Front body / air inlet channel integrated design method based on longitudinal segmented multi-stage compression

Abstract

The invention discloses a front body / air inlet channel integrated design method based on longitudinal segmentation multi-stage compression, and relates to the field of high-speed aircraft front body / air inlet channel design. The method comprises the steps that at least two-stage compressed cone-guided waverider front bodies are designed, and the incoming flow Mach number of an air inlet channel is determined; constructing an inner cone type axisymmetric reference flow field by adopting a characteristic line design method, and performing matching design on the axisymmetric reference flow field and a cone-guided waverider precursor shock wave to form an air inlet capture molded line; adopting a forward and reverse tracking design fusion mode to obtain an inward rotation type air inlet channel; and the inward rotation type air inlet channel and the cone-guided waverider precursor are fused, and the precursor and air inlet channel integrated configuration is obtained. After the forward and reverse fusion design is adopted for the air inlet channel, the air inlet channel is further fused with the front body, and compared with the existing independent design of the front body and the air inlet channel, the air inlet channel has the advantages that the pre-compression performance of the air inlet channel is improved, the length of the front body of the aircraft is shortened, the total pressure recovery and supercharge ratio of the air inlet channel is increased, and the starting performance of the air inlet channel is widened.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of high-speed aircraft forebody / inlet design. More particularly, the present application relates to a forebody / inlet integrated design method based on longitudinal segmented multi-stage compression. BACKGROUND

[0002] Air-breathing high-speed aircraft is powered by a scramjet engine, which can obtain oxygen from air, thus eliminating the need for additional oxidizer during flight, and enabling high-speed long-range flight. The forebody of the air-breathing high-speed aircraft is closely integrated with the inlet, and the forebody provides compressed air flow for the high-speed aircraft. The compression performance of the forebody plays a decisive role in the performance of the inlet and the scramjet engine. The waverider body, as a promising aerodynamic shape for high-speed aircraft, can pre-compress high-speed airflow to provide high-pressure uniform airflow for the inlet.

[0003] In order to improve the pre-compression performance of the forebody, the compression performance of the waverider body is closely related to the shock angle design. If single-stage compression is used, a larger shock angle is usually required to meet the demand for inlet air pressure ratio, but a larger shock will result in a larger lifting moment and reduce the lift-drag ratio of the forebody. Isothermal compression usually makes the forebody very long, which is not conducive to the starting performance of the inlet.

[0004] In the prior art, patent number 202210304326.8, a cone guide waverider forebody longitudinal segmented multi-stage compression design method, proposes a multi-stage compression technology that can achieve longitudinal segmentation. It can change the compression level, compression strength of different stages, compression length of different stages, and other design parameters according to the performance requirements of the inlet, which is very suitable for the forebody design field of hypersonic aircraft. Patent application number 202210304413.3, a tangent cone hypersonic waverider forebody longitudinal segmented multi-stage compression design method, can change the compression level, compression strength of different stages, and compression length of different stages according to the performance requirements of the inlet, and has great application potential in the forebody design field of hypersonic aircraft. However, the above two technologies are designed for the forebody and do not integrate the forebody with the inlet. The integrated design of the forebody and the inlet involves matching the shock contour, selecting the inlet flow, and selecting the capture contour, and is also closely related to the cruise state of the aircraft. If the integrated design is not good, the advantages of the multi-stage compression forebody may not be realized, and even the performance of the inlet may be worse. A simple multi-stage compression forebody cannot evaluate the integrated design performance of the high-speed aircraft. How to design the inlet matching for a specific multi-stage compression forebody is the core problem of improving the performance of the high-speed aircraft inlet. Only by realizing the coupling and integrated design of the multi-stage compression forebody and the inward turning inlet can the pre-compression performance of the forebody be better utilized and the overall performance of the aircraft be improved. SUMMARY

[0005] An object of the present application is to solve at least the above problems and / or defects, and to provide at least the advantages to be described later.

[0006] To achieve these objects and other advantages of the present application, a precursor / inlet duct integrated design method based on longitudinal segmentation multi-stage compression is provided, comprising: S1, a conical ducted waverider precursor design is completed based on a conical ducted waverider precursor design method, and a wave-after flow field parameter of the conical ducted waverider precursor is extracted through CFD calculation; S2, a free stream Mach number of the inlet duct is determined based on the wave-after flow field parameter; S3, an internal turning axisymmetric reference flow field is constructed based on the determined free stream Mach number, using a characteristic line design method; S4, the axisymmetric reference flow field is matched with a conical ducted waverider shock wave to form an inlet duct capture contour; S5, an internal turning inlet duct is obtained in a forward and reverse tracking design fusion manner according to the determined inlet duct capture contour; S6, the internal turning inlet duct is fused with the multi-stage compression conical ducted waverider precursor to obtain a precursor inlet duct, and the precursor inlet duct is locally modified.

[0007] Preferably, in S3, the axisymmetric reference flow field is constructed using a rotational characteristic line theory under a high-speed free stream, wherein a two-dimensional supersonic isentropic flow without viscosity and adiabatic needs to satisfy the following control equation: Characteristic line equation along the streamline: In the above formula, represents the characteristic line slope, represents the y-direction velocity, u represents the x-direction velocity, x represents the axial coordinate, and y represents the longitudinal coordinate; Compatibility equation along the streamline: In the above formula, p , p , a , V respectively represent density, pressure, sound speed, and flow velocity, represents the characteristic line slope; Characteristic line equations along left and right Mach lines are: In the above formula, represents the slope of the left and right Mach lines, θwhere a represents the angle of local streamline relative to the horizontal line, and a represents the Mach angle; The compatibility equations along the left-running and right-running Mach lines are: where M represents the Mach number, represents the flow type, and the subscript ± represents the parameter along the left-running or right-running Mach line.

[0008] Preferably, in S4, the inlet capture contour is determined based on the intersection of the multi-stage compression precursor and the symmetrical reference flow field.

[0009] Preferably, in S5, the design process of the inward-turning inlet includes: S50, according to the inlet capture contour, the inward-turning forward inlet surface is obtained by using the streamline tracing method based on the forward design; S51, based on the outlet contour design requirements of the inward-turning inlet, the fusion reference streamline is formed by the reverse tracing method, so as to obtain the inward-turning reverse design flow passage by using the reverse tracing design method; S52, the inward-turning forward inlet is fused with the inward-turning reverse design flow passage to form the inward-turning inlet.

[0010] Preferably, in S6, the content of the local modification of the precursor inlet includes: The lip of the inlet and the leading edge of the precursor are blunted; The modification design of the inlet and the fuselage is realized by the design of the spline curve.

[0011] Preferably, in S6, before the inward-turning inlet is fused with the multi-stage compression conical ducted waverider precursor, whether the inlet strength of the inward-turning inlet and the multi-stage compression performance of the precursor meet the requirements is judged by the following formula: In the above formula, is the pressure at the inlet of the inlet after the multi-stage compression of the conical ducted waverider precursor, is the Mach number at the inlet of the inlet after the multi-stage compression of the conical ducted waverider precursor, represents the empirical relationship between the pressure before and after the compression shock wave, and , is the empirical relationship between the Mach number before and after the compression shock wave, and , represents the incoming flow pressure, represents the incoming flow Mach number, , , and are constant parameters, i the value of is the number of compression stages; When the inlet entrance intensity and the precursor multi-stage compression performance meet the requirements, the internal turning inlet configuration is fused and designed with the multi-stage compression conical guide waverider precursor through the intersection line of the lower surface of the conical guide waverider precursor, so that the integrated precursor inlet is obtained.

[0012] The present application at least includes the following beneficial effects: Firstly, the integrated fusion design method of the precursor inlet provided by the present application can reduce the Mach number of the incoming flow into the inlet, so that the inlet has better total pressure recovery performance under the constraint of the same outlet Mach number at a low Mach number inlet. After the multi-stage compression of the precursor, the airflow entering the inlet can adapt to a larger range of flow and attack angle range, and has potential value in low Mach number starting. The multi-stage compression precursor can reduce the precursor length and the inlet length under the same pressure ratio, avoid the increase of the boundary layer thickness caused by the too long precursor, and improve the compression efficiency of the inlet, thereby providing a new design method for the optimization design of the inlet.

[0013] Secondly, the integrated fusion design method of the precursor inlet provided by the present application further fuses the precursor after the forward and reverse fusion design of the inlet, and has the advantages of improving the pre-compression performance of the inlet, shortening the length of the precursor of the aircraft, improving the total pressure recovery and pressure ratio of the inlet, and widening the starting performance of the inlet, compared with the independent design of the existing precursor and inlet.

[0014] Other advantages, objects and features of the present application will be partly embodied in the following description, and will be partly understood by those skilled in the art through research and practice of the present application. BRIEF DESCRIPTION OF DRAWINGS

[0015] Fig. 1 is a design schematic diagram of a conical guide waverider precursor in the prior art; Fig. 2 is a schematic diagram of the geometric relationship after two-stage compression in the prior art; Fig. 3 is a design schematic diagram of a precursor multi-stage compression of the present application; Fig. 4 is a pressure distribution diagram of different tangent surfaces of a conical guide waverider precursor after two-stage compression of the present application; Fig. 5 is a symmetric surface pressure cloud distribution diagram of a conical guide waverider precursor after two-stage compression of the present application; Fig. 6 is a three-dimensional internal turning inlet of the present application; Fig. 7 is a schematic diagram of the integrated design of the multi-stage compression precursor and the inlet of the present application; Fig. 8 is a schematic diagram of the inlet capture generatrix and the internal conical flow field of the present application; Fig. 9 is a schematic diagram of the throat outlet generatrix of the present application; Fig. 10 is a schematic diagram of a multi-stage compression precursor inlet integrated design configuration of the application; Fig. 11 is a grid diagram of the precursor inlet integrated structure of the application; Figure 12 Fig. 12 is a pressure cloud diagram of different longitudinal sections of the precursor inlet integrated design of the application; Fig. 13 is a symmetric plane pressure cloud diagram of the application; Fig. 14 is a symmetric plane density gradient cloud diagram of the application; Fig. 15 is a schematic diagram of density gradient of different sections along the wing span direction of the application; Fig. 16 is a schematic diagram of surface pressure when the Mach number is 6.0 and the attack angle is 0 degrees (viscous calculation result) of the application; Fig. 17 is a schematic diagram of the throat Mach number (inviscid) of the application; Fig. 18 is a schematic diagram of the throat Mach number (viscous) of the application; Fig. 19 is a schematic diagram of the throat total pressure recovery coefficient (inviscid) of the application; Fig. 20 is a schematic diagram of the throat total pressure recovery coefficient (viscous) of the application; Fig. 21 is a schematic diagram of the throat pressure ratio (inviscid) of the application; Fig. 22 is a schematic diagram of the throat pressure ratio (viscous) of the application. DETAILED DESCRIPTION

[0016] The application will be further described in detail below with reference to the accompanying drawings, so that those skilled in the art can implement the application according to the description and the drawings.

[0017] To further improve the compression performance of the waverider precursor and alleviate the compression demand of the inlet, a longitudinal segmented multi-stage compression precursor / inlet integrated design method is proposed. The high pressure area of the precursor is mainly concentrated at the middle position of the inlet entrance, avoiding edge pressure leakage, and achieving the design goals of shortening the length of the precursor, improving the lift-drag ratio, and improving the total pressure recovery of the inlet. The precursor / inlet integrated design based on two-stage cone guide compression is completed, and the aerodynamic performance, flow field distribution and total pressure distribution of the precursor inlet are analyzed. The analysis results show that the method can effectively improve the compression efficiency of the inlet and improve the flow field quality at the outlet of the inlet. In the field of precursor design of high-speed aircraft, the method has great application potential, and has not been seen in domestic and foreign public documents, and has high innovation and engineering application prospect.

[0018] A precursor / inlet integrated design method for air-breathing high-speed aircraft, comprising: S1, based on the cone guide waverider precursor design method, a two-stage cone guide waverider precursor is designed, and the wave behind flow field parameters of the two-stage cone guide waverider precursor are extracted through CFD calculation. In this step, a multi-stage compression precursor is designed based on the segmented proportion adjustable design method.

[0019] S2, determining the inlet flow Mach number based on the flow field parameters after the S1 wave.

[0020] S3, constructing an inner cone axisymmetric reference flow field based on the inlet flow Mach number determined above using a rotating characteristic line design method under high-speed incoming flow.

[0021] S4, matching the axisymmetric reference flow field with the shock wave of the three-stage conic wave-riding forebody to form an inlet capture contour by fusing the wave-riding body flow field with the inner cone flow field, i.e., intersecting the multi-stage compression forebody with the axisymmetric reference flow field to form the inlet capture contour.

[0022] S5, forming the inner turning inlet contour from the determined inlet capture contour by the method of stream tracing, and forming the reverse inner flow passage from the inlet outlet contour by the method of reverse stream tracing, and fusing the forwardly designed inlet with the reversely designed flow passage to form the required inlet; This step fuses the inner turning inlet captured from the leading edge contour with the reverse inner turning flow passage, thereby achieving the integrated inlet layout that meets the fusion design of the inlet and the multi-stage compression wave-riding forebody; S6, the forebody inlet configuration obtained by fusing the inner turning inlet with the conic wave-riding forebody is selected according to the cruise state of the aircraft, and the design parameters of the inlet are determined, which need to be combined with the spanwise velocity distribution after the multi-stage compression of the conic wave-riding forebody to determine the rotation parameters of the reference flow field, and in addition, the inlet contraction ratio needs to be determined according to the strength of the multi-stage compression of the conic wave-riding forebody, thereby maximizing the performance of the inlet, ensuring the comprehensive optimization of the performance parameters such as inlet flow, total pressure recovery and pressure ratio, and in addition, in order to reduce the problem of not starting caused by the thick boundary layer thickness and low flow energy caused by the too long length of the conic wave-riding forebody, the following calculation formulas for the inlet strength of the inner turning inlet and the multi-stage compression performance of the forebody are constructed: In the above formula, refers to the pressure at the inlet of the inlet after the multi-stage compression of the conic wave-riding forebody, refers to the Mach number at the inlet of the inlet after the multi-stage compression of the conic wave-riding forebody. represents the empirical relationship between the pressure before and after the compression shock wave, and the specific coefficient is fitted according to the specific configuration and shock angle design, represents the empirical relationship between the Mach number before and after the compression shock wave, and the specific coefficient is fitted according to the specific configuration and shock angle design. represents the incoming flow pressure, and Ma0 represents the incoming flow Mach number, , , and is a constant parameter, determined according to a specific configuration.

[0023] At the same incoming flow Mach number and the same initial shock wave angle, the outlet Mach number of the lower surface of the cone-drag waverider forebody is less than that of the cone-drag waverider body, which can reduce the design incoming flow Mach number of the inlet. When the incoming flow Mach number is 6, the outlet Mach number and the pressure ratio of the two-stage compression cone-drag waverider forebody are 4.5 and 5.0, and the outlet Mach number of the single-stage compression cone-drag waverider forebody is 5.1 and 1.9, respectively. The multi-stage compression cone-drag waverider forebody integrated design method can effectively improve the pre-compression performance of the cone-drag waverider forebody, reduce the shock wave intensity in the inlet, and thus improve the total pressure recovery coefficient of the inlet. If an isentropic compression forebody is used, when the outlet pressure ratio of the cone-drag waverider forebody is 5.0, the length of the two-stage compression cone-drag waverider forebody can be reduced by 30% compared with the isentropic compression length. On the one hand, the shortening of the cone-drag waverider forebody reduces the low-energy flow entering the inlet and improves the starting ability of the inlet. On the other hand, the shortening of the cone-drag waverider forebody is conducive to reducing the influence on the vertical tail directional stability of the aircraft at high angles of attack.

[0024] It is known through analysis that, when the outlet Mach number of the cone-drag waverider forebody is ensured to be the same, the volume ratio of the single-stage compression cone-drag waverider forebody is increased by 43% compared with that of the two-stage compression cone-drag waverider forebody, resulting in a decrease of 20.4% in the lift-drag ratio of the cone-drag waverider forebody.

[0025] When the forebody-inlet integration is performed, the inlet Mach number and the flow direction of the cone-drag waverider forebody after compression are determined according to the compression strength of the cone-drag waverider forebody, the reference inner cone flow field is designed, the rotation angle of the inward-turning inlet is selected, the inward-turning inlet after rotation (scaled according to the inlet flow demand) and the cone-drag waverider forebody are intersected to determine the capture profile, the inward-turning inlet configuration is obtained based on the obtained capture profile, and then the inward-turning inlet configuration and the multi-stage compression cone-drag waverider forebody are intersected through the lower surface to perform integration design, so as to obtain an integrated forebody-inlet configuration meeting the requirements.

[0026] The forebody-inlet is also locally modified, and the local modification includes: performing blunting treatment on the inlet lip and the forebody leading edge to meet the processing and structural bearing requirements; and performing modification design on the inlet and the body through spline curve design to meet certain volume requirements, so as to form an aerodynamic layout with certain application capability.

[0027] Embodiment: A method for designing an integrated air-breathing high-speed vehicle forebody / inlet, based on a waverider layout, through multi-stage compression waverider forebody and three-dimensional internal turning inlet flow field coupling, can reduce the aerodynamic interference between internal and external flows, moderate the compression requirement inside the inlet, reduce the generation of drag, and is very beneficial to improving the vehicle lift-drag ratio and inlet flow capture performance. The shock wave used by the three-dimensional internal turning inlet is in a contraction state, with relatively small total pressure loss and small air overflow. Compared with two-dimensional compression and side pressure inlets, it has better compression capacity and efficiency. The design method mainly includes: I. Multi-stage compression forebody design Based on the design method of a conical waverider longitudinal segmented multi-stage compression forebody, forebody design is also the premise of internal and external flow integrated design. The number of compression stages can be designed along the longitudinal direction, and the forebody design mainly includes the upper surface of the waverider (it should be noted that the upper surface can or can not be a free flow, and the free flow is a typical design state), the first compression surface, the second compression surface, etc. (the lower surface of the forebody is divided into two-stage compression, three-stage compression, and more stages of compression according to the design requirements, and the length of the two-stage compression and three-stage compression of the lower surface of the forebody can be combined in multiple proportions. Through multi-stage compression of the forebody, the inlet design Mach number can be adjusted. The upper surface of the waverider is designed using the same design method as the free flow. In order to match the inlet, the forebody is based on the given lower surface shock wave outlet profile (ICC, Inlet Capture Curve) and upper surface outlet profile (Flow Capture Curve), according to the flight Mach number Ma and the first stage compression shock angle Determine the compression strength (the Mach number and shock angle in different tangent surfaces can be designed as needed on the premise of ensuring continuous change), use the streamline tracking method to obtain the corresponding lower surface profile and leading edge curve, and use the free flow surface method to obtain the upper surface of the waverider based on the leading edge curve. This embodiment takes a two-stage compression forebody as an example for illustration.

[0028] Since the conical waverider forebody is the design basis of the multi-stage compression conical waverider forebody, the design principle of the conical waverider forebody is described as shown in Figure 1 Figure 1 ​In the formula, COC represents the center of curvature of the shock radius, mark 1 represents the trailing edge curve of the lower surface, and mark 2 represents the radius of the shock. The zero angle of attack conical basic flow field can be obtained by designing the shock angle given the flight Mach number, and the wave rider trailing edge upper surface curve FCC can also be the trailing edge lower surface ICC by giving the wave rider trailing edge upper surface curve FCC. Here, FCC is taken as an example, the upper surface streamline is obtained by reverse tracking in a manner parallel to the incoming flow, the leading edge line is obtained by intersecting the upper surface streamline with the shock surface, and the wave rider lower surface streamline is obtained by tracking the leading edge line, so that the conical wave rider forebody is obtained. The design parameters mainly include the expansion angle , the shock angle β , the Mach number Ma and the relative axial distance R 0, etc. Generally, the trailing edge curve equation of the upper surface is as follows, wherein, R 0 represents the distance between the symmetric position of the trailing edge curve and the axis, represents the expansion angle of the curve, represents the tangent angle of the curve at the end point. Given the parameters R 0, , , the equation of the curve can be obtained.

[0029] In the formula, a f represents the quadratic term coefficient, b f represents the quartic term coefficient, x and y represent the horizontal and vertical coordinates, and the unit is meter; Further, as shown in Figure 2 and Figure 3 , the multi-stage compression wave rider leading edge curve is divided into the first stage leading edge line and the second stage edge line according to the design requirements, so that the longitudinal length of the wave rider can be obtained.

[0030] For a two-stage wave rider, in order to ensure that the first stage compression and the second stage compression shock wave just match at the outlet, since the streamline behind the conical flow shock wave is curved, and the Mach number and direction corresponding to each point of the first stage compression lower surface are different, in order to ensure that the first stage shock wave and the second stage shock wave can match at the outlet, the second stage compression shock angle corresponding to each first stage node is different. As shown in Figure 2 , the left side is a conical two-stage wave rider compression schematic diagram Figure 2In the figure, mark 3 represents the first stage compression, and mark 4 represents the second stage compression. The leading edge line of the wave rider in the figure is obtained based on an axisymmetric reference flow field. The position corresponding to the first stage compression is determined according to the longitudinal length ratio. For example, the upper red line E-A-F in 2 is the leading edge line of the first stage compression, and the middle green line E-L-F is the lower surface curve at the outlet of the first stage compression, which is also part of the leading edge line of the second stage compression. The lower two green lines E-P and F-Q are the leading edge lines of the two sides of the second stage compression. The corresponding lower surfaces of the second stage compression are obtained by using stream tracing. The red line P-G on the top left is the trailing edge line obtained by the second stage compression of E-P. The red line G-M-H in the middle part on the top is the trailing edge line obtained by the second stage compression of E-L-F. The red line H-Q on the top right is the trailing edge line obtained by the second stage compression of F-Q. In order to better illustrate the geometric relationship that needs to be met by the local tangent plane, Figure 3 l 1 represents the length of the first stage compression, and M represents the outlet point of the lower surface stream line. Ma1 represents the Mach number of the incoming flow, and Ma2 represents the Mach number after the first stage compression, represents the direction of the velocity of the gas after the first stage compression. It is given Figure 2 The local tangent plane in which O-O'-B-C is located is illustrated. The flow field needs to meet the following relationship: In the above formula, β 2 is the shock wave angle of the second stage compression, β 1 is the shock wave angle of the first stage compression, d 1 is the distance from the local horizontal axis, d A is the distance from the horizontal axis of point A, l 2 is the length of the second stage compression, is the angle between the velocity at the outlet of the first stage compression and the local horizontal axis; Further, the first stage compression surface is designed from the leading edge line. The first stage lower surface stream line is obtained by using stream tracing. The stream line is truncated so that its edges on both sides of the corresponding leading edge line along the longitudinal coordinate are not greater than L1. All flow field parameters of the first stage stream line cutoff points are extracted.

[0031] ​Further, the first stage compression surface streamline cut-off point, that is, the outlet of the first stage compression surface, constitutes the internal front edge line of the second stage compression, and the outlet line of the first stage compression surface and the second stage edge line jointly constitute the front edge line of the second stage compression. The second stage edge line corresponds to the lower surface streamline obtained by streamline tracking of the axisymmetric reference flow field, and the flow field parameters corresponding to the first compression outlet line are equivalent to the internal front edge line inflow parameters of the second stage compression. The Mach number, speed, direction and pressure at each feature point are different. The corresponding axisymmetric reference flow field center, inflow direction, Mach number and shock angle need to be re-established according to the geometric relationship of the second stage compression. The Mach number, speed direction and relative position corresponding to the second stage compression front edge are different. In order to ensure that the shock wave of the second stage compression coincides with the shock wave of the first stage at the inlet of the inlet, the relationship between the length of the second stage and the shock angle corresponding to the second stage compression needs to be found in the two-dimensional section corresponding to the different expansion angles.

[0032] Further, when each expansion angle corresponds to the second stage compression shock angle β 2Determined, the Mach number, shock angle and speed direction of the front edge point can be used to construct a new conical flow field, so as to solve the lower surface streamline in the section corresponding to different expansion angles. The streamlines obtained by all the second stage front edge points jointly constitute the lower surface of the second stage compression. The streamlines obtained by the first stage compression are used to loft the first stage compression surface, and the streamlines of the second stage compression can be used to obtain the second stage compression surface, so as to obtain the entire waverider configuration. By analogy, a three-stage compression or even more stage precursor compression can be designed.

[0033] At this point, first, the two-stage compression precursor design is completed. The lower surface of the waverider is divided into multiple stages along the longitudinal direction according to the requirements. In order to ensure that the multiple stage compression shock waves in each section converge at the inlet of the inlet, a strict geometric relationship between the length and the corresponding shock angle is established. At the same time, the edge line remains the same as the single-stage tangent cone waverider during the multi-stage compression design process, which can ensure that the waverider precursor designed meets the waverider characteristics at the edge, without pressure leakage, and fully integrates the advantages of waverider characteristics and multi-stage compression. The numerical analysis results are shown in Figure 4 and Figure 5 , wherein P represents the local pressure, and the incoming flow pressure. In addition, this method can also change the shock angle at different positions according to the requirements to achieve the characteristics of changing the pressure distribution at the inlet of the inlet. The multi-stage compression precursor obtained by this method has high pre-compression ability, can provide high-pressure low-speed and relatively uniform airflow for the inlet, and the outlet profile of the inlet is flexible and adjustable, which is convenient for matching with the inlet. At the same time, the precursor also maintains a high lift-drag ratio, which has obvious advantages compared with the traditional single-stage compression or isentropic compression.

[0034] II. Design of internally turning inlet based on multi-stage compression precursor The three-dimensional inward turning inlet design is actually a kind of "reverse design" method, which firstly constructs an inward turning reference flow field based on the inward turning conical flow field, and then obtains an inlet curve according to the inlet capture shape, and generates flow lines based on the leading edge curve in the inward turning reference flow field through the stream line tracking technology, and all the flow lines are closed to form the inward turning inlet surface.

[0035] Firstly, the axisymmetric reference flow field is constructed by the rotational characteristic line theory. Under high-speed incoming flow, the axisymmetric reference flow field is constructed by the rotational characteristic line theory. The lower surface of the wave-multiplying and the precursor shock wave surface are intersected with the shock wave surface of the axisymmetric reference flow field to form the capture generatrix of the inlet. Then, different flow lines are obtained by stream line tracking, and the flow surface is formed by lofting to constitute the inward turning inlet surface. Since the solution of the axisymmetric reference flow field is obtained by using the characteristic line theory, the characteristic line theory is only applicable in the supersonic flow field. Therefore, when the axisymmetric reference flow field is constructed, a small part of the wedge is designed to generate the appendage shock wave.

[0036] The solution of the axisymmetric reference flow field is obtained by using the characteristic line theory, and the normal derivative of the characteristic line may be discontinuous. In the flow field, the stream line and the Mach line are weak discontinuous lines, and the flow field parameters satisfy the compatibility relationship along the characteristic line direction. The two-dimensional supersonic isentropic flow with no viscosity and adiabatic satisfies the following control equation.

[0037] The characteristic line equation along the stream line is: In the above formula, represents the slope of the characteristic line, represents the y-direction velocity, u represents the x-direction velocity, x represents the axial coordinate, and y represents the longitudinal coordinate; The compatibility equation along the stream line is: In the above formula, p , p , a , V respectively represent the density, pressure, sound speed and flow velocity, represents the slope of the characteristic line; The characteristic line equation along the left and right Mach lines is: In the above formula, represents the slope of the left and right Mach lines, θ represents the angle of the local stream line relative to the horizontal line, and a represents the Mach angle; The compatibility equation along the left and right Mach lines is: where M represents Mach number, where the subscript ± indicates the parameter along the left or right Mach line.

[0038] The above equation is rewritten as a finite difference equation, and then solved by iterative promotion. There are mainly four forms: inner point element, direct wall point element, inverse wall point element and shock boundary point element. The axisymmetric reference flow field is solved by the method of characteristics, and a small part of the wedge is designed to generate the shock wave of the appendage when the initial section of the body is generated. According to the inlet capture contour, the internal turning inlet can be obtained by stream tracing, as shown in Figure 6 .

[0039] III. Fusion design of precursor inlet When constructing the intersection line of the inlet leading edge, the inward converging shock wave is usually rotated at a certain angle. As shown in Figure 7 , the inward cone shock wave 5 is rotated at a certain angle of positive attack angle, intersects with the precursor reference shock wave 6 and the compression surface 7 (i.e. the lower surface of the waverider), and then the whole is rotated at the same angle of negative rotation. Then the intersection curve is extracted as the capture contour 8 (as shown by the red line part in Figure 8 ), and the inlet surface is obtained by stream tracing. The designed inlet is connected and fused with the precursor. Finally, the overall shape is rotated at a positive attack angle. In order to ensure the closure of the capture contour, the lower half of the capture contour is modified so that a section of the capture contour of the inlet is located above the reference center body 9 of the internal turning inlet.

[0040] Because the capture contour obtained by stream tracing of the inward cone shock wave intersecting with the waverider is irregular, the matching and fusion effect with the isolator is not good, which may lead to the decrease of flow field uniformity, therefore, by giving the outlet contour, the reverse stream tracing technology is used to generate the internal flow passage again, the inlet captured from the leading edge contour is fused with the internal flow passage traced in reverse, and the outlet contour of the throat is as shown in Figure 9 . Figure 9 (the red line with a ring in represents the positive design outlet, and the blue line with a triangle represents the given outlet contour), so as to meet the fusion design of the inlet and the waverider body to ensure the flow capture, and also meet the design requirements of the isolator.

[0041] In this embodiment, through the combination of multi-stage compression precursor and internal turning inlet, the total pressure recovery performance and pressure ratio of the inlet can be further improved.

[0042] Firstly, the forebody / inlet integration design scheme is completed based on the two-stage compression forebody and the internal turning inlet. Firstly, the multi-stage compression forebody is designed. The cruise Mach number is given as 6.0, the reference shock wave angle is given as 12°, the upper surface curve FCC of the afterbody of the wave rider is given, and the two-stage compression wave rider design method is used to complete the forebody design.

[0043] Secondly, the inlet flow Mach number and the streamline direction are determined according to the flow field of the two-stage compression forebody. Herein, the ratio of the center body radius to the inlet shock wave radius is 0.2 when the reference internal cone flow field is designed, the inlet flow Mach number of the reference internal turning flow field is designed as 4.5, and the outlet Mach number is 2.57. Since the internal turning inlet is designed based on the uniform inlet flow, the compressed air flow direction changes after the intersection of the internal cone shock wave and the forebody shock wave. When the intersection line of the inlet leading edge is constructed, the reference internal turning flow field is usually rotated by a certain angle. At the same time, it is ensured that the intersection point of the capture profile of the inlet is located above the reference center body of the internal turning inlet. According to the design characteristics of the multi-stage compression wave rider, the discontinuous breakpoint is generated when the internal cone shock wave intersects with the lower surface of the wave rider forebody. Herein, the intersection line is projected onto the plane perpendicular to the inlet flow, then the projected line is discretized and fitted by a spline as the capture profile of the inlet. The forebody / inlet obtained by the case design is shown in Figure 10

[0044] Secondly, the design method is verified by numerical simulation. The calculation region is discretized by using the structured grid (see Figure 11 ), the inlet flow Mach number is 6.0, the inlet flow static pressure is 1197.03 pa, and the inlet flow temperature is 226.509 K. Under the condition of no viscosity, the outlet Mach numbers at the attack angles of 0° and 2° are 2.743 and 2.59 respectively. Since the design Mach number of the forebody is 4.5, the calculated Mach number of the forebody is slightly greater than 4.5, so the outlet Mach number is slightly increased, which basically meets the design requirements.

[0045] The no-viscosity calculation results are shown in Table 1. When the attack angle is 0°, the total pressure recovery coefficient of the inlet throat is 0.831, and the pressure rise ratio is 54.28. When the attack angle is 2°, the total pressure recovery coefficient of the throat is 0.7873, and the pressure rise ratio is 70.23.

[0046] Table 1: No-viscosity calculation results Table 2 shows that under the viscous condition, the total pressure recovery coefficient of the throat at the attack angle of 0° is slightly lower than that at the no-viscosity design point, the Mach number changes from 2.743 to 2.3215, which decreases by 15.37%, the total pressure recovery coefficient changes from 0.831 to 0.686, which decreases by 17.45%, and the pressure rise ratio changes from 54.28 to 74.98, which increases by 38.14%.

[0047] Table 2: Viscous calculation results​ As shown in Figure 12 and Figure 16 , the pressure coefficient distribution of the waverider forebody along the longitudinal direction of different sections is given, it can be seen that the edge of the waverider body has a slight leakage of pressure due to blunting and viscous calculation, and the high pressure area is generally on the lower surface of the vehicle. Through the analysis of the pressure cloud distribution of the lower surface, the forebody compression surface appears a relatively obvious layered compression, which is basically consistent with the multi-stage compression forebody design theory. There is a slight pressure leakage at the lip of the inlet, the main reason is that the inlet is designed based on the inviscid characteristic line theory, and the average value of the incoming Mach number of the basic flow field is also based on the average value of the Mach number behind the wave. When the multi-stage compression forebody is integrated with the internal turning inlet, the interference problem occurs when the designed inlet and waverider body are matched, which may be related to the selection of the incoming Mach number when designing the inlet, and at the same time, the compression surface of the lower surface of the multi-stage compression forebody presents a gradient change, resulting in a lower compression degree of the inlet than that of the forebody.

[0048] In order to further analyze the shock wave distribution of the symmetric plane lip, Figure 13 the pressure ratio distribution of the symmetric plane is given, it can be seen that the first and second compression shocks of the forebody are basically hit on the inlet lip position, and the pressure ratio at the inlet entrance reaches about 5. As shown in Figure 14 , the density gradient change of the symmetric plane can also be seen, the multi-stage compression forebody appears two compression shocks, and the reflected shock wave at the lip position is just hit on the upper surface throat position, which is basically consistent with the design flow field. The compression region of the multi-stage compression waverider forebody appears two high-density sections, which is consistent with the design of the multi-stage waverider forebody, and in the first compression region, only the high-density gradient appears near the waverider body wall, and in the second compression region, two high-density regions appear, and the two compression regions converge at the inlet lip position, ensuring that the multi-stage compression high-pressure flow field of the forebody enters the interior of the inlet, realizing the efficient capture of the multi-stage compression flow field and improving the performance of the inlet.

[0049] Figure 15 The density gradient change under different sections along the spanwise direction of the waverider body is given, in the vicinity of the inlet, the two compression shocks of the forebody are hit on the lip position, indicating that the designed three-dimensional internal turning inlet is well matched with the shock wave of the multi-stage compression forebody. Through the shock wave distribution of the symmetric plane lip, the multi-stage compression forebody appears two compression shocks, and the reflected shock wave at the lip position is just hit on the upper surface throat position, which is basically consistent with the design flow field. Figures 17-22 The performance of the multi-stage compression forebody inlet under a wide range of Mach numbers is given, it can be seen that the inlet integration design based on the multi-stage compression forebody can well improve the performance of the inlet, widen the starting range of the inlet, increase the throat pressure ratio and total pressure recovery coefficient, and has obvious gain for improving the performance of the vehicle.

[0050] The above-described arrangements are merely illustrative of the application and should not be construed as limiting the same. There can be many variations of the application and the generic principles defined herein can be applied to other types of embodiments and implementations without departing from the scope of the application.

[0051] The application is not limited to the foregoing specific embodiments, and multi-stage compression of the conicoguide wave rider can be extended to three-stage compression or even more stage compression rider design. The multi-stage compression wave rider can reduce the incoming flow Mach number into the inlet, and under the constraint of the same exit Mach number, the inlet has better total pressure recovery performance at low Mach number inlet. The integrated design of the multi-stage compression rider / inlet can reduce the length of the rider and the inlet under the same pressure ratio, avoid the increase of the boundary layer thickness caused by the long rider, and avoid the influence of low energy flow on the inlet start. The combination of multi-stage shock compression of the rider and Mach wave compression of the internal turning inlet has potential application value in improving the compression performance and total pressure recovery performance of the inlet.

[0052] While the embodiments of the application have been disclosed as above, it is not limited to the use as set forth in the specification and embodiments. It can be fully applied to various fields suitable for the application. Those skilled in the art can easily make other modifications. Therefore, the application is not limited to specific details and figures shown and described herein without departing from the general concept defined by the claims and their equivalent scope.

Claims

1. A precursor / inlet duct integrated design method based on longitudinal segmentation multi-stage compression, characterized in that, The method comprises the following steps: S1. Designing a conical-boattail wave-riding forebody for at least two-stage compression based on a conical-boattail wave-riding forebody design method, and extracting wave-after flow field parameters of the conical-boattail wave-riding forebody through CFD calculation; S2. Determining the incoming flow Mach number of the inlet based on the wave-after flow field parameters; S3. Constructing an axisymmetric reference flow field based on the determined incoming flow Mach number by using a characteristic line design method; S4. Matching the axisymmetric reference flow field with the conical-boattail wave-riding forebody shock wave to form an inlet capture profile; S5. Obtaining an internal turning inlet by using a forward and reverse tracing design fusion method according to the determined inlet capture profile; S6. Fusing the internal turning inlet with the multi-stage compression conical-boattail wave-riding forebody to obtain a forebody-inlet configuration, and locally modifying the forebody-inlet.

2. The precursor / inlet duct integrated design method based on longitudinal segmentation multi-stage compression of claim 1, wherein, In S3, the axisymmetric reference flow field is obtained by using a rotational characteristic line theory under a high-speed incoming flow, wherein a two-dimensional supersonic isentropic flow without viscosity and adiabatic needs to satisfy the following control equations: Characteristic line equation along the streamline: In the above formula, denotes the characteristic line slope, denotes the y-direction velocity, u denotes the x-direction velocity, x denotes the axial coordinate, and y denotes the longitudinal coordinate; Compatibility equation along the streamline: In the above formulae, Characteristic line equation along the left and right Mach lines: , p , a , V respectively represent density, pressure, sound speed and flow velocity, represent the characteristic line slope; Compatibility equation along the left and right Mach lines: In the above formula, denotes the slope of the left and right Mach lines, In S4, the inlet capture profile is determined based on the intersection line of the multi-stage compression forebody and the axisymmetric reference flow field. denotes the angle of the local streamline relative to the horizontal, a denotes the Mach angle; In S5, the design process of the internal turning inlet comprises: where M represents the Mach number, denotes the flow type, and the subscript ± indicates the parameter along the left or right running Mach line.

3. The precursor / inlet duct integrated design method based on longitudinal segmentation multi-stage compression of claim 1, wherein, S50. Obtaining an internal turning forward inlet profile based on forward design by using a flow line tracing method according to the inlet capture profile; 4. The precursor / inlet duct integrated design method based on longitudinal segmentation multi-stage compression of claim 1, wherein, S51. Forming a fusion reference flow line by using a reverse tracing method to obtain an internal turning reverse design flow passage by using a reverse tracing design method based on the outlet profile design requirement of the internal turning inlet; S52. Fusing the internal turning forward inlet with the internal turning reverse design flow passage to form an internal turning inlet. In S6, the local modification of the forebody-inlet comprises: Blunting the inlet lip and the forebody leading edge; 5. The precursor / inlet duct integrated design method based on longitudinal segmentation multi-stage compression of claim 1, wherein, Modifying the inlet and the body by using a spline curve design. Before fusing the internal turning inlet with the multi-stage compression conical-boattail wave-riding forebody in S6, whether the inlet entrance intensity and the multi-stage compression performance of the forebody meet the requirements is determined by the following formula: When the inlet entrance intensity and the multi-stage compression performance of the forebody meet the requirements, the internal turning inlet configuration and the multi-stage compression conical-boattail wave-riding forebody are fused by the intersection line of the lower surface, so as to obtain an integrated forebody-inlet configuration meeting the requirements.

6. The precursor / inlet duct integrated design method based on longitudinal segmentation multi-stage compression of claim 1, wherein, ​ In the above formula, is the pressure at the inlet of the inlet duct after multi-stage compression of the conical ducted waverider forebody, is the Mach number at the inlet of the inlet duct after multi-stage compression of the conical ducted waverider forebody, represents an empirical relationship of the pressure before and after the compression shock wave, and , is an empirical relationship of the Mach number before and after the compression shock wave, and , represents the incoming flow pressure, represents the incoming flow Mach number, , , and are constant parameters, i the value of is the number of compression stages; ​

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