51 results about "Transonic" patented technology

The invention discloses an aerodynamic configuration structure of a precision guidance assembly and a design method, and belongs to the technical field of precision guidance, an assembly precursor adopts continuous three-section curve design: a supersonic speed section adopts a parabola profile, is provided with ribs on the surface, and is used for optimizing shock wave control under a supersonic speed working condition; the transonic speed section adopts a hyper-ellipsoid curved surface, ribs are arranged on the surface of the transonic speed section, and metamaterials are attached to the ribs and used for restraining transonic speed separation; the subsonic speed section adopts arc transition, and ribs are arranged on the surface and used for improving the lift force efficiency of the control surface; wherein the curvatures at the joints of the three sections of curves are continuous. According to the method, supersonic velocity-transonic velocity-subsonic velocity three-section collaborative design is realized, and meanwhile, a boundary layer is improved from a microscopic level by combining a collaborative drag reduction design of a microscopic rib and metamaterial technology, and the performance of a guidance assembly is optimized and improved.

The invention provides a wing-body fusion wide-speed-range aircraft layout design method based on consideration of an afterbody effect. The method comprises the following steps: firstly, generating a cone-guided waverider precursor of a wide-speed-range aircraft for supersonic flight; then, based on the basic airfoil profile, performing pneumatic optimization on the basic wing profile and the strake wing profile of the wide-speed-domain aircraft to obtain a multi-section wide-speed-domain waverider airfoil profile; then calculating the length of a waverider afterbody and designing a symmetry plane of the wide-speed-range aircraft; and finally, based on an osculating cone method and transonic speed strake wing constraints, obtaining three-dimensional layout parameters by using the waverider precursor parameters and the waverider afterbody length, and assembling the basic wing optimized airfoil profile, the strake wing optimized airfoil profile and the symmetry plane to generate the layout of the wide-speed-domain aircraft. According to the method, the cone-guided waverider generation method, the wide-speed-range airfoil optimization method and the osculating cone double-sweepback waverider generation method are coupled, and the aerodynamic design of the afterbody is considered, so that the aerodynamic performance of the wide-speed-range aircraft under the subsonic / transonic / supersonic speed is improved.

The invention provides an atmospheric data analysis method based on observability analysis, belongs to the field of aircraft navigation, guidance and control, and aims at solving the problem that the precision of traditional atmospheric data measurement is reduced under the condition of a complex flow field or a sensor fault. In the prior art, due to lack of strict theoretical support, estimation errors and convergence problems may occur in a transonic stage. According to the method, an atmospheric data state space model is established through flight dynamics, and atmospheric data estimation is carried out by adopting an extended Kalman filter based on the model; in the estimation process, observability analysis is executed to verify the reliability of an estimation result; and finally, optimizing the system model according to an analysis result. According to the method, the dependence of an existing method on an empirical model is reduced, the convergence and the stability of an estimation result are ensured, and the adaptability and the robustness of a system in a complex environment are remarkably improved.

The application discloses a high-pressure-ratio transonic axial flow compressor, which comprises a hub, front-row rotor blades, rear-row rotor blades and stator blades; the rear-row rotor blades are arranged between the front-row rotor blades and the stator blades; a plurality of front-row rotor blades and a plurality of rear-row rotor blades are uniformly arranged on the hub along a circumferential direction; and the front-row rotor blades and the rear-row rotor blades one-to-one correspond to form tandem rotors. The application realizes shock wave supercharging through the front-row rotor blades, and the rear-row rotor blades complete large airflow deflection; meanwhile, the application adopts an adsorption type stator to reduce or prevent airflow separation in a blade channel, thereby significantly improving a compressor stage pressure ratio and improving aerodynamic performance of the compressor. The application can realize a higher stage pressure ratio, and in the case of high inlet Mach number, high load and large airflow deflection angle, the application overcomes the problem that a high stage pressure ratio and efficiency are difficult to realize due to the fact that flow separation in the compressor cannot be controlled in the prior art.

This paper discloses an empirical formula blade generation system for designing subsonic and transonic turbine stages. The values of the design variables used in the parameterized system are calculated by the relevant equations. Iteration is required to modify the design variables. The parameterized system then generates the geometries and blocks of the stator and rotor. Using these geometries and blocks, a mesh is automatically formed using an ICEM CFD script. The turbine stage performance is then evaluated using a CFD solver. The calculation results are analyzed to identify future improvements. The final blade design is obtained through multiple rounds of repeated training and residual iteration. Finally, a highly efficient subsonic and transonic turbine stage is produced.

The invention discloses an asymmetric pulsating pressure measuring device and method suitable for a transonic flight fairing, and belongs to the technical field of information perception and recognition. The pressure sensitive circuit is used for collecting the external pulsating pressure of the fairing in the flight process, eliminating the internal pressure of the fairing through a through hole formed in the body, obtaining the pulsating pressure representing the internal and external pressure difference of the fairing, converting the pulsating pressure into an analog voltage signal and outputting the analog voltage signal to the signal conditioning circuit; the signal conditioning circuit receives the analog voltage signal, amplifies the analog voltage signal, and outputs the analog voltage signal to the filter circuit after zero setting; the filter circuit is used for filtering the signal and then outputting the signal; and the output interface circuit is used for conditioning the input signal to a voltage signal meeting a preset requirement and carrying out amplitude limiting so as to protect a rear-end acquisition circuit. The sensitive assembly based on the dry ceramic thick film capacitance pressure sensing principle is utilized, and the problems that in the prior art, pulsating pressure measurement is limited, and adaptability to industry and mining with static pressure changes is poor are solved.

The present invention relates to an improved sparrow search method based on penalty function constraint optimization. This method is an improved sparrow search method based on penalty function constraint optimization that integrates the sine-cosine algorithm and the light-sensing strategy of fireflies. By generating the initial population in combination with Cubic chaotic mapping in the initial population initialization stage, the accuracy and distribution of the initial population are improved; in the iterative process of updating the position of the followers, the sine-cosine algorithm is added to improve the position update strategy of the followers, making the search mechanism of the followers more flexible, expanding the global exploration range, increasing the population diversity, increasing the probability of finding the optimal solution, and accelerating the convergence speed; finally, the light-sensing idea of the firefly algorithm is used to refine and optimize the method, helping the method to jump out of local extrema, providing a better parental population for the next iteration, providing an optimization direction for the method, avoiding blind search, accelerating convergence, and thus improving the optimization quality of the method.

The present application discloses a tail wing control mechanism, which relates to the field of aircraft. A power piece is installed on the fuselage, and its output end is fixedly connected to the first rocker arm assembly and the second rocker arm assembly; a rotating mechanism, one end of which is rotatably connected to the first rocker arm assembly, and the other end of which is fixedly connected to the stabilizer surface of the tail wing; a locking mechanism is arranged between the first rocker arm assembly and the rotating mechanism; wherein, when the rotating mechanism is in a locked state, the rotating mechanism transmits the driving force of the first rocker arm assembly to the stabilizer surface, so that the deflection direction of the stabilizer surface is always opposite to the deflection direction of the rudder surface; the end of the second rocker arm assembly away from the output end of the power piece can be rotatably passed through the stabilizer surface on the front side of the rudder surface, and is fixedly connected to the rudder surface. The present application replaces the simple increase of the rudder surface area with mechanical linkage, avoids the problem of rudder effect overload during low-speed flight, takes into account the full-speed range control requirements, and effectively solves the contradiction between high-speed or low-speed rudder effect of transonic aircraft.

The application discloses a kind of launch vehicle fairing inner noise environment fast prediction method, comprising: obtaining the total noise of the core first stage engine nozzle center jet of launch vehicle;The distance from the characteristic interface of fairing to the interface where the core first stage engine nozzle is calculated;The jet outer noise at the characteristic interface of fairing is calculated;Using the aerodynamic noise empirical formula, the maximum aerodynamic outer noise of fairing in transonic flight phase is calculated;Based on the jet outer noise at the characteristic interface of fairing and the maximum aerodynamic outer noise of fairing in transonic flight phase, the maximum sound pressure level spectrum envelope of fairing is calculated;Based on the maximum sound pressure level spectrum envelope of fairing, combined with the sound insulation amount of fairing cabin section, the inner noise of fairing is calculated.The method can quickly predict the inner noise environment of new research launch vehicle fairing, solve the problem of new research launch vehicle fairing inner noise environment condition design.

An aircraft capable of transonic flight is equipped with a slotted airfoil to significantly improve the cruise performance during flight. The aircraft includes a plurality of slotted airfoils uniquely configured to maximize natural laminar flow across the aircraft wings. In addition to reduced wing drag, the unique configuration minimizes shocks at various conditions over the course of a flight. Each slotted airfoil includes a fore element, an after element, and a slot disposed between the fore element and aft element. The aft element is maneuverable relative to the fore element. The slot is configured to promote a substantially constant air flow field onto the aft element during all flight angles of attack.

The application discloses a kind of aerodynamic configuration structures and design methods of precision guidance assembly, belong to the field of precision guidance technology, assembly precursor uses continuous three curve design: supersonic section, uses parabolic profile, surface is provided with rib, for optimizing shock control under supersonic working condition;Transonic section, uses hyperelliptic surface, surface is provided with rib and is pasted on rib with super material, for inhibiting transonic separation;Subsonic section: using circular arc transition, surface is provided with rib, for improving rudder surface lift efficiency;Wherein, curvature is continuous at three curve junctions.The application realizes supersonic-transonic-subsonic three-section coordinated design, simultaneously combines the coordinated drag reduction design of micro rib and super material technology, improves boundary layer from micro level, optimizes and improves the performance of guidance assembly.

The invention provides a method and device for evaluating the comprehensive cooling efficiency of a turbine guide vane by using reverse heat transfer, and the method comprises the following steps: measuring a local convective heat transfer coefficient of the outer surface of a pneumatic test piece blade, and defining the local convective heat transfer coefficient as an external convective heat transfer coefficient without the influence of an internal cooling flow; acquiring a convective heat transfer coefficient between an internal cooling flow of the heat transfer test piece and a wall surface, and defining the convective heat transfer coefficient as an internal cooling coefficient; in the transonic speed blade grid wind tunnel, introducing mainstream gas into the outer surface of the heat transfer test piece and introducing heated internal cooling gas into an internal cooling cavity at the same time, so as to form a temperature difference through reverse heat transfer; measuring the convective heat transfer coefficient of the outer surface of the blade under the condition; and combining the Bit number determined by the external convective heat transfer coefficient and the wall thickness-heat conduction ratio, the ratio of the external convective heat transfer coefficient to the internal cooling coefficient, and the ratio of the correction coefficient to the internal cooling coefficient according to a preset function relationship to obtain the dimensionless comprehensive cooling efficiency. The problem that a traditional high-temperature test is difficult to implement is solved.

In order to solve the problems that in the prior art, calculation precision and efficiency are difficult to consider at the same time during supersonic jet machining, and detached shock waves cannot be accurately captured, the invention provides a finite element simulation method for the stagnation flexible constraint flow field characteristics of a jet impact area. The finite element simulation method comprises the steps that firstly, a geometric model of a jet area and an outer watershed is established, and preliminary solving is conducted; extracting jet flow axis pressure gradient characteristics to determine detached shock wave interval coordinates; then reconstructing a block geometric model comprising an upper jet flow region, a detached shock wave region, a stagnation region and an outer watershed according to the coordinates, and implementing differential local grid encryption on the shock wave region and the stagnation region; and finally, carrying out second iterative solution based on the Sutherland viscosity law to obtain final flow field distribution. According to the method, the thickness and the position of the shock wave layer are accurately captured while the number of the grids is controlled through the block encryption strategy, the contradiction between the calculation economy and the numerical precision under the transonic flow condition is solved, and the method is suitable for numerical simulation of the complex supersonic gas jet flow field containing the detached shock wave and the stagnation region.

The application discloses a method for suppressing transonic buffeting load and response based on agent cooperation, and relates to the technical field of aeroelasticity, and aims at solving the problem of transonic buffeting of a wing of an aircraft. The method comprises the following steps: firstly, calculating the elastic wing tip acceleration response based on a one-way fluid-solid coupling method; secondly, configuring a multi-target reinforcement learning agent, taking the variable-camber trailing edge and the telescopic micro-bulge as control mechanisms, taking the maintenance of the aerodynamic lift-drag ratio and the suppression of the vibration response of the structure as targets, and letting the agent adopt a proximal policy optimization algorithm to be trained to obtain an optimal control law capable of cooperatively driving the telescopic micro-bulge and the variable-camber trailing edge; and thirdly, deploying the trained cooperative control law on the target wing, and observing the suppression effect of the wing buffeting through quantitative evaluation of the slowing rate. The application trains the agent through a deep reinforcement learning method, so that the agent can guide the telescopic micro-bulge to suppress the diffusion in the separation zone, reduce the shock motion amplitude, drive the variable-camber trailing edge to regulate the flow at the trailing edge and compensate the overall lift fluctuation, and cooperatively suppress the buffeting load and response.

The invention relates to the technical field of aircrafts, in particular to an aircraft propulsion system and an aircraft, the aircraft propulsion system comprises a cartridge receiver, a propulsion device, a control device and an energy storage device; the propelling device comprises at least one stage of magnetic levitation rim motor set, a combustion chamber and an exhaust nozzle. The combustion chamber is located at the exhaust end of the magnetic levitation rim motor set and communicates with the exhaust end. The exhaust nozzle is located at the air outlet end of the combustion chamber and communicates with the air outlet end. The magnetic levitation rim motor set comprises at least one magnetic levitation rim motor, all the magnetic levitation rim motors are electrically connected with the energy storage device, and the control device is electrically connected with the energy storage device and all the magnetic levitation rim motors. And the control device is used for controlling each magnetic levitation rim motor to switch between a power generation state for supplying power to the energy storage device and an electric propulsion state for pressurizing air entering from the air inlet end according to the flight speed of the aircraft, so that the thrust requirement of the aircraft during high-speed flight is met, and the transonic flight capability of the aircraft is realized.

The present invention proposes a layout design method for a wing-body blended wide-speed range aircraft based on consideration of the rearbody effect. The method first generates a tapered waverider forebody for supersonic wide-speed range aircraft. Then, based on the basic airfoil, the basic wing and strake airfoils of the wide-speed range aircraft are aerodynamically optimized to obtain a multi-section wide-speed range waverider airfoil. The length of the waverider rearbody is then calculated and the wide-speed range aircraft symmetry plane is designed. Finally, based on the osculating cone method and transonic strake constraints, the waverider forebody parameters and the waverider rearbody length are used to obtain three-dimensional layout parameters. The optimized basic wing airfoil, strake optimized airfoil, and symmetry plane are assembled to generate the wide-speed range aircraft layout. The present invention couples the tapered waverider, wide-speed range airfoil optimization, and osculating cone double-swept waverider generation method, taking into account the rearbody aerodynamic design, thereby improving the aerodynamic performance of the wide-speed range aircraft at subsonic, transonic, and supersonic speeds.

A four-hole supersonic probe that does not require test calibration relates to a supersonic probe. The probe body adopts a seamless steel pipe with a bend near the bottom end, and four channel tubes extending with the probe body are arranged in the middle position inside. The four channel tubes are arranged at equal angles along the axis of the probe body and are solidly welded to the probe body. The front edge of the probe body is polished to form a conical probe for the measuring end. The four channel tubes are vertically connected with the conical probe busbar to form 4 holes. The 4 holes are all located at 50% of the busbar of the conical probe. The probe uses the slender cone theory under supersonic conditions to derive various aerodynamic parameters. The shock wave loss is small, and it helps to ensure the static pressure of the leading edge of the measured cone. The calibration coefficients are obtained through the slender cone theory, which greatly shortens the experimental time.

The application belongs to the technical field of wind tunnel test, and particularly relates to a transonic fan blade engine wind tunnel test system. By introducing a high-precision data synchronous recording device, it is ensured that the displacement signals captured by the vibration sensing component and the pressure data obtained by the dynamic pressure acquisition unit can be completely aligned in time. At the same time, the system also integrates an environmental parameter monitor, which can track and compensate the changes of physical quantities such as temperature, humidity and air pressure in real time, thereby significantly improving the accuracy of experimental data. Finally, the addition of an analysis platform and a result visualization interface realizes in-depth analysis of the blade vibration mode, provides intuitive and detailed data support for researchers, and greatly improves the understanding of the dynamic behavior of the blade under transonic conditions.

The application discloses a transonic buffet control structure based on a wing trailing edge air vent, and belongs to the technical field of aircraft flow control. The buffet control structure is an air vent arranged in the wing trailing edge. One end of the air vent is communicated with the upper surface of the airfoil, and the other end is communicated with the blunt trailing edge of the airfoil. The aperture of the air vent and the thickness of the trailing edge of the airfoil are of the same order of magnitude. The application can control the buffet while reducing the influence on the lift-drag characteristics of the original wing.

The invention relates to a method of operating a wind turbine. The wind turbine (1) comprises a rotor (4) comprising at least one blade (2) with an adjustable pitch angle (β). The method comprises adjusting the rotor speed (Ω) of the rotor (4) and / or the pitch angle (β) along a transonic safe trajectory. The transonic safe trajectory is based on an operational map comprising transonic flow risk information. The transonic flow risk information provides an indication of the risk of transonic air flows occurring around at least one blade (2) for a plurality of combinations of rotor speed (Ω) and pitch angle (β).

The invention provides a vehicle nose cone transonic speed reverse design method based on a bending shock wave theory and a vehicle nose cone. The method comprises the steps that incoming flow parameters are determined and designed according to a flight envelope; determining a curved shock wave surface shape function and a discrete shock wave molded line according to head cone geometric constraints, and solving discrete point aerodynamic parameters and first-order derivatives based on a curved shock wave theory; solving a subsonic velocity flow field control equation by adopting a time domain characteristic line method; and obtaining external flow field information by simultaneous equations, and reversely generating a nose cone wall surface. According to the method, through the innovative logic of'presetting the shock wave form-reversely deriving the wall surface ', the bending shock wave theory and the time domain characteristic line method are fused, the transonic shock wave structure is accurately controlled, the wall surface pressure distribution is optimized, the flow separation and shock wave oscillation are inhibited, the design efficiency is improved while the design precision is ensured, and the method is suitable for large-scale popularization and application. Reliable technical support is provided for rapid optimization design of the nose cone of the vehicle, and the method is suitable for aerodynamic design of nose cones of various transonic vehicles.

The invention provides a treatment casing for an ultrahigh-load transonic-speed serial rotor and a gas compressor, and relates to the technical field of pneumatic design of gas compressors. The serial rotor comprises a front-row rotor and a rear-row rotor which are sequentially arranged in the axial direction. The treatment casing comprises a casing body and a plurality of casing channels. The casing body is arranged outside the serial rotor in a surrounding mode. The multiple casing channels are arranged on the outer wall face of the casing body at intervals in the circumferential direction of the casing body, and each casing channel comprises a suction section, a connecting section and a spraying section. The suction section extends outwards from the outer wall face of the casing body and is suitable for sucking high-pressure air flow at the tail edge of a rear row rotor, the connecting section extends in the axial direction, and one end of the connecting section is connected with the suction section. The injection section extends from the other end of the connecting section to the axis of the casing body and is suitable for spraying the high-pressure airflow from the suction section to the front edge of the rear row rotor, and the included angle between the airflow sprayed out of the injection section and the inner wall face of the casing body is larger than 55 degrees and smaller than 70 degrees.

The present invention relates to a turbofan engine having a low-pressure turbine with at least four stages, wherein at least one of the at least four stages has a transonic blade cascade.

The invention discloses an airfoil profile stability optimization method and system for inhibiting transonic buffeting, and relates to the technical field of aircraft safety. According to the method, the geometric morphology of the airfoil profile is accurately described through CST parameterization, the grid quality and the calculation precision after the airfoil profile is deformed are guaranteed by combining the radial basis function dynamic grid technology, the flow stability is quantified by relying on the pre-solution gain, efficient iteration is conducted through the nonlinear conjugate gradient method, transonic buffeting can be remarkably restrained, and the flow stability can be remarkably improved; and the aerodynamic performance can be optimized on the premise that engineering constraints such as no reduction of the airfoil thickness and no reduction of the lift force are met, experience trial and error and violent search of traditional optimization are avoided, the efficiency and reliability of transonic airfoil design are greatly improved, and the method is suitable for high-performance design requirements of core components such as aerospace vehicle airfoils.

The invention belongs to the technical field of aviation aircraft design, and particularly relates to an aerodynamic configuration structure of a supercritical wing, the wing is generated by supercritical airfoil profile control of five specific stations in the spanwise direction, and 45% of the stations are reference airfoils. The front edge radius, the thickness and the mounting angle of the inner side (0% and 20% station) airfoil profile are sequentially larger than those of the reference airfoil profile; the leading edge radius of the outboard (70%, 100% station) airfoil profile is significantly larger than that of the reference airfoil profile, the mounting angle is negative, and the absolute values are sequentially increased. And front and rear edge points of each section are connected by using secondary splines to form front and rear edges of the wing, so that a three-dimensional molded surface is generated. Through differential design of spanwise airfoil parameters, a large-area separation area under a large angle of attack can be actively controlled to appear in the middle of the airfoil firstly, control failure caused by stalling of an outer wing and sudden lift drop caused by stalling of an inner wing are effectively avoided, and the stability of the airfoil is improved while the transonic cruise efficiency is maintained. And the stall safety of the upper single-wing transport plane is obviously improved.

The purpose of this invention is to provide a one-dimensional prediction method for the characteristics of a multi-stage axial-flow transonic turbine based on dimensionless numbers, comprising the following steps: turbine characteristic calculation; angle-of-attack loss calculation; total pressure loss calculation using a loss model; lag angle calculation using a lag angle model; transonic position determination; transonic cross-sectional characteristic parameter calculation; iterating the above process until the required accuracy is met, thus completing the characteristic prediction at the operating point. This invention's one-dimensional prediction method for the aerodynamic characteristics of a multi-stage axial-flow transonic turbine based on dimensionless numbers can plot turbine characteristic curves using turbine inlet velocity, inlet mass flow rate, and total expansion ratio as the abscissas; it can predict turbine design and non-design operating conditions; the method simultaneously considers the influence of angle of attack and lag angle, and incorporates a loss model; this method can simultaneously predict the characteristics of subsonic and transonic turbines; furthermore, this invention improves the accuracy of turbine characteristic prediction and shortens the design cycle of axial-flow turbines.

Apparatus and articles of manufacture are disclosed of an aircraft comprising: a brace to support a wing of the aircraft, the brace including a first end coupled to the wing and a second end coupled to a fuselage of the aircraft, the brace including: a strut extending between the first end and the second end; and an airfoil having a cavity containing the strut.

The invention discloses a data assimilation method for transonic airfoil profile streaming. The method comprises the following steps: S1, constructing a reduced-order model; s2, acquiring an input flow state and corresponding flow field observation data; s3, generating a flow initial set; s4, constructing a data joint space containing the flow state and the corresponding flow field simulation data, and filling the flow initial set into the data joint space; s5, performing simulation by adopting a reduced-order model to obtain flow field simulation data of the transonic airfoil in a flow state recorded in the data joint space, and updating the data joint space; s6, updating the reduced-order model by adopting an active learning method, and updating the data joint space; s7, calculating a mean value and a covariance of a data joint space, and calculating a Kalman gain; s8, updating a data joint space in combination with flow field observation data and Kalman gain; and S9, repeatedly executing the steps S5 to S8 until convergence. According to the method, the calculation cost is remarkably reduced while the data assimilation precision is ensured.

The present invention discloses a transonic stator and a centripetal turbine applied to a centripetal turbine. The transonic stator adopts a three-stage radial air flow passage and a "helmet-shaped" blade profile design. The width of the first-stage passage gradually decreases along the flow direction, greatly improving the air flow acceleration effect and significantly reducing the flow loss at the blade inlet. The width of the second-stage radial air flow passage remains unchanged, facilitating the inspection of the throat area of the stator. The change in the width of the third-stage radial air flow passage is coupled according to the air flow Mach number at the outlet of the stator blade, effectively reducing the flow loss in the vane-less area when the air flow Mach number at the outlet of the stator blade is in different situations, and further improving the acceleration performance of the stator. The "helmet-shaped" blade profile adopts a leading edge with a small curvature radius and a blade back with zero curvature design, which can effectively improve the leading edge load of the blade and the acceleration ability of the cascade passage, reduce the blade profile loss, and reduce the size and weight of the stator.

The invention provides a transonic unsteady aerodynamic force order reduction method based on global mapping, and the method comprises the steps: employing a filtered Gaussian white noise random signal as an input signal which needs to cover a wide frequency and amplitude range; carrying out numerical calculation on the designed input signal by adopting a CFD solver, constructing a multi-input-multi-output mapping relation between generalized displacement and aerodynamic force, and taking the multi-input-multi-output mapping relation as training data of the model; a Fourier neural operator model based on data driving is adopted to predict the unsteady aerodynamic force, and the Fourier neural operator model is trained; and inputting generalized displacement data into the trained Fourier neural operator model to predict a corresponding aerodynamic coefficient. The method is reasonable in conception, compared with a previous order reduction method, the time spent on training the same data is shorter, the precision and efficiency of aerodynamic coefficient prediction are higher, and dynamic linear characteristics and nonlinear characteristics can be captured at the same time.