82 results about "Steady flight" patented technology

The invention relates to a position attitude control method of a four-rotor unmanned aerial vehicle with unbalance loads, and belongs to the technical field of control of unmanned aerial vehicles, which aims at solving the problem of the unmanned aerial vehicle failing to continue to stably run due to unknown center of gravity and uncertainty in installation position in the process of loading unbalanced load. The position attitude control method comprises the following steps of firstly, deriving a kinematics model and a dynamics model before and after the unbalance load is loaded on the four-rotor unmanned aerial vehicle by coordinate conversion and dynamics analysis; obtaining the measuring data of a sensor in the flight process, utilizing a Kalman filtering method to obtain the high-precision and non-lagging flight state, and distinguishing the approximate location of center of gravity by the flight states; designing a position attitude controller and a position controller of the four-rotor unmanned aerial vehicle, and improving the controller by compensating the rotation speed of a propeller, so as to offset the additional rotary movement and linear movement due to location change of center of gravity. The position attitude control method is suitable for estimating the location of the center of gravity in the takeoff and hovering process of the four-rotor unmanned aerial vehicle, and establishing a compensation controller.

The invention provides a steady flight control method of a multi-rotor unmanned aerial vehicle. The method comprises the following steps: 1), obtaining flight real-time operation data of the aerial vehicle through an airborne sensor, performing corresponding analytical processing on kinematic problems of the aerial vehicle through an airborne processor, and establishing an aerial vehicle dynamical model; 2), designing a variable parameter convergence differential neural network solver of the multi-rotor aerial vehicle dynamical model according to a variable parameter convergence differential neural dynamical design method; 3), solving the output control variable of an aerial vehicle motor through the variable parameter convergence differential neural network solver by utilizing obtained flight real-time operation data and target attitude data of the aerial vehicle; 4), transmitting solving results of the third step to an aerial vehicle motor speed regulator to control the motion of the unmanned aerial vehicle. The method is based on the variable parameter convergence differential neural dynamical method, can rapidly and accurately approach the correct resolution of the problems in real time, and can properly solve various time dependent problems, such as matrixes, vectors, algebra and optimization.

A method of optimizing the operation of left and right propellers disposed on either side of the fuselage of a rotorcraft including a main rotor. Left and right aerodynamic surfaces include respective left and right flaps suitable for being deflected, yaw stabilization of the rotorcraft being achieved via first and second pitches respectively of the left and right propellers, and the deflection angles of the left and right flaps are adjusted solely during predetermined stages of flight in order to minimize a differential pitch of the left and right propellers so as to optimize the operation of the left and right propellers, the predetermined stages of flight including stages of flight at low speed performed at an indicated air speed (IAS) of the rotorcraft that is below a predetermined threshold, and stages of yaw-stabilized flight at high speed performed at an indicated air speed of the rotorcraft greater than the predetermined threshold.

The presently disclosed embodiments relate to vertical takeoff and landing (VTOL) aircraft that have the capability of hovering in both a “nose forward” and a “nose up” orientation, and any orientation between those two. The disclosed aircraft can also transition into wing born (non-hovering) flight from any of the hovering orientations. In addition, certain of the disclosed embodiments can, if desired, use only vectored thrust control to maintain stable flight in both hover and forward flight. No control surfaces (e.g. ailerons, elevators, rudders, flaps) are required to maintain a stable vehicle attitude. However, the disclosure contemplates aircraft both with and without such control surfaces.

The present invention relates to the field of methods and systems for optimizing the locating of an aircraft in airports and, more particularly, in the take-off and landing phases, and notably in the air-ground and ground-air transition phases. The method of optimizing the locating of an aircraft during the take-off or landing phase comprising a transition step (21), the latter being defined between two events (A0, AS), the first event being a condition of contact between the aircraft and the runway and the second event being a threshold condition defining a stabilized flight phase, is wherein the transition step (21) comprises the determination of at least one “transition position” of the aircraft by a weighting between the “ground position” (LOC_SOL) determined by the ground locating system and the “flight position” (LOV_VOL) determined by the in-flight locating system.

Embodiments are directed to obtaining at least one measurement of blade bending associated with at least one blade of at least one rotor of an aircraft, processing, by a device comprising a processor, the at least one measurement to obtain moment data, obtaining data from an inertial measurement unit (IMU), and processing, by the device, the moment data and the data from the IMU to generate a command configured to stabilize the aircraft.

The invention relates to an estimation and self-healing control method for a blade injury of a multi-rotor unmanned aerial vehicle autopilot. The estimation and self-healing control method comprises the steps of: firstly, calculating a moment acted on a machine body by an instruction output by an autopilot control system under a non-fault condition and a mathematical model between a force and moment which are acted on the machine body; then constructing a machine body attitude control system state equation, and estimating a moment disturbance variable generated for the machine body after the blade injury by adopting a time domain disturbance observer; then inverting and resolving a parameter estimated value of a blade injury model according to an inverse model of the blade injury mathematical model; and finally, according to the parameter estimated value of the blade injury model, designing a control distribution reconstruction matrix to implement control distribution reconstruction ofthe autopilot. Compared with the prior art, the estimation and self-healing control method for the blade injury of the multi-rotor unmanned aerial vehicle autopilot, which is disclosed by the invention, has the following advantages that a blade injury degree can be estimated online; and by reconstructing a control object mathematical model online, control self-healing of the multi-rotor unmannedaerial vehicle autopilot under the blade injury is implemented, so as to ensure safe and stable flight of an unmanned aerial vehicle under the blade injury condition.

The embodiment of the invention relates to the technical field of flight simulation, in particular to a flight simulation system, a flight simulation method and a storage medium, which simulate the situation of moment imbalance caused by aerodynamic configuration and speed change in the flight process of an aircraft by simulating aerodynamic force and engine acting force received by the aircraft during flight. An automatic balancing unit is additionally arranged so that a balancing system in the real flight process is simulated, wherein flight balancing is conducted on the aircraft through the automatic balancing unit, in other words, an operation face (ailerons, elevators and rudders) is finely adjusted, unbalanced moment and rod force in the steady state are eliminated, and the purpose of stabilizing the posture and heading of the aircraft is achieved. Therefore, the whole system of the aircraft can be restored more truly, and the flight state of the aircraft can be completely simulated, so that the state and operation data of the aircraft in flight simulation training are more suitable for a real flight scene.

The invention provides a high-stability control method for continuous high-overload maneuver of a supersonic high-maneuver target, which is a high-stability control method for controlling the target of a pitching and rolling channel only by an elevator and an aileron in the process of continuous high-overload maneuver in a horizontal plane. According to the method, firstly, overload nHc in the y direction of a transmitting system required by target height stability control is analyzed, under the condition that a preset overload instruction nyc is non-adjustable, in order to ensure that the height is kept stable in the horizontal maneuvering process, an online roll angle instruction planning algorithm is designed, and according to the algorithm, a roll angle instruction is planned in real time, so that the target height stability is guaranteed. Therefore, the component of the predetermined overload instruction nyc in the y direction of the launch system is equal to the highly stable required overload nHc, and the highly stable target in the flying process of the horizontal maneuver process is ensured. Flight test results prove the effectiveness of the method.

The invention provides a low-altitude aircraft which can be used as an unmanned aerial vehicle and can also be used as a manned craft. The aircraft is mainly characterized in that the aircraft is a large or ultra-large air craft formed by a plurality of multi-axis aircraft bodies and is uniformly controlled by a controller. Each multi-axis aircraft body is an independent unit and can be used for normal flight, and the multi-axis aircraft bodies do not interfere with one another after being assembled to a main body rack; the main body rack is further provided with one or more balance rods which are used for maintaining steady flight; besides, a safety umbrella is installed above the main body rack and used for maintaining balance and safety of landing; foot stools are arranged below the main body rack and used for supporting an aircraft body and play a role of buffering when the aircraft descends and lands on the ground. The low-altitude aircraft is easy to install and operate and capable of taking off and landing safely and flying steadily and has better application prospect.

The invention relates to an intelligent control four-axis aircraft with multiple cameras. The aircraft comprises a sensor 5 and the cameras A, B, C and D, and includes four propellers which are arranged in criss-cross intersection. A flight power system of independent research and development is integrated in the aircraft, the aircraft is configured with a professional radio remote control system, and the aircraft is capable autonomous navigation flight. Four rotors are mounted in four corners of the aircraft respectively, and driven by a motor, blades can rotate forwardly and backwardly, an inertial navigation module composed of gyroscopes of three directions and a three-axis acceleration sensor is mounted on the four-axis aircraft to maintain stable flight of the aircraft, and an electronic regulator is used to ensure rapid flight of the aircraft. The four cameras are respectively mounted on the four rotors, and connected with the sensor at the center of the aircraft to transmit information and monitor a moving object. According to the four-axis aircraft, monitoring is rapider and more comprehensive, and the monitoring efficiency is improved.

The invention discloses a four-rotor aircraft fault tolerance control method based on a switching adaptive algorithm, belongs to the field of aircraft control, and is used for solving the problem thata high-maneuverability four-rotor aircraft cannot well track an expected signal under the conditions that an input fault exists and kinetic parameters are unknown. The method is technically characterized by comprising the following steps of constructing a four-rotor aircraft segmented ray-imitating linear system containing unknown input faults and unknown kinetic parameters, a reference system and a controller; obtaining an error system model according to the segmented radioactive system, the reference system and the controller; designing a residence time constraint-based switching signal according to the segmented radioactive system and the reference system; and obtaining an adaptive law of control parameters in the controller according to the error system model and the switching signal.The method can achieve the good tracking performance of the high-maneuverability four-rotor aircraft for an expected signal, and can be used for the flight control of the four-rotor aircraft so as toguarantee the stable flight of the four-rotor aircraft under the conditions that an input fault and a kinetic parameter are unknown.

The invention discloses a helicopter with an adjustable empennage attitude, relates to the technical field of traffic tools, and aims to solve the problem of low landing and take-off safety of a conventional helicopter. An empennage of the helicopter with the adjustable empennage attitude comprises a first empennage part fixedly connected with a helicopter body and a second empennage part rotatably connected with the first empennage part; an attitude adjustment driving motor is used for driving the second empennage part to rotate, so that a main rotor wing and a tail rotor wing are positioned on the same horizontal plane, and influence, which is caused by high-speed rotation of the main rotor wing, on the tail rotor wing is avoided; a steady flight of the helicopter is guaranteed, and the service life of the helicopter is effectively prolonged; furthermore, a main engine is connected with the attitude adjustment driving motor through a generator and a processing conversion device; the processing conversion device is also connected with a storage battery; electric energy generated by the main engine can be stored in the storage battery; when the main engine has fault, the electric energy stored in the storage battery can be used for driving the second empennage part to adjust the attitude and stabilize the flying state, so that a pilot has enough time to escape.

The invention discloses a foldable variable-structure unmanned aerial vehicle which comprises a fuselage, two pairs of wings rotatably mounted on the fuselage, a front propeller mounted at the head of the fuselage, an empennage mounted at the tail of the fuselage and ducted fans arranged on the wings. The wings can be driven to rotate to be approximately parallel to the fuselage to form a folded state or approximately perpendicular to the fuselage to form an unfolded state or obliquely intersect with the fuselage to form a semi-unfolded state. According to the unmanned aerial vehicle, the landing mode and the flight mode can be switched in a self-adaptive mode according to the terrain, the unmanned aerial vehicle can adapt to complex terrains, stable flight in various application scenes can be achieved, and operation is more flexible; when the wings of the structure are folded, the wings are horizontally stored to the two sides of the aircraft body, and the landing occupied area of the whole aircraft can be greatly reduced; the unmanned aerial vehicle with the structure is relatively simple in structure, lightweight design is favorably realized, energy consumption is reduced, the fuselage size of the unmanned aerial vehicle is favorably reduced, and a high space utilization rate is realized.

Improved systems and methods for dynamic readouts on a primary flight display are provided. The system includes a computer system configured to receive and process aircraft status data and flight plan data to command a display system to overlay on the image, (i) an analog indicator of the airspeed, (ii) an analog indicator of altitude, (iii) a digital airspeed indicator, and (iv) a digital altitude indicator. The system continuously determines whether steady flight conditions are occurring. Steady flight conditions include one or both of: constant airspeed and constant rate of change of airspeed. When the steady flight conditions are occurring, (i) and (ii) are removed from the image. When the steady flight conditions are no longer occurring, (i) and (ii) again overlaid on the image.

This invention can take-off like a helicopter and then roll forward as a single unit to fly faster and further than any conventional helicopter. Just like the seats in a Ferris wheel, the payload of passengers or freight can remain upright and comfortable as this aircraft rotates from take-off to fast forward and back to landing mode. Prior inventions claiming similar performance generate lift for take-off and landing by relative rotation of major aircraft components or by vectoring a thrust through large angles of deflection which produces inherently unstable flight characteristics. This invention requires neither rotation of major assemblies nor highly vectorable thrust, making it simpler and more reliable than the aircraft with major rotating assemblies and more stable and efficient than aircraft relying on vectored thrust for lift.

The invention provides an aircraft and aircraft control method, and relates to the aircraft technical field. The aircraft includes at least four rotor arms; one end of each rotor arm is connected witheach other to form a connection point, and the multiple rotor arms are distributed along the circumferential direction of the connection point in a horizontal plane. Each rotor arm is provided with arotor. The rotor can rotate around the axis of the rotor arm in a vertical plane for adjustment of the working surface of the rotor. When the rotor on any one or more rotor arms fails, the rotor armon which a non-failure rotor is can rotate in the horizontal plane around the connection point, and is fixed at a position to ensure the aircraft can stably fly. In actual use, the aircraft adopting the air-changing technology can rapidly recover to a stable flight state through deformation when power failure causes instability, air crash and hazard to the ground can be avoided, and the safety ofthe aircraft can be improved.

The invention relates to a model airplane PID algorithm control method based on gravity center dynamics, and the method comprises the steps: firstly building a model airplane gravity center coordinatesystem and coordinate systems at all rotor wings, and solving a no-load rotor wing model airplane moment model and a loaded rotor wing model airplane moment model; calculating the gravity center momentum change rate of the model airplane through a gravity center kinetic model; and finally, dividing the resultant external force borne by the model airplane into acceleration calculation and inertiacalculation, and meanwhile, fusing gravity center dynamics and a PID control algorithm to realize the model airplane PID algorithm control method based on the gravity center dynamics, thereby reducingthe influence of gravity center shift on stable flight of the model airplane when the model airplane bears a load, and improving the flight stability of the model airplane.

The invention discloses a near space vertical launching single-channel stability augmentation control method and system. The method comprises the following steps: obtaining a rolling channel torque instruction and related parameters, wherein the related parameters comprise a rolling air torque, a flight pressure, an aircraft reference area, a rolling channel reference length and a design steering coefficient in a current flight state; and obtaining a differential control surface control instruction through calculation, and enabling the differential rudder to generate rolling torque by executing the differential control surface control instruction, and the aircraft flies in the target flight attitude. According to the embodiment of the invention, a pair of small differential control surfaces is added on the rolling channel of the aircraft, the stable flight effect of three channels of the aircraft is achieved only through rolling control, meanwhile, the influence of rapid change of atmospheric density in the process that the height is rapidly reduced after the aircraft is thrown in a near space is overcome, and the flight path can be adjusted in the preset flight vertical plane through a single channel, so that the requirement of a flight test for an airspace range is reduced, and the difficulty of searching and recovering an aircraft is also reduced.

The invention discloses a four-wing flapping-wing air vehicle capable of adjusting an attack angle. The four-wing flapping-wing air vehicle comprises a vehicle body, wing devices, a run-up device, a motor and a battery, the run-up device is mounted on the lower side of the body; the four sets of wing devices are installed on the two sides of the fuselage, the wing devices and the run-up device aredriven by motors, the multiple motors are powered by batteries, and an auxiliary lift device is arranged above the top of the fuselage. The wing device adopts a flapping wing form and simulates birdflight, the four groups of wings can effectively improve the flight stability and are combined with the rotor wings on the basis, the rotor wing blades are driven to rotate while the driving mechanismflaps, the rotation directions of the blades are the same during downward flapping and rising, continuous and stable lift force is provided, stable flight of the aircraft is guaranteed, and the flight efficiency is improved.