280 results about "Turboshaft" patented technology

The embodiment of the invention provides a digital simulation modeling method and device of a tandem type hybrid electric propulsion system, computer equipment and a medium, and relates to the technical field of simulation modeling, and the method comprises the following steps: constructing a turboshaft prime mover model, constructing a generator model, connecting the turboshaft prime mover model and the generator model in series to generate a generator set, the power generation groups are connected in parallel to construct a turbine power generation system; constructing a propulsion system; constructing a battery energy storage system; constructing a power transmission and distribution system, and respectively connecting the power transmission and distribution system with a turbine power generation system, a propulsion system and a battery energy storage system; and constructing a comprehensive control and energy management system, and outputting a digital simulation modeling result through a turbine power generation system, a power transmission and distribution system and a battery energy storage system. According to the scheme, a simulation modeling method based on multi-physics field coupling double-power-source dynamic response and energy flow regulation and control provides reliable modeling technical support for engineering application of a tandem type hybrid electric propulsion system.

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 turboshaft engine dynamic process performance parameter prediction method considering thermal inertia, and belongs to the field of aero-engine fault diagnosis. According to the method, the influence of thermal inertia on gas path parameters in the dynamic process of the turboshaft engine is analyzed, a thermal inertia coefficient is designed and introduced, a hot end component heat exchange model is established, a component characteristic diagram is corrected in combination with real test run data, and performance parameters in the dynamic process are predicted in real time through a neural network prediction algorithm. According to the method, the hysteresis phenomenon of gas temperature change in the dynamic process can be effectively described, the accuracy of dynamic modeling of the hot end component is improved, the capability of predicting dynamic thermodynamic parameters of the engine in real time is improved, and a scientific basis is provided for formulating a prediction health management strategy of the turboshaft engine.

The invention discloses a rapid dynamic response control method for a turboshaft engine. According to the method, a gas turbine rotation acceleration cascade control structure is used for controlling the turboshaft engine; a feed-forward compensation form with predicted required power as a feed-forward compensation amount is introduced into an inner ring fuel flow instruction control position of a gas turbine rotation acceleration cascade control structure; introducing a feed-forward compensation form which takes a required power prediction error as a feed-forward compensation amount into an outer ring gas turbine rotation acceleration instruction control position of the gas turbine rotation acceleration cascade control structure; and a recursive least square algorithm is used to carry out adaptive adjustment on feed-forward compensation parameters of the two loops. The invention further discloses a rapid dynamic response control system for the turboshaft engine. Compared with the prior art, the technical scheme can effectively improve the response speed of the engine control system.

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 the technical field of turboshaft engines, and provides a fuel oil compensation method and device for starting ignition of a turboshaft engine and a storage medium, the method comprises the steps that whether ignition is successful or not in an ignition interval is detected, if ignition fails, corresponding fuel oil compensation is increased, and the ignition interval is that ignition is conducted within the preset range of the rotating speed of the engine; whether the turbine outlet temperature rise meets a first condition or not is detected, and if the first condition is met, fuel compensation is ended. The ignition success rate is improved by detecting whether ignition in the ignition interval succeeds or not and performing fuel compensation outside the designed ignition interval after ignition fails, so that the purpose of reducing the starting time is achieved, and in order to further improve the ignition success rate, whether the turbine outlet temperature rise meets the first condition or not is continuously confirmed, so that the starting time is shortened. And the ignition success rate is further improved, and good adaptability is achieved.

The invention discloses a method, a device and equipment for monitoring the degradation state of a turboshaft engine and a medium, and aims at overcoming the defects of an existing method for sensing the degradation state of the turboshaft engine based on limited measurement parameters. A degeneration factor range linear solving method based on a degeneration factor fingerprint map theory is introduced on the basis of a traditional particle swarm optimization algorithm, and a more accurate and reasonable particle optimization limitation range is provided for the particle swarm optimization algorithm under the condition that available system input parameters are insufficient. For any sensor actual acquisition main performance parameter degradation situation combination, a corresponding specific optimization range can be provided, and the situation that degradation factor optimization falls into local optimization in a plurality of possible optimization spaces due to insufficient data complexity is avoided. Therefore, degradation state monitoring of the turboshaft engine is optimized, and the overall performance degradation trend of the engine is accurately obtained.

The invention discloses a turboshaft engine external pipeline vibration assessment method and system, equipment and a medium, and the method comprises the steps: firstly carrying out finite element calculation analysis on an external pipeline, correcting the connection rigidity through a knock frequency test, considering the influence of the connection rigidity, correcting a finite element calculation model, and carrying out the vibration assessment of the external pipeline; according to the method, the calculation resonance frequency and the knocking resonance frequency are kept consistent, so that accurate pipeline vibration characteristics can be obtained, the accuracy of pipeline risk assessment can be improved, a finite element calculation model is verified and corrected only by performing a simple test comparison frequency result, and a complex modal test does not need to be performed. And then, by adopting a mode of combining finite element calculation and a whole machine maximum unbalance test, the vibration condition of the external pipeline can be simulated more accurately based on an accurate finite element calculation result and combining a whole machine installation condition, and the accuracy of pipeline risk assessment can be further 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.

The invention provides a generator and battery power distribution method and device based on model prediction control, and belongs to the technical field of power distribution. The method provided by the invention comprises the following steps: loading a preset dynamic model library; generating and calibrating a total load power prediction sequence based on the flight plan and the rotor aerodynamic power model; calling the dynamic model library to predict and control the rolling optimization to output the optimal power instruction sequence of the generator; in the rolling optimization execution process, bus voltage dynamic characteristics are monitored in real time, a graded response mechanism is started based on a turboshaft engine dynamic model when load sudden change occurs, and the power of a generator and the power of a battery are adjusted; in the load abrupt change response process, rolling optimization is executed again based on the adjusted power and the corresponding system state, meanwhile, feed-forward compensation is carried out on the bus voltage deviation, and the voltage is controlled to be within a stable interval set by rolling optimization by adjusting the power of the generator; and circularly executing until the flight task is finished, and determining the generator power and the battery power.

The invention provides a fuel-based, single-engine multirotor drone, wherein the drone comprises a body, a turboshaft engine to generate rotational movement, a plurality of rotors, a plurality of pulleys, a plurality of transmission belts, and a gearbox assembly. The gearbox assembly comprises a pair of coaxial counter-rotating gears which are coupled to a pair of coaxial counter-rotating shafts, wherein the rotational movement of the engine is transferred synchronously from the counter-rotating shafts of the gearbox to the plurality of the rotors through the plurality of pulleys and transmission belts.

The invention relates to a hybrid vertical take-off and landing aircraft energy management method based on a model predictive control algorithm, which belongs to the technical field of aircraft energy management, and comprises the following steps: establishing a size model and a performance model of a multi-rotor configuration based on a top demand, a hybrid architecture and a propulsion configuration by using an iterative scheme of size design optimization; forming constraint of size iteration, and optimizing required overall size parameters of the aircraft through an fmincon function; a component-level turboshaft engine model and a battery internal resistance model are provided, a theoretical model is established based on gas-heat dynamic characteristics of each component of the turboshaft engine, the battery internal resistance model is established based on battery testing, and a battery peak current is obtained; and an energy management strategy based on an MPC algorithm, DP optimization information, supplementary CS and a charge strategy.

The invention relates to the technical field of aircraft trial, and provides a universal test flight platform of a turboshaft engine, which comprises an aircraft, a first test cabin, a first power pack, an independent reduction gearbox and a propeller, two wings are respectively arranged on two sides of the aircraft, the first test cabin is arranged on one wing, the first test cabin and the first power pack are positioned on the same side of the aircraft, and the independent reduction gearbox is arranged on the first test cabin. The first power pack is arranged at the joint of a wing and the aircraft, the first test chamber is provided with a containing cavity, a mounting part is arranged in the containing cavity and used for mounting an engine and enabling the axial direction of the engine to be consistent with that of the containing cavity, the engine is connected with the first power pack through the first test chamber, and the independent reduction gearbox is arranged in the containing cavity. The propeller is arranged outside the containing cavity and located at the end of the first test chamber, and the engine is in transmission connection with the independent reduction gearbox and the propeller in sequence. The turboshaft engine is well matched with the test flight platform, test flight verification of the whole flight envelope of the turboshaft engine is met, and the test flight platform has the characteristic of universality.

The invention discloses an engine accessory on-condition maintenance decision-making method considering inherent characteristics and repair characteristics, and belongs to the field of turbine engine maintenance, and the method comprises the following steps: classifying turboshaft engine accessories based on the inherent characteristics to obtain classified accessories, the classified accessories comprise a repairable accessory and an unrepairable accessory; performing fault damage analysis on the classified accessories to obtain an external field maintenance decision; determining factory return maintenance accessories based on the external field maintenance decision; and analyzing the factory return maintenance accessory based on the repair characteristics to determine a maintenance decision of the accessory in the overhaul. The method can be used for guiding the maintenance decision of accessories of the turboshaft engine in external field use and factory return overhaul stages, scientifically determining the maintenance mode and the disassembly depth, improving the reliability and maintainability of the engine, reducing the maintenance cost and improving the guarantee capability of the engine.

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 discloses a methanol / aviation kerosene dual-fuel chemical regenerative hybrid aero-engine, relates to the technical field of environment-friendly processing, and solves the problems that the electricity demand of turboshaft engine airborne equipment is increased and the carbon emission of the engine is too high. A core engine system comprises an aviation fuel storage tank, an air inlet channel, an air compressor, a combustion chamber, a turbine system, an exhaust nozzle and a rotor wing, wherein the aviation fuel storage tank is communicated with the combustion chamber; the air inlet channel, the air compressor and the combustion chamber communicate in sequence. One part of the methanol reforming product is introduced into a combustion chamber to be mixed with aviation fuel for combustion, and the other part of the methanol reforming product generates power through a fuel cell; the combustion chamber, the turbine system and the exhaust nozzle are communicated, and fuel gas generated by the combustion chamber is introduced into the turbine system to drive the turbine system; the turbine system drives the gas compressor and the rotor wing. The fuel cell is used for generating electricity to meet the electricity demand of airborne equipment, the shaft power on the engine main shaft is reduced, the engine output power is effectively improved, and the operation stability of a core engine system is improved.

The invention discloses a fault diagnosis system for an aviation turboshaft engine, and relates to the technical field of fault diagnosis. An acquisition module acquires horizontal, vertical and axial vibration signals; the processing module is used for generating real-time feature vectors by introducing dynamic short-time Fourier transform, variational mode decomposition and singular value decomposition of an unfixed-width Gaussian window; the incremental diagnosis module calculates the matching degree between a real-time feature vector or a prediction feature vector and a known class through a dynamic expansion network, selects to belong to the known class or triggers an expansion mechanism according to an adaptive resonance theory, constructs joint loss by using a Bayesian learning framework after expansion, and performs Gaussian process modeling; loss is determined through Bayesian optimization, and network parameters are updated; the prediction module generates a prediction feature vector when there is no fault; the maintenance module sends or stores a maintenance scheme according to needs, self-adaptive recognition of known classes and new fault classes is achieved, the state can be pre-judged in advance, the operation safety of the engine is effectively guaranteed, and the maintenance cost is reduced.

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 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 heating turbomachine for a fuel-conditioning system, which is configured to supply an aircraft turboshaft engine with fuel from a cryogenic tank. The heating turbomachine comprising: compressor; a turbine; a combustion chamber which is configured to discharge a flow of exhaust air loaded with heat energy and to rotate the turbine; at least one fluid circuit in which a heat-transfer fluid circulates in the downstream direction; and at least one post-combustion heat exchanger which is mounted in the fluid circuit and is configured to extract heat energy from the exhaust air chamber flow between the combustion and the turbine.

The invention relates to the technical field of aero-engine tests, and discloses a heat insulation cabin device and a temperature adjusting method thereof.The heat insulation cabin device comprises a heat insulation cabin body which comprises an upper cabin cover and a lower cabin body, and the upper cabin cover and the lower cabin body jointly define a heat insulation cabin; the lower cabin body is provided with a first wall body and a second wall body, and the first wall body and the second wall body are spaced to form an air flow channel; a runner inlet and a runner outlet are formed in the lower cabin body, one end of the runner inlet is communicated with the air runner, and the other end of the runner inlet is communicated with the heat insulation cabin; one end of the runner outlet is communicated with the air runner, and the other end is communicated with the air blower; one end of the cold air compensation interface is communicated with the heat insulation cabin, and the other end of the cold air compensation interface is communicated with a cold air source. The heat insulation cabin device provided by the invention can ensure the normal work of the power measurement system in the high and low temperature environment test of the turboshaft engine.

The utility model discloses a sleeve repair tool for a fuel manifold of a turboshaft engine, which comprises a support with the top end provided with a penetrating column with threads on the inner wall, and further comprises a threaded pressing rod with the top end provided with a pulling rod, and the threaded pressing rod is in threaded connection with the penetrating column through the threads on the outer wall; a semi-arc-shaped pressing plate is arranged at the bottom of the T-shaped sliding block, and the T-shaped sliding block is movably connected with the threaded pressing rod; a shape correcting arc plate is arranged on the positioning column opposite to the semi-arc pressing plate, and the positioning column is located at the middle section of the support; and the gland is connected to the positioning column through a bolt. According to the sleeve repairing tool for the turboshaft engine fuel main pipe, the problems that in the prior art, manual positioning is not accurate, and pipe wall cracks and internal threads are prone to being crushed to be scrapped due to repeated knocking are solved.

Disclosed in the present invention are a turboshaft engine, an aircraft, and a method, used for solving the problem of a traditional turboshaft engine failing to simultaneously meet reasonable matching of high-altitude takeoff power and sea-level takeoff power. The turboshaft engine comprises air inlet pipes, a driving mechanism, an impeller, a rotating shaft, an air compressor, an air compressor rotating shaft, a rotor braking mechanism, a clutch, combustion chambers, a turbine assembly, and a tail spray pipe; the driving mechanism is arranged on an output end of the rotating shaft; the impeller is electrically connected to the driving mechanism; the impeller and the rotor braking mechanism are arranged on the output end of the rotating shaft; the air compressor is arranged on the air compressor rotating shaft; the clutch is sleeved on the rotating shaft, and is arranged between the rotating shaft and the air compressor rotating shaft; the turbine assembly is arranged on a driving end of the rotating shaft. The turboshaft engine, the aircraft and the method provided by the present invention are used for matching the takeoff power of the aircraft under different working conditions.

An aviation turbine shaft engine over-rotation protection execution mechanism detection device comprises an oil way system and a control circuit, and the oil way system comprises an oil tank, a direct-current electric pump and a two-position three-way electromagnetic valve; a normally open loop of the electric pump and the electromagnetic valve and the oil tank form a circulating oil path through an oil pipe; a normally closed loop of the electromagnetic valve is connected with an oil inlet of the aviation turboshaft engine over-rotation protection execution mechanism test piece, and an oil outlet of the over-rotation protection execution mechanism test piece is connected with the oil tank through an oil pipe; the control circuit comprises a power switch, a detection switch and a two-position three-way electromagnetic valve; the power switch controls power supply of the direct-current electric pump, and the detection switch controls power supply of a control coil of the two-position three-way electromagnetic valve and a test piece of the aviation turbine shaft engine over-rotation protection executing mechanism. According to the utility model, an integrated box type design is adopted, the size is small, the structure is simple, an onboard direct current 27V power supply is used for supplying power, the external field use requirement of a helicopter can be met, and the working condition of an aviation turbine shaft engine over-rotation protection executing mechanism can be conveniently detected.

The invention relates to an adaptive starting method for a turboshaft engine (10) comprising a gas generator (11) to which an electric machine (21) is mechanically connected and which is electrically connected to one or more electrical energy sources (26-28). According to at least one characteristic of each electrical energy source (26-28), a maximum electrical strength of the current that can be delivered by the one or more electrical energy sources (26-28) is determined, and a maximum engine torque supplied by the electric machine (21) powered by each electrical energy source (26-28) at said maximum electrical strength is determined. The maximum engine torque is compared with a specified engine torque envelope (70) permitted by the gas generator (11). The maximum electrical strength of the current supplied by the one or more electrical energy sources (26-28) is adjusted to comply with the prescribed envelope (70) in order to start the gas generator (11).

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.

An adaptive starting method for a turboshaft engine comprising a gas generator, an electric machine connected mechanically to the gas generator and electrically to one or more electrical energy sources. A maximum electrical intensity of an electric current that can be delivered by the electrical energy hgsource or sources is determined as a function of at least one characteristic of each electrical energy source, a maximum engine torque supplied by the electric machine powered by each electrical energy source with the maximum electrical intensity is determined. The maximum engine torque is compared with a required engine torque envelope permitted by the gas generator. The maximum electrical intensity of the electric current supplied by the electrical energy source or sources is adapted in order to comply with the required envelope in order to start the gas generator.

The turboshaft engine comprises a speed reducer, a gas compressor, a combustor and an exhaust nozzle, the speed reducer and the combustor are both fixedly connected with the gas compressor, the input end of the combustor is connected with the output end of the gas compressor, and the exhaust nozzle is fixedly connected with the output end of the combustor. The speed reducer comprises a case, a transmission assembly and an auxiliary assembly are arranged on the inner wall, the transmission assembly is connected with the combustor, and the auxiliary assembly is connected with the gas compressor. And the output end of the lubricating assembly is communicated with the interior of the case. The turboshaft engine has the core advantages that a transmission assembly and an auxiliary assembly are integrated in the same case, the overall size is greatly reduced, and the compact installation requirement of the turboshaft engine is met; the case provides reliable protection for internal parts, and interference of the external environment on the transmission precision is reduced. By optimizing a power transmission path, the transmission assembly specially receives power transmission of the combustor, the auxiliary assembly concentrates on driving the gas compressor and related auxiliary parts, power distribution precision is achieved, and the power transmission stability of the whole machine is remarkably improved.

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 present invention provides a method, device, equipment and medium for controlling the NG speed in a turboshaft engine test. The method includes: collecting the NP speed, NG speed and intake air temperature of the engine; inputting the target value of the NG speed and the NP speed into a speed model to obtain the target torque value and throttle lever angle value, inputting the target torque value into a dynamometer to keep the output shaft torque of the engine constant at the target torque, and inputting the throttle lever angle value into an electronic controller to adjust the NG speed; when the difference between the actual value and the target value of the NG speed is within a preset range, adjusting the actual value of the NG speed, obtaining a speed output value when the adjustment is stable, comparing it with the expected difference and inputting it into a self-tuning PID, using the intake air temperature as a feedforward to obtain a self-tuning output value to correct the speed model until the NG speed converges to the target value. Considering the temperature and the performance changes accumulated in long-term tests in the NG speed control improves the adaptive ability and the dynamic response performance of long-term stability.