Model Scramjet Combustor Research Articles

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.

The impact of wedge and wavy-wall combustor side walls on shock-induced mixing and combustion in a hydrogen-fuelled model scramjet combustor

The mixing and combustion processes are linked and challenging for a supersonic combustion engine. Supersonic air doesn't spend much time in the combustor, making it challenging to blend fuel with supersonic air. The expansion of shear-mixing layer with shock waves and their interactions are the primary characteristics of mixing augmentation. This paper investigates the same methodology for a model scramjet combustion chamber with wedge and wavy sides. The novel scramjet combustion models include wedge and way-wall sidewalls based on the basic scramjet model. By combining the finite volume approach with a second-order upwind discretization method, the Reynolds-averaged Navier-Stokes equations are numerically analyzed. The interplay between wedge and wavy-wall effects and free stream flows is taken into consideration by using the shear stress transport (SST) k−ω turbulence model. The effect of the wedge and wavy wall combustor has been evaluated by analysing the shock production and their interactions with the mixing layer and internal flow structure. The turbulence characteristics are also examined to analyse the degree of flow disturbance for wedge and wavy-wall combustion. The wavy wall combustor model results in a 14.5% increase in turbulence intensity, which raises air-fuel mixing and there by improves mixing and combustion efficiency by 12.59% and 11.23%, compared to the base wedge type combustor.

Strut-assisted injection of liquid fuel in a supersonic combustor

This paper investigates mixing and combustion of liquid fuel in a model scramjet combustor with strut-assisted injection adjacent to a cavity flame holder. Fuel/air mixing in a Mach 2.2 cross-flow and stagnation temperature of 1300 K was studied using filtered IR imaging. The insight gained by this method is discussed along with the method limitations. The impact of the strut assisted injection on the mixing of the liquid fuel was seen by the shorter vaporization distance, as the fuel vaporized upstream of the cavity, in contract to a liquid jet that expanded above most of the cavity when the strut was absent. Moreover, the fuel jet was seen to propagate laterally to the tip of the strut, which lead to lateral dispersion of the fuel up the strut height. The combustion experiments were conducted by piloting the cavity with ethylene and were imaged using CH* chemiluminescence. At stagnation temperature of 1300 K thermal choking and stable combustion occurred at equivalence ratio of 0.43, with the flame stabilizing at the strut wake, above the cavity. Intermittent thermal choking occurred at leaner condition. The flame stabilized on the cavity shear layer when the shock train moved downstream of the cavity and the flame stabilized at the strut wake when the shock train moved upstream. Thermal choking with the strut guided injection relative to a combustion only in the cavity without the strut is the result of the observed higher mixing efficiency, that enabled flame spreading into the supersonic flow and higher heat release. At a higher stagnation temperature of 1500 K the shock train moved downstream of the cavity, and the flame stabilized on the cavity shear layer. Finally, it is shown that the combustion efficiency could be further improved by injecting the fuel from the strut side walls.

Combustion instability analysis in an ethylene-fueled scramjet combustor under various fuel penetration height conditions using an image-based nonlinear dimensionality reduction method

The present study investigates the effects of fuel penetration height on combustion instabilities in an ethylene-fueled scramjet model combustor with a cavity flameholder. Experiments were performed at the stagnation temperature of 1900 K, the stagnation pressure of 0.37 MPa, and the Mach number of 2. Three fuel injection orifice diameters:2.3, 2.4, and 2.5 mm, were tested to elucidate the effects of ethylene penetration height on combustion instabilities. Furthermore, high-speed measurements of CH* chemiluminescence and shadowgraphs were performed with high-speed video cameras. To examine the dynamics of the combustion instabilities, time-resolved CH* chemiluminescence images and shock parameters, extracted from snapshots of shadowgraphs, were analyzed using a non-linear dimensionality reduction algorithm: Gaussian Process Dynamical Model (GPDM). Thereafter, further analyses on the acquired latent variables were conducted with Recurrence Plot (RP). The experimental results showed that the fuel equivalence ratio (ϕ) range for cavity shear-layer combustion mode expanded as d decreased. Furthermore, for ϕ = 0.18, the instability behavior remarkably changed at around d = 2.4 mm. Therefore, the instability dynamics for ϕ = 0.18 were investigated using GPDM and RP including results from our previous study (d = 2, 3, and 4 mm), revealing differences in the instability behaviors. For d = 3 and 4 mm, jet-wake combustion and ram combustion modes were established alternately with a frequency of about 1600 Hz. In contrast, at d = 2.4 and 2.5 mm, although a similar instability in the d = 4 mm case was present at almost the same oscillation frequency, a different instability behavior was also confirmed. This additional instability exhibited an intermediate state between jet-wake combustion and ram combustion modes. These two instabilities emerged aperiodically. For an unstable combustion in shear layer observed for d = 2 and 2.3 mm, corresponding RPs exhibited black patches, indicating that the oscillation amplitude diminished substantially.

Reconstruction of Supersonic Flowfield Using Physical Neural Network Based on Channel Interaction

The evolution of shock waves in a combustor is a complex process comprising multiple time series, which results in a long computational cycle for unsteady numerical simulations. Current data-driven methods integrating physical information provide novel opportunities to rapidly and accurately predict the flowfield in the combustor. In this study, we construct a dataset under Mach numbers from 2 to 5 based on a self-designed scramjet combustor model and propose a physical convolutional neural network flowfield reconstruction method based on channel interaction (PICNNCI). This method embeds Euler equations that characterize the flow characteristics of inviscid fluids as physical constraints to solve the low-accuracy problem of traditional convolutional neural networks arising due to insufficient prior information. The model first performs dimensional transformations on temporal and spatial information input and converts a one-dimensional column vector into a two-dimensional vector to meet the convolution operation rules. Then, convolution is used for feature extraction to output the predicted flowfield, including velocity, pressure, and density fields. Finally, the predicted flowfield is optimized using the Euler equation and compared with the results obtained by numerical calculation and traditional neural network methods. The average peak signal-to-noise ratio of PICNNCI reaches 34.01 dB, and the linear correlation reaches 0.994.

Numerical investigations of supersonic combustion characteristics of thermally cracked fuels at typical pyrolysis temperatures

Insight into supersonic combustion characteristics of thermally cracked fuels in a scramjet combustor is crucial for the development of next-generation supersonic flight vehicles. In the present study, four typical thermally cracked fuels at pyrolysis temperatures of 450, 550, 600, and 650 ℃ (denoted as ckori, ck550, ck600, and ck650) were selected. Improved delayed detached eddy simulation (IDDES) modeling with a compact skeletal mechanism and a transported multi-regime flamelet model was carried out to simulate the supersonic combustion of the target fuels in a full-scale scramjet combustor. Supersonic flow, mixing, and combustion characteristics of these fuels were quantitatively compared. Discrepancies in the combustion process are mainly at the cavity region, where the mixing efficiencies and combustion intensities of the first three fuels increase monotonically with the increasing pyrolysis temperature, while that of ck650 locates between ck550 and ck600. The increase in pyrolysis gas content will enhance the mixing process, while the increase in injection pressure will suppress the heat/mass transport between fuel and incoming air. The mixing process dominates the distributions of intermediate radicals and main products. The first three fuels tend to combust under fuel-rich condition, while ck650 tends to combust under the stoichiometric condition, which explains the distributions of OH, CO, and CO2 concentrations. The premixed and non-premixed combustion modes coexist in the model scramjet combustor, and the non-premixed combustion mode dominates the combustion process of thermally cracked fuel.

Large eddy simulation of shock-combustion-film interaction at hydrogen injection in a model scramjet combustor

Supersonic film cooling, which utilizes fuel onboard, is a promising method for protecting the wall of a scramjet combustor from intense heat and friction loads during high-speed flight. Large eddy simulation (LES) is conducted to study the coupled interactions among shock, combustion, and film. First, reference solutions for individual effects of shock and combustion on film performance are computed. Then the coupled effects are analyzed based on the reactive hydrogen film impinged by an oblique shock at a deflection angle of 2° The results show that the incident shock will enhance the transport processes after the impingement point by an increased turbulence production in the separation bubble. On the contrary, boundary layer combustion has the effect of constraining the transport processes in the mixing layer due to the low-density region caused by thermal expansion. This leads to a decrease in the mixing rate and wall friction. Although the heat flux transported to the film is also reduced by combustion, the wall temperature still increases due to the excessive compensation of heat from intense endothermic chemistry. Besides, the boundary layer combustion of hydrogen film exhibits self-limited characteristic since the mixing process is further delayed under combustion circumstance. Therefore, an incident shock will break the mixing bottleneck of the boundary layer combustion, leading to a more intense reaction with a more wrinkled flame front. The current findings indicate that the additional gaining of cooling and friction reduction performance is exclusive, but also point out the direction for simultaneous improvement of the two performances.

Direct numerical simulation of supersonic internal flow in a model scramjet combustor under a non-reactive condition

A Mach 1.5 non-reactive flow in a cavity-stabilized combustor of a model scramjet is studied via a direct-numerical simulation approach, and the analysis is focused on the interaction among boundary layer, free shear-layer above the cavity and shock wave. It is found that the impingement of the free shear-layer on the aft wall of the cavity leads to strong turbulence kinetic energy, high local pressure, and a fan of compression waves. The compression waves evolve into an oblique shock, which reflects between the upper and lower walls and interacts with the boundary layers attached to the two walls. The analysis of the turbulence production reveals that the amplification of turbulence in the core of the shear-layer and around the reattachment point is mainly due to the shear production, but the deceleration production mechanism presents a significant impact in the regions above the aft wall of the cavity and around the shock interaction points. The very low frequency commonly observed in shock wave/boundary layer interactions is not observed in the present research, which might be due to the low Reynolds number of the studied case.

Experimental investigation on the ignition limits of a kerosene-fueled cavity in model scramjet combustor

The ignition limits of liquid kerosene injected upstream of a cavity flameholder using a spark plug are investigated in a model scramjet combustor with the isolator entrance Mach number of 2.52, the total pressure of 1.6 MPa, and stagnation temperature of 1486 K. The effects of injection pressure and mixing distance on ignition are studied. Flame dynamics and the evolution of the flow field during the ignition process are acquired through high-speed photography and Schlieren photography. For near-field injection, successful ignition of kerosene can only be achieved when the injection pressure is extremely low or in a higher injection pressure range. The ignition is failed due to the fuel-rich local equivalence ratio in the cavity recirculation zone when the injection pressure is in the middle of injection pressure range. For far-field injection, the ignition trend is totally opposite compared with near-field injection. The ratio of the fuel that the cavity can react completely to that consumed by the mainstream is extremely low due to the minimal mass exchange rate between the cavity and the mainstream. The local equivalence ratio in cavity is the controlling factor determining whether successful ignition could be achieved. Only local ignition with low combustion efficiency is achieved in the present single-cavity combustor. The flame could not spread to the mainstream due to the lower turbulence level weakening the interaction of combustion products in the recirculation zone with unburned reactants in the supersonic flow.

A method for optimizing reaction progress variable and its application

This paper presents a practical method for optimizing reaction progress variables for FPV models, and the potential of this method is demonstrated in the application of a scramjet model combustor simulation. Reaction progress variable is commonly employed in flamelet-based models as a reaction tracker, while it is normally given with less verifications. Focusing on the global reaction process, this optimization method is proposed with respect to the constraint of monotonicity and the target of the heat release rate, using a hybrid PSO/GA algorithm to quickly give the weighting factors. Consequently, the optimized progress variable together with another three progress variable expressions are utilized and compared in the simulations of a supersonic combustor with vaporized RP-3 injection upstream of dual parallel cavities. Noting that, a pressure-related FPV model is presented and adopted here to consider the effect of pressure pulsations in the combustion flow field. It is shown that the progress variables appear to be specifically coupled with the heat release process and the extent of reaction proceeding, thus affecting the temperature, pressure, and components. Significant improvements brought by the optimized progress variable over the conventional expressions are presented, and it is noteworthy that this method is also applicable to the original FPV model.

Experimental study on the flame stabilization mode in the rear-wall-expansion cavity

The present study experimentally investigates the transition of flame stabilization mode at Ma = 2.92 using a model scramjet combustor. The combustion mode is analyzed by pressure measurement and quasi-one-dimensional analysis. The effects of both equivalence ratio and aft wall offset on flame behaviors are further characterized by CH* chemiluminescence. The results show that the transition from cavity stabilized combustion mode to jet-wake stabilized combustion mode is realized by increasing the equivalence ratio. And the quasi-one-dimensional analysis reveals that all cases work in the Early Scram mode and approach thermal chocking after increasing equivalence ratio. As the equivalence ratio increases, the amplitudes of the flamebase and heat release gradually increase, and their distributions gradually migrate downstream. The amplitude of flamebase and heat transfer remains prominent with the consistent rise in equivalence ratio. At the same time, the spatial distribution of heat release tends to spread out upstream and occasionally appear upstream of the leading edge of the cavity, indicating that the flame stabilization mode transforms from cavity stabilized combustion mode to jet-wake stabilized combustion mode. The spatial variation of local heat release rate is responsible for the transition of flame stabilization mode. A slight aft wall offset promotes the transition from cavity stabilized combustion mode to jet-wake stabilized combustion mode at a higher equivalence ratio.

Enhancement of blowout limit in a Mach 2.92 cavity-based scramjet combustor by a gliding arc discharge

Plasma-assisted combustion (PAC) has shown excellent performance in the ignition and combustion enhancement of scramjet combustors. A multi-channel gliding arc (MCGA) plasma placed on the sidewall surface of a scramjet combustor was employed to enhance combustion near the flame blowout limit (BL) in a C2H4-fueled and cavity-based model scramjet combustor. Wall static pressure measurements, simultaneous 10 kHz CH* chemiluminescence imaging from the top and side views, optical emission spectroscopy, and discharge waveform measurements were used to elucidate the combustion physics throughout the PAC events. For the current cavity configuration and fueling condition, a general estimate of the flame BL in a rich-fuel environment was provided, in which a fuel flow rate (FFR) was 7.5 g/s. Once the FFR was exceeded, the flame was unstable and blown out, whereas the MCGA plasma demonstrated an excellent flame stabilization ability. The flame BL of the scramjet combustor in the presence of the MCGA plasma has an increase of ∼29%. Two PAC processes, including enhancement mode and re-ignition mode, were distinguished to play a vital role in extending the flame BL. The two PAC processes showed that the spark-type discharge generated by the MCGA could ignite the local fuel-rich mixture and increase the flame BL. The local ignition is dominated by the continuous accumulation of the massive flame kernels with the local fuel-rich mixture, as well as more heat and species exchanges.

Experimental investigation on the flame stabilization modes of liquid kerosene in a cavity-based scramjet combustor

The flame stabilization modes of liquid kerosene is investigated in a model scramjet combustor with flight condition of Ma 6. The effects of equivalence ratio and cavity configuration schemes on the flame stabilization modes are studied. Flame propagations and the evolution of the flow field from the ignition to the establishment of stable flame in the supersonic flow are acquired through high-speed photography and schlieren photography. Only cavity-recirculation-region stabilized combustion can be obtained in the single-cavity scramjet combustor. The intensity of combustion is weak and is termed as local ignition with low combustion efficiency. In the tandem-cavity scramjet combustor, the flame stabilization mode is largely determined by the total equivalence ratio. The flame could spread into the mainstream and even propagate against the supersonic flow as the equivalence ratio increases. The positive feedback mechanism between combustion and precombustion shock train is responsible for the evolutions of the flame stabilization modes, where the interactions between shock waves and turbulent boundary layer play a key role in the flame propagation process. The large-scale eddies in the shear layer of the upstream cavity shed at the cavity aft wall and could enhance the turbulent levels of the incoming flow of the downstream cavity. Consequently, an intense kerosene flame can be ignited in the tandem-cavity combustor. The flame could not spread to the mainstream in the single-cavity combustor due to the lower turbulence level weakening the interaction of combustion products in the recirculation zone with unburned reactants in the supersonic flow.

The numerical investigation of combustion performance of scramjet combustor with variation in angle of attack

The present article involves computational investigation of parallel fuel injection based scramjet combustor. The wedge shaped strut face is used as a fuel injector. Selected geometry is first validated and then further investigation is performed. Steady-state, two dimensional, scramjet combustor model has been chosen to complete the numerical convergence through ANSYS Fluent software. Grid independence analysis has also been performed. Reynolds Average Navier Stokes equation in addition with k-epsilon turbulence modelling have been utilised to reach the convergence at a lower computational cost. Finite rate eddy-dissipation based Species transport modelling is chosen to solve the chemical kinetics between hydrogen and air. The incoming boundary condition of free-stream air has been optimized to improved combustion efficiency. There are three selected model is utilised for comparison i.e. zero degree, positive five and negative five degree air angle of attack model. To change the angle of attack of the incoming air, a Modified isolator is added ahead of the combustor with constant length. It is observed that the behaviour of shock waves and flow properties are dependent on the angle of attack. Comparative observation has been analysed with monitoring the combustion efficiency graph. Maximum combustion efficiency reaches up to 93% in negative five degree air angle of attack model moreover, mixing is also improved by 4%. It can be summarized that the scramjet combustor performance is primarily highly influenced by the geometry configuration and the nature of flow of incoming air.

Effect of the injector flow field on the performance of a model scramjet combustor

Effect of the injector flow field on the performance of a model scramjet combustor

Effect of the injector flow field on the performance of a model scramjet combustor

Effect of the injector flow field on the performance of a model scramjet combustor

Partially premixed combustion simulation using a novel transported multi-regime flamelet model

To account for partially premixed process in combustion, a novel transported multi-regime flamelet model, which gives local contributions of premixed and non-premixed regimes, is proposed with a theoretical derivation of the governing equation of the premixed weighting coefficient. The model intrinsically ensures mass conservation and physically incorporates transport properties. Thus, it avoids artificial treatment that has to be employed in algebraic multi-regime flamelet models. An ethylene-fueled combustion in a scramjet model combustor is simulated using the transported multi-regime flamelet model. The pressure distribution that characterizes the flame anchoring position and wave distribution is well predicted by the model. It is observed that the premixed regime is distributed mainly in the recirculation zone upstream of the fuel orifice, inside the cavity and close to the lower wall. Besides, there is an apparent instantaneous distribution of the premixed regime in the fuel jet wake, which is attributed to the enhanced premixing by transient vortices. However, the overall contribution of the premixed regime is not as great as that of the non-premixed one, indicating that combustion is non-premixed dominated. The evolution of the OH concentration shows that the OH first emerges close to the cavity ramp. After that, it propagates upstream into the cavity and meanwhile downstream along the lower wall of the expanding section. The isoline of the premixed weighting coefficient used to mark the temperature interface reveals that the premixing can enhance combustion. In contrast, numerical oscillations and a pseudo regime distribution indicating a premixed-dominated combustion that conflicts the physics is observed in a reference simulation using an algebraic multi-regime flamelet model. It is thus demonstrated that the transported multi-regime flamelet model is superior over the algebraic one in the prediction of partially premixed combustion with multi-regimes involved.

Reduction of surface friction drag in scramjet engine by boundary layer combustion

This study investigates the effects of boundary layer combustion on skin friction reduction in a model scramjet combustor. The 4-equation RANS model (Transition SST model) is employed as the turbulence model and one experimental case which involves supersonic turbulent boundary layer combustion is used for validating the numerical simulation method. Then a wall jet device which can be used to inject hydrogen into boundary layer is designed and added to the lower wall of the combustor in the model scramjet engine. The numerical results showed that the optimal ratio between the primary fuel, producing thrust, and the wall jet fuel, used for drag reduction, for getting the largest skin-friction drag reduction when installing the wall jet device in the combustor, is 3:1. The study of modifying the location of the wall jet device showed that setting the wall jet position too close to the upstream or downstream could not get the maximum overall drag reduction. The installation position in the flow path depends on the coupling of primary fuel and wall jet fuel combustion. In practical scramjet design, the distribution ratio of primary to wall jet fuel and the location of wall jet device should be weighed to improve skin-friction reduction efficiency and overall engine performance.

Effect of fuel injection distance and cavity depth on the mixing and combustion characteristics of a scramjet combustor with a rear-wall-expansion cavity

The mixing and combustion characteristics of a model scramjet combustor with a rear-wall-expansion cavity under a Mach 2.0 inflow are experimentally and numerically investigated. Two different cavity depths and two fuel injection distances are compared to show the effect of these parameters. High-speed imaging system is used to capture the oscillation of flamebase and the flame oscillation of cavity shear layer in the transversal direction. Numerical simulation is carried out to reveal the jet mixing process around the cavity. The experimental results show that the combustor with longer fuel injection distance and smaller cavity depth has weaker flame and lower wall pressure due to the significant oscillation. The oscillation of flamebase is found to be stronger in the shallower cavity due to its smaller volume. Cavity with short fuel injection distance and large depth could suppress the oscillation, and the combustor can achieve a robust and strong combustion. The simulation results show that under the shorter injection distance, the center of cavity shear layer is rolled up to the main stream; thus the shear layer can interact with the fuel jets more intensively, which improves the mixing efficiency and promotes strong combustion. Compared to the cavity with larger depth, the shear layer in the smaller cavity has a larger thickness and displays more significant shear layer oscillation in the transverse direction. The oscillation of the cavity shear layer can promote the mixing efficiency but is detrimental to combustion. Therefore, a proper cavity depth should be carefully considered during the design of a cavity flameholder.

Kerosene Combustion in a Pseudoshock with Varied Conditions at the Scramjet Combustor Model

Kerosene Combustion in a Pseudoshock with Varied Conditions at the Scramjet Combustor Model