Search Results (16)

The hydrocarbon liquid rocket engine working environment is harsh; the thrust chamber needs to withstand high temperatures and a high-pressure working environment, and the thrust chamber wall material is difficult to bear, so it is necessary to design the cooling structure to reduce the gas damage to the chamber wall. Liquid film cooling is a common cooling method for hydrocarbon rocket engines, and numerical simulation is an important method for studying liquid film cooling. Most of the liquid film cooling numerical simulation is for a fixed model. This paper proposes a liquid film cooling numerical calculation method for a variable-configuration hydrocarbon liquid rocket engine, based on the secondary development of Fluent software (ANSYS Fluent 2022) to form a high-energy hydrocarbon liquid rocket engine design software, which can be realized on the Qt platform. The visualization interface can be for different engine injection port locations, numbers, angles, mass flow rates, and other parameters, to calculate and improve design efficiency and reduce operating difficulty. 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

To investigate the wall cooling of the thrust chamber in an engine, two film-cooling structures, namely, a circular hole structure and a slot structure, were designed. Numerical simulations were performed to study the coupled flow and regenerative cooling heat transfer in thrust chambers with different structures. The influences of parameters such as the film mass flow rate and film hole size on wall cooling were analyzed. Experiments were conducted in a thrust chamber to validate the accuracy of the numerical calculation method. The results indicate that the slot-structured film adheres better to the wall than the circular-hole-structured film, and the film closely adhering to the wall provides better insulation against hot gas, resulting in a reduction of approximately 6% in wall temperature. When the film hole size changes, the change in circumferential wall temperature in the upstream region of the slot-structured film is more pronounced. This paper aims to provide a reference for the design of the cooling structure at the head of the thrust chamber in engineering and suggests directions for optimization and improvement. 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

This paper presents the results of numerical simulation of a model combustion chamber of a pulse detonation engine using the authors’ developed software package. The main goal of the present study is to numerically investigate the effects of cyclic operation of pulse detonating chambers, as the former studies have been limited to simulating one cycle. To achieve this goal, a new mathematical model for heavy gas was applied simulating condensed fuel phase, which made it possible to accelerate computations and simulate multi-cycle operation of the device. Distributions of such characteristics as temperature, pressure, velocity, concentrations of reagents, intensity of reactions, and thrust force are obtained. A two-stage kinetic model of propellant combustion is proposed. Attention is paid to the main stages of PDE operation: filling of the chamber with reagents, ignition and transition to detonation, products exhaust, purification, and cooling the chamber with a neutral gas. The simulation of the working cycle with the shortest period for the specified system parameters was carried out, the execution time of each stage was obtained, and an assessment was carried out to minimize the main stages of the work cycle. Numerical results demonstrated that the characteristics of the engine cycle are stabilized already in the second cycle: the thrust in the first cycle differs from the thrust in the second by 5%, in the third from the second by 1%. Moreover, details of thrust dynamics in the second and third cycles were studied. 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

To better understand the characteristics of coupled heat transfer in liquid rocket engines, a calculation scheme is proposed in this paper. This scheme can simulate the coupled heat transfer processes, including combustion and flow in the thrust chamber, radiation heat transfer, heat conduction in the wall, heat transfer of coolant flow in the cooling channel, and gas film cooling in the thrust chamber wall. The numerical method used in each physical area, the data transfer method between each computing module, the strategy of data transfer on the coupling interface, the calculation process, and the convergence criterion are all introduced in detail. The calculation scheme was verified by analyzing a water-cooled nozzle. Then, the coupled heat transfer calculation was carried out for a liquid rocket engine using a propellant composed of unsymmetrical dimethylhydrazine and dinitrogen tetroxide. Two working conditions were analyzed: whether the gas film cooling was performed or not. The results showed that the algorithm successfully indicated the protective effect of the gas film on the wall surface, and the calculation results were reasonable. It played a guiding role for the coupled heat transfer of the liquid rocket engine using a composite cooling method. Full article

An essential part of the design of a liquid rocket engine is the thermal analysis of the thrust chamber, which is a component whose operative life is limited by the maximum allowable wall temperature and heat flux. A simplified steady-state thermal analysis model for regeneratively cooled rocket engine thrust chambers is presented. The model is based on semi-empirical correlations for the hot-gas and coolant convective heat transfer and on an original multi-zone approach for the wall conduction. The hot-gas heat transfer is calibrated with experimental data taken from an additively manufactured water-cooled nozzle that is connected to a combustion chamber either fed with decomposed hydrogen peroxide or decomposed hydrogen peroxide and automotive diesel. The thrust chamber (i.e., combustion chamber and nozzle) is designed to produce about 450 N of thrust when operating with a chamber pressure of 11 bar. For this application, the calibrated model predicts the total wall heat transfer rate very accurately and the temperature distribution within the wall structure with an uncertainty of a few tens of kelvins. This level of accuracy can be considered more than adequate for the design, and generally for engineering-type thermal analysis, of similar thrust chambers. Full article

Liquid rocket engines with hydrogen peroxide and kerosene have the advantages of high density specific impulse, high reliability, and no ignition system. At present, the cooling problem of hydrogen peroxide engines, especially with regenerative cooling, has been little explored. In this study, a realizable k-epsilon turbulence model, discrete phase model, eddy dissipation concept model, and 10-step 10-component reaction mechanism of kerosene with oxygen are used. The increased rib height of the regenerative cooling channel causes the inner wall temperature of the engine increases, the average temperature of the coolant outlet decreases slightly, and the coolant pressure decreases. The overall wall temperature decreases as the rib width of the regenerative cooling channel increases. However, in the nozzle throat area, the wall temperature increases, the average coolant outlet temperature decreases, and the coolant pressure drop increases. A decrease in the inner wall thickness of the regenerative cooling channel results in a significant decrease in the wall temperature and a small increase in the average coolant outlet temperature. These findings contribute to the further development of the engine with hydrogen peroxide and can guide the design of its regenerative cooling process. Full article

Rotating (also termed continuous spin) detonation technology is gaining interest in the global research and development community due to the potential for increased performance. Potential performance benefits, thrust chamber design, and thrust chamber cooling loads are analyzed for propellant applications using liquid oxygen or high-concentration hydrogen peroxide oxidizers with kerosene, hydrogen, and methane fuels. Performance results based on a lumped parameter treatment show that theoretical specific impulse gains of 3–14% are achievable with the highest benefit coming from hydrogen-fueled systems. Assessment of thrust chamber designs for notional space missions shows that both thrust chamber length and diameter benefits are achievable given the tiny annular chamber volume associated with the rotating detonation combustion. While the passing detonation front drastically increases local heat fluxes, global energy balances can be achieved if operating pressures are limited to be comparable to existing or prior space engines. Full article

Rocket kerosene plays an important role in the regenerative cooling process of rocket thrust chambers. Its thermal conductivity determines the cooling efficiency and the tendency to coke in rocket kerosene engines. In this paper, graphene nanoplatelets (GNPs) were introduced into rocket kerosene to improve its thermal conductivity. Molecular dynamics simulation was used to investigate the thermal conductivity of the composite system and its underlying thermal conductivity mechanism. Firstly, by studying the effect of the mass fraction of GNPs, it was found that, when the graphene mass fraction increases from 1.14% to 6.49%, the thermal conductivity of the composite system increases from 4.26% to 17.83%, which can be explained by the percolation theory. Secondly, the influence of the size of GNPs on the thermal conductivity of the composite system was studied. Basically, the thermal conductivity was found to increase by increasing the aspect ratio of GNPs, indicating that GNPs with a higher aspect ratio are more conducive to improving the thermal conductivity of rocket kerosene. By carefully analyzing the effect of the size of GNPs on thermal conductivity, it was concluded that the thermal conduction enhancement by adding GNPs is determined by the combined effect of the percolation theory and the Brownian motion. The results of the temperature effect study showed that the ratio of thermal conductivity to rocket kerosene increased from 1.16 to 1.26 and from 1.07 to 1.11 for the composite systems, with graphene sizes of 41.18 Å × 64.00 Å and 24.14 Å × 17.22 Å in the temperature range of 293 K to 343 K, respectively. It is further proved that the Brownian motion of GNPs has a non-negligible effect on the thermal conductivity of the composite system. This work provides microscopic insights into the thermal conduction mechanism of GNPs in nanofluids and will offer practical guidance for improving the thermal conductivity of rocket kerosene. 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

In the present study, a hydrogen and oxygen heat-sink engine thrust chamber and the corresponding injection faceplate with discrete slot orifices are devised to study the cooling performance near the faceplate region. Moreover, a set of experiments and numerical simulations are conducted to evaluate the effects of various factors on combustion performance and film cooling efficiency. According to the obtained result, the circumferential cooling efficiency has an M-shaped distribution in the near-injector region. Furthermore, it has been discovered that when the film flow ratio increases, so does the cooling efficiency. This is especially more pronounced in the range of 30–80 mm from the faceplate. The cooling efficiency is found to be proportional to the film flow rate ratio’s 0.4 power. Compared with the slot thickness, the reduction in the slot width is more beneficial in improving the cooling efficiency, and the advantage is more prominent for small film flow ratios. In addition, when the amount of coolant is not enough, the cooling effect of the discrete slot film orifice is better than that of the common cylindrical orifice. The present article demonstrates that setting the area ratio of the adjacent film orifices is an effective way to reduce the uneven circumferential distribution of the wall surface temperature. 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

Ablative-cooled hybrid rockets could potentially combine a similar versatility of a liquid propulsion system with a much simplified architecture. These characteristics make this kind of propulsion attractive, among others, for applications such as satellites and upper stages. In this paper, the use of hybrid rockets for those situations is reviewed. It is shown that, for a competitive implementation, several challenges need to be addressed, which are not the general ones often discussed in the hybrid literature. In particular, the optimal thrust to burning time ratio, which is often relatively low in liquid engines, has a deep impact on the grain geometry, that, in turn, must comply some constrains. The regression rate sometime needs to be tailored in order to avoid unreasonable grain shapes, with the consequence that the dimensional trends start to follow some sort of counter-intuitive behavior. The length to diameter ratio of the hybrid combustion chamber imposes some packaging issues in order to compact the whole propulsion system. Finally, the heat soak-back during long off phases between multiple burns could compromise the integrity of the case and of the solid fuel. Therefore, if the advantages of hybrid propulsion are to be exploited, the aspects mentioned in this paper shall be carefully considered and properly faced. Full article