Thermal Throat Research Articles

Investigation of Unsteady Combustion Regimes in a Kerosene-Fueled Scramjet with Air Throttling

This paper describes an experimental study investigating unsteady combustion regimes in a kerosene-fueled scramjet. The results are obtained under inflow conditions of a 2.9 MPa stagnation pressure, 1900 K stagnation temperature, and a Mach number of 3.0. The air throttling position is 240 mm downstream of the combustor entrance, with an air throttling flow rate (ratio of air throttling mass flux to inflow mass flux) of 38% and a fuel equivalence ratio of 0.37. Combustion is relatively stable when air throttling is applied and is dominated by auto-ignition. When air throttling is turned off, the combustion becomes more unsteady and is dominated by flame propagation. At the same time, the combustion mode changes, and the frequency of the combustion mode transition is 286 Hz. Schlieren images and one-dimension analysis show that the effect of air throttling is the coupling of cold throat (aerodynamic throat) and hot throat (thermal throat). The proper orthogonal decomposition and dynamic mode decomposition analysis present that when air throttling is applied or removed, the frequencies of injector–flame feedback are almost the same, while the frequencies of shock–flame feedback exhibit considerable variation, which is caused by the location of the precombustion shock affected by air throttling.

Effect of kerosene injection states on mixing and combustion characteristics in a cavity-based supersonic combustor

It has been found that the static pressure distribution along the axial direction of liquid kerosene is lower than that of the gaseous kerosene under the same flow condition and overall equivalent ratio from previous studies. To further investigate this phenomenon, a compressible two-phase parallel simulation method is utilized to analyze the mixing and combustion characteristics of gaseous and liquid kerosene jets in a cavity-based supersonic combustor. The numerical results are consistent with the experiments and demonstrate that gaseous injection leads to a cavity shear layer that dives deeper into the cavity, forming two recirculation zones in the front and rear of the cavity. In contrast, the cavity shear layer is closer to the mainstream during liquid injection, and only a large recirculation zone is formed in the rear of the cavity. As a result of the cavity shear layer and the recirculating flow, the fuel vapor of gaseous injection accumulates in the front of the cavity, while for the liquid injection, the fuel vapor disperses in the cavity, cavity shear layer, and the region above, and the rear of the cavity has a higher fuel vapor concentration than the front. This unique fuel distribution causes the combustion area to be concentrated in the cavity during the gaseous injection but dispersed inside and downstream of the cavity during the liquid injection. As a result, forming a thermal throat under the same conditions is more challenging during liquid injection, and the generated static pressure distribution is lower than that during the gaseous injection.

Investigation of the influence mechanism of geometric throats in variable-geometry rocket-based combined-cycle combustors

To explore the reaction mechanism a geometric throat and its influence on the performance of the combustor, this paper studies the working process of a rocket-based combined-cycle (RBCC) variable geometry combustor with geometric throats under Ma4 inflow conditions by both numerical simulation and ground direct-connect testing. The main research contents and conclusions are as follows: (1) The geometric throat can accurately control the choking position. The fuel in the variable geometry combustor can complete the concentrated heat release and bring higher combustor pressure. The thrust of the geometric throat combustor is 8.7% higher than that of the thermal throat combustor. Therefore, the geometric throat combustor provides a better combustor performance than the thermal throat combustor. (2) When the throat area of the Ma4 configuration is adopted, the thrust and specific impulse in the combustor are at their the highest, reaching 6.0% and 0.4% higher than those of the Ma5 and Ma3 throat configuration combustors, respectively. This shows that a certain amount of combustor heat release needs to match the corresponding geometric throat area to ensure the optimal performance of the combustor. (3) In the Ma4 combustor configuration, after closing the primary rocket, the thrust in the combustor decreases by approximately 10.8%, but its specific impulse performance greatly improved by 39.3%. Therefore, the geometric throat can be used in the RBCC mode transition process in combination with the secondary fuel injection strategy to avoid the thrust trap in the mode transition process.

Supersonic combustion diagnostics with dual comb spectroscopy

Supersonic engine development requires accurate and detailed measurements of fluidic and thermodynamic parameters to optimize engine designs and benchmark computational fluid dynamic (CFD) simulations. Here, we demonstrate that dual frequency comb spectroscopy (DCS) with mode-locked frequency combs can provide simultaneous absolute measurements of several flow parameters with low uncertainty across a range of conditions owing to the broadband and ultrastable optical frequency output of the lasers. We perform DCS measurements across a 6800–7200 cm−1 bandwidth covering hundreds of H2O absorption features resolved with a spectral point spacing of 0.0067 cm−1 and point spacing precision of 1.68 × 10−10 cm−1. We demonstrate 2D profiles of velocity, temperature, pressure, water mole fraction, and air mass flux in a ground-test dual-mode ramjet at Wright-Patterson Air Force Base. The narrow angles of the measurement beams offer sufficient spatial resolution to resolve properties across an oblique shock train in the isolator and the thermal throat of the combustor. We determine that the total measurement uncertainties for the various parameters range from 1% for temperature to 9% for water vapor mole fraction, with the absorption database/model that is used to interpret the data typically contributing the most uncertainty (leaving the door open for even lower uncertainty in the future). CFD at the various measurement locations show good agreement, largely falling within the DCS measurement uncertainty for most profiles and parameters.

Cavity flameholding subjected to an expansion fan in a rocket-based combined cycle combustor

The effect of expansion fan on cavity flameholding was investigated in a rocket-based combined cycle combustor fueled by ethylene. The inflow of the combustor was at a condition where the Mach number, stagnation temperature, and total pressure were 2.92, 1650 K, and 2.6 MPa, respectively. The global equivalence ratio was 0.25. A backward-facing step (BFS) was used to simulate a rocket that was shut down and create an expansion fan. The dimensionless distance between the BFS and the cavity (d/H) varied from 1.5 to 7.5, where d = 60–300 mm was the actual distance and H = 40 mm was the height of the combustor entrance. The result indicated that the cavity shear layer reattached at the cavity floor in the combustor with 2.5 ≤d/H≤ 3.5 because of the pressure gradient resulting from the expansion fan. The expansion fan also significantly decreased the static pressure in the cavity. As a consequence, the combustor with 2.5 ≤d/H≤ 3.5 failed in ignition. In the combustors with d/H = 4.0 and 6.0, the cavity flows were open-type. However, the ignition attempt failed because the expansion fan and the reattachment shock wave might worsen the local equivalence ratio in the cavity. Successful ignition and flame stabilization were achieved in the combustors with d/H = 1.5, 2.0, 4.5, and 7.5. Three combustion modes and three kinds of combustion oscillation were identified. In the combustor with d/H = 1.5, only a small part of the cavity was exposed to the expansion fan. The combustor operated in the cavity stabilized scramjet mode because the flame front was anchored at the cavity leading edge. The cavity shear layer impingement onto the cavity ramp lead to the combustion oscillation with a frequency ranging from 300 to 400 Hz. The combustion mode and combustion oscillation in the combustor with d/H = 2.0 were the same as their counterparts in the combustor with d/H = 1.5. Unsteady jet-wake stabilized scramjet mode was witnessed in the combustor with d/H = 4.5. In this combustor, the fuel jet and the front of the cavity were exposed to the expansion fan. The reattachment shock wave lifted the cavity shear layer. The combustion oscillation induced by shock-flame interaction showed a dominant frequency at 107.1 Hz. The expansion fan and the reattachment shock wave were upstream of the cavity in the combustor with d/H = 7.5. Both of them made the boundary layer more sensitive to adverse pressure gradient. The flame front was upstream of the cavity because the heat release resulting from vigorous chemical reaction choked the combustor and induced large-scale boundary layer separation. The combustor operated in ramjet mode. The unsteady shock train and thermal throat induced the combustion oscillation with a frequency of approximately 800 Hz.

Investigation of Combustion Characteristics in a Kerosene-Fueled Supersonic Combustor with Air Throttling

The paper presents the results of an experimental and numerical study on the combustion characteristics of a supersonic combustor operation enhanced by pilot hydrogen and air throttling, including ignition and combustion. Wall pressure measurements, schlieren images, and hydroxyl-radical planar-laser-induced-fluorescence (OH-PLIF) are used to obtain information on the flowfield and flame development. With the aid of air throttling, the kerosene is ignited by the pilot flame; and the kerosene flame remains stable even after the pilot hydrogen is switched off. The speed of the shock train generated by the throttling air and combustion is about in the expansion section but only about in the cavity region. The stable flame and flame blowout cases can be separated by a curve in a plot of the coexistence time of pilot hydrogen and throttling air versus the air throttling flux ratio. There is a special case lying on this curve for which the flame is unstable and will be blown out before the throttling air is switched off. It is also found that the position of the thermal throat is the same for all the subsonic combustion cases: about 595 mm downstream of the scramjet entrance.

Experimental research on the thermal throat of rocket based combined cycle combustor

Based on a previous one-dimensional theoretical analysis, direct-connection experiments of thermal throat generation and the factors that influence it in the Rocket Based Combined Cycle combustor are performed relative to the fuel injection position and equivalence ratio. The Mach 5.5 airflow experimental results indicate that the performance is improved by 3.0–8.3% by optimizing the fuel injection position, and at an airflow of Mach 4.0, the increase in the fuel equivalence ratio gradually shifts the generation of the thermal throat backward. When a thermal throat is generated at the end of the combustor, the thrust and specific impulse performance are optimal. The effectiveness of employing a thermal adjustment to improve engine performance is validated by the experimental studies. The radiation spectrum method is employed to measure the temperature along the flow field of the thermal throat. The temperature and pressure measurements along the flow path indicate that the Mach number variation trend calculated in the experiment is the same as that observed in the numerical simulation results in the thermal throat flow field. The generation process of the thermal throat is characterized by the acceleration of airflow from being subsonic to supersonic.

The mode transition characteristics in a dual-mode combustor at different total temperatures

The mode transition experiments in a dual-mode combustor were carried out under the conditions of total pressure 1.05 MPa, Mach number 2.0 and total temperature 1285 K, 1085 K and 885 K. The processes of mode transition of the combustor under different total temperature conditions were studied. The connection between the fuel equivalence ratio, the ratio of the peak pressure of the combustor to the inlet total pressure of the isolator, and different combustion modes was analyzed. The study showed that the lower the total temperature of the incoming flow and the higher the fuel equivalence ratio, the higher the dimensionless wall pressure, the more easily the shock train to disturb the upstream of the isolator, the more backward position of the choked thermal throat, and the closer the combustion tended to the ramjet mode; The results showed that there were pure scramjet mode, dual-mode scramjet mode, dual-mode ramjet mode and the combustion state caused the pressure disturbance that spread to the entrance of the isolator with the increase of the fuel equivalence ratio. The dimensionless peak pressures when mode transition occurred were 0.25, 0.40 and 0.5, and did not change with the change of the total temperature.

Numerical Study on Thermal Choke Behaviors Driven by Various Rocket Operations in an RBCC Engine in Ramjet Mode

Abstract Thermal choke is commonly employed in a fixed geometry RBCC combustor to eliminate the need for physically variable exit geometry. This paper proposed detailed numerical studies based on a two-dimensional integration model to characterize thermal choke behaviors driven by various embedded rocket operations in an RBCC engine at Mach 4 in ramjet mode. The influences of different embedded rocket operations as well as the corresponding secondary fuel injection adjustment on thermal choke generation process, the related thermal throat feature, and the engine performance are analyzed. Operations of embedded rocket bring significant effects on the thermal choke behaviors: (1) the thermal throat feature becomes much more irregular influenced by the rocket plume; (2) the occupancy range in the combustor is significantly lengthened; (3) the asynchrony of the flow in different regions accelerating to sonic speed becomes much more significant; (4) as the rocket throttling ratio decreases, the thermal choke position constantly moves upstream integrally, and the heated flow in the top region that is directly affected by the rocket plume reaches sonic speed more rapidly. Finally, we can conclude that appropriate secondary fuel injection adjustment can provide a higher integration thrust for the RBCC engine with the embedded rocket operating, while the thermal choke is stably controlled, and the increased heat release and combustion pressure are well balanced by the variations of pre-combustion shocks in the inlet isolator.

Investigation of flame flashback phenomenon in a supersonic crossflow with ethylene injection upstream of cavity flameholder

In the present study, a flame flashback phenomenon inside an ethylene-fuelled scramjet combustor equipped with a cavity flameholder is investigated under flight Mach 5.5 condition. Experimental results exhibit the quasi-periodic combustion oscillation between the fuel injectors and the leading edge of the cavity under a specified condition. As an indispensable key sub-process of combustion oscillation, the flame flashback from the boundary layer downstream of the cavity is responsible for unsteady combustion process in scramjet combustors. Numerical simulation has been carried out in the specified condition. Attributed by (i) thicker boundary layer, (ii) closer thermal disturbance and (iii) improved local mixing degree, the flame flashback phenomenon can be induced by thermal throat which is generated by interaction between separated boundary layer and intense combustion. Quantitative analysis also indicates that the flame flashing is more sensitive to temperature fluctuation downstream of the cavity. In addition, a simplified combustion opening system model has been established to analyse combustion oscillation mechanisms, which theoretically demonstrates that the three factors mentioned above can destroy the balance of heat release and dissipation, causing the system cannot self-stabilise once certain temperature fluctuation thresholds in sensitive areas are exceeded. At the same time, the auto-ignition model excludes the possibility of flame flashback generated by auto-ignition effect.

Experimental investigations of cavity parameters leading to combustion oscillation in a supersonic crossflow

The effects of cavity parameters on combustion oscillation inside an ethylene-fueled scramjet combustor equipped with a cavity flameholder are experimentally investigated for Mach 5.5 flight conditions. Three certain cases, such as i) longer length-to-depth ratio, ii) sharper aft degree of cavity, iii) closer air throttling downstream of the cavity, exhibit quasi-periodic combustion oscillation, which can be separately attributed to i) larger recirculation volume in the cavity and more mass and heat exchange between cavity shear layer and the core flow, ii) the stronger impinging shock wave in the cavity acting on shear layer, iii) improved fuel/air mixing owing to the interaction between separated boundary layer and combustion. High-speed and schlieren images demonstrate that the cavity and downstream of it act as the most sensitive areas, at the same time, the factors mentioned above can form a thermal throat further triggering flame flashback, which is an indispensable key sub-process of combustion oscillation. The quantitative analysis results obtained from iso-luminosity contour have shown the different distribution trends of flame front and distinct differences of quasi-periodic oscillation frequencies, whereas similar flame propagation speed distributions. In addition, a simplified combustion opening system model has been established to analyze combustion oscillation mechanisms, which theoretically demonstrates that above factors can destroy the balance of heat release and dissipation, causing the system cannot self-stabilize once certain temperature fluctuation thresholds in sensitive areas are exceeded.

Investigations of injection parameters on combustion oscillation in a supersonic crossflow

The influence of injection parameters on combustion oscillation inside an ethylene-fueled scramjet combustor equipped with a cavity flameholder is experimentally investigated for Mach 5.5 flight conditions. Experiments focusing on the influence of the fuel equivalence ratio, the premixing distance, the injection angle, and the number of jets have been carried out. Under certain conditions, high-speed camera and schlieren images show that flame oscillation occurs between the fuel injectors and the leading edge of the cavity. It is also demonstrated that those factors mentioned above can significantly affect the flow-shock interaction which generates low flow speed region and the mixing status which correlates with combustion intensity closely. The strong interaction between combustion and the separated boundary layer downstream of the cavity forms a thermal throat inducing thermal chocking to cause flame flashback. The distributions of the flame front, flame propagation speeds and quasi-periodic oscillation frequencies obtained by the iso-luminosity contour method show distinct differences in different cases. The strong interactions of flow and shock, combustion and separated region can easily trigger flame flashback, and the flame front can be maintained around the injector for long time in cases of deeper penetration depth and longer premixing distance. A simplified combustion opening system model predicts that the flame under sensitive conditions may acquire an exponential amplification once thresholds of the thermal fluctuation are exceeded.

Investigation of performance and mode transition in a variable divergence ratio dual-mode combustor

To operate in a wide range of incoming flow Mach number, a variable divergence ratio dual-mode combustor is designed. Experimental and numerical investigation of the dual-mode combustor has been conducted in this paper. By regulating the divergence ratio in combustor, in a high incoming flow Mach number or low equivalence ratio, a small divergence ratio could increase pressure peak and boost combustor performance. Under a low flight Mach number or a high equivalence ratio, a large divergence ratio has a benefit to accommodate more heat release and prevent the inlet unstart. The results of experiments and numerical investigations have been testified that the variable divergence ratio dual-mode combustor could operate in a wide range of incoming flow Mach number. Another aspect, the combustor wall ramp makes an additional effect on the combustion zone. The combustion zone distribution is close to the upper wall because of the pressure differentials of the main flow aroused by lower wall ramp compressing acting on the supersonic airflow. Under the effecting of incoming flow Mach number and heat release, Mach number near the leading edge of the ramp decides the location of the combustion zone. Comparing with a common combustor, it is harder to form a thermal throat in the combustor because of the big divergence angle in the variable divergence ratio combustor. Based on the flow field analysis, a conclusion is drawn that the combustion mode transition is dominated by the combustion heat release in the variable divergence ratio dual-mode combustor configuration.

Study on the effects of thermal throat on flame stabilization in a kerosene fueled supersonic combustor

The effects of thermal throat on flame stabilization in a kerosene fueled supersonic combustor were numerically and experimentally studied in the present paper. The results were obtained under the inflow condition of stagnation temperature 1100 K, stagnation pressure 1.0 MPa and Mach number 2.0. Wall pressure measurements and flame emission images were made during the experiments in an attempt to better understand the flame development, results showed the combustion organization scheme had a great effect on flame stabilization in the combustor. The flame development process in each case was different according to the flame luminosity pictures. When the pilot hydrogen was removed, the flame of case 1 and case 2 was still stable, and that of case 3 was blown off by the high-speed air flow, but it was surprised to find that the flame of case 4 was blowout even the pilot hydrogen existed. A shock train had generated in the isolator due to the thermal throat existing, which decreased the high-speed main stream, and increased the temperature and pressure, also enhanced the fuel mixing efficiency in the combustor, this was why the flame stabilization was achieved. But if the thermal throat disappeared, the shock train would be pushed out of the combustor by the high-speed air flow. The residence time for mixing, ignition and combustion was getting shorter, the flame was blown off finally.

Effect of continuous Mach number variation of incoming flow on ram–scram transition in a dual-mode combustor

Experimental and numerical investigation for a strut dual-mode combustor has been conducted in this paper. A Laval nozzle of alterable throat area is developed to change the incoming flow Mach number in isolator entrance. The Mach number of isolator entrance could be altered from 2.5 Ma to 2.0 Ma. By changing incoming flow Mach number, we obtained different pressure distribution variation at constant equivalence ratio. At 0.53 equivalence ratio, the pressure distribution in isolator and combustor has a different variation trend comparing with equivalence ratio 0.4. With incoming flow Mach number decreasing, the length of the shock train is shortened and pressure peak elevates. During the Mach number variation, the combustion heat release zone has a shift to strut. There is a balance for the combustion velocity and airflow velocity. The higher static temperature is also benefit for promoting combustion speed. The combustion zone shift makes more pressure increment act on the expansion segment to produce extra thrust under low incoming flow Mach number. Meanwhile, a ram–scram transition has taken place when the incoming flow Mach number variation. Further a subsonic discontinuity gap appears at constant segment and the thermal throat is chocked. The mass flow rate reaches up limits and leads to a two-stage pressure variation. The location of pressure variation would affect the combustor thrust.

Effect of Mach number and equivalence ratio on the pressure rising variation during combustion mode transition in a dual-mode combustor

Experiments of equivalence ratio linear increasing are utilized to employ to investigate the characteristics of combustion mode transition. For the experiments with kerosene of different increasing slope, it is found that the fuel equivalence ratio has a significant impact on pressure rising slope variation during combustion mode transition. In the experiments, the big equivalence ratio slope would not lead to slope variation of pressure rising. Similarly, this phenomenon also could not occur under some high Mach number of incoming flow experiments. Different equivalence ratio slope would lead to different border shape of the boundary layer. And the border shape of the boundary layer and normal shock wave at thermal throat restrains the main flow is an important factor to alter the pressure. Gas dynamics analysis is employed to explain the effect of the boundary layer on the pressure rising slope variation. Further, the analysis also reasonably explains that incoming flow Mach number impacts on pressure rising slope variation.

Pressure rising slope variation accompanying with combustion mode transition in a dual-mode combustor

Direct-connect experiment with equivalence ratio linear increasing in a dual-mode combustor with a strut is conducted at free stream Mach number of 2.0. Based on the pressure histories on side wall, pressure rising slope variation accompanying with combustion mode transition is found and discussed. The pressure rising has different slopes vary as equivalence ratio linearly increasing. Meanwhile, this characteristic of pressure slope variation leads to the change of combustor thrust. By analyzing the combustion heat release zone and the geometric configuration combustor, a typical simplified geometric model is proposed and studied by using numerical simulation. The phenomenon of the pressure rising slope variation is obtained and explained by the shock train movement in numerical simulation. In the process of shock train movement, a normal shock is established at thermal throat to make the airflow of throat position reach critical state, which could limit the mass flow rate of the supersonic mainstream. Further, inviscid and isentropic flow analysis is employed to illustrate the variation of pressure rising resulted from the occurrence of a normal shock wave. The analysis indicates that the cross-sectional area of mainstream is bounded by the subsonic boundary layer plays a key role to alter the pressure rising slope.

Flow field characteristics analysis and combustion modes classification for a strut/cavity dual-mode combustor

Experimental and numerical study of a strut/cavity dual-mode combustor has been conducted in this paper. Under different fuel equivalence ratio and allocation proportion conditions, the pressure distribution and flow field structure of combustor show distinct characteristics. For strut fuel injecting at a low equivalence ratio, the luminosity images show that combustion zone distributes in the shear layer behind the strut. The wall fuel injecting before strut would change the starting point of pressure rising. Based on the flow field structure, the dual-mode combustor operation process is classified into three combustion modes, including scramjet mode, weak ramjet mode and strong ramjet mode. Because of a strong interaction of the shock wave with the boundary layer, weak ramjet mode has a stronger isolator compression effect and higher combustion efficiency than scramjet mode. With heat release increasing, the thermal throat formation is an indication of the strong ramjet mode, which has a subsonic gap in the isolator. Further, by judging the pressure from dominant pressure sensor before the strut, the three different combustion modes could be classified. Comparing the specific impulse of combustor, it has an obvious distinction in the different combustion modes.

Effect of thermal choking on ejection process in a rocket-based combined cycle engine

Rocket-based combined cycle (RBCC) engines show great potential for use on future reusable launch vehicles. By combining traditional rocket and ramjet into a single engine, RBCC engines can make use of the oxygen in atmosphere and reduce the amount of onboard oxidizer required. The induced air reacts with the secondary injection fuel in combustor and the heat release may develop a thermal throat, which has an impact on the upstream ejection process. Combustor-ejector interaction plays an important role in the performance analysis of RBCC engines, but it is often ignored in previous studies. In this paper, a quasi-one-dimensional model focused on the influence of thermal choking is developed. A fixed-geometry engine operating in ejector mode is simulated to investigate the ejector mode of a hydrogen fueled RBCC engine. With a series of results, the ejector mode is classified into four distinctive types. By analyzing the operating conditions and the performance, the relation between thermal choking and ejection process is revealed. A deeper understanding of the working principle and the mode transition of RBCC engines can be gained, which may be useful for design and optimization of RBCC engines.

Mode transition of fuel control test in a dual-mode combustor

Ignition combustion test in dual-mode combustor with fuel flow rate change was carried out in a direct-connected wind tunnel facility that supplied airflow with total temperature of 880K, total pressure of 8atm and Mach number of 2.0. According to the wall pressure of combustor measured through the test, the distribution of other airflow parameters in the combustor was obtained by calculating with a one-dimensional analysis method, and the mode transition process in the condition of fuel flow rate change was studied. In addition, the position that the airflow in the isolator accessed to the subsonic velocity and the location of “choked thermal throat” at the combustor downstream were determined by analyzing of pressure distribution in the combustor. The results indicated that sole hydrogen combustion brought a supersonic combustion mode, then combined combustion of hydrogen and kerosene induced a subsonic combustion mode and combustion-induced pressure disturbance had been spread to the entrance to the combustor. In cases of sole kerosene combustion and the fuel equivalence ratio being kept at 0.53 and 0.61 respectively, both were in the subsonic combustion mode. Gradual increment of the fuel equivalence ratio was accompanied by enlarging subsonic combustion area. In case of combined combustion of hydrogen and kerosene, and improving kerosene flow rate, supersonic combustion mode was transformed to a subsonic one in a gradual manner.