65 results about "Stabilator" patented technology

The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.

The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.

The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.

The invention discloses a vertical take-off and landing aircraft with improved linked wing tilting ducted fan layout, and belongs to the technical field of pneumatic and power systems of vertical take-off and landing aircrafts. The aircraft comprises an improved joined wing structure layout and a tilting ducted fan system integrally designed with the improved joined wing structure layout. The improved joined wing structure layout comprises a front wing, a rear wing, a vertical connecting wing surface, a vertical stabilizer and a fuselage. The front wing is divided into a front wing inner side wing surface and a front wing outer side wing surface by a vertical connecting wing surface, and adopts a sweepback wing belt dihedral structure; two ends of the rear wing are respectively arranged above the front wing through a vertical connecting wing surface, the middle part of the rear wing is arranged above the tail part of the fuselage through a vertical stabilizer, and the rear wing adopts a forward swept wing structure; the tilting ducted fan system comprises a front wing integrated ducted fan flap 7, a rear wing integrated ducted fan lifting control surface 9 and a tail propelling ducted fan 16. The core problems that pitching of a traditional connected wing is out of control, the weight of a tilting structure is large, and the energy efficiency is low are solved.

The embodiment of the invention provides a method and device for stably controlling an airplane in the auxiliary landing process, computer equipment and a medium, and the method comprises the following steps: judging whether the airplane is in a gliding stage, if so, activating a gliding stage control law, and if not, controlling the airplane to stably control the airplane through the gliding stage control law. The aircraft is controlled through the elevator centering instruction and the horizontal stabilizer automatic trimming instruction until the aircraft quits the gliding stage, and the gliding stage control law is quitted; if the aircraft enters the pull-up stage from the glide stage, activating the leveling stage control law, generating an elevator control instruction based on the steering column displacement and the steering column displacement nonlinear gain through the leveling stage control law, generating a rudder control instruction based on the pedal displacement signal and the pedal displacement nonlinear gain, and controlling the elevator control instruction and the rudder control instruction according to the pedal displacement signal and the pedal displacement nonlinear gain. And controlling the aircraft through the elevator control instruction and the rudder control instruction until the aircraft is grounded. According to the scheme, the landing quality and safety are improved through the staged aircraft auxiliary landing control law.

An electric vertical take-off and landing aircraft comprises a fuselage and a tilt wing. The tilt wing is tiltable relative to the fuselage in a range of tilt wing angles. Control surfaces are arranged at the tilt wings. The aircraft comprises a horizontal stabilizer with an elevator, in particular a stabilator, a vertical stabilizer with a rudder, in particular a fully movable stabilizer, at least one electric engine for powering main propellers arranged at the tilt wing and at least one software control device. The control device is configured to control at least the control surfaces , the main propellers and the tilt wing angle during hover and transition between hover and cruise flight.

Disclosed are devices, systems and methods for mission-adaptable aerial vehicle. In some aspects, a mission-adaptable aerial vehicle includes a configuration having swappable, manipulatable, and interchangeable sections and components connectable by a connection and fastening system able to be modified by an end-user in the field. In some embodiments, a mission-adaptable aerial vehicle can be configured to include a main center body extending along a longitudinal direction, a wing with a lateral cross-sectional airfoil shape, and / or stabilizer and control surface structures with corresponding cross-sectional airfoil shapes.

The invention discloses a comprehensive test system for testing the performance of a horizontal stabilizer actuator, and relates to the technical field of performance test systems. The device comprises a rack, a lifting platform is arranged on the left side of the rack, a control platform is arranged below the lifting platform, a hydraulic driving system and a loading system are sequentially installed at the top of the lifting platform from front to back, a hydraulic connecting assembly is arranged on the right side of the hydraulic driving system, and a hoisting system is installed at the top of the rack; a hoisting clamp is arranged below the hoisting system, a manual driving system is arranged at the inner bottom of the rack, and a displacement speed detection system and a visual positioning system are installed on the inner wall of the rack. Through the hoisting system, the horizontal stabilizer actuator can be automatically connected with the testing system in an assembly line passing mode, the horizontal stabilizer actuator comprehensive testing system can achieve automatic testing without manual intervention in the whole process, and the production efficiency is greatly improved.

The present application discloses a tail wing control mechanism, which relates to the field of aircraft. A power piece is installed on the fuselage, and its output end is fixedly connected to the first rocker arm assembly and the second rocker arm assembly; a rotating mechanism, one end of which is rotatably connected to the first rocker arm assembly, and the other end of which is fixedly connected to the stabilizer surface of the tail wing; a locking mechanism is arranged between the first rocker arm assembly and the rotating mechanism; wherein, when the rotating mechanism is in a locked state, the rotating mechanism transmits the driving force of the first rocker arm assembly to the stabilizer surface, so that the deflection direction of the stabilizer surface is always opposite to the deflection direction of the rudder surface; the end of the second rocker arm assembly away from the output end of the power piece can be rotatably passed through the stabilizer surface on the front side of the rudder surface, and is fixedly connected to the rudder surface. The present application replaces the simple increase of the rudder surface area with mechanical linkage, avoids the problem of rudder effect overload during low-speed flight, takes into account the full-speed range control requirements, and effectively solves the contradiction between high-speed or low-speed rudder effect of transonic aircraft.

Disclosed are a method and device for prompting a horizontal stabilizer trim state, which add a corresponding trim display when a "horizontal stabilizer failure" and "current trim position inconsistent with optimal trim position" alarm is triggered, so as to better meet the requirements of aviation safety regulations for alarms.

An aircraft structure includes a fuselage, a tail structure, a hinge support structure, and at least one internal actuator. The tail structure is pivotally attached to the fuselage and includes a tailcone and horizontal stabilizers fixed to the tailcone. Each of the horizontal stabilizers have a stabilizer leading edge and a stabilizer trailing edge and extend in opposite directions from the tailcone, forming an H-tail or a U-tail configuration. The hinge support structure is fixed at least partially within the fuselage, pivotally coupling the fuselage and the tail structure. The at least one internal actuator actuates the tail structure, rotating the stabilizer leading edge up and down relative to the fuselage, thereby trimming the horizontal stabilizers for aircraft stabilization.

1. The name of the design product: airplane horizontal stabilizer control intelligent teaching equipment. 2. The use of the design product: for the control principle teaching of airplane horizontal stabilizer in airplane maintenance training course. 3. The design points of the design product: in shape. 4. The picture or photo that can best show the design points: perspective view. 5. The bottom of the design product is the part that is not easy to see or cannot be seen during use, so the top view is omitted.

The embodiment of the invention provides a simulation method and device for a fault of a horizontal stabilizer of a medium-sized propeller-driven aircraft, computer equipment and a medium, and the method comprises the following steps: setting a fault mode of the horizontal stabilizer, and setting a deflection rate proportion and control surface target skewness; the control surface skewness of each horizontal stabilizer is output through a control surface actuator model; estimating six-component aerodynamic parameters of the single horizontal stabilizer; when the fault mode is trim rudder shifting, the lift coefficient and the pitching moment coefficient of the six-component aerodynamic parameters are corrected; according to the control surface skewness and the six-component aerodynamic parameters of each horizontal stabilizer, the state parameters of the airplane at the next moment are obtained through simulation calculation of the state parameters of the airplane at the previous moment by means of the airplane six-degree-of-freedom simulation calculation model. According to the scheme, through the fault simulation of the horizontal stabilizer, the authenticity of the command response of the horizontal stabilizer of the airplane is improved.

A rotorcraft with distributed propulsion system having the capability to transition between two configurations suitable for thrust borne VTOL (vertical take-off and landing) flight and partial wing borne airplane flight. The rotorcraft includes a fuselage, an empennage, a vertical stabilizer, a rudder, a pair of wings, a pair of flaps, a pair of canards, a pair of pylons, a pair of tiltable rotor booms and a plurality of rotors. The tiltable rotor booms are associated with an actuator to change collectively the direction of the rotor thrust vector with respect to the fuselage. The rotor can be open rotor or ducted rotor. The rotor booms oriented in the canted position is the first configuration suitable for partial wing borne flight and the rotor booms oriented in the horizontal position is the second configuration suitable for thrust borne flight.

The vertical stabilizer provided by the application comprises a rudder surface and a side end plate, wherein the rudder surface comprises a sandwich structure, the side end plate comprises a frame beam structure and a sandwich structure, the rudder surface is movably connected with the side end plate, and the side end plate is fixedly connected with an end rib of a horizontal stabilizer.

The invention relates to a folding empennage of a patrol flight unmanned aerial vehicle, and belongs to novel unmanned aerial vehicle structures and mechanisms. The device is composed of a mounting base, a folding and unfolding mechanism, a locking device and an airfoil structure. The mounting seat is arranged on the side surface of the rear part of the unmanned aerial vehicle body and consists of a mounting shaft, a flat spiral spring, a ball bearing, a bearing bush, an end cover, a bottom plate and an empennage nut; the airfoil structure comprises a vertical stabilizer and a rudder; the locking device comprises a locking device 1 and a locking device 2; the locking device 1 comprises a spring clip-clamping block external locking device and a limiting surface limiting-damping device; the locking device 2 comprises a flat spring sliding block-rotating shaft convex angle built-in locking device and a fuselage stop block-vertical fin rear edge limiting-damping device.

The unmanned aerial vehicle foldable empennage with the optimized aerodynamic performance comprises an empennage frame, the empennage frame is of an inverted-U-shaped double-vertical-fin structure and comprises a horizontal-tail stabilizing plane and vertical-tail stabilizing planes on the two sides, the vertical-tail stabilizing planes are perpendicularly connected with the horizontal-tail stabilizing plane, the vertical-tail stabilizing planes and the horizontal-tail stabilizing plane are provided with an elevator and a rudder respectively, and the horizontal-tail stabilizing plane and the vertical-tail stabilizing plane are connected with the elevator. The empennage frame is connected with a connecting rod, the end of the connecting rod is connected with a folding component matched with the fuselage, the folding component is hinged to the fuselage and provided with a locking component, and overturning and folding of the empennage are achieved. The problems that an existing fixed-wing unmanned aerial vehicle empennage is redundant in structure, poor in pneumatic performance due to a detachable scheme and the like are solved, the empennage can avoid underwashing turbulent flow interference of a main wing, tremor is reduced, the empennage has foldability, rapid folding and efficient space utilization are facilitated, and the portability and high efficiency of the fixed-wing unmanned aerial vehicle are enhanced.

The invention discloses an aircraft vertical fin damping device based on shape memory alloy net-shaped non-planar weaving and an integrated compression molding method of the aircraft vertical fin damping device, and relates to the technical field of aviation composite material structures, the aircraft vertical fin damping device comprises a composite material combined thin-wall plate, and the appearance of the composite material combined thin-wall plate is matched with a curved surface structure of a vertical stabilizer of a new energy aircraft; the non-planar net-shaped woven shape memory alloy steel wire rope vibration reduction unit is formed by weaving horizontal memory alloy steel wire ropes and vertical memory alloy steel wire ropes in a transverse and longitudinal staggered mode. The integral forming mould pressing system comprises an inner surface forming mould, an outer surface forming mould, a bottom plate and a steel wire rope clamping device; the epoxy resin adhesive is used for bonding and curing the vibration reduction unit and the composite material thin-wall plate, the vibration reduction device and the composite material vertical fin are integrally formed through a mold pressing process, the additional weight and installation space occupation of a traditional externally-hung device are avoided, and additional energy supply is not needed based on the passive vibration reduction characteristic of the shape memory alloy.

The application discloses a vertical stabilizer conical bolt dismounting device and a dismounting force measuring method, which comprises a main plate, a back plate, a pressure sensor assembly, a connecting plate, a ball head bolt, a guide bolt, a wrench, and an industrial computer. The main plate and the back plate are fixedly connected, and an airplane tail wing product provided with the vertical stabilizer conical bolt is fixed between the main plate and the back plate. The ball head bolt is arranged on the main plate. The connecting plate is connected to the main plate through the guide bolt on the side close to the back plate. The pressure sensor assembly is fixed on the connecting plate on the side close to the back plate. The back plate is provided with a stepped through hole. The industrial computer is connected to the pressure sensor assembly. The wrench clamps the nut end of the ball head bolt and applies a torque to the nut end. The ball head bolt rotates to drive the connecting plate to move close to the airplane tail wing product, pushes the pressure sensor assembly to tightly press the vertical stabilizer conical bolt, and then pushes the vertical stabilizer conical bolt out of the mounting hole and makes the vertical stabilizer conical bolt fall into the stepped through hole on the back plate. Meanwhile, the thrust value of the process is recorded and transmitted to the industrial computer for reading and analysis.

Systems and methods for controlling the motion, flight, and stability of vertical takeoff and landing (VTOL) aircraft are disclosed. These systems employ four or more stabilators with airfoil cross-sections to redirect airflow from rotor blades, allowing independent adjustment for hover, roll, yaw, pitch, and forward or reverse movement. In one configuration, the VTOL aircraft transitions between helicopter and fixed-wing modes during flight, with stabilators augmenting lift in fixed-wing mode. Additional yaw control surfaces redirect thrust from engines to enhance maneuverability. The design enables low-cost, easily maintained VTOL aircraft capable of achieving forward speeds greater than those of traditional helicopters.

The application belongs to the technical field of aircraft design, and particularly relates to a high-lift-coefficient layout aircraft and a use method thereof. The aircraft comprises a fuselage, a hump arranged on an upper middle part of the fuselage, a lower wing connected to the hump, an upper wing located above the lower wing and detachably connected to the lower wing, the upper wing being connected to the lower wing through an actuator and a four-bar linkage mechanism, the separation or adhesion of the upper wing relative to the lower wing being controlled by the actuator, a trailing edge flap and aileron arranged on a trailing edge of the lower wing, a V-tail stabilizer connected to a rear part of the fuselage and a V-tail control surface connected to a tail part of the V-tail stabilizer, and an engine connected to the fuselage. The application can greatly increase wing area and effective angle of attack, and significantly improve take-off and landing lift.

The invention relates to the technical field of aircrafts, in particular to a horizontal omni-directional movement jet helicopter. Comprising a fuselage, a left wing, a right wing, a vertical wing, a left lower vertical stabilizer, a right lower vertical stabilizer, a left turboshaft engine, a right turboshaft engine, a single-shaft turbojet engine, a gas collection chamber on a shell of the turbojet engine, a left front lift fan, a left rear lift fan, a right front lift fan, a right rear lift fan, a right front lift nozzle and a right rear lift nozzle, a left front lifting force spraying pipe, a left rear lifting force spraying pipe, a left front horizontal spraying pipe, a left rear horizontal spraying pipe, a front left horizontal spraying pipe, a front right horizontal spraying pipe, a right front horizontal spraying pipe, a right rear horizontal spraying pipe, a rear right horizontal spraying pipe and a rear left horizontal spraying pipe. The horizontal omni-directional motion jet-propelled helicopter not only inherits the advantages of vertical lifting and front-back and left-right lateral flight of the helicopter, but also overcomes the defect that the stability of the helicopter is not better than that of a fixed-wing aircraft, and has the advantages of high flight speed and stable flight of the fixed-wing aircraft.

The invention discloses an aircraft stabilizer device, and belongs to the technical field of aircraft stabilizers, the aircraft stabilizer device comprises a C-shaped support, the outer side of the C-shaped support is rotatably connected with an angle measuring plate through a rotating shaft, the outer side surface of the angle measuring plate is fixedly connected with a gravity block, and the angle measuring plate is provided with two lightweight grooves. A plurality of scale marks are further arranged on the outer side surface of the angle measuring plate, and a synchronous driving assembly is installed on the C-shaped support; according to the aircraft stabilizer device, when the aircraft stabilizer device is used, a hand screw is manually screwed to rotate, due to threaded connection of a connecting block, a C-shaped transmission plate moves on two transmission columns, two clamping plates can move oppositely in cooperation with the design of an inclined groove, and therefore the aircraft stabilizer device is rapidly clamped and fixed to an aircraft sample to be detected; during measurement, under the action of the gravity block, the angle measuring plate can swing along with the rotating shaft due to the arrangement that the spline housing is matched with the spline bar, so that the vertical angle between a sample to be detected and the ground can be measured.

A B737NG aircraft horizontal stabilizer actuator secondary brake function testing and measuring tool relates to the technical field of aircraft maintenance tools, and is characterized in that after a torque wrench is adjusted to a target torque value required by a Boeing manual or a work card, the torque wrench is clamped on an upper stop block at the upper end of an actuator screw rod, and a torque is anticlockwise applied to the actuator screw rod and kept; the actuator lead screw rotates to read the distance or angle value through the rotation angle measuring instrument. Then, the same working procedure is executed, torque is applied to the actuator lead screw clockwise and clockwise and kept, at the moment, the distance or angle number is read through the rotation angle measuring instrument, the difference value of the two times of reading is the measured value needed by the manual or the work card, and a test result is obtained after the measured value is compared with the standard value in the manual or the work card.

The utility model discloses a fixed-wing unmanned aerial vehicle based on flexible perovskite solar cells, which adopts a wing body fusion design and comprises an equipment compartment in the middle of a vehicle body and wings on two sides, and a pair of ailerons with rolling and pitching control functions is arranged on the rear edge of the outer section of each wing; a pair of vertical stabilizers are arranged at the tips of the wings so as to enhance the course stability during flight; two flexible perovskite solar cells which accord with the shapes of the wings and are tightly attached to the wings are mounted on the surfaces of the fuselage and the wings, and power is provided for the unmanned aerial vehicle. According to the utility model, the sweepback wing flying wing type layout is adopted, so that the effective lift area of the wings and the laying area of the solar cells are increased while the structural weight is reduced; the battery is laid on the surface of the wing, damage to the battery in the laying process can be reduced, and the installation difficulty is lowered; meanwhile, the flexible perovskite solar cell has the advantages of low cost, good folding and bending performance and the like, so that the flight efficiency of the unmanned aerial vehicle is improved, the effective load is increased, and the endurance is prolonged.

A supply glider is disclosed herein. The supply glider includes a body having a first end and a second end, the body being configured to be coupled to a host aircraft, a container disposed between the first end and the second end of the body, wings coupled to the body between the first end and the second end, and stabilators coupled to the body adjacent the second end.

The application discloses a kind of dual floaters system based on sail navigation and its regional residence control strategy, system includes the upper buoyancy ball and lower float boat connected by traction rope;The vertical tail of the capsule body tail of lower float boat is provided with, the symmetrical stabilizer of the upper and lower of capsule body is provided with, can cause Bernoulli effect, energy cabin is connected outside capsule body, and adjusting device is arranged in energy cabin to realize the retraction of traction rope;Solar panel is installed on the left and right sides of capsule body to power supply adjusting device, and auxiliary air bag is built-in in capsule body;The upper buoyancy ball and float boat are filled with quantitative lifting gas in capsule body, and air inlet and air outlet are arranged in auxiliary air bag, compressor is installed at air inlet to supplement air for auxiliary air bag, and air release opening is opened to release air for auxiliary air bag;The capsule body of float boat is suspended with load cabin under side.This application solves the problem of unpowered floater regional / point residence, solves the key problem of long-time ground reconnaissance monitoring, regional communication relay and other application needs of floater.

The present invention discloses an impact mitigation device and method for an electric horizontal stabilizer actuator. In the device, a pawl angle sensor is coaxially mounted on each pawl to monitor the pawl's angle in real time. A load sensor is mounted on the axial load-bearing structure of the electric horizontal stabilizer actuator to detect the load signal applied to the actuator during operation. A controller is connected to the load sensor and each pawl angle sensor, respectively, and is configured to perform an impact mitigation operation after the reverse drive of the electric horizontal stabilizer actuator stops, based on signals collected from the load sensor and the pawl angle sensor, to prevent the pawl mounting seat from bearing the load of the corresponding pawl. Embodiments of the present invention mitigate the impact between the ratchet and pawl when the reverse drive of the horizontal stabilizer actuator stops.

An aircraft tail section comprises: - a vertical tail; - an aft fuselage section attached to the vertical tail and comprising a skin and internal stiffening members; - a horizontal tail comprising two lateral torsion boxes and a frame structure between the two lateral torsion boxes, the frame structure comprising a front spar, a rear spar and two ribs extending between the front spar and the rear spar, and each rib being adjacent a lateral torsion box. The frame structure encloses a portion of the vertical tail along a spanwise direction of the vertical tail. The aircraft tail section comprises an attachment assembly attaching the frame structure to the aft fuselage section, the attachment assembly crossing the skin and extending between the internal stiffening members of the aft fuselage section and the frame structure.

A method for determining the position of an anti-stabilizer surface of a canard missile comprises the following steps: obtaining original information; determining a first state, calculating a first position of a vortex core of a downwash vortex of the anti-stabilizer surface and a main canard rudder in the first state; obtaining parameter information in a second state, calculating a second position of a vortex core of a downwash vortex of the anti-stabilizer surface and a main canard rudder in the second state, wherein the second position is a position of the first position when a coupling condition is satisfied during movement of the first position in the second state; and calculating a distance position x from a root chord leading edge of the main canard rudder to a root chord trailing edge of the anti-stabilizer surface based on the first position and the second position. k , thereby determining the position of the anti-stabilizer at different sound speeds. By determining the first and second states, the first and second positions of the anti-stabilizer downwash vortex core and the main canard are calculated. The anti-stabilizer position is quickly determined by utilizing the unique geometric positional relationship between the anti-stabilizer downwash vortex core and the main canard when the anti-stabilizer and the main canard are in close proximity coupling.