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
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
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 ablative material is supposed to be one of good candidates for LRE （liquid rocket engine） combustion chamber to achieve both high reliability and low cost and a numerical analysis for the ablator is considered t...The ablative material is supposed to be one of good candidates for LRE （liquid rocket engine） combustion chamber to achieve both high reliability and low cost and a numerical analysis for the ablator is considered to be a potentially efficient tool to reduce cost as well. So far, ablators have been successfully applied for many SRM （solid rocket motors）, but the application to LRE is still quite limited in Japan. The authors believe that this is primarily because of the unpredictable nature of the heat load from combustion gases to the combustor wall. Indeed, reliable thermal design of ablative combustion chamber, namely reliable prediction of thermal performance, needs both reliable heat load model and reliable ablator response model. This paper elaborates our research activities and our recent research findings.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
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
Turbulent two-phase reacting flow in the chamber of LOX/RP-1 bipropellant liquid rocket engine is numerically investigated in this paper. The predicted pressure and mean axial velocity are qualitatively consistent wit...Turbulent two-phase reacting flow in the chamber of LOX/RP-1 bipropellant liquid rocket engine is numerically investigated in this paper. The predicted pressure and mean axial velocity are qualitatively consistent with the experimental measurements. The self-excited pressure oscillations are obtained without any disturbance introduced through the initial and boundary conditions. It is found that amount of abrupt pressure peaks appear frequently and stochastically in the head regions of the chamber, which are the important sources to drive and strengthen combustion instability. Such abrupt pressures are induced by local constant volume combustion, because local combustible gas mixtures with high temperature are formed and burnt out suddenly due to some fuel droplets reaching their critical state in a rich oxygen surrounding. A third Damkhler number is defined as the ratio of the characteristic time of a chemical reaction to the characteristic time of a pressure wave expansion to measure the relative intensity of acoustic propagation and combustion process in thrusters. The analysis of the third Damkhler number distributions in the whole thrust chamber shows that local constant volume combustion happens in the head regions, while constant pressure combustion presents in the downstream regions. It is found that the combustion instability occurs in the head regions within about 30 mm from the thruster head.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
A design method for a kerosene fuel-rich gas-generator of a liquid rocket engine using turbopumps to supply propellant was performed at a conceptual level. The gas-generator creates hot gases, enabling the turbine to ...A design method for a kerosene fuel-rich gas-generator of a liquid rocket engine using turbopumps to supply propellant was performed at a conceptual level. The gas-generator creates hot gases, enabling the turbine to operate the turbopumps. A chemical non-equilibrium analysis and a droplet vaporization model were used for the estimation of the burnt gas properties and characteristic chamber length. A premixed counter-flow flame analysis was performed for the prediction of the burnt gas properties, namely the temperature, the specific heat ratio and heat capacity, and the chemical reaction time. To predict the vaporization time, the Spalding model, using a single droplet in convective condition, was used. The minimum residence time in the chamber and the characteristic length were calculated by adding the reaction time and the vaporization time. Using the characteristic length, the design methods for the fuel-rich gas-generator were established. Finally, a parametric study was achieved for the effects of the O/F ratio, mass flow rate, chamber pressure, initial droplet temperature, initial droplet diameter and initial droplet velocity.展开更多
The pintle injector used for a liquid rocket engine is a newly re-attracted injection system famous for its wide throttle ability with high efficiency. The pintle injector has many variations with complex inner struct...The pintle injector used for a liquid rocket engine is a newly re-attracted injection system famous for its wide throttle ability with high efficiency. The pintle injector has many variations with complex inner structures due to its moving parts. In order to study the rotating flow near the injector tip, which was observed from the cold flow experiment using water and air, a numerical simulation was adopted and a verification of the numerical model was later conducted. For the verification process, three types of experimental data including velocity distributions of gas flows, spray angles and liquid distribution were all compared using simulated results. The numerical simulation was performed using a commercial simulation program with the Eulerian multiphase model and axisymmetric two dimensional grids. The maximum and minimum velocities of gas were within the acceptable range of agreement, however, the spray angles experienced up to 25% error when the momentum ratios were increased. The spray density distributions were quantitatively measured and had good agreement. As a result of this study, it was concluded that the simulation method was properly constructed to study specific flow characteristics of the pintle injector despite having the limitations of two dimensional and coarse grids.展开更多
An integrated program was established to design a combustor for a liquid rocket engine and to analyze regenerative cooling results on a preliminary design level.Properties of burnt gas from a kerosene-LOx mixture in t...An integrated program was established to design a combustor for a liquid rocket engine and to analyze regenerative cooling results on a preliminary design level.Properties of burnt gas from a kerosene-LOx mixture in the combustor and rocket performance were calculated from CEA which is the code for the calculation of chemical equilibrium.The heat transfer of regenerative cooling was analyzed by using SUPERTRAPP code for coolant properties and by one-dimensional correlations of the heat transfer coefficient from the combustor liner to the coolant.Profiles of the combustors of F-1 and RS-27A engines were designed from similar input data and the present results were compared to actual data for validation.Finally,the combustors of 30 tonf class,75 tonf class and 150 tonf class were designed from the required thrust,combustion chamber,exit pressure and mixture ratio of propellants.The wall temperature,heat flux and pressure drop were calculated for heat transfer analysis of regenerative cooling using the profiles.展开更多
In this paper, transient phenomenon during start up process of a pump fed liquidrocket engine is investigated through numerical simulation. The engine studied in this workis designed such that engine systems are not w...In this paper, transient phenomenon during start up process of a pump fed liquidrocket engine is investigated through numerical simulation. The engine studied in this workis designed such that engine systems are not wetted with propellant until the engine is com-manded to start. This is achieved by positioning the valves for propellant admission at the inter-face of test stand/flight stage and the engine. To evaluate engine performance during starttransient for such systems, unsteady flow simulation was conducted using Method of Charac-teristics and equations for priming. The same has been reported in this work. The results indi-cated a brief period of abrupt pressure rise at pump upstream after opening of the propellantadmission valves, during the process of priming of engine systems at valve downstream.The peak pressure obtained was significantly higher than the propellant tank pressure as wellas the steady state pump suction pressure. The transitory pressure rise was found to occurdue to flow resistance at impeller inlet caused by formation of a forced vortex for orientingthe flow through impeller blades during off design transient regime. The maximum pressureat pump upstream, as computed from start transient simulation, was used as a design inputfor pump inlet feed lines. The engine was realized and subsequently qualified in a ground test facility. Hot test data obtained for pressure and flow rate during transient regime were found tobe in good agreement with the simulation results.展开更多
During the experiment of gas/liquid coaxial swirl injector conducted with air and water under atmosphere environment, it is observed that the injector may selfoscillate. The self oscillation periodically occurs and va...During the experiment of gas/liquid coaxial swirl injector conducted with air and water under atmosphere environment, it is observed that the injector may selfoscillate. The self oscillation periodically occurs and vanishes with the increasing velocity of the gas flow.A theoretical model is presented based on the experiment investigation. Simulation of the acoustic process has been performed and conclusions consistent with the experiment can be drawn from the theoretical model, which explains the exPeriment phenomena quite well. At last, the comparison between phenomena of the self oscillation and some experiments of LRE indicates that some instability phenomena in oxygen/hydrogen propellant rocket engine may be the related to self oscillation in coaxial injectors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
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
Combustion process inside kerosene-GOx rocket combustor with kerosene Alm cooling is studied,and a modeling approach is proposed.The paper suggests to use the Lagrangian particle tracking technique to model fuel film ...Combustion process inside kerosene-GOx rocket combustor with kerosene Alm cooling is studied,and a modeling approach is proposed.The paper suggests to use the Lagrangian particle tracking technique to model fuel film behavior while the continuous fluid is simulated via the Navier-Stokes system of Favre-averaged equations.The approach is validated over the 12 experimental regimes by the criterions of characteristic velocity and pressure,ence on the adiabatic wall temperatures and relatively low impact on the pressure.In general,phenomena,the calculation of operational processes becomes fast and robust yet precise en-the design process.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
The atomization dynamic characteristics of a simplex swirl injector was investigated experimentally by using a hydrodynamic mechanical pulsator and the shadow photography technique. The frequency response characterist...The atomization dynamic characteristics of a simplex swirl injector was investigated experimentally by using a hydrodynamic mechanical pulsator and the shadow photography technique. The frequency response characteristics of the fluid film and atomization fluctuations and their correlations with pressure fluctuations were obtained by using an in-house code of image processing. It is demonstrated that the klystron effect induced by periodic pressure fluctuations results in periodic liquid film fluctuation with large amplitudes, periodic superposition of droplets and reduction of the breakup length. It was found that the atomization of the simplex swirl injector only responds to the pressure fluctuation in frequency range approximately from 0 to 300 Hz, and it is particularly sensitive to pressure fluctuations at frequencies from 100 to 200 Hz. According to this experiment, the responsive frequency limitation is merely affected by injector configuration, rather than the supply line.展开更多
An experimental study on the Klystron effect of periodic injection modulated by pressure drop fluctuations on subsequent atomization is conducted. Time-resolved atomization backlit images and atomization Mie scatter i...An experimental study on the Klystron effect of periodic injection modulated by pressure drop fluctuations on subsequent atomization is conducted. Time-resolved atomization backlit images and atomization Mie scatter images are captured by using the high speed camera. It is found that periodicity of forced atomization relies on pressure drop fluctuation amplitude and phase differences between atomization and pressure drop fluctuations relate to fluctuation frequencies. This feature of periodic atomization induced by Klystron effect corresponds to periodicities and high amplitudes of pressure fluctuations in unstable combustion chambers and chaos and low amplitudes of pressure fluctuations in stable combustions chambers. Drastically periodic varying of gross surface area of droplets with time was shown in Mie scatter images. The importance of periodic impinging jet atomization modulated by pressure drop fluctuations for acoustic liquid propellant combustion instabilities is illustrated.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.
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.
文摘The ablative material is supposed to be one of good candidates for LRE （liquid rocket engine） combustion chamber to achieve both high reliability and low cost and a numerical analysis for the ablator is considered to be a potentially efficient tool to reduce cost as well. So far, ablators have been successfully applied for many SRM （solid rocket motors）, but the application to LRE is still quite limited in Japan. The authors believe that this is primarily because of the unpredictable nature of the heat load from combustion gases to the combustor wall. Indeed, reliable thermal design of ablative combustion chamber, namely reliable prediction of thermal performance, needs both reliable heat load model and reliable ablator response model. This paper elaborates our research activities and our recent research findings.
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 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.
基金supported by the National Natural Science Foundation of China(Grant No.50706021)
文摘Turbulent two-phase reacting flow in the chamber of LOX/RP-1 bipropellant liquid rocket engine is numerically investigated in this paper. The predicted pressure and mean axial velocity are qualitatively consistent with the experimental measurements. The self-excited pressure oscillations are obtained without any disturbance introduced through the initial and boundary conditions. It is found that amount of abrupt pressure peaks appear frequently and stochastically in the head regions of the chamber, which are the important sources to drive and strengthen combustion instability. Such abrupt pressures are induced by local constant volume combustion, because local combustible gas mixtures with high temperature are formed and burnt out suddenly due to some fuel droplets reaching their critical state in a rich oxygen surrounding. A third Damkhler number is defined as the ratio of the characteristic time of a chemical reaction to the characteristic time of a pressure wave expansion to measure the relative intensity of acoustic propagation and combustion process in thrusters. The analysis of the third Damkhler number distributions in the whole thrust chamber shows that local constant volume combustion happens in the head regions, while constant pressure combustion presents in the downstream regions. It is found that the combustion instability occurs in the head regions within about 30 mm from the thruster head.
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.
基金supported by the Output-oriented Project of the Collaborative Research Program with the Higher Education Partners of the Korea Aerospace Research Institute in 2009
文摘A design method for a kerosene fuel-rich gas-generator of a liquid rocket engine using turbopumps to supply propellant was performed at a conceptual level. The gas-generator creates hot gases, enabling the turbine to operate the turbopumps. A chemical non-equilibrium analysis and a droplet vaporization model were used for the estimation of the burnt gas properties and characteristic chamber length. A premixed counter-flow flame analysis was performed for the prediction of the burnt gas properties, namely the temperature, the specific heat ratio and heat capacity, and the chemical reaction time. To predict the vaporization time, the Spalding model, using a single droplet in convective condition, was used. The minimum residence time in the chamber and the characteristic length were calculated by adding the reaction time and the vaporization time. Using the characteristic length, the design methods for the fuel-rich gas-generator were established. Finally, a parametric study was achieved for the effects of the O/F ratio, mass flow rate, chamber pressure, initial droplet temperature, initial droplet diameter and initial droplet velocity.
基金supported by Advanced Research Center Program(NRF-2013R1A5A1073861)through the National Research Foundation of Korea(NRF)grant funded by the Korea government(MSIP)contracted through Advanced Space Propulsion Research Center at Seoul National University
文摘The pintle injector used for a liquid rocket engine is a newly re-attracted injection system famous for its wide throttle ability with high efficiency. The pintle injector has many variations with complex inner structures due to its moving parts. In order to study the rotating flow near the injector tip, which was observed from the cold flow experiment using water and air, a numerical simulation was adopted and a verification of the numerical model was later conducted. For the verification process, three types of experimental data including velocity distributions of gas flows, spray angles and liquid distribution were all compared using simulated results. The numerical simulation was performed using a commercial simulation program with the Eulerian multiphase model and axisymmetric two dimensional grids. The maximum and minimum velocities of gas were within the acceptable range of agreement, however, the spray angles experienced up to 25% error when the momentum ratios were increased. The spray density distributions were quantitatively measured and had good agreement. As a result of this study, it was concluded that the simulation method was properly constructed to study specific flow characteristics of the pintle injector despite having the limitations of two dimensional and coarse grids.
基金supported by the Output-oriented Project of Collaborative Research Program with Higher Education Partners of Korea Aerospace Research Insti-tute in 2009
文摘An integrated program was established to design a combustor for a liquid rocket engine and to analyze regenerative cooling results on a preliminary design level.Properties of burnt gas from a kerosene-LOx mixture in the combustor and rocket performance were calculated from CEA which is the code for the calculation of chemical equilibrium.The heat transfer of regenerative cooling was analyzed by using SUPERTRAPP code for coolant properties and by one-dimensional correlations of the heat transfer coefficient from the combustor liner to the coolant.Profiles of the combustors of F-1 and RS-27A engines were designed from similar input data and the present results were compared to actual data for validation.Finally,the combustors of 30 tonf class,75 tonf class and 150 tonf class were designed from the required thrust,combustion chamber,exit pressure and mixture ratio of propellants.The wall temperature,heat flux and pressure drop were calculated for heat transfer analysis of regenerative cooling using the profiles.
文摘In this paper, transient phenomenon during start up process of a pump fed liquidrocket engine is investigated through numerical simulation. The engine studied in this workis designed such that engine systems are not wetted with propellant until the engine is com-manded to start. This is achieved by positioning the valves for propellant admission at the inter-face of test stand/flight stage and the engine. To evaluate engine performance during starttransient for such systems, unsteady flow simulation was conducted using Method of Charac-teristics and equations for priming. The same has been reported in this work. The results indi-cated a brief period of abrupt pressure rise at pump upstream after opening of the propellantadmission valves, during the process of priming of engine systems at valve downstream.The peak pressure obtained was significantly higher than the propellant tank pressure as wellas the steady state pump suction pressure. The transitory pressure rise was found to occurdue to flow resistance at impeller inlet caused by formation of a forced vortex for orientingthe flow through impeller blades during off design transient regime. The maximum pressureat pump upstream, as computed from start transient simulation, was used as a design inputfor pump inlet feed lines. The engine was realized and subsequently qualified in a ground test facility. Hot test data obtained for pressure and flow rate during transient regime were found tobe in good agreement with the simulation results.
文摘During the experiment of gas/liquid coaxial swirl injector conducted with air and water under atmosphere environment, it is observed that the injector may selfoscillate. The self oscillation periodically occurs and vanishes with the increasing velocity of the gas flow.A theoretical model is presented based on the experiment investigation. Simulation of the acoustic process has been performed and conclusions consistent with the experiment can be drawn from the theoretical model, which explains the exPeriment phenomena quite well. At last, the comparison between phenomena of the self oscillation and some experiments of LRE indicates that some instability phenomena in oxygen/hydrogen propellant rocket engine may be the related to self oscillation in coaxial injectors
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.
基金Financial support was provided by the Russian Ministry of Education and Science(Project 13.7418.2017/8.9).
文摘Combustion process inside kerosene-GOx rocket combustor with kerosene Alm cooling is studied,and a modeling approach is proposed.The paper suggests to use the Lagrangian particle tracking technique to model fuel film behavior while the continuous fluid is simulated via the Navier-Stokes system of Favre-averaged equations.The approach is validated over the 12 experimental regimes by the criterions of characteristic velocity and pressure,ence on the adiabatic wall temperatures and relatively low impact on the pressure.In general,phenomena,the calculation of operational processes becomes fast and robust yet precise en-the design process.
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.
基金supported by the National Natural Science Foundation of China(Nos.11502186 and 51606138)the National Key Basic Research Program of China(973 Program)and National Key Scientific Instrumentthe Equipment Development Projects of China(No.2012YQ04016408)
文摘The atomization dynamic characteristics of a simplex swirl injector was investigated experimentally by using a hydrodynamic mechanical pulsator and the shadow photography technique. The frequency response characteristics of the fluid film and atomization fluctuations and their correlations with pressure fluctuations were obtained by using an in-house code of image processing. It is demonstrated that the klystron effect induced by periodic pressure fluctuations results in periodic liquid film fluctuation with large amplitudes, periodic superposition of droplets and reduction of the breakup length. It was found that the atomization of the simplex swirl injector only responds to the pressure fluctuation in frequency range approximately from 0 to 300 Hz, and it is particularly sensitive to pressure fluctuations at frequencies from 100 to 200 Hz. According to this experiment, the responsive frequency limitation is merely affected by injector configuration, rather than the supply line.
基金supported by the National Natural Science Foundation of China (Nos. 11502186 and 51606138)the National Key Basic Research Program of ChinaNational Key Scientific Instrument and the Equipment Development Projects of China (No. 2012YQ04016408)
文摘An experimental study on the Klystron effect of periodic injection modulated by pressure drop fluctuations on subsequent atomization is conducted. Time-resolved atomization backlit images and atomization Mie scatter images are captured by using the high speed camera. It is found that periodicity of forced atomization relies on pressure drop fluctuation amplitude and phase differences between atomization and pressure drop fluctuations relate to fluctuation frequencies. This feature of periodic atomization induced by Klystron effect corresponds to periodicities and high amplitudes of pressure fluctuations in unstable combustion chambers and chaos and low amplitudes of pressure fluctuations in stable combustions chambers. Drastically periodic varying of gross surface area of droplets with time was shown in Mie scatter images. The importance of periodic impinging jet atomization modulated by pressure drop fluctuations for acoustic liquid propellant combustion instabilities is illustrated.
