Showing papers in "Journal of Aircraft in 1988"

Progress in analysis and prediction of dynamic stall

Abstract: Nomenclature c = chord CD =drag coefficient, D/qA CL =lift coefficient, L/qa CP = pressure coefficient, P-PW /q CM = pitching-moment coefficient, M/qA D =drag, Ib k = reduced frequency, wc/2U L =lift, Ib M =Mach number; pitching moment, ft-lb P = pressure, lb/in q = dynamic pressure, 1 /2p U R, Re, Re, Rn = Reynolds number t time U,V = freestream velocity, ft/s x = distance a = angle of attack, deg f =nondimensional chord length, x/c $tr transition location co = rotational frequency, rad/s p = density, lb/ft A = sweep angle, deg

Flight dynamics of aeroelastic vehicles

Abstract: The nonlinear equations of motion for an elastic airplane are developed from first principles. Lagrange's equation and the Principle of Virtual Work are used to generate the equations of motion, and aerodynamic strip theory is then employed to obtain closed-form integral expressions for the generalized forces. The inertial coupling is minimized by appropriate choice of the body-reference axes and by making use of free vibration modes of the body. The mean axes conditions are discussed, a form that is useful for direct application is developed, and the rigid-body degrees of freedom governed by these equations are defined relative to this body-reference axis. In addition, particular attention is paid to the simplifying assumptions used during the development of the equations of motion. Since closed-form, analytic expressions are obtained for the generalized aerodynamic forces, insight can be gained into the effects of parameter variations that is not easily obtained from numerical models. An example is also presented in which the modeling method is applied to a generic elastic aircraft, and the model is used to parametrically address the effects of flexibility. The importance of residualizing elastic modes in forming an equivalent rigid model is illustrated, but as vehicle flexibility is increased, even modal residualization is shown to yield a poor model.

Validation of approximate indicial aerodynamic functions for two-dimensional subsonic flow

Abstract: Approximations for two-dimensional indicial (step) aerodynamic responses due to angle of attack and pitch rate are obtained and generalized to account for compressibility effects up to a Mach number of 0.8. Using the Laplace transform method, these indicial functions are manipulated to produce explicit solutions for idealized harmonic forcings. These explicit solutions are subsequently compared with experimentally obtained pitch and plunge aerodynamic data in the reduced frequency domain. The results of this comparison are used to relate back and substantiate the generalization of the compressible indicial lift and moment functions.

Boundary-layer measurements on an airfoil at low Reynolds numbers

Abstract: Etude experimentale des caracteristiques des bulles de decollement de transition par anemometrie laser et a fil chaud

Airfoil dynamic stall at constant pitch rate and high Reynolds number

Abstract: An experiment has been performed to study the aerodynamics of dynamic stall penetration at constant pitch rate and high Reynolds number, in an attempt to model more accurately conditions during aircraft poststall maneuvers and during helicopter high-speed forward flight. An airfoil was oscillated at pitch rates, A = ac/2U between 0.001 and 0.020, Mach numbers between 0.2 and 0.4, and Reynolds numbers between 2-4 x 10. Surface pressures were measured using 72 miniature transducers, and the locations of transition and separation were determined using 8 surface hot-film gages. The results demonstrate the influence of the leading-edge vorticity on the unsteady aerodynamic response during and after stall. The vortex is strengthened by increasing the pitch rate and is weakened by increasing the Mach number and by starting the motion close to the steady-state stall angle. A periodic pressure oscillation occurred after stall at high pitch angle and moderate Reynolds number; the oscillation frequency was close to that predicted for a von Karman vortex street. A small supersonic zone near the leading edge at M = 0.4 was found to reduce significantly the peak suction pressures and the unsteady increments to the airloads. These results provide the first known data base of constant-pitch-rate aerodynamic information at realistic combinations of Reynolds and Mach numbers.

Control of Vortical Lift on Delta Wings by Tangential Leading-Edge Blowing

Abstract: An experiment has been performed to examine the feasibility of vortex control by tangential mass injection at the leading edge of a 60 deg delta wing. The initial results indicate that direct control of the primary separation allows significant control of the vortex flow up to angles of attack of 60 deg. At lower angles of attack, the vortical flow may be removed entirely from the surface of the wing, recovering the fully attached flow case. The effects of the mass injection have been shown to be decoupled from the geometric angle of attack, allowing the possibility for controlling lift without changing attitude. Nomenclature b =wing semispan c =wing root chord Cg' = spanwise sectional rolling moment coefficient CL' = wing rolling moment coefficient Cn = spanwise sectional normal force coefficient CN = wing normal force coefficient Cp = pressure coefficient CM = bio wing momentum coefficient CM * = crossflow blowing momentum coefficient h = slot height Vj =jet velocity Fw = freestream velocity ae = effective angle of attack oig = geometric angle of attack e = wing semi apex angle

Modern wing flutter analysis by computational fluid dynamics methods

Abstract: The application and assessment of the recently developed CAP-TSD transonic small-disturbance code for flutter prediction is described. The CAP-TSD code has been developed for aeroelastic analysis of complete aircraft configurations and was previously applied to the calculation of steady and unsteady pressures with favorable results. Generalized aerodynamic forces and flutter characteristics are calculated and compared with linear theory results and with experimental data for a 45 deg sweptback wing. These results are in good agreement with the experimental flutter data which is the first step toward validating CAP-TSD for general transonic aeroelastic applications. The paper presents these results and comparisons along with general remarks regarding modern wing flutter analysis by computational fluid dynamics methods.

Viscous transonic airfoil workshop compendium of results

Abstract: Results from the Viscous Transonic Airfoil Workshop held at the AIAA 25th Aerospace Sciences Meeting at Reno, NV in January 1987, are compared with each other and with experimental data. Test cases used in this workshop include attached and separated transonic flows for three different airfoils: the NACA 0012 airfoil, the RAE 2822 airfoil, and the Jones airfoil. A total of 23 sets of numerical results from 15 different author groups are included. The numerical methods used vary widely and include: 16 Navier-Stokes methods, 2 Euler/boundary-layer methods, and 5 full-potential/boundary-layer methods. The results indicate a high degree of sophistication among the numerical methods with generally good agreement between the various computed and experimental results for attached or moderately-separated cases. The agreement for cases with larger separation is only fair and suggests additional work is required in this area.

Coupled flow, thermal, and structural analysis of aerodynamically heated panels

Abstract: A finite element approach to coupling flow, thermal and structural analyses of aerodynamically heated panels is presented. The Navier-Stokes equations for laminar compressible flow are solved together with the energy equation and quasi-static structural equations of the panel. Interactions between the flow, panel heat transfer and deformations are studied for thin stainless steel panels aerodynamically heated by Mach 6.6 flow.

Low-Reynolds-number airfoil design for the M.I.T. Daedalus prototype- A case study

Abstract: The rationale used for the aerodynamic wing design of the prototype long-range human-powered aircraft Light Eagle is presented. Three different airfoils, designed for chord Reynolds numbers of 500,000,375,000, and 250,000 were used across the wingspan. The airfoil design rationale centered on minimizing the losses in the transitional separation bubbles typically occurring on airfoils at Reynolds numbers of less than 1 million. Structural and manufacturing constraints were also a consideration in the airfoil design, although to a lesser extent. Airfoil performance prediction during the design process was done entirely through numerical simulation. The numerical model employs the Euler equations to represent the inviscid flow, and an integral boundary-layer formulation to represent the viscous flow. Strong viscous-inviscid coupling and an amplification transition criterion included in the overall equation system permit calculation of transitional separation bubbles and their associated losses. Flow visualization tests performed on the Light Eagle at various lift coefficients in towed flight revealed transition occurring very near the intended position on the wing surface except within a few chords of the tip, where the flow appeared to be turbulent over most of the upper surface. Total drag aircraft polars obtained from the measured aircraft energy time history in glide contained too much scatter to be used as quantitative test data but did reproduce the basic trends of the calculations, including maximum lift coefficient levels.

Moving wall effects in unsteady flow

Abstract: L'effet de paroi mobile est present aussi bien dans les ecoulements bidimensionnels que tridimensionnels. Influence sur le decollement des couches limites laminaires et turbulentes et sur la transition des couches limites

Integrated aerodynamic/structural design of a sailplane wing

Abstract: Using lifting-line theory and beam analysis, the geometry (planiform and twist) and composite material structural sizes (skin thickness, spar cap, and web thickness) were designed for a sailplane wing, subject to both structural and aerodynamic constraints. For all elements, the integrated design (simultaneously designing the aerodynamics and the structure) was superior in terms of performance and weight to the sequential design (where the aerodynamic geometry is designed to maximize the performance, following which a structural/aeroelastic design minimizes the weight). Integrated designs produced less rigid, higher aspect ratio wings with favorable aerodynamic/structural interactions.

Response of airline pilots to variations in flight simulator motion algorithms

Abstract: The use of physical motion in flight simulation is still a much debated topic. This paper investigates the more narrow issue of its application in commercial jet transport simulators. We have attempted to quantify the perceptions of airline pilots about the quality of motion possible when a number of different motion-drive algorithms are tested on a simulator employing a state-of-the-art six-degrees-of-freedom motion-base. Four broad categories of algorithms were tested: classical washout, optimal control, coordinated adaptive, and no-motion. It was found that although there was little impact of algorithm type on performance arid control activity, there was a definite effect on how the pilots perceived the simulation environment. Based on these findings, it appears that the coordinated adaptive algorithm is generally preferred by the pilots over the other algorithms tested, there was almost unanimous dislike of the no-motion case.

Efficient algorithm for solution of the unsteady transonic small-disturbance equation

Abstract: A time accurate approximation factorization (AF) algorithm is formulated for solution of the three-dimensional unsteady transonic small-disturbance equation. The AF algorithm consists of a time linearization procedure coupled with a Newton iteration technique. Superior stability characteristics of the new algorithm are demonstrated through applications to steady and oscillatory flows at subsonic and supersonic freestream conditions for an F-5 fighter wing. For steady flow calculations, the size of the time step is cycled to achieve rapid convergence. For unsteady flow calculations, the AF algorithm is sufficiently robust to allow the step size to be selected based on accuracy rather than on stability considerations. Therefore, accurate solutions are obtained in only several hundred time steps yielding a significant computational cost savings when compared to alternative methods.

Unsteady low-speed aerodynamic model for complete aircraft configurations

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

F-111 generic weapons bay acoustic environment

Abstract: A wind tunnel test was conducted on a generic weapons bay model to define the acoustic environment in the bay and how it is affected by various missiles, suppressors, and flow conditions. Two different depth bays were tested. Three different missile designs were installed in the bay with the capability of putting one, two, or three in the bay at a time. One of the missiles was tested at three vertical positions to measure the effect it has while being deployed. Two leading-edge sawtooth suppressors were evaluated along with a slanted rear bulkhead. Overall levels as high as 163 dB were measured in the bay. Narrowband tones were strongly excited. The sawtooth spoilers were partially effective in suppressing the tones. In some cases the levels were amplified. The slanted aft bulkhead was very effective in suppressing the levels. The missiles generally reduced the acoustic levels in the bay, with the most reduction when a missile was partially out of the bay interacting with the shear layer. The measured levels were of high enough intensity to result in severe fatigue problems for sensitive missiles and components.

Experimental Study of Effects of Forebody Geometry on High Angle-of-Attack Stability

Abstract: A series of low-speed wind-tunnel tests on a generic airplane model with a cylindrical fuselage are made to investigate the effects of forebody shape and fineness ratio, and fuselage/wing proximity on static and dynamic lateral/directional stability. During the stability investigation ten forebodies were tested including three different cross-sectional shapes with fineness ratios of 2,3, and 4. In addition, the wing was tested at two longitudinal positions to provide a substantial variation in forebody/wing proximity. Conventional force tests were conducted to determine static stability characteristics, and single-degree-of-freedom free-to-roll tests were conducted to study the wing rock characteristics of the model with the various forebodies. Flow visualization data were obtained to aid in the analysis of the complex flow phenomena involved. The results show that the forebody cross-sectional shape and fineness ratio and forebody/wing proximity can strongly affect both static and dynamic (roll) stability at high angles of attack. These characteristics result from the impact of these factors on forebody vortex development, the behavior of the vortices in sideslip, and their interaction with the wing flowfield.

Numerical simulation of wings in steady and unsteady ground effects

Abstract: A numerical simulation of steady and unsteady ground effects is developed. The simulation is based on the general unsteady vortex-lattice method and is not restricted by planform, angle of attack, sink rate, dihedral angle, twist, cross wind, etc., as long as stall does not occur. The present computed results are generally in close agreement with limited exact solutions and experimental data. The present results show the influences of various parameters on the aerodynamic coefficients for both steady and unsteady flows. Generally, the aerodynamic coefficients increase with proximity to the ground, the greater the sink rates the greater the increases, increasing the aspect ratio increases both the steady and unsteady ground effects for both rectangular and delta planforms. The steady ground effect increases the rolling moment and the side force. The present results serve to demonstrate the potential of the present approach.

Combined translation/pitch motion: a new airfoil dynamic stall simulation

Abstract: The effects of simultaneous velocity and incidence fluctuations on the 2-D aerodynamic behaviour of a NACA 0012 airfoil are investigated in this paper. A new mechanical system allows driving the airfoil in pitching and in fore and aft motions, as well as in a simultaneous combination of these two basic unsteady motions. In response to the simultaneous velocity and incidence variations, the time-dependent lift and drag fluctuations are measured for increasing values of the reduced frequency and amplitude parameters, including dynamic stall conditions. Complementary information on the dynamic stall occurring in combined motion is provided by skin friction and pressure measurements along the airfoil surface.

Research on aircraft/vortex-wake interactions to determine acceptable level of wake intensity

Abstract: An evaluation of the literature on large aircraft wake-vortex encounters in flight and in flight simulators has furnished an estimate of the level to which the vortex-induced rolling moments must be reduced in order to be perceived as nonhazardous at a 2-n.mi. separation distance. The criteria are based on the ratio of the vortex-induced acceleration in roll to the aileron-induced roll acceleration. A wake is acceptably alleviated if the ratio of vortex-to-aileron rolling moments is less than about 0.5. When a satisfactory alleviation scheme is identified, the alleviated vortex structure should be inserted into a simulator to ascertain whether the maximum bank angles induced are within tolerable limits.

Stability of airplanes in ground effect

Abstract: In ground effect, the aerodynamic coefficients vary considerably with height. Longitudinal motion in ground effect reveals modes and stability conditions remarkably different from out-of-ground characteristics, a fact that has a large impact on a successful concept of "wing-in-ground" vehicles. Using calculated aerodynamic coefficients and a linearized approach, static and dynamic stabilities were studied for a chosen more conventional configuration. In addition, flare maneuvers were studied for an aircraft equipped with a simple two-term controller. It is shown that the performance of the controller is considerably influenced by the ground effect. The interrelation between height-dependent aerodynamic coefficients and longitudinal stability is explained and shows particular importance in flare maneuvers.

Multilevel Decomposition Approach to the Preliminary Sizing of a Transport Aircraft Wing

Abstract: A multilevel/multidisciplinary optimization scheme for sizing an aircraft wing structure is described. A methodology using nonlinear programming in application to a very large engineering problem is presented. This capability is due to the decomposition approach. Over 1300 design variables are considered for this nonlinear optimization task. In addition, a mathematical link is established coupling the detail of structural sizing to the overall system performance objective, such as fuel consumption. The scheme is implemented as a three level system analyzing aircraft mission performance at the top level, the total aircraft structure as the middle level, and individual stiffened wing skin cover panels at the bottom level. Numerical show effectiveness of the method and its good convergence characteristics.

Determination of the aerodynamic characteristics of the mission adaptive wing

Abstract: The Advanced Fighter Technology Integration F-lll flight research program was conducted to develop and demonstrate the potential technology enhancements of the mission adaptive wing. This wing incorporated smooth contour, variable-camber leadingand trailing-edge surfaces that could modify wing contour in flight by means of an internal linkage system and flexible skins. Extensive wind-tunnel and flight test data were gathered during the course of the program to define the aerodynamic performance benefits attributed to the mission adaptive wing. Full-scale aerodynamic characteristics and predicted performance were initially based on a wind-tunnel data base. Flight testing was conducted to determine lift, drag, buffet, and wing upper and lower surface pressures. The flight test data served to verify the wind-tunnel predictions and to provide a data base for follow-on analyses.

General formulation for the aeroelastic divergence of composite swept-forward wing structures

Abstract: This paper formulates a simple algorithm that allows for the determination, in a closed form, of the divergence instability of advanced composite swept (back and forward) wing structures The warping restraint effect is incorporated into the analysis and its influence on the associated static instability condition is put into evidence In this sense, it is shown that, in contrast to the case of conventional metallic wings where the warping restraint effect has a stabilizing influence only (which is strongly manifested in the case of small-aspect-ratio wings and declines for moderate AR wings), its influence becomes more complex in the case of anisotropic composite wing structures The principal goal of this study is the divergence behavior of swept composite-type wings to the warping restraint effect being included in the analysis The numerical examples illustrate the complex role played by the warping restraint effect on the divergence instability of composite wings

Integrated aeroservoelastic tailoring of lifting surfaces

Abstract: An approach to the integration of two design activities, structural design and active control design, for a highly idealized aeroservoelastic system is presented. The particular design goal for this study is the maximization of the stable airspeed envelope of an idealized model of an aeroservoelastic system through rational and systematic variation of structural and control design parameters. The steady-state linear quadratic regulator is used to model the control subsystem; the structural subsystem is assigned characteristic design parameters, such as shear center position. The application of the procedure described here produces optimally controlled structures with stability characteristics superior to those of open-loop and initial closed-loop designs.

Wind-tunnel investigation of wing-in-ground effects

Abstract: Wing-in-Ground effects from a wind-tunnel study of a NACA 4415-profile wing model with an aspect ratio of 2.33 are described. The wing model contains a 20%-chord, full-span, adjustable flap and removable end and enter plates. Ground boards are used in the wind tunnel to simulate the ground. In this study, the ground effects are expressed as variations to the aerodynamic coefficients (lift and drag) and lift-to-drag ratio. The ground effects are described in terms of angle of attack, flap angle, wing height above ground and use and size of end and center plates.

Fluid dynamics of airfoils with moving surface boundary-layer control

Abstract: Effectiveness of the moving surface boundary-layer control is assessed with reference to a symmetrical Joukowsky airfoil modified with a leading-edge rotating cylinder. Results of the test program and the numerical models suggest the following: 1) The surface singularity method is essential in modeling the complicated flow. With the inclusion of the boundary-layer correction scheme, it becomes an effective tool for obtaining useful information concerning moving surface boundary-layer control. The predicted pressure distributions are in good agreement with experiment almost up to the point of complete separation from the airfoil surface, except in the separation region, where the prediction of separated boundary layers with flow reversal would require the solution of the full Navier-Stokes equations. 2) The concept of moving surface boundary-layer control appears quite promising. The tests showed a significant improvement in maximum lift and stall characteristics. With cylinder rotation, the flow never separated completely from the upper surface for angles of attack as high as 48 deg. The higher rates of rotation (UC/U>1, Uc = cylinder surface velocity, U = freestream velocity) promoted reattachment of the partially separated flow, giving an increase in lift coefficient by as much as 150% for Ue/U = 4.

Visualization and Measurement of the Tip Vortex Core of a Rotor Blade in Hover

Abstract: Detailed measurements with a laser Doppler velocimeter (LDV) have been performed in the tip region and in the tip vortex core of a single-bladed model rotor in hover. The testing was conducted at a rotor tip speed of 32 m/s, a Reynolds number of 269,000, and at two values of the rotor thrust coefficient, 0.0022 and 0.0057. Strobed laser sheet flow visualization was used to verify the steadiness of the tip vortex trajectory in the near wake and quantify the vortex trajectory to guide LDV surveys of the vortex core. A remotely aligned off-axis receiving optics system enabled measurement of vortex core velocity profiles at large focal lengths. The core self-induced velocity components extracted from these data are presented. The data exhibit evidence of secondary structure even inside the rotational core of the vortex, the axial velocity profile along the core has been extracted and presented in the wake of a spinning rotor. It is seen that the tip vortex of a rotating blade differs considerably in structure from a fixed-wing vortex.

Improved exponential time series approximation of unsteady aerodynamic operators

Abstract: A procedure has been developed for approximating unsteady aerodynamic operators as truncated exponential series in the time domain. The approximation is accomplished using a least squares minimization fit to aerodynamic data in the frequency domain. The procedure extends previous methods by including the pole locations as unknown parameters of the least squares minimization. In addition, the error associated with both the real and imaginary parts of the Fourier transform of the approximation is minimized. A Newton-Raphson search algorithm is used to find the minimum of the weighted square error in the parameter space of the approximation while constraining the poles to be in the left half-plane. By freely varying the poles of the approximation during the numerical least squares minimization, the representation of the unsteady aerodynamics is improved and is comparable to existing higher-order Fade approximations. Hence, the method offers the aeroelastic designer a more direct method of finding approximate aerodynamic states. However, because the minima of the square error in the cost function found are not necessarily global and depend on the number of poles in the approximation, the initial trial minimum, and the details of the cost minimization algorithm, the poles found in the search do not necessarily correspond to the theoretical poles of the aerodynamic transfer function. Example exponential time series approximations of the Theodorsen function are presented and compared with a Fade approximation and other exponential time series approximations. Nomenclature A = curvature matrix of the cost function J A =true aerodynamic impulse response spectrum A' — approximate aerodynamic impulse response spectrum a - coefficient defined in Eq. (14) an = coefficients of the approximation bn - pole locations of the approximation c = cost minimization step length C = Theodorsen circulation function d - cost minimization search direction vector F = real part of A F' = real part of A' g = gradient of the cost function / G = imaginary part of A G' = imaginary part of A' J = weighted square error cost function k = reduced frequency ^max - frequency corresponding to the peak in IGI M = number of data points to be fitted N — number of terms in the series t = npndimensional time s — Laplace domain variable x — parameter vector of the coefficients and poles am = weighting factors for the real part of the approximation error 0m = weighting factors for the imaginary part of the approximation error = true step response ' = approximate step response