38 results about "Landing performance" patented technology

The invention belongs to the field of airplane structure dynamic strength design, and particularly relates to a dynamic load static equivalent coefficient determination method and a structure parameter design method.The method comprises the steps that S1, a relation curve of a water pressure load at the bottom of an airplane body and a time history is obtained according to a water forced landing dynamic load measurement test, and the peak value Ps of the water pressure load is determined according to the relation curve; s2, carrying out a fuselage bottom elastic model water forced landing test, and determining critical failure parameters of an elastic structure; s3, performing a static mechanical property test according to the critical failure parameters of the elastic structure to obtain the maximum static pressure Pd which can be borne by the elastic structure; and S4, calculating the static equivalent coefficient of the dynamic load of the forced landing on the water. The static equivalent coefficient of the dynamic load of the water pressure of the forced landing on the water is obtained through a scale model test. The objective of the invention is to solve the problem of lack of design load in airplane water forced landing performance design.

The present disclosure provides an unknown unmanned system data-driven multi-step kernel prediction control method and system. The method is aimed at the problem of unknown non-parametric unmanned system model. Only the off-line collected system input and output noise data are used. The Hankel matrix is constructed to extract the time sequence characteristics. The kernel matrix is constructed by combining the renewable kernel function. The nonlinear system rule is converted into high-dimensional linear representable data rule. The multi-step kernel prediction optimization problem based on the kernel matrix is solved on-line and rolling. The optimal control input is output to realize closed-loop control. The present invention does not need to construct an accurate mathematical model and does not need the full-dimensional state of the system. The problem of poor adaptability and long scene error accumulation of the traditional method is solved. The robustness and engineering landing performance are strong.

The application discloses a kind of aeroplane three-dimensional lift-increasing device aerodynamic and mechanism integrated optimization design method, belong to aircraft technical field, comprising: parameterization modeling is carried out to aeroplane three-dimensional lift-increasing device;According to the value range of design variable defined according to parameterization modeling, a plurality of sample points are randomly generated, and an initial optimization population is formed;According to mechanism size and flap collision interference state, sample points are marked as feasible design or infeasible design;The take-off and landing performance of the sample point marked as feasible design is calculated;The initial optimization population is optimized using a multi-objective evolutionary algorithm, with the maximum lift coefficient of the take-off configuration at a given angle of attack and the maximum lift coefficient of the landing configuration at a given angle of attack as the optimization objectives, and the mechanism size and flap collision interference state as the constraint conditions, to find the optimal design variables iteratively. The application uses the above method to ensure the feasibility of the mechanism of the three-dimensional lift-increasing device while improving the take-off and landing performance of the aircraft.

The invention provides an air-drop system intelligent control method and device based on GRPO reinforcement learning, and belongs to the technical field of air-drop system intelligent control. In the method, a reinforcement learning controller optimized by a group relative strategy is introduced, the autonomous decision-making ability, the dynamic environment adaptability and the high-precision landing performance in the air-drop process are remarkably improved, and the reliability of the air-drop system is improved. Manual intervention is not needed, a control strategy can be dynamically adjusted according to environment changes, control precision is high, and safety is good.

A system for monitoring impact on an electric aircraft is presented. The system includes a sensor communicatively connected to a flight component, wherein the sensor is configured to detect a measured force datum and generate an impact datum as a function of the measured force datum. The system further includes a computing device configured to simulate a landing performance model output as a function of the impact datum, generate a landing performance datum as a function of a comparison between the landing performance model output, determine an alert datum as a function of the landing performance datum, and display the landing performance datum on a user device.

The invention provides an aircraft performance parameter input system and method, and the system comprises a data collection module which is disposed on an installation type EFB device, and is used for collecting the aircraft system data of a flight management system of an aircraft and the environment data of a ground server; and the performance calculation module is deployed on the portable EFB equipment and the mounted EFB equipment, and is used for calculating and evaluating whether the take-off, cruise and landing performances of the airplane meet the safe operation standard or not according to the airplane system data and the environment data provided by the data acquisition module. According to the aircraft performance parameter input system and the input method thereof, the automation degree of the aircraft performance calculation module can be improved, and the consistency and the correctness of unit inspection data are helped; by applying the system and the input method, the situation that a unit spends a lot of time for data input and proofreading work can be avoided, the workload of the unit is reduced, and the possibility of human errors is reduced, so that the flight safety is improved.

The invention discloses a power lifting assisting device which is suitable for the technical field of aviation flight equipment such as unmanned aerial vehicles, manned vehicles, hovercars and flying saucers. The technical problems that existing aviation equipment is insufficient in lifting force, high in energy consumption, poor in lifting stability and low in power redundancy are solved. According to the technical scheme, the lifting assisting device comprises a lifting assisting device body shell, a lifting force auxiliary assembly, a flow guide structure, a power input interface, a fixed connecting base and a control interface; the power input interface is used for being externally connected with an airborne power source, and the lift force auxiliary assembly generates auxiliary lift force under cooperation of the flow guide structure and is in modular connection with an unmanned aerial vehicle or a manned vehicle or a hovercar body through the fixed connecting base. The invention has obvious lift force gain, and is characterized in that the lift force comes from the reaction force of fan rotation to air, the low pressure area formed by air flow above and the impact force of the air flow flowing through the turning of the pipeline, so that the take-off and landing performance and the loading capacity of the aircraft can be obviously improved, and the aircraft is widely applicable to various low-altitude flight equipment.

The invention discloses an airport runway calibration method based on autonomous take-off and landing of an unmanned aerial vehicle, and the method comprises the steps: dividing feature points on an airport runway used for autonomous take-off and landing of the unmanned aerial vehicle in advance, and then carrying out the calibration of the divided feature points through differential calibration equipment; the method comprises the following steps: firstly, calibrating the positions of two runway original points according to a reference original point and a reference coordinate original point determined by differential calibration equipment, calculating a runway course, and comparing the runway course with an airport runway course; then other feature points on the runway are calibrated in sequence, after calibration is completed, obtained runway feature point data are processed, feature point data are obtained, and therefore airport runway calibration work is completed. By adopting the calibration method, through the calibrated feature points of the airport runway and in combination with the take-off and landing performance of the unmanned aerial vehicle, the flight route planning accuracy and the take-off and landing autonomous controllability of the unmanned aerial vehicle can be improved, so that wider prospects are provided for expanding the application scene and intelligent development of the unmanned aerial vehicle, and the method is suitable for popularization and application.

The invention provides an aircraft take-off and landing performance evaluation method based on a taxiing video, which comprises the following steps: calculating the displacement, speed and acceleration of an aircraft by using the taxiing video, and predicting a take-off and off-land position or a taxiing stopping position of the aircraft in real time by using the obtained speed and acceleration. And finally, obtaining a ground takeoff skating distance LTO, a ground clearance speed VLOF, a grounding speed VTD and a total skating distance LLR during landing. Compared with the prior art, personnel on the ground can master the taxiing speed and acceleration of the aircraft, clearly know the takeoff and off-land position or the taxiing stopping position of the aircraft, quantitatively evaluate the takeoff and landing performance of the aircraft, and provide technical support for thrust evaluation of an engine and performance evaluation of a speed reduction device.

The utility model discloses a manned aircraft safety intelligent identification guiding buffer device which comprises a double-C-opening landing frame fixed to one side of the bottom end of a manned aircraft body in a bolted mode and a supporting plate arranged at the bottom end of the double-C-opening landing frame. The top end of the vertical rod extends into the double-C-opening lifting frame and is provided with a top beam, a first shock absorption spring is installed between the supporting plate and the double-C-opening lifting frame, and inverted-U-shaped frames are fixed to the two sides of the bottom of the double-C-opening lifting frame. By means of the design of the double-C-opening landing frame, the supporting plate, the shock-proof spring and the horizontal double-rod elastic buffering structure, a more comprehensive buffering effect can be provided in the landing process of the aircraft, effective elastic buffering can be provided in the Z-axis direction, buffering can be synchronously carried out in the vertical direction and the horizontal direction, and the safety of the aircraft is improved. Therefore, the landing performance of the aircraft is enhanced, and the safety and stability of the aircraft are guaranteed.

The invention discloses a distributed tilt rotor solar unmanned aerial vehicle and a design method thereof. The method may include: calculating solar radiation intensity; calculating the total weight of the unmanned aerial vehicle by combining energy balance and power balance according to the design input and the initial value of the flight weight, and iteratively optimizing the flight weight; component parameters of the distributed tilt-rotor solar unmanned aerial vehicle are designed. According to the solar unmanned aerial vehicle, short-distance take-off and landing can be achieved, and due to the good transverse stability, part of crosswind can be resisted, so that the solar unmanned aerial vehicle can have good take-off and landing performance and take-off and landing safety.

The invention relates to the technical field of engineering surveying and mapping, and discloses a planning scheme real-time optimization method and system based on engineering surveying and mapping, and the method comprises the steps: obtaining engineering surveying and mapping data and standard surveying and mapping data; performing spatial registration on actual measurement elements in the engineering surveying and mapping data and design elements in standard surveying and mapping data to obtain element pairs of the working condition; extracting a measurement feature point set from the element pairs, and calculating turning angle difference values of the actual measurement elements and the design elements at corresponding feature vertexes; measuring a direction vector and a space distance from the space gravity center of the design element to the space gravity center of the actual measurement element to obtain a point position deviation of the working condition; fusing the point position deviation and the turning angle difference value to obtain a surveying and mapping deviation result of the working condition; generating a surveying and mapping adjustment instruction of the working condition based on the surveying and mapping deviation analysis result; according to the method, the landing performance of quantitative analysis and accurate adjustment operation of terrain and element deviation in the engineering surveying and mapping field can be improved.

The invention provides an intelligent isotope labeling system with concentration prediction control capability. The intelligent isotope labeling system comprises a data acquisition subsystem, a rate prediction subsystem, a concentration target trajectory calculation subsystem, a control instruction generation subsystem and a periodic updating subsystem. According to the method, the problems that an existing method is inaccurate in speed prediction, uncontrollable in track, split in control strategy and the like are solved, and through structural design and module linkage, the system has good engineering landing performance, adaptability and sustainable optimization capacity.

This invention discloses a time-series-based method for detecting the landing performance of a standing long jump. The method decomposes the long jump motion and generates scatter plots for each decomposed motion based on the time series. The landing situation is determined based on the scatter plots, and the method identifies the frame containing the landing moment in the long jump image where the jump distance can be accurately calculated. The final long jump distance is then calculated quickly and accurately. The calculation process does not require setting complex image processing thresholds, determining the jump foot posture, or identifying skeletal key points. It can handle more landing situations, has greater versatility, is simple to calculate, has a fast processing speed, and high accuracy. Furthermore, the long jump image is processed before calculation, segmenting the human body and background, minimizing interference from the background during the calculation of the segmented image.

The invention relates to the technical field of aircrafts, and discloses a high-stability cracking type control surface ground effect aircraft and a control method thereof. Multi-dimensional active adjustment of the flight attitude is realized through independent and accurate control of the cracking type turbulent flow devices on the upper surfaces of the wings: the front / rear cracking type turbulent flow devices cooperatively act to intelligently manage the pitch attitude and the water leaving height, and the stability of the lift center is effectively maintained; the cracking type turbulent flow devices on the same side synchronously start and drive rolling and yawing integrated maneuvering, and the risk that wingtips make contact with water is avoided in combination with complete protection of a lower surface integrated structure on a ground effect area; the full-cracking type turbulent flow device is started to provide adjustable downward pressure, and the landing performance is optimized; and the structural durability, the flight stability and the safety controllability of the ground effect aircraft in a complex water surface environment are remarkably improved.

This application belongs to the field of large solar-powered unmanned aerial vehicles (UAVs), and describes an aerodynamic layout for such a UAV, including a fuselage and deformable wings. The deformable wings include outer wings and inner wings. When all the pivots on the fuselage are deflected to their maximum angle, this is the ground taxiing state of the solar-powered UAV, which facilitates taxiing and other operations. When all the pivots on the fuselage are deflected to an appropriate angle, the wing dimension of the solar-powered UAV is adjustable, which can make full use of the runway, take into account lift characteristics, and improve take-off and landing performance. This state is called the take-off and landing taxiing state of the solar-powered UAV. When all the pivots on the fuselage are deflected to their minimum angle, the overall flight length of the aircraft is controlled to the maximum, which is the aerial flight state of the solar-powered UAV. When the solar-powered UAV is in different stages of flight, it can ensure the optimal state for the current stage of flight, thereby achieving efficient flight of the solar-powered UAV.

The invention belongs to the technical field of shield launching construction, and particularly relates to a shield negative ring dismantling opportunity dynamic judgment method based on duct piece circular seam strain monitoring. Resistance strain gauges are arranged on the spliced-1 ring segment and the circular seam surface in contact with the 0 ring segment, and the resistance strain gauges are connected with a strain acquisition instrument; measuring reflectors are arranged on the inner arc surfaces of the-1 ring pipe pieces; firstly, whether the negative ring reaction frame is allowed to be dismantled to release supporting stress is judged by judging that the strain value of the resistance strain gauge is stabilized at a target threshold value; judging that the negative ring duct piece is allowed to be dismantled by judging that the daily variation of the horizontal displacement value and the elevation value of the-1 ring duct piece is smaller than a target threshold value; through'stress + deformation 'two-dimensional monitoring and quantitative threshold judgment, the construction safety is improved, the space of a starting well is released 10-15 days in advance, the reuse rate of negative ring segments is improved, the technical feasibility and the landing performance are high, and the method is suitable for various shield construction scenes.

The invention belongs to the technical field of flight performance of aircrafts, and particularly relates to a flight performance calculation method of a high-safety amphibious rotorcraft. The method comprises the steps that the flight performance of the aircraft at least comprises the cruise speed, the flight time range, the practical lift-off, the taxiing takeoff performance and the autorotation landing performance; the method is high in universality, performance calculation can be carried out on the safe amphibious rotor aircraft, and research and development investment of the amphibious rotor aircraft is effectively reduced.

The application discloses a helicopter multi-landing gear performance joint simulation method, comprising the following steps: step 1, taking the body coordinate system as a reference, the dynamics balance equation of the body is established by calculating the coordinates of each force point on the body; step 2, the numerical value of each force in the dynamics balance equation obtained in step 1 is calculated; step 3, the input quantity is given to the constructed dynamics balance equation, and iterative calculation is carried out, the whole machine landing performance calculation is completed, or the motion data results of the body in each direction are obtained. The technical scheme provided by the application takes the body as a rigid body, solves the landing load and work under the conditions of different weight gravity centers, different attitudes and different landing speeds of the body through the dynamics balance equation, so as to obtain the landing gear landing performance data, and obtains the taxiing response of the corresponding road after inputting the road spectrum.

The invention provides a method and a device for acquiring a landing load of a semi-rocker arm type undercarriage. The method comprises the following steps: acquiring basic parameters and an initial motion state of the designed semi-rocker arm type undercarriage; the basic parameters and the initial motion state are input into a preset differential equation set to be solved, and the vertical displacement of the mass center height of the undercarriage and the ground vertical load in the landing process of the undercarriage are obtained through a numerical iteration algorithm; the preset differential equation set is a mass center height change equation, a wheel fork rotation angle equation, a tire course deformation equation and a tire rolling motion equation established based on a Lagrange second-class equation; according to the vertical displacement and the ground vertical load in the landing process of the undercarriage, the energy absorption efficiency and the landing overload are obtained; and evaluating the landing performance according to the energy absorption efficiency and the landing overload. A differential equation set is derived by utilizing a Lagrange second-class equation, and the characteristics of parts such as a buffer and a tire are combined, so that the landing load is calculated, and a data basis is provided for design optimization of an undercarriage.

The present application belongs to the technical field of shield launching construction, and particularly relates to a dynamic determination method for shield negative ring dismantling opportunity based on segment ring joint strain monitoring; resistance strain gauges are arranged on the joint surface of the -1 ring segment which is in contact with the 0 ring segment after the -1 ring segment is assembled, and the resistance strain gauges are connected with a strain acquisition instrument; reflective sheets are arranged on the inner arc surface of the -1 ring segment; firstly, the strain value of the resistance strain gauge is judged to be stable at a target threshold value to determine that the negative ring reaction frame can be dismantled to release the supporting stress; then, the daily change amount of the horizontal displacement value and the elevation value of the -1 ring segment is judged to be less than a target threshold value to determine that the -1 ring segment can be dismantled; through the "stress + deformation" two-dimensional monitoring and quantitative threshold determination, the construction safety is improved, the launching shaft space is released 10-15 days in advance, the negative ring segment reuse rate is improved, the technical feasibility and landing performance are strong, and the method is suitable for various shield construction scenes.

The application provides a solar unmanned aerial vehicle gliding test method and a method for inversely calculating aerodynamic parameters. The method comprises the following steps: driving a propeller to drive a solar unmanned aerial vehicle to accelerate and run on a runway by using a motor at a preset rotating speed; when the solar unmanned aerial vehicle reaches a preset height, the rotating speed of the motor is reduced from the preset rotating speed to zero, and a theoretical sliding center line of the solar unmanned aerial vehicle is offset according to a preset lateral deviation; and the solar unmanned aerial vehicle is caused to land from the preset height until landing on the ground to test the lateral deviation landing performance of the solar unmanned aerial vehicle. The technical scheme of the application is used to solve the technical problem that it is difficult to verify the lateral deviation landing performance of the unmanned aerial vehicle in the prior art.

The invention discloses a Mars rotor take-off and landing performance test system under a near-earth effect, which belongs to the technical field of aviation test measurement and comprises a lifting mechanism, a force measuring system and a rotor mechanism. The force measuring system is installed on the lifting mechanism, and the rotor wing mechanism is installed on the force measuring system. The lifting mechanism can drive the force measuring system and the rotor wing mechanism to ascend and descend, and the force measuring system is used for measuring the lifting force and the torque when the rotor wing rotates. The test system is simple in structure, the performance of the rotor wing can be measured after rapid assembly, and the defects that an existing rotor wing test bed is limited in function, complex in mechanical structure, heavy in structural weight and complex in system, and a special environment is needed for testing are overcome.

The invention relates to the technical field of intelligent traffic and artificial intelligence crossing, in particular to a GCN multi-dimensional spatial-temporal feature OD path flow prediction method and device. According to the method, six types of multi-source data are accessed, preprocessing is carried out through a time sequence branch (cleaning, long tail processing, OD identification, flow statistics and data fusion) and a space branch (graph structure construction and matrix conversion), Conv-LSTM and an improved GCN are adopted to extract spatial and temporal features in parallel, and after Transform adaptive fusion, a full connection layer outputs the OD path flow of the next day. The device comprises a data access module, a data processing module, a feature extraction module, a model application module and a hardware support unit. According to the method, through multi-dimensional feature fusion and improved GCN adaptive variable input, the long tail effect and space sparsity problems are solved, the prediction precision is remarkably improved, generalization is high, traffic planning and resource allocation can be supported, and engineering landing performance is good.

Provided is a method for landing an aircraft, etc., in which the landing performance of a directional aircraft is improved by heading in the nose direction in which the lift and drag are in balance, based on wind speed data and wind direction data related to the landing site. In the method for landing an aircraft according to the present invention, the aircraft is configured to generate lift in response to wind from the nose direction of the fuselage, and the nose direction of the fuselage is controlled based on wind speed data and wind direction data related to the landing site, and the descent of the fuselage is started.

The application provides a helicopter field ground landing feasibility evaluation method, which comprises the following steps: determining the diameter of a helicopter landing area based on the overall length of the helicopter; wherein the landing area is a circular area; selecting three measuring points on the boundary of the landing area; measuring the ground bearing capacity at the measuring points based on a ground bearing capacity detection device to obtain test data; inputting the test data into a general ground soil layer model to obtain a landing area ground soil layer model; setting a maximum landing gear allowable settlement amount; performing ground model dynamic loading simulation by using Abaqus finite element simulation software based on tire characteristics and the landing area ground soil layer model, and calculating the ground settlement amount; if the ground settlement amount is equal to the landing gear allowable settlement amount, determining the maximum ground allowable landing load; and determining the landing state by using landing gear landing performance calculation software based on the maximum ground allowable landing load.

The invention discloses an embedded duct structure and a wing of a vertical take-off and landing aircraft. The embedded duct structure comprises a wing body, a duct, an upper surface system, a lower surface system and a power system, the upper surface system comprises an upper surface skin and a baffle; the lower surface system comprises a supporting driving module, a lower surface skin and a locking assembly. The locking assembly comprises a third driving piece, a third sliding groove, a second sliding block which is driven by the third driving piece and can slide in the third sliding groove, and a locking rod which is arranged in the third sliding groove and can penetrate through the second sliding block and the second framework. When the lock rod is inserted into the sliding block and the second framework, the lower surface skin is in a closed state and is locked and cannot rotate; and when the locking rod is separated from the sliding block and the second framework, the lower surface skin can be opened and closed. Through cooperative opening and closing of the upper surface system and the lower surface system, efficient switching of the aircraft between a fixed wing cruising mode and a multi-rotor vertical take-off and landing mode is achieved, and the vertical take-off and landing performance and aerodynamic efficiency are integrally improved.

This application belongs to the field of unmanned aerial vehicle (UAV) control technology, and provides a method and apparatus for automatic landing control of UAVs. The method includes: collecting operational state parameters and environmental disturbance data of the UAV during the automatic landing phase; determining the dynamic model and landing constraints of the UAV landing process based on the operational state parameters and environmental disturbance data; constructing a multi-objective fitness function for evaluating landing performance based on the landing constraints; globally optimizing the landing control parameters of the UAV using a particle swarm optimization algorithm, with the goal of optimizing the multi-objective fitness function, to determine the target landing control parameters; and controlling the UAV to complete the automatic landing based on the target landing control parameters. This application improves the landing accuracy and attitude stability of UAVs in disturbed environments, and achieves safe, accurate, and efficient autonomous landing of UAVs in complex scenarios such as strong winds and turbulence without relying on precise mathematical models.

The application discloses a controllable energy-absorbing CH-type corrugated stand column and an aircraft fuselage structure, and belongs to the technical field of aircraft structures. A H-type corrugated section is arranged in the middle section of the CH-type corrugated stand column, and upper and lower ends are C-type connecting sections, which are connected with a floor beam and a fuselage frame through corner pieces, respectively. The aircraft fuselage structure comprises the CH-type corrugated stand column and the fuselage frame, and a plurality of induction openings are arranged at the upper edge strips of the fuselage frame. Through the combined action of the CH-type corrugated stand column and the fuselage frame, a plurality of plastic hinges are generated at the induction openings, so that stable progressive failure of the CH-type corrugated stand column under pressure is ensured, controllable failure and energy absorption stability of the aircraft fuselage frame section are realized, the energy absorption efficiency of the structure and the landing performance are greatly improved, and the problems of uncontrollable deformation of the thin-walled structure of the aircraft fuselage lower opening section and low energy absorption efficiency are solved.

The invention provides a method and a device for calculating autonomous landing performance of an unmanned aerial vehicle. The method comprises the following steps of: 1, respectively determining a guiding process landing performance index of guiding process safety and a ship touching safety landing performance index of ship touching safety; 2, according to an unmanned aerial vehicle performance calculation starting point Dis0, a stable glide starting point Dis1, a ship wake flow disturbance starting point Dis2 and a ship tail point Dis3, stages of the guiding process are divided, and guiding process ship landing performance indexes of different stages of the guiding process are determined respectively; step 3, determining a safe carrier landing performance index of carrier touch; 4, evaluating the guide process according to the carrier landing performance index of the guide process to obtain a carrier landing performance calculation result of the guide process; evaluating the ship touch safety according to the ship touch safety landing performance index to obtain a ship touch safety performance calculation result; and 5, calculating the autonomous landing performance of the unmanned aerial vehicle through weight setting according to the landing performance calculation result of the guiding process and the landing safety performance calculation result.