4952 results about "Gas turbines" 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.

The invention belongs to the field of data processing, and particularly relates to a dynamic cooperative control system and method for a gas turbine and a micro-grid, and the method comprises the steps: constructing a micro-grid real-time monitoring module, continuously collecting distributed energy real-time output, controllable load demands, bus voltage frequency and equipment state parameters, and carrying out the filtering and noise reduction through a preprocessing unit, thereby guaranteeing the data precision; calculating a real-time power difference value based on the preprocessed data, calling an adaptive neural fuzzy inference system, taking the power difference value, the bus voltage deviation and the frequency deviation as input, and judging whether the power difference value, the bus voltage deviation and the frequency deviation exceed a preset threshold value by means of a fuzzy rule base and a neural network model; if the threshold values are not exceeded, the current states of the gas turbine and the energy storage system are maintained; if any one exceeds the threshold value, a dynamic cooperative control instruction is triggered, precise cooperative control of the gas turbine and the micro-grid is achieved, and the operation stability, the operation efficiency and the reliability of the micro-grid are improved.

The invention provides a gas turbine exhaust gas temperature prediction and fault early warning system and method, and relates to the technical field of gas turbine state monitoring and fault diagnosis. The system comprises a data acquisition and preprocessing module used for acquiring and preprocessing data; the dynamic graph construction module is used for constructing graph structure data; the graph neural network module is used for extracting spatio-temporal features through learning; the improved auto-encoder module is used for realizing feature reconstruction by learning normal working condition data and then outputting reconstruction errors to the fault early warning module; the temperature prediction module is used for predicting a future exhaust temperature value; the fault early warning module is used for performing fault judgment and decision; and the result display and output module is used for summarizing analysis results. According to the system, the prediction precision of the exhaust temperature is remarkably improved, the fault detection and early warning capabilities are enhanced, the system has better feature representation capability and adaptability to complex working conditions, and close integration and mutual enhancement of prediction and early warning are realized.

The invention provides a gas turbine exhaust temperature sensor fault diagnosis system and method, and relates to the technical field of industrial equipment state monitoring and diagnosis. The system comprises a sensor signal acquisition module for acquiring an original temperature signal; the signal preprocessing module is used for acquiring and preprocessing original signals and working condition parameters; the feature extraction and windowing module is used for acquiring data and calculating features; the data driving diagnosis module is used for receiving the time sequence characteristics and evaluating the health state; and the decision fusion module is used for comprehensively analyzing the multi-source information and making a final fault diagnosis judgment. According to the system, the accuracy and the reliability of fault diagnosis can be remarkably improved, the false alarm rate and the missing report rate are reduced, early warning and accurate identification of early weak faults of the sensor are realized, the adaptability and the robustness of the diagnosis system to variable working conditions of the gas turbine are improved, and the safety, the economical efficiency and the operation and maintenance efficiency of operation of the gas turbine are improved.

An aircraft gas turbine engine includes a combustor, with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include first and second subsets. The first subset of nozzles is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A lean cruise nvPM emissions index ratio isEIcruise⁡(lean),SAFEIcruise⁡(lean),FF,where EIcruise(lean),SAF isEImaxTO,SAF+EIclimb,SAF2and EIcruise(lean),FF isEImaxTO,FF+EIclimb,FF2.EImaxTO,SAF is nvPM emissions index operating at 100% available thrust and EIclimb,SAF is the nvPM emissions index operating at 85% available thrust if fuel includes sustainable aviation fuel. EImaxTO,FF is the nvPM emissions index operating at 100% available thrust and EIclimb,FF is the nvPM emissions index at 85% available thrust if fuel is a fossil-based hydrocarbon. The lean cruise nvPM emissions index ratio is less than 1.

A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. A thrust nvPM emissions index ratio is defined as:EImaxTO / FmaxTOEIidle / Fidlewhere: EIidle is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; EImaxTO is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; FmaxTO is the thrust of the gas turbine engine at around 100% available thrust in kN for the given operating conditions; and Fidle is the thrust of the gas turbine engine at around 7% available thrust in kN for the given operating conditions. The thrust nvPM emissions index ratio is greater than 0.001. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.

The invention relates to the technical field of power equipment control, in particular to an Agent-driven intelligent AI depth prediction method and system for a power system, and provides the following scheme that real-time operation data of a gas turbine is collected through a sensor, historical data and a preset mechanism model are combined, an intelligent AI model is constructed, and key parameters which are difficult to directly measure are predicted. And through deviation analysis, abnormity and deviation sources in the operation state are identified, reinforcement learning and a particle swarm optimization algorithm are utilized to dynamically update model parameters, optimization control parameters are generated, and the operation state of the gas turbine is adjusted. According to the method, the difficult-to-measure parameters can be accurately and reversely deduced from the easy-to-measure data, the operation efficiency and stability are continuously optimized through closed-loop control, the dynamic adjustment requirement under the complex working condition is met, and the economical efficiency and reliability of the gas turbine are improved.

Systems and methods for providing lubrication to a gas turbine engine are described. An example gas turbine engine includes a plurality of bearings disposed along an axial length of a rotatable shaft to provide rotational support between a static structure and at least one of a compressor and a turbine of the gas turbine engine. The gas turbine engine further includes a fuel system operable to deliver a mixture of fuel and lubricant sequentially to at least one of the plurality of bearings and the combustor so that during operation of the gas turbine engine, at least the lubricant within the mixture is in a liquid phase prior to being received by the at least one of the plurality of bearings, and in a vapor phase once the lubricant is delivered to the at least one of the plurality of bearings.

A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A MTO nvPM emissions index ratio is defined as:EImaxTO,SAFEImaxTO,FFwhere: EImaxTO,SAF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EImaxTO,FF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed are methods of operating the gas turbine engine.

A gas turbine engine for an aircraft. The gas turbine engine comprising: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. An MTO nvPM emissions index ratio-modified fuel flow is defined as:EImaxTO,SAFEImaxTO,FF×Wf,maxTOwhere: EImaxTO,SAF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); EImaxTO,FF is the system loss corrected nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and Wf,maxTO is the mass flow rate of fuel provided to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at around 100% available thrust for the given operating conditions. The MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine in kg / s is less than 2. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.

A gas turbine engine includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The nozzles include a first and second subset. Each of the nozzles of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to nozzles in the second subset from 1:3 to 1:6. A fuel-flow nvPM emissions index ratio isEIidle×Wf,idleRImaxTO×Wf,maxTO.EIidle and EImaxTO are the system loss corrected nvPM emissions index in mg / kg operating at around 7% and 100% available thrust for the given operating conditions, respectively. Wf,idle and Wf,maxTO are the rate of fuel flow to the fuel spray nozzles in kg / s at around 7% and 100% available thrust for the given operating conditions, respectively. The fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.3.

A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in the range 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An nvPM emissions index ratio is defined as:EIidle×Wf,idleEImaxTO×Wf,maxTOwhere: EIidle and EImaxTO are respectively the nvPM emissions index in mg / kg of gas turbine engine if operating at around 7% or 100% available thrust for given operating conditions; Wf,idle and Wf,maxTO are respectively the rate of fuel flow to fuel spray nozzles in kg / s at around 7% or 100% available thrust for given operating conditions. The fuel-flow nvPM emissions index ratio is less than 0.08. The gas turbine engine is configured to provide fuel including sustainable aviation fuel to fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.

A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

A gas turbine engine is provided. The gas turbine engine includes: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches; wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (FnTotal) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: FnTotal×EGT / (AHPCExit2×1000).

A gas turbine engine is provided, having: a bifurcation; and a turbomachine further including a cooled cooling air (CCA) system, the CCA system having: a cold side bleed assembly defining an inlet positioned to be in fluid communication with an airflow over the bifurcation; a CCA heat exchanger in thermal communication with the cold side bleed assembly downstream of the inlet of the cold side bleed assembly; and a hot side bleed assembly defining an inlet in fluid communication with a working gas flowpath through a compressor section, at a compressor discharge cavity, or both, the hot side bleed assembly in thermal communication with the CCA heat exchanger to cool an airflow through the hot side bleed assembly, the hot side bleed assembly further in thermal communication with a hot component of the turbomachine to cool the hot component of the turbomachine.

A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A MTO nvPM emissions index ratio is defined as:EImaxTO,SAFEImaxTO,FFwhere: EImaxTO,SAF is the nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EImaxTO,FF is the nvPM emissions index in mg / kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed are methods of operating the gas turbine engine.

The invention relates to the technical field of gas turbine performance monitoring, and provides a gas turbine transient model automatic correction method based on digital object fusion, which comprises the following steps: step 1, carrying out data preprocessing on operation data of a gas turbine; 2, establishing a transient model of the gas turbine; step 3, determining key influence parameters having significant sensitivity to thermodynamic efficiency and hydrodynamic characteristics, and determining correction factors according to the key influence parameters; step 4, calculating the correction factor of the transient model of the gas turbine determined in the step 3; and 5, adjusting the gas turbine transient model by using the correction factor obtained by calculation to obtain a corrected gas turbine transient model. According to the invention, the accuracy and response reliability of dynamic adjustment of the model can be improved.

A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in range of 14-22 or a number of fuel spray nozzles per unit engine core size in range 2 to 6. An MTO nvPM emissions index ratio is:EImaxTO,SAFEImaxTO,FF×Wf,maxTOwhere: EImaxTO,SAF and EImaxTO,FF are respectively the nvPM emissions index in mg / kg when operating around 100% available thrust for given operating conditions if fuel provided to fuel spray nozzles includes either sustainable aviation fuel (SAF) or fossil-based hydrocarbon fuel; Wf,maxTO is mass flow rate of fuel provided to fuel spray nozzles in kg / s when gas turbine engine is operating at around 100% available thrust for given operating conditions. MTO nvPM emissions index is less than 2. The gas turbine engine is configured to provide fuel including SAF to fuel spray nozzles. Also disclosed is method of operating gas turbine engine.

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.

The invention belongs to the technical field of gas turbine maintenance, and particularly relates to a solid-state repair method for a hot-end component of a gas turbine, which comprises the following steps: S1, acquiring damage information of the surface of a to-be-repaired matrix; s2, roughening treatment is conducted on a damaged area on the surface of the to-be-repaired base body, and electrochemical activation treatment is conducted on the roughened damaged area; s3, supersonic speed cold spraying treatment is conducted on the treated damaged area, and gradient materials are deposited on the damaged area layer by layer till a compact repairing layer is deposited on the surface of the base body; s4, selective laser remelting treatment is conducted on the compact repairing layer of the damaged area; s5, carrying out laser shock on the dense repair layer subjected to the remelting treatment; and S6, the base body subjected to laser shock treatment is machined till the standard geometric dimension of the base body is recovered. The method can solve the problems of large heat input, matrix damage, insufficient bonding strength and the like in the process of repairing the damaged area of the matrix in the traditional hot working technology.

The invention discloses a stator blade shape-following frequency modulation design method based on a bending vibration mode and application, belongs to the field of vibration design of stator blades of aero-engines and gas turbines, and is suitable for vibration frequency modulation design of non-cantilever type stator structure blades (such as fans and gas compressor stators) with inner rings and outer rings in the aero-engines or the gas turbines. According to the method, firstly, static frequency and dynamic frequency modal analysis is carried out through finite element modeling, and dangerous resonance points and corresponding vibration modes in a working rotating speed range are recognized by combining a Campbell diagram; based on a vibration mode amplitude normalization result, setting a vibration degree parameter and a scale factor, selecting a high response node and carrying out directed disturbance, and updating a structure model to improve a target vibration mode frequency; and through frequency amplification verification and iterative optimization, geometric reconstruction and frequency modulation confirmation are finally completed in combination with aerodynamic performance and processing requirements. The method can realize rapid and accurate frequency modulation of a specific bending vibration mode, avoids introduction of new resonance, and has the advantages of high efficiency, strong adaptability and the like.

A systematic framework for optimizing multi-stage axial compressor disk designs in gas turbine engines. By combining finite element analysis (FEA), Design of Experiments (DOE), and optimization algorithms of multi-objective genetic algorithm (MOGA), the method balances stress, deformation, and mass to enhance structural performance. The six-step process includes blade modeling, parameterizing disk geometry, structural analysis using FEA, developing functional relationships, applying optimization algorithms, and generating manufacturable 3D disk models. This approach reduces weight, improves fuel efficiency, and adapts to various compressor designs and materials, enhancing the overall performance of gas turbine engines.

A gas turbine engine includes a turbomachine comprising compressor, combustion, and turbine sections. The gas turbine engine defines a maximum exhaust gas temperature, a maximum drive turbine shaft torque, and a corrected specific power. The gas turbine engine includes a controller configured to autonomously regulate performance of the gas turbine engine in response to at least one of: a thrust demand, an energy efficiency target, or a flight profile condition.

A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:(TOUT / √{square root over (ADTExit)})2*EGT / AHPCExit*10−11;wherein CSP is greater than 0.0001194×EGT2−0.103×EGT+22.14 and less than 0.0003294×EGT2−0.306×EGT+77.91; and wherein EGT is greater than 525 degrees Celsius and less than 1250 degrees Celsius.

The invention relates to the technical field of ignition systems of gas turbines, and discloses a special ignition system for a gas turbine, which comprises a monitoring control layer, an energy distribution layer and a combustion optimization layer. The monitoring control layer monitors parameters such as combustion chamber pressure fluctuation in real time through multiple types of sensors and executes phase synchronization and injection control. The energy distribution layer is provided with a multi-stage energy release strategy and a self-adaptive time sequence matching mechanism, a stable ignition energy field is constructed, and physical and digital space linkage regulation and control are achieved through a multi-modal energy interface; and the combustion optimization layer establishes a multi-dimensional mapping combustion state control model based on a dynamic combustion algorithm and real-time data, and adopts a collaborative optimization algorithm to calibrate an ignition time sequence, regulate and control a fuel mixing ratio and maintain flame stability. According to the system, through multi-dimensional energy fusion, model parameter dynamic correction and intelligent algorithm application, the accuracy, stability and multi-working-condition adaptability of the gas turbine ignition process are remarkably improved, and the system is suitable for gas turbine efficient combustion control in the fields of aerospace, electric power and the like.

A blade for a fan of a gas turbine engine includes an airfoil section that spans along a radial direction between first and second radial ends, and a dovetail at the second end of the airfoil section. The airfoil section and the dovetail are comprised of a fiber-reinforced composite that includes inter-section fiber plies that each extend from the airfoil section into the dovetail and wedges that are interleaved with the inter-section fiber plies in the dovetail. The dovetail is asymmetric about the radial direction and has first and second circumferential sides that are sloped to face radially outwardly and form non-zero angles with the radial direction. One of the non-zero angles is 10° or less and a sum of the non-zero angles is from 35° to 45°.

A gas turbine engine includes: a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath; a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio during operation of the gas turbine engine in a cruise operating mode; and a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan; wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8.

The invention discloses a key sensor short-time abnormal distribution drift detection method in a unit start-stop process, and belongs to the technical field of gas turbine power plant financial supervision and artificial intelligence, and the method comprises the steps: synchronously triggering multi-channel signal collection through a main clock, and achieving noise suppression and data pre-screening through the combination of first-order difference and threshold filtering; constructing a nonlinear weighted feature matrix, and fusing a time attenuation coefficient and a shafting acceleration factor to enhance the transient feature expression capability; generating a sensor association graph based on double-threshold determination of weighted Pearson's correlation coefficients and mutual information, and dividing stable subgroups by using an incremental label propagation algorithm; designing a double-layer Cluster-GCN model, aggregating subgroup internal characteristics in the first layer, introducing a fuel valve position-acceleration comparison gating mechanism in the second layer to correct a global edge weight, and generating a node embedding vector sensitive to working condition change; gaussian kernel density estimation and an instantaneous deviation index of embedding similarity are fused, and a historical sliding mean value and subgroup connectivity analysis are combined, so that sensor faults and working condition abrupt changes are distinguished.

The invention relates to the technical field of gas turbine performance prediction, and provides a thermodynamics-based gas turbine performance lightweight modeling method, which comprises the following steps of: 1, acquiring actual operation data of a gas turbine, and preprocessing the operation data; 2, establishing a thermodynamic model of the gas turbine; 3, establishing a gas turbine agent model; and step 4, training the gas turbine proxy model to obtain a trained gas turbine proxy model. According to the method, the influence of external environment conditions and power requirements on working condition operation conditions can be eliminated, and key gas turbine performance parameters can be quickly obtained.

A combustor for a gas turbine (GT) system includes a combustor body including a non-additively manufactured (non-AM) part having corrosion at a first level and a replacement additively manufactured (AM) part coupled to the non-AM part. The replacement AM part has corrosion at a second level less than the first level. The replacement AM part includes at least an aft section of a tapered transition portion and an aft frame of the combustor body, and a plurality of shared parallel, sintered metal layers extending into the aft section of the tapered transition portion and the aft frame. The replacement AM part may also include a receiving element in a forwardmost end thereof configured to receive a rearwardmost remaining end of the non-AM part. The replacement AM part may also optionally include an AFS injector mount and part of a cylindrical portion of a combustion liner.