Experimental demonstration of an end-burning swirling flow hybrid rocket engine

International audienceWithin the framework of the European H2020 HYPROGEO program, an innovative hybrid engine combustion chamber compatible with satellite requirements (constant thrust over very long burn-times) had to be developed. A first test rig was designed and tested in order to better understand the functioning of this innovative combustion chamber, and to help the design of a breadboard to demonstrate the efficiency of this new engine with respect to the mission requirements. Two test campaigns, the first on the test rig with 87.5% hydrogen peroxide, and the second on the hybrid engine breadboard with 98% hydrogen peroxide, were performed under various operating conditions to demonstrate the catalytic ignitability of this new hybrid engine, and the sustainment of a stable combustion over firing durations up to 180 s. The test campaigns also enabled the identification of the main influencing parameter on the fuel regression rate for this innovative combustion chamber

Modélisation de la régression des combustibles liquéfiables dans un moteur hybride

Le dimensionnement préliminaire d un propulseur hybride passe par une phase d essais à échelle réduite afin de caractériser entre autre la loi de régression du couple oxydant/combustible envisagé pour remplir les besoins de la mission en terme de performances, durée de fonctionnement, etc. Afin de limiter le recours à ces campagnes expérimentales onéreuses et génératrices de délais pour les industriels, il est nécessaire de développer des outils numériques fiables permettant de prévoir rapidement, sous différentes conditions de fonctionnement et géométries de chambre de combustion, la loi de régression d un couple oxydant/combustible. L objectif de cette thèse est de proposer une modélisation monodimensionnelle du mécanisme de régression des combustibles liquéfiables. Cette classe de combustibles offre des vitesses de régression trois à cinq fois plus élevées que celles rencontrées avec les combustibles généralement utilisés en propulsion hybride (PBHT par exemple). Ce modèle se base alors sur le transport de la phase gazeuse et du film liquide se développant sur le combustible solide, la vitesse de régression dépendant des transferts de masse et d énergie entre ces trois phases. Afin de valider cette approche et l architecture du code Hydres conçu pour la résolution de ce modèle et la prévision des performances propulsives d un moteur hybride, des campagnes expérimentales ont été réalisées sur les bancs d essais Hycarre et Hycom. Ces essais ont également permis de développer une technique de mesure permettant l obtention de la vitesse de régression instantanée du combustible, conduisant à la restitution de la loi de régression instantanée du couple oxydant/combustible.The preliminary design of a hybrid rocket engine requires lab-scale tests to characterize the regression law of the oxidizer/fuel pair intended to fulfil the mission needs in terms of performances, etc. To limit these costly and potentialy delaying experimental campaigns, it is necessary to develop reliable numerical tools to quickly predict the regression law of the oxidiser/fuel pair under different operating conditions and with different combustion chamber geometries. The objective of the thesis is to develop a one-dimensional model of the regression mechanism of liquefying fuels. These particular fuels offer regression rates three to five times higher than those found with classic polymers used in hybrid propulsion (eg. HTPB). The model is based on the transportof the gaseous flow and the liquid film which is developing along the solid fuel grain. The regression rate depends on mass and energy transfers between these three phases. To validate this approach and the Hydres numerical tool, specifically designed to solve this model and forecast the performances of a hybrid engine, experimental tests were performed with the Hycarre and Hycom facilities. These tests also allowed for the development of a technique to measure the instantaneous regression rate of the solid fuel, providing directly the instantaneous regression law of the oxidizer/fuel pair.TOULOUSE-ISAE (315552318) / SudocSudocFranceF

Experimental and numerical study of aeroacoustic phenomena in large solid propellant boosters

The present research is an experimental and numerical study of aeroacoustic phenomena occurring in large solid rocket motors (SRM) as the Ariane 5 boosters. The emphasis is given to aeroacoustic instabilities that may lead to pressure and thrust oscillations which reduce the rocket motor performance and could damage the payload. The study is carried out within the framework of a CNES (Centre National d'Etudes Spatiales) research program. Large SRM are composed of a submerged nozzle and segmented propellant grains separated by inhibitors. During propellant combustion, a cavity appears around the nozzle. Vortical flow structures may be formed from the inhibitor (Obstacle Vortex Shedding OVS) or from natural instability of the radial flow resulting from the propellant combustion (Surface Vortex Shedding SVS). Such hydrodynamic manifestations drive pressure oscillations in the confined flow established in the motor. When the vortex shedding frequency synchronizes acoustic modes of the motor chamber, resonance may occur and sound pressure can be amplified by vortex nozzle interaction.Original analytical models, in particular based on vortex sound theory, point out the parameters controlling the flow-acoustic coupling and the effect of the nozzle design on sound production. They allow the appropriate definition of experimental tests.The experiments are conducted on axisymmetric cold flow models respecting the Mach number similarity with the Ariane 5 SRM. The test section includes only one inhibitor and a submerged nozzle. The flow is either created by an axial air injection at the forward end or by a radial injection uniformly distributed along chamber porous walls. The internal Mach number can be varied continuously by means of a movable needle placed in the nozzle throat. Acoustic pressure measurements are taken by means of PCB piezoelectric transducers. A particle image velocimetry technique (PIV) is used to analyse the effect of the acoustic resonance on the mean flow field and vortex properties. An active control loop is exploited to obtain resonant and non resonant conditions for the same operating point.Finally, numerical simulations are performed using a time dependent Navier Stokes solver. The analysis of the unsteady simulations provides pressure spectra, sequence of vorticity fields and average flow field. Comparison to experimental data is conducted.The OVS and SVS instabilities are identified. The inhibitor parameters, the chamber Mach number and length, and the nozzle geometry are varied to analyse their effect on the flow acoustic coupling.The conclusions state that flow acoustic coupling is mainly observed for nozzles including cavity. The nozzle geometry has an effect on the pressure oscillations through a coupling between the acoustic fluctuations induced by the cavity volume and the vortices travelling in front of the cavity entrance. When resonance occurs, the sound pressure level increases linearly with the chamber Mach number, the frequency and the cavity volume. In absence of cavity, the pressure fluctuations are damped.Doctorat en sciences appliquéesinfo:eu-repo/semantics/nonPublishe

Characterisation of the ONERA / CNES High Performance Green Monopropellants Towards a Thruster Application

International audienceThis paper describes the evolution of surface tension of three green monopropellants according to temperature to facilitate the design of a green monopropellant thruster prototype. During the testing, reactions with some of the materials in contact with these monopropellants have been observed leading to a material compatibility testing (60°C for 10 days). Finally, a preliminary testing of the combustion of an isolated droplet and a monopropellant atomisation have been made and filmed to begin the visualisation of the phenomena taking place

Numerical issues in the application of an amiet model for spanwise-varying incoming turbulence

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Prediction of incoming turbulent noise using a combined numerical / semi-empirical method and experimental validation

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Numerical prediction of the noise emitted by an airfoil in a turbulent jet with an aeroacoustic hybrid method

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Parametric Study of a 1.5-D Combustion Chamber Model on the Hybrid Rocket Engine Performances

International audienceA sensitivity analysis of a 1.5-D combustion chamber model of a hybrid rocket engine is performed in this paper. The goal is to assess the impact that some of the most important engine parameters have over its performances. Studies are carried out by considering three categories of parameters: the aerodynamic characteristics at the inlet of the chamber, the thermochemical quantities involved in the gas-surface interaction model, and the geometrical properties of the fuel block. Simulations have been made at steady-state regime for a cylindrical, lab-scale combustion chamber with a 1-D nozzle model, using mainly gaseous oxygen as oxidizer and high density polyethylene as fuel. The fundamental reference variables used for the sensitivity studies have been the regression rate, the averaged chamber pressure, the radial profiles of temperature and species mass fractions, and the thrust and specific impulse of the engine. Fuel regression rate results have shown a high dependence upon the oxidizer mass flux, the motor size, the number of ports of the fuel block, and the variables intervening directly on the energy balance at the fuel surface, such as the radiative heat flux source and the composition of the solid fuel. The retrieved influence of these parameters on engine performances has been found to be in agreement with the literature data, being in some cases of the same intensity

Effect of boundary conditions and source truncation in the prediction of flow-generated sound

info:eu-repo/semantics/publishe