169 results about "AxSTREAM" patented technology

The invention discloses an evtol tilt rotor aerodynamic noise wind tunnel test method, and belongs to the technical field of wind tunnel tests. In order to establish an evtol tilt rotor aerodynamic noise wind tunnel test method, the method comprises the following steps: carrying out a hovering empty propeller hub background noise test: rotating a rotor table, enabling the axis of a rotor to be perpendicular to the axis of a wind tunnel, carrying out an empty propeller hub aerodynamic and noise test, and then carrying out aerodynamic data acquisition and noise data acquisition at each rotating speed to obtain hovering empty propeller hub background noise data; performing a tilt rotor hovering noise test: installing blades, setting a rotor axis to be perpendicular to a wind tunnel axis, then performing pneumatic data acquisition and noise data acquisition at each total pitch and rotating speed, and correcting the obtained tilt rotor hovering noise data by using hovering empty propeller hub background noise data to obtain a tilt rotor hovering noise test result; obtaining static tension and static power under each total pitch and rotating speed of tilt rotor hovering; carrying out a transition state and axial flow forward flight empty propeller hub noise characteristic test; and carrying out a transition state and axial flow forward flight tilt rotor noise characteristic test.

The invention discloses an integrated safety monitoring control method and system for a jet flow axial flow fan, and relates to the field of fan monitoring. The method comprises the steps that fan vibration data of a current fan are collected in real time through a vibration sensor, loosening state data of fan hardware are detected in real time through a displacement sensor, and motor temperature data are collected in real time through a temperature sensor; waveform analysis is carried out on the vibration data to obtain abnormal vibration characteristics, spectral analysis is carried out on the loosening state data to obtain a specific loosening part, and analysis is carried out on the motor temperature data to obtain a temperature abnormal part; carrying out correlation analysis on the abnormal vibration characteristic, the specific loose part and the temperature abnormal part to obtain an overall fault state of the fan; feature matching is conducted on the overall fault state of the draught fan and a preset fault feature library, and when the feature matching degree exceeds a preset matching threshold value, the fault type and the fault danger level of the current draught fan are determined. By implementing the method, the accuracy of fault diagnosis can be improved.

The invention provides a multi-airfoil combined axial flow fan blade design method and system based on CST parameterization, and relates to the technical field of axial flow fan blade optimizing.The multi-airfoil combined axial flow fan blade design method comprises the following steps that S1, a plurality of characteristic sections are arranged in the radial direction of a blade, and the airfoil corresponding to each characteristic section is determined; s2, geometric modeling is conducted on the airfoils of all the characteristic sections through a CST parameterization method, and description parameter sets corresponding to all the airfoils are obtained; s3, performing aerodynamic performance simulation on the airfoils through the description parameter group to determine the attack angle of each airfoil; s4, the airfoils are combined at the corresponding attack angles, smooth transition cambered surfaces are generated between the airfoils, and an initial blade model is obtained; and S5, optimizing the initial blade model to obtain a target blade model by adjusting the description parameter group of each airfoil by taking the maximization of the lift-drag ratio as a target. Compared with a traditional single arc thin plate blade profile design, the pneumatic performance of the blade is more optimized, and the problem that a single blade profile is difficult to cope with complex pneumatic conditions where different sections are located is effectively solved.

The invention discloses an experiment table system for an axial flow compressor of an aero-engine, and aims to break through the bottleneck by designing single-stage test and surge simulation, so that the experiment table system can measure experiment parameters such as pressure, flow, temperature and rotating speed of the axial flow compressor. Comprising an axial flow compressor suite, a sensor suite, a PLC, an asynchronous motor, an electric control valve and an exhaust device. A driving shaft of the asynchronous motor is connected with the axial flow compressor suite; the exhaust device is located at the rear end of the axial flow compressor suite, and the electric control valve is arranged on an airflow channel of the axial flow compressor suite and the exhaust device. The sensor suite is used for detecting pressure, temperature and air flow signals of the axial flow compressor and sending the collected signals to the PLC, and the PLC sends the signals to the industrial personal computer; the industrial personal computer controls the opening degree of a throttle valve of the electric control valve through the PLC, and the flow of air flowing through the axial flow air compressor is changed. The industrial personal computer controls the rotating speed of the asynchronous motor through the frequency converter.

The invention provides a one-dimensional pneumatic design method and device for an axial flow compressor, and belongs to the field of axial flow compressor pneumatic design. The method comprises the following steps: taking values of a plurality of design variables as the state of a reinforcement learning agent; the intelligent agent outputs the adjustment amount of the design variable according to the state, and the adjustment amount serves as the action executed by the intelligent agent; applying the action to the value of the current design variable to obtain an updated design variable; utilizing a one-dimensional inverse problem design program to generate a one-dimensional pneumatic design scheme of the gas compressor; utilizing a one-dimensional normal problem analysis program to output performance indexes; constructing a reward function according to the performance indexes; based on feedback provided by the reward function, a depth deterministic strategy gradient algorithm is adopted to train the intelligent agent; and performing one-dimensional pneumatic design of the gas compressor by using the trained intelligent agent. According to the method and device provided by the invention, the problems that the one-dimensional design of the axial flow compressor depends on experience, the efficiency is low, automatic optimization is difficult and the response is slow can be solved.

The embodiment of the invention provides a multistage axial flow turbine pneumatic optimization design method, system and equipment and a storage medium. The method comprises the steps that design parameters of the axial flow turbine are obtained; initializing the flow coefficient, the load coefficient, the rotating speed and the whole machine efficiency of the axial flow turbine; the enthalpy drop of each stage of the axial flow turbine is distributed, pneumatic parameters of each stage of stationary blades and movable blades are calculated, and structural parameters of each stage of stationary blades and movable blades are calculated according to the pneumatic parameters; inputting the aerodynamic parameters and the structural parameters into a pre-established efficiency evaluation model, and outputting to obtain evaluation efficiency; and when the evaluation efficiency is converged and optimal, outputting a design result. According to the embodiment of the invention, by introducing the efficiency evaluation model, the design result and the design scheme are prevented from excessively depending on the traditional model, and the accuracy of the optimization design result is improved; in the optimization process, poor design schemes are quickly screened out, the defect that the global search speed of a genetic algorithm is low is overcome, the optimization design speed is effectively increased, and the method has wide applicability to different engineering problems.

The invention provides an axial flow compressor integrated aerodynamic design method based on reinforcement learning, and belongs to the field of compressor design, and the method comprises the following steps: in combination with a depth deterministic strategy gradient algorithm, carrying out one-dimensional inverse problem-based design and one-dimensional forward problem analysis in sequence on the basis of a given compressor index parameter to obtain a geometric parameter meeting a design index; through-flow inverse problem calculation is conducted on the given design parameters of the meridian flow surface of the gas compressor based on a streamline curvature method, aerodynamic parameters on the meridian flow surface of the gas compressor are obtained, and a three-dimensional model of the gas compressor is constructed; two-dimensional aerodynamic parameter checking is conducted on the three-dimensional model of the gas compressor through meridian flow surface / blade-to-blade flow surface direct problem analysis, and design parameters are verified through three-dimensional CFD numerical simulation; and adopting a proxy-assisted particle swarm optimization algorithm based on a committee to carry out multiple rounds of iterative optimization, and outputting an optimal three-dimensional model. According to the method, collaborative optimization of aerodynamic performance and geometric parameters is achieved, and a new technical approach is provided for intelligent aerodynamic design of the gas compressor.

The invention provides a design method of a multi-stage axial flow gas compressor without adjustable stationary blades, and belongs to the field of pneumatic design of multi-stage axial flow gas compressors, the design method of the multi-stage axial flow gas compressor without adjustable stationary blades comprises the following steps: step 1, selecting an average stage load coefficient of the multi-stage axial flow gas compressor; 2, according to the average stage load coefficient of the selected multi-stage axial flow compressor, the stage load coefficient of the first stage is determined; 3, according to the results of the step 1 and the step 2, selecting the average Mach number of a first-stage stator blade inlet, and calculating the radial Mach number of a first-stage stator according to the average Mach number of the first-stage stator blade inlet; 4, determining the relative chord-span ratio of the first-stage stator according to the radial Mach number of the first-stage stator in the step 3; and 5, determining the large negative attack angle of the first-stage stator blade according to the relative chord-span ratio of the first-stage stator.

An axial flow compressor rapid prediction method based on self-adaptive correction comprises the following steps that a one-dimensional performance rapid prediction model of an axial flow compressor is constructed based on compressor geometric parameters and a one-dimensional average pitch diameter theory; based on a one-dimensional performance rapid prediction model, taking a relative error between a calculation result and experimental performance as an optimization target, introducing a correction factor into a lag angle and total pressure loss empirical model as an optimization parameter, and carrying out optimization of the correction factor; establishing a mapping relation among the compressor inlet speed coefficient, the inlet relative Mach number and the optimal correction factor under different working conditions; based on an existing compressor experimental performance result, a correction factor optimization method is adopted, a mapping relation between the multiple compressor inlet working condition parameters and the optimal correction factor is constructed, and a data set is formed; a BP neural network is utilized to train a data set, so that the data set has generalization ability; and calling the trained BP neural network in the one-dimensional performance rapid prediction model to realize adaptive correction of the empirical model and rapid prediction of the performance.

The invention relates to the technical field of aero-engine design, in particular to an axial flow compressor-diffuser coupling design method and an ultra-compact coupling structure.The axial flow compressor-diffuser coupling design method comprises the steps that flow channels corresponding to a rotor and a stator of an original compressor are expanded to obtain a target compressor; performing expansion processing on the original diffuser to obtain a target diffuser; and an ultra-compact coupling structure of the axial flow compressor-diffuser is obtained. According to the axial flow compressor-diffuser, the expansion degree of an airflow channel of the compressor is increased, so that the airflow circulation area is increased, the actual axial length of the target diffuser is remarkably shortened while the speed reduction and diffusion performance of the diffuser is guaranteed, and then the coupling structure of the axial flow compressor-diffuser is more compact; and meanwhile, the adaptive capacity and the matching relation of the diffuser to the air flow at the outlet of the air compressor are optimized.

The invention discloses an axial flow compressor blade anti-erosion optimization method facing a sand and dust environment, and belongs to the technical field of compressor blade anti-erosion design and optimization. According to the method, on the basis of a gas-solid two-phase flow erosion mechanism in a sand and dust environment, links such as working condition analysis, parameterization platform building, constrained multi-objective optimization method design, model building and optimization and index verification are integrated, and an integrated anti-erosion optimization process is formed. CFturbo, AutoGrid, CFX and Matlab are called in a unified mode through a Python script, automatic coupling of geometric modeling, grid division, gas-solid coupling simulation and an intelligent optimization algorithm of the blade is achieved, and collaborative optimization of aerodynamic performance and erosion resistance can be achieved under the constraint of erosion wear. Compared with a traditional method which depends on experience modification or single simulation, the method has the advantages that the optimization efficiency is remarkably improved while the simulation precision is guaranteed, the personal error risk is reduced, and the method can flexibly adapt to different compressor models and diversified operation conditions.

The invention discloses a multi-target and multi-component integrated optimization method and equipment for a counter-rotating axial flow fan, and belongs to the technical field of turbomachinery optimization design. According to the method, parametric modeling is carried out on multiple components such as two rows of blades, an inlet current collector and an outlet diffuser of the contra-rotating axial flow fan, geometric parameter combinations of all the components are selected as multi-dimensional decision variables, and a multi-dimensional decision space is determined; obtaining decision sample sets according to the multi-dimensional decision space, solving a plurality of aerodynamic performances of the counter-rotating fan corresponding to different decision sample sets through numerical simulation, and forming a multi-target aerodynamic sample set; combining a genetic algorithm to optimize a weight and a threshold to train a neural network agent model; performing global search and nonlinear optimization through a genetic algorithm based on a neural network agent model by taking the optimal multi-target aerodynamic performance as an optimization target, and outputting a Pareto optimal solution set; and determining a global optimal solution and carrying out numerical simulation verification to obtain an optimized geometric model of the counter-rotating fan. The pneumatic performance of the contra-rotating axial flow fan can be effectively improved.

The invention relates to a batch automatic design method and system for turbine end wall modeling, belongs to the technical field of turbine mechanical design, and provides non-axisymmetric end wall modeling design for an axial flow turbine by taking a trigonometric function and a Bezier curve as axial modeling functions and taking a sine function as a circumferential modeling function. Meanwhile, the non-axisymmetric end wall curved surface is accurately drawn through programming, and a point cloud file of the three-dimensional curved surface is generated. And then the point cloud file is imported into three-dimensional modeling software for three-dimensional curved surface modeling, and automatic batch design of the non-axisymmetric end walls of the axial flow turbine is achieved. The end wall near the pressure surface is convex, and the end wall near the suction surface is concave, so that the airflow of the end wall near the pressure surface is accelerated to flow, the static pressure is reduced, the airflow of the end wall near the suction surface is decelerated, the static pressure is increased, and the non-uniformity of circumferential distribution of the static pressure in the end area is reduced; flow separation among the blades is effectively improved, vortex shedding is reduced, pressure loss is reduced, and turbine performance and efficiency are improved.

The invention discloses an axial flow compressor blade profile optimization method based on main and auxiliary agent models and a related device, and relates to the field of compressor blade design, and the method comprises the steps: generating a plurality of design parameter groups based on to-be-optimized design parameters and a parameter adjustable range of a to-be-optimized axial flow compressor blade by using a data sampling algorithm, outputting a blade profile performance prediction value corresponding to each design parameter group by using the main and auxiliary combination agent model, screening a plurality of alternative design parameter groups based on a comparison result of each blade profile performance prediction value and an initial performance reference, and performing computational fluid dynamics simulation; and determining the alternative design parameter group corresponding to the optimal simulation performance parameter as a blade profile optimization parameter group. According to the method, the invalid risk of the design parameter group is reduced by using the adjustable range of the parameters, the performance prediction precision is improved based on the flow field characteristic parameters by using the main and auxiliary combination agent model, finally, the blade profile optimization parameter group is determined based on the simulation result of the screened alternative design parameter group, and the prediction precision is further improved.

The purpose of the present invention is to provide a multi-dimensional coupled design method for the through-flow aerodynamic layout of a gas turbine transition section-high pressure compressor. In the present invention, the transition section and the high-pressure compressor are regarded as a system for integrated coupled design, and full integration with the upstream transition section is achieved in all dimensions and links of the high-pressure compressor's aerodynamic design. This realizes the integrated and collaborative design of the through-flow layout of the transition section and the high-pressure compressor, enabling the aerodynamic design of the transition section-high pressure compressor system in different dimensions to enter a stage of systematization, parameterization and refinement. It effectively improves the aerodynamic performance of the high-pressure compressor, enhances the accuracy of aerodynamic design, saves a large amount of design iteration time, shortens the design cycle, and is highly suitable for engineering design applications. The present invention is not limited to the high-pressure compressor of gas turbines; it is also applicable to the aerodynamic design process of aero-engine high pressure compressors and various industrial axial compressors with transition through-flow structures.

The invention relates to the technical field of axial flow turbine aerodynamic configuration design, in particular to a turbine cascade runner and blade modeling method based on an NURBS curved surface, and the turbine cascade runner and blade modeling method comprises the steps that the geometrical shape of an original cascade runner is generated; extracting the runner section of the original cascade runner; representing the original cascade flow channel as a flow channel control parameter; and generating a target flow channel geometric shape. According to the invention, through rich and limited flow channel control parameters, the accuracy and flexibility of flow channel geometric control are improved, so that the previous design difficulty is reduced. By utilizing the advantages of the NURBS curved surface, the smooth continuity of the flow channel curved surface can be realized under the condition of not introducing additional constraint conditions; the design process is simplified, it is ensured that the aerodynamic performance can be effectively improved when the flow channel is optimized, and potential problems caused by complex constraints are avoided.

The utility model relates to the technical field of steam power generation turbines, in particular to a mixed flow type saturated steam power generation turbine which comprises a cylinder, the cylinder comprises a vertical inlet cavity and a horizontal outlet cavity which are communicated, and reinforcing ribs are arranged at the positions, corresponding to the joint of the two cavities, of the inner wall of the cylinder. A protruding structure at the bottom of the side, away from the horizontal outlet cavity, of the outer wall of the air cylinder is fixedly connected with a gearbox. The stationary blade assembly comprises a first-stage radial flow stationary blade fixedly arranged on the inner wall of the vertical inlet cavity and a second-stage axial flow stationary blade fixedly arranged on the reinforcing ribs; the rotor assembly comprises a rotating part connected with the high-speed shaft of the gear box, a first-stage radial flow vane impeller and a second-stage axial flow vane impeller, wherein the periphery of the rotating part is sleeved with the first-stage radial flow vane impeller and the second-stage axial flow vane impeller. The radial flow turbine and the axial flow turbine are combined to form the mixed-flow turbine, so that the working capacity of a single cylinder is improved, the power generation efficiency is improved, the cost is saved, and the unit size is reduced.

The application provides a high-reliability one-dimensional inverse problem aerodynamic design method of an axial turbine, which is used for providing a better design method for the turbine in a one-dimensional inverse problem design link. The technical scheme is as follows: on the basis of the original, non-simplified processing is performed on core design parameters in turbine dimensionless parameters, such as inverse force degree, load coefficient, flow coefficient and the like, new inverse problem design calculation is performed on the processing, and a new inverse problem calculation process is reconstructed according to new definition formula.

The application discloses a high-pressure-ratio transonic axial flow compressor, which comprises a hub, front-row rotor blades, rear-row rotor blades and stator blades; the rear-row rotor blades are arranged between the front-row rotor blades and the stator blades; a plurality of front-row rotor blades and a plurality of rear-row rotor blades are uniformly arranged on the hub along a circumferential direction; and the front-row rotor blades and the rear-row rotor blades one-to-one correspond to form tandem rotors. The application realizes shock wave supercharging through the front-row rotor blades, and the rear-row rotor blades complete large airflow deflection; meanwhile, the application adopts an adsorption type stator to reduce or prevent airflow separation in a blade channel, thereby significantly improving a compressor stage pressure ratio and improving aerodynamic performance of the compressor. The application can realize a higher stage pressure ratio, and in the case of high inlet Mach number, high load and large airflow deflection angle, the application overcomes the problem that a high stage pressure ratio and efficiency are difficult to realize due to the fact that flow separation in the compressor cannot be controlled in the prior art.

The invention discloses a measuring and calculating system and method for online analysis of the performance of an axial flow fan, and belongs to the technical field of environmental protection. According to the technical scheme, the measuring and calculating system for online analysis of the performance of the axial flow fan comprises a first annular pipe installed at an inlet flange of the axial flow fan, and a plurality of first measuring points are evenly distributed on the first annular pipe; the second ring pipe is mounted in front of the impeller casing, a plurality of second measuring points are uniformly distributed on the second ring pipe, and the pressure difference between the inlet of the fan and the front of the impeller casing is measured through the pressure guide pipes of the first ring pipe and the second ring pipe; through an innovative measurement and calculation mode, real-time precise monitoring of performance parameters of the axial flow fan is achieved, the air volume measurement precision is greatly improved, and the problems that traditional measurement lags behind, and devices are prone to being damaged are solved; by means of multi-component cooperation and data processing optimization, the system reliability is enhanced, visual data are provided for operators, fan operation parameters can be adjusted in time, and the fan operation efficiency is remarkably improved.

The invention discloses a variable geometry turbine guide vane lag angle determination method and system, and relates to the technical field of axial flow turbine pneumatic analys.The method comprises the steps that an original vane type nominal outlet geometric angle is corrected based on an outlet structure angle, a trailing edge turning angle and a trailing edge wedge angle of a variable geometry turbine guide vane, and a corrected vane type nominal outlet geometric angle is obtained; correcting based on the original variable geometry turbine guide vane initial relief angle and substituting into the corrected blade profile nominal outlet geometric angle to obtain the corrected initial relief angle of the variable geometry turbine guide vane; obtaining an undetermined corrected backward angle model based on the corrected initial backward angle and the aerodynamic parameters; optimizing an undetermined coefficient in the undetermined correction lag angle model based on a genetic algorithm to obtain an optimal coefficient, and substituting the optimal coefficient into the undetermined correction lag angle model to obtain a final correction lag angle model; and based on the related parameters of the to-be-measured variable geometry turbine guide vane and the final correction lag angle model, obtaining the lag angle of the to-be-measured variable geometry turbine guide vane. And the lag angle of the variable geometry turbine guide vane can be quickly and accurately determined.

The present invention relates to a tool for repairing the flow path of an axial flow casing of an aero-engine that is easy to assemble and its usage method. The tool for repairing the flow path of the axial flow casing of the aero-engine includes a first protective ring adapted to the second-stage stator blade mounting groove, a second protective ring adapted to the third-stage stator blade mounting groove, a spigot protective ring, and a plurality of plugs adapted to the stator blade mounting holes. The outer wall of the first protective ring is in small clearance fit with the inner wall of the second-stage stator blade mounting groove. The outer wall of the second protective ring is in small clearance fit with the inner wall of the third-stage stator blade mounting groove. The spigot protective ring is sleeved on the left-end spigot of the axial flow casing. During assembly, the first protective ring and the inner wall of the second-stage stator blade mounting groove, and the second protective ring and the inner wall of the third-stage stator blade mounting groove are adhesively fixed with glue. The plugs and the stator blade mounting holes are adhesively fixed with glue.

This application discloses a method for optimizing the hub profile of an axial flow fan blade. The method includes: real-time acquisition of fan operating conditions to obtain the current fan speed and outlet back pressure, forming an operating condition vector; inputting the operating condition vector into a pre-trained adaptive prediction model, which outputs the optimal hub profile parameter vector corresponding to the current operating condition; the optimal hub profile parameter vector is set as a dynamic variable related to the operating condition vector, and the optimal hub profile parameter is pre-obtained based on a multi-objective comprehensive performance evaluation function with blade passing frequency sound pressure level and fan aerodynamic efficiency as optimization objectives, through adaptive optimization under different operating conditions; and dynamically adjusting the hub profile of the physical axial flow fan according to the optimal hub profile parameter. Thus, real-time adaptive adjustment of the axial flow fan hub profile to operating conditions is achieved.

The invention relates to the technical field of design of centripetal turbines of micro gas turbines, in particular to a design method for a volute and a nozzle ring of a centripetal turbine of a micro gas turbine. One-dimensional design of a volute and a nozzle ring is carried out by adopting a near-zero vortex design method, after a three-dimensional model is established, three-dimensional flow analysis and structure optimization are carried out, and the blade profile design of the nozzle ring adopts an integer operation mode in the design of an axial flow plane cascade gas turbine under the framework of the near-zero vortex design. The required annular blade grid is obtained through shape-preserving transformation, the blade profile can achieve a better pneumatic effect, and the energy conversion rate is increased; ribs are additionally arranged at an outlet in the volute design, and the problem that static strength of the volute is insufficient possibly is solved. Due to the design of the nozzle ring and the volute, the overall design requirement of the centripetal turbine of the micro gas turbine is met, the aerodynamic performance of the turbine is further improved, and the flowing loss of airflow is reduced.

The invention discloses a supersonic axial flow turbine blade profile design method for a turbine pump, and belongs to the field of turbine blade design. According to the method, a Laval spray pipe is replaced by a supersonic static blade designed based on a characteristic line method, so that an accelerated supersonic flow surface can be matched with an annular channel of a movable blade inlet, and circumferential uniform airflow is provided for the turbine movable blade inlet; and therefore, the pneumatic loss generated by air inlet degree configuration of similar parts in the turbine moving blades is reduced. The blade profile of the ultrasonic quick-acting blade designed based on the characteristic line and the free vortex does not depend on design experience, and the stability of the performance of the designed movable blade can be guaranteed. The blade profiles of the supersonic static blades and the moving blades designed through the method remarkably improve the performance of a supersonic turbine in a turbine pump, the fuel carrying amount of a rocket is reduced, and then the thrust-weight ratio of the rocket is increased.

The invention relates to the technical field of aero-engine gas compressor design, in particular to a multilevel optimization design method for a multilevel axial flow gas compressor based on a genetic algorithm, which comprises the following steps: determining a one-dimensional optimization variable of the gas compressor according to a design index, and performing performance optimization by combining the genetic algorithm with a one-dimensional characteristic calculation model of the axial flow gas compressor. A one-dimensional design scheme of the gas compressor is obtained; and further, according to an equal circulation design principle, determining a two-dimensional optimization variable of the gas compressor, performing performance optimization by combining a genetic algorithm with the two-dimensional characteristic calculation model of the axial flow gas compressor, finally determining a two-dimensional design scheme of the gas compressor which is designed to reach the standard, and inputting the two-dimensional design scheme of the gas compressor into a three-dimensional modeling calculation model of the axial flow gas compressor. And obtaining a designed three-dimensional geometric model of the gas compressor, and carrying out three-dimensional numerical simulation verification on the scheme. According to the method, the gas compressor design process is shortened, the investment of designers is reduced, and the optimization design efficiency is improved.

This invention discloses a rotor blade of a segmented stacked axial compressor and its design method. The design method includes: determining the axial position of the stacking line; determining the circumferential position of the stacking line: the stacking line is S-shaped in the circumferential direction, and the stacking line includes four control points A, B, C, and D sequentially from the blade root to the blade tip along the height direction; the coordinates of the centroid positions of each basic airfoil section in the first interval from point B to point C conform to a specific formula; and determining the shape of the stacking line based on the determined axial and circumferential positions. This invention, by setting the stacking line of the rotor blade to be S-shaped in the circumferential direction, generates a reverse pressure gradient in the radial direction, causing the gas at both ends to migrate towards the blade center, thereby improving the flow in the end region. This reduces the influence of secondary flow in the blade tip region; and by setting a specific function, the circumferential offset of the stacking line can be controlled by constructing a simple equation, improving the computational efficiency of optimizing the airfoil.

The invention provides a multi-stage axial flow compressor stage flow coefficient design method based on inter-stage test data. The method comprises the following steps that firstly, compressor stages and experimental test parameters are defined; 2, according to the outer wall static pressure # imgabs0 # measured in the step 1, the stator front average static pressure P is calculated; 3, according to the total pressure, measured in the step 1, of the multiple measuring points, the stator front average total pressure # imgabs1 # is calculated; 4, according to the total temperature, measured in the step 1, of the multiple measuring points, the stator front average total temperature # imgabs2 # is calculated; step 5, determining a stator inlet and outlet axial speed correction coefficient # imgabs3 #; and step 6, determining a stage flow coefficient # imgabs4 # according to the step 2 to the step 5. According to the method, on the basis of the aerodynamic basic equation, the capacity of carrying out compressor interstage parameter matching analysis on the basis of interstage test data can be remarkably improved through the reasonable hypothesis and curve fitting conforming to the compressor aerodynamic basic principle.

The application provides a rotary ram shock wave supercharged gas turbine engine and belongs to the technical field of turbine engines, which comprises an outer casing and an inner casing, an outer channel is formed between the inner casing and the outer casing, a guide vane is installed in the outer casing and is used for guiding airflow, a first mechanism is installed in the inner casing and is used for supercharging and heating the airflow and doing work, a second mechanism is arranged on one side of the inner casing and is used for supercharging and heating the airflow for the second time. The application solves the problems of large windward area of the centrifugal compressor, many parts of the multistage axial flow compressor and low efficiency of small flow by using the rotary ram compressor, higher flow and pressure ratio are obtained in compact size, greater thrust is obtained under the condition of smaller outer size of the original engine, the outer size is reduced by about 32%, and the thrust is increased by 17.3% under the condition of the same fuel consumption rate when the engine is at rest.

The application aims to provide a compressor aerodynamic performance analysis method, an analysis system and a medium. The compressor aerodynamic performance analysis method comprises the following steps: obtaining a design value of the overall efficiency of the compressor; obtaining an over-limit value and a design value of the tip clearance of the rotor blade of any specified stage of the compressor; obtaining an efficiency difference value when the tip clearance is the over-limit value and the design value; calculating a third efficiency value when the tip clearance is the design value, so that the actual efficiency is equal to the difference between the third efficiency value and the efficiency difference value, and recording the outlet temperature calculation amount; obtaining the compressor outlet temperature calculation amount according to the outlet temperature calculation amount; and obtaining the actual overall efficiency of the compressor when the tip clearance of the rotor blade of the specified stage is the over-limit value according to the compressor outlet temperature calculation amount. The compressor aerodynamic performance analysis method can improve the efficiency and accuracy of analyzing the influence of the tip clearance change of the multistage axial flow compressor on the performance.