Showing papers on "Vortex lattice method published in 2017"

Experimental and Numerical Study of Micro-Aerial-Vehicle Propeller Performance in Oblique Flow

Abstract: This paper presents the modeling of the performance of small propellers used for vertical takeoff and landing micro aerial vehicles operating at low Reynolds numbers and in oblique flow. The blade element momentum theory, vortex lattice method, and momentum theory for oblique flow are used to predict propeller performance. For validation, the predictions for a commonly used propeller for vertical takeoff and landing micro aerial vehicles are compared to a set of wind-tunnel experiments. Both the blade element momentum theory and vortex lattice method succeed in predicting correct trends of the forces and moments acting upon the propeller shaft, although accuracy decreases significantly in oblique flow. For the dataset analyzed here, combining the available data of the propeller in purely axial flow with the momentum theory for oblique flow and applying a correction factor for the wake skew angle results in more accurate performance estimates at all elevation angles.

Application of Vortex Methods to Coaxial Rotor Wake and Load Calculations in Hover

Abstract: The coaxial rotor has distinct advantages in hover performance and forward flight compared to a conventional isolated rotor. The viscous vortex particle approach based on the Lagrangian formulation...

Higher-Order Free-Wake Method for Propeller–Wing Systems

Abstract: A higher-order free-wake (HOFW) method has been developed to enable conceptual design-space explorations of propeller–wing systems. The method uses higher-order vorticity elements to represent the ...

Continuous-Time State-Space Unsteady Aerodynamic Modeling for Efficient Loads Analysis

Abstract: The present paper proposes a continuous-time state-space formulation of the unsteady vortex lattice method, which is derived through a discretization of the governing advection equation for transpo...

Vortex Lattice simulations of attached and separated flows around flapping wings

Abstract: Flapping flight is an increasingly popular area of research, with applications to micro-unmanned air vehicles and animal flight biomechanics. Fast, but accurate methods for predicting the aerodynamic loads acting on flapping wings are of interest for designing such aircraft and optimizing thrust production. In this work, the unsteady vortex lattice method is used in conjunction with three load estimation techniques in order to predict the aerodynamic lift and drag time histories produced by flapping rectangular wings. The load estimation approaches are the Katz, Joukowski and simplified Leishman–Beddoes techniques. The simulations’ predictions are compared to experimental measurements from wind tunnel tests of a flapping and pitching wing. Three types of kinematics are investigated, pitch-leading, pure flapping and pitch lagging. It is found that pitch-leading tests can be simulated quite accurately using either the Katz or Joukowski approaches as no measurable flow separation occurs. For the pure flapping tests, the Katz and Joukowski techniques are accurate as long as the static pitch angle is greater than zero. For zero or negative static pitch angles, these methods underestimate the amplitude of the drag. The Leishman–Beddoes approach yields better drag amplitudes, but can introduce a constant negative drag offset. Finally, for the pitch-lagging tests the Leishman–Beddoes technique is again more representative of the experimental results, as long as flow separation is not too extensive. Considering the complexity of the phenomena involved, in the vast majority of cases, the lift time history is predicted with reasonable accuracy. The drag (or thrust) time history is more challenging.

Flexible hydrofoil optimization for the 35th America's cup with constrained Ego Method

Abstract: This paper investigates the use of constrained surrogate models to solve the multi-design optimization problem of a flexible hydrofoil. The surrogate-based optimization (EGO) substitutes the complex objective function of the problem by an easily evaluable model, constructed from a limited number of computations at carefully selected design points. Associated with ad-hoc statistical strategies to propose optimum candidates within the estimated feasible domain, EGO enables the resolution of complex optimization problems. In this work, we rely on Gaussian processes (GP) to model the objective function and adopt a probabilistic classification method to treat non-explicit inequality constraints and non-explicit representation of the feasible domain. This procedure is applied to the design of the shape and the elastic characteristics of a hydrofoil equipped with deformable elements providing flexibility to the trailing edge. The optimization concerns the minimization of the hydrofoil drag while ensuring a non-cavitating flow, at selected sailing conditions (boat speed and lifting force). The drag value and cavitation criterion are determined by solving a two-dimensional nonlinear fluid-structure interaction problem, based on a static vortex lattice method with viscous boundary layer equations, for the flow, and a nonlinear elasticity solver for the deformations of the elastic components of the foil. We compare the optimized flexible hydrofoil with a rigid foil geometrically optimized for the same sailing conditions. This comparison highlights the hydrodynamical advantages brought by the flexibility: a reduction of the drag over a large range of boat speeds, less susceptibility to cavitation and a smaller angle of attack tuning range.

Induced Drag Calculations with the Unsteady Vortex Lattice Method for Cambered Wings

Abstract: A = panel area, m b = semichord, m c = chord, m Cd = sectional drag coefficient Cl = sectional lift coefficient D = panel drag, N F = force vector, N h = plunge displacement, m k = reduced frequency L = panel lift, N l = vortex segment vector M = number of chordwise panels N = number of spanwise panels n = panel normal vector P = orthogonal projection operator s = nondimensional time U = velocity vector, m · s−1 t = time, s α = angle of attack, rad ∆b = panel span, m ∆c = panel chord, m ∆p = pressure drop Γ = circulation, m · s ω = angular velocity, rad · s−1 *PhD Student, Aerospace and Mechanical Engineering Department, t.lambert@ulg.ac.be †Associate Professor, Aerospace and Mechanical Engineering Department, AIAA Senior Member, gdimitiradis@ulg.ac.be

Design and Testing of Aeroelastically Tailored Wings Under Maneuver Loading

Abstract: The goal of the present paper is to provide experimental validation data for the aeroelastic analysis of composite aeroelastically tailored wings with a closed-cell cross-sectional structure. Several rectangular wings with different skin thicknesses and composite layups are designed in order to minimize the root bending moment under maneuver loading using an aeroelastic analysis framework that closely couples a geometrically nonlinear beam model to a vortex lattice aerodynamic model. The globally convergent method of moving asymptotes is used to derive an optimized layup for the tailored wings. In addition, a quasi-isotropic wing is analyzed to serve as a reference. Both the tailored wings and the quasi-isotropic wing are manufactured and tested structurally and in the wind tunnel. In the wind tunnel, aerodynamic forces and moments and wing deformation are measured to provide experimental validation data.

Aeroelastic Optimization Design for High-Aspect-Ratio Wings with Large Deformation

Abstract: This paper presents a framework of aeroelastic optimization design for high-aspect-ratio wing with large deformation. A highly flexible wing model for wind tunnel test is optimized subjected to multiple aeroelastic constraints. Static aeroelastic analysis is carried out for the beamlike wing model, using a geometrically nonlinear beam formulation coupled with the nonplanar vortex lattice method. The flutter solutions are obtained using the - method based on the static equilibrium configuration. The corresponding unsteady aerodynamic forces are calculated by nonplanar doublet-lattice method. This paper obtains linear and nonlinear aeroelastic optimum results, respectively, by the ISIGHT optimization platform. In this optimization problem, parameters of beam cross section are chosen as the design variables to satisfy the displacement, flutter, and strength requirements, while minimizing wing weight. The results indicate that it is necessary to consider geometrical nonlinearity in aeroelastic optimization design. In addition, optimization strategies are explored to simplify the complex optimization process and reduce the computing time. Different criterion values are selected and studied for judging the effects of the simplified method on the computing time and the accuracy of results. In this way, the computing time is reduced by more than 30% on the premise of ensuring the accuracy.

Wing Optimisation for Tractor Propeller Configurations: Validation and Application of Low-Order Numerical Models Adapted to Include Propeller-Induced Velocities

Abstract: Even though propellers are the oldest form of propulsion, they are still a popular choice for unmanned aerial vehicles (UAVs) and passenger aircraft in certain market segments. If the propellers are mounted on the wing, strong propeller-wing interactions alter the aerodynamic efficiency of the aircraft. Using low-order numerical models, it is shown in literature that the wing chord and twist distribution can be changed to maximise this efficiency. These results are only theoretical. This thesis, therefore, aims to validate and apply a numerical model for optimisation of the wing design taking propeller-wing interactions into account. A vortex lattice method (VLM) was adapted to include the effects of propeller-induced velocities. Comparing the results of the adapted VLM with existing experimental data already validates the numerical model for predicting the lift distribution. To validate it for changes in lift distribution due to wing design changes, a wind-tunnel experiment is set up. Two wings are tested in a tractor propeller configuration. The only difference between the wings is in the twist distribution. To find the lift distribution, the circulation is evaluated in the flow around a wing at several stations along the wingspan. Particle-image velocimetry was used to obtain this flow field. Indeed, the lift distributions measured on both wings are matched by predictions from the adapted VLM, which proofs the numerical model is suitable for qualitative optimisation studies. For the wing and operating conditions used for the wind-tunnel experiment, an optimisation study is performed using the adapted VLM. It shows the drag can be reduced with 34% by adopting the optimal chord and twist distribution. Even though the operating conditions are not representative for full-scale aircraft or UAVs, it does show there is a great potential for taking propeller-wing interactions into account for the design of the wing.

Stochastic bifurcation analysis of an elastically mounted flapping airfoil in an inviscid fluid

Abstract: Biological flyers take the advantage of FSI to augment their propulsive efficiency by exploiting the coupling between the flexible wings and the surrounding unsteady flow-field. A proper understanding of the role of FSI in natural flights is essential for the design of biologically-inspired Micro Aerial Vehicles (MAVs). In field conditions, the flow is typically accompanied by irregular fluctuations which significantly affect the performance of very light weight flapping wing MAVs, that are primarily aimed for indoor applications. The present study focusses on carrying out a stochastic bifurcation analysis to understand the effects of fluctuating flow on the dynamical stability characteristics of the coupled dynamical system. In this study, a span-wise flexible wing has been modelled by an elastically mounted airfoil supported by translational and rotational cubic nonlinear springs along the plunge (bending) and pitch (torsion) degrees of freedom respectively. The nonlinear structural model has been coupled with an inviscid flow solver using a weak coupling strategy to build the FSI framework. The flow part is solved using the unsteady vortex lattice method (UVLM). A bifurcation analysis has been carried out considering the mean wind speed to be the bifurcation parameter. In sterile flow conditions, the system undergoes a Hopf bifurcation as the free-stream velocity is increased resulting in stable limit-cycle oscillation (LCO) from a fixed point response. However, when long time scale gusts are superimposed to the mean flow, the dynamics of the system undergoes qualitative changes with the appearance of an intermittency regime prior to the onset of noisy LCOs. A P-bifurcation analysis based on the transition in the topology associated with the structure of the joint pdf of the response variables reveals that the joint pdf corresponding to the stochastic fixed point response (Fig. 1(a)) possesses a Dirac delta function like structure with a sharp single peak around zero at low mean wind speed. As the mean wind speed increased, one can see in the joint pdf that along with the peak around zero value, a new weak attractor is born (Fig. 1(b)). In this regime, the system experiences on-off intermittency. As the mean wind speed is further increased, the joint pdf bifurcates to a crater-like structure corresponding to the random LCO (Fig. 1(c)). The present paper further focusses on the stochastic bifurcations of the coupled system in the post-flutter regime in the presence of an actuating force the details of which will be presented in the full paper.

Ready for Take-Off: Scaled Flight Testing

Abstract: Atmospheric free flight scaled flight testing is an affordable way to investigate the dynamic properties of an aircraft, while enabling a wider range of test possibilities than a windtunnel. This research is the first step into the development of a scaled flight testing model and as such will highlighted the difficulties and discrepancies which will be faced when scaling is performed. This study has the objective of investigating the effect of scaling on the aerodynamic properties and of the flow field at the tail location, with the focus on take-off conditions. The full scale case is a regional aircraft, the ATR72, with an unswept and slightly tapered wing. The scaled case is 14.7% geometric scaled The analysis of the aerodynamic properties is split between the clean wing and the wing with an additional surface in the form of a flap. The aerodynamic properties investigated were the lift, moment and drag curve, and the maximum lift coefficient. The clean wing was analysed using a quasi-3D analysis called Q3D, which is a combination of a vortex-lattice method (AVL) and a vortex panel method (XFOIL). A modification of XFOIL, called RFOIL, is used to analyse the difference in maximum lift coefficient between the airfoil of the scaled and full scale case. RFOIL is selected because of the better results near stall. The wing with flap is investigated using a vortex lattice method, in this case again AVL, where the airfoil with flap is investigated using the Euler-solver MSES. Finally the maximum lift coefficient for the wing with flap is analysed with the semi-empirical Pressure Difference Rule. The Pressure Difference Rule states that there is a ratio between the peak pressure and the trailing edge pressure, on a surface, at which maximum lift occurs. The methods are limited in their incorporation of viscous effects, the 2D-analysis tools (XFOIL, RFOIL and MSES) only include viscous effects in the small boundary layer region. AVL does not include viscous effects directly, but only via a correction factor for the lift-curve slope. However, the results found in literature using these methods are satisfactory. An analysis of the wake field is done using methods developed by the ESDU. The wake properties of both the scaled as well as the full scale are calculated using semi-empirical models and use a simplification where to place the vortex sheet. The dimensions and velocity loss of the wake itself can be calculated using either a method by ESDU or a method by Schlichting. Results showed that both these methods provided similar solutions. The scaling of the wing proves to change the aerodynamic properties and the similarity is no longer present for all the investigated aerodynamic properties. For both the wing with and without flap this is the case. These differences are mainly due to Reynolds number effects, the Mach number effects are only minimal on the lift coefficient. Due to the limits on the selected methods the exact magnitude of the difference found can not be guaranteed, but the trends in the found differences are certain. A basis for this is found in the different boundary layer properties, the scaled wing exhibits a thicker boundary layer, leading to a decambering effect, and laminar separation bubbles are found to occur. Both the clean and wing with flap shows more outboard loading for the scaled wing. The scaled wing exhibits a thicker wake, with a larger velocity loss. The difference between the full and scaled clean wing is thus larger than the difference between the full and scaled wing with flap. The reason for difference between the full and scaled wing can be found in the increased drag (coefficient) of the scaled wing. The reason behind the larger increase in wake for the clean wing is due to the fact that the difference in drag coefficient between the full and scaled case is larger for the clean wing than for the wing with flap. An investigation to minimize the differences between the full and scaled wing is done. It is decided to change the airfoil shape. Improvements are visible for the optimized airfoil, on all the aerodynamic properties. The optimized airfoil shapes tends towards a thinner airfoil and thus an investigation into solely optimizing the thickness of the airfoil is also done. This also proves to give better results than the original airfoil, however, not as good as the shape optimized airfoils.. This thinner profile was also investigated in the configuration with a flap, but here it proved to decrease the performance of the airfoil with flap. The difference between the full and scaled case can be reduced by an optimization. If an optimization is to be done, it must be done on the whole of configurations and flight conditions. As the next step into a development of a scaled flight test model, the exact extend of tests must be determined, only then can it be investigated how scaling affects the results of testing. No direct solutions exist to completely overcome the gap between the full and scaled aerodynamic properties of a flight testing model, however airfoil shape optimization does provide a better similarity.

High Fidelity Modelling for High Altitude Long Endurance Solar Powered Aircraft

Abstract: High Altitude Long Endurance (HALE) platforms are the aerial platforms capable of flying in the stratosphere for long periods of time. This master thesis presents aircraft system identification procedures geared towards such fixed wing platforms where aerodynamic forces and moments are parametrically modelled with so-called stability and control derivatives. The first part of the thesis addresses local System identification procedures intended for controller synthesis at low altitude flights whereas the second part of the thesis deals with a preliminary study on a new global system identification method. The local system identification procedure is based on the two step method, which offers flexibility regarding the aerodynamic structure. Therefore, it is suitable for the development of a system identification tool chain for various fixed wing platforms. Various system identification experiments have been conducted to collect flight test data. The parameters for the estimation of aerodynamic forces and moments are then found through an optimization procedure. Such parameters have been validated using a validation set from flight test data and their applicability for controller synthesis has been demonstrated. Global system identification typically requires the collection of flight test data at multiple points in the flight envelope and often, is combined with extensive Computational Fluid Dynamics (CFD) solutions as well as wind-tunnel experiments. Such an approach is time consuming and costly. This thesis presents a new method to overcome the limitations of the current methodology by applying a Parameter search on VLM-based (Vortex Lattice Method) dynamic simulations of aircraft System identification manoeuvres and correcting the estimated models with available flight test data. The current study shows improvements in fidelity with decrease in Root Mean Squared Error (RMSE) by factor 0.2 and 0.5 for x-axis and z-axis forces in body frame respectively, while reducing the effort for obtaining a model with similar fidelity.

High-Alfa Aerodynamics with Separated Flow Modeled as a Single Nascent Vortex

Abstract: A numerical iterative vortex lattice method is developed to study flow past wing(s) at high angles of attack where the separated flow is modelled using NY nascent vortex filaments. The wing itself is modelled using NX × NY bound vortex rings, where NX and NY are the number of sections along the chord and span of the wing respectively. The strength and position of the nascent vortex along the chord corresponding to the local effective angle of attack are evaluated from the residuals in viscous and potential flow, i.e. (Cl ) visc − (Cl ) pot and (Cm ) visc − (Cm ) pot . Hence, the 2D airfoil viscous Cl − α and Cm − α is required as input (from experiment, numerical analysis or CFD). Aerodynamic characteristics and section distribution along span are predicted for 3D wings at a high angle of attack. Effect of initial conditions and existence of multiple solutions in the post-stall region is studied. Results are validated with experiment.

Optimization design for aero-elastic morphing wings

Abstract: This paper presents an optimization design scheme for aero-elastic morphing regional jet wings. In terms of the requirement of a desired aircraft throughout the flight envelope involving altitude (e.g. 0–13000m) and Mach (e.g. M0.78–0.25), an inverse design based on vortex lattice method is practiced for the mean camber shapes and twist distribution of the wing, subject to the optimal aerodynamic performance, minimum lift-induced drag in both the cruise and off-cruise conditions. Subject to a range of different structural and aero-elastic constraints, a static aero-elastic optimization following a gradient based approach is developed to decide the stiffness distribution and the jig-shape required at cruise to achieve a minimum weight wing design. Based on the aerodynamic and structure optimization, morphing control strategy is discussed by using adaptive torsion wing structure scheme. Analysis results illustrate that torsional stiffness varies linearly as the deflection of the movable spars in condition of thin-wall and flat section.

Semi-Analytical Closed-Wing Weight Estimation during Conceptual Design

Abstract: One of the novel configurations that could revolutionize the aviation industry is the PrandtlPlane. The closed-wing design features a low front wing and a high rear wing, connected by a lateral surface. In this research, the rear wings of the configuration are attached to a set of fins. Previous conceptual closed-wing design studies have failed to accurately predict the wing weight of the configuration, by using empirical relations or semi-analytical methods designed for conventional wings. These methods fail to capture the three major structural characteristics of a closed-wing design: an over-constrained structure, significant secondary bending moments and shear forces, and a different lay-out of the constraints definition. Therefore, a semi-analytical closed-wing weight estimator was developed. The primary weight is estimated analytically, while the secondary weights are defined by empirical relations. The primary structure is sized using an equivalent beam method, which is designed to withstand the aerodynamic, inertial and fuel loads applied to the structure. The aerodynamic loads are estimated with a vortex lattice method, at a 2.5g pull-up manoeuvre. For the PrandtlPlane configuration, the fin is modelled as a support to the rear wing. Due to the over-constrained nature of the structure, a stiffness iteration loop is implemented in which the internal loads are determined using the displacement method and all cross-sections are sized. The cross-sectional design features four booms, which are cross-coupled and four skins, which are sized independently. A total weight estimation of a closed-wing design takes 20-30 seconds. Sensitivity analyses are performed to establish the main drivers of the total wing weight and its distribution. Wingspan, wing sweep angle, fin sweep angle and longitudinal center of gravity position were determined to have the most influence on the total wing weight and its distribution. The new tool was implemented in the conceptual design workflow of the Initiator. A comparison between a design study in the Initiator with the closed-wing weight estimator and the previous methodology showed large discrepancies. Three design studies were performed with missions varying in payload weight, number of passengers and range. The total offset in the wing weight between the two methods ranged from 4.5\% to 30.4\%, leading to discrepancies in the required fuel mass of 1.6% and 5.7% respectively. Apart from the large offsets in total wing weight, the previous methodology also failed to accurately predict the distribution of the weight between the front, rear and lateral wing. Furthermore, the offsets between the methodologies are case-dependent and no clear relationship between them can be distinguished. Future conceptual design studies should thus include a closed-wing weight estimator. Finally, a parametric study was performed to identify the effect of altering the wing area ratio between the front and rear wing on the required fuel mass. For a single aisle, medium range mission, a design with and area ratio of 1.25 is 2\% more fuel efficient compared to a design with an even area distribution between the front and rear wing.

Propulsion efficiency of flapping flight robots

Abstract: Flapping wing technologies have been successfully applied to robots either flying in the air or gliding through the water owing to the unique advantages of flapping flight under small Re conditions Propulsion efficiency of flapping-twisting compound motion wings are analyzed in the present study Actuator disc theory and undulating wave theory are utilized to model the wings' propulsion efficiency, and the present method is verified by a Lattice Boltzmann method(LBM) based code, on an 800mm wing-span virtual flapping wing model Compared with conventional methods including unsteady method, vortex lattice method or numerical method, the present method is much simpler and relates basic design parameters of flapping wings directly to the efficiency output In this way, the present method provides a macroscopic view towards propulsion mechanisms of flapping wings, and has the potential to greatly accelerates the process of flapping wing robots' first period design, which usually rely on experience or parameters adopted from flapping flight creatures

A Quick Computational Method for Improving Aerodynamic Shape of UAV Wing

Abstract: To obtain an optimal aerodynamic 3D shape of small-sized UAV wing at small flight speed and high lift coefficient, the optimization problem is set to minimize drag coefficient with fixed plane form and constant lift coefficient. The thickness of chordwise function is assumed to be given. The direct optimization problem must be solved by CFD methods with viscosity consideration in 3D flow that involves a great volume of computations which is feasible only to super computers [1, 6]. This chapter presents a combined direct inverse method that makes the optimization problem be feasible to ordinary PC.

Morphing wing application on Hydra Technologies UAS-S4

Abstract: This paper presents the aerodynamic results of a morphing wing study performed on the UAS S4 Ehecatl from Hydra Technologies. Only the cruise phase of the aircraft was considered (constant altitude and constant speed). The difference, from an aerodynamic point of view, between the morphing wing and the original wing was emphasized by computing and comparing their longitudinal aerodynamic coefficients (drag and lift). The computation of the aerodynamic characteristics was done using tornado with the Vortex Lattice Method.