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International audience; In this work, a novel approach to increase the performance of a Hybrid Rocket Engine (HRE) by approximating any idealized fuel port geometry by steps inside the fuel grain is proposed. This work serves as a first investigation into the effect of single steps inside the fuel grains of a typical HRE. In numerical simulations, it is shown that forward- and backward-facing steps have the potential to increase the average regression rate as they promote the turbulence and heat transfer inside the turbulent boundary layer. Both types of steps are more effective at lower oxidizer mass fluxes. Finally, as a first estimation, a set of best practices for the distribution of steps along a profile could be derived.

Unlike classical pyrotechnical solutions such as black powders, laser ignition is a non-intrusive, scalable method making it an interesting ignition technique for solid propellants in academia. When studies focus on the solid propellant gas flow, the shape of the burning surface becomes crucial. It is usually necessary to have a burning surface as flat as possible. The laser ignition delay of solid propellants follows a power law function of the laser heat flux. Because this law is not linear, ignition time varies strongly even for small laser irradiance variations. More precisely, the lower the laser irradiance, the more difficult it is to obtain a plane ignition. In this paper we present a calculation method to estimate the shape of the burning surface for a solid propellant sample based on the laser irradiance spatial distribution. For this purpose, a one-dimensional surface heating model is implemented together with a solid-propellant ignition model. This allows to calculate a local ignition delay as a function of the incident heat flux. The Gaussian spatial heat flux distribution has been discretised over the beam radius, and an ignition time calculation is performed for each spatial discretisation point. Considering that the solid propellant is a good thermal insulator, the ignition delay is independently calculated for each radius. Finally, the temporal evolution of the burning surface is determined, taking into account the flame propagation over the propellant surface. These results show the evolution of the surface with different laser parameters (i.e. beam size and power). This allows us to define, based on the laser irradiance distribution, a radius limit for which planar ignition can be assumed. An experimental setup is presented to measure the shape of the burning surface as a function of heat flux distribution for low laser irradiance. The calculation method is demonstrated with physical properties representative of a classical AP/HTPB research propellant composition. The approach is very promising to optimize ignition conditions for future experiments that require planar ignition.

A system design tool allowing to simulate the complete operation of a hybrid rocket engine is developed in this paper. The algorithm of the application is made in such a way that allows the simulation of an element, an ensemble of elements or all the elements constituting a hybrid space propulsion system. The objective is to simulate and predict the behavior of the engine at different conditions / configurations during the first pre-design phases. As the tool will be exclusively used during these phases, a balance between the duration of the simulation and its precision must be reached. Given these constraints, a maximum error of 30% between the experimental and the numerical results is considered adequate to validate the solution. The main parts of our HYCAT engine-characterized by a catalytic injection of oxidizer-are modeled and included in this tool: the mass flow rate regulator and the catalyst (feed and injection subsystems) through 0-D models; the combustion chamber with a 1.5-D model; and the nozzle through a 1-D model. An iterative method is employed to attain the pressure convergence in the combustion chamber between these three main parts of the engine whose corresponding system of equations is solved by a Newton-Raphson technique. Seven experiments performed on HYCAT are used to validate the results of the conceived tool. From all the performed simulations, only two of them present relative differences with the experiment above the defined threshold. The largest errors are produced for the fuel regression rate, the consumed mass of fuel, and the mixing ratio variables. The best agreements with the experiments are found for the simulations with the largest oxidizer mass fluxes (G ox >230 kg/m 2 /s) and where the mixing ratio is closer to the stoichiometric value, defining thus, the range of applicability of the system design tool.

Hybrid Rocket Engines (HREs) are a promising technology to achieve a cheaper, more sustainable and competitive access to space. So far, liquid rocket engines and solid rocket motors have been preferred over HREs for launcher applications. This is due to the low maturity of HREs, as well as challenges related to low regression rate, combustion inefficiency, oxidizer-to-fuel ratio (O/F) shift during operation and high residual inert mass. Nonetheless, in recent years, the space launcher market calls also for cheap micro- and nano-launchers. HREs could help to fill this niche, given their advantages (e.g. high performance, intrinsic safety, throttle-ability, flexibility and low cost), provided that some of the disadvantages are being solved. In this paper, the effect of Forward Facing Step (FFS) and Backward Facing Step (BFS) on the HRE performance (namely regression rate and combustion efficiency) is investigated through test firings. The application of diaphragms, resembling both a FFS and BFS at the same time, and other turbulence and heat exchange enhancing devices in the solid fuel grain have proven to increase both regression rate and combustion efficiency due to the formation of recirculation zones. These zones increase turbulence, mixing as well as heat transfer and, thus, leading to improved performance of the HREs. In order to gain further insight into this phenomenon, an experimental campaign is carried out on a lab-scale HRE using nitrous oxide (N2O) and High-Density Polyethylene (HDPE) as propellants. The cylindrical fuel grain with an external diameter of 130 mm and a length of 117 mm is separated halfway into two different internal diameters. Varying these diameters in the two sides allows creating different FFS and BFS geometries. The focus of the test campaign can be summarized as follows: First, the influence of the step height on the length of the recirculation zone is investigated. Therefore, the fuel grain is probed after the test for indications of the recirculation zone length. Further, the performance impact of BFS and FFS on the HRE is compared under identical conditions. This enables to distinguish the effects for BFS and FFS. Another measurement is the influence of the steps on regression rate and combustion efficiency by weighing the fuel grains before and after the tests and quantifying the thrust respectively. Lastly, the evolution of the steps over time can be assessed by comparing the final geometry of the stepped grain after different burn time lengths. To recapitulate, assessing the test campaign will allow to judge the practically of steps as performance enhancer in HREs.

International audience; Hybrid Rocket Engines (HREs) generate thrust by combining a liquid and a solid propellant. Given the hybrid nature, HREs possess several advantages over Liquid Rocket Engines (LREs) and Solid Rocket Motors (SRMs). Hybrid engines are cheaper and less complex than LREs while being safer and more sustainable than SRMs. HREs are throttleable as well as re-ignitable and their theoretical achievable specific impulses (Isp) lie between SRMs and LREs. However, hybrid engines suffer from several disadvantages that refrained them from reaching their full potential yet. Among other drawbacks, HREs have a low maturity, high fuel residuals and a low solid fuel regression rate. In fact, the low regression rate is considered the limiting characteristic for HRE applications. Around the globe, great research effort is put into enhancing the regression rate. Summarizing, the authors group the techniques together as a) adjusting the solid fuel properties, b) advanced injection methods and designs and c) improving the combustion chamber design. Each of these techniques comes with a different set of advantages and drawbacks, which need to be considered when designing a HRE.

International audience; The trend towards smaller satellites and mega-constellations has enormously changed the space sector and its utilisation in the last decades, allowing new players to enter the market and introducing stringent requirements to enable a variety of novel applications. Alongside, also the launcher market is undergoing a transformation epoch: the development, manufacturing, and integration of launcher systems is being shifted from the hands of governmental institutions to commercial industry. Moreover, nations like Unites States, China, India, and New Zealand are increasing the competition and pressure on Europe, urging the goal to ensure European autonomy in accessing and using space in a safe and secure environment. Europe does not only need innovations, but primarily a new generation of engineers, capable of understanding the full complexity of launcher development and trained to create and realise the necessary innovations. In this context, ASCenSIon is a multidisciplinary training programme involving 15 Early Stage Researchers (ESRs) from anywhere in the world, focused on several specific areas of cutting-edge space access research, particularly on launcher systems that are (partially) reusable and capable of injecting multiple payloads into multiple orbits. The network aims to identify and advance critical technologies to prove a feasibility of these concepts, and to advance the State of the Art in the field. ASCenSIon, whose acronym stands for “Advancing Space Access Capabilities –Reusability and Multiple Satellite Injection”, is a consortium of 11 beneficiaries and 17 partners across Europe, eager to contribute to the establishment of an ecologically and economically sustainable space access for Europe, oriented towards user needs. Unlike other single-aspect research projects, the core objective of ASCenSIon is not only to train 15 PhD students to become excellent specialists in their respective field, but also to provide them a thorough understanding of the complexity, multidisciplinary, and internationality of launcher development, in order to become leaders in the European effort of utilising space. This will be achieved through secondments, events, and lessons from experts, but mostly through strong interconnections among the ESRs, who will work on Individual Research Projects with a multi-disciplinal and multi-sectoral approach. This paper aims to provide an overview of ASCenSIon programme. Its values and core objectives will be introduced, together with the innovative aspects and content structure. An overview of the research methodology and recruitment strategy will be given, with a particular focus on the contributions and synergies of all participating organisations, core of such a novel training approach.

International audience; Thermodynamic properties as density and the speed of sound as a function of pressure and temperature are not very well-known for higher temperatures for typical storable propellants as Hydrazine, MMH or NTO. Most data is based on low temperature experiments that have been extrapolated or calculated from theoretical correlations. To our knowledge no experimental data exists up to the critical temperature except for a single experiment with Hydrazine that was carried out by Lobry de Bruyn [4] in 1896 and suffered from a catastrophic decomposition while trying to reproduce the results. This paper deals with the design, development and use of an experimental installation capable of measuring the density and the speed of sound of the liquid phase as well as the density of the gaseous phase for these storable propellants from (close to) the triple point up to the critical point. Utmost care has been given to the safety aspects of the remotely operated installation. Low cost solutions have been devised to respect a rather tight budget. The installation is currently in its trial and calibration phase using pure water, this paper contains some results of these tests. The real propellant tests are bound to start in the coming months.

International audience; Satellite attitude and orbit control manoeuvers are mostly realized by chemical thrusters, using catalytic decomposition of hydrazine. However, hydrazine is known for its carcinogenic and toxic effects and its use is thus threatened by the European chemical regulation REACh. A new monopropellant is therefore under development by CNES (the French Space Agency), and 3 prototypes are still in competition. The final choice of the best formulation will be based on real thruster tests. These new propellants are based on energetic ionic liquids, and the expected flame temperature is about 3000 K. To develop an optimized thruster demonstrator, numerical simulations, combining heat transfer and radiation in complex materials and diphasic behavior are necessary. Energetic ionic liquids have a very complex way of decomposing which explains why it is impossible to use current phase change models implemented in ONERA’s multiphysics code CEDRE. A new macroscopic combustion model is being developed and implemented in CEDRE following the phenomenology of solid homogeneous propellants. Simultaneously an experiment is prepared to observe the combustion of an isolated droplet, in order to measure its regression. This experiment is a proof of concept for the final experimental CNES’s setup GIMLI. This experiment will be validated on ethanol combustion and tested with the new family of green monopropellant.

International audience; Due to the regression of the burning fuel surface, the internal geometry of a hybrid rocket motor changes during operation. This implies a modification of the internal flow and the area of the wetted surface. The result is a change in mixing ratio and therefore a modification of the performances of the motor. This paper presents a genetic optimisation algorithm aiming to reduce the variation of the mixing ration while minimising the fuel residues by modifying the initial geometry of the grain. The use of a 3D printer allows for exotic geometries not feasible with more classic fabrication processes. Finally the optimised grain has been test fired and the results are analysed.

International audience; Oxidizer injection plays a major role in the behaviour of hybrid rocket engines, especially regarding combustion efficiency and stability. This study aims to compare two firing tests with associated numerical simulations based on gaseous oxidizer axial or swirl injection through a catalyzer. The oxidizer and fuel couple used was H2O2 (87.5 %)/HDPE. Numerical 3D MILES (Monotone Integrated Large Eddy Simulation) simulations of both configurations were performed and analyzed regarding on one side the combustion efficiency influenced by propellants mixing, and on the other side the pressure oscillations provoked by the formation of large scale vortices in the aft-combustion chamber.

Thermodynamic properties as density and the speed of sound as a function of pressure and temperature are not very well-known for higher temperatures for typical storable propellants as Hydrazine, MMH or NTO. Most data is based on low temperature experiments that have been extrapolated or calculated from theoretical correlations. To our knowledge no experimental data exists up to the critical temperature except for a single experiment with Hydrazine that was carried out by Lobry de Bruyn [4] in 1896 and suffered from a catastrophic decomposition while trying to reproduce the results. This paper deals with the design, development and use of an experimental installation capable of measuring the density and the speed of sound of the liquid phase as well as the density of the gaseous phase for these storable propellants from (close to) the triple point up to the critical point. Utmost care has been given to the safety aspects of the remotely operated installation. Low cost solutions have been devised to respect a rather tight budget. The installation is currently in its trial and calibration phase using pure water, this paper contains some results of these tests. The real propellant tests are bound to start in the coming months.

Hybrid rocket motors combine solid and bi-liquid chemical propulsion technologies and associate a solid fuel and a liquid oxidizer in its classical configuration. This architecture offers several advantages over liquid propulsion such as lower costs and a simplified architecture. The possibility of performing multiple extinctions and re-ignitions and a good theoretical specific impulse is also an improvement in regard to solid propulsion. Hybrid engines also have improved safety and a lower environmental impact than other chemical propulsion systems. As in all combustion chambers, hybrid engines suffer from pressure oscillations under specific operating conditions. These instabilities provoke thrust fluctuations that can damage the launcher and payloads. Various phenomena can induce the pressure oscillations observed in hybrid rocket engines. The objective of this thesis is to propose a model of hydrodynamics instabilities that appear in hybrid engines. A new exploitation of the database available at ONERA, and unsteady 2D and 3D numerical simulations were used for the modeling. The instabilities are provoked by the periodic formation of vortices in the combustion chamber that generate pressure fluctuations when passing through the nozzle throat. The originality of the model, which is based on the classical theory of vortices generation in a cavity, consists in taking into account the geometrical variations of the combustion chamber during operation. These variations have an effect on the flow velocity, on the recirculation area in the postchamber and on the vortices. Finally, several new firing tests of the hybrid engine HYCOM have been performed and compared to the model developed in this thesis.; Les moteurs-fusées hybrides combinent les technologies des deux autres catégories de moteurs à propulsion chimique, et associent un combustible et un oxydant stockés respectivement sous phase solide et liquide. Cette architecture offre un certain nombre d’avantages, comme par exemple des coûts plus faibles et une architecture simplifiée par rapport à la propulsion bi-liquide; la possibilité de réaliser de multiples extinctions et ré-allumages et une bonne impulsion spécifique théorique par rapport à la propulsion solide, et enfin une sécurité de mise en oeuvre accrue et un impact environnemental faible vis-à-vis de ces deux autres modes de propulsion. Comme toutes les chambres de combustion, celles des moteurs hybrides peuvent subir des oscillations de pression sous certaines conditions de fonctionnement. Ces instabilités se traduisent par des fluctuations de poussée qui peuvent dégrader la structure d’un lanceur ou d’un satellite. Des phénomènes divers peuvent être à l’origine des fluctuations de pression observées dans les moteurs hybrides. L’objectif de la thèse est de proposer une modélisation des instabilités d’origine hydrodynamique qui apparaissent dans les moteurs hybrides. Une exploitation nouvelle de la base de données disponible à l’ONERA a servi de support pour la modélisation, ainsi que des simulations numériques instationnaires 2D et 3D réalisées à l’aide du code CFD CEDRE. Les instabilités sont provoquées par la formation périodique de structures tourbillonnaires dans la chambre de combustion, qui génèrent des fluctuations de pression lors de leur passage dans le col de la tuyère. L’originalité du modèle, basé sur la théorie classique de génération tourbillonnaire dans une cavité, consiste à prendre en compte les variations géométriques de la chambre de combustion au cours des tirs. Ces variations ont un effet sur la vitesse de l’écoulement, sur la zone de recirculation dans la post-chambre, ainsi que sur les tourbillons eux-mêmes. Enfin, plusieurs nouveaux essais du moteur hybride HYCOM ont été effectués et confrontés au modèle développé dans le cadre de la thèse.