Showing papers on "Vortex lattice method published in 2007"

Design of fixed wing micro air vehicles

Abstract: The paper describes the methodology and computational design strategies used to develop a series of fixed wing micro air vehicles (MAVs) at the Ghent University. The emphasis of the research is to find an optimal MAV-platform that is bound to geometrical constraints but superior in its performance. This requires a multidisciplinary design optimisation but the challenges are mainly of aerodynamic nature. Key areas are endurance, stability, controllability, manoeuvrability and component integration. The highly three-dimensional low Reynolds number flow, the lack of experimental databases and analytical or empirical models of MAV-aerodynamics required fundamental research of the phenomena. This includes the use of a vortex lattice method, three-dimensional CFD-computations and a numerical propeller optimisation method to derive the forces and their derivatives of the MAV and propeller for performance and stability-related optimisation studies. The design method leads to a simple, stable and robust flying wing MAV-platform that has the agility of a fighter airplane. A prototype, the UGMAV25, was constructed and flight tests were performed. The capabilities of the MAV were tested in a series of successful flight manoeuvres. The UGMAV15, a MAV with a span of 15cm, is also developed to test flight-qualities and endurance at this small scale. With the current battery technology, a flight-time of at least one hour is expected.

Prediction of Sheet Cavitation on a Rudder Subject to Propeller Flow

Abstract: This paper presents two numerical methods, a vortex lattice method (MPUF-3A) coupled with a finite volume method (GBFLOW-3D) and a boundary element method (PROPCAV), which are applied to predict time-averaged sheet cavitation on rudders, including the effects of the propeller as well as of the tunnel walls. The coupled MPUF-3A and GBFLOW-3D determines the velocity field due to the propeller within the fluid domain bounded by tunnel walls. MPUF-3A solves the potential flow around the propeller by distributing the line vortices and sources on the blade mean camber surface and determines the pressure distributions on the blade surface. GBFLOW-3D solves Euler equations with the body force terms converted from the pressure distributions on the blade surface and determines the total velocity field inside the fluid domain. The tunnel walls are treated as a solid boundary by applying the slip boundary condition, and the propeller blades are modeled via body forces. The two methods are solved iteratively until the forces on the blade converge. The cavity prediction on the rudder is accomplished via PROPCAV, which can handle back and face leading edge or mid-chord cavitation, in the presence of the three-dimensional flow field determined by the coupled MPUF-3A and GBFLOW-3D. The present method is validated by comparing the cavity shapes and the cavity envelope with those observed and measured in experiment and computed by another method.

Potential of Articulated Split Wingtips for Morphing-Based Control of a Flying Wing

Abstract: This paper investigates a novel method for the control of ‘morphing’ aircraft. The concept consists of a pair of articulated split wingtips, independently actuated and mounted on a baseline flying wing. The general philosophy behind the concept was that adequate control of a flying wing about its three axes could be obtained through local modifications of the dihedral angle at the wingtips, thus providing an alternative to conventional control eectors such as elevons and drag rudders. Preliminary computations with a vortex lattice model and subsequent wind tunnel tests demonstrate the viability of the concept, with individual and/or combined wingtip deflections producing multi-axis, coupled control moments. Nomenclature b wing span Cl, Cm, Cn rolling, pitching and yawing moment coecients CDi lift-induced drag coecient CLx lift coecient derivative w.r.t. parameter x Clx , Cmx , Cnx rolling, pitching and yawing moment coecient derivatives w.r.t. parameter x D, CD drag, drag coecient g gravitational acceleration L, CL lift, lift coecient p, q, r aircraft rotation rates in body or stability axes R turn radius Vcg flight speed W aircraft weight CG centre of gravity LE leading edge TE trailing edge VLM vortex lattice method Symbols angle of attack LA, RA left and right aft wingtip dihedral angles LF, RF left and right fore wingtip dihedral angles turn rate bank angle air density

A numerical model for thin airfoils in unsteady compressible arbitrary motion

Abstract: A numerical method based on the vortex methodology is presented in order to obtain unsteady solution of the aerodynamic coefficients of a thin airfoil in either compressible subsonic or supersonic flows. The numerical model is created through the profile discretization in uniform segments and the compressible flow vortex singularity is used. The results of the proposed model are presented as the lift and the pressure coefficient along the profile chord as a function of time. The indicial response for the unit step change of angle of attack and unit sharp-edged gust response of the profile are also obtained numerically. The results yielded by the present methodology are also compared with solutions available in the literature.

Unsteady Numerical Simulation of Wings with Flaperon Flying Over Nonplanar Ground Surface

Abstract: A boundary-element method is developed for the conceptual design of a high-speed transportation system for flying over a nonplanar ground surface. The method is validated by comparing present results with experimental data and other numerical data. Unsteady aerodynamic characteristics of a tandem wing with the flaperon flying over the nonplanar ground surface are investigated using the present method. When a tandem wing with the flaperon flies inside the channel, the lift coefficients of the wings are increased further because the air trapped by the fence of the channel increases the ground effect On the other hand, the fence of the channel compensates for the lift decrement of the rear wing due to the wake generated by the front wing. Therefore, there is little change between the flat ground and the channel in the longitudinal stability of the wings. Moreover, because the lift increment due to the channel takes place on both sides of the wing with the same rate of increase, there is little difference between the flat ground and the channel in the lateral stability of the wings.

Compressible Unsteady Vortex Lattice Method for Arbitrary Two Dimensional Motion of Thin Profiles

Abstract: †A numerical method based to the vortex methodology is presented. It is a numeric model for unsteady solution of the aerodynamics coefficients of a thin profile in subsonic and supersonic compressible flows. The numeric m odel is made through the airfoil discretization in uniform segments and the used singularity is vortex in compressible flow. Results for the proposed model are presented as the lift and pressure coefficients along the profile chord for some instants of tim e. It is obtained numerically the indicial response (unit step function) of the profile. The method is also compared with solutions available in the literature. Nomenclature � a = sound speed, m/s M = dimensionless Mach number t = time, s p c = dimensionless pressure coefficient p c � = dimensionless pressure coefficient jump

Analysis of the hydrodynamics of a pectoral fin propulsor

Abstract: In this paper two-degree-of-freedom and three-degree-of-freedom motion models were established for a rigid pectoral fin.Also,a motion-time ratio model was constructed.Furthermore,the theoretical expression of the unsteady vortex lattice method(UVLM) was developed from the Hess-Smith panel method by using equivalence of normal dipole and vorticity distribution.The unsteady vortex lattice method was used to analyze the hydrodynamic performance of pectoral fin motion.The pectoral fin was suddenly accelerated to constant motion from rest.In each time step we obtained the vortex strength of the pectoral fin using the vortex lattice method.The wake was formed step by step from the edge of the pectoral fin.In this process the dissipation of the wake vortex strength was taken into account.The induced velocity of the vortex system was calculated with the Biot-Savart law and the pressure on the fin was obtained from the Bernoulli equation.In the way all parameters of the hydrodynamic performance of the fin were obtained.

Broadband Noise Analysis of Horizontal Axis Wind Turbines Including Low Frequency Noise

Abstract: This paper demonstrates a computational method in predicting aerodynamic noise generated from wind turbines. Low frequency noise due to displacement of fluid and leading fluctuation, according to the blade passing motion, is modelled on monopole and dipole sources. They are predicted by Farassat 1A equation. Airfoil self noise and turbulence ingestion noise are modelled upon quadrupole sources and are predicted by semi-empirical formulas composed on the groundwork of Brooks et al. and Lowson. Aerodynamic flow in the vicinity of the blade should be obtained first, while noise source modelling need them as numerical inputs. Vortex Lattice Method(VLM) is used to compute aerodynamic conditions near blade. In the use of program X-foil [M.Drela] boundary layer characteristics are calculated to obtain airfoil self noise. Wind turbine blades are divided into spanwise unit panels, and each panel is considered as an independent source. Retarded time is considered, not only in low frequency noise but also In turbulence ingestion noise and airfoil self noise prediction. Numerical modelling is validated with measurement from NREL [AOC15/50 Turbine) and ETSU [Markham's VS45] wind turbine noise measurements.

Aircraft Subsystems Modelling Using Different MBS Formalisms

Abstract: The paper presents three case studies of dynamic analysis of aircraft during landing manoeuvre using two basic formalisms encountered in rigid and flexible multibody system (MBS) modelling. In the first case a formulation in natural coordinates has been used to analyze the dynamics of a medium size aircraft. Equations of motion have been formulated and solved using velocity transformation method. The aircraft has been modelled as consisting of rigid bodies connected by universal joints with springs. Aerodynamic forces have been taken into account by applying the Vortex Lattice Method (VLM) to the calculations performed. The effect of ground proximity on the results (ground effect) has been analyzed. In the second case, a dynamic analysis of a glider during the landing manoeuvre has been carried out from the point of view of stress recovery by means of various methods. Body positions and orientations have been written in absolute coordinates with floating frame approach for flexible bodies. Finite element method (FEM) and component mode synthesis has been used to model the flexibility of the bodies. A comparison of stress results obtained for different computation methods has been carried out. In the third analysis a MBS model of the Su-22 military airplane main landing gear has been presented. The absolute coordinates and the differential algebraic equations (DAE) formulations were used in all calculations. The whole landing gear model includes individual models of hydraulic actuators, shock absorber, flexible tire and contacts between some landing gear parts. Several types of simulations like landing gear extension and selected ground manoeuvres were performed. On that basis values of the forces which will allow to assess fatigue and durability of landing gear in future experiments were obtained. The received results were compared to the experimental measurements which were carried out on a real military airplane. The key issues of that comparison and general remarks were formulated. In the final part of the paper general conclusions regarding application of various computation MBS methods to dynamical analyses of aircrafts have been presented.

Further developments in unsteady compressible vortex lattice method in two dimensional motion.

Abstract: Unsteady phenomena like flutter, buffeting, rapid maneuvers in flight and gust entry are usually modeled and studied by a theoretical treatment involving potential flow methods. The resulting equation from this approach is the governing differential equation for general non-steady, non-viscous, potential flow known as convected wave equation. The disturbance, represented in this equation by the velocity potential, is propagated as wave which spreads at a rate equal to the local speed of sound. Linearization on the basis of small disturbances in a uniform stream of compressible fluid is made upon the equation by the procedure of retaining first order terms. Elementary solutions for this simplified equation recognized as primary extension of the concepts of source, sink, vortex and doublet, used together with boundary conditions associated with the governing equation, enables proper treatment for understanding and tackling non-steady aerodynamic problems. This thesis presents a numerical solution for the aerodynamics lift coefficient of a thin airfoil in arbitrary motion in a uniform, compressible, subsonic flow field. Distribution of vortex type elementary solutions of the convected wave equation is used together with a time function that schedules the vortex strength in time to represent in effect the arbitrary vortex moving along a chosen path. A field point is then influenced by the continuous disturbances generated by the vortex with a delay relative to the time of action of the same vortex. A fixed coordinate system in space relative to the body is chosen. So the body is fixed in a moving flow. The analytical vortex solution is presented together with the appropriate transformation variables needed to treat the problem.