43 results about "Steady flight" patented technology

The invention relates to an engineering-oriented wave rider airfoil efficient design method. The method comprises a traditional wave rider theoretical aerodynamic shape cutting method and a wave rider airfoil shape and position layout design method for a wave rider wing, a vertical tail, a pelvic fin and other airfoils. Aiming at the problems that the traditional wave rider theoretical appearance is only suitable for a single design point and the stability and the controllability are insufficient during cross-speed-domain flight, a system solution is provided for cross-speed-domain stable flight of the hypersonic flight vehicle. According to the method, the design technology breakthrough is realized by combining the parameterized cutting technology of the traditional wave rider theoretical appearance and the multi-airfoil collaborative optimization design. Through a global parameter joint debugging algorithm, after airfoil design is completed, the aerodynamic center position variation of the waverider under different Mach numbers is reduced, trimming is easy, and engineering adaptability design of a theoretical shape is achieved.

The invention relates to the technical field of unmanned aerial vehicle control, and discloses an unmanned aerial vehicle rotor pulling force-based load weight measurement method, which comprises the steps of obtaining a plurality of original lift force measurement values generated by each rotor in a set sampling period when an unmanned aerial vehicle is in a stable flight state, and forming an original lift force value set corresponding to each rotor; performing data cleaning and validity verification on each lift original value set, and calculating to obtain an initial lift value of each rotor after removing abnormal values; and calculating a global lift reference value according to the initial lift values of all the rotors. According to the method, multi-sensor data can be effectively fused, and the influence of factors such as flight attitude and aerodynamic change is dynamically compensated in calculation, so that stable and reliable weight estimation can be provided when the unmanned aerial vehicle hovers, cruises and even executes certain maneuvering actions, and key data support is provided for flight control, endurance evaluation and task planning.

Disclosed herein are a vertical take-off and landing aircraft, a maneuvering method of the same, and a control apparatus for the same. The vertical take-off and landing aircraft includes: a fuselage; a lift unit including a first propeller disposed at a periphery of the fuselage to provide lift to the fuselage and a duct protecting the first propeller; and a thrust unit including a second propeller disposed at a rear end of the fuselage to provide thrust to the fuselage. Here, the duct includes: a propeller guard portion; a leading airfoil portion; a trailing airfoil portion; and a flow guide portion. The propeller guard portion is convexly formed toward the first propeller to have a central diameter smaller than a diameter of an inlet for air to flow in therethrough and a diameter of an outlet for air to flow out therethrough. The leading airfoil portion is disposed in a front region of the duct colliding with air during forward flight and has a shape corresponding to a shape of a front portion of an airfoil including a leading edge. The trailing airfoil portion is disposed in a rear region of the duct designed to allow air to flow thereover during forward flight and has a shape corresponding to a shape of a rear portion of the airfoil including a trailing edge. The flow guide portion protrudes from the propeller guard portion toward the first propeller and suppresses vortex generation at a tip of the first propeller upon rotation of the first propeller. The vertical take-off and landing aircraft according to the present invention can achieve high performance in terms of high-speed flight and stable flight while having excellent maneuverability and efficiency.

The invention discloses a wing body fusion layout vertical aircraft with T-shaped tilting power distribution. The vertical take-off and landing fixed-wing aircraft is mainly applied to a power system and aerodynamic layout design of the vertical take-off and landing fixed-wing aircraft, T-shaped power distribution is formed through cooperation of the wingtip tilting rotors and the tail ducted fans, and the composite power layout form of vertical take-off and landing, hovering, transition and horizontal cruise multi-working-condition stable flight is achieved. The technology can be widely applied to the fields of unmanned aerial vehicles, general aviation aircrafts, special task platforms and the like, and is suitable for efficient flight tasks in surveying and mapping inspection, logistics transportation, emergency rescue and complex terrain environments.

The invention discloses an aircraft control method and device, a storage medium and a program product, is suitable for the technical field of unmanned aerial vehicle control, and provides an unmanned aerial vehicle control party switching method, and the method comprises the steps: determining that an aircraft meets a control right switching condition; sending a control right switching instruction to the aircraft service platform, wherein the control right switching instruction is used for requesting the aircraft service platform to take over control of the aircraft; receiving a confirmation response sent by the aircraft service platform for the control right switching instruction; and sending a preset action execution instruction to the aircraft, wherein the preset action execution instruction is used for instructing the aircraft to execute a preset action. Through cooperative operation of the control party, the aircraft service platform and the aircraft, in the process of switching the control right, the control party sends instructions such as hovering, hovering or flying along a preset path to the aircraft, so that the aircraft automatically keeps a stable flight state in a switching gap; therefore, the out-of-control risk caused by interruption of the control instruction is thoroughly eliminated.

The invention discloses a flight unit recovery device and method based on a three-ring separation structure and a flight unit.The flight unit recovery device based on the three-ring separation structure is characterized in that a parachute cabin opening mechanism is mounted at a first set position of the flight unit and used for loading a parachute recovery system; the parachute recovery system is arranged in the parachute opening cabin mechanism and used for controlling the main parachute to pop out based on the three-ring separation structure, stabilizing the attitude of the flight unit and reducing the speed, and a sling structure in the parachute recovery system is used for bearing force and maintaining the relative attitude of the flight unit and the parachute system; the airbag cabin opening mechanism is mounted at a second set position of the flight unit and is used for loading an airbag system; the air bag system is arranged in the air bag cabin opening mechanism and used for buffering when the flight unit falls to the ground. The flying unit is stably decelerated through the parachute recovery system, the parachute opening stability is achieved through the three-ring separation structure and the sling structure, and then stable recovery of the flying unit is guaranteed through the air bag system.

The application provides a stability enhancement system for aircraft lateral layered control, belongs to the technical field of automatic flight control, and relates to the technical field of automatic flight control.The system meets the design standard of a civil aviation IIIa control system.The system adopts a layered control architecture combining outer loop PID track tracking control and inner loop lateral heading model predictive control (MPC).The outer loop PID control is responsible for basic track tracking, and the inner loop MPC finely adjusts the heading angle and the track to ensure that the aircraft accurately aligns with the runway and stably flies.The method not only improves the control precision and response speed, but also significantly reduces the calculation burden, enhances the robustness and anti-interference ability of the system by combining the simplicity of the PID control and the multivariable optimization capability of the MPC.The system can effectively improve the stability of the aircraft lateral control under complex flight environment, and ensures the safety and reliability of the aircraft during the automatic landing task.

This invention belongs to the field of reusable aircraft control technology, and relates to an intelligent fault diagnosis and control method for reusable aircraft. This method establishes a control-oriented aircraft dynamics model, integrates a neural network adaptive fault observer and a particle swarm optimization algorithm, to achieve high-precision tracking control of the aircraft's attitude and trajectory. When faced with faults such as actuator performance loss and unmeasurable deviations, it can quickly and accurately estimate fault information and integrate the reconstructed fault information into the control law design, ensuring stable flight even under severe fault conditions. Simultaneously, this method can effectively cope with complex disturbances such as aerodynamic uncertainties, elastic modal coupling, and external environmental interference. Through high-precision joint online estimation, it improves the real-time performance and accuracy of fault diagnosis. Furthermore, the particle swarm optimization algorithm further enhances the overall performance and adaptive capability of the control system, making the aircraft operate more efficiently and stably during flight.

The application provides a digital side-slip ball calculation method, and solves the problem of low accuracy of a current side-slip instrument in reflecting the side-slip condition of an aircraft, and comprises the following steps: step one, obtaining flight state information and flight parameter information of the aircraft; step two, obtaining parameter information of the digital side-slip ball; and step three, judging whether the aircraft is in steady flight or dynamic flight according to the flight state information, and selecting different parameter information to calculate corresponding digital side-slip angles in different stages through a dynamic equation of the digital side-slip ball under different flight states. The application establishes the dynamic equation of the digital side-slip ball under different flight states of the aircraft corresponding to the mechanical side-slip ball, calculates the digital side-slip angle corresponding to different flight states of the aircraft, and the calculated digital side-slip angle is basically consistent with the side-slip angle of the aircraft. The digital side-slip ball calculation of the application meets the accuracy requirement of the side-slip condition of the aircraft.

The invention belongs to the field of unmanned aerial vehicles, and provides a flight attitude compensation method and system for dynamic environment prediction. The flight attitude compensation method for dynamic environment prediction comprises the following steps: processing a standard data set by using a graph neural network model to obtain a spatial-temporal characteristic graph of an environment disturbance field; modeling the environmental disturbance field through topology invariance, representing the spatial-temporal characteristic pattern of the environmental disturbance field as topology invariant characteristics, and obtaining predicted distribution of the environmental disturbance field; calculating the required attitude compensation amount according to the predicted distribution of the environmental disturbance field and the attitude demand of the aircraft, and adjusting the attitude control input of the aircraft according to the attitude compensation amount to obtain a corresponding control signal so as to realize attitude compensation; and in the flight process, the predicted distribution of the environment disturbance field is updated in real time, and the attitude compensation amount of the aircraft is dynamically adjusted through a self-adaptive control strategy according to the predicted distribution so as to stabilize the flight attitude of the aircraft. The method can effectively guarantee the precision and robustness of aircraft attitude compensation in a complex dynamic environment.

The stabilizing surfaces of darts (flights) are designed to stabilize the darts in flight and enable accurate trajectory. If these stabilizing surfaces are damaged or bent, the geometry and aerodynamics are compromised, negatively impacting the dart's flight path. The invention achieves a rotating motion around the longitudinal axis of a dart in flight. This rotation creates a gyroscopic effect, which stabilizes the dart's trajectory. The rotation improves the dart's penetration into the target and reduces ricochets. The rotational movement allows the thrown darts to glide more smoothly past the darts already stuck on the dartboard. The rotational effect is achieved solely through deformation, without weakening the stabilizing surfaces with incisions or similar modifications. These deformations make the stabilizing surfaces stiffer, more dimensionally stable, and therefore more durable. The shape is created by bending the stabilizing surfaces (1) at the bending paths (2), thereby creating effective surfaces (3) that generate a rotational movement around the longitudinal axis during flight.

The application relates to a variant airplane steady flight deceleration maneuver decision method based on reinforcement learning and belongs to the technical field of aviation. The application comprises the following steps: establishing a kinematic model and a dynamic model of a variant airplane; establishing an aerodynamic force and an aerodynamic moment model of the variant airplane and considering the aerodynamic force and the aerodynamic moment change caused by the variant; designing an overall control structure of the steady flight deceleration maneuver, using PI control on the speed, using backstepping control on the attitude angle to keep the attitude stable, using PID control on the height in the outer loop of the pitch angle to keep the height unchanged; designing a reward and punishment function of the steady flight deceleration maneuver; and designing an intelligent agent based on a reinforcement learning DQN algorithm, discretizing the possible degrees of the left and right sweep angles and the V-tail upper anti-angle to adapt to the discrete action space of the DQN algorithm. The application uses the reinforcement learning control method to ensure that the airplane can achieve optimal aerodynamic force, flight efficiency and control efficiency when performing a specific maneuver.

The invention relates to bridge pier and bridge inspection, in particular to a bridge pier and bridge inspection system suitable for ultra-low-altitude flight of an unmanned aerial vehicle, an ultra-low-altitude flight path planning module, a Beidou grid code technology is adopted to divide an inspection area, an optimal flight path is generated, and the flight safety of the unmanned aerial vehicle below the side of a bridge pier, the side face of a box girder and the bottom of the girder is ensured; the airspace self-adaptive module is used for dynamically adjusting the flight height and the flight route of the unmanned aerial vehicle based on the real-time meteorological data and the bridge structure characteristics so as to adapt to the inspection requirements in different environments; the unmanned aerial vehicle flight control module performs multi-source data acquisition through multi-sensor fusion, ensures stable flight of the unmanned aerial vehicle in a narrow space by utilizing cloud side end cooperative control, and has an emergency flight mode; according to the technical scheme provided by the invention, the defects that the flight safety of the unmanned aerial vehicle in a complex environment is difficult to ensure, the flight stability in a narrow space is insufficient, and the accuracy and efficiency of disease identification need to be improved in the prior art can be overcome.

The application discloses a method, device and equipment for preventing geomagnetic interference of a UAV (Unmanned Aerial Vehicle) and a storage medium, and the method comprises the following steps: collecting sensor data through a sensor component configured on the UAV; obtaining a flight state variable of the UAV according to the sensor data; determining whether there is geomagnetic interference through a magnetic field anomaly monitoring algorithm based on the flight state variable; and if it is determined that there is geomagnetic interference, resetting the heading of the flight state variable of the UAV through a resetting strategy to make the UAV get rid of the geomagnetic interference. The application can accurately identify the geomagnetic interference condition in the flight process of the UAV, and reset the heading in the flight state variable of the UAV when detecting the interference, so that the deviation caused by the geomagnetic interference on the heading of the UAV is effectively overcome, the stability of the flight direction of the UAV is greatly improved, and the UAV can also keep steady flight in various complex electromagnetic environments.

The invention relates to an automobile flying container type aircraft and a control method thereof, belongs to the technical field of structural design of container type aircrafts for air transportation, and aims to solve the problems that an existing flying automobile scheme focuses on automobile self-transformation, is high in cost and technical difficulty and is not suitable for common automobiles, and the cost is low. And the problems that ground vehicles are not effectively and quickly converted into flight modes, and the practicability and the universality are limited are solved. The container type aircraft comprises a bearing fixing module, a power adjusting module, an energy supply module, a control navigation module and a safety guarantee module and can load and fix a common automobile, provide flight power and energy and achieve automatic flight and safety protection, and the control method covers the whole process of loading, lift-off, cruising and unloading. A common automobile can be conveniently loaded and fixed, stable flight power and emergency power supply are provided, automatic flight is achieved, a parachute and an escape channel can be triggered in danger, and safety is guaranteed.

This invention relates to an online identification method and medium for the center of gravity of an aircraft, as well as a control law gain scheduling method. It pertains to the field of aerospace technology. Under stable flight conditions, a preset multi-channel sinusoidal orthogonal excitation signal is input, and the aircraft's state variables, control variables, trim parameters, and flight state parameters are collected. Based on an augmented extended Kalman filter algorithm, the closed-loop system of the aircraft is identified online to obtain the stability matrix and control matrix, thereby obtaining the aircraft's longitudinal short-period modes. A subset of conditions closest to the obtained real-time flight state parameters is matched in a priori feature database. Within this subset, the real-time feature vectors are matched with feature vectors in the priori feature database to find the optimal match, thus obtaining an online estimate of the current center of gravity. This solution eliminates the need for dedicated sensors and fully utilizes existing flight control system hardware resources to identify the aircraft's center of gravity online, in real-time, and seamlessly during flight.

The invention belongs to the technical field of unmanned aerial vehicle systems, and particularly relates to a combined type large-load multi-rotor unmanned aerial vehicle system, which comprises a sensing module, a multi-rotor unmanned aerial vehicle control module and a multi-rotor unmanned aerial vehicle control module, the control module is used for performing algorithm processing on the data acquired by the sensing module to generate a corresponding control instruction; and the execution module is used for driving a motor and an electronic speed controller according to the control instruction to adjust the rotating speed and attitude of a propeller so as to realize flight control of the unmanned aerial vehicle. According to the invention, lift force and gravity unbalance caused by load increase can be avoided, the problem of unstable flight caused by external interference is effectively prevented, stable flight can be ensured under the condition of large load, and the load capacity and the flight efficiency are improved to a certain extent.

The invention belongs to the field of aircraft transportation, and particularly relates to a carrying platform for a low-altitude aircraft and a using method thereof.The carrying platform comprises a platform body, a guide plate, a base and a control device, unit holes for mounting balancing weights are formed in the platform body, and a plurality of display units are arranged on the guide plate; a weight sensor is arranged in the base and used for detecting weight data of the platform body. And the control device is used for determining an offset coordinate of the overall gravity center of the platform main body relative to the base coordinate origin O according to received weight data detected by the weight sensor, and controlling the display unit to display according to the determined offset coordinate so as to guide workers to load goods and / or balancing weights. According to the carrying platform, the overall gravity center can be simply and conveniently adjusted, and the stable flight and hovering capacity of an aircraft in the transportation process is improved.

The stabilizing surfaces of darts (flights) are designed to stabilize the darts in flight and enable accurate trajectory. If these stabilizing surfaces are damaged or bent, the geometry and aerodynamics are compromised, negatively impacting the dart's flight path. The invention achieves this by imparting a rotating motion around the longitudinal axis to a dart in flight. This rotation generates a gyroscopic effect, which stabilizes the flight path. The rotation facilitates better penetration of the dart into the target and reduces ricochets. Furthermore, the rotational movement allows the thrown darts to glide more smoothly past those already embedded in the dartboard. This rotational effect is achieved solely through deformation, without weakening the stabilizing surfaces with cuts or similar modifications.The deformations make the stabilizing surfaces stiffer, more dimensionally stable, and therefore more durable. The shape is created by bending the stabilizing surfaces (1) along the bending paths (2), thereby creating effective surfaces (3) that generate a rotational movement around the longitudinal axis during flight.

A deployable emergency transmitter device is configured to be launched from an aircraft in an initial passive flight state. The device includes an internal chassis, and an inertial mass member, rotatable relative to the internal chassis, for stabilising flight during the initial passive flight state. The device also includes a powered drive system, mounted to the internal chassis and configured for providing lift while the deployable emergency transmitter device is in an active flight state and a radio transmitter configured for transmitting an emergency signal after being launched from the aircraft. The device is configured to transition from the initial passive flight state to the active flight state, after being launched from the aircraft.

To provide a system for controlling an aircraft to enable stable flight when the aircraft is tethered by a cable. [Solution] The system of the present invention comprises a mobile body, a flying body connected to the mobile body via a cable, a control unit for controlling the flight of the flying body, and a tension sensor for detecting the tension T applied to the cable and outputting the detected tension T to the control unit. The control unit is configured to determine whether the detected tension T is greater than or equal to a threshold, and, in response to determining that the detected tension T is greater than or equal to the threshold, to control the flight of the flying body so that the detected tension T becomes less than the threshold.

The present application relates to the technical field of unmanned aerial vehicle control, and particularly relates to a I-shaped variant rotor aircraft and a dynamic reconfiguration control method, when flight environment parameters indicate that the aircraft encounters a narrow gap, the aircraft starts a configuration transformation mode, adjusts the included angle of two sections of the arm compared with the fuselage through a configuration rudder, and makes the two sections of the arm mirror and synchronously rotate along the horizontal normal direction of the fuselage at the same angular velocity, and the aircraft always keeps parallel to the ground during the configuration transformation process; the aircraft adjusts the flight configuration to a high-passability configuration, the differential of the brushless motor keeps the ability provided by the pitch rotation torque unchanged, the angle of the tilt rudder reaches 25 DEG, and the tilt is executed to provide the roll and yaw rotation torque. The beneficial effects of the present application include the reduction of the body size, the satisfaction of the accuracy and real-time requirement of the flight control on the body gravity center calculation, the simple flight control of the high-passability configuration, the full-cycle continuous mounting operation, and the strong omni-directional motion ability.

The application discloses a design method of an acceleration feedback autopilot based on a static unstable missile, and comprises the following steps: establishing a longitudinal linear dynamic equation; constructing an acceleration autopilot framework; setting acceleration autopilot control parameters to obtain an autopilot; and controlling the flight state of a flight vehicle by using the autopilot, wherein the acceleration autopilot framework is a three-loop autopilot structure, comprising a damping stable inner loop, a stability enhancement loop and an acceleration loop outer loop, the feedback signal of the damping stable inner loop is generated based on a pitch angular velocity, the feedback signal of the stability enhancement loop is generated based on a longitudinal acceleration, and the feedback signal of the acceleration loop outer loop is generated based on the longitudinal acceleration. The method disclosed by the application can stabilize and control a static unstable flight vehicle, and significantly improves the overload capacity of the static unstable flight vehicle.

The invention provides a flutter flight test method for a troposphere manned airship, and the method comprises the steps: taking an airship VD as a reference, and selecting a test speed point; the airship reaches a preset height and keeps stable flight; sending an instruction through ground command, enabling the airship to enter a test flight state point, balancing the airship through an auxiliary air bag, and keeping 80% VD constant-speed stable flight for first preset time; after completing a test flight state, the test flight person reports the airship response condition to a ground commander, if no abnormal condition exists, the airship continues to fly at the 90% VD speed point, if no abnormal condition exists, the airship continues to fly at the 100% VD speed point, and if abnormal conditions exist, test flight is stopped immediately; after the steps are completed, a shaking rod excitation flight test is carried out; a command is given out through ground command, the airship enters an 80% VD speed point, and the airship is trimmed through an auxiliary air bag to stably fly at a constant speed for a second preset time; the pilot implements rod shaking excitation, the airship continues to stably fly for a third preset time after responding, and the rod shaking process is repeated till three times of ending; the application fills the blank of the troposphere manned airship flutter flight test method.

The invention discloses a composite helicopter stable flight control method, which relates to the technical field of aviation flight control, and comprises the following steps: S100, synchronously acquiring a blade vibration trace, a helicopter body pitching swing trace and a tail section acceleration trace under a high-speed cruise steering working condition, pressing into a coupling time sequence band by using a unified time reference, and obtaining a coupling time sequence band; and continuously traceable resonance candidate segments are marked in the coupling time sequence band. According to the method, a multi-control-chain peak shifting admission mechanism is established by triggering an anchor point cable, time layering coordination of propeller pitch adjustment, pitching balancing and propelling thrust is achieved, modal energy superposition of rotors and an aircraft body is avoided, and the flight attitude is kept stable; and through breathing type phase turn-back gate control driven by the damping bill chain, dynamic circulation of energy absorption and recovery is realized, vibration accumulation is inhibited, flight immunity and structural stability are improved, and stable flight is guaranteed.

The invention discloses a single main rotor-partially inclined multi-slave rotor heterogeneous multi-rotor layout, and relates to the technical field of unmanned aerial vehicle aircraft design, the single main rotor-partially inclined multi-slave rotor heterogeneous multi-rotor layout comprises a large-size main rotor providing main lift force and n control rotors arranged around the large-size main rotor, part of the control rotors are horizontally mounted, and the rest are obliquely mounted at a fixed angle; during steady-state flight, the reverse torsional moment of the main rotor is statically balanced by the horizontal component moment of the pulling force of the inclined slave rotor, attitude control is carried out on the unmanned aerial vehicle by carrying out differential adjustment on the rotating speed of the control rotor, and attitude control comprises pitching control, rolling control and yaw control. Wherein the pitching control and the rolling control are realized by differentially adjusting the rotating speed of the control rotors at symmetrical positions, and the yaw control is realized by differentially adjusting the rotating speed between the horizontally mounted control rotor group and the obliquely mounted control rotor group; according to the invention, the integration of high aerodynamic efficiency and simple control is realized, and the cruising ability and flight performance of the unmanned aerial vehicle are effectively improved.

The application discloses a spinning angular velocity selection method of an aircraft based on a dynamic stability criterion and belongs to the technical field of aircraft control. The method is as follows: a dynamic equation of the aircraft is established by considering the introduction of a Magnus moment and gyroscopic effect under active spinning conditions; a general dynamic stability criterion of the spinning aircraft is obtained based on the dynamic equation of the aircraft and in combination with the Routh-Hurwitz method; and the dynamic stability criterion of the classical trajectory design and the coning motion is fused to select a corresponding spinning angular velocity, so that the aircraft stably flies in the spinning process. The application avoids the problems of attitude divergence and guidance precision attenuation caused by the out-of-bound coning motion parameters in the traditional design, improves the trajectory design precision, ensures that the aircraft keeps the gradual stability of the coning motion mode in a complex aerodynamic environment, reduces the attitude angle control error of the traditional method, and provides a systematic solution for the high-precision guidance and stable flight of the spinning aircraft.

The large lift and thrust enable stable flight, allowing for continued flight or a safe landing if one propeller stops rotating during flight, enabling the vehicle to land on a road with traffic and drive away, and stabilizing the lateral attitude even in the face of sudden crosswinds. [Solution] The amphibious mobile vehicle 1 comprises a mobile vehicle body 10, front wheels 31, 32 and rear wheels 33, 34 for driving, located in front of and behind the crew compartment 11 of the mobile vehicle body, and a pair of upper front propellers 421 and lower front propellers 441, and a pair of upper rear propellers 461 and lower rear propellers 481 for flight, located in the front and rear arrangement spaces of the mobile vehicle body. The front wheels, which are driven by electric motors, and the inner rotation shafts 411, 451 and outer rotation shafts 431, 471 of the propellers, which rotate in opposite directions to generate lift and thrust by electric motors 231, 241, 251, 261, are arranged parallel and coaxially with the mobile vehicle body in the vertical direction.