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A technique to obtain the sensitivity of the static aeroelastic response of a three dimensional wing model is designed and implemented. The formulation is quite general and accepts any aerodynamic and structural analysis capability. A program to combine the discipline level, or local, sensitivities into global sensitivity derivatives is developed. A variety of representations of the wing pressure field are developed and tested to determine the most accurate and efficient scheme for representing the field outside of the aerodynamic code. Chebyshev polynomials are used to globally fit the pressure field. This approach had some difficulties in representing local variations in the field, so a variety of local interpolation polynomial pressure representations are also implemented. These panel based representations use a constant pressure value~ a bilinearly interpolated value, or a biquadratic ally interpolated value. The interpolation polynomial approaches do an excellent job of reducing the numerical problems of the global approach for comparable computational effort. Regardless of the pressure representation used, sensitivity and response results with excellent accuracy have been produced for large integrated quantities such as wing tip deflection and trim angle of attack. The sensitivities of such things as individual generalized displacements have been found with fair accuracy. In general, accuracy is found to be proportional to the relative size of the derivatives to the quantity itself. / Ph. D.

The identification of linear structural dynamic systems has been dealt with extensively in past studies. Identification methods for nonlinear structures have also been introduced in previous articles, including procedures based on the method of multiple scales, iterative and noniterative direct methods, and state space mappings. Here, a procedure is introduced for the identification of nonlinear structural dynamic systems which is readily applicable to simple as well as more complex multiple degree of freedom systems. The procedure is based on multiple step integration methods for the solution of differential equations. The multiple step integration procedure and the iterative direct method are applied to a number of nonlinear single degree of freedom examples, and are applied to a simple two degrees of freedom example as well. RMS based noise is added to a simulated measured response in order to monitor the effects of measurement errors on the procedures. The input data is filtered before final processing in the identification algorithms. The multistep algorithm is compared to the iterative direct method on the basis of criteria such as accuracy, ease of use, and numerical efficiency. / Master of Science

Thesis (MTech (Mechanical Engineering))--Cape Peninsula University of Technology, 2016. / This work presents the computational uid dynamics (CFD) analysis of a light road vehicle. Simulations are conducted using the lattice Boltzmann method (LBM) with the wall adapting local eddy (WALE) turbulence model. Simulations include and compare the use of a rolling road, rotating wheels, adaptive re nement as well as showing comparison with a Reynolds-averaged Navier-Stokes (RANS) solver and the Spalart- Allmaras (SA) turbulence model. The lift coe cient of the vehicle for the most part was seen to show a much greater di erence and inconsistencies when compared to drag from the comparisons of solvers, turbulence models, re nement and the e ect of rolling road. Determining the drag of a road vehicle can be easily achieved and veri ed using multiple solvers and methods, however, the lift coe cient and its validation require a greater understanding of the vehicle ow eld as well as the solvers, turbulence models and re nement levels capable of correctly simulating the turbulent regions around a vehicle. Using the presented method, it was found that the optimisation of vehicle aerodynamics can easily be done alongside the design evolution from initial low-drag shapes to the nal detail design, ensuring aerodynamic characteristics are controlled with aesthetic change.

In this thesis an approximate method is developed which predicts the aerodynamic force on a cone-cylinder body in axially symmetric transonic now. The method places more emphasis on the physics of the now than on the mathematical rigors of solving the typical reduced non-linear transonic equation of motion. Under the assumption that the now is that of a steady, irrotational, inviscid, compressible gas, the body pressures are determined and the associated force defined. Recognizing that the transonic pressures are influenced by the character of the subsonic compressible pressures, which are obtained in this analysis through Gothert’s Rule, it is then mandatory that the incompressible case be defined with the best possible accuracy. Comparisons with experiments indicate that the classical method (axially distributed sources and sinks) does not provide this required accuracy. Thus the surface distributed vortex ring theory is used in the present analysis to obtain the incompressible body pressures. Gothert’s Rule, which represents a linear solution for the subsonic case, is known to be applicable up to a limit value of tree stream Mach number. An investigation is carried out herein to determine both the correct form of the rule and its limits of applicability. As a result of this investigation, it is concluded that the upper limit is the lower free stream critical Mach number. Also, at this Mach number, a solution is immediately available tor the lower limit of the transonic range of Mach number. In solving the transonic problem the law or stationarity of local Mach number is of fundamental importance. For an assumed isentropic flow over the body, and for sonic conditions being present at some point on the surface, the body pressures can be described in the ratio p_L/p*. Here p_L is the local surface pressure and p* is the sonic (body) pressure. Through the stationarity law, this ratio is recognized as an invariant for transonic speeds so long as the flow field remains essentially irrotational. Thus any change in local pressure is only a function of the free stream Mach number for any given body position. By this approach, the pressure distribution is defined for a range of Mach number from below to above the sonic stream value. The method is then capable of prediction for almost all of the transonic range of Mach number. It is only when the head shock baa significant curvature, causing the now adjacent to the body to be rotational, that the method fails. Though the procedure developed here is not capable of spanning the entire transonic range, it does provide a wider range of applicability than other known theories. Finally, for this problem, a correlation of transonic pressure drag data is formulated. This correlation is founded on physical interpretation and is not limited to the usual transonic similarity restrictions. In fact, to the author's knowledge, this is the first known such correlation tor axially symmetric flow covering the range of body sizes and Mach numbers considered in this investigation. In so far as is practicable the results obtained in this thesis have been compared to available experimental results. In particular, the drag data from this analysis compare closely with experimental transonic values. Experiment bears out the conclusion that the upper limit for linear theory is the lower critical tree stream Mach number. And, the pressures determined by the vortex ring theory agrees well with the low-speed experimental results obtained by the author. / Ph. D.

Two instruments mounted in a piggyback arrangement were developed for time-resolved measurements of dynamic flow quality in a transonic cryogenic wind tunnel. The first one is a dual hot-wire aspirating probe for measurement of stagnation pressure and temperature. The second is a miniature high-frequency response angle probe consisting of surface mounted pressure sensors. The aspirating probe was tested in the 0.3-m Transonic Cryogenic Tunnel (TCT) at NASA-Langley Research Center. Stagnation pressure and temperature measurements were taken in the free-stream of the settling chamber and test section. Data were also obtained in the unsteady wake shed from an airfoil oscillating at 5 Hz. The investigation revealed the presence of large stagnation pressure and temperature fluctuations in the settling chamber occurring at the blade passing frequency of the tunnel driving fan. The fluctuations in the test section are of a much more random nature and have amplitudes much lower than those in the test section. The overall results are consistent with previous tunnel disturbance measurements in the 0.3-m TCT. In the unsteady wake shed from the oscillating airfoil, stagnation temperature fluctuations as high as 42 K rms were observed. The high-frequency angle probe is a four sensor, pyramid type probe capable of simultaneously measuring time resolved stagnation and static pressures and two orthogonal flow angles. Using measurements from both probes, all flow parameters of interest can be deduced. Aerodynamic behavior of a full size model of the probe was established in an open air jet of known conditions. / M.S.

The recent resurgence in hypersonics is centered around the development of SCRAMJET engine technology to power future hypersonic vehicles. Successful flight trials by Australian and American scientists have created interest in the scramjet engine research across the globe. To develop scramjet engine, it is important to study heat transfer effects on the engine performance and aerodynamic forces acting on the body. Hence, the main aim of present investigation is the design of scramjet engine configuration and measurement of aerodynamic forces acting on the model and heat transfer rates along the length of the combustor. The model is a two-dimensional single ramp model and is designed based on shock-on-lip (SOL) condition. Experiments are performed in IISc hypersonic shock tunnel HST2 at two different Mach numbers of 8 and 7 for different angles of attack. Aerodynamic forces measurements using three-component accelerometer force balance and heat transfer rates measurements using platinum thin film sensors deposited on Macor substrate are some of the shock tunnel flow diagnostics that have been used in this study.