In liquid rocket engines,regenerative cooling technology is essential for preserving structural integrity under extreme thermal loads.However,non-uniform coolant flow distribution within the cooling channels often lea...In liquid rocket engines,regenerative cooling technology is essential for preserving structural integrity under extreme thermal loads.However,non-uniform coolant flow distribution within the cooling channels often leads to localized overheating,posing serious risks to engine reliability and operational lifespan.This study employs a three-dimensional fluid–thermal coupled numerical model to systematically investigate the influence of geometric parameters-specifically the number of inlets,the number of channels,and inlet manifold configurations-on flow uniformity and thermal distribution in non-pyrolysis zones.Key findings reveal that increasing the number of inlets from one to three significantly enhances flow uniformity,reducing mass flow rate deviation from 1.2%to below 0.3%.However,further increasing the inlets to five yields only marginal improvements indicating diminishing(<0.1%),returns beyond three inlets.Additionally,temperature non-uniformity at the combustion chamber throat decreases by 37%-from 3050 K with 18 channels to 1915 K with 30 channels-highlighting the critical role of channel density in effective thermal regulation.Notably,while higher channel counts improve cooling efficiency,they also result in increased pressure losses of approximately 18%–22%,emphasizing the need to balance thermal performance against hydraulic resistance.An optimal configuration comprising 24 channels and three inlets was identified,providing minimal temperature gradients while maintaining acceptable pressure losses.The inlet manifold structure also plays a pivotal role in determining flow distribution.Configuration 3(Config-3),which features an enlarged manifold and reduced inlet velocity,achieves a 40%reduction in velocity fluctuations compared to Configuration 1(Config-1).This improvement leads to a more uniform mass flow distribution,with a relative standard deviation(RSD)of less than 0.15%.Furthermore,this design effectively mitigates localized hot spots near the nozzle-where temperature gradients are most severe-achieving a reduction of approximately 1135 K.展开更多
This paper provides a detailed introduction to and analysis of the course of China's technological innovation in liquid hydrogen/liquid oxygen(LH2/LOX)rocket engines from a historical point of view.It starts with ...This paper provides a detailed introduction to and analysis of the course of China's technological innovation in liquid hydrogen/liquid oxygen(LH2/LOX)rocket engines from a historical point of view.It starts with the investigation of LH2/LOX rocket engines by relevant departments of the Chinese Academy of Sciences in the 1960s and their preliminary achievements.Then,the policy decision concerning LH2/LOX engine development,the project approval of the Long March-3(Chang Zheng-3,CZ-3)rocket,and the process of developing LH2/LOX engines are analyzed in detail,followed by an introduction to and summary of the development situation and technical innovation characteristics of China's LH2/LOX engines as they grew from 4 tons to 8 tons,and finally to 50 tons.Finally,the paper briefly analyzes the innovation experience connected with China's LH2/LOX engines.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
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
Liquid propellant rocket engines for a launch vehicle are an essential aerospace technology, representing the advanced level of hi-tech in a country. In recent years, China’s aerospace industry has made remarkable ac...Liquid propellant rocket engines for a launch vehicle are an essential aerospace technology, representing the advanced level of hi-tech in a country. In recent years, China’s aerospace industry has made remarkable achievements, and liquid rocket engine technology has also been effectively developed. In this article, the development processes of China’s liquid rocket engines are discussed. Then, the performance features of China’s new generation liquid rocket engines as well as the flight tests of the new-generation launch vehicles are introduced. Finally, the development direction and the most recent progress of the next generation large-thrust liquid rocket engine is presented.展开更多
Based on current research,the development trend of reusable liquid rocket engines was analyzed.Key technologies and research focuses of the reusable liquid rocket engine have been analyzed and summarized,and then sugg...Based on current research,the development trend of reusable liquid rocket engines was analyzed.Key technologies and research focuses of the reusable liquid rocket engine have been analyzed and summarized,and then suggestions on the development of future key technologies are proposed.展开更多
The reuse of liquid propellant rocket engines has increased the difficulty of their control and estimation.State and parameter Moving Horizon Estimation(MHE)is an optimization-based strategy that provides the necessar...The reuse of liquid propellant rocket engines has increased the difficulty of their control and estimation.State and parameter Moving Horizon Estimation(MHE)is an optimization-based strategy that provides the necessary information for model predictive control.Despite the many advantages of MHE,long computation time has limited its applications for system-level models of liquid propellant rocket engines.To address this issue,we propose an asynchronous MHE method called advanced-multi-step MHE with Noise Covariance Estimation(amsMHE-NCE).This method computes the MHE problem asynchronously to obtain the states and parameters and can be applied to multi-threaded computations.In the background,the state and covariance estimation optimization problems are computed using multiple sampling times.In real-time,sensitivity is used to quickly approximate state and parameter estimates.A covariance estimation method is developed using sensitivity to avoid redundant MHE problem calculations in case of sensor degradation during engine reuse.The amsMHE-NCE is validated through three cases based on the space shuttle main engine system-level model,and we demonstrate that it can provide more accurate real-time estimates of states and parameters compared to other commonly used estimation methods.展开更多
A genetic algorithm was used to develop optimal design methods for the regenerative cooled combustor and fuel-rich gas generator of a liquid rocket engine. For the combustor design, a chemical equilibrium analysis was...A genetic algorithm was used to develop optimal design methods for the regenerative cooled combustor and fuel-rich gas generator of a liquid rocket engine. For the combustor design, a chemical equilibrium analysis was applied, and the profile was calculated using Rao's method. One-dimensional heat transfer was assumed along the profile, and cooling channels were designed. For the gas-generator design, non-equilibrium properties were derived from a counterflow analysis, and a vaporization model for the fuel droplet was adopted to calculate residence time. Finally, a genetic algorithm was adopted to optimize the designs. The combustor and gas generator were optimally designed for 30-tonf, 75-tonf, and 150-tonf engines. The optimized combustors demonstrated superior design characteristics when compared with previous non-optimized results. Wall temperatures at the nozzle throat were optimized to satisfy the requirement of 800 K, and specific impulses were maximized. In addition, the target turbine power and a burned-gas temperature of 1000 K were obtained from the optimized gas-generator design.展开更多
Heat transfer in the thrust chamber is of great importance in the design of liquid propellant rocketengines. Regenerative cooling is an advanced method which can ensure not only the proper runningbut also higher perfo...Heat transfer in the thrust chamber is of great importance in the design of liquid propellant rocketengines. Regenerative cooling is an advanced method which can ensure not only the proper runningbut also higher performance of a rocket engine. The theoretical model is complicated, it relates to fluiddynamics, heat transfer, combustion, etc... In this papers a regenerative cooling model is presented.Effects such as radiation, heat transfer to environment, variable thermal properties and coking areincluded in the model. This model can be applied to all kinds of liquid propellant rocket engines aswell as similar constructions. The modularized computer code is completed in the work.展开更多
A nonlinear mathematical model of the low-frequency dynamics of the clustered multi-engine rocket propulsion system has been developed and the computations of the engine transient processes during the start-ups of the...A nonlinear mathematical model of the low-frequency dynamics of the clustered multi-engine rocket propulsion system has been developed and the computations of the engine transient processes during the start-ups of the four-engine propulsion system with a shared feed system have been made applied.Based on propulsion system start-up modeling the influence of the connectivity of engines in a cluster on the starting characteristics of individual engines is shown.In particular,an advanced nonlinear mathematical model of the pump cavitation phenomena is a distinctive feature of the mathematical model.The computation results showed that the asynchronous engines start-ups during rocket lift-off lead to severely nonlinear engine transients and clustered engine thrust misbalance.The influence of the rocket engines asynchronous start-ups on the clustered feed system transients depends on many factors,mainly on from the clustered feed system low-frequency dynamics,the magnitude of the disturbance and the phase difference between disturbances acting on different branches of the feed system.The deep lingering dips in the flow rate and pressure transients are possible due to the nonlinear dynamic interaction of the engines.In case of great pressure dips at the pump inlet(up to the pressure of saturated vapors during significant periods of start-up time)the cavitation breakdowns of the pumps of one or more engines from the cluster are possible.This can disrupt the operation of the entire propulsion system and leads to the failure of the launch vehicle mission.展开更多
Improved liquid rocket engine cycles were proposed and analyzed via comparison with existing staged combustion and gas-generator cycles.The key features of the proposed cycles are regenerative cooling of thrust chambe...Improved liquid rocket engine cycles were proposed and analyzed via comparison with existing staged combustion and gas-generator cycles.The key features of the proposed cycles are regenerative cooling of thrust chamber by oxygen and subsequent use of this oxygen for driving one or two oxygen pumps.The fuel pump(s)are driven in a conventional manner,for example,using a fuel-rich gas-generator cycle.Comparison with staged combustion cycle based on oxygen-rich pre-burner showed that one of the proposed semi-expander cycles has a specific impulse only on 0.4%lower while providing much lower oxygen temperature,more efficient tank pressurizing system and built-in roll control.This semi-expander cycle can be considered as a more reliable and cost-effective alternative of staged combustion cycle.Another semi-expander cycle can be considered as an improvement of gas-generator cycle.All proposed semi-expander cycles were developed as a derivative of thrust chamber regenerative cooling performed by oxygen.Analysis of existing oxygen/kerosene engines showed that replacing of kerosene regenerative cooling with oxygen allows a significant increase of achievable specific impulse,via optimization of mixture ratio.It is especially the case for upper stage engines.The increasing of propellants average density can be considered as an additional benefit of mixture ratio optimization.It was demonstrated that oxygen regenerative cooling of thrust chamber is a feasible and the most promising option for oxygen/kerosene engines.Combination of oxygen regenerative cooling and semi-expander cycles potentially allows creating the oxygen/kerosene propulsion systems with minimum specific impulse losses.It is important that such propulsion systems can be fully based on inherited and well-proven technical solutions.A hypothetic upper stage engine with thrust 19.6 kN was chosen as a prospective candidate for theoretical analysis of the proposed semi-expander cycles.The newly-developed software RECS was used for the comparative analysis of engine cycles.展开更多
The thermal protection of rocket engines is a crucial aspect of rocket engine design.In this paper,the gas film/regenerative composite cooling of the liquid oxygen/liquid methane(LOX/LCH4)rocket engine thrust chamber ...The thermal protection of rocket engines is a crucial aspect of rocket engine design.In this paper,the gas film/regenerative composite cooling of the liquid oxygen/liquid methane(LOX/LCH4)rocket engine thrust chamber was investigated.A gas film/regenerative composite cooling model was developed based on the Grisson gas film cooling efficiency formula and the one-dimensional regenerative cooling model.The accuracy of the model was validated through experiments conducted on a 6 kg/s level gas film/regenerative composite cooling thrust chamber.Additionally,key parameters related to heat transfer performance were calculated.The results demonstrate that the model is sufficiently accurate to be used as a preliminary design tool.The temperature rise error of the coolant,when compared with the experimental results,was found to be less than 10%.Although the pressure drop error is relatively large,the calculated results still provide valuable guidance for heat transfer analysis.In addition,the performance of composite cooling is observed to be superior to regenerative cooling.Increasing the gas film flow rate results in higher cooling efficiency and a lower gas-side wall temperature.Furthermore,the position at which the gas film is introduced greatly impacts the cooling performance.The optimal introduction position for the gas film is determined when the film is introduced from a single row of holes.This optimal introduction position results in a more uniform wall temperature distribution and reduces the peak temperature.Lastly,it is observed that a double row of holes,when compared to a single row of holes,enhances the cooling effect in the superposition area of the gas film and further lowers the gas-side wall temperature.These results provide a basis for the design of gas film/regenerative composite cooling systems.展开更多
As the core of the rocket system,the performance and quality of rocket engines are of paramount impor-tance.Currently,the production of aerospace model rocket engines does not differentiate the production and selectio...As the core of the rocket system,the performance and quality of rocket engines are of paramount impor-tance.Currently,the production of aerospace model rocket engines does not differentiate the production and selection of motors according to the importance of the mission,which is insufficient to ensure the high reliability requirements of important launch missions.To select rocket engines with better performance quality for more critical launch missions,this paper uses fuzzy comprehensive evaluation and TOPSIS methods based on the test value or assessment informa-tion of evaluation indicators.The method scientifically and accurately ranks the performance quality of rocket engines,choosing the engines with better performance quality for more strategic missions,and providing technical support for national management decisions.展开更多
Describes a new computer program (Regress-3D) to simulate the regression of complex 3D grain cavity and calculate the burning surface area. It has a large region of applicability in solid rocket motor design and has...Describes a new computer program (Regress-3D) to simulate the regression of complex 3D grain cavity and calculate the burning surface area. It has a large region of applicability in solid rocket motor design and has made new improvements compared with other available codes. User can easily and rapidly build his initial grain shapes and then obtain geometric information of his design. Considering with the calclulting results, redesigning can be performed as desire until reaching at the satisfied result. Advantages and disadvantages of this method are also discussed.展开更多
Reusable rocket engines are the core components of reusable launch vehicles, and have thus become a major focus of aerospace engineering research in recent years. In practice, subsystem design is based on the overall ...Reusable rocket engines are the core components of reusable launch vehicles, and have thus become a major focus of aerospace engineering research in recent years. In practice, subsystem design is based on the overall index allocation of an engine;therefore, a multidisciplinary optimization approach is necessary. In this study, design of a reusable methane/liquid oxygen(LOX/CH4)rocket engine with a gas generator cycle was investigated using multidisciplinary optimization. Two parameters were chosen as design variables: pressure and fuel mix ratio of the main combustion chamber. Optimization objectives were specific impulse, structural mass, and life cycle cost of the reusable rocket engine, and constraints were assigned to each discipline according to rocket design requirements. Then, an optimization model was developed, and optimal design parameters were acquired for the LOX/CH4 rocket engine. The proposed method is effective for designing the index allocation of reusable rocket engines and takes into account the multidisciplinary nature of complex systems.展开更多
The pulse detonation rocket engine （PDRE） requires periodic supply of oxidizer, fuel and purge gas. A rotary-valve assembly is fabricated to control the periodic supply in this research. Oxygen and liquid aviation k...The pulse detonation rocket engine （PDRE） requires periodic supply of oxidizer, fuel and purge gas. A rotary-valve assembly is fabricated to control the periodic supply in this research. Oxygen and liquid aviation kerosene are used as oxidizer and fuel respectively. An ordinary automobile spark plug, with ignition energy as low as 50 mJ, is used to initiate combustion. Steady operation of the PDRE is achieved with operating frequency ranging from 1 Hz to 10 Hz. Experimentally measured pressure is lower than theoretical value by 13% at 1 Hz and 37% at 10 Hz, and there also exists a velocity deficit at different operating frequencies. Both of these two phenomena are believed mainly due to droplet size which depends on atomization and vaporiza-tion of liquid fuel.展开更多
The PDRE test model used in these experiments utilized kerosene as the fuel, oxygen as oxidizer, and nitrogen as purge gas. The solenoid valves were employed to control intermittent supplies of kerosene, oxygen and pu...The PDRE test model used in these experiments utilized kerosene as the fuel, oxygen as oxidizer, and nitrogen as purge gas. The solenoid valves were employed to control intermittent supplies of kerosene, oxygen and purge gas. PDRE test model was 50 mm in inner diameter by 1.2 m long. The DDT （deflagration to detonation transition） enhancement device Shchelkin spiral was used in the test model. The effects of detonation frequency on its time-averaged thrust and specific impulse were experimentally investigated. The obtained results showes that the time-averaged thrust of PDRE test model was approximately proportional to the detonation frequency. For the detonation frequency 20 Hz, the time-averaged thrust was around 107 N, and the specific impulse was around 125 s. The nozzle experiments were conducted using PDRE test model with three traditional nozzles. The experimental results obtained demonstrated that all of those nozzles could augment the thrust and specific impulse. Among those three nozzles, the convergent nozzle had the largest increased augmentation, which was approximately 18%, under the specific condition of the experiment.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 predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray, the conventional uncoupled spray model for impinging injectors is extended by considering the coupling of the j...To predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray, the conventional uncoupled spray model for impinging injectors is extended by considering the coupling of the jet impingement process and the ambient gas field. The new coupled model consists of the plain-orifice sub-model, the jet-jet impingement sub-model and the droplet collision sub-model. The parameters of the child droplet are determined with the jet-jet impingement sub-model using correlations about the liquid jet parameters and the chamber conditions.The overall model is benchmarked under various impingement angles, jet momentum and offcenter ratios. Agreement with the published experimental data validates the ability of the model to predict the key spray characteristics, such as the mass flux and mixture ratio distributions in quiescent air. Besides, impinging sprays under changing ambient pressure and non-uniform gas flow are investigated to explore the effect of liquid rocket engine chamber conditions. First, a transient impingement spray during engine start-up phase is simulated with prescribed pressure profile. The minimum average droplet diameter is achieved when the orifices work in cavitation state, and is about 30% smaller than the steady single phase state. Second, the effect of non-uniform gas flow produces off-center impingement and the rotated spray fan by 38°. The proposed model suggests more reasonable impingement spray characteristics than the uncoupled one and can be used as the first step in the complex simulation of coupling impingement spray and combustion in liquid rocket engines.展开更多
Numerical simulation has been done for liquid film cooling in liquid rocket combustor.Multiple species of axial Navier-Stokes equations have been solved for liquid-film / hot-gas flow field,and k-εequations have been...Numerical simulation has been done for liquid film cooling in liquid rocket combustor.Multiple species of axial Navier-Stokes equations have been solved for liquid-film / hot-gas flow field,and k-εequations have been used for compressible turbulent flow.The results of the model agree well with the results of software FLUENT.The results show that :(1) Liquid film can decrease the wall heat flux and temperature effectively,and the cold border area formed by the film covers the whole combustor and nozzle wall.(2) The turbulent viscosity is higher than the physical viscosity,and its biggest value is in the border area of the convergent area in nozzle.The effect of turbulent flow on the whole simulation field can not be ignored.(3) The mass fraction of kerosene at the film inlet is 1,but it decreases along the nozzle wall and achieves its lowest value at the outlet.However,the mass fraction of kerosene near the wall is the biggest at any axial location.展开更多
基金supported by the Key project of Hunan Provincial Education Department(Grant Number:22A0485)The Natural Science Foundation of Hunan(Grant Number:2024JJ5293)The Key project of Hunan University of Arts and Science(Grant Number:23ZZ08).
文摘In liquid rocket engines,regenerative cooling technology is essential for preserving structural integrity under extreme thermal loads.However,non-uniform coolant flow distribution within the cooling channels often leads to localized overheating,posing serious risks to engine reliability and operational lifespan.This study employs a three-dimensional fluid–thermal coupled numerical model to systematically investigate the influence of geometric parameters-specifically the number of inlets,the number of channels,and inlet manifold configurations-on flow uniformity and thermal distribution in non-pyrolysis zones.Key findings reveal that increasing the number of inlets from one to three significantly enhances flow uniformity,reducing mass flow rate deviation from 1.2%to below 0.3%.However,further increasing the inlets to five yields only marginal improvements indicating diminishing(<0.1%),returns beyond three inlets.Additionally,temperature non-uniformity at the combustion chamber throat decreases by 37%-from 3050 K with 18 channels to 1915 K with 30 channels-highlighting the critical role of channel density in effective thermal regulation.Notably,while higher channel counts improve cooling efficiency,they also result in increased pressure losses of approximately 18%–22%,emphasizing the need to balance thermal performance against hydraulic resistance.An optimal configuration comprising 24 channels and three inlets was identified,providing minimal temperature gradients while maintaining acceptable pressure losses.The inlet manifold structure also plays a pivotal role in determining flow distribution.Configuration 3(Config-3),which features an enlarged manifold and reduced inlet velocity,achieves a 40%reduction in velocity fluctuations compared to Configuration 1(Config-1).This improvement leads to a more uniform mass flow distribution,with a relative standard deviation(RSD)of less than 0.15%.Furthermore,this design effectively mitigates localized hot spots near the nozzle-where temperature gradients are most severe-achieving a reduction of approximately 1135 K.
基金Projects funded by the National Social Science Fund (19BDJ064)
文摘This paper provides a detailed introduction to and analysis of the course of China's technological innovation in liquid hydrogen/liquid oxygen(LH2/LOX)rocket engines from a historical point of view.It starts with the investigation of LH2/LOX rocket engines by relevant departments of the Chinese Academy of Sciences in the 1960s and their preliminary achievements.Then,the policy decision concerning LH2/LOX engine development,the project approval of the Long March-3(Chang Zheng-3,CZ-3)rocket,and the process of developing LH2/LOX engines are analyzed in detail,followed by an introduction to and summary of the development situation and technical innovation characteristics of China's LH2/LOX engines as they grew from 4 tons to 8 tons,and finally to 50 tons.Finally,the paper briefly analyzes the innovation experience connected with China's LH2/LOX engines.
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.
文摘Liquid propellant rocket engines for a launch vehicle are an essential aerospace technology, representing the advanced level of hi-tech in a country. In recent years, China’s aerospace industry has made remarkable achievements, and liquid rocket engine technology has also been effectively developed. In this article, the development processes of China’s liquid rocket engines are discussed. Then, the performance features of China’s new generation liquid rocket engines as well as the flight tests of the new-generation launch vehicles are introduced. Finally, the development direction and the most recent progress of the next generation large-thrust liquid rocket engine is presented.
文摘Based on current research,the development trend of reusable liquid rocket engines was analyzed.Key technologies and research focuses of the reusable liquid rocket engine have been analyzed and summarized,and then suggestions on the development of future key technologies are proposed.
基金supported by the National Natural Science Foundation of China(Nos.62120106003 and 62173301)。
文摘The reuse of liquid propellant rocket engines has increased the difficulty of their control and estimation.State and parameter Moving Horizon Estimation(MHE)is an optimization-based strategy that provides the necessary information for model predictive control.Despite the many advantages of MHE,long computation time has limited its applications for system-level models of liquid propellant rocket engines.To address this issue,we propose an asynchronous MHE method called advanced-multi-step MHE with Noise Covariance Estimation(amsMHE-NCE).This method computes the MHE problem asynchronously to obtain the states and parameters and can be applied to multi-threaded computations.In the background,the state and covariance estimation optimization problems are computed using multiple sampling times.In real-time,sensitivity is used to quickly approximate state and parameter estimates.A covariance estimation method is developed using sensitivity to avoid redundant MHE problem calculations in case of sensor degradation during engine reuse.The amsMHE-NCE is validated through three cases based on the space shuttle main engine system-level model,and we demonstrate that it can provide more accurate real-time estimates of states and parameters compared to other commonly used estimation methods.
基金supported by the National Research Foundation of Korea grant funded by the Korean Government(MSIP)NRF-2012M1A3A3A02033146 and NRF-2013M1A3A3A02042434
文摘A genetic algorithm was used to develop optimal design methods for the regenerative cooled combustor and fuel-rich gas generator of a liquid rocket engine. For the combustor design, a chemical equilibrium analysis was applied, and the profile was calculated using Rao's method. One-dimensional heat transfer was assumed along the profile, and cooling channels were designed. For the gas-generator design, non-equilibrium properties were derived from a counterflow analysis, and a vaporization model for the fuel droplet was adopted to calculate residence time. Finally, a genetic algorithm was adopted to optimize the designs. The combustor and gas generator were optimally designed for 30-tonf, 75-tonf, and 150-tonf engines. The optimized combustors demonstrated superior design characteristics when compared with previous non-optimized results. Wall temperatures at the nozzle throat were optimized to satisfy the requirement of 800 K, and specific impulses were maximized. In addition, the target turbine power and a burned-gas temperature of 1000 K were obtained from the optimized gas-generator design.
文摘Heat transfer in the thrust chamber is of great importance in the design of liquid propellant rocketengines. Regenerative cooling is an advanced method which can ensure not only the proper runningbut also higher performance of a rocket engine. The theoretical model is complicated, it relates to fluiddynamics, heat transfer, combustion, etc... In this papers a regenerative cooling model is presented.Effects such as radiation, heat transfer to environment, variable thermal properties and coking areincluded in the model. This model can be applied to all kinds of liquid propellant rocket engines aswell as similar constructions. The modularized computer code is completed in the work.
文摘A nonlinear mathematical model of the low-frequency dynamics of the clustered multi-engine rocket propulsion system has been developed and the computations of the engine transient processes during the start-ups of the four-engine propulsion system with a shared feed system have been made applied.Based on propulsion system start-up modeling the influence of the connectivity of engines in a cluster on the starting characteristics of individual engines is shown.In particular,an advanced nonlinear mathematical model of the pump cavitation phenomena is a distinctive feature of the mathematical model.The computation results showed that the asynchronous engines start-ups during rocket lift-off lead to severely nonlinear engine transients and clustered engine thrust misbalance.The influence of the rocket engines asynchronous start-ups on the clustered feed system transients depends on many factors,mainly on from the clustered feed system low-frequency dynamics,the magnitude of the disturbance and the phase difference between disturbances acting on different branches of the feed system.The deep lingering dips in the flow rate and pressure transients are possible due to the nonlinear dynamic interaction of the engines.In case of great pressure dips at the pump inlet(up to the pressure of saturated vapors during significant periods of start-up time)the cavitation breakdowns of the pumps of one or more engines from the cluster are possible.This can disrupt the operation of the entire propulsion system and leads to the failure of the launch vehicle mission.
文摘Improved liquid rocket engine cycles were proposed and analyzed via comparison with existing staged combustion and gas-generator cycles.The key features of the proposed cycles are regenerative cooling of thrust chamber by oxygen and subsequent use of this oxygen for driving one or two oxygen pumps.The fuel pump(s)are driven in a conventional manner,for example,using a fuel-rich gas-generator cycle.Comparison with staged combustion cycle based on oxygen-rich pre-burner showed that one of the proposed semi-expander cycles has a specific impulse only on 0.4%lower while providing much lower oxygen temperature,more efficient tank pressurizing system and built-in roll control.This semi-expander cycle can be considered as a more reliable and cost-effective alternative of staged combustion cycle.Another semi-expander cycle can be considered as an improvement of gas-generator cycle.All proposed semi-expander cycles were developed as a derivative of thrust chamber regenerative cooling performed by oxygen.Analysis of existing oxygen/kerosene engines showed that replacing of kerosene regenerative cooling with oxygen allows a significant increase of achievable specific impulse,via optimization of mixture ratio.It is especially the case for upper stage engines.The increasing of propellants average density can be considered as an additional benefit of mixture ratio optimization.It was demonstrated that oxygen regenerative cooling of thrust chamber is a feasible and the most promising option for oxygen/kerosene engines.Combination of oxygen regenerative cooling and semi-expander cycles potentially allows creating the oxygen/kerosene propulsion systems with minimum specific impulse losses.It is important that such propulsion systems can be fully based on inherited and well-proven technical solutions.A hypothetic upper stage engine with thrust 19.6 kN was chosen as a prospective candidate for theoretical analysis of the proposed semi-expander cycles.The newly-developed software RECS was used for the comparative analysis of engine cycles.
基金supported by the National Science Fund Project(No.2019-JCJQ-ZQ-019)the Innovative Research Group Project of National Natural Science Foundation of China(No.T2221002).
文摘The thermal protection of rocket engines is a crucial aspect of rocket engine design.In this paper,the gas film/regenerative composite cooling of the liquid oxygen/liquid methane(LOX/LCH4)rocket engine thrust chamber was investigated.A gas film/regenerative composite cooling model was developed based on the Grisson gas film cooling efficiency formula and the one-dimensional regenerative cooling model.The accuracy of the model was validated through experiments conducted on a 6 kg/s level gas film/regenerative composite cooling thrust chamber.Additionally,key parameters related to heat transfer performance were calculated.The results demonstrate that the model is sufficiently accurate to be used as a preliminary design tool.The temperature rise error of the coolant,when compared with the experimental results,was found to be less than 10%.Although the pressure drop error is relatively large,the calculated results still provide valuable guidance for heat transfer analysis.In addition,the performance of composite cooling is observed to be superior to regenerative cooling.Increasing the gas film flow rate results in higher cooling efficiency and a lower gas-side wall temperature.Furthermore,the position at which the gas film is introduced greatly impacts the cooling performance.The optimal introduction position for the gas film is determined when the film is introduced from a single row of holes.This optimal introduction position results in a more uniform wall temperature distribution and reduces the peak temperature.Lastly,it is observed that a double row of holes,when compared to a single row of holes,enhances the cooling effect in the superposition area of the gas film and further lowers the gas-side wall temperature.These results provide a basis for the design of gas film/regenerative composite cooling systems.
文摘As the core of the rocket system,the performance and quality of rocket engines are of paramount impor-tance.Currently,the production of aerospace model rocket engines does not differentiate the production and selection of motors according to the importance of the mission,which is insufficient to ensure the high reliability requirements of important launch missions.To select rocket engines with better performance quality for more critical launch missions,this paper uses fuzzy comprehensive evaluation and TOPSIS methods based on the test value or assessment informa-tion of evaluation indicators.The method scientifically and accurately ranks the performance quality of rocket engines,choosing the engines with better performance quality for more strategic missions,and providing technical support for national management decisions.
文摘Describes a new computer program (Regress-3D) to simulate the regression of complex 3D grain cavity and calculate the burning surface area. It has a large region of applicability in solid rocket motor design and has made new improvements compared with other available codes. User can easily and rapidly build his initial grain shapes and then obtain geometric information of his design. Considering with the calclulting results, redesigning can be performed as desire until reaching at the satisfied result. Advantages and disadvantages of this method are also discussed.
文摘Reusable rocket engines are the core components of reusable launch vehicles, and have thus become a major focus of aerospace engineering research in recent years. In practice, subsystem design is based on the overall index allocation of an engine;therefore, a multidisciplinary optimization approach is necessary. In this study, design of a reusable methane/liquid oxygen(LOX/CH4)rocket engine with a gas generator cycle was investigated using multidisciplinary optimization. Two parameters were chosen as design variables: pressure and fuel mix ratio of the main combustion chamber. Optimization objectives were specific impulse, structural mass, and life cycle cost of the reusable rocket engine, and constraints were assigned to each discipline according to rocket design requirements. Then, an optimization model was developed, and optimal design parameters were acquired for the LOX/CH4 rocket engine. The proposed method is effective for designing the index allocation of reusable rocket engines and takes into account the multidisciplinary nature of complex systems.
基金National Natural Science Foundation of China (50976094)Doctoral Program Foundation of Education Ministry of China (20096102110022)+1 种基金Doctorate Foundation of Northwestern Polytechnical University (CX201112)Graduate Innovation Lab Center of Northwestern Polytechnical University (10006,10013)
文摘The pulse detonation rocket engine （PDRE） requires periodic supply of oxidizer, fuel and purge gas. A rotary-valve assembly is fabricated to control the periodic supply in this research. Oxygen and liquid aviation kerosene are used as oxidizer and fuel respectively. An ordinary automobile spark plug, with ignition energy as low as 50 mJ, is used to initiate combustion. Steady operation of the PDRE is achieved with operating frequency ranging from 1 Hz to 10 Hz. Experimentally measured pressure is lower than theoretical value by 13% at 1 Hz and 37% at 10 Hz, and there also exists a velocity deficit at different operating frequencies. Both of these two phenomena are believed mainly due to droplet size which depends on atomization and vaporiza-tion of liquid fuel.
基金National Natural Science Foundation of China (501- 06012, 50336030)Program for New Century Excellent Talents in Uni-versity(NCET-04-0960)
文摘The PDRE test model used in these experiments utilized kerosene as the fuel, oxygen as oxidizer, and nitrogen as purge gas. The solenoid valves were employed to control intermittent supplies of kerosene, oxygen and purge gas. PDRE test model was 50 mm in inner diameter by 1.2 m long. The DDT （deflagration to detonation transition） enhancement device Shchelkin spiral was used in the test model. The effects of detonation frequency on its time-averaged thrust and specific impulse were experimentally investigated. The obtained results showes that the time-averaged thrust of PDRE test model was approximately proportional to the detonation frequency. For the detonation frequency 20 Hz, the time-averaged thrust was around 107 N, and the specific impulse was around 125 s. The nozzle experiments were conducted using PDRE test model with three traditional nozzles. The experimental results obtained demonstrated that all of those nozzles could augment the thrust and specific impulse. Among those three nozzles, the convergent nozzle had the largest increased augmentation, which was approximately 18%, under the specific condition of the experiment.
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 predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray, the conventional uncoupled spray model for impinging injectors is extended by considering the coupling of the jet impingement process and the ambient gas field. The new coupled model consists of the plain-orifice sub-model, the jet-jet impingement sub-model and the droplet collision sub-model. The parameters of the child droplet are determined with the jet-jet impingement sub-model using correlations about the liquid jet parameters and the chamber conditions.The overall model is benchmarked under various impingement angles, jet momentum and offcenter ratios. Agreement with the published experimental data validates the ability of the model to predict the key spray characteristics, such as the mass flux and mixture ratio distributions in quiescent air. Besides, impinging sprays under changing ambient pressure and non-uniform gas flow are investigated to explore the effect of liquid rocket engine chamber conditions. First, a transient impingement spray during engine start-up phase is simulated with prescribed pressure profile. The minimum average droplet diameter is achieved when the orifices work in cavitation state, and is about 30% smaller than the steady single phase state. Second, the effect of non-uniform gas flow produces off-center impingement and the rotated spray fan by 38°. The proposed model suggests more reasonable impingement spray characteristics than the uncoupled one and can be used as the first step in the complex simulation of coupling impingement spray and combustion in liquid rocket engines.
文摘Numerical simulation has been done for liquid film cooling in liquid rocket combustor.Multiple species of axial Navier-Stokes equations have been solved for liquid-film / hot-gas flow field,and k-εequations have been used for compressible turbulent flow.The results of the model agree well with the results of software FLUENT.The results show that :(1) Liquid film can decrease the wall heat flux and temperature effectively,and the cold border area formed by the film covers the whole combustor and nozzle wall.(2) The turbulent viscosity is higher than the physical viscosity,and its biggest value is in the border area of the convergent area in nozzle.The effect of turbulent flow on the whole simulation field can not be ignored.(3) The mass fraction of kerosene at the film inlet is 1,but it decreases along the nozzle wall and achieves its lowest value at the outlet.However,the mass fraction of kerosene near the wall is the biggest at any axial location.
