Liquid rocket engine(LRE)fault diagnosis is critical for successful space launch missions,enabling timely avoidance of safety hazards,while accurate post-failure analysis prevents subsequent economic losses.However,th...Liquid rocket engine(LRE)fault diagnosis is critical for successful space launch missions,enabling timely avoidance of safety hazards,while accurate post-failure analysis prevents subsequent economic losses.However,the complexity of LRE systems and the“black-box”nature of current deep learning-based diagnostic methods hinder interpretable fault diagnosis.This paper establishes Granger causality(GC)extraction-based component-wise multi-layer perceptron(GCMLP),achieving high fault diagnosis accuracy while leveraging GC to enhance diagnostic interpretability.First,component-wise MLP networks are constructed for distinct LRE variables to extract inter-variable GC relationships.Second,dedicated predictors are designed for each variable,leveraging historical data and GC relationships to forecast future states,thereby ensuring GC reliability.Finally,the extracted GC features are utilized for fault classification,guaranteeing feature discriminability and diagnosis accuracy.This study simulates six critical fault modes in LRE using Simulink.Based on the generated simulation data,GCMLP demonstrates superior fault localization accuracy compared to benchmark methods,validating its efficacy and robustness.展开更多
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
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
This paper presents a method of thermal state calculation of combustion chamber in small thrust liquid rocket engine. The goal is to predict the thermal state of chamber wall by using basic parameters of engine: thrus...This paper presents a method of thermal state calculation of combustion chamber in small thrust liquid rocket engine. The goal is to predict the thermal state of chamber wall by using basic parameters of engine: thrust level, propellants, chamber pressure, injection pattern, film cooling parameters, material of wall and their coating, etc. The difficulties in modeling the startup and shutdown processes of thrusters lie in the fact that there are the conjugated physical processes occurring at various parameters for non-design conditions. A mathematical model to predict the thermal state of the combustion chamber for different engine operation modes is developed. To simulate the startup and shutdown processes, a quasi-steady approach is applied by replacing the transient process with time-variant operating parameters of steady-state processes. The mathematical model is based on several principles and data commonly used for heat transfer modeling: geometry of flow part, gas dynamics of flow, thermodynamics of propellants and combustion spices, convective and radiation heat flows, conjugated heat transfer between hot gas and wall, and transient approach for calculation of thermal state of construction. Calculations of the thermal state of the combustion chamber in single-turn-on mode show good convergence with the experimental results. The results of pulsed modes indicate a large temperature gradient on the internal wall surface of the chamber between pulses and the thermal state of the wall strongly depends on the pulse duration and the interval.展开更多
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
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
Safety analysis for liquid rocket engine has a great meaning for shortening development cycle, saving development expenditure and reducing development risk. The relationship between the structure and component of liqu...Safety analysis for liquid rocket engine has a great meaning for shortening development cycle, saving development expenditure and reducing development risk. The relationship between the structure and component of liquid rocket engine is much more complex, furthermore test data are absent in development phase. Thereby, the uncertainties exist in safety analysis for liquid rocket engine. A safety analysis model integrated with FMEA(failure mode and effect analysis) based on Bayesian networks (BN) is brought forward for liquid rocket engine, which can combine qualitative analysis with quantitative decision. The method has the advantages of fusing multi-information, saving sample amount and having high veracity. An example shows that the method is efficient.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
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
A brief review of nitrous oxide decomposition catalysts was presented.The features of catalyst operating conditions in low-thrust engines of space vehicles and requirements to monopropellant(hydrogen peroxide,hydrazin...A brief review of nitrous oxide decomposition catalysts was presented.The features of catalyst operating conditions in low-thrust engines of space vehicles and requirements to monopropellant(hydrogen peroxide,hydrazine,nitrous oxide)decomposition catalysts were considered.A scientific basis for development of a nitrous oxide decomposition catalyst and general principles for selection of efficient catalysts were formulated.The results of selecting catalyst systems for the development of decomposition catalysts for N2O as a monopropellant were presented.Preliminary selection of catalyst systems for the development of a catalyst designed for low-thrust rocket engines(LTREs)was carried out:supporter—Al2O3 and ZrO2;active substances—Co,Ni,Fe,Pd,Rh,Pt,Ru,Ir,NiO,Fe2O3,RuO2,Rh2O3,PdO,IrO2,PtO2,CoO,Al2O3,La2NiO4,Nd2NiO4,Pr 2NiO4,La2O3,TiO2,NiO,La2O3,TiO2,ZnO.With 71 references.展开更多
Recent development in space mission demands safer and more cost-effective space missions.Hybrid rocket engine technological advancements have prolonged a critical stage in their development and it is the better option...Recent development in space mission demands safer and more cost-effective space missions.Hybrid rocket engine technological advancements have prolonged a critical stage in their development and it is the better option for such space missions,as it has a lot of advantages over the solid rocket motor and liquid rocket engine.It is simple in design,has high thrust density,low weight,and is safer than a liquid rocket engine.It has restarted capability,safe,low explosion risk,and high specific impulse than a solid rocket motor.This paper shows the numerical analysis of a hybrid rocket engine.The paper highlights the initial boundary conditions in the analysis of a 300-N hybrid rocket engine.The process started with a chemical kinematic examination of engine-compatible fuels and oxidizers.This investigation provided the fundamental parameters required for the design and subsequent dimensioning of a hybrid rocket engine.It also produced a three-dimensional design model,performed numerical analysis using ANSYS software,and validated the findings using existing literature.Using the k-εturbulence model and transient solver on 8 mm port diameter for analyzing.The computational fluid dynamics model offered the qualities of a real hybrid rocket engine and it will be helpful to researchers and the scientific community in the future.展开更多
With the advent of new technologies in space science,numerous missions and programs to launch satellites of various scales have received significant attention in the scientific community.However,there are a few comple...With the advent of new technologies in space science,numerous missions and programs to launch satellites of various scales have received significant attention in the scientific community.However,there are a few complexities involved to encourage activities such as space tourism and exploration due to the technical constraints that the launch vehicle poses.To surmount these challenges,a novel hybrid rocket engine must be designed which allows us to tailor the characteristics during its endurance simulations.In the current work,we have investigated the propellant’s thermal and flow characteristics such as glass transition and melting temperature,specific impulse,characteristic velocity,regression rate,and thrust,respectively,without compromising the performance of likewise liquid propellant engine.In the current work,we investigated propellant characteristics for pure paraffin wax with oxidizers,such as liquid oxygen and nitrous oxide,and pure paraffin with additives of aluminum,magnesium,and oxidizers.It was observed that pure paraffin with liquid oxygen and additive aluminum(30%weight)resulted in the highest specific impulse among other combinations,whereas differential scanning calorimeter investigations reported a reduction in latent heat with the increase in weight%of paraffin and aluminum powder.It is also suggested that heat flow at 10%weight complies with the hybrid rocket engine.展开更多
文摘Liquid rocket engine(LRE)fault diagnosis is critical for successful space launch missions,enabling timely avoidance of safety hazards,while accurate post-failure analysis prevents subsequent economic losses.However,the complexity of LRE systems and the“black-box”nature of current deep learning-based diagnostic methods hinder interpretable fault diagnosis.This paper establishes Granger causality(GC)extraction-based component-wise multi-layer perceptron(GCMLP),achieving high fault diagnosis accuracy while leveraging GC to enhance diagnostic interpretability.First,component-wise MLP networks are constructed for distinct LRE variables to extract inter-variable GC relationships.Second,dedicated predictors are designed for each variable,leveraging historical data and GC relationships to forecast future states,thereby ensuring GC reliability.Finally,the extracted GC features are utilized for fault classification,guaranteeing feature discriminability and diagnosis accuracy.This study simulates six critical fault modes in LRE using Simulink.Based on the generated simulation data,GCMLP demonstrates superior fault localization accuracy compared to benchmark methods,validating its efficacy and robustness.
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.
文摘This paper presents a method of thermal state calculation of combustion chamber in small thrust liquid rocket engine. The goal is to predict the thermal state of chamber wall by using basic parameters of engine: thrust level, propellants, chamber pressure, injection pattern, film cooling parameters, material of wall and their coating, etc. The difficulties in modeling the startup and shutdown processes of thrusters lie in the fact that there are the conjugated physical processes occurring at various parameters for non-design conditions. A mathematical model to predict the thermal state of the combustion chamber for different engine operation modes is developed. To simulate the startup and shutdown processes, a quasi-steady approach is applied by replacing the transient process with time-variant operating parameters of steady-state processes. The mathematical model is based on several principles and data commonly used for heat transfer modeling: geometry of flow part, gas dynamics of flow, thermodynamics of propellants and combustion spices, convective and radiation heat flows, conjugated heat transfer between hot gas and wall, and transient approach for calculation of thermal state of construction. Calculations of the thermal state of the combustion chamber in single-turn-on mode show good convergence with the experimental results. The results of pulsed modes indicate a large temperature gradient on the internal wall surface of the chamber between pulses and the thermal state of the wall strongly depends on the pulse duration and the interval.
文摘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.
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Safety analysis for liquid rocket engine has a great meaning for shortening development cycle, saving development expenditure and reducing development risk. The relationship between the structure and component of liquid rocket engine is much more complex, furthermore test data are absent in development phase. Thereby, the uncertainties exist in safety analysis for liquid rocket engine. A safety analysis model integrated with FMEA(failure mode and effect analysis) based on Bayesian networks (BN) is brought forward for liquid rocket engine, which can combine qualitative analysis with quantitative decision. The method has the advantages of fusing multi-information, saving sample amount and having high veracity. An example shows that the method is efficient.
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.
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.
文摘A brief review of nitrous oxide decomposition catalysts was presented.The features of catalyst operating conditions in low-thrust engines of space vehicles and requirements to monopropellant(hydrogen peroxide,hydrazine,nitrous oxide)decomposition catalysts were considered.A scientific basis for development of a nitrous oxide decomposition catalyst and general principles for selection of efficient catalysts were formulated.The results of selecting catalyst systems for the development of decomposition catalysts for N2O as a monopropellant were presented.Preliminary selection of catalyst systems for the development of a catalyst designed for low-thrust rocket engines(LTREs)was carried out:supporter—Al2O3 and ZrO2;active substances—Co,Ni,Fe,Pd,Rh,Pt,Ru,Ir,NiO,Fe2O3,RuO2,Rh2O3,PdO,IrO2,PtO2,CoO,Al2O3,La2NiO4,Nd2NiO4,Pr 2NiO4,La2O3,TiO2,NiO,La2O3,TiO2,ZnO.With 71 references.
文摘Recent development in space mission demands safer and more cost-effective space missions.Hybrid rocket engine technological advancements have prolonged a critical stage in their development and it is the better option for such space missions,as it has a lot of advantages over the solid rocket motor and liquid rocket engine.It is simple in design,has high thrust density,low weight,and is safer than a liquid rocket engine.It has restarted capability,safe,low explosion risk,and high specific impulse than a solid rocket motor.This paper shows the numerical analysis of a hybrid rocket engine.The paper highlights the initial boundary conditions in the analysis of a 300-N hybrid rocket engine.The process started with a chemical kinematic examination of engine-compatible fuels and oxidizers.This investigation provided the fundamental parameters required for the design and subsequent dimensioning of a hybrid rocket engine.It also produced a three-dimensional design model,performed numerical analysis using ANSYS software,and validated the findings using existing literature.Using the k-εturbulence model and transient solver on 8 mm port diameter for analyzing.The computational fluid dynamics model offered the qualities of a real hybrid rocket engine and it will be helpful to researchers and the scientific community in the future.
文摘With the advent of new technologies in space science,numerous missions and programs to launch satellites of various scales have received significant attention in the scientific community.However,there are a few complexities involved to encourage activities such as space tourism and exploration due to the technical constraints that the launch vehicle poses.To surmount these challenges,a novel hybrid rocket engine must be designed which allows us to tailor the characteristics during its endurance simulations.In the current work,we have investigated the propellant’s thermal and flow characteristics such as glass transition and melting temperature,specific impulse,characteristic velocity,regression rate,and thrust,respectively,without compromising the performance of likewise liquid propellant engine.In the current work,we investigated propellant characteristics for pure paraffin wax with oxidizers,such as liquid oxygen and nitrous oxide,and pure paraffin with additives of aluminum,magnesium,and oxidizers.It was observed that pure paraffin with liquid oxygen and additive aluminum(30%weight)resulted in the highest specific impulse among other combinations,whereas differential scanning calorimeter investigations reported a reduction in latent heat with the increase in weight%of paraffin and aluminum powder.It is also suggested that heat flow at 10%weight complies with the hybrid rocket engine.
