Showing papers on "Lift-induced drag published in 2000"

Low Reynolds Number Aerodynamics of Low-Aspect-Ratio, Thin/Flat/Cambered-Plate Wings

Abstract: The design of micro aerial vehicles requires a better understanding of the aerodynamics of small low-aspect-ratio wings An experimental investigation has focused on measuring the lift, drag, and pitching moment about the quarter chord on a series of thin flat plates and cambered plates at chord Reynolds numbers varying between 60,000 and 200,000 Results show that the cambered plates offer better aerodynamic characteristics and performance It also appears that the trailing-edge geometry of the wings and the turbulence intensity in the wind tunnel do not have a strong effect on the lift and drag for thin wings at low Reynolds numbers Moreover, the results did not show the presence of any hysteresis, which is usually observed with thick airfoils/wings

Aerodynamic Measurements at Low Raynolds Numbers for Fixed Wing Micro-Air Vehicles

Abstract: : A description of the micro-air vehicle (MAV) concept and design requirements is presented These vehicles are very small and therefore operate at chord Reynolds numbers below 200,000 where very little data is available on the performance of lifting surfaces, ie, airfoils and low aspect-ratio wings This paper presents the results of a continuing study of the methods that can be used to obtain reliable force and moment data on thin wings in wind and water tunnels To this end, a new platform force and moment balance, similar to an already existing balance, was designed and built to perform lift, drag and moment measurements at low Reynolds numbers Balance characteristics and validation data are presented Results show a good agreement between published data and data obtained with the new balance Results for lilt, drag and pitching moment about the quarter chord with the existing aerodynamic balance on a series of thin flat plates and cambered plates at low Reynolds numbers are presented They show that the cambered plates offer better aerodynamic characteristics and performance Moreover, it appears that the trailing-edge geometry of the wings and the turbulence intensity up to about 1% in the wind tunnel do not have a strong effect on the lilt and drag for thin wings at low Reynolds numbers However, the presence of two endplates for two-dimensional tests and one endplate for the semi-infinite tests appears to have an undesirable influence on the lift characteristics at low Reynolds numbers

Topographically Induced Drag and Mixing at a Small Bank on the Continental Shelf

Abstract: Recent turbulence measurements over a small bank on the continental shelf off Oregon reveal a previously undetected site for intense mixing of the coastal ocean. The flow is hydraulically controlled and turbulence diffusivities over the bank are more than 100 times greater than estimates made on the shelf away from topography. The total drag exerted by the bank on the flow field is a combination of bottom friction plus form drag (analogous to mountain drag) and is comparable to the Coriolis force. This drag is sufficient to decelerate the flow over the bank in a matter of hours.

Design of Fuzzy Logic Guidance Law Against High-Speed Target

Abstract: An advanced guidance law is developed against very high-speed targets. Preliminary studies have shown that the aspect angle of the interceptor at lock-on near 180 deg is a fundamental requirement for achieving small miss distance against a very high-speed target. To meet this requirement, a fuzzy guidance law in midcourse phase that is more similar to human decision making is designed. In terminal phase, a proportional and derivative-type fuzzy terminal guidance law is explored. It is shown that the integrated guidance scheme offers a near head-on homing condition before the missile enters terminal phase and provides better e nal results (smaller miss distance and wider defensible volume ) than the conventional guidance law. A complete simulation study is performed to show the effects of the proposed design. Nomenclature CD = drag coefe cient CD0 = zero drag coefe cient CL = lift coefe cient CLa = @CL/@a D = drag h = vertical coordinate L = lift m = mass Q = dynamic pressure Stref = reference area of ballistic target s = reference area T = thrust v = missile speed vt = target speed W = target weight x = horizontal coordinate a = angle of attack b = ballistic coefe cient c = e ight-path angle c t = reentry angle of target d = velocity angle error h = inertial line-of-sight angle l = induced drag coefe cient q = atmospheric density r = heading error angle

Multidisciplinary optimization of transonic wing design based on evolutionary algorithms coupled with cfd solver

Abstract: Evolutionary Algorithms (EAs) were applied to multidisciplinary transonic wing design optimizations. Aerodynamic performances of the design candidates were evaluated by using the three-dimensional compressive Navier-Stokes equations to guarantee an accurate model of the flow field. The wing structure is modeled on a box-beam to estimate the wing thickness and wing weight. To overcome enormous computational time necessary for the optimization, the computation was parallelized on Numerical Wind Tunnel at NAL in Japan and NEC SX-4 computers at Computer Center of Tohoku University in Japan. First, a singleobjective wing design optimization was demonstrated by maximizing L/D with a structural constraint using a real-coded Adaptive Range Genetic Algorithm (ARGA). Because the structural constraint imposed a tradeoff between minimizations of the induced drag and the wave drag, the present ARGA found a compromised but reasonable design. Then, a multiobjective wing design optimization is performed by minimizing both drag and weight with a constraint on CL using a Multiobjective Evolutionary Algorithm (MOEA). Due to the tradeoff between minimization of aerodynamic drag and minimization of weight of wing structure, the solution to this problem is not a single point but a set of compromised designs. The present MOEA successfully captured these solutions that revealed the tradeoff information. These results showed that EAs were promising approach to multidisciplinary optimization problems.

Experiments on a 60-Degree Delta Wing with Rounded Leading-Edge Vortex Flaps

Abstract: Low-speed wind-tunnel measurements were done on a 1.15-m span 60-deg delta wing with rounded leading-edge vortex flaps. The purpose of the measurements is to assess the benefits of the rounded leading-edge vortex flaps in regard to improving the lift/drag ratio of delta wings. Force and surface pressure measurements were made at a Reynolds number based on a centerline chord of 2 x 10 6 . The increase in the radius of the rounded leading edge reduces the drag significantly both with and without flap deflection except in the minimum drag region. Deflecting the rounded leading-edge vortex flap improves the lift/drag ratio at relatively higher lift coefficients, when compared with the sharp-edged vortex flap. The largest improvement in the lift/drag ratio as compared with the sharp-edged delta wing with vortex flaps is more than 25% in the lift coefficient range between about 0.6 and 0.8 for the rounded-edge delta wing with flaps that were deflected 30 deg downward

Drag control system for flying machine, process for estimating drag of flying machine, boundary layer control system, and boundary layer control process

Abstract: The thrust T of an airplane is estimated by a thrust estimating device, and the motion state (the speed, the angular speed, the attitude angle and the elevation angle) of the airplane is detected by a motion state detecting device. Then, a drag estimating device estimates the drag D of the airplane, based on the thrust T and the motion state of the airplane. An operation-amount calculating device converges the steering angle into a target steering angle at which the drag of the airplane is minimized, by repeating the operation of varying the steering angle of a flap by a very small angle by an operating device, and further varying the steering angle by a very small angle, while monitoring the increase or decrease in drag resulting from such variation. Such a drag reducing control is carried out, while monitoring the actual drag and hence, is extremely effective, and also can exhibit an effectiveness, irrespective of the motion state of the airplane.

Three Surface Aircraft - A Concept For Future Transport Aircraft

Abstract: For the design and optimization of a canard for a three surface aircraft (TSA) a preliminary aircraft design code was extended by a higher-order panel method, a transonic data base and a trim routine suitable for a three surface aircraft. Preliminary aim of the canard study was the preliminary design of an optimum canard geometry for the DLR-F11 TSA - wind tunnel model under the design constraints of a realistic aircraft. Based on the conventional design for the wind-tunnel model, the retrofit of the geometrically fixed aircraft with a canard was simulated. With sensitivity studies the influence of basic canard geometry parameters was analyzed. Based on an optimization calculation the best set of geometry parameters which provide minimum fuel weight were determined. The analyses show that a swept back canard in low position located downstream the divergent nose part of the body, with high aspect ratio, low taper ratio and moderate canard span promises optimum performance. In addition, the canard lift of the TSA configuration allows to realize higher cruise lift coefficients for an improved induced drag at transonic speeds without the risk of additional wave drag. However, without the control concept of a free-floating canard, the static margin decreases significantly with the integration of a canard.

Aerodynamic and Flight Dynamic Simulations of Aileron Characteristics

Abstract: The subject of this investigation is the application of CFD computations to flows around airplane ailerons combined with flight mechanical simulations to study the impact on airplane rolling maneuvers and aileron dynamics. The practical application is on Saab 2000 commuter airplane. In the validation of CFD computations the low speed airfoils FX 61-163 and FX 6617AII-182 were investigated with the 2D Navier-Stokes code ns2d by comparing the computations with selected wind tunnel experiments. The medium speed MS(1)-0313 and the transonic DLBA032 airfoils with plain ailerons were investigated with ns2d and NSMB codes in selected wind tunnel cases representative for the ailerons of Saab 2000 aircraft. One algebraic and three k-e turbulence models were used in the calculations at different aileron deflections. The effects of local mesh refinement and grid convergence were studied on the aerodynamic coefficients. Two-dimensional CFD computations were made on Saab 2000 aileron to compare the hinge moment with flight test results, measured by disconnecting the left and right hand side ailerons. The local angles of attack were determined by using extended lifting line theory and the conversion to 3D coefficients was made with handbook methods. The airplane rolling moment was determined by inserting the CFD derived lift effectiveness into the calculations. The effects of aileron slot and tab slot gap sizes as well as aileron hinge axis position on the aerodynamic coefficients were computed with the ns2d code. The CFD derived aerodynamic coefficients were fed into a six degree of freedom flight mechanical simulation system to study the impact on airplane rolling maneuvers. Frequency analysis was performed on the response of aileron deflection, airplane roll rate and roll acceleration to applied wheel force using fast Fourier transform, spectrum analysis and system identification. A review was made on practical aileron design considerations with issues on maximum wheel force, aileron effectiveness, wind tunnel testing, induced drag and aileron control system.

Analysis of calculated and measured wake characteristics of a propeller-wing model

Abstract: A quantitative flow field survey based on an Augmented Betz Integral Method (ABI-method) is applied to both numerically and experimentally acquired flow field data behind a tractor propellerwing configuration. This technique, that is capable of splitting viscous drag and induced drag, has been used earlier to analyze experimental data. However its use with numerically determined data is limited. In the paper results of numerical calculations based on the Navier-Stokes equations and experimental data obtained from 5hole probe measurements are compared and used as input for the ABI-Method. Both the calculations and the experiments were performed on the same low aspect ratio propellerwing model equipped with a 4 bladed tractor propeller. Configurations with and without running propeller were investigated for angles of attack of 0” and 4”. The agreement between the experimentally and the numerically determined data turns out to be reasonable. The fineness of the grid, however, combined with the correct choice of viscous turbulence model and boundary conditions may cause serious problems for the accurate prediction of drag and lift characteristics using the ABIM if not tuned well. Although the numerical method still shows some shortfalls with respect to the accuracy of the flow field data in the wake, one important advantage is the appreciable detail of the flow around the entire configuration. Based on the numerical results im* Research ass., Department of Aerospace Engineering, Member AIAA ** Student Politechnico Torino Copyright 0 The American Institute of Aeronautics and Astronautics. All rights reserved portant sources of drag can be detected. This information further enhances the optimization strategy that is used on new propeller-wing layouts currently under investigation.

Unsteady Aerodynamics of a WIG Airfoil Flying over a Wavy Wall

Abstract: Aerodynamic characteristics of an airfoil, NACA 6409, flying over a wavy wall is investigated numerically. An Euler code based on the LU-factored algo rithm and higher-order upwind scheme is constructed and its accuracy is tested on two benchmark problems : NACA 4412 airfoil moving over a level ground and NACA 0012 airfoil in free-flight pitching oscillation. The calculated flow about NACA 6409 airfoil over the wavy ground represented by a moving sine function indicates that the aerodynamic property of the airfoil becomes sensitive if the wave number or amplitude of the wavy ground is increased, and/or if the proximity of the airfoil to the ground is smaller. Introduction When an aircraft flies near the ground, the aerodynamic performance of the wing is much changed from that of the free flight. In particular, a wing in proximity to the ground manifests reduced upwash, downwash, and tip vortices which cause enhancement of lift and pitching moment, and reduction of induced drag. These effects, in overall, are called ‘the ground effect’ whereas the wing taking advantage of these effects is called ‘wing in the ground effect’ or WIG, in short. There has been recently considerable interest in development of WIG crafts in such countries as Russia[1,2], Japan[3,4], Germany[5,6], and China[7] due to its energy saving feature as a means of passenger and cargo transportation. Preliminary design of a WIG airplane is under progress in South Korea[8] and the present research is motivated by this particular program. The gain in the lift-todrag ratio might be achieved by a WIG craft at the cost of stability due to the increased pitching moment. The flight properties of a WIG hence need to be investigated thoroughly during the aircraft development process. Recently, a few reports have appeared on the aerodynamic performance of WIG wings and airfoils[4-7,9] flying over a level *Doctoral candidate, Dept. of Aerospace Engineering. tprofessor, Dept. of Aerospace Engineering, Member AIAA. Copyright @ 2000 by YEHOO” Im. Kcun-Shik Chmg. Publidld by the Americ-n Institute of Aeronautica md Aatronautica. Inc. with permhion. ground. NACA Cdigit airfoils are, in general, known to have positive ground effect. However, a symmetric airfoil with large thickness, NACA 0012, for example, can exhibit negative ground effect at a small angle of attack[4,7]. The WIG craft operated over the sea is quite often expected to face with rough wavy surfaces. This rather periodic terrain will cause ground effect different from that of the level ground. In the literature, the unsteady flow past a twodimensional airfoil moving over a wavy ground has been investigated with the lifting surface theory by Ando et al.[lO], and with the unsteady panel method by Morishita and Ashihara[ll]. Mizutani and Suzuki[l2] used Rankine source and boundary element method to compute the wing aerodynamics over the coastal free surface. Their results have certainly offered some useful data for fundamental study of the lifting airfoils. However, the solutions of Euler and Navier-Stokes equations are still wanted in order to study the possible compressibility and viscosity effect of the WIG craft. In this paper we have numerically solved the Euler equations with the LU-factored algorithm[l3] and high-resolution upwind scheme for the unsteady WIG airfoil moving over a wavy wall. Fortunately, Mizutani and Suzuki[l2] have found that deformation of the free surface caused by the proximate flight of a WIG craft is only negligible. We verified accuracy of the present computer code by reproducing two earlier results : NACA 4412 airfoil over a level ground[l4], and the NACA 0012 airfoil in pitching oscillation[l5]. Various parameter effects on NACA 6409 airfoil flying over a wavy wall with a subsonic flight Mach number, M, = 0.3, is elaborated in this paper. Governing Equations We consider unsteady twodimensional Euler equations in the computational domain (t, 5, n), 1 OF 8 aQ ai aF z+-+-=O ac 877 with the flux vectors (1) AMERICAN INSTITUTE OF AERONAUTICS AND ASTRONAUTICS PAPER 2000-0657

Linear and non-linear potential-flow calculations of free-surface waves with lift and induced drag

Abstract: A potential-flow panel method is used to compute the waves and the lift force from surface-piercing and submerged bodies. In particular the interaction between the waves and the lift produced close to the free surface is studied. Both linear and non-linear free-surface boundary conditions are considered. The potential-flow method is of Rankine-source type using raised source panels on the free surface and a four-point upwind operator to compute the velocity derivatives and to enforce the radiation condition. The lift force is introduced as a dipole distribution on the lifting surfaces and on the trailing wake, together with a flow tangency condition at the trailing edge of the lifting surface. Different approximations for the spanwise circulation distribution at the free surface were tested for a surface-piercing wing and it was concluded that a double-model approximation should be used for low speeds while a single-model, which allows for a vortex at the free surface, was preferred at higher speeds. The lift force and waves from three surface-piercing wings, a hydrofoil and a sailing yacht were computed and compared with measurements and good agreement was obtained.

Wing planform optimization of human powered aircraft in low reynolds number range

Abstract: This study is concerned with a wing planform design of a high aspect ratio wing in a low Reynolds number range used by a human powered aircraft. In the low Reynolds number range, effect of the profile drag must be considered as well as the induced drag. For flight stability, an adequate deformation of the main spar is desirable to bring a dihedral effect. The structural problem is also incorporated into the optimization problem as a constraint concerning the tip displacement. Moreover, the spar sizing design has an effect on the weight and hence the power-required of the aircraft. Accordingly, the power-required minimized design is formulated in terms of the wing planform and the main spar dimensions. Through numerical calculations, effect of the profile drag on the wing planform design is demonstrated. The profile drag is approximated by polynomials in terms of the Reynolds number for an efficient calculation. Then, the importance of the multidisciplinary optimization is demonstrated for the wing design of the human powered aircraft.

Spiral cylindrical wing tip

Abstract: not available for EP1091870Abstract of corresponding document: WO0001580The invention concerns a device in the form of a spiral cylindrical cavity, mobile or fixed, capable of being adapted to the tip of all carrier profiles, in particular to aeroplanes and gliders or craft requiring lifting capacity in its displacement, for reducing induced drag, increasing thrust and reaching in flight or lift, low speed levels with wide angles of incidence and high speed levels in cruising degree. The invention enables to use for similar performances or for initial performances weaker traction systems for longer and faster flights using less fuel and/or using larger loads. Moreover, said device enables to considerably and artificially increase a wing extension while preserving the same profile and the initial aerodynamics of an aeroplane in flight or of a levitating craft using lifting capacity up to subsonic speed.

The lifting system with minimum induced drag

Abstract: The lifting system with minimum induced drag is a closed wing with a box shaped front view and was named Best Wing System by Prandtl; on the horizontal wings, the circulation results from the superposition of an elliptical and a constant distributions and, on the the side walls, it is butterfly shaped. Very recently, the Prandtl results (obtained by an approximate procedure) were confirmed by a closed form solution, briefly described in the paper. The concept of minimum induced drag can give rise to a family of aircraft, ranging from very small to very large (larger than the A3XX) aircraft and still compatible with the actual airport areas. These aircraft are named as PrandtlPlane, in honor of Prandtl. An example of a 600 seat category aircraft is presented.

Shock Control by Adaptive Elements for Transportation Aircraft Wings

Abstract: Different devices for the application of shock and boundary layer control on transportation aircraft wings have been investigated on 2-d airfoils and on a swept wing model. A cavity in the surface underneath the foot of the shock covered with a perforated plate reduces shock strength and hence wave drag, but viscous drag increases such that a net drag reduction can not be achieved in most cases. The application of additional boundary layer suction reduces the additional viscous drag, but not enough to result in a significant gain in total drag. On the contrary a contour bump underneath the shock, applied alone or in combination with suction, reduces very effectively wave drag without increasing viscous drag so that under off-design conditions up to 24% total drag reduction has been measured for a 2-d airfoil and somewhat lower values for the swept wing. This effect has been well predicted by numerical methods. Both devices, especially the perforation, have a positive influence on the buffet boundary. Trailing edge devices such as conventional and Gurney-type flaps also effect wave drag by redistributing the pressure on the wing or airfoil. Combining them with a contour bump has been investigated numerically. The results show that by careful optimization of the flap deflection together with the corresponding bump location and height a better performance can be achieved compared to the application of either device alone.

Design of wingtip vortex turbine

Abstract: During lift the pressure difference across upper and lower sides of the aircraft wing, forces the flow of air to curl around the wing tip resulting in wing tip vortices which in turn creates induced drag. Induced drag has been a great problem of all the aircraft and at times it reaches up to 70% of the total drag. The idea of wing tip turbine is to use the energy of the vortices to generate extra power thus reducing induced drag at the same time. However total drag would increase due to profile drag of the turbine. To cater it, this design has been done such that the drag created by turbine in a particular phase should not exceed the induced drag reduced by the turbine. In this way maximum energy is extracted with out paying the penalty of drag. The theoretical study is based upon drag type turbine installed on NACA 2418 airfoil whose specifications bear geometric similarity to the wing of T-37 aircraft. The experimental results are based upon the wind tunnel testing of the scale down model of the assumed wing.

Wing tip device of aircraft

Abstract: PROBLEM TO BE SOLVED: To provide a wing tip device of an aircraft acting effectively even on the aircraft with relatively small lift generated in a cruising state and reducing efficiently induced drag generated by wing tip eddy. SOLUTION: A leading edge (a straight line part) 11b of the wing tip device 10 in a sweepback wing shape mounted on a wing tip part 4a of a main wing 4 of the small aircraft has a predetermined sweepback angle θ. When the small aircraft is in the cruising state, air current in front of the leading edges 11a, 11b is added with an upward attack angle by upwash effect (phenomenon). Since the upward attack angle is added, the lift inclines forward in a flying direction and its horizontal component of force acts as a thrust. Since the straight line part 11b has the predetermined sweepback angle θ, a cord length of the wing tip 11c is short for reducing strength of the wing tip eddy. Since an upper half angle 8 of the wing tip device 10 for the main wing 4 is considerably small within a predetermined range, little interference drag is generated at a connecting portion of the wing tip device 10 and the main wing 4. COPYRIGHT: (C)2002,JPO

Studies on the optimum performance of tapered vortex flaps

Abstract: The effects of sweepback angles, leading-edge shapes and flap hinge-line positions over the performance of the leading-edge vortex flaps are discussed in this paper. As the sweepback angle decreases, improvements of the lift/drag ratio are attained over a wider lift coefficient range when compared with a slender delta wing. A rounded leading-edge vortex flap improves the lift/drag ratio at a relatively higher lift coefficient for both the 50° and 60° delta wings. The differences of the vortex flap hinge-line position also affect the performance of the vortex flap. The best lift/drag ratio is attained when the delta wing has vortex flaps with a relatively small spanwise length.

Reduction of the Induced Drag of a Supersonic Wing

Abstract: The oblique wing effect, i.e., a reduction in the wave drag for given lift, cannot be realized for a delta wing with supersonic leading edges owing to the lift reduction in the wing mid-section. To preserve the effect, the disturbances generated by the delta wing vertex must be eliminated by adding a body (wedge) to the wing by replacing the streamsurfaces behind the shock with rigid surfaces. Moreover, using wing tip deflection, and thereby reducing the wave drag to zero, makes it possible to obtain a lift- drag ratio close to that of the limiting, infinitely long flat plate.

Unsteady aerodynamics of a WIG airfoil flying over a wavy wall

Abstract: Aerodynamic characteristics of an airfoil, NACA 6409, flying over a wavy wall is investigated numerically. An Euler code based on the LU-factored algo rithm and higher-order upwind scheme is constructed and its accuracy is tested on two benchmark problems : NACA 4412 airfoil moving over a level ground and NACA 0012 airfoil in free-flight pitching oscillation. The calculated flow about NACA 6409 airfoil over the wavy ground represented by a moving sine function indicates that the aerodynamic property of the airfoil becomes sensitive if the wave number or amplitude of the wavy ground is increased, and/or if the proximity of the airfoil to the ground is smaller. Introduction When an aircraft flies near the ground, the aerodynamic performance of the wing is much changed from that of the free flight. In particular, a wing in proximity to the ground manifests reduced upwash, downwash, and tip vortices which cause enhancement of lift and pitching moment, and reduction of induced drag. These effects, in overall, are called ‘the ground effect’ whereas the wing taking advantage of these effects is called ‘wing in the ground effect’ or WIG, in short. There has been recently considerable interest in development of WIG crafts in such countries as Russia[1,2], Japan[3,4], Germany[5,6], and China[7] due to its energy saving feature as a means of passenger and cargo transportation. Preliminary design of a WIG airplane is under progress in South Korea[8] and the present research is motivated by this particular program. The gain in the lift-todrag ratio might be achieved by a WIG craft at the cost of stability due to the increased pitching moment. The flight properties of a WIG hence need to be investigated thoroughly during the aircraft development process. Recently, a few reports have appeared on the aerodynamic performance of WIG wings and airfoils[4-7,9] flying over a level *Doctoral candidate, Dept. of Aerospace Engineering. tprofessor, Dept. of Aerospace Engineering, Member AIAA. Copyright @ 2000 by YEHOO” Im. Kcun-Shik Chmg. Publidld by the Americ-n Institute of Aeronautica md Aatronautica. Inc. with permhion. ground. NACA Cdigit airfoils are, in general, known to have positive ground effect. However, a symmetric airfoil with large thickness, NACA 0012, for example, can exhibit negative ground effect at a small angle of attack[4,7]. The WIG craft operated over the sea is quite often expected to face with rough wavy surfaces. This rather periodic terrain will cause ground effect different from that of the level ground. In the literature, the unsteady flow past a twodimensional airfoil moving over a wavy ground has been investigated with the lifting surface theory by Ando et al.[lO], and with the unsteady panel method by Morishita and Ashihara[ll]. Mizutani and Suzuki[l2] used Rankine source and boundary element method to compute the wing aerodynamics over the coastal free surface. Their results have certainly offered some useful data for fundamental study of the lifting airfoils. However, the solutions of Euler and Navier-Stokes equations are still wanted in order to study the possible compressibility and viscosity effect of the WIG craft. In this paper we have numerically solved the Euler equations with the LU-factored algorithm[l3] and high-resolution upwind scheme for the unsteady WIG airfoil moving over a wavy wall. Fortunately, Mizutani and Suzuki[l2] have found that deformation of the free surface caused by the proximate flight of a WIG craft is only negligible. We verified accuracy of the present computer code by reproducing two earlier results : NACA 4412 airfoil over a level ground[l4], and the NACA 0012 airfoil in pitching oscillation[l5]. Various parameter effects on NACA 6409 airfoil flying over a wavy wall with a subsonic flight Mach number, M, = 0.3, is elaborated in this paper. Governing Equations We consider unsteady twodimensional Euler equations in the computational domain (t, 5, n), 1 OF 8 aQ ai aF z+-+-=O ac 877 with the flux vectors (1) AMERICAN INSTITUTE OF AERONAUTICS AND ASTRONAUTICS PAPER 2000-0657

Lift, drag and separate

Abstract: Helen Gavaghan finds that the storm-tossed life of a limpet is not made any easier by its shell.

Panel code estimates of wind tunnel wall and support interference effects on the hsct

Abstract: The Boeing internal version of the higher order panel code PANAIR, known as A502, was used to compute wall and model support interference effects on a Boeing/NASA High Speed Civil Transport, or HSCT, configuration. Computational results were compared to experimental data acquired on the 6% HSCT low speed model in the NASA-Langley 14'x22' wind tunnel. The numerical results indicated the influence of the model supports on the lift to be greater than that of the walls at low angles of attack and to decrease very slowly with incidence, in contrast to the increasing wall effect. The swept strut support had a much larger impact on the lift than the simple post mount and the impact of the model support systems on the lift curve slope was determined to be negligible. The magnitude of the induced drag correction from the walls and model supports increased with incidence. Pitching moment behavior indicated that the strut and post support influence decreased with angle of attack as wall influence increased. Interdependency between the walls and mounting system impact on the aerodynamic loads was observed. Application of the A502 corrections to experimental data obtained in the NASA 14' x 22' tunnel using different model supports reduced the differences between the experimentally measured lift curves. The results of this study indicated that while a panel code such as PANAIR cannot be used to predict absolute aerodynamic load levels, it can be used to estimate model support interference trends and increments and to correct differences in measured data.