Recent Advances in Cryogenic Solid Propellant Multiple Hybrid Combustion
- Paper number
IAC-05-C4.2.01
- Author
Dr. Roger Lo, AI: Aerospace Institute, Germany
- Year
2005
- Abstract
Roger E. Lo (corresponding author: Roger.Lo@aerospace-institute.com), Harry Adirim (both AI: Aerospace Institute, Berlin, Germany) Sascha Gläser, Volker Weiser, Norbert Eisenreich (FHG-ICT, Pfinztal, Germany) Ernst Noack (DLR, Bonn, Germany) Herman Schöyer (Schöyer Consultancy b.v., The Netherlands) Lev Filimonov, Serge Suchkov (MGTU Moscow State University, Moscow, Russia) “Recent Advances in Cryogenic Solid Propellant Multiple Hybrid Combustion” 56th Intenational Astronautical Congress, The paper describes recent experimental and theoretical advances concerning the combus-tion of cryogenic solid propellants (CSP) in hybrid rocket motors, sponsored by the German Aerospace Center DLR and the European Space Agency ESA at AI:Aerospace Institute, Fraunhofer ICT and Schöyer Consultancy b.v., NL. 1 kg (2,2 lbm) samples of solid hydrogen peroxide / polymer propellant were successfully hot-fired at FHG-ICT in the AI-X1 CSP test motor, a reusable test combustion chamber de-signed and built by AI within the framework of a current ESA TRP project that includes layout design, engineering and hot testing within a given time frame. The combustion chamber pro-vides means for igniting cylindrical CSP-charges under pressure and keeping them under self-sustained combustion. Using the NASA defined technology readiness levels (a scale from 1 to 9) it is actually a Level 4 device for component validation. In a typical experiment up to 40 pairs of ring-shaped fuel- (Polyethylene) and oxidizer- (frozen Hydrogen Peroxide) slabs form the propellant grain in what is called a disk-stack arrangement. Typical values of combustion pressure were 7.5 to 8.5 Mpa (75 to 85 bar) Combusion pressure increases from beginning to end due to the simple cylindrical combustion surface, that causes a pro-gressive burn. However, it is worth noting – other than what is seen with conventional hybrid combustion – that mixture ration is not affected by the changing combustion geometry. At present, the cylindrical shape is used for reasons of simplicity. However, in principle, any thrust-time shape that can be obtained with conventional solid propellants is just as easily obtained using multiple hybrid grains. The experiments described comprise the successful feasibility demonstration of multiple hy-brid CSP combustion under precisely defined conditions. Ignition conditions, pressure level, combustion duration and many other experimental parameters have been designed for fulfill-ing precisely defined feasibility demonstration conditions. Other propellant grains use sponge type propellant mixes (HTPB foam drenched with HP and frozen) The paper will present further details along with results of CFD simulations of hybrid disk stack and sponge combustion that were made by the MGTU group. By providing the means for selecting any desired rocket propellant combination in the form of an arrangement of slabs (rather than a fine grained mixture) CSP combustion bridges the gap between hybrid and solid rocket motors.
- Abstract document
- Manuscript document
IAC-05-C4.2.01.pdf (🔒 authorized access only).
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