Multi-phase hypergolic ignition model

Abstract

A novel approach for modeling hypergolic ignition is presented, validated, and used to successfully predict ignition delay times for a hybrid rocket propellant configuration. This configuration employs a hypergolic additive (sodium borohydride) that allows two non-hypergolic reactants (polyethylene and hydrogen peroxide) to gain hypergolic capabilities. The model considers multiple phases and multiple species, heterogeneous and homogeneous chemical reactions, mass transfer between phases, and heat transfer. The transient behavior of the various chemical and thermal properties involved in hypergolic ignition is studied. In addition, a parametric investigation is conducted to predict ignition delay times as functions of multiple variables such as additive loading, oxidizer concentration, and initial propellant temperatures, among others. The results are presented in the form of ignition delay time contour maps. The heat release rate is shown to be controlled mostly by hypergolic chemical reactions. Gas homogeneous reactions only take place during the last portion of the ignition process and are the ones responsible for ultimately leading to a gas thermal runaway. The proper inclusion of the vaporization and pyrolysis of the propellants is found to be crucial since these determine the formation of a gas phase, where ignition is achieved. It is found that the vaporization of the liquid oxidizer is the controlling mass transfer mechanism for the hybrid rocket configuration considered. The model successfully predicts the minimum hypergolic additive loading and the range of oxidizer-to-fuel ratios required for ignition. It is found that the optimal oxidizer-to-fuel ratio leading to the shortest ignition delay time is mostly a function of additive loading. In addition, it is found that pressure, propellant initial temperature, and oxidizer concentration, have a major influence on ignition delay times and that they barely affect the optimal oxidizer-to-fuel ratio. The presented model, although evaluated for the hybrid rocket configuration, is considered to be applicable for any hypergolic propellant configuration.