Large-incidence-angle aircraft capable of reducing asymmetry degree of flow field

Abstract

The invention belongs to the technical field of aircraft design, and particularly relates to a large-angle-of-attack aircraft capable of reducing the asymmetry degree of a flow field, two side wall surfaces which are symmetrical along the symmetry plane of the aircraft from the tip A of a nose to the leading edge B of a wing root are shaped, a control point C, a control point D and a control point E are sequentially arranged in the direction from the nose tip A to the wing root leading edge B. A first included angle between the side wall face from the control point C to the control point D and the plane of symmetry of the airplane is larger than a second included angle between the side wall face from the nose tip A to the control point C and the plane of symmetry of the airplane. And connecting from the control point D to the wing root leading edge B through a control point E by virtue of a spline curve sunken towards the plane of symmetry of the aircraft, wherein the control point E is a point closest to the plane of symmetry of the aircraft. According to the method, the propagation angle and form of the edging vortex can be effectively improved, so that the aircraft can avoid various non-commanded motions as much as possible during high-attack-angle flight, and various maneuvering actions can be completed with high quality.

Description

Technical Field

[0001] The present application belongs to the technical field of aircraft design, and in particular relates to a high-angle-of-attack aircraft capable of reducing flow field asymmetry. Background Art

[0002] Figure 1 Given the existing aircraft structure, the distance between the nose tip A and the wing root leading edge B is long, accounting for 45% to 48% of the fuselage length. The nose is relatively sharp, the angle between the front fuselage ridgeline and the symmetry plane is between 12° and 15°, and the wing leading edge is swept back widely, between 40° and 45°. At large angles of attack, large-scale flow separation will occur on the aircraft surface, such as Figure 2 As shown in the figure, after the leading edge vortex reaches the main wing surface, it merges with the leading edge vortex that rolls up from the leading edge of the main wing, and breaks up relatively early. These flow separation phenomena are often accompanied by highly unsteady and complex asymmetric vortex flows. The nonlinearity of the aerodynamic force and the longitudinal and lateral coupling become very serious, often accompanied by complex and even uncontrollable flight phenomena such as lateral deviation, side deviation, wing rock, and downstroke, which will have a significant impact on the aircraft's aerodynamic performance. For aircraft with higher maneuverability requirements, since the flight envelope usually needs to include a large angle of attack range, it is more important to pay attention to unsteady flow phenomena at high angles of attack, and to avoid or reduce the asymmetric flow that occurs at high angles of attack as much as possible. Summary of the Invention

[0003] In order to solve the above problems and change the strength, shape, spatial position, rupture position, etc. of the fringe vortex, the present application provides a high-angle-of-attack aircraft capable of reducing the asymmetry of the flow field, and reshapes the two sidewall surfaces symmetrical along the aircraft's symmetry plane from the nose tip A to the wing root leading edge B, wherein on each sidewall surface, control point C, control point D and control point E are sequentially arranged along the direction from the nose tip A to the wing root leading edge B, the first angle between the sidewall surface from the control point C to the control point D and the aircraft's symmetry plane is greater than the second angle between the sidewall surface from the nose tip A to the control point C and the aircraft's symmetry plane, and the control point D is connected to the wing root leading edge B via a spline curve concave to the aircraft's symmetry plane through the control point E, wherein the control point E is the point closest to the aircraft's symmetry plane.

[0004] Preferably, the first angle is 12° to 15°, and the second angle is 26° to 30°.

[0005] Preferably, the second angle is 28°.

[0006] Preferably, the distance LCA between the projection point of the control point C on the aircraft symmetry plane and the nose tip A is 15%-20% of the distance LBA between the projection point of the wing root leading edge B on the aircraft symmetry plane and the nose tip A; the distance LDA between the projection point of the control point D on the aircraft symmetry plane and the nose tip A is 55%-60% of the distance LBA; and the distance LEA between the projection point of the control point E on the aircraft symmetry plane and the nose tip A is 83%-85% of the distance LBA.

[0007] Preferably, the distance LCA is 18% of the distance LBA; the distance LDA is 58% of the distance LBA; and the distance LEA is 84% of the distance LBA.

[0008] Preferably, the width of the fuselage at the control point E is 80% of the original width of the fuselage at the axial position of the point.

[0009] The present application can effectively improve the propagation angle and shape of the side vortex, thereby changing the interference form between the side vortex and the leading edge vortex of the wing, delaying the vortex breakdown angle of attack, and reducing the asymmetry of the flow, so that the aircraft can avoid various non-command movements as much as possible during flight at a large angle of attack and complete various maneuvers with high quality. BRIEF DESCRIPTION OF THE DRAWINGS

[0010] Figure 1 This is a bird's-eye view of the existing aircraft.

[0011] Figure 2 It is a schematic diagram of the aerodynamic coupling phenomenon of the existing aircraft structure when flying at a high angle of attack.

[0012] Figure 3 This is a top view of an aircraft of a preferred embodiment of the present application for a high angle of attack aircraft capable of reducing flow field asymmetry.

[0013] Figure 4 It is a schematic diagram of the aerodynamic coupling phenomenon of the aircraft structure of the present application when flying at a high angle of attack.

[0014] Figure 5 This is a schematic diagram of the improvement effect of the aircraft provided in this application compared with existing aircraft in terms of lift coefficient.

[0015] Figure 6 This is a schematic diagram of the flight stability improvement effect in terms of rolling moment coefficient of the aircraft provided by this application compared with existing aircraft. DETAILED DESCRIPTION

[0016] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the implementation of this application will be described in more detail below in conjunction with the drawings in the implementation of this application. In the drawings, the same or similar numbers throughout represent the same or similar elements or elements with the same or similar functions. The described implementation is a part of the implementation of this application, not all of the implementations. The implementation described below with reference to the drawings is exemplary and is intended to be used to explain this application, and should not be understood as a limitation on this application. Based on the implementation in this application, all other implementations obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application. The implementation of this application is described in detail below in conjunction with the drawings.

[0017] The present application provides a high angle of attack aircraft capable of reducing the asymmetry of the flow field, such as Figure 3 As shown, two sidewall surfaces symmetrical along the aircraft symmetry plane from the nose tip A to the wing root leading edge B are modified, wherein on each sidewall surface, a control point C, a control point D and a control point E are sequentially set along the direction from the nose tip A to the wing root leading edge B, a first angle between the sidewall surface from the control point C to the control point D and the aircraft symmetry plane is greater than a second angle between the sidewall surface from the nose tip A to the control point C and the aircraft symmetry plane, and a spline curve concave from the control point D to the wing root leading edge B is connected to the wing root leading edge B via the control point E, wherein the control point E is the point closest to the aircraft symmetry plane.

[0018] This application presents a similar Figure 1 Aircraft structure design scheme in layout form, layout and Figure 1 When the vortex is similar and has the above characteristics, the effect of improving the high angle of attack characteristics is more obvious. In the improved scheme, the propagation angle and shape of the leading edge vortex of the front fuselage are significantly changed. After reaching the main wing surface, it does not merge with the leading edge vortex. Instead, the two vortex systems maintain independence and exist separately. Figure 4 As shown in the figure, the breaking point of the fringe vortex is greatly pushed back, the pressure distribution characteristics of the entire wing are significantly improved, the asymmetry of the flow field is effectively reduced at large angles of attack, the asymmetric flow phenomenon of the aircraft is improved, and the angle of attack with obvious asymmetry at large angles of attack is increased by more than 4°. Figure 5 The simulation analysis curve shows that the lift increases by more than 10% in the range of 20° to 30° angle of attack. Figure 6 The simulation analysis curve shows that the lateral instability angle is delayed by 2°, and the lateral instability after instability is reduced by more than 30%.

[0019] In some optional embodiments, the first angle is 12° to 15°, and the second angle is 26° to 30°.

[0020] In some optional embodiments, the second angle is 28°.

[0021] In some optional embodiments, the distance LCA between the projection point of the control point C on the aircraft symmetry plane and the nose tip A is 15%-20% of the distance LBA between the projection point of the wing root leading edge B on the aircraft symmetry plane and the nose tip A; the distance LDA between the projection point of the control point D on the aircraft symmetry plane and the nose tip A is 55%-60% of the distance LBA; the distance LEA between the projection point of the control point E on the aircraft symmetry plane and the nose tip A is 83%-85% of the distance LBA.

[0022] In some optional embodiments, the distance LCA is 18% of the distance LBA; the distance LDA is 58% of the distance LBA; and the distance LEA is 84% of the distance LBA.

[0023] In some optional embodiments, the fuselage width at the control point E is 80% of the original fuselage width at the axial position of the point.

[0024] The above parameter combination is a proven optimization combination. It precisely controls the vortex's generation location, development path, and interaction distance with the vortices at the leading edge of the main wing. This is the key geometrical guarantee for achieving the aforementioned excellent aerodynamic performance (delayed vortex breakup, reduced asymmetry, increased lift, and enhanced stability). These values represent the balance point found within the design parameter space that maximizes these advantages.

[0025] The above description is merely a specific embodiment of the present application, but the scope of protection of the present application is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in this application should be included in the scope of protection of the present application. Therefore, the scope of protection of the present application should be based on the scope of protection of the claims.

Claims

1. A high angle of attack aircraft capable of reducing flow field asymmetry, characterized in that: The two side wall surfaces symmetrical along the aircraft symmetry plane from the nose tip A to the wing root leading edge B are reshaped, wherein on each side wall surface, a control point C, a control point D and a control point E are sequentially set along the direction from the nose tip A to the wing root leading edge B, a first angle between the side wall surface from the control point C to the control point D and the aircraft symmetry plane is greater than a second angle between the side wall surface from the nose tip A to the control point C and the aircraft symmetry plane, and a spline curve concave from the control point D to the wing root leading edge B via the control point E is connected, wherein the control point E is the point closest to the aircraft symmetry plane.

2. The high angle of attack aircraft capable of reducing flow field asymmetry according to claim 1, characterized in that: The first angle is 12° to 15°, and the second angle is 26° to 30°.

3. The high angle of attack aircraft capable of reducing flow field asymmetry according to claim 2, characterized in that: The second angle is 28°.

4. The high angle of attack aircraft capable of reducing flow field asymmetry according to claim 1, characterized in that: The distance LCA between the projection point of the control point C on the aircraft symmetry plane and the nose tip A is 15%-20% of the distance LBA between the projection point of the wing root leading edge B on the aircraft symmetry plane and the nose tip A; the distance LDA between the projection point of the control point D on the aircraft symmetry plane and the nose tip A is 55%-60% of the distance LBA; the distance LEA between the projection point of the control point E on the aircraft symmetry plane and the nose tip A is 83%-85% of the distance LBA.

5. The high angle of attack aircraft capable of reducing flow field asymmetry according to claim 4, characterized in that: The distance LCA is 18% of the distance LBA; the distance LDA is 58% of the distance LBA; and the distance LEA is 84% of the distance LBA.

6. The high angle of attack aircraft capable of reducing flow field asymmetry according to claim 1, characterized in that: The fuselage width at the control point E is 80% of the original fuselage width at the axial position of the point.

Images