984results about "Aircraft stabilisation" patented technology

Load stability systems and methods for stabilizing swinging motions of suspended loads. The load stability systems include a fully automated, self-powered device that employs thrust to counteract and control lateral and rotational motion of an external load. The device is a temporary installment on the load, cable, or boom, and is agnostic to the platform from which it is suspended.

A craft is provided comprising: a propulsion system; a fuselage coupled with the propulsion system, wherein the fuselage is configured to carry a payload and fuel for the propulsion system; and a pre-assembled wing coupled with the fuselage; wherein the craft is designed to take-off without using a runway and does not have landing gear to land.

An assembly for maintaining the pitch angle of a flight control surface for an aircraft. The assembly includes a shaft comprising a screw thread and defining a shaft axis, and a nut. The nut includes a housing, a barrel located in the housing and comprising a screw thread for engaging with the screw thread of the shaft, and at least one locking element located in the housing. The housing comprises a radially inner cam surface adjacent to the locking element. The barrel is configured to rotate relative to the housing when the screw thread of the barrel is engaged and rotating with the screw thread of the shaft so as to move the locking element along the cam surface from a disengaged position in which the locking element is spaced from the screw thread of the shaft, to an engaged position.

The invention discloses a folding wing and rudder design method and an unfolding locking mechanism thereof. The locking mechanism comprises a connecting handle, an elastic torsional spring and a rotating shaft. A locking head and a connector are arranged on the two sides of the bottom of the folding wing or the folding rudder respectively; the design method comprises the steps of obtaining performance parameters, and determining specific indexes such as folding sizes and unfolding time of the folding wing and the rudder. And designing a structural model according to principles of generalization, reliability and the like, and determining geometric parameters of the wing and the rudder. Verifying the strength of parts of the locking mechanism, calculating a plurality of parameters of an elastic torsion spring, and analyzing the unfolding time and synchronism. Finally, consistency checking is conducted based on finite element simulation analysis, torque and bending moment are applied in ABAQUS software according to parameters, the structural strength of the folded rudder is analyzed, and it is ensured that the design meets the requirement. According to the folding wing and rudder design method, through the scientific design process and multi-principle comprehensive application, the high-performance requirement is met, meanwhile, the cost is effectively reduced, the reliability and maintainability of products are improved, and the remarkable beneficial effects are achieved.

An electric propulsion system for a vertical take-off and landing (VTOL) aircraft having a heat exchanger to cool fluids used in an electrical engine, the electric propulsion system comprising at least one electrical engine mechanically connected directly or indirectly to a fuselage of the VTOL aircraft and electrically connected to an electrical power source. The electrical engine may comprise an electrical motor having a stator and a rotor; a gearbox assembly comprising a sun gear; at least one planetary gear; a ring gear; and a planetary carrier. The electric engine may include an inverter assembly comprising a thermal plate and an inverter assembly housing; an end bell assembly that is connected to the thermal plate of the inverter assembly; and a heat exchanger comprising an array of cooling fins and tubes.

A vertical take-off and landing aircraft with a multi-element lifting system with integrated propulsion. The multi-element lifting system may have an upper wing element, a lower wing element, and a control wing element which is located below the upper wing element and above, and rearward of, the lower wing element. The system uses internal propulsion units, such as internal ducted fans, which flow air below the upper wing element and above the lower wing element such that the air flows through the multi-element lifting system. The control wing element may be articulated to route air vertically downward to allow for short or vertical take-off and landing. An aircraft with a multi-element wing assembly and a multi-element tail assembly raised above the wing assembly. An aircraft which resides on the ground with the wing assembly and the tail assembly pitched up.

Apparatus, systems, and methods are contemplated for electric powered vertical takeoff and landing (eVTOL) aircraft. Such are craft are engineered to carry safely carry at least 500 pounds (approx. 227 kg) using a few (e.g., 2-4) rotors, generally variable speed rigid (non-articulated) rotors. It is contemplated that one or more rotors generate a significant amount of lift (e.g., 70%) during rotorborne flight (e.g., vertical takeoff, hover, etc), and tilt to provide forward propulsion during wingborne flight. The rotors preferably employ individual blade control, and are battery powered. The vehicle preferably flies in an autopilot or pilotless mode and has a relatively small (e.g., less than 45′ diameter) footprint.

A solar powered plane configured to operate in the stratosphere includes a hinged vertical stabiliser that that pivots about a hinge; the stabiliser extends both above the fuselage and also below the fuselage, and the base of the vertical stabiliser includes a skid. The hinged vertical stabiliser is configured such that when the skid contacts the ground on landing, the top of the vertical stabiliser pivots forwards around the hinge point, minimising damage to the aircraft on landing. the skid includes a lightweight, replaceable sacrificial layer.

A vertical take-off and landing aircraft with a multi-element lifting system with integrated propulsion. The multi-element lifting system may have an upper wing element, a lower wing element, and a control wing element which is located below the upper wing element and above, and rearward of, the lower wing element. The system uses internal propulsion units, such as internal ducted fans, which flow air below the upper wing element and above the lower wing element such that the air flows through the multi-element lifting system. The control wing element may be articulated to route air vertically downward to allow for short or vertical take-off and landing. An aircraft with a multi-element wing assembly and a multi-element tail assembly raised above the wing assembly. An aircraft which resides on the ground with the wing assembly and the tail assembly pitched up.

The invention relates to an empennage structure of a flapping-wing robot, which comprises a yawing mechanism, a pitching mechanism and a connecting mechanism, the yawing mechanism comprises a yawing steering engine, a yawing steering engine rocker arm, a yawing rocker arm and a yawing control surface; the pitching mechanism comprises a pitching steering engine, a pitching steering engine rocker arm and a pitching surface; the connecting mechanism comprises a pitching steering engine fixing plate, an empennage surface plate and a fuselage. The pitching motion function of the empennage is achieved through the knuckle bearing and the connecting rod structure, the yawing motion function of the empennage is achieved through the flexible material, the steering engine rocker arm and the yawing rocker arm, the yawing structure and the pitching structure are combined through the empennage plane plate, and the whole empennage mechanism is formed. The yawing motion cannot be affected during pitching motion, decoupling of the pitching motion and the yawing motion is achieved, meanwhile, the yawing control steering engine and the pitching control steering engine are both installed on the fuselage, empennage structural parts are reduced, the loading capacity during empennage motion is reduced, and the beneficial effects of being compact in structure and high in adjusting capacity are achieved.

Apparatus, systems, and methods are contemplated for electric powered vertical takeoff and landing (eVTOL) aircraft. Such are craft are engineered to carry safely carry at least 500 pounds (approx. 227 kg) using a few (e.g., 2-4) rotors, generally variable speed rigid (non-articulated) rotors. It is contemplated that one or more rotors generate a significant amount of lift (e.g., 70%) during rotorborne flight (e.g., vertical takeoff, hover, etc), and tilt to provide forward propulsion during wingborne flight. The rotors preferably employ individual blade control, and are battery powered. The vehicle preferably flies in an autopilot or pilotless mode and has a relatively small (e.g., less than 45′ diameter) footprint.

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.

According to the modularized traffic tool with multiple take-off modes, through the innovative split modularized design, the form of the traffic tool can be flexibly changed between an automobile and an aircraft, the flight state and the land running state are switched, and more novel passenger and freight application scenes can be expanded through cooperation of the independent carrying cabin modules. By utilizing the tiltable ducted fan mechanism and by virtue of boosting of the autonomous traveling chassis, the composite wing flight module can realize vertical take-off and landing, and wing lift force can be provided for short-distance taxiing take-off and landing to reduce energy consumption, so that the purposes of large voyage and high load are finally realized. The foldable fixed wing, the horizontal tail and the telescopic tail wing structure ensure that the traffic tool is small in size when running and parking on the land, so that the requirements of existing urban and rural roads and traffic safety laws and regulations can be met, and actual popularization is more facilitated.

An unmanned aerial vehicle (UAV) having wings stowed against a fuselage of the UAV in a first arrangement is disclosed. Methods and systems for deploying the wings into a second arrangement are disclosed. For example, after a launch of the UAV, the UAV monitors for at least one precondition. The at least one pre-condition being a pre-condition associated with deploying wings of the UAV into the second arrangement. Upon detecting the at least one precondition, the wings of the UAV are deployed into a second arrangement. Deploying the wings comprises activating, in response to detecting the at least one precondition associated with the UAV, a gearbox configured to transition the wings from the first arrangement to the second arrangement. Roll control may be maintained throughout launch and deployment.

The present invention discloses a bionic hummingbird flapping-wing aircraft with an ultralight structure and a control system, comprising a frame and a wing system fixed on the frame, a power transmission system, a pitch servo control system, a roll servo control system and a micro-integrated flexible folding circuit board system; the frame comprises an upper frame and a lower frame, wherein the upper frame comprises an upper bottom plate pulley frame, a bottom plate connecting member and a lower bottom plate pulley frame; the lower frame comprises a suspension, a front support frame, a rear support frame and a base; the suspension is fixed to the lower surface of the lower bottom plate pulley frame through a first hanging ear; the wing system consists of a pair of left and right wings, each wing comprises a wing pulley, a wing rod and a wing root rod, a wing membrane is fixed on the wing rod and the wing root rod, and the wing membrane is provided with a wing membrane hole at the wing root; the power transmission system realizes the flapping motion of a pair of wings by driving the reciprocating rotation of the wing pulley. By using the present invention, flapping-wing flight can be realized without an external power cord and only relying on battery power supply.

An electric propulsion system for a vertical take-off and landing (VTOL) aircraft having a heat exchanger to cool fluids used in an electrical engine, the electric propulsion system comprising at least one electrical engine mechanically connected directly or indirectly to a fuselage of the VTOL aircraft and electrically connected to an electrical power source. The electrical engine may comprise an electrical motor having a stator and a rotor; a gearbox assembly comprising a sun gear; at least one planetary gear; a ring gear; and a planetary carrier. The electric engine may include an inverter assembly comprising a thermal plate and an inverter assembly housing; an end bell assembly that is connected to the thermal plate of the inverter assembly; and a heat exchanger comprising an array of cooling fins and tubes.

The invention belongs to the technical field of airplane structure design, and discloses a horizontal tail wing beam connecting structure, an airplane and an assembling method, and the horizontal tail wing beam connecting structure comprises a central rib plate, an upper T-shaped connector, a lower T-shaped connector, a left connecting plate, a right connecting plate and an outer side connecting plate. The upper T-shaped joint comprises a lower edge strip, a first left edge strip and a first right edge strip, the lower T-shaped joint comprises an upper edge strip, a second left edge strip and a second right edge strip, the central rib plate is arranged between the left beam and the right beam, and the left connecting plate, the right connecting plate and the outer side connecting plate are combined for use, so that the bearing capacity of a left beam web plate and a right beam web plate is improved; the first left edge strip and the first right edge strip can increase the contact area between the upper T-shaped connector and the left wing and the contact area between the upper T-shaped connector and the right wing, the second left edge strip and the second right edge strip can increase the contact area between the lower T-shaped connector and the left wing and the contact area between the lower T-shaped connector and the right wing, the connecting strength of the left beam and the right beam in the vertical direction is improved, and the lower edge strip, the upper edge strip and the central rib plate are matched with one another. And the connection strength of the left beam and the right beam in the horizontal direction is improved.

The invention belongs to the technical field of flight mechanics, and particularly relates to a V-shaped empennage control surface course control extension control method which comprises the steps that a left control surface deflection angle reaches an upper limit, a right control surface deflection angle continues to deflect leftwards, and a right empennage generates right lower lift increment perpendicular to the incoming flow direction; the right empennage can simultaneously generate a vertically downward force and a head-up torque relative to the gravity center of the whole aircraft; a downward head lowering moment is generated through the longitudinal periodic pitch change of the rotor wings and the tilt angle of the rotor wings, so that the pitching moment is balanced; the vertical force balance of the whole aircraft is achieved by adjusting the attack angle of the aircraft body.

An airframe includes a body, a first wing, a second wing and a vane. The body extends longitudinally along a longitudinal centerline from a forward end of the body to an aft end of the body. The first wing and the second wing are disposed to opposing lateral sides of the body. The vane projects vertically out from the body to a tip of the vane. The vane extends longitudinally along the body from a leading edge of the vane to a trailing edge of the vane. A first open rotor propulsion system is vertically above and mounted to the body. A second open rotor propulsion system is vertically above and mounted to the body. The vane is configured as a structural barrier laterally between the first open rotor propulsion system and the second open rotor propulsion system.

An assembly for maintaining a pitch angle of a flight control surface of an aircraft is disclosed. The assembly includes a shaft including a thread and defining an axis of the shaft; and a nut. The nut includes: a housing; a barrel located in the housing and including threads for engaging with the threads of the shaft; and at least one locking element, the at least one locking element being located in the housing. The housing includes a radially inner cam surface adjacent the locking element. The barrel is configured to rotate relative to the housing when the threads of the barrel engage with and rotate with the threads of the shaft, thereby moving the locking element along the cam surface from a disengaged position, in which the locking element is spaced from the threads of the shaft, to an engaged position, in which the locking element is wedged between the threads of the shaft and the housing, as a result, relative rotation between the shaft and the housing is prevented.

A vertical take-off and landing aircraft with a multi-element lifting system with integrated propulsion. The multi-element lifting system may have an upper wing element, a lower wing element, and a control wing element which is located below the upper wing element and above, and rearward of, the lower wing element. The system uses internal propulsion units, such as internal ducted fans, which flow air below the upper wing element and above the lower wing element such that the air flows through the multi-element lifting system. The control wing element may be articulated to route air vertically downward to allow for short or vertical take-off and landing. An aircraft with a multi-element wing assembly and a multi-element tail assembly raised above the wing assembly. An aircraft which resides on the ground with the wing assembly and the tail assembly pitched up.

The invention provides a wing type biological hybrid flapping wing aircraft with deformable wings. The wing type biological hybrid flapping wing aircraft comprises a flapping mechanism, a wing deformation mechanism, bird feather type wings, an aircraft body, an empennage deformation mechanism and a bird feather type empennage, the flapping mechanism and the empennage deformation mechanism are mounted on the fuselage; the flapping mechanism drives the wing deformation mechanism to do wing flapping motion; the bird feather type wings are arranged on the wing deformation mechanism; the empennage deformation mechanism comprises an empennage joint and a plurality of steering engines, the steering engines drive the empennage joint to do four-degree-of-freedom motion, and the bird feather type empennage is installed on the empennage joint. The multi-degree-of-freedom cooperative connecting rod transmission mechanism is adopted, high-fidelity reproduction of bird flapping wing motion is efficiently achieved, and independent and accurate regulation and control of the torsion angle, the torsion rate and the torsion direction of the empennage are achieved based on hierarchical steering engine control and the decoupling characteristic of the double-layer motion frame.

There is described an aircraft (1) comprising a fuselage (2, 2 ', 2' ') having a nose (4) and a tail (5) arranged on mutually opposite portions along a first longitudinal axis (Y); a pair of half wings (3) arranged on respective mutually opposite sides of the fuselage (2, 2 ', 2' '); first and second rotors (22a, 22b) carried by the respective half-wings (3), rotatable about second and third axes (F, G), respectively, tiltable relative to the fuselage (2, 2 ', 2' '), and operable independently of one another; the aircraft (1) being switchable between a first hover flight or takeoff / landing configuration in which the fourth and fifth axes (F, G) are arranged orthogonally to said first axis (Y) and a second forward flight configuration in which the fourth and fifth axes (F, G) are arranged parallel or inclined with respect to said first axis (Y); the aircraft (1) further comprises a tail portion (6) comprising a first aerodynamic surface (8) and third and fourth rotors (21a, 21b) rotatable about fixed fourth and fifth axes (F, G); and support means (31a, 31b) connected to the third and fourth rotors (21a, 21b) of the corresponding said half wings (3) and the corresponding said fins (52).

According to the rotary folding empennage of the aircraft and the using method, two or more empennages are arranged at the tail of a fuselage of the aircraft, each empennage is connected with a steering engine, each empennage is divided into a first empennage section and a second empennage section, the two empennage sections are connected through a folding and unfolding mechanism, and folding and unfolding of the empennages can be achieved. The first empennage is provided with a rotating shaft which penetrates through the fuselage tail skin to be connected with the steering engine, a gap is formed between the empennage and the fuselage tail, and when the empennage is in a folded state, the second empennage is close to the fuselage tail; when the empennage is in an unfolded state, the second empennage body rotates around the hinge to be flush with the first empennage body, after the empennage is unfolded, the empennage can rotate along with the rotating shaft under driving of the steering engine, and control over the flight attitude of the aircraft is achieved.

The invention provides a supersonic-speed jet adopting supersonic-speed double wings and variant design, and belongs to the field of novel supersonic-speed civil aircrafts, a large-area truncated delta wing with edgings on the inner side is adopted, an engine nacelle is arranged on the upper portion of the outer end of the large-area truncated delta wing, and the supersonic-speed double wings are installed on the outer side of the nacelle; the front edges and the rear edges of the upper wing and the lower wing of the supersonic double-wing adopt intelligent materials to realize a variant function; the empennage can rotate around a mounting shaft in the longitudinal direction of the airplane to change the inclination angle; the front edge and the rear edge of the supersonic double-wing are deformed, and the inclination angle of the empennage is changed to jointly form a supersonic state and a subsonic state. The problem that the conventional double-wing non-design-point performance is poor is solved, high aerodynamic performance can be achieved within the flight envelope range, and it is guaranteed that an aircraft achieves ideal supersonic speed environment-friendly performance such as acoustic explosion and noise and aerodynamic efficiency.

The invention discloses sound wave imaging electric flaw detection unmanned aerial vehicle equipment which comprises a fuselage and a rotor arm arranged on the fuselage, the fuselage is further provided with an ultrasonic flaw detection mechanism, one end of the rotor arm of an unmanned aerial vehicle is connected with the fuselage in a hinged mode, and the other end of the rotor arm is provided with a steering engine set. A shock absorber is arranged between the fuselage and each steering engine set; by arranging the shock absorbers, the rotor arms can be kept parallel to the fuselage when the rotors are not disturbed by wind power, and when the rotors are disturbed by the wind power, the rotor arms elastically swing under the action of the shock absorbers, so that the influence of the wind power on the whole unmanned aerial vehicle can be buffered and attenuated, and the flight stability of the unmanned aerial vehicle is improved.

An electrically powered multirotor aircraft that is adapted for vertical take-off and landing, comprising a fuselage, at least one aircraft lifting surface that is connected to the fuselage, at least one propeller that is connected to the at least one aircraft lifting surface, a sensor unit, a flutter reduction unit, and a settings adjustment device. The sensor unit is configured to generate sensor data that is indicative of flutter of the at least one aircraft lifting surface. The flutter reduction unit is configured to receive the sensor data to generate flutter reduction signals. The settings adjustment device is configured to receive the flutter reduction signals and to adjust current settings of the at least one propeller based on the flutter reduction signals to reduce flutter of the at least one aircraft lifting surface.

An electric propulsion system for a vertical take-off and landing (VTOL) aircraft having a heat exchanger to cool fluids used in an electrical engine, the electric propulsion system comprising at least one electrical engine mechanically connected directly or indirectly to a fuselage of the VTOL aircraft and electrically connected to an electrical power source. The electrical engine may comprise an electrical motor having a stator and a rotor; a gearbox assembly comprising a sun gear; at least one planetary gear; a ring gear; and a planetary carrier. The electric engine may include an inverter assembly comprising a thermal plate and an inverter assembly housing; an end bell assembly that is connected to the thermal plate of the inverter assembly; and a heat exchanger comprising an array of cooling fins and tubes.

Provided in the present application are a vertical take-off and landing aircraft and a control method therefor. The vertical take-off and landing aircraft comprises a fuselage and 2N tilt-rotors. Wings are provided on two sides of the fuselage, and an empennage is provided at the tail of the fuselage, an elevator rudder being provided on the empennage. The 2N tilt-rotors are symmetrically mounted on two sides of the fuselage, and some of the 2N tilt-rotors are located on the empennage, where N is a natural number greater than or equal to 2. In a vertical take-off and landing state, the projections of propellers of the 2N tilt-rotors on the horizontal plane are centrally symmetrical about point B, both the point B and the center of gravity G of the vertical take-off and landing aircraft are located in the symmetrical plane of the fuselage, and the point B is located on the side of the point G close to the empennage. During the transition of the vertical take-off and landing aircraft from the vertical take-off and landing state to a cruising state, both the point G and the point B move along the symmetrical plane towards the side close to the nose, and the point B is always located on the side of the point G close to the empennage. Such a layout enables the vertical take-off and landing aircraft to have a better pitch trim capability.

The invention provides a non-orthogonal tilt-rotor aircraft and a flight control method thereof, a tilt shaft at the lower end of a tilt nacelle is obliquely arranged, and the tilt shaft is parallel to an XZ plane, forms an angle of 45 degrees with an X axis and a Z axis, and is not orthogonal to an XY plane and a YZ plane of a fuselage coordinate system; the positive direction of the X axis is backward along the course, the positive direction of the Y axis is rightward along the course, and the positive direction of the Z axis is upward along the vertical direction. According to the invention, the transverse stability augmentation effect can be provided, the transverse stability is improved, and the problem of unstable pitching of the fuselage caused by tilting of the rotor nacelles can be avoided.