41 results about "Stabilator" patented technology

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.

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.

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.

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 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.

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.

The invention discloses an aircraft vertical stabilizer repeated impact resistance testing device, which relates to the technical field of performance testing equipment and comprises a supporting mechanism, a clamp mechanism and an impact mechanism. The supporting mechanism is provided with a base and a top cross beam, the base is provided with a containing groove, and the top cross beam is located above the containing groove; the clamp mechanism is fixed in the containing groove, two opposite clamping devices are arranged on the clamp mechanism, and the clamping devices are fixed to the clamp mechanism and provided with clamping gaps for clamping the test piece; a fixed flexible block is arranged at the lower end of the gap, and a movable flexible block is arranged at the upper end of the gap; the impact mechanism comprises a base, a drop hammer plate and two guide columns, the base is fixed above the containing groove, and the guide columns are fixed to the base in parallel; the sleeve assembly of the drop hammer plate vertically slides along the guide column, the hammer head is detachably fixed on the sleeve assembly, and the base is provided with an avoiding hole for the hammer head to penetrate through the test piece; the guide columns are sleeved with buffer springs, so that the hammer head can be prevented from impacting the test piece when the drop hammer plate rebounds. Damage to a test piece can be reduced, and the test piece is prevented from being interfered by secondary impact.

This utility model relates to the field of aircraft control surface clearance detection technology, specifically to an adaptive clamping device for detecting the clearance of aircraft main control surfaces. It includes a clamping mechanism, a loading clamping mechanism, a movable surface clamping mechanism, a control box, and a control host. The clamping mechanism and the loading clamping mechanism are connected to form a ring structure. The movable surface clamping mechanism is located inside the ring structure, and its inner side is equipped with several lifting block assemblies and pressure sensors that conform to the stabilizer surface. The loading clamping mechanism is equipped with a displacement adjustment assembly for adjusting the height of the movable surface clamping mechanism. The movable surface clamping mechanism is U-shaped, and its inner side is equipped with control surface clamping assemblies that conform to the movable surface. Pressure sensors are located on the contact surfaces. The displacement adjustment assembly and each pressure sensor are connected to the control host via the control box. This utility model uses force loading to cause the movable surface to rotate, and obtains the free clearance between the movable surface and the stabilizer surface by measuring the rotation angle, thereby improving the accuracy and reliability of the clearance measurement results.

The application provides a method for improving the docking efficiency of vertical stabilizer and rudder products, which is used for the assembly of tail wing large components made of composite materials of domestic large aircraft, realizes the rapid docking of the rudder assembly and the vertical tail assembly in the flight attitude under the rudder hanging point connection state, and belongs to the technical field of aircraft assembly. The pressure maintaining gas tank on the lifting clamp and the lifting connection mechanism ensure that the lifting joint can be connected, the design form that does not need to be connected to the gas source at all times ensures the flexibility of the lifting clamp movement, a storage battery is installed on the lifting device, and the angle of the product can be adjusted at any time by using a motor device and a wireless remote control device. The method of using the cylindrical bearing stop device improves the docking efficiency and protects the parts from damage, and the docking work can be completed by 4 people in 2 hours, and the efficiency is improved by more than 100%. It belongs to a semi-automatic docking scheme.

A fixing device for a disassembled general aircraft comprises a transportation frame, a fuselage fixing assembly, a wing fixing assembly arranged on a frame bottom plate, an engine fixing assembly arranged on the frame bottom plate and a horizontal stabilizer fixing assembly arranged on the frame bottom plate. The transportation frame is provided with a frame bottom plate. The machine body fixing assembly comprises a machine body fixing base part arranged on the frame bottom plate. The central wing fixing assembly comprises a first supporting piece arranged at one end of the fuselage fixing base piece at a first preset angle, and a second supporting piece arranged at the top of the first supporting piece. The central wing fixing piece is arranged at one end of the fuselage fixing base piece and forms a second preset angle with the fuselage fixing base piece; the fuselage tail boom fixing assembly is arranged at the other end of the fuselage fixing base piece and forms a third preset angle; and the fuselage tail boom fixing piece is arranged at the top of the second supporting piece and forms a fourth preset angle with the central wing fixing piece. And parallel fixation of the disassembly parts of the general aircraft is realized. And the occupied space of the fuselage of the general aircraft in the horizontal direction is reduced.

The invention discloses a test method for dynamically simulating the support stiffness of a control surface. The method comprises the following steps: S1, constructing a control surface-horizontal stabilizer finite element model; s2, carrying out a hinge constraint feasibility test based on the control surface-horizontal stabilizer finite element model; and S3, performing control surface test constraint and loading design based on a result of the hinge constraint feasibility test. According to the invention, accurate simulation of the elastic support of the body can be realized.

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 invention belongs to the field of flight control system design, and particularly relates to a modeling method for a horizontal stabilizer actuator of a large aircraft. The method comprises the following steps: S1, determining the working principle of the horizontal stabilizer actuator; s2, determining key parameters of the horizontal stabilizer actuator according to a working principle; s3, constructing a test environment, and obtaining test data of the key parameters; and S4, constructing a horizontal stabilizer actuator model according to the key parameters determined based on the test data. According to the modeling method for the horizontal stabilizer actuator of the large aircraft, it can be ensured that the horizontal stabilizer actuator model truly reflects the motion characteristics of the actuator, and the credibility of a flight control system simulation verification result is improved.

The utility model belongs to the technical field of aircraft manufacturing, and discloses an elevator positioning tool which comprises a cross beam and a plug block, the cross beam is provided with a plurality of positioning assemblies, the positioning assemblies are matched to enable the cross beam to be positioned relative to a horizontal stabilizer of an empennage, and one end of the cross beam extends to the lower portion of the edge of an elevator and is provided with a supporting piece. An inspection gap is formed between the supporting piece and the skin of the elevator; the plug block is in lap joint with the supporting piece and is contained in the inspection gap. Wherein the plug block has various height sizes. The deviation range of the elevator can be judged and adjusted by observing and adjusting the gaps between the elevator and the plug blocks with different heights, then the elevator is positioned within the allowable error range near the zero position, the positioning precision of the elevator is improved, meanwhile, the elevator positioning tool is simple in structure and convenient to operate, the labor cost of elevator positioning can be reduced, and the working efficiency is improved. The assembling efficiency is improved, and efficient and high-quality assembling is achieved.

This utility model discloses a rapid anti-sway and stabilization flexible carrier for multi-rotor drones, including a connecting unit. The connecting unit further includes a flexible connecting tube, a clamp assembly, a guide vane, and a locking body docking unit. The flexible connecting tube is multi-sectioned, and the docking portions of adjacent flexible connecting tubes are respectively equipped with an inner ring and a concave plate. This rapid anti-sway and stabilization flexible carrier for multi-rotor drones uses guide vanes to form a stabilizing surface (similar to the tail fin of a missile or arrow). The strong airflow generated by the multi-rotor drone's propellers passes through the tail fin, quickly stabilizing its attitude and achieving the purpose of anti-sway and stabilization.

The utility model discloses a kind of unmanned aerial vehicle for realizing vertical take-off and attitude control using small turbojet engine, small turbojet engine is installed at wing tip of wing, small turbojet engine is installed at the back of fuselage (or in front of fuselage) two sides;Small turbojet engine is made to produce upward thrust by vertical rotation 90 ° using turbine worm drive mechanism or motor with self-locking holding torque, and unmanned aerial vehicle vertical take-off is realized;Unmanned aerial vehicle has aileron installed at the trailing edge of left and right wing tips, and the differential deflection of aileron makes unmanned aerial vehicle tilt rotation around longitudinal axis, and the direction of thrust of small turbojet engine is deflected using turbine worm drive mechanism, to realize the flight state stabilization and flight attitude control of unmanned aerial vehicle, so that unmanned aerial vehicle has good maneuverability;Horizontal stabilizer and elevator are omitted, vertical stabilizer and rudder are omitted, so as to reduce flight resistance, which is conducive to improving the flight speed and effective load of unmanned aerial vehicle.

This invention relates to longitudinal trim control motion during takeoff and pull-up. This technology will allow takeoff from a single initial horizontal stabilizer or trim tab position while maintaining a satisfactory pull-up time, thus allowing for simpler aircraft operation and avoiding scenarios where the crew has not properly trimmed the aircraft (trim failure takeoff scenario) that could reduce safety margins.

The utility model discloses a double-slotted flap mechanism of a light unmanned aerial vehicle. The double-slotted flap mechanism comprises a main wing stabilizer, a steering engine, a connecting rod mechanism, a first-stage flap and a second-stage flap, the main wing stabilizer is fixedly connected with the steering engine in an embedded manner; the connecting rod mechanisms are distributed in a bilateral symmetry mode and are in transmission connection with the two ends of the main wing stabilizing plane, the steering engine, the first-stage flap and the second-stage flap in a hinged mode; the first-stage flap and the second-stage flap synchronously carry out retreating and downward-deviating composite motion along a preset track, a double-slotted pneumatic structure is formed, the boundary layer control capacity is effectively enhanced, airfoil airflow separation is delayed, the lift coefficient is increased, and especially under the low-speed short-distance take-off and landing working condition, the flow field regulation and control efficiency is remarkably superior to that of a traditional single-stage flap structure. By applying light materials and embedded installation design, the weight of the whole mechanism is much lighter than that of a traditional metal structure, meanwhile, the guide rails penetrating through the driving rods in the spanwise direction and the airfoil body are integrally arranged, the occupied space of the mechanism is greatly reduced, and the requirement for the compact aerodynamic shape of a light and small unmanned aerial vehicle is met.

The utility model discloses a kind of unmanned aerial vehicle for realizing vertical take-off and attitude control using rotor steering, there are two across installation in left and right wings in flight direction on hanger, hanger front end and rear end are equipped with rotors and power device that can be steered;When horizontal flight, front rotor generates pull, rear rotor generates thrust;Front rotor and power device vertically upward rotate 90 ° and generate upward pull, rear rotor and power device downward rotate 90 ° and generate upward thrust, realize unmanned aerial vehicle vertical take-off;Unmanned aerial vehicle flight attitude is controlled using deflection combination of front two rotors and rear two rotors, and using deflection combination of same side hanger front and rear rotors;Make unmanned aerial vehicle have good maneuverability, save horizontal stabilizer and elevator, save vertical stabilizer and rudder, thereby reduce flight resistance, it is favorable to improve the effective load of unmanned aerial vehicle.