To predict the thermal and structural responses of the thrust chamber wall under cyclic work,a 3-D fluid-structural coupling computational methodology is developed.The thermal and mechanical loads are determined by a ...To predict the thermal and structural responses of the thrust chamber wall under cyclic work,a 3-D fluid-structural coupling computational methodology is developed.The thermal and mechanical loads are determined by a validated 3-D finite volume fluid-thermal coupling computational method.With the specified loads,the nonlinear thermal-structural finite element analysis is applied to obtaining the 3-D thermal and structural responses.The Chaboche nonlinear kinematic hardening model calibrated by experimental data is adopted to predict the cyclic plastic behavior of the inner wall.The methodology is further applied to the thrust chamber of LOX/Methane rocket engines.The results show that both the maximum temperature at hot run phase and the maximum circumferential residual strain of the inner wall appear at the convergent part of the chamber.Structural analysis for multiple work cycles reveals that the failure of the inner wall may be controlled by the low-cycle fatigue when the Chaboche model parameter c3= 0,and the damage caused by the thermal-mechanical ratcheting of the inner wall cannot be ignored when c3〉 0.The results of sensitivity analysis indicate that mechanical loads have a strong influence on the strains in the inner wall.展开更多
A 500 N model engine filled with LO2/GCH4 was designed and manufactured.A series of ignition attempts were performed in it by both head spark plug and body spark plug.Results show that the engine can be ignited but th...A 500 N model engine filled with LO2/GCH4 was designed and manufactured.A series of ignition attempts were performed in it by both head spark plug and body spark plug.Results show that the engine can be ignited but the combustion cannot be sustained when head spark plug applied as the plug tip was set in the gaseous low-velocity zone with thin spray.This is mainly because flame from this zone cannot supply enough ignition energy for the whole chamber.However,reliable ignition and stable combustion can be achieved by body spark plug.As the O/F ratio increases from 2.61 to 3.49,chamber pressure increases from 0.474 to 0.925 MPa and combustion efficiency increases from 57.8%to 95.1%.This is determined by the injector configuration,which cannot produce the sufficiently breakup of the liquid oxygen on the low flow rate case.展开更多
To investigate the damage localization effects of the thrust chamber wall caused by combustions in LOX/methane rocket engines, a fluid-structural coupling computational methodology with a multi-channel model is develo...To investigate the damage localization effects of the thrust chamber wall caused by combustions in LOX/methane rocket engines, a fluid-structural coupling computational methodology with a multi-channel model is developed to obtain 3-demensioanl thermal and structural responses.Heat and mechanical loads are calculated by a validated finite volume fluid-thermal coupling numerical method considering non-premixed combustion processes of propellants. The methodology is subsequently performed on an LOX/methane thrust chamber under cyclic operation. Results show that the heat loads of the thrust chamber wall are apparently non-uniform in the circumferential direction. There are noticeable disparities between different cooling channels in terms of temperature and strain distributions at the end of the hot run phase, which in turn leads to different temperature ranges, strain ranges, and residual strains during one cycle. With the work cycle proceeding, the circumferential localization effect of the residual strain would be significantly enhanced. A post-processing damage analysis reveals that the low-cycle fatigue damage accumulated in each cycle is almost unchanged, while the quasi static damage accumulated in a considered cycle declines until stabilized after several cycles. The maximum discrepancy of the predicted lives between different cooling channels is about 30%.展开更多
This paper presents a method of thermal state calculation of combustion chamber in small thrust liquid rocket engine. The goal is to predict the thermal state of chamber wall by using basic parameters of engine: thrus...This paper presents a method of thermal state calculation of combustion chamber in small thrust liquid rocket engine. The goal is to predict the thermal state of chamber wall by using basic parameters of engine: thrust level, propellants, chamber pressure, injection pattern, film cooling parameters, material of wall and their coating, etc. The difficulties in modeling the startup and shutdown processes of thrusters lie in the fact that there are the conjugated physical processes occurring at various parameters for non-design conditions. A mathematical model to predict the thermal state of the combustion chamber for different engine operation modes is developed. To simulate the startup and shutdown processes, a quasi-steady approach is applied by replacing the transient process with time-variant operating parameters of steady-state processes. The mathematical model is based on several principles and data commonly used for heat transfer modeling: geometry of flow part, gas dynamics of flow, thermodynamics of propellants and combustion spices, convective and radiation heat flows, conjugated heat transfer between hot gas and wall, and transient approach for calculation of thermal state of construction. Calculations of the thermal state of the combustion chamber in single-turn-on mode show good convergence with the experimental results. The results of pulsed modes indicate a large temperature gradient on the internal wall surface of the chamber between pulses and the thermal state of the wall strongly depends on the pulse duration and the interval.展开更多
In order to measure the instantaneous thrust of a certain attitude-control solid rocket motor, based on the analysis of the measurement principles, the difference between the instantaneous thrust and steady thrust mea...In order to measure the instantaneous thrust of a certain attitude-control solid rocket motor, based on the analysis of the measurement principles, the difference between the instantaneous thrust and steady thrust measurements is pointed out. According to the measurement characteristics, a dynamic digital filter compensation method is presented. Combined the identification-modeling, dynamic compensation and simulation, the system's dynamic mathematic model is established. And then, a compensation digital filter is also designed. Thus, the dynamic response of the system is improved and the instantaneous thrust measurement can be implemented. The measurement results for the rocket motor show that the digital filter compensation is effective in the instantaneous thrust measurement.展开更多
The solid rocket motor driven system is one of the common ways for submarines to launch underwater missiles. It has significant advantages in improving the missile’s water exit speed, anti-interference capability, an...The solid rocket motor driven system is one of the common ways for submarines to launch underwater missiles. It has significant advantages in improving the missile’s water exit speed, anti-interference capability, and enemy striking power. The prediction of the underwater loading is a preliminary factor for the power system design of the underwater vehicle. This paper presents a rapid prediction method and validated by the experimental study for the underwater thrust of the solid rocket motor. Based on the potential flow assumption of the water field, a model of the bubble and a one-dimensional quasi-steady model of the nozzle are established to directly solve the flow status of the nozzle. The aerodynamic thrust and hydrodynamic thrust have been calculated and analyzed. The calculation results are within 5% error of the experimental results. Moreover, a design platform to predict the underwater thrust of the solid rocket motor has been developed based on Python and the PyQt library, which shows excellent system adaptability and computational efficiency.展开更多
To evaluate the structural failure risk of the regenerative cooling thrust chamber cylinder segment,a Finite Element Method(FEM)based on experimental data was developed.The methodology was validated and utilized to re...To evaluate the structural failure risk of the regenerative cooling thrust chamber cylinder segment,a Finite Element Method(FEM)based on experimental data was developed.The methodology was validated and utilized to reveal the thermal response and the nonlinear deformation behavior of the cylinder segment phase by phase.The conclusions of the research are as follows:The 2 D heat flux distribution caused by the injector determines the uneven temperature distribution on the gas-side wall and leads to the temperature disparity between various cooling channels;The reason for the accumulation of residual strain is that the tensile strain generated in the post-cooling phase is greater than the compressive strain produced in the hot run phase;Through the single-cycle simulation,two potential failure locations with conspicuous deformations were found,but it is difficult to determine which point is more dangerous.However,the multi-cycle thermo-structural analysis gives the evolution of the stress-strain curve and gradually discloses that the low-temperature corner of a particular channel is the most likely location to fail,rather than the maximum residual strain point of the gas-side wall.The damage analysis for dangerous point indicates that the quasistatic damage accounts for the majority of the total damage and is the main factor limiting the service life.展开更多
On August 2,a twin-segment solid rocket motor of the largest diameter,grain mass and thrust in China completed its ground test firing with success.The3 m solid motor was independently developed by the Academy of Aeros...On August 2,a twin-segment solid rocket motor of the largest diameter,grain mass and thrust in China completed its ground test firing with success.The3 m solid motor was independently developed by the Academy of Aerospace Solid Propulsion Technology（AASPT）under CASC.展开更多
A hybrid rocket motor combines components from both solid fuel and liquid fuel rocket motors. The fuel itself is a solid grain, (often paraffin or hydroxyl-terminated polybutadiene, known as HTPB) while the oxidizing ...A hybrid rocket motor combines components from both solid fuel and liquid fuel rocket motors. The fuel itself is a solid grain, (often paraffin or hydroxyl-terminated polybutadiene, known as HTPB) while the oxidizing agent is liquid (often hydrogen peroxide or liquid oxygen). These components are combined in the fuel chamber which doubles as the combustion chamber for the hybrid motor. This review looks at the advances in techniques that have taken place in the development of these motors since 1995. Methods of testing the thrust from rocket motors and of measuring the rocket plume spectroscopically for combustion reaction products have been developed. These assessments allow researchers to more completely understand the effects of additives and physical changes in design, in terms of regression rates and thrust developed. Hybrid rocket motors have been used or tested in many areas of rocketry, including tactical rockets and large launch vehicles. Several additives have shown significant improvements in regression rates and thrust, including Guanidinium azotetrazolate (GAT), and various Aluminum alloys. The most recent discoveries have come from research into nano-particle additives. The nano-particles have been shown to provide enhancements to many parameters of hybrid rocket function, and research into specific areas continues in the sub-field of nano-additives for fuel grains.展开更多
This study focuses on the stepwise procedure involved in the development of a numerical model of a bi-propellant hypergolic chemical propulsion system using key features and performance characteristics of existing and...This study focuses on the stepwise procedure involved in the development of a numerical model of a bi-propellant hypergolic chemical propulsion system using key features and performance characteristics of existing and planned (near future) propulsion systems. The study targets specific impulse of 100</span></span><span><span><span style="font-family:""> </span></span></span><span style="font-family:Verdana;"><span style="font-family:Verdana;"><span style="font-family:Verdana;">N delivery performance of thrust chambers which is suitable for primary propulsion and attitude control for spacecraft. Results from numerical models are reported and validated with the Rocket Propulsion Analysis (RPA) computation concept. In the modelling process, there was proper consideration for the essential parts of the thruster engine such as the nozzle, combustion chamber, catalyst bed, injector, and cooling jacket. This propulsion system is designed to be fabricated in our next step in advancing this idea, using a combination of additive manufacturing technology and commercial off the shelf (COTS) parts along with non-toxic propellants. The two non-toxic propellants being considered are Hydrogen Peroxide as the oxidiser and Kerosene as the fuel, thus making it a low-cost, readily available and environmentally-friendly option for future microsatellite missions.展开更多
文摘To predict the thermal and structural responses of the thrust chamber wall under cyclic work,a 3-D fluid-structural coupling computational methodology is developed.The thermal and mechanical loads are determined by a validated 3-D finite volume fluid-thermal coupling computational method.With the specified loads,the nonlinear thermal-structural finite element analysis is applied to obtaining the 3-D thermal and structural responses.The Chaboche nonlinear kinematic hardening model calibrated by experimental data is adopted to predict the cyclic plastic behavior of the inner wall.The methodology is further applied to the thrust chamber of LOX/Methane rocket engines.The results show that both the maximum temperature at hot run phase and the maximum circumferential residual strain of the inner wall appear at the convergent part of the chamber.Structural analysis for multiple work cycles reveals that the failure of the inner wall may be controlled by the low-cycle fatigue when the Chaboche model parameter c3= 0,and the damage caused by the thermal-mechanical ratcheting of the inner wall cannot be ignored when c3〉 0.The results of sensitivity analysis indicate that mechanical loads have a strong influence on the strains in the inner wall.
基金Project(613239)supported by the National Basic Research Program of China
文摘A 500 N model engine filled with LO2/GCH4 was designed and manufactured.A series of ignition attempts were performed in it by both head spark plug and body spark plug.Results show that the engine can be ignited but the combustion cannot be sustained when head spark plug applied as the plug tip was set in the gaseous low-velocity zone with thin spray.This is mainly because flame from this zone cannot supply enough ignition energy for the whole chamber.However,reliable ignition and stable combustion can be achieved by body spark plug.As the O/F ratio increases from 2.61 to 3.49,chamber pressure increases from 0.474 to 0.925 MPa and combustion efficiency increases from 57.8%to 95.1%.This is determined by the injector configuration,which cannot produce the sufficiently breakup of the liquid oxygen on the low flow rate case.
文摘To investigate the damage localization effects of the thrust chamber wall caused by combustions in LOX/methane rocket engines, a fluid-structural coupling computational methodology with a multi-channel model is developed to obtain 3-demensioanl thermal and structural responses.Heat and mechanical loads are calculated by a validated finite volume fluid-thermal coupling numerical method considering non-premixed combustion processes of propellants. The methodology is subsequently performed on an LOX/methane thrust chamber under cyclic operation. Results show that the heat loads of the thrust chamber wall are apparently non-uniform in the circumferential direction. There are noticeable disparities between different cooling channels in terms of temperature and strain distributions at the end of the hot run phase, which in turn leads to different temperature ranges, strain ranges, and residual strains during one cycle. With the work cycle proceeding, the circumferential localization effect of the residual strain would be significantly enhanced. A post-processing damage analysis reveals that the low-cycle fatigue damage accumulated in each cycle is almost unchanged, while the quasi static damage accumulated in a considered cycle declines until stabilized after several cycles. The maximum discrepancy of the predicted lives between different cooling channels is about 30%.
文摘This paper presents a method of thermal state calculation of combustion chamber in small thrust liquid rocket engine. The goal is to predict the thermal state of chamber wall by using basic parameters of engine: thrust level, propellants, chamber pressure, injection pattern, film cooling parameters, material of wall and their coating, etc. The difficulties in modeling the startup and shutdown processes of thrusters lie in the fact that there are the conjugated physical processes occurring at various parameters for non-design conditions. A mathematical model to predict the thermal state of the combustion chamber for different engine operation modes is developed. To simulate the startup and shutdown processes, a quasi-steady approach is applied by replacing the transient process with time-variant operating parameters of steady-state processes. The mathematical model is based on several principles and data commonly used for heat transfer modeling: geometry of flow part, gas dynamics of flow, thermodynamics of propellants and combustion spices, convective and radiation heat flows, conjugated heat transfer between hot gas and wall, and transient approach for calculation of thermal state of construction. Calculations of the thermal state of the combustion chamber in single-turn-on mode show good convergence with the experimental results. The results of pulsed modes indicate a large temperature gradient on the internal wall surface of the chamber between pulses and the thermal state of the wall strongly depends on the pulse duration and the interval.
文摘In order to measure the instantaneous thrust of a certain attitude-control solid rocket motor, based on the analysis of the measurement principles, the difference between the instantaneous thrust and steady thrust measurements is pointed out. According to the measurement characteristics, a dynamic digital filter compensation method is presented. Combined the identification-modeling, dynamic compensation and simulation, the system's dynamic mathematic model is established. And then, a compensation digital filter is also designed. Thus, the dynamic response of the system is improved and the instantaneous thrust measurement can be implemented. The measurement results for the rocket motor show that the digital filter compensation is effective in the instantaneous thrust measurement.
文摘The solid rocket motor driven system is one of the common ways for submarines to launch underwater missiles. It has significant advantages in improving the missile’s water exit speed, anti-interference capability, and enemy striking power. The prediction of the underwater loading is a preliminary factor for the power system design of the underwater vehicle. This paper presents a rapid prediction method and validated by the experimental study for the underwater thrust of the solid rocket motor. Based on the potential flow assumption of the water field, a model of the bubble and a one-dimensional quasi-steady model of the nozzle are established to directly solve the flow status of the nozzle. The aerodynamic thrust and hydrodynamic thrust have been calculated and analyzed. The calculation results are within 5% error of the experimental results. Moreover, a design platform to predict the underwater thrust of the solid rocket motor has been developed based on Python and the PyQt library, which shows excellent system adaptability and computational efficiency.
文摘To evaluate the structural failure risk of the regenerative cooling thrust chamber cylinder segment,a Finite Element Method(FEM)based on experimental data was developed.The methodology was validated and utilized to reveal the thermal response and the nonlinear deformation behavior of the cylinder segment phase by phase.The conclusions of the research are as follows:The 2 D heat flux distribution caused by the injector determines the uneven temperature distribution on the gas-side wall and leads to the temperature disparity between various cooling channels;The reason for the accumulation of residual strain is that the tensile strain generated in the post-cooling phase is greater than the compressive strain produced in the hot run phase;Through the single-cycle simulation,two potential failure locations with conspicuous deformations were found,but it is difficult to determine which point is more dangerous.However,the multi-cycle thermo-structural analysis gives the evolution of the stress-strain curve and gradually discloses that the low-temperature corner of a particular channel is the most likely location to fail,rather than the maximum residual strain point of the gas-side wall.The damage analysis for dangerous point indicates that the quasistatic damage accounts for the majority of the total damage and is the main factor limiting the service life.
文摘On August 2,a twin-segment solid rocket motor of the largest diameter,grain mass and thrust in China completed its ground test firing with success.The3 m solid motor was independently developed by the Academy of Aerospace Solid Propulsion Technology（AASPT）under CASC.
文摘A hybrid rocket motor combines components from both solid fuel and liquid fuel rocket motors. The fuel itself is a solid grain, (often paraffin or hydroxyl-terminated polybutadiene, known as HTPB) while the oxidizing agent is liquid (often hydrogen peroxide or liquid oxygen). These components are combined in the fuel chamber which doubles as the combustion chamber for the hybrid motor. This review looks at the advances in techniques that have taken place in the development of these motors since 1995. Methods of testing the thrust from rocket motors and of measuring the rocket plume spectroscopically for combustion reaction products have been developed. These assessments allow researchers to more completely understand the effects of additives and physical changes in design, in terms of regression rates and thrust developed. Hybrid rocket motors have been used or tested in many areas of rocketry, including tactical rockets and large launch vehicles. Several additives have shown significant improvements in regression rates and thrust, including Guanidinium azotetrazolate (GAT), and various Aluminum alloys. The most recent discoveries have come from research into nano-particle additives. The nano-particles have been shown to provide enhancements to many parameters of hybrid rocket function, and research into specific areas continues in the sub-field of nano-additives for fuel grains.
文摘This study focuses on the stepwise procedure involved in the development of a numerical model of a bi-propellant hypergolic chemical propulsion system using key features and performance characteristics of existing and planned (near future) propulsion systems. The study targets specific impulse of 100</span></span><span><span><span style="font-family:""> </span></span></span><span style="font-family:Verdana;"><span style="font-family:Verdana;"><span style="font-family:Verdana;">N delivery performance of thrust chambers which is suitable for primary propulsion and attitude control for spacecraft. Results from numerical models are reported and validated with the Rocket Propulsion Analysis (RPA) computation concept. In the modelling process, there was proper consideration for the essential parts of the thruster engine such as the nozzle, combustion chamber, catalyst bed, injector, and cooling jacket. This propulsion system is designed to be fabricated in our next step in advancing this idea, using a combination of additive manufacturing technology and commercial off the shelf (COTS) parts along with non-toxic propellants. The two non-toxic propellants being considered are Hydrogen Peroxide as the oxidiser and Kerosene as the fuel, thus making it a low-cost, readily available and environmentally-friendly option for future microsatellite missions.
