Resonance igniters are a promising alternative to conventional ignition devices for rocket engines using non-hypergolic propellants. This paper presents the development and analysis of a resonance igniter using gaseous oxygen and methane, supported by experimental measurements and numerical modelling. The effect of nozzle gap distance on acoustic resonance heating is investigated using oxygen and nitrogen as driving gases. Microphone data are used to determine the operating mode of the igniter; thermocouple data acquired on the outside of the resonator tip are used to evaluate heating performance across various nozzle pressure ratios and nozzle gap distances. A numerical model based on the open-source CFD software SU2 is developed and validated against resonance heating experimental data. This non-reacting flow model accurately captures the transition from the high-frequency Jet Screech Mode to the lower-frequency Jet Regurgitant Mode. Furthermore, it identifies the operational parameters leading to the highest rates of resonance heating observed in the experiments. Ignition attempts in non-premixed conditions, using gaseous oxygen and methane, show that the separate injection of methane in cross-flow into the combustion chamber causes severe disruption of resonance heating, preventing ignition.

With Mars colonisation becoming a reality for the near future, it is of importance to analyse how crew and cargo can be transported between Earth and a colony on Mars. This article is a feasibility and design study of a launch vehicle whose mission is to shuttle crew and cargo from Low Mars Orbit to a colony on the Martian surface. A single-stage reusable rocket has been selected to fulfil this mission, code-named Charon. The mission profile of such a vehicle was created, leading to a Maximum Growth Allowance (MGA) Delta-V budget of 6.2 km/s. With the mission profile in mind, each subsystem underwent a preliminary design. With reliability and maintainability in mind, subsystems were designed for redundancy and modularity, and an abort system was included for an added level of safety. The iterative design process resulted in a vehicle with a MGA mass of 198.7 tons, capable of transporting 1200 kg of cargo and a crew of 6 people to a 500 km orbit and back. The preliminary design of the vehicle is deemed safe. Following a fault tree analysis, the Single Launch Loss of Mission, Vehicle and Crew (SL-LOM, SL-LOV, SL-LOC) probabilities are computed to be of 0.975%, 0.12%, and 0.079%. Finally, from the vehicle’s constraints on the base, the feasibility of the project has been reflected upon. It is deemed that such a concept is of high interest only when the base is already operational, due to the launch and maintenance infrastructure that it requires, as well as the power it requires from the Martian base.