Characterisation and Simulation of Reusable Single-Stage-To-Orbit Vehicles Ascent Phase during Conceptual Design

The increasingly larger request for access to space, driven by the constant growth of the space economy, leads to the request for new tools capable of performing fast and reliable mission designs. In this context, this master’s thesis is part of a research project of Politecnico di Torino whose purpose is to develop a software tool to support the conceptual design of access-to-space missions with reusable vehicles. In particular, this tool will be able to guide researchers and engineers throughout the very first design phases, allowing the user to carry out quick but reliable analysis of alternative mission concepts. This thesis focuses on the conceptual design of the ascent phase of reusable Single Stage to Orbit (SSTO) launchers. Particular attention is paid to horizontal take-off and landing (HTOL) vehicles, given the growing interest and potential in the field of reusable launchers. The thesis has been carried out in collaboration with other two colleagues, whose work is focused on the sizing methodology and tool integration, and the analysis of the descent phase. At first, high-level requirements for the ascent phase of a reusable single-stage access to space vehicle are elicited. Subsequently, the problem of aerodynamic and propulsive characterisation of the vehicles is tackled, using a multi-fidelity approach. From the aerodynamic perspective, various analytical models already available in literature are analysed to identify the most suitable for a conceptual design of the ascent phase, in case numerical or experimental databases are not available. Taking vehicle geometrical data as input (e.g. the total volume of the vehicle, the wetted surface, the wing span), these analytical models shall provide as output the aerodynamic coefficients as a function of flight conditions (i.e., Mach number, angle of attack), spanning from subsonic to hypersonic regimes. The most promising models have been published by Curran, Williams and Torenbeek, while the Raymer model is considered not applicable, due to its limitation to configurations with a clear distinction between the fuselage and the delta wing, which made it not appliable for blended-body aircraft. From the propulsive perspective, the methodology supports the preliminary estimation of propulsive performance for all possible propulsive systems combinations and combined cycle engines that might support future reusable launchers. To this purpose, analytical formulations of thrust (or specific thrust) and specific impulse of each propulsion system are studied, required for the creation of a propulsive database. This analytical modelling is complemented by a statistical analysis of the available thrust for ramjet/scramjet engines as well as an insight on the ramjet inlet sizing. Following the aero-propulsive characterization, a revision of the Matching Chart is pursued to extend its applicability beyond the aeronautical sector, including access to space missions. For this purpose, the Multiple Matching Chart tool approach is upgraded adding a new requirement representing the minimum thrust-weight ratio as a function of wing loading necessary to reach the desired orbit. Moreover, the geometrical and aero-propulsive characterization pave the way to nominal and out-of-nominal mission analysis. In this case, the commercial software ASTOS is used. Finally, this thesis shows the application of the developed methodology to the SKYLON vehicle, a reusable single-stage-to-orbit spaceplane developed by the British company Reaction Engines Limited.