Showing papers on "Pitching moment published in 2015"
TL;DR: In this article, a series of aerodynamic shape optimization problems based on the Common Research Model wing benchmark case defined by the Aerodynamic Design Optimization Discussion Group are solved with 720 shape variables using a 28.8-million-cell mesh, reducing the drag by 8.5%.
Abstract: Despite considerable research on aerodynamic shape optimization, there is no standard benchmark problem allowing researchers to compare results. This work addresses this issue by solving a series of aerodynamic shape optimization problems based on the Common Research Model wing benchmark case defined by the Aerodynamic Design Optimization Discussion Group. The aerodynamic model solves the Reynolds-averaged Navier–Stokes equations with a Spalart–Allmaras turbulence model. A gradient-based optimization algorithm is used in conjunction with an adjoint method that computes the required derivatives. The drag coefficient is minimized subject to lift, pitching moment, and geometric constraints. A multilevel technique is used to reduce the computational cost of the optimization. A single-point optimization is solved with 720 shape variables using a 28.8-million-cell mesh, reducing the drag by 8.5%. A more realistic design is achieved through a multipoint optimization. Multiple local minima are found when starting...
246 citations
TL;DR: In this paper, the authors quantified the aerodynamic performance benefits of a morphing trailing-edge wing using aerodynamic design optimization, where the baseline geometry is optimized using a multipoint formulation and 192 shape design variables.
Abstract: Adaptive morphing trailing-edge wings have the potential to reduce the fuel burn of transport aircraft. However, to take full advantage of this technology and to quantify its benefits, design studies are required. To address this need, the aerodynamic performance benefits of a morphing trailing-edge wing are quantified using aerodynamic design optimization. The aerodynamic model solves the Reynolds-averaged Navier–Stokes equations with a Spalart–Allmaras turbulence model. A gradient-based optimization algorithm is used in conjunction with an adjoint method that computes the required derivatives. The baseline geometry is optimized using a multipoint formulation and 192 shape design variables. The average drag coefficient is minimized subject to lift, pitching moment, geometric constraints, and a 2.5g maneuver bending moment constraint. The trailing edge of the wing is optimized based on the multipoint optimized wing. The trailing-edge morphing is parameterized using 90 design variables that are optimized i...
117 citations
05 Jan 2015
TL;DR: The approach aims to introduce greater complexity and accuracy only when necessary to improve the design, and also support a greater degree of automation.
Abstract: Results are presented for four optimization benchmark problems posed by the AIAA Aerodynamic Design Optimization Discussion Group. The benchmarks are intended to exercise optimization frameworks on representative airfoil and wing design problems. All problems involve drag minimization subject to geometric and aerodynamic constraints. Our design approach involves two forms of adaptation. First, the shape parameterization is gradually and automatically enriched from an initially coarse search space. Second, adjoint solutions are used to drive adaptive mesh refinement to control discretization error. The error threshold is tailored so that the nest meshes, with the greatest accuracy, are used only when nearing the optimum. On the inviscid airfoil design problem, while reducing the drag by a factor of 10, we show how the combination of progressive parameterization and tiered discretization error control can dramatically accelerate the optimization. On the viscous airfoil design problem, we use inviscid analysis-driven optimization to reduce the total drag by a factor of two. Next, we improve the span efficiency factor of a wing by performing twist optimization. Finally, we optimize the Common Research Model wing, managing to hold drag roughly fixed, while targeting an initially-violated pitching moment constraint. Our approach aims to introduce greater complexity and accuracy only when necessary to improve the design, and also support a greater degree of automation.
50 citations
TL;DR: In this paper, a linearized version of the Goman-Khrabrov model is derived and shown to capture the dynamic hysteresis characteristics when the pitching amplitude is small.
Abstract: Dynamic hysteresis appearing in the lift force during pitching maneuvers is distinctly different from conventional static hysteresis. The size and shape of dynamic hysteresis loops are dependent on the degree of flow attachment, the dimensionless pitching frequency, and two time delays associated with the flow separation process. A linearized version of the Goman–Khrabrov model is derived and shown to capture the dynamic hysteresis characteristics when the pitching amplitude is small. Closed-loop control using a linearized version of the Goman–Khrabrov model is demonstrated, which incorporates a disturbance model into the feed-forward controller. The controller is shown to reduce the dynamic hysteresis during periodic pitching, step-up and step-down maneuvers, and quasi-random pitching maneuvers.
40 citations
TL;DR: It is found that the aerodynamic moment model became suitable when the rotational force component was assumed to act on the half-chord, which implies that the approximation borrowed from the conventional airfoil concept, i.e., the 'C.P. at the quarter-Chord' may lead to an incorrect moment prediction.
Abstract: A quasi-steady aerodynamic model in consideration of the center of pressure (C.P.) was developed for insect flight. A dynamically scaled-up robotic hawkmoth wing was used to obtain the translational lift, drag, moment and rotational force coefficients. The translational force coefficients were curve-fitted with respect to the angles of attack such that two coefficients in the Polhamus leading-edge suction analogy model were obtained. The rotational force coefficient was also compared to that derived by the standard Kutta–Joukowski theory. In order to build the accurate pitching moment model, the locations of the C.Ps. and its movements depending on the pitching velocity were investigated in detail. We found that the aerodynamic moment model became suitable when the rotational force component was assumed to act on the half-chord. This implies that the approximation borrowed from the conventional airfoil concept, i.e., the 'C.P. at the quarter-chord' may lead to an incorrect moment prediction. In the validation process, the model showed excellent time-course force and moment estimations in comparison with the robotic wing measurement results. A fully nonlinear multibody flight dynamic simulation was conducted to check the effect of the traveling C.P. on the overall flight dynamics. This clearly showed the importance of an accurate aerodynamic moment model.
38 citations
TL;DR: In this paper, a comparison of flow between an FRS and a typical FRS was conducted to evaluate the effect of rotational moment on the FRS's aerodynamic behavior, and the results showed that the aerodynamic force depends on nondimensional variables, including flapping amplitude, mean angle of attack, pitching amplitude, ratio of period of flapping to rotation motion n, and Reynolds number.
Abstract: A flapping rotary wing is a novel layout for micro air vehicle design. A computational fluid dynamics method is employed to understand the unsteady aerodynamic behavior of such a layout at a low Reynolds number. A comparison of flow between an flapping wing and a typical flapping rotary wing is conducted to evaluate the effect of rotational moment. Although the mean lift of the flapping wing is close to zero, a large mean rotational moment can drive the wing to rotate. A large mean lift coefficient can be obtained when the wing begins to rotate, but the mean rotational moment coefficient starts to decrease. The leading-edge vortex is attached to the wing surface until it moves to the trailing edge, despite the negative spanwise flow near the tip. The aerodynamic force depends on nondimensional variables, including flapping amplitude, mean angle of attack, pitching amplitude, ratio of period of flapping to rotation motion n, and Reynolds number. An analysis of these nondimensional variables shows that only...
37 citations
TL;DR: In this article, the aeroelastic behavior of a joined-wing PrandtlPlane configuration is investigated and the energy transfer between fluid and structure is carried out; the areas in which energy is extracted from the fluid are identified to gain insights on the mechanism leading to the instability.
34 citations
05 Jan 2015
TL;DR: In this article, a flow control strategy for the delay of unsteady separation and dynamic stall on a pitching NACA 0012 airfoil is explored by means of high-fidelity large-eddy simulations.
Abstract: A flow control strategy for the delay of unsteady separation and dynamic stall on a pitching NACA 0012 airfoil is explored by means of high-fidelity large-eddy simulations. The flow fields are computed employing a high-fidelity large-eddy simulation (LES) approach. The flow parameters are freestream Mach number M∞ = 0.1 and chord Reynolds numbers Rec = 5× 10. Both constant-rate and oscillatory pitching motions are considered. For the baseline cases, dynamic stall is initiated with the bursting of a contracted laminar separation bubble (LSB) present in the leading-edge region. This observation motivated a flow control approach employing high-frequency pulsed actuation imparted through a zero-net mass flow blowing/suction slot located on the airfoil lower surface just downstream of the leading edge. For the constant-rate pitching case, both pulsed and harmonic spanwise-nonuniform forcing are considered with a maximum non-dimensional frequency Stf = fc/U = 50.0 which corresponds to a sub-harmonic of the dominant natural LSB fluctuations for a baseline static case used for reference purposes. A significant delay in the onset of dynamic stall is demonstrated with pulsed forcing at high frequencies (Stf = 25.0, 50.0), however, control effectiveness diminishes with decreasing frequency. At Stf = 12.5, pulsed actuation is shown to be superior to harmonic forcing suggesting that the higher harmonic content present in the pulsed mode is still capable of energizing the LSB. For the oscillatory pitching motion, pulsed high-frequency flow control with Stf = 50.0) is considered for two cases exhibiting light and deep dynamic stall respectively. For light dynamic stall, flow actuation is capable of maintaining an effectively attached flow during the entire pitching cycle thereby inhibiting the formation of large-scale leading-edge and shear-layer vortical structures. For deep dynamic stall, control is found to also be very effective in eliminating leading-edge separation and the formation of a dynamic stall vortex. Nonetheless, trailing-edge separation eventually occurs at high incidence. For both cases, actuation provided a significant reduction in the cycle-averaged drag and in the force and moment fluctuations. In addition, the negative (unstable) net-cycle pitch damping found in the baseline cases was eliminated.
33 citations
TL;DR: In this article, a rotational method of characteristics computer program which takes into account the non-uniform inflow effects has been developed for designing asymmetric scramjet nozzles.
32 citations
TL;DR: To obtain the stability derivatives, trim conditions for linearization are numerically searched to find the exact trim trajectory and wing kinematics using an algorithm that uses the gradient information of a control effectiveness matrix and fully coupled six-degrees of freedom nonlinear multibody equations of motion.
Abstract: We show that the forward flight speed affects the stability characteristics of the longitudinal and lateral dynamics of a flying hawkmoth; dynamic modal structures of both the planes of motion are altered due to variations in the stability derivatives. The forward flight speed u e is changed from 0.00 to 1.00 m s(-1) with an increment of 0.25 m s(-1). (The equivalent advance ratio is 0.00 to 0.38; the advance ratio is the ratio of the forward flight speed to the average wing tip speed.) As the flight speed increases, for the longitudinal dynamics, an unstable oscillatory mode becomes more unstable. Also, we show that the up/down (w(b)) dynamics become more significant at a faster flight speed due to the prominent increase in the stability derivative Z(u) (up/down force due to the forward/backward velocity). For the lateral dynamics, the decrease in the stability derivative L(v) (roll moment due to side slip velocity) at a faster flight speed affects a slightly damped stable oscillatory mode, causing it to become more stable; however, the t(half) (the time taken to reach half the amplitude) of this slightly damped stable oscillatory mode remains relatively long (∼12T at u(e) = 1 m s(-1); T is wingbeat period) compared to the other modes of motion, meaning that this mode represents the most vulnerable dynamics among the lateral dynamics at all flight speeds. To obtain the stability derivatives, trim conditions for linearization are numerically searched to find the exact trim trajectory and wing kinematics using an algorithm that uses the gradient information of a control effectiveness matrix and fully coupled six-degrees of freedom nonlinear multibody equations of motion. With this algorithm, trim conditions that consider the coupling between the dynamics and aerodynamics can be obtained. The body and wing morphology, and the wing kinematics used in this study are based on actual measurement data from the relevant literature. The aerodynamic model of the flapping wings of a hawkmoth is based on the blade element theory, and the necessary aerodynamic coefficients, including the lift, drag and wing pitching moment, are experimentally obtained from the results of previous work by the authors.
31 citations
TL;DR: In this article, the authors evaluate lattice Boltzmann and finite element methods for wind load estimation of parabolic trough solar collectors using large-eddy simulation and compared with experimental results obtained in a boundary layer wind tunnel.
TL;DR: In this article, the behavior of a submarine model with a non-axisymmetric nose in pitch maneuver was investigated in a wind tunnel at Reynolds number of 6.6×10 6, based on the model length, and a range of pitch angles −10°≤ α ≤+27°.
TL;DR: In this article, the aerodynamic simulation for an airfoil section with an adaptive trailing edge flap, which deflects following a smooth deformation shape, is presented and compared with three substantially different methods: a Reynolds-averaged Navier-Stokes solver, a viscous-inviscid interaction method and an engineering dynamic stall model suitable for implementation in aeroelastic codes based on blade element momentum theory.
Abstract: The study presents and compares aerodynamic simulations for an airfoil section with an adaptive trailing edge flap, which deflects following a smooth deformation shape. The simulations are carried out with three substantially different methods: a Reynolds-averaged Navier–Stokes solver, a viscous–inviscid interaction method and an engineering dynamic stall model suitable for implementation in aeroelastic codes based on blade element momentum theory. The aerodynamic integral forces and pitching moment coefficients are first determined in steady conditions, at angles of attack spanning from attached flow to separated conditions and accounting for the effects of flap deflection; the steady results from the Navier–Stokes solver and the viscous–inviscid interaction method are used as input data for the simpler dynamic stall model. The paper characterizes then the dynamics of the unsteady forces and moments generated by the airfoil undergoing harmonic pitching motions and harmonic flap deflections. The unsteady aerodynamic coefficients exhibit significant variations over the corresponding steady-state values. The dynamic characteristics of the unsteady response are predicted with an excellent agreement among the investigated methods at attached flow conditions, both for airfoil pitching and flap deflection. At high angles of attack, where flow separation is encountered, the methods still depict similar overall dynamics, but larger discrepancies are reported, especially for the simpler engineering method. Copyright © 2014 John Wiley & Sons, Ltd.
22 Jun 2015
TL;DR: In this article, the aerodynamic environment of a high speed coaxial compound helicopter in steady level forward flight using CFD-CSD coupling is examined, and the model is validated using public-domain information and flight test data obtained for the Sikorsky X2 Technology Demonator.
Abstract: This paper examines the aerodynamic environment of a high speed coaxial compound helicopter in steady level forward flight using CFD-CSD coupling. Numerical simulations have been performed to study trim configurations of an eight-bladed coaxial compound configuration (four blades per rotor) using comprehensive analysis coupled to a vortex wake model. The model is validated using public-domain information and flight test data obtained for the Sikorsky X2 Technology Demonstrator. Rotor-rotor interference follows the expected momentum theory type model, with the lower rotor experiencing larger induced inflow. Thrust sharing between rotors remains nominally constant, with at most ±5% bias over the speed range studied. CFD analysis for the rotors is then introduced at select forward flight speeds to identify key aerodynamic modeling refinements. In particular, impulsive normal force due to rotor blade crossings are not captured accurately by the vortex wake model. The impulsive 8/rev pitching moment is completely missed by the reduced order model. While significant, this aerodynamic pitching moment does not excite a noticeable torsion response from the stiff rotor considered. For future designs with lower stiffness blades, these impulsive airloads may induce additional vibratory hub loads and drive the design of the rotor pitch control system.
TL;DR: In this paper, the authors proposed a process to obtain an optimal helicopter rotor blade shape including both planform and airfoil shape for helicopter aerodynamic performance in forward flight using an advanced geometry representation algorithm.
TL;DR: In this paper, a genetic-algorithm optimization of an objective function which maximizes the aerodynamic performance of airfoils having a larger thickness, providing better structural stiffness compared to more slender NACA design is presented.
Abstract: The present study addresses the experimental verification of the performance of a new airfoil design for lift-driven vertical-axis wind turbines (VAWTs). The airfoil is obtained through a genetic-algorithm optimization of an objective function which maximizes the aerodynamic performance of airfoils having a larger thickness, providing better structural stiffness compared to more slender NACA design. The work presents an experimental analysis of such improved performance of a 26% thick VAWT-optimized airfoil (DU12W262). The 2D flow velocity, pressure and aerodynamic loads are measured by combined use of particle image velocimetry, wall-pressure sensors and wake rakes. Additionally, the airfoil surface pressure is determined by integrating the pressure equation from the experimental velocity field. Results are initially obtained with the airfoil in steady conditions, at Reynolds 3.5*105, 7.0*105 and 1.0*106 with both free and forced (1%c) boundary layer transition. Xfoil simulations are employed for comparison with the experimental results, showing a good agreement in the linear range of angle of attack and a consistent lift/drag overestimation in the separated one. The airfoil performance is further assessed under pitching conditions (oscillatory, ramp up), at Reynolds 7.0*105 with reduced frequencies ranging from 0.07 to 0.11 and the aerodynamic load behaviour compared with the steady case. The experimental data are used as input for a numerical simulation of a 2D VAWT; while the performance are compared with those for NACA 4-series airfoils commonly used for VAWTs, showing a significantly higher maximum power coefficient for the optimized airfoil. All the data presented in the manuscript are made available in the Supporting information. Copyright © 2014 John Wiley & Sons, Ltd.
TL;DR: In this article, the feasibility of active flow control via arrays of synthetic jet actuators to mitigate hysteresis was investigated experimentally on a dynamically pitching finite span S809 blade.
Abstract: The feasibility of active flow control, via arrays of synthetic jet actuators, to mitigate hysteresis was investigated experimentally on a dynamically pitching finite span S809 blade. In the present work, a six-component load cell was used to measure the unsteady lift, drag and pitching moment. Stereoscopic Particle Image Velocimetry (SPIV) measurements were also performed to understand the effects of synthetic jets on flow separation during dynamic pitch and to correlate these effects with the forces and moment measurements. It was shown that active flow control could significantly reduce the hysteresis in lift, drag and pitching moment coefficients during dynamic pitching conditions. This effect was further enhanced when the synthetic jets were pulsed modulated. Furthermore, additional reduction in the unsteady load oscillations can be observed in post-stall conditions during dynamic motions. This reduction in the unsteady aerodynamic loading can potentially lead to prolonged life of wind turbine blades. Copyright © 2014 John Wiley & Sons, Ltd.
TL;DR: In this paper, an experimental investigation of aerodynamic control on a 35° swept flying wing by means of nanosecond dielectric barrier discharge (NS-DBD) plasma was carried out at subsonic flow speed of 20-40m/s.
TL;DR: In this paper, a quasi-three-dimensional method for wing aerodynamic analysis for drag prediction is presented, which can predict the wing drag with a level of accuracy similar to higher fidelity three-dimensional CFD analysis, with a much lower computational cost.
TL;DR: In this paper, the simulation of a two-rotor aircraft in different geometric configurations during hover flight was performed using an implementation of the unsteady vortex-lattice method (UVLM).
Abstract: This paper presents the simulation of a two-rotor aircraft in different geometric configurations during hover flight. The analysis was performed using an implementation of the unsteady vortex-lattice method (UVLM). A description of the UVLM is presented as well as the techniques used to enhance the stability of results for rotors in hover flight. The model is validated for an isolated rotor in hover, comparing numerical results to experimental data (high-Reynolds, low-Mach conditions). Results show that an exclusion of the root vortex generates a more stable wake, without affecting results. Results for the two-rotor aircraft show an important influence of the number of blades on the vertical thrust. Furthermore, the geometric configuration has a considerable influence on the pitching moment.
TL;DR: In this paper, the static ground effect of a 65° sweep delta wing (VFE-2) with sharp leading edge at α = 20 ° is investigated by the Delayed Detached Eddy Simulation (DDES) grounded on Spalart-Allmaras (S-A) turbulence model.
TL;DR: In this article, Spark discharge plasma synthetic jets (SPJ) have been used for the active flow control study on an NACA 0021 straight-wing model in a wind tunnel.
Abstract: Spark discharge plasma synthetic jets (SPJs) have been used for the active flow control study on an NACA 0021 straight-wing model in a wind tunnel. The model forces and moments were measured using a six-component sting balance at a 20 m/s wind speed. The aim was to explore the SPJ’s effect on airfoil aerodynamic by examining SPJ generators’ position along the chordwise and the jet flow direction about the chord. Near the wing leading edge, two SPJ generators raised the stall angle by 2° and increased the maximum lift coefficient by 9%. The drag coefficient was decreased by 33.1%, and the lift-drag ratio was increased by 104.2% at an angle of attack above 16°. The rolling-moment coefficient was modified by 0.002, and the yawing-moment coefficient was changed by 0.0007 at angles of attack in the range of 0°–16°. The results showed that SPJs can control wing aerodynamic forces at a high angle of attack and moments at a low angle of attack.
TL;DR: In this article, the authors used numerical simulation based on synthetic jet flow control technology to analyze the effects of typical flow control parameters and flow control mechanisms and showed that the periodic flow field disturbance introduced by synthetic jet control technology improves the blending of boundary layers, strengthens the momentum transport inside and outside the boundary layers and slows down the process of flow separation.
TL;DR: In this paper, the influence of the airfoil thickness on aerodynamic loads is investigated numerically for harmonically pitching airfoils at low incidence, under the incompressible and inviscid flow approximation.
Abstract: The influence of the airfoil thickness on aerodynamic loads is investigated numerically for harmonically pitching airfoils at low incidence, under the incompressible and inviscid flow approximation. Force coefficients obtained from finite-volume unsteady simulations of symmetrical 4-digit NACA airfoils are found to depart from the linear Theodorsen model of an oscillating flat plate. In particular, the value of the reduced frequency resulting in the inversion – from clockwise to counter-clockwise – of the lift/angle-of-attack hysteresis curve is found to increase with the airfoil thickness. Both the magnitude and direction of the velocity vector due to pitching over the airfoil surface differ from their flat-plate values. During the upstroke, namely nose-up rotation, phase, this results in a decrease (increase) of the normal velocity magnitude over the upper (lower) surface of the airfoil. The opposite occurs during the downstroke phase. This is confirmed by comparing the computed pressure distribution to the flat-plate linear Kussner model. Therefore, beyond the inversion frequency, the lift coefficient of a finite-thickness airfoil is higher during upstroke and lower during downstroke than its flat-plate counterpart. A similar dependence is also found for the quarter-chord moment coefficient. Accordingly, a modification to the classical Theodorsen model is proposed to take into account the effects of the airfoil thickness on unsteady loads. The new model is found to accurately predict the unsteady aerodynamics of a thick symmetric and a slightly cambered airfoil with a maximum thickness in the range 4–24 %. The limits of the present inviscid flow analysis are assessed by means of numerical simulation of high Reynolds number ( ) flows.
01 Jan 2015
TL;DR: In this paper, the authors presented the modeling and simulating processes of computational fluid dynamic (CFD) problem on a aircraft wing model, using typical section as NACA 2412 airfoil.
Abstract: Aerodynamic problems in general are often difficult to solve by analytics analysis. Experimental or numerical simulation can be used to analyze these computational models. However, due to the large expenses required in the experimental method, the numerical method is more preferred. This paper presents the modeling and simulating processes of computational fluid dynamic (CFD) problem on a aircraft wing model, using typical section as NACA 2412 airfoil. This wing model might be chosen in the future experimental design. ANSYS Fluent is used to analyze the pressure and velocity distribution on the surface of wing. The lift and drag forces are also determined by ANSYS Structural. Additionally, the coefficients of lift and drag forces can be calculated through the data obtained when the relative velocity inlet between the airflow and airfoil changes from 0 to 50 m/s. The numerical results shown are compatible with those of the theory, thus suggesting a reliable alternative to predict the aerodynamic characteristics of the tested wing model in fabricating the Unmanned Aircraft Vehicles (UAVs).
TL;DR: In this article, the effects of pitch phase angle and pitch amplitude on the aerodynamic properties of a two-dimensional (2D) flapping wing executing simple harmonic motion in hovering mode were investigated.
Abstract: This paper reports on a fundamental investigation of the effects of pitching phase angle (ϕ) and pitching amplitude (αA) on the aerodynamics of a two-dimensional (2D) flapping wing executing simple harmonic motion in hovering mode. A force sensor and digital particle image velocimetry were employed to obtain the time-dependent aerodynamic forces acting on the wing and the associated flow structures, respectively. Pitching phase angle ranging from 0° to 360° at three different pitching amplitudes, that is, 30°, 45° and 60°, was studied. Our experimental results revealed that the largest lift and lift/drag ratio were achieved under the condition of advanced pitching (ϕ > 90°). However, further increasing ϕ beyond a certain value would not enhance the average lift any more. In contrast, the delayed pitching (ϕ < 90°) would cause the average lift to decrease and generally the averaged drag to increase, compared to the normal pitching (ϕ = 90°), overall reducing the lift/drag ratio greatly. Our experimental results also supported the findings of Lua et al. (J Exp Fluids 51:177–195, 2011) that there are two kinds of wing–wake interactions, and they can either enhance or reduce lift on the wing depending on the wing motion and the timing of the reverse stroke. Our results show that wing–wake interaction of the first kind normally has an adverse effect on lift generation when the wing is undergoing delayed pitching but has a positive effect on the lift when the wing is undergoing advanced pitching motion. When the ϕ became larger, the second kind of wing–wake interaction, that is, sliding of the leading edge vortex under the wing, will cause the concurrent fall in lift and drag.
TL;DR: In this article, a weak coupling program was developed to simulate the interaction between the variable camber plunge airfoil and fluid, and the flow field, aerodynamic force and energy efficiency of different camber airfoils under different wind gust conditions were investigated.
Abstract: Variable camber deformation is observed during the flight of some insects and bird species; however, the effect of this special airfoil shape motion on the aerodynamic characteristics of the airfoil is not well understood, especially for the airfoil under gust wind. We did a numerical study to investigate the aerodynamic characteristics of a variable camber plunge airfoil under wind gusts. A weak coupling program was developed to simulate the interaction between the variable camber plunge airfoil and fluid, and the flow field, aerodynamic force and energy efficiency of different camber airfoils under different wind gust conditions are investigated. It was found that camber deformation influences the aerodynamic characteristics of the airfoil greatly. If the airfoil has an appropriate camber deformation, the deformation can increase the mean thrust and the propulsive efficiency of the airfoil. Moreover, the aerodynamic characteristics of the appropriate camber airfoils are not significantly affected by the gust frequency, and there exists a range of gust amplitude where the aerodynamic characteristics of the airfoils are also not significantly affected by the gust amplitude, which may be beneficial for aerodynamic stability of the airfoil. The results also show that appropriate camber deformation can suppress leading edge vortex separation, which improves the aerodynamic characteristics of the airfoil.
22 Jun 2015
TL;DR: In this article, a hybrid 2D/3D aerodynamic approach was used by combining 3D linear vortex lattice method with known 2D airfoil viscous data for multidisciplinary optimization of flexible wings in high-lift configurations, such as takeoff and landing.
Abstract: A rapid conceptual 3D aerostructural tool was developed for multidisciplinary optimization of flexible wings in high-lift configurations, such as takeoff and landing. In order to allow fast optimization while still capturing nonlinear aerodynamic phenomena such as the maximum lift near the stall, a hybrid 2D/3D aerodynamic approach was used by combining 3D linear vortex lattice method with known 2D airfoil viscous data. A two-variable ”decambering” method was employed to match the sectional lift and pitching moment coefficients along the wing span with that of the corresponding 2D viscous data, in a multivariable Newton-Raphson iterative scheme. Validation was done against experimental data of a straight wing with known both 3D and 2D experimental data, demonstrating successful prediction of lift, stall progression along the wing span, and decrease in pitching moment due to stall. The aerodynamic tool was coupled with a structural finite element method code to compute static aeroelastic deflections of a given wing, and subsequently integrated with an optimization routine. The resulting aeroelastic optimization framework was then applied to a low-Reynolds simplified version of the Generic Transport Model with variable camber continuous trailing-edge flaps, in high-lift configuration, to predict the optimal flap schedule for maximum lift. Results showed that for a rigid wing, the optimal flap deflection corresponded to a lift distribution similar to the chord distribution, while a flexible wing was dependent on its torsional and bending stiffness. For the wing stiffness and dynamic pressure utilized, the optimal flap deflection schedule showed to be similar to the rigid wing results, differing slightly by having a less loaded wing tip region, to avoid the observed outboard nose-up torsion tendency that would prematurely stall the wingtip.
TL;DR: In this paper, the authors investigate the mechanism that underlies the development of unsteady loads by a novel L-shaped Gurney flap conceived to perform vibration control on rotorcraft blades.
Abstract: The purpose of the work is to investigate the mechanism that underlies the development of unsteady loads by a novel L-shaped Gurney flap conceived to perform vibration control on rotorcraft blades. The device is combination of a spoiler with a Gurney flap. Exploiting the capabilities of a Reynolds-averaged Navier–Stokes flow solver employing the overset mesh approach, several numerical simulations are carried out at low Mach number. These simulations are used to develop a physically based linear reduced-order model in the frequency domain for the unsteady lift and pitching moment of a NACA 0012 airfoil, considering as input the pitch and plunge harmonic oscillations of the airfoil, together with the oscillations of the L-shaped Gurney flap. The aerodynamic assessment of the L-tab shows that the behavior of the loads can be predicted using an equivalent flat-plate model to represent the airfoil composed by three segments: the first representing the fixed part of the airfoil, the second representing the lon...
TL;DR: A computational analysis of the pitching dynamics by incorporating the realistic wing kinematics to determine the inertial effects shows that like many insects, pitch reversal of the hummingbird is, to a large degree, caused by the wing inertia.
Abstract: In hovering flight, hummingbirds reverse the angle of attack of their wings through pitch reversal in order to generate aerodynamic lift during both downstroke and upstroke. In addition, the wings may pitch during translation to further enhance lift production. It is not yet clear whether these pitching motions are caused by the wing inertia or actuated through the musculoskeletal system. Here we perform a computational analysis of the pitching dynamics by incorporating the realistic wing kinematics to determine the inertial effects. The aerodynamic effect is also included using the pressure data from a previous three-dimensional computational fluid dynamics simulation of a hovering hummingbird. The results show that like many insects, pitch reversal of the hummingbird is, to a large degree, caused by the wing inertia. However, actuation power input at the root is needed in the beginning of pronation to initiate a fast pitch reversal and also in mid-downstroke to enable a nose-up pitching motion for lift enhancement. The muscles on the wing may not necessarily be activated for pitching of the distal section. Finally, power analysis of the flapping motion shows that there is no requirement for substantial elastic energy storage or energy absorption at the shoulder joint.