127 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.

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 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 relates to a dual-motor input horizontal stabilizer balancing system which comprises a channel 1 motor assembly, a channel 2 motor assembly, a differential gear train, a final-stage gear, a reduction gear, an output gear, a ball screw and a lower connector. The dual-motor input horizontal stabilizer balancing system further comprises a main control electronic component PCE which is configured to be used for generating a corresponding balancing rate instruction according to a flight control law based on external data; and the motor control unit MCU comprises an MCU1 used for a channel 1 and an MCU2 used for a channel 2, and is configured to be used for generating a corresponding control instruction according to the balancing rate instruction and providing the control instruction to the channel 1 motor assembly (1) and the channel 2 motor assembly (2).

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.

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 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.

The invention relates to the technical field of aircrafts, and discloses an aircraft and an arm thereof. The vehicle arm is applied to the aircraft and comprises a vehicle arm body and an external extending part, the vehicle arm body and the external extending part are integrally connected, the external extending part extends outwards and is of a structure gradually reduced in the first direction, and the external extending part comprises a streamline surface structure so as to reduce air resistance when the aircraft flies. The machine arm main body and the vertical fin are integrally formed and connected, so that the stabilizer rigidity of the vertical fin can be improved, the loosening risk of the vertical fin is reduced, the number of parts is reduced, and the weight is reduced.

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.

The utility model discloses a full-lifting-body perovskite solar unmanned aerial vehicle which comprises a large-chord-length lifting body fuselage and wings on the two sides, ailerons are arranged on the rear edges of the outer sections of the wings to serve as rolling control faces, and an elevator is arranged on the rear edge of the fuselage to serve as a pitching control face. A pair of vertical stabilizers are arranged on the outer side surface of the rear edge of the fuselage to enhance the course stability; perovskite solar cells with flexible substrates are distributed on the surfaces of the fuselage and the wings in an array mode and provide power for the unmanned aerial vehicle. The large-chord-length lifting body fuselage is adopted, the lifting area of the unmanned aerial vehicle is increased, meanwhile, the wingspan length of the unmanned aerial vehicle is reduced, loads borne by wings of the unmanned aerial vehicle are reduced, the weight of the wing structure is reduced, and effective loads are increased; the flexible perovskite solar cells are adopted for array, the damage rate of the cells in the installation process is reduced, and meanwhile due to the advantages of low cost, excellent flexibility, foldability, light weight and mass production of the perovskite solar cells, the service life of the unmanned aerial vehicle is prolonged, the cost is reduced, and mass production of the unmanned aerial vehicle is facilitated.

An embodiment aircraft boom (66, 76) includes a boom member (34, 42), a vertical stabilizer (40, 50), a rear landing gear (114, 116), and a front landing gear (110, 112). The boom member (34, 42) may extend in a first direction and have a front internal structural member (154) disposed therein, where a portion of the front internal structural member (154) is a first attachment point for a first wing. The vertical stabilizer (40, 50) may extend from the boom member (34, 42), where the vertical stabilizer (40, 50) has a rear internal structural member (120) disposed therein and extends from the boom member (34, 42) and through an interior of the vertical stabilizer (40, 50), and where a portion of the rear internal structural member (120) is a second attachment point for a second wing. The rear landing gear (114, 116) may be coupled to the rear internal structural member (120). The front landing gear (110, 112) may be coupled to the front internal structural member (154) of the boom member (34, 42).

The invention relates to the technical field of aircrafts, and discloses an arm construction method. The arm construction method is applied to the aircraft, the arm comprises an arm body and a direction control part, and the arm construction method comprises the steps that S1, the arm body and the direction control part are integrally connected, the direction control part extends outwards and is of a structure gradually reduced in the first direction, and the direction control part comprises a streamline surface structure; therefore, air resistance during flight of the aircraft is reduced. According to the vehicle arm construction method, the vehicle arm body and the vertical fin are integrally formed and connected, the stabilizer rigidity of the vertical fin can be improved, the loosening risk of the vertical fin is reduced, the number of parts is reduced, and the weight is reduced.

The utility model belongs to an airplane structural strength optimization design technology, and particularly relates to an H-shaped winglet integral vertical stabilizer box section which comprises a front beam, a rear beam, an upper end rib, a lower end rib, a middle rib and an outer side wall plate, wherein the upper end rib and the lower end rib are used for connecting the two ends of the front beam and the two ends of the rear beam, the middle rib is used for connecting the middle positions of the front beam and the rear beam, and the outer side wall plate covers one side of the integral vertical stabilizer box section. Wherein three rudder suspension joints used for being hinged to a rudder are arranged behind the rear beam, a vertical stabilizer mounting joint used for being fixedly connected with a horizontal stabilizer main beam and a horizontal stabilizer auxiliary beam is arranged on one side face of the middle rib, and a rudder actuator mounting joint used for being connected with an actuator is arranged on one side face of the middle rib. The problems that a plurality of parts are difficult to arrange in a small space and structural parts are difficult to connect are fundamentally solved, the number of the parts and fasteners is reduced, the structural weight is reduced, and the structural efficiency is improved.

The invention relates to the field of aircraft design, in particular to an aerodynamic layout structure of a light sports aircraft with ultra-short take-off and landing, which comprises a nose, a fuselage, wings, a vertical tail, a horizontal tail and an elevator, and when the fuselage extends from the nose to the tail, the shape of the fuselage changes from a streamline curved surface to a rectangle. The wings comprise a left main wing and a right main wing which are symmetrically designed and fixed on the fuselage, leading edge slats are also arranged at the leading edges of the left main wing and the right main wing, a fly unfolding aileron is also suspended below the trailing edges of the left main wing and the right main wing, a horizontal stabilizer of the horizontal empennage is arranged at the top, the elevator is rotationally arranged at the trailing edge of the horizontal empennage, and a balancing sheet is also arranged at the trailing edge of the elevator. The horizontal empennage is inverted, so that the pressure of the lower airfoil surface of the horizontal empennage is smaller than that of the upper airfoil surface, a downward pressure difference is further generated, the clockwise couple moment borne by the whole aircraft is increased, and the minimum ground clearance speed is reduced. The aircraft wing combines the advantages of the slats and the flaps, the landing appearance stall speed of an aircraft is effectively reduced, and therefore the landing taxiing distance is reduced.

The utility model discloses a connecting and positioning device for a horizontal stabilizer and an elevator hinge fulcrum of a wing-in-ground ship, which belongs to the technical field of wing-in-ground ship ground support systems and comprises two positioning mechanisms which are symmetrically arranged, the two positioning mechanisms are connected through a rotating shaft, each positioning mechanism comprises a supporting rod, a fixing clamp and a hinge fulcrum positioning piece, the upper parts of a plurality of first supporting rods which are arranged in parallel are vertically connected with a second supporting rod, the second supporting rod is fixedly connected with the first supporting rods through the fixing clamps, and the lower part of the second supporting rod is provided with a plurality of hinge fulcrum positioning pieces. According to the utility model, the purpose that multiple external hinge fulcrums of elevators on two sides of a horizontal stabilizer of a wing-in-ground-effect vehicle are coaxial after being assembled can be achieved.

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 invention discloses an aircraft stabilizer trim motor test processing method and system and a medium, and the method comprises the steps: obtaining trim motor test information, extracting operation parameter monitoring information and basic performance detection information, and respectively extracting locked-rotor torque data and vibration amplitude data according to the operation parameter monitoring information, judging whether the locked-rotor torque and the vibration amplitude of the balancing motor meet requirements or not, obtaining a motor performance index according to the basic performance detection information, performing processing in combination with a historical motor performance average index, obtaining a motor performance correction index, and determining whether the motor performance correction index meets the requirements or not. And finally, comparing with a preset performance index threshold value to judge whether the motor performance meets the requirement or not. According to the invention, important operation parameters of the balancing motor are monitored emphatically, deviation can be found in time and corresponding measures can be taken once deviation occurs, and meanwhile, the performance of the balancing motor is monitored and evaluated regularly, so that potential faults can be prevented, and the overall working reliability is improved. The invention also discloses a processing system and a medium.

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.

The invention provides a rudder connecting structure, and belongs to the field of aircraft structure design, the rudder connecting structure comprises: a box section suspension support arm fixed on a fuselage structure; the rudder support is fixed on the rudder; the connecting pin shaft is used for connecting the box section suspension support arm and the rudder support and is provided with a boss, one side of the boss is of a polished rod structure, the other side of the boss is of a threaded structure, the polished rod structure is connected with the box section suspension support arm, and the threaded structure is connected with the rudder support; the stabilizer suspension joint is fixedly arranged on the stabilizer structure, a stepped hole is formed in the stabilizer suspension joint, and a second bearing is arranged in the stepped hole; and the rudder hollow rotating shaft is fixed on the rudder, the side part of the rudder hollow rotating shaft is provided with a rudder hollow rotating shaft support, the end part of the rudder hollow rotating shaft axially extends to form a threaded structure, and the threaded structure penetrates through the second bearing and then is locked through a nut, so that the connection between the rudder hollow rotating shaft and the stabilizer suspension joint is realized.

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 invention relates to a load loading device for installing a movable airfoil deflection assessment test on a stabilizer. The load loading device comprises a force bearing frame, a test piece, a force bearing wall, a lateral supporting beam, a bottom supporting stand column, a rotary loading frame, a position control actuator cylinder and a movable airfoil loading actuator cylinder. The bearing frame is rectangular, the top of the bearing frame is connected with the stabilizer loading actuator cylinder, a test piece is arranged in the bearing frame and comprises a stabilizer and a movable wing surface, and the stabilizer loading actuator cylinder is downwards connected with the stabilizer and is perpendicular to the stabilizer; the load-bearing wall is connected to one end of the load-bearing frame, and the stabilizing plane is fixed on the load-bearing wall; the lateral supporting beam is connected to the stand column on one side of the force bearing frame, and the position control actuator cylinder is connected to the lateral supporting beam. The bottom end of the bottom supporting stand column is connected to the bottom of the force bearing frame and located below the test piece; the rotary loading frame is arranged in the force bearing frame, the bottom of the rotary loading frame is hinged to the top end of the bottom supporting stand column and connected with the position control actuator cylinder, the top of the rotary loading frame is connected with the movable wing surface loading actuator cylinder, and the movable wing surface loading actuator cylinder is downwards connected with the movable wing surface.

The utility model belongs to the field of aviation aircraft structures, and relates to a control surface rotation limiting structure of a freight unmanned aerial vehicle. The limiting structure comprises an elevator right connector, an elevator left connector, a horizontal stabilizer right support connector and a horizontal stabilizer left support connector. The elevator right joint and the elevator left joint are respectively arranged at two ends of the elevator front beam; the horizontal stabilizer right support connector and the horizontal stabilizer left support connector are both connected with an airplane horizontal stabilizer. The elevator right joint is rotatably connected with the horizontal stabilizer right support joint, and the elevator left joint is rotatably connected with the horizontal stabilizer left support joint; limiting ribs are arranged on the elevator right connector and the elevator left connector, and the limiting ribs limit rotation between the elevator and a horizontal stabilizer based on a limiting plane. The limiting structure is high in limiting capacity, the structure can be integrally machined and formed, and follow-up disassembly, assembly and maintenance are facilitated.

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.