Showing papers on "Lift-induced drag published in 2001"

Tight Formation Flight Control

Abstract: Thetightformatione ightcontrolproblemisaddressed.Theformationconsistsofa lead andwingaircraft,where the wing e ies in tight formation with the lead, such that the lead’ s trailing vortices aerodynamically couple the lead and the wing, and a reduction in the formation’ s induced drag is achieved. A controller (i.e., a formation-hold autopilot for the wing aircraft ) is designed such that the formation’ s geometry is maintained in the face of lead aircraft maneuvers. In the formation e ight control system, the wing and lead aircraft dynamics are coupled due to kinematic effects, and, in the case of tight formations, additional aerodynamic coupling effects are introduced. These additional aerodynamic coupling effects are properly modeled. The most signie cant aerodynamic coupling effect introduced by tight formation e ight entails the coupling of the lateral/directional channel into the altitudehold autopilot channel. It is shown that formation-hold autopilots designed ignoring the aerodynamic coupling effect yield satisfactory performance in tight formation e ight. Nomenclature b = wingspan of wing CLL = lift coefe cient of the lead aircraft S = surface area of wing VSW = sidewash W = wash vector WUW = upwash

DRAG DUE TO LIFT: Concepts for Prediction and Reduction

Abstract: ▪ Abstract This article describes some of the fundamental ideas underlying methods for induced-drag prediction and reduction. A review of current analysis and design methods, including their development and common approximations, is followed by a survey of several approaches to lift-dependent drag reduction. Recent concepts for wing planform optimization, highly nonplanar surfaces, and various tip devices may lead to incremental but important gains in aircraft performance. Focusing on relatively high-aspect-ratio subsonic wings, the review suggests that opportunities for new concepts remain, but the greatest challenge lies in their integration with other aspects of the system.

Performance Analysis of a Wing With Multiple Winglets

Abstract: This effort examined the potential of multi-winglets for the reduction of induced drag without increasing the span of aircraft wings. Wind tunnel models were constructed using a NACA 0012 airfoil section for the untwisted, rectangular wing and flat plates for the winglets. Testing of the configurations occurred over a range of Reynolds numbers from 161,000 to 300,000. Wind tunnel balances provided lift and drag measurements, and laser flow visualization obtained wingtip vortex information. The Cobalt60 unstructured solver generated flow simulations of the experimental configuration via solution of the Euler equations of motion. The results show that certain multi-winglet configurations reduced the wing induced drag and improved L/D by 15-30% compared with the baseline 0012 wing. A substantial increase in lift curve slope occurs with dihedral spread of winglets set at zero incidence relative to the wing. Dihedral spread also distributes the tip vortex. These observations supplement previous results on drag reduction due to lift reorientation with twisted winglets set at negative incidence.

Multidisciplinary Design Optimization of a Transonic Commercial Transport with Strut-Braced Wing

Abstract: The multidisciplinary design optimization of a strut-braced wing (SBW) aircraft and its benefits relative to a conventional cantilever wing configuration are presented. The multidisciplinary design team is divided into aerodynamics, structures, aeroelasticity, and the synthesis of the various disciplines. The aerodynamic analysis uses simple models for induced drag, wave drag, parasite drag, and interference drag. The interference drag model is based on detailed computational fluid dynamics analyses of various wing-strut intersections. The wing structural weight is calculated using a newly developed wing bending material weight routine that accounts for the special nature of SBWs. The other components of the aircraft weight are calculated using a combination of NASA's flight optimization system and Lockheed Martin aeronautical systems formulas. The SBW and cantilever wing configurations are optimized using design optimization tools (DOT) software

A computational model for estimating the mechanics of horizontal flapping flight in bats: model description and validation

Abstract: We combine three-dimensional descriptions of the movement patterns of the shoulder, elbow, carpus, third metacarpophalangeal joint and wingtip with a constant-circulation estimation of aerodynamic force to model the wing mechanics of the grey-headed flying fox (Pteropus poliocephalus) in level flight. Once rigorously validated, this computer model can be used to study diverse aspects of flight. In the model, we partitioned the wing into a series of chordwise segments and calculated the magnitude of segmental aerodynamic forces assuming an elliptical, spanwise distribution of circulation at the middle of the downstroke. The lift component of the aerodynamic force is typically an order of magnitude greater than the thrust component. The largest source of drag is induced drag, which is approximately an order of magnitude greater than body form and skin friction drag. Using this model and standard engineering beam theory, we calculate internal reaction forces, moments and stresses at the humeral and radial midshaft during flight. To assess the validity of our model, we compare the model-derived stresses with our previous in vivo empirical measurements of bone strain from P. poliocephalus in free flapping flight. Agreement between bone stresses from the simulation and those calculated from empirical strain measurements is excellent and suggests that the computer model captures a significant portion of the mechanics and aerodynamics of flight in this species.

Measurements of polymer induced drag reduction and polymer scission in Taylor flow using standard double-gap sample holders with axial symmetry

Abstract: The rheological properties of 1–50 ppm poly α-olefins in Varsol 80 were studied using laminar and Taylor flows. In Taylor flow, all the studied poly α-olefins displayed polymer induced drag reduction and polymer scission. The measurements were carried out using a commercial rheometer equipped with a standard double-gap sample holder with axial symmetry. We find that this particular instrument even at the maximum obtainable geometrically averaged shear rate of 15,000 s −1 yields, an accuracy and a reproducibility better than ±2.5%. This unique precision for measurements in the presence of Taylor vortices, the small amounts of sample and time needed to carry out the measurements, and the general availability of the rheometer, suggests that these previously unexplored properties of this instrument will be of significant value for future investigations of polymer induced drag reduction.

America's Cup downwind sailsFvertical wings or horizontal parachutes?

Abstract: Analysis of the boat speed polar for a typical International America’s Cup Class yacht shows that the normal downwind apparent wind angles are between 901 and 1351 in winds over 5 m/s (10 knot). This means that the lift force generated by the sails is more aligned with the thrust direction than the drag force. Under such conditions the thrust is maximised by trimming the spinnaker to give its maximum lift while still providing significant drag and the mainsail to provide approximately equal amounts of lift and drag. It is therefore concluded that America’s Cup downwind sails need to act both as vertical wings and horizontal parachutes. r 2001 Elsevier Science Ltd. All rights reserved.

Optimum Spanloads Incorporating Wing Structural Weight

Abstract: The classic minimum induced drag spanload is not necessarily the best choice for an aircraft. Here, a discrete vortex method which finds the minimum induced drag in the Trefftz plane has been used to calculate optimum spanloads for non-coplanar multisurface configurations. The method includes constraints for lift coefficient, pitching moment coefficient and wing root bending moment. The wing root bending moment constraint has been introduced so that by holding wing geometry fixed, changes in wing weight can be related to variations in spanload distributions. Changes in wing induced drag and weight were converted to aircraft total gross weight and fuel weight benefits, so that the spanloads that give maximum take-off gross weight reduction can be found. Results show that a reduction in root bending moment from a lift distribution that gives minimum induced drag leads to more triangular spanloads, where the loads are shifted towards the root, reducing wing weight and increasing induced drag. A slight reduction in root bending moment is always beneficial, since the initial increase in induced drag is very small compared to the decrease in wing weight. Total weight benefits were studied for a B-777 type configuration, obtaining take-off gross weight improvements of about 1% for maximum range missions. When performing reduced-range missions, improvements can almost double. A long range, more aerodynamically driven aircraft like the B-777 will experience lower benefits as a result of increasing drag. Short to medium range aircraft will profit the most from more triangular lift distributions.

Aerodynamics of Low Aspect Ratio Wings at Low Reynolds Numbers with Applications to Micro Air Vehicle Design and Optimization

Abstract: : The recent interest in the development of small UAVs and micro air vehicles has revealed a need for a more thorough understanding of the aerodynamics of small airplanes flying at low speeds. In response to this need, the present work provides a comprehensive study of the lift, drag, and pitching moment characteristics of wings of low aspect ratio operating at low Reynolds numbers. Wind tunnel tests of wings with aspect ratios between 0.5 and 2.0 and with four distinct wing planforms have been conducted at chord-Reynolds numbers in the range of 70,000 to 200,000. In addition, the effect of leading edge shape and fuselage bodies has been studied. As an example of an application of this wind tunnel data, the experimental results are used as part of an aerodynamic analysis procedure. This procedure is incorporated into a genetic algorithm optimization program that generates optimum MAV configurations given certain requirements and constraints. Results obtained by use of this optimization procedure have revealed that useful and accurate design-optimization tools can be developed based on the experimental data presented within this report.

Wind-Tunnel Investigations of Blunt-Body Drag Reduction Using Forebody Surface Roughness

Abstract: This paper presents results of wind-tunnel tests that demonstrate a novel drag reduction technique for blunt-based vehicles. For these tests, the forebody roughness of a blunt-based model was modified using micomachined surface overlays. As forebody roughness increases, boundary layer at the model aft thickens and reduces the shearing effect of external flow on the separated flow behind the base region, resulting in reduced base drag. For vehicle configurations with large base drag, existing data predict that a small increment in forebody friction drag will result in a relatively large decrease in base drag. If the added increment in forebody skin drag is optimized with respect to base drag, reducing the total drag of the configuration is possible. The wind-tunnel tests results conclusively demonstrate the existence of a forebody dragbase drag optimal point. The data demonstrate that the base drag coefficient corresponding to the drag minimum lies between 0.225 and 0.275, referenced to the base area. Most importantly, the data show a drag reduction of approximately 15% when the drag optimum is reached. When this drag reduction is scaled to the X-33 base area, drag savings approaching 45,000 N (10,000 lbf) can be realized.

Aerodynamic drag reduction technologies : proceedings of the CEAS/DragNet European Drag Reduction Conference, 19-21 June 2000, Potsdam, Germany

Abstract: Session 1 The Challenge of Drag Reduction.- Perspectives for the Future of Aeronautics Research (invited).- The Importance of Aerodynamics in the Development of Commercially Successful Transport Aircraft (invited).- Drag Reduction: A Major Task for Research (invited).- Session 2 Achievements of Technology Demonstration.- Airbus A320 HLF Fin Flight Tests (invited).- Lessons Learned from Dassault's Falcon 900 HLF Demonstrator (invited).- Flight Testing of a HLF Wing with Suction, Ice-Protection and Anti-Contamination Systems.- Session 3 HLF Design & Suction System Integration.- New Aerodynamic Approach to Suction System Design.- Retrofit Studies based on Airbus A310 for HLFC Application on Aircrafts.- Light Transmission Control Technique and Correlation with Pressure Loss Characteristics of Perforated Panels for Hybrid Laminar Flow Applications.- DNS Study of Suction Through Arrays of Holes in a 3-D Boundary-Layer Flow.- Session 4 HLF Operational Aspects.- Saab 2000 In-Service Test of Porous Surfaces for HLFC.- Flight Operational Assessment of Hybrid Laminar Flow Control (HLFC) Aircraft.- Application of HLF Technology to Civil Nacelle (invited).- Session 5 Supersonic Flow Control Aspects.- European Research to Reduce Drag for Supersonic Transport Aircraft (invited).- Attachment Line Transition in Supersonic Flow.- Laminar Design for Supersonic Civil Transport.- Session 6.1 Transition Prediction.- Linear Stability Theory Applied to Natural and Hybrid Laminar Flow Experiments.- Modern Transition Prediction Techniques Based on Adjoint Methods.- Receptivity Processes and Transition Scenarios for Swept-Wing Flows with HFL Technology.- Influence of Acoustic Excitation on 3D Boundery Layer Instabilities.- Session 6.2 HLF Experimental Techniques.- Study of Wind Tunnel Simulation Methodology for HLFC Wings.- Transitional Flow Physics and Flow Control for Swept Wings: Experiments on Boundary-Layer Receptivity, Instability Excitation and HLF-Technology.- New Developments in Surface Flow Sensor Technology within the Framework of AEROMEMS.- Session 7 Realisation of Adaptive Wing Concepts.- Aspects of Shock Control and Adaptive Wing Technology (invited).- Drag Reduction on Gurney Flaps and Divergent Trailing Edges.- Next Steps Envisaged to Improve Wing Performance of Commercial Aircraft.- Session 8 Future Prospects.- Industrial Perspectives of Drag Reduction Technologies (invited).- Session 9 Turbulent Drag Reduction Methods.- Current Status and Prospects for Turbulent Flow Control (invited).- Yaw Angle Effects on Optimized Riblets.- Effects on the Resistance and on the Separation of V Shapes Passive Manipulators in a Turbulent Boundary Layer.- Session 10 Separation Control & Induced Drag Reduction.- Active Control of a Laminar Separation Bubble.- Induced Drag Reduction with the WINGGRID Device.- The Lifting System with Minimum Induced Drag.- Poster Session 1.- Assessment of the eN Method as a Transition Prediction Tool for Zero Pressure Gradient Flows with and without Boundery Layer Suction.- Active Control of Nonlinear Disturbances in 2-D Boundary Layers.- Control of the Coherent Vortical Structures of a Boundery Layer.- Novel Approaches to Combat Insect Contamination on HLFC Wings.- Evolutionary Search Algorithms for Boundary Layer Control.- Poster Session 2.- Experimental Investigations on Active and Dynamic Instability Control of Separated Turbulent Wing/Cylinder Flows.- Drag Reduction in Airfoils Using Control Devices in the Shock Wave-Boundary Layer Interaction Region.- Numerical Study of Separation Control by Movable Flaps.

Methods for creating large scale focused blade deflections

Abstract: Methods are disclosed to design resilient hydrofoils ( 164 ) which are capable of having substantially similar large scale blade deflections under significantly varying loads. The methods permit the hydrofoil ( 164 ) to experience significantly large-scale deflections to a significantly reduced angle of attack under a relatively light load while avoiding excessive degrees of deflection under increased loading conditions. A predetermined compression range on the lee portion of said hydrofoil ( 164 ) permits the hydrofoil ( 164 ) to deflect to a predetermined reduced angle of attack with significantly low bending resistance. This predetermined compression range is significantly used up during the deflection to the predetermined angle of attack in an amount effective to create a sufficiently large leeward shift in the neutral bending surface with the load bearing portions of the hydrofoil ( 164 ) to permit the hydrofoil ( 164 ) to experience a significantly large increase in bending resistance as increased loads deflect the hydrofoil ( 164 ) beyond the predetermined reduced angle of attack. The shift in the neutral bending surface causes a significant increase in the elongation range required along an attacking portion of the hydrofoil ( 164 ) after the predetermined angle of attack is exceed. Methods are also disclosed for designing the hydrofoil ( 164 ) so that it has a natural resonant frequency that is sufficiently close the frequency of the reciprocating strokes used to attain propulsion in an amount sufficient to create harmonic wave addition that creates an amplified oscillation in the free end of the reciprocating hydrofoil ( 164 ). Methods are also disclosed for focusing energy storage and blade deflections along focused regions of load bearing members and the hydrofoil ( 164 ). Methods are also disclosed for reducing induced drag vortex formation along the lee surface of the hydrofoil ( 164 ), reducing drag and increasing the formation of lift forces.

Analytical Prediction of Trajectories of Ice Pieces after Release in an Airstream

Abstract: lf ice accretes on an airframe ahead of the turbine engines, it may cause engine damage if it enters the inlet after being shed. An analytical method was developed to predict the trajectories of ice pieces that are shed from an airframe into the surrounding flow field to determine if there is a probability that the ice will enter the engine. The approach taken in this study was to model the aerodynamics and kinematics of a piece of ice. After the ice enters the flow field, the aerodynamic forces and moments will cause it to tumble, accelerate across the streamlines, and accelerate to the freestream velocity. Gravity will also influence the trajectory. By systematically varying the size, thickness, initial angle of attack, and rotational damping, it is possible to predict a maximum locus of trajectories for a given geometry and flight condition. NOMENCLATURE C Pitch damping coefficient CD Drag coefficient CL Lift coefficient D Drag L Lift m Mass of ice piece S Area of ice piece V Relative speed VT Airplane true airspeed Vx Speed in the longitudinal direction VY Speed in the lateral direction Copyright © 2000 The American Institute of Aeronautics and Astronautics, Inc. AH rights reserved. Vz Speed in the vertical direction X Location along the longitudinal axis y Incidence angle of the ice piece p Air density 9 Angular position of the ice piece 9 Angular velocity of the ice piece 9 Angular acceleration of the ice piece INTRODUCTION When ice accretes on an airplane surface upstream of an engine inlet, it is possible for pieces of ice to damage the engine if the ice is shed, usually by a deicing system or an increase in ambient temperature above the freezing temperature. No damage will occur if the pieces of ice are small or if the trajectory is such that the ice pieces miss the engine inlet. When designing airplanes, it often necessary to predict the trajectories of released pieces of ice. This paper presents a method of predicting the trajectories of individual pieces of ice after they leave the surface upon which they accrete. It is well known that when ice breaks off a surface, an individual piece of ice will traverse an apparently random path downstream. Slow motion views of ice leaving a surface in flight show that the pieces often begin rotating as soon as they leave the surface and continue rotating in a nearly stable fashion until they disappear from view. Sometimes the pieces appear to randomly tumble upon release. Even though the movement of an individual piece of ice appears random, the locus of the trajectories is confined to a relatively small region of space (References 1 and 2). This locus can be determined by computing the 1 American Institute of Aeronautics & Astronautics c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. trajectories of individual ice pieces that represent the sizes of concern and the range of probable aerodynamic characteristics. The approach taken in this study was to create a simple model of the aerodynamic and kinematic characteristics of a generic square piece of ice with area and thickness as the independent variables. This shape was chosen because it is capable of generating significant lift forces. A random shape that approximates a spherical lump will be influenced primarily by drag and will therefore be accelerated down stream and fall due to gravity. There will be little lateral dispersion. The square piece of ice is released into a uniform flow field with an initial angle of attack such that the lift is in the lateral direction and the axis of rotation remains fixed in the vertical direction. Rotation of the ice shape is calculated based on the aerodynamic pitching moment, aerodynamic damping, and the moment of inertia. The initial forces acting on the piece will be lift, drag, pitching moment, and gravity. The square pieces will be unstable in pitch because the location of the center of pressure will be ahead of the center of mass. Therefore, the piece of ice will begin to rotate. As it rotates, the angle of incidence between the ice piece and the relative wind changes as it accelerates and moves downstream. This entire dynamic process was modeled to calculate individual trajectories. The assumption that rotational motion is confined to the vertical axis is conservative because if the piece of ice begins to roll or yaw, as it is likely to do, it will decrease the forces perpendicular to the assumed axis of rotation and reduce the dispersion distance in that direction. AERODYNAMIC FORCES AND MOMENTS The aerodynamic forces that act on the ice shape are lift and drag. The distribution of these forces along the surface of the ice shape determines the pitching moment applied to the piece of ice. Lift and drag are calculated using the standard equations Lairplane. As the piece of ice moves downstream, its relative velocity decreases and changes direction. The lift coefficient was based on the lift produced by a low aspect ratio flat plate as described by Hoerner in Reference 1. At zero angle of attack, the lift produced is zero. Low aspect ratio flat plates typically have a low lift curve slope and high stalling angle of attack because the flow around the edges of the plate dominates the flow field. As the angle of attack increases, separation on the suction side occurs, but the plate still produces some lift. At an angle of attack of 90 degrees, no lift is produced. For this study, the following was chosen to describe the lift curve: CL = 1.31 tan^, for y 35° Drag is clearly related to angle of attack. At a low angle of attack, the drag due to lift dominates; at a high angle of attack, pressure drag dominates. To model these characteristics with a continuous relationship, the following relationship was used: CD = 1.25(0.01 + 0.531CL), for y 35° These equations are based on fundamental drag relationships from Hoerner in Reference 2. There also exists experimental data on tumbling objects. Raglan, et al, give overall drag values for objects injected into a furnace in Reference 4. To match these data, a multiplier (1.25 in this case) is applied to these equations so that continuous rotation of a simulated object will give the same drag as the comparable shape as reported by Raglan, et al. The pitching moment calculation depends on the location of the center of pressure on the piece of ice. At low angles of attack, the center of pressure is located at the quarter chord point, but as the angle of attack increases, the center of pressure moves aft. The following equations were used to describe the pitching moment relation: When the piece of ice is initially released into the airstream, its relative velocity is the speed of the American Institute of Aeronautics & Astronautics c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

Wind tunnel tests to reduce aerodynamic drag of trains by smoothing the under-floor construction

Abstract: As a train runs at a higher speed, aerodynamic drag increases. On long train-sets such as Shinkansen trains, the aerodynamic drag is mainly generated by intermediate vehicles. In the previous researches, we proved that smoothing the under-floor construction reduces the aerodynamic drag. To investigate the mechanism of this effect, we performed wind tunnel tests with train models consisting of three vehicles (representing head, intermediate and tail vehicles) and measured the aerodynamic drag and the pressure distribution on the intermediate vehicle. Test results show that the reduced aerodynamic drag is mainly the effect of decreases in the pressure drag around bogies.

Computational Study of Influences of a Seam Line of a Ball for Baseball on Flows

Abstract: Flows around a ball used in baseball games are calculated using third-order upwind-difference method with various seam positions determined by two rotation angles Those are four-seam rotation with an angle: a and two-seam rotation with an angle: b The computed results of the four-seam rotation are compared with experimental data measured in a wind tunnel and computed drag coefficients qualitatively agree well with experiments However, lift coefficients do not agree well The computed results and geometrical symmetry suggest that a supporting rod in the wind tunnel would have strong influence on the accuracy of the measurement Flow changes in two-seam rotation are also simulated It is found that the lowest drag force is observed at b=90 and that the value is less than half of the largest drag force at a=30 and 60 degrees The largest lift force is observed at b=20 degree In this case, a projection of the seam line on the top causes a large separation while smooth surface without the seam at the bottom dose not separate the flow A pair of longitudinal vortices are found in the wake, which make wake slant and generate large lift force

Physical Properties of Maskell's Induced Drag Integral

Abstract: This study investigates the physical properties of Maskell's induced drag integral, which is commonly used in the wake survey analysis method to predict induced drag of airplane models. This study also examines the reasons as to why Maskell's induced drag integral performs better than the well-known induced drag integral from the classic lifting line theory. Although the induced drag integral from the classic lifting line theory can be derived from Maskell's induced drag integral with a planar wake approximation, it does not necessarily mean that the integrand of Maskell's induced drag integral is identical to the induced drag distribution of the lifting line theory. For elliptically loaded wings, the induced drag distribution (from the lifting line theory) along the wing span is proportional to the local circulation strength F(t/). Maskell's integrand is proportional to the first moment of the slope of circulation, ydT/dy. Because of this physical property of Maskell's induced drag integral, the impacts on the induced drag calculation by the inaccurate lift distribution near the model symmetry plane (y ~ 0) are minimized. In general, it is difficult to obtain an accurate lift distribution of the model near its symmetry plane, because of the interference from the model's fuselage and support struts. Some of the wake survey results of a high-speed model test at the Boeing Transonic Wind Tunnel are presented to demonstrate the superiority of Maskell's induced drag integral over the induced drag integral from the lifting line theory. flow variables measured at a downstream survey station. The problem with the kinetic energy form of induced drag, which is derived from the integral momentum equations, is that the integral must be evaluated over the entire downstream cross section of the wind tunnel. Once Maskell [1] demonstrated that an integral formulation for induced drag can be obtained from flow variables measured within the model wake region, the practical application of quantitative wake surveys became possible [2,3,4,5]. After Maskell introduced his new induced drag integral, however, the clean physical meaning behind the original kinetic energy form was lost. In this brief paper we investigate the physical properties of Maskell's induced drag integral, and we seek the reasons as to why Maskell's induced drag integral performs better than that from the lifting line theory. It is worth noting that the well-known induced drag formula can be derived from Maskell's induced drag integral by applying the planar wake approximation [6,7]. To demonstrate the superiority of Maskell's induced drag integral over the induced drag integral from the lifting line theory, a few selected results from a highspeed model test (for the study of the effects of flap trap fairings on the airplane drag) at Boeing Transonic Wind Tunnel are used. Because this paper focuses on the properties of Maskell's integral, a detailed explanation of the experimental setup, including the dataacquisition and -reduction systems, are not discussed. Introduction Maskell's Integrand Generally, the forces acting upon a model can be derived by applying the integral form of the momentum equations to a control volume containing the model. Model drag and lift can be written as integrals of the Copyright ©2001 by The Boeing Company. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. Maskell's induced drag integral wA (1) which is commonly used in wake survey analysis method [1,2,3,4,5,6,7,8] can be derived from the kiAmerican Institute of Aeronautics and Astronautics c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. netic energy form of induced drag [1,9,10] (v +w)dydz (2) Here, we assumed the x— and z—axes to be aligned with the undisturbed freestream direction and lift direction, respectively. The remaining y—axis is normal to both x— and z—axes. 52 and WA represent transverse plane located downstream of model and model wake region on it, respectively (Figure 1). Wake Survey • Region Fig. 1 Control volume and coordinate system. For the calculation of Maskell's scalar function ^(y, z), a Poisson equation , dy dz (3) should be solved over the entire transverse plane (S2) with the boundary condition V = 0 (4) along the tunnel intersection [1,9,10] (see Figure 1). f represents the x-component of the vorticity vector, which is defined as (5) and £ = 0 outside of the model wake region WAFor the case of a half model test (including a full model test case without yaw), the boundary condition, ip = 0, will be also satisfied at the model symmetry plane (y — 0) [7]For the "free-air" case (without interference of tunnel walls) the boundary condition (4) should be evaluated at y + z -»> oo, and Maskell's scalar function, T/>, can be expressed using a Green's function in the following analytical form [6,7,8]. 1 It is apparent that the induced drag integral in kinetic energy form (which is given by equation (2)) has a clear physical meaning by itself: induced drag is determined by the total amount of kinetic energy loss that has been consumed to generate cross-flow velocities in the wake field. On the other hand, it is difficult to see clear physical meanings in Maskell's induced drag integral (equation (1)). In particular, when the integrand of Maskell's equation is plotted along the wake span, it is not clear what the integrand signifies. For the purpose of examining the physical properties of the integrand of Maskell's induced drag integral, we define "Maskell's integrand" as (7) ZWA denotes an integral region of z (for a given value of y) on the model wake WA> Then Maskell's induced drag integral can be simplified to -r JyL diM(y)dy (8) where yi and y# represent the minimum and maximum values of y on the model wake WA, respectively (see Figure 2).

Aerodynamic study for the ground effect of ski jumping

Abstract: We investigated the aerodynamic forces just before taking telemark of the landing phase. The full size model was employed to measure the lift area, the drag area and the moment volume, which was mounted in a 3-meter low speed wind tunnel. The ground plate was set in the test section of the wind tunnel. The height between the ground plate and the toe of the model was from 0.4 m to 1.0 m. In the case of the V style flight, the lift area with the ground plate is always larger than that without the ground plate, though the drag area with the ground plate is comparable to that without the ground plate. The ground effect of V style flight contributes to making the larger lift in the latter half of the flight. In the case of the parallel style, the lift and the drag areas with the ground plate are comparable to that without the plate.

A Model Rotor in Axial Flight

Abstract: A model rotor was mounted horizontally in the settling chamber of a wind tunnel to obtain performance and wake structure data under low climb conditions. The immediate wake of the rotor was carefully surveyed using 3-component particle image velocimetry to define the velocity and vortical content of the flow, and used in a subsequent study to validate a theory for the separate determination of induced and profile drag. Measurements were obtained for two collective pitch angles intended to render a predominately induced drag state and another with a marked increase in profile drag. A majority of the azimuthally directed vorticity in the wake was found to be concentrated in the tip vortices. However, adjacent layers of inboard vorticity with opposite sense were clearly present. At low collective, the close proximity of the tip vortex from the previous blade caused the wake from the most recent blade passage to be distorted. The deficit velocity component that was directed along the azimuth of the rotor blade was never more that 15 percent of the rotor tip speed, and except for the region of the tip vortex, appeared to have totally disappeared form the wake left by the previous blade.

Virtual Experiments on Drag Reduction

Abstract: This paper presents numerical simulations of external flow over a flat porous surface with blowing cross-flow thatresult in significant drag reduction at the interface between the boundary layer and the solid wall. Throughparametric numerical experiments it is demonstrated that it is fundamentally possible to achieve a reduction in theshear force on the wall, if in the blowing (injection) process the velocity of the flow is controlled in conjunction withthe free stream velocity and the nature of surface porosity. The results presented here demonstrate that the blowingflow uniformity is very critical to the drag reduction magnitude and even appearance. Non-uniformity is possibledue to interaction of cross flow and free-stream flow. Further based on the analysis of the parametric results, theauthors offer an explanation of the mechanism that apparently leads to the drag reduction phenomenon. Introduction Skin friction and separated flows overexternal surfaces penalize the performance andeconomics of airplanes, ships, cars, and in internalflow affect pressure drops and pump penalties in oiland gas pipelines. Generally they can affect theperformance of any manufacturing process thatemploys long piping runs, or fluid flows that becomeunstable. For aircraft, the reduction of turbulent drag isa long sought goal. In particular, it has been estimatedthat aircraft fuel costs per-mile could drop up to 40%,if the flow around aircraft could be 'smoothed out'.In the last few years there has beensignificant renewed interest in finding new andeconomic ways to control and reduce skin friction andassociated aerodynamic forces exerted on movingvehicles and in particular aircraft, surfaces ships andsubmarines. A variety of

Numerical simulation of magnus lift on a coal particle

Abstract: Magnus lift on a coal particle has been investigated numerically. An empirical equation has been obtained to describe both effects of the rotational speed and Reynolds number on the lift force on a sphere. It can be found from the comparison between both of drag and Magnus forces, that,as to a coal particle with high rotation speed, such as 1800 r/min, the magnitude of Magnus lift is only 1% of drag force. Therefore, neglecting Magnus force in the calculation of coal particle is really reasonable.

A Numerical and Experimental Aerodynamic Analysis of an Inboard-Wing/Twin-Fuselage Configuration

Abstract: In this paper, we present and compare numerical and experimental studies of the flowfield around a configuration consisting of an inboard wing mounted between twin fuselages. The results of both studies show that behind the doublefuselage configuration a "virtual wing-tip vortex system" forms. The trailing vortex system for the twin-fuselage configuration is shed from both the fuselages as well as the trailing edge of the wing. The vorticity shed from the fuselages combines with the vorticity from the trailing edge in a manner that is very similar to what happens in a conventional single-fuselage configuration. The numerical and experimental results are in qualitative agreement. Introduction The aviation industry has been urged to develop faster and/or bigger airplanes and to fly them closer together as part of a large effort to meet the needs of an ever-increasing air transportation system. One of the innovative ideas for a large airplane comes from the NASA Langley Research Center. Spearman and Feigh suggested a configuration that differs from a conventional wing-body-tail design. Instead of one fuselage located in the center of the configuration, twin fuselages are placed at the tips of an inboard wing. The intent is "to provide an increase in payload capacity without an increase Graduate Research Assistant Currently Prof., Univ. Nacional, Rio Cuarto, Argentina N. Waldo Harrison Prof., Assoc. Fellow AIAA Fred D. Durham Chair, Fellow AIAA Copyright© American Institute of Aeronautics And Astronautics, All Rights Reserved in overall length and width when compared to current designs and to achieve two-dimensional flow on the wing by eliminating free wing tips so that the wing tip flow that produces an induced drag and a hazardous trailing vortex would not exist." One aerodynamic model' has been tested in Virginia Tech's Stability Wind Tunnel to study this inboard-wing, twin-fuselage concept. The configuration is represented schematically in Figure 1. In addition, the configuration represented in Figure 1 has been modeled aerodynamically. Here we present some comparisons between the numerical and experimental data and respond to some of Spearman's and Feigh's comments. To model the flowfield, we use (1) the wellknown code PMARC, ignoring the deformations of the wake but accounting for the aerodynamic interference, and (2) the general unsteady vortex-lattice method (VLM), accounting for the deformations of the wakes and the aerodynamic interference among the various components of the configuration. The version of VLM that we use in this work has been developed at Virginia Tech over several years 16 to solve for the incompressible flow over both lifting and nonlifting bodies. The method utilizes a lattice of discrete vortex lines distributed over the surfaces of the configuration. The circulations around the discrete vortex segments are obtained by imposing the no-penetration condition on the body surfaces. Imposing the unsteady "Kutta condition" along the edges where separation occurs provides the vorticity-shedding rate, and convecting the vorticity in the wake with the local particle velocity renders the wake forcefree. The advantages of this method are that it also satisfies the no-slip condition and readily provides the surface velocity and pressure, and it

Drag Reduction on Gurney Flaps and Divergent Trailing Edges

Abstract: Miniflaps at the trailing edges of airfoils (e.g., Gurney flaps or divergent trailing edges) change the Kutta condition and thus produce higher lift. Unfortunately, however, the drag is also increased due to the flow separation downstream of this particular type of trailing edge. Therefore, the trade-off between beneficial and detrimental effects is considered in this paper. Various aspects of the flow on airfoils with Gurney flaps are addressed: (1) Transonic flow. Wind tunnel experiments have been carried out with a CAST 10–2/DOA2 airfoil with Gurney flap and at high subsonic flow speed. The lift to drag ratio is improved and the test wing behaves like one having a 20% larger surface area. In addition, buffeting becomes less critical. (2) Selection of flap size. Detailed wind tunnel studies have been carried out with a low-drag glider wing at low subsonic velocities. The Gurney flap height was varied in six steps so that the most relevant parameter regime was covered. For the lift increase and for the device drag, simple empirical laws were obtained. Subsequently, for practical applications, a procedure for the selection of a suitable flap height providing a beneficial effect has been devised. (3) Drag reduction by wake stabilization. In the separation regime downstream of a Gurney flap, an absolute instability, i.e., a Karman vortex street occurs, even if the incident boundary layers are turbulent. Therefore, one approach towards drag reduction is to stabilize the wake and hence eliminate the Karman vortex street. This can be achieved with a variety of trailing edge modifications, e.g., with slits or holes in the Gurney flap. A particular structure which exhibits spade-like protrusions at the trailing edge produces also good results. Actually, the latter structure has been adopted from the trailing edge of dragonfly wings. Eliminating the Karman vortex street with these various trailing edge modifications reduces the Gurney flap device drag by about 22 – 30% without a perceivable change of the enhanced lift. Obviously, vibration and noise radiation are also reduced together with the suppression of the Karman vortex street.

Aerodynamic performance improvement by divergent trailing edge modification to a supercritical airfoil

Abstract: A computational study has been performed to determine the effects of divergent trailing edge (DTE) modification to a supercritical airfoil in transonic flow field. For this, the computational result with the original DLBA 186 supercritical airfoil was compared to that of the modified DLBA 283. A Navier-Stokes code, Fluent 5. 1, was used with Spalart-Allmaras’s one-equation turbulence model. Results in this study showed that the reduction in drag due to the DTE modification is associated with weakened shock and delayed shock appearance. The decrease in drag due to the DTE modification is greater than the increase in base drag. The effect of the recirculating flow region on lift increase was also observed. An airfoil with DTE modification achieved the same lift coefficient at a lower angle of attack while giving a lower drag coefficient. Thus, the lift-to-drag ratio increases in transonic flow conditions compared to the original airfoil. The lift coefficient increases considerably whereas the lift slope increases just a little due to DTE modification.

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.