Showing papers on "Vortex lattice method published in 2004"

Object-oriented unsteady vortex lattice method for flapping flight

Abstract: The unsteady vortex lattice method is used to model the oscillating plunging, pitching, twisting, and flapping motions of a finite-aspect-ratio wing. Its potential applications include design and analysis of small unmanned air vehicles and in the study of the high-frequency flapping flight of birds and other small flyers. The results are verified by theory and, in the plunging and pitching cases, by experimental data. The model includes free-wake relaxation, vortex stretching, and vortex dissipation effects and is implemented using object-oriented computing techniques

Comparison of Predicted and Measured Formation Flight Interference Effects

Abstract: Results from a wind-tunnel test of two delta-wing aircraft in close proximity are presented and compared with predictions from a vortex lattice method. Large changes in lift, pitching moment, and rolling moment are found on the trail aircraft as it moves laterally relative to the lead aircraft. The magnitude of these changes is reduced as the trail aircraft moves vertically with respect to the lead aircraft. Lift-to-drag ratio of the trail aircraft is increased when the wing tips are slightly overlapped. Wake-induced lift is overpredicted slightly when the aircraft overlap in the spanwise direction. Wake-induced pitching and rolling moments are well predicted. A maximum induced drag reduction of 25% is measured on the trail aircraft, compared with a 40% predicted reduction. Three positional stability derivatives, change in lift and pitching moment with vertical position and change in rolling moment with lateral position, are studied. Predicted boundaries between stable and unstable regions were generally in good agreement with experimentally derived boundaries.

Computational-Fluid-Dynamics-Based Enhanced Indicial Aerodynamic Models

Abstract: Direct Euler calculations were performed to verify the accuracy of linearized airload models and to generate a numerical database for the determination of generalized sharp-edged vertical gust response functions. The field velocity approach was used in the computational fluid dynamics (CFD) code to incorporate the impulsive boundary conditions that were characteristic of the problems studied. Linearized airload models were constructed in three dimensions by coupling the indicial models obtained in two dimensions with a Weissinger-L vortex panel method. The vortex panel method was suitably modified to include the shed near wake effects and impulsive boundary conditions. Studies conducted on step change in angle of attack and step change in pitch rate for a finite wing showed that the linearized models are sufficiently accurate in predicting the sectionwise air loads. Generalized gust functions were determined with a constrained minimization of least-square error with a CFD-generated database

Aerodynamic Interference of a Large and a Small Aircraft.

Abstract: The influence of wake from a leading aircraft, Boeing 757, on a following aircraft, Cessna 750 Citation X, is investigated using the unsteady vortex lattice method. The effect of the wakes from both wings and each other is considered and the new position of the wakes from both aircraft calculated. As the midspan of the Cessna wing approaches the Boeing 757 trailing vortex core, the roll moment increases to its maximum value. Strong changes appear in aerodynamic loads of the Cessna as it moves near the Boeing 757’s wake. Rapid change occurs as the Cessna moves in transverse or vertical directions near the Boeing’s wake. The unsteady aerodynamic loads on wings moving along different paths are calculated. The Boeing 757’s wake strongly affects the aerodynamic loads and the wake of the Cessna. Also included in the present results are numerical simulations of wakes and velocity vectors in the Trefftz plane detailing the influence of the wake from the leading aircraft, the Boeing 757.

Numerical model for the simulation of fixed wings aeroelastic response

Abstract: A numerical model for the simulation of fixed wings aeroelastic response is presented. The methodology used in the work is to treat the aerodynamics and the structural dynamics separately and then couple them in the equations of motion. The dynamic characterization of the wing structure is done by the finite element method and the equations of motion are written in modal coordinates. The unsteady aerodynamic loads are predicted using the vortex lattice method. The exchange of information between the aerodynamic and structural meshes is done by the surface splines interpolation scheme, and the equations of motion are solved iteratively in the time domain, employing a predictor-corrector method. Numerical simulations are performed for a prototype aircraft wing. The aeroelastic response is represented by time histories of the modal coordinates for different airspeeds, and the flutter occurrence is verified when the time histories diverge (i.e. the amplitudes keep growing). Fast Fourier Transforms of these time histories show the coupling of frequencies typical of the flutter phenomenon.

Limit-cycle oscillations of aircraft caused by flutter-induced drag

Abstract: It has been shown in earlier work that the energy for flutter comes from propulsion, that is, flutter leads to an increase in drag. The present paper seeks to explain the effect of flutter-induced drag on aircraft response. It is shown that the aircraft can undergo limit-cycle oscillations even if the oscillations are in the linear structural and aerodynamic lift range. These limit-cycle oscillations occur as a result of the inability of the engine to support exponentially increasing oscillations. The limit-cycle oscillations are composed of variations in flight speed and amplitude of vibration. This modality of limit-cycle oscillations has not been presented in earlier literature and leads to better understanding of the limit-cycle oscillation problem.

Wind-tunnel measurements of hazard posed by lift-generated wakes

Abstract: The influence of wake from a leading aircraft, Boeing 757, on a following aircraft, Cessna 750 Citation X, is investigated using the unsteady vortex lattice method. The effect of the wakes from both wings and each other is considered and the new position of the wakes from both aircraft calculated. As the midspan of the Cessna wing approaches the Boeing 757 trailing vortex core, the roll moment increases to its maximum value. Strong changes appear in aerodynamic loads of the Cessna as it moves near the Boeing 757's wake. Rapid change occurs as the Cessna moves in transverse or vertical directions near the Boeing's wake. The unsteady aerodynamic loads on wings moving along different paths are calculated. The Boeing 757's wake strongly affects the aerodynamic loads and the wake of the Cessna. Also included in the present results are numerical simulations of wakes and velocity vectors in the Trefftz plane detailing the influence of the wake from the leading aircraft, the Boeing 757

A free-wake aerodynamic model for helicopter rotors

Abstract: The accurate computation of the induced velocity on helicopter rotor plane is a prerequisite for the precise evaluation of the angle of attack distribution along the rotor blades, which leads to improved air loads prediction for aerodynamics and aeroacustics design and performance analysis. The Vortex Lattice Method (VLM) provides a transparent investigation concerning the role of various physical parameters which influence the aerodynamic problem of rotor downwash calculation. This paper presents a method for the calculation of the nonuniform induced downwash of a helicopter rotor using the vortex ring model for a thin lifting surface coupled with a free-wake model. During the last few years, considerable research effort has been made with respect to the various aspects of rotors. In terms of modeling, there is no doubt that significant progress has been achieved during the last ten years. This progress was mainly based on the existing experience of other engineering fields. As regards the design of helicopter rotors a number of methods for structural, aerodynamic an aeroacoustic analysis used in aeronautical engineering have been applied. However, in many cases only simply models were used. This is true especially as far as Complete Design Tools are concerned. Of course, in order to devise a flexible and practical design tool, simple models constitute to some extent an obligatory choice. However, the choice of simple models means that only approximate results can be obtained. Therefore we need to determine their limitations. On the other hand during the last few years, within the numerous activities, a number of more elaborate models has been developed. In most cases the corresponding work was related to the analysis of only one aspect of the whole design problem of rotor (e.g. advanced aerodynamic models were applied to rigid rotors and for steady inflow conditions). Though we cannot expect that the advanced models will be applied to practical design problems at least in the near future, we can use them in order to validate and improve the simple models. The collection of the a fore mentioned problems defines the setting of a complete design tool summarized in the following diagram.