Structure for adjusting zero lift moment of aircraft and design method

Abstract

The invention belongs to the field of aerodynamic configuration design of aerospace aircrafts, and relates to a structure and a design method for adjusting the zero-lift torque of an aircraft, the structure comprises a non-circular-section fuselage, the tail of the non-circular-section fuselage is contracted, the tail contraction section is provided with a vertical tail and a horizontal tail, and a tail fin is arranged between the horizontal tail and the fuselage and used for adjusting the zero-lift torque of the whole aircraft. The requirement for the aerodynamic performance of the aircraft can be met, and meanwhile the requirement for the installation space of the steering engine can be met through the integrated design of the tail fin and the steering engine.

Description

Technical Field

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

[0002] In the aerodynamic shape design of aircraft, non-circular cross-section fuselage and curved airfoil are used to improve the stealth and aerodynamic performance of aircraft. As the aircraft shape scheme becomes more complex, the zero-lift moment of the shape scheme at each Mach number increases sharply. At present, the adjustment of zero-lift pitch moment of aircraft is mainly achieved by adjusting the airfoil of the wing surface and presetting the installation angle of the tail wing, such as Figure 1 shown.

[0003] The results of numerical simulation analysis and wind tunnel tests show that Figure 1 As shown, the design of changing the airfoil profile and the preset installation angle of the tail reduces the aerodynamic performance of the aircraft. The preset installation angle of the tail reduces the available rudder deflection angle range of the control surface, affects the controllability of the aircraft's control surface, and affects the flight performance of the aircraft. Summary of the invention

[0004] Purpose of the present invention:

[0005] Aiming at the problem that the zero-lift moment of the whole aircraft increases when the shape scheme of the aircraft with non-circular fuselage is complicated, a structure and design method for adjusting the zero-lift moment of the aircraft are proposed, which can not only meet the requirements of the aerodynamic performance of the aircraft, but also meet the requirements of the servo installation space through the integrated design of the tail fin and the servo.

[0006] The tail fin design requirements for adjusting the zero lift moment of the aircraft are: to meet the aircraft's requirements for zero lift moment of the entire aircraft, to meet the requirements for the aircraft's lift-drag characteristics after adding the tail fin, and to meet the requirements for the installation space of the servo in the side fin, thereby improving the utilization of the aircraft's tail space.

[0007] The technical solution of the present invention:

[0008] A structure for adjusting the zero lift moment of an aircraft comprises a non-circular cross-section fuselage, wherein the tail of the non-circular cross-section fuselage is contracted, a vertical tail and a horizontal tail are arranged in the contracted tail section, and a tail fin is arranged between the horizontal tail and the fuselage for adjusting the zero lift moment of the whole aircraft.

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

[0010] Furthermore, the tail fin is arranged symmetrically with respect to the symmetry plane of the fuselage.

[0011] Furthermore, 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 requirement for installing the tail rudder control servo internally; the chord length of the tail fin is determined by the length of the contracted section of the fuselage and the magnitude of the zero-lift moment to be adjusted; the maximum thickness of the tail fin is determined by the space requirement of the tail fin control servo for the tail fin; the maximum curvature and installation angle of the tail fin meet the requirements of the tail fin for the amount of adjustment required for the zero-lift moment; the span of the tail fin is designed according to the overall requirements for the outer envelope of the aircraft to meet the requirement that the span of the tail fin is not greater than the outer envelope of the aircraft; the design of the area size of the tail fin meets the requirement 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 layout of the tail servo and the layout requirements of the normal position of the tail fin.

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

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

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

[0015] Furthermore, the position of the maximum curvature of the tail fin is where the tail fin is 35% to 65% of the fuselage along the heading of the aircraft.

[0016] A design method for adjusting the zero-lift moment of an aircraft comprises the following steps:

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

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

[0019] The third step 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 the range of ±4 degrees according to the aircraft's requirements for adjusting the zero lift moment.

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

[0021] Step 5: Design the outer surface of the tail fin according to the aerodynamic performance of the aircraft and the space requirements of the tail fin;

[0022] Step 6: Perform CFD calculation or wind tunnel test verification on the designed tail fin shape of the aircraft;

[0023] The seventh step is to analyze whether the aerodynamic performance of the tail fin shape of the aircraft meets the requirements. If it meets the design requirements, complete the structural design of adjusting the zero-lift moment of the aircraft; if it does not meet the design requirements, repeat the above steps 3 to 7 to optimize the design until it meets the design requirements.

[0024] Beneficial effects of the present invention

[0025] The present invention proposes a structure and a design method for adjusting the zero-lift moment of an aircraft, and in particular, the design for the tail fin satisfies the demand for adjusting the zero-lift moment of the aircraft, with little influence on the lift-drag characteristics and handling characteristics of the aircraft. At the same time, the integrated design with the tail servo meets the requirements for the servo installation space, and effectively solves the space requirements for installing components at the tail of the aircraft.

[0026] The tail fin design requirements for adjusting the zero lift moment of the aircraft are: to meet the aircraft's requirements for zero lift moment of the entire aircraft, to meet the requirements for the aircraft's lift-drag characteristics after adding the tail fin, and to meet the requirements for the installation space of the servo in the side fin, thereby improving the utilization of the aircraft's tail space. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] Figure 1 It is a schematic diagram of the traditional design by changing the airfoil profile and the preset installation angle of the tail wing;

[0028] Figure 2 1 is a schematic diagram (side view) of a tail fin design for adjusting the zero lift moment of an aircraft according to the present invention;

[0029] Figure 3 1 is a schematic diagram of the tail fin design for adjusting the zero lift moment of the aircraft according to the present invention (bottom view);

[0030] Among them, 1. Fuselage, 2. Vertical tail, 3. Horizontal tail, 4. Tail fin. DETAILED DESCRIPTION

[0031] Aiming at the problem of large zero-lift moment of aircraft, a structure for adjusting the zero-lift moment of aircraft is proposed; specifically, a tail fin for adjusting the zero-lift moment is added at the contraction correction position of the tail section of the fuselage, which can meet the demand for adjusting the zero-lift moment and reduce the influence on the control characteristics of the rudder surface and the aerodynamic performance of the aircraft. At the same time, the requirement of the servo installation space can be met through the integrated design of the tail fin and the servo.

[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 of the non-circular cross-section fuselage is contracted, a vertical tail and a horizontal tail are arranged at the contracted tail section, and a tail fin is arranged between the horizontal tail and the fuselage for adjusting the zero lift moment of the whole aircraft, wherein the tail fin meets the following design requirements:

[0033] 1) A tail fin that can adjust the zero lift moment of the entire aircraft is designed in the contraction section of the tail of the fuselage. 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 installing the tail rudder control servo inside.

[0034] 2) The tail fin is located on both sides of the contracted section at the rear of the fuselage. The axial front-to-back position of the fuselage matches the tail wing direction, and is adaptively adjusted according to the requirements of the zero-lift moment to be adjusted.

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

[0036] 4) The chord length of the tail fin is adaptively designed according to the length of the retracted section of the fuselage and the size of the zero-lift moment that needs to be adjusted.

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

[0038] 6) The maximum curvature and installation angle of the tail fin are designed to adapt to the requirements of the tail fin to meet the amount of adjustment required to meet the zero lift moment.

[0039] 7) The length of the tail fin is designed according to the overall requirements of the outer envelope of the aircraft to ensure that the tail fin length is not greater than the outer envelope of the aircraft.

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

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

[0042] The specific design parameters are:

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

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

[0045] The maximum thickness position of the tail fin is at a position where the tail fin is 30% to 50% of the fuselage along the heading of the aircraft.

[0046] The position of the maximum curvature of the tail fin is where the tail fin is 35% to 65% of the fuselage along the heading of the aircraft.

[0047] Through specific parameter design, the demand for adjusting the zero-lift moment is met, and the impact on the lift-drag and handling characteristics of the aircraft is relatively small. At the same time, the demand for installation space for the tail servo is solved.

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

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

[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 designed Mach number aircraft requires a trim control surface greater than 3 degrees, which does not meet the requirements; specifically, a tail fin shape scheme to adjust the zero lift moment.

[0051] The third step 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 the range of ±4 degrees according to the aircraft's requirements for adjusting the zero lift moment.

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

[0053] Step 5: Design the outer surface of the tail fin according to the aerodynamic performance of the aircraft and the space requirements of the tail fin;

[0054] Step 6: Perform CFD calculation or wind tunnel test verification on the designed tail fin shape of the aircraft;

[0055] The seventh step is to analyze whether the aerodynamic performance of the tail fin shape of the aircraft meets the requirements. If it meets the design requirements, complete the structural design of adjusting the zero-lift moment of the aircraft; if it does not meet the design requirements, repeat the above steps 3 to 7 to optimize the design until it meets the design requirements.

[0056] The design of the tail fin of the present invention meets the demand for adjusting the zero-lift moment of the aircraft, and has little effect on the lift-drag characteristics and handling characteristics of the aircraft. At the same time, the integrated design with the tail servo meets the requirements for the servo installation space, effectively solving the space requirements for installing components at the tail of the aircraft.

Claims

1. A structure for adjusting the zero lift moment of an aircraft, characterized in that: The invention comprises a non-circular cross-section fuselage, wherein the tail of the non-circular cross-section fuselage is contracted, a vertical tail and a horizontal tail are arranged in the contracted tail section, and a tail fin is arranged between the horizontal tail and the fuselage for adjusting the zero lift moment of the whole aircraft.

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

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 with respect to the symmetric plane of the fuselage.

4. The structure for adjusting the zero-lift moment of an aircraft according to claim 1, characterized in that: 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 requirement for installing the tail rudder control servo inside; the chord length of the tail fin is determined by the length of the contracted section of the fuselage and the magnitude of the zero-lift moment to be adjusted; the maximum thickness of the tail fin is determined by the space requirement of the tail fin control servo for the tail fin; the maximum curvature and installation angle of the tail fin meet the requirements of the tail fin for the amount of adjustment required for the zero-lift moment; the span of the tail fin is designed according to the overall requirements for the outer envelope of the aircraft to meet the requirement that the span of the tail wing is not greater than the outer envelope of the aircraft; the design of the area size of the tail fin meets the requirement 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 layout of the tail servo and the layout requirements of the normal position of the tail wing.

5. The structure for adjusting the zero-lift moment of an aircraft according to claim 4, characterized in that: The thickness of the tail fin is 6% to 10% of the length of the tail fin.

6. The structure for adjusting the zero-lift moment of an aircraft according to claim 4, characterized in that: The installation angle of the tail fin is within the range of ±4°.

7. The structure for adjusting the zero-lift moment of an aircraft according to claim 4, characterized in that: The maximum thickness position of the tail fin is at a position where the tail fin is 30% to 50% of the fuselage along the heading of the aircraft.

8. The structure for adjusting the zero-lift moment of an aircraft according to claim 4, characterized in that: The position of the maximum curvature of the tail fin is where the tail fin is 35% to 65% of the fuselage along the heading of the aircraft.

9. A method for designing a structure for adjusting zero-lift moment of an aircraft according to any one of claims 1 to 8, characterized in that: The following steps are involved: The first step is to obtain zero-lift moment data of the aircraft at different design Mach numbers through CFD calculation or wind tunnel test; The second step is to design and adjust the structural shape of the zero-lift moment of the aircraft if the zero-lift moment data does not meet the requirements; The third step 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 the range of ±4 degrees according to the aircraft's requirements for adjusting the zero lift moment. The fourth step is to design and determine the maximum thickness of the tail fin to be 30% to 50% of the fuselage along the aircraft heading, and the maximum curvature to be 35% to 65% of the fuselage along the aircraft heading, according to the design requirements of the overall structure and rudder system of the aircraft. Step 5: Design the outer surface of the tail fin according to the aerodynamic performance of the aircraft and the space requirements of the tail fin; Step 6: Perform CFD calculation or wind tunnel test verification on the designed tail fin shape of the aircraft; Step 7: Analyze whether the aerodynamic performance of the tail fin shape of the aircraft meets the requirements. If it meets the design requirements, complete the structural design of adjusting the zero lift moment of the aircraft. If the design requirements are not met, repeat the above steps 3 to 7 to optimize the design until the design requirements are met.

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