Supersonic Combustion Of Hydrogen Research Articles

Design of a plasma-assisted injector: Principle, characterization and application to supersonic combustion of hydrogen

A new design of plasma-assisted injector to ignite or stabilize combustion is presented. This design is intended to produce discharges that penetrate the combustion chamber instead of staying at the wall. This plasma-assisted injector, simple and robust, allows compact designs for a weakly intrusive implementation in engines, and is compatible with either air or fuel, depending on the process or application. A prototype was manufactured, implemented and tested with quasi-DC discharges on the LAPCAT2 Dual Mode Ramjet combustor at LAERTE facility in ONERA. Combustion tests showed earlier ignition and increased pressure when the plasma was turned on with a measured plasma power of 1.4 kW.

Self-Adaptive Turbulence Eddy Simulation of Supersonic Combustor Considering Thermal Radiation

The newly developed self-adaptive turbulence eddy simulation (SATES) method coupled with three turbulent combustion models, i.e., the finite-rate model, eddy-dissipation model, and steady laminar flamelet model, is used to numerically study the non-premixed supersonic hydrogen combustion in a DLR (German Aerospace Center) model scramjet combustor with a wedge flame holder. The study investigates the interactions of the shock waves and the flowfields as well as the flame structures. The SATES-based combustion simulation results demonstrate satisfactory agreement with the available experimental data. The study also explores the effect of fuel injection temperature on the combustion process and reveals that an increase in hydrogen temperature improves combustion efficiency. Additionally, thermal radiation effects are considered with the discrete ordinates method/weighted-sum-of-gray-gases method. The results indicate that thermal radiation heat transfer reduces the flame temperature and significantly affects the flame shape and temperature fluctuations. Also, radiation heat flux is an important factor and should be included in supersonic combustion simulations. The study demonstrates the potential of the SATES method for complex supersonic combustion and also provides a reference for thermal radiation in supersonic combustion.

Numerical Modeling of Supersonic Combustion of Hydrogen in the Wake Region of a Strut

Numerical modeling of combustion of a sonic hydrogen stream injected into a supersonic stream of air using a strut is presented. Results are obtained by numerically solving unsteady, compressible, turbulent Navier-Stokes equations, using Unstructured Finite Volume Method (UFVM) incorporating RNG based κ−ε two equation model and time integration using three stage Runge Kutta method. Reaction is modeled using an eight step H2-air chemistry. The developed numerical procedure is based on the implicit treatment of chemical source terms by preconditioning and solved along with unsteady turbulent Navier Stokes equations explicitly. Code is validated against a standard wall jet experimental data and is successfully used to model the reacting flow field resulting due to combustion of hydrogen in the wake region of strut placed in heated supersonic air stream.

Computational study of hydrogen and air co-flow jets mixing in dual-combustor ramjet in presence of ramp shock generator

Flame stabilization distribution and efficient fuel mixing are highly significant in the performance of ramjet engine. In this paper, numerical studies are done to investigate the supersonic combustion of hydrogen and air co-flow in presence of the ramp edge. In this study, flame maintenance and combustion formation in a dual-combustion ramjet engine are fully investigated. The primary emphasis of this work is to disclose the impact of produced shocks due to the existence of ramp edge on the structure of fuel and air-jet immediately downstream of injector. The shock interactions of two different angles (10 deg and 30 deg) of the ramp edge are investigated. Favre-averaged conservation equations are considered for the simulation of compressible flows inside the combustor. Various mechanisms expressing the flame spreading characteristics are also examined. Depending on the angle of the ramp edge, the produced shock waves influence on size and strength of vortices. Our findings also show that high angles of ramp edge augment the fluctuation of the vortices by the splitter and consequently, fuel mixing enhances in the combustion chamber of ramjet engine. Our findings also confirm that the similar jet pressure would be more effective on the vortex production downstream of the splitter. The application of ramp edge with angle of 30 deg would increase average momentum thickness of the mixing layer up to 55%.

Evaluation of flamelet/progress variable model for the applications in supersonic combustion using hybrid RANS/LES approach

The Flamelet/Progress Variables (FPV) model coupled with large eddy simulation (LES) has been performed to appraise its applicability and performance for supersonic hydrogen combustion, based on the simulations of strut based scramjet and cavity based combustors. For the two typical schemes involved in the present work, the validation of the numerical solver compiled by in-house code is realized by the comparisons with the experimental measurements including temperature, velocity, and wall pressure. The methods for calculating mixture fraction variance in the framework of LES are also investigated and examined in both cases. It can be observed that the diffusion combustion mode is in the dominant state for the whole reactive zone in both cases, where the FPV model is physically compatible. Through the comparisons of the scale similarity method and the transport equation method for calculating mixture fraction variance, the importance of calculating this sub-grid turbulent scalar flux is emphasized. The results show that the mixture fraction variance calculated by the scale similarity model is locally high, but its diffusion is insufficient, which makes the local temperature of the combustion flow field overpredicted and concentrated. On the contrary, the transport equation of the mixture fraction variance can overcome this problem and improve the combustion prediction accuracy, thus, it is more recommended if the computing resources are guaranteed. Qualitatively, the Flamelet/Progress Variables (FPV) model with the large eddy simulation approach can accurately capture the structure and characteristics of the supersonic combustion flow field.

The Effect of Fuel Injection Location on Supersonic Hydrogen Combustion in a Cavity-Based Model Scramjet Combustor

Supersonic combustion experiments were performed using three different hydrogen fuel-injection configurations in a cavity-based model scramjet combustor with various global fuel–air equivalence ratios. The configurations tested were angled injection at 15° to the flow direction upstream of the cavity, parallel injection from the front step, and upstream injection from the rear ramp. Planar laser-induced fluorescence of the hydroxyl radical and time-resolved pressure measurements were used to investigate the flow characteristics. Angled injection generated a weak bow shock in front of the injector and recirculation zone to maintain the combustion as the equivalence ratio increased. Parallel and upstream injections both showed similar flame structure over the cavity at low equivalence ratio. Upstream injection enhanced the fuel diffusion and enabled ignition with a shorter delay length than with parallel injection. The presence of a flame near the cavity was determined while varying the fuel injection location, the equivalence ratio, and total enthalpy of the air flow. The flame characteristics agreed with the correlation plot for the stable flame limit of non-premixed conditions. The pressure increase in the cavity for reacting flow compared to non-reacting flow was almost identical for all three configurations. More than 300 mm downstream of the duct entrance, averaged pressure ratios at low global equivalence ratio were similar for all three injection configurations.

Supersonic combustion of hydrogen using an improved strut injection scheme

Numerical investigation of mixing is performed at Mach 2.0 model Scramjet combustor employing parallel strut injection schemes for fuel. In the present investigation, basic strut injector is modified in such a way to produce additional vortices in streamwise direction and improve fuel-air mixing. Air is injected at Mach 2.0 at the combustor inlet and fuel is injected at sonic speed from the blunt end of the strut. The flow field involving high-speed turbulent mixing and heat addition was modeled by three-dimensional Reynolds averaged Navier-Stokes equations. A realizable k-ε model was chosen to close the turbulence problem with the default model constants. Non-premixed combustion of hydrogen and air is modeled using the mixture fraction β-pdf framework. Turbulence-chemistry interactions are handled by a strained flamelet model. Comparisons of numerical results with experimental results have demonstrated the accuracy and applicability of computational grid and a numerical scheme for hot and cold flow solutions. The shock-shear layer interaction present within the combustor increases the local turbulent intensity and has a positive effect on mixing. The mixing efficiency obtained with improved strut injector is compared with the basic strut. Improved strut injection scheme showed a mixing efficiency of >95% with a 45% reduction in length. Further combustion efficiency is calculated in the streamwise direction and plot follows the similar trend as the mixing efficiency. The proposed modification of strut geometry showed improved mixing and combustion performance.

Differential Turbulent Supersonic Combustion of Hydrogen, Methane, and Ethylene, Without Assisted Ignition

Although hydrogen has the desired ignition properties for supersonic combustion in a scramjet, it has the disadvantage of low energy density, thereby motivating the interest in alternate, mostly hydrocarbon, fuels, such as methane, ethylene, and kerosene. Because the hydrocarbon fuels do not ignite easily, their use in scramjet combustion, where the strain rate is large and flame stability is difficult to maintain, depends on assisted ignition. However, a comparative evaluation of the combustion characteristics of these fuels under realistic turbulent supersonic conditions in a scramjet engine, and within the framework of advanced and highly efficient combustion modeling with detailed chemical mechanism, has not received enough attention. This study has the objective of addressing this issue. The effects of superposing ignition assistance can then be clearly delineated in subsequent studies. Three questions pertaining to the comparative combustion characteristics of hydrogen, ethylene, and methane under laminar flame condition, the significance of progress variable modeling, and turbulent supersonic combustion conditions are answered in this paper. To accomplish this, detailed analysis of opposed-jet flame is carried out for the fuels for the purpose of selecting the kinetic mechanisms, comparing the combustion characteristics, and studying the bifurcation curves for the three fuels as a function of pressure. Turbulent supersonic combustion simulations are then carried out using large-eddy simulations, from which many interesting results have been obtained. For example, of the three fuels, only hydrogen exhibits the backpressure phenomenon, which is traced to combustion-generated choking, rather than to high mechanical pressure from injection.

Numerical study on supersonic combustion of hydrogen and its mixture with Ethylene and methane with strut injection

In this paper, supersonic combustion and flow field of hydrogen and its mixture with ethylene and methane from strut injections in a Mach 2 supersonic flow are studied numerically. The fuel mixture of hydrogen, methane and ethylene represents the major products of pyrolysis of hydrocarbon fuels with large molecules such as kerosene as it acts as coolant and flows through cooling channels and absorbs heat. Detached Eddy Simulation with a reduced kinetic mechanism and steady flamelet model are applied to simulate turbulent combustion. The calculated temperature profiles of hydrogen are compared to the experimental results of DLR supersonic combustor for validation of the present numerical method. The supersonic combustion flows associated with shock waves, turbulent vortices and flame structures are studied. With addition of methane and ethylene, the flame zone moves further downstream of the strut and the maximum flow temperature at chamber exit decreases by 200 K. With analysis of total temperature ratios, it is found that combustion efficiency for hydrogen combustion is 0.91 and it decreases to 0.78 for the fuel mixture. The calculation of ignition delay time and flame speed reveals that fuel mixture of hydrogen and hydrocarbons has considerably larger delay time and smaller flame speed, that contributes to the weakened flame zone and lower combustion efficiency.

Research on effect of pollution gas components based on computational simulation and computational fluid dynamics technology on supersonic combustion

Based on computational simulation and computational fluid dynamics (CFD), this paper analyzes and compares the effects of incoming flow with clean air and air containing CO2/H2O pollution components on the performance of supersonic combustion of hydrogen fuel and ethylene fuel in the combustor through simulation and calculation. The research conclusion shows that the pressure change in the combustor obtained by the numerical simulation results and the theoretical calculation result are basically the same, indicating that both can distinguish the effect of pollution components on the performance of supersonic combustion. For hydrogen fuel, numerical simulation and calculation results show that the monomer and mixed pollution components have a strong inhibiting effect on the increase of pressure in the combustor, and the decline of pressure is nonlinear. Besides, the presence of pollutants can change the operating model in the combustor. For ethylene fuel, pollutants can lead to a decrease in the mass fraction of the chemical reaction product— H2O. Both hydrogen fuel and ethylene fuel will lead to a decrease in the temperature in the combustor and a reduction in combustion efficiency.

Large-Eddy Simulation of Supersonic Hydrogen/Oxygen Combustion: Application to a Rocketlike Igniter

The three-dimensional simulation under consideration is related to the supersonic combustion of hydrogen with oxygen. The application under study corresponds to the early phase of ignition in a rocketlike engine. The configuration is a sonic jet of hot burned gases underexpanding into a combustion chamber filled with gaseous hydrogen. Two- and three-dimensional simulations are performed to assess the numerical strategy, define the mesh requirements, and to assist in the choice of the kinetic scheme for hydrogen/oxygen combustion. The large-eddy simulation tool is used with a reduced chemistry (five species and three global steps) on a mesh resolution fine enough to use the laminar model assumption (, 680 million points). The Pope criterion and a subgrid-scale Damköhler number are provided with various definitions for the Damköhler number to validate the retained procedure for three-dimensional simulations. The sequence of ignition is detailed, revealing the flame behavior and shock arrangements. A diffusion flame develops in the supersonic mixing layer formed by the igniter flow and hydrogen injections. The flame structure is recovered with a counterflow flame configuration. A comparison with a nonreacting case shows that the driving physical mechanism is mainly linked to the aerodynamics of the igniter jet.

Investigation on supersonic combustion of hydrogen with variation of combustor inlet conditions

The present work numerically investigated the effect of variation in the inlet Mach number and stagnation temperature on the mixing of fuel with the oxidizer and the subsequent stabilization of a flame in a combustor at supersonic conditions. Dimensions of the studied combustor were taken from literature. It had a 10° wedge located at the top wall of the combustor. The combustor was modeled and analyzed using ANSYS FLUENT software. Three-dimensional, compressible, reacting flow calculations with a detailed chemistry model were performed. Turbulence was modeled using SST k-ω model. Necessary grid refinement was done to capture the incident oblique shock formed at the 10° wedge. Hydrogen was injected through the fuel inlet port. The computations were performed for Mach numbers of 2.0, 2.5 and 3.0 at the combustor inlet for a combustion inlet stagnation temperature of 1500 K. Later, the combustor inlet Mach number was kept constant at 2.5 and the combustor inlet stagnation temperature was varied as follows: 1500 K, 1700 K, and 1900 K. The results indicated that as the combustor inlet Mach number increased, the location of incidence of the oblique shock shifted to the downstream of the fuel inlet and it resulted in the better mixing of the fuel with cross flow stream of air and led to better degree of combustion of hydrogen. The contours of mole fraction of OH radical and hydrogen also corroborated the improvement in the mixing of fuel with the cross flow air and the subsequent flame stabilization at higher Mach numbers. The flow pattern, mixing of fuel with air and flame stabilization was not affected significantly till 1700 K whereas for 1900 K, combustion of hydrogen was more uniform.

On the Supersonic Combustion of Hydrogen around a Conical Obstacle Considering Variable Specific Heats with Application to Scramjets

One proposes a combined analytical-numerical method for the supersonic combustion around a conical obstacle, considering variable specific heats with temperature. One important aspect is to avoid the dissociation what is not possible if normal detonation waves (of Chapman-Jouguet type) occur. The Clarke model where the detonation wave is separated in a shock wave and a deflagation wave is able to reduce the temperature. Here conical waves are used. In order to characterize the combustion speed and intensity, new parameters are proposed. A comparison with the Chapman-Jouguet combustion is also presented.

CFD Analysis of Scramjet Combustor with Non-Premixed Turbulence Model Using Ramp Injector

This paper presents the supersonic combustion of hydrogen using strut injector along with two-dimensional turbulent non-premixed combustion model with air inlet temperature of 750 0k and vitiated Mach number of 2. In this process, a PDF approach is created and this method needs solution to a high dimensional PDF transport equation. As the combustion of hydrogen fuel is injected from the strut injector, it is successfully used to model the turbulent reacting flow field. It is observed from the present work that, the maximum temperature of 2096 0k occurred in the recirculation area which is produced due to shock wave-expansion and the fuel jet losses concentration and after passing successively through such areas, temperature decreased slightly along the axis. From the maximum mass fraction of OH, it is observed that there is very little amount of OH around 0.0017 were found out after combustion. By providing strut injector, expansion wave is created which causes the proper mixing between the fuel and air that results in complete combustion.

Hydrogen/air supersonic combustion for future hypersonic vehicles

In this work, supersonic hydrogen combustion in the Hyshot II scramjet engine is investigated. In particular, fundamental physics of mixing, combustion and vorticity generation as well as the interaction between shock waves, boundary layer and heat release are analyzed by means of 3D Large Eddy Simulations (LES) using detailed chemistry. Results show very complex structures due to the interaction between the four sonic H 2 crossflow injections and the airstream flowing at M = 2.79. A bow shock forms ahead of each H 2 injector: the interaction between bow shocks and boundary layers leads to separation zones where H 2 recirculates. In these recirculation zones, OH radicals are produced, indicating that a flame already starts upstream of the injectors and downstream of the flow separation. The formation of barrel shocks due to the H 2 expansion and recompressions is also predicted. Comparison of pressure distribution along the wall centreline at 1.3 ms shows agreement with experimental results, mostly in the first part of the combustor, where the grid is very fine. The combustion is very fast and efficient: only 12.35% of hydrogen is found unburned at the combustor exit. This confirms that burning hydrogen is efficient and feasible also in supersonic flows and therefore it is a good candidate for hypersonic airbreathing applications.

Numerical Analysis of Combustor Flow Fields in Supersonic Flow Regime with Spalart-Allmaras and k-ε Turbulence Models

In this numerical study, supersonic combustion of hydrogen has been presented for Mach 2. The combustor has a single fuel injection parallel to the main flow from the base. Finite rate chemistry model with k-e model and Spalart-Allmaras (S-A) Model have been used for modeling of supersonic combustion. The coupled phenomena of mixing and burning cm only be numerically modeled with the inclusion of a finite-rate chemical kinetic mechanism. The main issue in supersonic combustion is proper mixing within short burst of time. Attention is paid to the local intensity of heat release, which determines, together with the duct geometry, techniques for flame initiation and stabilization, injection techniques and quality of mixing the fuel with oxidizer, the gas-dynamic flow regime. The two-dimensional mathematical model involves the k-e and Spalarts-Allmaras turbulence model, modified for supersonic compressibility, and a detailed kinetic mechanism of mixture combustion. The five main parameters were considered like Mach number stagnation temperature, mass fraction, stagnation pressure and velocity. The result shows the better mixing of fuel and the flame speed increases almost linearly. The stagnation temperature in the combustion reaches up to 2820K. Fluctuation in pressure and Mach number was due to shock train. Supersonic combustion is the key enabling technology for sustained hypersonic flights. In scramjet engines of current interest, the combustor length is typically of the order of 1 m, and the residence time of the mixture is of the order of Milliseconds. Due to the high supersonic flow speed in the combustion chamber, problems arise in the mixing of the reactants, flame anchoring and stability and completion of combustion within the limited combustor length. The flow field in the scramjet combustor is highly complex. It is shown that when the flight speed is low, the kinetic energy of the air is not enough to be used for the optimal compression. Further compression by machines is needed in order to obtain a higher efficiency. For example, a turbojet employs a turbine machine for further compression. When the flight speed is higher than a certain value, the air flow entering a combustor will remain to be supersonic after the optimal compression. With a further compression (i. e. deceleration), the efficiency

Physics and Regimes of Supersonic Combustion

Understanding the physics of supersonic combustion is the key to design a performing engine for scramjet-powered vehicles. Despite studies on supersonic combustion dating back to the 1950s, there are still numerous uncertainties and misunderstandings on this topic. The following questions need to be answered: How does compressibility affect mixing, flame anchoring, and combustion efficiency? How long must a combustor be to ensure complete mixing and combustion while avoiding prohibitive performance losses? How can reacting turbulent and compressible flows be modeled? Experimental results in the past have shown that supersonic combustion of hydrogen and air is feasible and takes place in a reasonable distance, which is a necessary requirement in actual hypersonic vehicles powered by supersonic combustion ramjets. These results are explained based on a theoretical analysis of the physical mechanisms driving mixing and combustion in supersonic airstreams, where they are found to be different from those in the incompressible regime. In particular, the classic Kolmogorov scaling is shown to be no longer strictly valid, and the flame regime is predicted to be significantly affected by compressibility and different from that of subsonic flames. This analysis is also supported by the results of the numerical simulations presented, showing that by generating sufficiently intense turbulence, a supersonic combustion flame is short and can indeed anchor within a small distance from fuel injectors, with the flame typically burning in the so-called flamelets-in-eddies regime.

Investigation of the effect of chemistry models on the numerical predictions of the supersonic combustion of hydrogen

In this numerical study, the influence of chemistry models on the predictions of supersonic combustion in a model combustor is investigated. To this end, 3D, compressible, turbulent, reacting flow calculations with a detailed chemistry model (with 37 reactions and 9 species) and the Spalart–Allmaras turbulence model have been carried out. These results are compared with earlier results obtained using single step chemistry. Hydrogen is used as the fuel and three fuel injection schemes, namely, strut, staged (i.e., strut and wall) and wall injection, are considered to evaluate the impact of the chemistry models on the flow field predictions. Predictions of the mass fractions of major species, minor species, dimensionless stagnation temperature, dimensionless static pressure rise and thrust percentage along the combustor length are presented and discussed. Overall performance metrics such as mixing efficiency and combustion efficiency are used to draw inferences on the nature (whether mixing- or kinetic-controlled) and the completeness of the combustion process. The predicted values of the dimensionless wall static pressure are compared with experimental data reported in the literature. The calculations show that multi step chemistry predicts higher and more wide spread heat release than what is predicted by single step chemistry. In addition, it is also shown that multi step chemistry predicts intricate details of the combustion process such as the ignition distance and induction distance.

A comparison of the numerical predictions of the supersonic combustion of Hydrogen using the S-A and SST k – ω models

In the present work, the influence of the turbulence model on the numerical predictions of supersonic combustion of hydrogen in a model combustor is investigated. Three dimensional, compressible, turbulent, reacting flow calculations using the Shear Stress Transport (SST) k – ω model have been carried out. The results are compared with earlier results obtained using the Spalart-Allmaras model. The effect of detailed chemistry on the predictions of heat release and combustion efficiency is also investigated for one injection scheme. The calculations show that the mixing and hence the heat release is over predicted by the SA model. Whereas, the peak values for pressure and temperature are predicted better by the SST k – ω model. This investigation demonstrates the importance of the use of a two equation turbulence model over a one equation model for studying such complex flow phenomena.

Gas–Dynamic Structures in Supersonic Combustion of Hydrogen in a System of Plane Jets in a Supersonic Flow

Results of numerical and theoretical studies of supersonic diffusive combustion of a system of plane hydrogen jets in a supersonic air flow are described. It is shown that large–scale vortex structures appear in the mixing zone of the system of hydrogen jets and the cocurrent flow. These vortex structures affect the mechanism of turbulent exchange between the fuel and the oxidizer.