800 results about "Aerodynamic force" patented technology

The invention relates to an autonomous energy-saving soaring flight path planning method for a small-sized low-cost aircraft, belongs to the technical field of aircraft trajectory planning, solves the problem of low-cost wind field energy acquisition of the small-sized low-cost aircraft in the prior art, and comprises the following steps: S1, configuring a sensor for the aircraft, and measuring through the sensor to obtain observation parameters; s2, establishing a state vector of the aircraft; s3, establishing an aerodynamic force model, introducing a dynamic equation and a state transition equation, and performing accurate modeling on aerodynamic force; s4, performing multi-source data fusion by adopting extended Kalman filtering, establishing an extended Kalman filter of a nonlinear system, and executing real-time wind vector high-precision sensing; and S5, performing global wind field modeling, estimating a wind field environment, and performing energy-obtaining flight path planning to obtain an optimal energy-obtaining soaring flight path planning scheme.

The invention provides a new class of transportation with a partially aerostatically supported ram air cushion ship (PASRACS) that can provide safe, fast, efficient global transport services with extraordinary comfort, luxury and amenities. The PASRACS employs inventive synergistic combinations of lift from aerostatic, aerodynamic, hydrostatic and hydrodynamic forces for different modes of operation ranging from a stationary floating configuration on a water surface, to flight in ground effect at high speed. A propulsion system can utilize hydrogen as an energy source driving fluid-dynamic thrusters, to enable zero carbon emissions operations. A transition method is provided for a PASRACS to transition from a floating mode to a takeoff mode to a flight in ground-effect mode in an inventive optimized manner. A transport method for multimodally transporting payload is provided with PASRACS vehicles and systems, with quick turn time operations enabled by the use of payload transfer transport modules and transfer vehicles.

The invention belongs to the technical field of multi-component aerodynamic load measurement sensors, and discloses a method for applying a constant force to a half-mode balance calibration system. The method for applying the constant force comprises the following steps: installing the constant force value loading device; accurate force value loading is carried out; carrying out displacement real-time compensation; constant force value output is kept at all the force application device point positions; and completing the ground static calibration of the half-mode balance. According to the method for applying constant force, the displacement of the electric cylinder is calculated according to the value of the applied force, after the corresponding displacement of the electric cylinder is completed through closed-loop control of the driving electric cylinder, the steel spring inevitably generates creep displacement in the process of waiting for arrival of a plurality of force application device point positions, and the creep displacement is changed into the creep displacement. The applied force value corresponding to the creep deformation is synchronously measured by the high-precision force sensor, the compensation displacement of the piezoelectric ceramic is correspondingly calculated, the creep deformation displacement is dynamically compensated in real time, the constant force value loading device is ensured to be always at the required constant force value, and the high-precision calibration of the half-mode balance is realized.

The invention belongs to but not limited to the technical field of interception nets, and particularly relates to a multi-point traction type interception net launching device which comprises a base, a power device, a launcher and an interception net. The base can be adjusted and positioned at multiple angles, the borne impact force is at least not less than 15 KN, the base is designed into a folding structure, the packaging size is reduced, and transportation and storage are facilitated; the power device adopts compressed air as a power source, an air storage chamber is filled with high-pressure air, when a switch is turned on, energy of the high-pressure air is released instantly, and huge aerodynamic force acts on the pressed area of an object, so that the object moves in an accelerated mode from a static state and is thrown out. The power device body is composed of a launching tube, an air chamber, a valve element assembly, a firing device, a trigger, a carbon fiber air cylinder, a shell and the like, the launching tube, the carbon fiber air cylinder and the shell are distributed on the two sides of the air chamber, the firing device is arranged above the air chamber, and the trigger is pressed on the firing device; the launcher is in a horn shape, and four cylindrical storage bins connected with the launching tube are arranged on the edge of the launcher.

The invention relates to the technical field of helicopter aerodynamic force simulation, and discloses a sea rescue helicopter rotor aerodynamic force simulation method and system. The method comprises the following steps: acquiring a rotor wing downwash airflow historical speed data set, a historical sea surface wave data set and a historical aerodynamic pressure data set monitored within a preset historical time range; executing a dynamic volatility analysis process on the three types of historical data sets to generate a speed volatility index, a sea surface wave volatility index and an aerodynamic pressure volatility index; taking the three types of volatility indexes as guide parameters, performing adaptive search operation in a rotor aerodynamic force simulation parameter database, and determining an optimal simulation time interval; and performing numerical calculation simulation on pneumatic coupling between the rotor wing and the sea surface according to the optimal simulation time interval to generate a simulation data sequence, performing feature extraction and classification on the simulation data sequence by using an aerodynamic force recognition model, and outputting a rotor wing aerodynamic force simulation result.

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 ball is a thin-walled hollow sphere that includes an array of apertures with dimples in a pattern that are concentric to the aperture array. Dimples are provided on the exterior surface of the ball between the apertures with the pattern and dimensions of the dimples varying based on proximity to the apertures. Internal protuberances are provided on the interior surface of the hollow ball such that the interior surface also manages aerodynamic airflow through the ball. Channels are also provided around each of the apertures within the ball which may intersect proximate dimples or be spaced therefrom. The ball also includes apertures having converging and diverging sidewalls that form a venturi shape to increase inflow air into the apertures and outflow air out of the apertures. Finally, the ball also maintains a weight:diameter ratio within a fixed range.

The invention discloses a small sample aerodynamic modeling method based on multi-task learning. Comprising the following steps: constructing a multi-task prediction model, integrating an auxiliary task network and a target task network, and enhancing the feature extraction capability of an encoder through an SE layer and an attention layer; acquiring multi-source aerodynamic force data and carrying out standardized preprocessing on the multi-source aerodynamic force data; a dynamic weight mechanism is designed, the influence of auxiliary task prediction on target task output is adaptively adjusted according to input features, and effective fusion of low-fidelity data and high-fidelity data is achieved; two-stage training is carried out to ensure the efficiency and stability of multi-task learning; the prediction capability of the model under the limited sample condition is verified; through the synergistic effect of multi-task knowledge migration, feature enhancement and dynamic task integration, in combination with the feature optimization characteristics of an SE layer and an attention mechanism, high-precision aerodynamic prediction is realized under the condition of limited samples, and an efficient and accurate aerodynamic prediction method is provided for aircraft design.

The invention relates to a CFD (Computational Fluid Dynamics) method for nearly continuous hypersonic high-temperature thermochemical unbalanced flow numerical simulation, which comprises the following steps of: 1, introducing a Park double-temperature model into an NS equation to describe a thermochemical unbalanced effect at high temperature, and closing a shear stress tensor, a heat flow vector and a mass diffusion flux by adopting an equivalent constitutive model to obtain a closed shear stress model; capturing speed slippage and temperature jump phenomena on the wall surface by adopting corresponding slippage boundary conditions to obtain a macroscopic control equation and boundary conditions suitable for near-continuous hypersonic high-temperature thermochemical non-equilibrium flow; 2, discretely solving the macroscopic control equation and the boundary conditions given in the step 1 by adopting a finite volume method; and step 3, performing post-processing on a solution result obtained in the step 2 to obtain a streaming flow field, a wall surface aerodynamic force and a wall surface heat flow of the calculated appearance.

The invention provides a rotating speed determination method in aero-engine complete machine three-dimensional simulation, and belongs to the technical field of aero-engine simulation, and the method specifically comprises the steps: constructing a fluid-structure interaction control equation on each grid unit of an aero-engine complete machine geometric model; forming an integral coupling matrix, and solving the integral coupling matrix; simulating the calculated thermal state geometric model; the gas compressor and the turbine can only rotate around the rotating shaft; calculating aerodynamic forces acting on the compressor blades and the turbine blades; in each time step, the rotating speed is updated according to the angular acceleration; and after a plurality of iterations, when the difference value of the aerodynamic torque borne by the gas compressor and the aerodynamic torque borne by the turbine accords with a preset range, the corresponding rotating speed is the steady-state working rotating speed of the engine under the thermal state geometry. According to the processing scheme, the engine can automatically match and dynamically determine the working rotating speed in the simulation process, the inaccuracy of the preset rotating speed is avoided, and the precision and reliability of a simulation result are improved.

The invention discloses a large flexible wing static aeroelasticity analysis method based on a physical information neural network, and belongs to the field of aircraft design. The method comprises the following steps: for the wing with the super-high aspect ratio, firstly, dispersing the wing into n spanwise sections based on a strip theory, and initializing a reference angle-of-attack parameter of each spanwise section; then, a physical information neural network (PINN) is built, unsteady flow field solving is carried out on each section, for the current ith iteration, the physical information neural network is used for sequentially traversing and extracting pressure distribution of the airfoil surface of each section, respective aerodynamic load distribution is calculated, and the generalized force of the airfoil is obtained; and solving a structural dynamics equation based on the generalized force and structural information of the wing, calculating the deformation of the large-deformation wing by using a modal rotation method, and finally, carrying out solving iterative circulation on aerodynamic force and structural deformation until the attack angle is converged, so as to obtain the actual lift force L of the wing. According to the method, the calculation precision and efficiency are remarkably improved, and different flight conditions and wing design can be adapted.

The invention provides a self-adaptive cross damping flutter suppression method for blades of an offshore wind turbine, and belongs to the technical field of wind power generation based on computer data processing. The method comprises the following steps: firstly, establishing an aeroelastic model containing a non-diagonal damping item, and coupling sinking and floating motion and pitching motion through a cross damping coefficient; secondly, designing a self-adaptive control law based on a differential Riccati equation, and dynamically adjusting a damping parameter to compensate the time-varying aerodynamic force; and finally, proving closed-loop stability through a Lyapunov theory, and verifying a control effect by utilizing numerical simulation. An external pneumatic executing mechanism is not needed, the structure is simple, response is fast, the vibration convergence time can be shortened to be within 4 seconds at the wind speed of 8 m / s, the steady-state error approaches zero, and the method is suitable for flutter suppression of the blades of the large offshore wind turbine.

The invention relates to an anti-disturbance high-maneuverability flight control method for a miniature four-rotor unmanned aerial vehicle, and the method comprises the steps: introducing a linear time-varying actuator model on the basis of a standard angular velocity dynamical model of the four-rotor unmanned aerial vehicle, so as to more accurately consider the aerodynamic effect, and improve the maneuverability of the four-rotor unmanned aerial vehicle. The anti-interference effect of the algorithm provided by the invention in a gust environment is further verified; the controller provided by the invention is insensitive to noise and actuator delay, and can ensure good transient and steady-state performance in the high-maneuverability flight process of the quad-rotor unmanned aerial vehicle; according to the adaptive enhanced control module structure provided by the invention, the adaptive law and the control law are clearly separated. This separation allows better adjustment of performance and does not reduce the robustness of the closed loop system. The method shows that the tracking effect of the real angular velocity on the reference angular velocity is good, the phenomena of obvious overshoot and angular velocity oscillation do not exist, dynamic changes and pneumatic coupling in a complex flight environment can be effectively dealt with, and the accuracy and safety of flight control are ensured.

The invention discloses a point cloud neural network aerodynamic heat rapid prediction method considering a variable geometric shape, and belongs to the field of aerodynamics, and the method comprises the steps: determining a research object, and constructing a data set; then, a three-dimensional point cloud neural network is adopted to extract global shape features, local shape features and incoming flow condition features of different variable geometric shapes under the same basic configuration to serve as point cloud neural network input parameters, and the same model is adopted to predict wall surface pressure and heat flux density; wherein the training data of the three-dimensional point cloud neural network adopts the wall surface heat flux density and pressure under the non-uniform wall temperature boundary condition. According to the method, efficient prediction of the wall surface pressure distribution and the heat flux density under the same basic configuration and under the non-uniform wall temperature boundary condition of different variable geometric shapes can be realized, and the prediction accuracy is improved.

The invention relates to a stress calculation method, device, equipment and medium for a floating type draught fan foundation structure, and the method comprises the steps: carrying out the frequency domain potential flow water power analysis of a wet surface element model of a foundation floating body of a floating type draught fan, and obtaining the water power pressure distribution and a water power coefficient; boundary conditions are applied to the simulation model of the floating type wind turbine foundation structure, various working conditions are loaded, finite element analysis is carried out based on hydrodynamic pressure distribution, and various structural stress transfer functions are obtained; based on the hydrodynamic coefficient, performing linear processing on the fan aerodynamic force, the impeller pneumatic torque, the floating body hydrodynamic force and the mooring restoring force of the floating fan and the influence of a fan servo controller on the floating fan, and substituting a linearization result into a motion equation of the floating fan for calculation to obtain a plurality of dynamic response spectrums; and multiplying the structural stress transfer functions by the corresponding dynamic response spectrums, and adding a plurality of multiplied results to obtain a total stress spectrum. The method can improve the efficiency.

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 relates to the field of underwater pressurization, and discloses an underwater contra-rotating compressor blade tip leakage flow vibration prediction method and system based on fluid-solid coupling simulation, and the method comprises the steps: carrying out the division of flow field grids and the division of solid domain grids based on a constructed underwater contra-rotating compressor impeller three-dimensional model; establishing a blade flow field CFD model and a blade solid domain FEA model; aNSYS Fluent software is used for conducting underwater contra-rotating compressor three-dimensional unsteady blade flow field simulation so as to obtain unsteady blade tip leakage flow characteristics of a concerned flow field area; an external data module in ANSYS Workbench software is used for loading the aerodynamic force, obtained through unsteady flow simulation, of the surface of the blade to the surface of a blade solid domain FEA model, transmission of the aerodynamic force load between different blade surface grids is achieved, and one-way fluid-solid coupling is achieved; aNSYS Workbench software is used for conducting underwater counter-rotating compressor blade vibration simulation, vibration response of the blade under the combined action of aerodynamic force and centrifugal force is obtained, and the fatigue life of the blade is predicted based on a fatigue model.

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 discloses a multi-degree-of-freedom variant wing aeroelastic coupling test measurement method. The method comprises the following steps: S1, arranging a sensor network; s2, a PIV flow field is measured, laser sheet light is adopted to irradiate the surface of the wing, a high-speed camera captures tracer particle displacement, and a flow field velocity field is calculated; s3, constructing an aeroelastic inverse problem model, and inverting the unsteady aerodynamic force by combining the structural deformation, the vibration speed and the flow field speed; s4, correcting a model error by using a deep learning network; s5, constructing a Kalman filtering-data driven hybrid model; and S6, self-adaptive threshold flutter early warning is carried out, and the system stability is judged based on a Lyapunov index. According to the measurement method based on distributed optical fiber sensing, high-frequency PIV and laser Doppler vibration measurement fusion, high-precision and full-dimension synchronous acquisition of aeroelastic parameters can be realized; in addition, in combination with PIV flow field data, structural deformation and vibration data, the unsteady aerodynamic force is inverted through an aeroelastic inverse problem model.

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.

The invention discloses a brush type sealing fluid-solid-heat-grinding multi-physics field efficient coupling method, and particularly relates to the field of aero-engine sealing. Comprising the following steps: establishing an interdigitated tube bundle solid model, taking the interdigitated tube bundle solid model as a first model, and executing a division step; cFD simulation is carried out on the first model after grid division, and flow field distribution is obtained; determining a three-dimensional aerodynamic load of the brush wires based on the flow field distribution; determining the three-dimensional pneumatic deformation result of the brush wires according to the three-dimensional pneumatic load, and determining the wear volume of each brush wire; according to the pneumatic deformation result of the brush wires and the abrasion volume of each brush wire, the forked row tube bundle solid model is updated, and a deformation geometric model under aerodynamic force is obtained; and taking the deformation geometric model as a first model, starting to execute a division step, and stopping iteration until a brush wire three-dimensional pneumatic deformation result reaches a first condition, so as to finish accurate simulation of a fluid-solid-heat-grinding coupling effect.

The invention relates to the technical field of fatigue testing, and discloses a blade fatigue testing device and method based on a vortex-induced vibration principle, the blade fatigue testing device based on the vortex-induced vibration principle comprises a test bench foundation and at least one excitation unit, the test bench foundation is suitable for fixing a blade root of a tested blade; the excitation unit comprises a blade cover body and a pneumatic loading module, the blade cover body is of an annular structure fixedly arranged on the periphery of a test blade in a sleeving mode, the pneumatic loading module is arranged below the blade cover body, and the pneumatic loading module comprises a frame and at least one air blowing piece installed on the frame; wherein the air blowing piece is configured to blow air to the blade cover body, and turbulent flow is formed at the blade cover body. Non-contact excitation is achieved through aerodynamic coupling, the additional mass influence caused by mechanical connection is avoided, and full-size fatigue verification of the blade is completed under the condition that the natural vibration characteristic is kept.

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 an endogenous safety self-adaptive fault-tolerant control method for an overdrive vector rotor aircraft, solves the problem of how to improve the reliability of the vector rotor aircraft, and belongs to the field of rotor aircraft control. The method comprises the steps that in order to ensure that the vector rotorcraft can achieve a good fault-tolerant effect under different fault mode conditions, the fault state is described as the actuator aerodynamic force loss; the deviation of an expected control distribution matrix and an actual control distribution matrix at the current moment in a dynamic system and reference model output error of the control distribution system is progressively converged to a self-adaptive law of a zero design error, so that bounded uncertainty does not need to be estimated; the method only depends on the IMU feedback information of the aircraft, and no additional sensor needs to be added for fault detection. The vector rotor function can be realized under the conditions of complete failure of a single rotor, complete failure of two interphase and opposite rotors and complete failure of three interphase rotors.

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 invention discloses an aerodynamic force / thermal agent model-based rapid calculation method for the thermo-dynamic elasticity of a thermal protection plate of a high-speed aircraft, and the method comprises the steps: building an unsteady aerodynamic force / thermal agent model of a thermal protection wall plate in a high-speed flow based on a Kriging agent model method; a structural dynamics method based on a finite element method is used for establishing an aeroelastic module, a transient heat conduction equation is used for establishing an aerothermal module, and a thermal protection wallboard thermodynamic elasticity analysis frame is built. Finally, the aerodynamic force / thermal agent model is used for carrying out thermal protection wall plate thermodynamic elasticity calculation and analysis. Compared with a traditional high-order method, the calculation efficiency is improved by four orders of magnitude.

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.