Rocket Engine Chamber Research Articles

Development of a mathematical model for the cooling channel of a liquid propellant rocket engine’s chamber with respect for variations in coolant density

This study focuses on the processes occurring within the cooling channels of a liquid rocket engine chamber. It is important to note that, from a design-stage efficiency prediction perspective, the cooling system is the most critical part of the engine chamber. This is because it cannot be tested without costly and labor-intensive fire tests. Therefore, mathematical models of heat transfer and fluid dynamics must describe accurately all processes taking place in the chamber. The study emphasizes accounting for the change in the density of the propellant component within the cooling channels. To verify the importance of this issue, an analysis was conducted on the changing parameters of the propellant components in the cooling channels of an engine. The assessment revealed that even when using high-boiling propellant components and moderate heating in the cooling channels, density changes can exceed 25 %. This paper presents the results of developing a model for the flow of propellant in the cooling channels of a liquid rocket engine chamber, taking density changes into account. The model builds on a cooling channel model previously developed by the authors. An equation that accounts for density variations was derived using established principles of fluid mechanics. Using the developed mathematical model, test calculations were performed, and the simulation results were compared with and without considering density changes. Furthermore, a comparison was conducted with calculated data available in the literature on heat transfer in the RD107 engine chamber, revealing an error margin of no more than 1.5 %. The resulting mathematical model may be recommended for use in the design of new rocket engine chambers with regenerative cooling

Results of experimental evaluation of the possibility of using an annular nozzle with a flat central body as part of a stand liquid rocket engine chamber

One of the ways to increase the efficiency of a rocket engine chamber is to ensure the fixed nozzle operation in the design mode over the entire range of altitudes (ambient pressure) of engine operation. The paper presents the results of a full-scale experiment of the annular nozzle model with a flat central body. The purpose of the experiment was to evaluate the efficiency in terms of the specific impulse of the annular nozzle with a flat central body in comparison with the Laval nozzle. The tests were carried out at the “hot” bench of the Department “Engines and Power Plants of Aircraft” BSTU “VOENMEH” named after D.F. Ustinov. The generated high-temperature steam-gas was used as a working fluid, the experiment was carried out under terrestrial conditions. During the stand testing the readings of operating parameters (pressure, mass flow, thrust, components ratio) were recorded in automated mode. The parameters were changed by reconfiguring the appropriate elements of the pneumohydraulic system. Based on the processed results of the tests the efficiency of the annular nozzle with a flat central body was evaluated in comparison with the Laval nozzle, and the values of the experimental data were verified with the theoretically expected ones.

Simulation of the processes of spraying and combustion of kerosene and liquid oxygen in the chamber of a liquid-propellant rocket engine

The paper presents the results of simulation of spray and combustion processes in the liquid-propellant rocket engine chamber using the z77 reduced kinetic mechanism of chemical reactions and a hybrid method for determining spray parameters of a swirl atomizer. Simulation was carried out for the nominal operational conditions in a three-dimensional domain using the ANSYS Fluent software. The process of spraing liquid fuel components (T-1 kerosene and oxygen) by monopropellant injectors was simulated using a model of discontinuous phases. The simulation results (pressure, temperature and velocity) were compared with the data of analytical thermogasdynamic calculation results and experimental results for pressure in chamber and gas temperature near the wall. The difference between the simulation results and the experimental results does not exceed 8%. Thus, it was shown that it is possible to use the mechanism z77 and hybrid method for determining spray parameters of a swirl atomizer presented in this paper to obtain accurate simulation results of the kerosene T-1 and oxygen spray and combustion processes of the investigated liquid-propellant rocket engine.

Development and research testing of a low thrust on gaseous-propellant rocket engine chamber

The article presents preliminary results of the development and research testing of a rocket engine chamber with a thrust of 100 N operating on fuel components: gaseous oxygen-gaseous hydrogen, designed to be used as the main engine of a small upper stage for launching a payload weighing up to 150 kg into target orbits. The main features of the engine chamber are its fabrication by selective laser melting from 12X18N10T steel powder and its regenerative cooling with gaseous hydrogen. Calculation and experiments confirmed the possibility of regenerative cooling of the chamber in the nominal mode. In addition, the current results of the development of a technique for optical recording of the process of structural material removal from the chamber during tests using different optical filters are presented. It is shown that there is a correlation between the brightness of the obtained frames and the hydrogen flow rate. It is also shown that the afterburning of the removed material particles, in contrast to high thrust engines, occurs mainly in the tail of the jet.

Comparative results of computational and theoretical study of the annular nozzle with a flat central body

One of the ways to improve specific characteristics of the power plant of a launch vehicle for ladeploying payload to the near-Earth space is to provide the possibility of operation of a fixed nozzle in the design mode over the whole active leg of the flight trajectory. The nozzle should be compact, lightweight, well-cooled. For detailed testing of the possibility of introducing a nozzle into the rocket engine chamber it is necessary to be able to quickly assess the true value of the thrust and the specific impulse the chamber with such a nozzle can achieve. This article presents the results of comparison of the thrust and specific impulse, obtained during calculations using engineering methods, numerical modeling for the atmospheric section and high-altitude sections of the trajectory. The results of the calculation are compared with the experimental values of the specific impulse obtained on the rocket engine test-bed under atmospheric operating conditions. These results can be effectively applied both to evaluate new and to improve existing nozzle designs of wide-range rocket engines.

DEVELOPMENT OF A DIFFERENTIAL MODEL FOR COOLING AN LPRE CHAMBER BY AN INCOMPRESSIBLE FLUID

Abstract. The combustion chamber constitutes a crucial component of a liquid rocket engine (LPRE), and approximately 30% of serial engine failures are attributed to issues within it. Primarily, challenges encountered during engine chamber operation stem from the inadequacy of the current cooling systems to manage thermal loads effectively. These loads, unique to liquid-propellant rocket engine chambers, result of high pressures, temperatures, and flow rates of dissociating combustion products. Regenerative cooling stands as the primary method to safe the chamber walls from overheating and subsequent deterioration. Other methods serve as supplementary measures, often utilized in conjunction with regenerative cooling (e.g., internal film cooling), or are tailored to specific applications, such as radiation cooling of upper-stage engine nozzles. However, the advancement of heat transfer theory within liquid-propellant rocket engine chambers has been relatively stagnant, with widely accepted calculation methods established as far back as the 1970s. Consequently, these methods fail to harness the considerable computing power enhancements since their inception, neglect new experimental data, and lack adaptation to modern engine manufacturing technologies like 3D printing. Concurrently, the burgeoning number of startups in the rocket and space sector, constrained by limited budgets, necessitates the swift development of viable designs without extensive and costly testing. Hence, the need for validated calculation methods for regenerative cooling is paramount. This study endeavors to address this need by formulating a system of differential equations, grounded on the mass, momentum, and energy conservation laws, to describe the processes within the cooling circuit of the liquid-propellant rocket engine chamber. Additionally, the model is dimensionally reduced for simplification. Comparative analysis against established engineering methodologies showcases the advantages of the proposed approach, underscoring its potential for enhancing cooling system design efficacy.

FEATURES OF THE DEVELOPMENT OF ADDITIVE MANUFACTURING METHODS IN APPLICATION TO LIQUID PROPELLANT ROCKET ENGINES

Abstract. Modern development of additive manufacturing methods makes it necessary to seriously consider an alternative type of production for the manufacture of liquid rocket engine structures. On the one hand, additive manufacturing has a number of undeniable advantages compared to classical methods traditionally used in the rocket and space industry. On the other hand, products obtained by this method have characteristic features that must be taken into account at the early stages of design: relatively high surface roughness, the need to adapt the geometry of existing technical solutions, material properties, etc. Moreover, a full-fledged quality control system for the resulting products is currently the moment is missing, and the standardization of the production class is just emerging. It is also important to understand the direction of development of additive manufacturing, which in the future can lead to a significant increase in the characteristics of the resulting products. This paper discusses options for classifying additive manufacturing methods, provides definitions and classification of metal 3D printing methods, according to the existing ISO/ASTM 52900 standard, as well as the main features of the technology. The main advantages and disadvantages of additive manufacturing methods are considered, and a structural decomposition of the typical design of a liquid propellant rocket engine chamber is carried out. The work aimed at manufacturing structural elements for further integration into the production cycle of liquid-propellant rocket engines is considered. The work is considered, the results of which were a combination of additive technology methods, which made it possible to apply the concept of bimetallic additive production. Such products were produced without the use of complex technological equipment and belong to the product of “evolving” additive manufacturing. Based on the results of the work, it was concluded that it is possible to use additive technologies to improve the performance of new liquid-propellant rocket engine designs.

Experimental and numerical study on the process of removing kerosene residues in rocket engine chamber

This article conducts experimental and numerical research on the process of removing residual kerosene from the wall of the rocket engine chamber after test run, and proposes a cleaning method of high-pressure nitrogen blowing and low-pressure suction intermittent cycle. The research idea is to simulate the concentration changes of residual kerosene liquid and steam during the removal process using CFD method to analyze the mass transfer mechanism during the process, and to verify the accuracy of the numerical model through experiments. It can be concluded that the intermittent cycle cleaning method proposed in this article can efficiently remove residual kerosene from the wall of the rocket engine chamber. By using this cleaning method for 120 min, the maximum cleaning efficiency reaches 97.86%. At the same time, it was found that the high-pressure nitrogen blowing effect was the most obvious during the cleaning process, and switching between high and low pressure during the circulation process was beneficial for the removal of residual kerosene.

Serpentine and furrowed wavy channels-based enhancement in heat transfer for a rocket engine chamber

The thermal protection of the rocket engine combustion chamber is one of the fundamental challenges in supersonic flight. This paper investigates serpentine and furrowed wavy cooling channels with a rectangular cross-section to enhance heat transfer performance and hydrogen turbulent fluid flow characteristics. Ansys fluent was used to investigate the effects of configuration, wavy amplitude, and channel wavelength on the heat wall temperature, Nusselt number, coolant velocity, pressure drop, turbulent kinetic energy, and thermal stratification phenomena. Four different wave amplitudes and three wavelengths are compared to determine the optimal structure required to achieve the best heat transfer enhancement with acceptable pressure drop. Compared to the straight channel, the maximum temperature of the furrowed channel (Case 6) was decreased by 15.38% at an amplitude of 0.2 mm. Furthermore, the thermal performance factor is evaluated for eight furrowed wavy structures. The result is shown that the structure with an amplitude of 0.2 mm and a wavelength of 1 mm (Case 6) can demonstrate better thermal performance, which is increased by 31.55% compared to the straight channel. The wavy surfaces cause separation in diverging zones and secondary flows in converging zones, which disturbs the boundary layer, enhances particle mixing and improves convective heat transfer.

Numerical study on enhanced heat transfer and flow characteristics of supercritical hydrogen rocket engine's chamber wall using cylindrical ribs structure

One of the most critical issues in supersonic flight is the thermal protection of the rocket engine combustion chamber. The ribs' construction can improve heat transfer and coolant flow characteristics significantly. Ansys fluent was used to investigate the supercritical hydrogen fuel flow and heat transfer characteristics. Ribs spacing of 10 mm, 5 mm, and 2.5 mm were compared to determine the optimal spacing required to achieve maximum cooling influence and lower pressure drop to safeguard the overheated structure and lower the temperature of the walls. Compared to the smooth channel, the thermal performance factor and pressure drop of the 5 mm spacing increased by 32.5% and 16.64%, respectively, while the maximum temperature reduced by 12.07%. Therefore, the cylindrical rib construction with a 5 mm spacing is perfect for improving heat transfer, has acceptable pressure drops, and reduces thermal stratification inside the channel.

Thermal enhancement and entropy investigation in dissipative ZnO-SAE50 under thermal radiation: a computational paradigm

Exploring heat transport mechanism and entropy generation in new heat transfer fluids generation are imperative and broadly applicable in many industrial and engineering domains. Therefore, the study of heat and entropy generation inside a rocket engine chamber filled with ZnO-SAE50 nanolubricants is designed, and the main version is attained by exercising similarity equations. After that, the numerical computation was performed under varying physical constraints and explored fascinating outcomes. The results revealed that the fluid motion drops inside the chamber by strengthening Re and ZnO fraction factors, respectively. Furthermore, the entropy generation for ZnO-SAE50 is very high than regular liquids, and this characteristic allows nanofluids for effective industrial and manufacturing purposes. Implementation of thermal radiations over the chamber is a prominent factor for enhancing the Nusselt number at both ends of the chamber. Finally, the reported study would be valuable for applied thermal engineering problems.

Heat transfer enhancement of hydrogen rocket engine chamber wall by using V-shape rib

The cooling capacity of regenerative cooling channels is still insufficient for the liquid rocket engine due to the high temperature of the combustion chamber and the little refrigerant mass flow. To improve the thermo-hydraulic performance of the combustion chamber wall, a V-shape rib is proposed, and its effect on heat transfer and pressure drop is investigated. Numerical models are built of a regenerative cooling channel with V-shape rib using CFD commercial software. Simulation results showed that V-shape ribs enhance convection heat transfer and improve thermo-hydraulic performance by 23 % compared to smooth cooling channels. Although the maximum temperature was reduced by 20.9 % , the flow resistance increased slightly. The overall heat transfer coefficient of the cooling channel with 30 ∘ V-shape rib is superior to that of the 60 ∘ and 45 ∘ V-shape ribs, and 30 ∘ V-shape rib with 10 m m pitch has the best cooling capacity. • A low temperature of the rocket engine chamber wall is essential to ensure safe operation. • Three V-shape ribs are investigated and compared with the smooth regenerative cooling channel. • 30° V-shape rib has a better thermo-hydraulic performance than the 60° and 45° V-shape ribs.

Metal powder ignition characteristics in a hybrid rocket engine chamber

Metal powder ignition characteristics in a hybrid rocket engine chamber

Enhanced thermal resistance and ablation properties of ethylene-propylene-diene monomer rubber with boron-containing phenolic resins

With the development of aerospace industry, higher requirements have been put forward for the thermal properties of internal insulation materials for heat protection of solid rocket engine chamber. In this study, boron-containing phenolic resin (BPR) was chosen as the thermal-resistant filler of ethylene-propylene-diene monomer (EPDM). By adjusting the amount of BPR, a series of BPR/EPDM were prepared and their thermal, mechanical, and ablation properties were studied. The results showed that the addition of BPR significantly improved the char yield of EPDM. The char yield of BPR/EPDM was 21.6%, which increased by 17.9% (800 °C) relative to EPDM. By studying the structural evolution of BPR/EPDM during pyrolysis, boron oxide (B2O3) was formed, which avoided the formation of volatile carbon oxides and reduced the carbon loss. The molten B2O3 penetrated into the pores of carbonization products, which can protect the unstable edges of graphitic carbon and improve the ablation performance of BPR/EPDM. The linear ablation rate of BPR/EPDM was 0.045 mm/s (the lowest reported in current references). Additionally, the addition of BPR improved the graphitization degree of carbonization product. This research is expected to provide a resin matrix with excellent comprehensive performance and large-scale production for combustion chamber of the solid engine.

Численное исследование влияния физико-химических свойств жидкого горючего на эффективность рабочего процесса ракетного двигателя малой тяги

Low-thrust rocket engines are widely used in rocket and space technology for correcting the position of a spacecraft in orbit, for controlling motion along a trajectory, etc. Their number in the propulsion system can be from one to tens of units. Accordingly, the efficiency of their work significantly affects the perfection of the propulsion system as a whole. The object of the study was the low-thrust rocket engine combustion chamber operating according to the gas-liquid scheme. There were performed computational and parametric studies of various factor effects on the characteristics of the working process in the combustion chamber. The dependences of the coefficient of the consumable complex and parameters of the working process of the low-thrust rocket engine chamber on the influencing factors when using ethanol and kerosene as a fuel were calculated. A comparative analysis of the results of using these two components under similar conditions was carried out, which made it possible to reveal the influence of the physicochemical properties of the combustible component on the efficiency of the working process organization. The results obtained can be used in the design of low-thrust engines operating on the kerosene–oxygen and ethanol–oxygen propellants.

Obtaining a composite material based on quartz woven filler and pyrolysis matrix of organosilicon resin

The relevance of the study is conditioned by the fact that the most popular and irreplaceable materials that have found wide application in the aerospace industry are composites based on quartz materials. These materials are distinguished by their high mechanical and electrical strength, chemical and corrosion resistance. In this regard, it is of interest to obtain a composite material that combines a low specific gravity, processability of polymers, and thermal stability of ceramics. The aim of this work was to study the effect of the temperature of thermal oxidative destruction of a polymer binder, which is a semi-finished product of a pyrolysis matrix, on the electrophysical parameters of a composite material. The paper investigates a composite material based on woven quartz material with a pyrolysis matrix of an organosilicon binder and functional additives. This composite was considered as a material for creating an electric rocket engine chamber. Thermogravimetric analysis was used to evaluate the effect of the temperature of thermooxidative degradation of the polymer binder on the electro-physical parameters of the obtained material. The tests were carried out according to standard test methods on an Instron 5969 universal testing machine with Bluehill software until the samples failed. In the course of the study, it was found that the processes that occur up to 400°C are mainly associated with the course of the reaction for non-entered functional groups, the telomerisation reaction, intramolecular rearrangement of macromolecules and the removal of low-boiling substances. According to the results of the study, the obtained characteristics of the test material turned out to be suitable for its use in structural elements of electric propulsion engines.

Electrophysics of the Combustion of Hydrocarbon Fuel in the Liquid Propellant Rocket Engine Chamber

Electrophysics of the Combustion of Hydrocarbon Fuel in the Liquid Propellant Rocket Engine Chamber

Математическое моделирование течения охладителя в тракте охлаждения камеры ЖРД с предельно высокой степенью оребрения

The paper presents a computational analysis of coolant distribution in the cooling channel of a liquid rocket engine combustion chamber, performed in order to develop a set of practical guidelines towards increasing efficiency of a cooling system featuring an extremely high degree of ribbing. We created a three-dimensional mathematical model comprising a closed system of hydrodynamic equations as well as initial and boundary conditions for an element of the liquid rocket engine chamber we modelled, the chamber featuring longitudinal cooling channel arrangement manufactured via additive technology. We computed velocity and pressure fields in characteristic cooling channel regions for various levels of coolant mass flow rate, which confirmed the feasibility of the layout proposed in terms of uniform coolant distribution in the cooling channel of the liquid rocket engine modelled. We obtained the friction loss ξ as a function of coolant mass flow rate and particle size of the powder used in the additive technology to manufacture the combustion chamber wall and cooling channel.

К ВОПРОСУ ТЕРМОДИНАМИЧЕСКОГО ПРОЕКТИРОВАНИЯ КАМЕРЫ ЖРД

Analysis of the thermodynamic and thermophysical properties of combustion products (CP) in the liquid rocket engine (LRE) chamber shows that their dissociation degree depends on temperature T, gas expansion degree ε, etc. Practically, CP’s are always chemically active working fluid, therefore the number of moles N of the products varies along the length of the LRE chamber in the entire reaction mixture. The local values of the parameters T and N depend on the specific physical conditions. Therefore, the distribution of local numbers of moles of the components of the gas mixture and its heat capacities can be represented as dependencies N~f(T) and c~g(T). For this purpose based on the numerical values of the moles and the heat capacities of the gas mixture components in the main sections of the LRE chamber are formed as corresponding empirical functions through interpolation. Analysis of changes in moles and weight fractions of gaseous and condensed СP’s components shows that, depending on the specific conditions (α, Km, pc, ε), the number of moles of one group of individual substances increases, while these parameters of the other group decrease. These changes are alternating in nature and lead to the formation of new centers -sources of chemical and thermal energy along the length of the LRE nozzle. Thus, for different conditions (α, Km, pc, ε), the design of the LRE chamber should be carried out taking into account the nature of the change in N, cp, cv, and γ. Therefore, from energy conversion, the number of moles of the i-th CP component can be represented as a function Ni = f(Ti) or Ni = f(x,y). Numerical studies show that, based on the Ni values in the main sections of the LRE chamber with a given length, it is possible to form linear or nonlinear empirical functions in the form Ni= f(x) by interpolation. Depending on the specific tasks, one of the interpolation functions can be taken into account in the formulas for calculating the specific heats of CP. In this case, to form the refined geometry of the LRE chamber, the thermo-gas-dynamic calculation is repeated taking into account new indicated dependencies. Consequently, the system of equations for the thermodynamic calculation of an LRE is solved taking into account new functions. This approach allows forming the optimal contour of the LRE chamber at the preliminary stage of engine design and improving the results of gas-dynamic calculation and profiling of the nozzle using a modified method of characteristics. In the framework of the presented studies, to obtain an optimal geometry for the LRE nozzle, are compared values of the velocities, which obtained using the solutions of the direct and inverse problems. Thus, the correct consideration of changes in the basic parameters along the nozzle length allows us to organize the correct operation of the LRE chamber by changing the thermal properties of CP along the nozzle length in all flight conditions of the flight vehicle. This circumstance requires some improvement of the principles and schemes of regulation systems of the LRE operation, which leads to the conduct of extensive researches in this direction.

A method of rib-bed plate enhancing heat transfer in hydrogen rocket engine chamber wall

The hydrogen fuel rocket engine combustion chamber wall is very small and suffers high combustion temperature and heat rates. To effectively reduce the temperature of these overheated structures and improve the cooling performance for the entire channel, different from traditional smooth channel, the rib-bed plate structure is established. The several rib structures are investigated to find optimum height and spacing. The results indicate that the optimum cooling structure with rib-bed plate on the internal wall improves cooling performance. Comparied with the smooth channel, the rib enhances convection heat transfer coefficient and reduces the heat wall maximum temperature. The comprehensive comparisons of cooling performance for different rib structures indicate that the cross-distribution structure with a height 0.2 mm and spacing 2.5 mm is better to improve the cooling performance in a 2 mm*2 mm cross-section channel. Finally, the maximum heat wall temperature is decreased by 14.7%, (Nu/Nunofins)/(f/fnofins)1/3 is increased by a maximum over 25% than traditional smooth channel and cooling capacity is increased to 34.8%.