An Algorithm for the Calculation of the Potential Flow about an Arbitrary Two Dimensional Aerofoil. G.U. Aero Report 8502

A vortex panel method has been developed to calculate the potential flow about an arbitrary two dimensional aerofoil or axisymmetric shape at fixed incidence in a steady, uniform, irrotational, incompressible flow. The procedure replaces the contour by a suitably inscribed polygon, on which surface vorticity varies linearly and continuously along the panel and is piecewise continuous at the panel corners. The Neumann boundary condition is satisfied at control points situated at the midpoint of each panel and the classical Kutta condition is specified at the trailing edge by setting the net vorticity there equal to zero. This particular algorithm offers much flexibility in the treatment of a greater range of aerofoil geometries and at higher incidence than other surface singularity methods. Programme flow charts and FORTRAN code listings are given in the User Guide (1)

Modelling of Trailing Edge Separation on Arbitary Two-Dimensional Aerofoils in Incompressible Flow Using an Inviscid Flow Algorithm. G.U. Aero Report 8202

An algorithm for estimating the lift, moment and pressure distribution on arbitary two dimensional aerofoils in incompressible flow is presented. The procedure uses an inviscid analysis of the physics of the real flow, which invokes the application of a linear vortex panel model. The separated wake geometry is determined iteratively, starting from an initial assumption. A boundary layer analysis is not performed, hence the upper surface separation point is a necessary input to the algorithm. Lower surface separation is assumed to occur at the trailing edge. A selection of results and comparison with experimental data is presented. The scatter in the calculated and experimental data values is attributed mainly to the lack of boundary layer displacement and compressibility effects. A fortran code listing of the algorithm is given in the Appendix

Modelling of Trailing Edge Separation on Arbitary Two-Dimensional Aerofoils in Incompressible Flow Using an Inviscid Flow Algorithm. G.U. Aero Report 8202

An algorithm for estimating the lift, moment and pressure distribution on arbitary two dimensional aerofoils in incompressible flow is presented. The procedure uses an inviscid analysis of the physics of the real flow, which invokes the application of a linear vortex panel model. The separated wake geometry is determined iteratively, starting from an initial assumption. A boundary layer analysis is not performed, hence the upper surface separation point is a necessary input to the algorithm. Lower surface separation is assumed to occur at the trailing edge. A selection of results and comparison with experimental data is presented. The scatter in the calculated and experimental data values is attributed mainly to the lack of boundary layer displacement and compressibility effects. A fortran code listing of the algorithm is given in the Appendix

Aerodynamics of Pitching Wings: Theory and Experiments

Peer Reviewedhttps://deepblue.lib.umich.edu/bitstream/2027.42/140444/1/6.2014-2881.pd

Optimizing performance variables for small unmanned aerial vehicle co-axial rotor systems

The aim of this project was to design and build a test-rig that is capable of analyzing small unmanned aerial vehicles (SUAV) co-axial rotor systems. The intention of the test-rig development was to highlight important aeromechanical components and variables that dictate the co-axial units flight performance, with the intention of optimizing the propulsion systems for use on HALO® a co-axial SUAV designed by the Autonomous Systems Lab at Middlesex University. The major contributions of this paper are: an optimum COTS co-axial configuration with regards to motor and propeller variations, a thorough review and validation of co-axial rotor systems inter-rotor spacing which in turn identified an optimum H/D ratio region of between (0.41–0.65)

Collected Data for Sinusoidal Tests on a NACA 23012 Aerofoil. Volume 2: Pressure Data for a Reynolds Number of 1.5 Million. G.U. Aero Report 8600.

Summary: Herein is presented the collected data for tests in which a NACA 23012 aerofoil was subjected to oscillatory displacements in pitch about the quarter chord position. The data clearly illustrates the effect of reduced frequency on the aerofoil characteristics and the chordal pressure distribution at the midspan position of the aerofoil model. The data are presented in two volumes. This volume contains the test data for a Reynolds number of 1.5 x 10 6. For a full introduction and description of test procedure, see volume 1