321 results about "Power turbine" patented technology

The invention discloses a starter over-rotation protection control method and system, electronic equipment and a storage medium, and the method comprises the steps: firstly collecting a power turbine transmission chain tail end rotation speed signal and a gas generator rotation speed signal, and then judging whether the gas generator rotation speed reaches a first preset rotation speed or not; and if the rotating speed of the gas generator reaches a first preset rotating speed, the decreasing rate of the rotating speed at the tail end of the power turbine transmission chain is larger than a rotating speed decreasing rate threshold value, and the duration time exceeds preset time, or if the decreasing amplitude of the rotating speed at the tail end of the power turbine transmission chain exceeds a preset value, parking protection is started. If the rotating speed of the gas generator does not reach the first preset rotating speed all the time, it can be judged that the starter is not really ignited, the starter is in the test run test process of starting (oil) sealing, false starting, cold running and the like, at the moment, the power turbine does not have the over-rotating problem, and even if the decreasing rate or the decreasing amplitude of the rotating speed of the tail end of a power turbine transmission chain meets the preset conditions, the engine cannot be ignited. And the parking protection is not started, so that the normal test run test is not influenced.

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 maximum unbalance test method and system for a double-rotor turboshaft engine, and the method comprises the steps: analyzing the vibration mode and working speed of a rotor, determining the test content of each rotor, formulating the design requirements of a vibration test scheme of a whole machine, a rotor / shaft system, an accessory and a pipeline, and determining the maximum unbalance of the whole machine according to the test content. A fuel gas generator rotor acceleration and deceleration test and a power turbine rotor acceleration and deceleration test are respectively carried out by adopting an Np rotating speed open-loop control mode, and finally vibration test data are analyzed to verify the vibration level of each part. Due to the adoption of Np rotating speed open-loop control, follow-up change of the power turbine rotor under the action of airflow due to the change of the rotating speed of the fuel gas generator rotor is avoided, and the rotating speed of the fuel gas generator rotor does not need to be changed when the rotating speed of the power turbine rotor is changed; therefore, independent verification of the fuel gas generator rotor and the power turbine rotor can be realized, and interference is avoided.

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 invention belongs to a part surface treatment technology, and relates to a multi-runner blade inner cavity and air film hole abrasive particle flow polishing and grinding device and a using method. The device is composed of an upper hydraulic material grinding bin, a lower hydraulic material grinding bin and a blade workbench of abrasive particle flow equipment. The blade workbench can prevent pressure loss caused by abrasive leakage in abrasive flow polishing and grinding and meanwhile avoid flow offset caused by simultaneous entering of the abrasive at the communicating positions of inner cavity flow channels of a multi-flow-channel blade, so that all the flow channels of the blade have a good polishing and grinding effect. The defects that the inner cavity of the power turbine blade interferes with the air film hole wall structure, and the surface strengthening mode is unreachable are overcome, the grinding materials are made to move repeatedly in the inner cavity of the blade through repeated pressurization of the hydraulic cylinder, and the uniform strengthening effect is achieved at the inner cavity of the blade and the inner and outer hole openings of the air film hole. According to the method, only a small amount of grinding materials are consumed and are far less than vibration finishing grinding materials, the machining cost is lower, the dust amount is smaller, and noise is lower. According to the device, the inner cavity of the whole guide vane is strengthened within 20 min through abrasive flow, and the surface strengthening machining efficiency is high.

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 powerplant is provided for an aircraft. This powerplant includes an engine housing, an engine assembly, an electric generator, an electric motor and a propulsor rotor. The engine assembly is arranged in the engine housing. The engine assembly includes a core flowpath, a core compressor section, a core combustor section, a core turbine section and a power turbine section. The power turbine section includes a power turbine rotor. The electric generator is arranged in the engine housing. The electric generator includes a generator rotor. The power turbine rotor is configured to drive rotation of the generator rotor. The electric motor is arranged in the engine housing. The electric motor includes a motor rotor. The electric generator is configured to power operation of the electric motor. The motor rotor is configured to drive rotation of the propulsor rotor.

The invention provides a starter test run control method for an aero-engine, and relates to the technical field of engine test run control, in a test run starting time period, an output shaft of a starter is kept at a low rotating speed, a power turbine rotor is prevented from rotating over, and after the test run starting time period is over, the output shaft of the starter is kept at a low rotating speed. Taking a standard target value of the rotating speed of the output shaft of the starter as a set value, entering a rotating speed control mode, increasing the rotating speed of the output shaft of the starter, and triggering the starter to stop at a stop triggering moment; increasing the set value by considering that the rotational inertia of the output shaft of the starter is greater than the rotational inertia of a rotor of the starter and the output shaft of the starter is in an inertia uprush state; the rotating speed of the output shaft of the starter is continuously increased, the load on the output shaft of the starter is reduced, and if the rotating speed of the output shaft of the starter in an inertia uprush state cannot reach a set value, the load applied to the output shaft of the starter is continuously reduced on the basis of reducing the load applied to the output shaft of the starter; the overrun clutch inner sleeve and the overrun clutch outer sleeve cannot be meshed under the condition of high rotating speed until the rotating speed reduction rate of the starter output shaft is smaller than the rotating speed reduction rate of the starter rotor, vibration impact is avoided, and damage to the overrun clutch is reduced.

The invention discloses a method for assembling a power turbine transmission shaft assembly of a novel torque measuring structure. The method comprises the following steps that S1, pin holes in a power turbine transmission shaft, a torque measuring reference shaft and any torque measuring tooth are subjected to single-piece machining; s2, adjusting the position of the torsion measuring reference shaft in the long shaft by using an angle adjusting tool, and fixing the large end of the torsion measuring reference shaft with any torsion measuring tooth by using a centering tool; s3, aligning a hole in any torque measuring tooth with a hole in the large end of the torque measuring reference shaft, and combining a pin; s4, designing a positioning ring, and dividing inner holes of the positioning ring into five groups according to the excircle size of the torque measuring teeth; s5, connecting the hole of the torque measuring tooth and the hole of the torque measuring reference shaft through a step pin, and fixing the torque measuring tooth in the angular direction and the axial direction; the machining and assembling method is high in assembling precision, capable of preventing part deformation and good in stability.

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 discloses a turboshaft engine complete machine surge boundary determination method, system and device, and belongs to the technical field of turboshaft engines, the turboshaft engine complete machine surge boundary determination method comprises the following steps: determining whether an engine power turbine overrotates; if the power turbine is over-rotating, over-rotating protection is carried out, and the fuel flow is adjusted to the set flow; and after the fuel flow is adjusted, power turbine over-rotation oil drainage reignition tests under different engine powers are carried out, and the engine surge boundary is determined according to test results. According to the method, the complete surge boundary of the whole engine state can be obtained through the dynamic fuel step method based on the over-running oil drainage reignition test, the engine and a test bed do not need to be greatly transformed, only the over-running oil drainage time sequence needs to be adjusted in the test process, and the method is easy to implement.

A turboshaft engine for an air-craft including a gas generator including a compressor, a combustion chamber and an expansion turbine; a power turbine rotating a power take-off by a reduction gear; a heat exchanger including a first circuit and a second circuit. The compressor includes a first shaft rotated by a second shaft of the expansion turbine by a trans-mission mechanism, the transmission mechanism and the reduc-tion gear forming part of a gearbox which is arranged axially at a front end of the turboshaft engine, such that the compressor is arranged axially between the gearbox and the power turbine.

The invention belongs to the field of aero-engines, and particularly relates to a three-rotor mechanism with front interstage supports, an engine and a blade profile optimization method. According to the three-rotor mechanism with the front interstage support, a turbine meridian inner end wall is arranged in a turbine meridian outer end wall to form an annular channel; a high-pressure turbine assembly, a low-pressure turbine assembly and a power turbine assembly are sequentially arranged in the annular channel from the inlet end to the outlet end of the annular channel. The low-pressure turbine assembly comprises interstage supporting guide vanes and working vanes, the interstage supporting guide vanes are arranged close to the high-pressure turbine assembly, and the working vanes are arranged close to the power turbine assembly. By integrally designing the low-pressure turbine guide vane and the interstage support, the preposition of the interstage support is realized, the coordination compactness of the three-rotor mechanism and the whole engine structure is improved, the axial span of a power turbine rotor fulcrum in a front output shaft form is effectively shortened, and the robustness of the power turbine rotor structure is enhanced.

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.

A turboshaft engine for an air-craft including a gas generator including a compressor, a combustion chamber and an expansion turbine; a power turbine rotating a power take-off by a reduction gear; a heat exchanger including a first circuit and a second circuit. The compressor includes a first shaft rotated by a second shaft of the expansion turbine by a trans-mission mechanism, the transmission mechanism and the reduc-tion gear forming part of a gearbox which is arranged axially at a front end of the turboshaft engine, such that the compressor is arranged axially between the gearbox and the power turbine.

The invention provides an inter-turbine casing deflation structure with a closed channel, and belongs to the technical field of gas turbines. The power turbine case is connected with the bearing outer case; the deflation pipeline is mounted on the force-bearing outer casing; the guide blades are mounted on the outer ring; the outer ring is connected with the force-bearing outer casing and the power turbine casing, a cooling cavity is formed between the force-bearing outer casing and the outer ring, a deflation hole is formed in the front section of the outer ring, a boss structure is arranged on the outer ring, the deflation supporting ring and the deflation mounting base are fixed to the boss structure after being combined, and the mouthpiece is mounted in the deflation mounting base and the deflation pipeline to form a closed channel; the sealing ring is arranged at the front end of the outer ring, a deflation flow channel is formed between the sealing ring and the outer ring, and gas in the main flow channel flows out from the deflation pipeline along the deflation flow channel and the closed channel; the hollow support plate is arranged in the force-bearing outer casing on the front side of the deflation flow channel, and cooling holes for forming a cooling flow path are formed in the hollow support plate, the sealing ring and the outer ring.

The invention discloses a self-detection method and system for an over-rotation protection system, equipment and a storage medium, which adopt the rotating speed of a power turbine as a trigger condition, and can respectively perform self-detection on a hardware circuit and a control switch of an over-rotation logic processing unit of the over-rotation protection system, so as to quickly position a fault when an over-rotation protection function fails, and improve the reliability of the over-rotation protection system. Moreover, the self-detection process of the hardware circuit and the control switch of the overrun logic processing unit does not cause real fuel cut-off, so that an engine driving detection program does not need to be specially researched and developed, dynamic detection can be carried out in an actual installation and use environment, a special ground detection device is not needed, and the detection efficiency is improved. Therefore, the over-rotation protection function can be self-detected in the normal engine test and working process, and it is ensured that the over-rotation protection function is intact in the working process of the engine.

The invention provides a fuel oil control method for the high-working-condition sudden load unloading process of a power generation type gas turbine, and belongs to the technical field of gas turbine control. Continuously monitoring the power generation power of the power generation type gas turbine, when the tripping signal is effective and the power generation power is greater than or equal to the preset load operation, judging that the gas turbine has high-working-condition sudden load unloading, quitting the rotating speed closed-loop control mode and executing the fuel oil open-loop control mode at the moment; in the fuel oil open-loop control mode process, constant-speed-rate oil cutting is started to be executed, meanwhile, the minimum oil supply amount corresponding to the minimum oil pressure ratio limiting curve is executed, and the higher one of the fuel oil supply amount corresponding to the constant-speed-rate oil cutting and the minimum oil supply amount corresponding to the minimum oil pressure ratio limiting curve is selected for fuel oil control; the rotating speed of the power turbine is continuously monitored, and when the rotating speed of the power turbine is reduced to a first preset interval, the fuel oil open-loop control mode is quitted; and when the rotating speed of the power turbine continues to be reduced to a second preset interval, normal control is recovered.

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 discloses a cross-third-order bending critical power turbine rotor test device and method, and belongs to the technical field of aero-engines, the cross-third-order bending critical power turbine rotor test device comprises a supporting mechanism used for supporting a simulation rotor, a power mechanism, a control system used for controlling the power mechanism, a lubricating oil mechanism and a test system, the power turbine first-stage wheel disc and the power turbine second-stage wheel disc are used for being connected with a rotor shaft, are in wheel disc shapes and are connected with each other, the bearing mechanism comprises a test platform, a front support, a rear support and a vacuum box, the vacuum box is arranged on the test platform, and the vacuum mechanism is connected with the vacuum box. The vacuum mechanism is used for vacuumizing the vacuum box, the front support and the rear support are arranged on the test platform in a spaced mode and located in the vacuum box, the front support and the rear support are used for clamping and supporting a rotor shaft, and the rotor shaft is connected with the power mechanism. The simulation rotor replaces a real rotor to carry out the test, the processing period is greatly shortened, the test cost is reduced, and the test risk is reduced.

The invention discloses a method for repairing an armored thermocouple for measuring the temperature of a steam turbine cylinder of a thermal power unit, which relates to the technical field of steam turbine maintenance, and includes the following steps: cleaning the exposed thermocouple wire; grinding the surface of the thermocouple wire with 300 to 500 mesh sandpaper; dispersing the ground thermocouple wire; coating the thermocouple wire; wrapping and fixing the thermocouple wire after coating with a glass ribbon; and covering the repaired part with a thin aluminum plate until the thermocouple wire works normally. Compared with the traditional replacement and repair of armored thermocouples, the invention is simple to operate and low in cost, and can realize online rapid recovery of the insulation and resistance of the thermocouple after the explosion, so that the damaged armored thermocouple can be used normally again, which greatly reduces the cost of thermocouple repair; while brushing, the mixed liquid can be crushed and stirred, which avoids the waste of the mixed liquid while ensuring the uniformity and stability, and greatly improves the utilization efficiency of the mixed liquid.

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.