386 results about "Aerodynamic force" patented technology

The invention provides a hypersonic transition and forced motion coupling dynamic derivative prediction method, and belongs to the technical field of aerospace data processing, and the method comprises the steps: constructing an initial transition model based on a standard Reynolds shear stress transport turbulence model framework by using an intermittent factor transport equation and a transition initial momentum thickness Reynolds number transport equation; carrying out primary correction on a transportation equation of the initial transition model, and carrying out compressibility and transverse flow effect correction on the initial transition model according to a compressible boundary layer theory and transverse flow correlation to obtain a corrected transition model; carrying out coupling calculation on the corrected transition model and forced vibration by utilizing a rigid dynamic grid technology to obtain unsteady aerodynamic force and torque; and identifying the unsteady aerodynamic force and torque by using an integral method to obtain a predicted combined dynamic derivative. The method has a full-dimension dynamic derivative analysis capability, breaks through speed domain limitation, and realizes full coverage from subsonic velocity or transonic velocity or supersonic velocity to hypersonic velocity working conditions.

The embodiment of the invention provides an aircraft aerodynamic model training method and an aircraft aerodynamic data determination method. The aircraft aerodynamic model training method comprises the steps of determining a training set according to multiple sample data of an aircraft, and performing model training according to the training set to obtain an aircraft aerodynamic model. Wherein the aircraft comprises a distributed propulsion system, the distributed propulsion system comprises a propeller, and the sample data is simulation data or test data of the aircraft and at least comprises the speed of the aircraft, the rotating speed of the propeller, aerodynamic force, torque and chamfer; variables of the training set include the speed of the aircraft and the rotating speed of a propeller, and target variables of the training set include aerodynamic force, torque and chamfer of the aircraft; the aircraft aerodynamic model is used for determining the aerodynamic force, torque and chamfer of the aircraft according to the speed of the aircraft and the rotating speed of the propeller. According to the technical scheme, the accuracy and comprehensiveness of obtaining the pneumatic data are improved.

The invention discloses a performance simulation calculation method for an aviation hybrid electric power system. The method comprises the following steps: firstly, constructing a sub-component calculation model covering an aerodynamic thermal component, a power generation subsystem, a power conversion and transmission subsystem, a battery subsystem, a motor and an aerodynamic system; establishing a reverse logic calculation framework driven by a load side demand, and realizing hierarchical reverse parameter solution from a load end to a power source end; and establishing a power balance equation at a bus end, carrying out iterative solution by adopting a Newton-Rafson algorithm, and realizing dynamic distribution and coordination control of multi-source energy in combination with an energy constraint module. According to the method, through a reverse solution and bus end balance strategy, the simulation calculation efficiency and convergence of the complex aviation hybrid electric power system are remarkably improved, and meanwhile, the adaptability and expandability of the model to different topological architectures and task profiles are enhanced; and an efficient and high-credibility calculation tool is provided for design, performance evaluation and energy management strategy optimization of the hybrid electric propulsion system of the aircraft.

The invention discloses a blade flutter monitoring method and device, a monitoring system, a storage medium and a computer program product, and the method comprises the steps: carrying out the modal feature analysis of a blade to determine a flutter monitoring region, laying a flexible sensing element in the flutter monitoring region, and enabling the flexible sensing element to be used for collecting the strain signal of the blade; in the running process of the blade, real-time running parameters of the blade and real-time strain signals collected by the flexible sensing element are obtained; determining a reference strain signal corresponding to the inertial centrifugal force of the blade according to the real-time operation parameters and a preset mapping relation; the mapping relation is the mapping relation between the operation parameters of the blade and the reference strain signal; deducting the reference strain signal from the real-time strain signal to obtain a target strain signal corresponding to the aerodynamic force; and according to the target strain signal and a preset formula, calculating pneumatic work of the blade in a set vibration period, and according to the pneumatic work, judging whether the blade flutters or not. According to the scheme, the accuracy and the real-time performance of flutter judgment are improved.

The invention belongs to the field of aerodynamic analysis of unmanned aerial vehicles, and particularly relates to a method and system for predicting aerodynamic force of a rotor wing based on a physical information neural network, and the method comprises the following steps: determining key parameter variables and variation ranges which affect the aerodynamic performance of the rotor wing; then, a Latin hypercube sampling method is adopted, and a sampling space is generated in the parametric variable interval; obtaining flow field data through CFD simulation; constructing and training a physical information neural network reduced-order model, and constructing a physical information neural network model by taking the sample point coordinate information of the sampling space as input data and taking the corresponding flow field aerodynamic force data as a training target; in the training process of the model, network parameters are optimized through a composite loss function; by minimizing the composite loss function, training to obtain a reduced-order model of the predicted flow field; and solving the aerodynamic force based on the reduced-order model. According to the method, a physical information network is adopted, the precision of a CFD calculation method is reserved, and the efficiency of aerodynamic force estimation under parameter change is improved.

An electric aircraft includes: one or two or more electric propulsion systems each including a propeller or fan for propulsion disposed to contribute to a lift of a main wing and an electric motor that drives the propeller or fan; and a controller that adjusts the electric propulsion system on the basis of a relationship between a variable relating to an operating state of the electric propulsion system and an aerodynamic force generated on the main wing such that a total thrust by the electric propulsion systems or the aerodynamic force on the main wing has a predetermined value or falls within a predetermined range.

An air-moving device may include an aerodynamic rotor in the form of a bladed disk. The aerodynamic rotor may include a physical geometry that constrains a noise level of the air-moving device.

The invention discloses a wind turbine generator full-blade load monitoring method based on tower wall vision measurement, and the method comprises the steps: S1, arranging an observation platform, and installing a camera; s2, arranging a blade mark group; s3, synchronizing image acquisition and time; s4, camera calibration and distortion correction; s5, feature detection and tracking; s6, performing three-dimensional reconstruction; s7, centrifugal force and gravity compensation calculation; s8, blade net displacement calculation; and S9, carrying out aerodynamic force load inversion and result output. According to the invention, monitoring of the load in the edge direction of the whole blade of the wind turbine generator can be effectively realized, and reliable technical support is provided for safe operation, optimal design and economical efficiency improvement of the blade of the large-scale wind turbine generator.

The disclosure relates to a motor vehicle, to a driver assistance system, and to a method for operating the driver assistance system. A relevant slipstream region of a further motor vehicle driving in a predetermined surrounding region of the motor vehicle is detected. A current slipstream efficiency is ascertained based on whether the motor vehicle is driving in the slipstream region and, if so, based on which of the slipstream partial regions of the slipstream region the motor vehicle is driving in. An overall aerodynamic efficiency is ascertained based on a current driving speed of the motor vehicle and the current slipstream efficiency. An individual overall aerodynamic efficiency potential that characterizes an overall aerodynamic efficiency that will be achieved by the motor vehicle when same drives in the slipstream region of the further motor vehicle is assigned to the further motor vehicle based on the driving speed and slipstream region.

The invention discloses a design method of an array type silencer adaptive to a main ventilation air path between a stator iron core of a hydraulic generator and an air cooler and the array type silencer, and relates to the technical field of suppression or reduction of noise of the hydraulic generator. The method comprises the steps of S1, obtaining aerodynamic noise data and airflow parameters of a main ventilation air path, S2, determining a silencing thought and a micro-perforated plate array core structure according to broadband vortex, discrete rotation, vortex falling and surge noise, S3, determining the boundary dimension and preliminarily calculating the airflow parameters, S4, designing plate parameters in combination with a micro-perforated resonance formula, and S5, calculating the noise of the micro-perforated plate array. S5, performing simulation optimization and S6, accurately accounting airflow parameters, and locking a scheme meeting requirements; the silencer comprises a shell and a plurality of micro-perforated plates, and the multiple layers of micro-perforated plates are arranged in series. The invention has the advantages of no sound absorption material, low wind resistance (the air volume attenuation is less than or equal to 5%, and the pressure loss is less than or equal to 200Pa), the noise of the main wind path is reduced by 5-7dB (A), the service life is long, the operation and maintenance cost is low, and the requirements of the main wind path are met.

The invention discloses a method and a device for correcting hole wall interference after wing type stall of a high-lift rotor wing and a medium, and belongs to the field of wing type wind tunnel test data correction. The method is based on a single-parameter linearization wall pressure information method, and upper and lower tunnel wall pressure and airfoil aerodynamic force data in a wind tunnel test are measured; obtaining a vorticity field through numerical simulation or a PIV test, identifying a separation vortex, and calculating a vorticity center of the separation vortex to determine a vortex core center position; a source is arranged in the center of the vortex core of each separation vortex, and the intensity of the source is the same as or integral multiple of the vorticity; the original singular point distribution is kept unchanged, and the influence of the superposition source realizes hole wall interference correction after stall, so that the corrected lift force, resistance and pitching moment coefficients are obtained. The method is suitable for a rotor wing type wind tunnel test with the incoming flow Mach number ranging from 0.1 Ma to 0.9 Ma, the aerodynamic data correction accuracy can be improved after stall, and the correction result is highly consistent with reference data.

The invention discloses a digital twin platform of an aircraft flight control system, a data processing method of the digital twin platform and a medium, and relates to the technical field of aircraft simulation training, and the platform comprises a parameter setting unit which is used for setting aircraft basic information and current flight state parameters; the control signal acquisition unit is used for acquiring an electric signal generated when a user operates the control device; the calculation unit is used for calculating the initial control surface aerodynamic force according to the control surface aerodynamic force calculation model, calculating the control surface deflection angle according to the control surface aerodynamic-control device force transmission model, and recalculating the control surface aerodynamic force according to the control surface aerodynamic force calculation model; and the force sensing feedback unit is used for calculating joystick feedback force in the current aircraft flight state and the current control state according to the control surface aerodynamic force-control device force transmission model, and feeding back generated corresponding control resistance to the control device. The platform not only saves the economic cost and occupied area for establishing an entity iron bird test bench, but also can test the performance of a flight control system of an airplane in different flight states.

The invention discloses a wind tunnel test aerodynamic force and displacement signal phase difference correction method based on self-excited vibration characteristics. The method comprises the following steps: independently collecting structure surface wind pressure and vibration displacement signals; integrating the wind pressure signal to obtain a self-excitation aerodynamic force time history; identifying the vibration frequency and the total damping ratio of the system from the displacement signal; filtering the aerodynamic force time history to extract a fundamental frequency component; based on a structural dynamic balance principle, establishing an analytic relational expression between an aerodynamic force out-of-phase component coefficient and total damping of the system; and carrying out iterative time translation on the aerodynamic force time history by taking the displacement signal as a reference until the out-of-phase component of the aerodynamic force time history meets the analytic relationship, so as to determine and correct the system delay. The invention further discloses a wind tunnel test aerodynamic force and displacement signal phase difference correction system based on the self-excited vibration characteristic and a storage medium. According to the method, electric triggering and physical transmission delay can be compensated with high precision without additional hardware, and the pneumatic parameter identification reliability is effectively 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 discloses a parachute flight performance prediction method based on follow-up coupling flow field analysis, and the method is suitable for a parafoil and a traditional round parachute. Parachute flight flow field aerodynamic characteristics and flight dynamic characteristics are coupled by using a body-attached dynamic grid and a self-adaptive flow field boundary technology, real-time changes of a parachute streaming flow field along with body motion are accurately reflected, and the attitude, the speed and the flight trajectory of the parachute are automatically solved and obtained. According to the method, on one hand, real-time aerodynamic characteristics and motion characteristics of the parafoil are coupled, the coupling influence of flight attitude speed change on aerodynamic force and aerodynamic torque of the parafoil is reflected, and the calculation precision is guaranteed; and meanwhile, the change of the streaming flow field of the parachute system is decoupled into the change of the relative incoming flow and the change of the posture, so that the redundancy of flow field grids and the complex grid reconstruction are avoided, and the simulation calculation efficiency is improved.

The invention discloses a tilt rotor aircraft depending on aerodynamic force, and relates to the technical field of aircrafts. The four-axis unmanned aerial vehicle comprises a fuselage, wing assemblies, four-axis unmanned aerial vehicle modules and empennage assemblies. The wing assembly comprises a left wing, a right wing and a wing cross beam which are all provided with ailerons; the four-axis unmanned aerial vehicle module is I-shaped, is hinged to a wing cross beam through a bearing, and can tilt bidirectionally; the empennage assembly comprises a tail support, a horizontal electric variable pitch propeller, a horizontal empennage with a control surface and a vertical empennage. The four-axis unmanned aerial vehicle module serves as a tilting unit, and the structure is simplified; attitude control is realized by virtue of aerodynamic difference, and the efficiency is improved; the tail horizontal electric variable-pitch propeller realizes rapid pitching adjustment; vertical take-off and landing, low-altitude maneuvering and high-speed cruising capabilities are integrated, the application scene is wide, and safe forced landing can be achieved when faults occur.

The invention discloses an order reduction method for aerodynamic reconstruction in a wallboard ground flutter simulation test, and relates to the technical field of aircraft design, and the method comprises the following steps: employing a hypothesis modal method to carry out spatial discretization on a wallboard structure of an aircraft, and obtaining a first expression; determining a second expression based on the intrinsic mode matrix; substituting the second expression into the first expression to obtain a third expression; substituting the second expression into the first aeroelastic equation to obtain a second aeroelastic equation under the intrinsic mode coordinates; physical aerodynamic force is determined based on the third expression and the second aeroelastic equation; the physical aerodynamic force is subjected to order reduction through the displacement interpolation matrix and the force interpolation matrix, and the physical aerodynamic force after order reduction is obtained. According to the method, the number of loading points is reduced, it is guaranteed that the limited number of loading points are equivalent to the distributed aerodynamic force, the method is not limited by the structure size, test implementation is facilitated, time domain fitting is not needed, and the calculation precision and efficiency are higher.

A system for controlling at least one active aerodynamic appendage of a road vehicle is described, comprising an actuation control unit operating to implement a predictive control of the active aerodynamic appendage, based on an estimation, performed dynamically while the road vehicle is in motion, of an aerodynamic load demand of the road vehicle in a route section subsequent to one currently travelled.

The invention discloses a high-speed aircraft aerodynamic configuration agile optimization method based on an artificial intelligence model. The method comprises the following steps: A1, establishing an aerodynamic force agent model; a2, performing a global optimization algorithm; a3, establishing an intelligent autonomous decision-making model; by constructing a multi-target adaptive weight distribution model, weight dynamic collaborative optimization of proxy model confidence, a constraint penalty function and target function key elements is achieved, the mathematical modeling problem of dynamic matching of weight parameters and optimization targets in the iteration process is solved, a closed-loop optimization system with the autonomous perception-decision-execution capacity is developed, and the optimization efficiency is improved. Therefore, the convergence speed is improved in a complex engineering optimization scene, local optimum is avoided, and the robustness of a global optimal solution is ensured; and A4 pneumatic optimization iteration: based on the proxy model for predicting indexes such as the lift-drag ratio, performing global optimization of design parameters by using an artificial intelligence optimization algorithm, performing verification through CFD, and after verification, putting a CFD calculation result into a training set of the model to re-train the model and continuing optimization until optimization design is converged.

Disclosed is an aerodynamic system for an underside of a vehicle, the system comprising: an aerodynamic chassis, the aerodynamic chassis comprising a main compartment bounded by a plurality of walls and a floor extending between the walls; and an aerodynamic device, the aerodynamic device comprising: a mounting structure, the mounting structure being fixedly mounted to the floor of the aerodynamic chassis; and an aerodynamic body, the aerodynamic body comprising a leading edge, a trailing edge and at least two outboard edges, the aerodynamic body being fixable to the floor of the aerodynamic chassis by the mounting structure, the mounting structure being located wholly inboard of the outboard edges of the aerodynamic body.

The invention discloses an inflatable returner structure safety analysis method and system considering a thermo-pneumatic elastic effect, and the method comprises the steps: building a structural dynamics sub-model, a fluid domain calculation sub-model and a structural thermodynamics sub-model of an inflatable returner, and comprehensively considering the structural thermal deformation caused by aerodynamic heat and the thermal expansion effect of an internally inflated gas; and a bidirectional coupling relationship among aerodynamic force, aerodynamic heat, structural deformation, temperature and internal charging pressure is constructed through the feedback effect of structural deformation on a flow field and an aerodynamic load. And carrying out multi-physics field simultaneous solution by adopting a time step propulsion method, and finally realizing comprehensive safety evaluation of the temperature, stress and deformation characteristics of the structure. According to the method, aerodynamic load distribution, temperature field change, stress response and vibration characteristics of the inflatable returner in the reentry process can be accurately predicted, so that the thermal protection performance and structural safety of the inflatable returner are comprehensively evaluated, and reliable analysis means and engineering basis are provided for optimization design and structural stability improvement of a thermal protection system.

The invention discloses a method and a device for generating a simulation motion image of a fixed wing equipment model, and belongs to the technical field of computer simulation. The method comprises the following steps: acquiring model parameters, motion state parameters and airflow parameters of a fixed wing equipment model; calculating aerodynamic data of the fixed wing equipment model by using an aerodynamic model according to the model parameters, the motion state parameters and the airflow parameters; the motion state of the fixed wing equipment model is updated according to the aerodynamic force data, and updated motion state parameters are obtained; and generating a simulation motion image of the fixed wing equipment model according to the updated motion state parameters. The generation scheme of the simulation motion image of the fixed wing equipment model disclosed by the invention is high in simulation precision, and the simulation efficiency can be remarkably improved.

This invention discloses a method for correcting the phase difference between aerodynamic and displacement signals in wind tunnel tests based on self-excited vibration characteristics. The method includes: independently acquiring wind pressure and vibration displacement signals from the structural surface; integrating the wind pressure signal to obtain the self-excited aerodynamic time history; identifying the vibration frequency and the total system damping ratio from the displacement signal; filtering the aerodynamic time history to extract the fundamental frequency component; establishing an analytical relationship between the aerodynamic out-of-phase component coefficient and the total system damping based on the structural dynamic equilibrium principle; and iteratively shifting the aerodynamic time history using the displacement signal as a reference until the out-of-phase components satisfy the analytical relationship, thereby determining and correcting the system delay. This invention also discloses a system and storage medium for correcting the phase difference between aerodynamic and displacement signals in wind tunnel tests based on self-excited vibration characteristics. This invention requires no additional hardware and can accurately compensate for electrical triggering and physical transmission delays, effectively improving the reliability of aerodynamic parameter identification.

An active downforce control method for a vehicle may include determining an index of phase-shifted interactions between a driver of the vehicle and a controller. The controller is configured to control one or more aerodynamic actuators. Furthermore, the method may include determining one or more inputs to the active downforce control, at least partially, based on the index of phase-shifted interactions. The method may also include controlling one or more aerodynamic actuators, at least partially, based on the one or more inputs to the active downforce control.

This invention relates to the field of wind tunnel aerodynamic measurement technology for hypersonic vehicles, and discloses a method for evaluating and optimizing the dynamic load testing accuracy of a wind tunnel force measurement system. The evaluation method involves: equating the force measurement system to a multi-degree-of-freedom spring and mass model; constructing a dynamic differential equation and performing a Laplace transform; then, combining the dynamic stiffness matrix to construct a frequency response function matrix; defining the output of the force measurement channel of a rod-type strain balance and solving for the transfer function from the excitation point to the force measurement channel; combining the reference transfer function and the actual load, calculating the identified load and performing a Fourier transform to obtain the relative identification error of the force measurement system at the excitation point; finally, calculating the power spectrum weighted average identification error at the excitation point and comparing it with the engineering allowable error to determine whether the system has undergone stiffness optimization. This invention not only accurately identifies the dynamic load of the force measurement system but also achieves stiffness optimization, improving the accuracy of unsteady aerodynamic force measurement.

The invention provides a rotating shaft balance structure based on topological optimization, the rotating shaft balance structure comprises an inner ring, measuring beams and an outer ring, the inner ring is provided with a key hole, and the inner ring and the outer ring are connected through a plurality of groups of measuring beams; a strain gauge is attached to the measuring beam and can sense deformation of the measuring beam, and six-component aerodynamic force borne by a measured object is converted into an electric signal to be output. The outer ring is of a hollow structure composed of beams with irregular surfaces and holes, and a plurality of mounting holes used for being connected with a tested device are formed in the outer ring in the circumferential direction. A topological optimization design is adopted for the outer ring of the rotating shaft balance to form a hollow structure consisting of beams and holes with irregular surfaces, so that the rotating shaft balance greatly reduces the mass of the outer ring while ensuring the design rigidity, and solves the problem that the resistance measurement error is relatively large due to rigidity hardening of a measurement beam caused by centrifugal force during high-speed rotation.

This invention relates to a method and system for rapidly calculating the steady-state periodic response of a rigid body flapping rotor, belonging to the field of rotor flapping response calculation technology. It utilizes a rotor region mesh to iteratively calculate the rigid body flapping angle, flapping velocity, and flow field velocity, and then calculates the rotor's aerodynamic forces based on blade element theory. The aerodynamic forces are then used to calculate the rotor momentum source term. This term is added to the momentum source term in the Navier-Stokes equations, and the rotor region mesh is updated. Finally, the aerodynamic forces are used to calculate the Faroe Islands momentum source term. n This invention updates the rotor rigid body flapping response in a loosely coupled manner. It calculates the rotor's steady-state periodic flapping response to the collective pitch control input, longitudinal periodic pitch control input, and lateral periodic pitch control input. By loosely coupling the rotor momentum source term and the rotor rigid body flapping response, this invention comprehensively considers the influence of aerodynamic components on the flow field, effectively improving computational efficiency.

This provides a rear vehicle structure that can further enhance the aerodynamic performance provided by the rear spoiler. [Solution] The vehicle rear structure 10 is a structure disposed at the rear of the vehicle body 12 that generates an aerodynamic effect. The vehicle rear structure 10 comprises a rear spoiler 20 and a vehicle-side facing member 21. The rear spoiler 20 is a member that extends along the width direction of the vehicle body 12. The vehicle-side facing member 21 is a member disposed on the side of the vehicle body 12. The vehicle-side facing member 21 faces the rear spoiler 20 and is configured to protrude toward the rear spoiler 20.

The invention discloses a buffeting load multi-point synchronous measurement device and method for a large-span arch bridge wind tunnel test. The device comprises a base, a wind direction angle adjusting mechanism, an arch bridge truss model, a model supporting system and a multi-channel synchronous data acquisition system. The arch bridge truss model comprises at least two test segments embedded with six-component balances, the six-component balances are located in the geometric centers of the segments and fixedly connected with the main framework through rigid force transmission structures, and gaps are formed between the adjacent segments; the wind direction angle adjusting mechanism rotates around a vertical shaft, and the model supporting system is matched with the wind attack angle adjusting mechanism through a supporting stand column to achieve overall pitching of the arch bridge truss model so as to adjust the wind attack angle of the arch bridge truss model. And the multi-channel synchronous data acquisition system synchronously acquires time history signals of the three-dimensional aerodynamic force borne by each test section. According to the invention, double-degree-of-freedom independent adjustment, multi-point synchronous force measurement and load field reconstruction can be realized, the method is suitable for arch ribs with different curvatures, and meanwhile, the buffeting load identification precision of the large-span arch bridge can be remarkably improved.

The invention particularly relates to a direct force carrier landing control method based on predefined time active disturbance rejection control, and belongs to the technical field of aviation. Aiming at the problems of high landing accident rate and large wake flow disturbance influence of a shipboard aircraft, uncertain convergence of a traditional control method and the like, the method comprises the following steps: firstly, establishing a six-degree-of-freedom nonlinear model, an aerodynamic force and an aerodynamic moment model of the shipboard aircraft, and fusing the wake flow disturbance influence; constructing a torque control channel and a direct force control channel, and converting the torque control channel and the direct force control channel into a control-oriented model; and then designing PTADRC-based attitude angle, angular rate and track angle control laws, and realizing total disturbance estimation and compensation through PTESO. Simulation verification shows that the method has no overshoot and no steady-state error tracking instruction in the predefined time, is high in anti-interference capability, can guarantee the stability and accurate control of the landing attitude of the shipboard aircraft, and remarkably reduces the landing risk.