| Abstract [eng] |
In this thesis, a rocket thruster generating a thrust of 20 N, powered by high concentration (98 %) hydrogen peroxide, is designed. To evaluate the internal ballistics of the rocket engine, a mathematical model is applied, which is based on the works of Koopmans et al. and Fogler. The Python programming language is used to apply the mathematical model. The mathematical model of internal ballistics evaluates the influence of two catalyst bed layers with different catalyst granule sizes on the decomposition of hydrogen peroxide. Using the mathematical model, the reaction rate of hydrogen peroxide decomposition, the released heat and the pressure drop over the entire length of the catalyst bed are calculated. Based on the results of the internal ballistics mathematical model, a rocket engine is designed according to operational requirements. During the design, each component of the rocket engine is defined separately. Considering the compatibility with hydrogen peroxide, the structural materials of the rocket engine are selected. When comparing solenoid valves offered by different manufacturers, a Parker pulse valve is selected. A single capillary tube is used to feed fuel from the valve to the hydrogen peroxide decomposition chamber, which is selected based on the flow rate of the propellant and the pressure drop. After selecting the length of the capillary tube, the elongation of the tube due to temperature and the pressure drop due to the shape of the capillary tube are calculated. When designing the hydrogen peroxide decomposition chamber, a suitable catalyst granule material was selected – aluminium oxide coated with platinum. After evaluating the materials of the catalyst layer, pressure losses through the catalyst bed retainers were calculated. The possibilities of sealing the rocket engine were also evaluated. Swagelok mechanical and special threaded connections were selected for this purpose. Metal sealing rings were also selected. After describing all the components of the designed rocket engine, a rocket nozzle was designed, adapted to atmospheric pressure at sea level, with a throat diameter of 3,5 mm and a nozzle end diameter of 7,6 mm. The total pressure drop over the entire length of the rocket engine was calculated to be 28,4 bar, with such a total pressure drop, the rocket engine inlet pressure was calculated to be 43,54 bar. Based on the results of the mathematical model and engine design, strength calculations were performed using the finite element analysis method in the SolidWorks environment. After simplifying the spatial models, the strength properties of the engine housing, the rear catalyst bed retainer and the capillary tube were evaluated separately. At the end of the work, possible tests are proposed that would help confirm the design results. |
