Rocket Engine Combustion Chamber Research Articles

Regenerative Cooling Comparison of LOX/LCH4 and LOX/LC3H8 Rocket Engines Using the One-Dimensional Regenerative Cooling Modelling Tool ODREC

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.

Thermophysical properties of additively manufactured (AM) GRCop-42 and GRCop-84

Additive manufacturing (AM) has enabled new opportunities of high-performance alloys for use in complex components across the aerospace, industrial, automotive, and energy industries. These industries have growing interest in copper alloys and AM for high-performance heat exchangers such as rocket engine combustion chambers. GRCop copper alloys, such as GRCop-84 (Cu-8Cr-4 Nb at%) and GRCop-42 (Cu-4Cr-2 Nb at%), have high thermal conductivity and high mechanical properties at elevated operating temperatures. In this paper, thermophysical properties for GRCop-84 and GRCop-42 produced using laser powder bed fusion (L-PBF) are examined. While L-PBF parameters and resulting microstructures can be achieved in a laboratory environment, those parameters do not always translate well into a production environment, and the variations across multiple machines and powder chemistries must be understood. The repeatability and reproducibility of samples across the commercial supply chain is critical for designers to ensure parts meet operational requirements. Thermophysical properties, such as thermal conductivity and thermal expansion are key design attributes. Thermal conductivities for eight GRCop-42 blocks and two GRCop-84 blocks were measured from room temperature up to 700 ℃. Thermal expansion was also measured from 20 °C to 1000 °C. These GRCop alloy parts were made with gas atomized powders from different vendors and with different L-PBF systems and parameters. Equations for the thermal conductivity and expansion are given. The variation of thermal conductivities for L-PBF GRCop parts is estimated to be less than + /− 4%. Based on phase analysis, compositional analysis, and thermal conductivity measurements, the small variations in thermal conductivities for the L-PBF GRCop parts are elucidated. The thermophysical property variations can be used to establish process capabilities and design guidelines for ongoing commercial use and further research.

Enhancement of heat transfer and hydrogen fuel flow characteristics in a wavy cooling channel with secondary branch design

The heat transfer and hydrogen fuel flow characteristics in the novel fin structure for the branched-wavy cooling channel of the rocket engine combustion chamber where three different gap widths (wz = 0.4, 0.6, and 0.8 mm) and three different amplitudes were examined numerically. Nusselt number, heat wall temperature, pressure drop, and thermal stratification were studied using Ansys fluent to investigate how configuration, wavy amplitude, and transfer gap width influenced these variables. The findings show that the heat wall temperatures and the average Nusselt number for the original symmetric configuration are better than the parallel one. The modified symmetric wavy design results show that the secondary branch flow considerably enhances Nusselt numbers due to the promoted fluid mixing between neighboring channels; the average Nusselt number of the modified symmetric configuration channel (A = 0.3 mm) enhanced up to 28.23% compared to the original straight configuration when adopting transverse gaps of 0.8 mm. The modified configuration (amplitude of 0.1 mm with a 0.8 mm transfer gap) achieved the highest thermal performance factor, which was enhanced by 21% compared to the original wavy design. Modified symmetric designs can slightly reduce the pressure drop penalty by creating a transverse gap.

Разработка технологии механической обработки для получения турбулизаторов в камерах сгорания жидкостных ракетных двигателей

The paper considers design of the liquid-propellant rocket engine combustion chamber. Artificial roughness application area in the chamber shells is shown, and the existing methods to obtain artificial roughness in the rocket propulsion engineering are analyzed. The paper presents advantages and disadvantages of the existing technologies used in mechanical engineering for surface treatment of products being a part of the liquid-propellant rocket engine combustion chamber. Promising methods for obtaining artificial roughness on the rocket engine parts surfaces are demonstrated. Progressive means of technological equipment are considered to obtain artificial roughness by the mechanical machining on the surfaces of those parts that ensure integrity of the surface layer of the critical structural elements in the liquid-propellant rocket engines. The results obtained make it possible to significantly expand the production technological capabilities, as well as to seriously improve technical characteristics of the special equipment products in the aerospace and space industries.

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.

RANS based numerical simulation of a GCH4/GO2 rocket engine combustion chamber with film cooling and improvement of wall heat flux prediction

Methane/Oxygen rocket engine is becoming one of the most promising rocket engines today due to its cost-effectiveness and reusability. In the design process of rocket engines, cooling system is a crucial part and film cooling is a very important method. The accurate prediction of heat transfer characteristics is crucial for the design and development of rocket engines using film cooling. In this paper, a numerical framework based on the Reynolds Averaged Navier-Stokes (RANS) method and the Eddy Dissipation Concept (EDC) reaction model is established, verified and applied to simulations of single- and multi-injector combustion chamber with film cooling. Besides, a single injector combustion experiment with film cooling is carried out to verify the numerical framework. The investigation indicates that flow and chemistry reactions near the wall coupled influence the wall heat load significantly, and the coupled wall function exploited by Direct Numerical Simulation (DNS) is modelled and embedded on the numerical frame in order to consider these coupling effects. The results of single injector chamber investigation show that by considering the chemical reactions near the wall the wall heat flux reduced 50% and agree much better to the experimental data, which indicates that coupled wall function is more effective at predicting wall heat flux than general wall functions in a chamber with film. In addition, the results also denote that the coupled wall function only acts in the near-wall region and has no effect on the main flow. Furthermore, after being verified in the single injector combustion chamber experiment, the numerical framework is applied to a multiple-injector case. The results indicate that the wall heat flux in multi-injector combustion chamber is 75% lower than the general wall functions, Afterwards, the effect of the film on the chemical enthalpy term near the wall, as well as the coupling effect of the turbulent flow and the temperature gradient near the wall are discussed. Finally, the analysis of the vorticity in the multi-injector chamber shows that the film weakens the vorticity in the front section of the combustion chamber, and subsequently affects the expansion of the flame.

Numerical method for calculating the nozzle thrust of a wide-range rocket engine

A numerical method for calculating the thrust of a wide-range rocket engine nozzle has been developed. This type of engine is equipped with an annular nozzle with a flat central body and is designed to operate in the upper atmosphere and in vacuum. The nozzle forms a jet converging to the axis of symmetry due to which a more compact torch of the working fluid is formed. Nozzles of this type have important design advantages over conventional external expansion nozzles. They are more compact, simpler in terms of cooling, but they have increased losses in the bottom region due to the presence of a flat bottom near the central body. Therefore, the design of such engines needs parametric optimization. Currently, for engines equipped with an annular nozzle with a flat central body, there are no validated methods that would allow parametric optimization. The characteristics of the jet, the loss of specific impulse, and the magnitude of thrust for a given type of nozzle depend on three main parameters: the area of the bottom region of the central body, the area of the throat section, and the angle of rotation of the inner edge of the nozzle to the axis of symmetry. The gas flow in the bottom region is accompanied by complex shock-wave processes that require a lot of time for numerical calculations. To optimize the design of the nozzle, it is required a simple engineering method to calculate the thrust of the nozzle according to its main parameters. The calculation of the nozzle thrust is based on the integral distribution of pressure forces over its surface obtained by performing numerical calculations in a wide range of external pressure. The Reynolds-averaged Navier-Stokes equations closed by the SST-modification of the k-ω turbulence model are solved. Based on the results of numerical simulation, the calculated coefficients for one-dimensional engineering dependencies are determined; they make it possible to calculate the speed and pressure in a random section of the combustion chamber and engine nozzle. A simple engineering method for calculating the thrust of the chamber nozzle of a wide-range rocket engine has been developed. The technique is verified by comparison with the results of a numerical experiment. The problem of parametric optimization of the rocket engine combustion chamber, capable to operate in a wide range of altitudes, was solved, and it is of interest to the space industry. The developed method of calculation makes it possible to carry out a wide range analysis of the influence of the ratio of geometric dimensions, regime parameters on the thrust of the combustion chamber and nozzles of a wide-range rocket engine, and to estimate the thrust value at different engine operating heights.

Расчетное исследование особенностей рабочего процесса в камере сгорания кислородно-метанового жидкостного ракетного двигателя, работающего по схеме с дожиганием восстановительного генераторного газа

We performed a numerical investigation of cumulative efficiency and the structure in detail concerning the working process in the combustion chamber of a lox/methane liquid-propellant rocket engine operating in steady-state, boosted and throttled modes. In order to do it, we used tools developed by JSC SSC "Center Keldysh", that is, physical and mathematical models, numerical methods and software packages for numerical simulation of two-phase turbulent flows with combustion in liquid-propellant engine combustion chambers. The paper presents numerical simulation and investigation results concerning the specifics of fuel component flows, their mixing and combustion in the combustion chamber of a lox/methane liquid-propellant rocket engine using staged combustion cycle with reductant gas in steady-state, boosted (117 % by thrust) and throttled (30 % by thrust) operation modes. We performed a comparative analysis of work cycle parameters in combustion chambers at different fuel component consumption rates and pressure levels. The paper shows that the boosted mode increases the interaction of fuel jets, which intensifies mixing and burnout processes, while the deep throttling mode decreases the mixing and fuel burnout amplitudes as compared to the steady-state mode. The numerical simulation results may be used to investigate fuel combustion processes in combustion chambers of promising liquid-propellant rocket engines at the stages of development, design and refinement

Coaxial-Injector Surrogate Modeling Based on Reynolds-Averaged Navier–Stokes Simulations Using Deep Learning

Facing the need to increase the accuracy of rocket engine design tools, the present work introduces an innovative methodology for the design and optimization of rocket engine combustion chambers using numerical simulations and deep learning. An experimental test case of a single coaxial injector is taken as a reference point, and a design of experiments is generated by varying nine parameters (geometrical and operative conditions). Reynolds-averaged Navier–Stokes simulations are carried out to generate the data set. The data are used to train surrogate models of different fidelities, from low-dimensional outputs (zero-dimensional and one-dimensional) toward the two-dimensional temperature field. Attention is given on the selection of the proper machine learning technique. For low-dimensional outputs, results show that deep neural networks outperform other standard machine learning tools, namely, radial basis function and kriging. Regarding high-dimensional outputs, convolutional neural networks with gradient-based loss functions are found effective to capture the large and smooth temperature variations, as well as the thin and sharp temperature gradients at the flame front. Eventually, the models are used in the framework of an optimization problem. Results highlight the benefits of new design and optimization tools based on deep learning, capable of real-time predictions of complex flowfields.

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.

Remaining Useful Life Prediction with Uncertainty Quantification of Liquid Propulsion Rocket Engine Combustion Chamber

Reduction of spaceflight costs calls for development of new technologies that render rockets reusable. This new requirement and the continuous improvement of rocket engines require pro-active approach towards the possibility of integrating health monitoring systems on-board. These health monitoring strategies should also take into consideration the state of degradation and the remaining useful life prediction. In this paper, an Extended Kalman Filter is used to estimate the state of health and the dynamics of the degradation, and the remaining useful life is predicted with respect to failure thresholds pre-set by the user. The first-order inverse reliability method is employed to assess the quality of the remaining useful life prediction by quantifying the associated uncertainty. The overall method is validated using simulation study involving degradation data provided by Centre National d'Etudes Spatiales (CNES) applied to liquid propulsion rocket engine (LPRE) combustion chamber.

МАЛОЭМИССИОННЫЕ КАМЕРЫ СГОРАНИЯ И СПОСОБЫ ОХЛАЖДЕНИЯ

A modern gas turbine engine must meet a large list of requirements that are included in the parameters, resource and per-formance indicators. To increase the service life of a gas turbine engine at elevated temperatures of the gas flow, it is expendable to use thermal barrier protection on explosive structural materials. Cyclic tests of materials and thermal barrier coatings of gas tur-bine engines at temperatures above 1500 ºС are proposed to be carried out on a stand in which a hot gas flow is generated by an air-methane burner. In order to reduce the emission standards for nitrogen and carbon oxides, it is necessary to develop and use in stationary gas turbine engines fundamentally new technologies for organizing combustion and, as a result, designs of combustion chambers. From a detailed analysis of the current requirements, it follows that the newly designed low-emission combustion chamber for advanced gas turbine engines and installations should be accompanied by an increase in gas temperature by 200–300 K, an increase in the durability of the flame tube by 3–4 times, with a twofold decrease in the proportion of air for cooling the walls, a twofold or more reduction in the emission of harmful substances. In this article, heat-resistant coatings of structural elements of gas turbines are considered. The concepts of low-emission fuel combustion are described by organizing the working process according to the "DLE" - Dry Low Emission scheme. As an alter-native method for organizing low-emission combustion, stoichiometric combustion is proposed, which also makes it possible to provide the required temperature of the gas jet. A review of low-emission combustion chambers has been carried out. The existing methods of cooling the combustion chambers of gas turbine and liquid rocket engines are described. The analysis of the collected information made it possible to determine the concept of designing a high-temperature air-methane burner.

Reducing the Aerodynamic Drag of the Mating Sections of the Combustion Chambers of Rocket Engines

Technological methods for manufacturing cooling elements of modern rocket engines are considered. They are developed taking into account the possibility of reusable use, which reduces the cost of manufacturing similar products. It is shown that in this case, the thermal load on the walls of combustion chambers of liquid rocket engines increases significantly. This required the creation of new ways to protect the surface layer of the hot zone from the thermal effects of the flame in the fuel combustion zone.

Advanced active-gas 3D printing of 436 stainless steel for future rocket engine structure manufacture

During this study a novel means to fabricate a rocket engine combustion chamber was developed. Here the shape of this rocket engine component has been fabricated using prototype 3D metal active gas deposition technology. The manufacturing process is based upon an argon gas fused deposition process for producing complex component geometries. This process uses multi-pass deposition employing stainless steel 436 alloy as the feed material. The structures created during this process were metallographically characterised to understand the microstructural properties produced during fabrication. The engine combustion chamber casing was fabricated and later machined to produce a smooth precision finish. The chamber casing features high resolution details with fine control of the process resulting in micron-scale precision. This allowed for the creation of wall structures and hollow internal sections. It was by allowing the structure to cool precisely between each layer that resulted in no loss in precision due to imposed cooling stresses on this large-scale structure. It is this factor which is of significant importance towards the formation of complex components. The process developed was a paramount means towards the fabrication of robust structures.

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.

Contribution to the physical description of supercritical cold flow injection: The case of nitrogen

While increased pressure and temperature contribute to an overall efficiency gain in the mixing of propellants and oxidizers, characteristic of conditions in the combustion chambers of liquid rocket engines, they also propel mixtures to trans- and supercritical conditions. In these conditions, the engine flow exhibits a gas jet-like behavior that may be described using an approach developed for variable density incompressible flows. The present study focuses on an approach using the Reynolds-averaged Navier–Stokes equations to evaluate the jet topology for different injectors’ conditions. Based on the so-called ’thermal breakup mechanism concept’ proposed in the literature, the axial density decay in supercritical nitrogen jets is predicted for a wide range of conditions. The results show the influence of thermal breakup, providing a better insight of the available experimental data.

Chemical modeling for methane oxy-combustion in Liquid Rocket Engines

Methane–oxygen burning is considered for many future rocket engines for practicality and cost reasons. As this combustion is slower than hydrogen–oxygen, flame ignition and stability may be more difficult to obtain. To address these questions, numerical simulation with realistic chemistry is appropriate. However the high pressure and turbulence intensity encountered in rocket engines enhance drastically the stiffness of methane oxy-combustion. In this work, Analytically Reduced Chemistry (ARC) is proposed for accurate chemistry description at a reasonable computational cost. An ARC scheme is specifically derived for typical rocket engine conditions. It is validated by comparison with its parent skeletal mechanism on a series of laminar flames. Then the numerical stiffness of chemistry is overcome with an original approach for time integration, allowing to run simulations close to the acoustic time step whatever the chemical stiffness. It is demonstrated on laminar cases that the flame structure is well preserved, and that numerical stability is ensured while decreasing significantly the computational cost. The performance of ARC with the fast time integration method is finally demonstrated in a 3D Large-Eddy Simulation of a lab-scale Liquid Rocket Engine combustion chamber, where a detailed flame analysis is conducted.

Simulation of non-stationary gas dynamics of solid propellant rockets launch

The article presents the results of the development of methodology for the calculation of non-stationary gas dynamics processes occurring in gas dynamic paths of rocket engines and environment at the launch of rockets. The method takes into account the change of geometry of solid fuel combustion surface during the operation of the engine and the change in the geometry of the computational domain taking into account the dynamics of rocket launch. Numerical simulation of gas-dynamic processes of the launch of model solid-fuel rocket was done. The unsteady gas-dynamic flow pattern was investigated. The pressure curve in the solid propellant rocket engine combustion chamber, the speed of movement and overload were determined.

High-temperature diffusion in couple of electrodeposited iridium/rhenium

Ir coated Re (Ir/Re), combining the outstanding high-temperature mechanical property of Re and excellent high-temperature oxygen permeability resistance of Ir, is a promising high-temperature material in high temperature applications, such as rocket engine combustion chambers and crucibles for crystal growth, etc. The elemental diffusion behaviors in the Ir/Re material are closely related to its failure mode and lifetime. Therefore, the diffusion behavior investigation on the Ir/Re material is essential for understanding the failure mechanism and building the lifetime prediction model of this material. In this work, a three-layered CVD Re/ED Ir/ED Re sample was fabricated and its diffusion behaviors at the temperatures ranging from 1400 to 2000 °C were investigated. The diffusion coefficients and activation energy of ED Re diffusing into Ir using a semi-infinite diffusion model were obtained and compared with those of the CVD Re. The obviously lower diffusion activation energy of ED Re (0.98 eV) compared with the CVD Re is due to their microstructure difference. The higher defect concentration, thermodynamically unstable grain structure and preferred orientation of the ED Re layer promotes the dominant diffusion of Ir into Re, resulting in a thicker diffusion layer with a frame-shaped diffusion front.

An efficient modeling framework for wall heat flux prediction in rocket combustion chambers using non adiabatic flamelets and wall-functions

In this work an efficient numerical framework for the prediction of wall heat loads in Liquid Rocket Engine combustion chambers is presented. The proposed framework is based on a new version of the non-adiabatic flamelet model and on wall functions for turbulent boundary layer modeling. Different wall function models are applied to 2D and 3D wall heat flux simulations of an experimental single-element gaseous oxygen-gaseous methane combustor in an Unsteady Reynolds Averaged Navier Stokes context. A systematic analysis and a comprehensive comparison of the selected wall models is carried out. The role of the constant or variable properties assumption on the near-wall turbulent quantities affecting the wall heat flux is assessed and the resulting friction velocity scaling investigated. When the skin friction velocity based on the local turbulent kinetic energy is defined by considering constant properties across the boundary layer, the equilibrium boundary layer assumption is not fulfilled and a significant overestimation of the wall heat flux is observed. Results obtained with the corrected near-wall turbulence modeling, on the other hand, showed a substantial improvement in terms of wall heat flux when compared with both experimental data and higher fidelity simulations results.