Showing papers on "Pitching moment published in 2002"

An Active Flow Circulation Controlled Flap Concept for General Aviation Aircraft Applications

Abstract: A recent focus on revolutionary aerodynamic concepts has highlighted the technology needs of general aviation and personal aircraft. New and stringent restrictions on these types of aircraft have placed high demands on aerodynamic performance, noise, and environmental issues. Improved high lift performance of these aircraft can lead to slower takeoff and landing speeds that can be related to reduced noise and crash survivability issues. Circulation Control technologies have been around for 65 years, yet have been avoided due to trade offs of mass flow, pitching moment, perceived noise etc. The need to improve the circulation control technology for general aviation and personal air-vehicle applications is the focus of this paper. This report will describe the development of a 2-D General Aviation Circulation Control (GACC) wing concept that utilizes a pulsed pneumatic flap.

Navier-Stokes Optimization of Supersonic Wings with Four Objectives Using Evolutionary Algorithm

Abstract: The design optimization of wings for supersonic transport by means of Multiobjective Evolutionary Algorithms is presented. Three objective functions are first considered to minimize the drag for transonic cruise, the drag for supersonic cruise and the bending moment at the wing root at the supersonic condition. The wing shape is defined by planform, thickness distributions and warp shapes in total of 66 design variables. A Navier-Stokes code is used to evaluate the aerodynamic performance at both cruise conditions. Based on the results, the optimization problem is further revised. The definition of the thickness distributions is given more precisely by adding control points. In total 72 design variables are used. The fourth objective function to minimize the pitching moment is added. The results of the revised optimization are compared with the three-objective optimization results as well as NAL’s design. Two Pareto solutions are found superior to NAL’s design for all four objective functions. The planform shapes of those solutions are “Arrow wing” type.

Experimental Investigation of Self-Actuating, Upper-Surface, High-Lift-Enhancing Effectors

Abstract: The increase in the maximum lift of an airfoil caused by small, movable tabs mounted on its upper surface has been explored in low-speed, wind-tunnel experiments at a chord Reynolds number of 1 :0 ££ 10 6 . These devices, herein called lift-enhancing effectors, have a chord that is 9% that of the airfoil and deploy passively at angles of attack approaching stall.Compared to theclean airfoil, themaximum lift coefe cientis increased by approximately 20%withthesesimpledevices.Theliftincreaseismainlycausedbytheeffectorsactingas“ pressuredams,” allowing lower pressures upstream of their location than would occur otherwise. At an effector the pressure recovers in a stepwise manner and continues downstream toward a trailing-edge value that is higher than that of the clean airfoil. This higher trailing-edge pressure also contributes to the increase in lift by allowing higher pressures over much of thelowersurface. It has been shown that, in the absence of separation, properly installed effectors will lay e ush on the surface and allow the airfoil to have the same performance in the low-drag range as the clean airfoil.

Numerical Simulations on Separation of Scaled Supersonic Experimental Airplane from Rocket Booster at Supersonic Speed

Abstract: The present paper describes the numerical methods to compute flows around the NAL Scaled Supersonic Airplane in ascending flight and its separation process from the rocket booster. The Euler equations are solved on unstructured tetrahedral grid. A seamless procedure from CAD data to grid generation is developed to treat the complex configuration of the NAL airplane piggybacked on a rocket for ascending flight. The separation process of the airplane from the rocket is efficiently simulated by the overset unstructured grid method. Comparisons of the computed results with the experiments show good agreements in the lift and pitching moment histories. It is also shown that the small parts attached between the rocket and the airplane significantly change the aerodynamic coefficients of the airplane at transonic regime.

Use of aerodynamic forces to assist in the control and positioning of aircraft control surfaces and variable geometry systems

Abstract: The current invention relates primarily to the control of rotary wing and fixed wing aircraft where a small electric actuator (7, 104) changes the pitch on a small aerodynamic surface (5, 66) and the resulting airloads on said small aerodynamic surface are used to change the pitch on a significantly larger aerodynamic surface (3, 60). In the preferred embodiment the resulting airloads on said larger aerodynamic surface is then used to change the pitch on a still larger aerodynamic surface (1, 62). As a result small electric actuators are capable of moving and controlling large aircraft control surfaces with an effective two step amplification of power utilizing the energy in the airstream. The current invention also discloses means to control and prevent undesirable motion of said aerodynamic surfaces.

Experimental investigation of trailing-edge devices at transonic speeds

Abstract: The influence of trailing-edge devices such as Gurney flaps and divergent trailing edges of different height on the aerodynamic performance of an airfoil at transonic speeds has been investigated experimentally. The investigation has been carried out in the Transonic Wind Tunnel Gottingen (TWG) using the two-dimensional airfoil model VC-Opt at freestream Mach numbers of M element of [0.755, 0.775, 0.790] and a Reynolds number of Re = 5.0 · 10 6. The results have shown that the trailing-edge devices increase the circulation of the airfoil leading to a lift enhancement and pitching-moment decrease as well as an increase in minimum drag compared to the baseline configuration. The maximum lift-to-drag ratio is considerably improved and the onset of trailing-edge flow separation is shifted to higher lift. Besides the increased rear-loading, a downstream displacement of the shock provides the main lift enhancement in transonic flow. The simple Gurney flap provides the largest additional circulation of all geometries tested. The smoother turning of the flow due to the additional ramp of the divergent trailing edge leads to a smaller increase of circulation. Slightly less lift but considerably less viscous (pressure) drag is generated enhancing the maximum lift-to-drag ratio compared to the Gurney flap. The negative affect of the Gurney flap on the pitching moment is also reduced. For the high divergent trailing edges, different ramp slopes have a significant influence on the aerodynamic performance whereas at low device heights the influence is considerably diminished. The results show that the divergent trailing edge proves to be the better trailing-edge device at transonic speeds. The application as an element for an adaptive wing is generally possible.

Dynamics of Slender Bodies Separating from Rectangular Cavities

Abstract: Vertical and pitching motions (two degrees of freedom) of a thin body of revolution separating from a rectangular cavity in a subsonic stream are investigated using combined asymptotic and numerical methods. The analysis is based on explicit analytical solutions for the lift force and pitching moment obtained in our previous studies. Body trajectory dependencies on initial conditions, body parameters, and freestream velocity are studied. The problem is divided into three phases of the motion. In phase 1, the body is inside the cavity. In phase 2, the body crosses the shear layer, and in phase 3, the body is outside the cavity. For phases 1 and 3, analytical solutions of the body dynamics are obtained for typical cases. This analysis provides insight into the separation process and identifies governing lumped nondimensional parameters relevant to the body dynamics as well providing a model that can provide quick, computationally non-intensive estimates of store separation with a personal computer. The role of the nondimensional parameters in the dynamic stability eigenvalues is identified and found particularly useful in this connection. These parameters implicitly contain the effect of the shear layer

Bubble-Induced Unsteadiness on A Wind Turbine Airfoil

Abstract: The effect of laminar separation bubbles on the surface pressure distribution and aerodynamic force characteristics of two quite different airfoils is studied numerically. The low-Reynolds-number Eppler E387 airfoil is analyzed at a chord Reynolds number of 1.0×105 whereas the NREL S809 airfoil for horizontal-axis wind turbines is analyzed at 1.0×106 . For all cases in the present study, bubble induced vortex shedding is observed. This flow phenomenon causes significant oscillations in the airfoil surface pressures and, hence, airfoil generated aerodynamic forces. The computed time-averaged pressures compare favorably with wind-tunnel measurements for both airfoils.Copyright © 2002 by Mayda, van Dam, Duque, and ASME

On the influence of spike shape at supersonic flow past blunt bodies

Abstract: In order to eliminate the appearance of a strong shock wave at a supersonic flight of a missile, which considerably increases the drag during its flight through the air, a spike is mounted on its nose. Presented paper offers the results of an experimental analysis of the influence of the spike's shape on the aerodynamic coefficients (drag, lift and pitching moment coefficient, as well as the location of the center of pressure) at supersonic flow past blunt-nosed body. The experiment was carried out in a wind tunnel, for one value of Mach and Reynolds numbers, and for one value of the angles of attack, α = 2 o . For the body without spike, and with four different spike shapes, the aerodynamic forces and moment were measured. Using only the photos obtained by Schlieren visualization of the flow, the paper proposes a criterion of estimating the aerodynamic effect of the spike shape. The best spike shape from the experimental set of spikes, selected by the measurement of the aerodynamic coefficients, coincided with this qualitative criterion.

Numerical Studies of the Jet Interaction Flowfield with a Main Jet and an Array of Smaller Jets

Abstract: During the present numerical study the jet interaction flowfield associated with the sonic injection of a gas into a high speed crossflow was simulated using the Reynolds Averaged Navier Stokes (RANS) equations. Turbulence was modeled using Wilcox's 1988 k-ω turbulence model. The computations made use of the finite volume GASP Version 4. Calculations were run for a number of jet interaction configurations consisting of a primary jet alone, a primary jet and one pair of secondary jets, and a primary jet and two pairs of secondary jets. Two flow conditions were considered: one with a Mach number of 2.4 and a pressure ratio of 14 and the other with a Mach number of 4.0 and a pressure ratio of 532. The numerical solutions were compared to the experimental results of the corresponding jet interaction tests run at Virginia Tech in order to assess the capability of RANS and of first- order turbulence models to properly simulate the complex flowfield. The k-ω turbulence model proved to be reliable and robust, and the results it provided for this type of flowfield were accurate enough from an engineering standpoint to make informed decisions about the configuration layout. In spite of the overall good performance, the k-ω turbulence model failed to correctly predict the flow in the regions of strong adverse pressure gradients. Comparisons with experimental results showed that the separation region was often under-predicted thus highlighting the need to employ second-order turbulence models. The RANS simulations was found accurate enough to provide physical mean-flow solutions. A large effort was dedicated to the development of an efficient computational grid that could capture most of the flow- physics with the least amount of cells. To this end Chimera or overset grids were employed in the simulation of the secondary injectors. The simulations showed that the innovative configuration with one primary jet and an array of smaller secondary jets effectively decrease the nose-down pitching moment by as much as 160%. In some cases it also increased the total normal force acting on the flat plate (namely the thrust) by as much as 3%. This effect was found to be caused by the reduction in size and intensity of the low- pressure region aft of the primary injector.

Envelope protection and atmospheric disturbances in icing encounters

Abstract: Research is reported on aircraft performance and control in icing, related to the development of Smart Icing Systems for improved flight safety. Microburst and gravity wave atmospheric disturbances were modeled, and their effects on the aircraft performance and control were compared to that of an icing encounter. Simulations were run using a six degree- of-freedom computational flight dynamics model. The study showed that microbursts could easily be differentiated from icing encounters. On the other hand gravity waves are more difficult to differentiate. A plan was formulated for developing an envelope protection system effective in icing conditions. Two dimensional airfoil data were analyzed and showed promising results for prediction of envelope limit exceedence. Changes in unsteady hinge moments were especially effective in predicting stall. NOMENCLATURE Cd Airfoil drag coefficient Ch Airfoil hinge moment coefficient Ch,RMS Airfoil unsteady hinge moment coefficient Cl

DARPA/AFRL Smart Wing Phase 2 wind tunnel test results

Abstract: Northrop Grumman Corporation built and twice tested a 30 percent scale wind tunnel model of a proposed uninhabited combat air vehicle under the DARPA/AFRL Smart Materials and Structures Development - Smart Wing Phase 2 program to demonstrate the applicability of smart control surfaces on advanced aircraft configurations. The model constructed was a full span, sting mounted model with smart leading and trailing edge control surfaces on the right wing and conventional, hinged trailing edge control surfaces on the left wing. Among the performance benefits that were quantified were increased pitching moment, increased rolling moment and improved pressure distribution of the smart wing over the conventional wing. This paper present an overview of the result from the wind tunnel test performed at NASA Langley Research Center's Transonic Dynamic Tunnel in March 2000 and May 2001. Successful results included: (1) improved aileron effectiveness at high dynamic pressures, (2) demonstrated improvements in lateral and longitudinal effectiveness with smooth contoured smart trailing edge over conventional hinged control surfaces, (3) chordwise and spanwise shape control of the smart trailing edge control surface, and (4) smart trailing edge control surface deflection rates over 80 deg/sec.

Prediction of Aerodynamic Coefficients Using Neural Networks for Sparse Data

Abstract: Basic aerodynamic coefficients are modeled as functions of angles of attack and sideslip with vehicle lateral symmetry and compressibility effects. Most of the aerodynamic parameters can be well-fitted using polynomial functions. In this paper a fast, reliable way of predicting aerodynamic coefficients is produced using a neural network. The training data for the neural network is derived from wind tunnel test and numerical simulations. The coefficients of lift, drag, pitching moment are expressed as a function of alpha (angle of attack) and Mach number. The results produced from preliminary neural network analysis are very good.

Determining Aerodynamic Loads Based on Optical Deformation Measurements

Abstract: A preliminary study is described for determining aerodynamic loads based on optical elastic deformation measurementsusing a videogrammetricsystem. Data reduction methods are developed and used to extract the normal force and pitching moment from beam deformation data. The axial force is obtained by measuring the axial translational motion of a movable shaft in a spring/bearing device. Proof-of-concept calibration experiments are conducted to assess the feasibility of the optical technique for measuring aerodynamic loads. The uncertainties in optical force and moment measurements are discussed. I. Introduction I NTERNAL strain gauge balances have been used for years as a standard technique for measuring the integrated aerodynamic forces and moments on models in wind tunnels. A variety of internal strain gauge balances have been developed, and the technical aspects of various balances have been studied in detail. 1 Generally speaking, the structure of an internal strain gauge balance is complicated, and the cost of fabrication is high. This paper presents an exploratory study for remotely measuring aerodynamic loads using a videogrammetic system. Unlike strain gauges, this method optically measures beam deformation to determine the normal force and pitching moment. The axial force is obtained by measuring the translational motion of a movable shaft in a spring/bearing device. Mathematical models for data reduction are developed to extract the aerodynamic forces and moments from the deformation data. Uncertainty analysis is given to evaluate the contributions from the elemental error sources and correlation terms. At this stage, the normal force, pitching moment, and axial force are the primary quantities to be determined. In principle, the side force, rolling moment, and yawing moment can be determined in a similar manner. Proof-of-concept laboratory experiments have been conducted to validate the proposed methodology for measuring the aerodynamic loads. Potentially, this optical method can be used as an alternative to strain gauge balances. In addition, the technique described in this paper can be integrated with optical model attitude and deformation measurement techniques. 2;3

Pareto Solutions of Multipoint Design of Supersonic Wings Using Evolutionary Algorithms

Abstract: Design datamining is demonstrated through multipoint aerodynamic design of supersonic wings. Tradeoff information for the design is gathered by Multiobjective Evolutionary Algorithms (MOEAs) and visualized by Self-Organizing Maps (SOMs). Four design objectives are considered: Aerodynamic drags are minimized at both supersonic and transonic cruise conditions under lift constraints. Bending and pitching moments are also minimized for structure and stability considerations. MOEAs are first performed by using 72 design variables. SOM is then applied to map the resulting Pareto solutions obtained in the four dimensional objective function space to two dimensions. This reveals global tradeoffs between four design objectives. Furthermore, the relations between design variables are mapped onto another SOM. The resulting SOMs are confirmed to perform aerodynamic design datamining properly.

Prediction of the Nonlinear Aerodynamic Characteristics of Tandem-Control and Rolling-Tail Missiles

Abstract: fin deflection angle, deg Predicted nonlinear aerodynamic characteristics of roll angle, deg several canard-body-tail missile models are presented fin set roll angle with respect to body-fixed and compared to wind tunnel data. Configurations with vertical axis, positive right wing down, deg both fin sets deflectable (tandem-control) are analyzed to investigate the effectiveness of canard-only, tail-only, INTRODUCTION and combined tandem-control effectiveness for both vertical translation and pitch attitude changes. Recently, experimental data obtained by NASA Configurations with canard control fins and free-rolling personnel has become available for tandem-control tail fin sections are investigated for their ability to missile configurations. These data exhibit many minimize vortex-induced lateral forces and moments nonlinear characteristics associated with vortical associated with canard control. Engineeringand interaction between fin sets. In addition, there are intermediate-level aerodynamic prediction codes are several sets of experimental data taken for canardused for the analysis. Results presented include high controlled missile models with fixed and free-rolling tail angle of attack aerodynamics, induced lateral forces, sections. These data also exhibit nonlinearities tandem-control fin deflections, estimates of free rotating associated with strong canard-tail vortical interference fin section performance, and rotational damping including induced lateral forces and moments. An initial estimates. Good agreement with experimental data is investigation of the ability of an engineering-level obtained for a variety of nonlinear and asymmetric flight aerodynamic prediction code to predict the conditions. characteristics of these configurations has been LIST OF SYMBOLS configurations in more detail, using both engineeringa body radius at fin mid-rootchord level and intermediate-level aerodynamic prediction AR aspect ratio (two fins joined at root) codes. C body crossflow drag coefficient dc C rolling moment/q S l TECHNICAL APPROACH l ∞ R R C roll-damping coefficient; ∂C /∂(pl /2V ) lp l R ∞ C pitching moment/q S l ; positive nose up This section summarizes the experimental data and the m ∞ R R C normal force/q S prediction methodology employed in this investigation. N ∞ R C fin normal force/q S The estimation of rolling-tail section properties is also NF ∞ R C body dC /d at =0 presented. N N D body diameter, maximum L body length DESCRIPTION OF MISL3 l ,L reference length R REF p,q,r rotational rates, rads/sec The engineering-level missile aerodynamic prediction s exposed fin span s fin semispan measured from body centerline m S reference area R x center of pressure CP x moment center MC included angle of attack, deg c _____________________ Senior Research Engineer, Senior Member AIAA *

High-Pressure Shock Tunnel Experiments and CFD Calculations on Spike-Tipped Blunt Bodies

Abstract: The requirements for the design of a new short-range high-velocity missile are the drag reduction and the acquisition of correct information for the optoelectronic sensors embedded in the hemispherical nose. High angles of attack must be studied to fulfill the maneuverability requirements of present and future missiles. A supersonic missile generates a bow shock around its blunt nose leading to high surface pressure and temperature and, as a result, the development of high drag and the damaging of embedded sensors. Pressure and temperature on the hemispherical nose surface can be substantially reduced if an oblique shock is generated by a forward facing spike. Both the experiments and the computations are carried out to study the flowfield around threedimensional blunt bodies equipped with forward facing spikes for a large range of angles of attack at a Mach number of 4.5. A blunt body, a classical disc-tipped spike, a sphere-tipped spike and a biconical-tipped spike are studied. The experiments involve high-pressure shock tunnel investigations using the shock tube facility of ISL. The differential interferometry technique is used to visualize the flowfield around the different missile spike geometries. The differential interferogram pictures and surface pressure measurements are compared with numerical results. Numerical simulations based on steady-state 3D Navier-Stokes computations are performed in order to predict the drag, the lift and the pitching moment for the blunt body and for each spike-tipped missile. The computations allow to bring out the advantages of each spike geometry in comparison with the blunt body.

Optimization of Vertical Plane Cobralike Pitch Reversal Maneuvers

Abstract: Four time-optimal, vertical plane, cobralike pitch maneuvers corresponding to different sets of boundary conditions have been studied for the F-18 high-angle-of-attack research vehicle aircraft. These maneuvers are poststall maneuvers that enable an aircraft that is initially followed by a pursuing aircraft to shift positions, that is, the initial evader turns into the pursuer. The initial pursuer is assumed to remain at constant speed and altitude. In addition an 80-deg pitch reversal maneuver has been considered that was optimized using a Chebyshev approach. A nonlinear programming and collocation method was successfully used to find the optimal vertical plane trajectories in open-loop form. All trajectories that were found have been put through a six-degree-of-freedom simulation utilizing a nonlinear inversion closed-loop control technique. The results showed that the optimal vertical plane trajectories can be flown quite accurately, using longitudinal and lateral thrust vectoring, along with the conventional aerodynamic controls.

Linearized Motion of a Fin-Stabilized Projectile Subjected to a Lateral Impulse

Abstract: An existing analytical theory for modeling the free-e ight motion of nonspinning, statically stable projectiles is extended to include the effect of a simple lateral impulse applied during e ight. The extended theory is based on the incorporation of generalized lateral translational and angular disturbances into the familiar equations of projectile free-e ight motion. The applied disturbances are then modeled using specie ed mathematical forms, and the modie ed equations are solved to obtain the angular and translational motion of the projectile over the trajectory. The various components of the translational motion of the projectile are extracted and characterized. An idealized application is presented for a large-caliber e nned projectile, representative of the class of 120-mmlong rod e nned projectiles e red from current tracked vehicle weapon systems, subjected to a single lateral control impulse in e ight. The closed-form analytical solutions are compared against results obtained using a numerical trajectory simulation code that incorporates generalized guidance and control commands. Nomenclature A = reference area, od 2 =4, m 2 CL® = derivative of aerodynamic lift force coefe cient with respect to angle of attack CM® = derivative of aerodynamic pitching moment coefe cient with respect to angle of attack

Development of a pendulum support rig dynamic wind tunnel apparatus

Abstract: This paper describes a novel form of low -cost dynamic wind tunnel testing that can overcome some of the limi tations of existing wind tunnel testing methods. In its simplest form, this Pendulum Support Rig apparatus offers a complementary technique for establishing mathematical models of manoeuvring aerodynamic loads. The 2 degree -of -freedom system in operation at Bristol is described and some initial test results are provided here for an approximate 1/16 th scale BAe Hawk model. The influence of taking measurements during manoeuvres driven by the control effectors themselves is highlighted. In the longer term, the actively controlled pendulum rig may provide a new paradigm to replace the current costly sequence of wind tunnel testing followed by control law development - particularly for small unmanned aircraft. Experimental results and computational modelling of the rig dynamics are also included, as well as modelling of lift and pitching moment parameters.

Aeroelastic analysis of helicopter rotor blades on deformable chimera grids

Abstract: Aeroelastic Reynolds-averaged Navier‐Stokes and Euler computations are presented for an articulated model rotor in hover and forward flight. Comparative rigid-blade simulations are carried out to assess the effects of blade dynamics and elasticity on the numerical results. The INROT flow solver operates on deformable structured overset grids and can be tightly coupled with a finite element model of the rotor blade structure (DYNROT) based on Timoshenko beam theory. The order of time accuracy of fluid and structure modules is maintained in the overall analysis by an appropriate staggered coupling scheme. At the investigated thrust setting, global hover performance values computed by the coupled fully turbulent Navier‐Stokes analysis agree fairly well with available experimental data. In forward flight, the aeroelastic results are in much better agreement with the measurements than those obtained from rigid-blade simulations with prescribed articulation. Apart from superior rotor power predicition, the local pitching moment coefficients computed by the viscous analysis are found to correlate better with wind-tunnel data than the corresponding Euler output.

A Study of Dynamic Stall Vortex Development Using Two-Dimensional Data from the AFDD Oscillating Wing Experiment

Abstract: : The purpose of this study is to examine the previously unpublished instantaneous pressure data of the Aeroflightdynamics Directorate Two-Dimensional (2D) and Three-Dimensional (3D) Oscillating Wing Experiment to better understand the process of dynamic stall vortex development on the NACA 0015 airfoil This report presents representative 2D instantaneous pressure data for the upper and lower surfaces of the airfoil at various chordwise locations obtained at specific angles of attack during upstroke and downstroke cycles Furthermore, the report contains a complete set of plots of instantaneous pressure distributions for the upper surface for all the 2D data sets obtained in the experiment First, the lift, drag and pitching moment data of various testing conditions are reviewed and analyzed to classify the data both with and without a boundary layer trip into "no stall," "moderate stall," and "deep stall" data Next, instantaneous pressure distributions on the upper surface of the airfoil are examined for the study of vortex development The lift and pitching moment data are analyzed to document the dynamic overshoot which delays the development of the stall on the airfoil Next, the range of angles of attack are selected where the lift and pitching moment data shows significant changes from unsteady flow behavior daring oscillation cycles Furthermore, based on the unsteady flow characteristics found in each classification of dynamic stall, analysis is continued to identify the conditions where the reduced frequency clearly affects the unsteady flow behavior of the airfoil during the oscillation This can result in a change of the dynamic stall classification of the airfoil response under various unsteady flow conditions These conditions are discussed in detail in the comparative studies

Prediction of aerodynamic coefficients for wind tunnel data using a genetic algorithm optimized neural network

Abstract: A fast, reliable way of predicting aerodynamic coefficients is produced using a neural network optimized by a genetic algorithm. Basic aerodynamic coefficients (e.g. lift, drag, pitching moment) are modelled as functions of angle of attack and Mach number. The neural network is first trained on a relatively rich set of data from wind tunnel tests of numerical simulations to learn an overall model. Most of the aerodynamic parameters can be well-fitted using polynomial functions. A new set of data, which can be relatively sparse, is then supplied to the network to produce a new model consistent with the previous model and the new data. Because the new model interpolates realistically between the sparse test data points, it is suitable for use in piloted simulations. The genetic algorithm is used to choose a neural network architecture to give best results, avoiding over-and under-fitting of the test data.

Numerical investigation of dynamic stall control via airfoil thickness variation

Abstract: Numerical experiments are conducted to investigate the effects of deformation control strategy in a dynamic stall control via active thickness variation. The aim is to clarify the dominant control parameters for an optimal thickness variation strategy. To achieve this goal, the thickness variation strategies are determined by superposing two components, a linear variation and a curved one. The effects of the control parameters on the lift stall angle, the moment stall angle, the maximum lift coefe cient, and the maximum negative pitching moment are thoroughly examined. The results indicate that dynamic stall characteristics have a strong dependency on the thickness variation strategies of each phase. From the detailed analyses, it is found that 1 ) the variation strategies in the upstroke have a signie cant effect on the vortex formation at the leading edge and the vortex shedding speed is dependent of the thickness variation in the downstroke, 2 ) the dominant parameters in this control strategy are the sign and the magnitude of the curved variation, and 3 ) there exists an optimal thickness variation strategy to control of lift characteristics or moment characteristics.

Design and development of a biomimetic device for micro air vehicles

Abstract: This paper presents the design and development of a pitching and plunging (flapping) mechanism for small-scale flight. In order to harness the unsteady lift mechanisms, used by most insects, a biologically inspired flapping/pitching device in conjunction with a rotary wing concept was developed and built. This mechanism attempts to replicate some of the aerodynamic phenomena that enhance the performance of small fliers, replacing the periodic translational motion with a unidirectional circular motion while actively flapping and pitching the rotor blades. In order to find the appropriate combination of phase, amplitude, frequency and rotational speed that leads to enhancement in lift, the device requires uncoupled independent pitch and flap actuation systems to permit the complete mapping of the parameter space. In the device under consideration the phase shift between the flapping and the pitching oscillations can be adjusted from 0 to 360 degrees over a wide range of rotational speeds. Maximum flapping and pitching amplitudes of +/- 23 degree(s) and +/- 20 degree(s) respectively can be attained. Linear displacements of two coaxial shafts are translated into the flapping and pitching motion of the rotor blades. The mechanism was designed to minimize the actuation stroke so that smart materials and conventional actuators such as motors and cams could be used. Kinematic analysis as well as experimental tests were performed. Using a customized test stand thrust and torque produced by the rotor were measured at different angles of attack, in steady-state and under periodical pitching actuation. The results showed that hover efficiency was considerably increased for a range of thrust coefficients. The device was developed based on the University of Maryland's rotary wing Micro Air vehicle (MAV) the MICOR (MIcro COaxial Rotorcraft), an electrically driven 100 g coaxial helicopter. It is anticipated that active flapping and/or pitching could be implemented in the prototype to improve its aerodynamic performance. The present paper will discuss the design and development process of a rotating/pitching/flapping mechanism for MAVs. Test results indicate that unsteady pitching motion can be used to include the aerodynamic effect of delayed stall. Performance measurements confirm that unsteady pitching motion improves efficiency in hover.