Search Results (19)

This study presents a novel one-dimensional solution method to demonstrate the effects of fuel composition and channel roughness on phase-change heat transfer in spiral regenerative cooling systems. The calculated models are grounded in an experimental correlation of liquefied natural gas (LNG) flow boiling, and their accuracy is validated through ignition experiments conducted on a 1 kg/s-class thrust chamber. The experimental data shows that the physical characteristics of LNG contribute to an extended reach within the two-phase region, resulting in a calculated pressure drop that exceeds that of pure liquid methane. Variations in surface roughness influence the pressure drop by altering the frictional coefficient. Specifically, an increase in surface roughness from 2 µm to 8 µm results in a 47.8% rise in pressure drop. The proposed model demonstrates high accuracy, with deviations in the coolant temperature rise and the pressure drop being less than 9.0% and 7.6%, respectively, when compared to experimental data. The findings serve as an engineering guide for designing and optimizing heat transfer in LOX/LNG rocket engine cooling systems. Full article

The engine system requirements for different engine cycles significantly influence the design of the mixing head. A literature review of fuel-injection technology for hydrogen and methane is presented. The literature review aimed to answer proposed questions specific to the liquid rocket engine fuel injector design. The current review methodology accounts for the engine system effect. Thus, a comprehensive literature review of the working principles of startup-staged-combustion-cycle engines based on original patents is provided. At the end of the review, the research gaps and suggestions for further work are summarised. At high mass flow rate and injection pressure in the supercritical regime (>50 MPa), experience is limited to the staged-combustion cycle developed in Russia and the US. It is necessary to consider a fluid-dynamic heat transfer coupling study for the multi-injection element design in the supercritical state. Cryogenic spray atomisation experiments need to be designed with research significance in mind. It is still needed to study how the similarity of the spray flow field to the combustion performance affects a liquid rocket engine problem. Moreover, scaling stoichiometric mixing theory needs to be expanded to different injector types, such as tricoaxial and pintle injectors, to validate the correlation between the non-reactive mixing length and flame length. Full article

Variable-thrust liquid rocket engines are essential for precision landing in deep-space exploration, reusable launch vehicle recovery, high-accuracy orbital maneuvers, and emergency obstacle evasions of space drones. However, with the increasingly complex space missions, challenges remain with the development of different technical schemes. In view of these issues, this paper systematically reviews the technology’s evolution through mechanical throttling, electromechanical precision regulation, and commercial space-driven deep throttling. Then, the development of key variable thrust technologies for liquid rocket engines is summarized from the perspective of thrust regulation and control strategy. For instance, thrust regulation requires synergistic flow control devices and adjustable pintle injectors to dynamically match flow rates with injection pressure drops, ensuring combustion stability across wide thrust ranges—particularly under extreme conditions during space drones’ high-maneuver orbital adjustments—though pintle injector optimization for such scenarios remains challenging. System control must address strong multivariable coupling, response delays, and high-disturbance environments, as well as bottlenecks in sensor reliability and nonlinear modeling. Furthermore, prospects are made in response to the research progress, and breakthroughs are required in cryogenic wide-range flow regulation for liquid oxygen-methane propellants, combustion stability during deep throttling, and AI-based intelligent control to support space drones’ autonomous orbital transfer, rapid reusability, and on-demand trajectory correction in complex deep-space missions. Full article

Over the last few years, the appeal for using methane as green fuel for rocket engines has been on an increasing trend due to the more facile storage capability, reduced handling complexity and cost-effectiveness when compared to hydrogen. The present paper presents an attempt to simulate the ignition and propagation of the flame for a 1 kN gaseous methane–oxygen rocket engine using a pintle-type injector. By using advanced numerical simulations, the Eddy Dissipation Concept (EDC) combined with the Partially Stirred Reactor (PaSR) model and the Shielded Detached Eddy Simulation (SDES) were utilized in the complex transient ignition process. The results provide important information regarding the flame propagation and stability, pollutant formation and temperature distribution during the engine start-up, highlighting the uneven mixing regions and thermal load on the injector. This information could further be used for the pintle injector’s geometry optimization by addressing critical design challenges without employing the need for iterative prototyping during the early stages of development. Full article

Equivalent ratio (ER) is an important factor affecting detonation characteristics and propulsion performance of rotating detonation rocket engine (RDRE). In this paper, the effects of different equivalent ratios detonation characteristics and thrust performance of methane-oxygen RDRE were studied by 2D numerical simulation. The premixed reactants were injected through the injection holes to simulate the discrete injection of reactants on the injection panel in actual RDRE, the number of injection holes was 60 and 120. The results show that there is hybrid detonation mode (HDM), co-direction multi-wave detonation mode (CMM) and unstable detonation mode (UDM) in detonation combustion due to the influence of equivalent ratio and the number of injection holes, and the co-directional multi-wave detonation mode is beneficial to the thrust stability of RDRE. At the last, the number of detonation waves in RDRE decreases with the increase in the equivalent ratio, and the specific impulse (I_sp) increases with the increase of the equivalent ratio. Full article

Improper film cooling design and positioning of an injector in the face plate cause thermal damage to the thrust chamber wall and lead to rocket engine failures. An experimental combustor with five shear coaxial injectors was damaged owing to inadequate film cooling distribution on the thrust chamber wall. The present study aimed to simulate the experimental test case and investigate the causes of the thermal damage. In the simulation, gaseous methane and oxygen were injected at the inner and outer inlets of the shear coaxial injectors and nitrogen, used as the coolant, was injected near the upstream of the chamber wall. The turbulent chemistry interaction was modeled using a reduced DRM-19 mechanism by incorporating the Eddy Dissipation Concept model. Numerical investigations were conducted to examine the cause of thermal damage. The temperature contours of the thrust chamber wall were compared with the experimental image of the damaged wall. Further, simulations of single-row (SR) and multi-row (MR) injector configurations were conducted to assess the effect on film cooling distribution. The adiabatic film cooling effectiveness and specific impulse were determined for all simulated cases. The results showed that MR simulations with narrow injector angles had poor film cooling performance, while wider angles led to lower specific impulse. The face plate with an angle of 15 degrees between the injector positions showed better performance in terms of considering both the film cooling and specific impulse. Full article

Due to the extreme temperatures inside the combustion chambers of liquid propellant rocket engines, the walls of the combustion chamber and the nozzle are cooled by either the fuel or the oxidizer in what is known as regenerative cooling. This study presents an indigenous computational tool developed for the analysis of heat transfer in regenerative cooling of such rocket engines. The developed tool incorporates a one-dimensional (1-D) combustion analysis to calculate the thermophysical properties of the combustion gas. Basic engine properties were calculated and used to generate a thrust chamber profile based on a bell-shaped nozzle. The hot gas side was analyzed using 1-D isentropic flow assumptions, along with heat transfer correlations. The coolant side was evaluated using the hydraulic analysis in the axial direction and the heat transfer analysis in the radial direction. Thermophysical properties and the phase of the coolant were determined using the given property tables and the instantaneous state of the coolant. This flexible and computationally less demanding tool was used to analyze two small-scale engines utilizing liquid hydrocarbon fuels, which are used in modern rocket propulsion. The wall cooling analyses of a liquid oxygen (LOX)/liquid methane (LCH₄) engine and a liquid oxygen (LOX)/liquid propane (LC₃H₈) engine are presented. Fuel and oxidizer were used separately as coolants for both engines, and both of them experienced phase change. Results reveal the advantage of the high mass flow rate of the oxidizer in cooling performance. In addition, the results of this study show that the cooling of the LOX/LC₃H₈ engine is somewhat more challenging compared to the LOX/LCH₄ engine. Full article

Direct numerical simulations of non-premixed fuel-rich methane–oxygen flames at 20 bar are conducted to investigate the turbulent mixing burning of gaseous propellants in rocket engines. The reacting flow is simulated by using an EBI-DNS solver within an OpenFOAM frame. The transport of species is resolved with finite-rate chemistry by using a complex skeletal mechanism that entails 21 species. Two different flames at low and high Reynolds numbers are considered to study the sensitivity of the flame dynamics to turbulence. Regime markers are used to measure the probability of the flow to burn in premixed and non-premixed conditions at different regions. The local heat release statistics are studied in order to understand the drivers in the development of the turbulent diffusion flame. Despite the eminent non-premixed configuration, a significant amount of combustion takes place in premixed conditions. Premixed combustion is viable in both lean and fuel-rich regions, relatively far from the stoichiometric line. It has been found that a growing turbulent kinetic energy is detrimental to combustion in fuel-rich premixed conditions. This is motivated by the disruption of the local premixed flame front, which promotes fuel transport into the diffusion flame. In addition, at downstream positions, higher turbulence enables the advection of methane into the lean core of the flame, enhancing the burning rates in these regions. Therefore, the primary effect of turbulence is to increase the fraction of propellants burnt in oxygen-rich and near-stoichiometric conditions. Consequently, the mixture fraction of the products shifts towards lean conditions, influencing combustion completion at downstream positions. Full article

The cooling jackets of liquid rocket engines are composed of narrow passages surrounding the thrust chambers and ensure the reliable operation of the engine. Critical conditions may also be encountered, since the cooling jackets of cryogenic engines, such as those using LO_X/LCH₄ propellants, are based on a regenerative strategy, where the fuel is used as a refrigerant. Consequently, deterioration modes near where pseudocritical conditions are reached or low heat transfer coefficients where the fuel becomes a vapour and must therefore be managed. The verification of the cooling jacket behaviour to consolidate the design solutions in all the extreme points of the operating box represents a very important phase. The present paper discusses the full characterization of the HYPROB (HYdrocarbon PROpulsion test Bench Program) first unit of the final demonstrator, (DEMO-0A), by considering the working points within the limits of the operating box and comparisons with the nominal conditions are given. In this way, a full understanding of the cooling system behaviour, affecting the working of the entire thrust chamber, is accomplished. Moreover, the design strategy and choices have been confirmed, since the verifications also include potentially even more extreme conditions with respect to the nominal ones. The investigation has been numerically performed and supported the thermo-structural analyses accomplished before the final firing campaign, completed in December 2022. Since little information is available in the literature on LO_X/LCH₄ engines, suggestions are given as to the organization of the numerical simulations, which support the design of such rocket engine cooling systems. Full article

This study presents an optimal design approach of a pintle injector for a deep throttlable liquid-propellant rocket engine (LPRE). Even though the pintle injector is used in rocket engines, it has become more important since reusable launch vehicles (RLVs) recently became a trend due to their economic and environmental benefits. However, since many variables must be determined to design a pintle injector, optimizing the pintle injector design is complicated. For this, we design a pintle injector to optimize the performance parameters; the spray angle, vaporization distance, and Sauter mean diameter (SMD). To confirm the approach, we design a pintle injector using an optimization method based on convex quadratic programming (CQP) for a 1000 N thrust and a throttle ability of 5 to 1 LPRE with liquid oxygen and gaseous methane. Then, we verify the performance using a numerical simulation. Through this work, we check the effectiveness of the optimization method for a pintle injector design. Full article

Methane (CH₄) is a promising rocket fuel for various future space mission scenarios. It has advantages in terms of cost, performance, and environmental friendliness. Currently, there is no clear definition on standards and specifications for liquefied methane or similar liquids such as liquefied natural gas (LNG) for their use as rocket fuel. However, those regulations are necessary for the commercial, safe, and proper operation of methane rocket engines. Composition and impurities of liquefied methane gas mixtures obtained from natural gas or biogenic sources depend on location of the natural gas source (Europe, Asia, or America), its extraction method and treatment, used cleaning methods or conditions of the gasification process, and biomass sources. In the present work, effects of impurities (N₂, CO₂, C₂H₆) within liquid natural gas/liquid methane on the methalox rocket engine operation behavior are analyzed. Regarding the cold cryogenic side, phase diagrams are discussed and critical temperatures for the fuel side are outlined. Carbon dioxide is identified as a rather problematic pollutant. The combustion processes are investigated with several numerical simulations (1D and 2D CFD). The results indicate a minor influence on the overall combustion temperature and a minor but potentially relevant influence on the pressure within the combustion chamber. Additionally, the results indicate that with respect to temperature and pressure, no complex NOx nitrogen chemistry is required. Full article

Hot-fire tests were performed with a single-injector research combustor featuring a large optical access (255 × 38 mm) for flame imaging. These tests were conducted with the propellant combination of liquid oxygen and compressed natural gas (LOX/CNG) at conditions relevant for main- and upper-stage engines. The large optical access enabled synchronized flame imaging using OH* and CH* radiation wavelengths covering an area of the combustion chamber from the injection plane to shortly before the contraction section of the nozzle for two sets of operating conditions. Combined with temperature, pressure and unsteady pressure measurements, these data provide a high-quality basis for validation of numerical modeling. Flame width and opening angle were extracted from the imaging in order to determine the flame topology. A two dimensional Rayleigh Index was calculated for an acoustically unexcited and excited interval. These Rayleigh Indices are in good agreement with the thermoacoustic state of the chamber. Full article

The successful design of a liquid rocket engine is strictly linked to the development of efficient cooling systems, able to dissipate huge thermal loads coming from the combustion in the thrust chamber. Generally, cooling architectures are based on regenerative strategies, adopting fuels as coolants; and on cooling jackets, including several narrow axial channels allocated around the thrust chambers. Moreover, since cryogenic fuels are used, as in the case of oxygen/methane-based liquid rocket engines, the refrigerant is injected in liquid phase at supercritical pressure conditions and heated by the thermal load coming from the combustion chamber, which tends to experience transcritical conditions until behaving as a supercritical vapor before exiting the cooling jacket. The comprehension of fluid behavior inside the cooling jackets of liquid-oxygen/methane rocket engines as a function of different operative conditions represents not only a current topic but a critical issue for the development of future propulsion systems. Hence, the current manuscript discusses the results concerning the cooling jacket equipping the liquid-oxygen/liquid-methane demonstrator, designed and manufactured within the scope of HYPROB-NEW Italian Project. In particular, numerical results considering the nominal operating conditions and the influence of variables, such as the inlet temperature and pressure values of refrigerant as well as mass-flow rate, are shown to discuss the fluid transcritical behavior inside the cooling channels and give indications on the numerical methodologies, supporting the design of liquid-oxygen/liquid-methane rocket-engine cooling systems. Validation has been accomplished by means of experimental results obtained through a specific test article, provided with a cooling channel, characterized by dimensions representative of HYPROB DEMO-0A regenerative combustion chamber. Full article

The Italian Ministry of University and Research (MIUR) funded the HYPROB Program to develop regeneratively cooled liquid rocket engines. In this type of engine, liquid propellant oxygen–methane is used, allowing us to reach very good performances in terms of high vacuum specific impulse and high thrust-to-weight ratio. The present study focused on the HYPROB final ground demonstrator, which will be able to produce a 30 kN thrust in flight conditions. In order to achieve such a thrust level, very high chamber pressures (up to 50 bar) and consequently high thermal fluxes and gradients are expected inside the thrust chamber. Very complex and high-fidelity numerical FEM models were adopted here to accurately simulate the thermal–mechanical behavior of the thrust chamber cooling channels, accounting for plasticity, creep, and low-cycle fatigue (LCF) phenomena. The aim of the current work was to investigate the main failure phenomena that could occur during the thrust chamber’s service life. Results demonstrated that LCF is the main cause of failure. The corresponding number of loading cycles to failure were calculated accordingly. Full article

In this study, a single injector methane-oxygen rocket combustor is numerically studied. The simulations included in this study are based on the hardware and experimental data from the Technical University of Munich. The focus is on the recently developed generalized k–ω turbulence model (GEKO) and the effect of its adjustable coefficients on the pressure and on wall heat flux profiles, which are compared with the experimental data. It was found that the coefficients of ‘jet’, ‘near-wall’, and ‘mixing’ have a major impact, whereas the opposite can be deduced about the ‘separation’ parameter Csep, which highly influences the pressure and wall heat flux distributions due to the changes in the eddy-viscosity field. The simulation results are compared with the standard k–ε model, displaying a qualitatively and quantitatively similar behavior to the GEKO model at a Csep equal to unity. The default GEKO model shows a stable performance for three oxidizer-to-fuel ratios, enhancing the reliability of its use. The simulations are conducted using two chemical kinetic mechanisms: Zhukov and Kong and the more detailed RAMEC. The influence of the combustion model is of the same order as the influence of the turbulence model. In general, the numerical results present a good or satisfactory agreement with the experiment, and both GEKO at Csep = 1 or the standard k–ε model can be recommended for usage in the CFD simulations of rocket combustion chambers, as well as the Zhukov–Kong mechanism in conjunction with the flamelet approach. Full article