Variable incidence tailplane configuration for high mach number aircraft

Abstract

The application belongs to the field of aircraft design, and particularly relates to a variable angle tail wing configuration for a high Mach number aircraft, comprising a left wing, a right wing, a left tail wing and a right tail wing; the left tail wing and the right tail wing are arranged in mirror symmetry along a fuselage symmetry plane, the root of the left tail wing and the root of the right tail wing are connected to a tail wing rotating shaft, and the tail wing rotating shaft is rotationally connected to a wing trailing edge; the rotation angle of the left tail wing and the right tail wing is 0° to ±90°; a power unit is connected to a flight control system of the aircraft, and the flight control system of the aircraft can control the left tail wing and the right tail wing to rotate by 0° to ±90° along the tail wing rotating shaft. The tail wing serves as a lifting surface at a low speed stage according to the aerodynamic characteristic requirement at different flight speeds, and increases the lift of the aircraft. The tail wing serves as a heading stability and control surface at a high speed stage, and solves the problem of insufficient heading stability at a high Mach number. A dihedral angle of the tail wing is reasonably selected, and the peak value of wing area caused by the requirement of lift and heading stability in traditional design is avoided.

Description

Technical Field

[0001] This application belongs to the field of aircraft design, and in particular relates to a variable tilt tail configuration for high Mach number aircraft. Background Technology

[0002] Existing aircraft tail fins have a fixed, non-adjustable canard angle, with lift primarily provided by the wings, and the physical area of ​​the wing's lifting surface remains constant. The determination of the wing area is constrained by various factors, primarily including the high lift requirements during takeoff and landing, the low drag requirements during high-speed cruise, transonic drag characteristics, and the impact of the aircraft's overall weight on performance. However, due to the high lift requirements during takeoff and landing, especially runway length limitations, the wing area design is often determined by the high lift requirements during these phases. Therefore, the wing area designed to meet takeoff and landing performance requirements becomes redundant during transonic and high-speed cruise, resulting in an oversized wing. This increases drag and structural weight, reducing fuel efficiency and cruise performance.

[0003] Once the flight Mach number enters the supersonic range, the efficiency of the tail fin decreases significantly, and the focal point shifts forward, leading to a sharp decline in the aircraft's directional static stability. To meet the directional stability requirements at high Mach numbers, the vertical tail area must be increased to improve the directional projection area behind the center of gravity, thereby enhancing directional static stability. However, this traditional fixed-tail design leads to the peak values ​​of the wing area and tail area overlapping, resulting in additional structural weight and drag on the aircraft, ultimately affecting its overall performance.

[0004] For high Mach number aircraft (maximum flight Mach number approximately 3.0), the takeoff and landing, transonic, and high-speed cruise phases place significantly different demands on aerodynamic characteristics:

[0005] a) Low-speed takeoff and landing phase (Ma<0.2): Since the aircraft needs to generate a large amount of lift to meet the performance requirements for takeoff and landing, this is usually achieved by increasing the wing area.

[0006] b) Transonic phase (Ma≈1.0): As flight speed increases, shock wave drag gradually becomes dominant, leading to a sharp increase in aircraft drag. Therefore, reducing drag in the transonic phase becomes a core task of aerodynamic design.

[0007] c) High Mach number cruise phase: As flight speed increases, tail efficiency decreases, focus shifts forward, and the aircraft's directional static stability drops sharply, significantly increasing the requirements for directional stability and aerodynamic efficiency.

[0008] Traditional fixed-tail designs have the following main problems in high Mach number aircraft:

[0009] a) The lifting surface function is not fully utilized: During low-speed takeoff and landing, the fixed tail cannot effectively participate in lift generation, forcing the wing area to be increased to meet lift requirements. However, a larger wing area will significantly increase flight drag and structural weight during transonic and supersonic cruise, which is detrimental to the aircraft's fuel efficiency and cruise performance.

[0010] b) Insufficient heading stability performance: such as Figure 1 The longitudinal parameter is the directional static stability coefficient, and the lateral parameter is the Mach number. At high Mach numbers, tail efficiency decreases, the focal point shifts forward, and directional static stability drops sharply. To meet directional stability requirements, the vertical tail area needs to be increased; however, this further increases the aircraft's structural weight and drag.

[0011] c) Aerodynamic mismatch: Due to the fixed-wing configuration of the existing design, in order to simultaneously meet the lift requirements at low speeds and the directional stability requirements at high Mach numbers, the peak values ​​of the wing area and tail area overlap. This design prevents the aircraft from achieving optimal aerodynamic performance in different flight phases, thus limiting the overall performance of the aircraft.

[0012] Therefore, how to achieve the optimal balance between lift efficiency, transonic drag reduction performance, directional stability and flight performance through the optimized design of the tail configuration, and solve the technical problem of matching aerodynamic characteristics of aircraft in high-speed and low-speed flight phases, has become a key issue in the design of high-speed aircraft. Summary of the Invention

[0013] To address the aforementioned issues, this application provides a variable tilt tail configuration for high Mach number aircraft, thereby resolving the technical challenge of matching aerodynamic characteristics of aircraft during high and low speed flight phases in the prior art.

[0014] The technical solution of this application is: a variable tilt tail configuration for high Mach number aircraft, including a left wing, a right wing, a left tail and a right tail;

[0015] The left and right tail fins are arranged mirror-symmetrically along the fuselage symmetry plane. The roots of the left and right tail fins are both connected to a tail fin pivot, which is rotatably connected to the trailing edge of the wing. The rotation angle of the left and right tail fins is 0° to ±90°. The aircraft is equipped with a power unit, which is connected to the aircraft's flight control system. The flight control system can control the left and right tail fins to rotate 0° to ±90° along the tail fin pivot.

[0016] Preferably, when the rotation angle of the left and right tail fins is 0°, the aircraft is in a horizontal tail configuration; when the rotation angle of the left and right tail fins is 90°, the aircraft is in a downward tail configuration.

[0017] Preferably, a flight data processing module is provided in the flight control system. The flight data processing module can obtain the current flight Mach number Ma and control the left and right tail fins to rotate to corresponding angles; when Ma ≤ 1.2 and the aircraft heading static stability coefficient is as low as -0.00065, control the rotation angles of the left and right tail fins to 0°; when Ma = 3.5 and the heading static stability coefficient is not lower than -0.0003, control the rotation angles of the left and right tail fins to 90°.

[0018] Preferably, when the flight data processing module determines that the current flight Mach number is in the transition range: 1.2 < Ma < 3.5, the optimal angle of the tail fin axis is calculated specifically as follows:

[0019] Let the tail fin inclination angle be θ, and the following constraints are set:

[0020] First, the tail fin inclination angle θ changes linearly and positively with the flight Mach number Ma;

[0021] Second, the aircraft heading static stability coefficient calculated in real time needs to satisfy the linear transition constraint. If not, the tail fin inclination angle θ is compensated and corrected;

[0022] Third, the yaw control efficiency η ≥ , is the design minimum value of the yaw control efficiency; if not satisfied, directly increase θ to the design minimum value that meets the η requirement;

[0023] Taking the flight Mach number Ma as the independent variable, establish a linear constraint relationship between the tail fin inclination angle θ and the flight Mach number Ma, and calculate the basic inclination angle of the tail fin inclination angle θ and the flight Mach number through linear interpolation ;

[0024] Compare the yaw control efficiency η with the design minimum value and perform inclination angle correction:

[0025] If η ≥ : η meets the requirements, remain unchanged;

[0026] If η < : η does not meet the yaw control requirement, increase the tail fin inclination angle θ to the corrected inclination angle , the corrected inclination angle is and sum, where is the efficiency compensation inclination angle, obtained by solving the fitting curve calibrated by the wind tunnel test;

[0027] Set the correction step size to 5°, and take... and The maximum value is used as the corrected backslope angle The camber angle will be corrected. The signal is sent to the power unit, where the rotation angles of the left and right tail fins are controlled by adjusting the step size.

[0028] Preferably, the power unit includes a stepper motor and a gear transmission mechanism, wherein the stepper motor is connected to the flight control system, and the gear transmission mechanism is connected between the stepper motor and the tail fin shaft.

[0029] Preferably, the left and right wings are provided with inner grooves at positions corresponding to the left and right tail wings. When the rotation angle of the left and right tail wings is 0°, the left tail wing is embedded in the inner groove of the left wing, and the right tail wing is embedded in the inner groove of the right wing.

[0030] The variable-cantilever tail configuration for high Mach number aircraft described in this application has the following advantages:

[0031] The tail fin serves as a lifting surface at low speeds to increase lift, based on the aerodynamic requirements at different flight speeds. At high speeds, it acts as a directional stabilization and control surface, addressing the issue of insufficient directional stability at high Mach numbers. By rationally selecting the horizontal stabilizer's canard angle to address the shortcomings in the aircraft's aerodynamic characteristics, the peak wing area caused by lift and directional stability requirements in traditional designs is avoided.

[0032] By utilizing the lift function of the tail fin, the area of ​​the fixed wing is reduced, and the reduction can be on par with the area of ​​the tail fin of existing aircraft.

[0033] By adjusting the tail fin tilt angle in real time, the superposition effect of wing area and tail area in traditional design is eliminated, reducing drag across the entire speed range. In particular, the horizontal tail configuration is adopted in the transonic stage, which integrates the tail fin with the wing, reducing the forward projected area and thus reducing drag during transonic flight. Attached Figure Description

[0034] Figure 1 This is a schematic diagram illustrating the variation of directional static stability with Mach number for a conventional layout fixed-wing aircraft in the background art.

[0035] Figure 2 This is a schematic diagram of the variable-angle tail fin horizontal tail configuration of this application;

[0036] Figure 3 This is a schematic diagram of the under-tail configuration of the variable-angle tail fin in this application;

[0037] Figure 4 This is a schematic diagram of the tail fin rotation of the aircraft in this application;

[0038] Figure 5This is a schematic diagram of the vertical tail configuration on the variable tilt tail fin of this application;

[0039] Figure 6 This is a schematic diagram comparing the lift characteristics of the variable tilt tail configuration of this application with those of a fixed-wing aircraft;

[0040] Figure 7 This is a schematic diagram comparing the zero-drag characteristics of the variable-angle tail fin horizontal tail configuration of this application with those of a conventional fixed-wing aircraft;

[0041] Figure 8 This is a schematic diagram illustrating the variation of the yaw control efficiency of the variable tilt tail fin with Mach number in this application.

[0042] Figure 9 This is a schematic diagram illustrating the static stability analysis and tail fin usage scheme of the variable tilt tail aircraft in this application.

[0043] 1. Left wing; 2. Right wing; 3. Left tail; 4. Right tail; 5. Tail pivot. Detailed Implementation

[0044] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are only some, not all, of the embodiments of this application. The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain this application, and should not be construed as limiting this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without inventive effort are within the scope of protection of this application. The embodiments of this application will be described in detail below with reference to the accompanying drawings.

[0045] The first aspect of this application provides a variable-angle tail configuration for high Mach number aircraft, such as Figures 2-4 It includes the left wing 1, the right wing 2, the left tail 3, and the right tail 4;

[0046] The left tail 3 and the right tail 4 are arranged symmetrically along the fuselage symmetry plane. The roots of the left tail 3 and the right tail 4 are both connected to the tail wing pivot 5, which is rotatably connected to the trailing edge of the wing. The rotation angle of the left tail 3 and the right tail 4 is 0° to ±90°. The aircraft is equipped with a power unit, which is connected to the aircraft's flight control system. The aircraft's flight control system can control the left tail 3 and the right tail 4 to rotate along the tail wing pivot 5 from 0° to ±90°.

[0047] Through the symmetrical layout of the left tail fin 3 and the right tail fin 4, shaft connection, and the rotational control design of the flight control + power unit, the tail fin can be precisely adjusted from 0° to ±90°, providing the core structure and control basis for switching the tail fin configuration in different flight stages, and breaking through the limitation of the single function of the fixed tail fin.

[0048] Preferably, when the rotation angles of the left tail fin 3 and the right tail fin 4 are 0°, the aircraft is in the horizontal tail configuration; when the rotation angles of the left tail fin 3 and the right tail fin 4 are 90°, the aircraft is in the lower vertical tail configuration.

[0049] By defining the configuration correspondence of the 0° horizontal tail and the 90° lower vertical tail, the tail fin can be precisely switched between the lift surface and the course stability surface, adapting to the core aerodynamic requirements of low-speed high lift and high-speed high stability, and alleviating the problem of mismatched aerodynamic characteristics at high and low speeds.

[0050] Preferably, a flight data processing module is set in the flight control system. The flight data processing module can obtain the current flight Mach number Ma and control the left tail fin 3 and the right tail fin 4 to rotate to the corresponding angles; when Ma ≤ 1.2 and the course static stability coefficient is as low as -0.00065, control the rotation angles of the left tail fin 3 and the right tail fin 4 to 0°; when Ma = 3.5 and the course static stability coefficient is not lower than -0.0003, control the rotation angles of the left tail fin 3 and the right tail fin 4 to 90°.

[0051] Implement the configuration control with dual-parameter linkage of the Mach number and the course static stability coefficient, precisely matching the aerodynamic constraints of the low / high-speed critical Mach number, and ensuring the low-speed lift efficiency and high-speed course stability.

[0052] Preferably, when the flight data processing module determines that the current flight Mach number is in the transition range: 1.2 < Ma < 3.5, the optimal angle calculation of the tail fin shaft 5 is carried out, specifically:

[0053] Let the tail fin inclination angle be θ, and set the following constraints:

[0054] First, the tail fin inclination angle θ has a positive linear correlation with the flight Mach number Ma;

[0055] Second, the course static stability coefficient of the aircraft calculated in real time needs to meet the linear transition constraint. If not, compensate and correct the tail fin inclination angle θ;

[0056] Third, the yaw control efficiency η ≥ , is the design minimum value of the yaw control efficiency; if not, directly increase θ to meet the design minimum value required by η;

[0057] Using the flight Mach number Ma as the independent variable, a linear constraint relationship between the tail fin cant angle θ and the flight Mach number Ma is established. The tail fin cant angle θ and the basic cant angle of the tail fin at the flight Mach number are calculated by linear interpolation. ;

[0058] The yaw control efficiency η is compared with the design minimum. Compare and correct the tilt angle:

[0059] If η≥ η meets the requirements. Remain unchanged;

[0060] If η < η does not meet the heading control requirements, and the tail fin cant angle θ needs to be increased to the corrected cant angle. Correcting the tilt angle for and The sum of, among which To compensate for the tilt angle in terms of efficiency, the angle calibrated by wind tunnel testing... Obtained by solving the fitted curve;

[0061] Set the correction step size to 5°, and take... and The maximum value is used as the corrected backslope angle The camber angle will be corrected. The signal is sent to the power unit, where the rotation angles of the left tail wing 3 and the right tail wing 4 are controlled by adjusting the step size.

[0062] A method for optimal inclination angle calculation in the transition range is proposed. By using linear interpolation and dual-index compensation correction, the method balances drag reduction and heading control / stability requirements in the transition range, achieves smooth inclination angle adjustment, avoids abrupt changes in aerodynamic characteristics, and improves flight performance across the entire speed range.

[0063] Preferably, the power unit includes a stepper motor and a gear transmission mechanism. The stepper motor is connected to the flight control system, and the gear transmission mechanism is connected between the stepper motor and the tail fin shaft 5. This, in conjunction with the flight control system, enables high-precision, high-response control of the tail fin tilt angle, ensuring the accuracy and timeliness of tilt angle adjustment and adapting to the real-time aerodynamic requirements of the aircraft's dynamic flight.

[0064] Preferably, the left wing 1 and right wing 2 are provided with recesses at positions corresponding to the left tail 3 and right tail 4, such as... Figure 3 When the rotation angle of the left tail 3 and the right tail 4 is 0°, the left tail 3 is embedded in the inner groove of the left wing 1, and the right tail 4 is embedded in the inner groove of the right wing 2.

[0065] By incorporating the internal groove design of the wing, the lower tail fin of the horizontal tail configuration is integrated with the wing, further reducing the forward / horizontal projected area of ​​the entire aircraft, significantly reducing transonic shock wave drag, and optimizing the structural layout to reduce aerodynamic interference during flight.

[0066] In summary, this application has the following advantages:

[0067] The tail fin serves as a lifting surface at low speeds to increase lift, based on the aerodynamic requirements at different flight speeds. At high speeds, it acts as a directional stabilization and control surface, addressing the issue of insufficient directional stability at high Mach numbers. By rationally selecting the horizontal stabilizer's canard angle to address the shortcomings in the aircraft's aerodynamic characteristics, the peak wing area caused by lift and directional stability requirements in traditional designs is avoided.

[0068] By utilizing the lift function of the tail fin, the area of ​​the fixed wing is reduced, and the reduction can be on par with the area of ​​the tail fin of existing aircraft.

[0069] By adjusting the tail fin tilt angle in real time, the superposition effect of wing area and tail area in traditional design is eliminated, reducing drag across the entire speed range. In particular, the horizontal tail configuration is adopted in the transonic stage, which integrates the tail fin with the wing, reducing the forward projected area and thus reducing drag during transonic flight.

[0070] As one specific implementation method, the following is an illustration using a concrete example:

[0071] Combination Figures 2-4 The aircraft employs a horizontal tail configuration for takeoff, landing, and transonic flight, and a downward-facing tail configuration for high-speed cruise flight. To illustrate the benefits of this application, the upward-facing tail configuration aircraft is considered a conventional fixed-wing aircraft, such as... Figure 5 Based on this, a comparative explanation is conducted.

[0072] Among them, the only difference between the upper tail configuration aircraft and the other two tail configuration aircraft is the tail deflection angle; all other components are exactly the same, thus illustrating the benefits brought about by the tail change.

[0073] The overall parameters of the variable-wing aircraft and the conventional fixed-wing aircraft in this case are compared in Table 1 below.

[0074] Table 1. Overall Parameters of Conventional Aircraft and Variable-Wing Aircraft

[0075]

[0076] Meanwhile, the variable-wing aircraft was verified in terms of takeoff and landing lift, transonic drag, and cruise heading stability as follows:

[0077] 1. Lift benefits during takeoff and landing

[0078] During takeoff and landing, the aircraft is analyzed based on typical flight conditions (Ma=0.2, α=10° (angle of attack)). The variable-wing aircraft with a horizontal tail configuration shows a significant increase in lift compared to the fixed-wing aircraft, with an increase of approximately 10%. Figure 6Based on this, the maximum takeoff weight of the aircraft can be increased by 10% equivalently. If the takeoff weight is kept constant, the wing area can be reduced accordingly, thereby reducing the structural weight of the aircraft and optimizing fuel efficiency and empty weight.

[0079] 2. Drag gain of supersonic flight

[0080] In transonic and supersonic flight, the horizontal tail configuration reduces the forward projected area by 8.2 square meters (6.6%) compared to a fixed-wing aircraft, effectively reducing shock wave drag. Flight data shows that at Ma=1.1, drag decreases by 6.5%; while at Ma>1.5, drag decreases by approximately 3.0%, such as... Figure 7 As shown in the figure. The above results verify the drag reduction performance advantages of the horizontal tail configuration in the transonic to supersonic flight range.

[0081] 3. Static stability during high-speed cruising still meets requirements.

[0082] By analyzing the variation of heading control capability with Mach number, such as Figure 8 As shown, in the low-speed range (Ma < 1.1), the horizontal stabilizer configuration exhibits high directional control efficiency; in the supersonic and above range (Ma > 1.1), the control efficiency gradually decreases with increasing Mach number. Based on this characteristic, segmented design constraints are proposed for the aircraft's directional static stability, such as... Figure 9 As shown:

[0083] When Ma=0.2, the permissible directional static stability coefficient of the aircraft is... As low as -0.00065;

[0084] When Ma=3.5, the required heading static stability coefficient is... Not less than -0.0003;

[0085] Within the range of Ma=0.2 to Ma=3.5, the heading static stability coefficient transitions linearly.

[0086] Experimental results show that the above design not only meets the directional static stability requirements of different flight stages, but also effectively avoids the problem of peak overlap of wing area and tail area in fixed-wing design, and significantly improves the aerodynamic performance matching capability of the aircraft in high and low speed flight states.

[0087] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A variable-angle tail configuration for high Mach number aircraft, characterized in that, It includes a left wing (1), a right wing (2), a left tail wing (3) and a right tail wing (4); The left tail wing (3) and the right tail wing (4) are arranged symmetrically with respect to the mirror image of the fuselage symmetry plane. The roots of the left tail wing (3) and the right tail wing (4) are both connected to the tail wing rotating shaft (5), and the tail wing rotating shaft (5) is rotatably connected to the trailing edge of the wing. The rotation angles of the left tail wing (3) and the right tail wing (4) are from 0° to ±90°. A power unit is arranged inside the aircraft, and the power unit is connected to the flight control system of the aircraft. The flight control system of the aircraft can control the left tail wing (3) and the right tail wing (4) to rotate from 0° to ±90° along the tail wing rotating shaft (5); When the rotation angles of the left tail wing (3) and the right tail wing (4) are 0°, the aircraft is in the horizontal tail configuration; when the rotation angles of the left tail wing (3) and the right tail wing (4) are 90°, the aircraft is in the lower vertical tail configuration; The flight control system is equipped with a flight data processing module, which can obtain the current flight Mach number Ma and control the left tail (3) and right tail (4) to rotate to the corresponding angle; when Ma≤1.2, and the yaw static stability coefficient of the aircraft is... When the value is as low as -0.00065, control the rotation angle of the left tail fin (3) and the right tail fin (4) to 0°; when Ma=3.5, and the directional static stability coefficient is... When the value is not lower than -0.0003, control the rotation angle of the left tail wing (3) and the right tail wing (4) to 90°.

2. The variable cantilever tail configuration for high Mach number aircraft as described in claim 1, characterized in that, When the flight data processing module determines that the current flight Mach number is in the transition interval: 1.2 < Ma < 3.5, the optimal angle calculation of the tail wing rotating shaft (5) is carried out, specifically: Let the tail wing inclination angle be θ, and the following constraints are set: First, the tail wing inclination angle θ changes linearly and positively with the flight Mach number Ma; II. Real-time calculation of the aircraft's heading static stability coefficient The linear transition constraint must be met; if it is not met, the tail fin tilt angle θ should be compensated and corrected. III. Yaw control efficiency η≥ , The minimum design value for yaw control efficiency; If not satisfied, directly increase θ to the design minimum value that meets the η requirement; Using the flight Mach number Ma as the independent variable, a linear constraint relationship between the tail fin cant angle θ and the flight Mach number Ma is established. The tail fin cant angle θ and the basic cant angle of the tail fin at the flight Mach number are calculated by linear interpolation. ; The yaw control efficiency η is compared with the design minimum. Compare and correct the tilt angle: If η≥ η meets the requirements. Remain unchanged; If η < η does not meet the heading control requirements, so increase the tail fin cant angle θ to the corrected cant angle. Correcting the tilt angle for and The sum of, among which To compensate for the tilt angle in terms of efficiency, the angle calibrated by wind tunnel testing... Obtained by solving the fitted curve; Set the correction step size to 5°, and take... and The maximum value is used as the corrected backslope angle The camber angle will be corrected. The signal is sent to the power unit, and the rotation angle of the left tail fin (3) and the right tail fin (4) is controlled by combining the correction step size.

3. The variable cantilever tail configuration for high Mach number aircraft as described in claim 1, characterized in that, The power unit includes a stepper motor and a gear transmission mechanism. The stepper motor is connected to the flight control system, and the gear transmission mechanism is connected between the stepper motor and the tail wing rotating shaft (5).

4. The variable cantilever tail configuration for high Mach number aircraft as described in claim 1, characterized in that, Inner grooves are provided at the positions of the left wing (1) and the right wing (2) corresponding to the left tail wing (3) and the right tail wing (4). When the rotation angles of the left tail wing (3) and the right tail wing (4) are 0°, the left tail wing (3) is embedded into the inner groove of the left wing (1), and the right tail wing (4) is embedded into the inner groove of the right wing (2).

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