349 results about "Turboshaft" patented technology

Disclosed is a method for structural damage monitoring for whole-machine life test of small turboshaft engine, comprising: determining a failure mode based on load borne by typical components of the small turboshaft engine in a service process and common failure forms, screening a structural damage index of a non-major component based on a coupling relationship between structural damages, selecting a destructive damage of components in the small turboshaft engine and an indirect damage causing changes in material properties of key component, planning a structural damage monitoring means based on the failure mode, the structural damage index, the destructive damage, and the indirect damage, designing an accelerated mission test platform of the small turboshaft engine, and designing a torque automatic loading scheme of an eddy current dynamometer, and testing the small turboshaft engine, and performing damage monitoring of the small turboshaft engine based on the planned structural damage monitoring means.

The invention relates to the technical field of aviation turboshaft engines, and provides a method for evaluating the influence of sand dust erosive wear on the performance of a turboshaft engine, which comprises the following steps: calculating the amount of sand dust sucked into the turboshaft engine; calculating and solving an air flow field in the air compressor channel and a movement track of sand dust particles by adopting computational fluid mechanics software; establishing a mathematical model of erosive wear of the air compressor blade of the turboshaft engine caused by sand dust; on the basis of the motion trail and the mathematical model, the blade profile of the blade after abrasion is obtained, the airflow field in the air compressor channel after abrasion is further calculated and solved, and the reduction amount of the efficiency of the air compressor after sand dust erosion abrasion is obtained; and on the basis of the reduction amount of the compressor efficiency, engine overall performance calculation software is adopted to calculate and evaluate the engine performance parameter change condition. According to the method, a mathematical model of compressor blade abrasion is constructed, the turbine shaft engine performance degradation evaluation method is provided, and quantitative analysis and evaluation of turbine shaft engine performance degradation in the sand and dust environment are achieved.

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 an integrated case of a turboshaft engine. The integrated case comprises an integrally-formed case body and a lubricating oil tank connected with the case body. And an air inlet flow channel is arranged in the casing body. A lubricating oil loop and an anti-icing gas flow path are further integrated in the casing body, the oil inlet end of the lubricating oil loop extends into the lubricating oil tank so as to be used for leading lubricating oil contained in the lubricating oil tank into the accessory transmission casing outwards, and the oil outlet end of the lubricating oil loop communicates with the lubricating oil tank so as to be used for enabling the lubricating oil of the accessory transmission casing to flow back to the lubricating oil tank. The air inlet end of the anti-icing air flow path extends into the lubricating oil tank and penetrates through the lubricating oil tank so as to be used for being connected with an external pipeline for supplying anti-icing air, so that hot anti-icing air enters the anti-icing air flow path, and then the inner ring face and the outer ring face of the air inlet flow path are subjected to heating anti-icing. According to the integrated case, the integration degree is high, additional parts do not need to be added for preventing ice and providing lubricating oil, the number of the parts is small, the structure is relatively simple, therefore, the number of connectors is small, air leakage and oil leakage are not likely to happen, and the reliability of an engine is improved.

The invention discloses a rapid switching emergency protection system for an aero-engine test, and the system comprises a rapid switching unit which is used for rapidly disconnecting a DC load according to a switching instruction transmitted by a controller, and switching the DC load to an emergency protection load; the measuring unit is used for measuring signals such as engine gas turbine rotating speed, free turbine rotating speed, alternating-current voltage and current, direct-current voltage and current and the like; the controller is used for carrying out fault judgment on the system according to the measurement data acquired by the measurement unit and correspondingly sending a switching instruction to the rapid switching unit when a fault occurs; the monitoring system is used for monitoring and displaying parameters and operation states of the rectifier, the bus and the direct-current load system; and the emergency protection load is in circuit connection with the rapid switching unit and is used for providing the emergency protection load when the hybrid power system is in a large state and the load is suddenly lost. According to the invention, overspeed protection of a turboshaft engine hybrid power system test is realized, so that accidents such as breakage and flying-out of turbine blades are avoided.

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 provides a turboshaft double-rotor inter-shaft rub-impact identification method and device based on VMD and a double attention mechanism TCN, and relates to the field of aviation turboshaft engine fault diagnosis. The method comprises the following steps: S1, carrying out vibration data acquisition on the turboshaft birotor test device through multiple sensors; s2, carrying out modal decomposition on the vibration data by using a TTAO optimized VMD algorithm; and S3, signal reconstruction is carried out by using a screening criterion of correlation kurtosis and 1.5-dimensional spectrum combination. S4, performing sliding window processing on the reconstructed signal, forming a multi-dimensional feature sequence diagram, and dividing a training set and a test set; and S5, constructing a TCN model combining a channel and a time attention mechanism, extracting short-term local information in a time sequence through convolution operation, and obtaining a channel weight and a global dependency relationship by using the attention mechanism. According to the method, the modal component containing the impact characteristic and the inter-axis rub-impact characteristic can be effectively extracted, identification of key fault information is enhanced, interference and noise are weakened, and efficient diagnosis of inter-axis rub-impact is achieved.

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 hybrid turboshaft engine optimization method and system, a hybrid turboshaft engine is installed on flight equipment, the optimization method comprises the step of introducing a motor to assist in providing takeoff and climbing power of the flight equipment, and the motor is selected by the following steps: S1, determining a P turboshaft: determining the P turboshaft by taking W fuel oil minimization as a target, meanwhile, W fuel oil further needs to meet the inequality that W turboshaft + W motor + W clutch + W accessory + W battery + W fuel oil is smaller than or equal to K, the W turboshaft, the W motor, the W battery and the W fuel oil are all converted into a relational expression related to a P turboshaft, and the P turboshaft is obtained; s2, determining a P motor: P motor = P takeoff-P turboshaft; according to the optimization method, the motor is selected, the requirement of the hybrid turboshaft engine is met, extra thrust is increased, high-power power output is provided for flight equipment in the take-off and climbing stages, fuel combustion operation of the hybrid turboshaft engine is conducted under a stable working condition, and the total fuel consumption of the hybrid turboshaft engine is reduced.

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 discloses a variable mixing ratio design method for an aviation hybrid electric aircraft and a propulsion system, and the method comprises the steps: S1, dividing an aircraft mission profile into stages of taking off, climbing and the like, and building a linkage integrated design model of the hybrid electric aircraft and the propulsion system; s2, determining parameters of the aircraft and the propulsion system under different application scene requirements by combining calculation formulas in the linkage integrated design model of the hybrid electric aircraft and the propulsion system; and S3, by changing the mixing ratio of the takeoff stage and the climbing stage in different application scenes, optimizing and determining the corresponding mixing ratio when the maximum takeoff weight and the unit commercial load per kilometer power cost are at the lowest level in different application scenes, realizing the optimal design of the aviation hybrid electric aircraft and the propulsion system, and providing a basis for aircraft power type selection. Hybrid electric system design can be carried out on the basis of an existing turboshaft engine, optimal design of an aviation hybrid electric aircraft and a propulsion system is achieved, and a reliable basis is provided for power type selection.

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.

The invention relates to a synchronous electric machine (10) for an aircraft, comprising a stator (13) comprising two sets of stator coils (47, 49) intended to be connected to different power converters and a wound rotor (38) inserted in the stator, and comprising a rotor shaft (11) and two rotor coils (39, 40), each rotor coil is intended to be supplied by a different supply current. The two sets of stator coils are arranged in the stator such that when a first set of stator coils (47) fails, a second set of stator coils (49) cooperates with at least a second rotor coil (40) powered by an associated supply current to generate a mechanical torque on the rotor shaft, and such that the power converter connected to the first set of stator coils does not deliver any electrical power.

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 provides a turboshaft engine starting torque measuring tool system and method, and relates to the technical field of torque measuring tools.The turboshaft engine starting torque measuring tool system comprises a base, main supporting plates are evenly and fixedly connected to the top of the base, a casing supporting plate is fixedly connected to the top of the base, and a gas compressor shaft positioning block is arranged on one side of the casing supporting plate; an upper front cover is arranged on one side of the main supporting plate, a bearing seat is arranged on one side of the upper front cover, when the rotating device is used, a round rod is manually controlled to rotate, then a first belt wheel is driven to rotate through a first connecting belt, and due to the fact that a clamping block is well clamped on the surface of a nut, the first belt wheel can well drive the nut to rotate; therefore, the two screw rods can be simultaneously controlled to be screwed out or screwed in, so that the two screw rods can be simultaneously adjusted to be screwed out or screwed in the torque transmission supporting plate, data can be replaced for different degrees of measurement, and the measurement accuracy is ensured.

The invention relates to the technical field of aviation turbine shaft engines, and provides an aviation turbine shaft engine design load spectrum compilation reference object analysis method, which comprises the following steps of: counting high-power states of engines corresponding to air typical mission profiles of various helicopters and high-power state use time of the engines according to information of the typical mission profiles; calculating a high-power state cumulative time ratio of an engine in a typical mission profile related to each type of helicopter; based on the high-power state cumulative time proportion of the engine, calculating the high-power state reduced time proportion of the engine in all typical mission profiles of various helicopters; and selecting a helicopter as a load spectrum compiling reference object according to the high-power state-to-time ratio of the engine, and compiling a load spectrum based on the helicopter. According to the method, through effective and simple engine power state comparative analysis, a proper helicopter platform is selected as a design object, and the obtained design load spectrum can meet the use requirements of all helicopter platforms.

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.