Solid Fuel Scramjet Combustor Research Articles

Influence of Axial Flow Field Parameter Fluctuations on Performance of Scramjet Combustor

Abstract When air-breathing aircraft is flying over a wide area, the upflow of the combustor central axis will change greatly, and the flow field parameters will fluctuate along the axis. Therefore, it is necessary to carry out a study on the influence of axis flow field parameter fluctuation on combustion chamber performance, so as to provide relevant theoretical support for the combustion chamber of air-breathing aircraft in wide-area flight design. Based on the N-S gas-phase control model, combined with combustion model, turbulence model, burning rate model and mass transfer model, a numerical simulation model of flow combustion in solid fuel scramjet combustor is established. Through this model, the influence of the upflow field parameter fluctuation of the central axis on the combustor performance is carried out. The results show that for a optimized combustor configuration, different inlet air flow rates will cause the Mach number of the combustor to oscillate along the flow direction. The greater the oscillation amplitude, the greater the total pressure loss. Too high or too low inlet air flow rate will both result in Mach number oscillation in the combustor. However, selecting a suitable inlet air flow rate can significantly reduce the Mach number oscillation of combustor flow. Therefore, for different combustor configurations, appropriate inlet air flow should be designed to reduce the Mach number oscillation of the flow field in the combustor, reducing the flow loss and improving the working performance of the combustor.

Mixing and heat release characteristics in the combustor of solid-fuel rocket scramjet based on DES

Based on the previous experimental results, in order to improve the understanding of the mixing and heat release process of the internal flow field in the solid-fuel scramjet combustor, and to provide theoretical reference for the layout of the combustor configuration in the next step, the numerical simulation method of DES combined with the gas mixing degree, mixing layer thickness and mixing fraction was used to analyze quantitatively the microscopic structures and the mixing combustion enhancement mechanism in the combustor at different lobe sweep angles based on the influence of the change of lobe sweep angle on engine performance, and then the influence law was explained from the microscopic point of view. Results demonstrate that 1) based on the engine configuration in this paper, in the design process of staggered support plate structure, the lobe sweep angle should be based on the width and the straight section length of the support plate. And the engine performance improves with the decrease of the lobe sweep angle. 2) The growth of the reaction mixture layer is in four typical stages: the initial growth zone, the rapid growth zone, the saturated growth zone and the full development zone. 3) The downstream part of combustor is in the full development zone of the reaction mixing layer. The full development zone as the main area determining the heat release performance of the engine, the mixing degree increases with the decrease of the lobe sweep angle in the behind part of the combustor. Therefore, the pressure in the downstream flow passage of the combustor and the space of combustor occupied by the high pressure area increase with the decrease of the lobe sweep angle.

Influence of lobe geometry on mixing and heat release characteristics of solid fuel rocket scramjet combustor

The highly efficient and organized combustion of combustors should be achieved to improve the performance of solid fuel rocket scramjets. This study investigated the influence of changes in lobe sweep angle and height on the mixing and heat release characteristics of a solid fuel scramjet combustor on the basis of a staggered support plate structure introduced in a previous experimental study. The goal was to determine the optimal parameters of lobe sweep angle and height for subsequent engine design. Numerical simulation was conducted to quantitatively analyze the mixing and heat release characteristics of the combustor at different lobe sweep angles and heights in terms of combustion efficiency, gas mixing degree, and oxygen–fuel ratio. Results demonstrate that 1) the lobe sweep angle should be based on the width and straight section length of the staggered support plate structure. Engine performance improves with the decrease of the lobe sweep angle. 2) There is an optimum lobe height in the range of 0–24 mm, and the optimum value point is approximately 14 mm according to the proposed engine configuration.

Numerical investigation of a combined solid fuel scramjet combustor

A new scramjet configuration using solid fuel as propellant is proposed, namely, the combined solid fuel scramjet. The combined solid fuel scramjet combustor has been investigated numerically to predict its characteristics. Two-dimensional axisymmetric Reynolds-averaged Navier–Stokes equations and species transport equations are solved numerically. Turbulence closure is achieved using the shear stress transport k-ω model. Three global one-step reaction mechanisms are used in this study. The experimental data in the published literature has been used to validate the accuracy of numerical model. The results of the numerical analysis show that secondary combustion occurs in the supersonic combustor. At the combustor outlet, the total pressure loss is about 0.65 and the combustion efficiency is about 48%. A subsonic recirculation zone is formed near the wall between the air inlet and the fuel-rich gas inlet, which is beneficial to mixing and flame holding. A normal shock wave is generated at the end of the expansion zone after the fuel-rich gas is injected into the combustor, and this is due to the strong pressure downstream induced by secondary combustion.

Numerical study on solid-fuel scramjet combustor with fuel-rich hot gas

Solid-fuel scramjet combustor with cavity faces challenges of ignition and flame holding. In the current study, a novel concept, solid-fuel scramjet combustor with fuel-rich hot gas, is proposed. The Reynolds-average Navier–Stokes (RANS) equations coupled with the SST k−ω turbulence model and the second-order spatially accurate upwind scheme are employed to calculate its flow field. The feasibility of the solid-fuel scramjet combustor with fuel-rich hot gas is studied based on the validation of numerical method. The comparison of the performance is made between the combustor with fuel-rich gas and the combustor with cavity. Various parameters, namely excess air coefficient of gas generator and mass flow rate of fuel-rich gas, are studied, and their effects on the combustion efficiency, total pressure recovery and fuel regression rate are analyzed. This is the basic study for experiments of a solid-fuel scramjet combustor with fuel-rich hot gas.

Flow Field Computational Analysis in a Solid Fuel Scramjet Combustor

According to the domestic and foreign institutions of solid fuel scramjet experimental data, adopt the numerical simulation method, design to construct the numerical model of the combustion chamber, using RNG k- turbulence model and eddy dissipation combustion model, simulation combustor working condition, analysis of the combustion chamber flow field data, concluded that the combustion phenomenon occurs mainly in the center of the flow field area. At the entrance of the combustion chamber can produce normal shock. The cavity can reduce the air velocity and it can guarantee the stability of the flame burning, and that to calculate the specific impulse is 126N·s/kg.

Numerical Investigation of Self-Ignition Characteristics of Solid-Fuel Scramjet Combustor

The self-ignition characteristics of a solid fuel under supersonic flow have been investigated theoretically and numerically. Time-dependent two-dimensional axisymmetric compressible Navier–Stokes equations and species transport equations are solved numerically. Turbulence closure is achieved using the shear stress transport model. Polymethylmethacrylate fuel and a global one-step reaction mechanism are used in this study. The reaction rate is determined by finite-rate chemical kinetics with the turbulence–chemistry interaction modeled by an eddy-dissipation model. The numerical results generally agree with the experimental data in the published literature. The flame spread and pressurization during the self-ignition of polymethylmethacrylate in the combustor have been studied. The effect of inlet flow conditions and the geometry of the combustor on self-ignition behavior have been analyzed. Three stages of flame spread (namely, heat accumulation, secondary recirculation zone self-ignition, and orderly flame spreading) are identified during the self-ignition transient. Pressurization in the combustor during the third stage is evident. Self-ignition of the solid fuel can be affected by three kinds of limits: namely, lean oxygen, rich oxygen, and total temperature inlet conditions. A relatively long and deep cavity is beneficial to establish self-ignition. It is suggested that a step-type cavity can enhance self-ignition performance.

Numerical and Experimental Study on the Solid-Fuel Scramjet Combustor

The specific thrust change rule during the working process of the solid-fuel scramjet combustor was researched using both numerical and experimental methods. A methane-burning vitiated air heater was designed and made to simulate the flight environment of Mach 6 at a 25 km altitude. Based on the two-dimensional numerical flowfield results, a new quasi-one-dimensional numerical method was established to simplify the unsteady combustion and flow of the solid-fuel scramjet. This method can be used to calculate the parametric changing process of the combustor quickly. Self-ignition and fuel regression rate characteristics in the current experiment were consistent with the results from previous works. The agreement between the numerical and the experimental results was generally good as well. The specific thrust can reach a level of in the designed flight environment. Moreover, it was found that the specific thrust decreased with the total pressure loss increasing during the working process.

Numerical investigation on cavity structure of solid-fuel scramjet combustor

The influences of miscellaneous combustor structures for solid fuel scramjet combustion on the performance are investigated, including a detailed interaction analysis between shocks/waves and combustion. Hydroxyl-terminated polybutadiene is chosen as the solid fuel with the non-premixed equilibrium probability density function combustion model. The results show combustion enhancement when structure of combustor is modified. The radical emphasis is to examine the sensitivity of the properties due to variations on the length-to-depth ratio of cavity, aft wall angle, and offset ratio. It is noted that there is an appropriate structure of cavity (L/D=4, θ=45°, and Dd/Du=1.25–1.5) regarding the combustion efficiency, total pressure loss and specific impulse. The observation of function for combustor components provides instructional insight into the design considerations for a combustor of a solid-fuel scramjet.

Quasi-One-Dimensional Numerical Method for Solid Fuel Scramjet Combustor Analysis and Design

AbstractTo research the performance of a solid fuel scramjet combustor, a solid fuel regression rate model was coupled into a quasi-one-dimensional flow equation, and so the flow parameters and fuel regression rate could be calculated together. According to the regression rate and the combustor diameter in the previous moment, the combustor diameter of the next moment could be obtained. The unsteady combustion and flow matter in the combustor was simplified into a steady calculation of every moment by solving the boundary condition at different moments independently. The numerical results were compared to actual experimental data from the literature, and they agreed well with the experiment. Taking the fuel/air ratio and Mach number as the optimization conditions, the initial size of the combustor could be obtained. On the basis of an optimized combustor, the relevant parameters’ variation rules were calculated and analyzed. It was found that during the working process, the combustor flow field could be k...

Numerical Investigation on Internal Regressing Shapes of Solid-Fuel Scramjet Combustor

The purpose of this paper is to investigate the combustion process and its change with time due to solid-fuel regression in a solid-fuel scramjet combustor. The combustion process is simplified in the sequence changing of three regression shapes. The solid fuel uses hydroxyl-terminated polybutadiene with a global one-step reaction mechanism. Numerical simulations based on the shear stress transport model are confirmed by experiments. The simulation data generally agree with the experimental data published in the literature. For simplification, three typical instants during the combustion process, implying three different internal shapes due to fuel regression, are considered. With the sequence of three regression shapes, the fuel diffusion from a solid-fuel surface becomes more extensive and, furthermore, enhances the global combustion efficiency, which is nearly up to 45% at the outlet of the combustor. The simulation results also reveal that the cavity in the combustor provides a fuel enrichment zone and a stable pool of hot reaction products. A thermal choke is caught in the flowfield for the initial geometry, whereas a shock train appears in the main flow and has a tendency to extend with the surface regression. The investigation of the combustion process in this work provides some insight into the design considerations for a combustor with a cavity for a solid-fuel scramjet.

Theoretical study of a solid fuel scramjet combustor

The combustion of a solid fuel under supersonic cross flow conditions was investigated theoretically. A two-dimensional, axisymmetric, turbulent ( k– ε), two-reaction, six-species reactive flow model was developed and solved numerically. The nominal case investigated simulated stagnation conditions resulting from flight Mach number of 5 at 15 km for a designed combustor entrance Mach number of 1.5. The results demonstrate the feasibility of solid fuel combustion under supersonic cross flow where the diffusion flame supplies a substantial heat addition to the flow. The sensitivity of the combustor performance to flight Mach number, combustor inlet Mach number, combustor size, step size and length of fuel grain was studied. The effects on flow pattern, regression rate profiles as well as flame temperature and location were evaluated. Typical computed combustion efficiencies of the order of 0.7–0.9 support potential use of solid fuel scramjet.

Investigation of a Solid Fuel Scramjet Combustor

The combustion of a solid fuel under supersonic crosse ow conditions in a scramjet cone guration has been studied experimentally. Self-ignition and sustained combustion of poly-methyl-methacrylate with no external aid (such as reactive gas injection or a pilot e ame ) were demonstrated in static tests simulating a e ight Mach number of about 5 at high altitude. The appropriate inlet conditions, i.e., stagnation temperature and pressure in excess of 1200 K and 16 atm, respectively, were provided by a vitiated air heater. The diverging combustion chamber included a fore-end e ame stabilization zone, whose e ameholding limits were determined experimentally. Flow and combustion phenomena were studied both by pressure measurements along the fuel grain and by video recording, taking advantage of the fuel transparency. Comparison between theoretical results of a one-dimensional e ow model with test data showed fair agreement, indicating the existence of a supersonic e ow regime within the combustor. The video pictures provided temporal and spatial fuel regression rate data, using a computerized image analysis system. Nomenclature