Aircraft low-wave drag aerodynamic configuration and design method

Abstract

The application belongs to the field of aircraft design and relates to a low-wave resistance aerodynamic configuration and design method of an aircraft, which comprises a body and a wing, and the body and the wing are both rotary bodies, the body comprises a front cone and a rear cone; incoming airflow is compressed for the first time at a first compression cavity on the front cone, the incoming airflow is compressed for the second time after reaching a second compression cavity, a front shock wave surface is formed, and then the incoming airflow forms a body expansion wave surface when reaching a decompression cavity at the rear cone, and the decompression is realized under the action of the shock wave expansion wave; in this way, the pressure acting on the front cone of the body only undergoes one compression, and the pressure acting on the rear cone of the body undergoes two compressions and one decompression; therefore, the pressure acting on the rear cone of the body is close to the pressure acting on the front cone of the body, that is, the pressure difference between the front cone and the rear cone of the body is relatively small, the pressure difference between the front cone and the rear cone of the body can be significantly reduced, and thus the shock wave resistance of the aerodynamic configuration is greatly reduced.

Description

TECHNICAL FIELD

[0001] The present application belongs to the field of aircraft design, and particularly relates to an aircraft low-wave drag aerodynamic configuration and a design method. BACKGROUND

[0002] New concept supersonic / hypersonic aircraft is one of the research hotspots in recent years. However, with the increase of Mach number, the shock wave drag of the aircraft is getting higher and higher. How to reduce the shock wave drag is one of the key problems restricting the development of supersonic / hypersonic aircraft at present. Low shock wave drag aircraft has faster flight speed and better aerodynamic performance. Therefore, reducing the shock wave drag has important academic significance and engineering value.

[0003] Among the existing technologies for reducing shock wave drag, there is a reverse jet flow technology. The action mechanism of this technology is to inject high-pressure gas from the head cone in the opposite direction, forcing the bow shock wave in front of the head cone to move away from the object surface, thereby reducing the pressure after the shock wave, and achieving the purpose of reducing the shock wave drag. Another technology for reducing shock wave drag displays an aerodynamic rod device. The action mechanism of this device is to convert the bow shock wave in front of the head cone of the aircraft into an oblique shock wave, thereby achieving the purpose of reducing the shock intensity. Although the reverse jet flow technology and the aerodynamic rod device have relatively good drag reduction effect, they can only be used to reduce the shock wave drag of the head cone of the aircraft.

[0004] Another technology for reducing shock wave drag is a design method of supersonic annular wing. The annular wing combines the characteristics of Busman wing and Litcher wing, has two inner and outer ring wings, and the wing type is a triangular wing type. The thickness of the inner ring wing type is relatively large. At zero angle of attack, the annular wing has lift. And with the increase of the angle of attack, the lift increases, and the lift-drag ratio becomes larger. However, the shock wave drag of this aerodynamic configuration is still relatively large.

[0005] Therefore, how to more efficiently reduce the shock wave drag of the aerodynamic configuration is a problem to be solved. SUMMARY

[0006] The purpose of the present application is to provide an aircraft low-wave drag aerodynamic configuration and a design method to solve the problem of large shock wave drag of the aircraft in the prior art.

[0007] The technical scheme of the present application is: a low-wave resistance aerodynamic configuration of an aircraft, comprising a fuselage and a wing, the fuselage and the wing are both rotary bodies, the fuselage comprises a front cone and a rear cone, the bottom surfaces of the front cone and the rear cone coincide, supports are arranged between the wing and the front cone and the rear cone, a first compression surface is arranged on the front cone, a second compression surface is arranged on the rear cone, the wing is a cylindrical structure, a first compression cavity is formed between the wing and the front cone, a second compression cavity is formed between the two ends of the wing and the bottom surface edges of the front cone and the rear cone, and a decompression cavity is formed between the wing and the rear cone.

[0008] Preferably, the front edge of the wing is located on the shock wave surface of the front cone of the fuselage.

[0009] Preferably, the rear edge of the wing is located on the expansion wave post-Mach number line of the fuselage.

[0010] Preferably, the outer edges of the bottom surfaces of the front cone and the rear cone are located on the shock wave surface of the front edge of the wing.

[0011] Preferably, the front cone, the rear cone and the wing are coaxially arranged.

[0012] As a specific embodiment, a design method of a low-wave resistance aerodynamic configuration of an aircraft, adopting the low-wave resistance aerodynamic configuration of the aircraft described above, characterized in that, comprising:

[0013] determining the cone angles of the front cone and the rear cone of the fuselage, and the radius of the bottom surface of the cone;

[0014] determining the shock wave angle of the shock wave of the front cone of the fuselage;

[0015] determining the position of the front edge of the wing;

[0016] determining the shock wave angle of the shock wave of the front edge of the wing;

[0017] judging whether the bottom surfaces of the front cone and the rear cone are on the shock wave of the front edge of the wing, if yes, performing the next step, if not, re-determining the position of the front edge of the wing and the shock wave angle of the shock wave of the front edge of the wing, until the bottom surfaces of the front cone and the rear cone are on the shock wave of the front edge of the wing;

[0018] determining the expansion wave post-Mach number line of the fuselage;

[0019] determining the supports between the wing and the fuselage;

[0020] carrying out numerical simulation or wind tunnel test to verify the aerodynamic characteristics of the aerodynamic configuration.

[0021] Preferably, the shock wave angle of the fore-cone shock wave of the fuselage is determined by solving the Taylor-Macoll equation or a numerical simulation method; the shock wave angle of the leading edge shock wave of the wing is determined by a numerical simulation method; and the Mach number line of the expansion wave of the fuselage is determined by a numerical simulation method.

[0022] The aircraft low-wave-resistance aerodynamic configuration and design method provided by the application comprises a fuselage and a wing, both of which are spin bodies, and the fuselage comprises a fore-cone and an after-cone; when the aircraft is flying, the incoming airflow is compressed for the first time at a first compression cavity on the fore-cone, compressed for the second time after reaching a second compression cavity to form a leading edge shock wave, and then the airflow is expanded after reaching a pressure-reducing cavity on the after-cone to form a fuselage expansion wave, and the pressure is reduced under the action of the shock wave and the expansion wave; in this way, the pressure acting on the fore-cone of the fuselage is compressed only once, while the pressure acting on the after-cone of the fuselage is compressed twice and reduced once; therefore, the pressure acting on the after-cone of the fuselage is close to the pressure acting on the fore-cone of the fuselage, that is, the pressure difference between the fore-cone and the after-cone of the fuselage is small, which can significantly reduce the pressure difference between the fore-cone and the after-cone of the fuselage, thereby greatly reducing the shock wave resistance of the aerodynamic configuration. BRIEF DESCRIPTION OF DRAWINGS

[0023] In order to more clearly illustrate the technical solutions provided by the application, the following will briefly introduce the drawings. Obviously, the drawings described below are only some embodiments of the application.

[0024] Figure 1 is an axial view of the overall structure of the application;

[0025] Figure 2 is a generatrix and design principle diagram of the fuselage and the wing of the application;

[0026] Figure 3 is an overall structure contour diagram of the application;

[0027] Figure 4 is a pressure contour diagram of the fuselage (without the wing) of the embodiment of the application at an attack angle of 0°;

[0028] Figure 5 is a pressure contour diagram of the embodiment of the application at an attack angle of 0°;

[0029] Figure 6 is a pressure contour diagram of the embodiment of the application at an attack angle of 4°;

[0030] Figure 7 is a pressure contour diagram of the embodiment of the application at an attack angle of 8°;

[0031] Figure 8 is a pressure contour diagram of the embodiment of the application at an attack angle of 12°;

[0032] Figure 9 is the change of the shock wave drag coefficient of the present application with the angle of attack.

[0033] 1. wing; 2. fuselage; 3. support; 4. first compression cavity; 5. front cone shock wave surface; 6. front shock wave surface; 7. fuselage expansion wave surface; 8. second compression cavity; 9. pressure reduction cavity. DETAILED DESCRIPTION

[0034] In order to make the purpose, technical scheme and advantages of the present application clearer, the technical scheme of the present application will be described in more detail below in combination with the drawings in the embodiments of the present application.

[0035] A low wave drag aerodynamic configuration of an aircraft, as shown in Figures 1-3 includes a fuselage 2 and a wing 1, both of which are spinners, the fuselage 2 includes a front cone and a rear cone, the bottom surfaces of the front cone and the rear cone coincide, supports 3 are arranged between the wing 1 and the front cone and the rear cone, a first compression surface is arranged on the front cone, a second compression surface is arranged on the rear cone, the wing 1 is a cylindrical structure, a first compression cavity 4 is formed between the wing 1 and the front cone, a second compression cavity 8 is formed between the two ends of the wing 1 and the edge of the bottom surface of the front cone and the rear cone, and a pressure reduction cavity 9 is formed between the wing 1 and the rear cone.

[0036] When the aircraft is flying, the Mach number of the incoming airflow is M1, the incoming airflow is compressed for the first time at the first compression cavity 4 on the front cone, forming a front cone shock wave surface 5, the Mach number of the air after the front cone shock wave surface 5 is M2, M2 < M1, the magnitude of the decrease in the Mach number depends on the radius of the bottom surface of the front cone and the cone angle, and the required M2 is determined according to actual requirements, the pressure acting on the front cone is equal to the post-wave pressure of the front cone shock wave of the fuselage 2.

[0037] The incoming airflow is compressed for the second time after reaching the second compression cavity 8, forming a front shock wave surface 6, the Mach number of the air after the front shock wave surface 6 is M3, M3 < M2.

[0038] The incoming airflow forms a fuselage expansion wave surface 7 when reaching the rear cone of the pressure reduction cavity 9, and the pressure is reduced under the action of the shock wave expansion wave, the magnitude of the decrease in the pressure depends on the radius of the bottom surface of the rear cone and the cone angle, at this time the pressure acting on the rear cone of the fuselage 2 is equal to the post-wave pressure of the expansion wave of the fuselage 2.

[0039] The overall force is as shown in Figure 2As shown in the figure, A is the front vertex of the wing 1 generatrix, B is the rear vertex of the wing 1 generatrix, C is the front vertex of the fuselage 2 generatrix, D is the rear vertex of the fuselage 2 generatrix, E is the highest point of the fuselage 2 generatrix, AB is the wing 1 generatrix, CE is the fuselage 2 front conical generatrix, and ED is the fuselage 2 rear conical generatrix. The wing 1 generatrix AB rotates 360° around the symmetry axis 4 to form the wing 1. The fuselage 2 front conical generatrix CE and the fuselage 2 rear conical generatrix ED rotate 360° around the symmetry axis 4 to form the fuselage 2.

[0040] Thus, the pressure acting on the fuselage 2 front conical surface is compressed only once, while the pressure acting on the fuselage 2 rear conical surface is compressed twice and decompressed once. Therefore, the pressure acting on the fuselage 2 rear conical surface is close to the pressure acting on the fuselage 2 front conical surface, i.e., the pressure difference between the fuselage 2 front conical surface and the fuselage 2 rear conical surface is small, which can significantly reduce the pressure difference between the fuselage 2 front conical surface and the fuselage 2 rear conical surface, thereby greatly reducing the shock wave resistance of the aerodynamic configuration.

[0041] Preferably, the front edge of the wing 1 is located on the front conical shock wave surface 5 of the fuselage 2 to ensure structural stability.

[0042] The change curve of the shock wave resistance coefficient at different attack angles is shown in the figure. Figure 9 As shown in the figure, 8 is the shock wave resistance coefficient curve of the fuselage 2 (without the wing 1), and 9 is the shock wave resistance coefficient curve of the embodiment of the present application. It can be clearly seen that the shock wave resistance coefficient of the present application is greatly reduced at any attack angle compared with the existing design.

[0043] Preferably, the rear edge of the wing 1 is located on the expansion wave rear Mach number line of the fuselage 2. Since the pressure difference between the front conical surface and the rear conical surface is approximately the same, the pressure borne by the front and rear of the wing 1 is approximately the same, which ensures stability and makes the wing 1 easier to design.

[0044] Preferably, the outer edge of the bottom surface of the front conical surface and the rear conical surface is located on the front edge shock wave surface 6 of the wing 1, so that there is no large difference in the pressure of the entire fuselage 2.

[0045] Preferably, the front conical surface, the rear conical surface, and the wing 1 are coaxially arranged to improve the stress stability.

[0046] As a specific embodiment, a method for designing a low wave resistance aerodynamic configuration of an aircraft includes the following steps:

[0047] Step S100, determining the conical angle of the fuselage 2 front conical surface and the fuselage 2 rear conical surface, and the radius of the bottom surface of the conical surface;

[0048] Step S200, determining the shock wave angle of the fuselage 2 front conical shock wave;

[0049] Preferably, the shock angle of the fore-cone shock wave of the fuselage 2 is determined by solving Taylor-Macoll equation or numerical simulation method.

[0050] In step S300, the position of the leading edge of the wing 1 is determined to ensure that the leading edge of the wing 1 extends to the fore-cone shock wave of the fuselage 2.

[0051] In step S400, the shock angle of the leading edge shock wave of the wing 1 is determined.

[0052] Preferably, the shock angle of the leading edge shock wave of the wing 1 is determined by numerical simulation method.

[0053] In step S500, it is determined whether the bottom surface of the fore-cone and the aft-cone is on the leading edge shock wave of the wing 1. If yes, the next step is performed, and if not, the position of the leading edge of the wing 1 and the shock angle of the leading edge shock wave of the wing 1 are determined again, i.e. steps S300-S400 are repeated until the bottom surface of the fore-cone and the aft-cone is on the leading edge shock wave of the wing 1.

[0054] In step S600, the expansion wave back-Mach line of the fuselage 2 is determined.

[0055] Preferably, the expansion wave back-Mach line of the fuselage 2 is determined by numerical simulation method.

[0056] In step S700, the support 3 between the wing 1 and the fuselage 2 is determined to ensure the structural strength.

[0057] In step S800, numerical simulation or wind tunnel test is carried out to verify the aerodynamic characteristics of the aerodynamic configuration. If the verification is consistent with the theoretical design, the design is completed.

[0058] Preferably, the shock angle of the fore-cone shock wave of the fuselage 2 is determined by solving Taylor-Macoll equation or numerical simulation method, the shock angle of the leading edge shock wave of the wing 1 is determined by numerical simulation method, and the expansion wave back-Mach line of the fuselage 2 is determined by numerical simulation method.

[0059] The following is described by a specific verification example: the cone angle of the fore-cone and the aft-cone is 14°, the radius of the bottom surface of the cone is 124.7 mm, the length of the fore-cone generatrix CE of the fuselage 2 and the length of the aft-cone generatrix ED of the fuselage 2 are both 515.3 mm, the length of the wing 1 generatrix AB is 393.1 mm, and the distance from point E to the wing 1 generatrix AB is 74.3 mm. Numerical simulation or wind tunnel test is carried out to collect test data, and the following results are obtained:

[0060] Figure 4The pressure contour of the body 2 (without the wing 1) at an attack angle of 0° is given. In the case of the wing 1, the pressure acting on the rear cone of the body 2 has only undergone one compression and one decompression. As can be seen from the figure, the pressure acting on the rear cone of the body 2 is much smaller than the pressure acting on the front cone of the body 2, which leads to a high shock wave resistance of the body 2 (without the wing 1).

[0061] Figure 5 The pressure contour of the body 2 at an attack angle of 0° is given. After the wing 1 is installed, the pressure acting on the rear cone of the body 2 has undergone two compressions and one decompression. As can be seen from the figure, the pressure acting on the rear cone of the body 2 is greatly improved. Therefore, the shock wave resistance of the embodiment of the present application is very low.

[0062] Figures 6-8 The pressure contour of the body 2 at an attack angle of 0° is given. After the wing 1 is installed, the pressure acting on the rear cone of the body 2 has undergone two compressions and one decompression. As can be seen from the figure, the pressure acting on the rear cone of the body 2 is greatly improved. Therefore, the shock wave resistance of the embodiment of the present application is very low.

[0063] Figure 9 The figure of the change of the shock wave resistance coefficient with the attack angle is given. As can be seen from the figure, at the same attack angle, the shock wave resistance coefficient of the embodiment of the present application is much smaller than the shock wave resistance coefficient of the body 2 (without the wing 1).

[0064] The above merely provides a specific implementation of the present application, but the protection scope of the present application is not limited thereto, any change or replacement within the technical range disclosed by the present application can be easily thought by those skilled in the art, which should be covered in the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection scope of the claims.

Claims

1. A low-wave drag aerodynamic configuration for an aircraft, comprising a fuselage (2) and a wing (1), characterized in that: The fuselage (2) and the wing (1) are both spin-formed bodies. The fuselage (2) includes a front cone and a rear cone. The bottom surfaces of the front cone and the rear cone overlap. The wing (1) is provided with supports (3) between the front cone and the rear cone. The front cone has a first compression surface, and the rear cone has a second compression surface. The wing (1) is a cylindrical structure. The wing (1) and the front cone form a first compression cavity (4). The two ends of the wing (1) and the bottom edges of the front cone and the rear cone form a second compression cavity (8). The wing (1) and the rear cone form a decompression cavity (9).

2. The low-wave drag aerodynamic configuration of the aircraft as described in claim 1, characterized in that: The leading edge of the wing (1) is located on the front conical shock surface (5) of the fuselage (2).

3. The low-wave drag aerodynamic configuration of the aircraft as described in claim 1, characterized in that: The trailing edge of the wing (1) is located on the Mach number line behind the expansion wave of the fuselage (2).

4. The low-wave drag aerodynamic configuration of the aircraft as described in claim 1, characterized in that: The outer edges of the bottom surfaces of the front and rear cones are located on the leading edge shock surface (6) of the wing (1).

5. The low-wave drag aerodynamic configuration of an aircraft as described in claim 1, characterized in that: The front cone, rear cone, and wing (1) are all coaxially arranged.

6. A method for designing a low-wave-drag aerodynamic configuration for an aircraft, employing the low-wave-drag aerodynamic configuration as described in any one of claims 1-5, characterized in that, include: Determine the cone angles of the front and rear cones of the body (2), as well as the radius of the cone base; Determine the shock angle of the conical shock wave in front of the body (2); Determine the position of the leading edge of wing (1); Determine the shock angle of the leading edge of the wing (1); Determine whether the bottom surfaces of the front and rear cones are on the leading edge shock wave of the wing (1). If yes, proceed to the next step. If no, re-determine the position of the leading edge of the wing (1) and the shock wave angle of the leading edge shock wave of the wing (1) until the bottom surfaces of the front and rear cones are on the leading edge shock wave of the wing (1). Determine the Mach number line after the expansion wave of the body (2); Determine the support (3) between the wing (1) and the fuselage (2); Conduct numerical simulations or wind tunnel tests to verify the aerodynamic characteristics of the aerodynamic configuration.

7. The aircraft low-wave drag aerodynamic configuration design method as described in claim 6, characterized in that: The shock angle of the front conical shock wave of the fuselage (2) is determined by solving the Taylor-Macoll equation or by numerical simulation; the shock angle of the leading edge shock wave of the wing (1) is determined by numerical simulation; and the Mach number line behind the expansion wave of the fuselage (2) is determined by numerical simulation.

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