Annamaria Russo Sorge

Role of Injection in Hybrid Rockets Regression Rate Behavior

A series of firing tests was carried out to investigate the influence of the oxidizer injection on the solid fuel regression rate behavior in a hybrid rocket engine. For this purpose, a conical subsonic nozzle as the injector of the gaseous oxidizer was selected to generate nonuniform conditions at the entrance of the fuel port. Gaseous oxygen and polyethylene fuel cylindrical grains were used. When the oxygen was fed by this kind of injector, the fuel regression in the region of the oxygen impingement on the grain’s surface was increased several times, which led to irregular fuel consumption with the characteristic afterburn port shape typical of solid fuel ramjets having a rearward-facing step at the air inlet. The local instantaneous regression rates, measured by means of the ultrasound pulse-echo technique, showed regression rate-time profiles strongly dependent on the geometric configuration and helped to explain the complex regression phenomenon deriving from the impingement zone displacement during the time. Empirical correlations for the prediction of the average regression rate were developed taking into account the influence of mass flux, pressure, and port diameter. Finally, a nondimensional semi-empirical correlation involving the effect of the blowing number yielded improved accuracy.

Basic Aspects of the Hybrid Engine Operation

In the framework of the Hybrid Propulsion Evaluation Program developed under the European Space Agency financial support, a series of static firing tests have been carried out on a lab-scale hybrid engine fed by gaseous oxygen and two different fuel compositions and grain geometries, to assess some basic aspects of the hybrid rocket internal ballistics and operation. Pure HTPB and HTPB with aluminum particles were tested; single bore cylindrical grains of HTPB have been burned by injecting oxygen either with an axial conical subsonic nozzle or a radial injector. The main purpose has been starting to investigate, on the one hand, the effect of aluminum particles on the fuel regression rate with a given injector and, on the other, the influence of oxygen injection on combustion efficiency, stability, and regression rate itself, with a given fuel formulation. In addition, multiport grains of pure HTPB have been tested. Stability issues have also been considered through the possible damping effect of the aluminum particles. Fundamental parameters such as average and instantaneous fuel regression rate, fuel consumption uniformity, combustion efficiency and combustion stability have been evaluated in order to gain information on the motor operating characteristics with the different fuel grains. Results are presented to compare the effects of the fuel formulation and geometry.

DESIGN OF A LAB -SCAL E COOLED TWO -DIMENSIONAL PLUG NOZZLE FOR EXPERIMEN TAL TESTS

Present space transportation systems though, on the one hand, are able to solve many economical and managerial problems, on the other have given rise to some new ones of purely propulsive nature. In fact, the engines of these vehicles have to perform high thrusts and have to operate at very different altitudes (0 400 km ), therefore their nozzles have to generate high expansion ratios and, as a consequence, have to work for long time in over -expansion. Furthermore, the use of highly energetic pr opellants (H 2-O2), but easily dissociable (H 2), leads to heavy losses for frozen effect. Unconventional nozzles seem to meet very well these requirements, ensuring small losses in terms of specific impulse in overexpansion conditions and the almost total and immediate recombination of dissociated species. However one of the most serious problems of unconventional nozzles is the higher thermal stress on the wall, particularly in the throat region whose dimensions are very small. Then a cooling system is cert ainly required to ensure a long operation time. Actual numerical models concerning hot flow in advanced nozzles have to be improved and they need experimental tests that can show the feasibility of the realization of this kind of configuration, even from a technological point of view. For this reason a two -dimensional plug nozzle has been designed and constructed so that the thermo -fluid dynamic behavior of truncated plug nozzles can be tested in order to try of reducing the related technological troubles. The nozzle shape has been chosen following some basic criteria for making easier the plug cooling. In this paper some numerical results of the flow -field in over -expanded condition with turbulent reacting flow and the heat flux distribution along the plug have been reported. The nozzle will be tested on the hybrid rocket set up in the Propulsion Laboratory of DISIS, that allows to perform tests safely and cheaply at different running pressures up to 30 atm and different mass flow rates up to 0.3 kg/s . The acq uisition system employed allows to measure the motor thrust and to calculate the specific

Catalytic Ignition in Hydrogen Peroxide-based Space Propulsion Systems

This paper describes an innovative catalytic system, able to decompose hydrogen peroxide. The development process has been carried out following a step-by-step approach. Firstly, a plug-flow reactor has been designed and used to perform a preliminary screening of catalysts prepared according to different techniques and identify the most promising preparation methods. A mono-propellant test bench has been then assembled to verify the behavior of a selected number of catalytic samples in a test environment closer to the actual operating conditions of a space propulsion system. The concept has been finally scaled-up to a hybrid configuration, based on the utilization of hydrogen peroxide as the oxidizer and high density polyethylene as the fuel.

Experimental evaluation of the radiative heat flux from rocket plumes on satellite structures

Abstract An experimental method capable of directly measuring the net radiant heat flux on complex geometry surfaces by non invasive methods has been devised, designed, realized and tested. The method is based on a computerized thermographic system capable of directly outputting the radiative heat fluxes absorbed by the surface to be examined. The calibration tests have shown the feasibility and the accuracy of the proposed method. Preliminary tests are presented which simulate the radiation flux from a jet exhaust.