245 results about "Power turbine" patented technology

A power turbine blade for a gas turbine engine includes at least one of a third-stage blade, a fourth-stage blade, or a fifth-stage blade. The power turbine blade includes a platform and an airfoil that extends in a spanwise direction from the platform to a tip shroud. The airfoil has an external airfoil surface formed in conformance with multiple cross-section airfoil profiles described by a set of Cartesian coordinates.

A gas turbine engine for an aircraft includes a fan and a turbo-engine having a combustor that combusts compressed air and fuel to generate combustion gases and a power turbine coupled to a shaft for driving the fan. The gas turbine engine includes a direct contact condenser for condensing steam in a first flow into liquid water droplets and to generate a second flow including the liquid water droplets, a water tank for collecting the liquid water droplets of the second flow, and a water pump for introducing the liquid water droplets from the water tank into the direct contact condenser to act as condensation nuclei for the steam in the first flow. The first flow is an exhaust-vapor flow extracted from the combustion gases and the second flow is an exhaust-condensate flow from the direct contact condenser.

The invention discloses a power turbine optimization method and system based on data driving, and the method comprises the steps: selecting a plurality of optimization working condition points, and determining turbine operation working condition parameters under each optimization working condition point; obtaining a blade profile optimization geometric parameter, and determining the dimension and the variation range of the blade profile optimization geometric parameter; forming a plurality of groups of geometric parameter schemes based on the blade profile optimization geometric parameters, and establishing a turbine three-dimensional model and dividing computational grids; cFD simulation is carried out on the turbine under each group of geometric parameter schemes, aerodynamic performance parameters of each geometric parameter scheme under each working condition are obtained, and a sample library is formed based on the blade profile optimization geometric parameters, turbine operation working condition parameters and the aerodynamic performance parameters so as to establish a neural network prediction model; and performing multi-objective optimization on the blade profile optimization geometric parameters by adopting a neural network prediction model to determine optimal blade profile geometric parameters. By means of the scheme, the optimal blade profile geometric parameter combination meeting the multiple performance targets at the same time can be efficiently searched in the design space.

The invention discloses a combined type coaxial high-speed helicopter flight / turboshaft engine comprehensive control system. The system comprises the following steps that according to control mechanism parameters output by a flight control system and a combined type coaxial high-speed helicopter fuzzy torque prediction model, a flight required torque advanced prediction value of a combined type coaxial high-speed helicopter system is output; the flight required torque advanced prediction value is input into a turboshaft engine system identification model after being superposed by a variable transmission ratio proportionality coefficient to obtain a fuel prediction value, and the fuel prediction value is further superposed with the output of an engine control system to obtain the fuel input of the turboshaft engine system; calculating a flight required torque advanced prediction value error, and if the error value is greater than a threshold value, updating the fuzzy torque prediction model by using the flight state parameters of the composite coaxial high-speed helicopter system; the relative error value of the current power turbine rotating speed is calculated, and if the relative error value is larger than a threshold value, primary parameter updating is conducted on the identification model of the turboshaft engine system through a recursive least square algorithm according to the real-time and historical torque input of the turboshaft engine system and the power turbine rotating speed. According to the method, under the condition that the constant rotating speed control target of the turboshaft engine is considered, the stability of a coaxial high-speed helicopter flight / turboshaft engine comprehensive control system is enhanced, the response speed of the turboshaft engine is increased, and the overshoot and droop amount of the rotating speed of a power turboshaft is reduced.

The utility model relates to a long cantilever double rail type gas turbine running guide rail device which comprises double guide rails and a guide rail supporting structure, the double guide rails are symmetrically installed on the two sides of a gas turbine, and the upper layer guide rail and the lower layer guide rail of the double guide rails are fixedly connected through a connecting assembly and connected with a supporting column. A front supporting point of the fuel gas generator is located on the axis horizontal plane and located at the front end of the generator; the rear supporting point is located on the intercooler, and the generator intermediate roller runs on the upper-layer guide rail; front and rear supporting points of the power turbine are located on the axis horizontal plane, and the power turbine and the front and rear rollers run on the upper-layer guide rail. The front roller and the rear roller are located in the two guide rail grooves respectively and run along different tracks, the double-rail type running guide rail device can reduce the structure interference risk, the guide rail supporting tool can be fixed to the two sides of an intercooler, the span of the supporting tool is effectively reduced, and the stability of the gas turbine in and out of a box for assembly is improved.

The invention discloses a turboshaft engine component characteristic dynamic optimization method, device and equipment and a medium, and the method comprises the steps: S1, calculating to obtain typical parameters of an engine in each flight stage, after multiple rounds of typical parameter calculation results are obtained through iteration by changing the relative conversion rotating speed increment, the gas compressor characteristic diagram with the lowest fuel flow weighted average value is selected as the optimal optimization design result of the gas compressor; s2, performing engine parameter calculation according to the optimal optimization design result of the gas compressor to determine the position relationship between the highest efficiency area of the power turbine on the turbine characteristic diagram and each stage of the aircraft; the power turbine characteristic diagram is adjusted through a double-Gaussian correction formula, the power turbine efficiency of the design point is adjusted, then iteration is carried out, typical parameter calculation results of the engine in the flight stages of multiple rounds are obtained, and the power turbine characteristic diagram with the lowest fuel flow weighted average value is selected as the optimal optimization design result of the power turbine. The method is low in optimization cost, efficient and accurate.

A small and medium-sized gas turbine power turbine rotor air system relates to the technical field of gas turbines. The problems that an existing gas turbine with a power turbine needs to bleed air from different positions of an air compressor, so that the using amount of cooling air is large, and the cooling efficiency is low are solved. The gas turbine comprises a front wall, a main flow gas channel, a first-stage turbine wheel disc, a rear wall and a shaft cavity, a first cavity is formed between the first-stage turbine wheel disc and the front wall, the first cavity and the shaft cavity communicate with a gas compressor at the same time, and the first cavity communicates with the main flow gas channel through a first gas outlet, a second gas outlet and a third gas outlet; only one strand of compressed air is led out from a certain part of the air compressor, one part of the compressed air flows into the first cavity, the other part of the compressed air flows into the shaft cavity and converges into the mainstream fuel gas channel after passing through the multi-stage blades, and the compressed air is shunted, so that the effects of cooling the wheel disc, sealing the disc cavity and sealing and pressurizing the bearing cavity can be completed, and the cold air consumption can be saved; the gas turbine unit efficiency is improved.

An electrically-powered turbine assembly includes a fixed housing; an air compressor including a compressor stator fixed to the housing; a turbine including a turbine stator fixed to the housing; a central shaft supported by the compressor stator and affixed to the turbine stator; an electric motor including an electric motor stator fixed to the central shaft, said electric motor having a rotating shell; an air compressor rotor affixed to the rotating electric motor shell; and a turbine rotor affixed to the rotating electric motor shell. The electrically-powered turbine assembly preferably includes one or more fuel cells as an electric power source. The electrically-powered turbine assembly may be used to power an aircraft.

The invention belongs to the technical field of helicopter control, and provides a turboshaft engine identification and prediction control method based on MRR-KELM in order to solve the problem that a prediction model constructed for a turboshaft engine is poor in generalization, an MRR-KELM algorithm is combined with a Huber loss function and a self-adaptive regularization strategy, noise interference and abnormal value influence are restrained, and the prediction model is optimized. The prediction error of the constructed turboshaft engine prediction model is greatly reduced, and adaptive adjustment can be performed through Gaussian kernel dynamic mapping and working condition parameters; according to the algorithm, a regularization item and a robust loss function are introduced, the kernel function mapping capability is combined, the generalization performance and the calculation efficiency of the model under the complex working conditions of small samples, noise interference and the like are remarkably improved, further, the turboshaft engine prediction model is embedded into a nonlinear model prediction control framework, and the prediction accuracy of the turboshaft engine is improved. And high-precision stable control over the rotating speed of the power turbine is achieved through rolling optimization and feedback correction.

A fully-submerged multi-stage Horizontally-oriented Conical-helical Hydrokinetic Turbine that captures water stream flows (river, tidal, deep ocean currents) and funnels this kinetic-rich concentrated flow into a conical turbine system consisting of a rotating arrangement of helical hydrofoil blades wherein the internal stream is forced to exit out through the frustum turbine blade assembly. With a higher Capacity Factor and Power Coefficient than conventional renewables, this turbine is designed to be an environmentally friendly baseload electricity generator to meet the ever-increasing electricity demands.

The invention discloses a power turbine rotor structure design method, which belongs to the technical field of aero-engines, and comprises the following steps: arranging an oil film damper according to strain energy distribution and vibration mode; performing structure optimization on the power turbine rotor according to the radial deformation condition of the connection interface; designing the sectional wall thickness of the power turbine shaft; a balance boss is arranged on the outer wall of the maximum displacement position in each order of vibration mode power turbine shaft; and the gap between the central pull rod of the power turbine rotor and the power turbine disc is designed according to the maximum vibration displacement of the power turbine disc. According to the invention, the oil film dampers are arranged at reasonable positions, so that the rotor has good vibration reduction; connection interface deformation coordination calculation is carried out, a connection structure is optimized, and the connection interface is not damaged under long-term work; and designing and optimizing the segmented shaft wall thickness, optimizing and designing the critical rotating speed margin, considering the influence of the limit unbalance on the gap, and realizing a good dynamic balance characteristic.

The invention relates to the technical field of aero-engines, and discloses a starting oil supply rule control method and device for an unmanned aerial engine, and the method comprises the steps: obtaining the rotating speed and acceleration at a target moment and a historical moment, and judging whether a first switching condition is met or not; if yes, the inlet temperature of the power turbine at the target moment and the historical moment is obtained, and whether a second switching condition is met or not is judged; if yes, the fuel given quantity and the fuel feedback quantity are obtained, and whether a third switching condition is met or not is judged; if yes, judging whether a fourth switching condition is met or not based on the first time consumption, the second time consumption and the third time consumption; if yes, whether a fifth switching condition is met or not is judged based on the first time consumption, the second time consumption and the fourth time consumption; if yes, switching to a third starting stage in advance, and executing the oil supply rule of the third starting stage. Through multi-stage judgment, oil supply rules are timely and effectively switched, suspension protection is prevented from being triggered, and the starting success rate of the engine is increased.

The invention relates to a method and system for predicting the maximum rotating speed of a power turbine after an engine loses load, and belongs to the field of aviation gas turboshaft engines. The method comprises the following steps that a pressure signal of a gas compressor outlet of the engine, a pressure signal of a main oil way of a fuel oil distribution device and a pressure signal of an auxiliary oil way of the fuel oil distribution device are obtained; performing a simulation over-rotation test under the maximum state of the engine, and obtaining a fuel oil rule of the maximum state over-rotation of the engine according to a test result; carrying out segmented iterative calculation on the rotating speeds of the power turbine at different time points in the maximum state over-rotating process of the engine to obtain a relationship between the rotating speeds of the power turbine and time; and according to the relationship between the power turbine rotating speed and time, the maximum power turbine rotating speed in the maximum state overrunning process of the engine is obtained. According to the method, the response time, the fuel attenuation rule and the torque change rule of the over-rotation protection system are fully considered, and the maximum rotating speed under the condition of shaft breakage of the engine can be accurately predicted.

An environmental control system (ECS) of an aircraft, having a heat exchanger; a compressor that receives a first airflow from a first air source, compresses the first airflow, and directs the first airflow to the heat exchanger; an axial flow turbine that receives a second airflow from a second air source, extracts energy from the second airflow, and directs the second airflow to the heat exchanger, wherein the heat exchanger directs the first airflow to a cabin of the aircraft and directs the second airflow overboard; and a shaft connected between the turbine and the compressor.

The invention relates to the technical field of gas turbine control, in particular to a gas turbine multi-parameter cooperative control and dynamic safety adjusting system and method. According to the system, PID outputs of high-pressure compressor rotating speed tracking, power turbine rotating speed tracking, active power tracking and parameter limit control loops are generated in parallel through a multi-parameter cooperative control module, and a minimum fuel demand signal is selected as a control reference based on competitive logic; through a dynamic safety adjusting module, a dynamic threshold curve linked with the inlet air temperature is preset, action allowance is superposed, and active intervention before parameter overrun is achieved; and through the fault tolerance module, the operation is maintained by adopting historical data fitting and an online compensation strategy when the sensor has a fault. According to the method, the problems of single parameter control response delay, insufficient limit protection, poor fault tolerance and the like are solved, and the control precision, response speed and operation reliability of the gas turbine under variable load, extreme environment and sensor abnormal working conditions are improved.

A turbine engine includes integrated electric machines in the compressor section and the turbine section to supplement power produced from fuel with electric power. The example compressor section includes a compressor electric motor that is coupled to a compressor generator. The example turbine section includes a turbine electric motor that is coupled to the geared architecture to supplement power driving the fan section. A turbine generator provides electric power to the turbine electric motor.

A lubricating system is disclosed. The lubricating system includes: at least one first to-be-lubricated component, wherein an inlet of each of the at least one first to-be-lubricated component is connected with a first lubrication oil inlet pipe, and an outlet of the each of the at least one first to-be-lubricated component is connected with a first lubrication oil outlet pipe; and at least one second to-be-lubricated component, wherein an inlet of each of the second to-be-lubricated component is connected with a second lubrication oil inlet pipe, and an outlet of the each of the at least one second to-be-lubricated component is connected with a second lubrication oil outlet pipe. An operating pressure of the each of the at least one first to-be-lubricated component is different from a working pressure of the each of the at least one second to-be-lubricated component.

The invention discloses a hybrid power turboshaft engine and a working method, the engine comprises a core machine system and a fuel cell power supply system, and the core machine system comprises an air inlet channel, an air compressor, a main combustion chamber, a gas turbine and a power turbine; the fuel cell power supply system comprises a liquid ammonia storage tank, a fuel pump, a liquid ammonia decomposer, a pressure reducing valve, a heat exchanger and a proton exchange membrane fuel cell. According to the invention, a hybrid power form of combining a turbine power system and a fuel cell power system is adopted, so that the output power of the turboshaft engine is ensured, and meanwhile, the proton exchange membrane fuel cell provides electric energy for airborne electric equipment of the engine; the problems that the electric power requirement of an existing turboshaft engine is gradually improved, power of storage batteries such as lithium batteries is insufficient, and electric power is difficult to extract from a power turbine are solved. Meanwhile, the ammonia working medium serves as power generation fuel, the energy conversion efficiency of the aviation turboshaft engine is improved, and energy consumption is reduced.

A turbomachine engine including an engine core including a high-pressure compressor, which has an exit stage having an exit stage diameter (DCORE), a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, a power turbine in flow communication with the high-pressure turbine, and a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft / sec)1 / 2 and three hundred (ft / sec)1 / 2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft / sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft. The low-pressure shaft has a length (LMSR) defined by an engine core length (LCORE) given by:LCORE=[m(20+m)*n(10+n)](1100)*DCORE+CIS.

The present application belongs to the field of aero-engine technology, and particularly relates to a bolt connection structure suitable for a power turbine casing, a power turbine casing, and a plurality of segmented power turbine outer rings mounted on the inner wall surface of the power turbine casing; wherein, the inner wall surface of the power turbine casing is provided with a plurality of annular stoppers and a plurality of annular mounting platforms; the annular mounting platform is composed of a ring having a plurality of circumferentially distributed U-shaped through grooves for accommodating bolts, the opening of the U-shaped through grooves faces the inner side of the power turbine casing, and the end face of the U-shaped through grooves has a U-shaped groove for accommodating bushing nuts; the power turbine outer ring has an annular groove clamped in the annular stopper and a mounting edge, the mounting edge has a plurality of through holes, the bolts pass through the through holes and are connected with the U-shaped through grooves and bushing nuts, so that the mounting edge is fixedly connected to the annular mounting platform.

This invention discloses a one-dimensional design method, system, device, and storage medium for a variable-speed power turbine. The one-dimensional design method for the variable-speed power turbine iteratively optimizes the design speed point by calculating the comprehensive performance of the variable-speed power turbine under all operating conditions. It takes into account the characteristics of the wide operating range of the variable-speed power turbine and establishes a comprehensive performance evaluation index and system applicable to the operating characteristics of the variable-speed power turbine. It accurately reflects the comprehensive performance of the variable-speed power turbine throughout the entire mission cycle and accurately evaluates and predicts the full-mission performance of the variable-speed power turbine. It can achieve the goal of efficient aerodynamic design of the variable-speed power turbine under a wide range of operating conditions and meet the design requirements of stable and efficient operation of the variable-speed power turbine within a wide range of operating conditions.

The present invention discloses a self-detection method, system, device and storage medium for an over-speed protection system. The method adopts the power turbine speed as a trigger condition and can respectively perform self-detection on the hardware circuit and control switch of the over-speed logic processing unit of the over-speed protection system, so as to facilitate rapid fault location when the over-speed protection function fails. Moreover, the self-detection process of the hardware circuit and the control switch of the over-speed logic processing unit will not cause actual fuel cut-off, thereby eliminating the need for special development of engine start-up detection programs. Moreover, dynamic detection can be performed in an actual installation and use environment without the need for special ground detection equipment. Therefore, the over-speed protection function can be self-detected during normal engine testing and operation, ensuring that the over-speed protection function is intact during engine operation.

The application discloses a turbo-shaft engine ground bench performance debugging method based on turbine guide vane area, comprising the following steps: obtaining different gas turbine guide vane areas A GT and power turbine guide vane areas A PT under the condition of same gas generator physical rotating speed t4.5 , corresponding gas temperature change data ΔT, power change data ΔP, and forming gas temperature change and power change tables respectively; obtaining different A GT and A PT under the condition of same take-off power g , corresponding specific fuel consumption change data Δsfc and rotating speed change data Δn, and forming specific fuel consumption change table and rotating speed change table respectively; obtaining A GT and A PT corresponding to the change data which simultaneously satisfy the related constraint conditions from the two-dimensional linear interpolation of each table, and taking A GT and A PT as candidate A GT and A PT ; and finding A GT and A PT with the minimum adjustment amount for performance debugging. The application can improve the debugging efficiency and reduce the debugging cost.

The invention discloses a power turbine rotor dismounting device and method, and the device comprises a first supporting piece, the bottom of which is provided with a first supporting part, and the first supporting part is used for extending between a previous stage turbine disc and a next stage turbine disc so as to abut against and cooperate with the upper edge of the previous stage turbine disc; a second supporting part is arranged at the bottom of the second supporting piece, and the second supporting part is used for stretching into the position between the previous-stage turbine disc and the next-stage turbine disc so as to be matched with the lower edge of the next-stage turbine disc in an abutting mode; the first limiting piece is used for being connected with the first supporting piece; the second limiting piece is used for being connected with a second supporting piece, and the second supporting piece and the first supporting piece are distributed in the circumferential direction at intervals; the limiting mechanism is used for limiting rotation of the first supporting piece and the second supporting piece; the adjusting mechanism is connected with the first limiting piece and the second limiting piece and used for adjusting the axial distance between the first limiting piece and the second limiting piece so as to adjust the axial distance between the first supporting part and the second supporting part.

A propulsion system for an aircraft includes a propulsor section that includes a fan with a plurality of fan blades that are rotatable about a fan axis, a core engine that is configured to generate a gas flow, a power turbine that is mechanically uncoupled from the core engine and rotatable independent of the core engine, the gas flow from the core engine is expanded through the power turbine to generate shaft power, a support structure where the power turbine is mounted relative the core engine, a fan drive gear system that is driven by the power turbine and at least partially mounted to the support structure, and a power turbine shaft that includes a first coupling to the fan drive gear system and a second coupling to the propulsor section.

The purpose of the present invention is to provide an exhaust system and design method based on a reversible power turbine, including an adaptive switching structure for turbine outlet working conditions, including an adaptive transition baffle structure, a support plate structure, a turbine section, and an exhaust volute section. The adaptive transition baffle structure is located at the outlet end of the turbine middle casing. The working state of the adaptive transition baffle structure rotates and changes automatically with the blowing of the airflow in different flow channels of the turbine. The support plate structure is located between the turbine section and the exhaust volute section. When the reversing turbine of the present invention is in the forward working condition, the baffle opens outward to block the forward flow channel airflow from flowing back into the reverse flow channel, effectively reducing flow losses; when the reversing turbine is in the reverse working condition, the baffle contracts inward, which can also reduce the backflow loss to a certain extent. The exhaust volute can further expand the turbine outlet airflow to slow down and cool it down, and rotate the airflow direction by 90° to discharge vertically upward to meet emission standards.

The present application provides a self-locking power turbine integrated output structure, which belongs to the field of gas turbine technology. The self-locking power turbine integrated output structure includes: a speed sensor fixedly connected to the bearing seat casing; a transmission flange that moves synchronously with the output shaft through a spline structure, the front end of the transmission flange is provided with a cantilevered speed measuring gear structure, and the speed measuring gear structure matches the speed sensor in the axial position; and a shaft end plug connected to the rear end of the transmission flange, through which the output shaft is blocked. The self-locking power turbine integrated output structure provided by the present application increases the internal space of the bearing seat casing by integrating the function of the speed measuring gear into the transmission flange, realizes a good layout space for the bearing and the sealing structure, and facilitates the disassembly and maintenance of the output shaft.

The invention relates to the technical field of gas turbine control, and discloses an intelligent fuzzy self-tuning PID optimization control method for the rotating speed of a gas turbine, which adopts the intelligent fuzzy self-tuning PID optimization control method for the rotating speed of the gas turbine based on an improved cloud drift optimization algorithm, is used for adaptively tuning the control parameters of fuzzy PID, can improve the fuzzy PID performance, and can improve the stability of the rotating speed of the gas turbine. The control effect of the gas turbine under complex conditions is improved, and the control precision and robustness of the gas turbine are improved. Compared with an existing classical intelligent optimization algorithm, the method is high in convergence speed and good in optimization effect, the precision of gas turbine power turbine rotating speed control is effectively improved, the adjusting time is shortened, and the overshoot is reduced.

The invention discloses a vertical torque measuring tooth adjusting device, the vertical torque measuring tooth adjusting device abuts against a power turbine shaft, a through cavity is arranged in the power turbine shaft along the length direction, a torque measuring shaft is arranged in the through cavity and is in clearance fit with the power turbine shaft, and the vertical torque measuring tooth adjusting device comprises a supporting assembly, a fixing assembly and an ultrasonic vibration assembly, the fixing assembly comprises a fixing base and a fixing piece, the fixing base is connected with the supporting assembly, the fixing piece is connected with the fixing base, the power turbine shaft is fixed to the fixing base through the fixing piece, and the power turbine shaft is in the vertical direction; the ultrasonic vibration assembly comprises a first ultrasonic vibration structure, the first ultrasonic vibration structure is installed on the supporting assembly, and the first ultrasonic vibration structure abuts against the periphery of the power turbine shaft. Ultrasonic waves emitted by the ultrasonic vibration assembly are applied to the power turbine shaft, so that micro motion is generated between the torsion measuring shaft and the power turbine shaft, the clamping stagnation or hysteresis phenomenon in the torsion measuring shaft adjusting process is reduced or eliminated, and adjustment is more accurate.

This application discloses a method, apparatus, equipment, and medium for calibrating the torque of an aero-gas turbine shaft engine based on a torque meter. The method includes the following steps: S1, engine steady-state performance testing: starting the engine to idle, then slowly increasing the engine to several takeoff torque states, recording the torque values ​​corresponding to the power turbine shaft torque sensor and torque meter at each takeoff torque state; S2, torque calibration: selecting N=2 takeoff torque states for the engine, the torque values ​​corresponding to the power turbine shaft torque sensor and torque meter, and obtaining the calibration function relationship through linear fitting; S3, post-calibration check: calculating the calibrated torque values ​​of the power turbine shaft torque sensor at several takeoff torque states based on the calibration function relationship. If the differences between each calibrated torque value and the torque value corresponding to the torque meter all meet the requirements, then the torque calibration meets the requirements. This application improves the accuracy of engine torque calibration.