Observing that the majority of SSTO programs failed due to the propulsion system being far ahead of its time. Furthermore, acknowledging the that the private launch markets is increasingly dominated by partly reusable TSTOs. This thesis aims to evaluate whether next generation rocket engines make an SSTO viable within a foreseeable economical time frame, for the increasingly privatised space industry. From the market analysis it follows that the VT SSTO RLV developed by the private launch industry, with a payload capability of 15- 20 tons to a LEO with an altitude of 200-400 km, is most likely. The evaluation is performed by a two level trade-off. Initially, a literature trade-off is preformed on elaborate range of potential next generation rocket engine, which are subdivided into pure rocket engine and breathing engine. All engines are evaluated on the market requirements, performance, and achievability. Only three engine types are selected. Upon evaluating the propulsion literature it was noticed that no literature cross evaluated advanced propulsion systems with high performance propellant pallets. Therefore, a literature propellant trade-off is performed. For the aerospike, the pulse detonation engine and the precooled hybrid airbreathing rocket engine with the propellant pallets H2/O2, CH4/O2, C2H2/O2, and C2H4/O2 a performance analysis is performed. The performance analysis is done by both optimising the engine configurations and the ascent trajectory for a 15 ton payload delivery to an orbit of 400 km altitude. To simulate each engine, three performance analysis are developed, which are all derived from the continuity equations to ensure a consistent comparison. The aerospike is simulated as a continuous optimal convectional via the frozen equilibrium method, although it is found an alternation is needed to account for directional losses. The open-end of the PDE is simulated via a CJ-detonation wave followed by an exponential decaying blowdown phase. The DC nozzle expansion is iteratively simulated via the frozen equilibrium method. The intake and HPC of the precooled hyrbid airbreathing engine is simulated with convectional aero engine performance analysis, while the precooler is simulated ass a counter-flow heat exchanger. The CC and CD nozzle are simulated via the frozen equilibrium method. From the optimised trajectory the representative performance of each engine is extracted and five engine configurations are identified as viable VT SSTO RLV engines. These are the H2/O2 powered aerospike, the pulse detonation engine and the precooled hybrid airbreathing rocket engine operating on an H2/O2 pallet, and the C2H2/O2 powered aerospike and pulse detonation engine. The H2/O2 powered PDE is found to be the most promising, as it offers the best performance gain relative to its achievability. Therefore, it advocates that future VT SSTO RLV research utilises the H2/O2 powered pulse detonation engine as the main propulsion system