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Sounding rockets find applications in multiple fields of scientific research including meteorology, astronomy and microgravity. Indigenous sounding rocket technologies are absent on the African continent despite a potential market in the local aerospace industries. The UKZN Phoenix Sounding Rocket Programme was initiated to fill this void by developing inexpensive medium altitude sounding rocket modeling, design and manufacturing capacities. This dissertation describes the development of the Hybrid Rocket Performance Simulator (HYROPS) software tool and its application towards the structural design of the reusable, 10 km apogee capable Phoenix-1A hybrid sounding rocket, as part of the UKZN Phoenix programme. HYROPS is an integrated 6–Degree of Freedom (6-DOF) flight performance predictor for atmospheric and near-Earth spaceflight, geared towards single-staged and multi-staged hybrid sounding rockets. HYROPS is based on a generic kinematics and Newtonian dynamics core. Integrated with these are numerical methods for solving differential equations, Monte Carlo uncertainty modeling, genetic-algorithm driven design optimization, analytical vehicle structural modeling, a spherical, rotating geodetic model and a standard atmospheric model, forming a software framework for sounding rocket optimization and flight performance prediction. This framework was implemented within a graphical user interface, aiming for rapid input of model parameters, intuitive results visualization and efficient data handling. The HYROPS software was validated using flight data from various existing sounding rocket configurations and found satisfactory over a range of input conditions. An iterative process was employed in the aerostructural design of the 1 kg payload capable Phoenix-1A vehicle and CFD and FEA numerical techniques were used to verify its aerodynamic and thermo-structural performance. The design and integration of the Phoenix-1A‟s hybrid power-plant and onboard electromechanical systems for recovery parachute deployment and motor oxidizer flow control are also discussed. It was noted that use of HYROPS in the design loop led to improved materials selection and vehicle structural design processes. It was also found that a combination of suitable mathematical techniques, design know-how, human-interaction and numerical computational power are effective in overcoming the many coupled technical challenges present in the engineering of hybrid sounding rockets. / Thesis (M.Sc.)-University of KwaZulu-Natal, Durban, 2012.

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The following dissertation describes an investigation of the structural response behaviour of a composite solid rocket motor nozzle subjected to thermal and pressure loading during the motor ignition period, derived on the basis of a multidisciplinary numerical simulation approach. To provide quantitative and qualitative context to the results obtained, comparisons were made to the predicted aerothermostructural response of the nozzle over the entire motor burn period. The study considered two nozzle designs – an exploratory nozzle design used to establish the basic simulation methodology, and a prototype nozzle design that was employed as the primary subject for numerical experimentation work. Both designs were developed according to fundamental solid rocket motor nozzle design principles as non-vectoring nozzles for deployment in medium sized solid rocket booster motors. The designs feature extensive use of spatially reinforced carbon-carbon composites for thermostructural components, complemented by carbon-phenolic composites for thermal insulation and steel for the motor attachment substructures. All numerical simulations were conducted using the ADINA multiphysics finite element analysis code with respect to axisymmetric computational domains. Thermal and structural models were developed to simulate the structural response of the exploratory nozzle design in reference to the instantaneous application of pressure and thermal loading conditions derived from literature. Ignition and burn period response results were obtained for both quasi-static and dynamic analysis regimes. For the case of the prototype nozzle design, a flow model was specifically developed to simulate the flow of the exhaust gas stream within the nozzle, for the provision of transient and steady loading data to the associated thermal and structural models. This arrangement allowed for a more realistic representation of the interaction between the fluid, thermal and structural fields concerned. Results were once again obtained for short and long term scenarios with respect to quasi-static and dynamic interpretations. In addition, the aeroelastic interaction occurring between the nozzle and flow field during motor ignition was examined in detail. The results obtained in the present study provided significant indications with respect to a variety of response characteristics associated with the motor ignition period, including the magnitude and distribution of the displacement and stress responses, the importance of inertial effects in response computations, the stress response contributions made by thermal and pressure loading, the effect of loading condition quality, and the bearing of the rate of ignition on the calculated stress response. Through comparisons between the response behaviour predicted during the motor ignition and burn periods, the significance of considering the ignition period as a qualification and optimisation criterion in the design of characteristically similar solid rocket motor nozzles was established. / Thesis (M.Sc.Eng.)-University of KwaZulu-Natal, Durban, 2009.

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Thesis (M.App.Sc.) - Carleton University, 2006. / Includes bibliographical references (p. 163-172). Also available in electronic format on the Internet.

Thesis (M.S.)--Virginia Polytechnic Institute and State University, 1992. / Vita. Abstract. Includes bibliographical references (leaves 111-113). Also available via the Internet.

Thesis (M.S.)--Florida State University, 2004. / Advisor: Dr. Steven Van Sciver, Florida State University, College of Engineering, Department of Mechanical Engineering. Title and description from dissertation home page (viewed Sept. 23, 2004). Includes bibliographical references.

Thesis (M.S.)--Rochester Institute of Technology, 2009. / Typescript. Includes bibliographical references (leaves 70-73).

Thesis (MTech (Mechanical Engineering))--Cape Peninsula University of Technology, 2017. / A framework for optimising a tactical solid rocket motor nozzle is established and investigated within the ANSYS Workbench environment. Simulated results are validated against thrust measurements from the static bench firing of a full-scale rocket. Grid independence is checked and achieved using inflation based meshing. A rocket nozzle contour is parametrized using multiple control points along a spline contour. The design of experiments table is populated by a central composite design method and the resulting response surfaces are used to find a thrust optimised rocket nozzle geometry. CFD results are based on Favre-mass averaged Navier-Stokes equations with turbulence closure implemented with the Menter SST model. Two optimisation algorithms (Shifted Hammersley Sampling and Nonlinear Programming by Quadratic Lagrangian) are used to establish viable candidates for maximum thrust. Comparisons are made with a circular arc, Rao parabolic approximation and conical nozzle geometries including the CFD simulation there-off. The effect of nozzle length on thrust is simulated and optimised within the framework. Results generally show increased thrust as well as demonstrating the framework's potential for further investigations into nozzle geometry optimisation and off-design point characterisation.