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
Sea-based rocket launches encounter significant challenges stemming from dynamic marine environmental interactions.During the hot launch phase,characterized by low-velocity ascent,the departure of the rocket from the ...Sea-based rocket launches encounter significant challenges stemming from dynamic marine environmental interactions.During the hot launch phase,characterized by low-velocity ascent,the departure of the rocket from the oscillatory platform exhibits heightened sensitivity to external disturbances.In the development stage,assessing the launch dynamics and the clearance between the rocket and framed launcher are crucial for improving the reliability of sea-based rocket launches in rough sea conditions.This study presents a high-fidelity dynamic model of maritime hot launch system,demonstrating 3.21%prediction error through rigorous validation against experimental datasets from comprehensive modal analyses and the full-scale rocket flight test.To mitigate collision risks,we develop a computational method employing spatial vector analysis for dynamic measurement of rocket-launcher clearance during departure.Systematic investigations reveal that in rough sea conditions,optimal departure dynamics are achieved at θ_(thrust)=270°nozzle azimuth configuration,reducing failure probability compared to conventional orientations.The developed assessment framework not only resolves critical safety challenges in current sea launch systems but also establishes foundational principles for optimizing adapter axial configuration patterns in future designs.展开更多
The rocket sled system is not only a high-speed dynamic ground test system,but also one of the future aerospace horizontal launch schemes.The winged load,as a common type of payload,has greater vibration and noise int...The rocket sled system is not only a high-speed dynamic ground test system,but also one of the future aerospace horizontal launch schemes.The winged load,as a common type of payload,has greater vibration and noise intensity than the wingless load.Due to the severe aerodynamic instability prior to separation,the head-up or head-down phenomena are more evident and the test accuracy significantly decreases.The high-precision computer fluid dynamics and aeroacoustic analysis are employed to explore the multifield coupling mechanism of a rocket sled with the winged payload in the wide speed range(Ma=0.5–2).The results show that as the incoming velocity increases,the cone angle of the shock wave of the rocket sled decreases,the shock pressure increases quickly,and the vortex between the slippers splits and gradually shrinks in size.The velocity of the rocket sled exerts little influence on the modal resonance frequency.The wing has a significant impact on aerodynamic noise,and as the sound pressure level rises,the propagation direction gradually shifts towards the rear and upper regions of the wing.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 study introduced an innovative numerical approach to examine combustion instability in Solid Rocket Motors(SRMs).The paper commenced with the derivation of a transient model for the solid propellant's condens...This study introduced an innovative numerical approach to examine combustion instability in Solid Rocket Motors(SRMs).The paper commenced with the derivation of a transient model for the solid propellant's condensed phase,followed by its numerical discretization.Subsequently,this model was integrated with gas phase computations of the chamber's internal flow field,encompassing fluid dynamics and combustion processes.The precision of the numerical method was validated by experimental data,and its reliability was confirmed through a grid independence analysis.The study then investigated the motor's stability under various operating conditions,revealing the impact of parameters such as the sensitivity coefficient of the burning rate to temperature and the nozzle throat diameter on the motor's stability.The results confirmed the bistable nature of combustion instability in specific regions.For instance,when the sensitivity coefficients of burning rate to ambient temperature(k_(1))ranged from 1.4 to 1.8,the SRM adopted in this study with a throat diameter of 0.12 m remained stable under small disturbances but triggered instability under large disturbances.Moreover,increasing the value of k_(1)and reducing the throat diameter can exacerbate combustion instability,leading to more pronounced nonlinear characteristics.The numerical method developed in this paper could effectively capture the nonlinear features of the combustion instability occurring in the motor,providing guidance for SRMs design.展开更多
Taking a C1x motor with a backward-facing step which can generate a typical corner vortex as a reference,a numerical methodology using large eddy simulation was established in this study.Based on this methodology,the ...Taking a C1x motor with a backward-facing step which can generate a typical corner vortex as a reference,a numerical methodology using large eddy simulation was established in this study.Based on this methodology,the position of the backward-facing step of the motor was computed and analyzed to determine a basic configuration.Two key geometrical parameters,the head cavity angle and submerged nozzle cavity height,were subsequently introduced.Their effects on the corner vortex motion and their interactions with the acoustic pressure downstream of the backward-facing step were analyzed.The phenomena of vortex acoustic coupling and characteristics of pressure oscillations were further explored.The results show that the maximum error between the simulations and experimental data on the dominant frequency of pressure oscillations is 5.23%,which indicates that the numerical methodology built in this study is highly accurate.When the step is located at less than 5/8 of the total length of the combustion chamber,vortex acoustic coupling occurs,which can increase the pressure oscillations in the motor.Both the vorticity and the scale of vortices in the downstream step increase when the head cavity angle is greater than 24°,which increases the amplitude of the pressure oscillation by maximum 63.0%.The submerged nozzle cavity mainly affects the vortices in the cavity itself rather than those in the downstream step.When the height of the cavity increases from 10 to 20 mm,the pressure oscillation amplitude under the main frequency increases by 39.1%.As this height continues to increase,the amplitude of pressure oscillations increases but the primary frequency decreases.展开更多
A two-dimensional large eddy simulation numerical model is proposed to study the transient vortex flow and pressure oscillation of a large-aspect-ratio solid rocket motor.The numerical model is validated through exper...A two-dimensional large eddy simulation numerical model is proposed to study the transient vortex flow and pressure oscillation of a large-aspect-ratio solid rocket motor.The numerical model is validated through experimental data,finite element analysis and cumulative error analysis.The numerical simulations are executed to obtain the characteristics of the vortex-acoustic and pressure oscillation.The results show that the burning surface regression decreases the motor aspect ratio,increasing the corresponding natural frequency from 260 Hz to 293 Hz.The pressure oscillation phenomenon is formed due to the vortex-acoustic coupling.Decreasing the corner vortex shedding intensity shows negative effects on the dimensionless amplitude of the pressure oscillation.The head cavity without the injection can decrease the vortex-acoustic coupling level at the acoustic pressure antinode.The modified motor with head cavity can obtain a lower dimensionless oscillating pressure amplitude 0.00149 in comparison with 0.00895 of the original motor.The aspect ratio and volume of the head cavity without the injection have great effects on the pressure oscillation suppression,particularly at the low aspect ratio or large volume.The reason is that the mass in the region around the acoustic pressure antinode is extracted centrally,reducing the energy contribution to the acoustic system.With the volume increasing,the acoustic energy capacity increases.展开更多
As a critical component of pulse solid rocket motors(SRMs),the soft pulse separation device(PSD)is vital in enabling multi-pulse propulsion and has become a breakthrough in SRM engineering applications.To investigate ...As a critical component of pulse solid rocket motors(SRMs),the soft pulse separation device(PSD)is vital in enabling multi-pulse propulsion and has become a breakthrough in SRM engineering applications.To investigate the opening performance of the PSD,an axial PSD incorporating a star-shaped prefabricated defect was designed.The opening process was simulated using peridynamics,yielding the strain field distribution and the corresponding failure mode.A single-opening verification test was conducted.The simulation results showed good agreement with the experimental data,demonstrating the reliability of the peridynamic modeling approach.Furthermore,the effects of the prefabricated defect shape and depth on the opening performance of the PSD were analyzed through simulation.The research results indicate that the established constitutive model and failure criteria based on peridynamics can reasonably predict the failure location and the opening pressure of the soft PSD.Under the impact loading,the weak zone of the soft PSD firstly ruptures,and the damaged area gradually propagates along with the prefabricated defect,eventually leading to complete separation.A smaller prefabricated defect depth or a wider prefabricated defect distribution can cause a reduction in opening pressure.These research results provide valuable guidance for the preliminary design and optimization of PSDs in pulse solid rocket motors.展开更多
As the performance of the box-type multiple launch rocket system(BMLRS)improves,its mechanical structures,particularly the plane clearance design between the slider on the rocket and the guide inside the launch canist...As the performance of the box-type multiple launch rocket system(BMLRS)improves,its mechanical structures,particularly the plane clearance design between the slider on the rocket and the guide inside the launch canister,have grown increasingly complex.However,deficiencies still exist in the current launch modeling theory for BMLRS.In this study,a multi-rigid-flexible-body launch dynamics model coupling the launch platform and rocket was established using the multibody system transfer matrix method and the Newton-Euler formulation.Furthermore,considering the bending of the launch canister,a detection algorithm for slider-guide plane clearance contact was proposed.To quantify the contact force and friction effect between the slider and guide,the contact force model and modified Coulomb model were introduced.Both the modal and launch tests were conducted.Additionally,the modal convergence was verified.By comparing the modal experiments and simulation results,the maximum relative error of the eigenfrequency is 3.29%.thereby verifying the accuracy of the developed BMLRS dynamics model.Furthermore,the launch test validated the proposed plane clearance contact model.Moreover,the study investigated the influence of various model parameters on the dynamic characteristics of BMLRS,including launch canister bending stiffness,slider and guide material,slider-guide clearance,slider length and layout.This analysis of influencing factors provides a foundation for future optimization in BMLRS design.展开更多
This study investigates the potential of metal additives in acrylonitrile butadiene styrene(ABS)polymer fuel to enhance hybrid rocket motor(HRM)performance through computational analysis,Chemical Equilibrium with Appl...This study investigates the potential of metal additives in acrylonitrile butadiene styrene(ABS)polymer fuel to enhance hybrid rocket motor(HRM)performance through computational analysis,Chemical Equilibrium with Applications(CEA),software.ABS was selected as the base fuel due to its thermoplastic nature,which allows for the creation of complex fuel geometries through 3D printing,offering significant flexibility in fuel design.Hybrid rockets,which combine a solid fuel with a liquid oxidiser,offer advantages in terms of operational simplicity and safety.However,conventional polymer fuels often exhibit low regression rates and suboptimal combustion efficiencies.In this research,we evaluated a range of metal additives-aluminium(Al),boron(B),nickel(Ni),copper(Cu),and iron(Fe)-at chamber pressures ranging from 1 to 30 bar and oxidiser-to-fuel(O/F)ratios between 1.1 and 12,resulting in 1800 unique test conditions.The main performance parameters used to assess each formulation were characteristic velocity(C^(*))and adiabatic flame temperature.The results revealed that each test produced a different optimum O/F ratio,with most ratios falling between 4 and 6.The highest performance was achieved at a chamber pressure of 30 bar across all formulations.Among the additives,Al and B demonstrated significant potential for improved combustion performance with increasing metal loadings.In contrast,Fe,Cu,and Ni reached optimal performance at a minimum loading of 1%.Future work includes investigating B-Al metal composites as additives into the ABS base polymer fuel,and doing experimental validation tests where the metallised ABS polymer fuel is 3D printed.展开更多
Accurate navigation is important for long-range rocket projectile＇s precise striking. To obtain stable and high-per- formance navigation result, a ultra-tight global positioning system/inertial navigation system （GP...Accurate navigation is important for long-range rocket projectile＇s precise striking. To obtain stable and high-per- formance navigation result, a ultra-tight global positioning system/inertial navigation system （GPS/INS） integration based nav- igation approach is proposed. The accurate short-time output of INS is used by GPS receiver to assist in acquisition of signal, and output information of INS and GPS is fused based on federated filter. Meanwhile, the improved cubature Kalman filter with strong tracking ability is chosen to serve as the local filter, and then the federated filter is enhanced based on vector sharing theory. Finally, simulation results show that the navigation accuracy with the proposed method is higher than that with traditional methods. It provides reference for long-range rocket projectile navigation.展开更多
The active control theory and methods of initial disturbances for rockets and missiles are investigated. The rocket or missile/launcher is simplified as a flexible beam excited by a moving varying velocity rigid body ...The active control theory and methods of initial disturbances for rockets and missiles are investigated. The rocket or missile/launcher is simplified as a flexible beam excited by a moving varying velocity rigid body which has two points in contact with the beam. The control force is applied at the supporting point on the beam. Active control strategies based on optimal control theory are proposed and computer simulation is carried out. Simulation results are consistent with the theoretical results, and show that the active control strategies proposed can accomplish the purpose to control the initial disturbances actively. The results show that active control of initial disturbances for rockets and missiles is feasible for application.展开更多
A hybrid rocket motor combines components from both solid fuel and liquid fuel rocket motors. The fuel itself is a solid grain, (often paraffin or hydroxyl-terminated polybutadiene, known as HTPB) while the oxidizing ...A hybrid rocket motor combines components from both solid fuel and liquid fuel rocket motors. The fuel itself is a solid grain, (often paraffin or hydroxyl-terminated polybutadiene, known as HTPB) while the oxidizing agent is liquid (often hydrogen peroxide or liquid oxygen). These components are combined in the fuel chamber which doubles as the combustion chamber for the hybrid motor. This review looks at the advances in techniques that have taken place in the development of these motors since 1995. Methods of testing the thrust from rocket motors and of measuring the rocket plume spectroscopically for combustion reaction products have been developed. These assessments allow researchers to more completely understand the effects of additives and physical changes in design, in terms of regression rates and thrust developed. Hybrid rocket motors have been used or tested in many areas of rocketry, including tactical rockets and large launch vehicles. Several additives have shown significant improvements in regression rates and thrust, including Guanidinium azotetrazolate (GAT), and various Aluminum alloys. The most recent discoveries have come from research into nano-particle additives. The nano-particles have been shown to provide enhancements to many parameters of hybrid rocket function, and research into specific areas continues in the sub-field of nano-additives for fuel grains.展开更多
Our study identifies a subtle deviation from Newton’s third law in the derivation of the ideal rocket equation, also known as the Tsiolkovsky Rocket Equation (TRE). TRE can be derived using a 1D elastic collision mod...Our study identifies a subtle deviation from Newton’s third law in the derivation of the ideal rocket equation, also known as the Tsiolkovsky Rocket Equation (TRE). TRE can be derived using a 1D elastic collision model of the momentum exchange between the differential propellant mass element (dm) and the rocket final mass (m1), in which dm initially travels forward to collide with m1 and rebounds to exit through the exhaust nozzle with a velocity that is known as the effective exhaust velocity ve. We observe that such a model does not explain how dm was able to acquire its initial forward velocity without the support of a reactive mass traveling in the opposite direction. We show instead that the initial kinetic energy of dm is generated from dm itself by a process of self-combustion and expansion. In our ideal rocket with a single particle dm confined inside a hollow tube with one closed end, we show that the process of self-combustion and expansion of dm will result in a pair of differential particles each with a mass dm/2, and each traveling away from one another along the tube axis, from the center of combustion. These two identical particles represent the active and reactive sub-components of dm, co-generated in compliance with Newton’s third law of equal action and reaction. Building on this model, we derive a linear momentum ODE of the system, the solution of which yields what we call the Revised Tsiolkovsky Rocket Equation (RTRE). We show that RTRE has a mathematical form that is similar to TRE, with the exception of the effective exhaust velocity (ve) term. The ve term in TRE is replaced in RTRE by the average of two distinct exhaust velocities that we refer to as fast-jet, vx<sub>1</sub>, and slow-jet, vx<sub>2</sub>. These two velocities correspond, respectively, to the velocities of the detonation pressure wave that is vectored directly towards the exhaust nozzle, and the retonation wave that is initially vectored in the direction of rocket propagation, but subsequently becomes reflected from the thrust surface of the combustion chamber to exit through the exhaust nozzle with a time lag behind the detonation wave. The detonation-retonation phenomenon is supported by experimental evidence in the published literature. Finally, we use a convolution model to simulate the composite exhaust pressure wave, highlighting the frequency spectrum of the pressure perturbations that are generated by the mutual interference between the fast-jet and slow-jet components. Our analysis offers insights into the origin of combustion oscillations in rocket engines, with possible extensions beyond rocket engineering into other fields of combustion engineering.展开更多
In order to measure the parameters of flight rocket by using radar,rocket impact point was estimated accurately for rocket trajectory correction.The Kalman filter with adaptive filter gain matrix was adopted.According...In order to measure the parameters of flight rocket by using radar,rocket impact point was estimated accurately for rocket trajectory correction.The Kalman filter with adaptive filter gain matrix was adopted.According to the particle trajectory model,the adaptive Kalman filter trajectory model was constructed for removing and filtering the outliers of the parameters during a section of flight detected by three-dimensional data radar and the rocket impact point was extrapolated.The results of numerical simulation show that the outliers and noise in trajectory measurement signal can be removed effectively by using the adaptive Kalman filter and the filter variance can converge in a short period of time.Based on the relation of filtering time and impact point estimation error,choosing the filtering time of 8-10 scan get the minimum estimation error of impact point.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the experimental technology support provided by the China Academy of Launch Vehicle Technology
文摘Sea-based rocket launches encounter significant challenges stemming from dynamic marine environmental interactions.During the hot launch phase,characterized by low-velocity ascent,the departure of the rocket from the oscillatory platform exhibits heightened sensitivity to external disturbances.In the development stage,assessing the launch dynamics and the clearance between the rocket and framed launcher are crucial for improving the reliability of sea-based rocket launches in rough sea conditions.This study presents a high-fidelity dynamic model of maritime hot launch system,demonstrating 3.21%prediction error through rigorous validation against experimental datasets from comprehensive modal analyses and the full-scale rocket flight test.To mitigate collision risks,we develop a computational method employing spatial vector analysis for dynamic measurement of rocket-launcher clearance during departure.Systematic investigations reveal that in rough sea conditions,optimal departure dynamics are achieved at θ_(thrust)=270°nozzle azimuth configuration,reducing failure probability compared to conventional orientations.The developed assessment framework not only resolves critical safety challenges in current sea launch systems but also establishes foundational principles for optimizing adapter axial configuration patterns in future designs.
基金supported by the National Natural Science Foundation of China(No.12104047)。
文摘The rocket sled system is not only a high-speed dynamic ground test system,but also one of the future aerospace horizontal launch schemes.The winged load,as a common type of payload,has greater vibration and noise intensity than the wingless load.Due to the severe aerodynamic instability prior to separation,the head-up or head-down phenomena are more evident and the test accuracy significantly decreases.The high-precision computer fluid dynamics and aeroacoustic analysis are employed to explore the multifield coupling mechanism of a rocket sled with the winged payload in the wide speed range(Ma=0.5–2).The results show that as the incoming velocity increases,the cone angle of the shock wave of the rocket sled decreases,the shock pressure increases quickly,and the vortex between the slippers splits and gradually shrinks in size.The velocity of the rocket sled exerts little influence on the modal resonance frequency.The wing has a significant impact on aerodynamic noise,and as the sound pressure level rises,the propagation direction gradually shifts towards the rear and upper regions of the wing.
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.
基金supported by the National Natural Science Foundation of China(No.U2241250)。
文摘This study introduced an innovative numerical approach to examine combustion instability in Solid Rocket Motors(SRMs).The paper commenced with the derivation of a transient model for the solid propellant's condensed phase,followed by its numerical discretization.Subsequently,this model was integrated with gas phase computations of the chamber's internal flow field,encompassing fluid dynamics and combustion processes.The precision of the numerical method was validated by experimental data,and its reliability was confirmed through a grid independence analysis.The study then investigated the motor's stability under various operating conditions,revealing the impact of parameters such as the sensitivity coefficient of the burning rate to temperature and the nozzle throat diameter on the motor's stability.The results confirmed the bistable nature of combustion instability in specific regions.For instance,when the sensitivity coefficients of burning rate to ambient temperature(k_(1))ranged from 1.4 to 1.8,the SRM adopted in this study with a throat diameter of 0.12 m remained stable under small disturbances but triggered instability under large disturbances.Moreover,increasing the value of k_(1)and reducing the throat diameter can exacerbate combustion instability,leading to more pronounced nonlinear characteristics.The numerical method developed in this paper could effectively capture the nonlinear features of the combustion instability occurring in the motor,providing guidance for SRMs design.
基金Sponsored by the Natural Science Foundation of Shaanxi Province (Grant No. S2025-JC-YB-0532)the Practice and Innovation Funds for Graduate Students of Northwestern Polytechnical University (PF2024044)
文摘Taking a C1x motor with a backward-facing step which can generate a typical corner vortex as a reference,a numerical methodology using large eddy simulation was established in this study.Based on this methodology,the position of the backward-facing step of the motor was computed and analyzed to determine a basic configuration.Two key geometrical parameters,the head cavity angle and submerged nozzle cavity height,were subsequently introduced.Their effects on the corner vortex motion and their interactions with the acoustic pressure downstream of the backward-facing step were analyzed.The phenomena of vortex acoustic coupling and characteristics of pressure oscillations were further explored.The results show that the maximum error between the simulations and experimental data on the dominant frequency of pressure oscillations is 5.23%,which indicates that the numerical methodology built in this study is highly accurate.When the step is located at less than 5/8 of the total length of the combustion chamber,vortex acoustic coupling occurs,which can increase the pressure oscillations in the motor.Both the vorticity and the scale of vortices in the downstream step increase when the head cavity angle is greater than 24°,which increases the amplitude of the pressure oscillation by maximum 63.0%.The submerged nozzle cavity mainly affects the vortices in the cavity itself rather than those in the downstream step.When the height of the cavity increases from 10 to 20 mm,the pressure oscillation amplitude under the main frequency increases by 39.1%.As this height continues to increase,the amplitude of pressure oscillations increases but the primary frequency decreases.
基金supported by the Natural Science Foundation of Hunan Province of China(No.2023JJ40672)the Innovation Science Fund Project of National University of Defense Technology,China(No.ZK2023-039)。
文摘A two-dimensional large eddy simulation numerical model is proposed to study the transient vortex flow and pressure oscillation of a large-aspect-ratio solid rocket motor.The numerical model is validated through experimental data,finite element analysis and cumulative error analysis.The numerical simulations are executed to obtain the characteristics of the vortex-acoustic and pressure oscillation.The results show that the burning surface regression decreases the motor aspect ratio,increasing the corresponding natural frequency from 260 Hz to 293 Hz.The pressure oscillation phenomenon is formed due to the vortex-acoustic coupling.Decreasing the corner vortex shedding intensity shows negative effects on the dimensionless amplitude of the pressure oscillation.The head cavity without the injection can decrease the vortex-acoustic coupling level at the acoustic pressure antinode.The modified motor with head cavity can obtain a lower dimensionless oscillating pressure amplitude 0.00149 in comparison with 0.00895 of the original motor.The aspect ratio and volume of the head cavity without the injection have great effects on the pressure oscillation suppression,particularly at the low aspect ratio or large volume.The reason is that the mass in the region around the acoustic pressure antinode is extracted centrally,reducing the energy contribution to the acoustic system.With the volume increasing,the acoustic energy capacity increases.
基金supported by the National Natural Science Foundation of China(No.12202011)the Youth Research fund of Shanghai Academy of Spaceflight Technology(KJW-KT-QNKYJJ-2022-25)China Postdoctoral Science Foundation(Nos.2024T170009,2022M710190).
文摘As a critical component of pulse solid rocket motors(SRMs),the soft pulse separation device(PSD)is vital in enabling multi-pulse propulsion and has become a breakthrough in SRM engineering applications.To investigate the opening performance of the PSD,an axial PSD incorporating a star-shaped prefabricated defect was designed.The opening process was simulated using peridynamics,yielding the strain field distribution and the corresponding failure mode.A single-opening verification test was conducted.The simulation results showed good agreement with the experimental data,demonstrating the reliability of the peridynamic modeling approach.Furthermore,the effects of the prefabricated defect shape and depth on the opening performance of the PSD were analyzed through simulation.The research results indicate that the established constitutive model and failure criteria based on peridynamics can reasonably predict the failure location and the opening pressure of the soft PSD.Under the impact loading,the weak zone of the soft PSD firstly ruptures,and the damaged area gradually propagates along with the prefabricated defect,eventually leading to complete separation.A smaller prefabricated defect depth or a wider prefabricated defect distribution can cause a reduction in opening pressure.These research results provide valuable guidance for the preliminary design and optimization of PSDs in pulse solid rocket motors.
基金supported by National Natural Science Foundation of China(Grant No.92266201).
文摘As the performance of the box-type multiple launch rocket system(BMLRS)improves,its mechanical structures,particularly the plane clearance design between the slider on the rocket and the guide inside the launch canister,have grown increasingly complex.However,deficiencies still exist in the current launch modeling theory for BMLRS.In this study,a multi-rigid-flexible-body launch dynamics model coupling the launch platform and rocket was established using the multibody system transfer matrix method and the Newton-Euler formulation.Furthermore,considering the bending of the launch canister,a detection algorithm for slider-guide plane clearance contact was proposed.To quantify the contact force and friction effect between the slider and guide,the contact force model and modified Coulomb model were introduced.Both the modal and launch tests were conducted.Additionally,the modal convergence was verified.By comparing the modal experiments and simulation results,the maximum relative error of the eigenfrequency is 3.29%.thereby verifying the accuracy of the developed BMLRS dynamics model.Furthermore,the launch test validated the proposed plane clearance contact model.Moreover,the study investigated the influence of various model parameters on the dynamic characteristics of BMLRS,including launch canister bending stiffness,slider and guide material,slider-guide clearance,slider length and layout.This analysis of influencing factors provides a foundation for future optimization in BMLRS design.
文摘This study investigates the potential of metal additives in acrylonitrile butadiene styrene(ABS)polymer fuel to enhance hybrid rocket motor(HRM)performance through computational analysis,Chemical Equilibrium with Applications(CEA),software.ABS was selected as the base fuel due to its thermoplastic nature,which allows for the creation of complex fuel geometries through 3D printing,offering significant flexibility in fuel design.Hybrid rockets,which combine a solid fuel with a liquid oxidiser,offer advantages in terms of operational simplicity and safety.However,conventional polymer fuels often exhibit low regression rates and suboptimal combustion efficiencies.In this research,we evaluated a range of metal additives-aluminium(Al),boron(B),nickel(Ni),copper(Cu),and iron(Fe)-at chamber pressures ranging from 1 to 30 bar and oxidiser-to-fuel(O/F)ratios between 1.1 and 12,resulting in 1800 unique test conditions.The main performance parameters used to assess each formulation were characteristic velocity(C^(*))and adiabatic flame temperature.The results revealed that each test produced a different optimum O/F ratio,with most ratios falling between 4 and 6.The highest performance was achieved at a chamber pressure of 30 bar across all formulations.Among the additives,Al and B demonstrated significant potential for improved combustion performance with increasing metal loadings.In contrast,Fe,Cu,and Ni reached optimal performance at a minimum loading of 1%.Future work includes investigating B-Al metal composites as additives into the ABS base polymer fuel,and doing experimental validation tests where the metallised ABS polymer fuel is 3D printed.
基金Project Funded by Chongqing Changjiang Electrical Appliances Industries Group Co.,Ltd
文摘Accurate navigation is important for long-range rocket projectile＇s precise striking. To obtain stable and high-per- formance navigation result, a ultra-tight global positioning system/inertial navigation system （GPS/INS） integration based nav- igation approach is proposed. The accurate short-time output of INS is used by GPS receiver to assist in acquisition of signal, and output information of INS and GPS is fused based on federated filter. Meanwhile, the improved cubature Kalman filter with strong tracking ability is chosen to serve as the local filter, and then the federated filter is enhanced based on vector sharing theory. Finally, simulation results show that the navigation accuracy with the proposed method is higher than that with traditional methods. It provides reference for long-range rocket projectile navigation.
文摘The active control theory and methods of initial disturbances for rockets and missiles are investigated. The rocket or missile/launcher is simplified as a flexible beam excited by a moving varying velocity rigid body which has two points in contact with the beam. The control force is applied at the supporting point on the beam. Active control strategies based on optimal control theory are proposed and computer simulation is carried out. Simulation results are consistent with the theoretical results, and show that the active control strategies proposed can accomplish the purpose to control the initial disturbances actively. The results show that active control of initial disturbances for rockets and missiles is feasible for application.
文摘A hybrid rocket motor combines components from both solid fuel and liquid fuel rocket motors. The fuel itself is a solid grain, (often paraffin or hydroxyl-terminated polybutadiene, known as HTPB) while the oxidizing agent is liquid (often hydrogen peroxide or liquid oxygen). These components are combined in the fuel chamber which doubles as the combustion chamber for the hybrid motor. This review looks at the advances in techniques that have taken place in the development of these motors since 1995. Methods of testing the thrust from rocket motors and of measuring the rocket plume spectroscopically for combustion reaction products have been developed. These assessments allow researchers to more completely understand the effects of additives and physical changes in design, in terms of regression rates and thrust developed. Hybrid rocket motors have been used or tested in many areas of rocketry, including tactical rockets and large launch vehicles. Several additives have shown significant improvements in regression rates and thrust, including Guanidinium azotetrazolate (GAT), and various Aluminum alloys. The most recent discoveries have come from research into nano-particle additives. The nano-particles have been shown to provide enhancements to many parameters of hybrid rocket function, and research into specific areas continues in the sub-field of nano-additives for fuel grains.
文摘Our study identifies a subtle deviation from Newton’s third law in the derivation of the ideal rocket equation, also known as the Tsiolkovsky Rocket Equation (TRE). TRE can be derived using a 1D elastic collision model of the momentum exchange between the differential propellant mass element (dm) and the rocket final mass (m1), in which dm initially travels forward to collide with m1 and rebounds to exit through the exhaust nozzle with a velocity that is known as the effective exhaust velocity ve. We observe that such a model does not explain how dm was able to acquire its initial forward velocity without the support of a reactive mass traveling in the opposite direction. We show instead that the initial kinetic energy of dm is generated from dm itself by a process of self-combustion and expansion. In our ideal rocket with a single particle dm confined inside a hollow tube with one closed end, we show that the process of self-combustion and expansion of dm will result in a pair of differential particles each with a mass dm/2, and each traveling away from one another along the tube axis, from the center of combustion. These two identical particles represent the active and reactive sub-components of dm, co-generated in compliance with Newton’s third law of equal action and reaction. Building on this model, we derive a linear momentum ODE of the system, the solution of which yields what we call the Revised Tsiolkovsky Rocket Equation (RTRE). We show that RTRE has a mathematical form that is similar to TRE, with the exception of the effective exhaust velocity (ve) term. The ve term in TRE is replaced in RTRE by the average of two distinct exhaust velocities that we refer to as fast-jet, vx<sub>1</sub>, and slow-jet, vx<sub>2</sub>. These two velocities correspond, respectively, to the velocities of the detonation pressure wave that is vectored directly towards the exhaust nozzle, and the retonation wave that is initially vectored in the direction of rocket propagation, but subsequently becomes reflected from the thrust surface of the combustion chamber to exit through the exhaust nozzle with a time lag behind the detonation wave. The detonation-retonation phenomenon is supported by experimental evidence in the published literature. Finally, we use a convolution model to simulate the composite exhaust pressure wave, highlighting the frequency spectrum of the pressure perturbations that are generated by the mutual interference between the fast-jet and slow-jet components. Our analysis offers insights into the origin of combustion oscillations in rocket engines, with possible extensions beyond rocket engineering into other fields of combustion engineering.
文摘In order to measure the parameters of flight rocket by using radar,rocket impact point was estimated accurately for rocket trajectory correction.The Kalman filter with adaptive filter gain matrix was adopted.According to the particle trajectory model,the adaptive Kalman filter trajectory model was constructed for removing and filtering the outliers of the parameters during a section of flight detected by three-dimensional data radar and the rocket impact point was extrapolated.The results of numerical simulation show that the outliers and noise in trajectory measurement signal can be removed effectively by using the adaptive Kalman filter and the filter variance can converge in a short period of time.Based on the relation of filtering time and impact point estimation error,choosing the filtering time of 8-10 scan get the minimum estimation error of impact point.
