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Research on the Secondary Combustion Chamber Operation Process of Boron-Based Propellant Ducted Rockets

Author: HuJianXin
Tutor: XiaZhiXun
School: National University of Defense Science and Technology
Course: Aeronautical and Astronautical Science and Technology
Keywords: Solid Rocket Ramjet Secondary Combustion Chamber Boron-Based Propellant Two Phase Flow Numerical Simulation Connected-Pipe Ramjet Test Temperature Measurement Response Surface Method
CLC: V435.12
Type: PhD thesis
Year: 2006
Downloads: 479
Quote: 25
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Abstract

Boron has been considered for many years as a prime candidate used for increasing the ducted rocket capabilities based on its high potential energy release on both a volumetric and gravimetric basis coupled with a high-energy of combustion, high c ombustion temperature, and low molecular weight products. These properties make boron an attractive material for use in ducted rocket propellants. In order for these advantages to be realized, however, the boron particles must ignite and burn completely within a very limited residence time. Since boron particles are generally initially coated with an oxide layer that inhibits combustion and since boron has an extremely high boiling point, which necessitates surface burning subsequent to oxide removal, this can become difficult, particularly under adverse operating conditions. Although the aforementioned work has provided considerable information to the understanding of flow and combustion process in secondary combustion chambers, many fundamental issues regarding the detailed mixing and combustion process, especially in the secondary combustion chambers of the ducted rocket, still need to be addressed.In view of this, a research program involving both experimental and numerical approaches is initiated. This thesis is aimed at providing the complete details of the combustion processes and flowfields in the secondary combustion chambers and seeking ways to improve the combustion efficiency in order to provide theoretical basis and direction for the design of secondary combustion chambers. The results attained are as follows.In the secondary combustion chamber, there always exists a relative speed between the boron particle and air stream. Hence, it is required to study the effect of forced convective on the ignition and the burning of the boron particle. Based on the boundary layer theory and aerodynamic stripping, a theoretical model for the ignition of the boron particle in high-speed flow has been established for the problem of the aerodynamic stripping, which is a result form the interaction between gas and the particle surface when the particle is in high-speed flow. The influence of temperature, pressure, initial particle radius, the mole fraction of oxygen on the ignition of the boron particle are investigated. An explicit expression for the burning rate of the boron particle under forced convection conditions is developed in terms of the Reynolds and mass transfer numbers. The special features of this study are that the partial differential equations for the two-dimensional boundary layer are simplified into a set of ordinary differential equations by a similarity transformation and then their numerical solution can be easily conducted.Reacting flow model in the secondary combustion chambers of the ducted rocket is established. The three-dimensional Favre-averaged compressible turbulent N-S equations are used as the governing equations of the reacting flow, the improved k～εtwo-equation turbulence model is used to simulate the turbulent flow, and Eddy-Dissipation Model is used to simulate the gas combustion. The particle-phase solution is obtained using a well-established boron particle ignition and combustion model. Boron particles are ejected from the exit of gas generator into secondary combustion chambers and their trajectories are traced through the reacting flow field using discrete phase models. Information gained from CFD modeling provides future engine designers valuable information regarding the design of the secondary combustion chambers.The secondary combustion in a cylindrical combustor for the ducted rocket is investigated preliminarily conducted on a direct-connect test bed. During the test, boron-based fuel-rich HTPB propellant is used. The effect of factors, such as air/fuel ratio, velocity of air injecting, dome height, angle of gas injector and others on the performance of the combustor and the engine is examined. The experimental results show that the reacting flow model established in this paper is correct.Based on the thermodynamics calculation, four kinds of vitiator fuel (hydrogen, kerosene, methane and alcohol) are investigated and the effect of four kinds of vitiator fuel on the performance in connected-pipe ramjet testing is investigated. The study shows that the effect of vitiated air on the performance of ramjet is petty.In order to get better insight in the combustion behavior of the secondary combustion chambers, a high-speed photography has been used to analyze the highly turbulent multiphase combustion flow. The portion of the primary plume is viewed through a quartz window and is recorded by photography. The geometry has been designed with a square section for the duct to facilitate optical access inside the combustor. That rebuilding the combustion flame’s temperature field by applying the digital image processing technology based on the principle of primary colors is introduced.The process of mixing and chemical reaction in the secondary combustion chambers is very complicated, and many factors affect the combustion efficiency. The new method combining computation fluid dynamic with orthogonal-polynomial based response surface methodology is applied successfully to optimize the effects of secondary combustion chambers. The solution determined by response surface methodology is based on the yield of the maximum combustion efficiency. And some meaningful result is derived, which shows the effectiveness of this new method.

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CLC: > Aviation, aerospace > Aerospace ( Astronauts ) > Propulsion system ( engine, propeller ) > Solid propellant rocket motor > Principles and calculation > Combustion process and combustion characteristics
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