A flow field reconstruction design method under the control of internal flow channel strong wave series interference

Abstract

The present application relates to a kind of flow field reconstruction design methods under the control of internal flow channel strong wave series interference, improve the aerodynamic heat distribution of inlet duct, overcome the limitation that characteristic line method cannot realize viscous calculation, highlight the advantage of CFD method to carry out fine design, can provide design support for internal flow channel aerodynamic heat optimization.The present application proposes and realizes the shock control design scheme of internal rotating inlet duct, compared with simple configuration inlet duct, shock control design can reasonably allocate overall shock and local shock distribution, reduce the intensity of shock interference under high Mach number condition, with good aerodynamic performance and comprehensive performance of heat reduction.

Description

Technical Field

[0001] This invention belongs to the field of combined propulsion aircraft technology and relates to a flow field reconstruction design method under strong wave interference control in the internal flow channel. Background Technology

[0002] For combined-propellant aircraft, the air intake system is one of the key components. Currently, air intake systems are mainly classified into two-dimensional planar compression, axisymmetric compression, three-dimensional side-pressure with top pressure, and various three-dimensional internal contraction layouts. These compression methods all utilize compression waves or shock waves generated by the compression surface to compress the airflow. How to rationally and efficiently organize and utilize shock waves or compression waves to compress the airflow has always been a goal pursued by people.

[0003] To achieve the thermal design goals for descent under wide-range flight conditions, local key geometries require heat reduction measures such as non-uniform leading-edge radius passivation, transition zones, and roughening zones. Therefore, the performance of the inlet design needs to be evaluated, and control measures to compensate for performance losses need to be sought. Currently, the main method for inlet performance evaluation is numerical simulation, which has high accuracy for shock wave interference and unsteady flow characteristics. However, research on inlet flow field control measures mainly focuses on variable structure approaches, involving heavy moving parts and large stroke variations. Therefore, the feasibility and rapid response capability in practical applications limit further application.

[0004] This project addresses the problems of strong shock wave / boundary layer interference and poor self-starting characteristics in geometrically constrained hypersonic inlets. It develops a flow field reconstruction design technology under the control of strong wave interference in the internal flow channel, which solves the problems of shock wave / boundary layer interference and weak starting capability in the internal flow channel of the inlet under geometrically constrained conditions. Summary of the Invention

[0005] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a flow field reconstruction design method under strong shock wave interference control in the internal flow channel, filling the technical gap in my country's design of shock wave control for aircraft inlet in the strong shock wave interference zone, and providing support for the aerodynamic and thermal protection of future high-speed aircraft and the aerodynamic and thermal environment analysis of strong shock wave interference.

[0006] The technical solution of this invention is:

[0007] This invention discloses a flow field reconstruction design method under strong wave interference control in an internal flow channel, comprising the following steps:

[0008] S1. Based on the aircraft design conditions and requirements, and using the two-dimensional surface design method of the forebody / inlet, important aerodynamic parameters that can control the inlet surface are proposed. The two-dimensional surface of the forebody / inlet is parametrically designed to obtain the basic two-dimensional surface model.

[0009] S2. Based on the basic two-dimensional surface model, establish a two-dimensional hypersonic three-dimensional curved surface compression precursor / inlet design model based on streamline tracing.

[0010] S3. Based on the binary hypersonic three-dimensional curved surface compression precursor / inlet design model, the multi-stage compression waverider configuration is obtained through the streamline tracing method.

[0011] S4. Import the multi-stage compression waverider configuration into the modeling software to generate a multi-stage compression waverider model;

[0012] S5. The aerodynamic characteristics of the multi-stage compression waverider model are obtained by using the flow field CFD numerical simulation method.

[0013] S6. Based on the multi-stage compression waverider model, the flow field control reconstruction method is used to calculate the pressure gradient at the shock wave incident point; the aerodynamic characteristics are adjusted until the pressure gradient at the shock wave incident point meets the requirements, and the adjusted multi-stage compression waverider model is obtained.

[0014] S7. Based on the regulated multi-stage compression waverider model, the interference type and flow field parameters of the incident shock wave are obtained using the polar curve method. The shock wave is regulated to obtain the aerodynamic characteristics of the regulated multi-stage compression waverider model. It is determined whether the aerodynamic characteristics meet the required values. If yes, the process exits; otherwise, the process proceeds to step S6.

[0015] In the above design method, the aircraft design conditions and requirements in step S1 include: flight altitude, flight Mach number, capture flow rate, flow coefficient, total contraction ratio, and total deflection angle of the external pressure section.

[0016] In the above design method, the two-dimensional surface design method for the forebody / intake duct in step S1 is specifically as follows:

[0017] (1) Determine the surface parameters of the external pressure section based on the geometric configuration of the aircraft forebody / inlet;

[0018] (2) Determine the inlet area of ​​the inner channel of the inner contraction section based on the profile parameters of the outer pressure section;

[0019] (3) Calculate the height and length of the isolation section.

[0020] In the above design method, the calculation of the height and length of the isolation section specifically involves:

[0021] The height of the isolation section is equal to the height of the throat.

[0022] The formula for the length of the isolation section is:

[0023]

[0024] Where x is the length of the isolation section, and M1 is the Mach number at the throat exit. Where θ is the momentum Reynolds number, D is the throat outlet diameter, θ1 is the throat outlet surface angle, P is the incoming static pressure, and P1 is the throat outlet pressure.

[0025] In the above design method, the length of the isolation section is 6 to 10 times the height of the isolation section.

[0026] In the above design method, the external pressure section profile parameters are specifically: initial compression angle, ratio of central body radius to reference flow field radius, dimensionless central radial distance Rc / Ri, capture profile aspect ratio, and throat section aspect ratio.

[0027] In the above design method, step S2 involves establishing a binary hypersonic three-dimensional curved surface compression forebody / inlet design model based on streamline tracing, specifically:

[0028] Determine the geometric control variables;

[0029] Based on the basic two-dimensional surface model, according to the three angles of the three-wave external pressure section and the incoming Mach number, the two-dimensional configuration inviscid flow field after flowing through three external pressure shock waves is obtained.

[0030] The upper surface of the aircraft forebody adopts a parallel inviscid flow field, while the lower surface of the aircraft adopts a two-dimensional inviscid flow field with an air intake configuration.

[0031] Different curves are selected on the cross section perpendicular to the flow direction at the air intake to perform streamline tracing and generate several different physical surfaces;

[0032] Based on several different physical surfaces, a three-dimensional curved surface compression precursor / intake design model is generated.

[0033] In the above design method, the geometric control variables include the front body width ratio, the front body angle, and the degree n of the control curve of the lower surface of the front body. down and the number of times the control curve of the upper surface of the precursor n up ;n up ≤n down .

[0034] In the above design method, the multi-stage compression waverider configuration is obtained in step S3 by the streamline tracing method. The specific method is as follows:

[0035] Step S31: Intersect a plane parallel to the cone axis with the first-stage shock surface to obtain a leading edge line, and discretize the leading edge line into a certain number of point sets; the number of points in the point sets n≥2;

[0036] Step S32: Starting from a point on the leading edge, perform streamline tracing in the first-stage conical flow field and generate an intersection point q1 with the second-stage shock wave surface. Obtain the airflow parameters of the intersection point q1 by solving the conical flow.

[0037] Step S33: Starting from the intersection point q1, use the airflow parameters of the intersection point q1 as the initial condition to perform streamline tracing, obtain the intersection point q2 of the streamline and the third-stage shock surface, and obtain the airflow parameters of the intersection point q2 by solving the conical flow.

[0038] Step S34: Starting from the intersection point q2, and using the airflow parameters of the intersection point q2 as initial conditions, perform streamline tracing to obtain a streamline in the three-stage conical basic flow field within a certain cross-section.

[0039] Step S35: Starting from each point on the leading edge, repeat steps S31 to S34 to obtain streamlines in each cross-section of the entire space, thus obtaining a multi-stage compression waverider configuration.

[0040] In the above design method, the CFD numerical simulation method in step S5 includes:

[0041] Select the high-resolution flux function as the RoeMAS format;

[0042] The shock wave stabilization method selected is a combination of the MUSCL scheme and the high-precision WENO scheme;

[0043] The radial basis function interpolation method was chosen as the unsteady coupled heat transfer simulation method.

[0044] The non-steady time propagation methods chosen are the explicit Runge-Kutta scheme and the implicit backdifference time scheme;

[0045] The two-equation SST turbulence model was selected as the turbulence model.

[0046] The governing equations are selected as the three-dimensional compressible Reynolds-averaged Navier-Stokes equations, specifically:

[0047]

[0048] in, It is a conserved variable; For inviscid vector flux in three directions; Let be the viscous vector flux in three directions.

[0049] In the above design method, step S7 uses the polar curve method to obtain the interference type and flow field parameters of the incident shock wave. The specific method is as follows:

[0050] The incident shock waves DS1 and DS2 respectively change the incoming flow to the back-shock state. The airflow parameters of the back-shock state and the polar curves of the incident shock waves DS1 and DS2 are obtained from the oblique shock wave relation.

[0051] The type of incident shock wave interference can be determined by the location of the intersection of the polar curves;

[0052] Starting with the airflow parameters after the incident shock waves DS1 and DS2, the flow field parameters of the transmitted shock waves TS1 and TS2 are determined using the following formula:

[0053]

[0054]

[0055]

[0056] Where P2 and Ma2 are the pressure and Mach number behind the shock wave, respectively, and P1 and Ma1 are the pressure and Mach number in front of the shock wave, respectively; θ is the airflow deflection angle, β is the shock wave angle, and γ is the gas constant.

[0057] In the above design method, the determination of the incident shock wave interference type by the intersection of the polar curves is as follows: if the polar curves of the transmitted shock waves TS1 and TS2 do not intersect, the interference type of the incident shock waves DS1 and DS2 is determined to be Mach reflection; if the polar curves of the transmitted shock waves TS1 and TS2 intersect, the interference type of the incident shock waves DS1 and DS2 is a regular reflection structure.

[0058] In the above design method, step S6 employs a flow field control reconstruction method to calculate the pressure gradient at the shock wave incident point. The specific method is as follows:

[0059] The pressure gradient at the shock wave incident point is given by the following formula:

[0060]

[0061] in, For the pressure gradient, Ma in P is the theoretical shock front Mach number or the expansion wavefront Mach number. in dθ / dx is the theoretical static pressure before the shock wave; dθ / dx is the expansion angle gradient; β is the shock wave angle; k is a constant; and Δl is the influence length of the introduced shock wave in the boundary layer.

[0062] In the above design method, the aerodynamic characteristics of the multi-stage compression waverider model include the total pressure recovery coefficient, wall pressure, temperature, Mach number, and pressure spatial distribution.

[0063] Compared with the prior art, the present invention has the following advantages:

[0064] (1) The present invention proposes and implements an internal rotating intake shock wave control design scheme. Compared with a simple configuration intake, the shock wave control design can reasonably allocate the overall shock wave and local shock wave distribution, reduce the shock wave interference intensity under high Mach number conditions, and has better aerodynamic performance and heat reduction performance.

[0065] (2) This invention corrects the simulation effect of shock wave control under high back pressure boundary by using a high-precision flow numerical simulation model;

[0066] (3) This invention establishes a simulation method and quantitative parameters for evaluating the intensity of shock wave interference under multi-wave interference through theoretical analysis of shock wave interference, and solves the problem of local shock wave interference configuration under different Mach numbers.

[0067] (4) This invention proposes a design method for the air intake of a wide Mach number air-breathing aircraft based on the heat reduction design concept. It includes a matching design method for local interference shock waves and overall compression shock waves, as well as a local shock wave control design method under complex background wave system conditions. This method is verified by CFD numerical simulation and provides technical support for the optimization design of the aircraft in the design stage. Attached Figure Description

[0068] Figure 1 This is a regular reflection diagram of the internal flow channel shock wave interference type described in this invention;

[0069] Figure 2 This is the Mach reflection diagram of the internal flow channel shock wave interference type described in this invention;

[0070] Figure 3 The following are the flow field contour maps of the inlet design state under Mach 12 of this invention; (a) is the Mach number contour map, and (b) is the pressure contour map.

[0071] Figure 4 This is a schematic diagram of the key geometric parameters for the cross-sectional design of the inner contraction section of this invention. Detailed Implementation

[0072] The present invention will now be described in further detail with reference to the accompanying drawings and specific embodiments.

[0073] This invention discloses a flow field reconstruction design method under strong wave interference control in an internal flow channel, comprising the following steps:

[0074] Step S1: Based on the aircraft design conditions and requirements, and using the two-dimensional surface design method of the forebody / inlet, propose important aerodynamic parameters that can control the inlet surface, perform parametric design on the two-dimensional surface of the forebody / inlet, and obtain the basic two-dimensional surface model.

[0075] Step S2: Based on the basic two-dimensional surface model, establish a two-dimensional hypersonic three-dimensional curved surface compression forebody / inlet design model based on streamline tracing;

[0076] Step S3: Based on the binary hypersonic three-dimensional curved surface compression precursor / inlet design model, obtain the multi-stage compression waverider configuration through the streamline tracing method;

[0077] Step S4: Import the multi-stage compression waverider configuration into the modeling software to generate a multi-stage compression waverider model;

[0078] Step S5: Obtain the aerodynamic characteristics of the multi-stage compression waverider model using the flow field CFD numerical simulation method; the aerodynamic characteristics of the multi-stage compression waverider model include the total pressure recovery coefficient, wall pressure, temperature, Mach number, and pressure spatial distribution.

[0079] Step S6: Based on the multi-stage compression waverider model, the flow field control reconstruction method is used to calculate the pressure gradient at the shock wave incident point; the aerodynamic characteristics are adjusted until the pressure gradient at the shock wave incident point meets the requirements, and the adjusted multi-stage compression waverider model is obtained.

[0080] Step S7: Based on the regulated multi-stage compression waverider model, the interference type and flow field parameters of the incident shock wave are obtained using the polar curve method. The shock wave is regulated to obtain the aerodynamic characteristics of the regulated multi-stage compression waverider model. It is determined whether the aerodynamic characteristics meet the required values. If yes, the process exits; otherwise, the process proceeds to step S6.

[0081] The design conditions and requirements for the aircraft in step S1 include: flight altitude, flight Mach number, capture flow rate, flow coefficient, total contraction ratio, and total deflection angle of the external pressure section.

[0082] The two-dimensional surface design method for the forebody / intake duct in step S1 is as follows:

[0083] Step S11: Determine the surface parameters of the external pressure section based on the geometric configuration of the aircraft forebody / inlet;

[0084] Step S12: Determine the inlet area of ​​the inner channel of the inner contraction section based on the profile parameters of the outer pressure section;

[0085] Step S13: Calculate the height and length of the isolation section.

[0086] The height and length of the isolation section are calculated as follows:

[0087] The height of the isolation section is equal to the height of the throat.

[0088] The formula for the length of the isolation section is:

[0089]

[0090] Where x is the length of the isolation section, and M1 is the Mach number at the throat exit. Where θ is the momentum Reynolds number, D is the throat outlet diameter, θ1 is the throat outlet surface angle, P is the incoming static pressure, and P1 is the throat outlet pressure.

[0091] The length of the isolation section is 6 to 10 times the height of the isolation section.

[0092] The external pressure section profile parameters are as follows: initial compression angle (α), ratio of central body radius to reference flow field radius, dimensionless central radial distance y / Ri (Rc / Ri), capture profile width-to-height ratio B1 / H1, and throat section width-to-height ratio B2 / H2.

[0093] In this embodiment, the two-dimensional profile design of the forebody / inlet is carried out from three parts: the external pressure section, the internal contraction section, and the isolation section. The important aerodynamic parameters that can control the profile of the inlet are extracted, and the parametric design of the two-dimensional profile of the forebody / inlet is realized, providing a basic profile for the three-dimensional forebody / inlet design in the following text.

[0094] (1) External pressure section

[0095] Based on the geometry of the aircraft's forebody / inlet, the wedge angles of each stage of the external pressure section can be determined. The shock wave angle β of each stage of the wedge surface... i Mach number M after wave i Pressure P i and total pressure recovery coefficient σ i Determined by the following formula:

[0096]

[0097]

[0098]

[0099]

[0100] In the formula, M0 is the Mach number of the incoming flow.

[0101] (2) Internal contraction segment

[0102] After obtaining the external pressure section profile through the calculations in the previous section, the inlet area of ​​the intake passage can be determined accordingly, such as... Figure 4 As shown, the coordinates of the midpoint A of the inlet section of the air intake are calculated according to the following formula:

[0103]

[0104] To calculate the coordinates of point B at the midpoint of the throat section, the centerline length L and the centerline deflection angle θ need to be input. Based on geometric relationships, the coordinates of point B are calculated using the following formula:

[0105]

[0106] Since point B is the midpoint of the throat section EF, the throat area H th According to the internal contraction ratio CR in It can be calculated Therefore, the coordinates of points E and F are calculated according to the following formula:

[0107]

[0108]

[0109] For the design of the upper wall CE and lower wall DF of the internal pressure section, on the one hand, it is necessary to consider the intensity of the lip shock wave to ensure that the lip shock wave is not too strong, which would cause a large-scale separation of the shoulder boundary layer and reduce the performance of the intake. On the other hand, the airflow is continuously compressed and gradually deflected in the internal pressure section. Studies have found that using an arc or curve transition at the shoulder can improve the stability of the airflow, improve the boundary layer's resistance to back pressure, and weaken the high pressure on the lip surface, which is beneficial to improving the performance of the intake. Therefore, many designs of the inner contraction section of the intake often adopt the form of a smooth curve transition.

[0110] In this invention, the upper and lower wall profiles of the inner channel are designed using cubic curves, and their functional form is as follows:

[0111]

[0112] in and Given dimensionless coordinates of magnitude 1, and knowing the coordinates of points A and B and their corresponding slopes, all the undetermined coefficients can be obtained.

[0113] (3) Isolation section

[0114] The isolation section of this invention is designed as a straight pipe, and the height of the isolation section is the same as the throat height. The length of the isolation section is given according to the empirical formula of Waltrup and Billig:

[0115]

[0116] The longer the isolation section, the higher the back pressure it can withstand. However, an excessively long isolation section increases the weight of the entire aircraft. Therefore, the length of the isolation section is generally taken as about 6 to 10 times the height of the isolation section. Unless otherwise specified in this method, the length of the isolation section is taken as 7 times the height of the throat.

[0117] In step S2, based on the basic two-dimensional surface model, a two-dimensional hypersonic three-dimensional curved surface compression forebody / inlet design model based on streamline tracing is established, specifically as follows:

[0118] Step S21: Determine the geometric control variables;

[0119] Step S22: Based on the basic two-dimensional surface model, according to the three angles of the three-wave external pressure section and the incoming Mach number, the two-dimensional configuration inviscid flow field after flowing through the three external pressure shock waves is obtained.

[0120] Step S23: The upper surface of the aircraft forebody adopts a parallel inviscid flow field, and the lower surface of the aircraft adopts a two-dimensional configuration inviscid flow field of the air intake.

[0121] Step S24: Select different curves on the cross section perpendicular to the flow direction at the air intake to perform streamline tracing and generate several different physical surfaces;

[0122] Step S25: Generate different binary hypersonic three-dimensional curved surface compression precursor / inlet design models based on streamline tracing according to several different physical surfaces.

[0123] Geometric control variables, including forebody width ratio, forebody angle, and the degree n of the forebody lower surface control curve. down and the number of times the control curve of the upper surface of the precursor n up ;n up ≤n down .

[0124] In this embodiment, based on the two-dimensional inlet model from step S1, the three angles of the three-wave external pressure section are known, and the Mach number of the incoming flow is known. Therefore, the two-dimensional inviscid flow field after the outflow passes through the three external pressure shock waves can be calculated. Based on this flow field, if it is considered as an infinitely wide three-dimensional flow field, then at each point on the cross-section perpendicular to the flow direction at the inlet of the inlet, a corresponding streamline can be found. If a line is taken on the cross-section, then a corresponding streamline can be found at each point on the line, thus forming a surface. Using this surface as the physical shape surface, the streamlines of the incoming flow passing through this physical shape surface should be consistent with the previous flow field.

[0125] Based on the above principles, the upper surface of the aircraft forebody adopts a parallel inviscid flow field, while the lower surface adopts the inviscid flow field of the two-dimensional configuration of the air intake mentioned above. Different curves are selected on the cross-section perpendicular to the flow direction at the air intake inlet for streamline tracing, thereby generating different physical surfaces and thus different three-dimensional configurations of the forebody / air intake. Since the original profiles of the upper and lower surfaces of the aircraft forebody are generated on the cross-section perpendicular to the flow direction at the air intake inlet, and then streamlines are traced in reverse based on streamline tracing, it is necessary to ensure that the generated surfaces intersect and stitch at the starting point of the forebody. Therefore, the selection of the number n of the control curves for the upper and lower surfaces must satisfy n... up ≤n down .

[0126] In step S3, the multi-stage compression waverider configuration is obtained using the streamline tracing method. The specific method is as follows:

[0127] Step S31: Intersect a plane parallel to the cone axis with the first-stage shock surface to obtain a leading edge line, and discretize the leading edge line into a certain number of point sets; the number of points in the point set n≥2;

[0128] Step S32: Starting from a point on the leading edge, perform streamline tracing in the first-stage conical flow field and generate an intersection point q1 with the second-stage shock wave surface. Obtain the airflow parameters of the intersection point q1 by solving the conical flow.

[0129] Step S33: Starting from intersection point q1, use the airflow parameters of intersection point q1 as initial conditions to perform streamline tracing, obtain the intersection point q2 of the streamline and the third-stage shock surface, and obtain the airflow parameters of intersection point q2 by solving the conical flow.

[0130] Step S34: Starting from intersection point q2, use the airflow parameters of intersection point q2 as initial conditions to perform streamline tracing and obtain a streamline in the three-stage conical basic flow field within a certain cross-section.

[0131] Step S35: Starting from each point on the leading edge, repeat steps S31 to S34 to obtain streamlines in each cross-section of the entire space, thus obtaining a multi-stage compression waverider configuration.

[0132] The CFD numerical simulation method in step S5 includes:

[0133] Step S51: Select the high-resolution flux function as the RoeMAS format;

[0134] Step S52: Select a shock wave stabilization method that combines the MUSCL scheme with the high-precision WENO scheme;

[0135] Step S53: Select the radial basis function interpolation method as the unsteady coupled heat transfer simulation method;

[0136] Step S54: Select the non-steady time propagation method as the explicit Runge-Kutta format and the implicit backdifference time format;

[0137] Step S55: Select the two-equation SST turbulence model;

[0138] Step S56: Select the three-dimensional compressible Reynolds-averaged Navier-Stokes equations as the governing equations, specifically:

[0139]

[0140] in, It is a conserved variable; For inviscid vector flux in three directions; Let be the viscous vector flux in three directions.

[0141] like Figure 3 , Figure 4 As shown, in step S7, the polar curve method is used to obtain the disturbance type and flow field parameters of the incident shock wave. The specific method is as follows:

[0142] Step S71: The incident shock waves DS1 and DS2 respectively change the incoming flow to the post-shock state. The airflow parameters of the post-shock state and the polar curves of the incident shock waves DS1 and DS2 are obtained from the oblique shock wave relation.

[0143] Step S72: Determine the type of incident shock wave interference by the location of the intersection of the polar curves; if the polar curves of the transmitted shock waves TS1 and TS2 do not intersect, then the interference type of the incident shock waves DS1 and DS2 is determined to be Mach reflection, and proceed to step S73; if the polar curves of the transmitted shock waves TS1 and TS2 intersect, then the interference type of the incident shock waves DS1 and DS2 is a regular reflection structure, and proceed to step S74.

[0144] Step S73, the method for obtaining the airflow parameters after Mach reflection is as follows: the airflow parameters in the regions after the transmitted shock waves TS1 and TS2 are the pressure values ​​corresponding to the intersection points of the incident shock wave DS1 and the transmitted shock wave TS1, and the incident shock wave DS2 and the transmitted shock wave TS2 on the polar curves; for example... Figure 2 As shown.

[0145] Step S74, the method for obtaining the airflow parameters after regular reflection is as follows: the airflow parameters after the transmitted shock waves TS1 and TS2 are the pressure values ​​corresponding to the intersection points of the incident shock waves DS1 and DS2, and the transmitted shock waves TS1 and TS2 on the polar curve; for example... Figure 1 As shown.

[0146] The method for obtaining the interference type and flow field parameters of the local bow shock wave is as follows: the flow parameters of the bow shock wave are obtained by using the supersonic airflow through the curved wall calculation method, and the polar curve of the bow shock wave is obtained. If the polar curves of the bow shock wave and the transmitted shock wave intersect, the interference type is determined to be opposite-side reflection; if there is no intersection, the interference type is determined to be same-side reflection.

[0147] Starting with the airflow parameters after the incident shock waves DS1 and DS2, the flow field parameters of the transmitted shock waves TS1 and TS2 are determined using the following formula:

[0148]

[0149]

[0150]

[0151] Where P2 and Ma2 are the pressure and Mach number behind the shock wave, respectively, and P1 and Ma1 are the pressure and Mach number in front of the shock wave, respectively; θ is the airflow deflection angle, β is the shock wave angle, and γ is the gas constant.

[0152] In step S6, the flow field control reconstruction method is used to calculate the pressure gradient at the shock wave incident point. The specific method is as follows:

[0153] The pressure gradient at the shock wave incident point is given by the following formula:

[0154]

[0155] in, For the pressure gradient, Ma in P is the theoretical shock front Mach number or the expansion wavefront Mach number. in dθ / dx is the theoretical static pressure before the shock wave; dθ / dx is the expansion angle gradient; β is the shock wave angle; k is a constant, taken as 1.4; and Δl is the influence length of the introduced shock wave in the boundary layer.

[0156] Example 1

[0157] The method provided in this embodiment has the following specific steps:

[0158] 1) Construct a conical shock wave flow field with three intersecting shock waves, and give the relationship between the angle of airflow turning through the oblique shock wave and the pressure ratio before and after the oblique shock wave through the oblique shock wave relation;

[0159] 2) Obtain the multi-stage compression waverider configuration using the streamline tracing method, import it into modeling software, and generate a multi-stage compression waverider model;

[0160] 3) The flight Mach number, altitude, angle of attack, atmospheric density, etc. are used as input conditions for flow calculation. The overall performance, including total pressure recovery coefficient, inlet pressure, temperature, Mach number, and pressure spatial distribution, is obtained through flow field CFD numerical simulation.

[0161] 4) Repeat steps (1) to (3) until the required intake aerodynamic performance is achieved;

[0162] 5) By adopting appropriate flow field intervention methods, the topological positional relationship between the incident shock wave and the local shock wave is controlled to realize the inlet flow field reconstruction under complex background environment with strong shock wave and high adverse pressure gradient;

[0163] 6) Shock wave control is performed on the high Mach number inlet. An inward rotating shape is selected, and the flow field is intervened in the shock wave interference area at the inlet lip and root to control the topological position relationship between the incident shock wave and the local shock wave.

[0164] This embodiment solves the following technical problem:

[0165] 1. Establish a geometrically constrained design method for a two-dimensional hypersonic three-dimensional curved surface compression forebody / inlet, which effectively improves the flow capture capability and total pressure recovery coefficient of the inlet in a wide velocity range, provided that the shock wave of the forebody does not enter the inlet flow channel.

[0166] 2. Based on the typical design scheme of a two-dimensional three-dimensional curved surface compression forebody / inlet, this paper addresses the high adverse pressure gradient wave system control problem in the high Mach number inlet across the local interference zone by adding shock wave control measures at key local locations. This ultimately forms a design method for reconstructing the flow field of the high Mach number inlet across the local zone under shock wave control. The flow field design results are as follows: Figure 3 (a) and Figure 3 As shown in (b).

[0167] 3. Based on the aerodynamic characteristics, flow field control and reconstruction of the trans-domain high Mach number inlet under shock wave regulation, wind tunnel test verification research was carried out, and flow field reconstruction control criteria under strong wave interference in the hypersonic internal flow channel were established.

[0168] This invention relates to a geometrically constrained design method for a two-dimensional hypersonic three-dimensional curved surface compression forebody / inlet. By analyzing and revealing the coupling flow mechanism between the shock wave morphology reflected at the inlet lip and the lip sweep angle, it presents the flow field structure variation law and design selection principles under the control of strong wave interference in the inner flow channel. Under the premise that the forebody shock waves do not enter the inner flow channel of the inlet, it can effectively improve the flow capture capability and total pressure recovery coefficient in the wide velocity range of the inlet.

[0169] The above description is only the best specific embodiment of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in the present invention should be included within the protection scope of the present invention.

[0170] The contents not described in detail in this invention are common knowledge to those skilled in the art.

Claims

1. A flow field reconstruction design method under strong wave interference control in an internal flow channel, characterized in that, Includes the following steps: S1. Based on the aircraft design conditions and requirements, and using the two-dimensional surface design method of the forebody / inlet, important aerodynamic parameters that can control the inlet surface are proposed. The two-dimensional surface of the forebody / inlet is parametrically designed to obtain the basic two-dimensional surface model. S2. Based on the basic two-dimensional surface model, establish a two-dimensional hypersonic three-dimensional curved surface compression precursor / inlet design model based on streamline tracing. S3. Based on the binary hypersonic three-dimensional curved surface compression precursor / inlet design model, the multi-stage compression waverider configuration is obtained through the streamline tracing method. S4. Import the multi-stage compression waverider configuration into the modeling software to generate a multi-stage compression waverider model; S5. The aerodynamic characteristics of the multi-stage compression waverider model are obtained by using the flow field CFD numerical simulation method. S6. Based on the multi-stage compression waverider model, the flow field control reconstruction method is used to calculate the pressure gradient at the shock wave incident point; the aerodynamic characteristics are adjusted until the pressure gradient at the shock wave incident point meets the requirements, and the adjusted multi-stage compression waverider model is obtained. S7. Based on the modulated multi-stage compression waverider model, the interference type and flow field parameters of the incident shock wave are obtained using the polar curve method. The shock wave is modulated to obtain the aerodynamic characteristics of the modulated multi-stage compression waverider model. It is then determined whether the aerodynamic characteristics meet the required values. If yes, exit; otherwise, proceed to step S6. In step S6, a flow field control reconstruction method is used to calculate the pressure gradient at the shock wave incident point. The specific method is as follows: The pressure gradient at the shock wave incident point is given by the following formula: in, For pressure gradient, dθ / dx is the theoretical pre-shock Mach number or the pre-expansion Mach number; β is the theoretical pre-shock static pressure; dθ / dx is the expansion angle gradient; β is the shock angle; k is a constant. To introduce the influence length of the shock wave in the boundary layer.

2. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 1, characterized in that: The design conditions and requirements for the aircraft in step S1 include: flight altitude, flight Mach number, capture flow rate, flow coefficient, total contraction ratio, and total deflection angle of the external pressure section.

3. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 1, characterized in that: The two-dimensional surface design method for the forebody / intake duct in step S1 is as follows: (1) Determine the surface parameters of the external pressure section based on the geometric configuration of the aircraft forebody / inlet; (2) Determine the inlet area of ​​the inner channel of the inner contraction section based on the profile parameters of the outer pressure section; (3) Calculate the height and length of the isolation section.

4. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 3, characterized in that: The calculation of the height and length of the isolation segment is specifically as follows: The height of the isolation section is equal to the height of the throat. The formula for the length of the isolation section is: in, M1 is the length of the isolation section, and M2 is the Mach number at the throat exit. where is the momentum Reynolds number. D The diameter of the throat outlet. θ 1 is the throat exit surface angle. For incoming static pressure, This refers to the pressure at the throat outlet.

5. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 3, characterized in that: The length of the isolation section is 6 to 10 times the height of the isolation section.

6. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 3, characterized in that: The external pressure section profile parameters are specifically: initial compression angle, ratio of central body radius to reference flow field radius, dimensionless radial distance Rc / Ri from the center, aspect ratio of the capture profile, and aspect ratio of the throat section.

7. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 1, characterized in that: In step S2, a binary hypersonic three-dimensional curved surface compression forebody / inlet design model based on streamline tracing is established based on the basic two-dimensional surface model. Specifically: Determine the geometric control variables; Based on the basic two-dimensional surface model, according to the three angles of the three-wave external pressure section and the incoming Mach number, the two-dimensional configuration inviscid flow field after passing through three external pressure shock waves is obtained. The upper surface of the aircraft forebody adopts a parallel inviscid flow field, while the lower surface of the aircraft adopts a two-dimensional inviscid flow field with an air intake configuration. Different curves are selected on the cross section perpendicular to the flow direction at the air intake to perform streamline tracing and generate several different physical surfaces; Based on several different physical surfaces, a three-dimensional curved surface compression precursor / intake design model is generated.

8. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 7, characterized in that: The geometric control variables include the forebody width ratio, forebody angle, and the number n of the control curves on the lower surface of the forebody. down and the number of times the control curve of the upper surface of the precursor n up ;n up ≤n down .

9. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 1, characterized in that: In step S3, the multi-stage compressed waverider configuration is obtained through streamline tracing. The specific method is as follows: Step S31: Intersect a plane parallel to the cone axis with the first-stage shock surface to obtain a leading edge line, and discretize the leading edge line into a certain number of point sets; the number of points in the point sets n≥2; Step S32: Starting from a point on the leading edge, perform streamline tracing in the first-stage conical flow field and generate an intersection point q1 with the second-stage shock wave surface. Obtain the airflow parameters of the intersection point q1 by solving the conical flow. Step S33: Starting from the intersection point q1, use the airflow parameters of the intersection point q1 as the initial condition to perform streamline tracing, obtain the intersection point q2 of the streamline and the third-stage shock wave surface, and obtain the airflow parameters of the intersection point q2 by solving the conical flow. Step S34: Starting from the intersection point q2, and using the airflow parameters of the intersection point q2 as initial conditions, perform streamline tracing to obtain a streamline in the three-stage conical basic flow field within a certain cross-section. Step S35: Starting from each point on the leading edge, repeat steps S31 to S34 to obtain streamlines in each cross-section of the entire space, thus obtaining the multi-stage compression waverider configuration.

10. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 1, characterized in that: The CFD numerical simulation method in step S5 includes: Select the high-resolution flux function as the RoeMAS format; The shock wave stabilization method selected is a combination of the MUSCL scheme and the high-precision WENO scheme; The radial basis function interpolation method was chosen as the unsteady coupled heat transfer simulation method. The non-steady time propagation methods chosen are the explicit Runge-Kutta scheme and the implicit backdifference time scheme; The two-equation SST turbulence model was selected. The governing equations are selected as the three-dimensional compressible Reynolds-averaged Navier-Stokes equations, specifically: in, It is a conserved variable; For inviscid vector flux in three directions; Let be the viscous vector flux in three directions.

11. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 1, characterized in that, In step S7, the polar curve method is used to obtain the interference type and flow field parameters of the incident shock wave. The specific method is as follows: Incident shock wave DS1 and DS2 The incoming flow is transformed into the post-shock state, and the airflow parameters of the post-shock state and the polar curves of the incident shock waves DS1 and DS2 are obtained from the oblique shock wave relation. The type of incident shock wave interference can be determined by the location of the intersection of the polar curves; With incident shock wave DS1 and DS2 Starting with the subsequent airflow parameters, determine the transmitted shock wave. TS1 and TS2 The flow field parameters are given by the following formula: in, , These are the pressure and Mach number after the shock wave, respectively. , These are the pressure before the shock wave and the Mach number, respectively. It is the airflow deflection angle. The shock angle; is the gas constant.

12. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 11, characterized in that: The method of determining the type of incident shock wave interference by the location of the intersection of the polar curves specifically involves: if the transmitted shock wave... TS1 and TS2 If the polar curves do not intersect, then the incident shock wave is determined to be... DS1 and DS2 The interference type is Mach reflection; if the transmitted shock wave TS1 and TS2 If the polar curves intersect, then the incident shock wave... DS1 and DS2 The interference type is a regular reflection structure.

13. The flow field reconstruction design method under strong wave interference control in an internal flow channel according to claim 1, characterized in that: The aerodynamic characteristics of the multi-stage compression waverider model include the total pressure recovery coefficient, wall pressure, temperature, Mach number, and pressure spatial distribution.

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