Showing papers in "Journal of Aircraft in 1993"

Application of nonlinear systems theory to transonic unsteady aerodynamic responses

Abstract: A methodology is presented for using the Volterra-Wiener theory of nonlinear systems in aeroservoelastic (ASE) analyses and design. The theory is applied to the development of nonlinear aerodynamic response models that can be defined in state-space form and are, therefore, appropriate for use in modern control theory. The Volterra-Wiener theory relies on the identification of nonlinear kernels that can be used to predict the response of a nonlinear system due to an arbitrary input. A numerical kernel identification technique, based on unit impulse responses, is presented and applied to a simple bilinear, single-input-single-output (SISO) system. The linear kernel (unit impulse response) and the nonlinear second-order kernel of the system are numerically identified and compared with the exact, analytically defined linear and second-order kernels. This kernel identification technique is then applied to the computational aeroelasticity program-transonic small disturbance (CAP-TSD) code for identification of the linear and second-order kernels of a NACA64A010 rectangular wing undergoing pitch at M = 0.5, M = 0.85 (transonic), and M = 0.93 (transonic). Results presented demonstrate the feasibility of this approach for use with nonlinear, unsteady aerodynamic responses.

Nonlinear aerodynamic modeling using multivariate orthogonal functions

Abstract: A technique was developed for global modeling of nonlinear aerodynamic coefficients using multivariate orthogonal functions based on the data. Each orthogonal function retained in the model was decomposed into an expansion of ordinary polynomials in the independent variables, so that the final model could be interpreted as selectively retained terms from a multivariable power series expansion. A predicted squared-error metric was used to determine the orthogonal functions to be retained in the model; analytical derivatives were easily computed. The approach was demonstrated on the Z-body axis aerodynamic force coefficient (Cz) wind tunnel data for an F-18 research vehicle which came from a tabular wind tunnel and covered the entire subsonic flight envelope. For a realistic case, the analytical model predicted experimental values of Cz very well. The modeling technique is shown to be capable of generating a compact, global analytical representation of nonlinear aerodynamics. The polynomial model has good predictive capability, global validity, and analytical differentiability.

Thermoacoustic loads and fatigue of hypersonic vehicle skin panels

Abstract: A thermo-vibro-acoustic analysis of skin panels for airbreathing hypersonic vehicles is made for a generic trajectory and vehicle design. Aerothermal analysis shows that impingement of the bow shock wave on the vehicle produces fluctuating pressures, and local heat fluxes greatly exceed those due to the attached turbulent boundary. Thermal analysis of carbon-carbon skin panels shows that maximum temperatures will exceed 2700°F (1480°C) at the top of the ascent trajectory. Engine acoustic analysis indicates that sound levels will exceed 170 dB. As a result, loads due to engine acoustics and shock impingement dominate the design of many transatmospheric vehicle skin panels.

Unsteady Subsonic Aerodynamics for Bodies and Wings with External Stores Including Wake Effect

Abstract: An unsteady subsonic method has been developed for aerodynamic computations of any elastic or rigid aricraft with external stores. The method consists of two integral parts: a body surface panel method (SPM) and a constant-pressure lifting surface method, which is the subsonic parallel of the HGM (or the ZONA51 code) for unsteady supersonics. The body considered can be flat-based or close-ended and its geometry input is amenable to any given fuselage or store configuration. The present method is considered an advancement over the past development at least in three aspects: (1) correct unsteady boundary condition on body, (2) a new wake model to account for the body/wake effect and (3) improved accuracy in wingbody interference. Various AGARD iifting surfaces, truncated blunt and pointed bodies and a number of NLR wing-storetiptank combinations were studied for method validation. The present method has shown substantial improvement in the pressures, stability derivatives and airloads on these configurations. For all cases considered, the present results, with or without the wake model, have consistently shown closest agreement with all measured data among existing methods. Therefore, we believe that an accurate and effective method is finally at hand for subsonic aeroelastic applications.

Sensitivity-Based Scaling for Approximating Structural Response

Abstract: This article presents a sensitivity-based linearly varying scale factor used to reconcile results from simple and refined models for analysis of the same structure. The improved accuracy of the linear scale factor compared to a constant scale factor, as well as the commonly used tangent approximation, is demonstrated. A wing-box structure is used as an example, with displacements, stresses, and frequencies correlated. The linear scale factor could permit the use of a simplified model in an optimization procedure during preliminary design to approximate the response given by a refined model over a considerable range of design changes. HE design optimization of an engineering system typi- cally requires hundreds of analyses of that system. The use of approximation s to the objective function and con- straints during portions of the design process is quite com- mon,1 because of the high computational cost of these detailed analyses. Such design approximations can be divided into two classes. First there are local, derivative-based approximations such as the linear approximation based on a Taylor-series expansion about a design point. These approximations are typically based on an accurate model to obtain the system response and its derivatives. Second, there are global ap- proximations that try to capture the behavior of the objective function or constraints over the entire design domain. Such approximations can be based on a response surface which is constructed on the basis of many analyses.2 However, global approximations are often based on a simplified theory, a coarser model, or both.3 Such global approximations are referred to here as simple-model approximations. Local approximations are typically very accurate near the design point where they are generated, but since they are based on an extrapolation procedure their accuracy can deteriorate catastrophically at a distance. Simple-model approximations are intended to cap- ture the physics of the problem at some lower, but acceptable level of accuracy over the entire design domain. Therefore, at a particular design point the simple-model approximations are generally less accurate than local approximations, but on the other hand, they typically do not experience the cata-

Plume and Wake Dynamics, Mixing, and Chemistry Behind a High Speed Civil Transport Aircraft

Abstract: The environmental perturbations caused by the exhaust of a high speed civil transport (HSCT) depend on the deposition altitude and the amount and composition of the emissions. The chemical evolution and the mixing and vortical motion of the exhaust need to be analyzed to track the exhaust and its speciation as the emissions are mixed to atmospheric scales. Elements of an analytic model of the wake dynamical processes are being developed to account for the roll-up of the trailing vorticity, its breakup due to the Crow instability, and the subsequent evolution and motion of the reconnected vorticity. The concentrated vorticity is observed to wrap up the buoyant exhaust and suppress its continued mixing and dilution. The chemical kinetics of the important pollutant species will be followed throughout the plume and wake. Initial plume mixing and chemistry are calculated using an existing plume model, standard plume flowfield (SPF), with additional H/C/O, OH/SO2, and NOj, chemistry and equilibrium H2O condensation included. The species tracked include those that could be heterogeneously reactive on the surfaces of the condensed solid water (ice) particles when condensation occurs, and those capable of reacting with exhaust soot particle surfaces to form active contrail and/or cloud condensation nuclei (ccn).

Vortex generators used to control laminar separation bubbles

Abstract: This study examines the effect of various forms of vortex generators on the laminar separation bubble of a two-dimensional low Reynolds number Liebeck LA2573A airfoil. The objective of this research was to determine the effects that different generator sizes and spacings have upon the separation bubble and the drag. Windtunnel measurements were made on several generator configurations at Reynolds numbers ranging from 200,000 to 600,000 at angles of attack less than the stall angle where the separation bubble can provide a significant contribution to the airfoil drag. The vortex generators used were constructed small enough to be contained completely within the laminar boundary layer. Wind-tunnel data included airfoil drag and mean and fluctuating velocity measurements in the laminar and turbulent boundary layers. Results have shown that the use of vortex generators provides a measurable decrease in airfoil drag at the lower range of Reynolds numbers tested. At the airfoil's design condition and Reynolds number of 235,000, the submerged vortex generators were shown to decrease the airfoil drag by a maximum of 38% at C/ = 0.572.

Optimization of Joined-Wing Aircraft

Abstract: The joined wing is an innovative aircraft configuration with a rear wing that has its root attached near the top of the vertical tail and a tip that sweeps forward to join the trailing edge of the main wing. This study demonstrates the application of numerical optimization to aircraft design and presents a quantitative comparison of joined-wing and conventional aircraft designed for the same medium-range transport mission. The computer program developed for this study used a vortex-lattice model of the complete aircraft to estimate aerodynamic performance, and a beam model of the lifting-surface structure to calculate wing and tail weight. Weight estimation depended on a fully stressed design algorithm that included a constraint on buckling and a correlation with a statistically based method for total lifting-surface weight. A variety of "optimum" joined-wing and conventional aircraft designs are compared on the basis of direct operating cost, gross weight, and cruise drag. Maximum lift and horizontal tail buckling were identified as critical joined-wing design issues. The addition of a buckling constraint is shown to decrease the optimum joined-wing span and increase direct operating cost by about 4%. The most promising joined-wing designs were found to have a joint location at about 70% of the wing semispan, a fuel tank in the tail to trim, and a flap spanning 70% of the wing. These designs are shown to cost 3% more to operate than a conventional configuration designed for the same medium-range mission.

Progress in high-lift aerodynamic calculations

Abstract: The current work presents progress in the effort to numerically simulate the flow over high-lift aerodynamic components, namely, multi-element airfoils and wings in either a take-off or a landing configuration. The computational approach utilizes an incompressible flow solver and an overlaid chimera grid approach. A detailed grid resolution study is presented for flow over a three-element airfoil. Two turbulence models, a one-equation Baldwin-Barth model and a two equation k-omega model are compared. Excellent agreement with experiment is obtained for the lift coefficient at all angles of attack, including the prediction of maximum lift when using the two-equation model. Results for two other flap riggings are shown. Three-dimensional results are presented for a wing with a square wing-tip as a validation case. Grid generation and topology is discussed for computing the flow over a T-39 Sabreliner wing with flap deployed and the initial calculations for this geometry are presented.

Active Control of Aerothermoelastic Effects for a Conceptual Hypersonic Aircraft

Abstract: This article describes the procedures and results of aeroservothermoelastic studies. The objectives of these studies are to develop the necessary procedures for performing an aeroelastic analysis of an aerodynamically heated vehicle and to analyze a configuration in the classical "cold" state and in a "hot" state. Major tasks include the development of the structural and aerodynamic models, open loop analyses, design of active control laws for improving dynamic responses, and analyses of the closed loop vehicles. The analyses performed focused on flutter speed calculations, short period eigenvalue trends, and statistical analyses of the vehicle response to controls and turbulence. Improving the ride quality of the vehicle, and raising the flutter boundary of the aerodynamically heated vehicle up to that of the cold vehicle, were the objectives of the control law design investigations.

Design tradeoffs on scramjet engine integrated hypersonic waverider vehicles

Abstract: Two classes of air breathing hypersonic vehicle concepts, one for primarily cruise missions and the other for accelerator type missions, are presented. Both are designed with waverider airframes and hydrogen-fueled scramjet engine modules. Cruise configurations are optimized for the product of Isp and LID while matching lift to weight, corrected for centrifugal force, and thrust to drag at some equivalence ratio. Accelerator configurations are optimized for effective specific impulse while matching lift to weight, corrected for centrifugal force, at an equivalence ratio of 1. The method and computer code developed to optimize the configurations is discussed. The features and design tradeoffs for each class of vehicles are described. Recently available weight estimates for all-body waveriders have had a significant impact on the integrated configurations. A 60-m Mach 8 vehicle, flying at 30.3-km altitude optimized for cruise, has a LID of 4.7 and an Isp of 2786 s. A 60-m Mach 14 accelerator, flying at 36.9-km altitude, has an Ispett of 531 s.

Potential hazard of aircraft wake vortices in ground effect with crosswind

Abstract: Aircraft wake vortices, evolving close to the ground in a cross wind, are a potential hazard to aircraft landing or taking off on the same or parallel runways. The objective of this work has been to study, by means of numerical simulation, the effect of aerodynamic and environmental conditions on the generation and transport of these vortices. The approach has been to use a computer code which solves the two-dimensional, timedependent, incompressible Navier-Stokes equations expressed in stream function-vorticity form, to study wake vortices in ground effect with crosswind. The code permits the specification of arbitrary atmospheric stability and wind profiles. A mixed no-slip/free-slip lower boundary condtion has been invoked to model the interaction of the vortices with the ground. Comparisons of code output with laboratory and field data have been used to validate the code. Simulation results have shown that, even after evolution times and cross-runway transport distances on the order of 3 min and 500 m, vortices generated by large aircraft close to the ground in a crosswind can carry sufficient average circulation to be a potential hazard to smaller aircraft. Additional full-scale data need to be acquired and additional numerical comparisons need to be performed to assess the significance of these new results.

Optimization of aircraft engine suspension systems

Abstract: An interactive computer program has been developed to design aircraft engine mounting systems used for vibration isolation. Mount design is largely driven by two competing criteria. Mounts must be soft enough to provide vibration isolation, yet stiff enough to support the engine without excessive motions. The constrained variable metric optimization technique is used to determine the mount design parameters which minimize the transmitted forces in the mounts, subject to constraints on the maximum allowable deflection of the engine to static forces. The design parameters are the stiffness and orientation of each individual engine mount. The aircraft engine is modeled as a rigid body that is mounted to a rigid base representing the nacelle. An example is used to show that the optimization technique is effective in designing engine mounting system.

Integrated Structure/Control/Aerodynamic Synthesis of Actively Controlled Composite Wings

Abstract: A new multidisciplinary optimization capability for integrated synthesis of actively controlled composite wings is reviewed. It is shown that the nonlinear programming/approximation concepts approach to design optimization, combined with appropriate simplified analysis techniques for the different disciplines, make multidisciplinary wing synthesis both feasible and practical for the conceptual and preliminary design stages. The composite wing of a remotely piloted vehicle is used for numerical experimentation. Synthesis studies with design variables and constraints that span the disciplines of structures, control, and aerodynamics are presented. These studies provide new insight into the complex nature of multidisciplinary interactions in wing design. [A], { [AD} D/(APPRX) {P} {P} {qc} {qc} Nomenclature aeroservoelastic system state space matrix and input vector in standard form aerodynamic matrix used for the calculation of induced drag control system design variables (coefficients in the transfer function of a control law) induced drag induced drag approximation number of terms in the thickness polynomial for t(x, y) number of terms in the Ritz polynomial series for elastic deformation w(x, y) number of terms in the polynomial series for jig shape structural generalized stiffness matrix aeroelastic (structural and aerodynamic) generalized stiffness matrix generalized static loads generalized loads in maneuver (including inertial, jig shape, and control surface contributions) static generalized displacements generalized displacements in maneuver (including control surface rotation needed) approximate generalized displacements in maneuver

Finite Element Nonlinear Flutter and Fatigue Life of Two-Dimensional Panels with Temperature Effects

Abstract: A frequency domain method for two-dimensional nonlinear panel flutter with thermal effects obtained from a consistent finite element formulation is presented. von-Karman nonlinear strain-displacement relation is used to account for large deflections, and the quasisteady first-order piston theory is employed for aerodynamic loading. The panel motion under a combined thermal-aerodynamic loading can be mathematicall y separated into two parts and then solved in sequence: 1) thermal-aerodynamic static deflection (time-independent equilibrium position), and 2) limit-cycle oscillations. The finite element frequency domain results are compared with numerical time domain solutions. In a limit-cycle motion, the panel frequency and stress can be determined, thus fatigue life can be predicted. The influence of temperature and dynamic pressure on panel fatigue life is presented. Endurance and failure dynamic pressures can be established at a given temperature from the present method.

Efficient Simulation of Incompressible Viscous Flow over Single and Multielement Airfoils

Abstract: The incompressible, viscous, turbulent flow over single and multielement airfoils is numerically simulated in an efficient manner by solving the incompressible Navier-Stokes equations. The solution algorithm uses the method of pseudocompressibility with an upwind-differencing scheme for the convective fluxes, and an implicit line-relaxation scheme. The motivation for this work includes interest in studying high-lift takeoff and landing configurations of various aircraft. In particular, accurate computation of lift and drag at various angles of attack up to stall is desired. Two different turbulence models are tested in computing the flow over a NACA 4412 airfoil; an accurate prediction of stall is obtained. The approach used for multielement airfoils involves the use of multiple zones of structured grids fitted to each element. Two different approaches are compared: 1) a patched system of grids, and 2) an overlaid Chimera system of grids. Computational results are presented for two-element, three-element, and four-element airfoil configurations. Generally, good agreement with experimental surface pressure coefficients is seen. The code converges in less than 200 iterations, requiring on the order of 1 min of CPU time on a CRAY YMP per element in the airfoil configuration.

Envelope function - A tool for analyzing flutter data

Abstract: In this article a new tool for flutter clearance is presented and a preliminary assessment of its capabilities undertaken. It is intended to be complementary to other approaches. The method is based on quantifying the change in the shape of the decay envelope associated with control pulse responses or impulse response functions. An indication of overall stability is obtained without curve fitting by way of a shape parameter which cjuickly shows whether there has been any significant change since the last test point. In addition, the envelope functions can be overlayed from different speeds. The method is illustrated with data from a simulated aeroelastic model of a multiengined transport aircraft, and the effects of turbulence are considered. Finally, the method is successfully applied to real flight test data.

Evaluation and Extension of the Flutter-Margin Method for Flight Flutter Prediction

Abstract: For a binary flutter the so-called flutter-margin method is a good way of extrapolating from subcritical flight test data to estimate the flutter speed; the best estimates are obtained with a linear extrapolation. Good estimates of the flutter speed can be obtained from data at speeds as low as 50% of the flutter speed. The flutter-margin is*shown to be relatively insensitive to errors in the damping measurements, but is very sensitive to errors in frequency measurements. It does not give good predictions of the flutter speed when the instability is dominated by a single degree-of-freedom mechanism. A new form of flutter-margin has been developed for a trinary flutter, which also varies in a sensibly linear manner with dynamic pressure; it is also relatively insensitive to errors in damping, but is very sensitive to errors in frequency.

Aerodynamic Performance Effects due to Small Leading-Edge Ice (Roughness) on Wings and Tails

Abstract: Systematic experimental studies have been carried out to establish the aerodynamic impact of very small leading-edge simulated ice (roughness) formations on lifting surfaces. The geometries studied include singleelement configurations (airfoil and three-dimensional tail) as well as multielement high-lift airfoil geometries. Emphasis in these studies was placed on obtaining results at high Reynolds numbers to insure the applicability of the findings to full-scale situations. It has been found that the well-known Brumby correlation for the adverse lift impact of discrete roughness elements at the leading edge is not appropriate for cases representative of initial frost formation (i.e., distributed roughness). It has further been found that allowing initial ice formations, of a size required for removal by presently proposed de-icing systems, could lead to maximum lift losses of approximately 40% for single-element airfoils. Losses in angle-of-attack margin-to-stall are equally substantial— as high as 6 deg. Percentage losses for multielement airfoils are not as severe as for single-element configurations, but degradations of the angle of attack-to-sta ll margin are the same for both.

Static aeroelastic control of an adaptive lifting surface

Abstract: Piezoelectric actuators are embedded in an idealized laminated composite wing structure that can adapt to changing flight conditions. A feedback control system senses wing roots strains and then applies a proportional voltage to active actuator layers to change wing lift or divergence dynamic pressure

Numerical simulation of unsteady flow induced by a flat plate movingnear ground

Abstract: An aerodynamic model based on the general unsteady two-dimensional vortex-lattice method and the method of images was developed to predict the unsteady ground effect on the aerodynamic characteristics of a flat plate. The wake is computed as part of the solution by allowing it to deform and roll up into its natural force-free position. The model is not restricted by angle of attack, sink rate, and camber. The results agree perfectly with available exact (steady) solutions. It is also shown that the increase in the magnitude of C, and CM as a result of unsteady ground effect is greater for high sink rates, in a general agreement with published results. On the other hand, the effect of ground on wake shape and position is greater for lower sink rates. For large sink rates, the wake becomes very close to the flight path with its position less dependent on the height above the ground.

Lateral Aerodynamic Interference Between Tanker and Receiver in Air-to-Air Refueling

Abstract: Wind-tunnel data have been obtained from a tapered tanker wing and receiver aircraft model at varying vertical separation. The data are presented in derivative form and compared with theory using a flat vortex sheet model of the tanker wing wake to determine the induced angle-of-attack variation on the receiver wing, fin, and tailplane due to the tanker wing. Aerodynamic loads on the receiver are obtained by the vortex lattice method, with an allowance made for the vertical displacement of the tanker wake in the estimation of the fin side force. In the experiments, the lateral aerodynamic interference between tanker and receiver was determined by banking the tanker wing and displacing it sideways, and by yawing the receiver aircraft model. Data were obtained from open and closed test sections in order to assess the significant boundary interference effect and corrections estimated from the image vortex system of the tanker and receiver wings in the test section. In general, the theory compares favorably with the experimental data. The most significant terms are the rolling moments due to sideways and bank displacements. Significant side forces are produced due to sidewash on the fin from the tanker and receiver wings and, when displaced in yaw, the receiver experienced a loss in directional stability.

Slender wing rock revisited

Abstract: Analysis of experimental results for slender delta wings reveals that several flow phenomena play a role in the observed wing rock motions. The relative importance of the various flow mechanisms is determined by leading edge sweep and angles of attack and sideslip. The underlying fluid mechanics are analyzed, laying the foundation for future development of means for predicting not only when wing rock will occur but also how wing rock charact~tistics, such as limit cycle amplitude and oscillation frequency, depend on wing geometry and flight conditions.

Aerodynamic Shape Optimization Using Sensitivity Analysis on Third-Order Euler Equations

Abstract: Previously, the authors have shown an aerodynamic optimization method with two design variables using sensitivity analysis on the first-order-accurate discretization of the Euler equations. Two advancements of this method are reported in this article. First, nonlinear fluid dynamic phenomena including flow discontinuities are better predicted by an improved flow prediction method which uses the third-order accurate discretization of the Euler equations. Using this method, the flowfield of a modified shape which generates shocks and other large gradients is predicted based on the shock-free flowfield of the original shape and without solving the flowfield equations. Secondly, every surface grid point is used as a design variable, which virtually eliminates all geometrical restrictions on the shape as it is optimized for the specified objective. This improved algorithm is demonstrated by optimizing the ramp shape of a scramjet-afterbody configuration for maximum axial thrust. Starting with totally different initial designs, virtually identical shapes are obtained as the optimum. The method is more efficient than the traditional design methods for a few reasons, which include the use of flow predictions and the elimination of a priori guessing of possible shapes from which the optimum is to be selected.

Schlieren studies of compressibility effects on dynamic stall of transiently pitching airfoils

Abstract: : Compressibility effects on the flowfield of an airfoil executing rapid transient pitching motion from 0-60 deg over a wide range of Mach numbers and pitching rates were studied using a stroboscopic schlieren flow visualization technique. The studies have led to the first direct experimental documentation of multiple shocks on the airfoil upper surface flow for certain conditions. Also, at low Mach numbers, additional coherent vortical structures were found to be present along with the dynamic stall vortex, whereas at higher Mach numbers the flow was dominated by a single vortex. The delineating Mach number for significant compressibility effects was 0.3 and the dynamic stall process was accelerated by increasing the Mach number above that value. Increasing the pitch rate monotonically delayed stall to angles of attack as large as 27 deg.

Simultaneous structure/control design optimization of a wing structure with a gust load alleviation system

Abstract: Simultaneous design optimization is considered for structure and control parameters of a wing with a gust load alleviation (GLA) control system The application of a goal programming (GP) formulation to the design synthesis of an aeroservoelastic system is carried out by the use of a simple mathematical model in conjunction with a wind-tunnel model having a GLA control system Numerical applications are based on a cantilever wing having an aileron surface controlled by wing-tip accelerometer feedback signals System equations are obtained in the form of state equations, thus enabling the statistical characteristics of both the gust-induced wing stress and the control surface deflection angle to be evaluated using their standard deviations The wing spar height and the controller feedback gain are simultaneously optimized to obtain the minimum spar weight while satisfying the following structure and control design constraints: 1) the spar stress is limited with the control system either on or off; 2) the control surface deflection angle is restricted; and 3) system stability should be guaranteed by incorporating a controller stability margin Numerical examples demonstrate the successful application of a GP formulation for the simultaneous structure/control design synthesis by specifying priorities to the conflicting design constraints

Vortex control technique for the attenuation of fin buffet

Abstract: The application of tangential leading edge blowing (TLEB) to reduce levels of single-fin buffeting has been studied. The tests were performed at the University of Bath 2.1 x 1.5 m wind tunnel, using two cropped 60deg delta wings. To measure the unsteady pressures on the fin surface, a rigid fin instrumented with miniature differential pressure transducers was used. A flexible fin of similar planform and size was used to measure the buffeting response. Steady-state static pressure data and laser light sheet flow visualization were employed to aid interpretation of the vortical flow over the wings, and hence, identify the causes of the buffeting. The profiles of the unsteady pressures and the buffeting response were found to match each other very closely. It was observed that symmetric leading-edge blowing modified the leading-edge vortices by reducing the "effective angle of attack" of the vortex. Blowing at a constant rate shifted the buffet excitation and response to higher angles of attack. Flow visualization confirmed that the mechanism at peak buffeting had not changed, but had been merely shifted. It has been shown that the use of an optimum blowing profile could completely suppress the buffeting response without impairing the wing lift characteristics.

Computation of induced drag for elliptical and crescent-shaped wings

Abstract: Recent interest in the induced drag characteristics of crescent-shaped wings has led to a closer look at the methods used for determination of induced drag from computational aerodynamic methods. Induced drag may be computed by integration of surface pressure or by evaluation of a contour integral in the Trefftz plane. A high-order panel method was used to study the induced drag of elliptical and crescent-shaped wings using both techniques. Induced drag computations using surface pressure integration were strongly affected by panel density and angle of attack. Drag computations for the crescent wing were especially sensitive to spanwise panel density because of the complex flowfield near the tip. Trefftz-plane results for the two wing planforms were not sensitive to panel density or angle of attack. The effect of correctly modeling the force-free, rolled-up wake geometry on the predicted span efficiency was demonstrated for both wing planforms. The span efficiency predicted from Trefftz-plane integration was about 0.97 for the elliptical wing and 0.99 for the crescent wing, both somewhat less than the classical theoretical maximum for planar wings. Most of the apparent drag reduction of the crescent wing claimed in previous studies was probably an artifact of the surface-pressure integration. The slightly higher span efficiency for the crescent wing was attributed to a more nearly elliptical spanwise lift distribution.

Prediction of fluctuating pressure in attached and separated turbulent boundary-layer flow

Abstract: A methodology is presented that predicts the fluctuating pressure and power spectra for attached zeropressure gradient and separated turbulent boundary-layer flow on smooth and rough surfaces. Attached flow conditions use a prediction technique that employs a transformation of boundary-layer properties from the compressible to the incompressible plane where a comparison can be made to an extensive data base. For a rough wall, it is shown that the rms pressure, which scales with wall shear, can be predicted by augmenting the smooth wall value by the rough/smooth skin-friction ratio. Relative to nonattached flow, represented by two-dimensional compression corner and three-dimensional fin-generated shock/boundary-layer interactions, the rms pressure is shown to scale with approach flow conditions and the oblique shock in viscid pressure rise. For this situation, a new in viscid angle has been defined as = a + /3 sin^U/M) where a is the shock generator angle and /3 is a parameter based on two-dimensional or three-dimensional interactions. Both rms pressure and power spectra have been correlated in terms of undisturbed approach flow boundary-layer parameters and modified inviscid shock strength relations to provide engineering solutions for the design resolution to complex flow problems.