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A Rocket Engine Under a Magnifying Glass L A Rocket Engine under a Magnifying Glass L. Vingert, G. Ordonneau, N. Fdida, P. Grenard To cite this version: L. Vingert, G. Ordonneau, N. Fdida, P. Grenard. A Rocket Engine under a Magnifying Glass. Aerospace Lab, Alain Appriou, 2016, pp.15. 10.12762/2016.AL11-15. hal-01369600 HAL Id: hal-01369600 https://hal.archives-ouvertes.fr/hal-01369600 Submitted on 21 Sep 2016 HAL is a multi-disciplinary open access L’archive ouverte pluridisciplinaire HAL, est archive for the deposit and dissemination of sci- destinée au dépôt et à la diffusion de documents entific research documents, whether they are pub- scientifiques de niveau recherche, publiés ou non, lished or not. The documents may come from émanant des établissements d’enseignement et de teaching and research institutions in France or recherche français ou étrangers, des laboratoires abroad, or from public or private research centers. publics ou privés. ARTICLE DE REVUE A Rocket Engine under a Magnifying Glass L. Vingert, G. Ordonneau, N. Fdida, P. Grenard AEROSPACELAB JOURNAL No 11, AL11-15, 13 pages TP 2016-559 Powered by TCPDF (www.tcpdf.org) Challenges in Combustion for Aerospace Propulsion L. Vingert, G. Ordonneau, N. Fdida, P. Grenard A Rocket Engine under (ONERA) E-mail: [email protected] a Magnifying Glass DOI : 10.12762/2016.AL11-15 ven though the technology of cryogenic rocket engines is well mastered today, Eand has been applied successfully in many launchers all over the world, research activities on the various elementary or coupled processes involved in these complex systems are still relevant and useful for future developments, cost reduction, and knowledge improvement. The Mascotte test facility was designed and built with this in mind twenty years ago. Since then, numerous configurations have been tested, enabling almost all of the phe- nomena involved in the operation of a rocket engine to be investigated, in a research lab environment, but nevertheless under representative conditions. Research items addressed on the Mascotte test rig include: injector concepts; liquid oxygen jet atomization and combustion, new propellant combinations (methane instead of hydrogen), ignition, heat transfer at the chamber walls, high frequency instabilities, flow-separation in an over-expanded nozzle, plume and infrared signature. Introduction manufacturers, even more so with reusability, which appears as the main driver for future applications [1]. In addition, in recent years, the Chemical propulsion relies on the principle that energy is stored in propellant combination liquid oxygen/methane (LOX/CH4) has attrac- the chemical reactants and supplied to the system through exother- ted considerable attention in the USA, Europe and Japan for attitude mic reactions. Despite the fact that reactants have a fixed amount of control, upper stage or booster engines. Methane has two advantages energy per unit mass, which limits the achievable exhaust velocity or over hydrogen, which compensates for the slight loss in specific im- specific impulse, and because the propellants are their own energy pulse: its higher specific mass and the proximity of its thermal cha- source, the rate at which energy can be supplied for propulsion is racteristic to those of oxygen, especially its liquefaction temperature. independent from the propellant mass. Thus, very high powers and Both permit cost reduction through simplification of the stage: smaller thrust levels can be achieved. Among the numerous available propel- tank volume and easier to handle cryogenic technology, thanks to the lants, the hydrogen/oxygen combination is the most efficient in terms higher temperature, around 100 K instead of 20. of specific impulse. Even though the development of a device often precedes detailed understanding of the phenomena prevailing there, Up to now, the standard practice in the design of space propulsion for instance liquid oxygen/hydrogen (LOX/H2) was envisaged for J2 systems has mostly relied on accumulated know-how and trial and and RL10 engines at the end of the fifties and used in the Apollo error methodologies, even though computational tools have been program, the best use and optimization of an engine performance progressively introduced in the design process, taking advantage of are possible only if the basic physical phenomena and their coupling the increase in computational power. Nevertheless, numerical codes are well understood and described. These various items require well need to be validated on detailed experimental results, gained under instrumented testing, modeling and research activities. It is the reason well controlled, but as representative as possible, operating condi- why many teams all over the world have worked on this subject for tions. With this objective in mind, ONERA designed the Mascotte test decades, and still do. bench to examine a broad range of processes controlling the com- bustion of cryogenic propellants, such as atomization, droplet vapori- It is indisputable that the technology of cryogenic rocket engines is zation, combustion at high pressure under subcritical and transcritical well known today, and that it has been applied successfully in many conditions, etc. Other topics of major importance in liquid rocket en- launchers (Saturn V, Ariane 1 to 5, Space Shuttle, H-II, etc.), but the gines, like ignition, combustion instabilities, heat transfer and nozzle low cost development of such complex systems, which have to be flow separation, are investigated too. It was initially decided to work increasingly performing and reliable, is still a big challenge for the on single injector configurations fed with cryogenic propellants under Issue 11 - June 2016 - A Rocket Engine under a Magnifying Glass AL11-15 1 representative pressure and temperature conditions, but the more This article describes the structure of cryogenic flames formed by recent studies of ignition and high frequency instabilities were car- injection of liquid oxygen and gaseous hydrogen at pressures excee- ried out on multiple injector configurations comprising up to five units ding the oxygen critical pressure of 5.04 MPa. The data was supple- arranged in a row or as a pack. The first years of Mascotte were mented a few years later with images obtained from PLIF of OH at devoted to the progress, in three directions in parallel. The first one a pressure of 6.3 MPa. This dataset is perhaps the only one today was to increase the operation domain by progressively exploring high that corresponds to laser induced fluorescence imaging at very high pressure conditions and reaching supercritical conditions of the type pressure [39]. prevailing in cryogenic engines. The second point was the develop- ment of advanced and non-intrusive optical diagnostics adapted to The extension of the Mascotte test bench, which took place in a second these extreme conditions. These included high resolution spectros- stage [43] allowed new studies of flames formed by liquid oxygen and copy, backlighting, OH* emission imaging, Planar Laser Induced methane [46]. This has led to some unique findings [37], with in par- Fluorescence (PLIF) of OH radicals, Raman scattering of oxygen and ticular the first detailed analysis of flames formed by liquid oxygen and Coherent Anti-Stokes Raman Spectroscopy (CARS) of H2 and H2O gaseous or liquid methane. The peculiar structure of liquid oxygen/liquid molecules. The third work package was the building of an experimen- methane flames will be described later in this article. Much of the effort tal database through intensive testing. carried out up to the year 2005 was summarized in [5] and in [14]. Fur- ther work on liquid oxygen /methane flames was carried out with laser A number of research projects were carried out in the basic confi- induced fluorescence imaging and emission imaging techniques. These guration of a coaxial injector fed with liquid oxygen and gaseous allowed a detailed analysis of the flame structure up to a pressure of 2 hydrogen. It was first important to examine the flame structure, sta- MPa [38], [40]. In parallel, the CARS technique was adapted to use the bilization process, operating parameter effects, pressure effects, pro- methane as probe molecule [2] and later, a unique coupling between cesses controlling transcritical combustion and geometrical effects two CARS systems sampling simultaneously H2 and H2O [13] was associated with the recess of the liquid oxygen channel. At the start operated within the framework of the European “In Space Propulsion” of this experimental program, it was not known whether the flame program[28]. The problem of flame stabilization was also revisited, by was stabilized aerodynamically at a distance from the injection unit comparing the near field structures of liquid oxygen/gaseous hydrogen or whether the flame was attached to the injector unit or close to that and liquid oxygen/methane flames [41]. This was used to verify a sta- unit. There were no data, at least in the open literature, that provided bility criterion derived in [18]. On a more practical level, a number of this information. While most experts believed that the flame was for- injector design issues have been considered. This is the subject of refe- med at a distance from the injector unit because of the very large rences [6] and [9] and will be used later on as an example of application velocities characterizing the hydrogen stream, the data gathered in the of the test bench. first experiments at low pressure (1, 5 and 10 bar) clearly indicated that the anchor point
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