Aerodynamic analysis and preliminary design tool

Thesis Overview

<p> </p><p>The aim of this study was to develop a potential flow calculation model which includes computation of flow around aircraft bodies (fuselage, engines) and a boundary layer method which calculates the viscous effects over the aircraft wings. The models developed will be merged with an already existing panel program developed by Saab, Linköping, Sweden.</p><p>Different methods have been studied but the basis of this work has been to develop a model using a panel method which can provide results from a simple geometry description, with short calculation time and hence be used in early design phases. In this thesis Matlab has been used as programming language, ensure that future development and maintenance is possible.</p><p>The body model uses a panel method where the flow domain is divided into an inner and an outer part where the outer problem uses a three dimensional panel description while the inner problem performs two dimensional calculations. The inner and outer problems are separated by an arbitrarily shaped reference box. The inner area is divided into a number of cross sections which are described by line segments. With the help of these the two dimensional cross flow is obtained. This result is connected to the outer part through boundary conditions and the entire three dimensional flow domain can be determined.</p><p>The resulting body program is limited to aircraft bodies with a slenderness ratio less than 1/5. Higher values violate the model assumption. The number of cross sections needed to describe a body of one unit length is between 80-150 and the number of line segments needed for one cross sections is 20 for the inner boundary and 40 line segments for the outer. This configuration gives results with acceptable accuracy within a computation time less than 15 seconds/body.</p><p>The viscous effects around the aircraft wings are modelled with a two dimensional boundary layer model where the boundary layer displacement thickness over the wing profile is calculated with two different methods depending on if the flow in the boundary layer is laminar or turbulent. The computed displacement thickness is then added to the wing profile geometry and new pressure distributions are computed on the modified geometry.</p><p>The computed pressure distributions including the viscous effects show better agreement with results from experimental wind tunnel tests than the inviscid without boundary layer contribution. Separation is not modelled and neither are the large effects this has on the pressure distribution. The model gives usable results up to 15-20 degrees angle of attack; at higher angles the separated regions are so large that the model is not valid anyway.</p> <br><p></p>