Numerical Study On The Effect Of Tail Nozzle On The Airfield Of Continuous Rotating Detonation Turbine Combined Engine

On the basis of traditional turbine engine,continuous rotating detonation turbine combined engine has the special advantages of simple structure,self-pressurization and high thermal cycle efficiency of continuous rotating detonation engine,which has broad application prospect in aerospace propulsion field and has been widely concerned in recent years.There are various combinations of continuous rotating detonation engines and turbine engines.In this paper,an outer ring continuous rotating detonation combustor surrounds the vortex jet flow in the inner cavity.At present,the nozzle flow field under this combination form still needs further research.The internal and external flow field of the nozzle of the continuous rotating detonation turbine combined engine based on the above combination form is studied.The main work is as follows:Three-dimensional numerical simulation of combustor and outer flow field of a continuous rotating detonation turbine engine was carried out.The finite volume method is used to solve the three-dimensional unsteady Euler governing equation based on density,ignoring the influence of heat transfer and viscosity.The reactants are kerosene vapor gas and oxygen-rich air premixed gas.The chemical reaction adopts the finite rate model of one-step chemical mechanism.The changes of internal and external field structures and parameters from detonation initiation to stable propagation were calculated in the case of equivalent ratio 0.6.The results show that the intensity of the detonation wave on the outer wall is higher than that on the inner wall,and the height of the detonation wave on the outer wall is lower than that on the inner wall due to the characteristics of the annular wall structure.It is found that the overexpansion of the detonation product in the outflow field will form a shock wave propagating backwards which will enter the combustion chamber and affect the stability of the flow field in the combustion chamber.The effect of vortex jet flow field on detonation parameters is not obvious.An combined nozzle structure was carried out,in which the outer ring detonation combustor and the inner cavity were designed to share a single nozzle.The combined nozzle with different configurations was designed and the internal and external flow fields of the detonation combustor,the inner cavity and the nozzle were numerically simulated.The results show that the peak value of detonation pressure and detonation wave velocity are reduced in the case of both the combined straight nozzle and the common convergence nozzle with smaller convergence ratio.Detonation temperature peak is not sensitive tothe nozzle configuration change.The vortex jet flow field is always in subsonic velocity under different nozzle configurations.In the combined laval nozzle,the mixing speed of detonation product and vortex jet is the fastest,but the flow separation occurs in the expansion section of the nozzle,and a small eddy zone is generated between the mixing flow field and the wall surface of the expansion section.A coaxial separate nozzle structure scheme was developed for the detonation combustor nozzle and vortex jet nozzle respectively.Separate nozzle structures with different configurations were designed and the internal and external fields of the detonation combustor,central flow passage and nozzle were numerically simulated.The results show that zigzag fluctuations are easily formed in the premixed gas deflagration zone in the detonation combustion chamber under the condition of separate nozzle,which reduces the stability of detonation propagation in the detonation combustion chamber under the condition of combined nozzle.The impact of detonation products on the inner cavity of the separated nozzle is less than that of the combined nozzle.The propulsion performance of combined engines without nozzles,combined nozzles and separate nozzles was compared.The results show that:(1)Comparing the engine thrust and fuel specific impulse with or without vortex jet,it is found that the total fuel specific impulse of the engine decreases by 36.5% with the addition of vortex jet compared with the single detonation engine,indicating the high specific impulse advantage of detonation.(2)The Combined straight nozzle and the Combined nozzle with smaller convergence ratio reduces the propulsion performance,while the Combined nozzle with larger convergence ratio and the Combined laval nozzle enhanced combined engine propulsion performance.Especially,he Combined laval nozzle increases the engine thrust and fuel ratio by 15.7% and 14.6% respectively,which is the most obvious in all combined nozzle configurations.(3)The thrust and fuel specific impulse of the combined engine under the condition of the Separate double straight nozzle decrease by 1.6%and 2.6% respectively.The Separate convergence-straight nozzle,the Separate laval-straight nozzle and the Separate laval-convergence nozzle can improve the propulsion performance of the combined engine.Among them,the Separate laval-convergence nozzle is the best which increases the thrust and fuel specific impulse by 23.9% and 23.1% respectively.Compared with the combined nozzle,the separated nozzle can further improve the propulsion performance of the combined engine.