Showing papers on "Wing root published in 2010"

Wingbeat Shape Modulation for Flapping-Wing Micro-Air-Vehicle Control During Hover

Abstract: A new method of controlling a flapping-wing micro air vehicle by varying the velocity profiles of the wing strokes is presented in this manuscript. An exhaustive theoretical analysis along with simulation results show that this new method, called split-cycle constant-period frequency modulation, is capable of providing independent control over vertical and horizontal body forces as well as rolling and yawing moments using only two physical actuators, whose oscillatory motion is defined by four parameters. An actuated bob-weight is introduced to enable independent control of pitching moment. A general method for deriving sensitivities of cycle-averaged forces and moments to changes in wingbeat kinematic parameters is provided, followed by an analytical treatment for a case where the angle of attack of each wing is passively regulated and the motion of the wing spar in the stroke plane is driven by a split-cycle waveform. These sensitivities are used in the formulation of a cycle-averaged control law that successfully stabilizes and controls two different simulation models of the aircraft. One simulation model is driven by instantaneous aerodynamic forces derived from blade-element theory, while the other is driven by an empirical representation of an unsteady aerodynamic model that was derived from experiments.

Transonic Aeroelastic Stability Predictions Under the Influence of Structural Variability

Abstract: DOI: 10.2514/1.46971 Flutter prediction as currently practiced is almost always deterministic in nature, based on a single structural model that is assumed to represent a fleet of aircraft. However, it is also recognized that there can be significant structural variability, even for different flights of the same aircraft. The safety factor used for flutter clearance is in part meant to account for this variability. Simulation tools can, however, represent the consequences of structural variability in the flutter predictions, providing extra information that could be useful in planning physical tests and assessing risk. The main problem arising for this type of calculation when using high-fidelity tools based on computational fluid dynamics is the computational cost. The current paper uses an eigenvalue-based stability method together with Euler-level aerodynamics and different methods for propagating structural variability to stability predictions.The propagation methodsare Monte Carlo,perturbation, andinterval analysis. Thefeasibility of this type of analysis is demonstrated. Results are presented for the Goland wing and a generic fighter configuration.

Prediction of Tiltrotor Vibratory Loads with Inclusion of Wing­-Proprotor Aerodynamic Interaction

Abstract: A numerical methodology for the prediction of vibratory loads arising in wing―proprotor systems is presented. It is applicable to tiltrotor operating conditions ranging from airplane to helicopter-mode flights. The aeroelastic formulation applied takes into account the aerodynamic interaction effects dominated by the impact between proprotor wake and wing, along with the mutual mechanical influence between elastic wing and proprotor blades. A boundary integral formulation suited for configurations where strong body―vortex interactions occur yields the aerodynamic loads, and beamlike models are used to describe the structural dynamics. A harmonic balance approach is applied to determine the aeroelastic solution. In the numerical investigation, first, the aerodynamic solver is validated by correlation with experimental and numerical results available in the literature, then the vibratory loads transmitted by the wing―proprotor system to the airframe are predicted, focusing the attention on the analysis of the different aerodynamic contributions.

Gust Response for Flexibly Suspended High-Aspect Ratio Wings

Abstract: A theoretical aeroelastic study for a flexibly suspended high-aspect ratio wing aeroelastic model excited by gust loads is presented along with a companion wind-tunnel test. A flexible support system at the wing root has been constructed to simulate a rigid body mode. Structural equations of motion based on a nonlinear beam theory are combined with the ONERA aerodynamic model. Also, a dynamic perturbation analysis about a nonlinear static equilibrium is used to determine the small perturbation flutter boundary and structural natural frequencies. The effects of the flexible support system (bi-beam system) on the wing structural dynamics, the static aeroelastic displacement at root, nonlinear flutter, and gust response to a harmonic or a frequency sweep excitation are discussed. Also the effect of gust distribution along the span on the response is described based upon computations. The fair to good quantitative agreement between theory and experiment demonstrates that the present analysis method has reasonable accuracy.

Wing-fuselage section of an aircraft

Abstract: A wing-fuselage section of an aircraft, which wing-fuselage section comprises a wing root (7, 8) at which the wing (1) of the aircraft is connected to the fuselage (2), a fuselage region (3) with fuselage frame elements (11-15, 21-25) that extent across the longitudinal direction of the aircraft, and a wing region (5, 6) with spars (16-19, 26-29) that extend in the direction of the wingspan. According to the invention, the spars (16- 19, 26-29) of the wing region (5, 6) and the fuselage frame elements (11-15, 21-25) of the fuselage region (3) form part of an integral assembly (40) that extends at least over a middle part of the wing (1) and the fuselage region (3), including the wing roots (7, 8).

Simulations of Vortex Formation Around a Blunt Wing Tip

Abstract: Wereport ourfindings froma computational study onwing-tip vortices. Themain emphasis of the simulations is to compute the formation of a tip vortex around a blunt wing tip and its interaction with the wing and tip side-edge surfaces. Comparisons of simulation results with available experimental data are done to assess the prediction capability of the simulations. The simulations are performed on a computational grid containing 110 million points total. The blunt tip geometry actually gives rise to the formation of two vortices. The primary vortex forms over the upper surface of the wing, and the secondary vortex forms off of the side edge. These two vortices merge together around the tip trailing edge to form the tip vortex. Simulations at the experimental Reynolds number of 1:8 10 are performed using several subgrid-scalemodels. It is shown that the simulation performedwithout an explicit subgridscale model, also known as implicit large eddy simulation, produces the best agreement with the experimental measurements. Although some differences between the simulation results and experimental data still exist, the overall agreement between the experiment and simulation is found to be satisfactory.

Flight Control Using Wing-Tip Plasma Actuation

Abstract: A concept for lift modification on a conventional aircraft wing for roll control at low angle of attack with dielectric barrier discharge plasma actuators is proposed and assessed through computational fluid dynamics simulations and preliminary wind-tunnel experiments. The concept consists of placing plasma actuators around the wing tip to add momentum in the direction opposite to that of the flow forming the tip vortex. Because of the limited strength of existing plasma actuators, the assessment is carried out for a relatively small two-dimensional wing (NACA 4418) with a rounded tip at zero angle of attack and 15 m/s for a Reynolds number in the range of 1.5 x 10 5 . Computational fluid dynamics simulations show a significant alteration of the vorticity field downstream of the trailing edge characterized by a more diffused vortex surrounded by zones of negative vorticity induced by the actuators and, but not necessarily, outboard displacement of the tip vortex. This leads to a reduced downwash that results in a change in lift of up to almost 20% for actuator strength levels that should be achievable in the short term with a new generation of dielectric barrier discharge actuators. The actuator placed on the suction side contributes the most to the lift increase, with its induced jet blocking the flow around the wind tip at the origin of the formation of the tip vortex. Wind-tunnel experimental results support the computational fluid dynamics predictions in both magnitude and trend. Furthermore, preliminary computational fluid dynamics simulations are carried out for a symmetric nonlifting wing (NACA 0018), representative of aircraft tail surfaces at zero angle of attack to generate lift for pitch and yaw control. Results indicate lift generation that increases and becomes larger than drag at higher actuator strengths. These promising results show a potential for the proposed concept to replace movable flight control surfaces on future aircraft wings and empennages.

Experimental Characterization of Limit Cycle Oscillations in Membrane Wing Micro Air Vehicles

Abstract: The idea of using small-scale vehicles, often termed micro air vehicles, for various surveillance applications has become increasingly popular in recent years. A micro air vehicle design of particular interest is the membrane wing micro air vehicles, in which the structural skeleton is covered with a thin membrane instead of conventional wing skin materials, developed in particular for its lightweight nature, static stability, and passive gust rejection. In the current work, membrane wing micro air vehicles are developed and tested experimentally in order to determine the structural response of batten-reinforced membrane wing micro air vehicles to varying conditions: small angles of attack, number of battens, and membrane pretension. A self-excited instability (flutter) was noted for each model with limit cycle oscillations occurring at postflutter flow velocities. Small angles of attack had little effect on the flutter velocity, frequency, and mode for a given configuration, while increasing the membrane pretension delayed flutter and reduced the magnitude of limit cycle oscillation experienced by the model at a given flow velocity. Increasing the number of structural battens for the membrane wing micro air vehicle models also delayed the flutter velocity and reduced the magnitude of limit cycle oscillation at a given flow velocity while altering the flutter mode.

Load Alleviation on a Joined-Wing Unmanned Aircraft

Abstract: In this paper a method of alleviating wing structural load of a flexible aircraft during a symmetric balanced maneuver is presented. An application on the unmanned aircraft in development at the Italian Aerospace Research Center, high-altitude performance demonstrator, characterized by a joined-wing configuration, is illustrated. This load alleviation technique enables a desired value of the bending moment on a fixed wing control station to be obtained. The load reduction is achieved by deflecting a suitable set of flight-control surfaces, by always keeping the vertical load factor constant to preserve the maneuvering performance. The main hypotheses are: significant aeroelastic effects, linear behavior of aerodynamics and structure, and unvarying tensor of inertia under structural deflections. High-altitude performance demonstrator is a scaled performance demonstrator of an 80m-wing span high-altitude and long endurance unmanned aircraft in a joined-wing configuration. The advantages in terms of performance, fatigue life extension, and weight reduction can be achieved from the integration of an onboard load alleviation system. The results show that the attainable value of load alleviation in terms of bending moment reduction at the wing root is 37%. Moreover, the test-case analyses show that the maximum value of the alleviation increaseswithrespecttothedynamicpressurealthoughtheloaddistributionvariesbecauseofsignificantaeroelastic effects.

Numerical Simulation of the Elastic and Trimmed Aircraft

Abstract: A simulation environment has been developed enabling the computation of the elastic and trimmed aircraft. It consists of a trim algorithm which is coupled with a procedure to account for the interaction between fluid and structure. The trim algorithm is based on a Newton method with a discretized Jacobian. It incorporates the six degreesof- freedom (DoF) flight-mechanics equations and thereby enables to compute different trimmed states. The fluid-structure interaction (FSI) procedure uses a partitioned approach to compute the flow around the configuration in static aeroelastic equilibrium. This simulation environment has been successfully applied to trim a rigid transportaircraft configuration in viscous flow as well as in inviscid flow with rigid and flexible wing.

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.

Performance and Stability of an Agile Tail-less MAV with Flexible Articulated Wings

Abstract: This paper considers the problems of (a) modelling the ight mechanics of a tail-less MAV equipped with exible articulated wings, and (b) the analysis of its turning performance. The wings are assumed to have two degrees of freedom - heave and twist. They are assumed to be actuated from the root, which is the abstraction of an experimental control mechanism being developed by the authors. The dihedral and twist angles at the wing root are controlled. A novel actuator concept of axial tension to control wing stiffness has been explored in this paper. It is shown that axial tension in the wing has a significant effect on the turning performance of the aircraft, although the effect is not uniformly beneficial in nature. The effect of exibility on the steady state turning performance of the aircraft has been demonstrated by comparing it with that of a rigid aircraft, and with that of a similar aircraft possessing a wing with different elastic properties.

Aero-structural Dynamics Experiments at High Reynolds Numbers

Abstract: The elastic wing model, its excitation and comprehensive high frequency measuring equipment for the High Reynolds Number Aero-Structural Dynamics (HIRENASD) tests in the European Transonic Windtunnel (ETW) are shortly described. Some of the stationary polars are presented in terms of wing deformation, as well as aerodynamic coefficients and pressure distributions. Then unsteady processes observed in the measurements of static aerodynamic coefficients, are regarded with focus on small amplitude pressure waves travelling upstream from the trailing edge and triggering periodically break-down and redeployment of the local supersonic domains with transonic shock waves to run upstream and to disappear. Another focus is on stochastic vibrations excitation while moving forward during nominally static experiments. Emphasis is put on measured variations of pressure distribution on the wing surface caused by defined vibration excitation applying internal force couples at the wing root, whereby the exciter frequencies were chosen close to natural frequencies of the wing model. Phase and magnitude of measured local lift fluctuations as well as real and imaginary parts of pressure distributions are presented.

Controllability analysis for a flapping-wing mav with power and control actuators

Abstract: The world's first robotic insect capable of vertical takeoff was developed and demonstrated by the Harvard Microrobotics Lab. This vehicle consisted of a single power actuator that drove both wings symmetrically. In this case, there was no method to generate body torques, only lift could be modulated. The Harvard Microrobotics Lab has modified this configuration to include not only a single power actuator, but also control actuators for each wing, with the ultimate goal of being able to generate body torques. The objective of this work is to analyze the controllability of a flapping-wing micro air vehicle equipped with a power actuator to provide the mechanical power for flapping as well as left and right control actuators. The control actuators cause a small displacement of the wing root which causes a change to the wing kinematics, such that pitch and yaw torques can be generated.

Primary Modeling and Analysis of Wing Based on Aeroelastic Optimization

Abstract: An optimization design method in the preliminary design stage using aeroelastic performance indexes as constraints is proposed to design wing stiffness and build primary model. A beam-frame structure model with a high aspect ratio wing as an example is built. Stiffness distribution of the main beam is designed by engineering estimation method and aeroelastic optimization method to obtain the estimation model and the optimization model. Because of comparative precise, a three-dimensional model of the wing is introduced as the reference model for the comparison of two methods. Static aeroelastic responses are calculated, and the advantage of our method is verified through comparing and analyzing. Effect of aeroelastic performance index on stiffness distribution is also studied under critical load and small deformation, and it indicates that the stiffness of the wing root is very sensitive to the tip vertical displacement constraint.

Aerodynamics of the Wing/Fuselage Junction at an Transport Aircraft in High-Lift Configuration

Abstract: This paper presents the numerical simulation (DLR TAU-code) and the analysis of viscous high-lift flow around a complex wing/body configuration (DLR ALVAST) in landing configuration. The investigations aim for a better understanding of the aerodynamics at the wing root and the lift breakdown for such a configuration.

Multiple Degree-of-Freedom Control of a Flapping-Wing Micro Air Vehicle with Power and Control Actuators

Abstract: Flapping wing micro air vehicles have received a great deal of interest from the research community due to their potential to achieve insect-like maneuverability. The ability to mimic the flight behavior of insects could enable such a vehicle to perform missions which larger, fixed wing vehicles are unable to perform, such as intelligence, surveillance, and reconnaissance in urban environments and confined locations. The Harvard Microrobotics Lab developed the world’s first robotic insect capable of vertical takeoff. This vehicle used a single piezoelectric actuator to drive the motion of both wings symmetrically in order to generate average lift which exceeded the vehicle’s weight. There was no mechanism to produce body torques or forces in the horizontal or lateral directions. Upon extending this work, the Harvard Microrobotics Lab has modified this configuration to include not only a single power actuator, but also control actuators for each wing, with the goal of generating body torques. The power actuator provides the mechanical power for flapping, while the control actuators cause a small displacement of the wing root which changes the wing kinematics, such that pitch and yaw torques can be generated. This work utilizes the power

Fixed wing of an aircraft

Abstract: A fixed wing of an aircraft, having a main wing extending over a half-span with a wing box comprised of spars extending along the spanwise direction of the main wing, a plurality of ribs arranged one in back of the other viewed in the spanwise direction, and an outer skin. The wing box is comprised of a wing box base section (K10) and a wing box adapter section (20), that forms the outer end area of the wing box, as viewed from the wing root, and is designed to secure a wing tip device (W). In its nominal state of construction, the dihedral angle of the wing box adapter section referring to the respective local spanwise direction continuously increases by at most 60 degrees from the outer rib (R3) of the wing box base section (K10) up to the outermost rib (R5) of the wing box adapter section.

Self-locking 90-degree full-wing variable sweepback transmission mechanism

Abstract: The invention discloses a self-locking 90-degree full-wing variable sweepback transmission mechanism The transmission mechanism comprises a connecting rod assembly A, a connecting rod assembly B, a wing root rotation assembly A, a wing root rotation assembly B, a driving assembly, an upper connecting plate, a lower connecting plate and four props, wherein the upper connecting plate, the lower connecting plate and the four props form a bearing rack body; the connecting rod assembly A and the wing root rotation assembly A form a left wing motion part; and the connecting rod assembly B and the wing root rotation assembly B form a right wing motion part The transmission mechanism of the invention drives a plurality of levers to reciprocate by using a screw rod, so that wing support bars connected to the two symmetrical wing root rotation assemblies can move forwards or backwards by 90 degrees

Inflatable type flexible wing flap

Abstract: The utility model relates to an inflatable type flexible wing flap which is characterized in that the lower surfaces of two sides of a wing are adhered with inflatable type flexible wing flaps; the rectangular air cell wing flaps (2) made of rubber materials are fixedly adhered on the lower surface of the wing (1); the chordwise size of the inflatable type flexible wing flap (2) is 1.5-8 percent of the chord length of the wing, and the spanwise size thereof is 22-100 percent of the spanwise length of the wing; the distance between the rear edge of the inflatable type flexible wing flap (2) and the rear edge of the wing is 0-10 percent of the root chord length of the wing; the distance between the right end of the inflatable wing flap and the wing root of the wing 1 is 0-10 percent of the spanwise length of the wing; and an air compressor 3 is positioned inside an aeroplane body and connected with the inflatable type flexible wing flap on the lower surface of the wing. An air cell inflated is protruded out of the lower surface of the wing and changes the surface shape of the wing, thereby influencing the streaming flow field of the wing, substantially changing the pneumatic property of the wing 1 and obtaining the posture of operating an aeroplane and controlling the pneumatic force of flying of the aeroplane; and the inflatable type flexible wing flap has the characteristics of light structural weight, smooth wing surface and convenient processing and maintenance.

Deployment Dynamics of Inatable Wings

Abstract: ination rates, dynamic pressures and angles of attack to vary the wing load and bending moment. A simple analysis was used to determine the sustainable aerodynamic load prior to wrinkling of the membrane at the wing root as a function of the deployment shape. Lift peaks prior to the complete deployment due to the deployment shape, which may be used to optimize the aerodynamic characteristics.

Multi-Objective Optimization Methods Applied for Manoeuvre Load Control on Combat Aircraft wing

Abstract: Manoeuvre Load Control (MLC) is a subset of Active Control Technology (ACT), deals with optimal rotation of control surfaces to effectively redistribute the forces and moments on airframe, resulting in structural benefits and performance improvements. The present formulation consists of flight control parameters as design variables; bending and twisting moment at the wing root as conflicting independent multiple objective functions, together with aircraft stability/trim equations as equality constraints and actuator hinge moments as inequality constraints. This work attempts to bridge the gap observed with computational approaches and contemporary flight test programs, by treating MLC problem as multi-objective functions. Several multi-objective optimization methods such as Goal programming method, Multi-objective Evolutionary Algorithm (MOEA) and its hybrid form, which is a combination of the above two methods are explored for their capabilities when applied to this real world problem. Another significan...

A Bendable Load Stiffened Wing for Small UAVs

Abstract: A bendable load stiffened wing, developed at the University of Florida, has the ability to load stiffen in the positive flight load direction while remaining compliant in the opposite direction, enabling UAV storage inside smaller packing volumes. The wing employs an under-cambered airfoil with a swept planform providing dissimilar stiffness in the flight load and the folding direction. A comparative experimental study is performed using two wing geometries; straight camber and swept camber. The load stiffening ability is tested by performing three point bend tests while monitoring the wing root airfoil shape change using a visual image correlation technique. For the wing utilizing a swept camber design, increase in the root airfoil camber with increased loading resulted in a load stiffening structure. Swept camber wing showed a higher load carrying capacity (7 g's load factor) over a straight camber wing design (2 g's load factor), still maintaining the compliant nature in the folding direction. Long ter...

Cowling Design for Wing Root Regions of the Airplane with High Wing Configuration

Abstract: This paper analyze the drag reduction principle of the aircraft′s cowling region,on the basis of the displacement effect,lift effect,asymmetric effect,viscous flow effect that result from the wing-body interference.Taking the wing-body with high wing as the example,It has been discussed that the vary configuration of cowling affected on the change of aerodynamic characteristic,by calculating the RANS equations.In the paper,six factors which control the shape of the cowling has been chosen,changing one of these factors alternately and calculating each configuration,and comparing the results.At last,it concludes the several principles about cowling design from the analysis of the results rely on the flow mechanism.

Three-fuselage aircraft

Abstract: PROBLEM TO BE SOLVED: To provide a three-fuselage aircraft capable of enhancing lift without using a long outer wing with high aspect ratio, performing taking-off flight in a short distance, performing turning in a short distance and performing hovering. SOLUTION: In the three-fuselage aircraft 1, the side fuselages 4 are arranged at left and right outer side parts of the main fuselage 2 through main wings 3 long in a front/rear direction respectively. It is provided with a horizontal empennage 9; a vertical empennage 5, a propulsion unit 7; a direction rudder 6; and a lifting rudder 10. In the left and right main wings 3, chord length of a wing end part is formed longer than chord length of a wing root end part on an upper surface. COPYRIGHT: (C)2010,JPO&INPIT

Manoeuvre load alleviation using multi-objective optimisation for combat aircraft wing

Abstract: The focus of the paper is to present the systematic approach for manoeuvre load alleviation (MLA) problem with multi-objective formulation and optimisation methods. MLA is a subset of active control technology (ACT) and deals with optimal rotation of control surfaces to effectively redistribute the forces on airframe. The mathematical model consists of flight control parameters as design variables, bending and twisting moment at the wing root as multiple independent conflicting fitness functions along with aircraft stability equations as equality constraints. By and large this MLA problem is addressed as single objective optimisation. The present work emphasises multi-objective problem formulation and presents the population based NSGA-II of multi-objective evolutionary algorithm (MOEA) as solution method. Symmetric manoeuvres with vertical acceleration load cases are optimised. The approach presented provides significant reduction of wing root bending and twisting moment through optimal ELEVator and aileron (ELEVON) rotations.

Cutting tool for food processing machine

Abstract: The utility model relates to an appliance for a kitchen, in particular to a cutting tool for a food processing machine The cutting tool for the food processing machine comprises a wing root and a tool wing, wherein the wing root is provided with a mounting hole, and the tool wing is provided with an edge portion The cutting tool for the food processing machine is characterized in that a torsion angle Alpha is formed between the tool wing and the wing root, and the torsion angle Alpha is reduced from the wing root to a wing tip The utility model provides a cutting tool for the food processing machine, which has high grinding efficiency, steady operation and simple processing technology

Supersonic aircraft (versions)

Abstract: FIELD: transport. ^ SUBSTANCE: invention relates to long-range executive aircraft. Proposed aircraft comprises airframe, sweptback wing, vertical tail unit, running gear and power plant made up of engines, air intakes and nozzles. Airframe front has flatted nose cone smoothly aligned with cockpit and passenger cabin with circular sections. Wing root front edge is rounded and smoothly aligned with airframe. Wing root rear edge has a break. Vertical rudder integrated with horizontal tail unit is arranged on tip of vertical tail extension. Wing features crosswise V angle. Supersonic air intakes are arranged above wing top surface on both sides of airframe, while, ahead of air intakes, both wing and airframe are a bit contracted. Ahead of air intakes, there are perforated sections for intake of boundary layer. Supersonic air intakes comprise mechanism of controlled air cross flows from boundary layer discharge channel into channel feeding air into engine. Supersonic nozzle critical section is arranged above airframe top surface between two vertical tail fins. Flat nozzle has rotary top flap. Airframe tail section changes into flat surface to smoothly terminate in elevation rudder. Tail elevation rudder comprises mechanism of down-displacement in take-off-landing conditions. Reverse rotary panel is arranged ahead of elevation rudder above airframe top surface. Channels for reverse lower jets are arranged below said panel. ^ EFFECT: minimised effects on ecology at high cruising speeds. ^ 14 cl, 5 dwg