2320results about "Jet-propulsion engine testing" patented technology

The invention provides an equipment supervision data real-time analysis method and system, and the method comprises the steps: calculating an energy value of a frequency band corresponding to a sub-band signal, determining the energy characteristics of the surface roughness and dust deposition distribution uniformity of a blade in each sub-band in a square sum form, and generating energy distribution vectors which are arranged according to a frequency band sequence; performing time-frequency analysis on the vibration signal by using the acquired real-time rotating speed information of the gas turbine, identifying a frequency component corresponding to the eccentric distance value of the rotor and the vibration frequency offset, and determining the amplitude of the frequency component; according to the classification result and the real-time rotating speed information, the comprehensive influence index of the material density change rate and the ash deposition layer thickness difference is calculated, the fault early warning level is determined through a weighted average method, and an early warning signal is generated; and according to the change trend of the early warning signal and the blade edge wear degree, a random forest algorithm is adopted to predict a development curve of the vibration frequency offset in the future target time, and real-time equipment supervision analysis data are obtained.

An engine for aeromechanical instability abatement includes a sensor configured to capture data from rotating blades of the engine system, a airflow effector device, and an engine controller. The engine controller is configured to control the airflow effector device according to a nominal schedule, detect, based on a signal from the sensor indicating a vibration amplitude of the rotating blades within a frequency band, an incipient instability condition, in response to the incipient instability condition being present, determine a modified control parameter for at least one of the airflow effector device, and control the airflow effector device according to the modified control parameter, deviating from the nominal schedule.

The invention provides an aero-engine state monitoring method and device based on sound and vibration fusion and a computer readable storage medium, and relates to the technical field of aero-engine state monitoring. The method comprises the following steps: acquiring signals respectively acquired by a vibration sensor and a sound sensor; respectively intercepting a vibration signal low-frequency component and a sound signal high-frequency component based on the frequency response characteristic difference of the sensor; performing normalization processing on the intercepted signal to eliminate amplitude difference; constructing a transition region through an interpolation method, and splicing the continuous mixed frequency spectrum to obtain a sound-vibration fusion spectrum; and establishing a health benchmark based on the normal state fusion spectrum, and realizing abnormity monitoring and alarm through characteristic difference analysis. Advantages and characteristics of a low-frequency band of the vibration sensor and a high-frequency band of the sound sensor are fully utilized, frequency response limitation of a traditional single sensor is broken through, deep fusion of sound and vibration signals is achieved, the effective detection frequency range is remarkably widened, and accuracy and robustness of recognition of early faults such as aero-engine blade breakage are improved.

The invention relates to the technical field of combustion monitoring, and discloses a combustion instability boundary prediction system based on complex environment condition analysis, and the system comprises a multi-source sensing module, a data fusion module, an anti-interference processing module, an instability feature library, an intelligent decision engine and a digital twinborn body. A dynamic environment interference compensation mechanism is constructed, spectrum correction parameters are configured according to different environment working conditions during real-time prediction of a combustion instability boundary, the problem of spectrum distortion caused by environment disturbance on combustion dynamic parameters is eliminated in real time, spectrum interference deviation of a combustion measurement position can be automatically sensed and corrected, and real-time prediction of the combustion instability boundary is realized. The accuracy of combustion instability sensitive feature extraction is guaranteed, the instability early warning error is reduced, a flame pose real-time tracking system is deployed, fuel control parameter self-adaptive adjustment is triggered immediately when it is detected that the pose is abnormal, autonomous correction of the combustion instability feature recognition position is achieved, and it is guaranteed that the feature capture space is accurately positioned.

The invention discloses a gas turbine high-temperature component fault early warning method, which comprises the following steps: S1, constructing and training to obtain an early warning model of a gas turbine high-temperature component for monitoring the performance change trend of the gas turbine high-temperature component; and S2, predicting the time series data of the real gas turbine parameters at the current moment by using the early warning model, and then judging the abnormal condition of the high-temperature part of the gas turbine by comparing the residual quantity of a predicted value and an actual value and using a preset threshold value. According to the method, normal data and abnormal data can be effectively distinguished, the sensitivity of early warning of faults of the high-temperature parts is improved, the dynamic threshold judgment technology based on the sliding time window is added in the prediction and use process of the early warning model, the method can well adapt to dynamic changes of the data, dependence on historical data is reduced, and the prediction efficiency of the early warning model is improved. Compared with a traditional Gaussian hypothesis method, errors are not required to obey specific distribution, and the robustness of the model in the real-time operation process is also improved.

A method for monitoring a rotating machine in order to detect a fault in a bearing, the method including acquiring, from the rotating machine, a vibration signal measured by a vibration sensor; determining a first-order spectrogram by first-order cyclostationary analysis of the vibration signal using a delta transform and spectral standardisation; determining a second-order spectrogram by second-order cyclostationary analysis of the vibration signal using averaged cyclic coherence, a delta transform and spectral standardisation; and detecting a vibration signature of the fault in the bearing on the basis of the first-order spectrogram and the second-order spectrogram.

The invention provides a multivariable decoupling control method and system for an air inlet system of a high-altitude simulation cabin, and the method comprises the steps: firstly building a precombustion chamber cavity thermodynamic differential equation, and building a state equation in combination with a regulating valve first-order inertia model and an engine model; constructing a decoupling matrix by using a differential flatness theory, and converting a pressure and temperature coupling system into an independent control channel; monitoring a system state in real time by using an extended state observer, uniformly estimating internal and external disturbances as total disturbances, and generating a compensation signal; the model prediction controller performs rolling optimization on future time domain control quantity according to deviation between a set value and an actual value, generates an initial instruction in combination with the compensation signal, and converts the initial instruction into an independent valve control signal through a decoupling matrix; the high / low-temperature airflow mass flow is controlled through the opening degree of the adjusting valve, the engine model outputs actual parameters to form closed-loop feedback, and the controller updates the control quantity in each cycle. The multi-variable strong coupling, the dynamic disturbance and the model uncertainty are effectively overcome, the control precision and the parameter robustness are both achieved, and the complex working condition requirements are met.

The invention relates to the technical field of aero-engines, in particular to an aero-engine thermal barrier coating working condition simulation experiment system which comprises a high-temperature thermal cycle module, a high-speed airflow scouring module, a high-pressure loading module, an atmosphere control module and a data monitoring and collecting module. The high-temperature thermal cycle module is used for realizing rapid heating and cooling so as to simulate a thermal environment; the high-speed airflow scouring module is used for generating high-speed airflow to simulate gas scouring; the coating performance can be accurately evaluated through the multi-working-condition high-precision simulation capability, research and development personnel can be assisted to optimize components and structures, coating defects and weak points can be rapidly and accurately found through strict working condition simulation in the aspect of quality detection, the product quality is guaranteed, the defective rate is reduced, and the method has good application prospects from cost benefit consideration. The test frequency and the part replacement cost caused by the coating problem in engine research and development are reduced, the research and development period is shortened, the performance and reliability of the aero-engine are greatly improved, the development of the aerospace industry is powerfully promoted, and a solid foundation is laid for advanced engine research and development.

The invention relates to a multi-dimensional high-resolution turbine blade heat-flow coupling field synchronous measurement method and system, and the method comprises the steps: obtaining a target spot image of known space distribution in a measurement space, and building a three-dimensional space mapping function; carrying out particle matching on the velocity field measurement original image based on a three-dimensional space mapping function, and reconstructing a three-dimensional particle trajectory by adopting a self-adaptive grid algorithm combined with machine learning to obtain a high-resolution three-dimensional velocity vector field; acquiring a temperature field measurement original image, performing three-dimensional space registration according to a three-dimensional space mapping function, further calculating voxel-level temperature distribution, and inverting a three-dimensional temperature field; and carrying out coupling analysis on the three-dimensional velocity vector field and the three-dimensional temperature field, and outputting a three-dimensional heat-flow coupling cloud picture. According to the invention, multi-dimensional and high-precision synchronous measurement of the temperature field and the velocity field in the complex flow field of the turbine blade is realized, the spatial resolution and dynamic response capabilities are remarkably improved, and a reliable experimental basis is provided for analysis and optimization of the cooling performance of the blade.

The invention discloses a creep test system of a turbine rotor blade film hole simulation piece and a test method of the creep test system. Comprising a clamping assembly used for fixing a test piece, a loading assembly used for applying a load to the test piece, a strain measuring assembly used for measuring the strain rate of the test piece, a heating assembly used for covering the test piece to form a heating channel, and a cooling assembly used for manufacturing and conveying cooling airflow to flow through the interior of the test piece, the strain measurement assembly is used for measuring the external temperature of the test piece and the internal temperature of the test piece, the control assembly is connected with the strain measurement assembly, the heating assembly, the cooling assembly and the temperature measurement assembly, and the heating assembly is used for manufacturing and conveying high-temperature airflow to flow through the heating channel so as to heat the outside of the test piece to a first set temperature; the cooling assembly is used for cooling the interior of the test piece to a second set temperature, and the heating assembly and the cooling assembly are used for forming a dynamic convection environment around the test piece.

The invention discloses an engine fault detection method, device and equipment based on a single sensor and a medium, and relates to the technical field of fault detection. The method comprises the following steps: acquiring an eddy current original signal and an optical fiber pulse signal of the optical fiber eddy current composite sensor; the optical fiber eddy current composite sensor is installed at the position of an aero-engine casing. Calculating a blade radial displacement signal according to the eddy current original signal, and calculating a blade circumferential displacement signal according to the optical fiber pulse signal; filtering the blade circumferential displacement signal to obtain a target circumferential displacement signal; resampling the radial displacement signal of the blade to obtain a time sequence blade tip clearance signal; and engine blade fault information is determined according to the target circumferential displacement signal and the time sequence blade tip gap signal. Thus, the eddy current original signal and the optical fiber pulse signal are directly measured through the single sensor to detect faults, the transmission path is shortened, the diagnosis sensitivity and accuracy and the measuring point utilization rate are improved, and the sensor arrangement complexity is reduced.

The present disclosure relates to a system and method for adjusting a tool with a plurality of segments for inserting into a cavity of a machine. The method includes determining an initial alignment measurement of at least a portion of the tool; comparing the initial alignment measurement to a target alignment measurement to determine an adjustment amount; and adjusting at least one of the plurality of segments based on the adjustment amount. The tool includes a plurality of segments moveably coupled, where one or more of the plurality of segments comprises a unique dimensional measurement affecting a fit tolerance.

The invention discloses a combustion oscillation on-line monitoring method and device based on recursive quantitative analysis. The method comprises the steps that dynamic signals near a combustion chamber are obtained; performing recursive analysis on the dynamic signals in a period of time window, reconstructing a phase space, calculating time delay and embedding dimensions, and constructing a recursive matrix to obtain a recursive plot; extracting features from the recursion plot to obtain recursion quantization features; taking the recursive quantization features as input variables of the neural network, and inputting the recursive quantization features into the trained neural network; and the neural network gives a judgment result of the combustion oscillation state, and triggers an early warning mechanism when an abnormality is detected. According to the method, recursive quantitative analysis and deep learning are combined, online real-time monitoring of the combustion oscillation state is achieved, the precision and real-time performance of combustion oscillation monitoring are improved, efficient recognition and early warning of combustion oscillation are achieved, operation of a combustion system is optimized, combustion stability and safety are improved, the operation and maintenance cost of equipment is reduced, and the working efficiency is improved. And the method has wide engineering application value.

The invention relates to the technical field of ultrasonic detection, and provides an aero-engine fuel control valve leakage amount ultrasonic detection method, which forms a coherent technical link through leakage ultrasonic signal acquisition, leakage signal separation, leakage feature extraction, leakage area positioning, offset compensation and cross validation of a sensor, and the links are mutually verified and optimized. A leakage signal is separated through a dynamic threshold value, the identification degree of a deep belief network is enhanced through multi-dimensional features, environment interference is compensated and corrected through working condition parameters, positioning deviation caused by single-link error accumulation in a traditional method is effectively restrained, accurate capturing and positioning of a micro leakage signal are achieved, and the method is suitable for large-scale popularization and application. The leakage thermodynamic diagram integrates leakage rate grading, coarse positioning area and micro leakage point information, the analysis and decision-making time of maintenance personnel is greatly shortened through a visual presentation mode, and rapid interpretation of fault positions and severity is achieved.

The invention belongs to the technical field of gas turbine blade test equipment, and relates to a multi-gas-path coordinated regulation high-temperature blade gas thermosetting rotation test system, which comprises a test shell, a transmission shaft, a rotation assembly, a high-temperature gas supply assembly, a cooling gas supply assembly and a detection assembly. The transmission shaft can be driven to rotate through the rotating assembly, centrifugal force load borne in the actual operation process is simulated, high-temperature gas is conveyed into the test shell through the high-temperature gas supply assembly, a high-temperature environment is provided for the tested turbine blade, and thermal load borne in the actual operation process is simulated; cooling gas is conveyed into the tested turbine blade through the transmission channel by the cooling gas supply assembly, the cooling environment in the actual operation process is simulated, and finally various data of the test are detected through the detection assembly. Therefore, the flow heat transfer characteristic and the strength characteristic of the turbine blade under the actual operation condition can be obtained as far as possible through tests.

The invention discloses a two-degree-of-freedom clamp capable of adjusting the bending moment and direction of a turbine blade and a test method. The two-degree-of-freedom clamp comprises a blade clamping component, an upper adjustable U-shaped head, a lower adjustable U-shaped head, an upper clamping rod and a lower clamping rod. The blade clamping component is in threaded connection with the upper adjustable U-shaped head and the lower adjustable U-shaped head, the other end of the upper adjustable U-shaped head and the other end of the lower adjustable U-shaped head are connected with the upper clamping rod and the lower clamping rod respectively, and the other end of the upper clamping rod and the other end of the lower clamping rod are connected with an upper clamping head and a lower clamping head of the fatigue testing machine respectively. The upper clamping rod and the upper adjustable U-shaped head are adjustably connected in pairs in the transverse direction intersecting with the vertical axis so as to adjust the eccentricity of the turbine blade load. The eccentric angle of the blade load can be adjusted by adjusting the rotation cycle of the blade clamping component. According to the clamp disclosed by the invention, by adjusting the eccentricity and the eccentricity angle during a turbine blade test, different bending moment directions and relative sizes of different sections during the turbine blade test can be realized, so that the aim of simulating a real load is fulfilled.

The invention belongs to the technical field of aero-engine high-temperature part test, and particularly relates to a thermal mechanical composite fatigue test device in a turbine blade rotation environment, which comprises a high-temperature gas system, a stator system, a rotor system, a support system, a test system, an auxiliary system and a turbine rotor blade test piece, the high-temperature fuel gas system is used for providing a fuel gas environment for a turbine rotor blade test piece and controlling the high-low temperature circulation state of fuel gas, a combustion chamber is arranged at the front end of the high-temperature fuel gas system and provided with a double-oil-way nozzle, a fuel gas flowing system is arranged at the rear end of the high-temperature fuel gas system, and a blade grid structure is designed for the fuel gas flowing system according to the geometrical shape of a turbine rotor blade. And a spray cooling section is arranged at the rear part of the test section. The device is composed of six core systems including a high-temperature fuel gas system, a stator system, a rotor system, a supporting system, a testing system and an auxiliary system, all the systems work cooperatively, and integrated simulation of the high-temperature fuel gas environment, rotating centrifugal loads and heat-force synchronous circulation is achieved.

An inspection method is provided during which a head of an inspection scope is inserted into an interior of a powerplant. The head of the inspection scope includes an actuator. The powerplant includes a component located within the interior of the powerplant. The head of the inspection scope is abutted against a surface of the component within the interior of the powerplant. The head of the inspection scope is preloaded against the surface of the component using a preload device located outside of the interior of the powerplant. Vibrations are induced in the component using the actuator. A vibratory response in the component excited by the vibrations is measured using a sensor to provide sensor data.

The invention discloses an engine fault diagnosis method, device and equipment and a storage medium, and relates to the technical field of fault diagnosis. The method comprises the following steps: synchronously acquiring a sound signal and a vibration signal of an engine in a normal operation state, and performing feature extraction on the sound signal and the vibration signal to obtain corresponding target features; the target features comprise time domain features, frequency domain features and time-frequency domain features; determining a statistical index of each target feature through statistical characteristic analysis, and performing feature fusion on all the target features according to a target fusion mode based on the statistical indexes to obtain a fusion index; and an early warning threshold value is determined according to the data statistical characteristics of the fusion index, so that fault monitoring is carried out on the engine by using the early warning threshold value. Fault diagnosis is carried out according to the early warning threshold value determined according to the sound vibration data of the engine in the normal operation state, high dependence on fault samples is avoided, and the comprehensiveness of fault analysis is improved.

The invention discloses a turbine for testing, an engine testing device and a method for determining a pneumatic instability boundary of a compression system, which can determine the pneumatic instability boundary of the compression system in a complete machine environment, do not cause extra burden to the turbine and compression system components, and do not need to modify a tester and surrounding equipment on a large scale. The turbine for testing is used for replacing an original turbine in a complete machine surge test of an engine, and the engine testing device is used for the complete machine surge test of the engine. Stator blades of the test turbine are arranged to be rotatable blades, so that the circulation sectional area of the test turbine is reduced by rotating the stator blades, and the compression system can enter a surge state.

The invention provides an aero-engine test bed. The aero-engine test bed comprises a rack and an auxiliary support, the auxiliary support comprises a mounting platform, a mounting seat, a connecting piece, a supporting arm, a field force limiting mechanism and a controller; the mounting platform is arranged on the rack; the two mounting seats are oppositely arranged on the mounting platform; the connecting piece is arranged between the two mounting seats and is opposite to the mounting platform; the connecting piece is configured to move in the vertical direction; the supporting arm corresponds to the mounting seat; the two ends of each supporting arm are hinged to the corresponding mounting base and the corresponding connecting piece correspondingly. The supporting arm is provided with an installation joint used for being connected to an aero-engine. The field force limiting mechanism comprises a first component and a second component which are oppositely arranged; the first component is arranged on the mounting platform, and the second component is arranged on the connecting piece; and the controller is electrically connected with the field force limiting mechanism. Through cooperative work of the field force limiting mechanism and the controller, the supporting rigidity of the auxiliary support can be dynamically adjusted according to the real-time rotating speed of the aero-engine in the test run process.

The present invention provides a design method for a dual-duct fan mixed exhaust test device, comprising the following steps: Step 1, determining the axial position of the outlet of the intermediate casing, the axial length of the mixed exhaust, and the radial height of the outlet of the mixed exhaust; Step 2, determining the shape of the mixed exhaust inner flow path based on the results of Step 1; Step 3, determining the shape of the inner-outer flow path according to the expansion ratio based on the shape of the mixed exhaust inner flow path in Step 2; Step 4, determining the shape of the mixed exhaust outer flow path and the mixing intersection point of the inner and outer ducts according to the outlet extension shape of the mode selection valve. The present invention can achieve the performance measurement of the dual-duct fan variable duct ratio test, improve the overall performance measurement of the fan component under variable duct ratio conditions, and has important guiding significance for the variable cycle whole machine test research.

An inspection system detects defects within a component using eddy current thermography. An infrared camera detects heat from a defect within the component generated by the eddy current interfering with the defect. The eddy current is generated in the component by a magnetic field applied to the component by an inductor assembly. The inductor assembly includes a magnetic core having magnetic core material and two arms, each arm including a coil to generate the magnetic field. The inductor assembly is configured to apply the magnetic field at an angle offset from a horizontal axis of the magnetic core such that the eddy current flows at an angle within the component relative to the offset angle. Further, the magnetic core material can be shaped to apply the magnetic field at the offset angle or to accommodate complex component shapes.

A design method for a blade with orientation structures, a blade and a performance test method for a blade are provided, which relate to the technical field of blade design, and are universal to blades whose parameters are completely or partially known. In the design method, orientation awls are provided at a trailing edge of a blade from a variable-pitch propeller; orientation awls are provided at a trailing edge of a blade from a fixed-pitch propeller or a fan blade from a jet engine, and orientation grooves are provided on a back of the blade. Meanwhile, during providing orientation awls for either blade, the blade is stretched in a chord direction to ensure that an area of a pressure surface of a modified blade is equal to an area of a pressure surface of its original.

The invention relates to a combustion oscillation fault rapid detection method and device, and the method comprises the steps: obtaining a key signal of a combustion chamber in a combustion process, and determining a plurality of combustion states; the key signal comprises time sequence data; performing symbol segmentation on the time sequence data based on a maximum entropy segmentation method to obtain a symbol time sequence; obtaining a reference state according to the symbol time sequence and the plurality of combustion states; determining a current combustion state in the plurality of combustion states according to the reference state; when the current combustion state is a combustion oscillation mode, an early warning mechanism is started; according to the method, the time sequence data and the combustion state of the combustion chamber in the combustion process through the maximum entropy segmentation method can be processed, and the current combustion state is determined through the combustion state and the reference state, so that the combustion oscillation is monitored in real time through the current combustion state, complex sensor arrangement and algorithm processing are not needed, and the reliability is high. The cost is reduced, the calculation amount is reduced, and the early warning efficiency is improved.

The invention discloses a method, a device and equipment for monitoring the degradation state of a turboshaft engine and a medium, and aims at overcoming the defects of an existing method for sensing the degradation state of the turboshaft engine based on limited measurement parameters. A degeneration factor range linear solving method based on a degeneration factor fingerprint map theory is introduced on the basis of a traditional particle swarm optimization algorithm, and a more accurate and reasonable particle optimization limitation range is provided for the particle swarm optimization algorithm under the condition that available system input parameters are insufficient. For any sensor actual acquisition main performance parameter degradation situation combination, a corresponding specific optimization range can be provided, and the situation that degradation factor optimization falls into local optimization in a plurality of possible optimization spaces due to insufficient data complexity is avoided. Therefore, degradation state monitoring of the turboshaft engine is optimized, and the overall performance degradation trend of the engine is accurately obtained.

The utility model belongs to the technical field of high-temperature flow field pressure testing, and discloses a water-cooling pressure probe for measuring a turbine interstage three-dimensional dynamic high-temperature flow field, which is characterized by comprising a probe head, a support rod, a mounting seat and a lock cover. The probe head comprises a cylinder and a cylinder beveled body which share the same bottom surface, and two dynamic pressure sensors are packaged in the probe head. The diagonal plane of the cylindrical beveling body and the side face of the cylinder are each provided with a pressure sensing hole, and the pressure sensing holes are communicated with a pressure sensor. A plug board with a vortex generating device is arranged in the probe head and the support rod, so that heat convection between cooling water and the pressure measuring channel is enhanced. Three-dimensional flow parameters such as a pitch angle, a deflection angle, total pressure, static pressure and Mach number in a turbine interstage flow field can be measured through calibration of a calibration wind tunnel. Compared with an existing probe, the probe can realize flow field measurement in a high-temperature environment, prevents a sensor at the head of the probe from temperature drift and even damage due to heating, and further ensures the accuracy of measured data.

The invention relates to the technical field of aero-engine rotor vibration tests, in particular to an aero-engine rotor unsteady-state thermally induced vibration test system which comprises a rotor experiment table, an air inlet and exhaust system, a lubricating oil system, a protection system and a heating system. The rotor experiment table comprises a first supporting system, a second supporting system, a rotor outer protection casing, a gas compressor disc, a gas compressor sealing disc, a turbine disc, a turbine front sealing disc, a front shaft diameter, an inter-disc drum, a rear shaft diameter, a first squirrel-cage elastic support, a ball bearing, a second squirrel-cage elastic support and a rolling rod bearing. The first supporting system is composed of a first squirrel-cage type elastic support and a ball bearing, meanwhile, an air inlet is formed in the first supporting system, and the ball bearing in the first supporting system is connected with the front shaft diameter in an interference fit mode. Rotor unsteady-state characteristics considering heating conditions are finally created, and vibration response and durability of a wheel disc structure under the unsteady-state thermal load condition are studied in combination with test conditions such as temperature vibration and the like.

The invention discloses a micro-scale turbine blade boundary layer heat-flow double-field synchronous test method and system. The test method comprises the steps that temperature field initial data and velocity field initial data are acquired; based on the initial data of the temperature field and the initial data of the velocity field, the distribution state of the microparticles in the boundary layer is diagnosed, and according to the diagnosis result, the optimal distribution of the microparticles in the boundary layer is obtained by feeding back and controlling the operation parameters of the reverse cyclone separation system; based on the optimal distribution, an improved PIV algorithm is adopted to calculate a boundary layer velocity field, and the improved PIV algorithm is obtained by introducing multistage grid adaptive encryption, physical constraint conditions and a transition area recognition mechanism into the PIV algorithm; a boundary layer temperature field reconstruction model is established, depth phase coupling is carried out on a temperature field reconstruction result and a boundary layer velocity field, synchronous analysis of heat-flow coupling behaviors in a turbine blade boundary layer is achieved, and finally a boundary layer coupling cloud picture reflecting heat-flow interaction characteristics is obtained.