An integrated design method for the forebody / inlet based on longitudinal segmented multi-stage compression

Abstract

This invention discloses an integrated design method for a forebody / inlet based on longitudinal segmented multi-stage compression, relating to the field of high-speed aircraft forebody / inlet design. The method includes: designing a conical waverider forebody with at least two stages of compression and determining the incoming Mach number of the inlet; constructing an inner conical axisymmetric reference flow field using a characteristic line design method, and matching the axisymmetric reference flow field with the shock wave of the conical waverider forebody to form the inlet capture profile; obtaining an inward-rotating inlet using a forward and reverse tracking design fusion approach; and fusing the inward-rotating inlet with the conical waverider forebody to obtain an integrated forebody / inlet configuration. This invention, after employing forward and reverse fusion design of the inlet, further integrates it with the forebody. Compared to existing independent forebody and inlet designs, this method offers advantages such as improved inlet pre-compression performance, shortened aircraft forebody length, increased total pressure recovery and pressure ratio of the inlet, and enhanced inlet start-up performance.

Description

Technical Field

[0001] This invention relates to the field of forebody / inlet design for high-speed aircraft. More specifically, this invention relates to an integrated forebody / inlet design method based on longitudinal segmented multi-stage compression. Background Technology

[0002] Air-breathing high-speed aircraft are powered by scramjet engines, enabling them to obtain oxygen from the air and thus eliminating the need for additional oxidizers during flight, allowing for high-speed, long-range flight. The forebody of an air-breathing high-speed aircraft is tightly integrated with the air intake, providing compressed airflow. The compression performance of the forebody plays a decisive role in the performance of both the air intake and the scramjet engine. Waveriders, as a promising aerodynamic shape for high-speed aircraft, can pre-compress the high-speed incoming airflow, providing a high-pressure, uniform airflow to the air intake.

[0003] To improve the pre-compression performance of the forebody, the compression performance of the waverider is closely related to the shock angle design. If single-stage compression is used, a larger shock angle is usually required to meet the pressure ratio requirement of the inlet airflow. However, a larger shock wave will lead to a larger pitching moment and will also reduce the lift-to-drag ratio of the forebody. Using isentropic compression usually results in a very long forebody, which is detrimental to the starting performance of the inlet.

[0004] In the prior art, for example, a longitudinal segmented and graded compression design method for a conical waverider forebody, patent number 202210304326.8, proposes a multi-stage compression technology that can realize longitudinal segmentation. It can change the number of compression stages, the compression intensity of different stages, and the compression length of different stages according to the performance requirements of the inlet, which is very suitable for the forebody design of hypersonic vehicles. Another example is a longitudinal segmented and multi-stage compression design method for a kissing cone hypersonic waverider forebody, patent application number 202210304413.3. This design method can change the number of compression stages, the compression intensity of different stages, and the compression length of different stages according to the performance requirements of the inlet, which has great application potential in the forebody design of hypersonic vehicles. However, both of the aforementioned technologies focus on the design of the forebody and do not integrate the forebody with the air intake. The integration of the forebody and air intake involves matching shock wave profiles, selecting airflow rates, and choosing capture profiles, and is closely related to the aircraft's cruise state. Poor integration may prevent the full utilization of the advantages of a multi-stage compression forebody, and could even lead to worse air intake performance. Furthermore, a simple multi-stage compression forebody cannot adequately assess the integrated design performance of a high-speed aircraft. Therefore, how to conduct air intake matching design for a specific multi-stage compression forebody is the core issue in improving the air intake performance of high-speed aircraft. Only by achieving an integrated design that couples the multi-stage compression forebody with the internal rotating air intake can the pre-compression performance of the forebody be better utilized, thereby improving the overall performance of the aircraft. Summary of the Invention

[0005] One object of the present invention is to solve at least the above-mentioned problems and / or defects, and to provide at least the advantages described below.

[0006] To achieve these objectives and other advantages of the present invention, an integrated forebody / inlet design method based on longitudinal segmented multi-stage compression is provided, comprising:

[0007] S1. Based on the design method of conical waverider forebody, complete the design of conical waverider forebody with at least two stages of compression, and extract the backflow field parameters of the conical waverider forebody through CFD calculation.

[0008] S2. Determine the incoming Mach number of the air intake based on the backflow field parameters;

[0009] S3. Based on a determined incoming Mach number, an internally rotating axisymmetric reference flow field is constructed using the characteristic line design method.

[0010] S4. Match the axisymmetric reference flow field with the cone-guided wave-riding forebody shock wave to form the intake trap profile;

[0011] S5. Based on the determined intake duct capture profile, an internal rotating intake duct is obtained by combining forward and reverse tracking designs.

[0012] S6. The inward rotating intake is fused with the multi-stage compressed cone-guided wave-riding forebody to obtain the forebody intake, and the forebody intake is locally modified.

[0013] Preferably, in S3, the axisymmetric reference flow field is constructed using the theory of rotatable characteristic lines under high-speed inflow, wherein the inviscid, adiabatic two-dimensional supersonic isentropic flow needs to satisfy the following governing equations:

[0014] Equations of characteristic lines along streamlines:

[0015]

[0016] In the above formula, Indicates the slope of the characteristic line. Represents the velocity in the y-direction. u This represents the velocity in the x-direction, where x represents the axial coordinate and y represents the longitudinal coordinate.

[0017] Compatibility equation along streamlines:

[0018]

[0019]

[0020] In the above formula, ρ , p , a , VThese represent density, pressure, sound velocity, and flow velocity, respectively. Indicates the slope of the characteristic line;

[0021] The equations of the characteristic lines along the left-hand and right-hand Mach lines are:

[0022]

[0023] In the above formula, Indicates the slope of the Mach lines moving to the left and right. θ This represents the angle of the local streamline relative to the horizontal line, where 'a' represents the Mach angle.

[0024] The compatibility equations along the left-hand and right-hand Mach lines are:

[0025]

[0026] Where M represents the Mach number, The subscript ± indicates the flow type, and the subscript ± indicates the parameter along the left or right Mach line.

[0027] Preferably, in S4, the inlet capture profile is determined based on the intersection line of the multi-stage compression precursor and the symmetrical reference flow field.

[0028] Preferably, in the S5, the design process for the internal rotary intake includes:

[0029] S50. Based on the intake duct capture profile, the internal rotating forward intake duct profile based on forward design is obtained by using streamline tracing.

[0030] S51. Based on the exit profile design requirements of the internal rotating intake, a blended reference streamline is formed by reverse tracking, and the internal rotating reverse design flow channel is obtained by adopting the reverse tracking design method.

[0031] S52. The internal rotating forward air intake and the internal rotating reverse design flow channel are integrated to form an internal rotating air intake.

[0032] Preferably, in S6, the partial modification of the forebody air intake includes:

[0033] The intake lip and the leading edge of the forebody are passivated.

[0034] The design of the air intake and the fuselage is achieved by using spline curves.

[0035] Preferably, in S6, before fusing the internal rotating intake with the multi-stage compression cone-guided waverider forebody, the following formula is used to determine whether the inlet strength of the internal rotating intake and the multi-stage compression performance of the forebody meet the requirements:

[0036]

[0037] In the above formula, This refers to the pressure at the intake of the air duct after multi-stage compression by the cone-guided waverider forebody. This refers to the Mach number at the inlet of the air intake after multi-stage compression by the cone-guided waverider forebody. This represents the empirical relationship between pressure before and after the compression shock wave, and , To compress the empirical relationship of Mach number before and after the shock wave, and , Indicates the incoming flow pressure. The Mach number of the incoming stream is represented. , , and For constant parameters, i The value is the compression level;

[0038] When the inlet strength of the intake duct and the multi-stage compression performance of the forebody meet the requirements, the internal rotating intake duct configuration and the multi-stage compression cone-guided waveriding forebody are integrated through the cross lines on the lower surface to obtain an integrated forebody intake duct that meets the requirements.

[0039] The present invention has at least the following beneficial effects:

[0040] Firstly, the integrated design method for the forebody and inlet duct proposed in this invention utilizes a multi-stage compression waverider forebody to reduce the Mach number of the incoming flow into the inlet. This results in better total pressure recovery performance for the inlet under the constraint of a low Mach number inlet and the same outlet Mach number. Through multi-stage compression of the forebody, the airflow entering the inlet can adapt to a wider range of flow and angle of attack, and also has potential value in low-Mach number start-up. The multi-stage compression forebody can reduce the forebody and inlet length at the same pressure ratio, avoiding the increased boundary layer thickness and low-energy flow affecting inlet start-up caused by an excessively long forebody, thus improving the compression efficiency of the inlet and providing a new design method for inlet duct optimization.

[0041] Secondly, the integrated design method of the forebody air intake proposed in this invention, after adopting forward and reverse fusion design of the air intake, further integrates it with the forebody. Compared with the existing independent design of the forebody and air intake, it has advantages such as improving the pre-compression performance of the air intake, shortening the length of the aircraft forebody, improving the total pressure recovery and pressure ratio of the air intake, and broadening the start-up performance of the air intake.

[0042] Other advantages, objectives and features of the present invention will become apparent in part from the following description, and in part from those skilled in the art through study and practice of the invention. Attached Figure Description

[0043] Figure 1 is a schematic diagram of the design of the infeeding cone waverider front body in the prior art;

[0044] Figure 2 is a schematic diagram of the geometric relationship after two-stage compression in the prior art;

[0045] Figure 3 is a schematic diagram of the precursor multi-stage compression design of the present invention;

[0046] Figure 4 shows the pressure distribution within different cross sections of the secondary compression of the cone-guided waverider precursor of the present invention.

[0047] Figure 5 is a pressure cloud distribution diagram of the second-stage compression symmetry plane of the conical waverider precursor of the present invention.

[0048] Figure 6 shows the three-dimensional internal rotating air intake of the present invention;

[0049] Figure 7 is a schematic diagram of the integrated design of the multi-stage compression precursor and the intake duct of the present invention;

[0050] Figure 8 is a schematic diagram of the intake trap profile and inner cone flow field of the present invention;

[0051] Figure 9 is a schematic diagram of the throat outlet profile of the present invention;

[0052] Figure 10 is a schematic diagram of the integrated design configuration of the multi-stage compression precursor intake duct of the present invention.

[0053] Figure 11 is a grid diagram of the integrated structure of the forebody air intake of the present invention;

[0054] Figure 12 Pressure cloud diagrams of different longitudinal sections for the integrated design of the forebody air intake of this invention;

[0055] Figure 13 is a pressure cloud diagram of the symmetry plane of the present invention;

[0056] Figure 14 is a density gradient cloud map of the symmetry plane of the present invention;

[0057] Figure 15 is a schematic diagram of the density gradient of different cross sections along the wingspan direction of the present invention;

[0058] Figure 16 is a schematic diagram of the surface pressure at Mach number 6.0 and angle of attack 0 degrees according to the present invention (viscosity calculation results);

[0059] Figure 17 is a schematic diagram of the throat Mach number (non-adhesive) of the present invention;

[0060] Figure 18 is a schematic diagram of the throat Mach number (viscosity) of the present invention;

[0061] Figure 19 is a schematic diagram of the total pressure recovery coefficient of the larynx (non-adhesive) of the present invention;

[0062] Figure 20 is a schematic diagram of the total pressure recovery coefficient (viscosity) of the larynx according to the present invention;

[0063] Figure 21 is a schematic diagram of the throat pressure ratio (non-adhesive) of the present invention;

[0064] Figure 22 is a schematic diagram of the throat pressure ratio (viscosity) of the present invention. Detailed Implementation

[0065] The present invention will now be described in further detail with reference to the accompanying drawings, so that those skilled in the art can implement it based on the description.

[0066] To further improve the compressibility of the waverider forebody and alleviate the compression requirements of the inlet, a longitudinally segmented multi-stage compression integrated design method for the forebody / inlet is proposed. The high-pressure zone of the forebody is mainly concentrated in the middle of the inlet, avoiding edge pressure leakage and achieving design goals such as shortened forebody length, increased lift-to-drag ratio, and high total pressure recovery in the inlet. An integrated forebody / inlet design based on two-stage cone-guided compression was completed. The aerodynamic performance, flow field distribution, and total pressure distribution of the forebody / inlet were analyzed. The analysis results show that this method can effectively improve the inlet compression efficiency and enhance the flow field quality at the inlet outlet. This method has significant application potential in the forebody design of high-speed aircraft. It has not yet been seen in publicly available literature both domestically and internationally, demonstrating high innovation and promising engineering applications.

[0067] An integrated design method for the forebody / inlet of an air-breathing high-speed aircraft, comprising:

[0068] S1. Based on the cone-guided waverider forebody design method, complete the design of the two-stage cone-guided waverider forebody. Extract the backflow field parameters of the two-stage cone-guided waverider forebody through CFD calculation. In this step, the multi-stage compression forebody design is mainly completed based on the segmented proportional adjustable design method.

[0069] S2. Based on the backflow field parameters of S1, determine the Mach number of the inlet flow.

[0070] S3. Under high-speed incoming flow, based on the aforementioned Mach number of the inlet flow, an internal conical axisymmetric reference flow field is constructed using a swirling characteristic line design method.

[0071] S4. The axisymmetric reference flow field is matched with the shock wave of the three-stage conical waverider forebody. The inlet capture profile is formed by merging the waverider flow field and the inner cone flow field. That is, the multi-stage compression forebody and the axisymmetric reference flow field intersect to form the inlet capture profile.

[0072] S5. Based on the determined intake duct capture profile, an internal rotating intake duct profile is formed by streamline tracing. Then, based on the intake duct outlet profile, a reverse internal flow channel is formed by reverse streamline tracing. The intake duct obtained by forward design and the flow channel obtained by reverse design are merged to form the required intake duct.

[0073] This step integrates the design of the inward-turning inlet captured from the leading edge profile with the reverse inward-turning flow channel, thereby achieving an integrated inlet layout that satisfies the design of the inlet and the multi-stage compression waverider forebody.

[0074] S6, the forebody inlet configuration obtained by fusing the internal rotating inlet with the conical waverider forebody, requires selecting inlet design parameters based on the aircraft's cruise state. This necessitates determining the rotational parameters of the reference flow field by considering the spanwise velocity distribution after multi-stage compression of the conical waverider forebody. Furthermore, the inlet contraction ratio needs to be determined based on the intensity of multi-stage compression of the conical waverider forebody to maximize inlet performance and ensure optimal overall performance of parameters such as inlet flow rate, total pressure recovery, and pressure ratio. Additionally, to mitigate the inability to start due to a thick boundary layer caused by an excessively long conical waverider forebody, which leads to low kinetic energy and easy flow separation in the inlet, the following formulas for calculating the inlet strength of the internal rotating inlet and the multi-stage compression performance of the forebody are constructed:

[0075]

[0076] In the above formula, This refers to the pressure at the intake of the air duct after multi-stage compression by the cone-guided waverider forebody. It refers to the Mach number at the intake of the air duct after multi-stage compression by the cone-guided wavefront forebody. This represents the empirical relationship between pressure before and after the compression shock wave. The specific coefficients are fitted based on a specific configuration and shock wave angle design. This represents the empirical relationship between the Mach number before and after the compression shock wave, with specific coefficients fitted according to the specific configuration and shock wave angle design. Ma represents the incoming flow pressure, and Ma0 represents the incoming flow Mach number. , , and These are constant parameters, determined based on the specific configuration.

[0077] Under the same incoming Mach number and the same initial shock angle, the exit Mach number of the lower surface of the conical waverider forebody is less than that of the conical waverider body, which can reduce the design incoming Mach number of the inlet. Based on the design parameters of the multi-stage compression forebody, when the incoming Mach number is 6, the exit Mach number and pressure ratio of the two-stage compression conical waverider forebody are 4.5 and 5.0, respectively, while the corresponding exit Mach numbers of the single-stage compression conical waverider forebody are 5.1 and 1.9. The integrated design method of the multi-stage compression conical waverider forebody inlet can effectively improve the pre-compression performance of the conical waverider forebody, reduce the shock wave intensity inside 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 conical waverider forebody is 5.0, the length of the second-stage compression conical waverider forebody can be reduced by 30% compared to the isentropic compression length. On the one hand, shortening the length of the conical waverider forebody reduces the low-energy flow entering the air intake and improves the start-up capability of the air intake. On the other hand, shortening the conical waverider forebody helps to reduce the impact on the directional stability of the aircraft's vertical tail at high angles of attack.

[0078] Analysis shows that, while maintaining the same Mach number at the exit of the conical waverider front, the volume ratio of the single-stage compressed conical waverider front is 43% higher than that of the two-stage compressed conical waverider front, resulting in a 20.4% decrease in the lift-to-drag ratio of the conical waverider front.

[0079] When integrating the forebody inlet, the Mach number and flow direction of the inlet after compression by the cone-guided waverider forebody are determined based on the intensity of multi-stage compression. The reference inner cone flow field design is completed. Then, the rotation angle of the inner rotating inlet is selected. The rotated inner rotating inlet (scaled proportionally according to the intake flow requirements) and the cone-guided waverider forebody are intersected by shock waves to determine the capture profile. Based on the obtained capture profile, the configuration of the inner rotating inlet is obtained by streamline tracing. Then, the configuration of the inner rotating inlet and the multi-stage compressed cone-guided waverider forebody are integrated through the intersection line on the lower surface to obtain an integrated forebody inlet configuration that meets the requirements.

[0080] The forebody air intake is partially modified, including: passivating the air intake lip and the leading edge of the forebody to meet the requirements of processing and structural load-bearing; and modifying the air intake and the fuselage through spline curve design to meet certain volume requirements and form an aerodynamic layout with certain application capabilities.

[0081] Example:

[0082] An integrated design method for the forebody / inlet of an air-breathing high-speed aircraft is presented. Based on a waverider layout, it integrates the design by coupling a multi-stage compression waverider forebody with the flow field of a three-dimensional internally rotating inlet. This reduces aerodynamic interference between internal and external flows, mitigates the compression requirements within the inlet, and lowers drag, which is beneficial for improving the aircraft's lift-to-drag ratio and inlet flow capture performance. The three-dimensional internally rotating inlet utilizes a contracting shock wave, resulting in relatively small total pressure loss and minimal air overflow. Compared to two-dimensional compression and side-pressure inlets, it offers better compression capacity and efficiency. The design method mainly includes:

[0083] I. Multi-stage compression precursor design

[0084] Based on the longitudinal segmented multi-stage compression forebody design method of the conical waverider, forebody design is also a prerequisite for integrated internal and external flow design. The number of compression stages of the forebody can be designed in segments along the longitudinal direction. The forebody design mainly includes the upper surface of the waverider (it should be noted that the upper surface can be free flow or not, with free flow being the typical design state), the first-stage compression surface, the second-stage compression surface, etc. (The lower surface of the forebody is divided into second-stage compression, third-stage compression, and more stages of compression according to design requirements. The lengths of the second-stage and third-stage compression on the lower surface of the forebody can be combined in various proportions. Through the multi-stage compression of the forebody, the design Mach number of the inlet flow of the inlet can be adjusted). The upper surface of the waverider forebody 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 exit curve (ICC, Inlet Capture Curve) and upper surface exit curve (Flow Capture Curve), according to the flight Mach number Ma and the first-stage compression shock angle. The compressibility strength is determined (the Mach number and shock angle within different cut surfaces can be designed as needed while ensuring continuous variation). The corresponding lower surface profile and leading edge curve are obtained using the streamline tracing method. The upper surface of the waverider is then obtained using the free-flow surface method based on the leading edge curve. This implementation case uses a two-stage compression precursor as an example.

[0085] Since the cone-guided waverider front is the design basis for multi-stage compressed cone-guided waverider fronts, the design principle of the cone-guided waverider front is explained here, such as... Figure 1 As shown ( Figure 1In this diagram, COC represents the center of curvature of the shock radius, 1 represents the trailing-edge curve of the lower surface, and 2 represents the radius of shock. By designing the shock angle given the flight Mach number, the basic flow field of the zero angle-of-attack cone can be obtained. This can be achieved by specifying the upper surface curve (FCC) of the waverider's trailing edge, or the lower surface curve (ICC) of the trailing edge. Taking FCC as an example, the upper surface streamlines are obtained by tracing the FCC in reverse, parallel to the incoming flow. The intersection of the upper surface streamlines with the shock surface yields the leading edge line. The leading edge line is then used to trace the lower surface streamlines of the waverider, thus obtaining the conical waverider forebody. The main design parameters include the expansion angle. Shock angle β Mach number Ma and distance relative to the axis R 0, etc. The equation for the trailing edge curve of the upper surface is typically shown below, where, R 0 represents the distance between the symmetrical position of the trailing edge curve and the axis. It represents the expansion angle of the curve. This represents the tangent angle of the curve at its endpoints. Given parameters... R 0、 , With these three parameters, the equation of the curve can be obtained.

[0086]

[0087] In the above formula, a f Denotes the coefficient of the quadratic term. b f The coefficients of the fourth-order term are represented by x and y, which represent the horizontal and vertical coordinates, respectively, in meters.

[0088] Furthermore, such as Figure 2 and Figure 3 As shown, the leading edge curve of the multi-stage compression waverider is divided into the first-stage leading edge line and the second-stage edge line according to the design requirements, so that the longitudinal lengths of the two segments of the waverider can be obtained.

[0089] For a second-order waverider, to ensure that the first-order and second-order compression shock waves coincide perfectly at the exit, the streamlines behind the conical flow shock wave are curved, and the Mach number and direction are different at each point on the lower surface of the first-order compression. Therefore, to ensure that the first-order and second-order shock waves coincide at the exit, the angle of the second-order compression shock wave corresponding to each first-order node is different. For example... Figure 2 The left side shows a schematic diagram of the compression of a second-order waverider body in a cone-guided system. Figure 2In the diagram, 3 indicates first-stage compression and 4 indicates second-stage compression. The leading edge of the waverider is obtained based on the axisymmetric reference flow field. The position corresponding to the first-stage compression is determined according to the longitudinal length ratio. For example, in diagram 2, the upper red line EAF is the leading edge of the first-stage compression, and the middle green line ELF is the lower surface curve of the first-stage compression outlet, which is also part of the leading edge of the second-stage compression. The lower green lines EP and FQ on both sides are the leading edge lines on both sides of the second-stage compression. The corresponding lower surfaces of the second-stage compression are obtained by streamline tracing. The red line PG on the top left is the trailing edge line obtained by EP after second-stage compression, the red line GMH in the middle of the top is the trailing edge line obtained by ELF after second-stage compression, and the red line HQ on the top right is the trailing edge line obtained by FQ after second-stage compression. To better illustrate the geometric relationships that the local cross-section needs to satisfy, Figure 3 ( l 1 Ma1 represents the first-stage compression length, point M represents the outlet point of the lower surface streamline, and Ma1 represents the first-stage compression length. Ma1 represents the incoming Mach number, and Ma2 represents the Mach number after first-order compression. (This indicates the direction of the gas velocity after the first stage of compression) gives Figure 2 The local cross section where OO′-BC is located is described, and the flow field needs to satisfy the following relationship:

[0090]

[0091] In the above formula, β 2 represents the second-order compression shock angle. β 1 For first-order compression shock angle, d 1 This is the distance from the local horizontal axis. d A Let A be the distance from the horizontal axis. l 2 represents the second-level compression length. The angle between the primary compression outlet velocity and the local horizontal axis;

[0092] Furthermore, the first-stage compression surface design starts from the leading edge line and uses streamline tracing to obtain the streamlines of the lower surface of the first stage. With the longitudinal length of the first stage as a constraint, the obtained streamlines are truncated so that they are extracted along the two edges of the leading edge line corresponding to the longitudinal coordinate not greater than L1. All flow field parameters of the first-stage streamline cutoff point are extracted.

[0093] Furthermore, the streamline cutoff point of the first-stage compression surface, which is also the outlet of the first-stage compression surface, constitutes the internal leading edge line of the second-stage compression. The outlet line of the first-stage compression surface and the edge line of the second stage together constitute the leading edge line of the second-stage compression. The streamlines of the lower surface corresponding to the edge line of the second stage are obtained by streamline tracing using an axisymmetric reference flow field. The flow field parameters corresponding to the outlet line of the first compression are equivalent to the incoming flow parameters of the internal leading edge line of the second stage. The Mach number, velocity, direction, and pressure at each characteristic point are different. It is necessary to re-establish the corresponding axisymmetric reference flow field center, incoming flow direction, Mach number, and shock angle based on the geometry of the second-stage compression. The Mach number, velocity direction, and relative position corresponding to the leading edge of the second-stage compression are different. In order to ensure that the shock wave of the second-stage compression matches the shock wave of the first stage at the inlet of the air intake, it is necessary to find the relationship between the lengths of the first and second stages and the shock angle corresponding to the second-stage compression in the two-dimensional cross-sections corresponding to different expansion angle positions.

[0094] Furthermore, when each expansion angle corresponds to the second-order compression shock angle β Once step 2 is determined, a new conical flow field can be constructed using the Mach number, shock angle, and velocity direction at the leading edge. This allows for the calculation of the lower surface streamlines within the corresponding cross-section at different expansion angles. The streamlines obtained from all second-stage leading edge points collectively constitute the lower surface of the second-stage compression. The streamlines obtained from the first-stage compression are used to loft the first-stage compression surface, and the streamlines from the second-stage compression can be used to obtain the second-stage compression surface, thus obtaining the entire waverider configuration. This process can be repeated to design three-stage compression or even more stages of forebody compression.

[0095] Thus, the design of the two-stage compression precursor was completed. The lower surface of the waverider was divided into multiple stages along the longitudinal direction according to requirements. To ensure that the multi-stage compression shock waves in each section converge at the inlet of the air duct, a strict geometric relationship between the segment length and the corresponding shock wave angle was established. Simultaneously, the edge line remained identical to that of the single-stage scissor-cone waverider during the multi-stage compression design process, ensuring that the designed waverider precursor strictly satisfies the waveriding characteristics at the edge, with no pressure leakage. This fully integrates the advantages of waveriding characteristics and multi-stage compression. The numerical analysis results are as follows: Figure 4 and Figure 5 As shown in the figure (P represents the local pressure). (This represents the incoming flow pressure). Furthermore, this method can also alter the shock wave angle at different locations to change the inlet pressure distribution of the inlet. The multi-stage compression precursor obtained using this method has high pre-compression capability, providing a high-pressure, low-speed, and relatively uniform airflow to the inlet. The inlet outlet profile is also flexible and adjustable, facilitating matching with the inlet. Simultaneously, the precursor maintains a high lift-to-drag ratio, offering significant advantages compared to traditional single-stage compression or isentropic compression.

[0096] II. Inlet Design Based on Multi-Stage Compression Precursor

[0097] The design of a three-dimensional internal rotating intake is actually a "reverse design" method. First, an internal rotating reference flow field is constructed based on the internal rotating conical flow field. Then, the inlet curve is obtained according to the intake capture shape. Streamlines based on the leading edge curve are generated in the internal rotating reference flow field through streamline tracing technology. All streamlines close to form the internal rotating intake profile.

[0098] First, an axisymmetric reference flow field is constructed using the theory of swirling characteristic lines. Under high-speed incoming flow, the axisymmetric reference flow field is constructed again using the theory of swirling characteristic lines. The intersection of the waverider's lower surface and the forebody shock surface with the shock surface of the axisymmetric reference flow field forms the capture profile of the inlet. Then, different streamlines are obtained through streamline tracing, and the flow surface is formed by lofting to construct the inward-rotating inlet profile. Since the solution of the axisymmetric reference flow field uses the theory of characteristic lines, which is only applicable in supersonic flow fields, a small wedge is designed in the initial section of the generator to produce an attached shock wave when constructing the axisymmetric reference flow field.

[0099] Solving the axisymmetric reference flow field utilizes the theory of characteristics, where the normal derivative of the characteristics may be discontinuous. In the flow field, streamlines and Mach lines are weakly discontinuous lines along the direction of the characteristics, and the flow field parameters satisfy compatibility relations. Inviscid, adiabatic two-dimensional supersonic isentropic flow satisfies the following governing equations.

[0100] Equations of characteristic lines along streamlines:

[0101]

[0102] In the above formula, Indicates the slope of the characteristic line. Represents the velocity in the y-direction. u This represents the velocity in the x-direction, where x represents the axial coordinate and y represents the longitudinal coordinate.

[0103] Compatibility equation along streamlines:

[0104]

[0105]

[0106] In the above formula, ρ , p , a , V These represent density, pressure, sound velocity, and flow velocity, respectively. Indicates the slope of the characteristic line;

[0107] The equations of the characteristic lines along the left-hand and right-hand Mach lines are:

[0108]

[0109] In the above formula, Indicates the slope of the Mach lines moving to the left and right. θ This represents the angle of the local streamline relative to the horizontal line, where 'a' represents the Mach angle.

[0110] The compatibility equations along the left-hand and right-hand Mach lines are:

[0111]

[0112] Where M represents the Mach number, The subscript ± indicates the flow type, and the subscript ± indicates the parameter along the left or right Mach line.

[0113] The above equations are rewritten as finite difference equations and then solved iteratively. They are mainly categorized into several forms: interior point elements, direct wall point elements, inverted wall point elements, and shock wave boundary point elements. The axisymmetric reference flow field is solved using characteristic line theory. When constructing the axisymmetric reference flow field, a small wedge is designed in the initial section of the generator to produce attached shock waves. Based on the inlet capture profile, the internal rotating inlet can be obtained through streamline tracing, such as... Figure 6 As shown.

[0114] III. Front Intake Channel Integration Design

[0115] When constructing the intersection line of the leading edge of the air intake, the inward contraction shock wave is usually rotated at a certain angle. For example... Figure 7 As shown, the inner cone shock wave 5 is first rotated at a certain angle of attack to intersect with the reference shock wave 6 and the compression surface 7 (i.e., the lower surface of the waverider). Then, the entire structure is rotated at the same angle of attack. Finally, the intersection curve is extracted as the capture profile 8 (e.g., ...). Figure 8 (As shown by the red lines in the diagram), the air intake profile is obtained through streamline tracing. The designed air intake is then docked and integrated with the front body. Finally, the overall shape is rotated at a positive angle of attack. To ensure the sealing of the capture profile, the lower half of the capture profile is modified so that a section of the capture profile of the air intake is located above the reference center body 9 of the inward rotating air intake.

[0116] Because the intake outlet section obtained by streamline tracing the capture profile generated by the intersection of the inner cone shock wave and the waverider is irregular, the matching and fusion effect with the isolation section is poor, which may lead to a decrease in flow field uniformity. Therefore, given the outlet profile, the inner flow channel is regenerated using reverse streamline tracing technology. The intake captured by the leading edge profile is then integrated with the reverse-traced inner flow channel in the design. The throat outlet profile is as follows: Figure 9 As shown ( Figure 9 The red line with a circle indicates the forward design outlet, and the blue line with a triangle indicates the given outlet profile. This achieves the goal of integrating the intake and the waverider body to ensure airflow capture, while also meeting the design requirements of the isolation section.

[0117] In this embodiment, the combination of a multi-stage compression precursor and an internal rotary intake can further improve the total pressure recovery performance and pressure ratio of the intake.

[0118] Example 2:

[0119] First, an integrated design scheme for the forebody / inlet was completed based on a two-stage compression forebody and an internal rotating inlet. The first step was to complete the design of the multi-stage compression forebody. Given a cruise Mach number of 6.0 and a reference shock wave with a design shock wave angle of 12°, and given the upper surface curve FCC of the trailing edge of the waverider, the forebody design was completed using a two-stage compression waverider design method.

[0120] Secondly, the Mach number and streamline direction of the inlet flow are determined based on the flow field after the second stage compression of the forebody. Here, in the design of the reference inner cone flow field, the ratio of the central body radius to the inlet shock wave radius is 0.2. The inlet Mach number for the inward-rotating reference flow field is designed to be 4.5, and the outlet Mach number is 2.57. Since the inward-rotating inlet is designed based on a uniform inlet flow, the direction of the compressed airflow changes after the inner cone shock wave intersects with the forebody shock wave. When constructing the intersection line of the inlet leading edge, the inward-rotating reference flow field is usually rotated by a certain angle. Simultaneously, it is ensured that the intersection point of the inlet's capture profile is located above the reference central body of the inward-rotating inlet. Considering the design characteristics of a multi-stage compression waverider, discontinuous points are generated when the inner cone shock wave intersects with the lower surface of the waverider forebody. In this case, the intersection line is projected onto a plane perpendicular to the inlet flow, and then the projected line is discretized and fitted using splines to obtain the inlet capture profile. The forebody inlet obtained from the case design is shown below. Figure 10 As shown.

[0121] II. Verification of the design method through numerical simulation. The computational domain is discretized using a structured mesh (see...). Figure 11 The incoming Mach number is 6.0, the incoming static pressure is 1197.03 Pa, and the incoming temperature is 226.509 K. Under non-viscosity conditions, the exit Mach numbers at 0° and 2° angles of attack are 2.743 and 2.59, respectively. Since the precursor design Mach number is 4.5, the calculated precursor Mach number is slightly greater than 4.5, thus slightly increasing the exit Mach number, which basically meets the design requirements.

[0122] The results of the non-viscosity calculation are shown in Table 1. At an angle of attack of 0°, the total pressure recovery coefficient of the intake throat is 0.831 and the pressure rise ratio is 54.28. At an angle of attack of 2°, the total pressure recovery coefficient of the throat is 0.7873 and the pressure rise ratio is 70.23.

[0123] Table 1. Calculation results for non-viscous properties

[0124]

[0125] Table 2 shows that, under viscous conditions, the total pressure recovery coefficient of the throat at an angle of attack of 0° is slightly lower than that at the non-viscous design point. The Mach number changes from 2.743 to 2.3215, a decrease of 15.37%, the total pressure recovery coefficient changes from 0.831 to 0.686, a decrease of 17.45%, and the pressure rise ratio changes from 54.28 to 74.98, an increase of 38.14%.

[0126] Table 2 Results of viscosity calculation

[0127]

[0128] like Figure 12 and Figure 16 As shown, the pressure coefficient distribution of the waverider forebody along different longitudinal sections is presented. It can be seen that the waverider edge experiences slight pressure leakage due to passivation and viscosity calculations, while the overall high-pressure zone is located on the lower surface of the aircraft. Analysis of the pressure cloud distribution on the lower surface reveals a clear hierarchical compression on the forebody compression surface, which is largely consistent with the design theory of multi-stage compression forebody. Slight pressure leakage occurs at the inlet lip, primarily because the inlet design is based on the inviscid characteristic line theory, and the incoming Mach number of the basic flow field is an average of the wave-after Mach number. When the multi-stage compression forebody is integrated with the internal rotating inlet design, interference occurs when matching the designed inlet and waverider due to the influence of the capture profile. This may be related to the incoming Mach number selected during inlet design. Additionally, the gradient change in the compression surface of the multi-stage compression forebody's lower compression degree results in a lower compression degree in the inlet compared to the forebody.

[0129] To further analyze the shock wave distribution at the lip of the symmetry plane, Figure 13 The pressure ratio distribution on the symmetrical plane is given. It can be seen that the first and second stage compression shock waves of the forebody basically hit the inlet lip, and the pressure ratio at the inlet reaches about 5. Figure 14 The diagram shows the density gradient change on the symmetry plane. It also reveals two compression shock waves in the multi-stage compression precursor. The reflected shock wave at the lip strikes the upper surface grate, consistent with the designed flow field. Two high-density dividing surfaces appear in the compression region of the multi-stage compression waverider precursor, consistent with the designed multi-stage waverider precursor. In the first-stage compression region, a high-density gradient appears only near the waverider wall. In the second-stage compression region, two high-density regions appear, and the two compression stages converge precisely at the inlet lip. This ensures that the entire high-pressure flow field of the multi-stage compression precursor enters the inlet, achieving efficient capture of the multi-stage compression flow field and improving inlet performance.

[0130] Figure 15The density gradient variations along the spanwise direction of the waverider are presented for different cross-sections. Near the inlet, both compression shock waves of the forebody strike the lip position, indicating that the designed three-dimensional internal rotating inlet and the multi-stage compression forebody shock waves are well matched. The shock wave distribution at the symmetry plane lip reveals two compression shock waves in the multi-stage compression forebody, with the reflected shock wave at the lip striking the upper surface throat position precisely, consistent with the designed flow field. Figures 17-22 The performance of a multi-stage compression forebody inlet is presented over a wide Mach number range. It can be seen that the integrated design of the inlet based on the multi-stage compression forebody can significantly improve the inlet performance, broaden the inlet start-up range, increase the throat pressure rise ratio and total pressure recovery coefficient, and has a significant benefit on improving aircraft performance.

[0131] The above solution is merely an illustration of a preferred example, but is not limited thereto. When implementing this invention, appropriate substitutions and modifications can be made according to the user's needs.

[0132] This invention is not limited to the specific embodiments described above. The multi-stage compression of the cone-guided waverider forebody can be extended to three-stage compression or even more-stage compression forebody designs. The multi-stage compression waverider forebody can reduce the incoming Mach number of the inlet flow. Under the constraint of a low Mach number inlet and the same outlet Mach number, the inlet exhibits better total pressure recovery performance. The integrated design of the multi-stage compression forebody / inlet allows for a reduction in forebody and inlet length at the same pressure ratio, avoiding the increased boundary layer thickness and the impact of low energy flow on inlet start-up caused by an excessively long forebody. The combination of multi-stage shock wave compression of the forebody and Mach wave compression in the inlet has potential application value in improving inlet compression performance and total pressure recovery performance.

[0133] Although embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the specification and embodiments. It is entirely applicable to various fields suitable for the present invention. Other modifications can be readily implemented by those skilled in the art. Therefore, without departing from the general concept defined by the claims and their equivalents, the present invention is not limited to the specific details and examples shown and described herein.

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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