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Current thinking on the best propulsion system for a next-generation supersonic cruising (Mach 2 to Mach 4) aircraft is a mixed-flow turbofan engine with afterburner. This study investigates the performance increase of a turbofan engine through the use of isothermal combustion inside the high-pressure turbine (High-Pressure Turburner, HPTB) as an alternative form of thrust augmentation. A cycle analysis computer program is developed for accurate prediction of the engine performance and a supersonic transport cruising at Mach 2 at 60,000 ft is used to demonstrate the merit of using a turburner. When assuming no increase in turbine cooling flow is needed, the engine with HPTB could provide either 7.7% increase in cruise range or a 41% reduction in engine mass flow when compared to a traditional turbofan engine providing the sane thrust. If the required cooling flow in the turbine is almost doubled, the new engine with HPTB could still provide a 4.6% increase in range or 33% reduction in engine mass flow. In fact, the results also show that the degradation of engine performance because of increased cooling flow in a turburner is less than half of the degradation of engine performance because of increased cooling flow in a regular turbine. Therefore, a turbofan engine with HPTB will still easily out-perform a traditional turbofan when even more cooling than currently assumed is introduced. Closer examination of the simulation results in off-design regimes also shows that the new engine not only satisfies the thrust and efficiency requirement at the design cruise point, but also provides enough thrust and comparable or better efficiency in all other flight regimes such as transonic acceleration and take-off. Another finding is that the off-design bypass ratio of the new engine increases slower than a regular turbofan as the aircraft flies higher and faster. This behavior enables the new engine to maintain higher thrust over a larger flight envelope, crucial in developing faster air-breathing aircraft for the future. As a result, an engine with HPTB provides significant benefit both at the design point and in the off-design regimes, allowing smaller and more efficient engines for supersonic aircraft to be realized. / Ph. D.

Since the late 19th century, the average temperature on Earth has risen by approximately 1.1 °C because of the increased carbon dioxide (CO2) and other man-made emissions to the atmosphere. The transportation sector is responsible for approximately 33% of the global CO2 emissions and 14% of the overall greenhouse gas emissions. Therefore, increasingly stringent regulations in the European Union require CO2 emissions to be lower than 95 gCO₂/km by 2020. In this regard, improvements in internal combustion engines (ICEs)must be achieved in terms of fuel consumption and CO2 emissions. Given that only up to 35% of fuel energy is converted into mechanical power, the wasted energy can be reused through waste heat recovery (WHR) technologies. Consequently, organic Rankine cycle (ORC) has received significant attention as a WHR technology because of its ability to recover wasted heat in low- to medium-heat sources. The Expansion machine is the key component in ORC systems, and its performance has a direct and significant impact on overall cycle efficiency. However, the thermal efficiencies of ORC systems are typically low due to low working temperatures. Moreover, supersonic conditions at the high pressure ratios are usually encountered in the expander due to the thermal properties of the working fluids selected which are different to water. Therefore, this thesis aims to design an efficient radial-inflow turbine to avoid further efficiency reductions in the overall system. To fulfil this aim, a novel design and optimisation methodology was developed. A design of experiments technique was incorporated in the methodology toexplorethe effects of input parameters on turbine performance and overall size. Importantly, performance prediction modelling by means of 1D mean-line modelling was employed in the proposed methodology to examine the performance of ORC turbines at constant geometries. The proposed methodology was validated by three methods: computational fluid dynamics analysis, experimental work available in the literature, and experimental work in the current project. Owing to the lack of actual experimental works in ORC-ICE applications, a test rig was built around a heavy-duty diesel engine at Brunel University London and tested at partial load conditions due to the requirement for a realistic off-high representation of the performance of the system rather than its best (design) point, while taking into account the limitation of the engine dynamometer employed. Results of the design methodology developed for this projectpresented an efficient single-stage high-pressure ratio radial-inflow turbine with a total to static efficiency of 74.4% and an output power of 13.6 kW.Experimental results showed that the ORC system had a thermal efficiency of 4.3%, and the brake-specific fuel consumption of the engine was reduced by 3%. The novel meanlineoff designcode (MOC) was validated with the experimental works from three turbines. In comparison with the experimental results conducted at Brunel University London, the predicted and measured results were in good agreement with a maximum deviation of 2.8%.

This diploma thesis deals with basic design and calculation of an aircraft power unit (APU). For desired shaft power the thermal cycle is calculated. Basic dimensions of compressor, turbine and exhaust system are calculated as well as thermodynamic parameters of air and hot gases along the motor. Corresponding compressor map, turbine map and characteristics of the exhaust system are designed. Based on these characteristics, an effective shaft power is being calculated in various off-design conditions. Dependency of the shaft power vs. altitude is also presented. All the calculations are made in MATLAB. Finally, simplified motor cross-section is created.