Design enhancement and analysis of a 20,000 lbf class annular aerospike engine utilizing LOX/propylene as propellants
Final Report Abstract
The research project dealt with the development and analysis of a 30kN annular, multi-chamber aerospike engine, which uses LOX and densified propylene as propellants. The engine is designed to be used as a first stage engine of a two-stage pressure-fed Nanosat Launch Vehicle (NLV) capable of lifting a payload of up to 10 kg into a low earth orbit. The engine development includes detailed designs of the injector and ignition system, thrusters and a truncated plug nozzle along with cooling technologies and materials used. The subsystem selection is based on reliability and a low-cost concept. Due to limited space for each thrust cell and its injector, hypergolic ignition is applied by placing a TEA/TEB cartridge into the engine’s feed lines, which leaves maximum injector space to be used for propellant ducts. Each injector is fed by three propellant lines, i.e., an outer fuel line, an oxidizer line and a small centered fuel line, in which the hypergolic cartridge is integrated. Each injector includes 72 fuel- and oxidizer orifices, which form an inner circle of 12 and an outer circle of 24 split-triplets to allow an even distribution of the propellant mixture into the chamber. 12 additional fuel orifices along the chamber wall are used for film cooling. The design of the aluminum injector follows common rules to minimize pressure losses and allows low-cost manufacturing. The plug nozzle using phenolic resin coating for ablation cooling is designed according to the common application of rules by Angelino, assuming constant gas composure and axisymmeric flow field. Adjacent thruster flow interactions are neglected. The plug is truncated at 34% of its theoretical isentropic length to reduce overall weight by simultaneously maintaining a high thrust level. A major enhancement compared to a classical multi-chamber design with straight thrusters is achieved by introducing an advanced thruster design with nozzle curvature in the circumferential and in the longitudinal direction. This allows the use of a larger extent of the limited available cross section and increases thrust. The cells are made of the refractory ceramic matrix composites (CMC) C/CSiC, which in addition to the fuel film cooling is radiatively cooled. A simple design allows manufacturing of a thrust cell by wrapping carbon fibers around two reusable molds, which keeps production costs low. The design work of the thrusters is supported by numerical analyses of the flow in a single thruster and along the whole engine, i.e., inside thrust cells and along the truncated plug, at various points along a typical ascent trajectory including slipstream effects. For the flow simulation, Reynolds- Averaged Navier-Stokes equations with a realizable k-ω turbulence model are solved. Results show that longitudinal curvature leads to a radial pressure stratification of the thruster exit flow, which is caused by internal expansion and compression waves, with a decreased pressure level at the plug nozzle side and an increased pressure level towards the outer thruster lip compared to a linear layout. Circumferential curvature leads to a more homogeneous flow in the lateral direction reducing pressure peaks at thruster intersections. The larger engine area realized with the dual curved thrusters leads to an improved performance, which is about 2% above the linear design and above an ideal bell nozzle of equal area ratio. Improvements result from increased thruster performances at higher pressure ratios and larger contributions of the plug, which comprises a larger area. In addition, engines with curved thrusters utilize greater portion of the launcher's cross section and thus reduce base drag.
Publications
- Advanced Design of a Multi-Thruster LOX/Propylene Aerospike Engine. 49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference (JPC), San Jose, California, July 2013
J.-H. Meiss, E. Besnard
