10 results about "AxSTREAM" patented technology

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.

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.

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.

This invention relates to the field of aerodynamic design technology for multi-stage axial compressors, and discloses a design method and system for passive stability enhancement of a multi-stage axial compressor through self-circulation. By calculating the aerodynamic performance of the multi-stage axial compressor and evaluating engine stability, the method identifies operating conditions where the multi-stage axial compressor has insufficient stability margin. Unsteady flow field simulations are then conducted under these conditions to determine the first stage stall. For the first stage rotor tip stall, a self-circulating casing is used for passive stability enhancement of the multi-stage compressor. The self-circulating casing structure is derived from sensitivity analysis of its design parameters, selecting the optimal design parameters to ensure an increase in the stability margin of the multi-stage axial compressor across the entire speed range. This satisfies the stability margin requirements of multi-stage axial compressors and avoids the problem of insufficient stability margin in high-load multi-stage axial compressors across the entire operating speed range.

This invention belongs to the technical field of compressor aerodynamic stability analysis. It proposes a deterministic learning-based method and system for analyzing the aerodynamic stability of axial compressors. Using pulsating pressure data during the aerodynamic instability development process of an axial compressor, a deterministic learning identification algorithm based on radial basis function neural networks is used to model the nonlinear dynamics of the compressor's internal flow field, thereby obtaining dynamic diagrams of normal and instability precursor states. Dynamic indices are extracted from these diagrams to construct dynamic feature vectors describing the compressor's aerodynamic stability. Based on the dynamic feature vectors corresponding to different operating states, the dynamic classification boundary between normal and instability precursor states is determined. The distance between the dynamic feature vector of the compressor data under test and the dynamic classification boundary is calculated to determine the compressor's operating state. This invention enables early warning of aerodynamic instability and has good interpretability and engineering application value.

The present application relates to the technical field of axial flow turbine, especially to a large expansion ratio high-efficiency axial flow turbine, which comprises a turbine hub and turbine blades, the turbine blade inlet tip wheel diameter is, the turbine blade outlet tip wheel diameter is, the turbine blade top axial chord length is, the relationship among the turbine blade inlet tip wheel diameter, the turbine blade outlet tip wheel diameter and the turbine blade top axial chord length is = 0~0.364, in this way, the traditional long blade pull wire design is abandoned, the blade height is limited, the blade root section is increased to reduce the root stress, the blade tip section area is greatly reduced, the intermediate section area is slightly smaller to improve the flow capacity, the wedge angle is optimized, the shock loss is reduced, the efficiency is improved, and the technical problem of the prior art that the turbine is designed as a long blade, the long blade needs to adopt a pull wire to ensure the strength, but the pull wire structure reduces the turbine flow area, thereby reducing the turbine efficiency and increasing the assembly difficulty of the axial flow turbine is solved.

The application relates to a design method for deriving and developing a downward flow path of an aero-engine axial flow compressor, which comprises the following steps: step one, downward pressing of compressor inlet and outlet hubs, namely, downward pressing of inner flow paths of the compressor inlet and outlet; step two, re-determination of the compressor hub flow path according to the downward pressing amount of the compressor inlet and outlet hubs; step three, re-determination of the compressor casing flow path while keeping the compressor flow path area unchanged; step four, re-determination of the design rotating speed of the compressor while keeping the compressor rotating speed inlet blade tip tangential velocity unchanged; step five, adjustment of the blade profile of each stage of rotor and stator blades while keeping the compressor stage pressure ratio, efficiency and radial distribution unchanged; step six, adjustment of the inlet pre-whirl angle of each stage while keeping the compressor rotor counterforce unchanged; and step seven, checking of the radial distribution of the diffusion factors of each stage of the compressor rotor and stator, and if the radial distribution exceeds the engineering applicable value, returning to step one, adjusting the downward pressing amount of the compressor inlet and outlet hubs and performing iterative optimization.

This application provides a support base for the stator vane adjusting ring of an ultra-large axial compressor. The support base is disposed between the housing and the adjusting ring to support the adjusting ring. The support base includes: a support frame disposed on the housing, with rollers disposed on the support frame and the rollers contacting the adjusting ring; and a fixing member passing through the support frame and connected to the housing to assemble the support frame onto the housing. This application, through the inclusion of a buffer assembly, allows the buffer assembly to undergo adaptive deformation under pressure when the housing expands. This deformation, combined with a buffer gap, absorbs and offsets the expansion, preventing excessive supporting stress due to thermal expansion differences. This effectively protects the thin-walled housing and adjusting ring of the ultra-large axial compressor from damage, while maintaining the concentricity of the adjusting ring and the housing.

A method for predicting leakage losses in axial compressor grates is disclosed. This method calculates the grate leakage flow rate based on the inlet and outlet aerodynamic parameters at the compressor's planar blade mid-diameter and the grate geometric parameters. Then, it calculates the compressor's mainstream loss caused by the grate leakage by comparing the grate leakage flow rate with the compressor's mainstream flow rate. Finally, it determines the influence range and radial distribution law of the grate leakage loss using the compressor's planar blade geometric parameters. This invention targets compressor stators with grate seals. During compressor design and analysis, it eliminates the need for experimental measurements. It predicts grate leakage flow losses using the aerodynamic parameters at the stator's mid-diameter and the geometric parameters of the grate cavity, considering the impact of the leakage flow on the compressor stator blade passages. This allows for the incorporation of the radial influence of grate leakage on compressor outlet performance into the compressor's flow path design and analysis, predicting the grate leakage flow rate, and the influence range and magnitude of the stator outlet loss caused by the grate leakage flow, thus achieving more refined flow path calculations for axial compressors.