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Regenerative Cooling System for a Hybrid Rocket Engine Nozzle

Filippo Landi

Regenerative Cooling System for a Hybrid Rocket Engine Nozzle.

Rel. Dario Giuseppe Pastrone. Politecnico di Torino, Corso di laurea magistrale in Ingegneria Aerospaziale, 2020


This paper present a study carried out in the frame of a collaborative research project carried out between California Polytechnic State University (Cal Poly) and Politecnico di Torino (PoliTO) regarding the development of an innovative, simple, low cost, regenerative cooling system for a convergent-divergent nozzle of a N2O-PMMA hybrid rocket engine. Regenerative rocket nozzle cooling technology is a well-developed technology for liquid rocket engines (LREs), but it has yet to be widely applied to hybrid rockets. Due to the high heat fluxes that a rocket experiences in its nozzle throat, most nozzles are predisposed to ablation, supporting the need to implement a cooling system. An effective solution is cooling the nozzle using a regenerative scheme: the propellant is first used as a coolant for the nozzle and then re-injected in the combustion chamber. In this case, the enthalpy of the propellant is increased, so that the specific impulse of the system may even be slightly improved. An oxidizer often used in hybrid rocket motors is nitrous oxide (N_2O). Because of its high vapor pressure at ambient condition, N_2O can be self-pressurizing, thereby making it a convenient choice of oxidizer for simple, low-cost and low hazard level applications. Several nozzles, including aerospike and convergent-divergent ones, were already designed, built and tested at Cal Poly using N_2O as a coolant, though never in a regenerative setup, making the venue a logical one for this project. The solid fuel is commonly made from polymers such as acrylic or paraffin wax, since that these materials are easy to manufacture, chemically inert and safe to store; for these reasons Plexiglas® was chosen to run the experiments. In the case of aerospike nozzle, a configuration obtained by means of CNC machining was used: an outer cylinder liner surrounds the combustion chamber wall with a helical internal surface that runs along the vertical axis of the rocket. For convergent-divergent nozzle test and verification of the concept, a simpler, cylindrical grain will be used. The detailed design of the test stand is based on the results of previous research involving experimental testing of oxidizer-cooled converging-diverging nozzles. In the hybrid rocket engine, nitrous oxide will be used as oxidizer and it will be routed to a cooling copper annulus surrounding the throat of the nozzle and then it will be directly re-injected in the head-end part of the combustion chamber after cooling.

Relators: Dario Giuseppe Pastrone
Academic year: 2020/21
Publication type: Electronic
Number of Pages: 79
Additional Information: Tesi secretata. Fulltext non presente
Corso di laurea: Corso di laurea magistrale in Ingegneria Aerospaziale
Classe di laurea: New organization > Master science > LM-20 - AEROSPATIAL AND ASTRONAUTIC ENGINEERING
Ente in cotutela: California Polytechnic State University (STATI UNITI D'AMERICA)
Aziende collaboratrici: UNSPECIFIED
URI: http://webthesis.biblio.polito.it/id/eprint/15724
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