Wing for balancing rotor reaction torque, and aircraft

Abstract

The present invention relates to the technical field of aircrafts. Disclosed are a wing for balancing rotor reaction torque, and an aircraft. Said wing comprises a support, a drive device (2) and a pair of blades (1) arranged in 180° central symmetry, wherein the support is configured to be mounted at the top of an aircraft (200) or close to the top; both the blades (1) are movably mounted at the end of the support, and both the blades (1) are located below a rotor (202) of the aircraft (200); and the drive device (2) is used to drive the two blades (1) to synchronously rotate in opposite directions, with the rotational axis of any one of the blades (1) being perpendicular to a rotary shaft of the rotor (202). The aircraft comprises a fuselage (201), a rotor (202), a flight control center, and said wing (100).

Description

Rotor anti-torque stabilizer and aircraft Technical Field

[0001] This invention relates to the field of aircraft technology, and in particular to a rotor anti-torque balance wing and an aircraft. Background Technology

[0002] Common aircraft, such as large manned and unmanned helicopters, small unmanned helicopters, rotorcraft flying toys, and models, are all equipped with rotor systems. Taking helicopters as an example, most helicopters use dual rotors, tail rotors, or tail nozzles to counteract rotor torque and ensure flight performance. Some small and medium-sized helicopters use a rear-mounted forward-pushing rotor combined with a vertical tail to increase forward speed while eliminating rotor torque. However, tail rotors, tail nozzles, and vertical tails all increase the aircraft's additional active energy consumption and add complex mechanical devices, which not only makes the aircraft structure complex and manufacturing cost high, but also results in significant power loss. To solve this problem, this invention proposes a novel rotor torque counterbalancing device. Summary of the Invention

[0003] The purpose of this invention is to provide a rotor anti-torque balance wing and an aircraft including the rotor anti-torque balance wing, which can directly utilize the downdraft generated by the rotor to counteract the anti-torque effect generated when the rotor rotates, so that the aircraft can get rid of dual rotors, tail rotors or tail nozzles, etc., which simplifies the aircraft structure, reduces manufacturing costs, and reduces flight energy consumption, thereby solving the problems existing in the background art.

[0004] To achieve the above objectives, the present invention provides the following solution:

[0005] This invention provides a rotor anti-torque balance wing, comprising a support frame, a drive device, and a pair of blades arranged 180° centrally symmetrically. The support frame is used to be installed on or near the top of an aircraft. Both sets of blades are movably mounted on the ends of the support frame, and both sets of blades are located below the rotor of the aircraft. The drive device is installed on the aircraft or the support frame, and the drive device is used to drive the two sets of blades to rotate synchronously and in opposite directions, and the rotation axis of any set of blades is perpendicular to the rotor axis.

[0006] In some embodiments, the support frame includes two main supports, which are symmetrically arranged on the left and right sides of the fuselage with the center line of the fuselage in the front-rear direction as the axis; two sets of blades are respectively installed at the mutually distant ends of the two main supports, and the rotation axes of the two sets of blades coincide.

[0007] In some embodiments, the outer contours of both main supports are designed with a streamlined shape.

[0008] In some embodiments, the drive device includes an active bevel gear, a drive linkage, and a rotary drive source. The active bevel gear is rotatably mounted on the aircraft or the support frame and is located on two main supports. A drive linkage is movably mounted in each of the two main supports, and the two adjacent ends of the two drive linkages mesh with the active bevel gear through a driven bevel gear. The two distant ends of the two drive linkages are connected to two sets of blades. The rotary drive source is connected to the active bevel gear and is used to drive the active bevel gear to rotate, so as to drive the two blades to rotate synchronously and in opposite directions through the two drive linkages.

[0009] In some embodiments, each of the two sets of blades is configured with one of the drive devices.

[0010] In some embodiments, any of the blades is a flat blade with a streamlined outer profile.

[0011] In some embodiments, any group of blades is a single-blade blade.

[0012] In some embodiments, any one of the blades is an aerodynamically adjustable grating, the grating including a grating rectification frame, a grating blade adjustment mechanism, and multiple strip-shaped grating blades. The grating rectification frame is movably mounted on the end of the support frame, and the middle part of the grating rectification frame is a hollow area. Multiple strip-shaped grating blades are installed in the hollow area, and both ends of any one strip-shaped grating blade are rotatably mounted on the grating rectification frame via a grating blade shaft. The multiple strip-shaped grating blades are arranged in parallel at intervals. The grating blade adjustment mechanism is disposed on the grating rectification frame, and the grating blade adjustment mechanism can adjust the tilt angle of each strip-shaped grating blade relative to the grating rectification frame to adjust the aerodynamic surface area of ​​the grating.

[0013] In some embodiments, the grating blade adjustment mechanism includes a linkage and a linkage drive mechanism. The linkage is located on one side of the grating blade rectifier frame and is parallel to the grating blade rectifier frame. The linkage is movably connected to the grating blade rectifier frame through multiple hinged links. One end of any hinged link is fixed to the corresponding grating blade shaft, and the other end of any hinged link is hinged to the linkage. The linkage drive mechanism drives the linkage to move parallel to the grating blade rectifier frame and to move the linkage closer to or away from the grating blade rectifier frame, so as to synchronously adjust the rotation angle of each grating blade shaft relative to the grating blade rectifier frame under the limiting action of the hinged links. The linkage drive mechanism is one of linkage drive mechanism one and linkage drive mechanism two, wherein:

[0014] The linkage drive mechanism includes a main drive rod, a rocker arm, and a rocker arm drive. The first end of the main drive rod is fixed to the corresponding grating blade shaft, and the second end of the main drive rod is hinged to the linkage. The rocker arm is fixedly connected to the second end of the main drive rod. The rocker arm drive is used to drive the rocker arm to rotate so as to drive the linkage to move parallel to the grating blade rectifier frame through the main drive rod.

[0015] The second linkage drive mechanism includes a lead screw, a fan-shaped vortex gear, and a lead screw drive. Multiple fixed sleeves are rotatably mounted on the lead screw, and any one of the fixed sleeves is fixed to the grating rectifier frame. The rotation center of the fan-shaped vortex gear is fixed to the corresponding grating blade shaft, and the fan-shaped vortex gear meshes with the lead screw. The lead screw drive is used to drive the lead screw to rotate, so as to drive the fan-shaped vortex gear to drive the linkage to move parallel to the grating rectifier frame.

[0016] The present invention also proposes an aircraft, including a fuselage, a rotor and a flight control center, and further including a rotor anti-torque balance wing as described in any of the above-mentioned embodiments. The support frame is installed at the top or near the top of the fuselage, and the orthographic projection of any of the blades on the rotor is located within the region of 1 / 2 to 3 / 4 of the radius length of the rotor from the inside to the outside.

[0017] The present invention achieves the following technical effects compared to the prior art:

[0018] The rotor anti-torque balancing wing proposed in this invention has a simple structure, few components, and advantages such as easy installation and low manufacturing cost. In application, while ensuring the stability of the aircraft's center of gravity, it is installed below the rotor. Under the downward wind pressure of the rotor, the adjusted tilted aerodynamic surface of the blades automatically eliminates the anti-torque effect of the rotor rotation on the flight device, achieving rotor anti-torque balancing without the need for tail rotors or tail nozzles. Furthermore, the two blades are centrally symmetrically arranged on the aircraft, and the downforce and forward drag generated by the two blades can cancel each other out. The aircraft does not need to make additional attitude adjustments, making piloting simpler and safer due to the absence of additional forces.

[0019] In some of the technical solutions disclosed in this invention, the entire device uses an electric motor as the drive source, which makes the structure and operation of the aircraft simpler, safer and more reliable. It not only reduces manufacturing costs, but also saves fuel and has the advantage of low energy consumption.

[0020] In some of the technical solutions disclosed in this invention, the main support and the outer contour of the blades are both designed with a flat and streamlined shape, which helps to reduce forward wind resistance and does not affect other functions of the aircraft.

[0021] In some of the technical solutions disclosed in this invention, the blades adopt gratings and are equipped with grating adjustment mechanisms, which can flexibly adjust the aerodynamic surface area of ​​the gratings to adapt to different aircraft or flight states.

[0022] In some of the technical solutions disclosed in this invention, the blades are arranged in a region of 1 / 2 to 3 / 4 of the rotor radius from the inside to the outside (but not limited to this region). This ensures that the blades are arranged close to the rotor without affecting the rotor's operation, so that the blades can receive effective wind pressure from the rotor regardless of the flight state of the aircraft, and amplify the leverage effect to ensure that the aircraft will not lose control.

[0023] The aircraft proposed in this invention includes the aforementioned rotor anti-torque balance wing and possesses all the features and advantages of the aforementioned rotor anti-torque balance wing, which will not be repeated here. Attached Figure Description

[0024] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0025] Figure 1 is a schematic diagram of the structure and installation of a rotor anti-torque balancer disclosed in one or more embodiments;

[0026] Figure 2 is a top view of Figure 1;

[0027] Figure 3 is a schematic diagram of the drive device in one or more embodiments of the rotor anti-torque balance wing;

[0028] Figure 4 is a schematic diagram of the structure of the first type of blade in a rotor anti-torque balancer disclosed in one or more embodiments;

[0029] Figure 5 is a schematic diagram of the linkage drive mechanism one of the blades in Figure 4;

[0030] Figure 6 is a schematic diagram of the second linkage drive mechanism for the blade in Figure 4;

[0031] Figure 7 is a schematic diagram of the structure of the second type of blade in a rotor anti-torque balancer disclosed in one or more embodiments;

[0032] Figure 8 is a schematic diagram of the structure of the third type of blade in a rotor anti-torque balancer disclosed in one or more embodiments;

[0033] Figure 9 is a schematic diagram of the balance principle of the rotor anti-torque balancer disclosed in one or more embodiments;

[0034] Figure 10 is a schematic diagram of the hydraulic drive system used in the grating blade adjustment mechanism.

[0035] In the diagram, the reference numerals are as follows: 100, rotor anti-torque balance wing; 200, aircraft; 201, fuselage; 202, rotor; 203, ECU controller; 1, blade; 11, single-blade; 12, grating; 121, grating rectifier frame; 122, strip grating blade; 1221, grating blade shaft; 123, grating blade adjustment mechanism; 1231, linkage rod; 1232, lead screw; 1233, fan-shaped spiral gear; 1234, transmission gear; 1235, fixed sleeve; 1236, main drive rod; 1237, articulated link; 2, drive device; 21, driving bevel gear; 22, drive link; 23, driven bevel gear; 24, rotary drive source; 3, main support; 4, forward and backward centerline; 5, traction wire. Detailed Implementation

[0036] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0037] One of the objectives of this invention is to provide a rotor anti-torque balancer that can directly utilize the downdraft generated by the rotor to counteract the anti-torque effect generated when the rotor rotates, so that the aircraft does not need to be equipped with additional counter-rotating wings, tail rotors or tail nozzles, which simplifies the aircraft structure, reduces manufacturing costs, and reduces flight energy consumption.

[0038] Another object of the present invention is to provide an aircraft comprising the above-described rotor anti-torque balance wing.

[0039] To make the above-mentioned objects, features and advantages of the present invention more apparent and understandable, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0040] Example 1

[0041] As shown in Figures 1 and 2, this embodiment provides a rotor anti-torque balance wing 100, including a support frame, a drive device 2, and a pair of blades 1 arranged 180° centrally symmetrically. The support frame is used to be installed on or near the top of the aircraft 200. The two sets of blades 1 arranged 180° centrally symmetrically are movably installed at the ends of the support frame, and the support frame and the two sets of blades 1 are located below the rotor 202 of the aircraft 200. The drive device 2 is installed in a parallel drive shaft of the aircraft 200 or the support frame. The drive device 2 is used to drive the two sets of blades 1 to rotate synchronously and in opposite directions, and the rotation axis of any blade 1 is perpendicular to the rotation axis of the rotor 202. It should be noted that the axis of rotation of rotor 202 can be defined as the height direction of fuselage 201. Although the aircraft 200 may have attitudes such as flipping or tilting during flight, the reference of the forward (tail) direction, height direction, and width direction (i.e. the left and right direction of fuselage 201) of the aircraft 200 itself remains unchanged. It is only that the orientation of the forward (tail) direction, height direction, and width direction (i.e. the left and right direction of fuselage 201) of the aircraft 200 will change under different attitudes.

[0042] In some embodiments, the aforementioned support frame includes two main supports 3, which are symmetrically arranged on the left and right sides of the fuselage 201 with the forward and backward centerline 4 of the fuselage 201 as the axis. Both main supports 3 are rod-shaped, and two blades 1 are respectively installed at the far ends of the two main supports 3, and the rotation axes of the two blades 1 coincide. In some specific embodiments, the two main supports 3 may have an angle of less than 180°, for example, both main supports are inclined towards the front of the fuselage 201 or both are inclined towards the rear of the fuselage 201; or, the two main supports 3 are coaxial, that is, they have an angle of 180° between them, as shown in Figure 2. In this case, the rotation axes of the two main supports 3 coincide with those of the blades 1, and the axes of the two main supports 3 and the rotation axes of the blades 1 are both perpendicular to the forward and backward centerline 4 and the rotation axis of the rotor 202. Generally, it is preferable that the two ends of the two main supports 3, which are far apart from each other, extend to the outside of the fuselage 201 to ensure that after the rotor anti-torque balance wing 100 is installed, its two blades 1 are located on the outside of the fuselage 201.

[0043] In some implementations, in order to ensure that the rotor anti-torque balance wing 100 does not affect the flight performance of the aircraft, it is preferable that the outer contours of both main supports 3 are streamlined, for example, the outer contours of the main supports 3 are set as circumferential surfaces with a smooth arc transition.

[0044] In some embodiments, to improve the integration of the rotor anti-torque balance wing 100 and simplify the equipment structure, it is preferable that the two blades 1 share a single drive device 2. Specifically, as shown in Figure 3, the drive device 2 includes an active bevel gear 21, a drive link 22, and a rotary drive source 24. The active bevel gear 21 is rotatably mounted on the aircraft 200 or the support frame, and the active bevel gear 21 is located between the two main supports 3 (i.e., at the midpoint of the support frame). A drive link 22 is movably installed in each of the two main supports 3, and the two ends of the two drive links 22 that are close to each other are respectively meshed with the active bevel gear 21 through a driven bevel gear 23. The two ends of the two drive links 22 that are far apart from each other are respectively connected to the two blades 1. In order to ensure the rotational stability of the drive link 22, it is preferable that each of the two main supports 3 is provided with a bearing seat to support the drive link 22, and a bearing is provided between the drive link 22 and the bearing seat to ensure the smoothness and reliability of the drive link 22 during rotation. The rotary drive source 24 is generally a motor, which can be directly connected to the driving bevel gear 21 via a coupling or via a reducer. The rotary drive source 24 drives the driving bevel gear 21 to rotate. After the driving bevel gear 21 rotates, it drives the two driven bevel gears 23 to rotate in opposite directions and synchronously. In turn, the two driven bevel gears 23 drive the two drive linkages 22 to rotate synchronously and in opposite directions, and the two drive linkages 22 drive the two blades 1 to rotate synchronously and in opposite directions. As shown in Figure 3, preferably, the driven bevel gears 23 and the driving bevel gear 21 are perpendicularly meshed. The bevel gear meshing assembly can not only realize power transmission, but also realize the change of power direction. As shown in Figure 3 of this embodiment, the driven bevel gears 23 and the driving bevel gears 21 are perpendicularly meshed, which can convert the rotational motion of the rotary drive source 24 with the vertical direction as the axis into the rotational motion with the horizontal direction as the axis (i.e., the direction of the rotation axis of the blade 1).

[0045] In some embodiments, the preferred support frame also includes a center seat for fixed installation with the fuselage 201. Two main brackets 3 are symmetrically installed on both sides of the center seat. The active bevel gear 21 can be installed on the center seat using a conventional bevel gear mounting method, which will not be described in detail here. The rotary drive source 24 can be directly installed on the center seat or installed at a corresponding position on the fuselage 201. As a preferred option, the rotary drive source 24 can be directly installed on the center seat and communicated with the controller (such as the ECU controller 203) of the flight control center of the aircraft. Directly installing the rotary drive source 24 on the center seat facilitates the quick and convenient disassembly and assembly of the rotor anti-torque balance wing 100 as a whole.

[0046] In some embodiments, an auxiliary support can be used to fix the main support 3 near the root position to the center seat, so that the main support 3, the auxiliary support and the center seat form a triangular support shape, ensuring that the installation structure of the main support 3 and the blade 1 is firm and stable.

[0047] In some embodiments, besides the structure described above where the two blades 1 share a single drive unit 2, each blade 1 can also be equipped with its own drive unit 2. In this case, the drive unit 2 can be a motor, which is embedded in the main support 3. It can be directly connected to the blade 1 via a coupling, or via a reducer, or via a chain drive mechanism, gear meshing mechanism, etc. Preferably, the blade 1 is rotatably mounted to the end of the main support 3, for example, through a bearing or roller rotatably engaging with the end of the main support 3.

[0048] In some embodiments, as shown in Figures 1 and 2, any blade 1 is preferably a flat blade with a streamlined outer contour. The overall shape of the flat blade includes, but is not limited to, a crescent shape, and its cross-sectional shape includes, but is not limited to, a curved willow leaf shape, a straight blade shape, etc., as long as it conforms to the morphology of aerodynamics.

[0049] Specifically, in some embodiments, any blade 1 is preferably an adjustable aerodynamic surface area grating 12, as shown in Figure 4. The grating 12 includes a grating rectifier frame 121, a grating blade adjustment mechanism 123, and multiple strip-shaped grating blades 122. The grating rectifier frame 121 is movably mounted on the end of the support frame, and the middle part of the grating rectifier frame 121 is a hollow area. Multiple strip-shaped grating blades 122 are installed in the hollow area, and both ends of any strip-shaped grating blade 122 are rotatably mounted on the grating rectifier frame 121 through a grating blade shaft 1221. Multiple strip-shaped grating blades 122 are arranged in parallel at intervals. The grating blade adjustment mechanism 123 is disposed on the grating rectifier frame 121. The grating blade adjustment mechanism 123 can adjust the tilt angle of each strip-shaped grating blade 122 relative to the grating rectifier frame 121 to adjust the aerodynamic surface area of ​​the grating 12.

[0050] In some implementations, each fin 12 requires multiple strip blades 122 to generate lateral force. The thickness of the fin 12 can be effectively reduced by increasing the number and spacing of the strip blades 122, thereby reducing the overall lateral drag of the fin 12. In lightweight, simple aircraft, drones, and flying toys where high flight speed is not a primary concern, a small number of strip blades 122, or even just one strip blade 122, can be used to achieve the rotor anti-torque balance effect.

[0051] In some embodiments, as shown in Figures 5 and 6, the aforementioned grating blade adjustment mechanism 123 includes a linkage 1231 and a linkage drive mechanism. The linkage 1231 is located on one side of the grating blade rectifier frame 121 and is parallel to the grating blade rectifier frame 121. The linkage 1231 is movably connected to the grating blade rectifier frame 121 through multiple hinged linkages 1237. Preferably, the number of hinged linkages 1237 is the same as the number of grating blades 12 (i.e., grating blade shafts 1221), and the two correspond one-to-one. One end of any hinged link 1237 is fixedly connected to the corresponding grating blade shaft 1221, and the other end of any hinged link 1237 is hinged to the linkage link 1231. The linkage drive mechanism is used to drive the linkage link 1231 to move parallel to the grating blade rectifier frame 121, and to move the linkage link 1231 closer to or further away from the grating blade rectifier frame 121, so as to synchronously adjust the rotation angle of each grating blade shaft 1221 relative to the grating blade rectifier frame 121 under the limiting action of the hinged link 1237. The above-mentioned linkage drive mechanism can be one of the following linkage drive mechanism one and linkage drive mechanism two, wherein:

[0052] (I) As shown in Figure 5, the linkage drive mechanism includes a main drive rod 1236, a rocker arm, and a rocker arm drive. The number of main drive rods 1236 is less than that of the grating blade 12 (i.e., the grating blade shaft 1221). They are evenly distributed along the axial direction of the linkage 1231. The first end of each main drive rod 1236 is fixed to the corresponding grating blade shaft 1221. The second end of each main drive rod 1236 is hinged to the linkage 1231. The second end of each main drive rod 1236 is fixedly connected to a rocker arm. The rocker arm drive is used to drive the rocker arm to rotate, so as to drive the linkage 1231 to move parallel to the grating blade rectifier frame 121 through the main drive rod 1236, thereby completing the rotation adjustment of the grating blade shaft 1221 relative to the grating blade rectifier frame 121. The aforementioned rocker arm drive can use a micro motor, such as a servo motor, which is mounted on the frame of the grid rectifier frame 121 or the outer wall of the main support 3. The micro motor can be directly connected to the end of the rocker arm through a coupling, or through a reducer, or through a chain drive mechanism, gear meshing mechanism, etc.

[0053] (II) As shown in Figure 6, the second linkage drive mechanism includes a lead screw 1232, fan-shaped vortex teeth 1233, and a lead screw drive. Multiple fixed sleeves 1235 are rotatably fitted onto the lead screw 1232, and the fixed sleeves 1235 are threadedly engaged with the lead screw 1232. Any one of the fixed sleeves 1235 is fixed to the grating blade rectifier frame 121. The number of fan-shaped vortex teeth 1233 is less than that of the grating blade 12 (i.e., the grating blade shaft 1221), and they are evenly distributed along the axial direction of the linkage 1231. Each fan-shaped vortex tooth... The rotation centers of the fan-shaped vortex teeth 1233 are all fixed to the corresponding grating blade shaft 1221, and each fan-shaped vortex tooth 1233 meshes with the lead screw 1232. The lead screw drive can be a micro motor, which is installed on the frame of the grating blade rectifier frame 121 or the outer wall of the main support 3. The micro motor can be directly connected to the end of the lead screw 1232 through a coupling, or through a reducer, or through a chain drive mechanism, gear meshing mechanism, etc. The motor is preferably a servo motor. When the motor is started, it can drive the lead screw 1232 to rotate and move the lead screw 1232 relative to the fixed sleeve 1235. Thus, the lead screw 1232 can drive the fan-shaped vortex teeth 1233 to rotate around their own rotation centers, and through the fan-shaped vortex teeth 1233, drive the linkage 1231 to move parallel to the grating blade rectifier frame 121, thereby completing the rotation adjustment of the grating blade shaft 1221 relative to the grating blade rectifier frame 121.

[0054] Both the aforementioned linkage drive mechanism one and linkage drive mechanism two are reliable and precise high-torque linkage devices, which can ensure the reliability of the angle adjustment of the strip grid blade 122 and the self-locking performance after the position adjustment is in place.

[0055] The following section uses a helicopter as an example to describe in detail the specific arrangement and operating principle of the aforementioned rotor anti-torque balance wing 100:

[0056] (I) Installation arrangement of rotor anti-torque balancer 100:

[0057] (1) The aerodynamic surface area and support frame shape of the required grid 12 can be calculated based on the weight of the helicopter, the rotor speed and the handling characteristics. The materials used for the grid 12 and the support frame are ensured to be safe after rigorous wind tunnel testing, and no vibration or resonance will be generated that will affect flight, so as to ensure that the maximum lateral moment generated by the grid 12 meets the standard. "Lateral" refers to the rotor rotation direction, i.e., the circumferential direction.

[0058] (2) In terms of the position arrangement, as shown in Figure 1, the center seat of the support frame can be set at the top center of the cockpit in front of the rotor shaft (i.e., the rotor shaft) (i.e., the central axis of the center seat intersects perpendicularly with the aforementioned front-rear center line 4), and the center seat of the support frame is close to the rotor shaft, so as to ensure that after the rotor anti-torque balance wing 100 is installed, the entire helicopter can achieve natural balance of the center of gravity under the sufficient downforce generated by the rotor in the hovering attitude.

[0059] (3) The length of the main support 3 in the support frame and the height difference between the rotor anti-torque balance wing 100 and the rotor after installation should ensure that the rotor anti-torque balance wing 100 can always be in the position with the best verticality of the rotor's downward wind and sufficient wind force, and will not interfere with the rotor in a stationary state. Based on this, it is generally preferred to be below the rotor radius from the inside to the outside 1 / 2 to 3 / 4, that is, the orthogonal projection of any blade 1 on the rotor 202 is located in the region from the inside to the outside 1 / 2 to 3 / 4 of the radius length of the rotor 202. Based on this, even if an unexpected vortex effect occurs, the rotor anti-torque balance wing 100 can still function normally, and this position also makes full use of the leverage effect, which can effectively amplify the lateral force and minimize the aerodynamic surface area of ​​the grid wing 12. However, it is necessary to prevent vibration and resonance effects, which can be assisted by setting traction steel wire 5 to assist in the stable installation of the support frame. As shown in Figure 2, the front and rear sides of the outer end of the main support 3 are connected to the machine body 201 by a traction steel wire 5 to ensure the stable installation of the support frame as a whole.

[0060] (II) Automatic adjustment logic of rotor anti-torque balancer 100:

[0061] (1) A minimum initial angle is usually set according to the force of the slat 12, which can automatically counteract the rotor counter-torque generated by the helicopter in the unloaded hovering state. This initial angle should not be too large, as shown in Figure 5, and can usually be kept within 30 degrees of the vertical line (the vertical line is perpendicular to the aerodynamic surface of the slat 12). If this angle is too large, it will restrict the adjustment space of each strip slat 122 and affect the maximum maneuverability of the helicopter. The aircraft will only adjust this initial angle in the opposite direction (i.e., increase the adjustment) when in an emergency.

[0062] (2) The helicopter's built-in flight control center uses conventional technology. The rotor anti-torque balance wing 100 can communicate with the flight control center through the ECU controller 203. The blade adjustment mechanism and drive device 2 of the slat wing 12 are generally connected to the ECU controller 203. Based on this, the ECU controller 203 can collect information from the helicopter's original orientation sensor (geomagnetic pointing system), load sensor (aircraft takeoff weight), speed sensor (rotor or engine output shaft), gyroscope sensor (electronic or non-electronic), control commands (pilot manual control system), etc., to comprehensively judge the helicopter's flight attitude or the flight attitude to be achieved. Then, through the drive device 2, it automatically adjusts the tilt angle of the two centrally symmetrical slat wings 12, so that the two slat wings 12 generate a force opposite to the rotor anti-torque, thereby counteracting the helicopter's anti-torque torque and making the helicopter run smoothly.

[0063] When the helicopter is ready to take off, the ECU controller 203 controls the drive unit 2 by collecting the detection signals from the load sensor and the speed sensor to adjust the overall tilt angle of the two slats 12. At the moment of takeoff, the ECU controller 203 finely adjusts the overall tilt angle of the two slats 12 by using the gyroscope sensor and the pull command of the rotor control stick to make the circumferential force of the helicopter completely balanced.

[0064] (3) Since the two slats 12 in the rotor anti-torque balance wing 100 are arranged in a centrally symmetrical manner, the downforce and forward drag of the two slats 12 can completely cancel each other out, and the helicopter does not need to make additional attitude adjustments. In practical applications, the rotor anti-torque balance wing 100 can even incorporate most of the existing tail rotor control logic for control. The rotor anti-torque balance wing 100 does not need to be adjusted most of the time during helicopter flight, and only adjusts rapidly when the rotor speed changes significantly, the rotor control stick is pulled up or down significantly, or the helicopter is rotating horizontally rapidly. For example, when the ECU controller 203 receives a command to pull up the rotor control stick significantly, it can release the brake on the motor in the drive unit 2 in advance and start the motor in the drive unit 2 to slowly adjust the overall tilt angle of the slats 12, so that the aircraft can maintain stability to the greatest extent. The rotor anti-torque balance wing 100 can determine the attitude of the helicopter based on the geomagnetic sensor when the helicopter loses its main force and participate in emergency adjustments to the greatest extent. A geomagnetic sensor is generally a precision gyro-shaped compass installed on the dashboard. Unlike ordinary compasses, it has a signal data output line and serves as the last line of defense in case other orientation devices fail.

[0065] In practical applications, the two grids 12 of the rotor anti-torque balance wing 100 can also be arranged in front of or behind the fuselage 201, or a single grid 12 can be placed in front of or behind the fuselage 201. However, such arrangements will pass through the engine and rotor shaft, affecting the internal structure of the fuselage, or have to deal with the high temperature problem of the engine. Setting a single grid 12 will also cause the fuselage to suffer unbalanced downforce during flight, making control more complicated and prone to loss of control.

[0066] In summary, the rotor anti-torque balancer 100 proposed in this solution has the following main advantages:

[0067] (1) The entire rotor anti-torque balance wing 100 has a simple structure and few parts, and has the advantages of easy installation and low manufacturing cost. When applied, under the premise of ensuring the stability of the aircraft's center of gravity, it is installed directly below the rotor and in front of the rotor main shaft. Under the action of the rotor's downward wind pressure, the adjusted inclined aerodynamic surface of the grid wing 12 can automatically eliminate the anti-torque effect of the rotor rotation on the flight device. There is no need to design a tail rotor or tail nozzle, and the purpose of counter-torque balance of the rotor can be achieved.

[0068] (2) This scheme adopts a pair of grid wings 12 and is arranged symmetrically on the left and right sides of the aircraft fuselage. In this way, the downforce and forward drag generated by the two grid wings 12 can cancel each other out, and the helicopter does not need to make additional attitude adjustments. Since no additional force is generated, the piloting is simpler and safer.

[0069] (3) The entire device uses ECU and servo motor as the drive source, which makes the aircraft structure and operation simpler, safer and more reliable. It not only reduces manufacturing costs, but also saves fuel and has the advantage of low energy consumption.

[0070] (4) The main support and outer contour of the grid wing in this scheme adopt a streamlined design, which helps to reduce forward wind resistance and does not affect other functions of the aircraft.

[0071] (5) A grating blade adjustment mechanism is configured in the grating 12, which can flexibly adjust the aerodynamic surface area of ​​the grating 12 to adapt to different aircraft or flight states.

[0072] (6) All moving parts of the transmission and linkage device are equipped with rollers or bearings to keep the operation smooth and quiet, which can effectively reduce the energy consumption of the drive.

[0073] (7) Arranging the grid 12 within 1 / 2 to 3 / 4 of the rotor radius from the inside out ensures that it is positioned close to the rotor without affecting rotor operation. This allows the grid 12 to receive effective wind pressure from the rotor regardless of the aircraft's flight status, amplifying the leverage effect and ensuring that the aircraft does not lose control. In addition, this arrangement minimizes interference with the loading and unloading of personnel and cargo, as well as the maintenance and upkeep of the aircraft's engines.

[0074] (8) The rotor anti-torque balance wing 100 can be configured with dual ECUs to ensure that if one ECU controller fails, the other ECU controller can take over in time, and an emergency manual rocker arm control system can be set up.

[0075] Example 2

[0076] As shown in Figure 7, this embodiment proposes a rotor anti-torque balance wing 100, which differs from Embodiment 1 only in that any blade 1 is a single blade 11. The single blade 11 is mostly a plate-shaped structure with a constant aerodynamic surface area and no hollowing.

[0077] The installation arrangement, working principle, and effect of the rotor anti-torque balance wing 100 in this embodiment are similar to those in Embodiment 1, and will not be repeated here.

[0078] Example 3

[0079] As shown in Figure 8, this embodiment proposes a rotor anti-torque balance wing 100, which differs from Embodiment 1 in that the main support 3 adopts a square frame, and the end of the support frame is provided with a grid wing 12.

[0080] Example 4

[0081] In Embodiment 1 or 2, the aircraft's existing hydraulic system can be used as the drive source on heavy or super-heavy helicopters. The hydraulic system drive is arranged on the wing frame, as shown in Figure 10. The electromagnetic hydraulic control valve system 31 is communicatively connected to the ECU controller 203. The electromagnetic hydraulic control valve system 31 is connected to a hydraulic inlet pipe 33 and a hydraulic distribution pipe 34. The hydraulic distribution pipe 34 is connected to a hydraulic telescopic rod 32. The hydraulic telescopic rod 32 is connected to either the aforementioned linkage drive mechanism one or linkage drive mechanism two, thereby realizing the adjustment of the rotation angle of the strip wing 122. Specifically:

[0082] When the hydraulic telescopic rod 32 is connected to the aforementioned linkage drive mechanism, the end 35 of the hydraulic telescopic rod 32 can be directly hinged to the second end of the main drive rod 1236 so as to control the rotation of the main drive rod 1236 through the extension and retraction of the hydraulic telescopic rod 32.

[0083] When the hydraulic telescopic rod 32 is connected to the above-mentioned linkage drive mechanism 2, a rack can be used to replace the lead screw 1232. The rack is slidably installed with the fixed sleeve 1235, and the rack meshes with the fan-shaped volute 1233. The end 35 of the hydraulic telescopic rod 32 is directly the same as the end of the rack, and is used to drive the rack to move in parallel, thereby driving each fan-shaped volute 1233 to rotate.

[0084] Example 5

[0085] This embodiment proposes a rotor anti-torque balance wing 100, which can be designed as a foldable structure and can be stored in the small hangar of aircraft such as shipborne helicopters when not in use.

[0086] Specifically, the main support 3 can be configured as a structure that can fold forward, backward, upward or downward relative to the aircraft fuselage, such as a bi-fold support or a tri-fold support. Its volume can be reduced by folding the rotor anti-torque balance wing 100, thereby achieving storage.

[0087] Example 6

[0088] As shown in Figures 1 and 2, this embodiment proposes an aircraft 200, including a fuselage 201, a rotor 202, and a flight control center. It also includes a rotor anti-torque balancer 100 disclosed in Embodiment 1 or Embodiment 2. The support frame is installed on or near the top of the fuselage 201, and the orthographic projection of any blade 1 onto the rotor 202 lies within the radius of the rotor 202, extending from the inside to the outside in 1 / 2 to 3 / 4 of its length. This design ensures that the rotor anti-torque balancer 100 is positioned close to the rotor without affecting rotor operation, allowing the blade 1 to receive effective wind pressure from the rotor regardless of the aircraft's flight state, amplifying the leverage effect and preventing the aircraft from going out of control. Furthermore, this arrangement minimizes disruption to personnel and cargo loading / unloading, and to activities such as engine maintenance.

[0089] Specific examples have been used to illustrate the principles and implementation methods of this invention. The descriptions of the above embodiments are only for the purpose of helping to understand the method and core ideas of this invention. Furthermore, those skilled in the art will recognize that, based on the ideas of this invention, there will be changes in the specific implementation methods and application scope. Therefore, the content of this specification should not be construed as a limitation of this invention.

Claims

1. A rotor counter torque balance wing characterized by, include: Support frame for mounting on or near the top of the aircraft (200); A pair of blades (1) arranged symmetrically at 180° center, both sets of blades (1) are movably mounted at the end of the support frame, and both sets of blades (1) are located below the rotor (202) of the aircraft (200); A drive device (2) is installed on the aircraft (200) or the support frame. The drive device (2) is used to drive the two sets of blades (1) to rotate synchronously and in opposite directions, and the rotation axis of any set of blades (1) is perpendicular to the rotation axis of the rotor (202).

2. The rotor anti-torque balancer according to claim 1, characterized in that, The support frame includes two main supports (3), which are symmetrically arranged on the left and right sides of the fuselage (201) with the center line (4) of the front and rear direction of the fuselage (201) as the axis; two sets of blades (1) are respectively installed at the two far apart ends of the two main supports (3), and the rotation axes of the two sets of blades (1) coincide.

3. The rotor anti-torque balancer according to claim 2, characterized in that, Both main supports (3) have a streamlined outer contour design.

4. The rotor anti-torque balancer according to claim 2 or 3, characterized in that, The driving device (2) includes: An active bevel gear (21) is rotatably mounted on the aircraft (200) or the support frame, and the active bevel gear (21) is located on the two main supports (3); Drive link (22), one drive link (22) is movably installed in each of the two main brackets (3), and the two ends of the two drive links (22) that are close to each other are respectively meshed with the active bevel gear (21) through a driven bevel gear (23), and the two ends of the two drive links (22) that are far apart from each other are respectively connected to two sets of blades (1). A rotary drive source (24) is connected to the active bevel gear (21). The rotary drive source (24) is used to drive the active bevel gear (21) to rotate so as to drive the two blades (1) to rotate synchronously and in opposite directions through the two drive linkages (22).

5. The rotor anti-torque balancer according to claim 1, characterized in that, Each of the two sets of blades (1) is equipped with a drive device (2).

6. The rotor anti-torque balancer according to claim 1, characterized in that, Any of the blades (1) described herein is a flat blade with a streamlined outer profile.

7. The rotor anti-torque balancer according to claim 6, characterized in that, Any set of blades (1) is a single-piece blade (11).

8. The rotor anti-torque balancer according to claim 6, characterized in that, Any one of the blades (1) is an aerodynamically adjustable slat (12), the slat (12) comprising: A grille rectifier frame (121) is movably mounted at the end of the support frame, and the middle part of the grille rectifier frame (121) is a hollow area; Multiple strip-shaped grating blades (122) are installed in the hollow area, and both ends of any one of the strip-shaped grating blades (122) are rotatably mounted on the grating wing rectifier frame (121) through the grating blade shaft (1221). The multiple strip-shaped grating blades (122) are arranged in parallel at intervals. A blade adjustment mechanism (123) is disposed on the grating wing rectifier frame (121). The blade adjustment mechanism (123) can adjust the tilt angle of each strip blade (122) relative to the grating wing rectifier frame (121) to adjust the aerodynamic surface area of ​​the grating wing (12).

9. The rotor anti-torque balancer according to claim 8, characterized in that, The grating blade adjustment mechanism (123) includes a linkage (1231) and a linkage drive mechanism, wherein: The linkage link (1231) is located on one side of the slat rectifier frame (121) and is parallel to the slat rectifier frame (121). The linkage link (1231) is movably connected to the slat rectifier frame (121) through multiple hinged links (1237). One end of any one of the hinged links (1237) is fixed to the corresponding slat blade shaft (1221), and the other end of any one of the hinged links (1237) is hinged to the linkage link (1231). The linkage drive mechanism is used to drive the linkage link (1231) to move parallel to the slat rectifier frame (121), and to move the linkage link (1231) closer to or further away from the slat rectifier frame (121), so as to synchronously adjust the rotation angle of each slat blade shaft (1221) relative to the slat rectifier frame (121) under the limiting action of the hinge link (1237); the linkage drive mechanism adopts one of linkage drive mechanism one and linkage drive mechanism two, wherein: The linkage drive mechanism includes a main drive rod (1236), a rocker arm, and a rocker arm drive. The first end of the main drive rod (1236) is fixed to the corresponding grating blade shaft (1221), and the second end of the main drive rod (1236) is hinged to the linkage rod (1231). The rocker arm is fixedly connected to the second end of the main drive rod (1236). The rocker arm drive is used to drive the rocker arm to rotate so that the linkage rod (1231) moves parallel to the grating blade rectifier frame (121) through the main drive rod (1236). The second linkage drive mechanism includes a lead screw (1232), a fan-shaped vortex tooth (1233), and a lead screw drive. Multiple fixed sleeves (1235) are rotatably mounted on the lead screw (1232), and any one of the fixed sleeves (1235) is fixed to the grating rectifier frame (121). The rotation center of the fan-shaped vortex tooth (1233) is fixed to the corresponding grating blade shaft (1221), and the fan-shaped vortex tooth (1233) meshes with the lead screw (1232). The lead screw drive is used to drive the lead screw (1232) to rotate, so as to drive the fan-shaped vortex tooth (1233) to drive the linkage (1231) to move parallel to the grating rectifier frame (121).

10. An aircraft, comprising a fuselage (201), rotors (202), and a flight control center, characterized in that, It also includes a rotor anti-torque balance wing (100) as described in any one of claims 1 to 9, wherein the support frame is installed at or near the top of the fuselage (201), and the orthographic projection of any one of the blades (1) on the rotor (202) is located within the region from the inside to the outside of the radius length of the rotor (202) in a range of 1 / 2 to 3 / 4.

Images