A waverider precursor / compression surface integrated configuration based on plane and conical shock waves and its design method

Abstract

The present invention provides a waverider forebody / compression surface integrated configuration based on plane and conical shock waves and a design method thereof, comprising the following steps: (1) obtaining a first-stage compression surface according to a plane shock wave flow field and a preset leading edge capture profile shape; (2) obtaining a second-stage compression surface according to a conical shock wave flow field; (3) designing a third-stage compression surface of isentropic compression; (4) the first-stage compression surface based on a plane shock wave flow field, the second-stage compression surface based on a conical shock wave flow field, and the third-stage compression surface of isentropic compression obtained in steps (1) to (3) are spatially connected in sequence to form a waverider forebody / compression surface integrated configuration. The forebody / compression surface integrated configuration designed by the method has compression surfaces with different three-stage compression modes, and realizes the displacement of the forebody boundary layer and the improvement of the inlet capture flow in the forebody / inlet integrated configuration of a hypersonic aircraft.

Description

Technical Field

[0001] The present invention relates to the field of hypersonic aircraft air inlets, and in particular to a waverider forebody / compression surface integrated configuration based on a plane and a conical shock wave and a design method thereof. Background Art

[0002] Hypersonic aircraft have the advantages of high flight speed and strong penetration capability, and have always been one of the focuses of equipment research and development in major countries around the world. As the "breathing duct" and "compressor" of the air-breathing hypersonic propulsion system, the air inlet is responsible for providing the engine with an appropriate amount of high-quality air, and its performance is directly linked to the engine thrust. In order to improve the performance of hypersonic aircraft, hypersonic aircraft often adopt an integrated aircraft / air inlet design. In the aircraft / air inlet integrated design configuration, the front body of the aircraft often serves as the compression surface of the air inlet, which is responsible for partially compressing the incoming flow and is closely related to the performance of the air inlet. Therefore, the integrated design of the forebody / compression surface is an important part of the integrated design of the aircraft / air inlet.

[0003] For hypersonic aircraft with an integrated aircraft / inlet configuration, the forebody may be relatively long because the forebody needs to decelerate and pressurize the incoming flow. The lower wall of the forebody of the US X-43A aircraft also serves as the compression surface of the inlet, using three-stage two-dimensional compression, with a total compression angle of 16°, and the total length of the forebody accounts for about 46% of the total length of the aircraft; the lower wall of the forebody of the US X-51A aircraft is also the compression surface of the inlet, with a compression angle of 6°, and the total length accounts for about 42% of the total length of the aircraft. It is not difficult to find that the forebody of such hypersonic aircraft is very long, and the long forebody is bound to develop a thicker boundary layer, which will increase the non-uniformity of the airflow at the inlet entrance, aggravate the shock wave / boundary layer interference phenomenon in the flow channel of the inlet, and cause flow separation in severe cases, resulting in the inlet not starting or even engine flameout. Therefore, in this type of configuration, trying to weaken the influence of the thick boundary layer of the forebody on the inlet is the key to improving the performance of the inlet and broadening its stable working boundary. In addition, the shape of the forebody is also closely related to the captured flow of the inlet. The forebody compression surfaces of the X-43A and X-51A are designed to be straight along the span direction, which will lead to obvious lateral overflow at the edge of the compression surface. The gas compressed by the compression surface cannot completely enter the interior of the inlet, which is not conducive to the improvement of the captured flow of the inlet and the thrust of the engine. Although directly widening the forebody can effectively weaken the influence of the expansion fans on both sides of the forebody and thus improve the inlet flow coefficient, widening the forebody will increase the overall weight of the aircraft, and the final overall benefit needs to be judged. In summary, a new forebody / compression surface integrated design method is needed to improve the displacement of the forebody boundary layer and the capture flow of the inlet in the forebody / inlet integrated configuration of hypersonic aircraft. Summary of the invention

[0004] Purpose of the invention: The purpose of the present invention is to provide a waverider precursor / compression surface integrated configuration based on plane and conical shock waves and a design method thereof, which can reduce the influence of the precursor boundary layer and increase the captured flow in the inlet.

[0005] Technical solution: A waverider precursor / compression surface integrated design method based on plane and conical shock waves, comprising the following steps:

[0006] S1 obtains the first-stage compression surface according to the plane shock wave flow field and the preset leading edge capture profile shape;

[0007] S2 obtains the second-stage compression surface based on the conical shock wave flow field;

[0008] S3 designs the third-stage compression surface of isentropic compression;

[0009] S4 The first-stage compression surface based on the plane shock wave flow field, the second-stage compression surface based on the conical shock wave flow field, and the third-stage compression surface of isentropic compression obtained in steps S1 to S3 are spatially connected in sequence to form an integrated waverider precursor / compression surface configuration.

[0010] Specifically, step S1 includes the following steps:

[0011] S11 determines the wedge-shaped reference flow field required for the first-stage compression surface design according to the incoming flow Mach number and attack angle under the design state;

[0012] S12 solves the oblique shock wave relationship according to the given incoming flow conditions and the first-order plane shock wave compression angle, obtains the airflow deflection angle of the plane shock wave and the remaining airflow parameters after the oblique shock wave, and determines the first-order plane shock wave wave surface; the remaining airflow parameters after the oblique shock wave include pressure, Mach number, density, speed, speed direction and temperature;

[0013] S13 determines the shape of the leading edge capture profile according to the overall design requirements of the aircraft, and projects the leading edge capture profile onto the first-stage plane shock wave surface to obtain the leading edge line of the forebody;

[0014] S14 selects a number of discrete points on the leading edge line of the forebody, performs streamline tracing according to the airflow parameters behind the oblique shock wave, and combines the obtained streamlines behind the wave to obtain the first-stage compression surface.

[0015] Specifically, step S2 includes the following steps:

[0016] S21 specifies the semi-cone angle of the cone generating the conical shock wave. According to the airflow parameters behind the oblique shock wave, the Taylor-Maccoll equation is used to calculate the semi-cone angle of the conical shock wave and other airflow parameters behind the conical shock wave.

[0017] S22 adjusts the position of the conical shock wave, and finds the intersection line between the conical shock wave and the first-stage compression surface as the starting point of the second-stage compression surface;

[0018] S23: Take a number of discrete points on the intersection line of step S22, rotate each discrete point to the same meridian plane, and use the airflow parameters after the conical shock wave to perform streamline tracking;

[0019] S24 reverses the streamlines obtained by the streamline tracing in step S23 back to the meridian plane where each discrete point is located, and combines them to obtain a second-level compression surface.

[0020] Specifically, step S3 includes the following steps:

[0021] S31, in each plane where the rear streamlines after the inversion in step S24 are located, starting from the leading edge point of the inlet lip, draw a left-going characteristic line, and take the intersection of the characteristic line and the rear streamlines after the conical shock wave as the starting point of the third-stage compression surface;

[0022] S32 specifies the isentropic compression angle of the third-level compression surface, and uses the axisymmetric characteristic line theory to obtain the isentropic compression profile, and the obtained isentropic compression profiles are combined to obtain the third-level compression surface.

[0023] Specifically, the leading edge capture profile is a concave curve, the first-stage compression surface is an inner concave profile; the second-stage compression surface is an outer convex profile; and the third-stage compression surface is an outer convex profile.

[0024] Specifically, in step S22, the axis of the conical shock wave is parallel to the direction of the airflow after compression by the first-stage plane shock wave, and moves to a position just incident on the inlet lip cover.

[0025] Specifically, the second-stage compression surface and the third-stage compression surface are completely covered by the first-stage compression surface.

[0026] To achieve the above-mentioned purpose of the invention, the present invention also provides an integrated waverider precursor / compression surface configuration based on plane and conical shock waves obtained by the above-mentioned design method, including: a first-stage compression surface, a second-stage compression surface extending backward from the first-stage compression surface, and a third-stage compression surface extending backward from the second-stage compression surface; the first-stage compression surface has an inner concave surface; the second-stage compression surface is an outer convex surface, and the radius of each point on the wall surface at the same flow direction position of the second-stage compression surface gradually increases from the symmetric surface to both sides; the third-stage compression surface is an outer convex surface.

[0027] Beneficial effect: Compared with the prior art, the present invention has the following significant effects: compared with the traditional forebody / compression surface integrated configuration, the forebody / compression surface integrated configuration designed by the technical solution of the present invention has three-stage compression surfaces with different compression modes. The first-stage compression surface adopts a waverider design and has an inwardly concave surface, which helps to improve the captured flow rate in the inlet. The second-stage compression surface uses the conical shock wave flow field as the reference flow field. The radius of each point on the wall surface at the same flow direction position gradually increases from the symmetrical surface to the two sides, and the conical reference flow field is beneficial to the displacement of the forebody boundary layer. The third-stage compression surface adopts isentropic compression, which improves the total pressure recovery of the inlet while meeting the forebody compression requirements, thereby realizing the displacement of the forebody boundary layer and the improvement of the captured flow rate in the inlet in the forebody / inlet integrated configuration of the hypersonic aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] Figure 1 It is a schematic diagram of the integrated configuration of the waverider precursor / air inlet according to Example 1 of the present invention.

[0029] Figure 2 It is a schematic diagram of the first-stage plane shock wave compression surface design method of Example 2 of the present invention.

[0030] Figure 3 It is a schematic diagram of the design method of the second-stage conical shock wave compression surface in Example 2 of the present invention.

[0031] Figure 4 It is a schematic diagram of the streamline tracking method of the second-stage conical shock wave compression surface in Example 2 of the present invention.

[0032] Figure 5 It is a schematic diagram of the third-stage isentropic compression surface design method of Example 2 of the present invention. DETAILED DESCRIPTION

[0033] The present invention is further explained below in conjunction with the accompanying drawings and specific implementation methods.

[0034] The primary purpose of the present invention is to provide a method for designing a waverider precursor / compression surface integration based on a plane and a conical shock wave, comprising the following steps:

[0035] S1: According to the incoming Mach number, angle of attack and total pressure recovery requirements of the inlet, the first-stage plane shock wave compression angle, the second-stage conical shock wave generation cone semi-cone angle and the third-stage isentropic compression angle are given;

[0036] S2: According to the incoming flow Mach number, angle of attack and the first-stage plane shock wave compression angle, the basic relationship of oblique shock wave is used to calculate the airflow parameters behind the wave;

[0037] S3: According to the overall design requirements of the aircraft, determine the shape of the leading edge capture profile, project the leading edge capture profile onto the first-stage plane shock wave surface, and obtain the leading edge line of the aircraft front body;

[0038] S4: taking a number of discrete points on the front body leading edge line, performing streamline tracing according to the wave-back airflow parameters in step S2, obtaining the wave-back streamlines intersecting with the front body leading edge line of the aircraft, and combining the wave-back streamlines into a first-level compression surface;

[0039] S5: according to the airflow parameters after the wave in step S2 and the cone half angle of the cone-shaped shock wave, the Taylor-Maccoll equation is used to solve the distribution of the parameters after the cone-shaped shock wave;

[0040] S6: Adjust the position of the conical shock wave so that it moves to the position where it just impacts the inlet lip cover, and find the intersection line between the conical shock wave and the first-stage compression surface as the starting point of the second-stage compression surface;

[0041] S7: taking a number of discrete points on the intersection line in step S6, rotating each discrete point to the same meridian plane, and performing streamline tracing using the conical shock wave post parameter distribution in step S5;

[0042] S8: Reverse the streamlines obtained by the streamline tracing in step S7 back to the meridian plane where each discrete point is located, and combine them to obtain a second-level compression surface;

[0043] S9: in each plane where the streamlines after the inversion in step S8 are located, a left-moving characteristic line is drawn from the leading edge of the inlet lip, and the intersection of the characteristic line and the streamlines after the conical shock wave is taken as the starting point of the isentropic compression, and the isentropic compression profile is obtained using the axisymmetric characteristic line theory;

[0044] S10: Combining the isentropic compression profiles in each plane obtained in step S9 to obtain a third-level isentropic compression surface.

[0045] The technical solution of the present invention is specifically described below through embodiments.

[0046] Embodiment 1:

[0047] Please refer to Figure 1 As shown, based on the above-mentioned waverider precursor / compression surface integrated design method, a waverider precursor / inlet integrated configuration based on plane shock wave and cone shock wave is designed, including a first-stage compression surface 1, a second-stage compression surface 2 extending backward from the first-stage compression surface 1, and a third-stage compression surface 3 extending backward from the second-stage compression surface 2. In this embodiment, according to conventional settings, a leakage gap 4 in the inlet channel and a leakage outlet 5 outside the inlet are provided behind the third-stage compression surface 3, and the second-stage compression surface 2 and the third-stage compression surface 3 are completely covered in the space where the first-stage compression surface 1 is located.

[0048] Specifically, the first-stage compression surface 1 adopts a waverider design and has an inwardly concave surface, which can improve the captured flow rate of the inlet duct; the second-stage compression surface 2 uses the conical shock wave flow field as the reference flow field, and the radius of each point on the wall surface at the same flow direction position gradually increases from the symmetrical surface to the two sides, and the conical reference flow field is beneficial to the displacement of the forebody boundary layer; the third-stage compression surface 3 adopts isentropic compression to maximize the recovery of the total pressure of the inlet duct while meeting the forebody compression requirements.

[0049] Embodiment 2:

[0050] With respect to the above-mentioned integrated design method of waverider precursor / compression surface based on plane and conical shock waves, an embodiment under specific parameter conditions is given:

[0051] 1. Please refer to Figure 2 As shown, under the reference conditions of design Mach number of 10.0, angle of attack of 0°, and first-stage plane shock wave compression angle δ of 4°, using the oblique shock wave relationship:

[0052]

[0053] Where M1 is the design Mach number, β is the shock wave angle, k = 1.4, and the shock wave angle β of the first-stage plane shock wave is 8.65°. At the same time, the flow field parameters after the remaining waves, such as pressure, Mach number, density, velocity and temperature, can be obtained from the relationship between the parameters before and after the oblique shock wave. For specific formulas, please refer to Zhang Kunyuan (Fluid Dynamics [M]. Beijing: Science Press, 2017).

[0054] 2. Determine the shape of the leading edge capture profile 6. In this embodiment, the leading edge capture profile 6 is a quadratic curve that opens upward and is tangent to the x-axis at the design origin, so as to ensure that the generated first-stage plane shock wave compression surface 10 has a concave characteristic.

[0055] 3. Project the leading edge capture profile line 6 onto the first-stage plane shock wave surface 7 to obtain the forebody leading edge line 8.

[0056] 4. Take several discrete points on the front edge line 8 of the forebody ( Figure 2 By taking the above discrete points as starting points and tracing the streamlines, a number of corresponding post-wave streamlines 9 can be obtained. By combining the above post-wave streamlines 9, the first-stage plane shock wave compression surface 10 can be obtained.

[0057] 5. Please refer to Figure 3 As shown, the semi-cone angle θ of the cone 11 generating the second-stage conical shock wave is taken as 6°, and the semi-cone angle of the corresponding conical shock wave 12 is obtained by solving the Taylor-Maccoll equation as 9.15°, and the airflow parameters after the residual wave are obtained at the same time. For the Taylor-Maccoll equation and its solution process, please refer to Tong Binggang (Gas Dynamics [M]. Beijing: Higher Education Press, 2012).

[0058] 6. Move the position of the conical shock wave in the reference flow field until the conical shock wave is just attached to the front end of the inlet lip cover under the design state, and at the same time, the axis 13 of the conical shock wave is parallel to the flow direction 14 of the airflow after the first-stage plane shock wave.

[0059] 7. After determining the position of the conical shock wave, find the intersection line 15 formed by the conical shock wave 12 and the first-stage plane shock wave compression surface 10 as the starting line of the second-stage conical shock wave compression surface 17.

[0060] 8. Take several discrete points on the intersection line 15 formed by the conical shock wave 12 and the first-stage plane shock wave compression surface 10 ( Figure 3 The streamlines are traced to obtain the streamlines 16 after the conical shock wave; please refer to Figure 4 As shown, taking a discrete point A′ on the above intersection line 15 as an example, the coordinates of A′ are (x a′ ,y a′ ,z a′ ), its coordinate after rotation to the y′Oz′ plane is (x a″ ,y a″ ,z a″ ), the angle of rotation of point A′ is θ A′ =arctan(x a′ / y a′ ). Starting from point A″, streamline tracing is performed in the cone-shaped reference flow field to obtain the streamline 18 after the wave passes through A″, and all points on the streamline are rotated by -θ A′ Returning to the original plane, we can obtain the streamline passing through point A′ behind the wave.

[0061] 9. The streamlines obtained by the above streamline tracing are reversed back to the meridian plane where each discrete point is located, and the second-stage conical shock wave compression surface 17 is obtained by combining them.

[0062] 10. Please refer to Figure 5 As shown, taking the flow field in the yOz plane as an example, AB in the figure is the intersection line of the first-stage plane shock wave compression surface 10 and the yOz plane, BC is the intersection line 19 of the second-stage cone shock wave compression surface and the yOz plane, and E is the intersection point of the leading edge line of the inlet lip and the yOz plane; the method for determining point C is as follows: starting from point E, with the help of the axisymmetric characteristic line theory, a left-running characteristic line of the fluid is made, and the intersection point of the left-running characteristic line and the second-stage compression surface is point C, which is also the starting point of the third-stage compression surface. CD is the isentropic compression section. In this embodiment, the isentropic compression angle is 6°. This section of the isentropic compression profile 20 is designed by the axisymmetric characteristic line theory and the isentropic compression wave and the second-stage cone shock wave generated by it converge at the inlet lip under the design state.

[0063] 11. Combine the isentropic compression profiles in each plane obtained in step 10 to obtain a third-level isentropic compression surface.

[0064] 12. Combine the first-stage plane shock wave compression surface 10, the second-stage conical shock wave compression surface 17, and the third-stage isentropic compression surface in sequence to complete the integrated design of the waverider forebody / inlet duct.

Claims

1. A method for designing an integrated waverider precursor / compression surface based on a plane and a conical shock wave, characterized in that: The following steps are involved: S1 obtains the first-stage compression surface according to the plane shock wave flow field and the preset leading edge capture line shape: S11 determines the wedge-shaped reference flow field required for the first-stage compression surface design according to the incoming flow Mach number and attack angle under the design state; S12 solves the oblique shock wave relationship according to the given incoming flow conditions and the first-order plane shock wave compression angle, obtains the airflow deflection angle of the plane shock wave and the remaining airflow parameters after the oblique shock wave, and determines the first-order plane shock wave wave surface; the remaining airflow parameters after the oblique shock wave include pressure, Mach number, density, speed, speed direction and temperature; S13 determines the shape of the leading edge capture profile according to the overall design requirements of the aircraft, and projects the leading edge capture profile onto the first-stage plane shock wave surface to obtain the leading edge line of the forebody; S14 selects a number of discrete points on the front edge line of the forebody, performs streamline tracing according to the airflow parameters behind the oblique shock wave, and combines the obtained streamlines behind the wave to obtain the first-stage compression surface; S2 obtains the second-stage compression surface based on the conical shock wave flow field: S21 Given the semi-cone angle of the cone generating cone of the conical shock wave, the semi-cone angle of the conical shock wave and the other airflow parameters behind the conical shock wave are calculated using the Taylor-Maccoll equation according to the airflow parameters behind the oblique shock wave. S22 adjusts the position of the conical shock wave, and finds the intersection line between the conical shock wave and the first-stage compression surface as the starting point of the second-stage compression surface; S23: Take a number of discrete points on the intersection line of step S22, rotate each discrete point to the same meridian plane, and use the airflow parameters after the conical shock wave to perform streamline tracking; S24: Invert the streamlines obtained by the streamline tracing in step S23 back to the meridian plane where each discrete point is located, and combine them to obtain a second-level compression surface; The third-stage compression surface of S3 design isentropic compression: S31, in each plane where the rear streamlines after the inversion in step S24 are located, starting from the leading edge point of the inlet lip, draw a left-going characteristic line, and take the intersection of the characteristic line and the rear streamlines after the conical shock wave as the starting point of the third-stage compression surface; S32 Given the isentropic compression angle of the third-stage compression surface, the isentropic compression profile is obtained by using the axisymmetric characteristic line theory, and the obtained isentropic compression profiles are combined to obtain the third-stage compression surface; S4 The first-stage compression surface based on the plane shock wave flow field, the second-stage compression surface based on the conical shock wave flow field, and the third-stage compression surface of isentropic compression obtained in steps S1 to S3 are spatially connected in sequence to form an integrated waverider precursor / compression surface configuration.

2. The waverider precursor / compression surface integrated design method according to claim 1, characterized in that: The leading edge capture profile is a concave curve, and the first-stage compression surface is an inwardly concave surface.

3. The waverider precursor / compression surface integrated design method according to claim 2, characterized in that: The second-stage compression surface is a convex surface.

4. The waverider precursor / compression surface integrated design method according to claim 2, characterized in that: The third-stage compression surface is a convex surface.

5. The waverider precursor / compression surface integrated design method according to claim 1, characterized in that: In step S22, the axis of the conical shock wave is parallel to the direction of the airflow after the first-stage plane shock wave is compressed, and moves to a position where it just impacts the inlet lip cover.

6. The waverider precursor / compression surface integrated design method according to claim 2, characterized in that: The second-stage compression surface and the third-stage compression surface are completely covered by the first-stage compression surface.

7. A waverider precursor / compression surface integrated configuration based on a plane and a conical shock wave obtained according to the design method of any one of claims 1 to 6, characterized in that: include: A first-stage compression surface (1), a second-stage compression surface (2) extending backward from the first-stage compression surface (1), and a third-stage compression surface (3) extending backward from the second-stage compression surface (2); the first-stage compression surface has an inner concave surface; the second-stage compression surface is an outer convex surface, and the radius of each point on the wall of the second-stage compression surface at the same flow direction position gradually increases from the symmetrical surface to both sides; the third-stage compression surface is an outer convex surface.

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