Aircraft aerodynamic force / thermal environment-material ablation response coupling calculation method
By employing a coupled calculation method that iterates through the entire trajectory, the problem of low accuracy and insufficient efficiency of existing algorithms in hypersonic vehicles is solved, achieving higher accuracy and efficiency in material ablation response simulation, thus ensuring the safety and lifespan of the vehicle.
Patent Information
- Application Number
- CN202511491960.7
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- Filing Date
- 2025-10-20
- Publication Date
- 2026-02-10
- Estimated Expiration
- 2045-10-20
AI Technical Summary
Existing aerodynamic/thermal environment-material ablation response coupling algorithms for aircraft suffer from low accuracy and insufficient computational efficiency. Especially under hypersonic flight conditions, existing algorithms fail to effectively consider the bidirectional coupling effect of material ablation response on the near-wall flow state and aerodynamic shape of the aircraft.
A coupled calculation method of iterative calculation across the entire trajectory is adopted. The aerodynamic and thermal environment information is processed by interpolation algorithm to realize the bidirectional coupling between the aerodynamic/thermal environment of the aircraft and the ablation response of the material. The thermal environment and thermal response of each trajectory point are calculated iteratively until the convergence condition is reached.
It improves the accuracy and efficiency of calculating the ablation response of materials in the entire ballistic trajectory of aircraft, enabling more accurate simulation of the material ablation process, reducing redundant thermal protection, and ensuring the safety and service life of aircraft.
Smart Images
Figure CN120974641B_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The present specification relates to the technical field of aerospace, and particularly relates to a method for calculating aerodynamic force / thermal environment-material ablation response coupling of a spacecraft. BACKGROUND
[0002] When a spacecraft flies at a hypersonic speed, the surrounding gas and the surface of the spacecraft rub fiercely, and the kinetic energy of the gas is largely converted into heat energy, causing the temperature of the surface of the spacecraft to rise sharply. For example, in a typical hypersonic flight working condition, the temperature of the stagnation point region of the spacecraft can reach 2000K or even higher. Such an extremely high temperature environment will generate a great thermal load on the thermal protection system of the spacecraft, and then cause ablation phenomena of the thermal protection material, including thermal chemical ablation of the surface of the material, high-temperature sublimation of the material, and pyrolysis of the material. The ablation process not only changes the geometric shape and physical properties of the material, but also can cause structural deformation, which seriously threatens the flight safety and service life of the spacecraft.
[0003] Existing aerodynamic force / thermal environment-material ablation response coupling algorithms mostly adopt a loose coupling strategy based on flow field and solid heat transfer partition calculation. However, the existing loose coupling algorithm strategy is mostly one-way coupling between the aerodynamic force / thermal environment and the material ablation response calculation, that is, the aerodynamic force / thermal environment corresponding to the ablation time is considered to be constant during the material ablation response calculation, and the influence of the material ablation response on the near-wall flow state and aerodynamic shape of the spacecraft is not considered. Although this strategy is simple, it violates the coupling principle of mutual correlation and mutual influence between the aerodynamic force / thermal environment and the material ablation response, so the algorithm accuracy is low, and redundancy thermal protection of the spacecraft is often caused. The existing loose coupling algorithm considering the two-way coupling relationship between the aerodynamic force / thermal environment and the material ablation response mostly adopts a calculation strategy of “single coupling and point-by-point advancement”, that is, the aerodynamic force / thermal environment at time T0 is calculated first and used as input for the material ablation response calculation, and the material ablation response calculation is advanced to the next ablation time T1, the material ablation response at time T1 is used as input for the calculation of the aerodynamic force / thermal environment at time T1, and the aerodynamic force / thermal environment at time T1 is used as input for the calculation of the material ablation response and advanced to time T2, and so on until the ablation termination time. Although this method considers the two-way coupling between the aerodynamic force / thermal environment and the material ablation response, the aerodynamic force / thermal environment is considered to be constant in the material ablation response calculation between adjacent two times, so the algorithm accuracy is insufficient; at the same time, the time step of the algorithm cannot be too large, so the calculation efficiency is low and the algorithm is not suitable for full-trajectory aerodynamic force / thermal environment and material ablation response coupling calculation of a spacecraft during reentry. SUMMARY
[0004] To address the aforementioned shortcomings in existing technologies, this invention provides a coupled calculation method for aerodynamic / thermal environment and material ablation response of aircraft, which solves the problem of difficulty in calculating the material ablation response of aircraft trajectories in real time and accurately.
[0005] To achieve the aforementioned objectives, the technical solution adopted by this invention is: a coupled calculation method for aerodynamic / thermal environment and material ablation response of an aircraft, comprising:
[0006] S1. Discretize the reentry trajectory of the spacecraft into multiple trajectory points, obtain the trajectory parameters of each trajectory point, and determine the temperature of the heat shield material and the wall temperature of the spacecraft at the start of the trajectory based on the trajectory parameters.
[0007] S2. Calculate the initial external flow field of each ballistic point based on the ballistic parameters of each ballistic point to form the initial thermal environment data along the ballistic trajectory.
[0008] S3. Using the initial thermal environment data as boundary conditions, and interpolating the wall heat flux and wall pressure between adjacent time points, the initial material thermal response calculation results are obtained.
[0009] S4. Based on the initial material thermal response calculation results, generate the material ablation response wall boundary conditions, transfer the material ablation retreat amount to the mesh deformation program for mesh update, and calculate the flow field at each ballistic point according to the ballistic parameters to obtain the thermal environment data at each ballistic point.
[0010] S5. Using thermal environment data as boundary conditions, interpolate the wall heat flux and pressure between adjacent time points to calculate the material thermal response and obtain the thermal response calculation results at each ballistic point.
[0011] S6. Based on the thermal environment data obtained in S4 and the thermal response calculation results obtained in S5, determine whether the relative deviation of heat flow between the Mth coupling iteration and the (M-1)th coupling iteration exceeds the threshold. If so, return to S4; otherwise, the coupling calculation of the aerodynamic / thermal environment-material ablation response of the aircraft's full ballistic trajectory has reached the convergence condition, the calculation is stopped, and the coupling calculation results of the material ablation response are obtained.
[0012] Further, S2 includes:
[0013] The wall temperature at the start of the trajectory is T0. w 0=T ini =300K, ablation ejection amount m dot 0=0, the wall temperature at times T1~TN is all set to T w 1~N = Radiation equilibrium wall temperature condition, excitation quantity m dot Set all numbers from 1 to N to 0;
[0014] Using an aerodynamic / thermal environment CFD calculation program, the initial external flow field of each ballistic point is calculated based on the ballistic parameters of each ballistic point, forming the initial thermal environment data along the ballistic trajectory. The thermal environment data at the ballistic start time T0 remains unchanged during the coupled iteration process.
[0015] Further, S3 includes:
[0016] Set the temperature of the vehicle's thermal protection material at the ballistic start point T0 as T. s 0=T ini =300K;
[0017] Using the initial thermal environment data as boundary conditions, and based on the temperature of the spacecraft's heat shield material at the start of the ballistic trajectory, the ablation response calculation program is used to interpolate the wall heat flux and wall pressure between adjacent time points to calculate the ablation response of the material, thus obtaining the initial material thermal response calculation result. The initial material thermal response calculation result remains unchanged during the coupled iteration process.
[0018] Furthermore, the step of using the initial thermal environment data as boundary conditions to interpolate the wall heat flux and wall pressure between adjacent time points, calculating the ablation response of the material, and obtaining the initial material thermal response calculation results includes:
[0019] The thermal environment data, i.e., the heat flux Q of the aircraft wall, is received from the initial flow field calculation by the material ablation response calculation program. w Wall pressure P w As boundary conditions, and by interpolating the wall heat flux and wall pressure between adjacent time points, the ablation response of the material is calculated, yielding the ablation wall temperature, ablation component mass fraction, ablation mass flow rate, and ablation retreat at each trajectory point. The material wall temperature T at the trajectory start time t0 is also calculated. w 0=T ini =300K, ejection mass m dot 0=0, mass fraction of wall ablation component S0=0, ablation retreat amount a bl 0 = 0, and remains unchanged in subsequent iterations.
[0020] By considering the bidirectional coupling between the aerodynamic / thermal environment of the aircraft and the ablation response of the material, and by using an interpolation algorithm to interpolate the aerodynamic and thermal information between adjacent ballistic points, the calculation of the ablation response of the material is closer to the result of the tightly coupled calculation. Therefore, the coupling algorithm provided by this invention has higher calculation accuracy than the existing unidirectional coupling algorithm and single coupling algorithm.
[0021] Further, S4 includes:
[0022] Based on the initial material ablation response calculation results, the material ablation response wall boundary conditions are generated, and the material ablation retreat is transferred to the mesh deformation program for mesh update. Based on the mesh update results, the flow field of each ballistic point is recalculated according to the ballistic parameters to obtain the thermal environment data of each ballistic point at any given time.
[0023] Furthermore, the process of updating the computational grid specifically includes:
[0024] The aerodynamic / thermal environment CFD calculation program receives the initial material thermal response calculation results, namely ablation wall temperature, mass fraction of ablation components on the wall, ablation mass flow rate, and ablation retreat, generates the material ablation response wall boundary conditions, and transmits the material ablation retreat to the mesh deformation program for mesh update calculation.
[0025] Further, S5 includes:
[0026] The material ablation response calculation program receives thermal environment data, namely wall heat flux and wall pressure, as boundary conditions and performs interpolation processing on the wall heat flux and pressure between adjacent time points to calculate the material thermal response and obtain the thermal response calculation results at each ballistic point.
[0027] Further, S6 includes:
[0028] The aerodynamic / thermal environment CFD calculation program receives the material thermal response calculation results of the M-1th iteration, namely the ablation wall temperature, the mass fraction of the wall ablation component, the ablation mass flow rate, and the ablation retreat, generates the wall boundary conditions of the material ablation response, and transmits the material ablation retreat to the mesh deformation program for calculation mesh update. It also calculates the flow field of each ballistic point according to the ballistic parameter conditions to obtain the thermal environment data of each ballistic point at the Mth iteration.
[0029] The material ablation response calculation program receives the thermal environment data of the Mth iteration, namely the wall heat flow and wall pressure, as boundary conditions and interpolates the wall heat flow and pressure between adjacent time points to calculate the material thermal response and obtain the thermal response calculation results of each ballistic point in the Mth iteration.
[0030] The relative deviation of heat flow between the Mth iteration and the (M-1)th iteration is analyzed. When the relative deviation of heat flow between the Mth iteration and the (M-1)th iteration does not exceed the threshold, the coupling calculation of the aerodynamic / thermal environment-material ablation response of the entire ballistic trajectory of the aircraft reaches the convergence criterion, the iteration is stopped, the coupling calculation result of material ablation response is obtained, and the coupling calculation of aerodynamic / thermal environment-material ablation response of the aircraft is completed.
[0031] The beneficial effects of the present invention are as follows: The present invention provides a coupling calculation method for aerodynamic / thermal environment-material ablation response of an aircraft. (1) It adopts a convergence method of iterative calculation of the entire trajectory, that is, it can realize the coupling calculation of aerodynamic / thermal-material ablation response of all ballistic points of the entire trajectory in one go, which can achieve higher calculation efficiency; (2) By coupling and iterating the thermal environment data and thermal response calculation results of each ballistic point, the coupling calculation of material ablation response of the aircraft trajectory is realized, which makes the construction of material ablation response coupling edge more comprehensive and has higher coupling accuracy. Attached Figure Description
[0032] This specification will be further described by way of exemplary embodiments, which will be described in detail with reference to the accompanying drawings. These embodiments are not limiting; in these embodiments, the same reference numerals denote the same structures, wherein:
[0033] Figure 1 This is an exemplary flowchart illustrating a coupled calculation method for aerodynamic / thermal environment-material ablation response of an aircraft, according to some embodiments of this specification.
[0034] Figure 2 This is an exemplary flowchart illustrating an iterative calculation method for the coupled aerodynamic / thermal-material ablation response of a hypersonic reentry vehicle throughout its entire trajectory, based on some embodiments of this specification.
[0035] Figure 3 This is an exemplary schematic diagram of the inputs and outputs of each calculation stage of a coupled calculation method for the full ballistic aerodynamic / thermal-material ablation response of a hypersonic reentry vehicle, as shown in some embodiments of this specification.
[0036] Figure 4 This is an exemplary schematic diagram of the reentry trajectory and computational grid division of the Stardust return capsule in the calculation working condition section, according to some embodiments of this specification.
[0037] Figure 5 This is an exemplary schematic diagram illustrating the calculation of heat flow on the ablation-iterative wall surface of each wheel at the T34 ballistic point of the Stardust return capsule, according to some embodiments of this specification.
[0038] Figure 6 This is an exemplary schematic diagram illustrating the calculation of the ablation iteration wall temperature of each wheel at the T34 ballistic point of the Stardust return capsule according to some embodiments of this specification;
[0039] Figure 7 This is an exemplary schematic diagram illustrating the calculation of heat flow on the ablation-iterative wall surface of each wheel at the T42 ballistic point of the Stardust return capsule, according to some embodiments of this specification.
[0040] Figure 8This is an exemplary schematic diagram illustrating the calculation of the ablation iteration wall temperature of each wheel at the T42 ballistic point of the Stardust return capsule according to some embodiments of this specification;
[0041] Figure 9 This is an exemplary schematic diagram illustrating the calculation of heat flow on the ablation-iterative wall surface of each wheel at the T48 trajectory point of the Stardust return capsule, according to some embodiments of this specification.
[0042] Figure 10 This is an exemplary schematic diagram illustrating the calculation of the ablation iteration wall temperature of each wheel at the T48 ballistic point of the Stardust return capsule according to some embodiments of this specification;
[0043] Figure 11 This is an exemplary schematic diagram illustrating the calculation of heat flow on the ablation-iterative wall surface of each wheel at the T54 ballistic point of the Stardust return capsule, according to some embodiments of this specification.
[0044] Figure 12 This is an exemplary schematic diagram illustrating the calculation of the ablation iteration wall temperature of each wheel at the T54 ballistic point of the Stardust return capsule according to some embodiments of this specification;
[0045] Figure 13 This is an exemplary schematic diagram illustrating the calculation of heat flow on the ablation-iterative wall surface of each wheel at the T60 trajectory point of the Stardust return capsule, according to some embodiments of this specification.
[0046] Figure 14 This is an exemplary schematic diagram illustrating the calculation of the ablation iteration wall temperature of each wheel at the T60 ballistic point of the Stardust return capsule according to some embodiments of this specification;
[0047] Figure 15 This is an exemplary schematic diagram illustrating the calculation of heat flow on the ablation-iterative wall surface of each wheel at the T66 ballistic point of the Stardust return capsule, according to some embodiments of this specification.
[0048] Figure 16 This is an exemplary schematic diagram illustrating the calculation of the ablation iteration wall temperature of each wheel at the T66 ballistic point of the Stardust return capsule according to some embodiments of this specification;
[0049] Figure 17 This is an exemplary schematic diagram illustrating the calculation of heat flow on the ablation-iterative wall surface of each wheel at the T76 ballistic point of the Stardust return capsule, according to some embodiments of this specification.
[0050] Figure 18 This is an exemplary schematic diagram illustrating the calculation of the ablation iteration wall temperature of each wheel at the T76 ballistic point of the Stardust return capsule according to some embodiments of this specification;
[0051] Figure 19This is an exemplary schematic diagram showing the ablation profile busbar receding over time according to some embodiments of this specification. Detailed Implementation
[0052] The specific embodiments of the present invention are described below to enable those skilled in the art to understand the present invention. However, it should be understood that the present invention is not limited to the scope of the specific embodiments. For those skilled in the art, various changes are obvious as long as they are within the spirit and scope of the present invention as defined and determined by the appended claims. All inventions utilizing the concept of the present invention are protected.
[0053] Example
[0054] Figure 1 This is an exemplary flowchart illustrating a coupled calculation method for aircraft aerodynamic / thermal environment-material ablation response, based on some embodiments of this specification. Figure 1 As shown, the process includes the following steps.
[0055] In some embodiments, the aerodynamic / thermal environment-material ablation response coupling calculation method for the full trajectory of the Stardust reentry capsule configuration can be used to perform aerodynamic / thermal environment-material ablation response coupling calculations on the Stardust reentry capsule configuration.
[0056] S1: Discretize the reentry trajectory of the spacecraft into multiple trajectory points, obtain the trajectory parameters of each trajectory point, and determine the temperature of the heat shield material and the wall temperature of the spacecraft at the start of the trajectory based on the trajectory parameters.
[0057] The ballistic parameters at each trajectory point are discrete parameters of the hypersonic reentry vehicle's entire trajectory. For example, the trajectory parameters may include altitude, Mach number, and angle of attack, and the trajectory may include a portion of the Stardust reentry capsule's reentry trajectory. Specifically, the calculation conditions (partial reentry trajectory) and mesh generation for the Stardust reentry capsule are as follows: Figure 4 As shown.
[0058] In some embodiments, the reentry trajectory of the aircraft can be discretized into N+1 trajectory points (i.e., T0~TN) to obtain the trajectory parameters of the N+1 trajectory points and the temperature T of the aircraft's heat shield material at the trajectory start time T0. S 0 and wall temperature T w 0.
[0059] In some embodiments, the wall ablation ejection amount m at the start of the trajectory can be considered as... dot 0 is negligible, that is, mdot0 = 0.
[0060] S2: Calculate the initial external flow field of each ballistic point based on the ballistic parameters of each ballistic point, forming the initial thermal environment data along the ballistic trajectory, and the thermal environment data at the start of the ballistic trajectory T0 remains unchanged during the coupling iteration process.
[0061] The initial thermal environment data along the trajectory is the thermal environment data of the initial external flow field at each trajectory point.
[0062] In some embodiments, the wall temperature T at the start of the trajectory can be... w 0=T ini =300K, ablation ejection amount m dot 0=0, the wall temperature at times T1~TN is all set to T w 1~N=300K, ejection quantity m dot All numbers from 1 to N are set to 0. Using the aerodynamic / thermal environment CFD calculation program, the initial external flow field of each ballistic point is calculated based on the ballistic parameters of each ballistic point, forming the initial thermal environment data along the ballistic trajectory. The thermal environment data at the ballistic start time T0 remains unchanged during the coupled iteration process.
[0063] In some embodiments, the wall temperature T at the start of the trajectory can be... w 0=T ini =300K, ablation ejection amount m dot 0=0, the wall temperature at times T1~TN is all set to T w 1~N=300K (or radiation equilibrium temperature), excitation quantity m dot All numbers from 1 to N are set to 0; the initial external flow field at each trajectory point is calculated by the aerodynamic / thermal environment CFD calculation program based on the trajectory parameters at each trajectory point, forming the initial thermal environment data (t0, Q) along the trajectory. w 0, P w 0) ini 、(t1、Q w 1. P w 1) ini ... (tN, Q) w N, P w N) ini , where (tN, Q) w N, P w N) ini This represents the initial wall heat flux and wall pressure of the reentry capsule at the Nth trajectory point, and the thermal environment data (t0, Q) at the start of the trajectory at time T0. w 0, P w 0) ini It remains unchanged in subsequent iterations.
[0064] S3: The initial thermal environment data is used as the boundary condition, and the wall heat flux and wall pressure between adjacent time points are interpolated to obtain the initial material thermal response calculation results. The initial material thermal response calculation results remain unchanged during the coupled iteration process.
[0065] The initial material thermal response calculation results are the calculation results of the initial ablation response of the material.
[0066] The external flow field of a ballistic trajectory is the external flow field data of the aircraft's trajectory.
[0067] In some embodiments, the temperature of the vehicle's thermal protection material at the ballistic start time T0 can be set to T. s 0=T ini =300K; Using the initial thermal environment data along the trajectory as boundary conditions, based on the temperature of the spacecraft's heat shield material at the start of the trajectory, the wall heat flux and wall pressure between adjacent time points are interpolated using a material ablation response calculation program to calculate the ablation response of the material and obtain the initial material thermal response calculation results.
[0068] In some embodiments, the temperature of the vehicle's thermal protection material at the ballistic start time T0 can be set to T. s 0=T ini =300K; The thermal environment data, i.e., the heat flux Q of the aircraft wall, is received from the initial flow field calculation program by the material ablation response calculation program. w Wall pressure P w As boundary conditions, and by interpolating the wall heat flux and wall pressure between adjacent time points, the ablation response of the material is calculated, yielding the ablation wall temperature, ablation component mass fraction, ablation mass flow rate, and ablation retreat (t1, T) at each trajectory point. w 1, S1, m dot 1. a bl 1) ini ... (tN, T) w N, SN, m dot N, a bl N) ini , where (tN, T) w N, SN, m dot N, a bl N) ini , respectively, represent the ablation wall temperature, ablation component mass fraction, ablation mass flow rate, and ablation retreat amount calculated from the ablation response of the return capsule at time N, and the material wall temperature T at the start of the trajectory t0. w 0=T ini =300K, ejection mass m dot 0=0, mass fraction of wall ablation component S0=0, ablation retreat amount a bl 0 = 0, and remains unchanged in subsequent iterations.
[0069] S4: Based on the initial material thermal response calculation results, generate the material ablation response wall boundary conditions, transfer the material ablation retreat amount to the mesh deformation program for mesh update, and calculate the flow field at each ballistic point according to the ballistic parameters to obtain the thermal environment data at each ballistic point.
[0070] In some embodiments, based on the initial material ablation response calculation results, material ablation response wall boundary conditions are generated, and the material ablation retreat is transmitted to the mesh deformation program for mesh update calculation; based on the mesh update results, the flow field of each ballistic point is recalculated according to the ballistic parameters to obtain the thermal environment data of each ballistic point at any given time.
[0071] In some embodiments, such as Figure 2 As shown, the initial material thermal response calculation results, namely ablation wall temperature, mass fraction of ablation components on the wall surface, ablation mass flow rate, and ablation retreat, can be received by the aerodynamic / thermal environment CFD calculation program. This generates the material ablation response wall boundary conditions, and the material ablation retreat is transmitted to the mesh deformation program for mesh updating. Then, the flow field at each trajectory point is recalculated based on the ballistic parameters, obtaining the thermal environment data (t1, Q) at each trajectory point during the first iteration. w 1. P w 1) iter1 ...(tN, Q) w N, P w N) iter1 .
[0072] In some embodiments, such as Figure 3 As shown, the aerodynamic / thermal environment CFD calculation program can receive the material thermal response calculation results from the first iteration, namely the ablation wall temperature, the mass fraction of ablation components on the wall surface, the ablation mass flow rate, and the ablation retreat, to generate the material ablation response wall boundary conditions. The material ablation retreat is then transferred to the mesh deformation program for mesh updating. Afterward, the flow field at each trajectory point is recalculated based on the ballistic parameters, yielding the thermal environment data (t1, Q) at each trajectory point in the second iteration. w 1. P w 1) iter2 ...(tN, Q) w N, P w N) iter2 .
[0073] S5: Using thermal environment data as boundary conditions, interpolate the wall heat flux and pressure between adjacent time points to calculate the material thermal response and obtain the thermal response calculation results at each trajectory point.
[0074] The thermal environment data at each ballistic point in the first iteration is the thermal environment-related data of the return capsule at the ballistic point after the first iteration. For example, the thermal environment data of the first iteration may include the wall heat flux and the wall pressure of the first iteration.
[0075] In some embodiments, such as Figure 2 As shown, the material ablation response calculation program can receive the thermal environment data from the first iteration, namely the wall heat flux and wall pressure, as boundary conditions, and interpolate the wall heat flux and pressure between adjacent time points to calculate the material thermal response, obtaining the thermal response calculation results for each trajectory point in the first iteration: ablation wall temperature, wall ablation component mass fraction, ablation mass flow rate, and ablation retreat (t1, T). w 1, S1, m dot 1. a bl 1) iter1 ... (tN, T) w N, SN, m dot N, a bl N) iter1 .
[0076] In some embodiments, the thermal environment data at each ballistic point can be used as boundary conditions to interpolate the wall heat flux and pressure between adjacent time points, calculate the material thermal response, and obtain the thermal response calculation results at each ballistic point.
[0077] In some embodiments, the material ablation response calculation program can receive the second iteration thermal environment data, namely wall heat flux and wall pressure, as boundary conditions and perform interpolation processing on the wall heat flux and pressure between adjacent time points to calculate the material thermal response, obtaining the thermal response calculation results at each trajectory point in the second iteration: ablation wall temperature, wall ablation component mass fraction, ablation mass flow rate, and ablation retreat amount (t1, T). w 1, S1, m dot 1. a bl 1) iter2 ... (tN, T) w N, SN, m dot N, a bl N) ite2 .
[0078] S6: Based on the thermal environment data and thermal response calculation results at each ballistic point, determine whether the relative deviation of heat flow between the Mth coupling iteration and the (M-1)th coupling iteration exceeds the threshold. If yes, return to S4; otherwise, stop the calculation when the coupling calculation of the aerodynamic / thermal environment-material ablation response of the entire ballistic trajectory of the aircraft reaches the convergence condition, obtain the coupling calculation result of the material ablation response, and complete the coupling calculation.
[0079] The thermal environment data obtained at each ballistic point moment through iteration is the thermal environment-related data of the return capsule at the ballistic point moment relative to the previous iteration process.
[0080] The thermal response calculation results obtained at each ballistic point are the results of the ablation response of the return capsule at the ballistic point relative to the previous iteration.
[0081] The relative deviation of heat flow is the deviation of heat flow calculated from thermal environment data between adjacent iterations.
[0082] In some embodiments, the processor can calculate the relative heat flux deviation based on the thermal environment data and thermal response calculation results at each ballistic point. For example, the formula for calculating the relative heat flux deviation can be: Relative heat flux deviation = |current heat flux - previous heat flux| / current heat flux; where current heat flux can include the thermal environment data and thermal response calculation results at each ballistic point in the current iteration, and previous heat flux can include the thermal environment data and thermal response calculation results at each ballistic point in the previous iteration.
[0083] The results of the coupled calculation of the material ablation response are the results of the coupled calculation of the ablation response of the return capsule at the final ballistic point.
[0084] In some embodiments, the aerodynamic / thermal environment CFD calculation program can receive the material thermal response calculation results of the (M-1)th iteration, namely the ablation wall temperature, the mass fraction of ablation components on the wall surface, the ablation mass flow rate, and the ablation retreat, generate the material ablation response wall boundary conditions, and transmit the material ablation retreat to the mesh deformation program for computational mesh update. The flow field at each trajectory point is then recalculated based on the ballistic parameter conditions to obtain the thermal environment data (t1, Q) at each trajectory point in the Mth iteration. w 1. P w 1) iterM ...(tN, Q) w N, P w N) iterM .
[0085] In some embodiments, the material ablation response calculation program can receive the thermal environment data of the Mth iteration, namely the wall heat flow and wall pressure, as boundary conditions, and perform interpolation processing on the wall heat flow and pressure between adjacent time points to calculate the material thermal response, obtaining the thermal response calculation results at each trajectory point in the Mth iteration: ablation wall temperature, wall ablation component mass fraction, ablation mass flow rate, and ablation retreat amount (t1, T). w 1, S1, m dot 1. a bl 1) iterM ... (tN, T) w N, SN, mdot N, a bl N) iterM .
[0086] In some embodiments, the relative deviation of heat flow between the Mth iteration and the (M-1)th iteration can be analyzed. When the relative deviation of heat flow between the Mth iteration and the (M-1)th iteration does not exceed 5%, the coupling calculation of the aerodynamic / thermal environment-material ablation response of the entire ballistic trajectory of the aircraft reaches the convergence criterion, the iteration is stopped, the coupling calculation result of the material ablation response is obtained, and the coupling calculation of the aerodynamic / thermal environment-material ablation response of the aircraft is completed.
[0087] In some embodiments, the initial CFD calculation results and the wall heat flux and wall temperature from each CFD-ablation coupled iteration are extracted and analyzed, such as... Figures 5 to 18 As shown. After four rounds of coupled iterations, the calculated results of the total wall heat flux and wall temperature for all working conditions all tended to stabilize. The maximum CFD heat flux error between the fourth and third iterations for each working condition was less than 5%, reaching the convergence criterion. Further, the ablation profile generatrix of the calculation model corresponding to each ballistic point working condition in the final convergence result (fourth iteration result) was extracted, and the results are shown below. Figure 19 As shown in the figure, the ablation profile of the model moves backward as the ablation time progresses, with the largest ablation at the stagnation point, approximately 5 mm; the ablation gradually decreases from the stagnation point to the shoulder of the return capsule.
[0088] In some embodiments of this specification, a method for coupled calculation of aerodynamic / thermal environment-material ablation response of an aircraft is provided. (1) A convergence method of iterative calculation of the entire trajectory is adopted, that is, the coupled calculation of aerodynamic / thermal-material ablation response of all ballistic points of the entire trajectory can be realized at one time, which can achieve higher calculation efficiency; (2) The coupled calculation of material ablation response of the aircraft trajectory is realized, which makes the construction of material ablation response coupling strip more comprehensive and has higher coupling accuracy.
Claims
1. A coupled calculation method for aerodynamic / thermal environment and material ablation response of an aircraft, characterized in that, include: S1. Discretize the reentry trajectory of the spacecraft into multiple trajectory points, obtain the trajectory parameters of each trajectory point, and determine the temperature of the heat shield material and the wall temperature of the spacecraft at the start of the trajectory based on the trajectory parameters. S2. Calculate the initial external flow field at each ballistic point based on the ballistic parameters of each ballistic point of the aircraft, and form the initial thermal environment data along the trajectory. At the start of the trajectory, the wall temperature at T0 is set to Tw0=Tini=300K and the ablation ejection quantity mdot0=0. At times T1~TN, the wall temperature is set to Tw1~N=radiation equilibrium wall temperature condition and the ejection quantity mdot1~N is set to 0. Using an aerodynamic / thermal environment CFD calculation program, the initial external flow field of each ballistic point is calculated based on the ballistic parameters of each ballistic point, forming the initial thermal environment data along the ballistic trajectory. The thermal environment data at the ballistic start time T0 remains unchanged during the coupled iteration process. S3. Using the initial thermal environment data as boundary conditions, and interpolating the wall heat flux and pressure, the initial material thermal response calculation results are obtained. Set the temperature of the aircraft's thermal protection material at the ballistic start point T0 to Ts0=Tini=300K; Using the initial thermal environment data as boundary conditions, interpolation is performed on the wall heat flux and wall pressure between adjacent time points to calculate the ablation response of the material, thereby obtaining the initial material thermal response calculation result. The initial material thermal response calculation result remains unchanged during the coupled iteration process. S4. Based on the initial material thermal response calculation results, generate the material ablation response wall boundary conditions, update the computational mesh, calculate the flow field at each trajectory point according to the ballistic parameters, and obtain the thermal environment data of each trajectory point at any time. S5. Using thermal environment data as boundary conditions, interpolate the wall heat flux and pressure to calculate the material thermal response and obtain the thermal response calculation results at each trajectory point. S6. Based on the thermal environment data obtained in S4 and the thermal response calculation results obtained in S5, determine whether the relative deviation of heat flow between the Mth coupling iteration and the (M-1)th coupling iteration exceeds the threshold. If so, return to S4. Otherwise, the coupled calculation of the aircraft's full-trajectory aerodynamic / thermal environment and material ablation response reaches the convergence condition, the calculation stops, and the results of the coupled material ablation response calculation are obtained; including: The aerodynamic / thermal environment CFD calculation program receives the material thermal response calculation results of the M-1th iteration, namely the ablation wall temperature, the mass fraction of the wall ablation component, the ablation mass flow rate, and the ablation retreat, generates the wall boundary conditions of the material ablation response, and transmits the material ablation retreat to the mesh deformation program for calculation mesh update. It also calculates the flow field of each ballistic point according to the ballistic parameter conditions to obtain the thermal environment data of each ballistic point at the Mth iteration. The material ablation response calculation program receives the thermal environment data of the Mth iteration, namely the wall heat flow and wall pressure, as boundary conditions and interpolates the wall heat flow and pressure between adjacent time points to calculate the material thermal response and obtain the thermal response calculation results of each ballistic point in the Mth iteration. The relative deviation of heat flow between the Mth iteration and the (M-1)th iteration is analyzed. When the relative deviation of heat flow between the Mth iteration and the (M-1)th iteration does not exceed the threshold, the coupling calculation of the aerodynamic / thermal environment-material ablation response of the entire ballistic trajectory of the aircraft reaches the convergence criterion, the iteration is stopped, the coupling calculation result of material ablation response is obtained, and the coupling calculation of aerodynamic / thermal environment-material ablation response of the aircraft is completed. Based on the thermal environment data and thermal response calculation results at each ballistic point, the relative deviation of heat flow is calculated. The formula for calculating the relative deviation of heat flow is: Relative deviation of heat flow = |current heat flow - previous heat flow| / current heat flow. Wherein, current heat flow includes the thermal environment data and thermal response calculation results at each ballistic point in this iteration, and previous heat flow includes the thermal environment data and thermal response calculation results at each ballistic point in the previous iteration.
2. The coupled calculation method for aerodynamic / thermal environment-material ablation response of an aircraft according to claim 1, characterized in that, The step of using the initial thermal environment data as boundary conditions to interpolate the wall heat flux and wall pressure between adjacent time points, calculating the ablation response of the material, and obtaining the initial material thermal response calculation results includes: The thermal environment data, i.e., the heat flux Q of the aircraft wall, is received from the initial flow field calculation by the material ablation response calculation program. w Wall pressure P w As boundary conditions, and by interpolating the wall heat flux and wall pressure between adjacent time points, the ablation response of the material is calculated, yielding the ablation wall temperature, ablation component mass fraction, ablation mass flow rate, and ablation retreat at each trajectory point. The material wall temperature T at the trajectory start time t0 is also calculated. w 0=T ini =300K, ejection mass m dot 0=0, mass fraction of wall ablation component S0=0, ablation retreat amount a bl 0 = 0, and remains unchanged in subsequent iterations.
3. The coupled calculation method for aerodynamic / thermal environment-material ablation response of an aircraft according to claim 1, characterized in that, S4 includes: Based on the initial material ablation response calculation results, the wall boundary conditions for the material ablation response are generated, and the material ablation retreat is transferred to the mesh deformation program for mesh update. Based on the mesh update results, the flow field at each ballistic point is recalculated according to the ballistic parameters to obtain the thermal environment data at each ballistic point.
4. The coupled calculation method for aerodynamic / thermal environment-material ablation response of an aircraft according to claim 3, characterized in that, The process of updating the computational grid specifically includes: The aerodynamic / thermal environment CFD calculation program receives the initial material thermal response calculation results, namely ablation wall temperature, mass fraction of ablation components on the wall, ablation mass flow rate, and ablation retreat, generates the material ablation response wall boundary conditions, and transmits the material ablation retreat to the mesh deformation program for mesh update calculation.
5. The coupled calculation method for aerodynamic / thermal environment-material ablation response of an aircraft according to claim 3, characterized in that, S5 includes: The material ablation response calculation program receives thermal environment data, namely wall heat flux and wall pressure, as boundary conditions and performs interpolation processing on the wall heat flux and pressure between adjacent time points to calculate the material thermal response and obtain the thermal response calculation results at each ballistic point.