Abstract
This thesis evaluates the performance of a small-scale liquid-propellant rocket engine equipped with a truncated graphite aerospike nozzle. Thermochemical calculations, simulated results, and CFD predictions established expected performance targets and validated the aerospike design prior to testing. Finite element analysis confirmed that components could withstand expected thermal and mechanical loads. Static hot-fire testing at the Friends of Amateur Rocketry site produced a maximum thrust of 907 N and a chamber pressure of 1.25 MPa, both lower than predicted due to injector oversizing and pressure losses. Despite reduced performance, the measured thrust coefficient of 1.51 exceeded the predicted conventional nozzle value of 1.47, experimentally demonstrating the aerospike’s efficiency advantage reported in prior literature. These results confirm the feasibility of a graphite aerospike nozzle for small-scale propulsion applications and identify injector design and bonding methods as key areas of improvement.