Combustion Instability In Rocket Engines Research Articles

Early detection of combustion instabilities in liquid rocket engines: Assessment of statistical, recurrence, and fractal analyses for sub- and supercritical pressure conditions

Safe shutdown of a liquid rocket engine in ground testing prior to the onset of damaging combustion instabilities through reliable detection of instability precursors would translate to time and cost savings in engine development programmes. Methods derived from statistical, recurrence, and fractal analysis have been successfully applied in the literature to detect precursors in unsteady pressure signals from canonical combustion experiments, gas-turbine combustion experiments, and sub-scale rocket combustion experiments operated at low pressures. In the present work, several such methods were applied to data from two cryogenic oxygen-natural gas rocket experiments operated at higher pressures than previously reported; both sub- and supercritical with respect to oxygen. The goal was to identify methods that can discern limit-cycle instabilities from intermittently unstable operation and are sufficiently responsive to be applied as emergency shut-down criteria in engine tests. Among the methods applied were the standard deviation, variance of the auto-correlation, the second spectral moment, the ratio between determinism and recurrence rate, the Hurst-exponent, and the multifractal range. The second spectral moment, the Hurst-exponent, and a measure derived from the multifractal spectrum all have short detection delays for instability onset and short-lived could be discerned from self-sustaining instabilities with an appropriate choice of threshold value. They also have moderate computation cost which makes them of interest for potential real-time implementation. The Hurst-exponent has the additional advantage of a common threshold value for all test cases addressed, demonstrating its potential for broader application independent of combustion device or operating conditions.

Design and Verification of Acoustic Excitation Device for Simulation of Liquid Rocket Engine Combustion Instability

Design and Verification of Acoustic Excitation Device for Simulation of Liquid Rocket Engine Combustion Instability

액체 로켓 엔진 연소불안정 현상 모사를 위한 비원통형 연소실 설계 및 검증

액체 로켓 엔진 연소불안정 현상 모사를 위한 비원통형 연소실 설계 및 검증

Multiphysics Modeling for Combustion Instability in Paraffin-Fueled Hybrid Rocket Engines

The use of paraffin-based fuels is a promising approach to a low regression rate in hybrid rocket engines, and the capability to describe and predict combustion instability in the presence of liquefying fuels becomes an enabling step towards the application of hybrid rockets in a wide range of space transportation systems. In this work, a multiphysics model having this purpose is presented and discussed. The model is based on a network of submodels in which chamber gas dynamics is described by a quasi-1D Euler model for reacting flows while thermal diffusion in the grain is described by the 1D heat equation in the radial direction. An artificial neural network is introduced to reduce the computational cost required by the chemical submodel. A sensitivity analysis is performed to identify the key parameters, which have the largest influence on combustion instability and to evaluate the predictive capability of the model despite the uncertainty introduced with the necessary modeling simplifications. Results are presented considering two test cases with different oxidizers: hydrogen peroxide and gaseous oxygen. The procedure shows good agreement with the experimental results available in the literature.

Numerical investigation of combustion instability in a liquid rocket engine: Interaction effect between hydrodynamics and acoustic mode

In this paper, we present the numerical evidence of multiple bifurcations in thermoacoustic instability as we decrease the equivalence ratio (ϕ) from fuel-rich (ϕ = 1.5) to fuel-lean (ϕ =0.59) within a single-element coaxial liquid rocket engine burning hydrogen peroxide and methane. An axisymmetric compressible Large Eddy Simulation solver in an open-source platform OpenFOAM combined with a global methane chemical kinetical mechanism was used to model the turbulent mixing and combustion process. After the validation of grid independence and simulation results, the synchronization phenomenon between hydrodynamic instability, acoustics and combustion is the main focus of this paper. The results here first revealed that in a direction of decreasing equivalence ratios, the first transition to limit cycle state from combustion noise state was observed via intermittency. A second bifurcation was also observed, leading to quasi-period thermoacoustic instability from the limit cycle state. Intrinsic vorticity modes and natural acoustic modes could be regarded as oscillators. The competition and cooperation between the two modes led to a phenomenon beyond limit cycle states. The vortex dynamics of thermoacoustic systems are dominated by the interaction of intrinsic vorticity modes and natural acoustic modes and exhibit an adjustment of vortex shedding frequencies, resulting in different stability regions. Herein, the flame index was performed in more detail to distinguish between the premixed flame and diffusion flame. Results indicated that the burning mode was an important driver in promoting instability whilst it also served as a medium for competition and cooperation between the natural acoustic modes and intrinsic vorticity modes.

Numerical investigation of combustion instability in a single-element liquid rocket engine: Intermittency routes before and after thermoacoustic instability

This paper provides the first numerical proof of intermittency routes existing before and after thermoacoustic instability, as we increase the oxidizer inlet temperature under fuel-lean (ϕ=0.91) and fuel-rich (ϕ=1.25) conditions in coaxial liquid rocket engine utilizing methane and hydrogen peroxide as propellant. Axisymmetric compressible Large Eddy Simulation algorithms combined with global chemical reaction mechanisms are used to simulate the non-premixed turbulent combustion process in high-pressure flows, where the one-equation eddy sub-grid turbulence model and the PaSR sub-grid combustion model are employed based on OpenFOAM. After verification of the grid independence as well as the global chemical mechanism against experimental measurements and existing simulations, the multi-bifurcation process of the system dynamics and the corresponding lock-in phenomenon are the main focus of this paper. The results suggest that with an increase in oxidizer inlet temperatures, the system dynamics at ϕ=0.91 conditions experience an initial shift from states of combustion noise to states of thermoacoustic instability through intermittent oscillation. A further rise in oxidizer inlet temperature leads to a second bifurcation from thermoacoustic instability states back to combustion noise states via intermittency, which is the first observation in the liquid rocket engine. Such an intermittency route before and after thermoacoustic instability is preserved at ϕ=1.25 conditions. Furthermore, a decrease in the equivalence ratio accelerates the initial bifurcation point for the onset of thermoacoustic instability. Flame index analysis is also employed to distinguish between diffusion and premixed flames, providing further insight into the underlying physical mechanism of the frequency-locking phenomenon between the pressure and combustion flow subsystems both in fuel-lean and fuel-rich conditions. Other details of the combustion flow fields also include the migration of the flame anchor point towards the injector with an increase in oxidizer inlet temperature.

The acoustic transfer impedance of baffled injectors in three-dimensional combustion chambers

Baffled injectors are typical passive control devices to restrain combustion instabilities in liquid rocket engines. However, there still has been no appropriate model to describe the acoustic characteristics of baffled injectors. In this work, the baffled injector gap is modeled as a narrow passageway with a cross-sectional surface area gradient. An acoustic transfer impedance model is derived from the momentum conservation equation to relate the acoustic disturbances between the inlet and outlet of injector gaps. Validation is conducted by comparing the first tangential mode frequency of a three-dimensional combustion chamber without flame from the detailed structures model and that using the proposed acoustic transfer impedance model. The proposed impedance model is embedded in a numerical simulation model of a three-dimensional combustion chamber considering the unsteady heat release; the effects of baffled injectors geometry parameters on the first tangential acoustic mode are then investigated. Results show that the presence of baffled injectors not only damps the acoustic energy due to the viscothermal effect, but also can longitudinalise the original transverse mode.

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.

Combustion Simulation of 82-Injector Rocket Engine

A computational study of the nonlinear rocket-engine combustion instability is presented for an engine with 82 coaxial methane–oxygen injector ports, a choked nozzle, and a combustion chamber that models a Rocketdyne experiment. Computations use a three-dimensional unsteady shear-stress transport delayed detached-eddy simulation method. Four different one-step chemical kinetic rates are simulated on the resolved scale to determine the effect of the burning rate. Application of kinetic rates on the resolved scale give faster rates than expected with turbulent combustion where the final mixing and chemical reaction in the cascade occurs on the smallest scales; thus, reduction of the kinetic rate is examined here. When the normalized kinetic rate is 1, implying the nominal Westbrook–Dryer rate, only the longitudinal-mode instability is observed. As the kinetic rate decreases, less of the combustion occurs upstream near the longitudinal-mode pressure antinode, the longitudinal-mode instability becomes weak, and the first-tangential-mode instability emerges, which fits better with experiment. The proximity of the domain of major heat release to the longitudinal mode pressure antinode is found to be critical. Evidence is presented for the need for an improved combustion model that will provide a more accurate and lower burning rate.

Algorithm for computational and experimental determination of the acoustic characteristics of solid propulsion engine

Combustion instability in rocket engines is accompanied by large-amplitude pressure fluctuations in the combustion chamber and (or) intense vibration of the load-bearing structural elements. Unstable operation leads to a decrease in thrust, specific impulse, a decrease in engine life, a violation of the normal operation of vibration-sensitive flight control system equipment. The study considers a method for determining the frequencies of the first and second tone natural resonances of the longitudinal mode of acoustic vibrations in the combustion chambers of solid propulsion engines. The gas path of the combustion chamber is divided into homogeneous sections, for which solutions of the wave equation are presented. To determine the natural frequencies and the distribution of vibrational pressures and velocities, the method of "stitching" acoustic fields at the boundaries of cavities was applied. In the course of the study, to determine the decrement from natural perturbations, the following methods were used: spectral, correlation, amplitude, and instantaneous period method. In the study, vibrations of a certain frequency in a given range were excited using an electrodynamic emitter. To obtain the acoustic characteristics of the camera at various moments of engine operation, experiments were carried out with several charge models having different sizes. The solution of the wave equation in the form of standing waves was considered. In addition, the study results present a description of the experimental setup, in particular, the distribution of the oscillatory pressure in the chamber gas cavity (according to the calculated data). In the course of the study, an algorithm was developed for calculating and experimentally determining the acoustic characteristics of the combustion chambers of solid propulsion engines.

A Hybrid Real/Ideal Gas Mixture Computational Framework to Capture Wave Propagation in Liquid Rocket Combustion Chamber Conditions

The present work focuses on the development of new mathematical and numerical tools to deal with wave propagation problems in a realistic liquid rocket chamber environment. A simplified real fluid equation of state is here derived, starting from the literature. An approximate Riemann solver is then specifically derived for the selected conservation laws and primitive variables. Both the new equation of state and the new Riemann solver are embedded into an in-house one-dimensional CFD solver. The verification and validation of the new code against wave propagation problems are then performed, showing good behavior. Although such problems might be of interest for different applications, the present study is specifically oriented to the low order modeling of high-frequency combustion instability in liquid-propellant rocket engines.

Analysis and Control of Combustion Instabilities in Rocket Engines

Sound is the transmission of energy that is produced when two particles or objects undergo collision. It is one of the forms in which energy modulates and travels in a medium. Vibration is a phenomenon in which an object of mass executes a periodic oscillatory displacement when certain energy is transferred to it. Combustion is a chemical reaction in which a lot of molecules collide. Combustion instabilities occur in a reacting flow. It is a physical phenomenon. Combustion instabilities have mostly been studied in a particular flow but they also occur in real life. Real engines often feature specific unstable modes such as azimuthal instabilities in gas turbines or transverse modes in rocket chambers. Eddy simulation has been the major tool to study this type of instabilities so far but recently it has been proved to be insufficient to completely understand the complex nature of these instabilities [3]. These instabilities involve large Reynolds number, high pressure, densities in real engines. Combustion instabilities can occur at any part of the rocket propulsion system like nozzle, combustion chamber, injectors, feed systems and lines. Theory plays an important role in understanding and analysing these instabilities and the amount of damage they can cause to a particular object. In this paper we will look at different types or modes of combustion instabilities and active and passive ways to control them in real situations.

Theoretical and Experimental Investigations on the Acoustic Absorptions of Baffled Injectors

Baffled injectors are widely used to restrain transverse combustion instabilities in liquid rocket engines. However, there is still a lack of theoretical research on the attenuation mechanisms. In this work, a theoretical model accounting for the viscothermal effect has been constructed upon the narrow gap between two neighboring injectors. The acoustic wave equation for the pressure perturbation is derived, and the perturbation method is then used to derive its analytical solutions. Preliminary validation of the analytical solutions is performed by comparing them to the results by numerically resolving the wave equation. An experimental rig to measure the acoustic absorption coefficient of one modeled baffled injector unit is constructed in order to further validate the proposed theoretical models and analytical solutions. Good agreements have been observed between the analytical and experimental results. The normalized minimum gap between neighboring injectors of and the normalized frequency of are proposed for the analyses, where , , , and are the minimum gap, the viscous boundary-layer thickness, the angular frequency, the injector diameter, and the speed of sound, respectively. As the Strouhal number exceeds 0.1, the optimal ratio of to achieve the maximum acoustic absorption approximately equals 1.7 and does not rely on the forcing frequency, injector diameter, ambient pressure, and temperature.

Combustion Dynamics Simulation of a 30-Injector Rocket Engine

ABSTRACT A computational study of the nonlinear rocket-engine combustion instability is presented for an engine with 30 coaxial methane-oxygen injector ports, a choked nozzle, and a combustion chamber. Three combustion chamber lengths of 18 cm, 24 cm, and 30 cm are simulated with differing instability modes depending on length. Computations using a three-dimensional unsteady k − ω shear-stress transport delayed detached-eddy simulation method provide detailed time-resolved information about the combustion instability. Our flame-index analysis explains why the premixed-flame structures dominate over diffusion flame-structures in the driving mechanism for the instability. This causes the engine to become more unstable as the mixture becomes more fuel rich. Spontaneous longitudinal-mode and tangential-mode instabilities are observed. The oscillation amplitude of the two instability modes vary in alternating fashion during the simulation time. The alternating behavior of the two instability modes can be modified by pulsing the injector mass-flux on the mixture ratio at a particular frequency.

The auto-ignition boundary of ethylene/nitrous oxide as a promising monopropellant

C2H4/N2O has recently been proposed as a promising monopropellant for hydrazine replacement. In this work, the auto-ignition capability of argon diluted C2H4/N2O was experimentally investigated firstly in a rapid compression machine by detecting the light emission and the pressure evolution profiles. The auto-ignition boundary temperature which separates the non-ignition and successful ignition cases was then determined and modeled as a function of pressure and mixture composition. Results show that the auto-ignition temperature decreases with the increase of pressure and increases with the argon dilution fraction and fuel oxidizer ratio. In addition, for successful auto-ignition cases, the ignition delay times were measured and used to validate several recently developed kinetic mechanisms. A mechanism was developed and showed good agreement with the measured ignition delay time as well as the auto-ignition boundary temperature at all test conditions. Finally, the auto-ignition boundary temperatures at practical rocket engine conditions (without dilution) were numerically predicted at various pressure and fuel oxidizer ratio conditions. The model is of merit for designing the cooling system and the local flow residence time to reduce the risk of fuel line explosion and the resonance acoustic induced combustion instability in practical rocket engine.

Damping device to reduce the risk of injection-coupled combustion instabilities in liquid propellant rocket engines

A new countermeasure against injection-coupled combustion instabilities in liquid propellant rocket engines is presented. Whereas the problem is usually addressed by adding damping elements such as baffles or resonators to the combustion chamber, this approach directly damps the acoustic eigenmodes of the injector instead. The principle of the damping method is described in this article, as well as the implementation of such a device in a sub-scale rocket thrust chamber operated with liquid oxygen and hydrogen at conditions representative of upper stage engines. Test results are presented which show that flame and pressure oscillations were successfully reduced by the modification. The absorbers had no measurable influence on thrust chamber performance, and so the solution lends itself to retrofitting in existing engines, as well as integration during the design phase.

Effects of self-pulsation on combustion instability in a liquid rocket engine

The flame oscillation characteristics as well as the relationship between self-pulsation and combustion instability in a combustor burning gaseous oxygen (GOX) and liquid ethanol (LET) through a liquid-centered swirl coaxial (LCSC) injector were investigated experimentally. Data were obtained at pressures from 0.53 to 0.6 MPa and mixture ratios from 0.16 to 0.3. The side-on images of spray and flame were captured synchronously through two high-speed cameras. With the increase of recess length, spray and flame transform from the stable behavior to the self-pulsated behavior with periodic intensity variations of flame emission. When self-pulsation occurs, flame together with the pressure of the LET manifold oscillates at the same frequency, and the peak-to-peak pressure amplitudes of the LET-manifold are greater than 26%. Due to the blocking actions of the conical liquid sheet, spray self-pulsation caused by the periodic fluctuation of pressure in recess chamber is found to be the essential reason of flame and LET manifold pressure oscillation. Furthermore, when the self-pulsation frequency is coupled with the first-longitudinal-mode (1L) natural acoustic frequency of combustor, combustor pressure will be induced to oscillate at the same frequency. This further indicates that self-pulsation may probably be a critical factor inducing combustion instability in liquid rocket engines.

Nonlinear Combustion Instability in a Multi-Injector Rocket Engine

A computational study is presented of the nonlinear combustion instability of a multi-injector rocket engine. The study addresses a choked nozzle and a combustion chamber with 10 and 19 coaxial injector ports. Computations using a three-dimensional unsteady shear-stress transport delayed detached-eddy simulation method provide detailed time-resolved information about the combustion instability. The spontaneous longitudinal-mode instability is observed for the 10- and 19-injector geometries with a combustion-chamber diameter of 28 cm. The triggered tangential and longitudinal instability modes are obtained for the 19-injector geometry with a combustion-chamber diameter of 43 cm by pulsing the injector mass flux. It is shown that an oscillating combustion-chamber flow can be triggered to a new mode with a larger disturbance amplitude. The streamwise vorticity created by the instability impacts the heat release in a manner that supports the instability. A preliminary study of injector-size scaling effects is performed by comparison of the 10- and 19-injector combustors.

Reduced-Order Modeling Framework for Combustor Instabilities Using Truncated Domain Training

A multifidelity framework is established and demonstrated for the prediction of combustion instabilities in rocket engines. The main idea is to adapt appropriate fidelity modeling approaches for different components in a rocket engine to ensure accurate and efficient predictions. Specifically, the proposed framework integrates projection-based reduced-order models (ROMs) that are developed using bases generated on truncated domain simulations. The ROM training is performed on truncated domains, and thus does not require full-order model solutions on the full rocket geometry, thus demonstrating the potential to greatly reduce training costs. Geometry-specific training is replaced by the response generated by perturbing the characteristics at the boundary of the truncated domain. This training method is shown to enhance the predictive capabilities and robustness of the resulting ROMs, including at conditions outside the training range. Numerical tests are conducted on a quasi-one-dimensional model of a single-element rocket combustor, and the present framework is compared to traditional ROM development approaches.

Reduced Order Modeling Approach to Combustion Instabilities of Liquid Rocket Engines

This paper describes one possible approach to a numerical framework for a reduced order tool that aims to simulate combustion instabilities in liquid rocket engines. The numerical framework relies on the projection of the pressure fluctuations on the eigenmodes of the system. The pressure fluctuations are solutions of the wave equation of the system. After projection on the eigenmodes, the wave equation takes the form of a series of second-order harmonic equations with source terms that drive combustion instabilities and damping terms that attenuate them. A test rig was developed to study cavity interactions, injector impedance, and damping effects. The damping rates that were measured on the test rig show a trend that is consistent with observations in liquid rocket engines. On the whole, the test rig can be used to validate simplified models of combustion instabilities. The global framework of the reduced order approach that was developed to predict combustion instabilities was first validated by comparing the data from simulations against experimental results from the test rig in a series of nonreacting experiments. Our approach was then applied to a case study of a full-scale rocket engine. This engine, under certain operating conditions, exhibits instabilities. Stable and unstable behaviors have been revealed by the temporal evolution of calculated pressure amplitudes.