Design method and design device for thickness of heat-proof coating for spaceflight launching device

Abstract

The invention provides a design method and a design device for the thickness of a heat-proof coating for a spaceflight launching device, and solves the technical problems that the design of the thickness of an existing heat-proof coating lacks technical means, and the reliability and the economical efficiency of coating arrangement are poor. The method comprises the steps of determining thermal environment parameters of a transmitting device; according to the thermal environment parameters and the ablation stripping temperature Tr, the coating ablation amount is determined through a heat-proof coating heat conduction model of the emission device, and the coating thickness l of initial design is determined according to the coating ablation amount. According to the coating thickness and the heat-proof coating heat conduction model, simulation is carried out to determine the overall temperature field of the emission device, and according to the overall temperature field of the emission device, the organic bottom layer temperature Td and the emission device temperature Tz in the ablation process are determined; and according to the comprehensive comparison result of the organic bottom layer temperature Td and the failure temperature Ts and the comprehensive comparison result of the device temperature Tz and the metal annealing temperature, iterative optimization of the coating thickness is formed, and the design thickness of the heat-proof coating is determined. And a reliable heat-proof coating thickness which is finely formulated according to a thermal environment can be formed.

Description

Technical Field

[0001] The present invention relates to the technical field of thermal protection, and particularly relates to a method and a device for designing the thickness of a heat insulation coating for a space launch device. Background Art

[0002] With the gradual increase of the rocket thrust, parameters such as the temperature, velocity, and pressure of the gas flow will increase accordingly. As a result, the ablation and erosion effects of the gas on the launch devices represented by the diversion device and the launch pad are more intense. To overcome the harmful phenomena such as ablation, a heat insulation coating is usually coated on the surface of the launch device to protect the device body.

[0003] In the prior art, the heat insulation coating mainly includes an inorganic surface layer and an organic bottom layer. The protection of the launch device is achieved through the high temperature and high pressure resistance characteristics of the inorganic surface layer, and the bonding between the inorganic surface layer and the device body is achieved through the strong bonding characteristics of the organic bottom layer. The heat of the high temperature gas flow is conducted to the heat insulation coating, and after the heat insulation coating absorbs part of the heat energy, the heat is conducted to the launch device body. After the launch device body absorbs part of the heat energy, the remaining heat is transferred to the air on the other side. The failure of the heat insulation coating is mainly reflected in that the temperature of the launch device body is too high during the launch process due to the insufficient thickness of the heat insulation coating, and the large-scale shedding of the inorganic surface layer due to the failure of the organic bottom layer, resulting in the direct ablation of the launch device body by the gas. Any failure of the heat insulation coating will reduce the launch safety.

[0004] Currently, the design of the thickness of the heat insulation coating for each model of the launch device is mainly based on the engineering use experience of the previous models of the device, lacking the theoretical support and refined design for predicting the reliability of the coating, and unable to quantify the influence of the thermal environment on the ablation characteristics of the inorganic surface layer and the bonding characteristics of the organic bottom layer. This makes the reliability of the coating guarantee lack technical guidance, and at the same time, it is impossible to balance economy and reliability. Summary of the Invention

[0005] In view of the above problems, the embodiments of the present invention provide a method and a device for designing the thickness of a heat insulation coating for a space launch device, which solve the technical problems that the existing design of the thickness of the heat insulation coating lacks technical means, and the reliability and economy of the coating setting are lacking.

[0006] The method for designing the thickness of the heat insulation coating for the space launch device in the embodiments of the present invention includes:

[0007] Determine the thermal environment parameters of the launch device;

[0008] According to the thermal environment parameters and the ablation peeling temperature T r Use the heat conduction model of the heat insulation coating of the launch device to determine the ablation amount of the coating, and determine the initial designed coating thickness l according to the ablation amount of the coating.

[0009] Determine the overall temperature field of the launch device according to the coating thickness and the simulation of the heat insulation coating heat conduction model, and determine the temperature T of the organic bottom layer during the ablation process according to the overall temperature field of the launch device d and the temperature T of the launch device z ;

[0010] According to the temperature T of the organic bottom layer d and the failure temperature T s , the device temperature T z and the comprehensive comparison result with the metal annealing temperature, an iterative optimization of the coating thickness is formed to determine the design thickness of the heat insulation coating

[0011] In an embodiment of the present invention, the thermal environment parameters include the ablation heat flux Q and the ablation time t

[0012] In an embodiment of the present invention, in the heat insulation coating heat conduction model, the heat conduction equation between the heat insulation coating and the launch device body is:

[0013]

[0014] where, C p is the specific heat capacity, with the unit J / (kg*K), ρ is the density, with the unit kg / m 3 , T is the temperature, with the unit K, x is the distance, with the unit m, t is the time, with the unit s, and λ is the thermal conductivity

[0015] The coating is ablated by the gas and changes with time. The coating thickness x changes with time, and the situation is:

[0016]

[0017] Assume that when the heat temperature T≥the ablation peeling temperature Tr, the coating is ablated and peeled

[0018] The initial temperature field t = 0s, T = T0, that is, at the initial moment, the temperatures of the device body and the protective coating are both at room temperature

[0019] In an embodiment of the present invention, during the ablation simulation process of the heat insulation coating heat conduction model, on the side where the high-temperature gas contacts the heat insulation coating, the heat conduction from the high-temperature gas to the coating is:

[0020]

[0021] where, Q is the heat of the gas, with a duration of t0

[0022] During the ablation simulation process, on the side where the launch device body contacts the air, the convective heat transfer between the surrounding gas and the device body is:

[0023]

[0024] where h is the convective heat transfer coefficient, with the unit of W / (m 2 *K), T f is the ambient gas temperature, with the unit of K.

[0025] In one embodiment of the present invention, the formation process of the ablation peeling temperature T r includes:

[0026] - Determine the ablation rate v of the coating tested under different working conditions, the initial estimated value T r0 of the peeling temperature, and the measurement standard deviation ε of the ablation amount of the tested coating;

[0027] - Determine the prior distribution where M is the number of tests, C is the covariance matrix, and Tr is the vector composed of the peeling temperature;

[0028] - Determine the likelihood function where v is the measured coating consumption rate, and v0 is the coating consumption rate calculated according to the ablation peeling temperature Tr;

[0029] - Determine the corresponding posterior distribution P(Tr / v) = P(Tr)P(v / Tr);

[0030] - When the ablation peeling temperature Tr corresponding to the maximum value of the objective function P(Tr / v) in the iterative process is the optimal solution.

[0031] In one embodiment of the present invention, the determination of the failure temperature T s includes:

[0032] Conduct a tensile load test on the launching device coated with the thermal protection coating at different temperatures, and determine the failure temperature T s of the organic bottom layer through the tensile breaking load and the breaking position; at the failure temperature T s , the adhesion of the organic bottom layer decreases, resulting in large-scale shedding of the coating.

[0033] In one embodiment of the present invention, the determination of the designed thickness of the thermal protection coating includes:

[0034] When:

[0035] The temperature T d of the organic bottom layer ≤ the failure temperature T s , and the temperature T z of the device ≤ the metal annealing temperature T h , the current coating thickness l of the thermal protection coating forms the designed thickness of the thermal protection coating according to the safety margin coefficient.

[0036] In one embodiment of the present invention, the safety margin coefficient is 1.2.

[0037] The heat - resistant coating thickness design device for a space launch device according to an embodiment of the present invention includes:

[0038] A memory for storing program codes during the processing of the heat - resistant coating thickness design method for a space launch device as described in any one of claims 1 to 8;

[0039] A processor for executing the program codes.

[0040] The heat - resistant coating thickness design device for a space launch device according to an embodiment of the present invention includes:

[0041] A heat - flow parameter acquisition module for determining the heat - environment parameters of the launch device;

[0042] A coating - thickness initial module for determining the ablation amount of the coating by using the heat - conduction model of the heat - resistant coating of the launch device according to the heat - environment parameters and the ablation - peeling temperature T r and determining the initially designed coating thickness l according to the ablation amount of the coating;

[0043] A coating - thickness estimation module for simulating and determining the overall temperature field of the launch device according to the coating thickness and the heat - conduction model of the heat - resistant coating, and determining the temperature T d of the organic bottom layer and the temperature T z of the launch device during the ablation process;

[0044] A coating - effect verification module for forming an iterative optimization of the coating thickness according to the comprehensive comparison results of the temperature T d of the organic bottom layer and the failure temperature T s , and the temperature T z of the device and the metal annealing temperature, and determining the designed thickness of the heat - resistant coating.

[0045] The heat - resistant coating thickness design method and device for a space launch device according to an embodiment of the present invention form a reliable heat - resistant coating thickness refined according to the heat environment by the heat - conduction simulation process of the heat environment of the corresponding - model rocket launch and the supporting launch device. It ensures that the launch device will not be directly ablated by the gas, the body of the launch device will not reach the annealing temperature, the organic bottom layer will not fail, and the coating will not fall off in a large area. This method can provide technical support for the future high - density space - launch thermal - protection scheme, and ensure the reliability of the coating while controlling the launch cost. BRIEF DESCRIPTION OF THE DRAWINGS

[0046] Figure 1 The figure shows a schematic flow chart of the heat - resistant coating thickness design method for a space launch device according to an embodiment of the present invention.

[0047] Figure 2 The figure shows a schematic flow chart of calculating the coating peeling temperature based on the Bayesian statistical method in the heat - resistant coating thickness design method for a space launch device according to an embodiment of the present invention.

[0048] Figure 3 The figure shows a comparison diagram of the measured ablation amount and the calculated ablation amount of the coating when calculating the coating peeling temperature in the method for designing the thickness of the heat insulation coating for a space launch device according to an embodiment of the present invention.

[0049] Figure 4 The figure shows a schematic structural diagram of a device for designing the thickness of the heat insulation coating for a space launch device according to an embodiment of the present invention. Detailed implementation manners

[0050] To make the objectives, technical solutions and advantages of the present invention clearer and more understandable, the present invention will be further described below in conjunction with the accompanying drawings and specific implementation manners. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all of the embodiments. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without making creative efforts shall fall within the protection scope of the present invention.

[0051] A method for designing the thickness of the heat insulation coating for a space launch device according to an embodiment of the present invention is as Figure 1 shown. In Figure 1 , this embodiment includes:

[0052] Step 100: Determine the thermal environment parameters of the launch device.

[0053] Those skilled in the art can understand that the launch device includes but is not limited to a diversion device, a launch pad, etc. During the rocket launch process, the heat insulation coating is heated by the ejected gas flow to reach a certain temperature and then softened, and then peeled off by the shear of the gas flow to form an ablation phenomenon. The ablation situation is related to various gas parameters, mainly including gas flow pressure, temperature, flow velocity, etc. Heat flux is the amount of heat transferred per unit area and per unit time on the surface, and this parameter fully considers the gas pressure, temperature, and flow velocity. Therefore, the main parameters of the thermal environment include ablation heat flux Q and ablation time t. By carrying out gas flow simulation for a specific model of engine in the launch mission, the surface thermal environment of the launch device can be quantified.

[0054] Step 200: According to the thermal environment parameters and the ablation peeling temperature T r Use the heat conduction model of the heat insulation coating of the launch device to determine the ablation amount of the coating, and determine the initial designed coating thickness l according to the ablation amount of the coating.

[0055] The heat conduction model of the heat insulation coating is a finite element heat conduction analysis model for the heat insulation coating and the launch device body. In the heat conduction model of the heat insulation coating, the heat conduction equation of the heat insulation coating and the launch device body is:

[0056]

[0057] Among them, Cp where \(c_p\) is the specific heat capacity with the unit of J / (kg·K), \(\rho\) is the density with the unit of kg / m 3 ³, \(T\) is the temperature with the unit of K, \(x\) is the distance with the unit of m, \(t\) is the time with the unit of s, and \(\lambda\) is the thermal conductivity.

[0058] The ablation of the coating by the combustion gas varies with time, and the coating thickness \(x\) varies with time. The situation is as follows:

[0059]

[0060] It is assumed that when the heat temperature \(T\geq\) the ablation peeling temperature \(T_r\), the coating is ablated and peeled off.

[0061] At the initial temperature field \(t = 0\ s\), \(T = T_0\), that is, at the initial moment, the temperatures of the device body and the protective coating are both at room temperature.

[0062] During the ablation simulation process, on the side where the combustion gas contacts the heat - resistant coating, the heat conduction from the high - temperature combustion gas to the coating is:

[0063]

[0064] where \(Q\) is the heat of the combustion gas and \(t_0\) is the duration.

[0065] During the ablation simulation process, on the side where the launcher body contacts the air, the convective heat transfer between the surrounding gas and the device body is:

[0066]

[0067] where \(h\) is the convective heat transfer coefficient with the unit of W / (m 2 ²·K), \(T\) f is the temperature of the surrounding gas with the unit of K.

[0068] The above partial differential equations are numerically solved by the finite - volume method, that is, the grid is divided according to the grid - independence result, and the temperature of each grid control unit is obtained through the finite - volume method. For the grid control unit with \(T\geq T\) r , it is discarded in the process of solving the next time step, simulating the process of the temperature of a certain layer of the grid in the coating being greater than the critical temperature and the peeling of this layer of the grid. Then, the entire temperature field and the grid peeling amount (corresponding to the ablation thickness of the coating) are calculated, so as to obtain the calculated ablation amount and the calculated ablation rate of the heat - resistant coating under different ablation conditions.

[0069] The calculated ablation data from the thermal - conduction model simulation of the heat - resistant coating and the experimentally measured ablation data are statistically analyzed by the Bayesian statistics method to form the verified ablation peeling temperature \(T\) r of the heat - resistant coating.

[0070] The process of calculating the coating peeling temperature based on the Bayesian statistics method is as Figure 2 shown. AtFigure 2 Among them, the simulated ablation data includes the calculated ablation amount and the calculated ablation rate. The experimental measured ablation data includes the actual ablation amount and the measured ablation rate in the experimental environment. The specific calculation process includes:

[0071] - Determine the ablation rate v of the tested coating and the initial estimated value T of the peeling temperature under different working conditions r0 , and the measurement standard deviation ε of the ablation amount of the tested coating;

[0072] - Determine the prior distribution where M is the number of tests, C is the covariance matrix, and Tr is the vector composed of the peeling temperature;

[0073] - Determine the likelihood function where v is the measured coating consumption rate, and v0 is the coating consumption (ablation) rate calculated according to the ablation peeling temperature Tr;

[0074] - Determine the corresponding posterior distribution P(Tr / v) = P(Tr)P(v / Tr);

[0075] - When the ablation peeling temperature Tr corresponding to the maximum value of the objective function P(Tr / v) is obtained during the iteration process, it is the optimal solution.

[0076] The comparison between the calculated ablation amount of the protective coating and the experimentally measured ablation amount is as Figure 3 shown. In Figure 3 , the calculated ablation amount and the measured ablation amount are in good agreement, proving that the ablation peeling temperature Tr calculated by simulation is in line with the actual situation.

[0077] Step 300: Determine the overall temperature field of the launch device by simulating according to the coating thickness and the heat protection coating heat conduction model, and determine the temperature T of the organic bottom layer during the ablation process d and the temperature T of the launch device z .

[0078] By using the heat protection coating heat conduction model of the launch device to determine the ablation simulation of the coating thickness, the overall and dynamic temperature field data of the launch device including the heat protection coating during the ablation process can be obtained, forming a quantification of the continuous heat conduction characteristics of the launch device, and determining the temperature T of the organic bottom layer during the ablation process d and the temperature T of the launch device z of the real-time simulation data.

[0079] Step 400: Based on the comprehensive comparison results of the temperature T of the organic bottom layer d and the failure temperature T s , and the temperature T of the device z and the metal annealing temperature, an iterative optimization of the coating thickness is formed to determine the design thickness of the heat protection coating.

[0080] The failure temperature Ts characterizes the critical temperature at which the adhesive force of the organic bottom layer decreases, resulting in the coating peeling off. The metal annealing temperature T h characterizes the critical temperature at which the strength and hardness of the emission device significantly decrease.

[0081] The failure temperature T of the organic bottom layer of the protective coating s is obtained through a tensile load test. Specifically, a tensile load test is carried out on the emission device coated with the thermal protection coating at different temperatures, and the failure temperature T of the organic bottom layer is determined by the breaking load and the breaking position s . At the failure temperature T s , the adhesive force of the organic bottom layer decreases, causing large-scale peeling of the coating.

[0082] By forming the condition threshold for iterative optimization of the thermal protection coating through the failure temperature and the metal annealing temperature, a targeted thickness design of the thermal protection coating on the emission device for the specific rocket thermal environment is formed.

[0083] In an embodiment of the present invention, when:

[0084] The temperature T of the organic bottom layer d ≤ the failure temperature T s , and the device temperature T z ≤ the metal annealing temperature T h , the coating thickness l of the current thermal protection coating forms the design thickness of the thermal protection coating according to the safety margin coefficient. For example, when the safety margin coefficient is 1.2, the design thickness is 1.2l.

[0085] The method for designing the thickness of the thermal protection coating for the space launch device in the embodiment of the present invention forms a reliable thermal protection coating thickness refined according to the thermal environment based on the thermal environment of the corresponding rocket launch and the thermal conduction simulation process of the supporting launch device. It ensures that the launch device will not be directly ablated by the gas, the launch device body will not reach the annealing temperature, the organic bottom layer will not fail, and the coating will not peel off on a large scale. This method can provide technical support for the future high-density space launch thermal protection scheme, and ensure the reliability of the coating while controlling the launch cost.

[0086] An apparatus for designing the thickness of the thermal protection coating for a space launch device in an embodiment of the present invention includes:

[0087] A memory for storing the program code in the process of the method for designing the thickness of the thermal protection coating for the space launch device in the above embodiment;

[0088] A processor for executing the program code in the process of the method for designing the thickness of the thermal protection coating for the space launch device in the above embodiment.

[0089] The processor can adopt a DSP (Digital Signal Processor) digital signal processor, an FPGA (Field-Programmable Gate Array) field programmable gate array, an MCU (Microcontroller Unit) system board, an SoC (system on a chip) system board, a minimum system of a PLC (Programmable Logic Controller) including I / O, or computing power remotely or in the cloud.

[0090] As shown in the thickness design device of the thermal protection coating for a space launch device according to an embodiment of the present invention Figure 4 shown. In Figure 4 this embodiment includes:

[0091] A heat flux parameter acquisition module 10 for determining the thermal environment parameters of the launch device;

[0092] A coating thickness initial module 20 for determining the ablation amount of the coating according to the thermal environment parameters and the ablation peeling temperature T r using the thermal conductivity model of the thermal protection coating of the launch device, and determining the initially designed coating thickness l according to the ablation amount of the coating;

[0093] A coating thickness estimation module 30 for simulating the overall temperature field of the launch device according to the coating thickness and the thermal conductivity model of the thermal protection coating, and determining the temperature T of the organic bottom layer d during the ablation process and the temperature T of the launch device z ;

[0094] A coating effect verification module 40 for forming an iterative optimization of the coating thickness according to the comprehensive comparison results of the temperature T of the organic bottom layer d and the failure temperature T s and the temperature T of the device z and the metal annealing temperature, and determining the designed thickness of the thermal protection coating.

[0095] As mentioned above, only the specific embodiments of the present invention are better, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily thought of by those skilled in the art within the technical scope disclosed by the present invention should be covered by the protection scope of the present invention. Therefore, the protection scope of the present invention should be subject to the protection scope of the claims.

Claims

1. A method for designing the thickness of a heat - resistant coating for a space launch device, characterized in that, Including: Determine the thermal environment parameters of the launch device; According to the thermal environment parameters and the ablation peeling temperature T r Determine the ablation amount of the coating using the heat conduction model of the thermal protection coating of the launch device, and determine the initial designed coating thickness l according to the ablation amount of the coating; Determine the overall temperature field of the launcher according to the coating thickness and the simulation of the heat insulation coating thermal conductivity model, and determine the temperature T of the organic bottom layer during the ablation process according to the overall temperature field of the launcher d and the temperature T of the launcher z ; According to the organic bottom layer temperature T d and the failure temperature T s , the device temperature T z and the comprehensive comparison result with the metal annealing temperature form an iterative optimization of the coating thickness to determine the designed thickness of the thermal protection coating.

2. The method for designing the thickness of the thermal protection coating for the space launch device according to claim 1, characterized in that, The thermal environment parameters include the ablation heat flux Q and the ablation time t.

3. The method for designing the thickness of the thermal protection coating for the space launch device according to claim 1, wherein In the thermal protection coating heat conduction model, the heat conduction equation between the thermal protection coating and the launch device body is: where C p is the specific heat capacity with the unit of J / (kg*K), ρ is the density with the unit of kg / m 3 , T is the temperature with the unit of K, x is the distance with the unit of m, t is the time with the unit of s, and λ is the thermal conductivity. The coating is ablated by the gas over time, and the coating thickness x changes over time, and the situation is: Assume that when the heat temperature T≥the ablation peeling temperature Tr, the coating is ablated and peeled. The initial temperature field is t = 0s, T = T0, that is, at the initial moment, the temperatures of the device body and the protective coating are both at room temperature.

4. The method for designing the thickness of the heat protection coating for the space launch device according to claim 3, characterized in that, In the ablation simulation process of the thermal protection coating heat conduction model, on the side where the gas contacts the thermal protection coating, the high-temperature gas conducts heat to the coating as: Where Q is the heat of the gas, and the duration is t0. During the ablation simulation process, on the side where the launch device body contacts the air, the convective heat transfer between the surrounding gas and the device body is: where h is the convective heat transfer coefficient, in W / (m 2 *K), and T f is the surrounding gas temperature, in K.

5. The method for designing the thickness of the heat insulation coating for a space launch device according to claim 1, characterized in that, The ablation stripping temperature T r The formation process thereof includes: - Determine the ablation rate v of the tested coating and the initial estimated value T of the peeling temperature under different working conditions r0 , and the measurement standard deviation ε of the ablation amount of the tested coating; - Determine the prior distribution where M is the number of experiments, C is the covariance matrix, and Tr is the vector composed of stripping temperatures; - Determine the likelihood function where v is the measured coating consumption rate, and v0 is the coating consumption rate calculated based on the ablation peeling temperature Tr; - Determine the corresponding posterior distribution P(Tr / v)=P(Tr)P(v / Tr); - When the ablation peeling temperature Tr corresponding to the maximum value of the objective function P(Tr / v) is obtained during the iteration process, it is the optimal solution.

6. The method for designing the thickness of the thermal protection coating for the space launch device according to claim 1, wherein, The determination of the failure temperature T s includes: Conduct tensile load tests on the emission device coated with a thermal protection coating at different temperatures, and determine the failure temperature T of the organic bottom layer based on the tensile breaking load and the breaking position. s ; At the failure temperature T s , the adhesion of the organic bottom layer decreases, resulting in large-scale shedding of the coating.

7. The method for designing the thickness of the heat protection coating for the space launch device according to claim 1, characterized in that, The determination of the design thickness of the thermal protection coating includes: When: Organic bottom layer temperature T d ≤ Failure temperature T s , and when the device temperature T z ≤ Metal annealing temperature T h , the coating thickness l of the current thermal protection coating forms the design thickness of the thermal protection coating according to the safety margin coefficient.

8. The method for designing the thickness of the heat protection coating for the space launch device according to claim 1, characterized in that, The safety margin coefficient is 1.

2.

9. A device for designing the thickness of a heat-resistant coating for a space launch device, characterized in that, Including: A memory for storing program codes during the processing of the method for designing the thickness of the thermal protection coating for a space launch device according to any one of claims 1 to 8; A processor for executing the program codes.

10. A device for designing the thickness of a heat - resistant coating for a space launch device, characterized in that, Including: A heat flux parameter acquisition module for determining the thermal environment parameters of the launch device; Coating thickness initial module, used to determine the initial designed coating thickness l according to the thermal environment parameters and the ablation peeling temperature T r Determine the coating ablation amount using the thermal protection coating heat conduction model of the launch device, and determine the initially designed coating thickness l according to the coating ablation amount; The coating thickness estimation module is used to determine the overall temperature field of the launcher according to the coating thickness and the simulation of the heat insulation coating thermal conductivity model, and determine the temperature T of the organic bottom layer during the ablation process according to the overall temperature field of the launcher d and the temperature T of the launcher z ; Coating effect verification module, which is used to form iterative optimization of the coating thickness based on the comprehensive comparison result of the organic bottom layer temperature T d and the failure temperature T s , the device temperature T z and the metal annealing temperature, and determine the designed thickness of the heat-resistant coating.

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