34 results about "Inlet distortion" patented technology

The invention relates to the technical field of flutter, and particularly provides a flutter stability prediction method and device, equipment, a storage medium and a program product. The flutter stability prediction method comprises the following steps: dividing a flow field grid and a solid grid based on a gas compressor geometric model; on the basis of the solid grid, finite element modal analysis is executed to obtain intrinsic mode data, and the intrinsic mode data comprises inherent frequency and intrinsic mode distribution thereof; based on the flow field grid, the intrinsic mode data and preset circumferential air inlet distortion parameters, executing unsteady numerical simulation so as to determine first aerodynamic work under single-node diameter vibration; based on the phase angle between the blades and the first aerodynamic work, second aerodynamic work under all node diameter vibration is determined, and aerodynamic damping under all node diameter vibration is determined based on the second aerodynamic work; based on the aerodynamic damping, flutter stability is predicted. According to the technical scheme, the prediction speed of the flutter stability can be increased.

The invention relates to a blade flutter evaluation method based on an approximate time domain harmonic balance method under air inlet distortion, which comprises the following steps of: calculating a vibration mode of an aero-engine blade, and respectively appointing a blade vibration rule and an air inlet distortion boundary condition so as to determine and analyze an unsteady flow field fundamental frequency and a harmonic wave caused by blade vibration and air inlet distortion; and then, a coupling flow field of blade vibration and air inlet distortion is solved based on an approximate time domain nonlinear harmonic method, then accumulated work in a vibration period is calculated, and the flutter stability of the aero-engine blade is evaluated according to positive and negative accumulated work. According to the method, the influence of multiple parameters and any form of air inlet distortion on the flutter stability of the aero-engine blade can be evaluated at the same time; the evaluation result is calculated based on the numerical value of the single channel, the calculation time consumption and the calculation resource demand are remarkably reduced, and the method has important significance on the anti-vibration design of the rotor and stator blades of the aero-engine.

The utility model discloses a high-speed large-maneuvering target drone, which relates to the field of unmanned aerial vehicles and comprises a drone body, wings, a horizontal tail, a vertical tail and an air inlet channel. The fuselage can reduce resistance to a great extent to realize high-speed flight; the wings adopt sweepback upper single wings to reduce the resistance during high-speed flight, the problem of overlarge transverse stability caused by sweepback and high placement of the wings can be avoided through the downward reverse design, and meanwhile, the wings can be integrally disassembled and assembled and have better operability and convenience; the control efficiency of a full-motion horizontal tail is high, and the heading stability under a large-angle-of-attack attitude can be improved through the design of the inverted-V-shaped tail; the vertical fin without the control surface can improve the course stability and ensure the matching property of the transverse course stability, and is light in structural weight and small in flight resistance; the bearing separation channel type air inlet channel can avoid landing damage of the air inlet channel in the parachute landing recovery process, and meanwhile the requirements for high total pressure recovery, low air inlet distortion, radar stealth and the like are met.

The invention belongs to the field of test fluid mechanics, and discloses a flow correction method considering air inlet distortion and a calibration device. The flow calibration device comprises a standard flow spray pipe, an air inlet section, a measurement section, a rectification section and an exhaust spray pipe which are sequentially connected from front to back. The flow correction method considering the air inlet distortion comprises the steps that distortion air inlet flow correction is conducted; flow and distortion intensity evaluation is carried out; training a neural network; carrying out an air inlet channel test; calculating an air inlet flow distortion parameter; and correcting a flow calculation result. According to the flow correction method and the calibration device considering the air inlet distortion, the air inlet distortion, the total pressure loss and the Reynolds number influence are comprehensively considered, technical support can be provided for high-precision flow monitoring of a large-curvature aircraft air inlet performance evaluation test, and the method and the calibration device have high engineering practical value.

The invention belongs to the field of turbine engines, and provides a thermodynamic cycle parameter obtaining method and system for a low-cost configuration turbine engine, and the method comprises the steps: determining an engine configuration through a cost design target of the turbine engine; according to the engine configuration, a thermodynamic cycle parameter feasible region is obtained with overall performance indexes and a fan inlet size threshold value as constraints, and the overall performance indexes comprise a thrust index and an oil consumption rate index; thermodynamic cycle parameters are determined in the feasible region of the thermodynamic cycle parameters through turbine engine performance requirements and turbine engine air inlet distortion resistance. According to the method, the thermodynamic cycle parameters can be rapidly obtained without repeated iteration, and the design period is greatly shortened.

The invention discloses an aero-engine air inlet distortion test measuring device which is applied to be connected between an inlet of an aero-engine and a distortion generator, the test measuring device comprises a switching part and an air inlet flow guide basin, the switching part is used for being connected to the inlet of the aero-engine, a mounting cavity is formed in the switching part in a penetrating mode in the first direction, and the air inlet flow guide basin is arranged in the mounting cavity. A plurality of branch pipes are arranged in the mounting cavity at intervals in the second direction, and the branch pipes are parallel to the third direction; a plurality of static pressure measuring points are arranged on the windward side of each branch pipe at intervals in the third direction, and a total pressure measuring point is arranged on one side, in the third direction, of each static pressure measuring point; a plurality of total temperature measuring points are arranged on at least two branch pipes which are relatively close to the outside in the second direction; according to the device and the method, quick response is achieved when the change of an air inlet flow field of an engine reaches a distortion point, and static pressure parameters and total pressure parameters of all positions of the section, perpendicular to the first direction, of the mounting cavity are accurately obtained.

The invention provides a ground vortex intensity prediction method under a crosswind condition, and the method comprises the following steps: S1, judging whether a ground vortex is generated or not according to a ground vortex intensity prediction model under the crosswind condition, and calculating a dimensionless circulation under a corresponding working condition; s2, according to an air inlet distortion index prediction model, calculating the size of an air inlet total pressure distortion index caused by suction of a ground vortex; and S3, according to a CFD calculation result, determining the model output parameter precision and aerodynamic characteristics reflected by the model output parameter precision so as to verify the reliability of the model. The invention provides a reliable ground vortex intensity prediction model for 90-degree crosswind conditions. The ground vortex intensity prediction model not only can qualitatively judge whether a ground vortex is generated, but also can quantitatively calculate ground vortex intensity and air inlet total pressure distortion.

The invention discloses a rapid and high-precision adjustable conical inserting plate type air inlet distortion device which comprises a control module, a distortion generator body, an air inlet flange, an air inlet pipeline, an air outlet pipeline and an air outlet flange. A control module is mounted on the top surface of the distortion generator main body; one end of the air inlet pipeline is connected with the front surface of the distortion generator main body; one end of the air outlet pipeline is connected with the back surface of the distortion generator main body; an upper blanking cap is installed on the top face of the valve body, a lower blanking cap is installed on the bottom face of the valve body, and a movable inserting plate is installed in the valve body. The lower end of the bracket is connected with the top surface of the upper blanking cap; the upper end of the lead screw is connected with the output end of the servo motor, and the middle part of the lead screw is provided with a lead screw nut; and the connecting sleeve sleeves the outer side of the lead screw. According to the invention, a traditional hydraulic mechanism can be replaced, rapid movement and accurate positioning of the plugboard can be realized, and the reliability of the measured distortion index can be improved.

The invention provides an engine air inlet distortion suppression device and method, the engine air inlet distortion suppression device is arranged on an engine air inlet pipeline, the air inlet pipeline extends in the axial direction of the engine air inlet pipeline, the air inlet pipeline is provided with a transverse section perpendicular to the axial direction of the engine air inlet pipeline, and the engine air inlet distortion suppression device is characterized in that the suppression device is arranged on the transverse section and comprises a flow expander, circumferential blades and radial blades; the center of the flow expander coincides with the center of the transverse section, the projection area of the flow expander on the transverse section is adjustable in the radial direction of the transverse section, and the radial direction vertically extends from the center of the transverse section to the peripheral side of the transverse section; the circumferential blades are distributed around the flow expander and can rotate around the circumferential direction, so that the projection area of the circumferential blades on the transverse section is adjustable in the radial direction, and the circumferential direction is the direction around the axial direction on the transverse section; the radial blades are distributed in the radial direction of the transverse section and can rotate in the radial direction, so that the projection area of the radial blades on the transverse section is adjustable.

The invention provides a multi-mode pressure distortion generating system capable of being automatically adjusted, and belongs to the field of aero-engine tests.The multi-mode pressure distortion generating system comprises a distortion generating device, a hydraulic system and a control system.The distortion generating device comprises a sealing box body, a distortion inserting plate and a driving device; pipeline connecting flanges used for being connected with an air inlet channel of an air compressor test piece are arranged on the front face and the rear face of the sealing box body, the distortion inserting plate is arranged in the sealing box body, and the driving device is connected with the distortion inserting plate; the control system is connected with the distortion generating device and the hydraulic system, the hydraulic system is connected with the driving device, and the distortion inserting plate is controlled by the control system to move in the sealing box body, so that the air inlet channel is blocked, and the change of the blocking ratio of the inlet of the compressor test piece is realized. According to the invention, a 0-100% blockage ratio state can be covered, so that different air inlet distortion conditions are formed, and a function of quickly opening the distortion generation device under special conditions such as surge and the like is achieved.

The invention relates to a blade flutter evaluation method based on a complete time domain harmonic balance method under air inlet distortion, and the method comprises the steps: calculating the vibration mode of an aero-engine blade, respectively appointing a blade vibration rule and an air inlet distortion boundary condition, thereby determining and analyzing the fundamental frequency and harmonic waves of a blade vibration flow field and an air inlet distortion flow field, and further calculating the blade flutter evaluation result. A coupling flow field of air inlet distortion and blade vibration is solved based on a complete time domain harmonic balance method, accumulated work in a vibration period is obtained, and the aeroelastic stability of the aero-engine blade is evaluated according to positive and negative accumulated work. According to the method, the influence of multiple parameters and any form of air inlet distortion on the aeroelastic stability of the aero-engine blade can be evaluated at the same time; and the evaluation result is calculated based on single-channel and high-fidelity numerical values, so that the calculation time consumption and the calculation resource demand are remarkably reduced, and the method has important significance on the anti-vibration design of the rotor and stator blades of the aero-engine.

The application provides a wing-body fusion back support engine reverse thrust device capable of inhibiting air intake distortion, which comprises a ring-shaped reverse thrust vane located on the inner side of a back support engine nacelle cover and reverse thrust flow guiding plates located on the two side surfaces of a back support engine nacelle support. When reverse thrust is needed, with the increase of the unfolding angle of the flow guiding plates, the hindering effect of the flow guiding plates on the forward movement of the reverse thrust air flow is more and more significant, so that the air intake quality of the turbofan engine during the reverse thrust work is ensured; in addition, the flow guiding plates also have the interference effect on the air flow similar to that of a spoiler, so that the form resistance is increased, which is beneficial to the faster deceleration of the airplane. The application effectively solves the adaptability problem of the reverse thrust device and the wing-body fusion back support engine, inhibits the occurrence of air intake distortion, strengthens the braking effect of the reverse thrust device, and guarantees the flight safety of the wing-body fusion back support engine layout.

The present invention relates to a blade flutter assessment method based on an approximate time-domain harmonic balance method under intake distortion, comprising calculating the vibration modes of an aircraft engine blade, and specifying the blade vibration law and intake distortion boundary conditions respectively, thereby determining the fundamental frequency and harmonics of the unsteady flow field caused by the analytical blade vibration and intake distortion; then, solving the coupled flow field of the blade vibration and intake distortion based on an approximate time-domain nonlinear harmonic method, and then calculating the accumulated work within a vibration cycle, and evaluating the flutter stability of the aircraft engine blade based on the positive or negative value of the accumulated work. The method of the present invention can support the simultaneous assessment of the effects of multiple parameters and any form of intake distortion on the flutter stability of aircraft engine blades; the assessment results are based on single-channel numerical calculations, which significantly reduces the computational time and computing resource requirements, and is of great significance to the anti-flutter design of aircraft engine rotor and stator blades.

The invention belongs to the technical field of engine nacelle production, and particularly relates to an automatic welding system for an inner supporting framework for engine nacelle production, which comprises a frame, a fixed clamping assembly, a movable clamping assembly, an outer seam welding assembly, a cover ring pushing assembly and an inner seam welding assembly. The frame is composed of a cylinder part and an end plate. The fixed clamping assembly is used for clamping the inner heat dissipation pipe section and driving the inner heat dissipation pipe section to rotate in the circumferential direction. The movable clamping assembly clamps the outer radiating pipe section and achieves axial displacement of the outer radiating pipe section. The outer seam welding assembly is responsible for welding reinforcement and outer seam welding of the spoiler and the middle supporting pipe section; the cover ring pushing assembly adsorbs the cover ring piece and axially pushes the cover ring piece; and the inner seam welding assembly clamps the middle supporting pipe section and performs axial displacement, and inner seam welding is completed. Automatic welding of the inner supporting framework is achieved, the welding efficiency is remarkably improved, meanwhile, the inner supporting framework has the air inlet fairing supporting function, the heating function and the low-speed air inlet distortion control function, and the problems that in the prior art, efficiency is low, and the function is single are solved.

The invention relates to the technical field of aero-engines, in particular to an intake-engine integrated rapid performance prediction method considering circumferential total pressure intake distortion, comprising: acquiring compressor performance diagram data under a uniform intake condition, and based on the compressor performance diagram data, acquiring slope data of compressor total pressure rise and an axial speed curve; acquiring initial velocity distribution data based on the inlet flow field distribution data; generating a fan inlet Mach number based on the initial speed distribution data; determining a volume force coefficient based on the fan inlet Mach number and the Mach number predicted by the pre-constructed fan inlet flow field prediction model; based on the volume force coefficient and the control equation, performance prediction data is generated. According to the method, the fan inlet flow field prediction model under the circumferential total pressure air inlet distortion is deduced and combined with the streamline volume force model, the relation between the Mach number and the volume force coefficient is reconstructed, and then the air inlet-engine integrated performance prediction method considering the circumferential total pressure air inlet distortion is constructed.

The invention relates to the technical field of air compressors, in particular to a forced response rapid analysis method under the condition of circumferential air inlet distortion, and the method comprises the following steps: S1, carrying out grid division on a flow field and a blade; s2, setting a first parameter, carrying out modal analysis on the research object to obtain an order inherent frequency and a modal distribution map, and determining an excitation order capable of causing forced vibration of the blade; wherein the modes of the research object comprise a first-order bending mode, a second-order bending mode and a first-order torsional mode; s3, carrying out full-cycle steady numerical calculation on the research object, and selecting a plurality of circumferential distortions with different distortion ranges for calculation; s4, key parameters in the flow field are extracted; and S5, calculating the excitation amplitude according to the key parameters. According to the method, the low-order excitation amplitude influencing the blade vibration can be quickly estimated through steady calculation.

The application belongs to the technical field of engine nacelle production, and particularly relates to an automatic welding system for an inner support framework for engine nacelle production, which comprises a frame, a fixed clamping assembly, a movable clamping assembly, an outer seam welding assembly, a cover ring pushing assembly and an inner seam welding assembly. The frame is composed of a cylinder part and an end plate. The fixed clamping assembly is used for clamping an inner heat dissipation pipe section and driving the same to rotate circumferentially. The movable clamping assembly clamps an outer heat dissipation pipe section and realizes axial displacement of the same. The outer seam welding assembly is responsible for welding and reinforcing of a spoiler and an intermediate support pipe section and outer seam welding. The cover ring pushing assembly adsorbs a cover ring and pushes the same axially. The inner seam welding assembly clamps the intermediate support pipe section and realizes axial displacement of the same, thereby completing inner seam welding. The application realizes automatic welding of the inner support framework, significantly improves welding efficiency, and enables the inner support framework to have the functions of air inlet rectifier cover support, heating and low-speed air inlet distortion control, thereby solving the problems of low efficiency and single function of the prior art.

A method for a gas turbine engine includes calculating an inlet airflow distortion value at a reference area within an airflow duct and in front of a fan of the gas turbine engine. The inlet airflow distortion value is indicative of an air pressure of an air vortex at the reference area caused by a set of crosswind conditions. The calculating is based on a ratio between an inlet airflow static air pressure in the reference area and a total ambient air pressure. The inlet airflow distortion value is compared to a threshold corresponding to a permissible amount of inlet airflow distortion at the reference area for the set of crosswind conditions, and, based on the comparison indicating that the inlet airflow distortion exceeds the threshold, an inlet airflow distortion notification is provided. A system for a gas turbine engine and a method for a gas turbine engine are also disclosed.

This invention belongs to the field of aero-engine performance simulation technology, specifically disclosing a data-driven method for predicting compressor flow field and performance. The method includes: acquiring operational data of the target equipment under different operating conditions using a volumetric force model and forming a snapshot matrix; decomposing the snapshot matrix and extracting the spatial mode matrix and modal coefficient matrix of the inlet and outlet flow fields, and constructing a training set using these matrices; inputting the training set into a BP neural network to continuously update the mapping relationship between the inlet and outlet flow field modal coefficient vectors; inputting the sampled data into the BP neural network to perform flow field reconstruction and predict the performance parameters of the target equipment. The method has the following advantages: the volumetric force model reduces data acquisition costs, and the POD method performs data dimensionality reduction and feature extraction, enabling accurate prediction of complex distorted flow fields and performance parameters even under small sample conditions; the data-driven flow field prediction method accurately captures high-dimensional flow field features, overcomes the problem of predicting nonlinear characteristics of the flow field under inlet distortion, improves prediction accuracy, reduces computational costs, and optimizes performance prediction efficiency.

The invention relates to the technical field of aero-engine control, in particular to an aero-engine direct thrust control method considering air inlet distortion, which comprises the following steps: constructing an aero-engine model with air inlet distortion; the real-time thrust of the air inlet distortion aero-engine model at each moment in the full-state test run process and related parameters related to the thrust are obtained; constructing a neural network model; and the output thrust of the aero-engine is obtained. According to the direct thrust control considering the air inlet distortion, the air inlet distortion of flight under the special working condition is considered, the control accuracy and the flight safety are improved, a traditional aero-engine control system achieves indirect thrust control by controlling parameters related to thrust, a large safety margin needs to be reserved, margin waste is reduced through direct thrust control, and the cost is reduced. The working potential of the aero-engine is fully explored, and meanwhile the engine efficiency is improved.

The invention discloses an oil-gas matching intelligent adjusting technology self-adaptive to inlet distortion, and the core idea is that under the condition of a distorted inlet, a plurality of temperature and pressure measuring points are arranged at the front edge of a support plate, the pressure and temperature parameters of incoming flow can be obtained, and inlet static pressure is combined, so that the inlet distortion is adjusted. The measured data are input into an algorithm program to obtain the radial speed and temperature change trend of the airflow, then the airflow flow of each stage is calculated according to the radial classification of the nozzle at the tail edge of the supporting plate, and the fuel oil flow of each stage is obtained according to oil-gas matching, namely that the oil-gas ratio of each stage is kept consistent with the total oil-gas ratio; and then the mass flow controller controls the nozzle to perform fuel injection. And fuel oil supply can be flexibly adjusted according to actual requirements through radial graded oil supply of the supporting plate, so that the fuel oil concentration distribution behind the supporting plate is more uniform, and the combustion efficiency and the combustion stability of the afterburner are improved.

The invention discloses an aero-engine air inlet distortion simulation device, and relates to the technical field of engine air inlet distortion simulation devices.The aero-engine air inlet distortion simulation device comprises an air duct and multiple sets of fan blades arranged in the air duct, a base ring is arranged in the air duct, the multiple sets of fan blades are located at the horizontal left side position of the base ring, and the base ring is provided with a speed pressing mechanism for adjusting the horizontal right side air speed of the base ring; the rapid pressing mechanism comprises a rotating assembly used for driving the base ring to freely rotate in the vertical direction, four groups of wind passing areas which are annularly symmetrical from the axis and are distributed in a divergent mode are arranged on the base ring, and wind adjusting assemblies used for adjusting the wind speeds of the right sides of the four groups of wind passing areas correspondingly are arranged on the base ring; according to the invention, by arranging the speed pressing mechanism, the purpose of adjusting and simulating the airflow speed at different positions of an engine air inlet channel can be achieved by independently controlling the rotating speed of a single set of fan blades and the independent adjusting process of the air outlet speed of four groups of air passing areas, and meanwhile, the response speed of the gust simulation process is further improved; the simulation experiment efficiency is improved.

The invention provides a high-degree-of-freedom distortion index simulation device, and belongs to the technical field of aero-engines, the high-degree-of-freedom distortion index simulation device comprises a total pressure distortion plate positioning switching assembly, a turbulent flow plugboard adjusting assembly and a fixed pipeline, the total pressure distortion plate positioning switching assembly comprises a movable motor base, a distortion plate socket, a distortion plate and a base, a gear on the movable motor base is connected with a rack below a distortion plate socket in a meshed mode, the distortion plate socket is connected to the base through a guide rail and a lead screw, a groove is formed in the distortion plate socket, and a distortion plate is connected into the groove in a sliding mode through a pin shaft. The turbulent flow inserting plate adjusting assembly comprises a fan-shaped inserting plate and sealing sliding blocks, the sealing sliding blocks are arranged at the two ends of the fan-shaped inserting plate, a T-shaped groove is formed in the fan-shaped inserting plate, a guide block is slidably connected into the T-shaped groove, and the guide block is fixed to an annular plate of a fixed pipeline. The high-degree-of-freedom distortion index simulation device is adopted, and the problems that a traditional air inlet distortion test is low in efficiency, high in cost, limited in evaluation accuracy and poor in engineering applicability are solved.

The invention provides an air inlet distortion device. The air inlet distortion device comprises a first pipeline and a second pipeline which are sequentially arranged in the axial direction of the distortion device. The first pipeline is sleeved with the first part of the flexible sealing sleeve, the second pipeline is sleeved with the second part of the flexible sealing sleeve, and the joint position between the first pipeline and the second pipeline is wrapped with the third part of the flexible sealing sleeve; at least one tightening piece is connected with the first part of the flexible sealing sleeve in a sleeving manner, so that the flexible sealing sleeve is tightened and fixed to the first pipeline; at least one tightening piece is connected with the second part of the flexible sealing sleeve in a sleeving mode so that the flexible sealing sleeve can be tightened and fixed to the second pipeline. And the flexible buffer gasket is arranged between the first pipeline and the second pipeline. In conclusion, through a flexible connection structure formed by the flexible sealing sleeve and the flexible buffer gasket, mechanical damage can be avoided while the air tightness of the distortion device is ensured.

The invention discloses an automatic adjusting device for a radial air inlet angle of an axial flow compressor, and belongs to the field of axial flow compressors. Comprising a large rotation module, a small rotation module, a telescopic adjustment module and a flow deflector rotation module. The large rotary module comprises a device base and an annular gear ring fixed on the device base; the small rotary module drives a circumferential angle positioning gear on the small rotary module to be meshed with the annular gear ring through a steering engine, so that the small rotary module can do circular motion along the large rotary module; the telescopic adjusting module is installed on the small rotation module and drives a telescopic plate to do radial reciprocating motion through gear-rack transmission so as to adjust the radial action range of the flow deflector. The flow deflector rotating module is installed at the tail end of the telescopic plate and drives the flow deflector to rotate through at least one air cylinder so as to change the flow area of the air inlet channel. According to the method, the core problem that local and dynamic air inlet distortion cannot be dealt with in the traditional technology is solved.

The invention provides an ice throwing device which comprises a support assembly and an ice throwing assembly, and the ice throwing assembly comprises an ice storage box arranged on the support assembly and provided with an ice storage cavity and an ice outlet communicated with the ice storage cavity; the ejection assembly is arranged in the ice storage cavity and is opposite to the ice outlet; the stop assembly is movably arranged in the ice storage cavity and located between the ejection assembly and the ice outlet; an ice storage area defined between the ejection assembly and the stop assembly is used for placing ice blocks; the stop assembly can be located at the stop position so as to stop ice blocks in the ice storage area. The ejection assembly is used for driving the ice blocks in the ice storage area to be ejected out through the ice outlet when the stop assembly leaves the stop position. Potential energy generated by compression of the ejection assembly is used for adjusting the initial speed when the ice blocks are ejected out of the ice outlet, then the flight path of the ice blocks is adjusted, the precision of the ice blocks projected into the engine is improved, the test uncertainty is reduced, and the risk of air inlet distortion of the engine is reduced.

The present application belongs to the field of engine control technology, and specifically relates to a method and device for determining the warm-up speed of a dual-shaft turbofan engine. The method includes step S1, determining the first maximum distortion index at each low-pressure conversion speed that can ensure the stable operation of the engine fan; step S2, determining the second maximum distortion index at each low-pressure conversion speed that can ensure the stable operation of the compressor; step S3, determining the maximum low-pressure conversion speed based on the inlet distortion index and the first and second maximum distortion indices; step S4, determining the maximum high-pressure conversion speed corresponding to the maximum low-pressure conversion speed; step S5, calculating the maximum high-pressure physical speed corresponding to the maximum high-pressure conversion speed; step S6, using the area between the warm-up minimum speed and the maximum high-pressure physical speed as the warm-up speed area. While ensuring the warm-up effect, the present application can ensure that the engine will not experience surge problems due to excessive warm-up state and excessive inlet air flow.

The application relates to a modeling method of an engine with inlet distortion based on parallel compressor theory, comprising an air path model of the engine, a matching method of the air path model and application. On the basis of an original double-shaft split-fan engine model, three calling models of a distortion splitter model, a parallel fan model and a mixing chamber model are added, and then a new air path model matching of the engine is carried out. The application realizes accurate control of an airplane under special working conditions. The application is based on a parallel compressor model, takes a double-shaft split-fan engine with a large bypass ratio as a research object, and establishes an aero-engine simulation model considering inlet distortion.

The present application belongs to the technical field of engine nacelle component production, in particular to an intelligent welding system and method for a combined air inlet fairing of an engine nacelle. The intelligent welding system is designed rationally. Inner and outer cutting assemblies are used to process inner and outer spoiler movable grooves on an air inlet passage casting blank carried by an air inlet passage positioning assembly, so that the air inlet passage casting blank is formed into an air inlet passage component. First, second, third and fourth clamping and feeding assemblies are used to feed an inner reinforcing member, a heating member, a noise reduction and sound absorption member preassembled with a spoiler assembly and a ring cover into a cavity of the air inlet passage component. A welding assembly is used to weld and fix the fed components in the cavity of the air inlet passage component. In this way, intelligent welding and assembly operation of the combined air inlet fairing can be realized. The combined air inlet fairing has the advantages of high structural strength, good noise reduction and sound absorption effect, automatic ice removal of the lip and control of air inlet distortion.