Emeric Daniau

Pulsed and Rotating Detonation Propulsion Systems : First Step Toward Operational Engines

Due to its thermodynamic cycle, a detonation wave engine has theoretically a higher performance than an other classical propulsion concept using the combustion process. Nevertheless, it still has to be proved that this advantage is not compensated by the difficulties which could be encountered to practically define a real engine and to implement it in an operational flying system. For space launcher, the application of detonation wave propulsion to the liquid rocket engine could strongly increase the specific impulse and improve the feasibility of a full rocket SSTO. But, obtaining this result probably needs to radically change the concept of the vehicle in order to take all the advantage of the detonation wave process but also to limit the technology needs and to mitigate the dry mass increase, due to the generated environmental conditions. For military application, and for Europe, it seems that the application to an aircraft can not be considered in the near or middle term future. On the contrary, the airbreathing detonation wave engine could be very attractive for low cost missiles and UAV. Indeed, for these applications, a detonation wave engine with simplified technology could provide an economic alternative to the turbo-engine, particularly in the field of very small caliber turbo-engine. If PDE seems promising in air- breathing mode, for rocket engine the use of rotating detonation wave (CDWRE) can also be considered to reduce the environmental conditions generated by PDE while reducing the importance of initiation issue and simplifying some integration aspects. Such a concept has been studied since a long time, particularly at Lavrentiev Laboratory in Novosibirsk. As it was done for PDE, specific experimental program is being performed by MBDA and Lavrentiev Institute to assess some key points for the feasibility of an operational rotating wave engine for space launcher : detonation of two-phase mixture, limitation of injection pressure, feasibility of thrust vectoring… and the main results of this experimental program will be summarized in the paper. On the basis of these results, a preliminary design of an operational engine is under progress taking into account all engine/airframe integration issues in order to optimize the benefit of detonation wave engine. The main options of this design will be presented and discussed in the paper before a global performance evaluation.

A Contribution to the Development of Actual Continuous Detonation Wave Engine

During past years, MBDA performed some theoretical and experimental works, mainly in cooperation with LCD laboratory at ENSMA Poitiers, on Pulsed Detonation Engine (PDE). These studies aimed at obtaining a preliminary demonstration of the feasibility of the PDE in both rocket and airbreathing modes and at verifying the interest of such a PDE for operational application : rocket and airbreathing mode experimental evaluation, effect of filling coefficient, effect of a nozzle, thermal, mechanical, acoustic and vibrations environment generated, evaluation of different fuels, performance code development. Due to its thermodynamic cycle, the pulsed detonation engine (PDE) has theoretically a higher performance than an other classical propulsion concept using the combustion process (+ 20 to 25% in term of thermal efficiency). Nevertheless, it is necessary to verify that this advantage is not fully compensated by the difficulties, which could be encountered for practical use of the PDE concept or by the complex technology, which could be needed to implement it in an operational flying system. Moreover, a PDE a priori generates a severe vibration environment, which can imply higher more severe requirement for all on-board vehicle equipments or subsystems. After some in house studies performed in national and international cooperation, MBDA is now focusing its efforts to the development of a demonstration engine led in cooperation with Singapourian DSO in order to really assess the feasibility and the interest of this propulsion concept.

Semi-empirical and CFD analysis of actively cooled dual-mode ramjets : 2006 status

Dual-mode ramjets could thrust future hypersonic vehicles such as Reusable Space Launchers, and in particular SSTO. If the propulsive performance have to be computed, optimised and if possible demonstrated, a major challenge is the capacity to build such an engine, and to estimate its robustness and its weight. System studies, technologies demonstration and computational techniques have been under concern at MBDA France from conventional ramjets to hypersonic engines. Aerodynamical, chemical, mechanical and thermal investigations of composite and metallic structures have been carried out, using both analytical tools and 3-D numerical simulations.

SFGP 2007 - Pyrolysis of Supercritical Endothermic Fuel: Evaluation for Active Cooling Instrumentation

Hypersonic flight is expected to be achieved in the coming years by use of Supersonic Combustion RAMJET (SCRAMJET). One of the main issues is the thermal management of the overall vehicle and more specifically the cooling of the engine. In order to simulate the behaviour of an actively cooled SCRAMJET by use of supercritical endothermic fuel, a one-dimensional transient numerical model has been developed with heat and mass transfer, fluid mechanics and detailed pyrolysis chemistry. A dedicated experimental test bench is now available since 2006 at the LEES laboratory of Bourges to study supercritical fuel pyrolysis under steady-state and transient conditions. It aims to provide understanding of coupled phenomena, validation data for the numerical code and evaluation of onboard and real-time measurement methods for industrial use. A brief overview of the numerical code and a presentation of the experimental bench are proposed in this paper. Experimental results are discussed and a comparison is provided between numerical and experimental data. Discrepancies are shown to be lower than a few percent in terms of molar chemical compositions. This is due to uncertainties on experimental temperature measurement and to 2-D effects, which are not taken into account by the modelling. The numerical code appears to be of great importance in accessing unmeasured data and providing new understanding of coupled phenomena. Experimental and numerical tools are proved to be efficient to test future measurement methods under extreme conditions, especially at supercritical states.

Design of a Continuous Detonation Wave Engine for Space Application

*† ‡ § ** The last years showed a renewed interest in space exploration with the announcement of new and ambitious space programs. The prospect of the near term return of manned missions to the moon or the development of clusters of small-size satellites imply that we need simple and affordable access to space. Detonation wave engines, thanks to their more efficient thermodynamic properties, are expected to exhibit a higher level of performance than more conventional propulsion system that rely on constant-pressure combustion processes. Nevertheless, it still has to be proved that this advantage is not superseded by the difficulties which could be encountered to practically define a real engine and to implement it in an operational flying system. Application of detonation wave propulsion to the liquid rocket engine could increase the specific impulse and improve the feasibility of a full rocket SSTO, but obtaining this result probably needs to radically change the concept of the vehicle in order to take all the advantage of the detonation wave process. During past years, MBDA performed some theoretical and experimental works, mainly in cooperation with the Lavrentyev Institute of Hydrodynamics in Novosibirsk. These studies aimed at obtaining a preliminary demonstration of the feasibility of a Continuous Detonation Wave Engine (CDWE) for air-breathing and rocket application. Compared to a Pulsed Detonation Engine, this design allows an easier operation in reducedpressure environment and an increase in engine mass flow rate and thrust-toweight ratio. Those studies were focused on global performance and understanding of the unsteady, three dimensional flow behind the detonation wave. Investigations were conducted to gain knowledge of the noise generated by a CDWE operating at several kiloHertz, heat fluxes (intensity, areas) and cooling strategies, composite materials (Carbon / Silicon Carbide) compatibility, engine thrust vectoring capability and pollution (for air-breathing application). Those results were used to validate 1-D and 2-D unsteady computations. Discrepancies between numerical results and analytical flow-solutions were found that highlight the difficulties of detonation waves simulations. On the basis of these results, a preliminary design of an operational engine is under progress taking into account all engine/airframe integration issues in order to optimize the benefit of detonation wave engine. The main options of this design will be presented and discussed in the paper before a global performance evaluation

Hydrocarbons Heterogeneous Pyrolysis: Experiments and Modeling for Scramjet Thermal Management

The last years saw a renewal of interest for hypersonic research in general and regenerative cooling specifically, with a large increase of the number of dedicated facilities and technical studies. In order to quantify the heat transfer in the cooled structures and the composition of the cracked fuel entering the combustor, an accurate model of the thermal decomposition of the fuel is required. This model should be able to predict the fuel chemical composition and physical properties for a broad range of pressures, temperatures and cooling geometries. For this purpose, an experimental and modeling study of the thermal decomposition of generic molecules (long-chain or polycyclic alkanes) that could be good surrogates of real fuels, has been started at the DCPR laboratory located in Nancy (France). This successful effort leads to several versions of a complete kinetic model. These models do not assume any effect from the material that constitutes the cooling channel. A specific experimental study was performed with two different types of steel (regular: E37, stainless: 316L). Some results are given in the present paper.

First steps for the development and testing of a Pulse Detonation Engine for UAV application

Due to its thermodynamic cycle, the pulsed detonation engine (PDE) has theoretically a higher performance than an other classical propulsion concept using the combustion process (+ 20 to 25% in term of thermal efficiency). Nevertheless, it is necessary to verify that this advantage is not fully compensated by the difficulties, which could be encountered for practical use of the PDE concept or by the complex technology, which could be needed to implement it in an operational flying system. Moreover, a PDE a priori generates a severe vibration environment, which can imply higher more severe requirement for all on-board vehicle equipments or subsystems. On the basis of knowledge acquired during the last ten years, MBDA and DSO started a collaboration devoted to the pre-development of a small-scale PDE demonstrator that could be flight tested within the next years. This demonstrator should use storable fuel, be throttle-able, provides a good specific impulse in a small-size engine with a good thrust-to-weight ratio and minimum maintenance cost. This demonstrator should also be able to operate in a complete airbreathing mode, without onboard oxidizer, so the design of the ignition device is one of the key points of this engine, the other being the inlet integration and the capability to provide continuous inlet flow even with a pulsed combustion chamber operation. Since the repetitive and direct initiation of a detonation inside a PDE requires large amount of energy and hence high power consumption, two other low-energy mechanisms are generally used, they are deflagration-to-detonation transition (DDT) and Shock-to-detonation transition (SDT). Pre-detonation tubes with Shchelkin spirals are commonly found on most PDE concept. The mechanism involved is both DDT and SDT. Inside the pre-detonation tube, the flame accelerates from low-velocity laminar (or turbulent in some cases) to medium velocity turbulent regime, up to the thermal blockage velocity, and ultimately up to real CJ detonation velocity. For fuel – air mixture, the pre-detonation tube diameter is generally marginally larger than one detonation cellsize, so when the detonation diffracts into the main chamber, the sudden expansion from confinement results in decay of the reaction zone and decoupling from the shock wave in the subcritical regime (complete propagation failure as shock decouples from the reaction zone). In this case, re-ignition of the detonation could only be achieved by shock – shock interactions (main shock and reflected shock waves from the walls), so after the DDT inside the pre-detonation tube we could observe a SDT inside the main chamber. During the detonation phase, the pressure inside the chamber is higher than the total inlet pressure so without careful design some parts of the hot gases could be exhausted thru the inlet, decreasing both impulse from the detonation and engine operating frequency. The paper describes the numerical and experimental investigations on shock waves generated by fast flames as well as some work related to engine integration, performed by DSO and MBDA, in order to develop a first version of an operational ignition device for the envisioned PDE.

Numerical and experimental validation of transient modelling for Scramjet active cooling with supercritical endothermic fuel

One of the main issues of hypersonic flight is the thermal management of the overall vehicle and more specifically the cooling of the engine. In order to simulate the behaviour of a complete actively cooled scramjet, a one-dimensional transient numerical model has been developed with heat and mass transfer in a cooling channel for supercritical fuel under pyrolysis. This model is called RESPIRE (French acronym for Scramjet Cooling with Endothermic Fuel., Transient Reactor Programming). A supplementary step by step validation of the model, based on 2-D numerical data from CFD-ACE, is presented in this paper. On stationary cases, fluid temperature profiles are in good agreement and values are comprised between those of centre fluid and those of near-wall fluid. Effects of one-dimensional semi-empirical correlations are shown and the boundary layer impact at the channel entrance is of great importance on wall temperatures. Heat fluxes conservation is verified and hydraulic behaviour too. On transient cases, temperature and velocity evolutions are well followed. Values after the change in mass flow rate are exactly those of stationary test case. RESPIRE is quantitatively validated under stationary and transient conditions and can be used to compare with experimental data. Qualitatively good agreement is found with experimental results on a chemical aspect.

PRELIMINARY WORK FOR A PULSED DETONATION ENGINE DEMONSTRATOR

§** †† ‡‡ , Due to its thermodynamic cycle, a detonation wave engine has theoretically a higher performance than other classical propulsion concept using the combustion process. Nevertheless, it still has to be proved that this advantage is not compensated by the difficulties which could be encountered to practically define a real engine and to implement it in an operational flying system. For military application, and for Europe, it seems that the application to an aircraft can not be considered in the near or middle term future. On the contrary, the airbreathing detonation wave engine could be very attractive for low cost missiles and UAV. Indeed, for these applications, a detonation wave engine with simplified technology could provide an economic alternative to the turbo-engine, particularly in the field of very small caliber turbo-engine. During past years, MBDAFrance and DSO (Singapore) started a common effort on pulse detonation engines, focusing on air-breathing application and engine operation, the final objective being the realization and testing of a complete engine demonstrator. This paper will focus on two key points regarding air-breathing PDE, the inlet integration and the ignition of the main chamber.

Numerical Simulations and Experimental Results of Endothermic Fuel Reforming for Scramjet Cooling Application

The last years saw a renewal of interest for hypersonic research in general and regenerative cooling specifically, with a large increase of the number of dedicated facilities and technical studies. In order to quantify the heat transfer in the cooled structures and the composition of the cracked fuel entering the combustor, an accurate model of the thermal decomposition of the fuel is required. This model should be able to predict the fuel chemical composition and physical properties for a broad range of pressure, temperature and cooling geometry. For this purpose, an experimental and modeling study of the thermal decomposition of generic molecules (long-chain or polycyclic alkanes) that could be good surrogates of real fuels, has been started at the DCPR laboratory located in Nancy (France). This successful effort leads to several version of a complete kinetic model. The last version for n-dodecane will be described in this paper and now include a detailed mechanism for Poly-Aromatics Hydrocarbons synthesis. Validation have been done for a wide range of temperature and pressure (6501500 K and 1-100 bar) and good agreement was generally found between the numerical model and available experimental results. In addition to this model, other molecules have been studied to investigate the differences between linear alkanes and cyclanes hydrocarbons. The unimolecular decomposition of cyclic hydrocarbons involves either the breaking of a C-H bond, which has a large bond energy, or the opening of the ring and the formation of a diradical. The behavior of the diradicals is very different from the one of linear radicals and should be investigated in order to gain a good knowledge of the thermal stability of such cyclanes.