An integrated design method of aircraft weakening the interference of internal and external flow

Abstract

The application relates to the technical field of aircraft design, and discloses an integrated design method of an aircraft capable of weakening internal and external flow interference, which comprises the following steps: S1: generating an internal rotation type axisymmetric reference flow field and internal rotation type shock waves; S2: generating an external compression axisymmetric reference flow field and external compression shock waves; designing an aerodynamic wall surface; S3: designing a projection line of an air inlet lip contour line and a projection line of a waverider leading edge line; S4: generating a waverider surface and a wing upper surface in the external compression axisymmetric reference flow field generated in step S2; S5: generating an internal rotation type air inlet in the internal rotation type axisymmetric reference flow field generated in step S1; S6: designing an air inlet fairing with a set thickness based on the generated internal rotation type air inlet; and S7: designing a fuselage upper surface and a bottom surface. The application improves the problems of the prior art, such as weakening of the shock wave capturing characteristics of an air inlet and weakening of the waverider characteristics of a fuselage during integrated design.

Description

Technical Field

[0001] This invention relates to the field of aircraft design technology, specifically an integrated aircraft design method for reducing internal and external interference. Background Technology

[0002] Currently, the high compression efficiency and high total pressure recovery coefficient of the three-dimensional internal rotating intake, along with the high lift-to-drag ratio of the waverider (aircraft body), make them the preferred targets for integrated design. However, the internal contraction flow field on which the internal rotating intake relies and the external compression axisymmetric reference flow field on which the conventional waverider relies are fundamentally different. This makes it difficult to design them simultaneously. The internal rotating intake captures the internal contraction shock wave, while the waverider captures the external compression shock wave. Due to the difference in the shape characteristics of the two shock waves in space, the conventional method of directly splicing the internal rotating intake and the waverider cannot achieve a good transition between the two shock waves. The mutual influence causes changes in the shape and position of the two shock waves, resulting in a weakening of the shock wave capture characteristics of the intake and a weakening of the waveriding characteristics of the aircraft body. Consequently, the performance of both cannot reach the excellent performance of separate designs. Summary of the Invention

[0003] To overcome the shortcomings of existing technologies, this invention provides an integrated design method for aircraft that reduces internal and external flow interference. It designs an aerodynamic wall to separate the internal and external compression axisymmetric reference flow fields, thereby improving the problems of weakened shock wave capture characteristics of the air intake and weakened wave-riding characteristics of the airframe in the integrated design of existing technologies.

[0004] The technical solution adopted by the present invention to solve the above problems is:

[0005] An integrated design method for aircraft that reduces internal and external flow interference includes the following steps:

[0006] S1: Generate an inward-rotating axisymmetric reference flow field, and generate an inward-rotating shock wave within the inward-rotating axisymmetric reference flow field;

[0007] S2: Generate an externally compressed axisymmetric reference flow field and generate an externally compressed shock wave within the externally compressed axisymmetric reference flow field; design aerodynamic walls based on existing internally rotating axisymmetric reference flow fields and externally compressed axisymmetric reference flow fields;

[0008] S3: Projection lines of the intake lip profile and the leading edge of the wave rider;

[0009] S4: Generate the waverider surface and the upper surface of the wing in the externally compressed axisymmetric reference flow field generated in step S2;

[0010] S5: Generate an internal rotating air intake in the internal rotating axisymmetric reference flow field generated in step S1;

[0011] S6: The thickness of the inlet fairing is determined by the outward design of the generated internal rotating air intake.

[0012] S7: Design the upper and lower fuselage surfaces based on the generated internal rotating air intake, waverider surface, upper wing surface, and air intake fairing;

[0013] The execution order of steps S1 and S2 is one of the following: execute step S1 first and then step S2, execute step S2 first and then step S1, or execute steps S1 and S2 simultaneously; the execution order of steps S4 and S5 is one of the following: execute step S4 first and then step S5, execute step S5 first and then step S4, or execute steps S4 and S5 simultaneously.

[0014] As a preferred technical solution, in step S1, the inward-rotating shock wave is symmetrical about the symmetry plane; where the symmetry plane refers to the assumed cross section when the inward-rotating shock wave is generated.

[0015] As a preferred technical solution, in step S2, the external compression shock wave is symmetrical about the symmetry plane; wherein, the symmetry plane refers to the assumed cross section when the external compression shock wave is generated, and this cross section is the same as the symmetry plane of the internal rotation shock wave.

[0016] As a preferred technical solution, in step S3, there is a bottom projection surface at the tail outlet of the internal rotating inlet that is perpendicular to the rotation axis of the internal rotating axisymmetric reference flow field, and there is a bottom projection surface at the tail of the waverider that is perpendicular to the rotation axis of the external compression axisymmetric reference flow field; the two bottom projection surfaces are the same plane, and the projection profile A′B′C′D′ of the lip profile ABCD of the internal rotating inlet is designed on the bottom projection surface, and the projection profile A′E′ of the leading edge line AE of the waverider is designed on the bottom projection surface.

[0017] As a preferred technical solution, in step S3, the projection line of the intake duct lip profile and the leading edge line of the wave rider are designed simultaneously on the same plane.

[0018] As a preferred technical solution, step S4 includes the following steps:

[0019] S41: The projection profile A′E′ of the waverider leading edge line AE is uniformly discretized into several points. The free streamline method is applied to draw free streamlines from the discrete points and intersect with the external compression shock wave to generate the leading edge line points of the waverider. The leading edge line points of the waverider are connected to obtain the wing leading edge line A′E′. The free streamline method refers to the method of drawing straight lines parallel to the direction of the incoming flow from the discrete points.

[0020] S42: In the externally compressed axisymmetric reference flow field generated in step S2, the streamlines of the waverider surface are generated by tracing the leading edge points of the waverider to the bottom projection plane. All the wall streamlines of the waverider surface are lofted to form the waverider surface.

[0021] S43: Using the free streamline method, free streamlines are drawn from the leading edge points of the waverider and intersect with the bottom projection surface to generate the upper surface of the wing.

[0022] As a preferred technical solution, step S5 includes the following steps:

[0023] S51: The projection profile A′B′C′D′ of the lip profile ABCD of the inward rotating inlet is uniformly discretized into several points. The free streamline method is applied to draw free streamlines from the discretized points and intersect with the inward rotating shock wave to generate the leading edge line points of the inward rotating inlet.

[0024] S52: In the internal rotating axisymmetric reference flow field generated in step S1, streamlines are traced from the leading edge point of the internal rotating intake to the bottom projection plane to generate the internal rotating intake wall streamlines. All internal rotating intake wall streamlines are lofted to form the internal rotating intake.

[0025] As a preferred technical solution, step S7 includes the following steps:

[0026] S71: The obtained internal rotating air intake, waverider surface, upper wing surface, and air intake fairing are simultaneously symmetrical about the fuselage.

[0027] S72: The upper surface of the fuselage is designed based on the generated internal rotating air intake, waverider surface, upper wing surface, and air intake fairing. The lower surface of the fuselage is then filled to finally generate an integrated aircraft configuration.

[0028] As a preferred technical solution, in step S72, a ridge-shaped protrusion structure is designed on the upper surface of the fuselage.

[0029] Compared with the prior art, the present invention has the following advantages:

[0030] This invention designs an aerodynamic wall to separate the inner and outer compression axisymmetric reference flow fields, which to a certain extent improves the flow field mismatch problem between the waverider body and the inlet. It largely preserves the excellent performance of both the waverider and the internal rotating inlet. Unlike the existing principle of directly separating the inner and outer compression axisymmetric reference flow fields by using the inlet wall, in this invention, the flow fields of the inner rotating axisymmetric reference flow field and the outer compression axisymmetric reference flow field are discontinuous at the plane of symmetry. The airflow in each flow field does not flow from one side of the plane of symmetry to the other. This design can separate the inner and outer compression axisymmetric reference flow fields in principle. The inner rotating axisymmetric reference flow field is designed on one side, and the outer compression axisymmetric reference flow field is designed on the other side. Thus, in principle, the inner and outer flow are designed without interference, thereby basically preserving the characteristics of the waverider and the internal rotating inlet, and achieving the expected effect of weak interference between the waverider body and the inlet. Attached Figure Description

[0031] Figure 1 The airflow characteristics in the internally rotating axisymmetric reference flow field and the externally compressed axisymmetric reference flow field;

[0032] Figure 2 A three-dimensional schematic diagram of the internal rotating air intake in the internal rotating axisymmetric reference flow field and the wave rider in the external compression axisymmetric reference flow field;

[0033] Figure 3 A schematic diagram of the gap between the internal rotating shock wave and the external compression shock wave at the split surface;

[0034] Figure 4 Design principle diagram for integrated design method;

[0035] Figure 5 A side view of the aircraft;

[0036] Figure 6 A top view of the aircraft;

[0037] Figure 7 A pressure lift ratio cloud map of a series of cross sections of the aircraft;

[0038] Figure 8 This is a pressure lift ratio cloud map of the aircraft at a cross-section of x = 1.3m. Detailed Implementation

[0039] The present invention will be further described in detail below with reference to the embodiments and accompanying drawings, but the embodiments of the present invention are not limited thereto.

[0040] Example 1

[0041] like Figures 1 to 8 As shown, the present invention provides an integrated design method for aircraft that reduces interference between internal and external flow. An aerodynamic wall is designed to overcome the flow field shock wave matching problem between the waverider body and the internal rotating air intake in the prior art, so that the internal and external shock waves in the flow field can be better connected and transitioned, reducing the mutual interference between the flow field of the waverider body and the air intake, thereby preserving the excellent characteristics of the internal rotating air intake and the waverider body respectively.

[0042] To achieve the above objectives, this invention proposes an integrated design method for the internal and external flow of a hypersonic vehicle, which is a "non-interference" integrated design method for the internal rotating air intake and the waverider airframe, comprising the following steps:

[0043] S1: Generate an inward-rotating axisymmetric reference flow field based on design conditions, and generate an inward-rotating shock wave within the inward-rotating axisymmetric reference flow field. The inward-rotating shock wave is symmetrical about the symmetry plane (segmentation plane) (the symmetry plane refers to the assumed cross section when generating the inward-rotating shock wave). The characteristic of the airflow in the inward-rotating axisymmetric reference flow field is that the airflow is in a contracting state and will not flow from one side of the symmetry plane to the other side.

[0044] S2: Generate an externally compressed axisymmetric reference flow field based on design conditions, and generate an externally compressed shock wave within the externally compressed axisymmetric reference flow field. The externally compressed shock wave is symmetrical about the symmetry plane (segmentation plane) (the symmetry plane refers to the assumed cross section when generating the externally compressed shock wave, which is the same as the symmetry plane of the internally rotating shock wave); the characteristic of the airflow in the externally compressed axisymmetric reference flow field is that the airflow is radial and does not flow from one side of the symmetry plane to the other.

[0045] S3: Assume a bottom projection plane at the tail outlet of the inward rotating inlet, perpendicular to the rotation axis of the inward rotating axisymmetric reference flow field. Also assume a bottom projection plane at the tail of the waverider, perpendicular to the rotation axis of the external compression axisymmetric reference flow field. Both projection planes are on the same plane. On the bottom projection plane, design the projection profile A′B′C′D′ of the inward rotating inlet lip profile ABCD, and the projection profile A′E′ of the waverider leading edge line AE. MP1 is the projection line of the inward rotating shock wave on the bottom projection plane, and NP2 is the projection plane of the external compression shock wave on the bottom projection plane. Due to the characteristics of the airflow in the internal and external compression axisymmetric reference flow fields and the external compression axisymmetric reference flow field, the airflow will not cross the dividing surface from one side to the other. Therefore, the dividing surface is equivalent to an aerodynamic wall. This aerodynamic wall is used to... The method of separating the outflow is different from the existing method of directly dividing the flow field using the wall. Because there is a small gap between the inner rotating shock wave and the outer compression shock wave at the dividing surface, the final shock wave shape is not the shape of the inner rotating shock wave and the outer compression shock wave directly connecting, but is slightly different. The projection line of the final shock wave on the bottom projection surface is MQN. It is worth noting that the small gap between the inner rotating shock wave and the outer compression shock wave at the dividing surface can be reduced by adjusting the flow field parameters to adjust the shock wave shape. Therefore, the gap has a relatively small impact on the shock wave connection. Figure 4 In the diagram, r represents the r direction on the bottom projection plane of the internally rotating axisymmetric reference flow field or the r direction on the bottom projection plane of the externally compressed axisymmetric reference flow field; z1 represents the z1 direction on the bottom projection plane of the internally rotating axisymmetric reference flow field; and z2 represents the z2 direction on the bottom projection plane of the externally compressed axisymmetric reference flow field. The r direction is orthogonal to the z1 direction, the r direction is orthogonal to the z2 direction, and the z1 direction is parallel to the z2 direction.

[0046] S4: In the externally compressed axisymmetric reference flow field generated in step S2, a waverider surface is generated by the free streamline method and the streamline tracing method, and the upper surface of the wing is generated by the free streamline method.

[0047] S41: The projection profile A′E′ of the waverider leading edge line AE is uniformly discretized into several points. The free streamline method is applied to draw free streamlines from the discrete points and intersect with the external compression shock wave to generate the leading edge line points of the waverider. The leading edge line points of the waverider are connected to obtain the leading edge line A′E′ of the wing. The free streamline method refers to the method of drawing straight lines parallel to the direction of the incoming flow from the discrete points. Under normal circumstances, the direction of the incoming flow is parallel to the axis of rotation of the axisymmetric reference flow field.

[0048] S42: In the externally compressed axisymmetric reference flow field generated in step S2, the streamlines of the waverider surface are generated by tracing the leading edge points of the waverider to the bottom projection plane. All the wall streamlines of the waverider surface are lofted to form the waverider surface.

[0049] S43: Using the free streamline method, free streamlines are drawn from the leading edge points of the waverider and intersect with the bottom projection plane to generate the upper surface of the wing;

[0050] S5: The internal rotating intake is generated by the projection profile A′B′C′D′ of the lip profile ABCD of the internal rotating intake through the free streamline method and the streamline tracing method in the internal rotating axisymmetric reference flow field generated in step S1.

[0051] S51: The projection profile A′B′C′D′ of the lip profile ABCD of the inward rotating inlet is uniformly discretized into several points. The free streamline method is applied to draw free streamlines from the discretized points and intersect with the inward rotating shock wave to generate the leading edge line points of the inward rotating inlet.

[0052] S52: In the internal rotating axisymmetric reference flow field generated in step S1, streamlines are traced from the leading edge point of the internal rotating inlet to the bottom projection surface to generate the internal rotating inlet wall streamlines. All internal rotating inlet wall streamlines are lofted (a streamline is a curve generated in the flow field, and lofting is an operation in modeling to generate a surface from a line. Streamline lofting means combining all the streamlines into a curved surface) to form the internal rotating inlet.

[0053] S6: Based on the generated internal rotating air intake, an air intake fairing of a certain thickness is designed to be formed outward;

[0054] S7: Design the upper and lower fuselage surfaces based on the generated internal rotating air intake, waverider surface, upper wing surface, and air intake fairing;

[0055] S71: Simultaneously symmetrically arrange the internal rotating air intake, waverider surface, upper wing surface, and air intake fairing obtained through the above steps according to the fuselage symmetry plane.

[0056] S72: The upper surface of the fuselage is designed based on the generated internal rotating air intake, waverider surface, upper wing surface, and air intake fairing. To increase the volume of the aircraft, the upper surface of the fuselage adopts a ridge-shaped protrusion design. The design of the upper surface of the fuselage will not be explained here. Then, the bottom surface of the fuselage (bottom surface of the tail of the fuselage) is filled to finally generate an integrated aircraft configuration.

[0057] The integrated design method for the internal and external flow of hypersonic aircraft provided by this invention improves the flow field mismatch problem between the waverider body and the air intake to a certain extent, and largely preserves the excellent performance of both the waverider and the internal rotating air intake. Unlike the existing principle of directly separating the internal and external flow fields by designing aerodynamic walls, in this invention, the flow fields of the internal rotating axisymmetric reference flow field and the external compression axisymmetric reference flow field are discontinuous at the dividing surface. The airflow in each flow field does not flow from one side of the symmetry plane to the other. This design can separate the internal and external flow fields in principle, thereby achieving a design without interference between the internal and external flow in principle. This basically preserves the characteristics of the waverider and the internal rotating air intake, and achieves the expected effect of no interference between the waverider body and the air intake.

[0058] Other notes:

[0059] The internal rotation axisymmetric reference flow field can be obtained from scientific research papers, scientific journals, academic conferences or scientific research competitions.

[0060] In this embodiment, the specific method for solving the internal rotation axisymmetric reference flow field based on the theory of rotational characteristic lines can be found in the patent application number 201910325410.6, "An integrated design method for the internal rotation waverider forebody air intake of a supersonic aircraft".

[0061] The inward-rotating shock wave in this embodiment was generated under atmospheric parameters of an incoming Mach number of 5 and an altitude of 25 km.

[0062] The externally compressed axisymmetric reference flow field can be obtained from research papers, research journals, academic conferences, or research competitions. For the specific method of generating the externally compressed axisymmetric reference flow field based on the theory of swirling characteristic lines in this embodiment, please refer to the following reference: Ding Feng. Research on the Theory and Method of Integrated Internal and External Flow "Full Wave Riding" Aerodynamic Design for Air-breathing Hypersonic Vehicles [D]. Changsha: National University of Defense Technology, 2016, pp. 67-68.

[0063] In this embodiment, the external compression shock wave is generated under atmospheric parameters of incoming Mach number 5 and altitude 25km.

[0064] In this embodiment, the projection line A′E′ of the waverider leading edge AE is designed as a straight line segment, but it is not limited to a straight line segment. Other line shapes can be designed as needed, such as multiple curves.

[0065] In the method of this invention, the rotation axis of the externally compressed axisymmetric reference flow field is parallel to the rotation axis of the internally rotating axisymmetric reference flow field. The bottom projection plane is perpendicular to the rotation axis of the internally rotating axisymmetric reference flow field, and at the same time, the bottom projection plane is perpendicular to the rotation axis of the externally compressed axisymmetric reference flow field.

[0066] It is worth noting that, unlike existing methods that directly separate the inner and outer compression axisymmetric reference flow fields using the inlet duct wall, this invention designs an aerodynamic wall. The flow fields of the inner rotating axisymmetric reference flow field and the outer compression axisymmetric reference flow field are, in principle, discontinuous at the dividing surface. The airflow in each flow field does not flow from one side of the symmetry plane to the other. This design can, in principle, separate the inner and outer compression axisymmetric reference flow fields, thus achieving a design where the inner and outer flow fields do not interfere with each other. In this embodiment, both the inner rotating axisymmetric reference flow field and the outer compression axisymmetric reference flow field use axisymmetric reference flow field design. However, the principle of discontinuity between the inner and outer flow fields described in this invention is not limited to this type of flow field; other flow fields can also be used for design. Furthermore, this embodiment uses a matching of the inner rotating axisymmetric reference flow field and the outer compression axisymmetric reference flow field; similar flow fields, such as both inner and outer being outer compression axisymmetric reference flow fields, can also be used for matching design.

[0067] Effect description:

[0068] Numerical simulations can yield a series of pressure rise ratio cloud maps. In this implementation case, the calculation conditions are an incoming Mach number of 5 and atmospheric parameters at an altitude of 25 km. The SST turbulence model is used for viscous calculations.

[0069] Figure 7 , Figure 8 In this context, P represents the local total pressure. ∞ P / P represents the total incoming pressure. ∞ This indicates the pressure rise ratio. (From...) Figure 7 It can be seen that the inward rotating shock wave fits the edge of the air intake and maintains its "concave" characteristic, while the outward compressive shock wave fits the edge of the waverider wing and maintains its "convex" characteristic. At the junction of the inward rotating air intake and the waverider wing, the shape transition between the inward rotating shock wave and the outward compressive shock wave is good and consistent with the original design intention, indicating that the mutual interference of the flow field between the inward rotating air intake and the waverider wing is small. Figure 8 This displays a pressure lift ratio contour map of the aircraft at a cross-section of x = 1.3m (x represents the coordinate along the axial direction of the aircraft, which is consistent with the direction of the rotation axis of the internally rotating axisymmetric reference flow field or the direction of the rotation axis of the externally compressed axisymmetric reference flow field). Figure 8 It can be seen that the shape and position of the actual shock wave are consistent with the design value, thus achieving the design objective.

[0070] It is worth noting that the numerical simulation results of this implementation case were obtained under viscous conditions. If the results were obtained under non-viscous conditions, the results would be better and would better meet the design requirements.

[0071] As described above, the present invention can be implemented well.

[0072] All features disclosed in all embodiments of this specification, or steps in all methods or processes implied in the disclosure, may be combined and / or extended or replaced in any way, except for mutually exclusive features and / or steps.

[0073] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention in any way. Based on the technical essence of the present invention, any simple modifications, equivalent substitutions, and improvements made to the above embodiments within the spirit and principles of the present invention shall still fall within the protection scope of the present invention.

Claims

1. An integrated design method for aircraft that reduces internal and external flow interference, characterized in that, Includes the following steps: S1: Generate an inward-rotating axisymmetric reference flow field, and generate an inward-rotating shock wave within the inward-rotating axisymmetric reference flow field; S2: Generate an externally compressed axisymmetric reference flow field, and generate an externally compressed shock wave within the externally compressed axisymmetric reference flow field; Design aerodynamic walls based on existing internal rotation axisymmetric reference flow fields and external compression axisymmetric reference flow fields; S3: Projection lines of the intake lip profile and waverider leading edge; S4: Generate the waverider surface and the upper surface of the wing in the externally compressed axisymmetric reference flow field generated in step S2; S5: Generate an internal rotating air intake in the internal rotating axisymmetric reference flow field generated in step S1; S6: The thickness of the inlet fairing is determined by the outward design of the generated internal rotating air intake. S7: Design the upper and lower fuselage surfaces based on the generated internal rotating air intake, waverider surface, upper wing surface, and air intake fairing; The execution order of steps S1 and S2 is one of the following: execute step S1 first and then step S2, execute step S2 first and then step S1, or execute steps S1 and S2 simultaneously; the execution order of steps S4 and S5 is one of the following: execute step S4 first and then step S5, execute step S5 first and then step S4, or execute steps S4 and S5 simultaneously. In step S2, the external compression shock wave is symmetrical about the symmetry plane; where the symmetry plane refers to the assumed cross section when the external compression shock wave is generated, and this cross section is the same as the symmetry plane of the internal rotation shock wave. In step S3, a bottom projection plane perpendicular to the rotation axis of the internal rotating inlet's tail outlet is provided, and a bottom projection plane perpendicular to the rotation axis of the external compression axisymmetric reference flow field is provided at the tail of the waverider; the two bottom projection planes are on the same plane, and the projection profile of the internal rotating inlet lip profile ABCD is designed on the bottom projection plane. The projection profile of the waverider leading edge AE is designed on the bottom projection plane. ; In step S3, the projection lines of the intake duct lip profile and the leading edge line of the waverider are designed simultaneously on the same plane.

2. The integrated design method for reducing internal and external flow interference of an aircraft according to claim 1, characterized in that, In step S1, the inward-rotating shock wave is symmetrical about the symmetry plane; where the symmetry plane refers to the assumed cross section when the inward-rotating shock wave is generated.

3. The integrated design method for reducing internal and external flow interference of an aircraft according to claim 1, characterized in that, Step S4 includes the following steps: S41: Projection profile of the waverider leading edge AE The airfoil is uniformly discretized into several points. Using the free streamline method, free streamlines are drawn from the discrete points and intersected with the external compression shock wave to generate waverider leading edge points. Connecting these waverider leading edge points yields the wing leading edge line. The free streamline method refers to the method of drawing straight lines parallel to the direction of incoming flow from discrete points. S42: In the externally compressed axisymmetric reference flow field generated in step S2, the streamlines of the waverider surface are generated by tracing the leading edge points of the waverider to the bottom projection plane. All the wall streamlines of the waverider surface are lofted to form the waverider surface. S43: Using the free streamline method, free streamlines are drawn from the leading edge points of the waverider and intersect with the bottom projection surface to generate the upper surface of the wing.

4. The integrated design method for reducing internal and external flow interference of an aircraft according to claim 3, characterized in that, Step S5 includes the following steps: S51: Projected profile of the inward-rotating intake lip profile ABCD The points are uniformly discretized into several points. The free streamline method is applied to draw free streamlines from the discrete points and intersect with the inward rotating shock wave to generate the leading edge line points of the inward rotating air intake. S52: In the internal rotating axisymmetric reference flow field generated in step S1, streamlines are traced from the leading edge point of the internal rotating intake to the bottom projection plane to generate the internal rotating intake wall streamlines. All internal rotating intake wall streamlines are lofted to form the internal rotating intake.

5. An integrated aircraft design method for reducing internal and external flow interference according to any one of claims 1 to 4, characterized in that, Step S7 includes the following steps: S71: The obtained internal rotating air intake, waverider surface, upper wing surface, and air intake fairing are simultaneously symmetrical about the fuselage. S72: The upper surface of the fuselage is designed based on the generated internal rotating air intake, waverider surface, upper wing surface, and air intake fairing. The lower surface of the fuselage is then filled to finally generate an integrated aircraft configuration.

6. The integrated design method for reducing internal and external flow interference of an aircraft according to claim 5, characterized in that, In step S72, a ridge-shaped protrusion structure is designed on the upper surface of the fuselage.

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