Experimental investigation of the regression rate in an end-burning swirling flow hybrid rocket engine
Creators: name:  T. PALACZ 
name:   J. CIESLIK 
Publisher: name: Proceedings of  the 9th European Conference for Aerospace Sciences. Lille, France, 27 June - 1 July, 2022
Year: 2022
Subjects: PROPULSION PHYSICS
HYBRID PROPULSION

description: In recent years, there is growing interest in satellites\' commercial use, especially in the smallsat segment ( 500 kg). The fast development of small satellites in large numbers resulted in high demand for dedicated propulsion systems to perform various maneuvers such as orbit change, constellation deployment, station-keeping, formation flying, CAMs, or attitude control. Miniaturized electric thrusters enable high-performance, low-thrust maneuvers, but traditional chemical propulsion cannot be easily adapted for small satellites\' needs. The major requirements are to be green, safe, reliable, low-cost, and, at the same time, have high performance. Hybrid rocket engines, which are usually used for sounding rockets, may offer these benefits. An end-burning hybrid rocket engine is a concept first proposed by Rice et al. [1], in which the fuel disk is placed on top of the combustion chamber, while the oxidizer is injected tangentially to create a so-called vortex flow field or swirling flow. The exposed fuel area is kept constant throughout the burn, which should translate to the constant thrust and OF ratio in theory. Some variations have been studied by Hayashi et al. [2] and Paravan et al. [3]. A significant amount of work has been done within the HYPROGEO project, whose goal was to develop a constant 250 N thrust, long burn hybrid engine using an end-burning, swirling-flow design [4]. In the above papers, the fuel regression rate\'s radial distribution was axisymmetric but far from uniform. The heat coefficients distribution may explain non-uniformities at the end surface related to the swirling flow field [5]. In this paper, we develop the end-burning, swirling-flow hybrid rocket engine to study internal dynamics. The engine is sized for the thrust of around 5 N for use in small satellites, and this results in a significantly smaller diameter of the combustion chamber than in previous studies. Low-regression rate fuel is used, and GOX is injected tangentially to induce swirling-flow. The engine is tested with different oxidizer flowrates and burn durations. Axial and radial regression rates are observed and linked to the oxidizer mass flux and swirl-flow. References [1] E. E. Rice, M. J. Chiaverini, C. P. St. Clair, W. H. Knuth, and R. J. Gustafson, \"Mars ISRU CO/O2 hybrid engine development status,\" 38th Aerosp. Sci. Meet. Exhib., no. c, 2000, doi: 10.2514/6.2000-1066. [2] D. Hayashi and T. Sakurai, \"A Fundamental Study of An End - Burning Swirling - Flow,\" in 51st AIAA/SAE/ASEE Joint Propulsion Conference, 2015, pp. 18. [3] C. Paravan, J. Glowacki, S. Carlotti, F. Maggi, and L. Galfetti, \"Vortex combustion in a hybrid rocket motor,\" 52nd AIAA/SAE/ASEE Jt. Propuls. Conf. 2016, pp. 121, 2016, doi: 10.2514/6.2016-4562. [4] J. Y. Lestrade, J. Anthoine, A. J. Musker, and A. Lecossais, \"Experimental demonstration of an end-burning swirling flow hybrid rocket engine,\" Aerosp. Sci. Technol., vol. 92, pp. 18, 2019, doi: 10.1016/j.ast.2019.05.057. [5] E. P. Volchkov, S. V Semenov, and V. I. Terekhov, \"Turbulent heat transfer at the forward face surface of a vortex chamber,\" J. Eng. Phys., vol. 56, no. 2, pp. 109115, 1989.