129 results about "Forward flight" patented technology

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 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.

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.

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.

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 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.

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.

Disclosed is an electric aircraft, which includes a fuselage and at least one wing coupled to the fuselage. The electric aircraft includes a plurality of tilting fans coupled to the at least one wing, the plurality of tilting fans being configured to move between a vertical lift position and a forward flight position. The electric aircraft includes a plurality of tilting mechanisms coupled with at least one tilting fan. The electric aircraft includes a first actuator coupled to a first subset of the plurality of tilting mechanisms. The first subset of tilting mechanisms are identified among the plurality of tilting mechanisms according to a coupling scheme. The first actuator tilts a first group of tilting fans coupled to the first subset of the plurality of tilting mechanisms simultaneously.

This utility model discloses a waterborne vertical takeoff and landing (VTOL) manned aircraft, comprising a fuselage, wings, a vertical tail, tail floats, outer wing floats, and several rotor propulsion units. This waterborne VTOL manned aircraft is a lightweight seaplane capable of vertical takeoff and landing on water at any time, providing convenience for applications such as water rescue and sightseeing. The multi-float design, combining wing floats and tail floats, ensures the aircraft's stability on water. The propellers are entirely enclosed within duct openings, increasing propeller efficiency while significantly enhancing safety and reducing the risk of injury to personnel on the ground. Furthermore, the multi-rotor and wing-based structure allows both the multi-rotor and wings to provide lift during forward flight, ensuring a certain level of safety while increasing cruising efficiency. The motors employ a dual-winding configuration (or coaxial dual propellers), ensuring a stable landing even if a single winding fails.

Embodiments provide an aircraft having one or more tilting fan assemblies 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 tilt fan assembly may be coupled to a fuselage or wing of an aircraft via one or more tilt mechanisms. A control system coupled to the aircraft may control the one or more tilt mechanisms to move the tilt fan assembly between a forward flight position and a vertical lift position. The tilt fan assembly may be coupled to one or more support structures that couple a fuselage or wings of the aircraft.

A tandem wing aircraft (10) comprises a fuselage (20) having a nose (30) and a tail portion (45). A first wing structure (40) extends away from a right side of the fuselage (20) including a first forward wing (50) and a first rearward wing (60). A second wing structure (70) extends away from a left side of the fuselage (20) including a second forward wing (80) and a second rearward wing (90). A plurality of electric motors (100) having rotors (110) mounted on each of the first and second wing structures (40, 70). The tail portion (45) of the fuselage (20) terminates at a longitudinally extending edge, the edge being defined by the junction of an upper surface (130) and a lower surface (160), the edge extending generally perpendicular to a direction of forward flight. The first rearward wing (60) and second rearward wing (90) are connected to define a continuous span which passes above the upper surface (130), the continuous span being supported by a support structure (300) which extends upwardly away from the fuselage (20).

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 aircraft can include: an airframe, a tilt mechanism, a payload housing, and can optionally include an impact attenuator, a set of ground support members (e.g., struts), a set of power sources, and a set of control elements. The airframe can include: a set of rotors and a set of support members. By utilizing a larger rotor blade area (and / or larger rotor disc area) and adjusting the blade pitch and RPM, the rotors can augment the lift generated by the aerodynamic profile of the aircraft in the forward flight mode in addition to providing forward thrust. Variants generating lift with the rotors can reduce or eliminate additional control surfaces (e.g., wing flaps, ailerons, ruddervators, elevators, rudder, etc.) on the aircraft since the thrust and motor torque is controllable (thereby indirectly controlling lift) at each rotor, thereby enabling pitch, yaw, and / or roll control during forward flight.

The invention belongs to the field of pneumatic and flight control of tilting rotorcrafts, and relates to a vortex ring boundary protection control method of a tilting rotorcraft. Comprising the steps that firstly, according to the maximum takeoff weight M and the rotor wing radius R of the tilt rotorcraft, the induced speed V1 of rotor wing downwash during hovering of the tilt rotorcraft is determined; 2, according to the induced velocity V1 of rotor wing downwash, the vortex ring boundary of a single rotor wing is determined, and the boundary is a binary quadratic curve composed of a forward flight velocity VX and a vertical velocity Vy; 3, the vortex ring boundary of the whole tilt rotorcraft is determined according to the vortex ring boundary of the single rotor and the flight characteristics of the tilt rotorcraft in the flight process, and the flight characteristics of the tilt rotorcraft in the flight process are that the upper half part of the aircraft entering the binary quadratic curve indicates that the aircraft enters the vortex ring boundary; therefore, whether the airplane enters the upper half part of the binary quadratic curve or not is considered in practice; and 4, reserving a preset safety margin on the basis of the boundary of the vortex ring of the whole tilt rotorcraft, and taking the preset safety margin as an early warning boundary of the vortex ring in the flight process.

A rotor mounting assembly for a vertical take-off and landing aircraft includes a boom configured for mounting to a wing of the aircraft; a mount for mounting a rotor assembly, the mount connected to the boom at a joint and tiltable about the joint from a forward thrust orientation in which the rotor assembly can provide forward thrust for forward flight to a vertical thrust orientation in which the rotor assembly can provide vertical thrust for vertical take-off and landing and hover; a multi-link assembly extending from the boom to the mount; and a rotary actuator for actuating the multi-link assembly to control tilting of the mount.

The utility model discloses a variable-speed transmission system and a crossed double-rotor helicopter, the variable-speed transmission system comprises a first input shaft, a first clutch and a first output shaft, and the first input shaft is connected with the first output shaft through the first clutch; the speed changing device comprises a first gear train mechanism, a second gear train mechanism and a second clutch, the first gear train mechanism is connected with the second gear train mechanism through the second clutch, the input end of the first gear train mechanism is connected with the first input shaft, and the output end of the second gear train mechanism is connected with the first output shaft; and at least one of the first gear train mechanism and the second gear train mechanism is a speed reducing mechanism. The power output path of the first input shaft is changed through connection or disconnection of the first clutch and the second clutch, and the rotating speeds output by the first output shaft are different. According to the high-speed helicopter adopting the variable-speed transmission system, the purpose of reducing the rotating speed of the rotor wing through the variable-speed transmission system is achieved, and the influence of blade shock waves on the forward side of the rotor wing is reduced.

An aircraft comprises a fuselage, first and second wing segments each having a leading edge and a trailing edge, a plurality of spokes coupling the fuselage to the first and second wing segments, one or more motors disposed within or attached to the plurality of spokes, and three or more propellers proximate to a leading edge of the plurality of spokes, distributed along the plurality of spokes, and operably connected to the motors to provide lift whenever the aircraft is in vertical takeoff and landing and stationary flight and provide thrust whenever the aircraft is in forward flight. When the aircraft is in vertical takeoff and landing and stationary flight, the fuselage is approximately vertical. When the aircraft is in forward flight, the fuselage is approximately in the direction of the forward flight and extends forward beyond the leading edges of the first wing segment and the second wing segment.

A hydrogen fuel cell powered electric vertical take-off and landing (eVTOL) aircraft with a high efficiency hydrogen fuel system. The eVTOL aircraft may utilized tilt-up rotors for hover flight, which then transition to a forward facing forward flight configuration. The fuel cell system may use one or more compressors to compress air to a sufficiently high pressure for the fuel cell. Liquid hydrogen may be compressed and then utilized in heat exchangers to cool the compressed air, maintaining the air at a temperature low enough for the fuel cell. The hydrogen may also be used to cool the fuel cell as it is also depressurized prior to its entry in the fuel cell cycle.

An aerial vehicle configured a main fuselage and an upper nacelle coupled to the top of the main fuselage. The aerial vehicle may have an 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 or a turbogenerator within the aircraft, which may reside withing the upper nacelle. 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 aircraft may have pivoting wings with rotor assemblies which may provide thrust in both a forward flight configuration and a vertical take-off and landing (hover) configuration.

The invention discloses a noise reduction design method for blades of a multi-rotor aircraft, and relates to the technical field of design and manufacturing of multi-rotor aircrafts. Comprising a blade body, the blade body is in smooth transition from the blade root to the blade tip in the radial direction, and a sawtooth structure is integrally formed on the rear edge of the blade from the 50% radius position to the blade tip. Through bionic extraction of Tyto-owl and hummingbird low-noise sawtooth characteristics, in combination with NACA6412 airfoil profile modification and torsional angle design and more depending on a sensor-control unit-piezoelectric drive closed-loop system, the sawtooth height-width ratio is dynamically adjusted, hovering / forward flight working conditions are distinguished with 1m / s as a threshold value, hovering calls an optimal noise reduction parameter, forward flight calls an optimal comprehensive performance parameter, and the optimal performance is obtained. And independent control can be implemented for asymmetric incoming flow of the left and right paddles, verification shows that the noise reduction amount in all working conditions is larger than or equal to 3dB, the aerodynamic efficiency loss is smaller than or equal to 5%, the problems that traditional fixed sawteeth only adapt to a single working condition and noise reduction and efficiency are difficult to consider are thoroughly solved, and the scene requirements of urban aerial photography, residential area operation and the like are met.

The utility model provides a tandem type double-rotor structure which comprises a first rotor assembly and a second rotor assembly, the first rotor assembly is arranged at the front end of an unmanned helicopter body, and the second rotor assembly is arranged at the rear end of the unmanned helicopter body; the bottom ends of rotor shafts of the first rotor assembly and the second rotor assembly are connected with the corresponding mounting bases correspondingly, the mounting bases are connected with the fuselage, and the mounting bases have preset forward inclination angles relative to the fuselage. The front rotor wing and the rear rotor wing incline forwards by a certain angle at the same time, so that when the unmanned helicopter flies forwards at a high speed, the helicopter body can be kept horizontal as much as possible, only the front propeller disc and the rear propeller disc incline forwards to generate forward propelling component force, and therefore the windward area of the helicopter body can still be kept minimum; according to the unmanned helicopter, the resistance of the helicopter body during forward flight can be reduced, the windward area of the propeller disc can be smaller, the forward flight resistance can be further reduced, the aerodynamic performance of the unmanned helicopter is improved, and meanwhile the applicability of the unmanned helicopter is improved.

The invention discloses an aircraft additional air injection structure capable of adjusting the lift ratio and an adjusting method, and belongs to the field of aerospace, the aircraft additional air injection structure comprises an air injection unit, when an aircraft is a complete machine body integrating the wing function, the air injection unit is arranged at the top or the bottom of the complete machine body, and when the aircraft is provided with independent wings, the air injection unit is arranged at the top or the bottom of the complete machine body. The air injection units are arranged at the tops or the bottoms of a fuselage and wings of the aircraft, the installation positions of the air injection units are designed according to the aerodynamic basis of the fuselage and the wings of the whole aircraft, and the air injection units do not interfere with an air injection channel of an original engine. According to the aircraft additional air injection structure with the adjustable lift ratio and the adjusting method, high-speed airflow is injected through the angle-adjustable air injection units installed on the top or the bottom of the whole fuselage or the wings, and the angle and the strength of the injected high-speed airflow are used for assisting and controlling the aircraft to increase the forward flight bearing capacity; the aircraft can also fly upside down, roll horizontally, roll in a barrel, rush upwards, dive and other maneuvers at the bottom of the aircraft.