The flame establishment and propagation processes in a supersonic combustor fueled by liquid kerosene at Mach 6 condition were observed by high speed photography. In this paper, a thin strut acted as the flame holder equipped in the center of the combustor. In the experiment condition, the Mach number in the inlet of the combustor is 2.8, with the stagnation state Tt=1680K, Pt=1.87MPa. The combustion establishment process and the flame characteristics in the supersonic combustor were well captured and reproduced by the high-speed camera, with the camera parameter of 5000 frames per second. Experimental results show that initial flame appears around the plasma jet torch in the recirculation zone at the tailing edge of the strut, and grows to form a steady flame in the center of main inflow. The attribute of the flame is partially premixed flame in the experimental conditions, and the flame could be divided into three main parts in accordance with the flame characteristics. By the analysis of the high speed images, different flame propagation patterns in the flame establishment processes are found in different equivalence ratios, the mechanism of which is explained in this paper.

As part of our efforts to study the effect of the deflection angle on combustor performance of a variable geometry dual mode combustor, the flow field characteristics and mechanisms of the combustor performance loss, which comprised of compression loss, combustion heat addition loss and expansion loss, were investigated in the variable geometry dual mode combustor numerically with a Mach number of 3, a divergence ratio of 1.76, a fuel equivalence ratio of 0.6, and a deflection angle ranging from 8° to 16°. Numerical results indicated that the total pressure recovery coefficient and combustion efficiency increased with the deflection angle and there was a maximum to be obtained at the deflection angle of 12° due to the interaction between the dominant shock resulted by combustion heat release and the additional shock caused by the wedge system. Irreversible entropy generation loss was analyzed specifically in this paper to clarify and describe the combustor performance loss for the variable geometry dual mode combustor. Moreover, thrust-to-drag ratio was utilized to assess the effect of the deflection angle on combustor performance. By taking into account the flow field characteristics and combustor performance characteristics, the high combustor performance of a variable geometry dual mode combustor can be improved by selecting and optimizing the deflection angle.

As part of our efforts to study the hysteresis characteristic and formation mechanism for a variable geometry dual mode combustor, a series of geometry path continuous adjustment experiments and numerical simulation were conducted in the variable geometry dual mode combustor with a Mach number of 3, a divergent ratio ranging from 1.6 to 2.54 and a fuel equivalence ratio ranging from 0.6 to 1.0. Experimental results indicated that the wall static pressure of feature points had an obvious hysteresis phenomenon with the geometry path continuous variation. For given feature points, the hysteresis becomes much smaller with the increasing of the divergent ratio. The interaction between the oblique shock train motion and combustion heat release distribution was adequately considered to explicate the mechanism of the hysteresis formation. It was found that the hysteresis phenomenon was produced from the unstable positive feedback effect of the oblique shock train motion in the isolator. Moreover, the effect of the hysteresis on the total pressure loss, irreversible entropy loss and combustion performance were investigated numerically and experimentally. It was therefore strongly believed that the study of the hysteresis characteristic and formation mechanism on the combustion performance could be very significant to improve combustion performance for the variable geometry dual mode combustor, especially for a wide range of flight Mach numbers.

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.

•A thermodynamic cycle analysis model is developed to evaluate the thrust of RBCC.•The thrust of the ejector mode depends on the inducted air and thermal efficiency.•The optimum ejecting ratio at which the thrust attains a maximum is presented.Rocket-based combined cycle (RBCC) Engine can significantly reduce the amount of onboard oxidizer required. This will decrease the weight of the vehicle and improve the performance of the RBCC engine. In this short communication, an idealized thermodynamic cycle analysis is carried out to evaluate the thrust performance of RBCC engine for the saturated supersonic regime. The thrust for the rocket-ejector mode not only depends on the inducted air flow rate, but also depends on the thermal efficiency. Initially, the results show that the engine thrust grows asymptotically with ejecting ratio, then reaches a maximum, and finally reduces rapidly for a given primary stream conditions. The optimum ejecting ratio at which the value of the thrust attains a maximum is also presented.

•Mal-distribution of coolant causes a waste of heat sink and even over-temperature.•Distribution of cracked hydrocarbon fuel in parallel pipes are studied.•Two flow rate deviation amplification mechanisms are found.•The mal-distribution mode varies with thermal deviation.•Higher pressure suppresses the mal-distribution.A model consisting of two parallel pipes with common inlet and outlet manifolds was established and used to run simulation and experimental study on the flow rate distribution of cracked hydrocarbon fuel in parallel pipes under supercritical pressure. Both simulation and experimental results indicated that mass flow rate and fuel temperature distribution of cracked hydrocarbon fuel in parallel pipes was closely related to the difference in fuel density in pipes. Two deviation amplification mechanisms were found. In addition, the mode of mal-distribution varies with thermal deviation and the distribution was effectively improved by the increase of pressure. And the total mass flow rate could hardly have any effects on the flow rate distribution. All these results could be used to help the full utilization of fuel heat sink and avoid over-temperature.

•A hydrogen fueled scramjet model is developed to identify combustion mode transition.•Combustion mode transition is a double steady-state behavior.•Thermal choking leads to the shift of combustion mode.The combustion mode transition will occur repeatedly during the real flight of a hydrogen fueled scramjet. Under the assumption of inviscid flow, a series of the equations have been developed to identify combustion mode transition in an analytical way. The processes of heating and accelerating induced mode transition are investigated, respectively. This paper presents an analytical solution to understand the conditions for the transition between ramjet-scramjet operation modes. Besides, this paper also presents a qualitative explanation of the mode transition processes in the physical point of view. The mode transition is closely related to the thermal choking effects and shock wave motion. The process of mode transition is discontinuous and very fast. There is a sudden change with the combustor entry Mach number in this process.

•A numerical study is conducted for the convection transfer of fuel in cooling panel.•The structural optimization is conducted to the blockage structure.•Properties variation of fuel affects the flow field around the blockage structure.•The blockage structure is suitably designed in the subcritical temperature region.The convection heat transfer of hydrocarbon fuel at supercritical pressure has a great influence on the regenerative cooling technology of a scramjet engine. A three-dimensional numerical simulation was conducted for the convection transfer of hydrocarbon fuel in the cooling panel of a combustion chamber wall. And the flow field around the local flow blockage structure and the outlet flow rate distribution characteristics of fuel in the cooling channels were analyzed in detail. The results of analyses indicate that with the optimized local flow blockage structure, the outlet flow rate distribution of fuel among the cooling channels become more uniform, as the area of local flow dead zone decreases. However, as the fuel temperature increases, the dramatic variation of thermodynamic physical properties of fuel has a strong influence on the flow field around the local flow blockage structure. Especially, a local flow dead zone can be easily formed in the supercritical temperature region. Meanwhile, transverse pressure gradient around the throat region of blockage structure and additional loss, which is caused by turbulence fluctuation and energy exchange of fluid in the downstream area, affect the outlet flow rate distribution of fuel among the coolant passages seriously. It can therefore be concluded that the local flow blockage structure is more suitably designed in the subcritical temperature region by taking above-mentioned factors into consideration.

•Combustion mode transition is the way to realize wide range operation of scramjet.•Combustion mode optimization can improve hydrogen fueled scramjet performance.•The optimal operation route of hydrogen fueled scramjet is given.Hydrogen fueled scramjet is a candidate for use as the engine of the aerospace plane for its high specific impulse. To further improve the specific impulse performance, analysis of combustion mode and operating route for a hydrogen fueled scramjet engine was investigated in this study. A scramjet engine with two-staged hydrogen injection was simulated by one-dimension numeric method within the acceleration from Mach 4 to 7. Three typical combustion modes (scramjet-mode, transitional mode and ramjet-mode) could be attained by changing the total amount of fuel added or adjusting the fuel distribution between two injectors. Simulation results show that better thrust performance can be achieved as more fuel injected at the upstream fuel injector as possible, while ensuring the engine safety. From a standpoint of specific impulse maximization, an optimal scramjet combustion mode database was presented and the boundary of the combustion mode transition was determined. Meanwhile, optimal operating route was also suggested for scramjet operation in this study.

Hydrogen fueled scramjet is a candidate for use as the engine of the aerospace plane. Many methods to increase the fuel heat sink (cooling capacity) are widely carried out to meet the cooling requirement of scramjet engine. Especially, recooling cycle (RCC) has been newly proposed for actively cooled scramjet to increase fuel heat sink for cooling. With the working process of fuel cooled scramjet defined as a recuperated cycle, the difference between regenerative cooling (RC) and RCC in the characteristics of heat recovery, heat extraction and heat reinjection are compared first, and the scramjet performance of under RC and RCC mode is compared through off-design comparison analysis at different Mach numbers and equivalence ratios to evaluate the effect of recooling cycle on the performance of scramjet using a coupled heat transfer and flow model. It can be seen through analysis and comparison that recooling cycle can save the fuel mass flow rate for cooling, and make maximum use of the recovered heat. And the specific impulse of a scramjet can be greatly promoted as well.Highlights► The fuel cooling of scramjet is a heat recuperation process. ► The working process of a scramjet with fuel cooling is a recuperative cycle. ► Scramjet performance with recooling cycle is superior to that with regenerative cooling. ► Recooling cycle can make maximum use of the heat recovered.

A closed Brayton cycle thermal management system is proposed for a regeneratively cooled scramjet to reduce the hydrogen fuel flow for cooling, through converting part of the heat from fuel to other forms of energy to decrease the heat that must be taken away by hydrogen fuel. Fuel heat sink (cooling capacity) is thus indirectly increased. Instead of carrying excess fuel for cooling or seeking for any new coolant, the fuel flow for cooling is reduced, and fuel onboard is adequate to satisfy the cooling requirement for the whole hypersonic vehicle. A parametric study of an irreversible closed Brayton cycle thermal management system for scramjet has been performed with external as well as internal irreversibilities. It is known through performance analyses that closed Brayton cycle thermal management system has excellent potential performance over conventional regenerative cooling, due to the reduction in fuel flow for cooling and additional power output.

A Re-Cooled Cycle is newly proposed for a regeneratively cooled scramjet to reduce the hydrogen fuel flow for cooling. Upon the completion of the first cooling, hydrogen fuel can be used for secondary cooling by transferring the enthalpy from fuel to work. Fuel heat sink (cooling capacity) is thus repeatedly used to indirectly increase the fuel heat sink. Instead of carrying excess fuel for cooling or seeking for any new coolant, the cooling fuel flow is reduced, and fuel onboard is adequate to satisfy the cooling requirement for the whole hypersonic vehicle. Two new performance parameters are defined combining the characteristics of scramjet, differences between RCC and expander cycle is discussed, and pressure loss in cooling passage is also considered. Performance comparison is carried among multiple RCCs, and preliminary application limit and real possibility of RCC for scramjet is discussed.

•A state observer-based method is developed to measure inner wall temperature.•A mathematical model is established to describe the heat transfer of the combustor.•Results of simulations and experiments indicate a good accuracy of the method.A state observer-based method is developed for the indirect online measurement of the inner wall temperature using the out surface temperature and pressure of a scramjet combustor. A mathematical model is established to describe the heat transfer from gas to combustor wall and inside combustor wall as well. A proportional integral observer is developed using the mathematical model for establishing the relationship between the observed inner wall temperature and the experimentally measurable parameters, including the out surface temperature and pressure. Numerical simulations and ground experiments are carried out with a direct-connect hydrocarbon fueled scramjet combustor to prove the validity of the proportional integral observer. Test results indicate the proportional integral observer method could be used to measure the inner wall temperature of the scramjet combustor.

•The staged-combustor with dual-strut is introduced to avoid inlet-combustor interaction.•Fuel injection scheme has a significant effect on the staged-combustor operating modes.•Thrust with different combustion mode is reported through staged-combustor experiments.To enable the scramjet operate in a wider flight Mach number, a staged-combustor with dual-strut is introduced to hold more heat release at low flight Mach conditions. The behavior of mode transition was examined using a direct-connect model scramjet experiment along with pressure measurements. The typical operating modes of the staged-combustor are analyzed. Fuel injection scheme has a significant effect on the combustor operating modes, particularly for the supersonic combustion mode. Thrust performances of the combustor with different combustion modes and fuel distributions are reported in this paper. The first-staged strut injection has a better engine performance in the operation of subsonic combustion mode. On the contrast, the second-staged strut injection has a better engine performance in the operation of supersonic combustion mode.

Highlights•A coordinated control strategy for ducted rocket regulation and protection is proposed.•The strategy ensures that the engine obtains the maximum performance while ensuring safety.•Inner/outer loop control structure decomposes the contradiction between performance and safety.•A limit protection controller design is developed utilizing a linear quadratic optimal control technique.•A hysteresis coordinated control logic is presented for thrust regulation and inlet buzz protection.This study is concerned with the coordinated control problem for regulation/protection mode-switching of a ducted rocket, in order to obtain the maximum system performance while ensuring safety. The proposed strategy has an inner/outer loop control structure which decomposes the contradiction between performance and safety into two modes of regulation and protection. Specifically, first, the mathematical model including the actuator (gas regulating system) and the plant (ducted rocket engine) is introduced. Second, taking the inlet buzz for instance, the ducted rocket coordinated control problem for thrust regulation and inlet buzz limit protection is formulated and discussed. Third, to solve the problem, based on the main inner loop, a limit protection controller (outer loop) design method is developed utilizing a linear quadratic optimal control technique, and a coordinated control logic is then presented. At last, the whole coordinated control strategy is applied to the ducted rocket control model, and simulation results demonstrate its effectiveness.

A new Re-cooled Cycle has been proposed for a regeneratively cooled scramjet to reduce the fuel flow for cooling, fuel heat sink (cooling capacity) is repeatedly used to indirectly increase the fuel heat sink, and fuel onboard is adequate to satisfy the cooling requirement for the whole hypersonic vehicle. Parametric analysis of the Re-cooled Cycle carried out in previous study shows the potential for the cycle to be optimized. A thermodynamic optimization analysis of Re-cooled Cycle for a scramjet is obtained by satisfying hydrodynamic, thermal, and Mach number constraints. The optimal cycle parameters and length allocation between the first and second cooling passages are obtained.

The supersonic combustion ramjet (scramjet) engine is expected to become the most efficient air breathing propulsion system in the hypersonic flight regime. The behavior of wall pressures propagation was examined using a direct-connect model scramjet experiment along with high-frequency pressure measurements. High-frequency pressure transducers are employed to record the history of pressure response during ignition process. A first order plus dead time (FOPDT) model is employed to model the S-shape step response curve. The time scale distributions of pressure propagation are reported in this paper. The characteristics of pressure propagation are closely related to the combustion mode of the scramjet.

•Combustion modes could be classified by dominant pressure point.•Positive feedback between isolator and combustor promotes combustion efficiency.•Positive feedback is a feature of weak ramjet different from scramjet mode.•Thermal throat with a subsonic gap is a sign of strong combustion mode.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.

•Shock train reveals the different characteristics under the effect of combustion.•High temperature gas in boundary layer leads to shock train shrinkage and stretch.•Shock train movement processes are classified into three stages.In this paper, the effect of heat release on movement characteristics of shock train is numerically investigated in an isolator. It is found that the combustion heat release has a distinct effect on the shock train movement characteristics in the isolator. With increasing heat release, a shock train gradually forms and then propagates toward isolator entrance. In process of shock train formation, separation bubbles before injection ports entrain the high temperature burning gas into the boundary layer, which causes the shock train to shrink and stretch, and changes in configuration and number of shock waves. At the same time, the system force fluctuates. In addition, the shock train movement is divided into three stages, which have different wall pressure distribution. It is believed that these findings have a help the better understanding of the effect of heat release on the movement characteristics of shock train in an isolator.