Rocket Combustion Chamber Research Articles

Combustion flow field and thermal protection performance of a liquid oxygen/kerosene engine with pintle injectors

A pintle injector is a convenient method for adjusting propellant combustion and has received extensive attention in recent years. However, few studies have examined the effects of the injection direction and momentum ratio (ratio of liquid oxygen momentum to kerosene momentum) on the flow field, temperature field, and dual-belt liquid film cooling effect in the combustion chamber. This study verified a new computational fluid dynamics (CFD) engine combustion model based on a pintle injector through a hot-test experiment. A simulation was conducted using this model. With different injection directions, when the momentum ratio was increased, the trends in the combustion efficiency and chamber pressure growth were entirely opposite. For the axial kerosene and radial liquid oxygen injection direction of liquid film cooling, when the momentum ratio was 0.73–0.9, the proportion of the first ring belt liquid film flow was selected as 5–10 %, and the second ring belt liquid film flow proportion was selected as 5–7%. When the momentum ratio was less than 0.73, the liquid film flow proportions for both the first and second ring belts were selected as 5 %. When the total liquid film flow proportion was 10 %, the chamber pressure decreased by 0.2–0.3 MPa, and the combustion efficiency decreased by 8–10 %. This numerical study provides guidance for studying the combustion characteristics and cooling performance of pintle-type rocket combustion chambers.

Influence of Aluminum Content and Agglomerates Initial Velocity on Erosion in Solid Rocket Motor

Abstract Despite the aluminized propellants offering a high specific impulse, the challenge of nozzle erosion adversely impacts the rocket's performance and its potential for reusability. This study presents a numerical model aiming to predict the mechanical erosion of the propulsion chamber nozzle. The model employs an Eulerian/Lagrangian approach to simulate the complexity of the flow field within the rocket combustion chamber and the interactions between the continuous phase and particles. The model also includes a simplified representation of the aluminum particle combustion process, besides the consideration of secondary breakup phenomena in liquid droplets. Experimental and numerical data from the literature were used to validate the numerical model. Subsequently, the model was utilized to explore the impacts of increasing propellant aluminum content and varying particles' injection velocities on the nozzle mechanical erosion. The outcomes indicated that higher aluminum content leads to a 4-10% increase in nozzle erosion compared to the 15% content case. Furthermore, the aluminum particles tend not to fully burn within the combustion chamber and contribute to nozzle erosion. Lastly, particles with higher initial velocity at the inlet of the combustion chamber increase the nozzle mechanical erosion despite the observed decrease in incident mass flux.

Study on Densification and Defects of Cu-Cr-Zr Alloy Formed by Selective Laser Melting

Abstract Selective laser melting (SLM) technology features high forming accuracy and a short preparation cycle, providing a significant advantage for shaping complex structural components. Copper-chromium-zirconium exhibits outstanding comprehensive performance and finds wide applications in rocket combustion chambers, automotive inductor coils, nuclear shielding materials, welding electrodes, and other areas. However, due to its low laser absorption rate and high thermal conductivity, the forming process of copper-chromium-zirconium is unstable and prone to forming defects. This study aims to address this issue by employing the response surface method to design process parameters and investigate the impact of these parameters on microstructure and properties, to obtain high-density Cu-Cr-Zr alloy.

Effect of laser power and deposition sequence on microstructure of GRCop42 - Inconel 625 joints fabricated using laser directed energy deposition

The joining of Inconel 625 and GRCop42 using additive manufacturing is required for thermal management of high operating temperature components such as rocket combustion chambers. Though prior efforts have been made to use laser directed energy deposition to join these materials, the impact of laser power and deposition sequence on microstructure of these joints is not well understood. In this study, Inconel 625 onto GRCop42 and GRCop42 onto Inconel 625 joints are fabricated by powder directed-energy-deposition at various laser powers and subsequently subjected to characterization in terms of present defects, grain morphology, and phases. Results show lack-of-fusion free Inconel 625 onto GRCop42 joints can be fabricated by increasing the laser power in the first layer. Substrate remelting in GRCop42 onto Inconel 625 joints is found to result in a melt pool composition which induces liquid-state immiscibility, resulting in a Cu-deprived liquid solidifying to form crack prone islands. Increasing laser power decreases the embrittlement of these islands due to the precipitation of a lower volume fraction of intermetallic phases.

Investigation and assessment of production- and dissipation-limited DDES methods for rocket combustion chamber simulations

This paper compares two different delayed detached-eddy simulation (DDES) techniques, namely iDDES and l2-ω-DDES. Their performance is investigated with and without combustion. While iDDES limits the dissipation term of the turbulent kinetic energy equation, l2-ω-DDES confines its production term. In this way, both models achieve that resolved turbulence is enhanced in regions of separated flow. On the other hand, the models show significant differences. In a first step this is investigated using three non-reactive test cases. While here the first-order moments are in a fairly good agreement, the second-order moments exhibit slight differences. Despite the similarities in first- and second-order moments, the instantaneous flow fields differ considerably with the iDDES displaying finer vortex structures. This is caused by different levels of eddy viscosity. These disparities are even more pronounced in case of a laboratory-scale rocket combustion chamber, where strong differences in quantities as the wall heat flux are observed. This clearly demonstrates an impact on the combustion process.

Combustion of Date Stone and Jojoba Solid Waste in a Hybrid Rocket-like Combustion Chamber

The performance of two solid biomass wastes, date stone and jojoba solid waste, was experimentally examined for their potential application in combustion and propulsion systems. The fuels were tested in a hybrid rocket-like combustion environment, and the test result was analyzed with combustion and propulsion parameters. The performance of both fuels was comparatively evaluated and compared with a conventional hydrocarbon fuel in a hybrid rocket, with paraffin wax serving as a baseline. A compression device was introduced to compress the solid biomass wastes into a circular-shaped fuel grain compatible with a hybrid rocket combustion chamber with a hot surface ignitor. Thermogravimetric analysis (TGA) and chemical equilibrium analysis (CEA) results revealed that the performance of the biomass fuel can be comparable to conventionally used hydrocarbon paraffin-wax-based propellant within a certain range of oxidizer-to-fuel ratio, in terms of theoretical specific impulse performance. Through experimental performance tests, it was found that the compressed biomass fuel grains were successfully ignited and produced thrust. Both biomass fuels tested in a hybrid rocket combustion chamber are expected to pave the way for further developments in biomass fuels in the waste-to-energy field for their application in combustion and propulsion systems, potentially replacing fossil fuels with renewable resources.

Microstructural Stability of a Metallic Thermal Barrier Coatings for Rocket Engine

Metallic thermal barrier coatings (TBCs) consisting of a bond coating and a top coating have been extensively utilized for protecting the walls of rocket combustion chambers. However, standard coating systems often encounter failures due to the significant differences in coating composition and thermal expansion coefficient compared to the substrate under high heat flux conditions. To protect liquid rocket combustion chamber walls, a novel metallic multilayer TBC system applied with atmospheric plasma spraying is developed in the present work. It attempted to deposit dense Ni-based alloy and Cu-based bonding coatings with low oxide contents achieved by introducing boron as a deoxidizer element through atmospheric plasma spraying. The structural stability of the TBC was assessed through high temperature thermal exposure experiments, while the thermal cycle life is evaluated using laser thermal shock. Results show that the NiCrCu2B and CuNi2B bonding coatings prepared through in situ deoxygenation effect of boron exhibit dense structures, low oxide content, and excellent bonding quality. The high temperature thermal exposure experiment reveals that the multilayer structural TBC can withstand 850 °C for 10 hours without the formation of Kirkendal effect pores. Moreover, the thermal cycling life results indicate that the multilayer structural TBC designed in this study, employing a composition gradient transition and the in situ deoxygenation effect of boron, possesses a significantly improved thermal cycle lifetime compared to traditional structural TBCs.

Improved Wall Temperature Prediction for the LUMEN Rocket Combustion Chamber with Neural Networks

Accurate calculations of the heat transfer and the resulting maximum wall temperature are essential for the optimal design of reliable and efficient regenerative cooling systems. However, predicting the heat transfer of supercritical methane flowing in cooling channels of a regeneratively cooled rocket combustor presents a significant challenge. High-fidelity CFD calculations provide sufficient accuracy but are computationally too expensive to be used within elaborate design optimization routines. In a previous work it has been shown that a surrogate model based on neural networks is able to predict the maximum wall temperature along straight cooling channels with convincing precision when trained with data from CFD simulations for simple cooling channel segments. In this paper, the methodology is extended to cooling channels with curvature. The predictions of the extended model are tested against CFD simulations with different boundary conditions for the representative LUMEN combustor contour with varying geometries and heat flux densities. The high accuracy of the extended model’s predictions, suggests that it will be a valuable tool for designing and analyzing regenerative cooling systems with greater efficiency and effectiveness.

Model for Combustion Instability in Hybrid Rocket Engines Burning Liquefying Fuels

Starting from the data measured in several static firings of two laboratory-scale hybrid rocket engines of different sizes burning gaseous oxygen with either a liquefying fuel or a pyrolyzing polymer, a mechanism for the combustion instability observed only with liquefying fuel is suggested. With the latter type of fuels, regression rate is increased for the formation of an unstable liquid layer on the grain surface, for which liquid fuel structures are entrained into the gas stream and burn, upon vaporization and mixing, far from the surface. Based on this circumstance, the hypothesis of a time delay to burn, with the resulting effect of triggering longitudinal acoustic modes of the system, is formulated. A simplified analytical acoustic model of a series of interconnected ducts with different cross sections and gas properties is preliminarily addressed, which allowed demonstrating the acoustic nature of the main pressure oscillation peaks measured in three firing tests along with the influence of the oxygen-feeding line on the system acoustics. Finally, by leveraging an existing acoustic model of the hybrid rocket combustion chamber including the effects of cross-sectional discontinuities and distributed fuel mass addition and combustion, the linear stability of the system is analyzed, whose results are apt to reproduce the experimental findings.

Effect of gas injection holes on mixing and combustion in solid propellant ducted rockets

Combustion efficiency is an important parameter of the combustion chamber. Gas injection holes have an important influence on the mixing and combustion in solid propellant ducted rockets, but there is currently little research on gas injection holes. This paper describes numerical simulations of solid propellant ducted rocket combustion chambers, with the numerical method verified through a direct-connected experiment. We investigate the effect of adding injection holes to the gas generator nozzle on the mixing and combustion performance, and examine the mixing degree, combustion efficiency, and flow field characteristics. This paper focuses on the number and the position of the gas injection holes. The results show that, when the total mass flow rate is constant, the mixing effect becomes better as the dispersion degree of gas injection increases and the mass flow rate of each hole decreases. For this study, the 5-hole scheme has the best performance, and when the total mass flow rate and the number of injection holes are fixed, a smaller mass flow rate of gas injected toward the outlet of each air inlet is more conducive to better mixing and combustion effects. When the total mass flow rate is constant and the shape of the injection holes are complete circles, increasing the number of holes is more effective than decreasing the mass flow rate of gas injected toward the outlet of the air inlets in improving combustion efficiency.

Experimental studies on combustion performance of beeswax-paraffin blended solid fuels in a hybrid rocket

The study intends to investigate the physical, chemical and thermal characteristics of paraffin blended fuels to determine their suitability as a solid fuel in a hybrid rocket. Wax fuels are a viable and efficient alternative to conventional rocket fuels, having excellent structural strength and thermal and mechanical properties. By utilizing both axial and swirl injection technique, the combustion performance of paraffin – beeswax blended fuels have been tested with a fabricated cylindrical grain in a laboratory-scale rocket setting along with oxygen. The test outcomes revealed solid fuel compositions of more beeswax content in paraffin wax on an oxygenated gaseous environment with a swirl-flow injection method has the highest average regression rate of 1.649 mm/sec at 181 kg/m2s mass flux. Axially injected oxygen with pure paraffin wax has the lowest value of 0.85 mm/sec at 96 kg/m2s. The regression rate comparisons revealed that oxygen injection by a swirl injector increased the regression rates by 40% for mass fluxes greater than 80 kg/m2s. Compared to other studies, the combustion efficiencies have been obtained in this study are good. Blended fuels can manage and increase combustion efficiencies for axial and swirl flow conditions. Swirl injectors outperform axial injectors for oxygen injection and allow for a higher proportion of Beeswax combined with paraffin. This study exclusively designed and manufactured an axial injector and swirl injector, according to the required dimensions of a lab-scale hybrid rocket's combustion chamber, injector, and exhaust nozzle, and their performances have been evaluated.

Investigation of reacting fuel jets in hot vitiated crossflow

Near-wall reacting jets can occur in different engineering applications, such as rocket combustion chambers and in gas turbine combustors and turbines. This work presents results of experimental investigations of reacting gaseous fuel jets (hydrogen, methane, propane) injected normally into a hot oxygen-rich cross-flow using OH laser-induced fluorescence and OH* chemiluminescence diagnostics. RANS CFD simulations with EDC combustion model are performed for comparison. To determine the surface heat fluxes, the inverse heat conduction method was applied using measured wall temperature data as input. It was observed, that hydrogen jets ignite directly at the injection location and caused a higher surface heat flux than the hydrocarbon jets, where jet-like diffusion flames emerged. With the latter fuels, a significant ignition delay length was observed. Operation with methane fuel showed the least stability, least reactivity and the lowest surface heat flux. Increasing the jet momentum ratio led to a reduced ignition delay length due to enhanced jet-crossflow interaction, and to a reduction of surface heat flux as the reacting jet detached more from the surface. CFD results using the Eddy-Dissipation Concept (EDC) combustion model predicted ignition delay lengths and reaction zone shapes in qualitative agreement with the experiment but with a trend of overestimating combustion temperatures.

Heat Transfer in Rocket Combustion Chambers Firing Plates: Role of Injector Confinement

In this work the effect of the injector lateral confinement on flowfields and thermal loads in the injection region of a rocket combustion chamber is investigated. The reference test case consists of a gaseous-methane/gaseous-oxygen turbulent, non-premixed flame emanating from a single shear coaxial injector. Two baseline configurations are defined in order to model different situations encountered within a hypothetical multi-injector plate: a wall-bounded configuration is chosen to investigate the injector–wall interaction, while a symmetric condition is used to mimic the mutual interaction between injectors located in the inner part of the injection plate. The effect of the lateral confinement length is assessed by means of a parametric analysis, varying this feature in a range of interest for both configurations. The analysis is performed using two-dimensional axisymmetric simulations, carried out by means of a low-Mach number, unsteady Reynolds-averaged Navier–Stokes framework employing a non-adiabatic flamelet-based turbulent combustion model. It is found that the confinement length mainly affects the recirculation region, causing wider and warmer recirculation with increasing confinement lengths and therefore increasing thermal loads in the injection plate region.

Selection Rules for Resonant Longitudinal Injector-Coupling in Experimental Rocket Combustors

This paper investigates different types of longitudinal mode coupling in subscale rocket combustion chambers using experimental data and numerical simulations. Based on a one-dimensional planar wave acoustic model of coupled cavity resonators with two acoustic inlet boundary conditions, mode selection rules are derived, providing a simple way of predicting which injector and combustion chamber modes have matching frequencies. Longitudinal mode coupling of an injector with an acoustically open inlet boundary condition has been reported in the literature for the start-up transient of a research combustor experiment. In this experiment, every third injector mode couples to a corresponding chamber longitudinal mode, which is explained in terms of the selection rules derived in this paper. Numerical simulation results for a different combustor experiment show an unexpected mode coupling behavior when an acoustically closed injector inlet is used. Theoretical analysis by using the one-dimensional wave model and applying the derived selection rules shows that in this setup, the injector acoustic mode can accommodate two different acoustic boundary conditions at the injector-chamber interface simultaneously. This results in different acoustic mode shapes in the injector, explaining the unexpected behavior for the resonant coupling with an acoustically closed injector inlet.

Efficient time-resolved thermal characterization of single and multi-injector rocket combustion chambers

In this work, an efficient methodology for the time-resolved thermal characterization of rocket combustion chambers at reasonable computational cost is presented. The multi-scale and multi-physics numerical framework tackles simultaneously an arbitrary number of contiguous domains, either fluid or solid, and takes advantage of several modeling solutions aimed at stiffness reduction. Non-premixed turbulent combustion is handled through a flamelet-based approach accounting for non adiabatic and non equilibrium effects, thermal wall functions adapted for rocket operating conditions are employed to overcome the stiffness induced by the boundary layer, and a coupling strategy is implemented to guarantee temperature and heat flux continuity across the interfaces. The coupling strategy is based on a Conjugate Heat Transfer (CHT) condition, yielding the interface temperature as a result of a heat flux continuity constraint, and is then reformulated for convection-dominated phenomena, allowing for a further reduction of the computational cost. This allows for the simulation of long time windows, of industrial and experimental relevance. In particular, the solution of the chemically reactive flow is initialized with a CHT condition, and replaced, upon attainment of a statistical fluid dynamic steady state, by an equivalent convective boundary condition. The numerical framework is validated and tested by means of several 2D and 3D cases, the latter consisting in both single-element and multi-element experimental combustor chambers operating in rocket-like conditions.

Probabilistic constrained Bayesian inversion for transpiration cooling

AbstractTo enable safe operations in applications such as rocket combustion chambers, the materials require cooling to avoid material damage. Here, transpiration cooling is a promising cooling technique. Numerous studies investigate possibilities to simulate and evaluate the complex cooling mechanism. One naturally arising question is the amount of coolant required to ensure a safe operation. To study this, we introduce an approach that determines the posterior probability distribution of the Reynolds number using an inverse problem and constraining the maximum temperature of the system under parameter uncertainties. Mathematically, this chance inequality constraint is dealt with by a generalized polynomial chaos expansion of the system. The posterior distribution will be evaluated by different Markov chain Monte Carlo based methods. A novel method for the constrained case is proposed and tested among others on two‐dimensional transpiration cooling models.

Pseudo-Two-Dimensional Modeling and Validation of a Hybrid Rocket Combustion Chamber

A pseudo-two-dimensional (pseudo-2-D) model for the cylindrical combustion chamber of a hybrid rocket is developed in this paper with the purpose of its integration in a system design tool used for engine predesign phases. A compromise of simplicity and accuracy is thus required. Integral transient boundary-layer equations are included and coupled with the mass and energy balances at the fuel surface, where fuel pyrolysis is described by an Arrhenius law. Studies assessing the model in the nonreactive case provide satisfactory results regarding the description of the boundary layer. Concerning the hybrid case, the computed steady-state fuel regression rates are found to be in the low region of the reference values coming from semiempirical laws for several oxidizer/fuel couples and oxidizer mass fluxes, producing errors up to 50% for certain sources. The effect of mass flux and motor size is correctly simulated. Eventually, the whole engine operation is assessed through the comparison of the computed results with three experimental tests realized at our laboratory and using an 87.5% hydrogen-peroxide/high-density-polyethylene couple. Corresponding results lead to regression rate errors varying in a wide range (from 10 to 40%). Such differences, although high, remain acceptable for our engine predesign phase application.

Characterization of a new ultra-high pressure shock tube facility for combustion and propulsion studies.

A new shock tube facility has been designed, constructed, and characterized at the University of Central Florida. This facility is capable of withstanding pressures of up to 1000 atm, allowing for combustion diagnostics of extreme conditions, such as in rocket combustion chambers or in novel power conversion cycles. For studies with toxic gas impurities, the high initial pressures required the development of a gas delivery system to ensure the longevity of the facility and the safety of the personnel. Data acquisition and experimental propagation were implemented with remote access to ensure safety, paired with a LabVIEW- and Python-based user interface. Thus far, test pressures of 270 atm, blast pressures of 730 atm, and temperatures approaching 10 000K have been achieved. The extreme limitations of this facility allow for emission spectroscopy to be performed during the oxidation of fuel mixtures, e.g., alkanes diluted in argon and carbon dioxide. Ignition delay times were determined and compared to simulations using chemical kinetic mechanisms. The design, experimental procedures, processes of analysis, and uncertainty determination are outlined, and typical pressure profiles are compared with a new gas dynamics solver and empirical correlations developed across multiple shock tube facilities. Preliminary reactive mixture analyses are included with further investigation of the mixtures outlined.

Assessment of thrust chamber stability margins to high-frequency oscillations based on mathematical modeling of coupled ‘injector – rocket combustion chamber’ dynamic system

High-frequency instability of a liquid-propellant rocket engine (LRE) during static firing tests is often accompanied by a significant increase in dynamic loads on the combustion chamber structure, often leading to a chamber destruction. This dynamic phenomenon can also be extremely dangerous for the dynamic strength of a liquid-propellant rocket engine. The calculation of acoustic combustion product oscillation parameters is important in the design and static firing tests of such rocket engines. The determination of the oscillation parameters (natural frequencies and stability margins such as oscillation decrement) is one of the problems solved in the LRE design period as part of the development of measures to ensure the engine stability. The main aim of the paper is to develop a numerical approach to determining the parameters of acoustic oscillations of combustion products in liquid-propellant rocket engines combustion chambers taking into account the features of combustion space configuration and the variability of gaseous medium physical properties depending on the axial length of the chamber, acoustic impedance in critical throat and dissipation effects (damping experimental values) in the shell structure and the gas media in the chamber. The approach is based on mathematical modeling of the coupled ‘chamber shell structure – gas’ dynamic system by using the finite element method and the CAE (Computer Aided Engineering) system. The developed approach testing and further analysis of the results for the RD 253 engine using nitrogen tetroxide and unsymmetrical dimethylhydrazine as a propellant pair were carried out. The dynamic system shapes and frequencies of longitudinal, tangential and radial modes are determined. The results of mathematical modeling of the dynamic system indicate a satisfactory agreement of the calculated decrements of the first longitudinal oscillation mode and third tangential oscillation mode with the experimental decrements obtained by hot-fire tests data. From system harmonic analysis of the thrust chamber, it follows that the dynamic pressure gain factor of the gas media in the chamber at the first longitudinal mode frequency is 1.6 times greater than the system dynamic gain in the tangential mode. At the same time, the oscillation decrement for the system tangential mode is 2 times smaller than that of the first longitudinal mode. This means that the thrust chamber tangential mode is more dangerous and can lead to rocket engine combustion instability. The effect of the injector on the high-frequency stability of the combustion chamber and the possibility of partial suppression of combustion chamber thermoacoustic oscillations by adjusting the high-frequency dynamics of the injector are shown theoretically.

Effect of Protrusion on Combustion Stability of Hybrid Rocket Motor

AbstractIn hybrid rockets, low regression rate and low combustion efficiency are major drawbacks restricting its utilization in practical applications. Inserting a protrusion in a hybrid rocket combustion chamber is one of the methods to enhance the regression rate and combustion efficiency. In the present study, while utilizing the protrusion method the combustion stability of the hybrid rocket was investigated. The protrusions used for this study were of two different types; one was made of graphite and another one was nylon. The nylon protrusion can regress during the combustion along with the fuel whereas, graphite cannot. Before using the protrusion method, the combustion stability of the hybrid rocket motor (base case) was characterized. During the combustion of the base case motor, intense pressure oscillations were observed for a definite period after the beginning of the combustion. These oscillations were found to be the result of the interaction between the vortex shedding and fundamental longitudinal acoustic modes of the chamber at the fuel grain exit. Upon inserting the protrusion near the fuel grain exit, significant suppression in the pressure oscillations was observed. A 96 % reduction in amplitude of acoustic modes relative to the base case was observed. There was no substantial difference in combustion stability of the hybrid rocket when using graphite or nylon protrusions. Further, it was realized that the combustion stability of the rocket motor remains unchanged even after varying the thickness of the protrusion at 0.8X/L location.