a parallel interaction. The vortex was created by pitching a second airfoil placed upstream and parallel to an instrumented test airfoil. Hot-wire anemometer measurements were taken to determine the vortex velocity distribution. The interaction tests were conducted for a counterclockwise vortex passing above a symmetrical airfoil at zero angle of attack

Effect of large amplitude pitching motions on the unsteady aerodynamic characteristics of flat-plate wings

Abstract: Large-amplitude unsteady motion effects on the aerodynamic force and stability characteristics of flat-plate wings were investigated in a wind tunnel for the cases of two delta wings with respective leading edge sweeps of 70 and 45 deg, and a rectangular wing whose aspect ratio was equal to that of the 70-deg delta wing. Attention was given to the effects of reduced frequency and mean angle of attack. It is found that lags in vortex burst location and separation/reattachment of flow on the upper surface of the wing produced large overshoots and hysteresis loops in normal force and pitching moment coefficients that were a strong function of mean oscillation angle and reduced frequency.

Experimental aerodynamic characteristics of an NACA 0012 airfoil with simulated glaze ice

Abstract: The effect of a simulated glaze-ice accretion on the aerodynamic performance of a NACA 0012 airfoil was studied experimentally. Two ice shapes were tested: one from an experimentally measured accretion, and one from an accretion predicted using a computer model given the same icing conditions. Lift, drag, and pitching moment were measured for the airfoil with both smooth and rough ice shapes. The ice shapes caused large lift and drag penalties, primarily due to large separation bubbles. Surface pressure distributions clearly showed the regions of separated flow. The aerodynamic performance of the two shapes compared well at positive, but not negative, angles of attack.

Numerical study of the vortex burst phenomenon for delta wings

Abstract: A flux-difference splitting scheme is employed to compute low-speed flows over a delta wing for angles of attack from 0 to 40 deg as steady-state solutions to the three-dimensional, Reynolds-averaged Navier-Stokes equations in their thin-layer approximation. The finite-difference scheme is made spatially second-order accurate by applying a total variation diminishing-like discretization to the inviscid fluxes and central differencing to the viscous shear fluxes. Using first-order accurate Euler backward-time differencing, an efficient implicit algorithm is constructed, which combines approximate factorization in cross planes with a symmetric planar Gauss-Seidel relaxation in the remaining third spatial direction. The geometry of the thin (maximum thickness is 0.021), slender (aspect ratio is unity), sharp-edged delta wing is taken from Hummel's (1967, 1978) wind tunnel model. Over the entire angle-of-attack range, the computed values of lift and pitching moment are in good agreement with the experimental data. Also details of the flow-fieldlike spanwise surface pressure distributions compare well with the experiment. Computed flow-field results with a bubble-type vortex burst are analyzed in detail.

Experimental study of three-lifting surface configuration

Abstract: The aerodynamic ramifications of utilizing three lifting surfaces as opposed to the conventional or canard lifting configurations have been studied on a theoretical basis by previous researchers. This paper presents an experimental investigation of various configuration modifications for an unyawed typical business jet at a Reynolds number of 1.3 million. The three surface has better lift and high-lift drag characteristics than either the canard or tail-aft configurations, but the cruise drag is more. The induced drag at cruise is highest for the canard and lowest for the tail-aft configuration. The pitching moment characteristics are somewhat between those of the canard and tail-aft configurations. A decrease in gap adversely affects the pitching moment characteristics. A smaller stagger leads to better aerodynamic and stability characteristics. A decrease in span of the forward wing gives better cruise drag and longitudinal stability characteristics, but has adverse effects on high-lift drag. A variation in the incidence angles of either or both the forward and aft wings changes the zero-lift moments of the configuration, while marginally affecting overall lift and drag. At cruise, the lift to drag ratio is highest for the conventional and lowest for the three surface. For high lift conditions, the order is reversed.

Unsteady Aerodynamic Forces at Low Airfoil Pitching Rates

Abstract: : Experiments were conducted on a NACA-0015 airfoil undergoing low constant pitch rates to study the effects of dynamic stall formation on the airfoil upper surface pressure field The airfoil was pitched about pivot locations of 025c, 05c, and 075c at nondimensional pitch rates below 02 Lift and drag coefficients were evaluated for all cases, and smoke flow visualization at low pitch rates was studied for the quarter chord pivot location Results indicate that the greatest increases in lift due to the pitching motion occur prior to the nondimensional pitch rate of 01 for all three pivot locations The effects of pitch rate on the maximum lift and drag values appear similar for the three pivot locations studied Lift to drag ratios show significant enhancement even at very low nondimensional rates Flow visualization indicates that the leading edge dynamic stall vortex is present even at very low nondimensional pitch areas Keywords: Aerodynamic lift/drag; Pitch motion; Airfoils; Leading edges; Stall vortices; Aerodynamic forces; Unsteady flow; Reprints

Theory of Idealized Two-Dimensional Ballute in Newtonian Hypersonic Flow

Abstract: A differential equation governing the geometry of a two-dimensional ballute in hypersonic flow and its constraining boundary conditions are derived under idealized assumptions. By solving these equations, the shape of the ballute is determined over a range of conditions. Lift, drag, pitching moment, and the allowed limit of center-of-gravity location for stability (meta-center) are then calculated using Newtonian hypersonic approximation. It is shown that the meta-center occurs near the forward end because of compliance of the ballute membrane to the shock layer pressures, especially at low free-stream densities. In order for the vehicle employing the ballute to be stable at all densities, the center of gravity must be within approximately the forward 20 percent of overall length of the vehicle. However, typical flight trajectories of an aeroassisted orbital transfer vehicle employing the ballute for aerobraking show that the vehicle may be able to complete its atmospheric flight without tumbling provided that the center of gravity is located within the forward 43 percent of the vehicle length because of the relatively short duration of flight through the destabilizing low-density regime.

An experimental investigation of the aerodynamic characteristics of slanted base ogive cylinders using magnetic suspension technology

Abstract: This paper reports on an experimental investigation of aerodynamic characteristics of slanted base ogive cylinders at zero incidence. The Mach number range is 0.05 to 0.3. In this investigation, magnetically suspending the wind tunnel models eliminates flow disturbances associated with mechanical supports. This paper reports on the drastic changes in lift, pitching moment, and drag for a slight change in base slant angle. Flow visualization with liquid crystals and oil is used to observe base flow patterns responsible for the sudden changes in aerodynamic characteristics. This paper also reports on hysteretic effects that are present and discusses computational results using VSAERO and SANDRAG.

Further development of pneumatic thrust-deflecting powered-lift systems

Abstract: Wind-tunnel evaluations of two pneumatic thrust-deflecting powered-lift systems have been conducted to develop the capabilities of interchangeab le thrust recovery and reversal as well as longitudinal pitch trim. A circulation control wing/vectored thrust configuration employed underwing Pegasus-type nozzles to redirect the horizontal thrust component as needed for STOL operation and to provide nose-up pitching moment for trim. Although they provided a vertical thrust component to lift, the vectoring nozzles were relatively ineffective in augmenting aerodynamic lift. A circulation control wing/over the wing blowing configuration pneumatically deflected engine thrust for additional high lift beyond that provided by CCW alone. It also allowed pneumatic conversion of the resultant force along the flight path from thrust recovery to thrust reversal as required for takeoff or approach. Both configurations thus offer possible solutions to STOL operational problems, one by pneumatic/mechanical means and the other primarily pneumatically.

F-15 SMTD low speed jet effects wind tunnel test results

Abstract: Key results from low speed wind tunnel testing of the F-15 STOL and Maneuver Technology Demonstrator (SMDT) with thrust reversers are presented. Longitudinally, the largest induced increments in the stability and control occur at landing gear height. These generally reflect an induced lift loss and a nose-up pitching moment, and vary with sideslip. Directional stability is reduced at landing gear height with full reverse thrust. Nonlinearities in the horizontal tail effectiveness are found in free air and at landing gear height.

Numerical computations of transonic critical aerodynamic behavior

Abstract: The determination of aerodynamic coefficients by shell designers is a critical step in the development of any new projectile design. Of particular interest is the determination of the aerodynamic coefficients at transonic speeds. It is in this speed regime that the critical aerodynamic behavior occurs and a rapid change in the aerodynamic coefficients is observed. Three-dimensional, transonic, flowfield computations over projectiles have been made using an implicit, approximately factored, partially flux-split algorithm. A composite grid scheme has been used to provide the increased grid resolution needed for accurate numerical simulation of three-dimensional transonic flows. Details of the asymmetrically located shock waves on the projectiles have been determined. Computed surface pressures have been compared with experimental data and are found to be in good agreement. The pitching moment coefficient, determined from the computed flowfields, shows the critical aerodynamic behavior observed in free flights. I. Introduction T HE flight of projectiles covers a wide range of speeds. The accurate prediction of projectile aerodynamics at these speeds is of significant importance in the early design stage of a projectile. The critical aerodynamic behavior occurs in the transonic speed regime, 0.9

VSAERO analysis of tip planforms for the free-tip rotor

Abstract: The results of a numerical analysis of two interacting lifting surfaces separated in the spanwise direction by a narrow gap are presented. The configuration consists of a semispan wing with the last 32 percent of the span structurally separated from the inboard section. The angle of attack of the outboard section is set independently from that of the inboard section. In the present study, the three-dimensional panel code VSAERO is used to perform the analysis. Computed values of tip surface lift and pitching moment coefficients are correlated with experimental data to determine the proper approach to model the gap region between the surfaces. Pitching moment data for various tip planforms are also presented to show how the variation of tip pitching moment with angle of attack may be increased easily in incompressible flow. Calculated three-dimensional characteristics in compressible flow at Mach numbers of 0.5 and 0.7 are presented for new tip planform designs. An analysis of sectional aerodynamic center shift as a function of Mach number is also included for a representative tip planform. It is also shown that the induced drag of the tip surface is reduced for negative incidence angles relative to the inboard section. The results indicate that this local drag reduction overcomes the associated increase in wing induced drag at high wing lift coefficients.

Aerodynamic Behavior of Inclined Circular Cylinders-Cable Aerodynamics

Abstract: This study attempts to clarify of the mechanism of rain-wind induced vibration of the stay cables of a cable stayed bridge and to find an effective aerodynamic stabilization method. Through a series of wind tunnel tests the fundamental aerodynamic characteristics of a yawed and/or inclined circular Cylinder with and without rain were investigated. An intense secondary axial flow was found to form in the early wake, playi resulted in an aerodynamic ng a similar role to that of a sputter plate submerged in the wake. This axial flow exciting force acting on the yawed and/or the inclined circular cylinder. The role of rain in the rain-wind induced vibration of the stay cable is shown to be that of an amplifier of the essential unstable aerodynamic exciting force acting on the yawed and/or the inclined circular cylinder. The role of rain the rain-wind induced vibration of the stay cable is shown to be that of an _amplifier of the essential unstable aerodynamic characteristics of the yawed and/or the inclined circuolanrdezry linfiow. der. Aerodynamic stabilization should depend essentially on controlling this characteristic sec-

SAMID, an interactive system for aircraft drag minimization studies. Mathematical models and methods

Abstract: An interactive computer program system has been developed that can be used in aircraft drag minimization studies It comprises algorithms for choosing the spanwise distributions of lift, pitching moment, chord and thicknessto-chord ratio of lifting elements The choices are optimal in that they minimize induced plus viscous drag while satisfying constraints of aerodynamic, flight-mechanical and structural nature The configurations that can be dealt with may consist of a number of segments representing, for instance, wings or parts of wings, horizontal tails or canards, winglets, flaprail-fairrings, etc Also the interaction between propellers and lifting elements may be included in the procedure The induced drag is computed using the Trefftz-plane Integral (farfield analysis), while the viscous drag follows from form factor methods Novel mathematical formulations of the constrained optimization problems are used, that are based on the calculus of variations The analysis and optimization method is fast, numerically stable and easy to use, and therefore suitable for interactive design purposes where rapid configuration trade-offs have to be made This report presents the theoretical models and methods underlying the anlysis and optimization capability, comparisons with other theories, and some examples of applications

Two-dimensional aerodynamic characteristics of the OLS/TAAT airfoil

Abstract: Two flight tests have been conducted that obtained extension pressure data on a modified AH-1G rotor system. These two tests, the Operational Loads Survey (OLS) and the Tip Aerodynamics and Acoustics Test (TAAT) used the same rotor set. In the analysis of these data bases, accurate 2-D airfoil data is invaluable, for not only does it allow comparison studies between 2- and 3-D flow, but also provides accurate tables of the airfoil characteristics for use in comprehensive rotorcraft analysis codes. To provide this 2-D data base, a model of the OLS/TAAT airfoil was tested over a Reynolds number range from 3 x 10 to the 6th to 7 x 10 to the 7th and between Mach numbers of 0.34 to 0.88 in the NASA Langley Research Center's 6- by 28-Inch Transonic Tunnel. The 2-D airfoil data is presented as chordwise pressure coefficient plots, as well as lift, drag, and pitching moment coefficient plots and tables.

Tether satellite potential for rarefied gas aerodynamic research

Abstract: Applications of the tether satellite in aerodynamic research under conditions of low-density hypervelocity flow are described. The satellite is envisioned as a spheroidal nose, slab-sided cylinder with variably swept wing, tethered to the shuttle orbiter and deployed to altitudes of about 95 km and above. Suggested experiments reflect requirements for new understanding of low-density flows and include the direct measurement of normal and tangential stress on representative aerospace surfaces, the measurements of vehicle lift, drag and pitching moment, ambient gas density and composition, surface temperatures, and wall tap and impact probe behavior. Also suggested are measurements of gas densities above the wing surfaces and on or near the stagnation line using electron-beam fluorescence methods and, where applicable, free-molecule orifice probes. Experiments proposed may be implemented without significant development of new technology.

A new method for designing natural laminar flow airfoils

Abstract: A New method for designing Natural Laminar Airfoils is13; described. It is based on three ideas: Firstly, a plausible13; pressure distribution, that can maintain extensive laminar flow on the airfoil, is guessed and an inverse design method is then used to obtain the airfoil that sustains the given pressure distribution. If such an airfoil does not exist the method then changes the input pressure distribution. Secondly, it incorporates an appropriate camber line to obtain target lift and pitching moment coefficients. Lastly, it tests the airfoil for performance by a viscous code, compares the same with the given specifications. If the target performance is not met it changes the pressure distribution and goes through the entire iteration again. Some examples are given to illustrate the method.

Aerodynamics of complex bodies of revolution

Abstract: The aerodynamics of five complex bodies of revolution are investigated experimentally and theoretically at a low Mach number (M f ^ 01) and over the range of angle of attack from 0 to 35 deg The geometrical forms of the bodies are generally complex with the discontinuities in the slope of the body surface The surface-flow visualization is performed by using the oil method The balance measurements were made and the results compared with the potential theory and the method based on the crossflow analogy It was observed that the discontinuities in the slope of the body surface make the flow separation and consequently the flowfield very complicated It was also found that the method of crossflow analogy is applicable not only to simple-type bodies of revolution but also the complex ones Nomenclature CA — axial force coefficient CD = drag coefficient CDO — drag coefficient at a = 0 deg Cd = crossflow drag coefficient cdc = crossflow drag coefficient of circular cylinder section CL = lift coefficient CM = pitching moment coefficient d — maximum body diameter ds = diameter of corresponding cylinder to body of revolution ( = S p /l) /(X) = normal force distribution / = total body length lr = reference length M^ = freestream Mach number MCoo = crossflow Mach number ( = M^ sina) q^ = freestream dynamic pressure (q^ = ^p R = local body radius Red = Reynolds number based on the maximum body diameter d (Re d — V^d/v) Recds = crossflow Reynolds number based on ds ( = F^ sina ds/v) Ret - Reynolds number based on / S = local cross-sectional area ( = nR2) Sb = body base area Sp = body planform area ( = Jo 2R d*) Sr = reference area S, = body flat-nose area V^ = freestream velocity W = body volume ( = &nR2 d*) x = body axis (axial distance from body nose) xc = axial distance from body nose to centroid of body planform area xcp = axial distance from body nose to center of pressure xm = pitching moment center (axial distance from body nose to pitching moment center) a = angle of attack p = density of air D = kinematic viscosity of air r\ — correction factor for influence of fineness ratio

A Hybrid Analytical/Numerical Computational Method for Supercritical Transonic Flow Fields

Abstract: Shock-boundary layer interaction can significantly alter not only the local transonic flow on missiles, wings and turbine blades but its influence can also extend significantly downstream within the boundary layer and thereby alter the global aerodynamic properties of lift, drag and pitching moment. It is therefore important that these interactions and their Reynolds and Mach number scaling be included in analysis of practical aerodynamic flow fields. This paper describes the features of a non-asymptotic triple deck theory of shock-turbulent boundary layer interaction and its application as an element in the overall viscous-inviscid flow analysis of the body (Fig. 1).

Transient aerodynamic forces on a fighter model during simulated approach and landing with thrust reversers

Abstract: Previous wind tunnel tests of fighter configurations have shown that thrust reverser jets can induce large, unsteady aerodynamic forces and moments during operation in ground proximity. This is a concern for STOL configurations using partial reversing to spoil the thrust while keeping the engine output near military (MIL) power during landing approach. A novel test technique to simulate approach and landing was developed under a cooperative Northrop/NASA/USAF program. The NASA LaRC Vortex Research Facility was used for the experiments in which a 7-percent F-18 model was moved horizontally at speeds of up to 100 feet per second over a ramp simulating an aircraft to ground rate of closure similar to a no-flare STOL approach and landing. This paper presents an analysis of data showing the effect of reverser jet orientation and jet dynamic pressure ratio on the transient forces for different angles of attack, and flap and horizontal tail deflection. It was found, for reverser jets acting parallel to the plane of symmetry, that the jets interacted strongly with the ground, starting approximately half a span above the ground board. Unsteady rolling moment transients, large enough to cause the probable upset of an aircraft, and strong normal force and pitching moment transients were measured. For jets directed 40 degrees outboard, the transients were similar to the jet-off case, implying only minor interaction.