197 results about "Forward flight" patented technology

The invention discloses an evtol tilt rotor aerodynamic noise wind tunnel test method, and belongs to the technical field of wind tunnel tests. In order to establish an evtol tilt rotor aerodynamic noise wind tunnel test method, the method comprises the following steps: carrying out a hovering empty propeller hub background noise test: rotating a rotor table, enabling the axis of a rotor to be perpendicular to the axis of a wind tunnel, carrying out an empty propeller hub aerodynamic and noise test, and then carrying out aerodynamic data acquisition and noise data acquisition at each rotating speed to obtain hovering empty propeller hub background noise data; performing a tilt rotor hovering noise test: installing blades, setting a rotor axis to be perpendicular to a wind tunnel axis, then performing pneumatic data acquisition and noise data acquisition at each total pitch and rotating speed, and correcting the obtained tilt rotor hovering noise data by using hovering empty propeller hub background noise data to obtain a tilt rotor hovering noise test result; obtaining static tension and static power under each total pitch and rotating speed of tilt rotor hovering; carrying out a transition state and axial flow forward flight empty propeller hub noise characteristic test; and carrying out a transition state and axial flow forward flight tilt rotor noise characteristic test.

A vertical take-off or landing (VTOL) aircraft is disclosed that includes a propulsion system comprising at least four tilting propulsion assemblies and two or more lifting propulsion assemblies. Actuation of the tilting mechanisms can convert the aircraft from a hover configuration, in which all propellers produce thrust generally upward to counteract the aircraft's weight, to a forward flight configuration, in which the four or more tilting propellers produce thrust generally in a forward direction parallel to the fuselage to overcome drag in cruise flight and the lifting propellers produce no thrust or a level of thrust generally less than that required in the hover configuration.

The invention belongs to the field of electric vertical take-off and landing aircrafts, and discloses a novel vertical take-off and landing remote high-speed manned electric tilt-rotor aircraft structure which comprises an aircraft body, wings, a V-shaped empennage, six supporting arms, six tilt-rotor units, an energy system, an avionics system and an undercarriage. The six tilting rotor wing units are symmetrically installed on the wings and the V-shaped empennage through the six supporting arms. The aircraft is mainly characterized in that the design of six tilting rotor units is adopted, so that the aircraft has the advantages of vertical take-off and landing, high-speed forward flight, no structural'dead weight 'and useless power consumption; the design that the supporting arms extend out of the wings and the empennage to exceed the radius length of the rotors is adopted, and the downward projection of a rotor disc is not overlapped with the wings and the empennage in the vertical flight state, so that downward washing flow of the rotors cannot generate downward loads on the wings and the empennage, and power consumption for overcoming the downward loads is avoided; the design of the wings with high aspect ratio and winglets is adopted, so that the forward flying lift-drag ratio and the cruising efficiency are improved, and the aircraft has high-speed forward flying and remote flying capabilities; and the six tilting rotor units adopt the same tilting mechanism, so that the complexity of design and manufacturing is reduced.

The invention belongs to the field of aircraft design and control, and particularly relates to a control method of a combined type unmanned aerial vehicle based on passive hinge tilting. The composite unmanned aerial vehicle carries a tilting device and a fixed wing, the tilting of the tilting device is controlled by differential speed of motors at the two ends of the tilting device, and vertical take-off and landing and tilting transition of the unmanned aerial vehicle are controlled by adopting a PID control law. The main control board controls the main power motor and provides main power for the vertical take-off and landing stage and the horizontal cruise stage of the unmanned aerial vehicle. The sub-control board controls tilting motors at the two ends of the tilting device and provides power for enabling the tilting device to tilt. When the unmanned aerial vehicle takes off and lands, the tilting device is in a locked state, and after the unmanned aerial vehicle reaches a cruise height, the unmanned aerial vehicle is controlled to fly horizontally through the tilting device to complete a cruise task; and after the cruise is finished, the unmanned aerial vehicle is controlled to enter a hovering state through the tilting device. According to the invention, decoupling of forward flight and pitch angle attitude is realized, multiple purposes are realized, and flight tasks can be completed in a complex environment.

An exemplary tiltrotor aircraft having a vertical takeoff and landing (VTOL) flight mode and a forward flight mode includes tiltable rotors located at forward boom ends, tiltable ducted fans located at wings aft of the forward boom ends, and aft rotors located on aft boom portions.

The embodiments of the present invention provide a power processing method and device for a flying car. The method includes: obtaining state parameters of the operating state of the flying car; performing iterative calculation of theoretical values based on the state parameters of the operating state and basic aerodynamic data to obtain the power value of the flying car in the operating state; generating the total power of the flying car under the state parameters of the current operating state according to the power value. By adopting the combined means of theoretical calculation and CFD, the basic aerodynamic data with fewer calculation cases is obtained by using the CFD method, and the data source of theoretical calculation is supplemented. While improving the accuracy of theoretical calculation, the power values of the whole machine of the flying car at different hover altitudes, different forward flight speeds, different climbing speeds, and different takeoff weights can be quickly calculated, and the power requirements corresponding to the complete speed-altitude envelope can be obtained, so as to provide parameter basis for power selection. Since the calculation workload of CFD is greatly reduced, the effects of low calculation cost and high timeliness are achieved.

The invention discloses a distributed jet active flow control-based rotor wind tunnel test method, which belongs to the technical field of aviation aerodynamic wind tunnel tests, and comprises the steps of rotor test bed dynamic characteristic test, manipulation matrix calibration, preposed preparation of rotor model dynamic balance and common taper adjustment, and dynamic balance and common taper adjustment. And then hovering and forward flight wind tunnel tests of distributed jet active flow control are carried out in sequence, test data are collected and processed, and the aerodynamic performance of the rotor wing under the condition that jet control exists or not is compared and analyzed. According to the invention, the industrial blank of the lack of a distributed jet active flow control rotor wind tunnel test method is filled, and key technical support is provided for theoretical research, technical optimization and engineering application of the rotor distributed jet active flow control method.

An aircraft comprising a fuselage with a nose and a tail arranged on opposite parts to each other along a first longitudinal axis is described; a pair of half-wings arranged on respective mutually opposite sides of the fuselage; a first and a second rotor carried by respective half-wings, respectively rotatable around a second and third axis inclinable with respect to said fuselage, and independently operable from each other; the aircraft is switchable between a first hovering flight or take-off / landing configuration wherein the fourth and fifth axis are arranged orthogonal to said first axis; and a second forward flight configuration wherein the fourth and fifth axis are arranged parallel or inclined with respect to said first axis; the aircraft further comprising a tail portion comprising a first aerodynamic surface, and a third and a fourth rotor rotatable around a fixed fourth and a fifth axis; and support means of the third and fourth rotor connected to a corresponding said half-wing and to a corresponding said fin.

An aircraft designed for vertical take-off and landing (VTOL) and fixed wing flight. The aircraft's structure includes a streamlined fuselage, forward swept wing and a canard, featuring a configuration of propulsors symmetrically embedded along the aircraft's longitudinal center of gravity and tilting propellers on the forward swept wingtips. Additionally, the fuselage includes doors to the propulsor ducts that can be opened for vertical flight and closed for forward flight.

The application discloses a helicopter flight dynamics model parameter error correction method, comprising the following steps: S1, parameterizing the discrete form of helicopter aerodynamic data to obtain multiple aerodynamic parameters and form a flight dynamics parameterization model; S2, performing trimming error analysis on the parameterization model, determining each main input parameter through sensitivity analysis, and using a trimming error value matrix and a trimming error fitting coefficient matrix about the forward flight speed to represent the model trimming error; and S3, taking the trimming error value matrix and the trimming error fitting coefficient matrix as the input of a neural network to identify each error. The application parameterizes the discrete aerodynamic data to obtain the flight dynamics parameterization model, makes the parameters more explicit, adopts the input parameter error cross combination mode to calculate and generate training samples, and identifies the error of each main input parameter by using the neural network, so that the calculation cost is effectively reduced, and the correction efficiency and accuracy are improved.

The invention provides a micro-miniature semi-tilting unmanned aerial vehicle and a control method thereof.The unmanned aerial vehicle comprises a main body structure, a coaxial double-paddle system, a tilting double-rotor system, a full-motion control surface assembly, a flight control system and a power supply assembly; the main body structure is composed of a carbon fiber carbon plate, a carbon tube and a 3D printing photosensitive resin part, and light weight and rigid supporting are achieved; the coaxial double-propeller system provides main lift force and counteracts reverse torsional moment; the double-rotor system is tilted, and a lift force / thrust mode is switched; the pneumatic efficiency of the tilting rotor wing is improved by the full-motion control surface assembly; the control method comprises a low-speed mode and a high-speed forward flight mode, the vertical take-off and landing, hovering and high-speed cruising are integrated, the structure is simplified, the stability is improved, and the cost is low.

A vertical take-off and landing aircraft which uses fixed rotors for both VTOL and forward flight operations. Some of the rotors are forward canted, while some of the rotors are forward facing. The forward canted rotors are tilted forward slightly from vertical and provide some forward propulsion during horizontal flight. The forward facing rotors may be fixed in a forward facing configuration. Forward swept wings allow for rotor placement which brings efficiency during hover operations.

Embodiments provide an aircraft with one or more tilting fan assemblies that are configured to tilt between a forward flight position and a vertical lift position. The aircraft may also include a plurality of lift fan assemblies for vertical movement. The tilting fan assemblies may be coupled to the fuselage or wings of the aircraft via one or more tilting mechanisms. A control system coupled to the aircraft may control the one or more tilting mechanisms to move the tilting fan assemblies between the forward flight position and the vertical lift position. The tilting fan assemblies may be coupled to one or more support structures that are coupled the fuselage or wings of the aircraft.

Embodiments provide an aircraft with one or more tilting fan assemblies that are configured to tilt between a forward flight position and a vertical lift position. The aircraft may also include a plurality of lift fan assemblies for vertical movement. The tilting fan assemblies may be coupled to the fuselage or wings of the aircraft via one or more tilting mechanisms. A control system coupled to the aircraft may control the one or more tilting mechanisms to move the tilting fan assemblies between the forward flight position and the vertical lift position. The tilting fan assemblies may be coupled to one or more support structures that are coupled the fuselage or wings of the aircraft.

The invention provides a foldable vertical take-off and landing type fire-fighting unmanned aerial vehicle. The foldable vertical take-off and landing type fire-fighting unmanned aerial vehicle comprises a vehicle body, ducted propellers, a ducted assembly, a folding mechanism and an undercarriage mechanism. Wherein folding mechanisms are connected to the two sides of the aircraft body correspondingly, each folding mechanism is connected with a set of duct assembly, and three duct type propellers are connected into each set of duct assembly; the bottom of the aircraft body is connected with an undercarriage mechanism; an avionics system is further installed at the front end of the aircraft body, and a flight control system and an energy system which are used in a matched mode are further arranged in the aircraft body. Lift force in a hovering state is provided through the ducted fan, forward flight thrust and lift force requirements are provided through the tilting duct during forward flight, the flight energy consumption is effectively reduced, the endurance time is prolonged, and the unmanned aerial vehicle can carry heavier fire extinguishing equipment and relief goods or carry loads to transport trapped persons.

The invention relates to a tilting mechanism for the middle part of a wing of an electric tilting rotorcraft, which can realize the fixation of a rotor wing at any tilting angle without an additional self-locking device. According to the tilting mechanism, a worm and gear transmission mode is mainly adopted, and the stable and reliable fixation of the rotor wing is realized through an accurate meshing design. Through the mechanism design, the tilting mechanism can be efficiently switched among different flight modes of vertical take-off and landing, tilting transition flight and cruise forward flight, meanwhile, the compactness and streamline of the structure are kept, and the operation flexibility and stability of the tilting rotorcraft are enhanced. In addition, the requirement for an additional locking device is eliminated through the self-locking function of the tilting mechanism, the complexity of the tilting mechanism is simplified, and maintenance convenience is improved. According to the innovative tilting mechanism scheme, an efficient, safe and economical solution is provided for the electric tilting rotorcraft, and the tilting mechanism is suitable for being applied to the electric tilting rotorcraft.

A tailsitter aircraft includes an airframe having first and second wings with first and second pylons extending therebetween, a thrust array attached to the airframe, payloads and payload saddle assemblies coupled to the pylons each configured to secure a respective payload. The thrust array includes propulsion assemblies configured to transition the airframe between a forward flight orientation for wing-borne lift and a VTOL orientation for thrust-borne lift. Each payload saddle assembly includes a latch assembly and a retainer configured to secure the respective payload against a respective pylon. A latch assembly is movable between various positions including an open position and a closed position and is configured to secure the respective payload in the closed position and release the respective payload in the open position. Each latch assembly is configured to move from the closed position to the open position to release the respective payload in the VTOL orientation.

The invention provides a control-surface-free unmanned aerial vehicle and a flight attitude control method, and the control-surface-free unmanned aerial vehicle comprises a vehicle body which is provided with wings on two sides as lifting surfaces; the first power mechanism and the second power mechanism are arranged on the wings on the two sides respectively, and the rotation directions of the first power mechanism and the second power mechanism face the outer sides of the wings on the two sides respectively; the driving mechanisms are arranged on the wings on the two sides respectively and used for driving the first power mechanism and / or the second power mechanism to swing upwards or downwards in the forward flight direction of the unmanned aerial vehicle so as to adjust the flight attitude of the unmanned aerial vehicle. The swing angle range of the first power mechanism and the second power mechanism is larger than 90 degrees. According to the unmanned aerial vehicle, flight attitude control in a fixed wing mode can be realized without setting a control surface, the dead weight of the whole structure is greatly reduced, and the load and cruising ability of the unmanned aerial vehicle are improved.

The invention discloses a high-speed level flight ten-rotor unmanned aerial vehicle. The high-speed level flight ten-rotor unmanned aerial vehicle comprises a flight control device, a fuselage frame, a plurality of rotor assemblies on the two sides of the fuselage frame and a tail thrust assembly. The motor bases of the rotor assemblies on the two sides have a certain inclination angle, and can drive the propellers to generate thrust in the vertical direction and the lateral direction when the propellers work. A tail thrust motor of the tail thrust assembly drives a tail thrust propeller to generate thrust. Through the design of the rotor wing assembly and the tail thrust assembly, rapid flight can be achieved, and forward flight of the fuselage under the horizontal state is kept; meanwhile, the unmanned aerial vehicle adopts a modular design, multiple unmanned aerial vehicles can be combined, and the carrying capacity is improved. In the control aspect, pitching, rolling, yawing, up-and-down movement and front-and-back movement of the unmanned aerial vehicle are controlled through different motor differential speeds based on a longitudinal eight-rotor structure, and height position and attitude control of the unmanned aerial vehicle can be achieved in combination with the PID control principle. Compared with a traditional unmanned aerial vehicle, the flight performance and the carrying capacity of the unmanned aerial vehicle are remarkably improved.

A convertible aircraft with an airframe defining a first longitudinal axis is described; a pair of half-wings; a first, a second, a third, a fourth, a fifth and a sixth rotor rotatable about respective first, second, third, fourth, fifth and sixth axis, and operable independently of each other so as to generate respectively a first, a second, a third, a fourth, a fifth and a sixth thrust value independent of each other; said fifth and sixth rotor being carried by respective said half-wings and inclinable with respect to said airframe; the aircraft being switchable between a first hovering flight configuration wherein the sixth and seventh axis are arranged orthogonal to the first axis; and a second forward flight configuration wherein the sixth and seventh axis are arranged parallel to or inclined with respect to the first axis; the aircraft comprises first supports adapted to support the fifth and sixth rotor with respect to the respective half-wings; the first supports being arranged spaced apart from respective free ends of respective half-wings according to an extension direction of respective half-wings.

The application provides a four-rudder control coaxial rotor system, which comprises a rotor shaft, an upper rotor and a lower rotor installed on the rotor shaft, a gear box arranged on the rotor shaft, the gear box being located below the lower rotor, a first control mechanism for controlling the movement of the upper rotor to make the aircraft produce pitching movement and a second control mechanism for controlling the movement of the lower rotor to make the aircraft produce aileron movement, the first control mechanism being located between the gear box and the aircraft, the second control mechanism being located between the gear box and the lower rotor, and the first control mechanism comprising an upper tilt disc, a first rudder set and an upper variable-pitch pull rod. The application has the functions of using two sets of rudders to make the aircraft achieve elevator, forward flight and left-right steering, the rotor control functions are distinct and do not interfere with each other, the rotor control efficiency is improved, the structural complexity is reduced, the overall reliability of the device is improved, and the device is easy to maintain.

The invention provides a method, a device and equipment for predicting the maximum forward speed of an aircraft based on linear programming, and a medium, and the method comprises the steps: building a low-speed motion model of a tandem type power layout electric vertical take-off and landing aircraft; performing equivalent conversion on the solving objective function according to the moment balance relation; solving the maximum pitch angle of the aircraft based on the linear programming of the phase plane of the moment balance relationship; and calculating the maximum forward flight speed of the aircraft to solve the problems that the existing balancing method adopting the aircraft has a multi-solution problem and cannot directly determine the maximum forward flight speed, and a numerical optimization method can solve the speed estimation problem but is long in time consumption.

The invention belongs to the field of aviation fuel systems, and provides an integrated multifunctional fuel feed pump with pressure self-sensing and mode self-adaption functions, which comprises an upper pump assembly, a lower pump assembly, a motor controller assembly and a fuel discharge switch assembly, and the upper pump assembly and the lower pump assembly are respectively mounted on the upper side and the lower side of the motor controller assembly; the motor controller assembly provides a power source for the upper pump assembly and the lower pump assembly, the oil drainage switch assembly is installed on a flange plate at the bottom of the lower pump assembly, and the flange plate of the lower pump assembly is installed at the bottom of an oil tank. When the oil liquid completely submerges the oil supply pump, the upper pump assembly and the lower pump assembly simultaneously participate in pressurization work in a self-adaptive manner; when the airplane flies forward and the lower pump assembly is only submerged by oil in the oil tank, the lower pump assembly participates in pressurization work in a self-adaptive manner; when the airplane flies upside down and oil in the oil tank only submerges the upper pump assembly, the upper pump assembly participates in pressurization work in a self-adaptive mode. By means of the built-in pressure sensor, high-precision pressure self-sensing is achieved, and oil supply modes of forward flight and reverse flight of the aircraft are switched and adapted in a self-adaptive mode.

A vector tilting coaxial twin-rotor UAV is provided, which is composed of a fuselage (101), a power device (102) and a vector tilting device (103), wherein the power devices (102) and the vector tilting devices (103) on the left and right sides of the UAV are symmetrical and have completely the same structure. The problem of the rearward blade stalling caused by the high-speed forward flight of the UAV can be solved by the characteristics that the upper and lower rotors rotate at opposite speeds and the rearward blades are symmetrically located relative to the body; the roll moment generated by one rotor is balanced by the roll moment of the other rotor.

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 application discloses a flight control method and device of an aircraft and a medium, and is suitable for the technical field of the aircraft. The current sinking rate and forward flight speed of the aircraft are used to complete control on the lift value and attitude trim moment of the aircraft, the lift value and attitude trim moment are superimposed and distributed to obtain total pitch values of power systems of each propeller, then when the aircraft height satisfies a preset condition, each total pitch value is adjusted to reduce the vertical speed, and when the speed and height satisfy certain conditions, the aircraft attitude is adjusted to a target attitude to complete the aircraft attitude preparation before landing for landing. After the power system of the aircraft fails, stable glide is maintained, the propeller rotating speed in the unpowered state is adjusted through variable total pitch, and then the pulling force of the propeller is changed, so that the sinking rate and attitude of the aircraft in the power failure state are controlled, the disadvantages that the glide handling strategy cannot be implemented in the low-speed and stall states are made up, and additional emergency handling devices such as parachutes are not needed.

The invention relates to the technical field of aircrafts, and discloses an aircraft flight control device which comprises a longitudinal arm and a vertical empennage, the vertical empennage is arranged on the lower side of the tail of the longitudinal arm, and a plurality of lift propellers are further arranged on the upper side of the longitudinal arm in the front-back direction. The longitudinal arms are arranged in the front-back direction of the aircraft; when the aircraft flies forward and lowers the head, the cross section of the longitudinal arm and the horizontal plane is a standard airfoil profile, the projection of the standard airfoil profile on the cross section is the cross section of the arm, and the connecting surface of the vertical empennage and the longitudinal arm is in a vertical empennage type. The aircraft has the advantages of reducing the resistance of the longitudinal arms during forward flight, reducing the course control burden of a power system during high-speed forward flight and the like, and solves the problems that the course control capability is weak, course control is easy to oscillate or diverge, and the maximum forward flight speed is limited.

An aerial vehicle configured with air intake in an otherwise higher drag location. In some aspects, the aerial vehicle is a vertical take-off and landing aircraft. The aircraft may intake the air to provide air to a hydrogen fuel cell system within the aircraft. The aircraft may have electric motor driven rotor assemblies which provide thrust for both vertical take-off and landing and forward flight operations. The electric motor driven rotor assemblies may be powered by electric power from the fuel cell system. The air intake may be on the forward portion of a nacelle placed in what would be a high drag location, such as at the junction of the wings and the fuselage, or the junction of the vertical stabilizers and the fuselage, for example.

The present invention provides a method for manufacturing a propulsion unit having a rim wing, and a propulsion unit and an aircraft manufactured by the method. The rim wing can protect a propeller from surrounding obstacles while significantly reducing drag during forward flight. To this end, the present invention provides a method for manufacturing a propulsion unit having a rim wing, and a propulsion unit and an aircraft manufactured by the method. The method comprises: a plate member forming step of forming an airfoil-shaped plate member whose outer contour in a side view forms an airfoil; a rim wing forming step of forming a through hole in the airfoil-shaped plate member to form a rim wing member whose outer contour in a side view forms at least a portion of an airfoil; and a propeller setting step of setting a propeller in the through hole.

The present invention discloses a control method and system for the complete failure of the tail rotor of an unmanned helicopter, and relates to the technical field of unmanned helicopters. The method comprises: when the unmanned helicopter is in a hovering state, judging whether the tail rotor is completely failed according to the sensor data of the unmanned helicopter in the hovering state, and when the tail rotor is completely failed, controlling the height of the unmanned helicopter to drop to a preset height value, and then controlling the unmanned helicopter to autorotate, glide and land; when the unmanned helicopter is in a forward flight state, judging whether the tail rotor is completely failed according to the sensor data of the unmanned helicopter in the forward flight state, and when the tail rotor is completely failed, controlling the height of the unmanned helicopter to drop to a preset height value, and decelerating the forward flight speed to a preset forward flight speed value, and then controlling the unmanned helicopter to autorotate, glide and land. The present invention can realize the control of the unmanned helicopter to land safely when the tail rotor of the unmanned helicopter is completely failed.