663 results about "Aero engine" patented technology

The application belongs to the field of turbine blade thermal barrier coating of an aero-engine. Specifically, it is a structure design, a preparation method and a prepared coating of a double-layer radiation-resistant thermal barrier coating with a high-reflection layer and a high-absorption layer. The coating comprises a high-absorption layer and a high-reflection layer. The high-absorption layer can absorb part of the energy of photon radiation and re-emit it into the environment, while relying on a cooling gas film to quickly remove heat. Part of the energy of the photon radiation is reflected by the high-reflection layer again to the high-absorption layer after passing through the high-absorption layer, and the process of absorption-emission-cooling gas film heat dissipation is repeated. Within a set temperature range, the absorption rate of the high-absorption layer is determined based on the temperature difference requirement between the inner and outer surfaces of the thermal barrier coating, and the high-absorption layer material is determined based on the absorption rate. The application greatly reduces the radiation heat flow penetrating the coating, avoids the direct heating of the high-temperature alloy by infrared radiation, and significantly improves the heat insulation capacity of the thermal barrier coating and the service life of the turbine blade.

The application discloses a kind of flight condition confusion under aero-engine fault diagnosis method based on conditional perception contrast learning, belongs to aero-engine fault diagnosis and deep learning technical field, first, vibration signal of aero-engine under multiple flight conditions is collected and segmented pretreatment is carried out;Enhanced one-dimensional residual convolutional encoder is then constructed to extract deep feature representation;Then, a conditional perception contrast learning loss function is designed to learn fault causal features;Then, a gradient reversal layer is introduced for domain adversarial training to make the features invariant to flight conditions;Then, frequency domain enhancement techniques are used to simulate engine speed changes;Consistency regularization loss is then designed to constrain the consistency of original and enhanced sample predictions;Finally, classification loss, contrast loss, adversarial loss, speed coding loss and consistency loss are jointly optimized to complete model training;The application effectively decouples fault causal features and condition pseudo-features, achieving robust aero-engine fault diagnosis performance in flight condition confusion scenarios.

This invention belongs to the field of oil supply, lubrication, and cooling technology for aero-engine bearings. It provides a method for temperature control and parameter optimization design of under-ring lubricated bearings. The method includes: generating sample data on total heat generation, bearing element temperature, and oil slingering temperature of the bearing through experiments or simulations based on bearing structural parameters and operating condition parameters; constructing a predictive model with the parameters as input and thermal-fluid performance as output; quantifying the influence of input parameters using global sensitivity analysis, screening key design parameters, and constructing a multi-objective optimization model with the objectives of minimizing bearing element temperature control accuracy and total heat generation; solving the model to obtain the Pareto optimal solution set, and determining the optimal combination of structure and oil supply parameters that satisfies the oil slingering temperature safety constraint and bearing thermal reliability requirements while achieving optimal overall performance. This invention is applicable to the refined design of high-speed bearings in rotating machinery such as aero-engines, enabling precise temperature control and coordinated energy consumption optimization.

This invention relates to the field of aero-engine maintenance engineering and quality assessment technology, and discloses an aero-engine overhaul assessment method based on a physical information Bayesian neural network. The method includes collecting multi-source heterogeneous data from maintenance and testing sites and converting it into key process indicator scores; extracting rotor dynamics, imbalance transmission, and empirical formula features to construct physical constraint regularization terms; establishing an initial Bayesian neural network containing deterministic and variational Bayesian layers; inputting the key process indicator scores into the network, and optimizing parameters by combining the physical constraint regularization terms with adaptive annealing and early stop mechanisms; performing multiple Monte Carlo samplings on the trained network to calculate the mean distribution of the output results and obtain distribution statistical characteristics; calculating the quality score, confidence interval, and attribution warning information based on the distribution statistical characteristics, and outputting an overhaul assessment report. This invention solves the problem that existing pure data models violate physical laws and cannot quantify confidence levels.

This invention relates to the field of aero-engine assembly technology, and in particular to an auxiliary device and method for installing elongated blind cavity adjustment pads, including a positioning ring assembly clamp and a push rod fixture. The positioning ring assembly clamp is used to clamp and insert the front positioning ring into the elongated blind cavity. By setting push rods, support rods, pull heads, sliders, positioning discs, several chuck connecting arms and chucks and their connection relationships in the positioning ring assembly clamp, the front positioning ring is clamped and fixed, preventing it from falling off during installation. The push rod fixture, through the setting of adapters, bushings and cones, achieves radial positioning of the low-pressure turbine shaft with the front positioning ring and the front end of the fan journal, thus installing the low-pressure turbine shaft in place. This device has an ingenious structural design, is simple to operate, highly efficient, and low in cost. It effectively solves the problem of front positioning ring detachment during aero-engine assembly, while reducing engine assembly cycle, improving assembly efficiency, and reducing assembly costs.

The application discloses a fuel step rapid injection device and injection method, and belongs to the field of aero-engine whole machine test, and aims at solving the problems of complex system, high cost, low injection precision and slow response. The device comprises a servo motor, a screw pair, an oil injector, first and second remote switch valves, first and second check valves and a safety valve. The servo motor drives the oil injector piston rod through the screw pair, the oil injector rod cavity is connected with the low-pressure fuel system through the second remote switch valve, and is connected with the engine fuel manifold through the first remote switch valve and the first check valve. A bypass connected with the fuel control device through the second check valve is arranged between the first check valve and the fuel manifold. In normal operation, the fuel control device supplies fuel to the fuel manifold through the second check valve; in step injection, the second remote switch valve is opened, the piston rod is extended to suck oil, then the piston rod is closed, the first remote switch valve is opened, and the piston rod is quickly retracted to inject fuel into the manifold. The application can be used in aero-engine surge test.

The present application relates to the technical field of aero-engine aerodynamic optimization and computer-aided design, and discloses a numerical optimization design method and structure for aero-engine low-pressure turbine blade leading edge modeling based on response surface modeling. The method realizes precise matching of Klebanoff stripe and large-scale vortex generation timing by adjusting flow or rotor speed to change wake injection angle, and promotes stripe impact and accelerates large-scale vortex breakup. Meanwhile, by constructing a blade leading edge modeling parameterization model and establishing a response surface proxy model for the combination of modeling design variable parameters and boundary layer velocity mean square deviation, the optimal design variable parameter combination is determined in the constraint domain with the maximum velocity mean square deviation in the boundary layer as the target. The present application effectively reduces the boundary layer momentum thickness and total pressure loss under the premise of ensuring constant blade load, improves the design efficiency and engineering applicability under low Reynolds number conditions, and does not need to introduce additional flow control structures, thus being simple in structure and easy to implement.

The present application relates to the technical field of complex equipment component repair, in particular to a device component performance and unit body maintenance level relationship mining method capable of effectively improving the use efficiency of an aero-engine, which first carries out expansion processing of repair samples, and then selects a support vector machine regression method which is better in the condition of small sample problems to solve the mapping relationship between the component performance before repair, unit body maintenance level and component performance after repair. Since a component is generally composed of multiple unit bodies, each component has multiple maintenance levels, and the mapping relationship between the component performance before repair, unit body maintenance level and component performance after repair is a many-to-one mapping relationship. In order to improve the accuracy of support vector machine regression, a hybrid kernel function method is used to optimize it, and a particle swarm algorithm is used to optimize the related parameters.

This invention belongs to the field of aero-engine vibration testing technology, and specifically relates to a method and system for aero-engine vibration assessment testing; it includes: determining the most severe vibration environment for the components; determining the engine rotor dwell speed and dwell time at the dwell speed; formulating a vibration assessment test spectrum based on the engine rotor dwell speed and dwell time; conducting a whole-engine test based on the most severe vibration environment for the components and referring to the vibration assessment test spectrum; generating an engine speed versus vibration curve based on the engine whole-engine test measurement data, and formulating component vibration limit values; conducting the component vibration assessment test on the whole engine, and the whole engine used for the component vibration assessment test simulates the maximum possible vibration environment and test temperature of the components across the entire operating cycle and speed range; making the installation and vibration environment more consistent with actual engineering conditions.

The application discloses a gas turbine, a gas turbine blade and a leaf-shaped film cooling hole structure of the gas turbine blade. The existing traditional cylindrical film cooling hole is reshaped into a leaf-shaped structure including a left half circular arc, a right upper half elliptical arc, a right middle half elliptical arc and a right lower half elliptical arc, so that the cooling gas can be divided into three paths when being sprayed out of the hole, and is sprayed out along the passage secondary flow same direction, the film hole center line direction and the passage secondary flow opposite direction respectively, thereby the blowing-off effect of the passage secondary flow on the cooling film is inhibited, the wall-attaching ability and the along-flow development ability of the cooling film are enhanced, the covering uniformity of the cooling film on the base surface is improved, and the cooling efficiency and the service life of the blade are improved. The leaf-shaped film cooling hole structure has the characteristics of strong functionality and wide applicability, and can be widely applied to the cooling technical field of the gas turbine blade of an aero-engine or a ground gas turbine.

The application can relate to the technical field of aero-engine and aero-engine control system, and particularly relates to a pressure pulse resistant pressurization shutoff device and method. The device comprises a main shutoff valve assembly and a pulse pressure relief valve assembly. The pulse pressure relief valve assembly is arranged in a main valve central cavity in the main shutoff valve assembly, a sensing cavity of the pulse pressure relief valve assembly is in communication with a control cavity of the main shutoff valve assembly, an axial blind hole is arranged in the pulse pressure relief valve close to the sensing cavity, a radial pressure relief through hole in communication with the axial blind hole is arranged in the pulse pressure relief valve, a radial main through hole in communication with the central cavity is arranged in the main valve of the main shutoff valve assembly, and the pulse pressure relief valve in the pulse pressure relief valve assembly can slide along the main valve central cavity. When the pressure relief through hole is in communication with the main through hole, pulse pressure relief is realized. The pressurization shutoff device of the fuel regulating mechanism can meet the pressurization shutoff function, can reduce metering fuel pressure pulsation during engine start-stop stage, and has little influence of control cavity fuel pressure pulsation on the function performance. The device has been used in a certain type of aero-engine and has good effect.

The application discloses a kind of identification methods of the influence factor of aero-engine power control device parts working process, it is related to aero-engine maintenance technical field, by the risk level and repair means effectiveness level of data quantization baseline unit are matched diagnosis, to judge whether the current maintenance means can be matched with the risk level of parts, effectively improve the control precision of influence factor in repair process, eliminate internal and external field product quality hidden danger, comprehensively improve the quality prevention ability, process control ability and quality control ability in product repair process, it plays a guiding role to the systematic method of process development and optimization.

The application belongs to the technical field of aero-engine fuel control and discloses a three-dimensional space oil passage modeling method for a fuel control device shell. The method comprises the following steps: taking a set cross section of a pipeline where an oil passage hole is located as a reference, taking a center of the cross section as a coordinate origin, and establishing a three-dimensional orthogonal coordinate system OXYZ with the cross section as an XOY plane; the oil passage hole is located above the set cross section; the XOY plane in the three-dimensional orthogonal coordinate system is offset upward to form a reference plane 1; a two-dimensional circle model is established on the reference plane 1; a reference axis with the center of the circle as an end point is established on the two-dimensional circle; the reference plane is rotated by a preset angle with the reference axis as a rotation axis to form a reference plane 2, a target hole is established on the reference plane 2, and the positioning reference of the target hole is the center of the two-dimensional circle, so that a modeling result of a three-dimensional oil passage hole is obtained, and accurate positioning of drawing marking is realized by using an orthogonal positioning method.

Embodiments of the present application provide an online water injection system and method for fuel icing test, relating to the field of aero-engine fuel icing test. It aims to improve the problem of unstable water content of fuel distribution in fuel icing test. It includes main oil supply path, steady-state water injection cylinder, steady-state nozzle, transient water injection cylinder, transient nozzle, bypass solenoid valve, oil supply pressure sensor and test piece; the steady-state water injection cylinder is connected with the steady-state nozzle, and the transient water injection cylinder is connected with the transient nozzle; the bypass solenoid valve is connected with the test piece in parallel, and the bypass solenoid valve is connected with the oil supply pressure sensor. Through the parallel connection scheme of the bypass electric control valve and the test piece, the transition state water-containing fuel oil supply is converted into steady-state supply, solving the problem that the water injection flow cannot quickly follow the change of fuel flow during the transition state of the test piece, and further providing stable water-containing fuel oil, realizing the continuous supply of stable water-containing fuel oil under the transition condition.

This invention relates to the field of acoustic emission signal feature enhancement for aero-engines, and discloses a method, system, medium, and device for denoising weak fault signals under strong background noise in aero-engines. The method includes: acquiring strong background noise signals under normal operating conditions of the aero-engine; performing wavelet packet decomposition and reconstruction to obtain multiple frequency band component signals; constructing a corresponding autoencoder for each frequency band component signal, and using the frequency band component signal as input to perform unsupervised training on the autoencoder; performing wavelet packet decomposition and reconstruction on the acoustic emission signal to be processed to obtain multiple frequency band component signals to be processed; inputting each frequency band component signal to be processed into the corresponding trained autoencoder for reconstruction; calculating the reconstruction error between the input and output signals of each autoencoder; determining whether each frequency band component signal contains outliers based on the reconstruction error; reconstructing the frequency band component signals containing outliers to obtain the denoised fault signal.

This invention discloses a self-cleaning aero-engine turbine blade grinding device. The invention relates to the field of turbine blade grinding technology and includes a base on which a robotic arm is rotatably connected. A servo motor is mounted on the end of the robotic arm furthest from the base, and a telescopic rod is connected to the output end of the servo motor. A vacuum cleaner is fixedly mounted on the side wall of the robotic arm. This self-cleaning aero-engine turbine blade grinding device achieves a smooth transition of the grinding head from one turbine blade to another through a simple rotational adjustment of the grinding head's orientation. This effectively shortens the overall grinding time, improves grinding efficiency, and automatically collects and processes dust and debris from the turbine blade surface, effectively preventing dust and debris from flowing back or splashing out from the air inlet, thus eliminating the risk of secondary scratches on the turbine blade surface caused by dust and debris.

This application provides an inlet support plate, an aero-engine, and an aircraft, relating to the field of aviation equipment technology. The inlet support plate includes: a plate body with a working fluid channel for circulating heat exchange working fluid; a working fluid inlet pipe disposed on the plate body, the inlet end of the working fluid channel being connected to the working fluid outlet end of an evaporator via the working fluid inlet pipe; and a working fluid outlet pipe disposed on the plate body, the outlet end of the working fluid channel being connected to the working fluid inlet end of an evaporator via the working fluid outlet pipe. The inlet support plate provided in this application can circulate the heat exchange working fluid in conjunction with the evaporator during use, absorbing heat from the heat exchange working fluid and reducing the amount of ice forming on the plate surface. While improving anti-icing capability, it eliminates the need for hot air supplied by the aero-engine compressor or the placement of heating devices on the plate surface, which helps reduce thrust loss and fuel consumption of the aero-engine and ensures the guiding capacity of the inlet support plate.

The application discloses a kind of based on TCN-BiLSTM-FCAM model's aero-engine residual life prediction method and system, it is related to aero-engine prediction and health management (PHM) technical field, the model includes TCN layer, BiLSTM layer, FCAM module and life prediction regression layer.This method is first obtained aero-engine multi-sensor operating data and pre-processes;Local degradation features are extracted using TCN;Global context dependence is captured by BiLSTM;Key time step context vector is obtained by FCAM weighting;Finally, the predicted value is output by the regression layer.The present application integrates the advantages of multiple models, reduces complexity, avoids overfitting and data pollution, and is suitable for aero-engine residual life prediction under different operating conditions and failure modes.

This invention discloses a fault diagnosis method for aero-engines based on an adaptive spatiotemporal decoupled graph convolutional network, comprising: acquiring and preprocessing multi-channel sensor signals to construct time-series samples; adaptively constructing an adjacency matrix to form a fault signal graph structure representing the spatial correlation of multi-source signals; inputting the fault signal graph structure into a graph convolutional network to extract spatial features; inputting the extracted spatial feature sequence into a bidirectional gated recurrent unit network to obtain temporal features; introducing a global attention mechanism to the temporal features to assign adaptive weights to features at different time steps; decoupling the spatiotemporal features using a variational autoencoder decoupling layer based on a Gaussian mixture distribution; and outputting the fault category of the aero-engine to complete the fault diagnosis. This invention significantly improves the accuracy, robustness, and engineering applicability of aero-engine fault diagnosis under complex operating conditions by jointly modeling the spatiotemporal characteristics of fault signals and effectively decoupling and fusing features.

The present invention relates to an assembly comprising a first component (110; 210; 310) and a second component (120; 220; 320; 420) for arrangement in the same turbine or compressor stage of a turbomachine, in particular an aircraft engine, wherein the first component has a first surface (111) and the second component has a second surface (121; 221; 321) opposite the first surface, wherein the first surface has a first contact area (112) for contacting the second surface and the second surface has a first contact contour (122; 222; 322) opposite this first contact area for contacting this first contact area, which has a recess (227; 327; 427; 428) or a convex edge (123; 223; 323) or a convex tip (124). The invention also relates to an assembly arrangement, a turbine and compressor stage and a turbomachine.

The application provides an aero-engine design method based on SysML activity diagram native driving and AI autonomous decision, and belongs to the technical field of aero-engines, and comprises the following steps: constructing a SysML model integrating engine architecture and behavior, wherein the SysML model is constructed based on a SysML activity diagram; loading the SysML model by a diagram explanation engine based on asynchronous event driving, and driving a simulation calculation process by managing the life cycle of a token; when the simulation calculation is performed, performing AI autonomous decision and pruning strategy based on Bayesian inference, including dynamic pruning based on a sensitivity matrix and autonomous optimization based on Gaussian process regression; and associating the simulation calculation result with the attributes of the SysML model by the system automation to complete integration. The scheme reduces the communication and maintenance cost of a complex system, improves the process observability and debugging efficiency, and realizes intelligent scheduling and data alignment based on an architecture context.

A high-temperature resistant lightweight composite material, its preparation method, and its applications are disclosed, relating to the field of composite material technology. This invention solves the problems of existing materials where it is difficult to simultaneously achieve high temperature resistance and lightweight, low structural efficiency due to the separation of heat protection and load-bearing design, insufficient temperature resistance and toughness of the sandwich layer, and difficulties and high costs associated with large-size molding. The material comprises an integrated co-cured outer heat-resistant layer, an intermediate transition layer, and an inner load-bearing layer. The outer heat-resistant layer uses a modified epoxy / cyanate copolymer resin as the matrix and a high-silica glass fiber and carbon fiber blended fabric as reinforcement, with fumed silica and carbon nanotube fillers incorporated into the matrix resin. The intermediate transition layer is polymethacrylamide foam. The inner load-bearing layer uses an epoxy-modified cyanate ester resin as the matrix and a carbon fiber biaxial plain weave fabric as reinforcement. This invention can be used in rocket fairings, hypersonic vehicle skins, spacecraft reentry capsule thermal protection structures, hypersonic missile shells, aircraft engine nacelles, or high-speed train front fairings.

The application discloses a kind of hysteresis reference window sliding window adaptive threshold surge detection system and method, belong to aero-engine surge detection field.The method includes: real-time acquisition aero-engine pressure sensor output analog voltage signal, and carry out partial pressure, filtering and analog-digital conversion, obtain discrete sampling sequence;With sliding window intercept analysis window and carry out linear fitting, calculate its slope and residual standard deviation;Established by historical analysis window eigenvalue, with current window maintains hysteresis distance reference window;Based on the statistics of eigenvalue sequence in reference window, dynamically calculate adaptive threshold;By comparing current window eigenvalue with threshold value, and when the eigenvalue of continuous multiple windows exceeds threshold value, determine surge.The detection scheme provided by the application significantly reduces the amount of calculation through local window processing and dynamic threshold adjustment, solves the problem that existing algorithms are difficult to be deployed in real time in embedded hardware, and improves the detection accuracy and working condition adaptability.

The application relates to the technical field of turbine blades, in particular to a centripetal turbine, an aero-engine and a use method. The centripetal turbine comprises turbine guide vanes, the turbine guide vanes comprise a guide vane hub and at least two guide vanes, the at least two guide vanes are arranged in a circumferential array on the guide vane hub, each of the guide vanes is respectively provided with a shock wave weakening tail edge, and the shock wave weakening tail edge is used for weakening the shock wave intensity blowing to the guide vanes. The application provides the centripetal turbine, the aero-engine and the use method, so as to solve the problem that the shock wave pulsation load is overlapped with the blade inherent frequency, high-cycle fatigue failure is easily caused, and the service life of the blade is shortened.

This invention discloses a device and method for inspecting the free swing of a rotary oil suction mechanism for aircraft accessories, belonging to the field of process tooling technology. This invention aims to solve the technical problem that existing manual methods cannot accurately pull the rotary oil suction mechanism to a position of (90±5)° to check the number of swings, leading to inaccurate test results. The device of this invention includes a base and a fixing mechanism and a positioning mechanism mounted thereon; the fixing mechanism is used to fix the rotary oil suction mechanism to be tested; the positioning mechanism includes a slider movably mounted on the base, with two face positioning blocks on the slider, forming an angle of (90±5)° between the two face positioning blocks. The method of use is as follows: fix the rotary oil suction mechanism on the fixing mechanism; move the slider so that its rotating component enters between the two face positioning blocks and is limited to a preset angle; move the slider in the opposite direction to release the rotary oil suction mechanism; observe and record its number of free swings. This invention can achieve precise positioning of the release angle of the rotary oil suction mechanism, ensuring the accuracy and consistency of the test results, is easy to operate, and effectively guarantees the working reliability of the aircraft engine accessory transmission system.