Showing papers on "Pitching moment published in 1998"

Wind Turbine Performance Under Icing Conditions

Abstract: The effects of rime ice on horizontal axis wind turbine performance were estimated. For typical supercooled fog conditions found in cold northern regions, four rime ice accretions on the S809 wind turbine airfoil were predicted using the NASA LEWICE code. The resulting airfoil/ice profile combinations were wind tunnel tested to obtain the lift, drag, and pitching moment characteristics over the Reynolds number range 1--2 {times} 10{sup 6}. These data were used in the PROPID wind turbine performance prediction code to predict the effects of rime ice on a 450-kW rated-power, 28.7-m diameter turbine operated under both stall-regulated and variable-speed/variable-pitch modes. Performance losses on the order of 20% were observed for the variable-speed/variable-pitch rotor. For the stall-regulated rotor, however, a relatively small rime ice profile yielded significantly larger performance losses. For a larger 0.08c-long rime ice protrusion, however, the rated peak power was exceeded by 16% because at high angles the rime ice shape acted like a leading edge flap, thereby increasing the airfoil C{sub l,max} and delaying stall.

Flutter of an airfoil with a cubic nonlinear restoring force

Abstract: In this paper, the effect of a cubic structural restoring force on the flutter characteristics of a twodimensional airfoil placed in an incompressible flow is investigated. The aeroelastic equations of motion are written as a system of eight first-order ordinary differential equations. Given the initial values of plunge and pitch displacements and their velocities, the system of equations is integrated numerically using a 4 order Runge-Kutta scheme. Results for softand hard-springs are presented for a pitch degree-of-freedom nonlinearity. The study shows the dependence of the divergence flutter boundary on initial conditions for a soft spring. For a hard spring, the nonlinear flutter boundary is independent of initial conditions for the spring constants considered. The flutter speed is identical to that for a linear spring. Divergent flutter is not encountered, but instead limit cycle oscillation occurs for velocities greater than the flutter speed. The behaviour of the airfoil is also analyzed using analytical techniques developed for nonlinear dynamical systems. The Hopf-bifurcation point is determined analytically and the amplitude of the limit cycle oscillation in postHopf-bifurcation for a hard spring is predicted using an asymptotic theory. The frequency of the limit cycle *Principal Research Officer and Head, Experimental Aerodynamics and Aeroelasticity Group. Also adjunct professor, Depl. of Mathematical Sciences, University of Alberta. Associate Fellow AIAA. "Research Associate, Experimental Aerodynamics and Aeroelasticity Group. 'Professor, Dept. of Mathematical Sciences. Copyright © 1998 by B.H.K. Lee. L.Y. Jiang and Y.S. Wong, Published by the American Inst i tute of Aeronautics and Astronautics Inc. with permission oscillation is estimated from an approximate method. Comparisons with numerical simulations are carried out and the accuracy of the approximate method is discussed. The analysis can readily be extended to study limit cycle oscillation of airfoils with nonlinear polynomial spring forces in both plunge and pitch degrees of freedom. NOMENCLATURE ah non-dimensional distance from airfoil midchord to elastic axis b airfoil semi-chord CL aerodynamic lift coefficient CM pitching moment coefficient h plunge displacement m airfoil mass R response amplitude of pitch motion r response amplitude of plunge motion ra radius of gyration about elastic axis t time U free stream velocity U non-dimensional velocity, U/bcoa UL non-dimensional linear flutter speed xa non-dimensional distance from elastic axis to centre of mass X system variable vector XE system equilibrium point y variable vector a pitch angle of airfoil OCA pitch angle amplitude of l imit cycle oscillation EI, £2 constants in Wagner's function Pa, p= coefficients of cubic spring in pitch and plunge C,a, viscous damping ratios in pitch and plunge H airfoil/air mass ratio, m/Kpb

Aerodynamic and aero acoustic optimization of airfoils via a parallel genetic algorithm

Abstract: A parallel genetic algorithm (GA) was used to generate, in a single run, a family of aerodynamical ly efficient, low-noise rotor blade designs representing the Pareto optimal set. The n-branch tournament, uniform crossover genetic algorithm operates on twenty design variables, which constitute the control points for a spline representing the airfoil surface. The GA takes advantage of available computer resources by operating in either serial mode or "manager/work er" parallel mode. The multiple objectives of this work were to maximize lift-to-drag of a rotor airfoil shape and to minimize an overall noise measure including effects of loading and thickness noise of the airfoil. Constraints are placed on minimum lift coefficient, pitching moment and boundary layer convergence. The program XFOIL provides the aerodynamic analysis, and the code WOPWOP provides the aeroacoustic analysis. The Pareto-optimal airfoil set has been generated and is compared to the performance of a typical rotorcraft airfoil under identical flight conditions.

Prediction and Validation of Mars Pathfinder Hypersonic Aerodynamic Data Base

Abstract: Postflight analysis of the Mars Pathfinder hypersonic, continuum aerodynamic data base is presented. Measured data include accelerations along the body axis and axis normal directions. Comparisons of preflight simulation and measurements show good agreement. The prediction of two static instabilities associated with movement of the sonic line from the shoulder to the nose and back was confirmed by measured normal accelerations. Reconstruction of atmospheric density during entry has an uncertainty directly proportional to the uncertainty in the predicted axial coefficient. The sensitivity of the moment coefficient to freestream density, kinetic models and center-of-gravity location are examined to provide additional consistency checks of the simulation with flight data. The atmospheric density as derived from axial coefficient and measured axial accelerations falls within the range required for sonic line shift and static stability transition as independently determined from normal accelerations.

Aerodynamic Performance of Supersonic Missile Body- and Wing Tip-Mounted Lateral Jets

Abstract: Forward, missile body-mounted, lateral jet thrusters in a windward orientation yield amplie cation factors that are well below unity ( 0.5). This deamplie cation is caused by the intense interaction between the oncoming freestream and the jet e ow, causing massive loss of the favorable pressure on the windward wings. Additionally, the jet e ow creates a blockage effect that extends to the windward rear tail panels causing reduced tail control for this missile orientation. A means to enhance the amplie cation factor and regain control of tail panels is addressed by studying several alternate locations of the lateral jet thruster with a e xed body/wing/tail missile geometry. It is shown that the jet thrusters mounted on the wing tips enhance the amplie cation factor to 1.06 as compared to 0.5 for the forward missile body-mounted lateral thruster. This approach also recovers the control power of the windward rear tail panels. Specie c generic missile studies are presented after extensive jet interaction validations of the computational method with the wind-tunnel data. Nomenclature AF = amplie cation factor, 1 CNjet CNno-jet T q S Clm = rolling moment coefe cient, Mx q S Xref Cm = pitching moment coefe cient, My q S Xref CN = normal force coefe cient, N q S , airframe only CY = side force coefe cient, FY q S CYm = yawing moment coefe cient, Mz q S Xref dp = pressure differential, Pjet Pno-jet Pinf FY = side force, N M = freestream Mach number Mx = rolling moment My = pitching moment Mz = yawing moment N = normal force, N P = pressure, N/m 2 q = dynamic pressure, 1 2 S = missile cross-sectional area, m 2 T = jet thrust, N = velocity, m/s = angle of attack, deg = ratio of specie c heats = density = azimuth angle, deg

An experimental and computational investigation of spanwise-step-ice shapes on airfoil aerodynamics

Abstract: The objective of this research was to study the effects of spanwise-step-ice accretions (resulting from large droplet icing conditions) on subsonic aircraft aerodynamics. The airfoil investigated was a modified NACA 23012 with a simple flap. An experimental and computational program was conducted using simulated ice accretions to determine the sensitivity of ice shape size and location on airfoil performance and control as a function of angle of attack and flap deflection. Focus is paid on the critical conditions where the aerodynamic performance, and the hinge moment in particular, changes rapidly and non-linearly. The experimental program included wake surveys, surface pressure taps, and force-balance measurements to obtain lift, drag, pitching moment, and hinge-moment coefficients for a large variety of geometry and flow conditions. The accompanying computational investigation was performed with a high-resolution full Navier-Stokes solution using a solution-adaptive unstructured grid for both non-iced and iced configurations. Results are presented for experiments and predictions of sectional aerodynamic characteristics where the quarter-round ice shape heights of 0.0083 and 0.0139 chords resulted in a dramatic decrease in maximum lift coefficients as well as significant reductions in hinge moments for positive angles of attack.

Effects of Ballistic Damage on the Aerodynamics of Helicopter Rotor Airfoils

Abstract: Experiments were conducted to estimate the effects of ballistic damage on the aerodynamics of helicopter rotor airfoil sections. The lift, pitching moment, and drag were measured on nominally twodimensional blade specimens with representative prescribed and actual ballistic damage. The measurements were made at chord Reynolds numbers between 1 3 10 6 and 3 3 10 6 and Mach numbers up to 0.28. Force balance measurements were complemented by chordwise and spanwise pressure measurements to assess the three-dimensional nature of damage on the aerodynamics. The quantitative data were supplemented by surface oil-e ow visualization. Generally, it was found that ballistic damage degraded the aerodynamic performance of the blade specimens, with a reduction in lift accompanied by signie cant increase in drag and change in the center of pressure. However, in some cases signie cant damage produced surprisingly mild effects. The results are useful in helping to model and assess the overall vulnerability of helicopters to ballistic damage.

Waverider Aerodynamics and Preliminary Design for Two-Stage-to-Orbit Missions, Part 2

Abstract: Osculating cones waveriders are evaluated for future reusable space transportation systems An integrated design program hasbeen developed by theInstitute of AircraftDesign and Structure Mechanics, which simulates the iterative design process including the main interactions between the involved disciplines Starting points are the dee nition of the design task (range, payload to orbit, mission and orbit proe le, etc )and a generic waverider shape with its aerodynamic characteristics The design program contains the geometric sizing of the aircraft shape, a realistic engine model with bookkeeping, aerodynamic load calculations, structure mass estimation with a e nite element method, thermal protection design, and a e ight-path simulation with trim control The predesign investigations focus on the design of a two-stage-to-orbit aerospace plane with a waverider-based e rst stage performing a cruise e ight from Europe to the equator with an orbiter transferring a maximum payload of 7 tons to a low Earth orbit The comparison with a conventional hypersonic cone guration indicates that it should be possible to fule ll the dee ned mission with 35% less fuel and an 11% lower takeoff weight These results underline the potential of the waverider concept Nomenclature FX, FZ = forces in longitudinal and vertical direction L D = lift over drag ratio, aerodynamic efe ciency M = pitching moment n = load factor S = wing reference area, m 2 V = maximum total aircraft volume, m 3

Aerodynamic Performance of an Airfoil with Step-Induced Vortex for Lift Augmentation

Abstract: Physical and numerical experiments on flow developments around an NACA-0012 airfoil were conducted to explore the possibility of enhancing the airfoil's aerodynamic performance by vortex lift augmentation. The paper focuses on the effects of the separated flow and subsequent vortex formation, generated by backward-facing steps on pressure distributions and corresponding flow occurrences around the airfoil. Various step configurations are examined to determine their effect on lift and on lift-to-drag ratios. A discussion of the effects of main geometrical parameters of upper and lower surface steps on the airfoil performance, based on computational and physical flow visualization experiments, are presented. The results suggest that incorporation of backward-facing steps on the lower surface that are located at the midchord and extend back to the trailing edge with 50 depth of the airfoil chord may lead to considerable enhancements in lift coefficients and lift-to-drag ratios. The data produced may serve as...

The aerodynamics of a chined forebody

Abstract: A low-speed experimental study of the effects of angle of attack on the flowfield of a chined forebody has been performed. These tests were conducted in the University of Illinois low speed wind tunnel, and in the Boeing Shear Flow Facility. The high fidelity, NC machined aluminum model was sting mounted and positioned in pitch from 0° to 52° angle of attack (a) without sideslip. The effects of Reynolds number (Rj) were also investigated by running a range of tunnel velocities. Reynolds number, based on the 3-inch base width of the forebody, was varied from 1.4xl0 to 2.8xl0. Surface pressures were measured at all conditions using an array of 91 static pressure taps. Normal force and pitching moment were measured by an internal strain gauge balance. Laser Doppler velocimetry (LDV) data and smoke flow visualization were utilized to determine the vortex system. Surface oil flow visualization was utilized to define the surface flowfield and help interpret the surface pressure data. Steady flow visualization revealed the importance of boundary-layer separation on the leeward surface of the chined forebody. Primary separation was caused by the sharp chine edge, and resulted in the formation of large primary vortices. Secondary separation was caused by a steep spanwise surface pressure gradient between the chine edge and the suction peak associated with the primary vortex. This secondary separation resulted in the formation of small secondary vortices. This separation and vortex formation drove the aerodynamics of the chined forebody.

Flight Control Law Using Nonlinear Dynamic Inversion Combined With Quantitative Feedback Theory

Abstract: A method of designing control laws for uncertain nonlinear systems is presented. Dynamic inversion is used to partially linearize the dynamics and then a nonlinear version of quantitative feedback theory (QFT) is applied to the resulting system which assures robustness to plant uncertainty. The design yields good performance with low bandwidth. An application to the design of flight control laws for a high performance aircraft is presented. The control laws demonstrate good performance by accurately following large angle of attack commands at flight speeds ranging from 53 to 150 m/s. Robustness is verified by including ±20 percent variations in pitching moment derivatives. The reduced bandwidth compared to a fixed-gain, linear design, leads to greatly reduced actuator transients, which should give improved reliability and longer life for the actuators and assaciated structure.

Lateral Jet Control of a Supersonic Missile: Computational and Experimental Comparisons

Abstract: Several three-dimensional, viscous, turbulent Navier‐ Stokes computations have been performed for a missile equipped with and without divert jet thruster for threedifferent wing planforms (having a e xed tail cone guration ) at a nominal e ow Mach number of 3.94, angles of attack ranging from 2 to 25 deg, and jet thrust ratios of one and four. Results are presented to show that the normal force and pitching moment coefe cients for all computed cases are predicted within 5% of the wind-tunnel data. Synthesis of all of the results show low amplie cation factor for all windward jet thruster cases. For this case the upstream favorable pressure zone created by the windward jet is insufe cient to compensate for the massive unfavorable pressure loss on the windward wings/tails due to the jet blockage and jet wraparound effects.

The Real-Time Wall Interference Correction System of the NASA Ames 12-Foot Pressure Wind Tunnel

Abstract: An improved version of the Wall Signature Method was developed to compute wall interference effects in three-dimensional subsonic wind tunnel testing of aircraft models in real-time. The method may be applied to a full-span or a semispan model. A simplified singularity representation of the aircraft model is used. Fuselage, support system, propulsion simulator, and separation wake volume blockage effects are represented by point sources and sinks. Lifting effects are represented by semi-infinite line doublets. The singularity representation of the test article is combined with the measurement of wind tunnel test reference conditions, wall pressure, lift force, thrust force, pitching moment, rolling moment, and pre-computed solutions of the subsonic potential equation to determine first order wall interference corrections. Second order wall interference corrections for pitching and rolling moment coefficient are also determined. A new procedure is presented that estimates a rolling moment coefficient correction for wings with non-symmetric lift distribution. Experimental data obtained during the calibration of the Ames Bipod model support system and during tests of two semispan models mounted on an image plane in the NASA Ames 12 ft. Pressure Wind Tunnel are used to demonstrate the application of the wall interference correction method.

An aerodynamic moment-controlled surface for gust load alleviation on wind turbine rotors

Abstract: A method is described for limiting transient gust loading on horizontal-axis wind turbine rotors. The technique, known as aerodynamic moment control, is implemented by enclosing a pitchable section of the blade in an active control loop, using the external aerodynamic load as feedback variable. The actuator operates within an outer control loop, typically based on electrical power output. The properties of the actuator have been investigated by linear analysis, based on a constant-speed 330-kW wind turbine with active power control, and pitchable blade tips. Two cases were compared, in which the tip actuator was first implemented using position feedback (position control), then subsequently using aerodynamic moment feedback (moment control). The disturbance rejection properties of the overall power controller were found to improve in the latter case. A prototype aerodynamic moment controller has been demonstrated in wind tunnel tests. The controller was configured for an inherently unstable wing section, representing the pitchable tip of a wind turbine blade, at approximately 1/3 full scale. The response to external disturbances was investigated by introducing harmonic perturbations into the upstream airflow. The system successfully demonstrated the principle of aerodynamic moment feedback, although the actuator exhibited somewhat modest gust response characteristics due to the use of velocity feedback to enhance damping. The results of the tests, and the design implications for a full-scale wind turbine, are discussed.

Two-Dimensional Airfoil Performance Degradation Because of Simulated Freezing Drizzle

Abstract: Wind-tunnel tests were conducted to determine the performance degradation of a scaled two-dimensional NACA 23012 airfoil (outboard wing section of the Wyoming King Air 200T ) resulting from 1 ) ice because of various liquid hydrometeor sizes, 2 ) simulated drizzle ice roughness, and 3 ) simulated drizzle ice accretions on a model spar strap. The Wyoming King Air is equipped with a Saunders Fail-Safe Spar Strap that protrudes roughly 6 cm below the wing and extends spanwise from just outside both engine nacelles and may collect ice in large drop regions. The wind-tunnel evaluation facilitated quantifying the effects on aircraft performance degradation because of the King Air spar strap. The airfoil evaluations show that the drizzle drop ice shape and simulated drizzle ice roughness resulted in the highest performance degradation. In general, the ice shapes and simulated freezing drizzle roughness increased proe le drag, reduced angle of attack for maximum lift coefe cient, reduced the maximum lift coefe cient, altered the pitching moment, reduced lift over drag ratio, and marginally changed the lift curve slope. These evaluations also show that the most sensitive surface location on an airfoil is on the suction side between 6 and at least 11% of chord. Ice contaminations in this area are beyond the protective de-icing boots of most aircraft and lead to severe degradations in lift and drag characteristics. In addition, these results suggest that an ice-contaminated spar strap will increase King Air drag by approximately 12% at angles of attack consistent with cruise. Furthermore, any ice that forms on the lower surface of the wing, forward of the spar strap, does not signie cantly increase proe le drag.

Launch Vehicle Trajectories with a Dynamic Pressure Constraint

Abstract: The problem of three-dimensional launch trajectoriesfortheadvanced launch system to minimize fuel orequivalently maximize payload isconsidered. Adynamicpressureinequality constraintis involved dueto itssignie cance on the resulting trajectories and structure of the launch vehicle. Highly accurate solutions and exact switching structures of the optimal trajectories are presented by using a multiple shooting method. The multiple shooting structure is modie ed to satisfy the internal boundary conditions and discontinuities in variables at entry and exit points of the boundary arc. A methodology is also developed to visualize the optimal controls in a geometric sense using hodograph analysis. Nomenclature Ab = cross-sectional area Ae = exit area of a engine a = speed of sound CD = drag coefe cient CL = lift coefe cient Cm = pitching moment coefe cient Cmcg = pitching moment coefe cient about the center of gravity c = number of engines operating D = drag e = eccentricity of orbit f = dynamic equations G = function of states, controls, and time g = Earth’ s gravitational force gs = Earth’ s gravitational force at sea level H = Hamiltonian function h = altitude above mean sea level ha = apogee altitude h f = altitude at tf h p = perigee altitude h1 = density scale height h2 = pressure scale height Isp = specie c impulse J = performance index L = lift l = total core vehicle length lT = distance from the center of mass to the exit plane of engine M = Mach number Maero = aerodynamic pitching moment MT = pitching moment due to thrust m = mass mref = reference mass, sum of the masses of the payload, payload margin, and payload fairing p = atmospheric pressure q = dynamic pressure r = distance from the center of the Earth to the vehicle center of gravity, rs C h ra = radius at apogee from the center of the Earth rp = radius at perigee from the center of the Earth rs = radius of the Earth at sea level S = state inequality constraint Sp = pth time derivative of S

The utilization of low speed facilities in transonic stability of reentry vehicles research - An evaluation

Abstract: The customary practice of utilizing low speed facilities to extract flow characteristics of reentry vehicles in transonic regimes is examined. Using the Apollo capsule as a generic reentry configuration, the similarity between its wake in incompressible and compressible regimes is assessed from static measurements of lift, drag, pitching moment, surface pressure, wake size and periodicity of the wake flow at four nominal Mach numbers ranging from that of essentially incompressible to Mach 0.9. The results to date indicate that, for vehicle orientations of practical interest in which the heat-shield is facing the freestream, sufficient similitude is noted to warrant such an extrapolation at least as an exploratory tool in devising strategies for wake modification. Additional utility of low speed tests in acquiring wake features of an oscillating vehicle is conceptually feasible and discussed. Introduction A reentry capsule is shaped to be blunt so as to survive the intense aerodynamic heating in the hypersonic portion of the flight. However, during the transonic and subsonic phases of the reentry in what might be considered as off-design conditions, the vehicle faces a more pronounced stability and control problem. The dynamic vortex formation typical of the blunt body wake flow at the aforementioned speed regimes produces unsteady loading on the vehicle, and can sometimes thereby affect its handling characteristics severely. It is therefore important to determine the stability characteristics of the vehicle so as to ensure proper and timely deployment of parachutes prior to the onset of instability, as well as having a well documented *NSF-NATO Postdoctoral Fellow, Member AIAA. Associate Professor. 'Doctoral Candidate, Student Member AIAA. Research Engineer. Copyright © 1998 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. description of the flow-field in order to understand the associated fluid mechanics. The importance and difficulties in characterizing the unsteady wake in studying the stability of a reentry vehicle are illustrated and discussed in further details in Ref. 1. In the context of vehicle stability and parachute deployment issues which take place in high subsonic and transonic regimes, the flow-field characterization around and in the near-wake of a reentry capsule is often extracted from low speed tests where both visualization and detailed diagnostics can be carried out more readily. The utilization of a low speed facility to extrapolate flow-field information of a reentry capsule at higher Mach numbers, although attractive in many perspectives and often implemented, is a practice that deserves a methodical examination. Using the Apollo capsule as a representative shape of planetary reentry vehicles, the present paper presents the results of a first series of experiments conducted at the von Karman Institute (VKJ) in assessing this issue of fundamental and practical importance by comparing the global aerodynamic quantities and the flow structures of a non-oscillating capsule in the incompressible, high subsonic and transonic regimes. These findings would lead to a disclosure of the degree of similarity in the flow-field of the various regimes, as well as contributing to the fundamental knowledge of three-dimensional bluff body flow with compressibility effect. In the case of a two-dimensional symmetric or three-dimensional axisymmetric body in free flow at zero lift, the associated global time-averaged wake characteristics are strongly dependent of the drag exerted by the body as this quantity represents an integral of the wake motion. In the event of a lifting body, it is expected that the lift would play its role and manifest itself ultimately in the wake as well. Similarly on a more refined scale, the details of the flow around the body, which could at least be partially revealed by surface pressure measurements, constitute the initial distribution of vorticity shed downstream of the body thereby 1 American Institute of Aeronautics and Astronautics Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc. influencing the microscopies of the wake both on a time-averaged and temporal basis. With the above in mind, the global aerodynamic quantities lift, drag and moment coefficients about the center of gravity were measured at four different Mach numbers. As for the comparison of flow quantities, the investigation included surface pressure distributions as well as optical diagnostics around and in the near-wake of the capsule. In addition, the temporal characteristic of the flapping wake motion was mapped out. All of the subsonic and transonic measurements had been conducted in the VKI-S1 facility whereas their incompressible counterparts were measured in three separate low speed wind tunnels: VKI-L7, L7+ and L2A. The details of the various facilities are found in Ref. 2 Experimental Setup The Apollo Command Module Block I configuration with no protuberances was selected to represent a generic reentry vehicle in this investigation. A drawing of this configuration, with the various dimensions normalized with the maximum diameter (i.e., heat-shield) of the model, is shown in Figure 1. The model is at a =180° as illustrated, with the freestream flow in the direction from right to left. A clockwise rotation corresponds to the direction of decreasing angle of attack. Depending on the particular test, the model was supported either one-sided cantileverly or on both sides with a transverse rod through the theoretical center of gravity location shown in Figure 1. In this mounting arrangement the pitching moment at the model's theoretical center of gravity is therefore measured directly. The additional pertinent details for the various tests are tabulated in Table 1. Fig. 1 A scaled drawing of the Apollo model. Table 1: Summary of Test and Mounting Conditions in Various Facilities L7+

Navier-Stokes Analysis of Subsonic Flowfields over a Missile Configuration

Abstract: Subsonic e owe elds over a missile cone guration are computed at high angles of attack ranging from 15 to 60 deg with the NASA Ames Research Center OVERFLOW Navier‐ Stokes solver. The computed e owe elds are presented intermsof particletraces,helicity contours, and surfacestreamlines. The e owseparation overthemissilebody, the development of leeward-side vortex patterns, and the canard and tail vortices and their interaction are identie ed. The computed normal force and pitching moment coefe cients compare well with the experimental data at an incidence of 15 deg. At higherincidences thenormal force is underpredicted by up to 15%. The computed pitching moments agree qualitatively well with the experimental data, but percentage deviations are substantial in the higher incidence range.

Neural-Network-Based Controller for Nonlinear Aeroelastic System

Abstract: Attenuation of vibratory response is an important design consideration in many aeroelastic systems, and active methods of vibration reduction have been studied extensively in this context. Synthesis of active controllers requires that a good analytical model of the system be available. In those problems in which the aeroelastic system is inherently nonlinear, a robust control scheme is difficult to implement, particularly in the presence of large uncertainties in the model. The use of artificial neural networks, with on-line learning capabilities, is explored as an approach for developing robust control strategies for such problems. In particular, the use of neural networks to mimic the behavior of a linear quadratic Gaussian controller that is applicable to nonlinear systems is presented. The helicopter rotor blade is a classic example of an aeroelastic system in which vibration reduction is an overriding concern, and in which the plant is both nonlinear and contains uncertainties. A simplified two-dimensional representation of this aeroelastic system, consisting of an airfoil with a trailing-edge control flap, is considered as a test case in the present work; both structural and aerodynamic nonlinearities are included in the problem. Nomenclature a = offset of elastic axis from the midchord b = sernichord of airfoil c = distance between midchord and flap hinge cmo = mean pitching-moment coefficient before stall Cms, cPs = static pitching-moment and lift coefficients, respectively Cma, cpa = slopes of linear pitching moment and lift curves, respectively Ia, Is = moments of inertia of airfoil and trailing-edge flap per unit span, respectively Ka, Kh = spring stiffnesses of torsional and translational restraints m = mass of airfoil per unit span Sa, Ss = static moments of airfoil and trailing-edge flap per unit span, respectively 8 = flap angle

Multiple time-scale effects for a pitching 65-deg delta wing

Abstract: Normal force and pitching moment data are presented for a 65° delta wing undergoing pitch motions. Critical states are identified in the static data by apparent discontinuities in the force and moment values or their slopes. These critical states form boundaries, across which the time constants of the aerodynamic responses change. For angles of attack with either vortex breakdown or fully stalled flow present over the planform, significant unsteady aerodynamic effects occur due to lags in the flow-field response. For motions that cross the stall angle-of-attack, a persistent transient effect is shown to occur as the flow field changes to the new equilibrium flow-state. NOMENCLATURE Cm = pitching moment coefficient (nondimensional w.r.t. mean aerodynamic chord, dynamic pressure and planform area) CN = normal force coefficient (nondimensional w.r.t. dynamic pressure and planform area) cr = root chord q = pitch rate (nondimensional w.r.t. freestream speed and root chord) t = time (nondimensional w.r.t. freestream speed and root chord) Tc = time constant (nondimensional w.r.t. freestream speed and root chord) a = aerodynamic angle of attack (corrected to account for pitch rate) a(t0) = angle of attack at motion onset a(tf) = angle of attack at motion cessation 'Aerospace Engineer, Member AIAA. This paper is declared a work of the U. S. Government and is not subject to copyright protection in the United States. Subscript max = maximum

Stall resistance features of lifting-body airplane configurations

Abstract: A generic lifting-body airplane model was tested in a lowspeed wind tunnel. The experimental data indicate that at lower angles of attack the lift-over-drag ratio is comparable to other high-efficiency designs. The high angle of attack aerodynamics of this configuration is influenced by the side-edge vortex system observed above the aft section of the lifting fuselage. Consequently, the total lift of the airplane model continues to increase beyond the angle of wing stall, accompanied by increasing nose-down pitching moments. In principle, such characteristics allow the tailoring of the configuration lift and pitching moment in a manner that lift-loss effects beyond wing stall are minimal. Furthermore, the sharp increase in the nosedown moment of the fuselage can be positioned within the angle-of-attack performance curve such that airplane stall can become unreachable (and the configuration becomes stall resistant). The present study investigates some of the geometrical parameters influencing these aerodynamic effects so that such inherent stall-resistant characteristics can be developed early in the design stage of a lifting-body airplane configuration.

Extrapolation From Wind Tunnel to Flight: Shuttle Orbiter Aerodynamics

Abstract: The paper reviews a combined numerical and experimental activity on the Shuttle Orbiter, first performed at NASA Langley within the Orbiter Experiment (OEX) and subsequently at ESA, as part of the AGARD FDP WG 18 activities. The study at Langley was undertaken to resolve the pitch up anomaly observed during the entry of the first flight of the Shuttle Orbiter. The present paper will focus on real gas effects on aerodynamics and not on heating. The facilities used at NASA Langley were the 15-in. Mach 6, the 20-in, Mach 6, the 31-in. Mach 10 and the 20-in. Mach 6 CF4 facility. The paper focuses on the high Mach, high altitude portion of the first entry of the Shuttle where the vehicle exhibited a nose-up pitching moment relative to pre-flight prediction of (Delta C(sub m)) = 0.03. In order to study the relative contribution of compressibility, viscous interaction and real gas effects on basic body pitching moment and flap efficiency, an experimental study was undertaken to examine the effects of Mach, Reynolds and ratio of specific heats at NASA. At high Mach, a decrease of gamma occurs in the shock layer due to high temperature effects. The primary effect of this lower specific heat ratio is a decrease of the pressure on the aft windward expansion surface of the Orbiter causing the nose-up pitching moment. Testing in the heavy gas, Mach 6 CF4 tunnel, gave a good simulation of high temperature effects. The facilities used at ESA were the lm Mach 10 at ONERA Modane, the 0.7 m hot shot F4 at ONERA Le Fauga and the 0.88 m piston driven shock tube HEG at DLR Goettingen. Encouraging good force measurements were obtained in the F4 facility on the Orbiter configuration. Testing of the same model in the perfect gas Mach 10 S4 Modane facility was performed so as to have "reference" conditions. When one compares the F4 and S4 test results, the data suggests that the Orbiter "pitch up" is due to real gas effects. In addition, pressure measurements, performed on the aft portion of the windward side of the Halis configuration in HEG and F4, confirm that the pitch up is mainly attributed to a reduction of pressure due to a local decrease in gamma.

Analysis of free-flight data for the CAN4 hypersonic research projectile

Abstract: Free flight tests were conducted in the DREV aeroballistic range on a cone-cylinder-flare configuration to study the aerodynamic characteristics. The projectiles were fired from a 110 mm smooth bore gun at muzzle velocities ranging between 1650m/sec (Mach 4.8) and 2150 m/sec (Mach 6.3). The projectiles were tested with and without a base cavity. All the main aerodynamic coefficients were determined, including the non-linear cubic pitching moment term resulting in an accurate data base for CFD validation. Aerodynamic predictions from three semiempirical models were compared with the experimental data.

Computation of the Aerodynamic Characteristics of a Subsonic Transport

Abstract: An advanced low-order panel method, VSAERO, has been applied to a full-configuration transport aircraft. Wing pressure distributions, as well as lift, drag, and pitching moment coefficients, calculated for the full configuration are compared with experimental data. In addition, the effect of selected vehicle components on aircraft forces and moments was investigated. The results for the complete configuration were compared with those from several partial model configurations. It is shown that the wing-surface pressure distributions and pitching moment coefficient are predicted well by VSAERO. The lift coefficient is consistently overpredicted by the code, as expected for inviscid calculations

Hinge Moment Coefficient Prediction for Nose-Mounted Canard Controls at Supersonic Speeds

Abstract: Theprediction of nose-mounted canard hinge momentsforsupersonicrockets poses a uniqueproblem for which the semiempirical methods utilized in rapid aerodynamic prediction codes may not provide sufe cient accuracy for preliminary design of the control actuation system. Although providing accurate predictions of canard normal force, such codes generally cannot predict hinge moments effectively due to both their empirical nature and their inability to address the local e owe eld conditions on the rocket nose. It is shown that the local e owe eld properties must be characterized to accurately determine the longitudinal center of pressure of the canards. A theoretical approachhasbeendevelopedtopredictnormalforcecoefe cient,longitudinalcenterofpressure,andhingemoment coefe cient for nose-mounted canards. The method is based on shock-expansion theory and airfoil strip theory and accounts for local e owe eld properties, tip pressure losses, and body carryover effects. In contrast to predictions by an industry standard semiempirical code, the new theoretical method consistently estimates canard longitudinal center of pressure with a higher degree of accuracy, resulting in good agreement with experimental hinge moment data for nose-mounted canards at Mach 1.25‐ 2.00.

Numerical Representation of Pitching and Nonpitching Airfoils Undergoing Linear Acceleration

Abstract: A numerical approach is presented for computing unsteady aerodynamic effects on an airfoil that is undergoing independent pitching and/or plunging motion while attached to an accelerating body. The acceleration of the body may be at any angle to the horizontal axis. Grid speed terms are incorporated into a e rst-order e nite volume representation of the unsteady Euler equations, along with the appropriate acceleration terms for the boundary conditions. Unstructured grid methodology is utilized, along with a moving grid algorithm, to model the pitching/plunging of the airfoil within the grid. A NACA 0012 airfoil is considered for all work. Comparisons are made with nonaccelerating numerical and wind-tunnel data to demonstrate the validity of the methodology. Results are then presented for pitching and nonpitching airfoil, accelerated body cases to demonstrate the effects of the linear acceleration on the unsteady aerodynamics of the airfoil. The quantitative effect of the body acceleration on the aerodynamic coefe cients is seen to be a function of the type of motion (pitching/nonpitching ) imposed upon the airfoil.

The Aerodynamic Forces On Misaligned Platoons

Abstract: This report summarizes wind tunnel experimental measurementson the aerodynamic interaction between members of misaligned platoons. Experiments are conducted at the University of Southern California's Dryden Wind Tunnel Facility. All experiments are made using 1/8 scale models of 1991 Chevy Lumina minivan. Models are placed above a ground plane with a porous surface, through which slight suction is applied to remove the boundary layer. Refurbishing of the ground plane surface, and its repositioning to a 1 degree angle of attack produce significantly improved air flow through the test section. Automatization of the testing procedures allow measurements of drag, side force and yawing moment with extremely fine position resolution. The measured quantities are presented in the form of coefficient ratios by ratioing the forces and moment with the value of drag experienced by a vehicle in isolation. The results of two separate experiments are presented in this report. First, aerodynamic forces on misaligned three-vehicle platoons are presented for all possible platoon configurations resulting from a longitudinal separation range of 0 to 0.72 vehicle lengths and a lateral displacement range of 0 to 1.1 vehicle widths for the middle vehicle. Results are presented in the form of color maps of the drag, side force and yawing moment coefficient ratios for each individual vehicle in the platoon. Experimental results from a complete set of symmetric configurations and five sets of non-symmetric configurations, associated with five fixed separations between the leading vehicle and the trailing vehicle, are presented in detail. A second experiment consists of a detailed investigation of aerodynamic forces on a two-vehicle platoon in back-to-back geometry. Following previous observations, the present experiment investigates a two-fold increase in drag force occurring at specific separations between the two vehicles. Color maps for the drag, side force and yawing moment coefficient ratios document the presence of a hysteresis loop-the drag on the leading vehicle as separation increases is different from drag as separation narrows. It is argued that the resonance with hysteresis represents a matching between the wavelength of turbulent flow structure and the spacing itself. A dimensional analysis relates the drag increase phenomenon to longitudinal separations between different types of vehicles.