Showing papers on "Wing root published in 2009"

Smoke visualization of free-flying bumblebees indicates independent leading-edge vortices on each wing pair

Abstract: It has been known for a century that quasi-steady attached flows are insufficient to explain aerodynamic force production in bumblebees and many other insects. Most recent studies of the unsteady, separated-flow aerodynamics of insect flight have used physical, analytical or numerical modeling based upon simplified kinematic data treating the wing as a flat plate. However, despite the importance of validating such models against living subjects, few good data are available on what real insects actually do aerodynamically in free flight. Here we apply classical smoke line visualization techniques to analyze the aerodynamic mechanisms of free-flying bumblebees hovering, maneuvering and flying slowly along a windtunnel (advance ratio: −0.2 to 0.2). We find that bumblebees, in common with most other insects, exploit a leading-edge vortex. However, in contrast to most other insects studied to date, bumblebees shed both tip and root vortices, with no evidence for any flow structures linking left and right wings or their near-wakes. These flow topologies will be less efficient than those in which left and right wings are aerodynamically linked and shed only tip vortices. While these topologies might simply result from biological constraint, it is also possible that they might have been specifically evolved to enhance control by allowing left and right wings to operate substantially independently.

High-speed stereo DPIV measurement of wakes of two bat species flying freely in a wind tunnel

Abstract: Previous studies on wake flow visualization of live animals using DPIV have typically used low repetition rate lasers and 2D imaging. Repetition rates of around 10 Hz allow ~1 image per wingbeat in small birds and bats, and even fewer in insects. To accumulate data representing an entire wingbeat therefore requires the stitching-together of images captured from different wingbeats, and at different locations along the wing span for 3D-construction of wake topologies. A 200 Hz stereo DPIV system has recently been installed in the Lund University wind tunnel facility and the high-frame rate can be used to calculate all three velocity components in a cube, whose third dimension is constructed using the Taylor hypothesis. We studied two bat species differing in body size, Glossophaga soricina and Leptonycteris curasoa. Both species shed a tip vortex during the downstroke that was present well into the upstroke, and a vortex of opposite sign to the tip vortex was shed from the wing root. At the transition between upstroke/downstroke, a vortex loop was shed from each wing, inducing an upwash. Vorticity iso-surfaces confirmed the overall wake topology derived in a previous study. The measured dimensionless circulation, Γ/Uc, which is proportional to a wing section lift coefficient, suggests that unsteady phenomena play a role in the aerodynamics of both species.

Design Overview of a Resonant Wing Actuation Mechanism for Application in Flapping Wing MAVs

Abstract: This paper shows the design and analysis of the actuation mechanism for a four winged flapping wing MAV. The design is set up to exploit resonant properties, as exhibited by flying insects, to reduce the energy expenditure and to provide amplitude amplification. In order to achieve resonance a significantly flexible structure has to be incorporated into the design. The elastic structure used for the body of the MAV is a ring type structure. The ring is coupled to the wings by a compliant amplification mechanism which transforms and amplifies the ring deflection into the large wing root rotation. After initial sizing, the structures are analyzed by finite elements (eigenvalue and transient analysis). Based on the initial analysis, the structures are realized to be tested later. The wings are first analyzed independent of the structure in order to tune wing hinge stiffness to efficiently generate lift, exploiting passive wing pitching. The wings are tuned by using a quasi-steady aerodynamic model. The tuned...

Implicit large-eddy simulation of swept-wing flow using high-resolution methods

Abstract: A numerical investigation of a fully three-dimensional swept-wing geometry featuring separation from a curved leading edge is presented. The implicit large-eddy-simulation strategy based on a third-order high-resolution method for discretizing the advective fluxes and a second-order Runge-Kutta time-stepping scheme with an extended stability region have been employed. No attempt to incorporate a wall model has been made. Instead, the boundary layer is fully resolved over the majority of the wing. Qualitative and quantitative comparisons with experimental oil-film visualizations and three-dimensional laser Doppler anemometry measurements show very good agreement between the experiment and the numerically predicted flow structures, as well as velocity and stress profiles near the wing. Furthermore, data from a hybrid Reynolds-averaged Navier-Stokes and large-eddy simulation have been included for comparison.

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.

Theoretical Predictions of F-16 Fighter Limit Cycle Oscillations for Flight Flutter Testing

Abstract: A computational investigation of the flutter onset and limit cycle oscillation behavior of various F-16 fighter weapons and stores configurations is presented. A nonlinear harmonic balance compressible Reynolds-averaged Navier–Stokescomputational fluiddynamic flowsolverisusedtomodeltheunsteadyaerodynamicsoftheF-16wing. Slender body/wing theory is used as an approximate method for accounting for the unsteady aerodynamic effects of wing-tip launchers and missiles. Details of the computational model are presented along with an examination of the sensitivity of computed aeroelastic behavior to characteristics and parameters of the structural and fluid dynamic model. Comparisons with flight-test data are also shown. I. Introduction T HE SEEK EAGLE Office at Eglin Air Force Base performs an essential task in clearing new aircraft/stores configurations through flight tests for safe and effective operation. Many of these flighttestsarefortheF-16aircraftwhichcontinuestobeaworkhorse for the U.S. Air Forcewith continually new stores (missiles, bombs, and fuel tanks) being considered for aircraft operations. Similar aeroelastic flight tests are expected for future fighter aircraft as they go into service in the coming years. The number of needed flight tests is projected to be well beyond the financial and staff resources available. Hence there is a pressing need to identify the most critical aircraft/store configurations for the limited flight-test resources available and also insofar as possibly reduce the number of flight tests needed. Virtual flight testing may be the answer. Using new improved computational capability that provides much more rapid solutions, computational simulation can help identify the most critical aircraft/ store configuration and also hasthe potential of reducingthe number ofneeded flighttestsifconfidencecanbeestablishedinthecapability of simulations to correlate with flight-test data. A new methodology has been developed to produce these computer simulations based upon the notion that because the response is periodic in time, the solution need only be obtained over a single period of oscillation in time. By avoiding the traditional time marching solution which computes the long transient before a steady-state periodic oscillation is reached, computational times are reduced by a factor of 10–100. This enables a sufficiently rapid solutiontomakesuchsimulationsapracticalrealityforthe flight-test engineer and support team. Future developments of this methodology hold the promise of further substantial reductions in computational cost and are being vigorously pursued. Also further refinements in the physical fidelity of the simulation models are being considered.

Experimental Investigation of Annular Wing Aerodynamics

Abstract: A low-speed wind-tunnel investigation was conducted to explore the behavior of annular (ring) wings. Effects of aspect ratio as well as gap were investigated. Ring wings using a low Reynolds number Eppler section and a NACA 0012 profile were manufactured and tested. Measurements were recorded using a six-component sting balance. Experimental and theoretical trend comparisons were effected using a vortex-lattice code. The experimental results indicate wing efficiency factors well above 1 are achievable. The effect of gap was to increase the wing lift-curve slope as well as efficiency. The large increases in aerodynamic efficiency were generally mitigated by the significant minimum drag coefficient. Pitching moment characteristics were unfavorable and were dominated by dissimilar stall behavior between the upper and lower wing sections.

Experimental Analysis of Propeller-Wing Interactions for a Micro Air Vehicle

Abstract: An experimental investigation into the effects of the propeller slipstream on a small, flexible-wing air vehicle with a 24 in. wingspan was conducted at typical flight speeds, resulting in Reynolds numbers based on a mean aerodynamic chord of about 1 x 105. A static test bench was used to characterize the forces and moments acting on the motor and, separately, on the airframe. The motor and propeller, of diameter D, were mounted on a torque cell that was itself affixed to an air-bearing table configured to measure thrust A model of the fuselage, capable of wing location adjustment, was affixed to a 6-degree-of-freedom balance. The resulting set of decoupled force and moment data was collected for axial wing placements ranging from 0.17 to 0.73 D and also for varied vertical wing placements. The results suggest that between 12 and 18% of propeller thrust translates into airframe drag, with the largest percentage occurring for the wing placement closest to the propeller (0.17 D). The measured reaction torque was found to be as high as 45% of the motor torque, and its dependence on wing placement was measured. Wind-tunnel data, collected using a six-component balance and a triaxial hot-wire anemometer mounted on a traversing system, yielded all three components of velocity maps for propeller-off and propeller-on conditions at typical operating airspeeds. The secondary flows generated by the propeller, propeller-wing interaction, and wing-tip vortices were captured and analyzed.

Aerodynamics of flapping-wing micro air vehicles

Abstract: The research study outlined in the paper addresses the aerodynamic features of flexible flapping wings used in micro air vehicles called ornithopters. Aerodynamic force measurements were conducted for the 25-cm and 74-cm-wing-span models at different airflow velocities and flapping frequencies. A series of experiments were conducted on the 25-cm flapping-wing model without free stream airflow. In order to study the effect of a dihedral on generated thrust and normal force, the model was modified to accommodate three values of dihedral angle. It appeared that the thrust force is higher for higher dihedral angle. Effects of a wing’s bending stiffness and of the wing root constraint on the generated thrust force and required power were investigated. The aerodynamic forces on flapping wings were studied with the stroke plane angle varied from horizontal to vertical. An important result was found that flapping wings do not exhibit a typical, abrupt stall seen with the fixed wings. Experimental results were analyzed in the framework of the momentum theory. The results of this study were realized in micro ornithopter designs that was successfully flight tested.

Dye Visualization of the Flow Structure over a Yawed Nonslender Delta Wing

Abstract: S TUDIES of aerodynamic structures and behaviors of the nonslender delta wings are invariably essential to develop a method to control the development of the vortex breakdown as well as the development of vortices. Unsteady aerodynamics of nonslender delta wings, consisting of shear layer instabilities, the structure of vortices, the occurrence of breakdown, and fluid/structure interactions were extensively reviewed by Gursul et al. [1]. They emphasized the sensitivity of the vortical flow structures varying the angle of attack of the deltawing.Yavuz et al. [2] studied thevortical flow structure on a plane immediately adjacent to the surface of nonslender delta wing, 38:7 deg. Yaniktepe and Rockwell [3] performed experimental investigations on the flow structures at trailing-edge regions of diamondand lambda-type wings. In both wings, vortical flow structures in the crossflowplanes of trailing edge vary rapidly with the angles of attack . Sohn et al. [4] visually investigated the development and interaction of vortices in crossflow planes at various locations on the delta wing with leading edge extension (LEX) using micro water droplets and a laser beam sheet. The range of angle of attack was taken as 12 24 deg at yaw angles of 0, 5, and 10 deg. It was indicated that, by introducing yaw angle , the coiling, merging, and diffusion of thewing and LEX vortices increased on the windward side, whereas they became delayed significantly on the leeward side. Their study confirmed that the yaw angle had a profound effect on the vortex structures. Taylor and Gursul [5] visualized leading-edge vortices of a 50 deg sweep angle, having angles of attack as low as 2:5 deg. Gursul et al. [6] report that combat air vehicles (UCAVs) and micro air vehicles have particularly dominant vortical flows having low sweep angles (25–55 deg), and future UCAVs are expected to be highly maneuverable and highly flexible. Yaniktepe and Rockwell [7] aimed at investigating the unresolved concepts, which included averaged structure of shear layer from the leading edge of the wing, unsteady features of separated layer adjacent to the surface of the wing, and control of flow structure by leading-edge perturbations. Elkhoury and Rockwell [8] have investigated to provide various measurements of the visualized dye patterns, including the degree of interaction of vortices, the onset of vortex breakdown, and effective sweep angle of the wing root vortex, as a function of both Reynolds number and angle of attack . Elkhoury et al. [9] had investigated the Reynolds number dependence of the near-surface flow structure and topology on a representative UCAV planform. The present investigation focuses on the formation and development of leading-edge vortices, vortex breakdown, and threedimensional separationandstallof thecomplexanddisorganizedflow structure over the delta wing. The leading-edge sweep angle was 40 deg. The angle of attack was varied within the range of 7 17 deg and the yaw angle was varied within the range of 0 15 deg.

Structural Analysis and Testing of an Ultralight Unmanned-Aerial-Vehicle Carbon-Composite Wing

Abstract: This paper presents the results of an investigation examining the strength and stiffness characteristics of a carbon composite wing of an ultralight unmanned aerial vehicle. The wing consists of foam-core sandwich skins and multiple spars with varying laminate ply patterns and wall thickness dimensions. A three-tier whiffletree system is designed and used to load the wing in a manner consistent with a pull-up-maneuver condition. Multiple strain and deflection gauges are used to measure the static response of the wing at several spanwise and chordwise locations, and the wing is loaded incrementally beyond the limit and design ultimate loads to the point of structural failure. A geometric nonlinear finite element model is developed that accounts for the variations in laminate geometry of each component as well as the properties of the adhesively bonded joints in the wing assembly. By carefully matching the boundary conditions with those of the experimental setup, the static response of the wing under a simulated whiffletree loading condition is obtained. The strain and deflection predictions from the finite element simulations are found to be in good agreement with the experimental observations. Despite its light weight, the wing is found to be very strong, with a strength-to-weight ratio of more than 40 at failure.

Transport Wing Flutter Model Transonic Limit Cycle Oscillation Test

Abstract: The model for aeroelastic validation research involving computation semispan wind-tunnel model, a transport wing-fuselage flutter model, was tested in NASA Langley's Transonic Dynamics Tunnel with the goal of obtaining experimental limit cycle oscillation behavior data at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wing-tip configurations: clean, tip store, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6-0.9, and maps of limit cycle oscillation behavior were made in the range of M = 0.85-0.95. The effects of dynamic pressure and angle of attack were measured. Testing in both R134a heavy gas and air provided unique data on the Reynolds number, transition effects, and the effect of speed of sound on limit cycle oscillation behavior. This report gives an overview of the test results, including experimental flutter boundaries, and the conditions involving shock-induced transonic flow separation onset at low wing angles, including maps of limit cycle oscillation behavior.

Aerodynamics of battle-damaged finite-aspect-ratio wings

Abstract: When an aircraft is aerodynamically or structurally damaged in battle, it may not able to complete the mission and the damage may cause its loss. The subject of aircraft battle survivability is one of critical concern to many disciplines, whether military or civil. This thesis considered and focused on Computational Fluid Dynamics [CFD] predictions and experimental investigations into the effects of simulated battle damage on the low-speed aerodynamics of a fmite aspect ratio wing. Results showed that in two-dimensional [2d] and three-dimensional [3D] CFD simulations, Fluent's® models work reasonably well in predicting jets flow structures, pressure distributions, and pressure-coefficient Cp's contours but not for aerodynamic coefficients. The consequences were therefore that CFD prediction was poor on aerodynamic-coefficients increments. The prediction of Cp's achieved good agreement upstream and near the damage hole, but showed poor agreement at downstream of the hole. For the flow structure visualisation, at both weak and strong jet incidences, the solver always predicted pressure-distribution-coefficient lower at upstream and higher at downstream. The results showed relatively good agreement for the case of transitional and strong jet incidences but slightly poor for weak jet incidences. From the experimental results of Finite Wing, the increments for Aspect-ratio, AR6, AR8 and ARIO showed that as damage moves out towards the tip, aerodynamic-coefficients increments i.e. lift-loss and drag-rise decreased, and pitching-moment-coefficient increment indicated a more positive value at all incidence ranges and at all aspect ratios. Increasing the incidence resulted in greater magnitudes of lift-loss and drag-rise for all damage locations and aspect ratios. At the weak jet incidence 4° for AR8 and in all of the three damage locations, the main characteristics of the weak-jet were illustrated clearly. The increments were relatively small. Whilst at 8°, the flow structure was characterised as transitional to stronger-jet. In Finite Wing tests and for all damage locations, there was always a flow structure asymmetry. This was believed to be due to gravity, surface imperfection, and or genuine feature. An 'early strong jet' that indicated in Finite Wing-AR8 at 'transitional' incidence of 8°, also indicated in twodimensional results but at the weak-jet incidence of 4°. For the application of 2d data to AR6, AR8, and ARIO, an assessment of 2d force results led to the analysis that the tests in the AAE's Low Turbulence Tunnel for 2d were under-predicting the damage effects at low incidence, and over-predicting at high incidences. This suggested therefore that Irwin's 2d results could not be used immediately to predict three-dimensional.

Three-Dimensional Aerodynamic Design of Low-Wave-Drag Supersonic Biplane Using Inverse Problem Method

Abstract: In supersonic flight, airfoils generate strong sonic boom and wave drag accompanied by shock waves and expansion waves. The Busemann biplane is representative of an airfoil that has the possibility of realizing low drag. With the aim of realizing a new concept of supersonic transport, aerodynamic design and analysis are discussed based on computational fluid dynamics. Traditional biplane airfoils were extended to three-dimensional wings with a design Mach number of 1.7. Euler simulations of three-dimensional biplane wings of several configurations were conducted. Because of the existence of wing tips, three-dimensional biplane wings do not perform as well as two-dimensional biplane airfoils. This is because the areas affected by Mach cones originated from the wing tips, which precludes the occurrence of an appropriate pressure wave interaction. Thus, the wing tip area has a large drag coefficient. To overcome these problems, a tapered wing is herein considered. After parametric studies, a wing planform was determined, the taper ratio and aspect ratio of which were 0.25 and 5.12, respectively. Aerodynamic design of wing section shapes of the three-dimensional tapered biplane wing was conducted using an inverse-problem method. The designed biplane wing shows a better lift-to-drag ratio performance than the two-dimensional flat-plate airfoil in the range where the lift coefficient is more than 0.17. Tapered wings were also found to have several span sections that achieve better performances than two-dimensional biplane airfoils.

A Comparison of Surrogate Models in the Framework of an MDO Tool for Wing Design

Abstract: The replacement of the analysis portion of an optimization problem by its equivalent metamodel usually results in a lower computational cost. In this paper, three dierent metamodels are compared against the conventional non-approximative approach: quadratic interpolation based response surfaces, Kriging and Articial Neural Networks (ANN). The results obtained from the solution of three dierent case studies based on aircraft design problems reinforces the idea that quadratic interpolation is only well suited to very simple problems. At higher dimensionality, the usage of the more the complex Kriging and ANN models may result in considerable performance benets. Nomenclature b=2 Wing semispan, m c;ci Coecients for polynomial interpolation cbs Wing breakstation chord, m croot Wing root chord, m ctip Wing tip chord, m f (x) Regression model (Kriging) g (x) Constraint function nDV Number of design variables ns Number of samples nt Number of terms in polynomial interpolation/regression approximation qk(x) Values of regression functions at sample locations (Kriging) R (w; x; ) Correlation model (Kriging) sk Vector of independent variable samples (Kriging)

Application of Model Predictive Control to Gust Loads Alleviation Systems

Abstract: In this paper, a new control framework is proposed to address the gust load alleviation for very flexible aircraft. The novelty of this framework is two folded: on one hand, the deformable dynamics is considered at all stages of design to address the structural interaction challenge; on the other hand, the designed control function does not distinguish between flight dynamics and aeroelasticity to make it possible to improve the control of very flexible aircraft. To demonstrate the capability and effectiveness of the proposed structure, the nonlinear equations of motion for a very flexible aircraft are developed. The performance of the system is evaluated with an aircraft passing through a discrete gust (1 - cos profile) as well as a continuous one. The measurement criteria are set as the maximum wing root bending moment (for a discrete gust) and the RMS value of the bending moment (in a continuous gust). On the control side, two control techniques are applied: model predictive control (MPC) and linear quadratic regulation (LQR). The proposed control framework is compared with an existing control approach to show that it improves control effectiveness.

Plasma Control Of Vortex Flow On Delta-Wing At High Angles Of Attack

Abstract: Nomenclature AR = wing aspect ratio b = wing thickness Cp = pressure coefficient, (p-p0)/q+1 c = wing root chord length Cxa = drag coefficient Cya = lift coefficient Cza = side force coefficient K = lift-drag ratio, mxa = roll moment coefficient mza = pitching moment coefficients p = static pressure on the model surface Tu = free-stream turbulence level, % of U∞ U∞ = mean free-stream velocity q = dynamic pressure x = vortex breakdown position along the wing chord z = transversal direction along the wing span α = angle of attack χ = delta-wing leading edge sweep angle

Blended wing aircraft

Abstract: Apparatus and methods provide for a blended wing passenger or cargo aircraft. Aspects of the disclosure provide an aircraft having wings with spars having a thickness at the wing root corresponding to a height of the payload space within the fuselage to which the wings are attached. The wing spars within the wings on each side of the aircraft may each be spliced into an aircraft frame that is part of the fuselage. The wing thickness provides mounting locations for aircraft engines and other components within the wing and passing through the wing spars. With this mid-wing configuration, the fuselage provides support for the various loads experienced by the wings without the use of a conventional structural wing box.

Downward pointing winglet design and assessment within the M-DAW research project

Abstract: The design and performance analysis of a wing tip device proposed within the M-DAW project by ONERA is presented. A proto-design process is described and the device was thoroughly assessed (mainly with Reynolds-Averaged Navier-Stokes simulations). The process was further explained through wind-tunnel tests at both low speed and high speed in the pressurised and cryogenic European transonic wind tunnel in Cologne. The device is a downward pointing winglet designed for a retrofit scenario (the wing could be modified only within the 96% – 100% bounds of the span). It was designed to keep the wing root bending moment of the clean wing at cruise unchanged so that the aerodynamic gains are the net gains provided by the device that can be directly installed without structural modifications of the wing.

Flight vehicle with flapping wings ("dragonfly")

Abstract: FIELD: aircraft engineering. ^ SUBSTANCE: proposed flight vehicle comprises airframe accommodating engine with reduction gear and two crank drives. The latter serve to drive two power shafts, which make, via chain drives, four symmetric wings waving. Aforesaid wings incorporate mechanism designed to allow wings to rotate about lengthwise axis. Said mechanism is arranged in the wing root on central spar mounted in yoke to revolve therein when wings flap driven by two ratchet-and-pawl gears. Each of said gears consists of two disks, one attached onto wing spar and another represents a pulley with pawl, and is furnished with control lever. ^ EFFECT: higher stability and better controllability. ^ 2 dwg

Influence of Structural Flexibility on Flapping Wing Propulsion

Abstract: : The influence of structural deformations on the aerodynamic response of a flapping wing configuration was examined using Navier-Stokes based simulation Two deformation modes, torsion and bending, were considered for an elastic axis along the leading edge of the wing Both deformation modes influence the velocity and acceleration profile of the wing surface, altering the unsteady aerodynamic phenomena produced by the dynamic wing motion The spanwise feathering rotation, or torsional response, alters the motion of the wing near the wing root This variation in the acceleration profile influences the non-circulatory aerodynamic response and the local wake structures produced near the wing root during pronation and supination Increased lifting forces and enhanced aerodynamic efficiencies were observed for a moderate increase in torsional exibility Peak bending deformations near the wing tip also occur during pronation and supination, altering the velocity and acceleration profiles of the wing as the circulatory aerodynamic phenomena undergo a transition as the wing changes direction of motion Because of the timing of the bending deformations, small tip deformations may have a significant influence on overall aerodynamic performance

Aerodynamic Characteristics of Forward and Aft Swept Arrow Wings

Abstract: D ESPITE aerodynamic advantages, aeroelastic complications have precluded the wide-scale adoption of the forward swept wing planform [1]. Forward swept wings may demonstrate liftdependent drag benefits due to spanwise load distributions that are closer to elliptic then an equivalent aft swept wing [2]. Fuselageinduced spanwise flow is also beneficial for forward swept wings; it reduces the effective sweep, whereas the opposite is true for aft swept wings [3]. The most apparent benefit of the forward sweep configuration is in lateral control; spanwise flow toward the root delays flow separation over the ailerons, whereas the opposite is true for aft swept wings [4]. Little information is available on the effect of highly swept wingswith forward sweepwhere flow separation is enforced. Consequently, an investigation has been undertaken using a 65 deg sweep arrow wing with fore and aft extensions.

Flow over Delta Wing with Low Dorsal Fin-PIV Study

Abstract: A recent theoretical and experimental work done on the vortex flow over slender flatplate delta wing with low dorsal fin is reviewed, and a particle image velocimetry (PIV) study is conducted to further the investigation. The distributions of the vorticity component normal to the cross-flow planes show significant flow asymmetry and non-conicity over the wing-fin combination at a high angle of attack and enhance the validation of the theoretical prediction that adding a low dorsal fin to the wing may destabilize the symmetric and conical vortex pair. Ideals for further work are suggested. Nomenclature b = wing span c0 = wing root chord Cn = yawing-moment coefficient, yawing moment aboutZ-axis/q∞Sb hL = local height of dorsal fin K = Sychev similarity parameter, tan α/tan ǫ q∞ = free-stream dynamic pressure S = wing area s = local semi-span of wing U∞ = free-stream velocity v, w = cross-flow velocity components X, Y, Z = balance body axes, Fig. 4 x, y, z = body axes of the wing α = angle of attack ǫ = semi-apex angle of wing β = sideslip angle ωx = axial vorticity, ∂w/∂y ∂v/∂z

CFD analysis of an oscillating wing at various reduced frequencies

Abstract: The effect of various reduced frequencies has been examined for an oscillating aspect ratio 10 NACA 0015 wing. An unsteady, compressible three-dimensional (3D) Navier–Stokes code based on Beam and Warming algorithm with the Baldwin–Lomax turbulence model has been used. The code is validated for the study against published experimental data. The 3D unsteady flow field is simulated for reduced frequency values of 0.1, 0.2 and 0.3 for a fixed mean angle of attack position and fixed amplitude. The type of motion is sinusoidal harmonic. The force coefficients, pressure distributions and flow visualization show that at the given conditions the flow remains attached to the wing surface even at high angles of attack with no clear separation or typical light-to-deep category of dynamic stall. Increased magnitude of hysteresis and higher gradients are seen at higher reduced frequencies. The 3D effects are even found at midspan locations. In addition, the rate of decrease in lift near the wing tips compared with the wing root is not much like in the static cases. Copyright © 2008 John Wiley & Sons, Ltd.

Knife blade for pulverization

Abstract: The utility model relates to a pulverizing cutter blade, in particular to a pulverizing and stirring cutter blade used in a soymilk machine or a paste mixer, which comprises a wing root (1) and a cutter wing (4) which is connected with the wing root (1); the geometrical central position of the wing root (1) is provided with a cutter hole (11) for the penetration of a rotary shaft, wherein the cutter wing (4) which is connected with the wing root (1) comprises a first cutter wings (2) and a second cutter wing (3); the geometrical center of the first cutter wing (2) and the geometrical center of the second cutter wing (3) are not in a same plane. Because the pulverizing and stirring cutter blade adopts the structure which comprises two cutter wings connected with the wing root and the geometrical centers of the two cutter wings are not in a same plane, and the cutter wings are bent upwards and downwards respectively corresponding to the wing root, a larger pulverizing space can be formed when the blade rotates, and thus better pulverizing effect can be achieved. The utility model is a convenient and practical pulverizing cutter blade with skillful design and good performance.

High-lift fuselage aircraft

Abstract: FIELD: aircraft engineering. ^ SUBSTANCE: invention relates to aircraft engineering. Proposed aircraft represents a unified (integral) 3D system made up of elliptical fuselage with the height-to-width ratio of 1:2 to 1:4, a wing with elongation ?=7 to 11 and wing taper ratio ?=3-4. Proposed system lengthwise sections consist of composite sections modified so that the fuselage section and wing root sections satisfy the conditions of maximum Mk* and Mzo at moderate Cy max, and the root section satisfies the condition of Cy max maximum. The sections feature negative aerofoil section camber f~0.015 to 0.02 varying from the aircraft axis to over with wing span from X=0.6 for root section to X=0.9 nearby the wing tip section. Relative fuselage section thickness does not exceed C~0.17 and that of wind sections does not exceed C~0.15. Angles of rear surface inclination nearby rear edge of any aircraft surface does not exceed ?<6 degrees. ^ EFFECT: higher stability and controllability. ^ 5 dwg

Wing-in-ground effect craft wing

Abstract: The utility model discloses a wing-in-ground effect craft wing, in particular to a wing-in-ground effect craft wing for generating aerodynamic force for a wing-in-ground effect craft so that the wing-in-ground effect craft can take off from a water surface and can fly in the sky. The wing-in-ground effect craft wing comprises a wing body and is characterized in that the front edge of the wing body is flat and is perpendicular to the medial axis of a craft body, the width of a wing root is larger than that of a wing top, and the wing body is in an S shape. The wing-in-ground effect craft wing has light weight and low production cost and can lower the operation difficulty of a driver.

CFD Development and Optimisation of the ADFA SAE Aeronautical Design UAV

Abstract: This thesis report details the 2D CFD analysis of the canard and wing root airfoil configuration of the ADFA SAE Aero Design UAV. The UAV has been designed against the specifications detailed for the 2008 SAE Aero Design Competition, in which maximizing the payload lifting capacity of the UAV with fixed geometric and thrust restraints is the design intention. The CFD analysis has been conducted in an attempt to increase the lifting characteristics of the current design of the ADFA SAE Aero Design UAV through selecting the optimum location and angle of attack of both the canard and wing. This analysis been achieved through analysing the flow interaction between the two lifting surfaces and considering the required flight profile of the UAV. This report provides a review of literature relating to the engineering application of CFD and flow visualisation techniques to validate CFD results. The generation and formulation of a suitable mesh grid and CFD model is discussed in detail, as is the validation of the model with both published data and the experimental results obtained from wool tuft flow visualisation conducted for a half scale model of the ADFA UAV. This report has concluded that the canard and wing configuration can be optimised through location the canard 60mm vertically above its current location, and increasing the angle of incidence of both the canard and wing from 40 to 6.10 and 7.20 respectively. The optimised design is predicted to increase the lifting characteristics of the UAV by an additional 14.4N.