Process optimization method for thermal anti-invasion analysis and test of turbine pump of liquid rocket engine

Abstract

The invention relates to a thermal state analysis method, in particular to a liquid rocket engine turbine pump thermal anti-invasion analysis and test process optimization method, and aims to solve the problem that the thermal state of a turbine pump is difficult to obtain through a test method in the prior art. The method comprises the following steps: 1) selecting a computational 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 heat state S2; (4) judging the temperature of each key component, if the conditions are met, ending the process, and if the conditions are not met, executing the step (5); 5) controlling the temperature of the key component which does not meet the requirement, and applying a third boundary condition C3 to the heat transfer model M to obtain a thermal state S3; and 6) judging the temperature of each key component, if the conditions are met, ending the process, and if the conditions are not met, returning to the step 5) to modify 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 anti-intrusion analysis and testing of a liquid rocket engine turbo pump. Background Art

[0002] The working principle of a pump-type liquid rocket engine is to use a turbopump to transport the propellant in the tank to the thrust chamber through a pipeline. The turbopump includes a turbine and a propellant pump, which specifically includes components such as a centrifugal wheel, bearings, seals and gears. The oxidizer pump (also known as an oxygen pump) and the fuel pump in the propellant pump are driven by a turbine, and the turbine is coaxial with the oxidizer pump and the fuel pump or is driven by gears. The working fluid of the turbine is provided by a gas generator, and the temperature of the gas generated by the gas generator is about 500-700K. When the gas generated by the gas generator is oxygen-rich gas, the oxygen pump is generally connected to the turbine end. The advantage of such a setting is that even if there is a leak in the seal at the oxygen pump end, the leaked liquid oxygen will not burn after entering the oxygen-rich gas chamber, and will not cause the gas temperature to rise and cause the turbine to burn out; however, since the oxygen pump and the turbine need to be installed on the same shaft (spindle), the oxygen pump housing is connected to the turbine housing, the temperature of the liquid oxygen is about 90K (-173℃), and there is a temperature difference of about 500-700K 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 precooled to meet the engine starting conditions (the outlet temperature of the precooling pipe is less than a certain value). At this time, the oxygen pump structure will be at a relatively low temperature level, and the end face seals, rubber rings, bearings and other components of the oxygen pump 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 shutdown, the high-temperature waste heat of the turbine structure is redistributed in the turbopump, and the heat is transferred from the turbine housing and the stator and rotor blades of the turbine to the oxygen pump end, causing the temperature of the bearings and seals of the oxygen pump close to 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, which involves many parts and materials and 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 component 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 prior art. 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 anti-intrusion analysis and test process optimization of a liquid rocket engine turbine pump is provided.

[0005] To achieve the above purpose, the technical solution provided by the present invention is:

[0006] 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 which are sequentially sleeved on a main shaft, the oxygen pump comprises a bearing chamber and three internal flow passages, wherein the three internal flow passages are a liquid oxygen main passage, a bearing auxiliary passage and a leakage passage, and a plurality of key components are arranged in the bearing chamber, wherein the plurality of key components comprise a plurality of bearings and a plurality of sealing rings; the gas turbine pipeline comprises a turbine, and the turbine comprises a stator and a rotor blade; the method is special 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, applying the first boundary condition C1 to the heat transfer model M, and then performing the first simulation calculation on the heat transfer model M to obtain the thermal state S1 in the calculation domain when the turbo pump works normally for the first time;

[0009] Step 3: Taking 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 the engine turbo pump has no risk of thermal back-intrusion and subsequent tests can be performed; if the temperature of at least one key component is not within its corresponding allowable temperature range, execute step 5;

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

[0012] Step 6, performing a third simulation calculation on the heat transfer model M to obtain a 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 the engine turbopump has no risk of thermal back-intrusion and subsequent tests can be carried out; if there is at least one key component whose temperature 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 calculation domain, take the gas turbine pipeline as the heat source, and establish the initial heat transfer model M0 of the turbopump;

[0016] Step 1.2, select the key components in the heat transfer model M0, and establish a material property database for each key component, wherein the material properties include density, specific heat capacity and thermal conductivity;

[0017] Step 1.3, name each key component in the heat transfer model M0 and assign material properties;

[0018] Step 1.4: Check and set the contact relationship of each key component in the heat transfer model M0, and eliminate the contact relationship that does not conform to the actual situation, so as to complete the establishment of the heat transfer model M of the turbine 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 main liquid oxygen circuit, set a convective heat exchange boundary condition for the bearing auxiliary circuit and the 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 shell is in a covered state, set adiabatic boundary conditions for the turbine shell; if the turbine shell is in an exposed state, set the heat transfer coefficient for the turbine shell 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] Setting adiabatic boundary conditions for the outer shell of the oxygen pump;

[0031] For gas turbine piping:

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

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

[0034] Condition I: For oxygen pump:

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

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

[0037] Condition II: For gas turbine piping:

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

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

[0040] Among them, condition I is the boundary condition of the precooling 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 anti-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 prior art. It has a wide range of applications and high safety. BRIEF DESCRIPTION OF THE DRAWINGS

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

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

[0045] The turbine pump used in this embodiment includes a fuel pump, an isolation chamber, an oxygen pump and a gas turbine pipeline which are sequentially sleeved on the main shaft. The oxygen pump includes a bearing chamber and three internal flow passages, wherein the three internal flow passages are respectively a liquid oxygen main passage, a bearing auxiliary passage and a leakage passage. Several key components are arranged in the bearing chamber, and the several key components include several bearings and several sealing rings; the gas turbine pipeline includes a turbine, and the turbine includes a stator and a rotor blade. During normal operation, the flow rate of the oxygen pump is about 300kg / s, wherein the liquid oxygen main passage accounts for more than 90% of the total flow rate; the flow rate of the gas turbine pipeline is about 400kg / s, and the oxygen-rich gas flows in the turbine. From the perspective of the turbine structure size, the flow area at the stator and rotor blades is small, and the flow velocity is also high. The key components in this embodiment include components such as bearings and sealing rings.

[0046] A method for thermal intrusion analysis and test process optimization of a liquid rocket engine turbo pump, 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 calculation domain, take 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 the material property database of each key component, including density, specific heat capacity and thermal conductivity;

[0050] Step 1.3, name each key component in the heat transfer model M0 and assign material properties;

[0051] Step 1.4, check and set the contact relationship of each key component in the heat transfer model M0, and eliminate the contact relationship that does not conform to the actual situation, and complete the establishment of the heat transfer model M of the turbo pump; here, the heat conductivity of the special contact relationship needs to be set according to the actual contact thermal conductivity. The above special contact relationship refers to the many small gaps in the oxygen pump; there is thermal resistance in the heat transfer between two components with gaps. If it is considered as full contact, the heat transfer will be overestimated. Generally, the ratio of the width of the small gap (about 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, applying the first boundary condition C1 to the heat transfer model M, and then performing the first simulation calculation on the heat transfer model M to obtain the thermal state S1 in the calculation domain when the turbo pump works 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 main liquid oxygen circuit, set a convective heat exchange boundary condition for the bearing auxiliary circuit and the 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 shell is in a covered state, set an adiabatic boundary condition for the turbine shell; if the turbine shell is exposed, set the heat transfer coefficient for the turbine shell according to natural convection heat transfer; because the temperature of the gas will decrease accordingly after the gas 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 of different engines is different), and will continue to decrease by 20-100K after passing through the rotor, so the heat exchange temperature in the turbine gas path can be applied in sections according to the area, and the temperature reduction range can be theoretically estimated according to the aerodynamic formula, and can also be 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: Taking 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; wherein the second boundary condition C2 is specifically:

[0063] For oxygen pumps:

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

[0065] For gas turbine piping:

[0066] If the turbine shell is in a covered state, an adiabatic boundary condition is set for the turbine shell; if the turbine shell is in an exposed state, a heat transfer coefficient is set for the turbine shell 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 the engine turbo pump has no risk of thermal back-intrusion and subsequent tests can be performed; 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 the corresponding allowable temperature range, select one or both of the two methods of precooling and blowing, control the temperature of the key components that are not within the corresponding allowable temperature range, make the temperature of the key components that are not within the corresponding allowable temperature range within the allowable temperature range, and 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] Setting adiabatic boundary conditions for the outer shell of the oxygen pump;

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

[0072] Condition II: For gas turbine piping:

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

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

[0075] Among them, condition I is the boundary condition of the precooling method; condition II is the boundary condition of the blowing method;

[0076] Step 6, performing a third simulation calculation on the heat transfer model M to obtain a 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 the engine turbopump has no risk of thermal back-intrusion and subsequent tests can be carried out; if there is at least one key component whose temperature 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 explain step 5 in detail: There are two main methods to reduce 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 certain flow rate to continuously cool key parts, but this method has limitations because the flow resistance of the liquid oxygen main circuit is much smaller than that of the bearing auxiliary circuit. If the total flow rate of liquid oxygen is small, the flow rate of the bearing auxiliary circuit is even smaller, and the pre-cooling target may not be achieved. Another method is to blow nitrogen from the gas circuit to take away the heat from the turbine structure to reduce heat transfer to the oxygen pump structure. This method is limited by the amount of nitrogen carried. The method of reducing the temperature of key components can be selected according to 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 which are sequentially sleeved on a main shaft, the oxygen pump comprises a bearing chamber and three internal flow passages, wherein the three internal flow passages are a liquid oxygen main passage, a bearing auxiliary passage and a leakage passage, and a plurality of key components are arranged in the bearing chamber, wherein the plurality of key components comprise a plurality of bearings and a plurality of sealing rings; the gas turbine pipeline comprises a turbine, and the turbine comprises a stator and a rotor blade; wherein ... respectively a liquid oxygen main passage, a bearing auxiliary passage and a leakage passage, and a plurality of key components are arranged in the bearing chamber, and the plurality of key components comprise a plurality of bearings and a plurality of sealing rings; the gas turbine pipeline comprises a turbine, and the turbine comprises a stator and a rotor blade; wherein the oxygen pump comprises a bearing chamber and three internal flow passages, wherein the three internal flow passages are respectively a liquid oxygen main passage, a bearing auxiliary passage and a leakage passage, and the bearing chamber comprises a plurality of key components, wherein the plurality of key components comprise a plurality of bearings and a plurality of sealing rings; the gas turbine pipeline comprises a turbine, and the turbine comprises a stator and a rotor blade; 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, applying the first boundary condition C1 to the heat transfer model M, and then performing the first simulation calculation on the heat transfer model M to obtain the thermal state S1 in the calculation domain when the turbo pump works normally for the first time; Step 3: Taking 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; 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 the engine turbo pump has no risk of thermal back-intrusion and subsequent tests can be performed; if the temperature of at least one key component is not within its corresponding allowable temperature range, execute step 5; Step 5, select the key components that are not within the corresponding allowable temperature range, select one or both of the two methods of precooling and blowing, control the temperature of the key components that are not within the corresponding allowable temperature range, make the temperature of the key components that are not within the corresponding allowable temperature range within the allowable temperature range, and apply the third boundary condition C3 to the heat transfer model M; Step 6, performing a third simulation calculation on the heat transfer model M to obtain a 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 the engine turbopump has no risk of thermal back-intrusion and subsequent tests can be carried out; if there is at least one key component whose temperature 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. A method for thermal back-intrusion analysis and test process optimization of a liquid rocket engine turbo pump 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 calculation domain, take the gas turbine pipeline as the heat source, and establish the initial heat transfer model M0 of the turbopump; Step 1.2, select the key components in the heat transfer model M0, and establish a material property database for each key component, wherein the material properties include density, specific heat capacity and thermal conductivity; Step 1.3, name each key component in the heat transfer model M0 and assign material properties; Step 1.4: Check and set the contact relationship of each key component in the heat transfer model M0, and eliminate the contact relationship that does not conform to the actual situation, so as to complete the establishment of the heat transfer model M of the turbine pump.

3. A method for thermal back-intrusion analysis and test process optimization of a liquid rocket engine turbo pump according to claim 2, characterized in that: The first boundary condition C1 in step 2 is specifically: For oxygen pumps: Set a constant temperature boundary condition for the main liquid oxygen circuit, set a convective heat exchange boundary condition for the bearing auxiliary circuit and the 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 shell is in a covered state, set adiabatic boundary conditions for the turbine shell; if the turbine shell is in an exposed state, set the heat transfer coefficient for the turbine shell 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 fluid flow pipe; when the fluid is heated, n=0.4, and when the fluid is cooled, n=0.

3.

4. A method for thermal back-intrusion analysis and test process optimization of a liquid rocket engine turbo pump according to claim 3, characterized in that: The second boundary condition C2 is specifically: For oxygen pumps: Setting adiabatic boundary conditions for the outer shell of the oxygen pump; For gas turbine piping: If the casing of the turbine is in a covered state, an adiabatic boundary condition is set for the casing of the turbine. If the casing of the turbine is in an exposed state, a boundary condition is set for the casing of the turbine according to natural convection heat transfer.

5. A method for thermal back-intrusion analysis and test process optimization of a liquid rocket engine turbo pump according to claim 4, characterized in that: The third boundary condition C3 is specifically: Condition I: For oxygen pump: Setting adiabatic boundary conditions for the outer shell of the oxygen pump; Convective heat transfer boundary conditions are set for the liquid oxygen main path, bearing auxiliary path and leakage channel; Condition II: For gas turbine piping: If the turbine shell is in a covered state, an adiabatic boundary condition is set for the turbine shell; if the turbine shell is in an exposed state, a boundary condition is set for the turbine shell according to natural convection heat transfer; For the inner surface of the gas turbine pipeline: set the convection heat transfer boundary conditions; Condition I is the boundary condition of the precooling method; Condition II is the boundary condition of the blowing method.

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