Aerodynamic appearance structure of supercritical wing

Abstract

The invention belongs to the technical field of aviation aircraft design, and particularly relates to an aerodynamic configuration structure of a supercritical wing, the wing is generated by supercritical airfoil profile control of five specific stations in the spanwise direction, and 45% of the stations are reference airfoils. The front edge radius, the thickness and the mounting angle of the inner side (0% and 20% station) airfoil profile are sequentially larger than those of the reference airfoil profile; the leading edge radius of the outboard (70%, 100% station) airfoil profile is significantly larger than that of the reference airfoil profile, the mounting angle is negative, and the absolute values are sequentially increased. And front and rear edge points of each section are connected by using secondary splines to form front and rear edges of the wing, so that a three-dimensional molded surface is generated. Through differential design of spanwise airfoil parameters, a large-area separation area under a large angle of attack can be actively controlled to appear in the middle of the airfoil firstly, control failure caused by stalling of an outer wing and sudden lift drop caused by stalling of an inner wing are effectively avoided, and the stability of the airfoil is improved while the transonic cruise efficiency is maintained. And the stall safety of the upper single-wing transport plane is obviously improved.

Description

Technical Field

[0001] This application belongs to the field of aircraft design technology, and specifically relates to a supercritical wing aerodynamic shape structure. Background Technology

[0002] Modern transport aircraft generally employ supercritical wings with swept angles to adapt to transonic flight and reduce shock wave drag. However, when the angle of attack is large, large-area separation occurs in the wings. When large-area separation occurs on the outer wings, the aircraft experiences roll and the ailerons fail, resulting in a loss of safety. To solve this problem, the traditional method is to control large-area separation to occur on the inner wings. However, for high-wing monoplanes, when large-area separation occurs on the inner wings, the aircraft will suffer a significant loss of lift, and the aircraft will drop altitude rapidly, which can easily lead to danger. Summary of the Invention

[0003] To address the aforementioned issues, this application provides a supercritical airfoil aerodynamic shape structure, comprising: an integral airfoil generated by supercritical airfoil profiles at five specific spanwise positions;

[0004] Among them, the positions of the five airfoil sections along the wing span are 0%, 20%, 45%, 70% and 100% half span, respectively, and the third airfoil section (3) located at 45% half span is the reference airfoil.

[0005] The first airfoil section (1) and the second airfoil section (2) located at 0% half span and 20% half span, relative to the reference airfoil (3), have increasing leading edge radius, relative thickness and local installation angle with the airfoil leading edge point as the base point;

[0006] The fourth airfoil section (4) and the fifth airfoil section (5) located at 70% half span and 100% half span, respectively, have a leading edge radius greater than that of the first airfoil section (1) and the second airfoil section (2) relative to the reference airfoil (3), and a relative thickness less than or equal to that of the reference airfoil (3). The local installation angle with the airfoil leading edge point as the base point is negative and its absolute value increases.

[0007] The leading edge points of the five airfoil sections are connected by a quadratic spline to form the leading edge line of the wing, and their trailing edge points are connected by another quadratic spline to form the trailing edge line of the wing.

[0008] The three-dimensional profile of the wing is generated by the five airfoil profiles, the leading edge line, and the trailing edge line as control lines.

[0009] Preferably, the airfoil parameters of the reference airfoil (3) are: leading edge radius 0.0122, relative thickness 11.5%, local installation angle 0.1°, maximum thickness at 40.4% chord length, relative camber 1.45%, and maximum camber at 79.4% chord length.

[0010] Preferably, the airfoil parameters of the first airfoil section (1) are: leading edge radius 0.0132, relative thickness 13%, local installation angle 3°, maximum thickness at 35.2% chord length, relative camber 1.24%, and maximum camber at 77.6% chord length.

[0011] Preferably, the second airfoil section (2) located at 20% half span has the following airfoil parameters: leading edge radius 0.0136, relative thickness 12%, local installation angle 1.3°, maximum thickness at 40.3% chord length, relative camber 1.44%, and maximum camber at 80.3% chord length.

[0012] Preferably, the fourth airfoil section (4) located at 70% half span has the following airfoil parameters: leading edge radius 0.0280, relative thickness 11.5%, local installation angle -2.2°, maximum thickness at 36.5% chord length, relative camber 1.84%, and maximum camber at 76.5% chord length.

[0013] Preferably, the fifth airfoil section (5), located at 100% half span and serving as the wingtip airfoil, has the following airfoil parameters: leading edge radius 0.0280, relative thickness 11.5%, local installation angle -3.8°, maximum thickness at 36.5% chord length, relative camber 1.84%, and maximum camber at 76.5% chord length.

[0014] Preferably, the spatial coordinates of the leading edge points of the five airfoil profiles are determined by the overall wing parameters, so that the final generated wing has a sweep angle of 25°, an aspect ratio of 8, and a root-to-tip ratio of 4.

[0015] Preferably, the coordinates of the leading edge points of the five airfoil sections are: (0, 0, 0), (0.1083R, 0.2R, -0.0105R), (0.2436R, 0.45R, -0.0236R), (0.3789R, 0.7R, -0.0367R), (0.5413R, 1.0R, -0.0524R), where R is the wing half-span.

[0016] Preferably, the overall twist angle of the wing is 6.8°, and there is no dihedral.

[0017] Preferably, the spanwise position of the wing root airfoil (1) is aligned with the plane of symmetry of the aircraft, with its upper wing surface protruding from the fuselage and its lower wing surface fused to the fuselage.

[0018] Preferably, by matching the airfoil parameters according to any one of claims 1 to 6 with the geometric layout according to any one of claims 7 to 9, the large airflow separation area of ​​the wing at high angles of attack is controlled to appear first in the middle of the span.

[0019] Preferably, the wing is used in a high-wing transport aircraft.

[0020] By designing the wing airfoil, spanwise twist distribution, and spanwise airfoil matching aerodynamic shape, the large angle-of-attack separation of the wing is controlled in the middle of the wing, which not only does not significantly reduce lift but also avoids the outer wing area, giving the aircraft good safety and maintaining the characteristics of a supercritical wing, resulting in excellent transonic cruise aerodynamic performance. Attached Figure Description

[0021] Figure 1 This is a schematic diagram of the wing shape of the present invention;

[0022] Figure 2 This is a schematic diagram of the airfoil at each spanwise section of the wing of the present invention;

[0023] Figure 3 This is a schematic diagram of the large-area separation in the middle of the wing of the present invention. Detailed Implementation

[0024] 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 creative 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. Figures 1-3 As shown, a supercritical wing suitable for mid-span stall in high-wing aircraft is described. The wing is connected to the upper part of the fuselage. The wing layout parameters are a sweep angle of 25°, an aspect ratio of 8, a root-to-tip ratio of 4, a twist angle of 6.8°, and no dihedral. The wing is formed by configuring five supercritical airfoil schemes.

[0025] The airfoil curves for all five airfoils were generated using the quadratic spline method.

[0026] Section 1 shows the airfoil at the wing root, which is a supercritical airfoil located at 0% of the wing's half-span. The wing chord length is 40% of the half-span. The airfoil's leading edge radius is 0.0132, trailing edge angle is 7.79°, local installation angle is 3°, relative thickness is 13%, maximum thickness is 35.2%, relative camber is 1.24%, maximum camber is 77.6%, and the coordinates of the leading edge point of the section are (0,0,0).

[0027] Section 2 shows a supercritical airfoil located at 20% of the wing's half-span. The wing chord length is 34% of the half-span. The airfoil has a leading edge radius of 0.0136, a trailing edge angle of 2.12°, a local installation angle of 1.3°, a relative thickness of 12%, a maximum thickness of 40.3%, a relative camber of 1.44%, and a maximum camber of 80.3%. The coordinates of the leading edge point in the section are (0.1083R, 0.2R, -0.0105R, where R is the half-span).

[0028] Section 3 shows a supercritical airfoil located at 45% of the wing's half-span. The wing chord length is 26.5% of the half-span. The airfoil's leading edge radius is 0.0122, trailing edge angle is 3.05°, local installation angle is 0.1°, relative thickness is 11.5%, maximum thickness is 40.4%, relative camber is 1.45%, maximum camber is 79.4%, and the coordinates of the leading edge point in the section are (0.2436R, 0.45R, -0.0236R, where R is the half-span).

[0029] Section 4 shows a supercritical airfoil located at 70% of the wing's half-span. The wing chord length is 19% of the half-span. The airfoil's leading edge radius is 0.0280, trailing edge angle is 5.06°, local installation angle is -2.2°, relative thickness is 11.5%, maximum thickness is 36.5%, relative camber is 1.84%, maximum camber is 76.5%, and the coordinates of the leading edge point in the section are (0.3789R, 0.7R, -0.0367R, where R is the half-span).

[0030] Section 5 shows the wingtip airfoil, a supercritical airfoil located at 100% of the wing's half-span. The wing chord length is 10% of the half-span. The airfoil's leading edge radius is 0.0280, trailing edge angle is 5.06°, local angle of installation is -3.8°, relative thickness is 11.5%, maximum thickness is 36.5%, relative camber is 1.84%, maximum camber is 76.5%, and the coordinates of the leading edge point in the section are (0.5413R, 1.0R, -0.0524R, where R is the half-span).

[0031] The wing twist angle distribution is obtained by rotating the leading edge points of each airfoil section.

[0032] The wing root airfoil spanwise position is aligned with the aircraft's plane of symmetry. The upper wing surface should protrude from the fuselage longitudinally, while the lower wing surface can be integrated with the fuselage.

[0033] The five leading edge points of the airfoil are connected by a quadratic spline to form the leading edge line of the wing, and the five trailing edge points of the airfoil are connected by a quadratic spline to form the trailing edge line of the wing.

[0034] The airfoil uses five airfoil curves to generate a multi-section surface, and uses leading and trailing edge control lines for surface control.

[0035] 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 supercritical airfoil aerodynamic shape structure, characterized in that, include: An integral wing generated by supercritical airfoil profiles at five specific spanwise positions; Among them, the positions of the five airfoil sections along the wing span are 0%, 20%, 45%, 70% and 100% half span, respectively, and the third airfoil section (3) located at 45% half span is the reference airfoil. The first airfoil section (1) and the second airfoil section (2) located at 0% half span and 20% half span, relative to the reference airfoil (3), have increasing leading edge radius, relative thickness and local installation angle with the airfoil leading edge point as the base point; The fourth airfoil section (4) and the fifth airfoil section (5) located at 70% half span and 100% half span, respectively, have a leading edge radius greater than that of the first airfoil section (1) and the second airfoil section (2) relative to the reference airfoil (3), and a relative thickness less than or equal to that of the reference airfoil (3). The local installation angle with the airfoil leading edge point as the base point is negative and its absolute value increases. The leading edge points of the five airfoil sections are connected by a quadratic spline to form the leading edge line of the wing, and their trailing edge points are connected by another quadratic spline to form the trailing edge line of the wing. The three-dimensional profile of the wing is generated by the five airfoil profiles, the leading edge line, and the trailing edge line as control lines.

2. The supercritical airfoil aerodynamic shape structure according to claim 1, characterized in that, The airfoil parameters of the reference airfoil (3) are: leading edge radius 0.0122, relative thickness 11.5%, local installation angle 0.1°, maximum thickness at 40.4% chord length, relative camber 1.45%, and maximum camber at 79.4% chord length.

3. The supercritical airfoil aerodynamic shape structure according to claim 2, characterized in that, The airfoil parameters of the first airfoil section (1) are: leading edge radius 0.0132, relative thickness 13%, local installation angle 3°, maximum thickness at 35.2% chord length, relative camber 1.24%, and maximum camber at 77.6% chord length.

4. The supercritical airfoil aerodynamic shape structure according to claim 2, characterized in that, The second airfoil section (2) located at 20% half span has the following airfoil parameters: leading edge radius 0.0136, relative thickness 12%, local installation angle 1.3°, maximum thickness at 40.3% chord length, relative camber 1.44%, and maximum camber at 80.3% chord length.

5. The supercritical airfoil aerodynamic shape structure according to claim 2, characterized in that, The fourth airfoil section (4) located at 70% half span has the following airfoil parameters: leading edge radius 0.0280, relative thickness 11.5%, local installation angle -2.2°, maximum thickness at 36.5% chord length, relative camber 1.84%, and maximum camber at 76.5% chord length.

6. The supercritical airfoil aerodynamic shape structure according to claim 2, characterized in that, The fifth airfoil section (5), located at 100% half span and serving as the wingtip airfoil, has the following airfoil parameters: leading edge radius 0.0280, relative thickness 11.5%, local installation angle -3.8°, maximum thickness at 36.5% chord length, relative camber 1.84%, and maximum camber at 76.5% chord length.

7. The supercritical airfoil aerodynamic shape structure according to claim 1, characterized in that, The spatial coordinates of the leading edge points of the five airfoil profiles are determined by the overall wing parameters, resulting in a final wing with a sweep angle of 25°, an aspect ratio of 8, and a root-to-tip ratio of 4.

8. The supercritical airfoil aerodynamic shape structure according to claim 7, characterized in that, The coordinates of the leading edge points of the five airfoil profiles are as follows: (0, 0, 0), (0.1083R, 0.2R, -0.0105R), (0.2436R, 0.45R, -0.0236R), (0.3789R, 0.7R, -0.0367R), (0.5413R, 1.0R, -0.0524R), where R is the wing's half-span.

9. The supercritical airfoil aerodynamic shape structure according to claim 7 or 8, characterized in that, The overall twist angle of the wing is 6.8°, and it has no dihedral.

10. The supercritical airfoil aerodynamic shape structure according to claim 1, characterized in that, The spanwise position of the wing root airfoil (1) is aligned with the plane of symmetry of the aircraft, with its upper wing surface protruding from the fuselage and its lower wing surface merging with the fuselage.

11. The supercritical airfoil aerodynamic shape structure according to claim 1, characterized in that, By matching the airfoil parameters according to any one of claims 1 to 6 with the geometric layout according to any one of claims 7 to 9, the large airflow separation area of ​​the wing at high angles of attack is controlled to appear first in the middle of the span.

12. The supercritical airfoil aerodynamic shape structure according to claim 1, characterized in that, The wing is used in high-wing transport aircraft.

Images