Investigation Of Characteristics And Boundary For Combustion Mode Transition In Central Strut Supersonic Combustor

In order to meet a high-performance flight mission,a dual-mode scramjet engine need to operate with different combustion modes in a wide range incoming flow Mach number.Therefore,a suitable operating combustion mode would have a significant influence on promoting the engine performance.Comparing with other supersonic combustor configurations,a central strut supersonic combustor has a unique operating pattern for coupling of flow and combustion.The combustion heat release zone is anchored in the centre of combustor channel,which would lead to some unique flow field characteristics in the combustion mode transition process.Aiming at the special flow characteristics during combustion mode transition,a series of investigations is necessary.Therefore,to investigate the process of combustion mode transition in a strut supersonic combustor configuration,some research work has been carried out as follows:Aiming at the characteristics of complicated dynamic flow field in a dual-mode supersonic combustor produced by strong coupling effect between combustion and flow,the evolution process and shock train characteristics in an isolator aroused by combustion heat release and incoming flow Mach number is investigated.The research shows that the characteristics of shock train movement resulting from combustion heat release in the isolator have an obvious distinction comparing with backpressure linear increasing at isolator exit.During the process of combustion heat release increasing,a phenomenon of flame flashback along with sidewalls is produced due to the effect of combustion and flow.The flame flash back has a close relationship with the recirculation zone near the upstream wall.Based on the process of shock train movement with coupling effect of combustion and flow,the evolution of shock train could be classified into three stages,which correspond to the stage of different combustion modes.By considering the path variation of incoming flow Mach number and equivalence ratio,it is found that the shock train patterns variation depend on the path variation,and the path variation would lead to a hysteresis of shock train location.By affecting the shock train structure,incoming flow Mach number has a large effect on the hysteresis phenomenon.The flow field characteristics under the coupling effect of combustion and supersonic flow,are investigated in the process of combustion mode transition by virtue of ground experiments of supersonic combustion.The research shows that the combustion mode transition could be obtained by linearly increasing equivalence ratio.And pressure rising slope variation on the sidewalls appears with combustion mode transition and the phenomenon could force the combustor thrust change.According to the numerical simulation results which have similar conditions with experiments,a same phenomenon that is pressure rising slope variation is obtained.Based on the flow field structure analysis and thermodynamic relation,it is clarified that the pressure rising slope variation is aroused by the mass flow rate chock and the thermal throat from boundary layer restricting effect on the main supersonic flow.Based on the flow field structure obtained from numerical simulation,modeling is built to clarify the phenomenon of pressure variation and predicted the experiment results.For the factors influencing combustion mode transition,the two factors of incoming flow Mach number variation and divergence ratio variation are investigated to combustion mode transition.By a series of results from ground experiments and numerical simulation,it is indicated that the incoming flow Mach number at isolator exit and divergence ratio is changed has a significant effect on the combustion mode transition.Mach number decreasing could concentrate the combustion heat release,and decelerate the velocity of the supersonic flow.Further,the velocity variation of the main flow decreases the local Mach number.Scramjet mode in combustor is forced to transform ramjet mode.The decreasing divergence ratio leads to the heat release space reduced,which also makes the combustion heat release concentrated and local static temperature increase.The local static temperature enhanced the local velocity of sound,which would make the scramjet mode in combustor is forced to transform ramjet mode.Investigating flow field structure and flow field characteristics in the strut supersonic combustor,the combustion mode of the thermal chock and the pneumatic chock definition and classification are proposed by two operational patterns,respectively.By analyzing the factors of combustion mode transition,an operating state space of supersonic combustor described by the momentum ratio,equivalence ratio and sectional area ratio is proposed.Operating state,combustion mode boundary and the combustion mode margin are obtained by disposing the experiments results.Multiple factors effect on the combustion mode boundary are analyzed and investigated.The results show that the divergence ratio of combustor could achieve a wide range combustor mode boundary controlled.