The internal flow fields within a three-dimensional inward-tunning combined inlet are extremely complex,especially during the engine mode transition,where the tunnel changes may impact the flow fields significantly.To...The internal flow fields within a three-dimensional inward-tunning combined inlet are extremely complex,especially during the engine mode transition,where the tunnel changes may impact the flow fields significantly.To develop an efficient flow field reconstruction model for this,we present an Improved Conditional Denoising Diffusion Generative Adversarial Network(ICDDGAN),which integrates Conditional Denoising Diffusion Probabilistic Models(CDDPMs)with Style GAN,and introduce a reconstruction discrimination mechanism and dynamic loss weight learning strategy.We establish the Mach number flow field dataset by numerical simulation at various backpressures for the mode transition process from turbine mode to ejector ramjet mode at Mach number 2.5.The proposed ICDDGAN model,given only sparse parameter information,can rapidly generate high-quality Mach number flow fields without a large number of samples for training.The results show that ICDDGAN is superior to CDDGAN in terms of training convergence and stability.Moreover,the interpolation and extrapolation test results during backpressure conditions show that ICDDGAN can accurately and quickly reconstruct Mach number fields at various tunnel slice shapes,with a Structural Similarity Index Measure(SSIM)of over 0.96 and a Mean-Square Error(MSE)of 0.035%to actual flow fields,reducing time costs by 7-8 orders of magnitude compared to Computational Fluid Dynamics(CFD)calculations.This can provide an efficient means for rapid computation of complex flow fields.展开更多
This study aims to investigate the intricate dynamic characteristics of the high-speed duct during the over-under Turbine-Based Combined Cycle(TBCC)inlet mode transition process while operating in an off-design state ...This study aims to investigate the intricate dynamic characteristics of the high-speed duct during the over-under Turbine-Based Combined Cycle(TBCC)inlet mode transition process while operating in an off-design state under throttled conditions.A typical over-under TBCC inlet,designed for a working Mach number range of 0–6 with a transition Mach number of 3.5,is examined through experimental studies in a supersonic wind tunnel with a freestream Mach number of2.9.The investigation focuses on the complex oscillatory flow and unique hysteresis observed in the mode transition process of the high-speed duct under the mildly throttled condition,utilizing highspeed schlieren and dynamic pressure acquisition system.The findings reveal that the high-speed duct undergoes four distinct oscillation stages akin to those in a higher throttled state during the mode transition,albeit with smaller dominant frequency and energy.Moreover,an irregular alternating“big/little buzz”mode is observed in the early stage of the large oscillation stage.Notably,the mildly throttled state exhibits three intriguing hysteresis properties compared to the unthrottled and higher throttled states.Firstly,hysteresis is observed in the shock train motion stage in the duct before unstart,along with the corresponding inverse process.Subsequently,hysteresis is noted in the unstart and restart of the high-speed duct,with a smaller hysteresis interval than in the unthrottled state.Finally,the hysteresis characteristics of oscillation mode switching and the corresponding inverse process are explored.Based on the analysis,the first two hysteresis phenomena are associated with the formation and dissipation of the separation bubble.The significant adverse pressure gradient constrains the cross-sectional capacity of the channel,rendering the disappearance of the separation bubble more challenging.The hysteresis in oscillation mode switching is linked to not only the channel cross-sectional capacity but also the state of the incoming boundary layer.展开更多
The utilization of Inlet Guide Vane (IGV) plays a key factor in affecting the instability evolution. Existing literature mainly focuses on the effect of IGV on instability inception that occurs in the rotor region. Ho...The utilization of Inlet Guide Vane (IGV) plays a key factor in affecting the instability evolution. Existing literature mainly focuses on the effect of IGV on instability inception that occurs in the rotor region. However, with the emergence of compressor instability starting from the stator region, the mechanism of various instability inceptions that occurs in different blade rows due to the change of IGV angles should be further examined. In this study, experiments were focused on three types of instability inceptions observed previously in a 1.5-stage axial flow compressor. To analyze the conversion of stall evolutions, the compressor rotating speed was set to 17 160 r/min, at which both the blade loading in the stator hub region and rotor tip region were close to the critical value before final compressor stall. Meanwhile, the dynamic test points with high-response were placed to monitor the pressures both at the stator trailing edges and rotor tips. The results indicate that the variation of reaction determines the region where initial instability occurs. Indeed, negative pre-rotation of the inlet guide vane leads to high-reaction, initiating stall disturbance from the rotor region. Positive pre-rotation results in low-reaction, initiating stall disturbance from the stator region. Furthermore, the type of instability evolution is affected by the radial loading distribution under different IGV angles. Specifically, a spike-type inception occurs at the rotor blade tip with a large angle of attack at the rotor inlet (−2°, −4° and −6°). Meanwhile, the critical total pressure ratio at the rotor tip is 1.40 near stall. As the angle of attack decreases, the stator blade loading reaches its critical boundary, with a value of approximately 1.35. At this moment, if the rotor tip maintains high blade loading similar to the stator hub, the partial surge occurs (0° and +2°);otherwise, the hub instability occurs (+4° and +6°).展开更多
To tackle the instability fault diagnosis challenges in wide-speed-range supersonic inlets,this study proposes an inlet fault decision fusion diagnosis algorithm based on attention mechanism feature fusion,achieving e...To tackle the instability fault diagnosis challenges in wide-speed-range supersonic inlets,this study proposes an inlet fault decision fusion diagnosis algorithm based on attention mechanism feature fusion,achieving efficient diagnosis of instability faults across wide-speed regimes.First,considering the requirement for wall pressure data extraction in mathematical modeling of wide-speed-range inlets,a supersonic inlet reference model is established for computational fluid dynamics(CFD)simulations.Second,leveraging data-driven modeling techniques and support vector machine(SVM)algorithms,a high-precision mathematical model covering wide-speed domains and incorporating instability mechanisms is rapidly developed using CFD-derived inlet wall pressure data.Subsequently,an inlet fault decision fusion diagnosis method is proposed.Pressure features are fused via attention mechanisms,followed by Dempster-Shafer(D-S)evidence theory-based decision fusion,which integrates advantages of multiple intelligent algorithms to overcome the limitations of single-signal diagnosis methods(low accuracy and constrained optimization potential).The simulation results demonstrate the effectiveness of the data-driven wide-speed-range inlet model in achieving high precision and rapid convergence.In addition,the fusion diagnosis algorithm has been shown to attain over 95%accuracy in the detection of instability,indicating an improvement of more than 5%compared to the accuracy of other single fault diagnosis algorithms.This enhancement effectively eliminates the occurrence of missed or false diagnoses,while demonstrates robust performance under operational uncertainties.展开更多
An experimental investigation is conducted to evaluate the performance and the stalling process of a fan subjected to inlet swirls,as well as the effectiveness of an Impedance Boundary-Controlled(IBC)Casing Treatment(...An experimental investigation is conducted to evaluate the performance and the stalling process of a fan subjected to inlet swirls,as well as the effectiveness of an Impedance Boundary-Controlled(IBC)Casing Treatment(CT)on the stall margin recovery.An operating cycle is proposed based on the hysteresis effect of harmonic flap oscillation of airfoils and parallel compressor theory to explain the pressure characteristic of the fan under twin swirl inlets.Twin swirls are observed to reduce the stall margin of the fan,and the circumferential location where the spike is detected turns to the intersection area of the twin swirl.The IBC CT is proven to extend the stall margin of the fan for 12.7%–22.3%when subjected to inlet swirls with an efficiency loss of around 1%.The IBC CT helps to reduce the size of the operating cycle of the fan by redistributing the blade loading and adding the system damping to dissipate the perturbation energy.展开更多
An investigation on the ventral diverterless high offset S-shaped inlet is carried out at Mach numbers from 0.600 to 1.534, angles of attack from -4° to 9.4°, and yaw angles from 0° to 8°. Results ...An investigation on the ventral diverterless high offset S-shaped inlet is carried out at Mach numbers from 0.600 to 1.534, angles of attack from -4° to 9.4°, and yaw angles from 0° to 8°. Results indicate： （1） a large region of low total pressure exists at the lower part of the inlet exit caused by the counter-rotating vortices in the S-shaped duct; （2） the performances of the inlet at Mach number 1.000 reach almost the highest, so the propulsion system could work efficiently in terms of aerodynamics; （3） the total pressure recovery increases slowly at first and then remains unvaried as the Mach number rises from 0.6 to 1.0, however, it does in an opposite manner in the conventional diverter-equipped S-shaped inlet; （4） the performances of the inlet are generally insensitive to angles of attack from -4° to 9.4° and yaw angles from 0° to 8° at Mach number 0.850, and angles of attack from -2° to 6° and yaw angles from 0° to 5° at Mach number 1.534.展开更多
A new internal waverider inlet with a rectangular shape of entrance and exit in front view is designed at Ma=6.0.The design is based on a better basic flowfield ICFC than traditional one and derived with the technolog...A new internal waverider inlet with a rectangular shape of entrance and exit in front view is designed at Ma=6.0.The design is based on a better basic flowfield ICFC than traditional one and derived with the technology of stream tracing and shock cutting.Comparison between the newly designed inlet and a typical sidewall compression inlet is given.The design Mach number and entrance shape of this new inlet are chosen according to the sidewall compression inlet.Numerical results show that most of the performance parameters of the internal waverider inlet are a bit higher than the sidewall inlet,such as the flow capture coefficient,total pressure recovery and the kinetic efficiency.The performances of these two inlets at off-design points are compared.The internal waverider inlet can capture more than 91% of incoming flow under all simulated conditions.Results show that internal waverider inlet using 3-D compression and high flow capture coefficient is a kind of fixed-geometry inlet with better performance.展开更多
An experimental study of the flow in a helicopter inlet with front output shaft and partial flow dynamic head is conducted in low speed wind tunnel. The flow characters of the inlet in the range of the yaw angle from ...An experimental study of the flow in a helicopter inlet with front output shaft and partial flow dynamic head is conducted in low speed wind tunnel. The flow characters of the inlet in the range of the yaw angle from 0~135°are presented in this paper. The static pressure distributions along the duct, distortions of the flow field at the outlet section and total pressure recovery coefficients are measured and analyzed. The results show that this type of inlet has high total pressure recovery coefficients at a wide range of yaw angle. The regions of local flow separation and distortion are closely related to the yaw angle. It′s also found that the outlet section has the best characteristics at sideslip, and sharply deteriorated characteristics at the yawed flight with a yaw angle of more than 90°展开更多
The uniform design and response surface methodology （RSM） are applied to the multi-objective optimization of a 2-D mixed compression scramjet inlet. The set of experimental design points on the design space is selec...The uniform design and response surface methodology （RSM） are applied to the multi-objective optimization of a 2-D mixed compression scramjet inlet. The set of experimental design points on the design space is selected by the uniform design, and the inlet performance is analyzed by computational fluid dynamics （CFD）. Then complete quadratic polynomial response surface approximation models are constructed based on the performance analysis results and then used to replace theoriginal complex inlet performance model. The optimization is conducted using a multi-objective genetic algorithm NSGA-Ⅱ, and the Pareto optimal solution set is obtained. Results show that the uniform design and RSM can reduce the computational complexity of numerical simulation and improve the optimization efficiency.展开更多
文摘The internal flow fields within a three-dimensional inward-tunning combined inlet are extremely complex,especially during the engine mode transition,where the tunnel changes may impact the flow fields significantly.To develop an efficient flow field reconstruction model for this,we present an Improved Conditional Denoising Diffusion Generative Adversarial Network(ICDDGAN),which integrates Conditional Denoising Diffusion Probabilistic Models(CDDPMs)with Style GAN,and introduce a reconstruction discrimination mechanism and dynamic loss weight learning strategy.We establish the Mach number flow field dataset by numerical simulation at various backpressures for the mode transition process from turbine mode to ejector ramjet mode at Mach number 2.5.The proposed ICDDGAN model,given only sparse parameter information,can rapidly generate high-quality Mach number flow fields without a large number of samples for training.The results show that ICDDGAN is superior to CDDGAN in terms of training convergence and stability.Moreover,the interpolation and extrapolation test results during backpressure conditions show that ICDDGAN can accurately and quickly reconstruct Mach number fields at various tunnel slice shapes,with a Structural Similarity Index Measure(SSIM)of over 0.96 and a Mean-Square Error(MSE)of 0.035%to actual flow fields,reducing time costs by 7-8 orders of magnitude compared to Computational Fluid Dynamics(CFD)calculations.This can provide an efficient means for rapid computation of complex flow fields.
基金funded by the National Natural Science Foundation of China(Nos.12025202,U20A2070,12172175)the National Science and Technology Major Project,China(No.J2019-Ⅱ-0014-0035)+2 种基金the Postdoctoral Fellowship Program of CPSF,China(No.GZB20230970)the Science Center for Gas Turbine Project,China(Nos.P2022-C-II-002-001,P2022-A-II-002-001)the Young Scientific and Technological Talents Project of Jiangsu Association for Science and Technology,China(No.TJ-2021-052)。
文摘This study aims to investigate the intricate dynamic characteristics of the high-speed duct during the over-under Turbine-Based Combined Cycle(TBCC)inlet mode transition process while operating in an off-design state under throttled conditions.A typical over-under TBCC inlet,designed for a working Mach number range of 0–6 with a transition Mach number of 3.5,is examined through experimental studies in a supersonic wind tunnel with a freestream Mach number of2.9.The investigation focuses on the complex oscillatory flow and unique hysteresis observed in the mode transition process of the high-speed duct under the mildly throttled condition,utilizing highspeed schlieren and dynamic pressure acquisition system.The findings reveal that the high-speed duct undergoes four distinct oscillation stages akin to those in a higher throttled state during the mode transition,albeit with smaller dominant frequency and energy.Moreover,an irregular alternating“big/little buzz”mode is observed in the early stage of the large oscillation stage.Notably,the mildly throttled state exhibits three intriguing hysteresis properties compared to the unthrottled and higher throttled states.Firstly,hysteresis is observed in the shock train motion stage in the duct before unstart,along with the corresponding inverse process.Subsequently,hysteresis is noted in the unstart and restart of the high-speed duct,with a smaller hysteresis interval than in the unthrottled state.Finally,the hysteresis characteristics of oscillation mode switching and the corresponding inverse process are explored.Based on the analysis,the first two hysteresis phenomena are associated with the formation and dissipation of the separation bubble.The significant adverse pressure gradient constrains the cross-sectional capacity of the channel,rendering the disappearance of the separation bubble more challenging.The hysteresis in oscillation mode switching is linked to not only the channel cross-sectional capacity but also the state of the incoming boundary layer.
基金support of the National Natural Science Foundation of China(No.52322603)the Science Center for Gas Turbine Project of China(Nos.P2022-B-II-004-001 and P2023-B-II-001-001)+1 种基金the Fundamental Research Funds for the Central Universities,Chinathe Beijing Nova Program of China(Nos.20220484074 and 20230484479).
文摘The utilization of Inlet Guide Vane (IGV) plays a key factor in affecting the instability evolution. Existing literature mainly focuses on the effect of IGV on instability inception that occurs in the rotor region. However, with the emergence of compressor instability starting from the stator region, the mechanism of various instability inceptions that occurs in different blade rows due to the change of IGV angles should be further examined. In this study, experiments were focused on three types of instability inceptions observed previously in a 1.5-stage axial flow compressor. To analyze the conversion of stall evolutions, the compressor rotating speed was set to 17 160 r/min, at which both the blade loading in the stator hub region and rotor tip region were close to the critical value before final compressor stall. Meanwhile, the dynamic test points with high-response were placed to monitor the pressures both at the stator trailing edges and rotor tips. The results indicate that the variation of reaction determines the region where initial instability occurs. Indeed, negative pre-rotation of the inlet guide vane leads to high-reaction, initiating stall disturbance from the rotor region. Positive pre-rotation results in low-reaction, initiating stall disturbance from the stator region. Furthermore, the type of instability evolution is affected by the radial loading distribution under different IGV angles. Specifically, a spike-type inception occurs at the rotor blade tip with a large angle of attack at the rotor inlet (−2°, −4° and −6°). Meanwhile, the critical total pressure ratio at the rotor tip is 1.40 near stall. As the angle of attack decreases, the stator blade loading reaches its critical boundary, with a value of approximately 1.35. At this moment, if the rotor tip maintains high blade loading similar to the stator hub, the partial surge occurs (0° and +2°);otherwise, the hub instability occurs (+4° and +6°).
基金supported by the Na tional Natural Science Foundation of China(No.62373185)the National Key R&D Program of China(No.2023YFB3307100).
文摘To tackle the instability fault diagnosis challenges in wide-speed-range supersonic inlets,this study proposes an inlet fault decision fusion diagnosis algorithm based on attention mechanism feature fusion,achieving efficient diagnosis of instability faults across wide-speed regimes.First,considering the requirement for wall pressure data extraction in mathematical modeling of wide-speed-range inlets,a supersonic inlet reference model is established for computational fluid dynamics(CFD)simulations.Second,leveraging data-driven modeling techniques and support vector machine(SVM)algorithms,a high-precision mathematical model covering wide-speed domains and incorporating instability mechanisms is rapidly developed using CFD-derived inlet wall pressure data.Subsequently,an inlet fault decision fusion diagnosis method is proposed.Pressure features are fused via attention mechanisms,followed by Dempster-Shafer(D-S)evidence theory-based decision fusion,which integrates advantages of multiple intelligent algorithms to overcome the limitations of single-signal diagnosis methods(low accuracy and constrained optimization potential).The simulation results demonstrate the effectiveness of the data-driven wide-speed-range inlet model in achieving high precision and rapid convergence.In addition,the fusion diagnosis algorithm has been shown to attain over 95%accuracy in the detection of instability,indicating an improvement of more than 5%compared to the accuracy of other single fault diagnosis algorithms.This enhancement effectively eliminates the occurrence of missed or false diagnoses,while demonstrates robust performance under operational uncertainties.
基金supported by the National Natural Science Foundation of China(Nos.52306035 and 52325602)the Science Center for Gas Turbine Project,China(Nos.P2022-A-II-002-001 and P2022-C-II-003-001)+1 种基金the National Science and Technology Major Project,China(Nos.Y2022-II-0002-0005 and Y2022-II-0003-0006)the Key Laboratory of Pre-Research Management Centre,China(No.6142702200101)。
文摘An experimental investigation is conducted to evaluate the performance and the stalling process of a fan subjected to inlet swirls,as well as the effectiveness of an Impedance Boundary-Controlled(IBC)Casing Treatment(CT)on the stall margin recovery.An operating cycle is proposed based on the hysteresis effect of harmonic flap oscillation of airfoils and parallel compressor theory to explain the pressure characteristic of the fan under twin swirl inlets.Twin swirls are observed to reduce the stall margin of the fan,and the circumferential location where the spike is detected turns to the intersection area of the twin swirl.The IBC CT is proven to extend the stall margin of the fan for 12.7%–22.3%when subjected to inlet swirls with an efficiency loss of around 1%.The IBC CT helps to reduce the size of the operating cycle of the fan by redistributing the blade loading and adding the system damping to dissipate the perturbation energy.
基金National Basic Research Program of China (5130802)
文摘An investigation on the ventral diverterless high offset S-shaped inlet is carried out at Mach numbers from 0.600 to 1.534, angles of attack from -4° to 9.4°, and yaw angles from 0° to 8°. Results indicate： （1） a large region of low total pressure exists at the lower part of the inlet exit caused by the counter-rotating vortices in the S-shaped duct; （2） the performances of the inlet at Mach number 1.000 reach almost the highest, so the propulsion system could work efficiently in terms of aerodynamics; （3） the total pressure recovery increases slowly at first and then remains unvaried as the Mach number rises from 0.6 to 1.0, however, it does in an opposite manner in the conventional diverter-equipped S-shaped inlet; （4） the performances of the inlet are generally insensitive to angles of attack from -4° to 9.4° and yaw angles from 0° to 8° at Mach number 0.850, and angles of attack from -2° to 6° and yaw angles from 0° to 5° at Mach number 1.534.
文摘A new internal waverider inlet with a rectangular shape of entrance and exit in front view is designed at Ma=6.0.The design is based on a better basic flowfield ICFC than traditional one and derived with the technology of stream tracing and shock cutting.Comparison between the newly designed inlet and a typical sidewall compression inlet is given.The design Mach number and entrance shape of this new inlet are chosen according to the sidewall compression inlet.Numerical results show that most of the performance parameters of the internal waverider inlet are a bit higher than the sidewall inlet,such as the flow capture coefficient,total pressure recovery and the kinetic efficiency.The performances of these two inlets at off-design points are compared.The internal waverider inlet can capture more than 91% of incoming flow under all simulated conditions.Results show that internal waverider inlet using 3-D compression and high flow capture coefficient is a kind of fixed-geometry inlet with better performance.
文摘An experimental study of the flow in a helicopter inlet with front output shaft and partial flow dynamic head is conducted in low speed wind tunnel. The flow characters of the inlet in the range of the yaw angle from 0~135°are presented in this paper. The static pressure distributions along the duct, distortions of the flow field at the outlet section and total pressure recovery coefficients are measured and analyzed. The results show that this type of inlet has high total pressure recovery coefficients at a wide range of yaw angle. The regions of local flow separation and distortion are closely related to the yaw angle. It′s also found that the outlet section has the best characteristics at sideslip, and sharply deteriorated characteristics at the yawed flight with a yaw angle of more than 90°
文摘The uniform design and response surface methodology （RSM） are applied to the multi-objective optimization of a 2-D mixed compression scramjet inlet. The set of experimental design points on the design space is selected by the uniform design, and the inlet performance is analyzed by computational fluid dynamics （CFD）. Then complete quadratic polynomial response surface approximation models are constructed based on the performance analysis results and then used to replace theoriginal complex inlet performance model. The optimization is conducted using a multi-objective genetic algorithm NSGA-Ⅱ, and the Pareto optimal solution set is obtained. Results show that the uniform design and RSM can reduce the computational complexity of numerical simulation and improve the optimization efficiency.
