Fuel Injection Scheme Research Articles

Influence of fluidic shock control on the combustion flow field of a dual-mode scramjet

To improve the resistance capability of a dual-mode scramjet isolator, fluidic shock control devices with different control parameters were designed. First, the effects of the control parameters on the position distribution, morphological structure, evolution process of the shock train in the isolator and the resistance capability of the isolator were numerically simulated. Then, cold- and hot-state tests of the scramjet through a direct-connected pulse combustion wind tunnel were carried out at a simulated flight Mach number of 4 to obtain the ignition and combustion process in the scramjet combustor under fluidic shock control through measurements by high-speed schlieren and pressure transducers. Finally, the influence of fluidic shock control on supersonic combustion was analysed by comparing the flow field structure and combustion process with and without a control device at different equivalent ratios and under different fuel injection schemes. The results showed that the fluidic shock control parameters with a great influence on the resistance capability of the isolator were the angle and width of the suction slot and the angle of the injection slot and that the resistance capability of the isolator could be effectively improved by reasonably designing the control parameters. Under fluidic shock control, the morphological structure and positional distribution of the shock train showed an obvious hysteresis evolution process near the injection slot and the suction slot in the isolator with increasing and decreasing back pressures, and this hysteresis effect could further improve the combustion performance of the combustor.

Investigation of hydrogen/air co-flow jet flame propagation mechanism in supersonic crossflow

The design of fuel injection schemes is crucial for improving the combustion performance of high-Mach number scramjet. To clarify the feasibility of the coaxial jet injection scheme, high fidelity Large Eddy Simulation of the supersonic coaxial jet flame is conducted. The simulations are in good agreement with the experimental results in terms of time-averaged velocity, temperature, and species distribution. Auto-ignition phenomenon and the characteristics of partially premixed flame are well captured. The introduction of co-flow air increases the vorticity magnitude close to the injection port and downstream near-wall region, which results in a 400 K rise in the time-averaged temperature on the downstream near-wall region and a 4% increase in the proportion of premixed combustion near the injection port. Moreover, the instantaneous distribution of hydroxy radical indicates that the spanwise width of the windward reaction shear layer is reduced utilizing the coaxial jet scheme. Chemical kinetic analysis is applied to reveal the propagation mechanism of partially premixed flames. Thermal explosion is the chemical explosion mode for all coaxial jet flame front, which are dominated by a high-temperature reaction path. The low-temperature reaction path mainly exists in the transverse jet injection port, downstream near-wall region of the single transverse jet and co-flow lifted flame base. These significant findings provide valuable insights for the potential engineering application of the coaxial jet injection scheme to a high Mach number scramjet.

Experimental investigation on the impact of fuel injection schemes on the fuel/air mixing characteristics in alternating wedge struts

To enhance the performance of scramjet engines, it is essential to understand the fuel/air mixing characteristics downstream of alternating wedge struts and the dynamics of streamwise vortices. Two optical measurement techniques, Mie scattering and acetone-particle laser-induced fluorescence, were employed to study the impact of fuel injection location and different injection gases on the fuel/air mixing characteristics downstream of alternating wedge struts at total temperature of 300 K and total pressure of 800 kPa. The results indicated that the particle laser-induced fluorescence measurements provided superior results compared to Mie scattering, due to the more uniform distribution and better following characteristics of acetone molecules as seeding particles. Different fuel injection locations resulted in varying fuel distributions and plume morphologies downstream. The plume sizes obtained from the three types of fuel orifices were generally consistent, but the mixing effect of the orifices on the first strut was better, whose injection locations being at the trailing edge intersections, with a larger plume area and more uniform fuel distribution. The fuel/air mixing process downstream of the alternating wedge struts was primarily dominated by streamwise vortices. Despite certain differences in the physical properties of nitrogen and hydrogen molecules, the measurement results showed that the plume morphologies obtained with both gases were quite similar. Additionally, a streamwise vortex model was developed to predict the dynamics of streamwise vortices and the evolution of fuel plume morphology, and it was compared with experimental results.

Sustainability of flameless combustion mode with hydrogen addition to liquid kerosene fuel at high-intensity conditions

The application of hydrogen for fuel enrichment and blending has recently gained importance. The sustainability and performance enhancement through hydrogen addition in flameless combustion mode for liquid fuels needs further investigation to enhance our understanding. Therefore, the present study reports the effect of hydrogen addition on the flameless combustion of a kerosene-fired 30 kW swirl combustor with hydrogen addition. A divergence-shaped conical combustor with tangential air injection and asymmetric fuel injection scheme is adopted. The liquid fuel was injected using a solid-cone swirl nozzle with 34 μm Sauter Mean Diameter. The hydrogen addition was varied from 0 to 20% on a thermal input basis, and the thermal input of liquid fuel was maintained at a 30 kW level. Computational studies helped understand the flow physics behavior within the combustor volume. The hydrogen addition was observed to improve the reactant dilution ratio (Rdil). The maximum values of the Rdil increased from 4.47 to 5.44 with hydrogen addition. Hydrogen addition further helped extend the flameless combustion regime to φ = 0.3 without any air preheating. A uniformly distributed reaction with a reduced-temperature gradient is observed with extended combustion stability limits. A marginal increase in NOx emissions was observed with hydrogen addition, albeit within permissible limits, accompanied by ∼8–20 dB reduction in acoustic emissions.

Combustion modes of kerosene spray in a supersonic combustor

The mixing and combustion characteristics of liquid kerosene spray in a cavity-based supersonic combustor under a Mach 2.52 inflow are experimentally and numerically investigated. The combustion mode transition of kerosene spray includes two aspects: 1. As the injection pressure increases, the combustion mode shows a transition of weak-intensive-transitional(weak to intensive)-weak-blowoff; 2. in the transitional combustion mode, there is a transition from weak to intensive combustion over time. In the weak combustion mode, the combustion chamber pressure increases slightly, and the flame is inside the cavity. In the intensive combustion mode, the pressure rises sharply from the isolation section and the flame is a typical cavity stabilized combustion mode. The weak combustion at low/high injection pressure is due to rich/lean fuel inside the cavity. In the transitional combustion mode, the flame on the downstream wall of the cavity gradually expands and joins with the flame inside the cavity, evolving from weak combustion to intensive combustion. The transition is accomplished in a very short time, on the order of milliseconds. This study deepens the understanding of the combustion mode of scramjet engines and provides a reference for the design of fuel injection schemes for scramjet engines.

A scheme of combustion organization with combination of cone-strut and plasma torch/ oxygen addition in round supersonic combustor

Round combustors as a very desirable kind of supersonic combustor, face the difficulty of combustion organization. The typical cone-strut structure can accomplish some combustion organization, but it has a very high congestion ratio and is hence inefficient. To solve this problem, a method of combining cone-strut configuration and plasma/oxygen addition is proposed in this paper, and its scheme, experiment and combustion mechanism are studied. In the scheme of combustion organization, the fuel injection scheme and plasma arrangement scheme of the approach are projected, accomplished with the advantages of reducing congestion ratio from 20%-26%–13%. In the experiments of combustion organization, the effect of realizing combustion organization of such approach is proved, and three typical combustion organization states, which may be due to low congestion ratio, are displayed one by one, and its transformation speeds between three typical combustion organization states are investigated. In the speculations of mechanism, a set of simplified equations are provided for the first time, and these equations indicate that why the approach is effective and why the three kinds of combustion organization states display one by one. This research has theoretical and technical implications for the development of cone-strut configuration in round combustors.

Experimental study on the effect of injection schemes on fuel spray and combustion characteristics in a compact combustor

Fuel injection schemes based on close-coupled injection are usually used in compact combustor design. In this paper, an experimental method is adopted to study the fuel spray Sauter mean diameter (SMD) and combustion characteristics when different fuel injection schemes, strut cavity structures, and fuel/air jet momentum flux ratios are applied. Four fuel injection schemes based on close-coupled injection schemes and premixed injection schemes are proposed for use in a compact integrated combustor to compare the effects of the fuel injection schemes. Three struts are designed to study the influence of the cavity on the close-coupled fuel injection scheme. The results show that the spray and combustion characteristics of the close-coupled injection scheme differ from those of premixed injection schemes. Specifically, the SMD in the close-coupled scheme is in the range of 70-150 μm, which is larger and more uneven than that of the premixed scheme (15-40 μm), and the close-coupled scheme has better ignition performance at a temperature of 450 K. The flame structure of the close-coupled injection schemes with reverse mainstream inside the cavity has better radial uniformity, but more fuel is attached to the strut surface, resulting in a lower flame intensity and increased diffusion flame. The cavity on the strut sidewall can effectively promote fuel atomization by approximately 25% when the jet momentum flux ratio is low, make the fuel more uniform in the strut radial direction, increase the flame intensity, and reduce diffusion flame zone. In the close-coupled injection scheme, the large jet momentum ratio will lead to the collision and fusion of multiple fuel streams downstream of the struts, thus increasing the fuel droplet size and limiting the promoting effect of the strut wall structure. In summary, this paper provides further understanding of the combustor spray and combustion characteristics of different fuel injection schemes.

Combustion Oscillations in Scramjet Combustor with Different Fuel Injection Schemes

This work experimentally studied combustion oscillations in a cavity-based combustor operating under high-enthalpy supersonic flow. Three fuel-injection distances and the corresponding combined fuel injection schemes were tested and compared with the global equivalence ratio unchanged. The results indicate that the combustions of longer and medium distances (35 and 21 mm, respectively) present stronger oscillations, which are mainly attributed to the strengthened oscillations of the mainstream flame (MF). It is also found that combined fueling schemes can help weaken the oscillations. The predominant frequencies of the combustion oscillations in different fueling schemes are all distributed at approximately 400 Hz, which are proven to be induced by the MF. The low-frequency range (0–50 Hz) is more dominant for the shorter injection distance (7 mm) than in the other cases, which is because the flames are strongly influenced by low-frequency disturbances with fluctuations in the fueling injection and inflow. The analysis reveals that these disturbances act primarily on the cavity flame.

Research on combustion performance improvement by strut/wall combined injection in scramjet combustor

Strut/wall combined fuel injection scheme was adopted to improve mixing and combustion efficiency in a scramjet combustor fueled with liquid kerosene in the condition of Mach 6. Injectors were placed on the front of the strut and the side wall of the combustor. A series of numerical simulations and experiments were carried out to improve the combustor performance under conditions of different incoming flow velocity, injection methods, and fuel distribution ratios. The value of pressure was obtained by pressure sensor and the flame images were captured by the high-speed camera in experiment. By processing and analyzing the basic data, characteristics of fuel mixing and combustion performance were discussed in this paper. Then, the influence mechanism of the strut/wall combined injection on the performance of the combustor was explained based on the performance with influence factors. Results indicated that the mixing and combustion efficiency was related to condition, injection method, and nozzle arrangement. The strut/wall combined injection dispersed the heat release, which could reduce the pressure rise and total temperature. The fuel distribution ratio between the strut injection and wall injection is also a key factor affecting the performance of the combustor. These results in this paper are valuable for the combustion organization in the supersonic combustor and the improvement of the combustor performance.

Powder fuel transport process and mixing characteristics in cavity-based supersonic combustor with different injection schemes

The transport process and mixing characteristics of powder fuel in the supersonic airflow are essential for the working performance of a powder fueled scramjet. In the current investigation, the effect of different injection schemes on the powder fuel jet flow field and mixing characteristics are numerically studied in the cavity-based supersonic combustion chamber. The simulation is employed by the Eulerian-Lagrangian scheme coupling with the gas-solid two-phase Reynolds-average Navier-Stokes (RANS) method. Verification of numerical method and grid-independent is conducted firstly, which demonstrated the correctness and credibility of the method in this research. The results show that the powder fuel injection into the combustor induces typical gas-solid two-phase jet flow characteristics. With respect to the fuel injection schemes upstream of the cavity, the powder fuel moves across the cavity section and is continuously mixed with the mainstream along the flow direction. The mixing intensity in the combustion chamber is associated with the motion and distribution patterns of the particles. For the cavity internal fuel injection scheme, the particles are mainly distributed inside the cavity. Some of the particles can be coiled and sucked into the supersonic mainstream. The cavity shear layer and expansion section near the bottom wall is the main region with better mixing performance.

Effect of inclined fuel injection in the strut wake region for a hydrogen fueled scramjet combustor

The supersonic airflow and injected fuel stream mixing process is complicated for a scramjet combustor because of little supersonic air resident time and high airflow stream momentum. The direction of fuel injection into a supersonic airflow stream plays a significant role in mixing enhancement with turbulence generation. This study investigates the effect of strut-based parallel (primary) and inclined fuel injection in the strut flame holder wake region. The primary strut with parallel injection is derived from the open literature. The base model strut consists of parallel injection, whereas in the new models, the fuel is injected at an angle of θ for both the upward and downward directions. By computing the Reynolds averaged Navier-Stokes equations, finite-rate eddy dissipation turbulence chemistry, and SST k-ω turbulence models, the mixing and combustion behavior for various fuel injections were investigated. The impact of different fuel injections on the downstream wake region is assessed by evaluating the distribution of H2 and H2O species in the radial and axial directions. The turbulence intensity and turbulent kinetic energy are also considered to evaluate the cross-impact of the air stream and fuel stream for inclined fuel injection. Numerical results found that inclined injection significantly impacts the mixing phenomenon with the development of enriched re-circulation zones in the vicinity of the flame keeper. The mixing and combustion efficiencies performance parameters are evaluated for different fuel injection schemes and found the average increased mixing efficiency of 13% and combustion efficiency of 12%, respectively, for inclined fuel injection. From the analysis of mixing efficiency, it is observed that complete mixing has been found after 235 mm, whereas for inclined injection, the complete mixing happened within a 175 mm distance. The maximum of 75% and 62% combustion efficiency has been observed at a 225 mm and 285 mm distance for inclined and parallel fuel injection, respectively.

Combustion Instability due to Combustion Mode Transition in a Cavity-Stabilized Scramjet

This study experimentally investigates flame dynamics during a combustion-driven mode transition in a cavity-stabilized scramjet model with an inlet Mach number of 2.2 using a split fuel injection scheme. Fueling into the cavity was gradually increased while keeping a constant global equivalence ratio of 0.3. A cavity-fueling percentage below 15% resulted in low-intensity combustion that corresponded to operation in scram mode. Redirecting 29% of the fuel from the upstream injection port to the cavity yielded flame spreading into the main flowpath, inducing intense combustion. A sharp pressure rise and an intense combustion region in the diverging section suggest dual-mode combustion, with thermal choking and a pseudoshock formation downstream of the cavity. An intermediate cavity-fueling fraction of 21% led to a low-frequency instability () with intermittent transitions between the combustion modes, indicating operation at the thermal choking limit. Spectral and modal analyses of chemiluminescence images and static-pressure measurement indicate peak response at 10–20 Hz and , originating from oscillations in the air heater and a fuel jet instability, respectively. Thermally chocked conditions significantly amplify oscillations at these frequencies. At the thermal choking limit, there was high-amplitude response at , preventing stable dual-mode operation.

Investigation on low total temperature combustion characteristics of kerosene-fueled supersonic combustor

In this paper, the combustion characteristics of kerosene-fueled supersonic combustor under the conditions of Mach number 2.0, the total temperature at 700 K and the total pressure at 520 kPa (simulated flight Mach number at 3.5) were studied by using the flame stabilizing method of cavity and strut from three aspects such as blockage ratios, kerosene equivalence ratios, location and quantity of injection holes. The results showed that: (A) The combustor with the strut realized the independent and stable combustion of kerosene. The combustion-induced back pressure in the block test with blockage ratio of 20% and 10% destroyed the inlet flow conditions; while the blockage ratio was 7.3% and 5%, the incoming flow conditions when kerosene was burned stably were not destroyed. (B) The kerosene Equivalence Ratio (ER) was more likely to be disturbed upstream than the induced back pressure when it rose, and an excessively high-ER would reduce the combustion efficiency; when the equivalence ratio was constant, the combustion efficiency of the blockage ratio of 7.3% was higher than the blockage ratio of 5%. The combustion efficiencies were 0.86 (ER = 0.13) and 0.78 (ER = 0.19) when the blockage ratio was 5%, respectively; the combustion efficiencies were 0.89 (ER = 0.16) and 0.82 (ER = 0.19) when the blockage ratio was 7.3%, respectively; the combustion efficiencies were 0.51 (ER = 0.25), 0.81 (ER = 0.3), 0.65 (ER = 0.34) and 0.62 (ER = 0.42) when the blockage ratio was 20%, respectively. (C) The porous injection provided behind the strut was beneficial to the atomization of kerosene and improved the combustion efficiency of kerosene; the second injection after the cavity would reduce the combustion efficiency due to insufficient oxygen and combustion space. This study expanded the working range of ramjet and provided a reference fuel injection scheme for Turbine-Based Combined Cycle (TBCC) engine.

Flame stabilization and local combustion modes in a cavity-based scramjet using different fuel injection schemes

This paper investigates the impact of fuel distribution on flame stabilization and local combustion modes in a cavity-based scramjet. Two injection schemes with a global equivalence ratio of 0.4 are comparatively studied. Ethylene is injected only from upstream of the cavity in the first case, whereas in the second case 75% of the fuel is injected from upstream of the cavity and 25% of the fuel is injected from the cavity floor. The reactive flow in the scramjet is numerically investigated using a hybrid RANS/LES method. The agreement between calculation results and experimentally measured pressure profiles and CH* chemiluminescence images is good. The simulation results are further employed to gain a fundamental understanding of local combustion modes from different aspects: supersonic vs. subsonic, and premixed (fuel-rich vs. fuel-lean) vs. diffusion by using filter functions. In the upstream-only fueling case, it is found that the flame is stabilized by the cavity shear layer and most of the heat is released by supersonic-premixed combustion. Moreover, subsonic-premixed combustion and supersonic-diffusion combustion are significant, and the leading mode changes along the combustor. In the dual fueling case, a cavity assisted jet-wake stabilized flame is observed with higher combustion efficiency. Diffusion combustion becomes the dominant mode, with similar heat release in subsonic and supersonic regions. Nevertheless, the flashback of a premixed flame along the walls plays an important role, leading to diffusion combustion around the fuel jet. The results elucidate the link between the fuel injection schemes, flame stabilization locations, and local combustion modes.

The heat flux research in hydrogen supersonic combustor at Mach number of 4

The work is devoted to the study of the intensity of heat transfer in a supersonic combustion chamber at a Mach number of 4 under conditions of ignition and transition to intense combustion, including the transition to choking the channel. The experiments were carried out on a combustion chamber model in the connected pipeline mode with flow parameters in the channel close to flight conditions at Mach numbers 6–8. The experimental model is a rectangular channel with a flame holder in the form of backward facing step (BFS). Fuel injection was carried out in front of BFS on the top and bottom walls of the model through 8 circular holes, which were situated under the angles of 45° or 90°. It has been revealed that the choice of the fuel injection scheme leads to an increase in the level and a change in the distribution of the heat flux along the length of the combustion chamber. A decrease in the angle of hydrogen injection makes it possible to significantly reduce the heat flux into the wall of the combustion chamber, while choking the channel is accompanied by a twofold increase in the heat flux.

Research on combustion performance optimization in scramjet combustor with strut/wall combined fuel injection scheme

A strut/wall combined fuel injection scheme was experimentally investigated to improve the combustion performance in a flush-wall scramjet combustor fueled with liquid kerosene. A thin strut was adopted as fuel injector and flame holder, and wall injectors were also placed on side wall of the combustor, downstream of the strut. With the injection of pure oxygen through the injectors in the trailing edge of the strut, a stable flame was achieved in the combustor, and the flame interaction characteristics were experimentally investigated. Experiments were conducted at the inlet Mach number of 2.0 with different strut/wall fuel distribution schemes. Flame was divided into two parts, namely the core flame and the wall flame, and interaction mechanism between the two parts was experimentally investigated. Experiment results show that core flame is enhanced with the increasing of fuel equivalence ratio. Core flame propagates to side wall of the combustor to ignite wall fuel, forming a wall flame. Wall flame leads to a thermal chocking downstream of strut to enhance core flame. Combustor performance and flame characteristics are influenced by fuel distribution schemes. The location of wall fuel injectors is also optimized to improve combustion efficiency in this paper. These results in this paper are valuable for the future optimization of combustion organization in scramjet combustor with multi-stages fuel injection scheme.

Analysis of the Impact of Split Injection on Fuel Consumption and NOx Emissions of Marine Medium-Speed Diesel Engine

The medium-speed diesel engine in diesel-electric propulsion systems is increasingly used as the propulsion engine for liquefied natural gas (LNG) ships and passenger ships. The main advantage of such systems is high reliability, better maneuverability, greater ability to optimize and significant decreasing of the engine room volume. Marine propulsion systems are required to be as energy efficient as possible and to meet environmental protection standards. This paper analyzes the impact of split injection on fuel consumption and NOx emissions of marine medium-speed diesel engines. For the needs of the research, a zero-dimensional, two-zone numerical model of a diesel engine was developed. Model based on the extended Zeldovich mechanism was applied to predict NOx emissions. The validation of the numerical model was performed by comparing operating parameters of the basic engine with data from engine manufacturers and data from sea trials of a ship with diesel-electric propulsion. The applicability of the numerical model was confirmed by comparing the obtained values for pressure, temperature and fuel consumption. The operation of the engine that drives synchronous generator was simulated under stationary conditions for three operating points and nine injection schemes. The values obtained for fuel consumption and NOx emissions for different fuel injection schemes indicate the possibility of a significant reduction in NOx emissions but with a reduction in efficiency. The results showed that split injection with a smaller amount of pilot fuel injected and a smaller angle between the two injection allow a moderate reduction in NOx emissions without a significant reduction in efficiency. The application of split injection schemes that allow significant reductions in NOx emissions lead to a reduction in engine efficiency.

Ignition and combustion characteristics of supersonic combustor under flight Mach number 2.6–3.8

This paper simulated the flying conditions of flight Mach number 2.6 ~ 3.8. Under the inflow state of the total temperature of 507 K ~ 820 K, the ignition test under low total temperature was carried out on the supersonic combustor by using the hydrogen-guided flame ignition method. When the total temperature was 820 K (simulated flight Ma = 3.8), the kerosene sprayed by the non-strut spray block could achieve a single stable combustion after removing the Pioneer hydrogen, and the fuel rich limit of the combustor was 0.4. When the total temperature was 700 K (simulated flight Ma = 3.4), when the non-strut spray block was used, the kerosene after removing Pioneer hydrogen could not be burned stably alone. In the test with blocking ratios of 20% and 10%, kerosene could be burned steadily alone, but the combustion-induced back pressure propagated upstream to the entrance of the isolator, destroying the inlet flow conditions. When the spray block with blocking ratio of 7.3% and 5% was used for the test, kerosene could be burned stably alone, and the incoming flow conditions were not destroyed. At a total temperature of 600 K (simulated flight Ma = 3.0) and 507 K (simulated flight Ma = 2.6), hydrogen and kerosene with a blocking ratio of 7.3% were used to achieve stable combustion, but the combustion-induced back pressure severely damaged the inlet flow conditions of the combustor. After the Pioneer hydrogen was removed, kerosene at both total temperature conditions could not be burned steadily. This study extended the lower limit of the working Mach number of the ramjet engine, and provided a reference flame stabilization method and fuel injection scheme for the kerosene combustion of the supersonic combustor, and provided some design suggestions for the low-Mach-Number combustor.

Predictive Numerical Analysis on the Fuel Homogeneity in a Rotating Detonation Engine (RDE) Implementing Radially-Entered Fuel Injection Scheme

The potential for a continuous detonation wave in rotating detonation engine (RDE) hugely rely on the homogenous fuel-air mixture within annular chamber. Using that motivation, a comprehensive numerical study on the non-reacting flow field was carried out to investigate the mixing of hydrogen (H2)and oxygen (O2) in a RDE, allowing this study to represent the flow behavior at the split second before ignition. A validation with the published experimental data was successfully done by comparing the predicted detonation front displacement, r and velocity, vD . Below than 10 % error was detected from both variables. New parameter representing fuel homogeneity was introduced which is the magnitude of maximum deviation from the H2 mean mass fraction, |S max|. A comparison in the |S max| generated from RDE implementing axially-entered fuel (AEF) injection with RDE implementing radially-entered fuel (REF) injection was carried out. Based on the numerical results, the |5max| value from AEF injection is lower than REF injection implying the capability of AEF injection to produce better fuel homogeneity. Via AEF injection, the injected fuel stream has a direct impingement with the injected oxidizer stream allowing proper mixing of the reactants. In REF injection, thedistancefrom fuel injectortooxidizerinjector, R was varied to foresee the impact of R in |S max | value. The outcome shows that as the value of Rincreases, the value of |S max| increases as well. Thisisduetotheformation of vortex between the reactant streams plus the increased in distance between both streams which further reduced the reactants interaction. In the cases of R3 and R4 where the values of R are 7 and 9 mm respectively, the homogeneity of fuel is severed in the upstream region. This is a major concern since the ignition happened in that particular area. Thus, it is crucial for the reactants to mix properly within the upstream region.

Experimental validation of liquid hydrocarbon based fuel rich gas generator for high speed propulsion systems

For long duration hypersonic flight vehicles, research on high speed air-breathing propulsion systems has been progressing across the world for the past few decades. Ramjets/Scramjets fueled with hydrogen/liquid hydrocarbon fuels are the propulsion systems considered for different regimes of flight envelope. Dual Combustion Ramjet (DCR) engine is a prospective candidate for one of such kind of applications, where the engine is expected to work for a wider range of flight Mach number regime. It has a fuel rich subsonic dump combustor and a supersonic combustor, which work in tandem. In the present work, extensive experimental and analytical work has been carried out to develop a fuel rich gas generator often called as dump combustor suitable for operation from 20 to 28 km altitude and in the range of 4–6 flight Mach number. Connect pipe mode testing has been carried out by varying the throat area of dump combustor for Mach 4 conditions. Tests have been carried out with Jet-A fuel at different equivalence ratio and with fuel injection in different combination and stages. Fuel injection scheme used here is of shower head type in intake arm and a single simplex swirler with film cooling holes in the dump combustor. Study has been also carried out by varying the location of swirler along the combustor length. Combustor air entry conditions are simulated using H2–O2 based air heater and a test bed nozzle. An exhaustive amount of tests have been carried out to establish condition suitable for consistent ignition & sustained combustion in this highly fuel rich condition. A One dimensional (1D) mathematical model with NASA CEA interface has been developed to predict & compare the test performance. Sustained ignition is achieved for fuel flow rate of 0.8 kg/s and at an equivalence ratio of 0.8–0.9.