Model Rocket Combustor Research Articles

Laser induced spark ignition of a gaseous methane–oxygen model rocket combustor

The spatio-temporal evolution of laser induced spark ignition of non-premixed gaseous methane and oxygen is investigated. The reactants are injected into an optically accessible combustion chamber from an oxidizer centered shear-coaxial injector. High-speed schlieren imaging and measurements of laser energy deposited are used to characterize ignition behavior at various spark locations throughout the chamber. A spatial map of ignition probability is generated from tests at multiple axial and radial locations from the point of propellant injection. The rate of pressure rise from successful tests reveals two modes of ignition dependent on the location of the laser-induced breakdown, direct and indirect, based on laser-induced breakdown occurring within reactant jet or in the recirculation region, respectively. Physical processes occurring over multiple timescales throughout the direct ignition method are first discussed. Ignition outcomes associated with the indirect method are determined by the hydrodynamic ejection protruding from the laser spark, the behavior of which is explored in detail using flow statistics extracted from images taken at a single spark location. The spatial–temporal progression of this jet is found to be dependent on deposited laser energy, leading to a relationship between the amount of energy added to the flow and ignition outcome. General bounds for spark location, deposited laser energy, and ejection behavior are identified to predict ignition outcomes. Cases that are an exception from these bounds are investigated in detail to understand the cause of unique behavior. These cases lie within regions of the variable space that are susceptible to stochastic elements of the flow.Novelty and Significance Statement: This work investigates laser induced spark ignition of O2 and CH4 issuing from a shear-coaxial injector in a model rocket combustor. Probability of ignition as a function of deposited energy and spatial location is studied with high-speed schlieren imaging and pressure measurements. Two distinct ignition modes that have previously not been identified, direct and indirect, are dependent on the laser deposition location with respect to the reactant jet. For indirect cases, where laser deposition is at a distance from the reactants, the hydrodynamic jet ejection generated from the breakdown of the plasma plays a critical role on ignition probability. The behavior of the ejection jet is characterized to provide a phenomenological description of the ignition process. The results of this research show the important interaction between chemical, fluid dynamic, and thermodynamic processes and their impact on laser ignition.

Ensemble predictions of laser ignition with a hybrid stochastic physics-embedded deep-learning framework

When investigating stochastic ignition phenomena, high-fidelity simulations and experiments for statistical characterization can be a laborious endeavor. Machine learning (ML), in particular deep-learning, offers an accurate and cost-effective approach for reduced-order modeling of forced ignition. In this work, we investigate the potential and limitations of ML in modeling ignition, especially in the presence of limited data and multiple ignition modes. To this end, we introduce a hybrid stochastic physics-embedded deep-learning framework that combines sparse experimental data from Schlieren measurements with inert large-eddy simulations for predicting the spatio-temporal evolution of ignition kernel location and morphology. This model combines a stochastic differential equation for modeling kernel dynamics and a deep-learning model for representing kernel morphology. We evaluate this modeling approach for laser ignition of a gaseous methane/oxygen (CH4/O2) model rocket combustor, where morphological effects of the ignition kernel significantly influence ignition behavior. Results demonstrate that this model can reasonably capture behavior associated with the three dominant ignition modes, namely direct, indirect, and failed ignition, along with statistics associated with kernel growth and position, at lower computational costs than high-fidelity reacting simulations. In addition, we demonstrate that this modeling framework can be employed for generating spatially resolved ignition probability maps by incorporating physics to represent kernel interaction with the turbulent jet. We note that limitations in accuracy can be observed when predicting with vastly out-of-distribution data. Nevertheless, these results demonstrate that this physics-embedded ML approach can statistically characterize forced ignition in a more cost-effective manner than reacting high-fidelity simulations, as long as sufficiently representative data is available.

Analysis of self-excited transverse combustion instability in a rectangular model rocket combustor

A methane/oxygen mixture is considered to be an appropriate propellant for many future rocket engines due to its practicality and low cost. To better understand the combustion instability in methane/oxygen-fed rocket engines, the spontaneous transverse combustion instability in a rectangular multi-element combustor (RMC) was analyzed both experimentally and numerically. Severe combustion instabilities occurred in the RMC during repeatable hot-fire tests. The physical mechanisms were systematically investigated through numerical simulations based on the stress-blended eddy simulation and flamelet-generated manifolds method with detailed chemical mechanisms (GRI Mech 3.0). The numerical results for the frequency spectrum and spatial modes agree well with the theoretical analysis and experimental data. The driven regions of the combustion instability were identified on both sides of the combustion chamber through a Rayleigh index analysis. The longitudinal pressure oscillations in the oxidizer post were found to be coupled with the transverse pressure waves in the combustion chamber and led to periodic oscillations of the mass flow rate of propellant. Moreover, the mixing was highly enhanced when the pressure wave interacted with walls of the combustion chamber. Therefore, a sudden release of heat occurred. The pressure oscillations were enhanced by pulsated heat release. A closed-loop system with positive feedback associated with periodic oscillations mass flow rate of the propellant, and sudden heat release, was believed to account for the present combustion instability.

Coupled interaction between acoustics and unsteady flame dynamics during the transition to thermoacoustic instability in a multi-element rocket combustor

Rocket engine combustors are prone to transverse instabilities that are characterized by large amplitude high frequency oscillations in the acoustic pressure and the heat release rate. In this work, we study the coupled interaction between the acoustic pressure and the CH* intensity (representative of heat release rate) oscillations in a 2D multi-element self-excited model rocket combustor during the transition from a stable state to thermoacoustic instability through intermittency. We show the emergence of synchronization between these oscillations from desynchronization through intermittent phase synchronization during the onset of thermoacoustic instability. We find substantial evidence that the intensities of the jet flames close to the end wall is higher than that observed near the center of the combustor as a result of its strong coupling to the local acoustic field. Using concepts from recurrence theory, we distinguish the type of synchronization between the acoustic pressure and the CH* intensity oscillations at the end wall and the center of the combustor during thermoacoustic instability. Analyzing the local CH* intensity oscillations, we observe that the longitudinally propagating jet flames experience substantial transverse displacement with flame merging effects during thermoacoustic instability. Furthermore, we differentiate the interaction of the jet flames with the shock wave during intermittency and thermoacoustic instability at both the end wall and the center of the combustor. We also discerned the change from stochastic to deterministic nature in the local CH* intensity oscillations during the transition to thermoacoustic instability. Our results demonstrate that only the first few transverse modes contribute to the generation of acoustic power from the reacting flow, in spite of the pressure oscillations featuring several harmonics. The insights gained from such coupled analysis during the transition from stable state to thermoacoustic instability can be leveraged to aid modeling efforts of thermoacoustic instability in rocket engines with multiple injectors.

Experimental investigation on effects of injection distribution on combustion instability in a model rocket combustor

Self-excited combustion instabilities of transverse modes were experimentally investigated in a rectangular multi-injector model combustor, operating with the bipropellants O2/CH4. The propellants were injected through a linear array of five oxidizer-centered shear coaxial injectors into the combustor. High-amplitude limit cycles obtained in hot-fire tests were analyzed in detail. Different combustion instability modes, including first and second width modes, were observed in cases with three different injection distribution schemes. Hence, the injection distribution strongly determined the combustion dynamics. One insight can be gained that the stable combustion could be achieved by properly designing the propellants' injection distributions.

Prediction of combustion instability by combining transfer functions in a model rocket combustor

The flame transfer function (FTF) and acoustic transfer function (ATF) are estimated experimentally and numerically, respectively, and both are combined to predict combustion instability of a model rocket combustor with a single coaxial jet–swirl injector. FTF of the injector is measured experimentally by imposing external velocity excitation on the injector. ATF, i.e., model-chamber response to heat-release fluctuation from flames, is difficult to measure, and alternatively, an idea of numerical evaluation method is suggested by applying an equivalence-ratio perturbation to flames. Then, ATF is calculated by numerical simulations with a three-dimensional model and compared with ATF from the one-dimensional analytic equation. Results show that ATF estimated by the 3-D numerical model improves accuracy in predicting resonant frequencies and growth rates of pressure oscillations in the model chamber. Finally, a closed-loop system transfer function describing dynamic behaviors of the model chamber is formulated by combining FTF with ATF and it is applied to predict combustion instability of the chamber. This prediction is compared with experimental observation, which shows good agreement between them.

Detailed Measurement of Oxidizer-Rich Staged Combustion Injector Dynamics in Model Rocket Combustors

This paper presents experimental results from a model rocket combustor that uses a gas-centered swirl-coaxial injector element injecting warm oxygen and ambient temperature RP-2 fuel with 900 psia chamber pressure, representative of an oxidizer-rich staged combustion cycle. Results from the two injector geometries are analyzed and compared. One was a flat-faced injector with the element ending at a sudden expansion into the chamber. The other used a truncated cone at the injector exit: a design feature of specific interest that appears in drawings for the RD-170 main-chamber injector. To improve spatial resolution of high-frequency pressure measurements, the model injector was scaled up from RD-170 dimensions to facilitate measurements in the injector and head end of the combustor. Measured pressure spectra of the flat-face combustor showed a well-organized first axial mode and harmonics. The truncated-cone geometry showed a more diffuse spectra with higher-frequency modes, with a reduction in amplitude further upstream in the cone. Intermittency in measured dynamics is characterized via multiscale analysis. Adopting concepts from fractal theory, the stability margin of the injector configurations was quantified using the Hurst exponent of pressure oscillations. Comparatively, a larger range of fractal dimensions with smaller fluctuations in the initial reacting zone were observed for the flat-face injector.

Formation mechanism of high-frequency combustion oscillations in a model rocket engine combustor

We study the formation mechanism of high-frequency combustion oscillations in a model rocket combustor from the viewpoints of symbolic dynamics and complex networks. The flow velocity fluctuations in the fuel injector generated by the pressure fluctuations in the combustor give rise to the periodic ignition of the unburnt fuel/oxidizer mixture, resulting in a significant change in the heat release rate fluctuations in the combustor. The heat release rate fluctuations drive the pressure fluctuations in the combustor before a transition state, while the pressure fluctuations in the combustor gradually begin to significantly affect the heat release rate fluctuations during the transition to combustion oscillations. The directional feedback process during the transition and subsequent combustion oscillations is identified by the directionality index of the symbolic transfer entropy. The thermoacoustic power network enables us to understand the physical mechanism behind the transition and subsequent combustion oscillations.

Transported PDF modeling of compressible turbulent reactive flows by using the Eulerian Monte Carlo fields method

Although the transported probability density function (PDF) method has been developed for decades, its application has been mainly focused on the low-Mach number flow problems. This work extends the transported PDF method to compressible flow problems. The Eulerian Monte Carlo fields (EMCF) solution method is employed to solve the transported PDF equation for compressible flow problems. A pseudo stagnation enthalpy is introduced and its stochastic partial differential equation is derived to ensure total energy conservation numerically. A new mixing model called interaction by partial exchange with mean (IPEM) is introduced to expand the available choices of mixing models for the EMCF method. The consistency of the EMCF method is examined for solving the transported PDF equation. Numerical implementation details are discussed, such as the density coupling between the compressible flow solver and the EMCF solver, discretization schemes for the mixing terms and the stochastic terms. The implemented compressible flow solver coupled with the EMCF solver is verified and validated in a series of test cases with increasing level of complexity, ranging from a statistically one-dimensional turbulent mixing layer to a self-excited resonance model rocket combustor. It is observed that in general with the increase of compressibility, there is an increase in the sensitivity of the modeling results to the different models and algorithms. This makes it necessary to develop a thorough understanding of the model sensitivity in order to develop a robust and accurate simulation solver for highly compressible turbulent reactive flows. The thermo-acoustic instability inside the model rocket combustor case is captured reasonably, which demonstrates the overall capability of the developed compressible turbulent combustion solver based on the transported PDF method.

Combustion instability modeling using multi-mode flame transfer functions and a nonlinear Euler solver

Modern methods for predicting combustion dynamics in high-pressure combustors range from high-fidelity simulations of sub-scale model combustors, mostly for validation purposes or detailed investigations of physics, to linearized, acoustics-based analysis of full-scale practical combustors. Whereas the high-fidelity simulations presumably capture the detailed physics of mixing and heat addition, computational requirements preclude their application for practical design analysis. The linear models that are used during design typically use flame transfer functions that relate the unsteady heat addition [Formula: see text] to oscillations in velocity and pressure ([Formula: see text] and [Formula: see text]) that are obtained from the wave equation. These flame transfer functions can be empirically determined from measurements or derived from theory and analysis. This paper describes a hybrid approach that uses high-fidelity simulations to generate flame transfer functions along with nonlinear Euler CFD to predict the combustor flowfield. A model rocket combustor that presented a self-excited combustion instability with pressure oscillations on the order of 10% of mean pressure is used for demonstration. Spatially distributed flame transfer functions are extracted from a high-fidelity simulation of the combustor and then used in a nonlinear Euler CFD model of the combustor to verify the approach. It is shown that the reduced-fidelity model can reproduce the unsteady behavior of the single element combustor that was both measured in the experiment and predicted by a high-fidelity simulation reasonably well.

Spatiotemporal dynamics and early detection of thermoacoustic combustion instability in a model rocket combustor.

We numerically study the spatiotemporal dynamics and early detection of thermoacoustic combustion instability in a model rocket combustor using the theories of complex networks and synchronization. The turbulence network, which consists of nodes and vertexes in weighted networks between vortices, can characterize the complex spatiotemporal structure of a flow field during thermoacoustic combustion instability. The transfer entropy allows us to identify the driving region of thermoacoustic combustion instability. In addition to the order parameter, a phase parameter newly proposed in this study is useful for capturing the precursor of thermoacoustic combustion instability.

Reduced Dimension Analysis on Combustion Oscillation in a Model Rocket Combustor Using a Deep Neural Network

Reduced Dimension Analysis on Combustion Oscillation in a Model Rocket Combustor Using a Deep Neural Network

Lagrangian transported MDF methods for compressible high speed flows

This paper deals with the application of thermochemical Lagrangian MDF (mass density function) methods for compressible sub- and supersonic RANS (Reynolds Averaged Navier–Stokes) simulations. A new approach to treat molecular transport is presented. This technique on the one hand ensures numerical stability of the particle solver in laminar regions of the flow field (e.g. in the viscous sublayer) and on the other hand takes differential diffusion into account. It is shown in a detailed analysis, that the new method correctly predicts first and second-order moments on the basis of conventional modeling approaches. Moreover, a number of challenges for MDF particle methods in high speed flows is discussed, e.g. high cell aspect ratio grids close to solid walls, wall heat transfer, shock resolution, and problems from statistical noise which may cause artificial shock systems in supersonic flows. A Mach 2 supersonic mixing channel with multiple shock reflection and a model rocket combustor simulation demonstrate the eligibility of this technique to practical applications. Both test cases are simulated successfully for the first time with a hybrid finite-volume (FV)/Lagrangian particle solver (PS).

Self-Excited Non-Linear Acoustic Wave in a Single-Element Model Rocket Combustor and Its Influence on Flame

Self-Excited Non-Linear Acoustic Wave in a Single-Element Model Rocket Combustor and Its Influence on Flame

Approaches for Comparing Numerical Simulation of Combustion Instability and Flame Imaging

Combustion dynamics are controlled by the coupling between heat addition and gas dynamic modes, and their direct measurement and comparisons with prediction are key to improving computational tools, as well as our fundamental understanding of the problem. This paper demonstrates several methods for characterizing and comparing dynamic combustion response from experiment and simulation. A model rocket combustor that exhibits self-excited instabilities is used for the study. Comparisons are made for two configurations: one stable and one unstable. High-speed chemiluminescence imaging from the experiment was first phase averaged, then treated with an Abel inversion routine to produce a dynamic two-dimensional distribution of heat addition. The distribution before inversion could be compared with line-integrated calculations of the heat release rate from three-dimensional large-eddy simulations, and after inversion to azimuthally averaged cross sections from simulations. Modal decomposition of the two-dimensional distributions was performed and compared. The applicability and limitations of each comparison approach are assessed.

Unsteady High-Order Simulation of a Liquid Oxygen/Gaseous Hydrogen Rocket Combustor

Coupled Euler–Lagrange simulations are performed for a subcritically operated model rocket combustor at 10 bar pressure (MASCOTTE A-10 test case). A Lagrangian spray model is used to simulate the liquid phase. This liquid oxygen/gaseous hydrogen single element combustor experiment has been the focus of many numerical studies. In contrast to previous investigations, time-accurate unsteady Reynolds-averaged Navier–Stokes simulations are performed. A good spatial resolution is achieved by using a fifth-order multidimensional limiting process scheme to discretize the inviscid fluxes. Instead of assuming axisymmetry, the three-dimensional geometry of the rectangular combustor is taken into account. However, only a quarter of the combustor is simulated, to limit the computational cost. Presented results indicate that three-dimensional effects are important, questioning the assumption of axisymmetry. A second important finding is that time-averaged unsteady Reynolds-averaged Navier–Stokes results agree much better with experimental data than those of a comparable steady-state simulation. Close to the injector, results show a strong unsteadiness of the turbulent flame, which has been observed in the experiment, too. Finally, pressure fluctuations are monitored at several positions in the combustor and analyzed by fast Fourier transform. The peak amplitudes in the pressure spectra could successfully be assigned to the first longitudinal acoustic mode and its higher harmonics.

Comparative Evaluation Between Experiment and Simulation for a Transverse Instability

A transverse combustion instability study composed of experimental and hybrid Reynolds-averaged Navier–Stokes/large eddy simulation results is described. The study is focused on gas-centered swirl coaxial injector elements like those used in the main chambers of oxidizer-rich staged-combustion engines. Experimental results are obtained from a self-excited, multi-element, high-pressure model rocket combustor. The simulation uses a velocity forcing technique to match the frequency and amplitude of the pressure oscillations that are measured in the combustor. Chemiluminescence is used to indirectly measure heat release rate in the experimental investigation. Detailed comparisons between the predicted and measured pressure field and the predicted and measured heat release modes are made. Modal decomposition is used to systematically compare the heat release modes. Results show successful computational replication of the experimental unsteady environment through the use of the velocity forcing technique and the successful qualitative matching of experimental CH* emission with computational heat release rate.

Combustion Instability Control Through Acoustic Modulation at the Inlet Boundary: Experiments

This research investigation encompasses experimental tests demonstrating the control of a high-frequency combustion instability by acoustically modulating the propellant flow. This investigation complements an accompanying theoretical study implementing linear modal analysis [Bennewitz, J. W., Rani, S. L., Cranford, J. T., and Frederick, R. A., Jr., “Combustion Instability Control Through Acoustic Modulation at the Inlet Boundary: Analysis,” Journal of Propulsion and Power, (to be published)]. A model rocket combustor burned gaseous oxygen and methane using a single-element pentad-style injector. Flow conditions were established that spontaneously excited a 2430 Hz first longitudinal combustion oscillation at an amplitude up to . An acoustic speaker was placed at the base of the oxidizer supply line to modulate the flow and alter the oscillatory behavior of the combustor. Two speaker modulation approaches were investigated: 1) bands of white noise, and 2) pure sinusoidal tones. The first approach adjusted 500 Hz bands of white noise ranging from 0–500 to 2000–2500 Hz, whereas the second approach implemented individual harmonic signals with arbitrary phase swept from 500 to 2500 Hz. The results show that, above a modulation signal amplitude threshold, both approaches suppressed % of the spontaneous combustion oscillation. By increasing the applied signal amplitude, a wider frequency range of instability suppression became present for these two acoustic modulation approaches. Thus, this work further supports the strategic application of acoustic modulation within an injector as a potential method to control high-frequency combustion instabilities for liquid rocket engine applications.

Influence of spatial discretization and unsteadiness on the simulation of rocket combustors

SummaryThis paper investigates some important numerical aspects for the simulation of model rocket combustors. Precisely, (1) a new high‐order discretization technique (multi‐dimensional limiting process (MLP), low diffusion, and MLPld) is presented and compared with conventional second‐order schemes with different flux limiters. (2) Time accurate unsteady Reynolds‐averaged Navier–Stokes (RANS) simulations are performed to assess possible improvements in comparison with steady‐state RANS simulations. (3) Fully 3D simulations of an axisymmetric rocket combustor are compared with 2D axisymmetric ones. All studies are based on the Penn State preburner combustor experiment, which uses gaseous oxygen and hydrogen. This comprehensive study offers unique insight into how the mentioned numerical influence factors change the flow field, flame, and wall heat fluxes in the model rocket combustor. Because wall heat fluxes are known from the experiment only, numerical results are compared with LES of other authors, too. It will be shown that the high‐order spatial discretization significantly improves the agreement with measured wall heat fluxes at low additional computational cost. In general the transition from simple to more complex numerical approaches steadily improves the qualitative agreement between simulation and experiment. Copyright © 2015 John Wiley & Sons, Ltd.

Influence of Ammonia Borane on the Stability of a Liquid Rocket Combustor

Single drop studies have shown that there are significant differences in the way liquid hydrocarbons containing additives burn, compared with neat fuels. An experimental study was performed to evaluate how this change in burning behavior might affect the stability behavior of a model rocket combustor. Ammonia borane was investigated as a potential additive using a continuously variable resonance combustor, which has been shown to generate self-excited instabilities. The neat fuel was ethanol. High-speed imaging was used to obtain approximate heat release and boron oxidation profiles near the injector. Whereas neat ethanol exhibited only one mode of unsteady heat addition, it was found that the addition of the ammonia borane resulted in a bimodal temporal and spatial unsteady heat addition. This bimodal behavior produced combustion dynamics that allowed the system to be unstable over a wider range of geometric configurations.