A structure and design method for adjusting the zero lift moment of an aircraft

Abstract

This invention belongs to the field of aerodynamic shape design of aerospace vehicles, and relates to a structure and design method for adjusting the zero lift moment of an aircraft. The structure includes a non-circular cross-section fuselage, the tail of which is tapered, and a vertical tail and a horizontal tail are provided in the tapered tail section. A tail fin is provided between the horizontal tail and the fuselage for adjusting the zero lift moment of the entire aircraft. This invention can meet the aerodynamic performance requirements of the aircraft, and at the same time, the integrated design of the tail fin and the servo motor can meet the requirements of the servo motor installation space.

Description

Technical Field

[0001] This invention belongs to the field of aerodynamic shape design of aerospace vehicles, and relates to a structure and design method for adjusting the zero lift moment of an aircraft. Background Technology

[0002] In aircraft aerodynamic design, non-circular cross-section fuselages and cambered airfoils are used to improve stealth and aerodynamic performance. As aircraft shape becomes more complex, the zero-lift moment at various Mach numbers also increases dramatically. Currently, adjusting the zero-lift pitch moment is mainly achieved by adjusting the airfoil shape and the pre-set angle of the tail fin, such as... Figure 1 As shown.

[0003] Numerical simulation analysis and wind tunnel test results show that Figure 1 The design shown reduces the aerodynamic performance of the aircraft by changing the airfoil profile and the pre-set installation angle of the tail fin. The pre-set installation angle of the tail fin reduces the range of available control deflection angles, affecting the controllability of the aircraft and thus impacting its flight performance. Summary of the Invention

[0004] The purpose of this invention:

[0005] To address the issue of increased zero-lift moment in aircraft with complex non-circular cross-section fuselage designs, a structural and design method for adjusting the zero-lift moment is proposed. This method satisfies the aerodynamic performance requirements of the aircraft while also meeting the space requirements for servo motor installation through the integrated design of the tail fin and servo motor.

[0006] The design requirements for the tail fin of an aircraft to adjust zero lift moment are as follows: to meet the requirements of the aircraft for zero lift moment of the whole aircraft, to meet the requirements of the lift and drag characteristics of the aircraft after adding the tail fin, to meet the requirements of the space for the servo motor to be arranged and installed in the side fin, and to improve the space utilization requirements of the aircraft's tail.

[0007] The technical solution of the present invention:

[0008] A structure for adjusting the zero lift moment of an aircraft includes a non-circular cross-section fuselage, the tail of which is tapered, and a vertical tail and a horizontal tail are provided in the tapered tail section. A tail fin is provided between the horizontal tail and the fuselage for adjusting the zero lift moment of the entire aircraft.

[0009] Furthermore, the tail fin is positioned between the fuselage surface on both sides of the tail section and the horizontal tail.

[0010] Furthermore, the tail fin is arranged symmetrically about the fuselage.

[0011] Furthermore, the overall envelope of the tail fin is no greater than the maximum envelope requirement of the fuselage, and the thickness of the tail fin meets the space requirements for the internal installation of the tail fin control servo. The chord length of the tail fin is determined by the length of the fuselage contraction section and the required zero-lift moment. The maximum thickness of the tail fin is determined by the space requirements of the tail fin control servo. The maximum curvature and installation angle of the tail fin meet the requirements of the amount of zero-lift moment adjustment. The span of the tail fin is designed according to the overall requirements for the aircraft's outer envelope, ensuring that the tail fin span is no greater than the aircraft's outer envelope requirement. The size of the tail fin area meets the requirements for zero-lift moment adjustment of the entire aircraft. The normal position of the tail fin at the tail section of the fuselage meets the overall requirements for the arrangement of the tail fin servo and the layout requirements of the normal position of the tail fin.

[0012] Furthermore, the thickness of the caudal fin is 6% to 10% of the caudal fin length.

[0013] Furthermore, the installation angle of the tail fin is within the range of ±4°.

[0014] Furthermore, the maximum thickness of the tail fin is located at 30% to 50% of the fuselage along the aircraft's heading.

[0015] Furthermore, the maximum curvature of the tail fin is located at 35% to 65% of the fuselage along the aircraft's heading.

[0016] A design method for a structure for adjusting the zero lift moment of an aircraft includes the following steps:

[0017] The first step is to obtain zero-lift moment data of the aircraft under different design Mach numbers through CFD calculations or wind tunnel tests;

[0018] The second step is to design a structural shape scheme to adjust the zero-lift moment of the aircraft if the zero-lift moment data does not meet the requirements.

[0019] The third step, based on the aircraft's requirements for adjusting zero lift torque, involves designing the tail fin's thickness to be 6%–10% of its length, its span to be no greater than the fuselage envelope, and its installation angle to be within ±4 degrees.

[0020] The fourth step, based on the design requirements of the overall aircraft, structure and rudder system, is to design and determine the maximum thickness of the tail fin at 30% to 50% of the fuselage along the aircraft's heading, and the maximum curvature at 35% to 65% of the fuselage along the aircraft's heading.

[0021] The fifth step is to design the external curved surface of the tail fin based on the aerodynamic performance of the aircraft and the space requirements of the tail fin.

[0022] The sixth step is to perform CFD calculations or wind tunnel tests to verify the designed tail fin shape scheme of the aircraft.

[0023] Step 7: Analyze whether the aerodynamic performance of the aircraft's tail fin shape meets the requirements. If it does, complete the structural design for adjusting the zero-lift moment of the aircraft. If it does not meet the design requirements, repeat steps 3 to 7 above to optimize the design until the design requirements are met.

[0024] Beneficial effects of the present invention

[0025] This invention proposes a structure and design method for adjusting the zero lift moment of an aircraft. In particular, the design of the tail fin not only meets the requirement of adjusting the zero lift moment of the aircraft, but also has little impact on the lift-drag characteristics and handling characteristics of the aircraft. At the same time, the integrated design with the tail fin servo also meets the requirements of servo installation space, effectively solving the space requirements of the aircraft tail installation components.

[0026] The design requirements for the tail fin of an aircraft to adjust zero lift moment are as follows: to meet the requirements of the aircraft for zero lift moment of the whole aircraft, to meet the requirements of the lift and drag characteristics of the aircraft after adding the tail fin, to meet the requirements of the space for the servo motor to be arranged and installed in the side fin, and to improve the space utilization requirements of the aircraft's tail. Attached Figure Description

[0027] Figure 1 This is a traditional design schematic diagram that modifies the airfoil profile and the pre-set installation angle of the tail fin.

[0028] Figure 2 This is a schematic diagram (side view) of the tail fin design for adjusting the zero lift moment of an aircraft according to the present invention;

[0029] Figure 3 This is a schematic diagram (bottom view) of the tail fin design for adjusting the zero lift moment of an aircraft according to the present invention;

[0030] Among them, 1. fuselage, 2. vertical tail, 3. horizontal tail, and 4. tail fin. Detailed Implementation

[0031] To address the issue of excessive zero-lift moment in aircraft, a structure for adjusting zero-lift moment is proposed. Specifically, a tail fin for adjusting zero-lift moment is added at the tail section of the fuselage. This satisfies the need for adjusting zero-lift moment while reducing the impact on control surface characteristics and aircraft aerodynamic performance. Furthermore, the integrated design of the tail fin and servo motor also meets the space requirements for servo motor installation.

[0032] One embodiment of the present invention is a structure for adjusting the zero-lift moment of an aircraft, comprising a non-circular cross-section fuselage, wherein the tail section of the non-circular cross-section fuselage is tapered, and a vertical stabilizer and a horizontal stabilizer are provided in the tapered tail section. A tail fin is provided between the horizontal stabilizer and the fuselage for adjusting the zero-lift moment of the entire aircraft. The tail fin meets the following design requirements:

[0033] 1) A tail fin is designed in the tapered section of the fuselage to adjust the zero lift moment of the entire aircraft. The overall envelope of the tail fin is not greater than the maximum envelope requirement of the fuselage, and the thickness of the tail fin meets the space requirements for the internal installation of the tail fin control servo.

[0034] 2) The tail fins are located on both sides of the tail section of the fuselage. The forward and backward positions of the fuselage axis match the direction of the tail fin, and are adaptively adjusted according to the required zero lift moment.

[0035] 3) The tail fin design is symmetrical about the fuselage.

[0036] 4) The chord length of the tail fin is adaptively designed based on the length of the fuselage retraction section and the required zero lift torque.

[0037] 5) The maximum thickness of the tail fin is designed according to the space requirements of the tail fin control servo.

[0038] 6) The maximum curvature and installation angle of the tail fin should be adjusted to meet the requirements of zero lift torque and adapted to the tail fin design.

[0039] 7) The span of the tail fin is designed according to the overall requirements of the aircraft's outer envelope, ensuring that the tail fin span does not exceed the requirements of the aircraft's outer envelope.

[0040] 8) The size of the tail fin is designed to meet the requirements of zero-lift torque adjustment for the entire aircraft.

[0041] 9) The tail fin's position in the normal direction of the fuselage section meets the overall requirements for the arrangement of the tail fin servo and the layout requirements for the normal position of the tail fin.

[0042] The specific design parameters are as follows:

[0043] The thickness of the caudal fin is 6% to 10% of its length.

[0044] The installation angle of the caudal fin is within ±4°.

[0045] The maximum thickness of the tail fin is located at 30% to 50% of the fuselage along the aircraft's heading.

[0046] The maximum curvature of the tail fin is located at 35% to 65% of the fuselage along the aircraft's heading.

[0047] Through specific parameter design, the requirement for adjusting zero lift torque was met, while having minimal impact on the aircraft's lift-drag and handling characteristics, and the requirement for tail servo installation space was also resolved.

[0048] Another embodiment of the present invention provides a design method for a structure for adjusting the zero lift moment of an aircraft, comprising the following steps:

[0049] The first step is to obtain zero-lift moment data of the aircraft under different design Mach numbers through CFD calculations or wind tunnel tests;

[0050] The second step is to design a structural shape scheme to adjust the zero-lift moment of the aircraft if the zero-lift moment of the aircraft does not meet the requirements of the trim control surface which requires a zero-lift moment of greater than 3 degrees. Specifically, a tail fin shape scheme to adjust the zero-lift moment is proposed.

[0051] The third step, based on the aircraft's requirements for adjusting zero lift torque, involves designing the tail fin's thickness to be 6%–10% of its length, its span to be no greater than the fuselage envelope, and its installation angle to be within ±4 degrees.

[0052] The fourth step, based on the design requirements of the overall aircraft, structure and rudder system, is to design and determine the maximum thickness of the tail fin at 30% to 50% of the fuselage along the aircraft's heading, and the maximum curvature at 35% to 65% of the fuselage along the aircraft's heading.

[0053] The fifth step is to design the external curved surface of the tail fin based on the aerodynamic performance of the aircraft and the space requirements of the tail fin.

[0054] The sixth step is to perform CFD calculations or wind tunnel tests to verify the designed tail fin shape scheme of the aircraft.

[0055] Step 7: Analyze whether the aerodynamic performance of the aircraft's tail fin shape meets the requirements. If it does, complete the structural design for adjusting the zero-lift moment of the aircraft. If it does not meet the design requirements, repeat steps 3 to 7 above to optimize the design until the design requirements are met.

[0056] The tail fin design of this invention not only meets the requirement of adjusting the zero lift torque of the aircraft, but also has a small impact on the lift-drag characteristics and handling characteristics of the aircraft. At the same time, the integrated design with the tail fin servo also meets the requirements of servo installation space, effectively solving the space requirements of the aircraft tail installation components.

Claims

1. A structure for adjusting the zero-lift moment of an aircraft, characterized in that, The fuselage includes a non-circular cross-section fuselage with a tapered tail section. A vertical stabilizer and a horizontal stabilizer are installed in the tapered tail section, and a tail fin is positioned between the horizontal stabilizer and the fuselage to adjust the zero-lift moment of the entire aircraft. The overall envelope of the tail fin is no greater than the maximum envelope requirement of the fuselage, and the thickness of the tail fin meets the space requirements for the internal tail fin control servo mechanism. The chord length of the tail fin is determined by the length of the tapered tail section and the magnitude of the zero-lift moment to be adjusted. The maximum thickness of the tail fin is determined by the space requirements of the tail fin control servo mechanism. The maximum camber and installation angle of the tail fin meet the requirements of the amount of zero-lift moment adjustment. The span of the tail fin is designed according to the overall requirements for the aircraft's outer envelope, ensuring that the tail fin span does not exceed the aircraft's outer envelope requirement. The size of the tail fin area meets the requirements for zero-lift moment adjustment of the entire aircraft. The normal position of the tail fin in the tail section of the fuselage meets the overall requirements for the tail fin servo arrangement and the layout requirements of the tail fin's normal position. The thickness of the caudal fin is 6% to 10% of the caudal fin length; The installation angle of the caudal fin is within ±4°; The maximum thickness of the tail fin is located at 30% to 50% of the fuselage along the aircraft's heading. The maximum curvature of the tail fin is located at 35% to 65% of the fuselage along the aircraft's heading.

2. The structure for adjusting the zero-lift moment of an aircraft according to claim 1, characterized in that, The tail fin is located between the fuselage surface on both sides of the tail section and the horizontal tail.

3. The structure for adjusting the zero-lift moment of an aircraft according to claim 2, characterized in that, The tail fin is arranged symmetrically about the fuselage.

4. A design method for a structure for adjusting the zero-lift moment of an aircraft as described in any one of claims 1-3, characterized in that, Includes the following steps: The first step is to obtain zero-lift moment data of the aircraft under different design Mach numbers through CFD calculations or wind tunnel tests; The second step is to design a structural shape scheme to adjust the zero-lift moment of the aircraft if the zero-lift moment data does not meet the requirements. The third step, based on the aircraft's requirements for adjusting zero lift torque, is to design the thickness of the tail fin to be 6% to 10% of its length, the span to be no greater than the fuselage envelope, and the installation angle to be within ±4 degrees. The fourth step, based on the design requirements of the overall aircraft, structure and rudder system, is to design and determine the maximum thickness of the tail fin at 30% to 50% of the fuselage along the aircraft's heading, and the maximum curvature at 35% to 65% of the fuselage along the aircraft's heading. The fifth step is to design the external curved surface of the tail fin based on the aerodynamic performance of the aircraft and the space requirements of the tail fin. The sixth step is to perform CFD calculations or wind tunnel tests to verify the designed tail fin shape scheme of the aircraft. The seventh step is to analyze whether the aerodynamic performance of the tail fin shape of the aircraft meets the requirements. If it does, the structural design for adjusting the zero lift moment of the aircraft is completed. If the design requirements are not met, repeat steps three through seven above to optimize the design until the design requirements are met.

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