182 results about "Aerodynamic load" patented technology

The invention discloses a floating type offshore wind turbine aerodynamic load calculation method based on a frequency domain method. The method comprises the steps that a state derivative is determined based on a blade element momentum theory quasi-steady-state assumption, and a dynamic inflow model is used for correction; establishing a pneumatic-elastic coupling two-degree-of-freedom linear dynamic model; establishing a wind wheel kinetic model; establishing a control model according to a floating type offshore wind turbine pitch angle control strategy, a torque control strategy and a floating body motion feedback control strategy; based on the pneumatic-elastic coupling two-degree-of-freedom linear dynamic model, the wind wheel dynamic model and the control model, pneumatic additional mass and pneumatic additional damping are obtained; pneumatic exciting force acting on the hub when the position of the cabin is fixed is obtained; and based on the pneumatic additional mass, the pneumatic additional damping and the pneumatic exciting force, an overall linear dynamic model of the fan is established and solved, and overall response is obtained. According to the method, the calculation result precision can be improved while the calculation efficiency is ensured, and a technical support is provided for accelerating the simulation efficiency of the offshore wind turbine.

The invention provides a strong nonlinear simulation method and device suitable for a large floating fan, and the method comprises the steps: dividing a target floating fan into a plurality of body structures according to the model parameters of the target floating fan, so as to construct a multi-body model containing a plurality of degrees of freedom, and building a whole machine load and a motion transmission path; the aerodynamic load input of each body structure is calculated, and the nonlinear dynamic response of the blade is calculated according to the aerodynamic load input of each body structure; according to the coordinate transformation relation, the nonlinear dynamic response of the blade is converted into a whole coordinate system and input into a multi-body dynamic model of the whole machine, generalized active force and generalized inertia force of each body structure in the target floating type draught fan are obtained through calculation, and the motion response of each body structure of the target floating type draught fan is obtained through simultaneous solution. According to the scheme, the motion response of each body structure of the target floating fan can be accurately determined, so that the running state of the floating fan in the deep and far sea complex environment can be accurately evaluated.

The invention belongs to the technical field of numerical simulation and analysis of unmanned aerial vehicles, and relates to a rigid-flexible coupled variable-camber-range deformable wing optimization design method, which comprises the following steps of: obtaining working conditions of an aircraft, determining wing types and basic parameters of wings of the aircraft, and establishing a three-dimensional aerodynamic model of the wings; variable-attack-angle pneumatic simulation is carried out on working conditions of the aircraft in take-off and cruise, optimal camber values and corresponding pneumatic loads of wings under multiple working conditions are determined, and the positions of rotating joints of the deformed wing ribs are adjusted to obtain the deformed wing ribs after primary optimization; dividing the optimization range of the deformed wing rib after primary optimization into a fixed area and an optimizable area, loading an aerodynamic load under the optimal camber value in the optimizable area to perform aerodynamic simulation to obtain a deformed wing rib after secondary optimization, and optimizing the structure; and a deformed wing rib sample is made, performance testing is conducted, and the optimizable area is adjusted till the wing performance meets the preset requirement. The aerodynamic efficiency of the full flight envelope can be improved.

The invention discloses an aerodynamic load determination method and device of a train and the train, and relates to the field of aerodynamics. The method comprises the steps that geometric shape data of the train with the load to be determined are obtained; inputting the geometric shape data of the train with the to-be-determined load into the prediction model to obtain the pressure and shear stress of the train with the to-be-determined load, which are output by the prediction model; and according to the pressure and the shear stress of the train with the to-be-determined load, determining the resistance, the lift force and the lateral force of the train with the to-be-determined load. The prediction model is trained through the data of the train with the known load, the trained prediction model only needs to input the geometric file of the train object to be calculated, the pressure and the shear stress can be obtained, the threshold in the professional field is not involved, and therefore all workers have the operable ability. Simulation calculation is replaced by the prediction model, the prediction model does not need to be subjected to simulation calculation again, the pressure and the shear stress are directly output, the calculation time is shortened, and the calculation efficiency is improved.

A method for reducing aerodynamic loads on an aircraft located on the ground includes moving a wing tip relative to a fixed wing to an on-ground configuration in which the span of the wing is reduced. The wing tip includes an airflow channel, which is in an open configuration in which airflow through the channel between upper and lower surfaces of the wing tip is enabled, when the wing tip is in the on-ground configuration. This has been found to reduce aerodynamic loads on the wing tip and / or parts of the aircraft connected thereto during this phase of the aircraft's use.

The invention belongs to the technical field of unmanned aerial vehicle aircraft design, and particularly relates to a longitudinal aerodynamic parameter identification calculation method in a closed-loop state and a large angle-of-attack range, and the method comprises the steps: building a flight simulation model; performing aerodynamic characteristic simulation based on a flight simulation model to obtain an aerodynamic load value, comparing the aerodynamic load value with data tested in an actual flight experiment, and calculating and minimizing an error sum of squares; dividing the whole angle-of-attack range into a plurality of sub-intervals, respectively obtaining the optimal aerodynamic parameters of each sub-interval based on the minimum sum of squares of errors, inputting each optimal aerodynamic parameter into the flight simulation model, and carrying out model accuracy verification; and drawing an aerodynamic characteristic curve according to each optimal aerodynamic parameter, and verifying the applicability of the flight simulation model. The method has universality and is suitable for longitudinal aerodynamic coefficient parameter identification of different types of unmanned aerial vehicles, the parameter identification efficiency is improved, and meanwhile the applicability of the model under the complex flight working condition is enhanced.

The invention discloses an aerodynamic load loading method, device and equipment and a medium, relates to the technical field of ocean engineering, is applied to a processor, and comprises the following steps: receiving scale motion data currently acquired by a scaling model of a floating wind turbine generator, and amplifying the scale motion data to obtain full-scale motion data; determining a first expected aerodynamic load borne by the floating wind turbine generator based on the full-size motion data; reducing the first expected aerodynamic load to obtain a second expected aerodynamic load borne by the scaling model; decomposing the second expected aerodynamic load into output thrust of each propeller in the loading device, and converting the output thrust into a PWM control signal; the loading device regulates and controls the rotating speed of each propeller based on the PWM control signal, and loads a target aerodynamic load generated by rotation to the scaling model; and skipping to the step of receiving the scale motion data again until a preset stop condition is met. And the loading of the aerodynamic load on the scale model is accurately realized.

The invention discloses a self-coupling PID (proportion integration differentiation) flight control method of a rudder-controlled transverse-row type rotor unmanned aerial vehicle based on dynamic multi-dimensional heterogeneous fuzzy logic, and belongs to the technical field of flight control of rudder-controlled transverse-row type rotor unmanned aerial vehicles. A flight control system architecture of the self-coupling PID flight control method is composed of three core modules: (1) a dynamic multi-dimensional fuzzy logic coupling PID controller (DMHFL-SCPID); (2) control instruction distribution; and (3) a six-degree-of-freedom dynamical model of the rudder-controlled transverse rotorcraft. According to the control algorithm, input and output dynamic parameters of a system are analyzed in real time, and four-dimensional flight control quantities (roll, pitch, yaw and height instructions) are generated. A control distribution strategy based on the aerodynamic characteristics of the rotor converts the instructions into actuating quantity of an actuating mechanism, and flight attitude adjustment is realized by changing aerodynamic load distribution of a rotor system, so that the unmanned aerial vehicle is guided to realize accurate tracking along a predetermined flight path.

The invention provides a friction stir tunnel forming device and method for enhancing bearing performance, relates to the field of machining and inner flow channel forming, and aims to overcome the defects of an inner flow channel prepared through an existing friction stir tunnel forming technology in the aspects of structural bearing and fatigue performance. The non-rotating part is of a hollow structure, the rotating part is coaxially arranged in the non-rotating part, the non-rotating part comprises a clamping part, a step part and a shaft shoulder part, the bottom of the shaft shoulder part is provided with a reinforcing rib manufacturing part, the reinforcing rib manufacturing part is used for containing an inner runner surface material and providing forging pressure, and an inner runner surface with a reinforcing rib structure is manufactured. And a rounded rectangular inner flow channel is manufactured at the stirring needle part of the rotating part, so that the bearing efficiency and rigidity of the rotating part in an aerodynamic load and vibration environment are remarkably improved. And through different sections of the stirring needle part, the manufacturing requirements of the aerospace field on the lightweight and integrated heat management structure in the aspects of complex forms and multifunctional integration can be met.

The invention relates to the technical field of fan equipment management, in particular to a wind turbine gearbox fault diagnosis method and system based on multi-modal data fusion, and the method comprises the steps: carrying out the key structure recognition of a wind turbine gearbox, and marking the key structure as a gearbox key structural part; vibration signals of key structural parts of each gearbox and impeller strain signals of a wind turbine are collected; extracting aerodynamic load characteristics from the impeller strain signal; inputting the aerodynamic load characteristics into a pre-constructed load transmission evaluation model to obtain a load influence transmission coefficient set; performing impeller influence correction on each vibration signal based on the load influence transfer coefficient set to obtain a vibration signal correction data set; performing multi-dimensional feature extraction on the vibration signal correction data set, and converting extracted vibration features into a three-dimensional vibration feature matrix; and performing fault identification on the three-dimensional vibration characteristic matrix by using a preset gearbox fault identification model. The pneumatic load interference can be eliminated, and the fault diagnosis accuracy is improved.

This invention discloses a dynamic pressure-based method for testing and analyzing hammer shock waves in aero-engines, belonging to the field of aero-engine testing and structural strength design. The method first uses a surge-induced pressure method to generate hammer shock waves at the engine inlet. Based on the flow field characteristics of the air intake outlet section (AIP section) during surge, dynamic total pressure and static pressure measuring points are arranged in a combination of circumferential and radial directions. Pressure signals are synchronously recorded using a high-speed data acquisition system, and four characteristic parameters—surge overpressure, overpressure ratio, loading time, and pressure rise rate—are defined to quantitatively analyze the intensity, spatial distribution, and dynamic process of the hammer shock waves. This invention achieves rapid, accurate, and comprehensive characterization of the aerodynamic load of hammer shock waves, providing reliable input loads directly for the strength design of complex structures such as S-curve air intakes, thereby improving aircraft safety.

This application discloses a method, device, medium, and product for active control of aerodynamic noise in the target area during tiltrotor transition flight, relating to the field of noise control. The method includes: during tilt transition conditions, constructing a target noise control area within a set region based on flight mission and environmental requirements, discretizing it into multiple spatial observation points and assigning weights; parameterizing the controllable aerodynamic loads on the blades to obtain control load variables, establishing the relationship between the control load variables and the acoustic response of the target noise control area, determining unknowns, and establishing a global acoustic optimization objective function; using the DDQN algorithm as a solver, optimizing the global acoustic optimization objective function under the premise of satisfying actuator physical constraints to obtain the optimal control law, performing feedforward execution and feedback correction to achieve active control of aerodynamic noise in the target area. This application can solve the main shortcomings and derived problems caused by focusing on single-point noise control.

The application relates to the technical field of blade monitoring, in particular to a long flexible blade real-time monitoring and fault early warning method, which comprises the following steps: collecting vibration signals of a long flexible blade in a running state, wherein the vibration signals are coupled by multiple factors including aerodynamic load, gravity and centrifugal force; constructing an over-determined equation set containing multi-source excitation characteristics based on a blade dynamics model; decomposing the vibration signals by using a sparse decomposition algorithm to obtain signal decomposition results; solving the over-determined equation set based on the signal decomposition results to obtain modal parameters of each coupled vibration source; and determining a state monitoring result of the long flexible blade according to the modal parameters of each coupled vibration source.

The present disclosure discloses a rotor blade aerodynamic load simulation device, comprising a base, an aerodynamic load module and a clamping module are arranged on the base, wherein the clamping module is used for clamping a rotor blade; the aerodynamic load module is used for aerodynamic load simulation experiment on the rotor blade clamped by the clamping module. The device has simple structure, low cost, adjustable multi-directional angle of simulated blade excitation, wide coverage of excitation frequency, and excitation form closer to actual rotation.

Turbomachine (10) comprising several guide vanes (36), each guide vane (36) comprising a blade (37), wherein the turbomachine (10) comprises: opposing walls (40, 42) defining a passage (38) in which a fluid flow is received to flow through the passage (38), wherein a constriction distribution (Do) is defined at a narrowest region in the passage (38) between adjacent guide vanes (36), in which adjacent guide vanes (36) extend transversely across the passage (38) between the opposing walls (40, 42) to interact aerodynamically with the fluid flow;wherein the constriction distribution (Do) is defined by the airfoil (37), wherein the constriction distribution (Do) is defined by values ​​specified in Table 1, and wherein the values ​​of the constriction distribution (Do) are within a tolerance of + / -10% of the values ​​specified in Table 1, wherein the constriction distribution (Do) is configured to reduce aerodynamic losses and improve aerodynamic loads on each airfoil (37); and wherein the airfoil (37) has at least one of: a thickness distribution (Tmax / Tmax_midspan) as defined by values ​​specified in Table 2; a dimensionless thickness distribution corresponding to values ​​specified in Table 3; and / or a dimensionless axial chord distribution corresponding to values ​​specified in Table 4.

The invention relates to an eVTOL nacelle tilting steering engine based on high reduction ratio transmission and a working method, and relates to the field of electric vertical take-off and landing aircrafts. According to the technical scheme, the design that the double-cycloid speed reducer is combined with the parallel shaft speed reduction gear set is adopted, the wiring harness passing ability problem from the aircraft end to the nacelle is solved, the single-point failure risk that a lead screw of a traditional linear steering engine is stuck is avoided, and the safety and reliability of the steering engine are improved; moreover, depending on the dual-redundancy winding, the dual-redundancy electromagnetic brake and the dual-redundancy rotating speed feedback of the dual-redundancy servo motor, the dual-redundancy design is realized from the electrical level, and the safety performance is further enhanced. In addition, external disturbance can be reduced through high-reduction-ratio transmission, the influence of aerodynamic load disturbance on nacelle actuation is reduced, and the tilting stability is greatly improved.

The utility model discloses an airfoil locking structure which is arranged on an aircraft. The structure comprises a power mechanism and a fixing mechanism, wherein the power mechanism is arranged between the folding wing and the fuselage and is used for driving the folding wing to unfold; the fixing mechanism comprises an in-place block, a mounting block, a limiting pin, a first compression spring, a limiting plate, a contact rod and a second compression spring, a vertical first mounting hole is formed in one side of the bottom of the mounting block and matched with a machine body limiting hole, and the limiting pin is movably mounted in the hole and connected through the first compression spring; a mounting groove and a second mounting hole which are communicated are transversely formed in the other side of the mounting block, the limiting plate belt moves in the mounting groove through the hole, a plate body of the limiting plate belt can block the limiting pin, the limiting pin is released through the hole, and the contact rod is movably connected in the second mounting hole, fixedly connected with the limiting plate and connected through the second compression spring. According to the structure, the high-reliability unfolding and locking process is achieved through pure mechanical design, it is ensured that the folding wing is stably fixed under the aerodynamic load, and the overall structure is reasonable, small in size and light.

This invention relates to a method, device, and medium for intelligent assisted driving instability early warning of vehicles in windy and rainy conditions. The method involves collecting rainfall intensity, crosswind speed, and vehicle driving state parameters; querying a CFD wind-rain coupled aerodynamic load database to obtain corresponding aerodynamic six-component force data; inputting the aerodynamic six-component force data as external disturbances into the vehicle's multibody dynamics model, simulating and outputting vehicle state response parameters; calculating environmental impact parameters based on rainfall intensity and crosswind speed, and dynamically correcting the preset normal thresholds of various instability indicators based on these environmental impact parameters to obtain dynamic instability thresholds under windy and rainy conditions; assigning weights to various instability indicators according to the vehicle's current driving conditions, and calculating a comprehensive instability index based on the vehicle state response parameters and dynamic instability thresholds; determining the instability level based on the comprehensive instability index and calculating the remaining stabilization time. Compared with existing technologies, this invention has advantages such as high accuracy, strong adaptability, and good real-time performance.

The invention discloses a vertical axis fan pneumatic-wake flow coupling calculation method based on an octree method. The vertical axis fan pneumatic-wake flow coupling calculation method comprises the following steps: presetting initial parameters and environmental conditions of a wake flow speed and a pneumatic load; the wake flow field is divided into a near field and a far field, the wake flow velocity caused by near-field particles is directly calculated by adopting a classical Biot-Savart law, the wake flow velocity caused by far-field particles is solved by adopting an octree improved algorithm, and the aerodynamic load is calculated; the relative speed and the attack angle of the vertical axis fan blade are obtained, then the aerodynamic load on the blade is calculated, and meanwhile the induced speed, obtained through calculation, of the blade node is used for calculating the wake flow field speed in the next time step; in the wake flow module, all vortex elements perform convective motion along with the local induced velocity field, and newly released vortex elements form a new wake flow field for calculation of the next time step. The invention aims to provide a pneumatic-wake flow coupling response numerical calculation method capable of giving consideration to both precision and efficiency, and an effective numerical tool is provided for engineering application of a vertical axis fan.

This invention discloses an aerodynamic optimization design method for scaled-down model blades of large vertical axis wind turbines based on DMST theory. The method includes: determining scaled-down parameters according to wind tunnel similarity criteria; selecting an optimized airfoil for aerodynamic characteristic matching at low Reynolds numbers based on the lift-to-drag ratio characteristics of the prototype blades; constructing an aerodynamic performance model based on dual-disk multi-flow theory, calculating and verifying the theoretical thrust coefficients of the prototype and optimized airfoils; conducting force-pressure tests in a wind tunnel, extracting the pure thrust of the blades by subtracting the total thrust of the turbine from the tower thrust vector; and comparing the measured thrust coefficients with the theoretical thrust coefficients to verify the aerodynamic load matching degree. This invention effectively eliminates the Reynolds number effect in scaled-down tests by combining airfoil optimization, DMST theory verification, and wind tunnel test verification, achieving high-precision matching of aerodynamic loads on scaled-down model blades of vertical axis wind turbines and improving the reliability of wind tunnel test data.

This application discloses a method, device, and storage medium for calculating the aerodynamic distributed loads of an entire aircraft. The method includes: extracting the positional characteristic information of each component of the aircraft; determining the structural positional network of each component based on the positional characteristic information; obtaining the flight state parameters of the aircraft; calculating the aerodynamic load distribution data of each component based on the structural positional network and the flight state parameters; and performing strength verification on each component based on the aerodynamic load distribution data. This invention, by establishing a structural positional network, achieves systematic calculation of the aerodynamic distributed loads of the entire aircraft, providing an accurate and reliable load data foundation for aircraft structural strength verification.

The application discloses a flexible film self-supply flow control device based on Tesla valve and belongs to the field of aircraft design and wind turbine. First, a metal-film mixed skin structure is designed in the application, and a gas cavity body connected with a film skin is arranged below the film skin. Further, a Tesla valve is arranged in front of and behind the gas cavity body. The inlet of the front Tesla valve is communicated with the front opening of the gas cavity body, and the outlet is communicated with the air blowing port on the metal skin of the wing leading edge, so that the gas flows from the cavity to the airfoil in the Tesla valve. The inlet of the rear Tesla valve is communicated with the air suction port on the metal skin of the wing trailing edge, and the outlet is communicated with the rear opening of the gas cavity body, so that the gas flows from the airfoil to the cavity in the rear Tesla valve. Finally, under the action of dynamic stall pulsating aerodynamic load, the film skin generates flow-induced vibration and provides energy for air blowing and suction, and the flow-induced vibration and air blowing and suction are coordinated to generate disturbance to the dynamic stall flow field, so that the purpose of dynamic stall control is achieved.

The invention belongs to the technical field of nacelle air inlet duct resistance reduction, and relates to a nacelle air inlet duct resistance reduction structure and method.The nacelle air inlet duct resistance reduction structure comprises an inner annular flexible film section embedded in the inner wall of a nacelle air inlet duct, and the front edge and the rear edge of the inner annular flexible film section are in sealed connection with adjacent inner wall steps through bonding layers; the outer annular flexible film section is embedded in the outer wall of the nacelle air inlet channel, and the front edge and the rear edge of the outer annular flexible film section are in sealed connection with the adjacent outer wall steps through bonding layers respectively; under the action of aerodynamic load, the flow velocity gradient and the adverse pressure gradient of a near-wall area of the surface are passively adjusted through the structural flexibility and deformation response of different positions of the surface of the inner and / or outer annular flexible film section, so that the shear stress and friction resistance of airflow to the surface are reduced.

The invention discloses a shape-preserving and resistance-reducing design method for a rotating mechanism under rotor / fuselage coupling, which comprises the following steps of: setting the slenderness ratio of a streamline shaft sleeve based on a shape resistance coefficient formula, and verifying the surface pressure distribution and the separation area of the shaft sleeve through numerical simulation to ensure that the shape resistance coefficient meets a resistance-reducing target; setting the area ratio of an outlet to an inlet of the seam based on a Venturi effect formula, so that a low-pressure area is formed in the seam to suck a wing root boundary layer and delay airflow separation; establishing an aerodynamic load and motion parameter coupling model based on an instantaneous resistance formula, analyzing flow fields under different rotation angular velocities and oblique angles through numerical simulation, and optimizing a rotation track to reduce a resistance peak value and aerodynamic torque fluctuation; the invention aims to solve the problems of airflow vortex interference, large shape resistance, high resistance peak and poor aerodynamic stability by optimizing the shape of the rotating shaft sleeve, the fuselage-wing gap structure and the wing rotating track, and improves the cruise fuel efficiency and flight stability of the aircraft.

The application discloses a kind of based on multiple type signal dry vacuum pump working state intelligent diagnosis method and system, specifically related to vacuum pump monitoring technical field, by synchronous acquisition encoder pulse and vibration acceleration signal, equal-angle resampling establishes angular domain sequence, by pulse interval calculation instantaneous angular velocity and angular acceleration and combined with moment of inertia to obtain actual inertia moment, based on screw profile and ideal compression model generates theoretical aerodynamic load moment and difference integration obtains leakage flux index, corrects rotor dynamics aerodynamic term and iteratively inverts unbalance eccentricity, exports health status and fault grade.The application adopts angular domain alignment to make criterion consistent under speed fluctuation and working condition change, constructs leakage flux index by difference integration of aerodynamic load and inertia moment, and iteratively inverts physical unbalance eccentricity with residual error, so that leakage and unbalance have distinguishable physical quantity representation, improve evaluation consistency and traceability.

The invention discloses a propeller system dynamics analysis method based on a shell unit, and the method comprises the steps: firstly determining the geometric parameters and design conditions of a propeller according to the design task and working condition of an aviation aircraft; and blade chord length distribution and mounting angle distribution are calculated based on an aerodynamic load rule, and an NACA airfoil profile is selected to generate a three-dimensional geometric model. Then, grid division is carried out according to the model, the connection relation between blade nodes and units is established, and a quadrilateral eight-node shell unit grid is formed; a finite element model is established by adopting an isoparametric element theory, a stiffness matrix and a mass matrix are calculated, and then a kinetic equation of the propeller is established. The inherent frequency and vibration mode characteristics of the blade are obtained through modal analysis, and on the basis, aerodynamic loads are applied for dynamic response analysis. According to the method, the parameterized modeling process of the propeller from geometric modeling to kinetic analysis is realized, the aerodynamic coupling vibration characteristics of the blade can be efficiently evaluated, and the method has the advantages of high modeling precision, high analysis efficiency and high adaptability.

This invention relates to the field of wind turbine structural health monitoring, and discloses an adaptive monitoring method for the attitude of hybrid-tower wind turbines based on operational status feedback. The method includes: synchronously acquiring the wind turbine operational data stream and the original tower attitude monitoring signal; determining the transient aerodynamic thrust based on the operational data stream, and using this thrust to determine the dynamic propagation velocity of the mechanical excitation wave generated by the stress stiffening effect; determining the propagation delay based on the dynamic propagation velocity, and performing a reverse time shift on the monitoring signal to achieve time-domain synchronization; reducing the feedback weight of the measurement update link of the solution model under the unit's operational adjustment state, and outputting the tower attitude data. This invention establishes a mapping relationship between aerodynamic loads and mechanical wave propagation characteristics, achieving precise decoupling between the unit's operational excitation and the tower's actual response, eliminating monitoring data distortion caused by spatiotemporal misalignment, and improving the reliability of safety monitoring for large-megawatt units.

The invention provides an aircraft structure load distribution determination method, and relates to the technical field of aircraft structure load design, and the method comprises the following steps: S1, generating an aircraft aerodynamic load conversion model node force; s2, generating structural finite element model node aerodynamic force; s3, generating acceleration and angular acceleration of the aircraft in three directions of the aircraft body coordinate system, and calculating to obtain structure node inertia force; s4, the structural node aerodynamic force and the structural finite element model node inertia force are combined to generate a total structural load, and shear bending moment distribution is generated; as important input of aircraft structure strength design, rapid processing and generation of aircraft structure loads are achieved. According to the processing method, finite element software is not needed for solving, the aircraft structure load processing process is processed and standardized, the calculation efficiency is high, and the aircraft structure load can be rapidly processed only through aerodynamic load input, aerodynamic load conversion model node and unit data and aircraft mass distribution.

The application discloses a kind of aerodynamic load reproduction devices for fan model pool test, including support, tower tube, cabin, near end cantilever, far end cantilever, folding mechanism, variable pitch mechanism, controller, power supply, the cabin is set on the tower tube, support and cabin are connected, power supply is arranged in cabin, the edge of support is provided with multiple near end cantilevers that are equally spaced along circumference, and the inner end of each near end cantilever is respectively connected with variable pitch mechanism, and variable pitch mechanism is located in support, the outer end of each near end cantilever is connected far end cantilever by folding mechanism, and fan main machine is arranged on each far end cantilever, controller is set in support, and controller and power supply are electrically connected, and folding mechanism, variable pitch mechanism, fan main machine are electrically connected with controller respectively.The application can reproduce real-time aerodynamic load caused by real wind field in the process of test, improve test precision.

The utility model provides a fan structure. The fan structure comprises a center hub and a plurality of blades. A plurality of reinforcing ribs which are arranged at intervals are arranged on the circumferential outer surface of the central hub, and the blades are arranged on the central hub through the reinforcing ribs; the protruding height of the reinforcing ribs on the circumferential outer surface is gradually reduced in the direction from the front edge portion of the corresponding blade to the rear edge portion of the corresponding blade. The technical problems that due to the fact that the blade root (especially the front edge area) bears the composite effect of centrifugal force and aerodynamic loads, front edge local stress concentration is likely to be caused, microcracks are induced, and fatigue fracture is accelerated are solved. Particularly, under the working conditions of high rotating speed or frequent start and stop, the joints of the blades and the hub become structural weak points, and the service life of the fan is remarkably shortened.