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In this study the primary objective was to determine the effect of aspect ratio in particular and in general the effect of three dimensionality on the flow around wings at low Reynolds numbers. It was seen that the effects observed at high Reynolds number are also present in this Re range. There is the usual increase in lift slope and this increase can even be predicted with reasonable accuracy using Prandtl's lifting line theory. In addition to the change in lift slope the zero lift angle of attack was also influenced by the aspect ratio. Through flow visualization it was ascertained that the wingtips have a rather restricted effect on the laminar separation bubble. The disappearance of the bubble extends only for a small distance inboard from the tips. The size of the hysteresis loop and the Reynolds number at which hysteresis starts was found to be influenced by the aspect ratio. The momentum deficit method was used to validate the data obtained by the strain gauge method and there was adequate agreement between the values found through the two methods. From the measurements of pressure done around the airfoil contour one could determine both the location of the laminar separation bubble and the regions were flow is separated. The pressure taps themselves were found to influence measurements somewhat in certain regions of angle of attack and Reynolds number. In the future it would be beneficial to continue strain gauge measurements on this airfoil with flaps and control surfaces to determine their effect on the formation of the laminar separation bubble. Also measurements on other shapes would give more insight into the phenomena occurring here. The effects of turbulence and noise will have to be investigated in detail to determine what performance to expect from an actual aircraft. Finally detailed measurements on boundary layer stability and its effect on the occurrence of reattachment should be studied in detail to gain insight into the reasons for the presence of a hysteresis loop in stall at these Reynolds numbers. / Ph. D.

M. Tech. (Engineering: Electrical; Department of Electronic Engineering, Faculty of Engineering and Technology )-- Vaal University of Technology / In the ever growing sport of model aircraft, pilots are challenged with many obstacles. In the division of gliders, one of the biggest problems is the loss of model aircraft. Pilots launch their aircraft off mountain tops and if the aircraft crashes below, the pilot must make use of his best estimates in order to locate the aircraft. This either takes several hours, or the aircraft is never recovered. Pilots are also at a loss with regard to real time data, such as, but not limited to, battery levels, fuel levels, altitude and speed. Model aircraft competitions are also limited to the best estimate of officials. In this work an attempt has been made to design and develop a remote tracking device for model aircraft. This device will retrieve Global Positioning System (GPS) co-ordinates from the aircraft and relay them to the pilot on the ground. In the event of a crash, the pilot will retrieve the last GPS co-ordinates and then proceed to the location to collect the aircraft. An attempt will also be made to design add-on telemetry components that will allow for measurement and transmission of battery levels, fuel levels, altitude, G-Force, orientation, acceleration, wind and ground speed. Some of the data retrieved from the Flight Data Recorder (FDR) in trial 1 are: maximum altitude above sea level of 2139.20 m, maximum speed over ground which was 57.34 m/s and the average battery voltage for transceiver and servos was 15.2 v.

A dissertation submitted to the Faculty of Engineering and the Built Environment, School of Mechanical, Industrial and Aeronautical Engineering, University of the Witwatersrand, in fulfilment of the requirements for the degree of Master of Science in Engineering. Johannesburg, May 2016 / The present work describes the practical implementation of systems identification techniques to the development of a linear aerodynamic model for a small low-cost UAV equipped with a basic navigational and inertial measurement systems. The assessment of the applicability of the techniques were based on determining whether adequate aerodynamic models could be developed to aid in the reduction of wind tunnel testing when characterising new UAVs. The identification process consisted of postulating a model structure, flight test manoeuvre design, data reconstruction, aerodynamic parameter estimation, and model validation. The estimators that were used for the post-flight identification were the output error maximum likelihood method and an iterated extended Kalman filter with a global smoother. SIDPAC and FVSysID systems identification toolboxes were utilised and modified where appropriate. The instrumentation system on board the UAV consisted of three-axis accelerometers and gyroscopes, a three-axis vector magnetometer and GPS tracking while data was logged at 25 Hz. The angle of attack and angle of sideslip were not measured directly and were estimated using tailored data reconstruction methods. Adequate time domain lateral model correlation with flight data was achieved for the cruise flight condition. Adequacy was assessed against Theil’s inequality coefficients and Theil’s covariance. It was found that the simplified estimation algorithms based on the linearized equations of motion yielded the most promising model matches. Due to the high correlation between the pitch damping derivatives, the longitudinal analysis did not yield valid model parameter estimates. Even though the accuracy of the resulting models was below initial expectations, the detailed data compatibility analysis provided valuable insight into estimator limitations, instrumentation requirements and test procedures for systems identification on low-cost UAVs. / MT2016