This paper presents the first numerical evidence of an intermittency route to period-2 thermoacoustic instability in a subcritical single-element liquid rocket engine burning hydrogen peroxide/kerosene as we decrease the equivalence ratio (ϕ) from fuel-rich to fuel-lean. To achieve this, three-dimensional compressible large eddy simulation algorithms combined with the Euler–Lagrangian framework are used. A one-equation eddy sub-grid turbulence model with a partially stirred reactor sub-grid combustion model is employed to simulate the spray turbulent combustion process in a high-pressure liquid-fueled combustor based on open-source platform OpenFOAM. This paper focuses on examining the transition process of the dynamical states in the thermoacoustic system and the synchronization between multiple subsystems. The results indicate that, as the equivalence ratio reduces continuously (1.5 ≤ ϕ ≤ 0.5), the system dynamics shift from period-1 oscillations (ϕ = 1.5) to period-2 oscillations (ϕ = 0.5) via intermittency (1.3 ≤ ϕ ≤ 0.9). Under the equivalence ratio of 0.7 (ϕ = 0.7), a transient mode switching between period-1 and period-2 was also observed. The synchronization processes between the pressure and combustion subsystems in terms of phase-locking and frequency-locking are responsible for the emergence of complex dynamical states. The cycle snapshots analysis also provides more details on the synchronization processes between the pressure and the multiple subsystems, such as vortex dynamics, mixture fraction, and combustion heat release. In summary, this paper sheds light on the complex non-linear thermoacoustic oscillations and the underlying physical mechanisms related to the two-phase flow of spray combustion in liquid rocket engines using three-dimensional large eddy simulations, paving the way for developing passive or active control methods.

The recent growth of private options in launch vehicles has substantially raised price competition in the space launch market. This has increased the need to deliver reliable launch vehicles at reduced engine development cost, and has led to increased industrial interest in reduced order models. Large-scale liquid rocket engines require high-speed turbopumps to inject cryogenic propellants into the combustion chamber. These pumps can experience cavitation instabilities even when operating near design conditions. Of particular concern is rotating cavitation, which is characterized by an asymmetric cavity rotating at the pump inlet, which can cause severe vibration, breaking of the pump and loss of the mission. Despite much work in the field, there are limited guidelines to avoid rotating cavitation during design and its occurrence is often assessed through costly experimental testing. This paper presents a source term based model for stability assessment of rocket engine turbopumps. The approach utilizes mass and momentum source terms to model cavities and hydrodynamic blockage in inviscid, single-phase numerical calculations, reducing the computational cost of the calculations by an order of magnitude compared to traditional numerical methods. Comparison of the results from the model with experiments and high-fidelity calculations indicates agreement of the head coefficient and cavity blockage within 0.26% and 5% respectively. The computations capture rotating cavitation in a 2D inducer at the expected flow coefficient and cavitation number. The mechanism of formation and propagation of the instability is correctly reproduced.

This study presents a dynamic simulation framework to optimize thrust regulation in an electric pump-fed rocket engine using liquid oxygen-liquid methane propellants. The engine operates across a 20-100% thrust range. A comprehensive system model simulates transient interactions among electric pumps and injectors. It focuses on actuation timing effects on thrust control efficiency. The model integrates pump dynamics, combustion, and two-phase flow. It uses a time-stepped numerical method. Results show that synchronized pump actuation minimizes regulation time. Delays extend it. The oxidizer pump predominantly influences chamber pressure and mixture ratio due to its faster response. Pre-actuation of components modulates performance. Oxidizer pump adjustments reduce mixture ratio during thrust decrease. Injector pre-actuation elevates pressure ratios. An optimized timing sequence mitigates fluctuations in mixture ratio and injector pressure during deep-thrust transitions. Validation against experimental data confirms accuracy. Errors are below 1.2% for chamber pressure and mass flow rate. This methodology elucidates electric pump-fed engine dynamics. It offers a scalable tool for refining thrust regulation. It advances propulsion system design through computational modeling. 1579 scirep Scientific Reports Sci Rep Nature Publishing Group PMC12474916 12474916 12474916 41006603 10.1038/s41598-025-18499-5 A dynamic simulation approach to optimize thrust regulation in electric pump-fed rocket engines Li Tianwen 1 He Haodong 2 Yu Nanjia 1 3 ✉ Sun Xixiang 1 Cai Guobiao 1 1 School of Astronautics, Beihang University, Beijing, 100191 China 2 Lanzhou Institute of Physics, China Academy of Space Technology(CAST), Lanzhoou, 730000 China 3 National Key Laboratory of Aerospace Liquid Propulsion, Xi’an, 710100 China ✉ Corresponding author. The images or other third party material in this article are included in the article’s Creative Commons licence, unless indicated otherwise in a credit line to the material. If material is not included in the article’s Creative Commons licence and your intended use is not permitted by statutory regulation or exceeds the permitted use, you will need to obtain permission directly from the copyright holder. To view a copy of this licence, visit http://creativecommons.org/licenses/by-nc-nd/4.0/ . Abstract This study presents a dynamic simulation framework to optimize thrust regulation in an electric pump-fed rocket engine using liquid oxygen-liquid methane propellants. Keywords: Dynamic simulation, Thrust regulation, Timing optimization, Propulsion modeling, LO X -LCH 4 engine Subject terms: Energy science and technology, Engineering, Mathematics and computing, Physics status released display-pdf yes is-olf no is-manuscript no is-preprint no is-journal-matter no is-scanned no is-retracted no Received 2025 Jul 8; Accepted 2025 Sep 2; Collection date 2025. Introduction Liquid rocket engines are currently the mainstream propulsion systems in aerospace. They are known for their reusability and adjustable thrust 1 , 2 . Previous studies of EP rocket engines focused on scheme demonstration and mass estimation. Scheme demonstration began in the aerospace field last century 5 . Advancements in technology have sparked numerous studies. These include miniaturized high-energy power supplies, accurate propellant control, and light high-speed motors. They focus on mass estimation of the EP system 6 . The EP system has potential application value under conditions of high thrust chamber pressure and long working time 7 , 8 . Lee et al. concluded that the EP system is lighter than the gas generator system in small liquid rocket engines 9 . It serves as the pressurization component of liquid rocket engines. This paper uses the Suter curve to comprehensively characterize the performance of the centrifugal pump during various operating conditions. The pressure head function and torque function determine the operational state of the centrifugal pump. They are expressed by characteristic angle and specific speed . These functions and variables play a crucial role in evaluating the pump’s performance under actual working conditions. Their definitions are outlined in Wan’s paper 18 . This study adopts a three-phase PMSM that combines the characteristics of permanent magnet motors and synchronous motors, leading to high power density, efficiency, and power factor 19 . Thanks to its ability to maintain high efficiency and power factor within 25% -120% of the rated load, PMSM is suitable for the variable operating conditions of liquid-propellant rocket engines. Figure 3 depicts the primary components of PMSM, which consist of a rotor core, permanent magnets, and stator winding coils. In PMSM, the fixed coordinate system is usually transformed into the rotating coordinate system. This differentiated approach—compressible for injectors influenced by cooling channels (where boiling may induce density variations) and incompressible for valves—is justified by the engine’s architecture, where valve flows remain subsonic and single-phase under operational conditions. Validation against experimental data confirms the approximation’s validity, with errors in mass flow rates below 0.95%, consistent with practices in liquid rocket engine modeling 29 . Based on the assumption of It used gaseous hydrogen and gaseous oxygen torch igniters. As shown in Fig. 7 a, the fire test system was consistent with the thrust chamber component model studied in this paper. The combustion performance of the thrust chamber component was tested. It also examined the coordination between components such as combustion chamber, nozzle, valves, and injectors. The chamber pressure ( ) was regarded as an important parameter for describing its combustion performance. It characterizes its dynamic characteristics. Figure 7 b compares the variation of over time obtained from the rocket engine fire test with its system simulation.

Simulations on Optimization of Liquid Spray Burners & Operating Parameters <p>Spray burners form an essential part of any liquid propulsion system as they are responsible for injecting, atomizing, mixing and combusting the liquid fuel. Spray combustion used in aerospace applications like the liquid rocket engines, gas turbines or any other controlled environment for that matter places a huge emphasis on safe and effective operations. These applications make use of relatively small amounts of propellant volumes to generate enormous amounts of energy through combustion for producing thrust. For such cases involving enormous energy interactions, combustion comes with its own set of challenges. The predominant challenge among them all is that of combustion instability. Instabilities are physical phenomena occurring in both reacting and non-reacting flows. In the context of combustion, they tend to emanate from very small perturbations in the flow field and amplify in magnitude to alter the macroscopic properties of the flow causing problems for the smooth functioning of engines.

The rocket-based combined-cycle (RBCC) engine is regarded as one of the most viable propulsion systems for single-stage-to-orbit launch vehicles. Because of the relatively low total temperature of the incoming flow, it is difficult to maintain sustained and efficient subsonic combustion when the rocket engine is turned off. Mode transition and its control have also become critical techniques in the RBCC study. In the current work, it is proposed for the first time to improve the performance of RBCC engines in mode transition by using plasma combustion support. The numerical simulation and validation were conducted on the full path configuration of the RBCC engine, which is suitable for wide range operation. The impact of multi-channel gliding arc (MCGA) plasma-assisted combustion technology on the flow field was investigated during the transition phases from RBCC ejector/ramjet mode to ramjet/scramjet mode. The results show that: Adding arc plasma into the cavities at low Mach numbers promotes the heat release of the fuel, expanding the high-temperature flame range in the combustor. Although it weakens the ability of some air inlet to capture air, it overall increases engine thrust and diminishes aerodynamic drag during the transition process from ejector/ramjet mode. At high Mach numbers, the fast incoming flow speed results in the inability of the fuel to mix and burn with air in a timely manner, resulting in poor heat release and work performance. However, the addition of arc plasma for combustion support using the same mode transition method reduces the thrust fluctuation of the RBCC engine during the ramjet/scramjet mode transition process, greatly reducing the time required for mode transition, reducing engine resistance, and significantly improving engine thrust. Because of the relatively low total temperature of the incoming flow, it is difficult to maintain sustained and efficient subsonic combustion when the rocket engine is turned off. Mode transition and its control have also become critical techniques in the RBCC study. In the current work, it is proposed for the first time to improve the performance of RBCC engines in mode transition by using plasma combustion support. The numerical simulation and validation were conducted on the full path configuration of the RBCC engine, which is suitable for wide range operation. Keywords: Mode transition, Plasma assisted combustion, Combustion enhancement, Transient simulation Subject terms: Energy science and technology, Engineering status released display-pdf yes is-olf no is-manuscript no is-preprint no is-journal-matter no is-scanned no is-retracted no Received 2024 Jul 20; Accepted 2024 Nov 25; Collection date 2024. Introduction The rocket-based combined-cycle (RBCC) engine is acknowledged as a highly promising propulsion system for single-stage-to-orbit launch vehicles 1 . 24 , 25 also conducted direct connection experiments in a scramjet model engine, and two combustion modes were obtained by changing the fuel injection position. MCGA plasma is another way to increase combustion intensity, it is evolved from the gliding arc plasma discharge technique 26 . MCGA plasma can generate a large amount of heat and active substances to enhance the combustion intensity in the combustor 27 , 28 . A. Klimov et al. 29 conducted a dual power plasma combustion experiment. The results implicate that plasma combustion can accelerate the fuel transportation speed and enhance the stability of fuel combustion. Rodney Meyer et al. For RBCC engines, mode transition is a sudden change in flow path pressure, posing a huge challenge to engine thrust and overall aircraft control. Many physical phenomena and parameters, such as aerodynamics, thermodynamics, chemical reaction, fuel and oxidant characteristics, need to be considered in the numerical simulation of smooth transition and combustion characteristics of RBCC mode transition process. In previous studies, changes in fuel heat release position, rocket mass flow rate, and engine configuration were commonly used to achieve a smooth transition of mode transition. It can be seen that although the mass flow rate of the fuel has been significantly reduced, there is only a slight change in However, in Case 4 with arc plasma addition, a small amount of CO 2 is still retained in the isolator and the recirculation area behind the rocket strut. The overall reduction of the products in the central flow of the combustor is small. Therefore, after adding plasma for combustion in the same period of time, the RBCC combustion intensity decreases less, and the engine performance is more stable. Fig. 22 CO 2 distribution contour on the plane of y = 1 H is in Case 3 (upper half) and Case 4 (lower half). The subsonic region near the fuel inlet in the isolator indicates that RBCC can still maintain small-scale combustion in the upstream part of the engine after shutting down the rocket and reducing the fuel equivalence ratio, which undoubtedly expands the heat release distribution area of the fuel. Figure 26 illustrates the thrust profiles during the RBCC mode transition for Case 3 (kerosene equivalent ratio reduced from 1 to 0.4, rocket total mass flow rate from 0.2 kg/s to 0 kg/s, without plasma) and Case 4 (kerosene equivalent ratio reduced from 1 to 0.4, rocket total mass flow rate from 0.2 kg/s to 0 kg/s, with arc plasma added to the cavities). Fig. Conclusion This research proposes to apply arc plasma to assist combustion in the mode transition process in RBCC engine. The detailed flow field and flame structure are obtained by numerical simulation. The following conclusions can be drawn: The working state of the air inlet at low Mach numbers has a close impact on the combustion situation in the combustor. When the Mach number is 3.0, the inlet achieves a transition from non start to start during the ejector/ramjet mode transition process by shutting down the rocket and reducing the fuel equivalence ratio.

Investigation of the cooling film distribution in liquid rocket engine This study presents the results of the investigation of a cooling method widely used in the combustion chambers, which is called cooling film, and it is applied to a liquid rocket engine that uses as propellants liquid oxygen and kerosene. Starting from an engine cooling, whose film is formed through the fuel spray guns positioned on the periphery of the injection system, the film was experimentally examined, it is formed by liquid that seeped through the inner wall of the combustion chamber. The parameter used for validation and refinement of the theoretical penetration of the film was cooling, as this parameter is of paramount importance to obtain an efficient thermal protection inside the combustion chamber. Cold tests confirmed a penetrating cold enough cooling of the film for the length of the combustion chamber of the studied engine. Published in Journal of Aerospace Technology and Management

Simulation aux grandes échelles multi-physiques de l'oxy-combustion du méthane dans les moteurs-fusée La combustion dans les moteurs de fusée a lieu dans des conditions extrêmes qui impliquent plusieurs phénomènes multi-physiques. Pour cette raison la simulation numérique est utilisée afin de prédire et ainsi optimiser les performances et la durée de vie du moteur. Ces travaux de thèse se concentrent particulièrement sur deux aspects : l’oxy-combustion turbulente du méthane dans des flammes de diffusion haute pression et la prédiction des transferts thermiques pariétaux. Le code de Simulation aux Grandes Echelles (SGE) AVBP du CERFACS est utilisé. Malgré ses performances moindres, le méthane est aujourd’hui favorisé par rapport à l’hydrogène pour les futurs moteurs-fusée en raison de ses coûts réduits et de sa praticité tant à l’utilisation que pour son stockage. En termes de simulation numériques, cet ergol amène de nouvelles questions concernant son allumage ou la stabilisation de sa combustion. Pour ce faire, le développement de modèles chimiques réalistes est une étape clé. Des schémas cinétiques réduits contenant une quinzaine d’espèces sont dérivés et testés dans des conditions de haute pression et haut étirement pour des flammes de diffusion à contre-courant. Cependant, ces chimies réduites restent coûteuses pour une utilisation industrielle de la SGE. Une nouvelle méthode d’intégration des termes sources chimiques est alors proposée avec pour objectif de faire fonctionner des simulations réactives avec un pas de temps proche du pas de temps de l’écoulement. Le coût de calcul est ainsi considérablement réduit, tout en gardant un résultat similaire à l’intégration classique. Enfin, avec pour futur objectif le développement d’un modèle de flamme de diffusion turbulente, une étude montrant l’impact de la résolution du maillage sur les flammes de diffusion est réalisée. Le développement de ces chimies réduites permet alors d’étudier avec précision l’influence des réactions chimiques dans la région proche paroi d’un moteur-fusée sur le flux thermique pariétal. Des canaux turbulents périodiques sont simulés afin de comparer une couche limite turbulente résolue ou non, avec ou sans réactions chimiques. Les résultats montrent que ces réactions proche paroi peuvent avoir un impact significatif sur le flux thermique, et que le modèle de loi de paroi devrait prendre en compte cet effet dans le cadre d’une SGE non résolue en paroi. De plus une étude est menée afin de comprendre l’impact du couplage entre le modèle de sous-maille et la loi de paroi sur la prédiction des flux en paroi. On montre que le niveau de viscosité turbulente en paroi a une influence importante sur les flux. Un modèle stochastique est alors proposé dans le cas de simulations isothermes afin d’améliorer les résultats pour deux modèles de sous-maille usuels en SGE, WALE et Sigma. Le développement de ces modèles et les résultats de ces analyses sont alors utilisés pour deux SGE de bancs d’essai : le cas supercritique à cinq injecteurs GCH4/GOx de l’ONERA et le cas sous-critique mono-injecteur GCH4/LOx de TUM. Ces études se concentrent en particulier sur le comportement de la flamme et la comparaison du flux de chaleur pariétal avec les résultats expérimentaux.

The information included here attempts to present a coherent summary of experimental achievements pertinent to liquid rockets, focusing only on the injection of nonreacting cryogenic liquids into a high-pressure environment surpassing the critical point of at least one of the propellants. Moreover, some implications of the results acquired under such an environment are offered in the context of the liquid rocket combustion instability problem. Published in International Journal of Aerospace Engineering

Rocket propulsion is of vital importance for space travel. New innovations are continuously developed in order to facilitate the demand of the rapidly evolving space sector. Recently a focus on reusable rockets has appeared due to the economical and environmental benefits they bring. When designing reusable launch vehicles the propellant injector becomes very important since it is a critical component when throttleabilty is desired. Which is a key element of landable rockets. Selecting an appropriate injector type therefore becomes crucial, a common injector type used for throttleable rockets is the pintle injector. Unfortunately the design process of the pintle injector is complicated due to the large amount of variables that must be determined. This thesis aims to solve this problem by developing a numerical method to design and optimise a pintle injector and then produce a preliminary design. The numerical method developed in this thesis is used to produce a preliminary design of a pintle injector designed to utilise a combination of liquid oxygen and gaseous methane, theoretically capable of a max thrust of 1000N and a throttleabilty of 5 to 1. The design had a focus on optimising the performance of the parameters sauter mean diameter, vaporisation distance and spray angle for the injector. The resulting injector showcases great performance and is deemed to show a successful preliminary design. Which shows that the numerical design and optimisation process that was develo

In order to develop a new generation of intelligent satellites, fast-response bipropellant thrusters are required to work in minimum impulse mode without limitation. When a valve is opening, the fluctuation affects downstream spray atomization at the injector, which determines the thruster's impulse performance, involving combustion efficiency and impulse repeatability. Accordingly, the spray atomization under impulse working condition was investigated to optimize the thruster's dynamic response. The effects of propellant property, switch speed, valve stroke, and throttle orifice layout are respectively compared in simulation cases using OpenFOAM. The fluctuating flowrate caused by valve opening was simulated and then used as boundary conditions for downstream spray. Among these factors, orifice layout plays the most significant roles in transient spray development. Compared with MMH spray, NTO spray from outer swirl injector is more sensitive to upstream fluctuation. When the upstream flowrate stabilizes faster, the atomization stability can also be enhanced, thereby improving the impulse repeatability of thrusters in combustion. This experimental result was in good agreement with the simulation, thereby showing that only when atomization of MMH spray and NTO spray both develop into a steady state within 5 ms after valve opening can the impulse performance be reliably achieved. When the upstream flowrate stabilizes faster, the atomization stability can also be enhanced, thereby improving the impulse repeatability of thrusters in combustion. This experimental result was in good agreement with the simulation, thereby showing that only when atomization of MMH spray and NTO spray both develop into a steady state within 5 ms after valve opening can the impulse performance be reliably achieved. valve opening process swirl injector dynamic spray atomization bipropellant thruster pmc-status-qastatus 0 pmc-status-live yes pmc-status-embargo no pmc-status-released yes pmc-prop-open-access yes pmc-prop-olf no pmc-prop-manuscript no pmc-prop-legally-suppressed no pmc-prop-has-pdf yes pmc-prop-has-supplement no pmc-prop-pdf-only no pmc-prop-suppress-copyright no pmc-prop-is-real-version no pmc-prop-is-scanned-article no pmc-prop-preprint no pmc-prop-in-epmc yes pmc-license-ref CC BY 1. Introduction The bipropellant thruster is a low-thrust liquid rocket engine used to stabilize or change the attitude of spacecrafts. Traditionally, multiple firing-test runs are needed for validation of the injector design at a high cost. However, spray simulation analysis uses a large amount of data and information, aiming for failure prevention and design optimization. Many researchers have been experimentally and numerically conducting work to study transient swirl spray [ 2 , 3 , 4 ]. Keller used phase Doppler anemometry and conducted Eulerian–Lagrangian computations to investigate transient spray field phenomena on droplet dynamics and dispersion of an isothermal flow [ 5 ]. Snyder used 2D Mie scattering images to test the transient cone angle of pressure swirl sprays from injectors intended for use in gasoline direct injection engines [ 6 ]. Baldwin ran a computational study to investigate the influence of transient needle motion on gasoline direct injection (GDI) internal nozzle flow and near-field sprays [ 7 ]. Yang analyzed the overall transient spray impingement structure and fuel film formation with different swirl ratios using the CONVERGE CFD code during the intake stroke [ 8 ]. It was concluded that increasing the ambient pressure increased the size of droplets in all of the spray regions, with decreased secondary atomization. Zhang et al. researched the overall process of double swirl atomization using the methods of large eddy simulation and volume of fluid [ 13 ]. In the simulation, they observed two wave patterns in the development of spray, with variable droplet size distribution. In previous research, the influence of opening characteristics of the solenoid valve on flow fluctuation and transient spray has not been studied. A K-H instability wave is generated on the surface of the liquid film due to ambient gas disturbance and the boundary layer effect. What needs to be emphasized is that there is more obvious surface fluctuation just from the nozzle exit than the spray of constant flow. When the surface wave develops to a certain extent with higher amplitude over the film thickness, the liquid film begins to break up into ligaments and droplets. After the first atomization, the liquid ligaments and droplets still move with a high velocity, and continue to split under the aerodynamic interaction, resulting in secondary atomization [ 17 ]. Therefore, the swirl injection with lower velocity significantly weakens the aerodynamic interaction, which causes the decrease of atomization quality, especially for NTO spray with the structural deficiency shown in Figure 14 . Under the condition of throttle at upstream, the spray angle gets smaller, droplets get bigger, and breakup length gets longer, showing negative effects on the spray and the following combustion. Therefore, throttle orifice layout at upstream is not recommended for fast-response bipropellant thrusters. 3.5. Low-speed switch seems to slightly strengthen the breakup under strong fluctuation, but the resulting instability and inconsistency are negative for the following combustion. In Figure 15 b, as for the case of NTO spray, it is sensitive to the change of different orifice throttle layouts. Using the downstream orifice layout can increase the surface area of spray by 2–3 times. Meanwhile, valve stroke and switch speed also have a certain effect on the spray development, but there is no unidirectional tendency. Regarding the above analysis, 1 m/s switch speed, 0.15 mm valve stroke, and downstream throttle layout should be a better combination for optimization.