Simulation Study On Ignition Process And Internal Flow Field Of Micro/Nano Satellite Solid Rocket Motor

Solid rocket motors are easy to realize integrated and modular design because of its advantages such as simple structure,rapid response ability and high working reliability.It has a wide application prospect in the large-scale orbital maneuver of micro/nano satellites.Therefore,this thesis is oriented to solid rocket motors for micro/nano satellites.In order to obtain a technical approach to improve the ignition performance of solid rocket motors and promote its engineering application in space,numerical simulation and experimental verification are conducted.The ignition characteristics and vacuum plume phenomenon under different parameters of the motors with fore-ignition and aft-ignition in vacuum are researched.By analyzing the ignition gas filling and pressure establishment process,the influence of structural parameters on ignition performance and the distribution of flow parameters are obtained,which has guidance for the structural optimization design of solid rocket motors and the study of grain integrity analysis.The main research work of this thesis includes:(1)The mass flow rate model of ignition chamber for the black powder is modified by introducing the combustion diffusion coefficient.A numerical simulation model of the ignition process in the solid rocket motors was established in vacuum.Prototype motors are designed with different ignition methods,the outlets of ignition chamber,grain positions,and expansion ratios.And the vacuum ignition test system is built,including a vacuum test device,a protection box,an image and data acquisition system.(2)Two dimensional axisymmetric model of the motor is established using front ignition method,the effects of design parameters such as igniter quantity,grain position,outlet size of ignition chamber,and the combustion process on the ignition performance are studied.Comparing with experimental results,the validity of the numerical simulation is verified.Meanwhile,the vacuum plume phenomenon is analyzed.It is concluded that increasing the amount of igniter quantity will increase the ignition pressure peak,which will adversely affect grain integrity.The ignition performance of black powder changes under vacuum conditions.0.2g ignition powder cannot break the end cap.A large number of particles ejected are recombusted in the air environment,while the particles in vacuum do not further burn.The grain position in the middle can improve the stability of flow field in ignition process,which is beneficial to the stable development of flame.Ignition flow loss will increase as the ignition chamber outlet size increasing.When considering combustion massing process,propellant combustion and ignition gas in chamber work together to accelerate the pressure establishment process.(3)The influence of design parameters such as expansion ratio with aft-end igniter on gas filling and pressure building process is analyzed,and the rationality of numerical simulation calculation is verified by comparing with experiments.It is concluded that at the same igniter quantity,the internal flow field impact on aft-end igniter is more severe than front ignition method,and there is an obvious pressure oscillation.The pressure difference reaches1.82MPa(0.4MPa in fore-ignition),which will do harm to grain structural integrity.With the expansion ratio increasing at the same throat diameter,the pressure in the combustion chamber will increase and it will delay the opening.Comparing with the atmospheric pressure test,the vacuum plume phenomenon is more obvious,the stability of the ignition performance is poorer,and the burning time is shorter.Through the research work in this thesis,the in-depth discussion of the ignition process under different methods and parameters was conducted.A theoretical and practical foundation is laid for improving the ignition performance of solid rocket motor and promoting their engineering applications in space.