A lattice-core regenerative cooling channel and its support rod cross-section parameterization design method

Abstract

This invention discloses a lattice-core regenerative cooling channel and its support rod cross-section parameterization design method, belonging to the aerospace field. The cooling channel includes an inner wall, an outer wall, and a lattice core layer. The cross-section of the lattice support rod is a segmented composite airfoil cross-section, wherein the leading edge segment is constructed based on the NACA thickness skeleton function, and the trailing edge segment is constructed by a rational quadratic Bézier curve. The parameterization design method includes: S1, inputting and initializing the target parameters of the cross-section; S2, performing constrained reconstruction of the leading edge segment and solving it; S3, determining the splicing point and constructing the trailing edge segment; S4, similarity scaling and geometric constraint verification. This method can achieve independent and precise control of multiple geometric parameters, and automatically iterates and updates when constraints are not met through an adaptive parameter search mechanism. The generated cross-section has both a streamlined shape and smooth trailing edge characteristics, which can significantly reduce the flow resistance of the lattice-core regenerative cooling channel and enhance heat transfer.

Description

Technical Field

[0001] This invention relates to the field of aerospace technology, and in particular to a lattice sandwich regenerative cooling channel and a parametric design method for its support cross-section. Background Technology

[0002] Regenerative cooling technology is one of the key technologies for thermal protection of the combustion chamber and nozzle of liquid rocket engines and scramjet engines. In recent years, in order to take into account the comprehensive performance requirements of lightweight, high structural rigidity and efficient heat transfer of regenerative cooling channels, rod-type lattice sandwich regenerative cooling channels formed by additive manufacturing process have become an important development direction in this field.

[0003] Currently, conventional lattice sandwich structures mostly use circular cross-section struts. This cross-section shape easily leads to significant inflow blockage and a low-velocity wake region within the cooling channel, and flow separation occurs early, increasing channel pressure drop and thus restricting further improvement in flow and heat transfer performance. Therefore, adopting streamlined cross-sections is an important direction for improving the overall channel performance. However, classic airfoil or teardrop-shaped cross-sections usually have sharp trailing edges, which not only pose challenges to the forming accuracy and surface quality control of additive manufacturing, but also easily induce stress concentration in the trailing edge region, affecting the reliability of the structure under high temperature and high pressure environments. In addition, in the actual engineering optimization and design of lattice sandwich structures, the strut cross-section usually needs to simultaneously meet multiple geometric constraints such as thickness ratio, maximum thickness location, leading edge bluntness parameter, chord length limit, and target cross-sectional area. In the current parameter system of mainstream classic airfoil or teardrop-shaped cross-sections, the various geometric features are coupled with each other, limiting the degree of design freedom and making it difficult to achieve independent and precise control of shape and key dimensions under multiple constraints. Summary of the Invention

[0004] In order to achieve synergistic optimization of low flow resistance and high heat transfer while ensuring that the cross-sectional design can simultaneously meet multiple geometric constraints and has good design flexibility and optimization space, this application provides a parametric design method for the cross-section of a lattice sandwich regenerative cooling channel and its support rod.

[0005] Firstly, this application provides a lattice-shaped sandwich regenerative cooling channel, employing the following technical solution:

[0006] A lattice sandwich regenerative cooling channel includes an outer wall, an inner wall, and a lattice sandwich layer disposed between the outer wall and the inner wall. The lattice sandwich layer is composed of multiple rod cells, each rod cell consisting of several support rods. The cross-section of the support rods is a segmented composite airfoil section, the edge contour of which is formed by splicing a leading edge segment and a trailing edge segment, which are spliced ​​together at the splicing point. The shape of the leading edge segment is defined by a half-thickness distribution function, which does not satisfy the functional form of the NACA thickness polynomial skeleton. The trailing edge segment is constructed by a rational quadratic Bézier curve.

[0007] Optionally, the rod unit is one or more of the following: body-centered cubic (BCC) unit, Kagome unit, Kelvin unit, octahedral truss unit, tetrahedral truss unit, pyramidal truss unit, single-column unit, or a topologically equivalent variant of the above units; the rod units are arranged in a staggered array, and the orientation of the segmented composite airfoil section support is: the trailing edge section faces the incoming flow direction, and the leading edge section faces away from the incoming flow direction.

[0008] Secondly, this application provides a parametric design method for segmented composite airfoil sections, employing the following technical solution:

[0009] A parametric design method for segmented composite airfoil sections includes the following steps:

[0010] S1. Input the target parameters for the cross-section design. The target parameters include at least the thickness ratio t and the location of the maximum thickness ξ. m chord length limit L limit Leading edge bluntness parameter R * and the target cross-sectional area A tar , where t=D m / L,D m Where L is the maximum thickness and L is the chord length; ξ m =x m / L,x m R represents the chord coordinate of the location of maximum thickness; * =R n / D m R n Define the equivalent radius of the leading edge; initialize the starting position parameters of the rational quadratic Bézier curve of the trailing edge segment. With curve weighting factor w t ;in Used to determine the location of the splicing point, w t Used to adjust the fullness of the tail edge and the rate of contraction;

[0011] S2, Based on chord-dimensional coordinates Construct the half-thickness distribution function of the symmetrical cross section Forward edge amplitude coefficient To actively control the quantity, the polynomial skeleton of the NACA thickness distribution function of the leading edge segment is reconstructed under constraints, so that the leading edge segment is fixed at the position of maximum thickness as ξ. m Under the premise of establishing R * With effective leading edge coefficient The constraints between them are determined through iterative solutions. This allows the leading edge segment to simultaneously satisfy t and ξ. m With R * Constraints;

[0012] S3, the leading edge segment at the splicing point ξ s The segment smoothly transitions to the tail edge segment, with the splicing point located at ξ. s From parameters Determined and satisfied ; in ξ=ξ s The coordinates and tangent direction of the splicing point are obtained, and the control point H of the tail edge segment is determined by the intersection of the splicing point tangent and the tail edge closing direction. A rational quadratic Bézier tail edge curve that is spliced ​​with the leading edge segment at the splicing point in a first-order geometric continuous manner is generated.

[0013] S4. Combine the leading edge segment and the trailing edge segment to form a closed profile, calculate the cross-sectional area of ​​the closed profile under the condition of unit chord length, and perform the operation based on the target area A. tar The constraint-based similarity scaling makes the final cross-sectional area equal to A. tar It outputs segmented composite airfoil cross-sectional geometry data that satisfies all parameter constraints.

[0014] Optionally, in step S2, the NACA thickness distribution function of the leading edge segment is... The polynomial skeleton is constrained and reconstructed to construct an unnormalized thickness skeleton function:

[0015] .

[0016] in The preset NACA thickness distribution function The skeleton constant; to ensure the thickness skeleton function is within... At the extreme value, for Applying extreme value constraints Thus, the coefficients of the linear terms are explicitly determined. for:

[0017] .

[0018] Optionally, an amplitude normalization coefficient is introduced in step S2. A normalized thickness skeleton function is defined to automatically satisfy the thickness ratio constraint:

[0019] .

[0020] Optionally, the effective leading edge coefficient in step S2 is defined as:

[0021]

[0022] And establish the leading edge bluntness parameter R * and Display constraints:

[0023]

[0024] The constant C is determined by the analytical relation of the NACA leading edge curvature; it is solved through fixed-point iteration or equivalent one-dimensional iteration, making... Satisfy objective R * Thus determining the unique .

[0025] Optionally, the splicing point position in step S3 satisfies:

[0026]

[0027] And in ξ=ξ s Take half the thickness at the splicing point and the slope of the tangent line s:

[0028] .

[0029] Optionally, the control point H of the trailing edge segment in step S3 is determined by the following geometric construction: defining the splicing point. With the trailing edge point Take the tangential direction of the splicing point With the tail edge closing direction Solve for the equation of the line:

[0030]

[0031] Where α and β represent the distance from point P, respectively. s Starting from T along their respective direction vectors t P With t T The scalar displacement coefficient; the intersection point H is the control point of the rational quadratic Bézier curve.

[0032] Optionally, in step S3, the trailing edge segment is weighted by a factor of w. t The rational quadratic Bézier curve is expressed as:

[0033]

[0034] By positioning the control point H along the tangent of the splicing point, the trailing edge segment can be geometrically continuous with the leading edge segment at u=0.

[0035] Optionally, in step S4, based on the target area A tar The similarity scaling of the constraints includes: within a unit chord length L unit Given a constant value of 1, a closed contour is discretized and its area A is calculated. unit The chord length is determined by the similarity scaling relationship:

[0036]

[0037] And perform on the closed contour The proportional scaling ensures that the final cross-sectional area satisfies A=A tar When the chord length L > L after scaling limit At that time, within the preset range, the trailing edge weight parameter w is adjusted. t or / and trailing edge starting point ratio parameter The process involves automatic searching and iterative updates, repeating the closed contour generation, area calculation, and similarity scaling steps after each update until the chord length constraint L≤L is satisfied. limit .

[0038] In summary, this application includes the following beneficial technical effects:

[0039] (1) By adopting a segmented composite airfoil closed section profile, the streamline characteristics of the support rod are significantly improved, which helps to reduce the flow blockage effect and the low-velocity region of the wake in the channel, delay or suppress flow separation, thereby enhancing convective heat transfer while reducing flow resistance, and achieving synergistic optimization of low flow resistance and high heat transfer.

[0040] (2) By using a rational quadratic Bézier curve to construct a smooth and closed trailing edge segment, the stress concentration problem caused by the sharp trailing edge of classic airfoils or teardrop-shaped sections is avoided. This design not only helps to improve the forming accuracy and surface quality of additive manufacturing, but also significantly improves the fatigue life and structural reliability of the support rod under high temperature and high pressure environment;

[0041] (3) By reconstructing the thickness distribution of the leading edge segment under constraints and connecting the leading and trailing edges using a parametric splicing method, independent control of multiple geometric parameters such as thickness ratio, maximum thickness location, leading edge bluntness parameter, chord length, and target cross-sectional area is achieved. This method overcomes the limitations of strong coupling of geometric parameters and insufficient design freedom in classic airfoils or teardrop-shaped airfoils, enabling designers to flexibly and accurately generate cross-sectional configurations that meet specific engineering requirements under multiple constraints.

[0042] (4) The parametric design method has a clear input and output interface and an iterative solution process, which is easy to integrate into optimization algorithms or computer-aided design systems, supports multi-objective performance optimization and automated design, and significantly improves design efficiency and repeatability. Attached Figure Description

[0043] Figure 1 This is a schematic diagram of the structure of the dot matrix sandwich regeneration cooling channel described in this invention;

[0044] Figure 2 A schematic diagram of the BCC cell structure, the support rod and its cross-section using a segmented composite airfoil cross-section support rod;

[0045] Figure 3A flowchart for the parametric design of segmented airfoil sections;

[0046] Figure 4 The segmented composite airfoil profile curves generated under different thickness ratios (t=0.3~0.8);

[0047] Figure 5 The curve shows the channel pressure drop as a function of thickness ratio t.

[0048] Figure 6 The curve shows the variation of the channel average Nusselt number with thickness ratio t;

[0049] Figure 7 The curve shows the variation of the channel integration coefficient (PEC) with the thickness ratio t.

[0050] Explanation of reference numerals in the attached figures:

[0051] 1. Inner wall; 2. Outer wall; 3. Lattice sandwich layer; 4. Rod cell; 5. Support rod; 6. Segmented composite airfoil section. Detailed Implementation

[0052] The following is in conjunction with the appendix Figure 1-7 This application will be described in further detail.

[0053] This application discloses a dot-matrix sandwich regenerative cooling channel, such as... Figure 1 As shown, this embodiment of the invention provides a lattice sandwich regenerative cooling channel structure based on segmented composite airfoil-shaped cross-section support rods. The structure, from the inside out, includes an inner wall 1, a lattice sandwich layer 3, and an outer wall 2.

[0054] The cooling medium (such as liquid hydrogen, kerosene, methane, etc.) flows in from the inlet at one end of the channel, passes through the lattice sandwich layer 3, and finally flows out from the outlet at the other end. During this process, the cooling medium continuously absorbs the high-temperature heat load conducted by the inner wall 1, thereby achieving effective thermal protection for the thrust chamber combustion chamber and nozzle wall.

[0055] like Figure 1 As shown, the lattice sandwich layer 3 fills the annular space between the inner wall 1 and the outer wall 2, serving the dual functions of structural support and flow disturbance. Figure 2A preferred embodiment is shown: the lattice sandwich layer 3 is composed of a plurality of rod cells 4 arranged periodically. In this embodiment, the rod cells 4 are body-centered cubic (BCC) cells. The BCC cells are arranged along the channel curve in the axial direction and are evenly distributed along the circumference in the circumferential direction. Adjacent cells are preferably arranged in an alternating array to enhance flow disturbance and heat transfer efficiency. The rod cells can also be one or more of the following: Kagome cells, Kelvin cells, octahedral truss cells, tetrahedral truss cells, pyramidal truss cells, single-pillar cells, or topological equivalent variants of the above cells.

[0056] The core of this invention lies in, as Figure 2 As shown, the cross-section of the BCC lattice strut 5 is a segmented composite airfoil section 6. This section has a smooth, closed profile, formed by the leading edge and trailing edge segments joined at the splicing point P in a first-order geometrically continuous manner. The leading edge segment is defined by a half-thickness distribution function that satisfies the functional form of the NACA thickness polynomial skeleton. The trailing edge segment is constructed using a rational quadratic Bézier curve, thus forming a smooth, closed tail without sharp features. This improves the quality of additive manufacturing and reduces the risk of stress concentration in the trailing edge region.

[0057] The core advantage of this segmented composite configuration lies in the fact that, through the decoupled design and parametric splicing of the leading edge and trailing edge segments, multiple geometric constraints such as thickness ratio, maximum thickness position, leading edge bluntness parameter, target area and chord length can be independently specified and coordinated to meet them. This facilitates the controllable adjustment and repeated generation of cross-sectional geometry, overcoming the limitations of strong parameter coupling and insufficient design freedom of classic airfoils or teardrop-shaped cross-sections.

[0058] Preferably, the support rod 5 with section 6 is oriented along the direction of the incoming flow of the cooling medium, so that the trailing edge faces the direction of the incoming flow and the leading edge faces away from the direction of the incoming flow, in order to reduce the obstruction of the incoming flow and delay the occurrence of flow separation, while increasing the effective contact and scouring area between the incoming flow and the surface of the support rod, thereby achieving synergistic optimization of flow resistance and heat transfer.

[0059] (2) A parametric design method for segmented composite airfoil cross sections

[0060] like Figure 3 As shown, this embodiment of the invention provides a parametric design method for segmented composite airfoil cross-sections, used to output closed cross-section contour coordinate data that satisfies preset geometric constraints, including the following steps:

[0061] S1. Input the target parameters for the cross-section design. The target parameters include at least the thickness ratio t and the location of the maximum thickness ξ. m chord length limit L limit Leading edge bluntness parameter R * and the target cross-sectional area Atar , where t=D m / L,D m Where L is the maximum thickness and L is the chord length; ξ m =x m / L,x m R represents the chord coordinate of the location of maximum thickness; * =R n / D m R n The radius of the leading edge is the equivalent circular arc radius.

[0062] Simultaneously, initialize the starting position parameters of the rational quadratic Bézier curve at the trailing edge. With curve weighting factor w t ;in Used to determine the location of the splicing point, w t Used to adjust the fullness of the tail edge and the rate of contraction.

[0063] To clearly explain the calculation process of the parametric design method for the segmented composite airfoil section of the support rod, and to demonstrate the key intermediate quantities and output results, this embodiment constrains the input target parameters as follows: t=0.6, ξ m =0.3、R * =0.2、A tar =0.58mm 2 L limit =1.8mm, and initialized =0.1、w t =0.9.

[0064] S2. Construct the half-thickness distribution function of the leading edge segment and reconstruct it through constraint to simultaneously satisfy: maximum thickness location constraint, thickness ratio constraint, and leading edge bluntness parameter constraint. Preferably, the leading edge segment is represented by the half-thickness basis function of the NACA thickness polynomial skeleton, and an adjustable parameter is introduced. Used to adjust the relationship between the curvature level of the leading edge segment and the scale mapping.

[0065] Preferably, the commonly used fitting constants for the thickness frame of a NACA four-digit symmetric airfoil with a closed trailing edge are: a0=0.2969, a2=−0.3516, a3=0.2843, a4=−0.1036, used to ensure that the leading edge segment... The surrounding area exhibits reasonable smoothness. The unnormalized half-thickness basis function of the leading edge segment is defined as follows: ,in For chord-normalized coordinates, To adjust the parameters:

[0066]

[0067] To make the leading edge segment in At the location where the maximum thickness is reached, ,get Constraint reconstruction expression:

[0068]

[0069] Therefore, given This can uniquely determine The shape of the skeleton and satisfying ξ m constraint.

[0070] To satisfy the dimensional constraint of the thickness ratio t, a normalization coefficient is introduced. And in The specific execution of the calibration is as follows: and define The normalized half-thickness distribution of the leading edge segment is then written as: .in, Used to calibrate the maximum half-thickness in the unit chord length coordinate system. This allows for the automatic satisfaction of the thickness ratio constraint under a unit chord length.

[0071] To further satisfy the leading edge bluntness parameter R * The constraints require the establishment of R * The mapping relationship with the leading edge segment parameters. Because... hour The dominant term is Substitute The equivalent radius of the circular arc corresponding to the local curvature of the leading edge can be obtained, thus yielding:

[0072] , ,

[0073] The constant C is determined by the following method: exist of The dominant term is written as The equivalent radius of the leading edge arc satisfies Combining and It can be deduced .

[0074] In this implementation, constant The effective leading edge coefficient of the target is obtained from the input parameters t=0.6 and R*=0.2. :

[0075]

[0076] Furthermore, in order to solve for the equation... of The preferred approach is to use a fixed-point iteration method: and take the initial value. After iterative convergence, the result can be obtained. Corresponding to verify This meets the requirements. Therefore, the half-thickness distribution of the leading edge segment... and its derivative Completely determined, and simultaneously satisfying ξ m t and R * Three types of constraints.

[0077] S3, Obtain the leading edge segment in S2. Then, by specifying the splicing position and constructing the trailing edge curve, the segmented splicing of the leading and trailing edge segments is achieved, thereby ensuring the continuity of the contour and tangential direction at the splicing point. Specifically, the splicing position parameters are first... Determine the chordal position of the splicing point, and then... The half-thickness and tangent slope of the splice point are calculated using its derivative. Therefore, the splice point can be used as the starting point of the trailing edge curve, and a rational quadratic Bézier curve satisfying the closure and tangential conditions can be constructed accordingly.

[0078] Preferably, the dimensionless position ξ of the splicing point chord. s Defined in linear extrapolation form as:

[0079]

[0080] In this embodiment, ξ m =0.3, Find ξ s =0.37.

[0081] ξ s Substituting the distribution of the leading edge segment obtained in step S2 And its derivative, to obtain the half thickness and tangent slope of the splicing point, respectively denoted as ,as well as Under the input parameters in this embodiment, the calculation is as follows: =0.292, s=−0.212, therefore, in the unit chord length coordinate system, the splicing point can be recorded as... .

[0082] Furthermore, when the chord length is calibrated to the actual chord length L, the coordinates of the splicing point in the actual coordinate system are determined by x=Lξ. Scaling, therefore Similarly, in the unit chord length coordinate system, the tail edge closure point is taken as T = (1, 0), and its corresponding actual coordinates are T. s = (L, 0).

[0083] To construct a rational quadratic Bézier curve for the trailing edge segment, control points H need to be determined such that the tangential direction of the trailing edge curve at point P is consistent with that of the leading edge segment, and it closes at point T with the tangential direction perpendicular downwards. Preferably, the tangential direction vector at point P is taken as... Let the tangential direction vector at point T be... Then the control point H can be represented as the intersection of two parametric lines:

[0084]

[0085] Where α and β represent the distance from point P, respectively. s Starting from T along their respective direction vectors t P With t T Scalar displacement coefficient.

[0086] From the above equation, the explicit inline expression for the control point can be obtained as follows: Substituting the parameter values ​​from this embodiment, we get H = (1, 0.159).

[0087] The trailing edge segment is expressed using a rational quadratic Bézier expression, with parameters set to... Weighting factor w t Using P, H, and T as control points, the upper half profile curve of the trailing edge segment is obtained as follows:

[0088]

[0089] In this embodiment, w is taken t =0.9, the curve satisfies C(0)=P, C(1)=T, and the tangential direction at u=0 is perpendicular to t. P Consistent, at u=1, tangentially parallel to t T This consistency ensures that the curve and the leading edge segment are geometrically continuous at the splicing point, and guarantees a smooth transition at the closing point of the trailing edge, avoiding sharp geometric features.

[0090] To form a closed cross-sectional profile, it is preferable to mirror the upper half of the profile about the chord to obtain the lower half. Furthermore, the lower half can be obtained by using a fixed or adaptive step size. and Discrete sampling is used to output a closed contour point set; when the chord length is calibrated to L, all points are processed according to... or Map to the actual coordinate system.

[0091] S4. This step aims to calculate the area and perform similar scaling on the closed cross-sectional profile obtained in the unit chord length coordinate system to ensure that it meets the target cross-sectional area A. tar And the chord length limit L limit Geometric constraints are checked; if they are not satisfied, parameter updates are triggered, and the iteration backtracks to form a closed loop. Specifically, the contour area A per unit chord length is first calculated. unitThe actual chord length L is then obtained using a similarity scaling factor, and the contour coordinates are scaled as a whole to obtain the actual cross-sectional dimensions. This ensures that the target area is met while maintaining the chord length limit L. limit The constraints were verified.

[0092] Preferably, the closed profile area A per unit chord length unit A can be obtained through numerical integration or the area formula of a discrete polygon. In this embodiment, numerical integration is used to obtain A. unit =0.426.

[0093] To meet the target cross-sectional area A tar The actual chord length L is obtained from the similarity scaling relationship:

[0094]

[0095] In this embodiment, A tar =0.58mm 2 The calculation yields L = 1.168 mm. Then, for any point on the unit chord length profile... or Scaling will give you the actual coordinate profile.

[0096] After obtaining L, the maximum thickness and the equivalent radius of the leading edge can be further obtained as dimensionless quantities. In this embodiment, , .

[0097] The constraints were checked; in this embodiment, L = 1.168 mm. <L limit =1.8mm, the upper limit of the chord length satisfies the constraint. If L>L limit The parametric design program will automatically trigger the constraint correction and parameter search process, automatically searching for the splicing position parameters. or / and weighting factor w t Repeat steps S3 to S4 until L ≤ L is satisfied. limit Constraints.

[0098] To verify the ability of the above-mentioned parametric design method to accurately control the geometry of the segmented composite airfoil section support under multiple geometric constraints and its automatic parametric generation capability, this embodiment focuses on the position ξ at the maximum constraint thickness. m =0.3, target cross-sectional area A tar =0.58mm 2 Remain unchanged, initialize splicing position parameters =0.1, trailing edge weighting factor w t Under the condition of 0.9, the thickness ratio t is continuously varied in the range of 0.3 to 0.8, and the corresponding closed section profiles are generated in batches accordingly. Figure 4The segmented composite airfoil profile curves obtained under different t values ​​are shown. The results show that this method can achieve stable and continuous control over the cross-sectional thickness distribution and trailing edge convergence morphology using a single parameter t, ensuring the consistency and controllability of cross-sectional generation while satisfying various constraints.

[0099] Based on this, to further illustrate the synergistic enhancement effect of the BCC lattice sandwich regenerative cooling channel of the segmented composite airfoil section support rod in terms of flow and heat transfer, this embodiment adopts a method based on... The CFD numerical simulation method of the turbulence model is used to compare the channel schemes corresponding to different thickness ratios t under the same geometric and operating boundary conditions. Meanwhile, to provide a unified performance reference, this embodiment introduces a BCC lattice sandwich regenerative cooling channel with a conventional cylindrical cross-section support as the benchmark channel under the same calculation settings.

[0100] Figure 5 The variation of channel pressure drop with thickness ratio t is shown; Figure 6 The variation of the average Nusselt number with thickness ratio t is shown; Figure 7 The composite coefficients calculated with reference to the aforementioned reference channel are shown. The variation of thickness ratio t with the thickness ratio is shown in each figure. The corresponding index of the benchmark channel is used as the reference line in each figure, so as to realize the quantitative comparison and trend analysis of different cross-section schemes under the same evaluation system.

[0101] Taking t=0.6 as an example, under the CFD calculation conditions and boundary conditions of this embodiment, the temperature cooling effect of the BCC lattice sandwich regenerative cooling channel of the segmented composite airfoil section support is comparable to that of the conventional cylindrical section support. With comparable cooling effects, the pressure drop of the BCC lattice sandwich regenerative cooling channel of the segmented composite airfoil section support is reduced by 49.07% compared to the reference channel, and the average Nusselt number is increased by 16.89%. Furthermore, calculated using the conventional cylindrical section support's BCC lattice sandwich regenerative cooling channel as a reference, the comprehensive coefficient of the segmented composite airfoil section support's BCC lattice sandwich regenerative cooling channel is 1.42, an improvement of 42% compared to the reference channel. This demonstrates that the parametric design method can not only achieve stable generation and controllable adjustment of the segmented composite airfoil section, but also significantly reduce flow resistance and improve heat transfer capacity while maintaining similar cooling effects at a typical parameter t=0.6. This reflects the synergistic enhancement advantage of the lattice sandwich regenerative cooling channel of the segmented composite airfoil section support in terms of flow and heat transfer performance.

[0102] The above are all preferred embodiments of this application and are not intended to limit the scope of protection of this application. Therefore, all equivalent changes made in accordance with the structure, shape and principle of this application should be covered within the scope of protection of this application.

Claims

1. A lattice-core regenerative cooling channel, comprising an outer wall, an inner wall, and a lattice-core layer disposed between the outer wall and the inner wall; the lattice-core layer is composed of multiple rod cells, each rod cell being composed of one or more support rods; characterized in that: The cross section of the support rod is a segmented composite airfoil section. The edge profile of the segmented composite airfoil section is formed by splicing the leading edge segment and the trailing edge segment. The leading edge segment and the trailing edge segment are spliced ​​together with first-order geometric continuity at the splicing point. The shape of the leading edge segment is defined by a half-thickness distribution function, which is constructed based on the classical polynomial method of NACA airfoil thickness distribution; the trailing edge segment is constructed by a rational quadratic Bézier curve.

2. The dot matrix sandwich regenerative cooling channel according to claim 1, characterized in that: The rod system cells are one or more combinations of body-centered cubic (BCC) cells, Kagome cells, Kelvin cells, octahedral truss cells, tetrahedral truss cells, pyramidal truss cells, and single-column cells, or topologically equivalent variants of the above cells; the rod system cells are arranged in a staggered array, and the orientation of the segmented composite airfoil-shaped section support rods is as follows: the trailing edge section faces the incoming flow direction, and the leading edge section faces away from the incoming flow direction.

3. A parametric design method for segmented composite airfoil-shaped cross-sections, used to design the cross-sectional shape of the support rod in a lattice sandwich regenerative cooling channel as described in claim 1, characterized in that, Includes the following steps: S1. Input the target parameters for the cross-section design. The target parameters include at least the thickness ratio t and the location of the maximum thickness ξ. m chord length limit L limit Leading edge bluntness parameter R * and the target cross-sectional area A tar , where t=D m / L,D m Where L is the maximum thickness and L is the chord length; ξ m =x m / L,x m R represents the chordal coordinate of the location of maximum thickness; * =R n / D m R n Set the equivalent radius of the leading edge; simultaneously, initialize the starting position parameters of the rational quadratic Bézier curve of the trailing edge segment. With curve weighting factor w t ;in Used to determine the location of the splicing point, w t Used to adjust the fullness of the tail edge and the rate of contraction; S2, Based on chord-dimensional coordinates Construct the half-thickness distribution function of the symmetrical cross section Forward edge amplitude coefficient To actively control the quantity, the polynomial skeleton of the NACA thickness distribution function of the leading edge segment is reconstructed under constraints, so that the leading edge segment is fixed at the position of maximum thickness as ξ. m Under the premise of establishing R * With effective leading edge coefficient The constraints between them are determined through iterative solutions. This allows the leading edge segment to simultaneously satisfy t and ξ. m With R * Constraints; S3, Leading edge segment at splicing point ξ s Smooth transition to the tail edge section, splicing point position ξ s From parameters Determined and satisfied ; in ξ=ξ s The coordinates and tangent direction of the splicing point are obtained, and the control point H of the tail edge segment is determined by the intersection of the splicing point tangent and the tail edge closing direction. A rational quadratic Bézier tail edge curve that is spliced ​​with the leading edge segment at the splicing point in a first-order geometric continuous manner is generated. S4. Combine the leading edge segment and the trailing edge segment to form a closed profile, calculate the cross-sectional area of ​​the closed profile under the condition of unit chord length, and perform the operation based on the target area A. tar The constraint-based similarity scaling makes the final cross-sectional area equal to A. tar And the chord length limit L limit Constraints are validated; if a constraint is not met, parameters κ and / or w are automatically searched. t Then re-execute the contour generation and scaling steps until the output of segmented composite airfoil cross-sectional geometry data that satisfies all parameter constraints.

4. The parametric design method for a segmented composite airfoil section according to claim 3, characterized in that: In step S2, the NACA thickness distribution function of the leading edge segment is... The polynomial skeleton is constrained and reconstructed to construct an unnormalized thickness skeleton function: in The preset NACA thickness distribution function The skeleton constant; to ensure the thickness skeleton function is within... At the extreme value, for Applying extreme value constraints Thus, the coefficients of the linear terms are explicitly determined. for: 。 5. The parametric design method for a segmented composite airfoil section according to claim 4, characterized in that: In step S2, an amplitude normalization coefficient is introduced. A normalized thickness skeleton function is defined to automatically satisfy the thickness ratio constraint: 。 6. The parametric design method for a segmented composite airfoil section according to claim 5, characterized in that: The effective leading edge coefficient in step S2 is defined as follows: And establish the leading edge bluntness parameter R * and Display constraints: The constant C is determined by the analytical relation of the NACA leading edge curvature; it is solved through fixed-point iteration or equivalent one-dimensional iteration, making... Satisfy objective R * Thus determining the unique .

7. The parametric design method for a segmented composite airfoil section according to claim 6, characterized in that: The splicing point position in step S3 satisfies: And in ξ=ξ s Take half the thickness at the splicing point and the slope of the tangent line s: 。 8. The parametric design method for a segmented composite airfoil section according to claim 7, characterized in that: In step S3, the control point H of the trailing edge segment is determined through the following geometric construction: Define the splice point. With the trailing edge point Take the tangential direction of the splicing point With the tail edge closing direction Solve for the equation of the line: Where α and β represent the distance from point P, respectively. s Starting from T along their respective direction vectors t P With t T The scalar displacement coefficient; the intersection point H is the control point of the rational quadratic Bézier curve.

9. The parametric design method for a segmented composite airfoil section according to claim 8, characterized in that: The trailing edge segment mentioned in step S3 is composed of a weighting factor w. t The rational quadratic Bézier curve is constructed as follows: The control point H is located on the tangent direction of the splicing point, so that the tail edge segment and the leading edge segment can be spliced ​​together in a first-order geometric continuous manner at u=0.

10. The parametric design method for a segmented composite airfoil section according to claim 3, characterized in that: In step S4, based on the target area A tar The similarity scaling of the constraints includes: within a unit chord length L unit Given a constant value of 1, a closed contour is discretized and its area A is calculated. unit The chord length is determined by the similarity scaling relationship: The closed contour is then scaled proportionally to ensure that the final cross-sectional area satisfies A=A tar When the chord length L > L after scaling limit At that time, within the preset range or / and w t Perform automatic search and iterative updates to ensure that the scaled closed contour simultaneously satisfies the chord length. Geometric constraints.

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