Design method of double-design-point inlet with wide speed range based on bending shock wave

Abstract

The design method of a double-design-point wide-speed domain retractable inlet based on curved shock waves involves the field of wide-speed domain inlet for aircraft. 1) Determine the high / low Mach number design points of the hypersonic retractable inlet according to the design requirements of the hypersonic aircraft; 2) Construct a double-design-point axisymmetric basic flow field; 3) Place the fan-shaped inlet in the high Mach number basic flow field for streamline tracing to obtain the basic compression profile of the inlet; 4) Design the high / low Mach number operating conditions actuation profile and actuation scheme. The design of a hypersonic wide-speed domain retractable inlet is achieved according to the high / low Mach number flow requirements at the same time. The hypersonic wide-speed domain retractable inlet designed based on this design method can ensure that the inlet can start at a low Mach number at the beginning of the design, and has the ability to capture low Mach number flow that meets the requirements. Develop a wide-domain design method for a wide-speed domain retractable inlet to improve the designability and design efficiency of a hypersonic wide-speed domain retractable inlet.

Description

Technical Field

[0001] The invention relates to the field of wide-speed range air inlets for aircraft, and in particular to a design method for a wide-speed range contraction air inlet with double design points based on curved shock waves. Background Art

[0002] The wide-speed range inlet is an important component of an air-breathing hypersonic vehicle. Its main function is to decelerate and pressurize the hypersonic flow to meet the working needs of the engine. The flow requirement is the primary performance indicator that the hypersonic inlet needs to meet. The flight conditions of a hypersonic vehicle vary within a certain range, so the hypersonic inlet needs to meet the flow requirements within a certain speed range, rather than the flow requirements within a single condition, which will bring great difficulties to the design of the wide-speed range inlet.

[0003] The design difficulty of a wide-speed range inlet lies in the following two contradictions: first, the increase in the frontal area of ​​the wide-speed range inlet is beneficial for the inlet to meet the flow requirements in a wide speed range, but it leads to an increase in headwind resistance; second, the stronger the deceleration and pressurization effect of the wide-speed range inlet on the airflow, the better the inlet performance, that is, the higher the airflow compression efficiency, but it increases the risk of the wide-speed range inlet not being able to work normally due to flow congestion under low Mach number conditions, that is, the inlet does not start.

[0004] In order to solve the above contradictions, domestic and foreign scholars have conducted a lot of theoretical research and design method exploration, and developed three common configurations of wide-speed range inlets: axisymmetric inlet, dual inlet, and inner-contracting inlet. The working principle is to compress the airflow by generating shock waves through aerodynamic surfaces (wedge surface, conical surface, inner-contracting surface). However, due to the different compression forms, the aerodynamic characteristics have obvious differences [Wang Jiangfeng, Wang Xudong, Li Jiawei, Yang Tianpeng, Li Longfei, Cheng Keming. A review of aerodynamic design of waveriding layout of hypersonic cruise vehicle [J]. Journal of Aerodynamics, 2018, 36(05): 705-728.].

[0005] The typical design process of a wide-speed range inlet is as follows: first, determine the design point of the basic flow field and the target airflow compression effect (usually expressed as the throat Mach number and the throat total pressure recovery) according to the design requirements of the wide-speed range inlet; then, construct an incident shock wave system to compress the airflow to the target performance; then, construct the inlet compression surface by tracing the streamlines of the basic flow field given the inlet shape; finally, construct the geometric adjustment scheme and actuation scheme according to the wide-range working requirements [You Yancheng, Huang Guoping, Guo Junliang, et al. Internal waverider hypersonic inlet based on arbitrary shock wave shape and design method: China, 101392685[P]. 2009-03-25.].

[0006] The aerodynamic performance of wide-speed range inlets is largely determined by the basic flow field. The axisymmetric inlet is designed based on the outer cone axisymmetric basic flow field, the binary inlet is designed based on the two-dimensional plane basic flow field, and the inward-contracting inlet is designed based on the three-dimensional inward-contracting basic flow field. It is worth mentioning that the inward-contracting inlet has been the focus of research by scholars at home and abroad due to its advantages such as high airflow compression efficiency and small frontal area, but its complex curved surface structure makes it difficult to design its geometric adjustment ability and wide-range starting ability.

[0007] The basic flow field in traditional design methods is designed based on a single design point. The wide-speed range inlet can only better match the aerodynamic characteristics of the basic flow field at a single design point. The aerodynamic requirements under other working conditions cannot be used as design input for the design. Such a design method cannot take into account multi-condition performance at the aerodynamic design level. Summary of the invention

[0008] The purpose of the present invention is to address the above-mentioned problems existing in the prior art and to provide a design method for a double-design-point, wide-speed range contraction inlet based on a curved shock wave, which can take double-design conditions as input.

[0009] The present invention comprises the following steps:

[0010] 1) Given the high / low Mach number incoming flow conditions of the basic flow field with dual design points, the center body height of the basic flow field with high / low Mach numbers, the low Mach number incident shock wave shape, and the low Mach number capture flow coefficient;

[0011] 2) Calculate the starting height of the lip streamline in the low Mach number basic flow field according to the low Mach number capture flow coefficient, and determine the coordinates of the starting point of the lip streamline in combination with the low Mach number shock wave shape;

[0012] 3) applying the characteristic line method to calculate the known flow field behind the wave according to the low Mach number shock wave shape, obtaining part of the known wall generatrix, and then combining the starting coordinates of the lip streamline in step 2) to perform streamline tracking, obtaining part of the known lip streamline in the low Mach number flow field;

[0013] 4) applying the characteristic line method to the wall generatrix obtained in step 3) and the high Mach number incoming flow condition in step 1) to calculate the partial incident shock wave shape and the post-wave flow field under the high Mach number condition;

[0014] 5) Based on the wall generatrix in step 3), the wall generatrix is ​​extended in the flow direction to form an extended section, and the high / low Mach number flow field corresponding to the extended section is obtained by applying the characteristic line method, and then the low Mach number lip streamline and the high Mach number incident shock wave shape are obtained, and the two intersect at the horizontal line corresponding to the height of the high Mach number center body, and the intersection point is the lip point; the extended section of the wall generatrix and the wall generatrix known in step 3) together constitute the low Mach number wall generatrix of the basic flow field;

[0015] 6) Based on the low Mach number wall busbar in step 5), the high Mach number flow field is obtained in combination with the high Mach number incoming flow condition, and the reflected shock wave parameters are given. The wavefront flow field and the wavefront flow field corresponding to the reflected shock wave are obtained by applying the characteristic line method, wherein the upper boundary of the wavefront flow field is connected to the low Mach number wall busbar to form the high Mach number busbar of the basic flow field, and the lower boundary of the wavefront flow field is the high Mach number lower wall busbar of the basic flow field; at this point, both the high Mach number basic flow field and the low Mach number basic flow field are constructed;

[0016] 7) placing the fan-shaped inlet shape into the high Mach number basic flow field of step 6) for streamline tracing, wherein the two arc edges of the fan-shaped shape are respectively fitted with the outer contour and the center body of the high Mach number basic flow field, and at this time, the two side edges of the fan-shaped shape coincide with the radial lines of the basic flow field; the profile obtained by streamline tracing is the basic compression profile of the contraction inlet duct in the hypersonic dual design point wide speed domain after scaling and viscosity correction;

[0017] 8) With the arc corresponding to the downstream end point of the low Mach number generatrix as the boundary, the upstream side of the basic compression profile in step 7) is the low Mach number compression profile, and the downstream side is the compression side actuation profile; the intersection line of the two compression side actuation profiles and the side wall surface is modified into a straight line as the rotation axis, and the actuation profile is divided into two small sectors along the symmetry plane, which rotate around the rotation axis respectively to achieve geometric adjustment;

[0018] 9) On the basis of the basic compression profile in step 7), the area surrounded by the leading edge profile of the walls on both sides of the basic compression profile and the low Mach number incident shock wave and the lip streamline is used as the leading edge actuation profile; the leading edge actuation profile is actuated by rotation, the rotation axis is located at the leading edge point of the compression side, and the rotation axis direction is the normal direction of the radial plane where the side wall surface is located;

[0019] 10) Under high Mach number conditions, the leading edge actuating profile in step 9) deflects toward the compression side, and at this time, the shape of the front edge of the side wall of the compression profile is the front edge of the side wall of the basic compression profile; under low Mach number conditions, the leading edge actuating profile deflects toward the lower wall surface, and at this time, the front edge of the side wall of the compression profile is the front edge of the leading edge actuating profile.

[0020] The present invention has the following advantages: the present invention realizes the design of a hypersonic wide-speed domain retractable inlet according to the needs of high / low Mach number flow. First, the controllable design of the captured flow of the low Mach number design point is realized through the design of the basic flow field with dual design points; second, the streamline tracking is performed through the fan-shaped inlet shape, so that the incident shock wave shape under the high / low Mach number design conditions is consistent with the high / low Mach number basic flow field; third, the low Mach number start of the compression surface is realized by constructing the leading edge actuation profile and the compression side actuation profile, and the streamline deviation caused by the pressure gradient after the incident shock wave under the low Mach number condition is avoided. The hypersonic wide-speed domain retractable inlet designed based on the design method can ensure that the inlet can start at a low Mach number at the beginning of the design, and has the low Mach number flow capture capability that meets the requirements. Develop a wide-domain design method for the retractable inlet in a wide speed domain to improve the designability and design efficiency of the hypersonic wide-speed domain retractable inlet. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Figure 1 The schematic diagram of the basic flow field structure of double design points axisymmetric according to an embodiment of the present invention is shown in FIG. In the figure, the solid line represents the coordinate axis and the wall surface, the dashed line represents the low Mach number flow field boundary, the dotted line represents the low Mach number flow field boundary, and the double dotted line is the auxiliary line for locating the coordinate point.

[0022] Figure 2 The present invention adopts Figure 1 The basic flow field performs streamline tracing on the fan-shaped inlet to obtain a schematic diagram of the basic compression profile.

[0023] Figure 3 It is a schematic diagram of the geometric adjustment of the compression profile of the retracted inlet duct in a wide speed range under low Mach number conditions according to the present invention.

[0024] Figure 4 It is a schematic diagram of the low Mach number profile of the compression profile of the retracted inlet duct in a wide speed range described in the present invention.

[0025] Figure 5 It is a schematic diagram of the geometric adjustment of the high Mach number working condition of the compression profile of the retracted inlet duct in a wide speed range as described in the present invention.

[0026] Figure 6 It is a schematic diagram of the high Mach number profile of the contracted inlet compression profile in a wide speed range described in the present invention.

[0027] The meanings of the marks in the figure are as follows: 1 represents the low Mach number incident shock wave in the basic flow field, 2 represents the lip streamline determined by the low Mach number incident shock wave, 3 represents the flow field area behind the wave determined by the low Mach number incident shock wave, 4 represents the high Mach number incident shock wave determined by the wall busbar, 5 represents the wall busbar determined by the low Mach number incident shock wave, 6 represents the high Mach number flow field area determined by the wall busbar, 7 represents the high Mach number incident shock wave determined by the extended section of the wall busbar, 8 represents the high Mach number flow field area determined by the extended section of the wall busbar, 9 represents the extended section of the wall busbar, 10 represents the wall busbar determined by the high Mach number reflected shock wave, 11 represents the high Mach number flow field area determined by the reflected shock wave, 12 represents the high Mach number reflected shock wave, 13 represents the flow field area behind the reflected shock wave, 14 represents the wall busbar under the basic flow field, 15 represents the lip point, 16 represents the extended section of the wall busbar 17 represents the low Mach number lip streamline determined by the extension section of the wall generatrix, 18 represents the outer contour of the basic flow field, 19 represents the side of the fan-shaped inlet shape, 20 represents the upper side of the fan-shaped inlet shape, 21 represents the leading edge of the compression side of the basic compression profile, 22 represents the intersection line of the side wall of the basic compression profile and the compression side, 23 represents the low Mach number compression profile, 24 represents the downstream boundary of the low Mach number compression profile, 25 represents the compression side actuating profile, 26 represents the intersection line of the modified side wall and the compression side actuating profile, 27 represents the throat boundary of the basic compression profile, 28 represents the lower wall of the basic compression profile, 29 represents the side wall of the basic compression profile, 30 represents the leading edge of the side wall of the basic compression profile, 31 represents the central body of the high Mach number basic flow field, 32 represents the leading edge of the low Mach number lip streamline, 33 represents the leading edge of the low Mach number incident shock wave, and 34 represents the leading edge actuating profile. DETAILED DESCRIPTION

[0028] In order to make the technical problems, technical solutions and beneficial effects to be solved by the present invention clearer and more understandable, the present invention is further described in detail below in conjunction with the accompanying drawings and embodiments.

[0029] like Figure 1 As shown, the dual design point basic flow field design of the embodiment of the present invention includes the following steps:

[0030] 1) Given low Mach number input conditions (including incoming flow conditions, incident shock wave shape, center body height H_low, and capture flow coefficient) and high Mach number input conditions (including incoming flow conditions and center body height H_high).

[0031] 2) According to the low Mach number incident shock wave 1 in the basic flow field, the characteristic line method is used to calculate the post-wave flow field area 3 determined by the low Mach number incident shock wave, and the corresponding wall generatrix 5 determined by the low Mach number incident shock wave is obtained. According to the starting point position of the lip streamline corresponding to the low Mach number capture flow coefficient, streamline tracking is performed to obtain the lip streamline 2 determined by the low Mach number incident shock wave.

[0032] 3) The wall generatrix 5 and the high Mach number incoming flow condition are calculated using the characteristic line method in the high Mach number flow field region 6 determined by the wall generatrix.

[0033] 4) Construct the wall generatrix extension section 9 and apply the characteristic line method to calculate the high Mach number flow field region 8 and the low Mach number flow field region 17 determined by the wall generatrix extension section. By iterating the shape of the wall generatrix extension section 9, the high Mach number incident shock wave 7 determined by the wall generatrix extension section and the low Mach number lip streamline 16 determined by the wall generatrix extension section obtained by continuing streamline tracing on the basis of the lip streamline 2 determined by the low Mach number incident shock wave intersect on the horizontal line corresponding to the high Mach number center body height, and the intersection point is the lip point 15.

[0034] 5) The wall generatrix 5 and the extended wall generatrix 9 together constitute the generatrix of the low Mach number compression profile. Based on the high Mach number flow field region 8, combined with the given high Mach number reflected shock wave 12, the characteristic line method is applied to obtain the high Mach number flow field region 11 determined by the reflected shock wave and the wall generatrix 10 determined by the high Mach number reflected shock wave. Subsequently, the flow field region 13 after the reflected shock wave is calculated based on the aerodynamic parameters of the high Mach number flow field region 11 determined by the reflected shock wave, and the wall generatrix 14 under the basic flow field is obtained.

[0035] At this point, the construction of the basic flow field with dual design points and axisymmetric is completed. Then the flow field can be applied to the design of hypersonic dual design points and wide speed range contraction inlet, such as Figures 2 to 6 The specific steps of designing a hypersonic dual-design-point wide-speed range retractable inlet are as follows:

[0036] 1. The inlet shape of the given air intake is fan-shaped, and the two arcs of the fan-shaped inlet are respectively fitted with the outer contour 18 of the basic flow field and the central body 31 of the high Mach number basic flow field, and the side edge 19 of the fan-shaped inlet shape is a radial line in the basic flow field.

[0037] 2. The basic compression profile of the air inlet can be obtained by tracing the streamlines, making viscosity corrections and geometric modifications to the fan-shaped inlet shape. Among them, the leading edge profile of the basic compression profile side wall 29 is the high Mach number incident shock wave 4 determined by the wall generatrix and the high Mach number incident shock wave 7 determined by the extended section of the wall generatrix, the lower wall profile is the lower wall generatrix 14, the downstream profile is the basic compression profile throat boundary 27, the basic compression profile lower wall 28 is the rotational curved surface of the basic flow field lower wall generatrix, and the compression side leading edge 21 of the compression profile is the upper side 20 of the fan-shaped inlet shape, which is an arc with the same curvature as the basic flow field. The compression profile is a rotational curved surface of the basic flow field wall generatrix, so the intersection 22 of the basic compression profile side wall and the compression profile also has the characteristics of the basic flow field wall generatrix. The compression surface is divided into two upstream and downstream regions by the arc corresponding to the downstream end point of the wall generatrix extension section (the downstream boundary 24 of the low Mach number compression surface). The upstream region is the low Mach number compression surface 23, and the downstream region is divided into two curved surfaces along the symmetry plane as the compression side actuating surface 25.

[0038] 3. The part of the intersection line 22 between the side wall of the basic compression profile and the compression side that is associated with the compression side actuating profile 25 is modified into a straight line by modification, and the intersection line 26 between the modified side wall and the compression side actuating profile is used as the rotation axis of the compression side actuating profile 25 to realize the rotation opening and closing of the compression side actuating profile 25. Under high Mach number conditions, the profile is in a closed state, such as Figure 5 As shown; Under low Mach number conditions, the profile is in an open state, such as Figure 3 shown.

[0039] 4. The area surrounded by the low Mach number incident shock wave 1, the lip streamline 2 determined by the low Mach number incident shock wave, the low Mach number lip streamline 16 determined by the extended section of the wall generatrix, the high Mach number incident shock wave 4 determined by the wall generatrix, and the high Mach number incident shock wave 7 determined by the extended section of the wall generatrix in the basic flow field is taken as the leading edge actuating profile 34 located on the side wall surface. The curve intercepted by the low Mach number incident shock wave 1 is the low Mach number incident shock wave leading edge 33, and the lip streamline 2 determined by the low Mach number incident shock wave and the low Mach number lip streamline 16 determined by the extended section of the wall generatrix are the low Mach number lip streamline leading edge 32.

[0040] 5. The leading edge actuating profile 34 is symmetrically arranged on the two side walls, and rotates around the upstream end point of the intersection line 22 of the two side walls and the compression side profile, and the rotation axis direction is the normal direction of the side wall. Under high Mach number conditions, the profile is deflected and retracted toward the low Mach number compression side profile 23, and at this time, the side wall leading edge of the inlet is the basic compression side wall leading edge 30; under low Mach number conditions, the profile is deflected toward the lower wall 28 of the basic compression profile, and serves as the new side wall leading edge.

[0041] The present invention uses dual design conditions as input to provide a new wide-range design method for a retractable inlet in a wide speed range. The present invention determines the high / low Mach number design points of the hypersonic retractable inlet according to the design requirements of the hypersonic aircraft; constructs a dual-design point axisymmetric basic flow field; places the fan-shaped inlet in the high Mach number basic flow field to perform streamline tracing to obtain the basic compression profile of the inlet; and designs the actuation profile and actuation scheme for high / low Mach number conditions. The above embodiments are only preferred embodiments of the present invention and cannot be considered to limit the scope of implementation of the present invention. All equivalent changes and improvements made within the scope of the application of the present invention should still fall within the scope of the patent coverage of the present invention.

Claims

1. A dual-design-point wide-speed range contraction inlet design method based on curved shock waves, characterized by The following steps are involved: 1) Given the high / low Mach number incoming flow conditions of the basic flow field with dual design points, the center body height of the basic flow field with high / low Mach numbers, the low Mach number incident shock wave shape, and the low Mach number capture flow coefficient; 2) Calculate the starting height of the lip streamline in the low Mach number basic flow field according to the low Mach number capture flow coefficient, and determine the coordinates of the starting point of the lip streamline in combination with the low Mach number shock wave shape; 3) applying the characteristic line method to calculate the known flow field behind the wave according to the low Mach number shock wave shape, obtaining part of the known wall generatrix, and then combining the starting coordinates of the lip streamline in step 2) to perform streamline tracking, obtaining part of the known lip streamline in the low Mach number flow field; 4) applying the characteristic line method to the wall generatrix obtained in step 3) and the high Mach number incoming flow condition in step 1) to calculate the partial incident shock wave shape and the post-wave flow field under the high Mach number condition; 5) Based on the wall generatrix in step 3), the wall generatrix is ​​extended in the flow direction to form an extended section, and the high / low Mach number flow field corresponding to the extended section is obtained by applying the characteristic line method, and then the low Mach number lip streamline and the high Mach number incident shock wave shape are obtained, and the two intersect at the horizontal line corresponding to the height of the high Mach number center body, and the intersection point is the lip point; the extended section of the wall generatrix and the wall generatrix known in step 3) together constitute the low Mach number wall generatrix of the basic flow field; 6) Based on the low Mach number wall busbar in step 5), the high Mach number flow field is obtained in combination with the high Mach number incoming flow condition, and the reflected shock wave parameters are given. The wavefront flow field and the wavefront flow field corresponding to the reflected shock wave are obtained by applying the characteristic line method, wherein the upper boundary of the wavefront flow field is connected to the low Mach number wall busbar to form the high Mach number busbar of the basic flow field, and the lower boundary of the wavefront flow field is the high Mach number lower wall busbar of the basic flow field; at this point, both the high Mach number basic flow field and the low Mach number basic flow field are constructed; 7) placing the fan-shaped inlet shape into the high Mach number basic flow field of step 6) for streamline tracing, wherein the two arc edges of the fan-shaped shape are respectively fitted with the outer contour and the center body of the high Mach number basic flow field, and at this time, the two side edges of the fan-shaped shape coincide with the radial lines of the basic flow field; the profile obtained by streamline tracing is the basic compression profile of the contraction inlet duct in the hypersonic dual design point wide speed domain after scaling and viscosity correction; 8) With the arc corresponding to the downstream end point of the low Mach number generatrix as the boundary, the upstream side of the basic compression profile in step 7) is the low Mach number compression profile, and the downstream side is the compression side actuation profile; the intersection line of the two compression side actuation profiles and the side wall surface is modified into a straight line as the rotation axis, and the compression side actuation profile is divided into two small sectors along the symmetry plane, which rotate around the rotation axis respectively to achieve geometric adjustment; 9) On the basis of the basic compression profile in step 7), the area surrounded by the leading edge profile of the walls on both sides of the basic compression profile and the low Mach number incident shock wave and the lip streamline is used as the leading edge actuation profile; the leading edge actuation profile is actuated by rotation, the rotation axis is located at the leading edge point of the compression side, and the rotation axis direction is the normal direction of the radial plane where the side wall surface is located; 10) Under high Mach number conditions, the leading edge actuating profile in step 9) deflects toward the compression side, and at this time, the shape of the front edge of the side wall of the compression profile is the front edge of the side wall of the basic compression profile; under low Mach number conditions, the leading edge actuating profile deflects toward the lower wall surface, and at this time, the front edge of the side wall of the compression profile is the front edge of the leading edge actuating profile.

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