Aircraft tail lift rotor system and pitch trim control method thereof

Abstract

This invention discloses a horizontal stabilizer lifting rotor system and its pitch trim control method for an aircraft, belonging to the field of aircraft overall layout. The system includes lifting rotors located on the left and right sides of the horizontal stabilizer. The lifting rotors are single-rotor or coaxial dual-rotor systems. The tail-mounted lifting rotor assembly uses the horizontal stabilizer as its mounting reference plane. Its lowest drive motor is located on the upper surface of the horizontal stabilizer or embedded within it. The mounting path does not pass through the mast, tailhook, or cantilever components of the vertical stabilizer. The leading edge of the horizontal stabilizer is not lower than its trailing edge, providing partial rear positive lift during cruise. The rotor shaft is tilted in the same direction relative to the fuselage longitudinal axis, and the mounting angle is set so that the rotor disk's angle to the ground is zero or close to zero during cruise trim. The mounting angle of the lifting rotor assembly relative to the fuselage remains unchanged during mode transitions. In cruise mode, the lifting rotor assembly remains operational, generating a pitch moment on the longitudinal lever arm of the rotor disk center relative to the center of gravity through synchronous increases and decreases in thrust from both sides, thus undertaking pitch trim and control.

Description

Technical Field

[0001] This invention relates to the field of aircraft overall layout and flight control technology, specifically to a horizontal stabilizer lifting rotor system for an aircraft in which the lifting rotor is mounted on the upper surface of the horizontal stabilizer as the mounting reference surface and is positioned on the left and right sides of the horizontal stabilizer, as well as a pitch trim control method for the aircraft. Background Technology

[0002] Composite aircraft with both vertical takeoff and landing (VTOL) and fixed-wing cruise capabilities typically have their lifting rotors mounted on the fuselage arms, upper wing surfaces, or above the motor nacelles. In cruise mode, these rotors are mostly stationary and do not perform pitch trim functions; pitch trim is handled by aerodynamic control surfaces such as the horizontal stabilizer and elevators. This presents two problems: First, if the tail section needs to provide both controllable lift and pitch trim torque simultaneously, separate lift and trim devices are required, increasing the weight and complexity of the tail section due to the two sets of structures and actuations. Second, if the tail lift device is mounted cantilevered behind or through the vertical stabilizer using a mast or tailhook, the motors are far from the load-bearing structure, resulting in a significant bending moment at the mounting base due to the rotor disk thrust, necessitating additional load-bearing components.

[0003] Furthermore, the conventional horizontal stabilizer has a downward angle of attack, with its leading edge pitched down, providing negative lift in cruise mode and resulting in a net loss of lift for the entire aircraft. In a tail configuration where the center of gravity is located behind the wings, the point of application of tail lift has a long longitudinal lever arm from the center of gravity, making the thrust required for pitch trim costly. While adding a tilt mechanism to change the direction of the propeller disk or using a pitch control mechanism can achieve changes in the direction of the propeller disk and thrust, these introduce actuators such as servos, tilt shafts, or pitch mechanisms, increasing structural weight, failure modes, and maintenance burden.

[0004] Existing technologies have disclosed configurations in which a propeller is mounted on a horizontal stabilizer. The lift propeller is positioned on the horizontal stabilizer and pitches along with it. The drive motor is supported behind the horizontal or vertical stabilizer by a mast, tailhook, or cantilever structure passing through the vertical stabilizer. In cruise mode, the propeller disk tilts along with the overall aircraft attitude, and the disk's angle to the ground is not horizontal. Existing technologies have also disclosed concepts of using the thrust of a lift rotor or propeller to perform pitch trim and control, replacing the elevator function, including multi-rotor aircraft that adjust the overall aircraft attitude by increasing or decreasing the thrust of rotors at different positions. Furthermore, prior applications have disclosed layouts such as a tilt propeller at the end of the horizontal stabilizer, a tail tilt propeller providing lift and pitch trim, and a T-tail with a tail tilt mechanism. Vertical takeoff and landing (VTOL) aircraft also have configurations where the VTOL power unit is mounted on a tail boom or inverted V-tail structure. Therefore, the concepts of placing the lifting rotor on the horizontal stabilizer and using the thrust of the lifting rotor to undertake pitch trim are already public knowledge; the difference of this invention lies not in the presence or absence of the above-mentioned functional concepts, but in the specific structural relationships described below and their coordination with the mode conversion method.

[0005] Furthermore, eliminating the ailerons and elevators simplifies the wing trailing edge and eliminates trailing edge hinge gaps, but the pitch and yaw channels need to be handled by other components, requiring a tail layout that can structurally support this channel allocation. Therefore, a tail-mounted lift rotor system is needed that is structurally simple, does not rely on tilt and pitch mechanisms for mode transitions, and whose configuration on the same mounting base simultaneously performs lift and pitch trim functions. Summary of the Invention

[0006] The technical problem to be solved by this invention is to provide a horizontal stabilizer lifting rotor system for an aircraft and a pitch trim control method for the aircraft. Based on the previously disclosed method of mounting the lifting rotor on the horizontal stabilizer and using thrust to handle pitch trim, this invention further addresses four points: First, the lifting rotor drive motor is mounted on the upper surface of the horizontal stabilizer as the mounting reference surface and supported thereon, eliminating the mast, tailhook, and cantilever mounting path through the vertical stabilizer; Second, the angle of attack above the horizontal stabilizer is reversed, so that the horizontal stabilizer provides part of the rear positive lift in cruise mode, changing the direction in which the conventional horizontal stabilizer provides negative lift during cruise; Third, the forward tilt angle of the rotor shaft is matched with the cruise trim angle of attack, so that the rotor disk angle to the ground is zero or close to zero in cruise trim state, eliminating the need for a tilt mechanism to handle mode transitions; Fourth, the mounting angle of the lifting rotor assembly relative to the fuselage remains unchanged during mode transitions, with the overall attitude change of the aircraft handling the mode transitions. This allows the same set of lifting rotors to provide lift in hovering or low-speed modes, and to handle pitch trim and pitch control by increasing or decreasing thrust in cruise mode. This eliminates the need for tilt and pitch mechanisms to handle mode switching, removes the elevator, and allows the horizontal stabilizer to function as a lifting surface, mounting platform, and structural reinforcement.

[0007] To solve the above-mentioned technical problems, the present invention adopts the following technical solution: a horizontal tail lift rotor system for an aircraft, the aircraft including a fuselage, a wing disposed on the fuselage, a propulsion power unit disposed on the wing, and a horizontal tail disposed at the tail of the fuselage, the horizontal tail lift rotor system including: a tail lift rotor assembly including at least two lift rotors disposed on the left and right sides of the horizontal tail, the lift rotor being a single rotor, or a coaxial dual rotor consisting of upper and lower rotors with coincident rotation axes and opposite rotation directions, wherein the lift rotor is the coaxial... In the dual-rotor configuration, the upper rotor and the lower rotor are each equipped with a drive motor. These drive motors are positioned vertically along the common propeller shaft of the coaxial dual rotors. The thrust of the upper rotor and the lower rotor is combined along the common propeller shaft and jointly contributes to the lift on the side where the coaxial dual rotors are located. The tail lift rotor assembly uses the horizontal stabilizer as its mounting reference surface. The lowermost drive motor in the tail lift rotor assembly is located on the upper surface of the horizontal stabilizer, or is embedded within the horizontal stabilizer with its propeller shaft extending from the upper surface of the horizontal stabilizer, or is located on a mounting base on the upper surface of the horizontal stabilizer. Or on a reinforcing beam; the installation load of the tail lift rotor assembly is transferred to the fuselage via the horizontal stabilizer in the form of surface support, and its installation path does not pass through the mast, tailhook, or cantilever components passing through the vertical stabilizer; the leading edge of the horizontal stabilizer is not lower than the trailing edge of the horizontal stabilizer, so that the aircraft receives part of the rear positive lift from the horizontal stabilizer in cruise mode; the rotor shaft of the tail lift rotor assembly is tilted in the same direction as the longitudinal axis of the fuselage, and the installation angle of the tail lift rotor assembly relative to the fuselage is set so that the rotor disk of the lift rotor is aligned with the fuselage in cruise trim mode. The ground angle is zero or close to zero; the tail lift rotor assembly is fixedly mounted on the horizontal stabilizer with its rotor shaft tilted forward relative to the longitudinal axis of the fuselage, or it is equipped with a tilt drive device, the maximum tilt stroke of which is less than the angle of rotation of the aircraft around the transverse axis during mode transition; the lift rotor is a fixed-pitch rotor, or it is equipped with a pitch adjustment mechanism; during the mode transition of the aircraft, the installation angle of the tail lift rotor assembly relative to the fuselage remains unchanged, or the change in its installation angle is less than the change in the angle of rotation of the aircraft around the transverse axis.

[0008] Furthermore, the trailing edge of the wing is not equipped with ailerons, and the trailing edge of the horizontal stabilizer is not equipped with elevators; in cruise mode, the tail lift rotor assembly remains operational, and the thrust of the left and right lift rotors increases and decreases synchronously, generating a pitching moment on the longitudinal lever arm of the rotor disk center with respect to the center of gravity of the entire aircraft. The tail lift rotor assembly undertakes the pitch trim and pitch control of the aircraft in cruise mode; when the lift rotor is the coaxial twin rotor, the thrust of the upper rotor and the lower rotor increases and decreases synchronously.

[0009] Furthermore, the mounting angle of the horizontal stabilizer ranges from 0° to 10°; the propeller shaft is tilted forward from the longitudinal axis of the fuselage by 0° to 20°, and the difference between the tilt angle and the cruise trim angle of attack of the aircraft is within ±5°.

[0010] Furthermore, the lifting rotor also includes a bushing and a hub, and the drive motor, the bushing, the hub and the disk are connected in sequence from bottom to top, with the drive motor located on the upper surface of the horizontal stabilizer.

[0011] Furthermore, upright plates are respectively provided at the left and right wingtips of the horizontal stabilizer, and the lateral position of the upright plates corresponds to the position of the drive motor. The upright plates constitute the mounting reinforcement of the drive motor and the end plate of the horizontal stabilizer.

[0012] Furthermore, the aircraft has two vertical tail plates symmetrically arranged on the top of its fuselage. The lower edge of the vertical tail plate is embedded in the top arc surface of the fuselage and merged with the fuselage. The rear part of the two vertical tail plates is respectively embedded with auxiliary rudders. The auxiliary rudders are trapezoidal inserts, and the leading edge of the auxiliary rudders forms their hinge line.

[0013] Furthermore, the tail end face of the fuselage is a positive end face; the rotor disk center of the lifting rotor is located behind the center of gravity of the entire aircraft, and the auxiliary rudder is located behind the center of gravity of the entire aircraft.

[0014] Furthermore, the number of blades of the lifting rotor is two, three, four or more; the ratio of the diameter of the rotor disk to the chord length of the horizontal stabilizer is 1.5 to 6; the diameter of the rotor disk is greater than the maximum width of the fuselage, and the distance between the centers of the rotor disks of the left and right sides of the lifting rotor is greater than the maximum width of the fuselage.

[0015] This invention also provides a pitch trim control method for an aircraft, the aircraft including the aforementioned horizontal tail lift rotor system, comprising the following steps: Step 1: acquiring the current flight mode, fuselage angle of attack, lift requirement, and pitch trim requirement of the aircraft; Step 2: in hovering mode or low-speed mode, the tail lift rotor assembly and the propulsion power unit jointly bear the weight of the entire aircraft, and distribute the thrust of the left and right lift rotors according to the pitch trim requirement; Step 3: in cruise mode, maintaining the installation angle of the tail lift rotor assembly relative to the fuselage unchanged, and keeping the tail lift rotor... The assembly operates by synchronously increasing or decreasing the thrust of the lifting rotors on both sides, changing the tail lift on the upper surface of the horizontal stabilizer. This generates a pitch moment on the longitudinal arm of the rotor disk center relative to the overall center of gravity. The tail lifting rotor assembly is responsible for the pitch trim and pitch control of the entire aircraft, and keeps the rotor disk angle to the ground at zero or close to zero. At the same time, the horizontal stabilizer provides some positive rear lift in cruise mode. Step 4: The heading is jointly borne by the deflection of the two auxiliary rudders at the tail of the aircraft and the differential thrust of the propulsion unit, or by the two auxiliary rudders alone.

[0016] Furthermore, during mode switching, the entire aircraft rigidly pitches around its center of gravity to the target angle of attack, and the tail lift rotor assembly moves together with the horizontal stabilizer. The change in the installation angle of the tail lift rotor assembly relative to the fuselage is less than the change in the angle of rotation of the aircraft around its transverse axis. In cruise mode, the horizontal stabilizer provides part of the rear positive lift to reduce the trim thrust requirement of the tail lift rotor assembly.

[0017] It should be noted that the use of the tail-mounted lifting rotor assembly for pitch trim and pitch control in this configuration is technically necessary: ​​the aerodynamic center of the aircraft is located ahead of the center of gravity, making it statically unstable with a negative static stability margin; the elevator authority and bandwidth of a conventional fixed-wing aircraft are insufficient to handle the pitch trim and pitch control of this configuration, while the tail-mounted lifting rotor assembly has a fast and high-authority thrust response, with a thrust authority and response speed greater than that of a conventional elevator. Furthermore, its rotor disk center is located behind the center of gravity and has a long longitudinal lever arm relative to the center of gravity. Therefore, it is feasible for the tail-mounted lifting rotor assembly to handle pitch trim and pitch control by synchronously increasing and decreasing thrust, considering its authority, bandwidth, and torque magnitude. Accordingly, this configuration does not have an elevator; instead, the tail-mounted lifting rotor assembly uses the rotor as a rudder.

[0018] The beneficial effects of this invention are as follows: First, the tail-mounted lifting rotor assembly provides lift in hovering or low-speed modes and generates pitching moment in cruise mode through the synchronous increase and decrease of thrust on both sides. The same set of motors and propellers performs two functions, eliminating the need for elevators. In the preferred embodiment, the tail section does not have a tilting mechanism, pitch control mechanism, or servo motor, minimizing the number of actuators. Second, the upper surface of the horizontal stabilizer directly serves as the mounting reference surface for the lifting rotor drive motor. The motor is supported on the horizontal stabilizer, eliminating the mast, tailhook, and cantilever mounting path through the vertical stabilizer. The force distribution at the mounting base changes from cantilever to surface support. The wingtip uprights of the horizontal stabilizer further reinforce the motor position. The horizontal stabilizer simultaneously serves as a lifting surface, mounting platform, and structural reinforcement, achieving three functions in one device. Third, the propeller shaft of the lifting rotor assembly is mounted with a fixed forward tilt angle relative to the longitudinal axis of the fuselage. The difference between the forward tilt angle and the cruise trim angle of attack is ±5°. Within this range, the rotor disk angle to the ground is zero or close to zero in cruise trim mode, eliminating the need for a tilt mechanism to handle mode transitions and achieving a level cruise rotor disk. This avoids the weight and failure modes associated with the tilt mechanism. Even with a tilt drive, its maximum tilt stroke is less than the rotation angle of the entire aircraft around the transverse axis during mode transitions. Mode transitions are still primarily handled by the attitude changes of the entire aircraft, and the installation angle of the tail lift rotor assembly relative to the fuselage remains unchanged during mode transitions. Fourth, the horizontal stabilizer and above are mounted at reverse angles of attack. In cruise mode, it provides some positive lift from the rear, which is the opposite of the negative lift provided by a conventional horizontal stabilizer during cruise, reducing the trim thrust requirement of the tail lift rotor assembly; fifth, the diameter of the rotor disks and the distance between the disk centers of the left and right lift rotors are both greater than the maximum width of the fuselage, and the rotor disks avoid the wake obstruction above the fuselage, obtaining a larger effective rotor disk area in hovering or low-speed modes; sixth, the two auxiliary rudders at the rear and the differential thrust of the forward propeller assembly jointly undertake heading, and the heading efficiency is higher than that of heading control by differential propeller alone, and the auxiliary rudders... The rudder is an embedded plate, which does not affect the tail shape; seventh, there are no ailerons on the trailing edge of the wing and no elevators on the trailing edge of the horizontal stabilizer. The only movable control surfaces are two auxiliary rudders, and there are no hinge gaps on the trailing edge of the wing, which is conducive to maintaining laminar flow and lift-to-drag ratio for a high aspect ratio wing; eighth, the tail-mounted lifting rotor assembly continues to operate in cruise mode and undertakes lift sharing. Its cruise power proportion is small, estimated to be about one-fifth of that of the forward propeller assembly, while its equivalent lift-to-drag ratio is higher than that of the wing. The lift component of the forward propeller assembly is a byproduct of thrust, therefore... It is energy-efficient for two propellers to share lift; ninth, the lift component coefficient of the forward propeller automatically increases with the angle of attack of the entire aircraft. This is because the forward propeller is installed at an angle of attack above the propeller shaft. This characteristic constitutes passive stabilization without relying on any actuators, which manifests as overload and oscillation rather than lift collapse under gusts; tenth, the lifting body characteristics of the fuselage allow the entire aircraft to continue to provide linearly increasing lift after the wing enters the stall angle of attack. The propeller and the fuselage lifting body jointly maintain the power load, and wing stall does not necessarily lead to lift collapse.

[0019] It should be noted that the claims of this invention and the applicant's related applications are as follows: the tail rotor dynamic horizontal stabilizer layout is claimed in the related applications, and this invention does not claim a tail rotor layout located behind or through the vertical stabilizer; the tail rotor tilting mechanism is claimed in the related applications, and the lifting rotor assembly of this invention is mainly fixed and tilted forward, and the maximum tilting stroke of the tilting drive device is less than the angle of rotation of the entire aircraft around the horizontal axis during mode transition, and the tilting mechanism and its tilting angle linkage are not claimed again; the tail oblique-cut attitude propeller is claimed in the related applications, and the tail of this invention has a positive end face, no oblique-cut opening at the tail, and a horizontally mounted attitude propeller; the internal flow integration is claimed in the related applications, and this invention does not claim the internal flow themes such as boundary layer suction, internal cavity flow path, and tail exhaust propulsion; the stall protection power protection is claimed in the related applications, and this invention does not claim the control method for the lift rotor stall protection; the U-shaped retractable landing gear is claimed in the related applications, and this invention only refers to its landing gear deployment condition and does not claim the landing gear structure. This invention advocates the configuration combination of the present case, namely, using the upper surface of the horizontal stabilizer as the mounting reference surface for the lift rotor drive motor and placing the motor on it with surface support; mounting the horizontal stabilizer at the angle of attack in reverse; keeping the difference between the propeller shaft tilt angle and the cruise trim angle of attack within ±5° so that the propeller disk angle to the ground is zero or close to zero in the cruise trim state; keeping the mounting angle of the lift rotor assembly relative to the fuselage unchanged during mode transition; and enabling the dual-function multiplexing of the same set of lift rotors in hover mode and cruise mode. Attached Figure Description

[0020] Figure 1 is a side view of the overall aircraft of the horizontal tail-mounted lifting rotor aircraft according to an embodiment of the present invention.

[0021] Figure 2 is a partially enlarged side view of the tail lift rotor assembly and horizontal stabilizer in an embodiment of the present invention.

[0022] Figure 3 is a top view of the tail horizontal stabilizer and lifting rotor assembly according to an embodiment of the present invention.

[0023] Figure 4 is a schematic diagram of a typical cross-section of the horizontal stabilizer and the wingtip uprights according to an embodiment of the present invention.

[0024] Figure 5 is a diagram showing the arrangement of the tail section vertical tail plate and auxiliary rudder in an embodiment of the present invention.

[0025] Figure 6 is a schematic diagram of four flight attitudes according to an embodiment of the present invention.

[0026] Figure 7 is a block diagram of the pitch trim and heading control principle of an embodiment of the present invention.

[0027] Figure 8 is a flowchart of the pitch trim control according to an embodiment of the present invention.

[0028] Among them, 10 is the flight control computer; 11 is the fuselage; 12 is the wing; 13a is the left forward propulsion assembly; 13b is the right forward propulsion assembly; 14 is the forward propulsion motor nacelle; 15 is the horizontal stabilizer; 16 is the horizontal stabilizer wingtip upright; 17a is the left tail lift propeller; 17b is the right tail lift propeller; 18 is the lift propeller motor; 19 is the lift propeller shaft sleeve and hub; 20a is the left vertical tail plate; 20b is the right vertical tail plate; 21a is the left auxiliary rudder; 21b is the right auxiliary rudder; 22 is the servo; 23 is the tail end face of the fuselage; 24 is the U-shaped landing gear; and 25 is the center of gravity calibration point. Specific implementation

[0029] The present invention will be further described below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only for explaining the present invention and are not intended to limit the present invention. The values ​​in the embodiments are all design values ​​at the prototype level, with lengths in millimeters and angles in degrees; the coordinate system is the fuselage coordinate system, with the origin located at the nose of the fuselage, the y-axis pointing towards the tail along the longitudinal axis of the fuselage being negative, the x-axis pointing horizontally to the right being positive, and the z-axis pointing vertically upwards being positive; the angle aperture is the mounting angle, i.e., the axis-to-ground angle and the disk-to-ground angle are equal to 90 degrees minus the axis-to-ground angle. The above values ​​are used to illustrate the relative positions and angular relationships between the components, and are not intended to limit the scope of protection of the present invention.

[0030] As shown in Figure 1, the aircraft in this embodiment includes a fuselage 11, wings 12 disposed on the fuselage 11, propulsion power units respectively disposed on the left and right wings (i.e., left front propulsion group 13a and right front propulsion group 13b), propulsion motor nacelles 14 disposed on the wings, and a horizontal stabilizer 15 disposed at the tail of the fuselage 11. The fuselage 11 is a hollow shell. In this embodiment, the fuselage length is 1096.19 mm, the maximum width is 503.2 mm, the total height is 191.52 mm, the shell wall thickness is 4.0 mm, and the tail end face 23 of the fuselage is a positive end face. The wings 12 are high aspect ratio wings with a half span of 1096.7 mm, of which 845.1 mm is exposed above the fuselage. The left front propulsion group 13a and the right front propulsion group 13b each include a propulsion motor and a three-bladed propeller driven by it. In this embodiment, there are two propellers with a diameter of 28 inches (711.2 mm). The center position of the propeller disk is ±1116.1 mm laterally, ±243.6 mm longitudinally, and ±62.8 mm vertically. The motor shaft is tilted upwards at 35° and outwards at 10° relative to the longitudinal axis of the fuselage. It is fixedly connected to the fuselage and cannot be tilted or have its pitch changed. The front propulsion motor nacelle 14 is an egg-shaped nacelle, which is composed of an egg-shaped hub cover and a motor cylinder cone. In this embodiment, the radius of the egg-shaped cover is 58.447 mm. The nacelle is enlarged by 1.25 times around the tail tip. The entire wingtip surface is embedded in the nacelle surface and fused to a depth of 0.3 mm. The flight control computer 10 is located inside the fuselage and is connected to the left front propulsion group 13a, the right front propulsion group 13b, the left tail lift propeller 17a, the right tail lift propeller 17b, the left auxiliary rudder 21a, and the right auxiliary rudder 21b via signal links.

[0031] As shown in Figures 3 and 4, the horizontal stabilizer 15 is located at the tail of the fuselage 11. In this embodiment, the horizontal stabilizer 15 is a constant chord horizontal stabilizer with a full span of 800.16, i.e., 400.08 on each side; the constant chord length is 167.42, the leading edge position is longitudinally -732.7, and the trailing edge position is longitudinally -900.12; the wingtips of the horizontal stabilizer 15 have angled retraction, and in this embodiment, the angled retraction radii are 20.47 and 321.525, respectively. The mounting angle of the horizontal stabilizer 15 in this embodiment is positive 2.74°, which is the reverse-mounted upper angle of attack, with its leading edge pitching upward relative to the trailing edge, which is opposite to the direction of the conventional horizontal stabilizer's leading edge pitching downward and providing negative lift during cruise; the upper surface of the horizontal stabilizer 15 is a plane, forming the mounting reference surface for the lift propeller drive motor.

[0032] As shown in Figure 2, the left tail lifter 17a and the right tail lifter 17b are respectively located on the left and right sides of the horizontal stabilizer 15. The mounting chains of each lifter, from bottom to top, are as follows: the upper surface of the horizontal stabilizer 15, the lifter motor 18, the lifter bushing and hub 19, and the rotor disk. In this embodiment, the lifter motor 18 is an external rotor motor with a diameter of 80 and a height of 40, located on the top surface of the horizontal stabilizer 15, with a vertical range of 319.1 to 359.1. The lifter bushing and hub 19 includes a bushing with a diameter of 20 and a length of 18.6 and a hub with a diameter of 40 and a height of 20. The rotor disk is a four-bladed rotor with a diameter of 24 inches, or 609.6, with the disk surface located at a vertical angle of 391.1. The center position of the rotor disk is ±395.08 laterally, ±841.6 longitudinally, and ±391.1 vertically. The upper surface of the horizontal stabilizer 15 serves as the mounting reference surface for the lift propeller motor 18. The motor is supported on the horizontal stabilizer. In this embodiment, there is no mast, no tailhook, and no shaft passing through the tail plate. The motors, shafts, and propellers of the left tail lift propeller 17a and the right tail lift propeller 17b form a single unit. This unit is tilted forward by 5° relative to the longitudinal axis of the fuselage, i.e., the angle between the propeller shaft and the longitudinal axis of the fuselage is 85°. The propeller disk tilts along with the propeller shaft, rather than just the blades tilting relative to the propeller hub. In this embodiment, the lift propeller assembly does not have a tilting mechanism, a pitch control mechanism, or a servo motor. The direction of its propeller disk is determined by a fixed forward tilting installation angle and the rigid pitch attitude of the entire aircraft around its center of gravity. As an alternative implementation, the lift propeller assembly may also be equipped with a small-stroke tilt drive device or a pitch adjustment mechanism. Its tilting stroke is much smaller than the angle of rotation of the entire aircraft around the transverse axis during mode transition. See the section on alternative implementations below for details. The overall center of gravity 25 is located at -548.8 in the longitudinal direction and +81.7 in the vertical direction. The center of the lift disk is located behind the center of gravity 25. In this embodiment, the longitudinal lever arm of the disk center to the center of gravity 25 is 292.8 and the vertical difference is 309.4. This longitudinal lever arm constitutes the lever arm of the pitching moment.

[0033] As shown in Figures 3 and 4, the left and right wingtips of the horizontal stabilizer 15 are respectively provided with wingtip uprights 16. In this embodiment, the thickness of the upright is 10 mm, the lateral position is 390.02 to 400.08 mm, the leading edge sweep radius is 121.45 mm, the longitudinal range is 753 to 919 mm, and the vertical range is 277 to 339 mm. The lateral position of the wingtip uprights 16 corresponds to the position of the lift propeller motor 18, and simultaneously constitutes the mounting reinforcement of the lift propeller motor and the end plate of the horizontal stabilizer 15; thus, the horizontal stabilizer 15 has the triple functions of a lifting surface, a lift propeller mounting platform, and a structural reinforcement.

[0034] As shown in Figure 5, two vertical tail plates 20a and 20b are symmetrically arranged on the top of the fuselage 11. In this embodiment, the thickness of the vertical tail plate is 10 mm, the lateral position is 175.3 to 185.3 mm, the longitudinal range is 638.07 to 907.49 mm, and the vertical range is 146.26 to 295.82 mm. The leading edge of the vertical tail plate is a long sweeping arc, and the rear part is provided with a rudder compartment opening. Its bottom edge is embedded in the arc surface of the top of the fuselage. In this embodiment, it is embedded by 2 mm and fused together, so that there is no splicing gap between the vertical tail plate and the fuselage shell. The rear portions of the two vertical tail plates 20a and 20b are respectively fitted with a left auxiliary rudder 21a and a right auxiliary rudder 21b. The auxiliary rudders are trapezoidal inserts with a thickness of 10 mm, located at the same column position as the vertical tail plates, with a longitudinal range of 831.33 to 898.01 mm and a vertical range of 200.44 to 280.04 mm. A hinge line is formed at the longitudinal position of its leading edge at 831.33 mm. The servo motor 22 is connected to the auxiliary rudder to drive its deflection. In this embodiment, the servo motor 22 is not modeled, and its installation position is indicated by a dashed line. The tail end face 23 of the fuselage is a positive end face. In this embodiment, an exhaust port is provided on this end face. The exhaust port and its exhaust thruster are claimed by the same family of applications and will not be elaborated here. The lower part of the fuselage 11 is provided with a U-shaped landing gear 24. This embodiment uses its retraction structure. The parking condition is a superposition of the hovering condition and the landing gear deployment condition.

[0035] As shown in Figure 6, the aircraft in this embodiment has four attitude states: sprint, cruise, hover, and stationary. In all four states, the entire aircraft rigidly pitches around its center of gravity 25°. The horizontal stabilizer 15 and the tail lift propellers 17a and 17b pitch along with the entire aircraft, maintaining their relative positions to the fuselage. Only the angle of the propeller disk relative to the ground changes with the overall angle of attack. With the propeller assembly fixed at a 5° forward tilt relative to the fuselage longitudinal axis, the propeller shaft angle to the ground equals 85° plus the fuselage angle of attack, and the propeller disk angle to the ground equals 5° minus the fuselage angle of attack. The four attitude states of this embodiment are shown in the table below.

[0036] Sprint 0° 5° (forward lean) 35° 55° cruise +5° 0° (horizontal) 40° 50° Hover +30° 25° 65° 25° Stop Similar to hovering, additional landing gear deployment is also included. 25° 65° 25°

[0037] As shown in the table above, the forward tilt angle of 5° in this embodiment is equal to the cruise trim angle of attack of ±5°. Therefore, the lift rotor disk angle to the ground is zero in cruise mode, the rotor disk is horizontal, and the thrust direction of the lift rotor is perpendicular to the longitudinal axis of the fuselage. This eliminates the need for a tilting mechanism to handle mode transitions, thus achieving a horizontal cruise rotor disk. It should be noted that the cruise trim angle of attack is not a fixed value; it varies with flight speed and takeoff weight, estimated to be within the range of approximately 0° to 12° according to design. The design cruise point in this embodiment is ±5°, so the equal values ​​in the table above are the result of selecting the design cruise point in this embodiment. The claims define the functional relationship between the rotor shaft forward tilt angle and the cruise trim angle of attack as being within ±5°, rather than limiting them to being fixedly equal. As long as the difference falls within the above range, the lift rotor disk angle to the ground in cruise mode is zero or close to zero. In sprint mode, the fuselage angle of attack is 0°, and the propeller disk tilts forward by 5°. In hover mode, the fuselage angle of attack is ±30°, and the absolute value of the propeller disk tilt is 25°, with the tilt direction opposite to that in sprint mode. The parking mode is hover mode with the U-shaped landing gear 24 deployed. In all four operating conditions, the center of gravity 25 is located at the same point with a longitudinal angle of -548.8° and a vertical angle of +81.7°.

[0038] As shown in Figure 7, the principle of pitch trim and heading control is as follows: The thrust of the left tail lift propeller 17a and the right tail lift propeller 17b generates a pitch moment through the lever arm of their propeller disk center about the center of gravity 25. The synchronous increase or decrease of the thrust on both sides changes the total tail lift at the upper surface of the horizontal stabilizer, thereby generating a pitch trim moment on the longitudinal lever arm 292.8 and forming a pitch control moment. In this embodiment, the lift propeller assembly is fixedly installed, and its propeller disk direction does not change with the control command. The pitch moment is entirely generated by adjusting the thrust magnitude. The pitch moment is obtained by cross-product of the thrust vectors on both sides and the position vector of the propeller disk center about the center of gravity 25. The contribution of the vertical difference 309.4 varies with the propeller shaft direction. In the hovering mode, the propeller shaft deviates from vertical by 25°. When calculating, the thrust needs to be projected according to the propeller shaft direction for each operating condition. The yaw moment is generated by the deflection of the left auxiliary rudder 21a and the right auxiliary rudder 21b, and the differential thrust of the left forward thrust group 13a and the right forward thrust group 13b. This embodiment does not limit the specific allocation of the roll channel; the coupling magnitude caused by the differential thrust of the forward thrust group on the left and right lift needs to be determined by aerodynamic and flight control calculations. The flight control computer 10 outputs control commands to the lift propeller motor 18, the propulsion motor of the forward thrust group, and the servo motor 22 of the auxiliary rudder, respectively.

[0039] As shown in Figure 8, the pitch trim control method of this embodiment includes the following process. Step S1: Status acquisition: Acquire the current flight mode, fuselage angle of attack, lift requirement, pitch trim requirement, and pitch angular velocity. Step S2: Lift sharing during hovering or low speed: The left tail lift rotor 17a and right tail lift rotor 17b, together with the left front thruster group 13a and right front thruster group 13b, share the weight of the entire aircraft, and distribute the thrust of the two lift rotors according to the pitch trim requirement. The trim distribution is coordinated with the longitudinal position of the center of gravity 25. Step S3: Cruise Pitch Trim: The lift propellers on both sides remain operational. By synchronously increasing or decreasing the thrust on both sides, the tail lift at the upper surface of the horizontal stabilizer is changed, generating a pitch moment on the 292.8 longitudinal lever arm of the propeller disk center about the center of gravity 25. This moment undertakes the pitch trim of the entire aircraft and serves as a dynamic pitch control surface. In this embodiment, the lift propeller assembly is fixedly installed, the tilting mechanism does not participate, and the pitch moment is entirely generated by adjusting the thrust magnitude. At the same time, the horizontal stabilizer 15 provides part of the rear positive lift in cruise mode with its positive angle of attack of 2.74°. Step S4: Heading Control: The left auxiliary rudder 21a and right auxiliary rudder 21b deflect, generating a yaw moment in combination with the differential thrust of the left forward thruster assembly 13a and the right forward thruster assembly 13b. The left and right lift difference coupling caused by the differential thrust of the forward thruster is compensated by the remaining roll channels. This embodiment does not limit the compensation method. Step S5 Mode Switching: With mode switching, the entire aircraft rigidly pitches around its center of gravity 25° to the target angle of attack. The sprint mode is 0°, the cruise mode is ±5°, and the hovering mode is ±30°. The horizontal stabilizer 15 and the tail lift propellers 17a and 17b follow suit. The upper angle of attack of the horizontal stabilizer 15 provides part of the rear positive lift in cruise mode, reducing the trim thrust requirement of the lift propeller assembly. Step S6 is the differential channel for the two lift propellers. This channel can be used to generate roll torque or compensate for roll coupling caused by the differential of the forward propeller. In this embodiment, the allocation of this channel is determined by aerodynamic and flight control calculations and is shown in the dashed box in Figure 8.

[0040] The main design parameters of this embodiment are shown in the table below.

[0041] Fuselage length / maximum width / total height 1096.19／503.2／191.52 fuselage casing wall thickness 4.0 Wings partially extended / Exposing fuselage section 1096.7／845.1 Forward propeller 2 x 28 inches (711.2) trefoil Forward propeller disk center position Lateral ±1116.1, Longitudinal −243.6, Vertical +62.8 Forward propulsion motor shaft tilt / outward 35°／±10° Tail lift propeller 2 x 24 inches (609.6) four-leaf Lift rotor disk center position Lateral ±395.08, Longitudinal −841.6, Vertical +391.1 Lift propeller motor With a diameter of 80 mm and a height of 40 mm, it sits atop the flat-tailed vessel, with a vertical dimension of 319.1 to 359.1 mm. Lift propeller bushing / hub Diameter 20, Length 18.6 / Diameter 40, Height 20 Lift propeller disk Vertical 391.1 Overall rake angle of propeller assembly Tilted forward 5° relative to the longitudinal axis of the fuselage (motor-shaft-propeller assembly) Cruise trim angle of attack Based on design estimates, the range is approximately 0° to 12°, with the design cruise point taken as +5°. Lift propeller operation This embodiment lacks a tilt mechanism, a pitch adjustment mechanism, and a servo motor; alternative embodiments may include a short-stroke tilt drive device or a pitch adjustment mechanism. Flat tail full span / chord length 800.16 (±400.08) / 167.42 (equichords) Front / back edge positioning of the flat tail Longitudinal -732.7 / -900.12 Flat-tail mounting angle +2.74° (reverse mounting, upward pitch of leading edge) horizontal tail wingtip angle Radius 20.47 / 321.525 horizontal tail wingtip uprights Thickness 10; Lateral ±390.02~400.08; Leading edge sweep radius 121.45; Longitudinal 753~919; Vertical 277~339 Tail vertical tail plate Thickness 10mm; lateral dimensions ±175.3~185.3mm; longitudinal dimensions 638.07~907.49mm; vertical dimensions 146.26~295.82mm; bottom edge inlaid 2mm along the top arc of the fuselage. auxiliary rudder Trapezoidal insert, thickness 10mm; longitudinal dimension 831.33~898.01mm; vertical dimension 200.44~280.04mm; leading edge longitudinal dimension 831.33mm is a hinge line. Overall center of gravity Longitudinal -548.8, Vertical +81.7 (same point for all four operating conditions) The lever arm of the lift disk center about the center of gravity Longitudinal deviation: 292.8; Vertical difference: 309.4 Four-condition propeller disk relative to ground angle Sprint 5° / Cruise 0° / Hover 25° (reverse) / Park as hover Channel configuration No ailerons, no elevators; pitch is driven by the lift propeller thrust; yaw is driven by the differential combination of the auxiliary rudder and the forward propulsion unit. Scrolling Channel The allocation method is determined by aerodynamic and flight control calculations, and this embodiment does not impose any limitations.

[0042] Alternative implementation: The tail-end lifting rotor assembly can be a single rotor on each side, or a coaxial twin-rotor configuration consisting of upper and lower rotors on each side. The upper and lower rotors of the coaxial twin-rotor configuration have their rotation axes coincident and rotate in opposite directions. Each rotor is equipped with a drive motor. The drive motor for the lower rotor is located on the upper surface of the horizontal stabilizer, and the drive motor for the upper rotor is located on the upper part of the shared propeller shaft. The drive motors are positioned vertically along the shared propeller shaft. The thrust of the upper and lower rotors is combined along the common rotor shaft direction and shares the lift on that side. The rotor disk diameters of the two rotors can be equal or unequal. In cruise mode, the thrust of the upper and lower rotors increases and decreases synchronously, generating a pitching moment on the longitudinal lever arm of the rotor disk center about the center of gravity of the entire aircraft. Its mechanism is the same as that of a single rotor. The number of blades of the lift rotor can be selected from two, three, four, or more blades according to the thrust matching requirements. The ratio of the rotor disk diameter to the horizontal stabilizer chord length can be 1.5 to 6. The range of values ​​can be selected, and large-diameter, low-blade-number propellers or multiple propellers in parallel can be used instead of a single propeller; the drive motor of the lift propeller can also be embedded inside the horizontal stabilizer, with its propeller shaft extending upward from the upper surface of the horizontal stabilizer, or it can be in the form of an internal rotor motor with a reduction gear; the lift rotor can also be mounted on a reinforcing beam or mounting base on the upper surface of the horizontal stabilizer, with the upper surface of the horizontal stabilizer and the mounting base together forming the mounting reference surface; the horizontal stabilizer can also be equipped with a movable elevator but normally locked in cruise mode, or other aerodynamic control surfaces such as flaps and ailerons can be used to supplement pitch trim; the installation angle of the horizontal stabilizer can be selected within the range of 0° to 10° according to the cruise trim requirements, when it is 0°, the chord of the horizontal stabilizer is parallel to the longitudinal axis of the fuselage, and it still provides some positive lift at the rear at the cruise trim angle of attack; its wingtips can also be swept back or swept forward; the forward tilt angle of the propeller assembly can be 0° to 20°. Within the range, the cruise trim angle of attack is selected; the two auxiliary rudders can also be one or more, and their hinge form can be axial hinge or flexible hinge; the propulsion power unit can also adopt the form of ducted fan or internal propeller to replace the exposed electric propeller; the combined heading distribution of the auxiliary rudder and the forward propulsion group can also be undertaken by the auxiliary rudder alone in cruise mode; all the above variants can be adaptively scaled according to the overall aircraft size, and enlarged to larger models such as logistics and transportation; all the above variants can realize the configuration function of using the horizontal stabilizer as the mounting reference plane for the lifting rotor, the lifting rotor group is installed with a fixed forward tilt angle or with small stroke tilt for trim fine adjustment, the leading edge of the horizontal stabilizer is not lower than the trailing edge, and the same set of lifting rotors can be reused in both hovering and cruise modes.

[0043] As an alternative implementation, the tail lift rotor assembly may be equipped with a short-stroke tilt drive device. The short-stroke tilt drive device is only used for trim fine-tuning or mounting angle fine-tuning. Its maximum tilt stroke is much smaller than the pitch angle change of the entire aircraft around the transverse axis during mode transition. Mode transition is mainly undertaken by the attitude change of the entire aircraft around the transverse axis, and not by the short-stroke tilt drive device. When this device is provided, the tail lift rotor assembly still uses the upper surface of the horizontal stabilizer as the mounting reference plane, and reuses the same set of lift rotors in hovering mode and cruise mode.

[0044] As an alternative implementation, the lifting rotor can be a fixed-pitch propeller or a variable-pitch propeller with a pitch adjustment mechanism, so that the thrust of the lifting rotor can be adjusted by combining the pitch and the rotational speed. Regardless of whether a fixed-pitch propeller or a variable-pitch propeller is used, the installation angle of the tail lifting rotor assembly remains unchanged during the mode transition of the aircraft, or the change in its installation angle is less than the change in the angle of the aircraft's rotation around the transverse axis. The mode transition is mainly undertaken by the attitude change of the whole aircraft around the transverse axis.

[0045] The specific embodiments of the present invention have been described above with reference to the accompanying drawings. However, these descriptions should not be construed as limiting the scope of the present invention. The scope of protection of the present invention is defined by the appended claims. Any modifications based on the claims of the present invention are within the scope of protection of the present invention.

Claims

1. A horizontal stabilizer lifting rotor system for an aircraft, the aircraft comprising a fuselage, a wing disposed on the fuselage, a propulsion unit disposed on the wing, and a horizontal stabilizer disposed at the tail of the fuselage, characterized in that, The horizontal stabilizer lifting rotor system includes: The tail-end lifting rotor assembly includes at least two lifting rotors located on the left and right sides of the horizontal stabilizer. The lifting rotor is a single rotor or a coaxial dual rotor consisting of two rotors with coincident rotation axes and opposite rotation directions. When the lifting rotor is the coaxial dual rotor, the upper rotor and the lower rotor are respectively equipped with drive motors. The drive motors of the two rotors are located vertically along the common propeller shaft of the coaxial dual rotor. The thrust of the upper rotor and the lower rotor is combined along the direction of the common propeller shaft and jointly shares the lift on the side where the coaxial dual rotor is located. The tail lift rotor assembly uses the horizontal stabilizer as the mounting reference surface. The lowest drive motor in the tail lift rotor assembly is located on the upper surface of the horizontal stabilizer, or is embedded in the horizontal stabilizer with its rotor shaft extending from the upper surface of the horizontal stabilizer, or is located on a mounting base or reinforcing beam on the upper surface of the horizontal stabilizer. The installation load of the tail lift rotor assembly is transmitted to the fuselage through the horizontal stabilizer in the form of surface support, and its installation path does not pass through the mast, tailhook, or cantilever components that pass through the vertical stabilizer. The leading edge of the horizontal stabilizer is not lower than the trailing edge of the horizontal stabilizer, so that the aircraft can be provided with a portion of the rear positive lift by the horizontal stabilizer in cruise mode; The propeller shaft of the tail lift rotor assembly is tilted in the same direction as the longitudinal axis of the fuselage. The installation angle of the tail lift rotor assembly relative to the fuselage is set so that the angle of the rotor disk relative to the ground is zero or close to zero when the aircraft is in cruise trim state. The tail lift rotor assembly is fixed to the horizontal stabilizer with its propeller shaft tilting forward relative to the longitudinal axis of the fuselage, or it is equipped with a tilt drive device. The maximum tilt stroke of the tilt drive device is less than the angle of rotation of the aircraft around the transverse axis during mode transition. The lift rotor is a fixed-pitch rotor, or it is equipped with a pitch adjustment mechanism. During the mode transition of the aircraft, the installation angle of the tail lift rotor assembly relative to the fuselage remains unchanged, or the change in its installation angle is less than the change in the angle of the aircraft's rotation around the transverse axis.

2. The horizontal stabilizer lifting rotor system according to claim 1, characterized in that, The trailing edge of the wing is not equipped with ailerons, and the trailing edge of the horizontal stabilizer is not equipped with elevators. In cruise mode, the tail lift rotor assembly remains operational, and the thrust of the left and right lift rotors increases and decreases synchronously, generating a pitching moment on the longitudinal arm of the rotor disk center with respect to the center of gravity of the entire aircraft. The tail lift rotor assembly is responsible for the pitch trim and pitch control of the aircraft in cruise mode. When the lift rotor is a coaxial dual rotor, the thrust of the upper rotor and the lower rotor increases and decreases synchronously.

3. The horizontal stabilizer lifting rotor system according to claim 1, characterized in that, The mounting angle of the horizontal stabilizer ranges from 0° to 10°; the propeller shaft is tilted forward from the longitudinal axis of the fuselage by 0° to 20°, and the difference between the tilt angle and the cruise trim angle of attack of the aircraft is within ±5°.

4. The horizontal stabilizer lifting rotor system according to claim 1, characterized in that, The lifting rotor also includes a bushing and a hub. The drive motor, the bushing, the hub and the disk are connected in sequence from bottom to top. The drive motor is located on the upper surface of the horizontal stabilizer.

5. The horizontal stabilizer lifting rotor system according to claim 1, characterized in that, The horizontal stabilizer has upright plates at the left and right wingtips, and the lateral position of the upright plates corresponds to the position of the drive motor. The upright plates constitute the mounting reinforcement of the drive motor and the end plate of the horizontal stabilizer.

6. The horizontal stabilizer lifting rotor system according to claim 1, characterized in that, The aircraft has two vertical tail plates symmetrically arranged on the top of its fuselage. The lower edge of the vertical tail plate is embedded in the top arc surface of the fuselage and merged with the fuselage. The rear part of the two vertical tail plates is respectively embedded with auxiliary rudders. The auxiliary rudders are trapezoidal inserts, and the leading edge of the auxiliary rudders forms their hinge line.

7. The horizontal stabilizer lifting rotor system according to claim 6, characterized in that, The tail end face of the fuselage is a positive end face; the rotor disk center of the lifting rotor is located behind the center of gravity of the entire aircraft, and the auxiliary rudder is located behind the center of gravity of the entire aircraft.

8. The horizontal stabilizer lifting rotor system according to claim 1, characterized in that, The lifting rotor has two, three, four, or more blades; the ratio of the diameter of the lifting rotor's rotor disk to the chord length of the horizontal stabilizer is 1.5 to 6; the diameter of the lifting rotor's rotor disk is greater than the maximum width of the fuselage, and the distance between the centers of the rotor disks on the left and right sides is greater than the maximum width of the fuselage.

9. A pitch trim control method for an aircraft, the aircraft comprising the horizontal stabilizer lifting rotor system of claim 1, characterized in that, Includes the following steps: Step 1: Obtain the current flight mode, fuselage angle of attack, lift requirements, and pitch trim requirements of the aircraft; Step 2: In hovering mode or low-speed mode, the tail lift rotor assembly and the propulsion power unit jointly bear the weight of the entire aircraft, and distribute the thrust of the left and right lift rotors according to the pitch trim requirements; Step 3: In cruise mode, maintain the installation angle of the tail lift rotor assembly relative to the fuselage, keep the tail lift rotor assembly in operation, and change the tail lift at the upper surface of the horizontal stabilizer by synchronously increasing or decreasing the thrust of the left and right lift rotors. This generates a pitching moment on the longitudinal arm of the rotor disk center about the center of gravity of the entire aircraft. The tail lift rotor assembly is responsible for the pitch trim and pitch control of the entire aircraft, and keeps the rotor disk angle with respect to the ground at zero or close to zero. At the same time, the horizontal stabilizer provides part of the rear positive lift in cruise mode. Step 4: The heading is borne by the deflection of the two auxiliary rudders at the tail of the aircraft and the differential thrust of the propulsion unit, or by the two auxiliary rudders alone.

10. The pitch trim control method according to claim 9, characterized in that, During mode switching, the entire aircraft rigidly pitches around its center of gravity to the target angle of attack. The tail lift rotor assembly moves together with the horizontal stabilizer. The change in the installation angle of the tail lift rotor assembly relative to the fuselage is less than the change in the angle of rotation of the aircraft around its transverse axis. In cruise mode, the horizontal stabilizer provides part of the rear positive lift to reduce the trim thrust requirement of the tail lift rotor assembly.

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