Performance Of Scramjet Combustor Research Articles

Investigation on the effects of operating variables on the performance of two-strut scramjet combustor

The two equation SST k−ω turbulence model and the two-dimensional compressible Reynolds-Averaged Navier–Stokes (RANS) equations along with single-step chemical reaction mechanism have been applied to analyze the effect of divergence angle, radius of the strut tip and the half angle of the strut on the combustion enhancement of two-strut scramjet combustor. Here all the simulations are carried out by using ANSYS 14-FLUENT code. Simultaneously, the validation of the present computational simulation for single strut injection has been completed by comparing its result with the available DLR experimental data which is available in the literature. The results are in satisfactory agreement with the experimental schlieren image and pressure, velocity and temperature distribution curve obtained experimentally for the case of single strut combustor. The presence of two struts enhances the mixing and combustion efficiency significantly compared to single strut and the variation of divergence angle, radius of the strut tip and the half angle of the strut also have a great impact on the performance of two-strut scramjet combustor. The present results show that there is an optimal divergence angle for which the performance of the two-strut combustor particularly increases and effective combustion phenomena occur within the combustion region. Again on increasing the radius of the strut tip, the formation of separation region takes place on both the walls of two-strut scramjet combustor which may deteriorate the performance of combustor. Further, combustion phenomena are also affected by the variation of half angle of the strut from 5° to 8°.

Effect of variation of length-to-depth ratio and Mach number on the performance of a typical double cavity scramjet combustor

The two equation standard k–ɛ turbulence model and the two-dimensional compressible Reynolds-Averaged Navier–Stokes (RANS) equations have been used to computationally simulate the double cavity scramjet combustor. Here all the simulations are performed by using ANSYS 14-FLUENT code. At the same time, the validation of the present numerical simulation for double cavity has been performed by comparing its result with the available experimental data which is in accordance with the literature. The results are in good agreement with the schlieren image and the pressure distribution curve obtained experimentally. However, the pressure distribution curve obtained numerically is under-predicted in 5 locations by numerical calculation. Further, investigations on the variations of the effects of the length-to-depth ratio of cavity and Mach number on the combustion characteristics has been carried out. The present results show that there is an optimal length-to-depth ratio for the cavity for which the performance of combustor significantly improves and also efficient combustion takes place within the combustor region. Also, the shifting of the location of incident oblique shock took place in the downstream of the H2 inlet when the Mach number value increases. But after achieving a critical Mach number range of 2–2.5, the further increase in Mach number results in lower combustion efficiency which may deteriorate the performance of combustor.

Computational simulation of multi-strut central lobed injection of hydrogen in a scramjet combustor

Summary Multi-strut injection is an approach to increase the overall performance of Scramjet while reducing the risk of thermal choking in a supersonic combustor. Hence computational simulation of Scramjet combustor at Mach 2.5 through multiple central lobed struts (three struts) have been presented and discussed in the present research article. The geometry and model used here is slight modification of the DLR (German Aerospace Center) scramjet model. Present results show that the presence of three struts injector improves the performance of scramjet combustor as compared to single strut injector. The combustion efficiency is also found to be highest in case of three strut fuel injection system. In order to validate the results, the numerical data for single strut injection is compared with experimental result which is taken from the literature.

Effect of variation of angle of attack on the performance of two-strut scramjet combustor

Computational simulation of Scramjet combustor at Mach 2.5 through two struts and with three angle of attack (α = −3°, α = 0° and α = 3°) have been presented and discussed in the present research article. Here the fuel which is injected through two struts is placed parallel to the air stream. The geometry and model used here is slight modification of the DLR (German Aerospace Center) scramjet model. Steady two and three dimensional (2D and 3D) Reynolds-averaged Navier–stokes (RANS) simulation and SST k-ω turbulence model is used to predict the shock structure and combustion phenomenon inside the scramjet combustor. All the simulations are done by using ANSYS 14-FLUENT code. Eddy dissipation along with finite rate chemistry model is used as combustion model for the present geometry because of the absence of Arrhenius calculations where turbulence controls the reaction rate. The effect of variation of angle of attack on the efficiency of two-strut scramjet combustor is also discussed in the research article. Present results show that the geometry with zero angle of attack i.e. α = 0°, have smallest ignition delay and it improves the performance of scramjet combustor as compared to geometry with α = −3° and α = 3°. The combustion phenomena as well as efficiency is also found to be highest in case of geometry with α = 0°. In order to validate the results, the numerical data for single strut injection is compared with experimental result which is taken from the literature.

Parametric Study of Fuel Distribution Effects on a Kerosene-Based Scramjet Combustor

Generally, the overall performance of scramjet combustor is greatly impacted by the fuel distribution scheme. The current paper mainly conducted a comprehensive parametric study of the impact of fuel distribution on the overall performance of a kerosene-based scramjet combustor. Herein, a 3D supersonic combustor with a recessed cavity and four injection orifices was taken into consideration. The combustor’s performance was analyzed by 3D RANS model. The fuel equivalence ratio for each injection port was taken as the design variables. And the combustion efficiency, the total pressure recovery coefficient, and the drag coefficient were chosen as the objective functions. Some novel data mining methods including DOE technique, Kriging approximation model, interaction analysis, and main effects analysis methods were employed to conduct the parametric study. The distributed fuel injection scheme was optimized by nondominated sorting genetic algorithm. The results show that three objective functions were remarkably affected by both of the total fuel equivalence ratio and the fuel distribution scheme. The objective functions cannot reach the optimal solution at the same time, and there must be a tradeoff among the objective functions.

Computational Fluid Dynamics Simulation of Tip-to-Tail for Hypersonic Test Vehicle

Computational fluid dynamics simulations are carried out for a complete hypersonic vehicle integrating both external (nonreacting) and internal (reacting) flow together to calculate the scramjet combustor performance and vehicle net thrust minus drag. Simulations are carried out for a flight Mach number of about six. Three-dimensional Navier–Stokes equations are solved along with the shear stress transport turbulence model and single-step chemical reaction based on fast chemistry. The Lagrangian particle tracking method for droplets is used for combustion of kerosene fuel. Flow is largely nonuniform at the inlet of the combustor. Mass flow of ingested air increases with increase in angle of attack. Because of more combustion of fuel, wall surface pressure is higher for compared with . Combustion efficiency and thrust achieved are found to increase with the increase in angle of attack. Considerable amount of thrust is obtained from a single expansion ram nozzle and achievement of positive thrust–drag for the whole vehicle is demonstrated. The computational analysis of the whole vehicle provides net forces and moments of the whole vehicle, which is very useful for the mission analysis of the vehicle.

Effect of ramp-cavity on hydrogen fueled scramjet combustor

Sustained combustion and optimization of combustor are the two challenges being faced by combustion scientists working in the area of supersonic combustion. Thorough mixing, lower stagnation pressure losses, positive thrust and sustained combustion are the key issues in the field of supersonic combustion. Special fluid mechanism is required to achieve good mixing. To induce such mechanisms in supersonic inflows, the fuel injectors should be critically shaped incurring less flow losses. Present investigations are focused on the effect of fuel injection scheme on a model scramjet combustor performance. Ramps at supersonic flow generate axial vortices that help in macro-mixing of fuel with air. Interaction of shocks generated by ramps with the fuel stream generates boro-clinic torque at the air & liquid fuel interface, enhancing micro-mixing. Recirculation zones present in cavities increase the residence time of the combustible mixture. Making use of the advantageous features of both, a ramp-cavity combustor is designed. The combustor has two sections. First, constant height section consists of a backward facing step followed by ramps and cavities on both the top and bottom walls. The ramps are located alternately on top and bottom walls. The complete combustor width is utilized for the cavities. The second section of the combustor is diverging area section. This is provided to avoid thermal choking. In the present work gaseous hydrogen is considered as fuel. This study was mainly focused on the mixing characteristics of four different fuel injection locations. It was found that injecting fuel upstream of the ramp was beneficial from fuel spread point of view.

Investigations on the influence of the in-stream pylon and strut on the performance of a scramjet combustor.

The influence of the in-stream pylon and strut on the performance of scramjet combustor was experimentally and numerically investigated. The experiments were conducted with a direct-connect supersonic model combustor equipped with multiple cavities. The entrance parameter of combustor corresponds to scramjet flight Mach number 4.0 with a total temperature of 947 K. The research results show that, compared with the scramjet combustor without pylon and strut, the wall pressure and the thrust of the scramjet increase due to the improvement of mixing and combustion effect due to the pylon and strut. The total pressure loss caused by the strut is considerable whereas pylon influence is slight.

Numerical Study on the Plasma Generated by Different Discharge Modes at the Upstream of Fuel Jet in Scramjet Combustor

How to enhance the combustion and reduce the total pressure loss in scramjet combustor are very critical for the practical application of hypersonic aircraft. Based on the dominant thermal mechanism of arc plasma, the plasma generated in combustor is regarded as a promising method to improve the combustion. As a result, the combustor model with transverse fuel jet and plasma generated by two discharge modes at the upstream of flameholding cavity is established and it is used to study the mechanism of fuel mixing enhancement through numerical investigation. The results show that an oblique shock wave would be formed at the upstream of the pseudo small plasma hump, and interact with the separation shock wave induced by the transverse jet. This results in the recirculation zone at the upstream of fuel jet being enlarged obviously. Besides that, under the non-reaction flow conditions, the total pressure recovery coefficient increases due to the plasma generated. However, the total pressure recovery coefficient varies apparently and the shear layer above the cavity is fluctuant when the plasma is generated by periodical discharge mode. While under the reaction flow conditions, the shear layer develops thicker and the total pressure recovery coefficient decreases. And due to the existing of plasma, the mole fraction of product water increases. But compared with the steady discharge mode, the level of water is lower and the total pressure recovery coefficient decreases more under the periodical discharge mode. Though the plasma generated by steady discharge mode shows better performance in assisting combustion and reducing the pressure loss, considering the energy saving and the use of different parameters of the periodical discharge, the same effects of enhancing the fuel mixing through enlarging the recirculation zone located at the upstream of fuel jet and promoting the mass exchange of cavity can be reached. More numerical experiments have to be done to optimize the parameters of periodical discharge plasma to receive a best improvement on the performance of scramjet combustor.

Numerical Study on the Plasma Generated by Different Discharge Modes at the Upstream of Fuel Jet in Scramjet Combustor

How to enhance the combustion and reduce the total pressure loss in scramjet combustor are very critical for the practical application of hypersonic aircraft. Based on the dominant thermal mechanism of arc plasma, the plasma generated in combustor is regarded as a promising method to improve the combustion. As a result, the combustor model with transverse fuel jet and plasma generated by two discharge modes at the upstream of flameholding cavity is established and it is used to study the mechanism of fuel mixing enhancement through numerical investigation. The results show that an oblique shock wave would be formed at the upstream of the pseudo small plasma hump, and interact with the separation shock wave induced by the transverse jet. This results in the recirculation zone at the upstream of fuel jet being enlarged obviously. Besides that, under the non-reaction flow conditions, the total pressure recovery coefficient increases due to the plasma generated. However, the total pressure recovery coefficient varies apparently and the shear layer above the cavity is fluctuant when the plasma is generated by periodical discharge mode. While under the reaction flow conditions, the shear layer develops thicker and the total pressure recovery coefficient decreases. And due to the existing of plasma, the mole fraction of product water increases. But compared with the steady discharge mode, the level of water is lower and the total pressure recovery coefficient decreases more under the periodical discharge mode. Though the plasma generated by steady discharge mode shows better performance in assisting combustion and reducing the pressure loss, considering the energy saving and the use of different parameters of the periodical discharge, the same effects of enhancing the fuel mixing through enlarging the recirculation zone located at the upstream of fuel jet and promoting the mass exchange of cavity can be reached. More numerical experiments have to be done to optimize the parameters of periodical discharge plasma to receive a best improvement on the performance of scramjet combustor.

Insight in the Performance of Scramjet Combustor Based on Orthogonal Experimental Design

A numerical insight was accomplished to optimize the scramjet combustor configuration based on orthogonal experimental design. Parametric modeling of combustor configurations was performed by the orthogonal array with 13 factors at 3 levels. Numerical simulations were proceeded by k-ε standard turbulence model and eddy-dissipation model in the combustion process. The performance indexes of combustion efficiency, total pressure recovery coefficient and thrust gain coefficient were evaluated. Detailed comparison with the effect of the factors on the performance was also carried out to demonstrate the main factors and determine the optimal configuration. The analysis of the extreme differences of the factors indicates that the main factors affecting combustion efficiency were the length of the wedge, the length depth ratio of the cavity, the depth of the cavity, and the length of the expanding section; The main factors affecting total pressure recovery coefficient are the angle of the primary combustor, the length of the expanding section, and the thickness of the strut; The main factors affecting thrust gain coefficient are the thickness of the strut, the length of the expanding section, and the angle of the secondary combustor. Validation of the optimal configuration is then confirmed that its performance is higher than the rest of the configurations, with the combustion efficiency of 0.915 and the total pressure recovery coefficient of 0.486, which are 31.5% and 65.9% higher than the experimental results, respectively.

Measurement of Water Vapor Levels for Investigating Vitiation Effects on Scramjet Performance

C OMBUSTION heated wind-tunnel facilities are required for many different aspects of ground-based testing of hypersonic airbreathing propulsion systems. This is particularly true for largescale scramjet testing in facilities with run times on the order of tens of seconds to minutes at equivalent flight Mach numbers above 4. Unfortunately, hydrogen and hydrocarbon combustion-heated facilities both result in water vapor levels in the test flow that are significantly greater than for atmospheric air. Hydrocarbon heated facilities additionally introduce carbon dioxide into the test flow. The effect of test gas vitiation due to combustion heating on scramjet combustor performance is an area of active research. Several recent studies have shown that the presence of elevated levels ofwater vapor may significantly affect combustor performance in scramjet engine test articles [1–4]. The experimental results reported in [1,2] were obtained in an electrical resistance heated wind tunnel where steam was artificially added to the flow to simulate a combustion-heated facility. Although the amount of steam addition can be calculated, an independent measurement is useful to confirm the presence of water vapor in the scramjet combustor at the reported levels. To that end, this note reports on results of the application of an instream optical measurement technique whereby water vapor levels in the test flow through the scramjet were independently measured and verified using a tunable diode laser absorption diagnostic. This paper first presents and describes the method of steam addition, and the associated makeup oxygen and optional carbon dioxide vitiation systems. Next, the tunable diode laser absorption technique that was applied to the facility is described. Finally, the added and measured water vapor levels are compared to validate the water vapor vitiation technique and to provide confidence in the similarmakeup oxygen and carbon dioxidevitiation techniques. This study represents the first time steam addition levels have been independently verified for the facility of [1,2] and it provides confidence in the claims made by the investigators in [1,2] with respect to the effects of test gas vitiation on scramjet performance and operation.

Effect of Connected Pipe Test Conditions on Scramjet Combustor Performance

Scramjet engine combustion chamber is normally ground tested in connected pipe mode and has limitations in simulating all combustor entry parameters as in flight engine. Effect of variation in combustor entry pressure, test gas medium due to vitiation and combustor entry flow profile on combustor performance has been brought out through CFD simulations. Computations are carried out using in-house CFD software “PARAS-3D” which solves the Reynolds Averaged Navier-Stokes equations and uses 7 species 7 reaction model for combustion. Entry pressure is having substantial influence on combustor performance at pressures lower than a certain value and this pressure limit reduces for higher fuel equivalence ratios. In the case of vitiated medium, presence of CO2 and the excess H2O result in lower total temperatures leading to reduced pressures in the combustion chamber. Non-uniform flow profile at combustor entry is also found to lower the scramjet combustor performance due to reduced mixing.

Liquid-Fueled Strut-Based Scramjet Combustor Design: A Computational Fluid Dynamics Approach

Computational-fluid-dynamics-based design and analysis is presented for a full-scale scramjet combustor with kerosene fuel injected from struts placed in the combustor flowpath. Three-dimensional Navier-Stokes equations are solved with a K-e turbulence model using commercial computational-fluid-dynamics software. Combustion is modeled based on infinitely fast chemical kinetics. Lagrangian dispersed-phase analysis is considered for fueldroplet evaporation and mixing in the supersonic stream. Parametric studies are carried out to investigate the effect of combustor-inlet Mach number and total pressure on the flow development process. A higher combustor-entry Mach number and distributed-fuel-injection system will ensure the existence of predominant supersonic flow in the combustor. Simulations are also carried out to investigate two different kinds of fuel injection struts in the scramjet combustor performance. A distributed-fuel -injection system, required to avoid thermal choking, increases the three-dimensionality of the flowfield.

Experimental Study on Effects of Fuel Injection on Scramjet Combustor Performance

In order to investigate the effects of fuel injection distribution on the scramjet combustor performance, there are conducted three sets of test on a hydrocarbon fueled direct-connect scramjet test facility. The results of Test A, whose fuel injection is carried out with injectors located on the top-wall and the bottom-wall, show that the fuel injection with an appropriate close-front and centralized distribution would be of much help to optimize combustor performances. The results of Test B, whose fuel injection is performed at the optimal injection locations found in Test A, with a given equivalence ratio and different injection proportions for each injector, show that this injection mode is of little benefit to improve combustor performances. The results of Test C with a circumferential fuel injection distribution displaies the possibility of ameliorating combustor performance. By analyzing the effects of injection location parameters on combustor performances on the base of the data of Test C, it is clear that the injector location has strong coupled influences on combustor performances. In addition, an inner-force synthesis specific impulse is used to reduce the errors caused by the disturbance of fuel supply and working state of air heater while assessing combustor performances.

Dynamics of Supersonic Droplets of Volatile Liquids

T HE breakup and vaporization of liquid droplets in supersonic flow is an interesting research problem with potentially important implications for supersonic combustion ramjets (scramjets). Noncryogenic, liquid hydrocarbons have substantial benefits as scramjet fuel [1], including higher energy density, lower cost, and ease of handling compared to liquid hydrogen fuel. The complications and time associated with the atomization and vaporization of liquid hydrocarbon scramjet fuel can, in principle, be avoided by injecting the fuel in the vapor phase. In some situations, however, such as a “cold start” in which the fuel is not preheated, hydrocarbon fuel may necessarily be injected while still in liquid phase. In this case, the rates and physical mechanisms associated with the disruption and vaporization of liquid droplets under supersonic flow conditions become critical issues to scramjet combustor performance. One possible technique to increase the dispersion of liquid fuels is to exploit the accelerated vaporization made possible by superheating the liquid. Investigation of the vaporization of superheated droplets and sprays have to date been largely confined to incompressible flows [2,3], with the physics of superheated liquid droplet disruption and vaporization in supersonic flow not yet well established. Studies of droplet disruption in compressible flows that have been conducted [4–6] have generally not considered the effects of superheating. The numerical study of Joseph et al. [7] did suggest a flash vaporization mechanism in the disruption of liquid drops in steady, supersonic flow, arguing that superheating may occur over small regions of the droplet surface due to a local combination of low pressure and frictional heating. The disruption of droplets in high-speed flow has often been studied by the sudden application of aerodynamic loads though the use of shock tubes [4,6,8–10]. Though this technique can produce liquid droplets under locally supersonic conditions, this is accomplished only after the passage of a shock wave through the droplets. Other droplet disruption studies have been conducted at subsonic speeds [11,12], for example, by droplets falling across a high-speed gas jet [13], droplet-bearing jets in cross flow [14], and in drop tubes [15]. In any case, these techniques do not typically result in the droplets achieving a significant degree of superheating. The research presented here investigated the dynamics of droplets consisting of volatile fluid smoothly accelerated to supersonic Mach numbers without passage of shock waves through the droplet. This was accomplished over a range of liquid vapor pressures using a compact, underexpanded supersonic jet configuration. This is an extension of a previous study of superheated droplet disruption with compressible, but subsonic, flow relative to the droplets [16]. One challenging aspect in the study of liquid drops in high-speed airstreams is the measurement of the drop velocity and acceleration [17]. Double-pulsed, planar laser imaging was employed here to determine the velocity of the droplets for various test fluids.

Effects of injector geometry on scramjet combustor performance

An experiment was conducted to investigate the effect of injector/combustor geometry on combustion-induced peak wall pressure and associated upstream influence, as well as on mixing/combustion characteristics at an entrance Mach number of 2.5. The length of the constant area section downstream of injection orifices had a strong influence on the above-mentioned characteristics. However, the sweep of the rearward-facing steps on both side walls had little effect on these characteristics, nor did reversing them have any effect. The peak wall pressure and the length of the upstream influence agreed qualitatively with predictions of an analytical model and an empirical formula developed at Johns Hopkins University. Fuel jets injected from the model with the longest constant area section and the fuel equivalence ratio of unity, coalesced at a very early stage downstream of the fuel injection orifices. This coalescence led to a decrease in mixing rate downstream, despite the higher degree of mixing near the injection orifices. The combustion efficiencies were higher than those obtained at NASA Langley in the upstream region due to the higher mixing rate near the injection orifices.

Engineering model for analysis of scramjet combustor performance with finite-rate chemistry

An engineering model of diffusion and reaction-limited supersonic combustion is described. The key feature of this model is that it divides the flowfield into three parallel, one-dimensional streams: 1) the fuel stream, 2) the oxidizer stream, and 3) the product stream. Fuel and oxidizer are continuously fed into the product stream in accordance with an empirical mixing model and allowed to react at a finite rate. Comparisons with experimental data for hydrogen-air combustion are presented. The model is clearly oversimplified compared to the complex two- and three-dimensional turbulent mixing and induced shock train processes known to exist in supersonic combustors. Nevertheless, it includes sufficient physics to provide reasonably good agreement with the data and enables rapid interpretation of the global features of the experiments. Accordingly, it may be useful in conceptual design studies.