Showing papers on "Lift-induced drag published in 2009"

Plasma Flaps and Slats: An Application of Weakly Ionized Plasma Actuators

Abstract: The experimental validation of an application of weakly-ionized plasma actuators for improved aerodynamic performance of multi-element wings and wings with movable control surfaces is presented. Two spanwise arrays of plasma actuators, configured to produce a wall-jet effect, were applied on the suction surface of a two-dimensional NACA 0015 wing model, one at the leading edge and the other near the trailing edge to mimic the effects of a wing leading-edge slat and a trailing-edge flap, respectively. Flow control tests were conducted at chord Reynolds numbers, corrected for blockage, of 0.217 x 10 6 and 0.307 x 10 6 in a low-speed wind tunnel at the University of Notre Dame. The leading-edge-separation control resulted in an increase in both the maximum lift coefficient and the stall angle of attack and a lift-to-drag improvement of as much as 340%. An optimum frequency was found to exist for unsteady excitation of the leading-edge separation. Under this condition, the power to the actuator was estimated to be only 2 W. The trailing-edge actuator was found to produce the same effect as a plain trailing-edge flap. This included a uniform shift at all angles of attack of the lift coefficient and a shift toward higher lift coefficients of the drag bucket. In addition, there was a slight decrease in the minimum drag coefficient. The obvious advantages of this approach are its simplicity, as there are no moving parts, and its lack of hinge gaps, which add drag. An example of their use as ailerons for roll control produces a comparable roll moment coefficient to a sample general aviation aircraft.

Power Balance in Aerodynamic Flows

Abstract: A control volume analysis of the compressible viscous flow about an aircraft is performed, including integrated propulsors and flow-control systems. In contrast to most past analyses that have focused on forces and momentum flow, in particular thrust and drag, the present analysis focuses on mechanical power and kinetic energy flow. The result is a clear identification and quantification of all the power sources, power sinks, and their interactions, which are present in any aerodynamic flow. The formulation does not require any separate definitions of thrust and drag, and hence it is especially useful for analysis and optimization of aerodynamic configurations that have tightly integrated propulsion and boundary-layer control systems. Nomenclature b, c = wingspan and chord CD = dissipation coefficient Cf = skin friction coefficient Di = induced drag Dp = profile drag Dw = wave drag dS = surface element of control volume dV = volume element of control volume _ Ea = axial kinetic energy deposition rate _ Ep = pressure-work deposition rate _ Ev = transverse (vortex) kinetic energy deposition rate _ Ew = lateral wave-outflow energy deposition rate Fn = streamwise force from lateral outflow velocity Vn Fu = streamwise force from axial velocity u Fv = streamwise force from transverse velocities v, w Fx, Fz = total streamwise, normal aerodynamic forces

Plasma-induced force and self-induced drag in the dielectric barrier discharge aerodynamic plasma actuator

Abstract: §Through high-time-resolution laser interferometry, we observe the motion of a test article under the influence of a dielectric barrier discharge aerodynamic plasma actuator, and from this motion we deduce the time history of the force produced by the actuator to a resolution substantially smaller than the period of the actuator’s AC cycle. We find that the negative- and positive-going half cycles of the plasma discharge produce a force on the surrounding neutral air in the same direction, but that only the negative-going half cycle produces a force sufficient to substantially overcome the drag induced by accelerating the air in the immediate vicinity of the aerodynamic surface (within the boundary layer under normal circumstances).

Best wing system: an exact solution of the Prandtl’s problem

Abstract: In 1924, Ludwig Prandtl published a fundamental paper in which he showed that the lifting system with minimum induced drag is a box wing. The results were obtained through an approximate procedure. In the present paper we obtain an exact solution of Prandtl’s problem. In particular, we prove that the lift on the horizontal wings results of the superposition of a constant and an elliptical distributions and, on the vertical wings, it is butterfly shaped, as already shown by Prandtl. The discrepancies between the two solutions are discussed.

Formation mechanism of aerodynamic drag of high-speed train and some reduction measures

Abstract: Aerodynamic drag is proportional to the square of speed. With the increase of the speed of train, aerodynamic drag plays an important role for high-speed train. Thus, the reduction of aerodynamic drag and energy consumption of high-speed train is one of the essential issues for the development of the desirable train system. Aerodynamic drag on the traveling train is divided into pressure drag and friction one. Pressure drag of train is the force caused by the pressure distribution on the train along the reverse running direction. Friction drag of train is the sum of shear stress, which is the reverse direction of train running direction. In order to reduce the aerodynamic drag, adopting streamline shape of train is the most effective measure. The velocity of the train is related to its length and shape. The outer wind shields can reduce train’s air drag by about 15%. At the same time, the train with bottom cover can reduce the air drag by about 50%, compared with the train without bottom plate or skirt structure.

Ideal Lift Distributions and Flap Angles for Adaptive Wings

Abstract: An approach is presented for determining the optimum flap angles and spanwise loading to suit a given flight condition. Multiple trailing-edge flaps along the span of an adaptive wing are set to either reduce drag in rectilinear flight conditions or to limit the wing bending moment at maneuvering conditions. For reducing drag, the flaps are adjusted to minimize induced drag, while simultaneously enabling the wing sections to operate within their respective low-drag ranges. For limiting wing bending moment, the flaps are used to relieve the loading near the wing tips. An important element of the approach is the decomposition of the flap angles into a distribution that can be used to control the spanwise loading for induced-drag control and a constant flap that can used for profile-drag control. The problem is linearized using the concept of basic and additional lift distributions, which enables the use of standard constrained-minimization formulations. The results for flap-angle distributions for different flight conditions are presented for a planar and a nonplanar wing. Postdesign analysis and aircraft-performance simulations are used to validate the optimum flap-angle distributions determined using the current approach.

Drag Characteristics for Optimally Span-Loaded Planar, Wingletted, and C Wings

Abstract: This paper focuses on the drag characteristics of optimally span-loaded planar, wingletted, and C wings. The span load is optimized resulting in minimum induced or total drag. The wing-root bending moment is kept constant for all analyzed wings to ascertain that different wings have comparable weight. The optimum span loadings for the different types of wing are calculated using a fast and simple numerical method. The wings are analyzed in the Trefftz plane, infinitely far behind the wing. Lagrange multipliers are used to calculate the optimum span loading resulting in minimum induced or total drag, with the wing-root bending moment and/or the lift coefficient as constraint. The induced drag can be calculated using the optimum span loading. The profile drag is assumed to be a function of the local lift coefficient. The results indicate that the C wing does not have real aerodynamic performance advantages compared to a wingletted wing. For wings with span and/or aspect ratio constraints, a winglet offers a drag reduction relative to a planar wing.

Aerodynamic and Aeroacoustic Properties of a Flatback Airfoil: An Update

Abstract: Results from an experimental study of the aerodynamic and aeroacoustic properties of a atback version of the TU Delft DU97-W-300 airfoil are presented for a chord Reynolds number of 3 10 6 . The data were gathered in the Virginia Tech Stability Wind Tunnel, which uses a special aeroacoustic test section to enable measurements of airfoil self-noise. Corrected wind tunnel aerodynamic measurements for the DU97-W-300 are compared to previous solid wall wind tunnel data and are shown to give good agreement. Aeroacoustic data are presented for the atback airfoil, with a focus on the amplitude and frequency of noise associated with the vortex-shedding tone from the blunt trailing edge wake. The effect of a splitter plate attachment on both drag and noise is also presented. Computational Fluid Dynamics predictions of the aerodynamic properties of both the unmodied DU97-W-300 and the atback version are compared to the experimental data. Technical risks associated with the use of atback airfoils for the inboard region of wind turbine blades include increased aerodynamic noise and increased aerodynamic drag. Both of these penalties are the result of the blunt trailing edge shape and the wake that is produced by this shape. The relatively low pressure at the blunt base results in a much larger drag force than for a conventional airfoil shape. The effect of this drag penalty on rotor thrust and torque coefcient for typical inboard twist angles is not severe, and in fact can be offset by the additional lift that a atback airfoil generates. 3 Consideration of drag reducing devices such as splitter plates or trailing edge serrations may be desireable to further boost performance, however. The increased noise from the atback is due primarily to the vortex shedding phenomenon associated with bluff- body wakes. The vortex shedding often leads to tonal noise, similar to the Aeolian tones of o w past circular cylinders. The intensity of bluff-body vortex shedding tones at low Mach number scales with the sixth power of the relative o w velocity. Broadband aeroacoustic noise sources associated with turbulent boundary layer-trailing edge interaction scale with the fth power of the relative o w velocity. Since outboard o w velocities are much higher than those encoun- tered inboard, the overall aerodynamic noise levels of a rotor incorporating inboard atback shapes will likely continue to be dominated by outboard trailing edge noise. However, two aspects of the atback noise source may be cause for concern. First, the vortex-shedding noise from atbacks is likely to be contained in a relatively low-frequency band (50-200 Hz). Some community noise regulations have separate low-frequency noise standards apart from considera- tion of A-weighted sound, which emphasize higher frequencies to which the human ear is more sensitive. Second, the

The effect of chord-wise flexibility on the aerodynamic force generation of flapping wings: Experimental studies

Abstract: Wings of insects are flexible structures. Although there has been much recent progress in the area of insect flight aerodynamics, very little is known about how wing flexibility influences aerodynamic forces during flapping flight. We investigated this question using a dynamically scaled mechanical model of insect wings. Using a suite of wings with varying flexural stiffness (EI) values, we generated aerodynamic polar plots to characterize the force coefficients of flexible wings. These polar plots showed that the aerodynamic performance of the wings varied with wing flexibility. In general, aerodynamic force production decreased with increasing flexibility. Both lift and drag coefficients of wings are greater when wings are more rigid. However, at very high angles of attack, flexible wings generated greater lift than a rigid wing. In addition, the ratio of lift-to-drag also decreased with increasing flexibility. These data show that flexible wings offer no aerodynamic advantage over a rigid wing under steady state circumstances. Because wing material in insects is usually flexible but reinforced by wing veins, we tested the hypothesis that wing veins enhance the aerodynamic performance of wings by increasing their effective stiffness. Our data suggests that even a very basic framework of appropriately placed wing veins can substantially increase the functional rigidity of the wings thereby enhancing its aerodynamic performance.

Efficient Algorithms for Future Aircraft Design: Contributions to Aerodynamic Shape Optimization

Abstract: EFFICIENT ALGORITHMS FOR FUTURE AIRCRAFT DESIGN : CONTRIBUTIONS TO AERODYNAMIC SHAPE OPTIMIZATION Jason Edward Hicken Doctor of Philosophy Graduate Department of Aerospace Science and Engineering University of Toronto 2009 Advances in numerical optimization have raised the possibility that efficient and novel aircraft configurations may be “discovered” by an algorithm. To begin exploring this possibility, a fast and robust set of tools for aerodynamic shape optimization is developed. Parameterization and mesh-movement are integrated to accommodate large changes in the geometry. This integrated approach uses a coarse B-spline control grid to represent the geometry and move the computational mesh; consequently, the mesh-movement algorithm is two to three orders faster than a node-based linear elasticity approach, without compromising mesh quality. Aerodynamic analysis is performed using a flow solver for the Euler equations. The governing equations are discretized using summation-by-parts finite-difference operators and simultaneous approximation terms, which permit C mesh continuity at block interfaces. The discretization results in a set of nonlinear algebraic equations, which are solved using an efficient parallel Newton-Krylov-Schur strategy. A gradient-based optimization algorithm is adopted. The gradient is evaluated using adjoint variables for the flow and mesh equations in a sequential approach. The flow adjoint equations are solved using a novel variant of the Krylov solver GCROT. This variant of GCROT is flexible to take advantage of non-stationary preconditioners and is shown to outperform restarted flexible GMRES. The aerodynamic optimizer is applied to several studies of induced-drag minimiza-

Basic Induced Drag Study of the Joined-Wing Aircraft

Abstract: J OINED-WING configuration was considered for the first time by Prandtl in 1924 [1] The first attempt to build a joinedwing airplane was undertaken by the Moscow Institute of Aviation in the 1940s, when they redesigned a classic biplane called the Polikarpow Po-2 to create the joined-wing airplane [2] More recently, the configuration was studied by Wolkovitch [3] and Kroo and Gallman [4] in the 1980s A project described in [5] was inspired by these works It consisted of a series of wind-tunnel tests with models of a small transport airplane (see Fig 1) The most interesting result obtained in this project was that a greater L=Dwas achieved for negative angles of attack than for positive ones (see Fig 2) This observation suggests that the front wing of the aircraft in a joined-wing configuration should be located on top of the fuselage and the rear wing at the bottom to obtain greater aerodynamic efficiency This complicates the design because it is not possible to use the fin as a pylon for the rear wing and fuselages usually are not particularly high Anyway, this front-high-joinedwing configuration can be considered in cases of small 2–4 seat general aircraft, small transport aircraft, and very large commercial double-deckers Later, in the 1990s, a few papers on the joinedwing configuration were published [6], most of them focusing on mass-strength analysis

Streamlining the time trial apparel of cyclists: The Nike swift spin project

Abstract: This paper documents the development of aerodynamic apparel for the Tour de France individual time trial (TT), the Olympic TT, and track cycling races. A wind tunnel and metric balance were used to measure the drag force (Fd) and wind tunnel air velocity on cylinders, limb models, and live cyclists clad in samples or suits sewn with one or more of 200 stretch fabrics. A concurrent measurement of model dimensions and frontal areas provided the non‐dimensional drag coefficient (Cd) and Reynolds Numbers (Re) that characterized the ability of the various fabrics and suits to reduce frictional drag and induce a drag crisis (DC) or premature flow transition. DC defines a critical air velocity over the body segments at which the airflow transitions from laminar to turbulent, yielding a smaller wake behind the body segment and a corresponding decrease in Fd. A number of fabrics triggered DC on cylinders and limb segments, reducing cylinder and limb Cd by over 40 per cent. Several methods of lowering the Fd of cyc...

Aerodynamic Characteristics of Airfoil with Perforated Gurney-Type Flaps

Abstract: A low-speed wind-tunnel experiment was conducted to examine the aerodynamic and near-wake characteristics of a NACA 0012 airfoil equipped with perforated Gurney-type flaps of different heights and porosities at Re = 2.32 × 10 5 . The perforation-generated jet flows disrupted the periodic wake behind the flap, reducing the effective camber effects, and led to a reduced lift and drag coefficient as well as a less negative peak C m , compared with the solid flap. The magnitude of C l,max , C d , and C m,peak decreased with increasing amount of perforation. The decrease in drag, however, outweighed the loss in lift and rendered an improved lift-to-drag ratio. The reduction of drag alongside the lowered lift was found to be closely coupled to the characteristics of the unsteady wake. A significantly narrowed extent of the unsteady wake with a weakened fluctuating intensity was observed.

Drag reducing deflector

Abstract: A drag reducing deflector system is disclosed. The drag reducing deflector system provides aerodynamic drag reduction during crosswind flow (CF) conditions including zero crosswind flow conditions in which the crosswind flow angle, α, is 0.0. The crosswind flow angle α is measured from the vehicle longitudinal axis A that also defines the vehicle direction of motion. Embodiments of the drag reducing deflector system 100 may also be utilized to assist braking of a vehicle.

Aerodynamic Investigations of an Advanced Over-the-Wing Nacelle Transport Aircraft Configuration

Abstract: The transonic aerodynamic characteristics of an advanced, Over-the-Wing Nacelle (OWN), subsonic transport configuration are assessed using both inviscid Euler and viscous Navier-Stokes CFD. Results of the assessments are compared to a similar configuration with a more traditional Under-the-Wing Nacelle (UWN) installation and a similar reference Wing-Body (WB). The engine installation and inboard wing section of the OWN are designed to create a channel between the nacelle and fuselage that accelerates the upper surface wing flow, increases leading edge suction, and subsequently reduces the overall drag. CFD results were used to examine the span loadings of the three configurations and allow normalization of their wing twist distributions. Qualitative observations and quantitative drag computations are performed for the three configurations at a cruise Mach number of 0.78, an approximate altitude of 35,000 ft, and a CL of 0.44. It was found that the OWN’s inboard wing channel section is effective in producing favorable leading edge suction but that the overall drag, compared to the UWN, is higher by 32 counts. Although the source of this drag excess could not be precisely identified, it was isolated in the nacelle region. CFD solutions were obtained for several transonic Mach numbers to assess the drag rise characteristics of the three configurations. Inviscid Euler results showed similar trends to previous studies where the OWN is higher in drag at lower Mach numbers but enjoys a much milder drag rise than the UWN or WB and has lower drag at higher Mach numbers. Viscous Navier-Stokes results however showed that the effect of shock-induced flow separation significantly increases the drag rise but that the OWN still likely experiences it less severely than the UWN. Overall, the results of this study indicate that the Over-theWing Nacelle concept under consideration appears to be an aerodynamically competitive engine installation to the more traditional Under-the-Wing pylon installation.

Curved wing tip

Abstract: A wing incorporating a wing tip with a curved leading edge and a curved trailing edge to minimize induced drag for a given wing form is disclosed. The curve of the leading and trailing edges of the wing tip may be described generally as parabolic, elliptic, or super elliptic. A finite tip segment may be included with a sweep angle between the end of the curved leading edge and the end of the curved trailing edge. The wing tip may also include a spanwise camber.

Drag Effect of a Dented Surface in a Turbulent Flow

Abstract: ow conditions. Shallow dents were found to reduce the drag over the entire velocity regime, while deep lead to an increasing drag. Dents with intermediate depths cause drag reduction for low velocities only. The analysis of the CFD calculation shows that an asymmetrical ow exists leading to a skewed vortex that leaves the dents in lateral directly. The presence of this vortex may be of key importance for possible drag reduction of dents surfaced as published by other researchers.

Numerical Implications of Spanwise Camber on Minimum Induced Drag Configurations

Abstract: Lifting line and Euler techniques have indicated that an elliptical planform cambered in the spanwise direction appears to have a lower induced drag than the uncambered, planar configuration, when both of then have identical arc lengths and the same wetted areas. In other words, a wing geometry with a lower projected span than the planar wing, only because of spanwise cambering, has been found to have a lower induced drag than the corresponding planar case. It was also noted that in the Euler solutions the theoretical minimum induced drag was not achieved on a planar configuration without applying a twist distribution. The increased efficiency of the non-planar configuration is due to induced lift and without accounting for this induced lift, the induced drag is larger then that of the corresponding unfolded planar wing. Induced lift and aerodynamic efficiency of the spanwise cambered configuration was observed to increase with reduced aspect ratio.

Flight control method and apparatus to produce induced yaw

Abstract: A method of controlling an aircraft in a turn without the use of a rudder by producing induced yaw. Induced yaw being produced by creating a net induced drag differential between an inboard wing to the turn and an outboard wing to the turn, whereby the net induced drag differential overcomes adverse yaw produced by the outboard wing.

Effects of Gurney Flaps on an Annular Wing

Abstract: F OR a constrained wing span, nonplanar wings exhibit significant aerodynamic benefits. These are generally indicated through increased lift curve slopes and reduced drag due to lift. These benefits are accrued through thewing capturing a larger volume of air to generate the lift impulse, and consequently the downwash at a point is reduced [1]. A simple biplane is a clear embodiment of nonplanarity. As the spacing between the upper and lower wings tends to infinity, the lift curve slope of the biplanewing tends to twice that of the equivalent monoplane, and the lift dependent drag tends to half that of the monoplane. Naturally, an infinitely spaced biplane or even that with a significant gap is not a practical design solution. As the gap between the biplanewings reduces, mutual wing interference effects reduce the nonplanar benefits. An annular (or ring) wing achieves the same aerodynamic benefits as an infinitely spaced biplane but with a maximum vertical wing spacing of the diameter of the ring. However, annular wings are hindered by their significant minimum drag coefficient that is largely due to their wetted area, which is 1.57 times greater than an equivalent projected span biplane of equal aspect ratio. Annular wings also suffer from pitch-up at stall due to dissimilar stall characteristics between the upper and lower wing halves [2]. Consequently, an active design space for annular wings is where flight operations at low speed (high loading) for extended duration are a design driver, as may be required for a small reconnaissance unmanned air vehicle [3]. Gurneyflaps have received considerable attention of late as simple devices capable of augmenting the lift of a wing or airfoil [4–7]. The flap itself is usually composed of a small metal tab, typically nomore then 1–2% of the chord attached perpendicular to the trailing edge on the pressure side (generally within the pressure side’s boundary layer). The flaps essentially work by violating the Kutta condition at the trailing edge; thus greater loading is carried over the aft section as final pressure recovery occurs in thewake. This also has the benefit of reduced pressure recovery demands on the upper surface boundary layer [5]. The flaps have also been observed to increase base suction; beneficial in delaying stall but also causing a drag increase. Flap inclination has been indicated to be favorable in reducing drag, and placement at the trailing edge is generally indicated as being themost advantageous [5]. Segmented flaps have been studied as a means to attenuate the drag penalty of the flaps by introducing instabilities into the wake, which culminate in the breakdown of the two-dimensional vortex street that has been observed aft of theflap [6]. Liftmodulation of an annular wing can be complicated due to the wing’s geometric shape; flaps or ailerons may require large end gaps for clearance. Consequently, a Gurney flap may prove an effective device for force andmomentmodulation on an annular wing. Lift augmentation of an annular wing usingGurney flaps is apparently not documented in the literature. Consequently, a study has been undertaken to elucidate the effects of various Gurney flap configurations and layouts on the aerodynamic characteristics of an annular wing.

Wingtip eddy diffusion device

Abstract: The invention relates to a wingtip eddy diffusion device, comprising an upper winglet and a lower winglet which play the role of an end plate to block bottom wing surface air current from flowing to the upper wing surface; the upper winglet and the lower winglet are both of symmetrical airfoil profile; the symmetrical airfoil profile can avoid intense shock wave under supercritical air current condition, extra wave drag; certain deflection exists in wingtips of corresponding wings of the two winglets, thus improving lifting force of wings. The invention has the advantages of weakening wingtip eddy diffusion and trailing vortex intensity, downwash field of wingtips of wings and induced drag.

Numerical Investigation of Turbulent Flow Around a Stepped Airfoil at High Reynolds Number

Abstract: This study presents the numerical simulation of flow development around NACA-2412 airfoil which utilized the backward facing step to explore the possibility of enhancing airfoil aerodynamic performance by trapped vortex lift augmentation. This article concentrate on the effect of separated flow and following vortex formation which is created by backward facing step on pressure distribution and subsequently on lift and drag coefficient. Reynolds number that based on the free stream velocity and airfoil chord is 5.7×106 . The two-equation shear stress transport (SST) k-ω turbulence model of Menter is employed to determine accurately turbulent flow, as well as the recirculation pattern along the airfoil. The Reynolds-averaged Navier Stokes (RANS) equations are solved numerically using finite volume based solution with second-order upwind Roe’s scheme. Steps are located on both suction side and pressure side of the airfoil, at different locations, different lengths and various depths in order to determine their effects on lift, lift to drag ratio and near stall behavior. The modeling results showed that all stepped airfoil cases studied experienced higher drag compared to the base airfoil. Considerable lift enhancement was found for airfoil with backward facing step on pressure side at all values of angle of attack because of trapped vortex. The results suggest that the steps on the lower surface that extended back to trailing edge can lead to more enhancement of lift to drag ratio for some angles of attack; while the rear locations for the step on upper surface was found to have negative effect on lift to drag ratio. Based on this study, the backward facing step on suction surface offers no discernable advantages over the conventional airfoil but showed some positive effect on delaying stall.© 2009 ASME

Simulation on Flow Control for Drag Reduction of Revolution Body Using Bionic Dimpled Surface

Abstract: Numerical simulations on drag reductions for the dimpled and the smooth revolution bodies were performed and compared with SST k-ω turbulence model,to explain the reasons of friction and base drag reductions on the bionic dimpled surface and the control behaviors to boundary layer near wall of the revolution body.The simulated results show that the dimpled surface arranged on the rearward of the revolution body reduces the skin friction drag by 8.05%,the base drag by 1.9% and the total drag by 6.24% at Mach number 0.4;the dimpled surface reduces the skin friction drag through reducing the velocity gradient and turbulent intensity,and reduces the base drag through weakening the pumping action on the flow behind the revolution body caused by the external flow;the flow control behavior on boundary layer produced by dimpled surface displays that the low speed rotating vortexes in the dimples like vortex cushions,which segregate the external flow and the revolution body;and the low speed rotating vortexes forming in the bottom of dimples can produce negative skin friction against to the other area,which can be considered a accessional impetus.

Wave drag reduction by means of aerospikes on transonic wings

Abstract: Wave drag and shock induced boundary layer separation are important issues of transonic wings. The negative effect of the transonic flow regime can be mitigated by controlling the shock terminating the supersonic region above the wing. In the past many different concepts based, for example, on passive ventilation, active suction, contour bumps or on adaptive walls have been pursued (see references). These approaches have in common that measures for controlling the shock are applied directly at the surface of the wing. However, a control of the shock wave is also possible by external devices placed above the surface of the wing in the supersonic flow regime. Experiments related to the latter concept that is related to the one of aerospikes on blunt bodies, will be presented in the present contribution. In a test series the effectiveness of a variety of different spike-shaped bodies placed above a transonic wing was tested in the transonic wind tunnel DNW-TWG, Gottingen. In addition to pressure measurements a colour schlieren system was set up for providing information about the influence of spikes on the flow field. The drag reduction mechanism of spike-shaped bodies that are placed in the supersonic flow above a transonic wing is based on the generation of wake flows and oblique shock waves interfering with the normal shock terminating the supersonic region. In this manner the pressure increases more gradually thus limiting losses. Since the spike is located above the surface of the wing the boundary layer on the wing is not directly disturbed. In the streamwise direction the exact position of the spike is of less importance than in the case of measures applied at the surface of the wing. The height at which the spike is arranged above the surface is chosen so that the shock is especially weakened in its lower part where the shock strength is greatest. Typical dimensions depend on the chord length and the Mach angle. Similarly, in order to weaken the shock over the whole span width several spikes are placed next to each other in spanwise direction. Bodies of various geometries that are acting as wake and shock inducing spikes, have been studied. Results are reported that were obtained with a cylindrical body having a needle-like tip, a punctured pipe that was open at its leading edge and a cone. A 400mm-chord model of the transonic airfoil VC-opt was mounted in the 1m x 1m adaptive walls test section of the TWG. Initially, a single spike was placed on the suction side of the VC-Opt model, the tip of the spike being located about in the middle of the chord. A colour schlieren system was set up for observing the flow field. A comparison of colour schlieren pictures of the flow about a clean airfoil and the flow about an airfoil with a conical spike shows only little differences. This is due to the span width (1m) being much greater than the size of the spike (diameter about 12 mm). However, a shock wave and Mach lines originating from the tip and the surface of the spike, respectively, can well be seen. Lift and drag were determined by pressure measurements. On the wing the static pressure was obtained via pressure taps that were arranged in a slightly diagonal manner thus avoiding interferences between the taps. The drag was calculated from total pressure data obtained by wake-rake measurements one and a half chord lengths behind the trailing edge. Initially, the rake was laterally displaced with respect to the spike by ten percent of the chord length. Tests were performed at a Reynolds number of Re ≈ 5⋅106 and at two Mach numbers, M = 0.775 and M = 0.795, respectively. Lift and drag polars were obtained for different configurations such as a clean airfoil and for an airfoil with shock inducing bodies. At certain angles of attack a gain in lift is observed and the drag is clearly reduced. This effect is most pronounced for a conical spike, i.e., for a body which produces a notable displacement in the flow. Hence, the concept of using aerospikes on transonic wings clearly shows a potential for reducing wave drag.

Aerodynamic Analysis of the Unmanned Aerial Vehicle for Ecological Conservation

Abstract: The present work has as objective to do a comple te aerodynamic analysis of the Unmanned Aerial Vehicle for Ecological Conservation. The panel method code PAN AIR is used to compute the inviscid flowfield. The viscous effects in drag are estimated by the classic Hoerner method, and the maximum lift coeff icient via the classic Valarezo and Chin method. The numerical aerodynamic forces of the complete airplane are compared to experimental data for validation. The spanload and wing pressure distribution are estimated for four configurations: wing, wing -body, wing -body -tail, and wing -body -tail with wing twist. The sources of induced drag for all configurations are achieved graphically via Trefftz plane. All the data were estimated at cruise flight, Reynolds number equal to 1.413×10 6