24 results about "Power turbine" patented technology

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 application discloses a power turbine rotor combination blade root X-ray detection method. The power turbine rotor combination comprises a long rod and a wheel disc located on two sides of a blade respectively, and the wheel disc has a bottom recess at the center. The X-ray directional penetration is implemented on all blade roots of the power turbine rotor combination in the circumferential direction, and a film is tightly attached to the circular arc surface of the annular structure during the X-ray directional penetration. The detection method creatively regards the blade root as an annular structure, reduces the penetration thickness, can significantly reduce energy loss and improve detection sensitivity, realizes complete image of the detected part, improves the contrast of the film, and guarantees stable and reliable detection quality of the blade root. The directional penetration device is used for auxiliary detection, can quickly clamp and stabilize the power turbine rotor combination, guarantees that the ray beam is perpendicular to the blade root, and tightly attaches the film and the effective penetration area, and improves the image contrast and the detection sensitivity.

The application relates to the technical field of aero-engines and discloses a starting fuel supply law control method and device for an aero-unmanned engine, which comprises the following steps: obtaining the rotating speed and acceleration of a target moment and a historical moment, judging whether a first switching condition is met, obtaining the power turbine inlet temperature of the target moment and the historical moment if the first switching condition is met, judging whether a second switching condition is met, obtaining the fuel given quantity and the fuel feedback quantity if the second switching condition is met, judging whether a third switching condition is met, judging whether a fourth switching condition is met based on a first time consumption, a second time consumption and a third time consumption if the third switching condition is met, judging whether a fifth switching condition is met based on the first time consumption, the second time consumption and a fourth time consumption if the fourth switching condition is met, turning into a starting third stage in advance if the fifth switching condition is met, and executing the fuel supply law of the starting third stage. Through multi-stage judgment, the application realizes timely and effective switching of the fuel supply law, avoids triggering the suspension protection, and improves the engine starting success rate.

This invention relates to the field of aero-engine technology, specifically to an engine gearbox and an aero-engine having the same. The engine gearbox includes a gearbox body, within which a power output shaft assembly is integrated. The input end of the power output shaft assembly is adapted for transmission with the engine body, and the output end is adapted for transmission with a load mechanism. A central gear is mounted on the power output shaft assembly. An accessory drive gear assembly meshes with the central gear and is adapted for transmission with an accessory mechanism. This allows the load mechanism to rotate synchronously at high speed with the engine's power turbine rotor. Simultaneously, a portion of the power is transmitted to the accessory mechanism via the accessory drive gear assembly through the central gear, enabling the power turbine rotor's rotational speed to drive the accessory mechanisms at their respective suitable speeds after gear shifting.

The application relates to a high-rotating-speed small-power turbine performance test method and belongs to the technical field of engine testing. The application improves a test process, utilizes power consumption characteristic test under one-to-one corresponding rotating speed conditions of turbine output characteristic test, accurately measures the change relation of the comprehensive power consumption of bearings, seals and cooling flow paths with test rotating speed, and corrects the turbine output torque and power according to the change relation, realizes the technical effect of correcting the output characteristic by considering the rotating speed change and the influence of test times, and significantly improves the test precision; meanwhile, the application sets one power consumption characteristic test before and after the turbine output characteristic test, adopts a quadratic curve to simulate the change rule of the comprehensive power consumption with running time (wear) in the test product, obtains the actual power consumption of each output characteristic test, realizes the technical effect of correcting the output characteristic by considering the influence of the running time (wear), and accurately obtains the turbine efficiency under different pressure ratio conditions.

The application discloses a torsional moment load loading device for a toothed sleeve connecting structure of an aero-engine power turbine rotor, which comprises a test-bed base, a power turbine rotor test piece, a front supporting seat, a rear supporting seat, a torsional moment load loading structure, an L-shaped support, a torsional load loading threaded support, an angular limiting spline shaft adapter plate and an angular limiting spline, wherein the front supporting seat and the rear supporting seat are used for supporting the power turbine rotor test piece; the torsional moment load loading device is installed between a compressor and a turbine of the power turbine rotor test piece, the L-shaped support is used for restraining the torsional moment load loading structure; and the angular limiting spline shaft adapter plate and the angular limiting spline exert circumferential restraint on a power output shaft of the power turbine rotor. The application can exert torsional moment load on the toothed sleeve connecting structure rotor test piece, and simulate the aerodynamic torque effect of the power turbine on the rotor in a working state.

This application provides a method and apparatus for designing control laws for a thermally managed combined power unit, belonging to the field of aerospace electromechanical systems. The method includes: establishing mathematical models for the core components of the thermally managed combined power unit; constructing an initial steady-state simulation model by combining flow rate, pressure, and energy conservation relationships; revising the initial steady-state simulation model to obtain a faithful steady-state simulation model; building a transient simulation model based on the faithful steady-state simulation model, incorporating mode switching logic, fuel supply strategy, control target selection rules, and system safety boundaries, and designing a basic control law; employing a particle swarm optimization algorithm, using the fuel control PID parameters of the transient simulation model as the optimization object, and using mode switching time, system overshoot, and power turbine exhaust temperature as optimization indicators to optimize the control parameters of the transient simulation model; updating the optimized parameters to the transient simulation model to obtain a transient simulation model of the thermally managed combined power unit.

This invention provides a bypass system for an intercooled cycle gas turbine, comprising a three-shaft intercooled cycle gas turbine, a low-pressure compressor, a high-pressure compressor, a combustion chamber, a high-pressure turbine, a low-pressure turbine, a power turbine, and an intercooler. An intercooler inlet transition section is provided between the low-pressure compressor and the intercooler, and an intercooler outlet transition section is provided between the intercooler and the high-pressure compressor. By adding a bypass system with a control valve to the intercooler inlet transition section and the intercooler outlet transition section of the three-shaft intercooled cycle gas turbine, before the intercooled cycle gas turbine starts, an electric actuator first opens the control valve, placing the control valve on the bypass flow path in the open position. After successful and stable combustion in the combustion chamber, once the high-pressure rotor speed reaches a certain value, the control valve is gradually closed, eventually reaching the closed position. This technology enables the rapid start-up of the three-shaft intercooled cycle gas turbine.

A first core turbine rotor of a first propulsion system is disposed in a first core turbine section and is configured to rotate in a first rotational direction. A first power turbine rotor is disposed in a first power turbine section and is configured to rotate in the first rotational direction. The first power turbine rotor is operatively coupled to a first open propulsor rotor. A second core turbine rotor of a second propulsion system is disposed in a second core turbine section and is configured to rotate in the first rotational direction. A second power turbine rotor is disposed in a second power turbine section and is configured to rotate in a second rotational direction that is opposite the first rotational direction. The second power turbine rotor is operatively coupled to a second open propulsor rotor.

This invention relates to the field of gas turbine engine technology, specifically to a turboshaft engine aerodynamic adjustment system and its control method. The turboshaft engine aerodynamic adjustment system includes: a compressor; a particle separator with a first air intake on its inner flow channel wall; a first air conditioning flow path, one end of which is connected to the compressor and the other end to the first air intake, the first air conditioning flow path being adapted to adjust the flow rate of compressed air in the introduced airflow channel to adjust the flow field of the particle separator; a power turbine guide vane with a second air intake; and a second air conditioning flow path, one end of which is connected to the compressor and the other end to the second air intake, the second air conditioning flow path being adapted to adjust the flow rate of compressed air introduced into the inter-vane passage of the power turbine to adjust the flow function of the power turbine. The turboshaft engine aerodynamic adjustment system provided by this invention can actively adjust component characteristics according to changes in flight conditions and environmental factors, improving engine performance and environmental adaptability.

This invention pertains to the field of engine technology and proposes an overall design method for upgrading aero-engine components. The method includes: ensuring the ratio of intake loss to reference intake pressure loss is within a preset range; meeting the performance requirements of the compressor (pressure performance), the combustion-related performance of the combustion chamber (combustion-related performance), the rotational inertia of the gas turbine (rotational inertia), the inlet temperature, and the performance requirements of the gas turbine; meeting the rotational inertia requirements of the reference power turbine and the turbine flow area requirements of the power turbine; and meeting the pressure loss requirements of the reference exhaust section and the exhaust velocity requirements of the exhaust section. This allows for the simultaneous or individual installation of improved components on a reference engine to enhance engine power and lifespan, among other technical indicators.

A hydrokinetic turbine system with dynamic tuning capabilities is disclosed. Individual hydrokinetic turbine units are dynamically tuned to accommodate changes in height and flow velocity corresponding to water in a waterway. Dynamically tuning the turbine units to accommodate waterway changes optimizes power generation output. Dynamically tuning a turbine system includes raising or lowering turbine blade height, extending or retracting turbine blade length, and narrowing or widening a turbine mouth, channel, and exit through which water flows. The hydrokinetic turbines may be arranged in an array along a waterway, and each hydrokinetic turbine in the array is connected over a controls system configured to adjust turbine characteristics at each turbine unit in the array for optimizing power generation output for the waterway in which the turbine array is installed.

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.

The application discloses a power turbine and gas generator butt joint assembly method, and the butt joint assembly method comprises the following steps: sequentially assembling and balancing a power turbine rotor; measuring a guide vane throat area of the power turbine; sequentially performing general assembly of the power turbine; disassembling a first guide vane from the power turbine, hoisting the first guide vane and adjusting the first guide vane to be horizontal, and then butt joining the first guide vane with a gas generator; and butt joining the gas generator and the power turbine through a first guide vane rear mounting edge of the power turbine, leveling a part of the power turbine after the first guide vane is disassembled, butt joining the part with the first guide vane of the gas generator which has been butt joined, and completing the assembly. The method is simple and easy to implement, has no safety hidden danger, is convenient for workers to quickly master the butt joint technology of the power turbine and the gas generator, and can effectively improve assembly efficiency and qualification rate.

The application discloses an aero-engine lubricating oil interruption test method, sets monitoring parameters in the S1 step, records the monitoring parameters of the engine in the S2 step to the S4 step, obtains monitoring parameter alarm values through the monitoring parameters, interrupts the lubricating oil supply of the engine in the S7 step, and monitors whether the second current gas generator rotating speed, the second current output power, the second current power turbine inlet temperature and the second current ball bearing wall temperature exceed the corresponding first alarm values within the second preset time. Since the signs of bearing failure are directly judged through the monitoring parameters, the situation that the signs of bearing failure are judged by artificial through abnormal phenomena is avoided, the efficiency and accuracy of judging the signs of bearing failure are improved, the monitoring parameters of the engine are obtained through the S2 step to the S4 step, the monitoring parameter alarm values are obtained according to the monitoring parameters, the monitoring parameter alarm values correspond to the actual situation of the engine, and the inaccurate test results caused by the differences between different engines are avoided.

The present application relates to a kind of engine loses load power turbine maximum speed prediction method and system, belong to aviation gas turbine engine field including the following steps: obtaining the pressure signal of the compressor outlet of engine, the pressure signal of the main oil line of fuel distribution device and the pressure signal of the auxiliary oil line of fuel distribution device;Maximum state simulation overspeed test is carried out to engine, and the fuel law of engine maximum state overspeed is obtained according to test result;The power turbine speed of different time points in the maximum state overspeed process of engine is calculated by section iteration, and the relationship between power turbine speed and time is obtained;According to the relationship between power turbine speed and time, the maximum power turbine speed of engine maximum state overspeed process is obtained.The present application fully considers the response time of overspeed protection system, fuel attenuation law and torque variation law, and can accurately predict the maximum speed of engine broken shaft.

This invention relates to the field of gas turbine technology and discloses a gas turbine fuel regulation and condition protection control device, including a gas turbine unit and a fuel controller. The gas turbine unit includes a gas generator and a power turbine. The gas generator includes a compressor. The output end of the gas turbine unit is connected to the input end of the fuel controller. The output end of the fuel controller is connected to both the input end of the gas turbine unit and the input end of the fuel controller. The fuel controller includes an over-temperature protection limiting module, an over-pressure protection limiting module, an over-speed protection limiting module, a fuel quantity boundary limiting module, a transition state boundary limiting module, a fuel main control module, and an optimal selection logic module. The invention also discloses a gas turbine fuel regulation and condition protection control method. This invention can adapt to the full operating conditions of gas turbines and achieve precise fuel quantity regulation and multiple protections, providing a guarantee for the safe and efficient operation of gas turbines throughout their entire life cycle.

The application provides a turboshaft engine and belongs to the technical field of engines.The turboshaft engine comprises a gas turbine, a power turbine, a transmission device, a second speed reducer, a main transmission shaft and a transmission sleeve.The transmission device comprises an accessory gear train and a first speed reducer.The first speed reducer, the power turbine and the second speed reducer are sequentially connected in transmission through the main transmission shaft.The transmission sleeve is rotatably installed on the outside of the main transmission shaft, and the main transmission shaft is coaxial with the transmission sleeve.The accessory gear train and the gas turbine are fixedly installed on the transmission sleeve, and the accessory gear train is located on the side of the gas turbine close to the first speed reducer.The main transmission shaft and the transmission sleeve of the application adopt concentric output, which reduces the aircraft profile size, reduces the windward area and is beneficial to the aerodynamic shape design of the aircraft.The main transmission shaft and the transmission sleeve make the structure layout of the turboshaft engine compact compared with the same power level turboshaft engine with the existing configuration, which is beneficial to reducing the weight of the engine.

The present application relates to starter over-speed protection technical field, specifically relates to a kind of gas turbine starter over-speed protection control device and control method.A kind of gas turbine starter over-speed protection control device, comprising: control box, it is suitable for being located in the outside of gas turbine starter, control box is equipped with control box, gas turbine starter includes gas generator unit body and power output unit body;First rotational speed sensor, it is located on the gas generator casing of gas generator unit body, first rotational speed sensor is suitable for measuring the first rotational speed of gas generator rotor;Second rotational speed sensor, it is located on the power output casing of power output unit body, second rotational speed sensor is suitable for measuring the second rotational speed of sky axis gear, to indirectly reflect the rotational speed of power turbine.The present application provides a kind of gas turbine starter over-speed protection control device and control method, to solve the problem of power turbine acquisition failure when transmission chain is abnormal, power turbine over-speed breakage.

The application provides a parallel hybrid optimization control method and device, and relates to the technical field of vehicle hybrid control. The method comprises the following steps: determining the range of a first output variable and a second output variable; collecting a hybrid control data set under different throttle instructions according to the range and a hybrid system; training a hybrid simulation model according to the data set; solving an optimal control variable based on the simulation model and a first part of data in the data set; training an optimization control model with the first part of data as input and the optimal control variable as output, wherein the optimal control variable comprises the first output variable for controlling a power turbine and the second output variable for controlling a motor, and the power turbine and the motor are connected in parallel; and controlling the power turbine and the motor based on the trained optimization control model. The application can realize the optimal matching of engine and motor control variables, reduce the interference of motor response on the engine, reduce the risk of over-speed and over-temperature, and improve the system response performance and stability.

This invention relates to the field of blade cleaning technology and discloses a blade cleaning device for thermal power turbines, comprising: a support base, a dry ice insulation hopper on the upper part of the support base, the dry ice insulation hopper being connected to a Venturi solid-gas two-phase mixer, the air inlet end of the Venturi solid-gas two-phase mixer being connected to a high-pressure air compressor; the discharge end of the Venturi solid-gas two-phase mixer being connected to a high-pressure flexible hose, the discharge end of the high-pressure flexible hose being connected to a dry ice cleaning gun and a nozzle; a fixing block being installed at the center of a valve seat, a through-type material passage chamber being opened at the center of the fixing block; a guide rod being axially arranged at the center of the valve seat, a first sealing valve core and a second sealing valve core being slidably sleeved on both sides of the guide rod; and a reset element being sleeved at the center of the guide rod. This device forms a bidirectional hard seal when the pipeline pressure fluctuates, blocking the reverse flow of high-pressure airflow, preventing accelerated sublimation of dry ice particles and instability in the feeding process, and ensuring the stability of the solid-gas two-phase flow transportation.

A turbo-compressor system is disclosed comprising a multi-shaft gas turbine for generating power. The gas turbine comprises a gas producer producing a flow of exhaust gas, a power turbine or low pressure turbine and an internal coupling shaft linking these elements. The operation of the gas turbine is defined by flow rate and pressure difference. The system further comprises a rotary compressor with rotating blades for carrying out gas compression and connected to the power turbine via a main coupling shaft. The turbo-compressor system is characterized by the provision of inlet guide vanes at the compressor inlet and a variable area turbine nozzle between the gas producer and the power turbine to enable adjustment of both the compressed gas flow and the exhaust gas flow. The design of the system enables control of the operating point of the gas turbine by coordinated manipulation of the inlet guide vanes and the variable area turbine nozzle.