A process optimization method for thermal intrusion analysis and testing of liquid rocket engine turbopumps

Abstract

The present invention relates to a thermal state analysis method, and in particular to a method for thermal intrusion analysis and test process optimization of a liquid rocket engine turbopump, the purpose of which is to solve the problem that it is difficult to obtain the thermal state of a turbopump through experimental methods in the prior art. The present invention comprises the following steps: 1) selecting a calculation domain and establishing a heat transfer model M; 2) applying a first boundary condition C1 to the heat transfer model M to obtain a thermal state S1; 3) applying a second boundary condition C2 to the heat transfer model M to obtain a thermal state S2; 4) determining the temperature of each key component, and ending if all conditions are met, and executing 5 if not; 5) controlling the temperature of the key components that do not meet the requirements, and applying a third boundary condition C3 to the heat transfer model M to obtain a thermal state S3; 6) determining the temperature of each key component, and ending if all conditions are met, and returning to 5 if not; and modifying the third boundary condition C3 until the temperature of each key component meets the requirements.

Description

Technical Field

[0001] The invention relates to a thermal state analysis method, in particular to a process optimization method for thermal back-intrusion analysis and testing of a liquid rocket engine turbopump. Background Art

[0002] The operating principle of a pump-type liquid rocket engine is to use a turbopump to transport propellant from the tank to the thrust chamber via a pipeline. The turbopump consists of a turbine and a propellant pump, specifically including components such as a centrifugal wheel, bearings, seals, and gears. The oxidizer pump (also known as the oxygen pump) and the fuel pump in the propellant pump are driven by a turbine, which is coaxial with the oxidizer and fuel pumps or driven by gears. The working fluid of the turbine is provided by a gas generator, which produces gas at a temperature of approximately 500 to 700 K. When the gas generator produces oxygen-enriched gas, an oxygen pump is generally connected to the turbine. This arrangement has the advantage that even if the oxygen pump seal leaks, the leaked liquid oxygen entering the oxygen-enriched gas chamber will not burn, thereby preventing the gas temperature from rising and causing turbine burnout. However, since the oxygen pump and turbine must be mounted on the same shaft (main shaft), and the oxygen pump housing is connected to the turbine housing, the liquid oxygen temperature is approximately 90 K (-173°C), resulting in a temperature difference of approximately 500 to 700 K between the oxygen pump and the turbine. Liquid oxygen is usually used as a cryogenic propellant. Before the engine is started, it needs to be pre-cooled to meet the engine starting conditions (the outlet temperature of the pre-cooling pipe is less than a certain value). At this time, the oxygen pump structure will be at a relatively low temperature level, and the oxygen pump's end face seals, rubber rings, bearings and other components are all made of low-temperature resistant materials. When the engine has a need to be reusable, before the second start of the engine after the first use and shutdown, the high-temperature waste heat of the turbine structure is redistributed in the turbopump. The heat is transferred from the turbine housing and the turbine's stator and rotor blades to the oxygen pump end, causing the temperature of the oxygen pump bearings and seals near the turbine end to rise. This process is called turbine thermal intrusion. This thermal intrusion will cause the temperature of key components such as bearings and seals to exceed the operating temperature range, making the second start of the engine risky.

[0003] Due to the complex structure of the turbopump, involving many parts and materials, complex boundary conditions, and considering the existence of rotating parts and high internal pressure inside the turbopump, it is difficult to obtain the structural thermal state of the bearing cavity of the key components of the turbopump through experimental methods. Even during the engine hot test, only temperature measurement points can be set on the outer casing of the turbopump, and the temperature obtained cannot reflect the temperature of the bearing or sealing ring. Summary of the Invention

[0004] The purpose of the present invention is to solve the problem that it is difficult to obtain the structural thermal state of the turbine pump bearing cavity through experimental methods in the existing technology. Even during the engine hot test, only temperature measuring points can be set on the turbine pump outer shell, and the obtained temperature cannot reflect the temperature problem at the bearing or sealing ring. Instead, a method for thermal back-intrusion analysis and test process optimization of a liquid rocket engine turbine pump is provided.

[0005] To achieve the above objectives, the technical solutions provided by the present invention are:

[0006] A process optimization method for thermal intrusion analysis and testing of a liquid rocket engine turbine pump is disclosed. The turbine pump comprises a fuel pump, an isolation chamber, an oxygen pump, and a gas turbine pipeline, which are sequentially sleeved on a main shaft. The oxygen pump comprises a bearing chamber and three internal flow paths, wherein the three internal flow paths are a main liquid oxygen path, an auxiliary bearing path, and a leakage channel. The bearing chamber is provided with several key components, including several bearings and several sealing rings. The gas turbine pipeline includes a turbine, which includes a stator and rotor blades. The method is unique in that it comprises the following steps:

[0007] Step 1: Select the calculation domain of the turbo pump and establish the heat transfer model M of the turbo pump;

[0008] Step 2: Apply a first boundary condition C1 to the heat transfer model M, and then perform a first simulation calculation on the heat transfer model M to obtain a thermal state S1 in the calculation domain when the turbo pump operates normally for the first time;

[0009] Step 3: Using the thermal state S1 as the initial condition, applying the second boundary condition C2 to the heat transfer model M, and then performing a second simulation calculation on the heat transfer model M to obtain the thermal state S2 of the turbo pump from the first shutdown to the second startup;

[0010] Step 4: Analyze the thermal state S2 to determine whether the temperature of each key component in the calculation domain is within its corresponding allowable temperature range. If the temperature of each key component is within its corresponding allowable temperature range, it is determined that there is no risk of thermal intrusion in the engine turbopump and subsequent tests can be carried out. If the temperature of at least one key component is not within its corresponding allowable temperature range, execute step 5.

[0011] Step 5: Select key components that are not within their corresponding allowable temperature ranges, select one or both of the two methods of pre-cooling and blowing, control the temperature of the key components that are not within their corresponding allowable temperature ranges, and make the temperature of the key components that are not within their corresponding allowable temperature ranges within the allowable temperature range, and apply the third boundary condition C3 to the heat transfer model M;

[0012] Step 6: Perform a third simulation calculation on the heat transfer model M to obtain the thermal state S3 of the new turbo pump from the first shutdown to the second startup;

[0013] Step 7: Analyze the thermal state S3 to determine whether the temperature of each key component in the calculation domain is within its corresponding allowable temperature range. If the temperature of each key component is within its corresponding allowable temperature range, it is determined that there is no risk of thermal intrusion in the engine turbopump and subsequent tests can be carried out. If the temperature of at least one key component is not within its corresponding allowable temperature range, return to step 5, modify the third boundary condition C3, and then re-execute the subsequent steps until the temperature of each key component is within its corresponding allowable temperature range.

[0014] Furthermore, step 1 is specifically as follows:

[0015] Step 1.1: Select the oxygen pump and gas turbine pipeline in the turbopump as the computational domain, use the gas turbine pipeline as the heat source, and establish the initial heat transfer model M0 of the turbopump;

[0016] Step 1.2: Select key components in the heat transfer model M0 and establish a material property database for each key component, including density, specific heat capacity, and thermal conductivity.

[0017] Step 1.3: Name the key components in the heat transfer model M0 and assign material properties;

[0018] Step 1.4: Check and set the contact relationships of the key components in the heat transfer model M0, and eliminate the contact relationships that do not conform to the actual situation to complete the establishment of the heat transfer model M of the turbo pump.

[0019] Furthermore, the first boundary condition C1 in step 2 is specifically:

[0020] For oxygen pumps:

[0021] Set a constant temperature boundary condition for the liquid oxygen main line, set a convective heat transfer boundary condition for the bearing auxiliary line and leakage channel, and set an adiabatic boundary condition for the outer shell of the oxygen pump; the heat exchange temperature of the liquid oxygen in the oxygen pump is set according to the test temperature during the actual process test;

[0022] For gas turbine piping:

[0023] Set constant temperature boundary conditions for the stator and rotor blades of the turbine, and set convective heat transfer boundary conditions for other areas of the turbine except the stator and rotor blades. If the turbine casing is in a covered state, set adiabatic boundary conditions for the turbine casing. If the turbine casing is in an exposed state, set the heat transfer coefficient for the turbine casing according to natural convection heat transfer. The heat transfer temperature of the gas turbine pipeline is applied in sections according to the area.

[0024] The boundary conditions of the convective heat transfer include the convective heat transfer coefficient and the heat transfer temperature. The convective heat transfer coefficient is estimated according to the flow heat transfer formula. When calculating the convective heat transfer coefficient h in the bearing auxiliary path and the leakage channel, according to the Dittus-Boelter formula and calculate;

[0025] The Dittus-Boelter formula is:

[0026] Nu=0.023Re 0.8 Pr n

[0027] Where Nu is the Nusselt number, Re is the Reynolds number, and Pr is the Prandtl number. The applicable condition is that the fluid Re>10 4 And 0.7<Pr<160, the qualitative temperature is Where t1 is the inlet temperature of the fluid, t2 is the outlet temperature of the fluid, d is the diameter of the pipe through which the fluid flows, d is taken as the characteristic length, and λ is the thermal conductivity of the fluid flow pipe; when the fluid is heated, n = 0.4, and when the fluid is cooled, n = 0.3.

[0028] Furthermore, the second boundary condition C2 is specifically:

[0029] For oxygen pumps:

[0030] Set adiabatic boundary conditions for the outer shell of the oxygen pump;

[0031] For gas turbine piping:

[0032] If the turbine casing is in a covered state, an adiabatic boundary condition is set for the turbine casing. If the turbine casing is in an exposed state, a boundary condition is set for the turbine casing according to natural convection heat transfer.

[0033] Furthermore, the third boundary condition C3 is specifically:

[0034] Condition I: For oxygen pump:

[0035] Set adiabatic boundary conditions for the outer shell of the oxygen pump;

[0036] Set convective heat transfer boundary conditions for the liquid oxygen main line, bearing auxiliary line and leakage channel;

[0037] Condition II: For gas turbine piping:

[0038] If the turbine casing is in a covered state, an adiabatic boundary condition is set for the turbine casing. If the turbine casing is in an exposed state, a boundary condition is set for the turbine casing according to natural convection heat transfer.

[0039] For the inner surface of the gas turbine tube: set the convection heat transfer boundary condition;

[0040] Condition I is the boundary condition of the pre-cooling method; Condition II is the boundary condition of the blowing method.

[0041] Compared with the prior art, the present invention has the following beneficial effects:

[0042] The liquid rocket engine turbopump thermal back-intrusion analysis and test process optimization method provided by the present invention is mainly used to analyze the thermal state of the bearing cavity of the rocket engine turbopump before the turbopump is reused, and to determine whether the temperature of the bearing and the sealing ring is lower than the allowable temperature of the material before the turbopump is repeatedly started. If there is a starting risk, this method can be used to predict the thermal state of the improvement plan, optimize the test system process, and ensure the normal start-up of the engine turbopump. It adopts a simulation analysis method to solve the problem that it is difficult to obtain the thermal state of the turbopump through experimental methods in the existing technology. It has a wide range of applications and high safety. BRIEF DESCRIPTION OF THE DRAWINGS

[0043] Figure 1 This is a flow chart of an embodiment of a method for optimizing the thermal intrusion analysis and test process of a liquid rocket engine turbopump according to the present invention; DETAILED DESCRIPTION

[0044] The present invention will be further described below with reference to the accompanying drawings and specific embodiments.

[0045] The turbopump employed in this embodiment comprises a fuel pump, an isolation chamber, an oxygen pump, and a gas turbine pipeline, all sequentially mounted on a main shaft. The oxygen pump includes a bearing chamber and three internal flow paths: a main liquid oxygen path, an auxiliary bearing path, and a leakage channel. The bearing chamber houses several key components, including bearings and seals. The gas turbine pipeline includes a turbine, which includes a stator and rotor blades. During normal operation, the oxygen pump's flow rate is approximately 300 kg / s, with the main liquid oxygen path accounting for over 90% of the total flow rate. The gas turbine pipeline's flow rate is approximately 400 kg / s. Oxygen-enriched fuel gas flows within the turbine. Due to the turbine's structural dimensions, the flow area around the stator and rotor blades is relatively small, resulting in higher flow velocities. Key components in this embodiment include bearings and seals.

[0046] A liquid rocket engine turbo pump thermal intrusion analysis and test process optimization method, the flow chart of which is shown in Figure 1 , including the following steps:

[0047] Step 1: Select the calculation domain of the turbo pump and use TMG, ANSYS THERMAL or other heat transfer software to establish the heat transfer model M of the turbo pump;

[0048] Step 1.1: Select the oxygen pump and gas turbine pipeline in the turbopump as the computational domain, use the gas turbine pipeline as the heat source, and establish the initial heat transfer model M0 of the turbopump;

[0049] Step 1.2: Select the key components in the heat transfer model M0 and establish a material property database for each key component, including density, specific heat capacity, and thermal conductivity.

[0050] Step 1.3: Name the key components in the heat transfer model M0 and assign material properties;

[0051] Step 1.4: Check and set the contact relationships of each key component in the heat transfer model M0, and eliminate any contact relationships that do not conform to reality to complete the establishment of the turbopump heat transfer model M. Here, the heat conductivity of the special contact relationships needs to be set according to the actual contact thermal conductivity. The above special contact relationships refer to the many small gaps in the oxygen pump. There is thermal resistance between two components with gaps. If the heat transfer is considered as full contact, the heat transfer will be overestimated. Generally, the ratio of the width of the small gap (approximately 0.1-3mm, depending on the situation) to the thermal conductivity of the medium in the gap is considered as the contact thermal resistance.

[0052] Step 2: Apply a first boundary condition C1 to the heat transfer model M, and then perform a first simulation calculation on the heat transfer model M to obtain a thermal state S1 in the calculation domain when the turbo pump operates normally for the first time;

[0053] The first boundary condition C1 is specifically:

[0054] For oxygen pumps:

[0055] Set a constant temperature boundary condition for the liquid oxygen main line, set a convective heat transfer boundary condition for the bearing auxiliary line and leakage channel, and set an adiabatic boundary condition for the outer shell of the oxygen pump; the heat exchange temperature of the liquid oxygen in the oxygen pump is set according to the test temperature during the actual process test;

[0056] For gas turbine piping:

[0057] Set constant temperature boundary conditions for the stator and rotor blades of the turbine, and set convective heat transfer boundary conditions for other areas of the turbine except the stator and rotor blades; if the turbine casing is in a covered state, set an adiabatic boundary condition for the turbine casing; if the turbine casing is in an exposed state, set the heat transfer coefficient for the turbine casing according to natural convection heat transfer; since the temperature of the gas will decrease accordingly after it expands and does work in the turbine, that is, the gas will decrease by 20-100K after passing through the stator (the temperature drop range is different for different engines), and will continue to decrease by 20-100K after passing through the rotor. Therefore, the heat exchange temperature in the turbine gas path can be applied in sections according to the region. The temperature reduction range can be theoretically estimated based on the aerodynamic formula, or obtained by conducting three-dimensional CFD calculations of the gas path;

[0058] The boundary conditions of convective heat transfer include the convective heat transfer coefficient and the heat transfer temperature. The convective heat transfer coefficient is estimated according to the flow heat transfer formula. When calculating the convective heat transfer coefficient h in the bearing auxiliary path and the leakage channel, according to the Dittus-Boelter formula and calculate;

[0059] The Dittus-Boelter formula is:

[0060] Nu=0.023Re 0.8 Pr n

[0061] Where Nu is the Nusselt number, Re is the Reynolds number, and Pr is the Prandtl number. The applicable condition is that the fluid Re>10 4 And 0.7<Pr<160, the qualitative temperature is Where t1 is the inlet temperature of the fluid, t2 is the outlet temperature of the fluid, d is the diameter of the pipe through which the fluid flows, d is taken as the characteristic length, and λ is the thermal conductivity of the pipe through which the fluid flows; when the fluid is heated, n = 0.4, and when the fluid is cooled, n = 0.3;

[0062] Step 3: Using the thermal state S1 as the initial condition, apply the second boundary condition C2 to the heat transfer model M, and then perform a second simulation calculation on the heat transfer model M to obtain the thermal state S2 of the turbo pump from the first shutdown to the second startup; wherein the second boundary condition C2 is specifically:

[0063] For oxygen pumps:

[0064] Set adiabatic boundary conditions for the outer shell of the oxygen pump;

[0065] For gas turbine piping:

[0066] If the turbine casing is in a covered state, an adiabatic boundary condition is set for the turbine casing. If the turbine casing is in an exposed state, a heat transfer coefficient is set for the turbine casing according to natural convection heat transfer.

[0067] Step 4: Analyze the thermal state S2 to determine whether the temperature of each key component in the calculation domain is within its corresponding allowable temperature range. If the temperature of each key component is within its corresponding allowable temperature range, it is determined that there is no risk of thermal intrusion in the engine turbopump and subsequent tests can be carried out. If the temperature of at least one key component is not within its corresponding allowable temperature range, execute step 5.

[0068] Step 5: Select key components that are not within their corresponding allowable temperature ranges, select one or both of the pre-cooling and blowing methods, and control the temperature of the key components that are not within their corresponding allowable temperature ranges so that the temperature of the key components that are not within their corresponding allowable temperature ranges is within the allowable temperature range. Apply the third boundary condition C3 to the heat transfer model M. The third boundary condition C3 is specifically:

[0069] Condition I: For oxygen pump:

[0070] Set adiabatic boundary conditions for the outer shell of the oxygen pump;

[0071] Set convective heat transfer boundary conditions for the liquid oxygen main line, bearing auxiliary line and leakage channel;

[0072] Condition II: For gas turbine piping:

[0073] If the turbine casing is in a covered state, an adiabatic boundary condition is set for the turbine casing. If the turbine casing is in an exposed state, a boundary condition is set for the turbine casing according to natural convection heat transfer.

[0074] For the inner surface of the gas turbine tube: set the convection heat transfer boundary condition;

[0075] Condition I is the boundary condition of the pre-cooling method; Condition II is the boundary condition of the blowing method;

[0076] Step 6: Perform a third simulation calculation on the heat transfer model M to obtain the thermal state S3 of the new turbo pump from the first shutdown to the second startup;

[0077] Step 7: Analyze the thermal state S3 to determine whether the temperature of each key component in the calculation domain is within its corresponding allowable temperature range. If the temperature of each key component is within its corresponding allowable temperature range, it is determined that there is no risk of thermal intrusion in the engine turbopump and subsequent tests can be carried out. If the temperature of at least one key component is not within its corresponding allowable temperature range, return to step 5, modify the third boundary condition C3, and then re-execute the subsequent steps until the temperature of each key component is within its corresponding allowable temperature range.

[0078] To elaborate on step 5: There are two main methods for reducing the temperature of key components. One method is to continuously pre-cool the oxygen pump, especially the bearing auxiliary circuit. Liquid oxygen flows continuously at a constant flow rate to continuously cool key parts. However, this method has limitations. The flow resistance of the main liquid oxygen circuit is much smaller than that of the bearing auxiliary circuit. If the total liquid oxygen flow rate is low, the flow rate through the bearing auxiliary circuit is even lower, and pre-cooling may not be achieved. Another method is to purge nitrogen from the gas circuit to remove heat from the turbine structure and reduce heat transfer to the oxygen pump structure. This method is limited by the amount of nitrogen carried. The method for reducing the temperature of key components can be selected based on actual needs.

Claims

1. A process optimization method for thermal intrusion analysis and testing of a liquid rocket engine turbine pump, wherein the turbine pump comprises a fuel pump, an isolation chamber, an oxygen pump, and a gas turbine pipeline sequentially sleeved on a main shaft, the oxygen pump comprising a bearing chamber and three internal flow passages, wherein the three internal flow passages are a main liquid oxygen passage, an auxiliary bearing passage, and a leakage passage, the bearing chamber being provided with a plurality of key components, the plurality of key components comprising a plurality of bearings and a plurality of sealing rings; the gas turbine pipeline comprising a turbine, the turbine comprising a stator and rotor blades; and characterized in that: The following steps are involved: Step 1: Select the calculation domain of the turbo pump and establish the heat transfer model M of the turbo pump; Step 2: Apply a first boundary condition C1 to the heat transfer model M, and then perform a first simulation calculation on the heat transfer model M to obtain the thermal state S1 in the calculation domain when the turbopump operates normally for the first time. The first boundary condition C1 is specifically: For oxygen pumps: Set a constant temperature boundary condition for the liquid oxygen main line, set a convective heat transfer boundary condition for the bearing auxiliary line and leakage channel, and set an adiabatic boundary condition for the outer shell of the oxygen pump; the heat exchange temperature of the liquid oxygen in the oxygen pump is set according to the test temperature during the actual process test; For gas turbine piping: Set constant temperature boundary conditions for the stator and rotor blades of the turbine, and set convective heat transfer boundary conditions for other areas of the turbine except the stator and rotor blades. If the turbine casing is in a covered state, set adiabatic boundary conditions for the turbine casing. If the turbine casing is in an exposed state, set the heat transfer coefficient for the turbine casing according to natural convection heat transfer. The heat transfer temperature of the gas turbine pipeline is applied in sections according to the area. The boundary conditions of the convective heat transfer include the convective heat transfer coefficient and the heat transfer temperature. The convective heat transfer coefficient is estimated according to the flow heat transfer formula. When calculating the convective heat transfer coefficient h in the bearing auxiliary path and the leakage channel, according to the Dittus-Boelter formula and calculate; The Dittus-Boelter formula is: Now=0.023Re 0.8 Per n Where Nu is the Nusselt number, Re is the Reynolds number, and Pr is the Prandtl number. The applicable condition is that the fluid Re>10 4 And 0.7<Pr<160, the qualitative temperature is Where t1 is the inlet temperature of the fluid, t2 is the outlet temperature of the fluid, d is the diameter of the pipe through which the fluid flows, d is taken as the characteristic length, and λ is the thermal conductivity of the pipe through which the fluid flows; when the fluid is heated, n = 0.4, and when the fluid is cooled, n = 0.3; Step 3: Using the thermal state S1 as the initial condition, apply the second boundary condition C2 to the heat transfer model M, and then perform a second simulation calculation on the heat transfer model M to obtain the thermal state S2 of the turbo pump from the first shutdown to the second startup. The second boundary condition C2 is specifically: For oxygen pumps: Set adiabatic boundary conditions for the outer shell of the oxygen pump; For gas turbine piping: If the turbine casing is in a covered state, an adiabatic boundary condition is set for the turbine casing. If the turbine casing is in an exposed state, a boundary condition is set for the turbine casing according to natural convection heat transfer. Step 4: Analyze the thermal state S2 to determine whether the temperature of each key component in the calculation domain is within its corresponding allowable temperature range. If the temperature of each key component is within its corresponding allowable temperature range, it is determined that there is no risk of thermal intrusion in the engine turbopump and subsequent tests can be carried out. If the temperature of at least one key component is not within its corresponding allowable temperature range, execute step 5. Step 5: Select key components that are not within their corresponding allowable temperature ranges, select one or both of the pre-cooling and blowing methods, and control the temperature of the key components that are not within their corresponding allowable temperature ranges so that the temperature of the key components that are not within their corresponding allowable temperature ranges is within the allowable temperature range. Apply a third boundary condition C3 to the heat transfer model M; the third boundary condition C3 is specifically: Condition I: For oxygen pump: Set adiabatic boundary conditions for the outer shell of the oxygen pump; Set convective heat transfer boundary conditions for the liquid oxygen main line, bearing auxiliary line and leakage channel; Condition II: For gas turbine piping: If the turbine casing is in a covered state, an adiabatic boundary condition is set for the turbine casing. If the turbine casing is in an exposed state, a boundary condition is set for the turbine casing according to natural convection heat transfer. For the inner surface of the gas turbine tube: set the convection heat transfer boundary condition; Condition I is the boundary condition of the pre-cooling method; Condition II is the boundary condition of the blowing method; Step 6: Perform a third simulation calculation on the heat transfer model M to obtain the thermal state S3 of the new turbo pump from the first shutdown to the second startup; Step 7: Analyze the thermal state S3 to determine whether the temperature of each key component in the calculation domain is within its corresponding allowable temperature range. If the temperature of each key component is within its corresponding allowable temperature range, it is determined that there is no risk of thermal intrusion in the engine turbopump and subsequent tests can be carried out. If the temperature of at least one key component is not within its corresponding allowable temperature range, return to step 5, modify the third boundary condition C3, and then re-execute the subsequent steps until the temperature of each key component is within its corresponding allowable temperature range.

2. The method for thermal intrusion analysis and test process optimization of a liquid rocket engine turbopump according to claim 1, characterized in that: Step 1 is as follows: Step 1.1: Select the oxygen pump and gas turbine pipeline in the turbopump as the computational domain, use the gas turbine pipeline as the heat source, and establish the initial heat transfer model M0 of the turbopump; Step 1.2: Select key components in the heat transfer model M0 and establish a material property database for each key component, including density, specific heat capacity, and thermal conductivity. Step 1.3: Name the key components in the heat transfer model M0 and assign material properties; Step 1.4: Check and set the contact relationships of the key components in the heat transfer model M0, and eliminate the contact relationships that do not conform to the actual situation to complete the establishment of the heat transfer model M of the turbo pump.

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