572 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 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 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 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 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.

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.

In the present disclosure, the side spoiler is provided on a lower side of a mobility vehicle so that aerodynamic performance is improved, and a deployment angle of the side spoiler is controlled according to driving conditions and road conditions so that desired aerodynamic performance is realized, and damage to the side spoiler is prevented. In addition, introduced is a side spoiler device for the mobility vehicle, in which the side spoiler is divided into a plurality of side spoilers so that aerodynamic performance for each driving situation is optimized through individual control of each side spoiler.

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.

The invention relates to the technical field of low-orbit aircrafts, in particular to an airfoil auxiliary control air-breathing electric propulsion upper-layer atmosphere aircraft which comprises a main aircraft body, a deflection folding assembly, an electric driving system and a power supply system. The deflection folding assembly is used for locking the position of the folding wing folded on the main wing before the aircraft takes off, and is used for actively unlocking the main wing to unfold the folding wing and locking the position of the unfolded folding wing after the aircraft enters an orbit; when the folding wings of the aircraft are in the unfolded state, the folding wings are driven to deflect so as to adjust the deflection angles of the folding wings relative to the surfaces of the main wings; the power supply system is used for supplying power to the electric driving system and the deflection folding assembly. According to the device, the folding wings are folded to reduce the size of the whole aircraft, so that the requirement on the space size of the carrier is reduced; the aerodynamic lift-drag ratio of the aircraft is adjusted by adjusting the deflection angle of the folding wings, and then aerodynamic force control capacity is provided for the aircraft.

The invention relates to the technical field of unmanned aerial vehicle flight control, in particular to a track tracking control method for a tandem dual-rotor unmanned helicopter, which comprehensively considers factors such as aerodynamic force and inertia force and is completed through structural parameter measurement, aerodynamic force analysis, kinetic equation derivation and model verification and correction. Designing a self-adaptive sliding mode controller, defining a sliding mode surface, designing a control law and a self-adaptive law, ensuring the robustness of the system, and optimizing a control strategy; in combination with model prediction control, establishing a prediction model, defining a target function, improving the track tracking precision through rolling optimization and feedback correction, and correcting prediction errors and interference influences through sliding mode control; a multi-task switching control strategy is designed, task stages are divided, parameters are set, switching conditions are determined, and a fuzzy logic reasoning or neural network algorithm is adopted to realize rapid adjustment of the control strategy. The technology can accurately track the track, enhance the robustness, improve the real-time performance, flexibly adapt to multiple tasks and have a good application prospect.

The invention discloses an unmanned aerial vehicle control method combining sequence modeling and predictive control, and the method comprises the steps: collecting the state information, including speed information, attitude information, position information and PWM motor control signals, of an unmanned aerial vehicle; obtaining a state observation value through a Transform model according to the collected state information, and predicting future residual aerodynamic force according to the state observation value and the current residual aerodynamic force to obtain predicted residual aerodynamic force; dynamically adjusting the predicted residual aerodynamic force by adopting a least square estimation method; and according to the adjusted residual aerodynamic force and the target trajectory, an optimal flight trajectory is generated, and unmanned aerial vehicle control is completed. According to the method, a sequence modeling mechanism based on a Transform structure and a model prediction control strategy are fused, and the control stability, precision and adaptive capacity of the unmanned aerial vehicle in a complex wind field are remarkably improved.

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.