A high angle of attack aircraft capable of reducing flow field asymmetry

Abstract

The application belongs to the technical field of aircraft design, and particularly relates to a high-angle-of-attack aircraft capable of reducing flow field asymmetry, wherein two side wall surfaces along the aircraft symmetry plane from a nose tip A to a wing root front edge B are modified, wherein on each side wall surface, a control point C, a control point D and a control point E are sequentially arranged along the direction from the nose tip A to the wing root front edge B, a first included 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 included angle between the side wall surface from the nose tip A to the control point C and the aircraft symmetry plane, the control point E is connected to the wing root front edge B through a spline curve which is recessed to the aircraft symmetry plane from the control point D, and the control point E is a point closest to the aircraft symmetry plane. The application can effectively improve the propagation angle and form of the side stripe vortex, so that the aircraft can avoid various non-instructional movements as much as possible and complete various maneuvering actions with high quality.

Description

Technical Field

[0001] This application belongs to the field of aircraft design technology, and specifically relates to a high angle-of-attack aircraft that can reduce flow field asymmetry. Background Technology

[0002] Figure 1 The existing aircraft structure is presented, with a relatively long distance between the nose tip A and the wing root leading edge B, accounting for 45%–48% of the fuselage length. The nose is relatively sharp, and the angle between the forward fuselage edge and the plane of symmetry is between 12° and 15°. The wing leading edge has a large sweepback, between 40° and 45°. At high angles of attack, large-scale flow separation phenomena will occur on the aircraft surface, such as… Figure 2 As shown, the leading-edge vortex along the front fuselage merges with the leading-edge vortex generated from the leading edge of the main wing after reaching the main wing surface, and breaks up relatively early. These flow separation phenomena are often accompanied by high unsteadiness and complex asymmetric vortex flows. The nonlinearity of aerodynamics and longitudinal and lateral coupling become very severe, often accompanied by complex and even uncontrollable flight phenomena such as lateral deviation, side deviation, wing rocking, and downthrow, which can significantly affect the aerodynamic performance of the aircraft. For aircraft with high maneuverability requirements, since the flight envelope usually needs to cover a large angle of attack range, it is even more necessary to pay attention to unsteady flow phenomena at large angles of attack and avoid or reduce asymmetric flows at large angles of attack as much as possible. Summary of the Invention

[0003] To address the aforementioned issues and modify the intensity, shape, spatial location, and breakup location of the leading edge vortex, this application provides a high angle-of-attack aircraft capable of reducing flow field asymmetry. The two sidewalls symmetrical along the aircraft's plane of symmetry, from the nose tip A to the wing root leading edge B, are reshaped. On each sidewall, control points C, D, and E are sequentially positioned along the direction from the nose tip A to the wing root leading edge B. The first angle between the sidewall from control point C to control point D and the aircraft's plane of symmetry is greater than the second angle between the sidewall from the nose tip A to control point C and the aircraft's plane of symmetry. A spline curve concave towards the aircraft's plane of symmetry connects control point D to the wing root leading edge B via control point E, where control point E is the point closest to the aircraft's plane of symmetry.

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

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

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

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

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

[0009] This application can effectively improve the propagation angle and morphology of leading-edge vortices, thereby changing the interference pattern between leading-edge vortices and wing leading-edge vortices, delaying vortex breakup angle of attack, and reducing flow asymmetry. This allows the aircraft to avoid various uncommanded movements as much as possible during high angle-of-attack flight and complete various maneuvers with high quality. Attached Figure Description

[0010] Figure 1 It is an existing top view of the aircraft.

[0011] Figure 2 This is a schematic diagram of the aerodynamic coupling phenomenon of an existing aircraft structure during high angle of attack flight.

[0012] Figure 3 This is a top view of an aircraft of a preferred embodiment of a high angle-of-attack aircraft that can reduce flow field asymmetry, according to this application.

[0013] Figure 4 This is a schematic diagram of the aerodynamic coupling phenomenon of the aircraft structure in this application during high angle of attack flight.

[0014] Figure 5 This is a schematic diagram illustrating the improvement in lift coefficient between the aircraft provided in this application and existing aircraft.

[0015] Figure 6 This is a schematic diagram illustrating the improvement in flight stability of the aircraft provided in this application compared to existing aircraft in terms of roll moment coefficient. Detailed Implementation

[0016] 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.

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

[0018] This application provides a similar Figure 1 Aircraft structural design schemes with different layout forms, layout and Figure 1 When similar characteristics are present, the improvement in high angle-of-attack characteristics is more significant. In the improved design, the propagation angle and morphology of the leading-edge vortex are significantly altered. Upon reaching the main wing surface, it does not merge with the leading-edge vortex; instead, the two vortex systems remain independent and exist separately. Figure 4 As shown, this significantly delays the rupture point of the leading-edge vortex, substantially improves the pressure distribution characteristics across the entire airfoil, effectively reduces the asymmetry of the flow field at high angles of attack, and mitigates asymmetric flow phenomena in the aircraft, increasing the angle of attack at which significant asymmetry is observed by more than 4°. Figure 5 Simulation analysis curves show that lift increases by more than 10% in the large angle of attack range of 20° to 30°. Figure 6 Simulation analysis curves show that delaying the lateral instability angle by 2° reduces the lateral instability by more than 30%.

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

[0020] In some alternative implementations, the second included angle is 28°.

[0021] In some alternative implementations, the distance LCA between the projection point of control point C on the aircraft's plane of symmetry 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's plane of symmetry and the nose tip A; the distance LDA between the projection point of control point D on the aircraft's plane of symmetry and the nose tip A is 55%-60% of the distance LBA; and the distance LEA between the projection point of control point E on the aircraft's plane of symmetry and the nose tip A is 83%-85% of the distance LBA.

[0022] In some alternative implementations, the distance to LCA is 18% of the distance to LBA; the distance to LDA is 58% of the distance to LBA; and the distance to LEA is 84% ​​of the distance to LBA.

[0023] In some alternative implementations, the fuselage width at control point E is 80% of the original fuselage width at that point's axial position.

[0024] The combination of the above parameters is an optimized and validated set. They precisely control the vortex generation location, development path, and interaction distance with the leading-edge vortex of the main wing, providing crucial geometric assurance for achieving the aforementioned superior aerodynamic performance (delaying vortex breakup, reducing asymmetry, increasing lift, and enhancing stability). These values ​​represent a balance point found in the design parameter space that maximizes these advantages.

[0025] 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 high angle-of-attack aircraft capable of reducing flow field asymmetry, characterized in that, The two sidewalls symmetrical along the aircraft's plane of symmetry, from the nose tip A to the wing root leading edge B, are reshaped. On each sidewall, control points C, D, and E are sequentially set along the direction from the nose tip A to the wing root leading edge B. The first angle between the sidewall from control point C to control point D and the aircraft's plane of symmetry is greater than the second angle between the sidewall from the nose tip A to control point C and the aircraft's plane of symmetry. From control point D, a spline curve concave towards the aircraft's plane of symmetry is connected to the wing root leading edge B via control point E, where control point E is the point closest to the aircraft's plane of symmetry.

2. The high angle-of-attack aircraft as described in claim 1, characterized in that, The first included angle is 12° to 15°, and the second included angle is 26° to 30°.

3. The high angle-of-attack aircraft capable of reducing flow field asymmetry as described in claim 2, characterized in that, The second included angle is 28°.

4. The high angle-of-attack aircraft as described in claim 1, characterized in that, The distance LCA between the projection of control point C onto the aircraft's plane of symmetry and the nose tip A is 15%-20% of the distance LBA between the projection of the wing root leading edge B onto the aircraft's plane of symmetry and the nose tip A; the distance LDA between the projection of control point D onto the aircraft's plane of symmetry and the nose tip A is 55%-60% of the distance LBA; and the distance LEA between the projection of control point E onto the aircraft's plane of symmetry and the nose tip A is 83%-85% of the distance LBA.

5. The high angle-of-attack aircraft as described in claim 4, characterized in that, The distance to LCA is 18% of the distance to LBA; the distance to LDA is 58% of the distance to LBA; and the distance to LEA is 84% ​​of the distance to LBA.

6. The high angle-of-attack aircraft as described in claim 1, characterized in that, The fuselage width at control point E is 80% of the original fuselage width at that point's axial position.

Images