Rocket engine nozzle temperature measurement method and system

Abstract

The application provides a rocket engine nozzle temperature measuring method and system, comprising the following steps: step S1: setting a standard temperature Te under a non-shielding condition; step S2: preparing a test piece, forming a base, fixing a platinum-rhodium thermocouple in contact with the base, and configuring a plurality of ceramic protection tubes on a thermocouple measuring end; collecting platinum-rhodium thermocouple temperature data and simultaneously measuring temperature by using a two-color infrared thermometer; step S3: taking the temperature measured by the two-color infrared thermometer in step S2 as a reference, and establishing a platinum-rhodium thermocouple calibration curve; step S4: removing the outer surface of a throat part of a rocket engine nozzle by machining to form a base; fixing a platinum-rhodium thermocouple in contact with the base, and connecting the thermocouple to a temperature data collection system; and step S5: obtaining a throat temperature Tt by interpolation calculation according to the platinum-rhodium thermocouple calibration curve. The application solves the problem that the throat wall temperature of a key parameter of a nozzle cannot be measured by an infrared thermometer.

Description

Technical Field

[0001] This invention relates to the field of spacecraft liquid rocket engine testing technology, specifically to a method and system for measuring the temperature of rocket engine nozzles, and more particularly to a method for accurate temperature measurement of high-temperature rocket engine nozzles in confined space environments, especially suitable for high-altitude simulated hot-fire testing and verification of rocket engines installed in recesses or sleeves. Background Technology

[0002] The outer wall temperature of the nozzle throat of a spacecraft's attitude and orbit control liquid rocket engine reaches as high as 1400–1550°C. The nozzle substrate is a niobium alloy, and both the inner and outer surfaces are coated with a high-temperature anti-oxidation silicon compound coating. Current technology primarily relies on monochromatic or dual-color infrared thermometers for non-contact measurement. When the spacecraft's overall assembly layout is compact, the engine nozzle must be embedded in a recess, shielding the throat and above. Furthermore, reusable spacecraft, to withstand the aerodynamic loads and thermal environment during their round-trip journey to Earth, must employ a sleeve-type structure to conceal the rocket engine's high-temperature nozzle within the sleeve, while simultaneously isolating the nozzle from the thermal impact on surrounding heat-sensitive components. To ensure flight reliability, it is essential to conduct test firings and verifications of the engine on the ground, simulating the vacuum and installation environments of the spacecraft in orbit. However, the infrared light path is blocked by the recessed or sleeve mechanical structure, making it impossible to obtain the accurate temperature of the throat outer wall. In addition, the temperature of the engine nozzle outer wall is high, and the conductivity of the silicon coating on the outer surface is poor, making it impossible to directly spot weld thermocouples for temperature measurement. Conventional K-type thermocouples (maximum temperature measurement 1200℃) and T-type thermocouples (maximum temperature measurement 400℃) have insufficient temperature resistance, and existing contact measurement methods cannot eliminate the measurement error caused by the thermal conduction of the coating material.

[0003] Patent document CN119572375A discloses an automatic temperature measurement device and its usage method after a solid rocket engine test. The device includes a robotic arm and a temperature measuring device connected to the robotic arm via a flange. The temperature measuring device comprises a temperature measuring rod, a temperature measuring rod insulation layer, and a high-temperature thermocouple. The temperature measuring rod has a wiring groove and multiple through holes. The wiring groove is used to arrange the high-temperature thermocouple wiring harness, and the through holes are used to fix the high-temperature thermocouple. The temperature measuring rod insulation layer covers the outside of the temperature measuring rod. However, this patent document still has the drawback of not being able to eliminate measurement errors caused by the thermal conduction of the coating material. Summary of the Invention

[0004] To address the shortcomings of existing technologies, the purpose of this invention is to provide a method and system for measuring the temperature of rocket engine nozzles.

[0005] A method for measuring the temperature of a rocket engine nozzle according to the present invention includes the following steps:

[0006] Step S1: Under unobstructed conditions, use a dual-color infrared thermometer to measure the nozzle throat temperature of the rocket engine nozzle under standard conditions, and set it as the standard temperature Te.

[0007] Step S2: Prepare a test piece and prepare a high-temperature anti-oxidation coating on both sides. The material of the test piece and the coating preparation process and requirements are the same as those of the rocket engine nozzle.

[0008] The high-temperature anti-oxidation coating in the middle of the test piece is removed by machining to form an exposed area of ​​a preset width, while retaining a coating transition layer of a preset thickness to form the nozzle substrate. The surface roughness of the nozzle substrate is controlled within a preset threshold.

[0009] The platinum-rhodium thermocouple is fixedly in contact with the nozzle substrate, and multiple ceramic protective tubes are configured at the thermocouple measuring end.

[0010] The test piece was heated by an electric heating test platform under vacuum environment test conditions, and the temperature data of the platinum-rhodium thermocouple was collected. At the same time, the temperature was measured by a dual-color infrared thermometer.

[0011] Step S3: Using the temperature measured by the dual-color infrared thermometer in step S2 as a reference, establish the calibration curve of the platinum-rhodium thermocouple;

[0012] Step S4: The high-temperature anti-oxidation coating on the outer surface of the nozzle throat of the rocket engine nozzle is removed by machining to form an exposed area of ​​a preset width, while retaining a coating transition layer of a preset thickness to form the nozzle substrate. The surface roughness of the nozzle substrate is controlled within a preset threshold.

[0013] The platinum-rhodium thermocouple is fixed in contact with the nozzle substrate, and then connected to the measurement and control system to collect temperature data.

[0014] Step S5: Measure the temperature of the nozzle throat of the rocket engine under standard operating conditions, obtain the temperature Tr of the platinum-rhodium thermocouple, and calculate the mapping from the measured value to the true value based on the calibration curve of the platinum-rhodium thermocouple to obtain the temperature Tt of the nozzle throat.

[0015] Preferably, in step S2, a test piece with dimensions of 10mm × 70mm × 1.5mm is prepared;

[0016] Form an exposed area with a width of 3-6 mm, retain a coating transition layer of 10-20 μm, and control the surface roughness of the nozzle substrate to no more than 3.2 μm;

[0017] Adjust the test temperature to cover the standard temperature value Te and the temperature values ​​Te±50℃, Te±100℃, Te±150℃, and Te±200℃.

[0018] Preferably, in step S4, an exposed area with a width of 3 to 6 mm is formed, a coating transition layer of 10 to 20 μm is retained, and the surface roughness of the nozzle substrate is controlled to be no greater than 3.2 μm.

[0019] Preferably, the platinum-rhodium thermocouple is an S-type exposed thermocouple or a B-type exposed thermocouple.

[0020] Preferably, the temperature range of the platinum-rhodium thermocouple is not less than 1400°C.

[0021] Preferably, the material of the nozzle substrate of the rocket engine nozzle is Nb521 niobium-tungsten alloy.

[0022] Preferably, the high-temperature antioxidant coating is a silicon-chromium-titanium-hafnium silicide coating of grade 056.

[0023] Preferably, the ceramic protective tube is made of alumina or silicon carbide.

[0024] Preferably, the method for fixing the platinum-rhodium thermocouple to the nozzle substrate is as follows: first, a titanium sheet is used to press the platinum-rhodium thermocouple, and then the titanium sheet and the nozzle substrate are spot welded together using pulsed microbeam plasma arc welding.

[0025] The present invention also provides a rocket engine nozzle temperature measurement system, comprising the following modules:

[0026] Module M1: Under unobstructed conditions, a dual-color infrared thermometer is used to measure the nozzle throat temperature of the rocket engine nozzle under standard conditions, and this temperature is set as Te.

[0027] Module M2: Prepare test pieces and prepare high-temperature anti-oxidation coatings on both sides. The materials, coating preparation process and requirements of the test pieces are the same as those of rocket engine nozzles.

[0028] The high-temperature anti-oxidation coating in the middle of the test piece is removed by machining to form an exposed area of ​​a preset width, while retaining a coating transition layer of a preset thickness to form the nozzle substrate. The surface roughness of the nozzle substrate is controlled within a preset threshold.

[0029] The platinum-rhodium thermocouple is fixedly in contact with the nozzle substrate, and multiple ceramic protective tubes are configured at the thermocouple measuring end.

[0030] The test piece was heated by an electric heating test platform under vacuum environment test conditions, and the temperature data of the platinum-rhodium thermocouple was collected. At the same time, the temperature was measured by a dual-color infrared thermometer.

[0031] Module M3: Using the temperature measured by the dual-color infrared thermometer in Module M2 as a reference, establish the calibration curve of the platinum-rhodium thermocouple;

[0032] Module M4: The high-temperature anti-oxidation coating on the outer surface of the nozzle throat of the rocket engine nozzle is removed by machining to form an exposed area of ​​a preset width, while retaining a coating transition layer of a preset thickness to form the nozzle substrate. The surface roughness of the nozzle substrate is controlled within a preset threshold.

[0033] The platinum-rhodium thermocouple is fixed in contact with the nozzle substrate, and then connected to the measurement and control system to collect temperature data.

[0034] Module M5: Measures the temperature of the nozzle throat of a rocket engine under standard operating conditions, obtains the temperature Tr of the platinum-rhodium thermocouple, and calculates the temperature Tt of the nozzle throat by interpolation based on the calibration curve of the platinum-rhodium thermocouple to realize the mapping from the measured value to the true value.

[0035] Compared with the prior art, the present invention has the following beneficial effects:

[0036] 1. This invention solves the problem that the throat wall temperature, a key parameter of the nozzle, cannot be measured by an infrared thermometer when a rocket engine installed in a confined space such as a recess or sleeve undergoes a high-altitude simulated hot-fire test.

[0037] 2. This invention achieves a balance between high-temperature oxidation resistance and accurate temperature measurement through localized coating treatment; the measurement error caused by thermal conduction of the coating material is eliminated through electric heating calibration test, thereby improving the measurement accuracy of the nozzle throat wall temperature. Attached Figure Description

[0038] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0039] Figure 1 This is a schematic diagram of high-temperature nozzle measurement for a rocket engine in a confined space environment, according to an embodiment of the present invention.

[0040] Figure 2 for Figure 1 A magnified view of a portion of the throat of the central nozzle.

[0041] Figure 3 The calibration curve of the platinum-rhodium thermocouple on the test piece to retain the coating transition layer.

[0042] The diagram shows:

[0043] Dual-color infrared thermometer 1; rocket engine nozzle 2

[0044] Nozzle throat 21 High-temperature anti-oxidation coating 22

[0045] Coating transition layer 221 Nozzle substrate 23

[0046] Platinum-rhodium thermocouple 3 Ceramic protection tube 4

[0047] Concave compartment 5, titanium sheet 6 Detailed Implementation

[0048] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0049] Example 1

[0050] like Figures 1 to 3 As shown in the figure, this embodiment provides a method for measuring the temperature of a rocket engine nozzle, including the following steps:

[0051] Step S1: Under unobstructed conditions, use a dual-color infrared thermometer 1 to measure the temperature of the nozzle throat 21 of the rocket engine nozzle 2 under standard conditions, and set it as the standard temperature Te.

[0052] Step S2: Prepare a test piece and prepare a high-temperature anti-oxidation coating 22 on both sides. The material, coating preparation process, and requirements of the test piece are the same as those of the rocket engine nozzle 2. Remove the high-temperature anti-oxidation coating 22 in the middle of the test piece by machining to form an exposed area of ​​a preset width. Retain a coating transition layer 221 of a preset thickness to form the nozzle substrate 23. The surface roughness of the nozzle substrate 23 is controlled at a preset threshold. Fix the platinum-rhodium thermocouple 3 in contact with the nozzle substrate 23. Configure multiple ceramic protective tubes 4 at the thermocouple measuring end. Heat the test piece through an electric heating test platform under vacuum environment test conditions, collect the temperature data of the platinum-rhodium thermocouple 3, and measure the temperature with a dual-color infrared thermometer 1.

[0053] In step S2, a sample with dimensions of 10mm × 70mm × 1.5mm is prepared;

[0054] An exposed area with a width of 3 to 6 mm is formed, and a coating transition layer 221 of 10 to 20 μm is retained. The surface roughness of the nozzle substrate 23 is controlled to be no greater than 3.2 μm.

[0055] Adjust the test temperature to cover the standard temperature value Te and the temperature values ​​Te±50℃, Te±100℃, Te±150℃, and Te±200℃;

[0056] Step S3: Using the temperature measured by the dual-color infrared thermometer 1 in step S2 as a reference, establish the calibration curve of the platinum-rhodium thermocouple 3.

[0057] Step S4: The high-temperature anti-oxidation coating 22 on the outer surface of the nozzle throat 21 of the rocket engine nozzle 2 is removed by machining to form an exposed area of ​​a preset width. The coating transition layer 221 of a preset thickness is retained to form the nozzle substrate 23. The surface roughness of the nozzle substrate 23 is controlled within a preset threshold. The platinum-rhodium thermocouple 3 is fixedly contacted with the nozzle substrate 23. The platinum-rhodium thermocouple 3 is connected to the measurement and control system to collect temperature data.

[0058] In step S4, an exposed area with a width of 3 to 6 mm is formed, a coating transition layer 221 of 10 to 20 μm is retained, and the surface roughness of the nozzle substrate 23 is controlled to be no greater than 3.2 μm.

[0059] Step S5: Measure the temperature of the nozzle throat 21 of the rocket engine nozzle 2 under standard operating conditions, obtain the temperature Tr of the platinum-rhodium thermocouple 3, and calculate the mapping from the measured value to the true value based on the calibration curve of the platinum-rhodium thermocouple 3 to obtain the temperature Tt of the nozzle throat 21.

[0060] The platinum-rhodium thermocouple 3 is fixed in contact with the nozzle substrate 23 by first pressing the platinum-rhodium thermocouple 3 with a titanium sheet, and then spot welding the titanium sheet 6 to the nozzle substrate 23 using pulsed micro-beam plasma arc welding. The platinum-rhodium thermocouple 3 is an S-type exposed thermocouple or a B-type exposed thermocouple. The temperature resistance range of the platinum-rhodium thermocouple 3 is not lower than 1400℃. The material of the nozzle substrate 23 of the rocket engine nozzle 2 is Nb521 niobium-tungsten alloy. The high-temperature anti-oxidation coating 22 is a silicon-chromium-titanium hafnium silicide coating of grade 056. The ceramic protective tube 4 is made of alumina or silicon carbide.

[0061] The present invention also provides a rocket engine nozzle temperature measurement system, which can be implemented by executing the process steps of the rocket engine nozzle temperature measurement method. That is, those skilled in the art can understand the rocket engine nozzle temperature measurement method as a preferred embodiment of the rocket engine nozzle temperature measurement system.

[0062] Example 2

[0063] This embodiment provides a rocket engine nozzle temperature measurement system, including the following modules:

[0064] Module M1: Under unobstructed conditions, a dual-color infrared thermometer 1 is used to measure the temperature of the nozzle throat 21 of the rocket engine nozzle 2 under standard conditions, and set it as the standard temperature Te.

[0065] Module M2: Prepare a test piece and prepare a high-temperature anti-oxidation coating 22 on both sides. The material, coating preparation process, and requirements of the test piece are the same as those of the rocket engine nozzle 2. Remove the high-temperature anti-oxidation coating 22 in the middle of the test piece by machining to form an exposed area of ​​a preset width. Retain a coating transition layer 221 of a preset thickness to form the nozzle substrate 23. The surface roughness of the nozzle substrate 23 is controlled within a preset threshold. Fix the platinum-rhodium thermocouple 3 in contact with the nozzle substrate 23. Configure multiple ceramic protective tubes 4 at the thermocouple measuring end. Heat the test piece through an electric heating test platform under vacuum test conditions, collect the temperature data of the platinum-rhodium thermocouple 3, and measure the temperature with a dual-color infrared thermometer 1.

[0066] In module M2, a sample with dimensions of 10mm × 70mm × 1.5mm is prepared;

[0067] An exposed area with a width of 3 to 6 mm is formed, and a coating transition layer 221 of 10 to 20 μm is retained. The surface roughness of the nozzle substrate 23 is controlled to be no greater than 3.2 μm.

[0068] Adjust the test temperature to cover the standard temperature value Te and the temperature values ​​Te±50℃, Te±100℃, Te±150℃, and Te±200℃;

[0069] Module M3: Using the temperature measured by the dual-color infrared thermometer 1 in module M2 as a reference, establish the calibration curve of the platinum-rhodium thermocouple 3;

[0070] Module M4: The high-temperature anti-oxidation coating 22 on the outer surface of the nozzle throat 21 of the rocket engine nozzle 2 is removed by machining to form an exposed area of ​​a preset width, while retaining a coating transition layer 221 of a preset thickness to form the nozzle substrate 23. The surface roughness of the nozzle substrate 23 is controlled within a preset threshold. The platinum-rhodium thermocouple 3 is fixedly contacted with the nozzle substrate 23, and the platinum-rhodium thermocouple 3 is connected to the measurement and control system to collect temperature data.

[0071] In module M4, an exposed area with a width of 3 to 6 mm is formed, a coating transition layer 221 of 10 to 20 μm is retained, and the surface roughness of the nozzle substrate 23 is controlled to be no greater than 3.2 μm;

[0072] Module M5: Measure the temperature of the nozzle throat 21 of the rocket engine nozzle 2 under standard operating conditions, obtain the temperature Tr of the platinum-rhodium thermocouple 3, and calculate the mapping from the measured value to the true value based on the calibration curve of the platinum-rhodium thermocouple 3 to obtain the temperature Tt of the nozzle throat 21.

[0073] The platinum-rhodium thermocouple 3 is fixed in contact with the nozzle substrate 23 by first pressing the platinum-rhodium thermocouple 3 with a titanium sheet, and then spot welding the titanium sheet 6 to the nozzle substrate 23 using pulsed micro-beam plasma arc welding. The platinum-rhodium thermocouple 3 is an S-type exposed thermocouple or a B-type exposed thermocouple. The temperature resistance range of the platinum-rhodium thermocouple 3 is not lower than 1400℃. The material of the nozzle substrate 23 of the rocket engine nozzle 2 is Nb521 niobium-tungsten alloy. The high-temperature anti-oxidation coating 22 is a silicon-chromium-titanium hafnium silicide coating of grade 056. The ceramic protective tube 4 is made of alumina or silicon carbide.

[0074] Example 3

[0075] Those skilled in the art can understand this embodiment as a more specific description of Embodiment 1 and Embodiment 2.

[0076] This embodiment provides a method for accurate temperature measurement of high-temperature nozzles of rocket engines in confined spaces, including the following steps:

[0077] Step 1: Under unobstructed conditions, use a dual-color infrared thermometer to measure the temperature of the rocket engine nozzle throat under standard conditions, and set it as the standard temperature Te;

[0078] Step 2: Prepare a specimen with dimensions of 10mm × 70mm × 1.5mm, and apply a high-temperature anti-oxidation coating to both sides. The material, coating preparation process, and requirements of the specimen are consistent with those of the rocket engine nozzle. Remove the high-temperature anti-oxidation coating from the middle of the specimen by machining, forming an exposed area with a width of 3–6mm, retaining a 10–20μm coating transition layer. The surface roughness of the substrate is controlled to be no greater than 3.2μm. Fix the platinum-rhodium thermocouple in contact with the substrate, and equip the thermocouple measuring end with multiple ceramic protective tubes. Heat the specimen using an electric heating test platform under vacuum conditions, collect the temperature data of the platinum-rhodium thermocouple, and simultaneously measure the temperature using a dual-color infrared thermometer. Adjust the test temperature to cover the standard temperature value Te and the temperature values ​​Te±50℃, Te±100℃, Te±150℃, and Te±200℃.

[0079] Step 3: Using the temperature measured by the dual-color infrared thermometer in Step 2 as a reference, establish a platinum-rhodium thermocouple calibration curve;

[0080] Step 4: The high-temperature anti-oxidation coating on the outer surface of the rocket engine nozzle throat is removed by machining, forming an exposed area 3–6 mm wide, retaining a 10–20 μm coating transition layer. The surface roughness of the substrate is controlled to be no greater than 3.2 μm. A platinum-rhodium thermocouple is fixed in contact with the substrate, and the thermocouple is connected to the measurement and control system to collect temperature data.

[0081] Step 5: In a space-constrained environment, measure the temperature of the rocket engine nozzle throat under standard operating conditions to obtain the temperature Tr of the platinum-rhodium thermocouple. Based on the calibration curve of the platinum-rhodium thermocouple, interpolate to realize the mapping from the measured value to the true value and obtain the accurate throat temperature Tt.

[0082] The platinum-rhodium thermocouples are of type S (Pt-10%Rh / Pt) or type B (Pt-13%Rh / Pt or Pt-30%Rh / Pt-6%Rh) exposed type, with a temperature range of not less than 1400℃.

[0083] The rocket engine nozzle substrate is made of Nb521 niobium-tungsten alloy, the high-temperature anti-oxidation coating is a silicon-chromium-titanium-hafnium silicide coating of grade 056, and the ceramic protective tube is made of alumina or silicon carbide.

[0084] The method for fixing the platinum-rhodium thermocouple to the substrate involves first pressing the thermocouple with a titanium sheet, and then spot welding the titanium sheet to the nozzle substrate using pulsed micro-beam plasma arc welding to ensure a secure connection.

[0085] This embodiment provides a method for accurate temperature measurement of high-temperature rocket engine nozzles in a space-constrained environment, solving the problem that the throat wall temperature, a key parameter of the nozzle, cannot be measured using an infrared thermometer. The method includes the following steps: Step 1: Under unobstructed conditions, the throat temperature of the rocket engine nozzle under standard operating conditions is measured using a dual-color infrared thermometer; Step 2: A test piece and a high-temperature anti-oxidation coating are prepared. The high-temperature anti-oxidation coating in the middle of the test piece is removed by machining, forming an exposed area while retaining the coating transition layer. A platinum-rhodium thermocouple is fixed in contact with the substrate, and the test piece is heated using an electrically heated test platform under vacuum conditions. The temperature data of the platinum-rhodium thermocouple is collected, and the temperature is simultaneously measured using a dual-color infrared thermometer; Step 3: A platinum-rhodium thermocouple calibration curve is established; Step 4: The high-temperature anti-oxidation coating on the outer surface of the rocket engine nozzle throat is removed by machining using the method in Step 2; Step 5: Under space-constrained conditions, the throat temperature of the rocket engine nozzle under standard operating conditions is measured. Based on the platinum-rhodium thermocouple calibration curve, interpolation is performed to map the measured value to the true value, obtaining an accurate throat temperature.

[0086] Example 4

[0087] Those skilled in the art can understand this embodiment as a more specific description of Embodiment 1 and Embodiment 2.

[0088] like Figures 1-3 As shown, this embodiment provides a method for accurate temperature measurement of high-temperature nozzles of rocket engines in a confined space environment, including the following steps:

[0089] Step 1: Under unobstructed conditions, use a dual-color infrared thermometer to measure the temperature of the rocket engine nozzle throat under standard conditions, set as the standard temperature Te, and the measured value of the standard temperature Te is 1350℃.

[0090] Step 2: Prepare a test piece with dimensions of 10mm × 70mm × 1.5mm, and apply a high-temperature anti-oxidation coating to both sides. The material, coating preparation process, and requirements of the test piece are consistent with those of the rocket engine nozzle. Remove the high-temperature anti-oxidation coating from the middle of the test piece by machining, forming a 5mm wide exposed area, retaining a 15μm coating transition layer. The surface roughness of the substrate should not exceed 3.2μm. Fix the platinum-rhodium thermocouple to the substrate in contact. Multiple ceramic protective tubes are configured at the thermocouple measuring end to effectively prevent oxidation and corrosion of the thermocouple under high-temperature conditions, and to prevent the thermocouple from touching other areas, ensuring measurement accuracy. The appropriately thick coating transition layer forms an interdiffusion layer between the coating and the substrate, providing both high-temperature anti-oxidation properties and facilitating current breakdown during spot welding, thus fixing the platinum-rhodium thermocouple to the substrate. Under vacuum environmental testing conditions, the test piece was heated via an electrically heated test platform, and temperature data from the platinum-rhodium thermocouples were collected. Simultaneously, a dual-color infrared thermometer was used to measure the temperature, and the test temperature was adjusted. The temperatures measured by the dual-color infrared thermometer were: 1150℃, 1200℃, 1250℃, 1300℃, 1350℃, 1400℃, 1450℃, 1500℃, and 1550℃. Correspondingly, the temperatures measured by the platinum-rhodium thermocouples were: 792℃, 852℃, 896℃, 949℃, 1008℃, 1070℃, 1117℃, 1192℃, and 1190℃.

[0091] Step 3: Using the temperature measured by the dual-color infrared thermometer in Step 2 as a reference, establish a platinum-rhodium thermocouple calibration curve, such as... Figure 3 As shown;

[0092] Step 4: Machining is used to remove the high-temperature anti-oxidation coating from the outer surface of the rocket engine nozzle throat, creating a 5mm wide exposed area. A 15μm coating transition layer is retained, and the substrate surface roughness is no greater than 3.2μm. A platinum-rhodium thermocouple is fixed in contact with the substrate, and the thermocouple is connected to the measurement and control system to collect temperature data.

[0093] Step 5: With the recess 5 installed, measure the temperature of the rocket engine nozzle throat under standard operating conditions. The temperature Tr of the platinum-rhodium thermocouple is obtained as 1039℃. Based on the interpolation calculation of the platinum-rhodium thermocouple calibration curve, the measured value is mapped to the true value, and the accurate throat temperature Tt is obtained as 1375℃.

[0094] The platinum-rhodium thermocouple used is an S-type (Pt-10%Rh / Pt) exposed thermocouple with a long-term operating temperature limit of 1400℃ and a short-term operating temperature limit of 1600℃, which meets the test requirements.

[0095] The rocket engine nozzle substrate is made of Nb521 niobium-tungsten alloy, the high-temperature anti-oxidation coating is a silicon-chromium-titanium-hafnium silicide coating of grade 056, and the ceramic protective tube is made of alumina or silicon carbide.

[0096] The method for fixing the platinum-rhodium thermocouple to the substrate involves first pressing the thermocouple with a titanium sheet, and then spot welding the titanium sheet to the nozzle substrate using pulsed micro-beam plasma arc welding to ensure a secure connection.

[0097] This invention solves the problem that the throat wall temperature, a key parameter of the nozzle, cannot be measured using an infrared thermometer.

[0098] Those skilled in the art will understand that, besides implementing the system and its various devices, modules, and units provided by this invention in the form of purely computer-readable program code, the same functions can be achieved entirely through logical programming of the method steps, making the system and its various devices, modules, and units of this invention function in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers. Therefore, the system and its various devices, modules, and units provided by this invention can be considered as a hardware component, and the devices, modules, and units included therein for implementing various functions can also be considered as structures within the hardware component; alternatively, the devices, modules, and units for implementing various functions can be considered as both software modules implementing the method and structures within the hardware component.

[0099] In the description of this application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application.

[0100] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A method for measuring the temperature of a rocket engine nozzle, characterized in that, Includes the following steps: Step S1: Under unobstructed conditions, the temperature of the nozzle throat (21) of the rocket engine nozzle (2) under standard conditions is measured using a dual-color infrared thermometer (1) and set as the standard temperature Te. Step S2: Prepare a test piece and prepare a high-temperature anti-oxidation coating on both sides (22). The material of the test piece and the coating preparation process and requirements are the same as those of the rocket engine nozzle (2). The high-temperature anti-oxidation coating (22) in the middle of the test piece is removed by machining to form an exposed area of ​​a preset width, and a coating transition layer (221) of a preset thickness is retained to form a nozzle substrate (23). The surface roughness of the nozzle substrate (23) is controlled at a preset threshold. The platinum-rhodium thermocouple (3) is fixedly in contact with the nozzle substrate (23), and multiple ceramic protective tubes (4) are configured at the thermocouple measuring end; The test piece was heated by an electric heating test platform under vacuum environment test conditions, and the temperature data of the platinum-rhodium thermocouple (3) was collected. At the same time, the temperature was measured by a dual-color infrared thermometer (1). Step S3: Using the temperature measured by the dual-color infrared thermometer (1) in step S2 as a reference, establish the calibration curve of the platinum-rhodium thermocouple (3); Step S4: The high-temperature anti-oxidation coating (22) on the outer surface of the nozzle throat (21) of the rocket engine nozzle (2) is removed by machining to form an exposed area of ​​a preset width, and a coating transition layer (221) of a preset thickness is retained to form a nozzle substrate (23). The surface roughness of the nozzle substrate (23) is controlled at a preset threshold. The platinum-rhodium thermocouple (3) is fixed in contact with the nozzle substrate (23), and the platinum-rhodium thermocouple (3) is connected to the measurement and control system to collect temperature data; Step S5: Measure the temperature of the nozzle throat (21) of the rocket engine nozzle (2) under standard operating conditions, obtain the temperature Tr of the platinum-rhodium thermocouple (3), and calculate the mapping from the measured value to the true value based on the calibration curve of the platinum-rhodium thermocouple (3) to obtain the temperature Tt of the nozzle throat (21).

2. The method for measuring the temperature of a rocket engine nozzle according to claim 1, characterized in that, In step S2, a test piece with dimensions of 10mm × 70mm × 1.5mm is prepared. An exposed area with a width of 3-6 mm is formed, and a coating transition layer of 10-20 μm is retained (221). The surface roughness of the nozzle substrate (23) is controlled to be no greater than 3.2 μm. Adjust the test temperature to cover the standard temperature value Te and the temperature values ​​Te±50℃, Te±100℃, Te±150℃, and Te±200℃.

3. The method for measuring the temperature of a rocket engine nozzle according to claim 1, characterized in that, In step S4, an exposed area with a width of 3 to 6 mm is formed, a coating transition layer (221) of 10 to 20 μm is retained, and the surface roughness of the nozzle substrate (23) is controlled to be no greater than 3.2 μm.

4. The method for measuring the temperature of a rocket engine nozzle according to claim 1, characterized in that, The platinum-rhodium thermocouple (3) is an S-type exposed thermocouple or a B-type exposed thermocouple.

5. The method for measuring the temperature of a rocket engine nozzle according to claim 1, characterized in that, The temperature range of the platinum-rhodium thermocouple (3) is not less than 1400℃.

6. The method for measuring the temperature of a rocket engine nozzle according to claim 1, characterized in that, The material of the nozzle substrate (23) of the rocket engine nozzle (2) is Nb521 niobium-tungsten alloy.

7. The method for measuring the temperature of a rocket engine nozzle according to claim 1, characterized in that, The high-temperature antioxidant coating (22) is a silicon-chromium-titanium-hafnium silicide coating of grade 056.

8. The method for measuring the temperature of a rocket engine nozzle according to claim 1, characterized in that, The ceramic protective tube (4) is made of alumina or silicon carbide.

9. The method for measuring the temperature of a rocket engine nozzle according to claim 1, characterized in that, The method for fixing the platinum-rhodium thermocouple (3) to the nozzle substrate (23) is as follows: first, use a titanium sheet to press the platinum-rhodium thermocouple (3), and then use pulsed micro-beam plasma arc welding to spot weld the titanium sheet (6) to the nozzle substrate (23).

10. A rocket engine nozzle temperature measurement system, characterized in that, Includes the following modules: Module M1: Under unobstructed conditions, a dual-color infrared thermometer (1) is used to measure the temperature of the nozzle throat (21) of the rocket engine nozzle (2) under standard conditions, and set it as the standard temperature Te; Module M2: Prepare test pieces and prepare high-temperature anti-oxidation coatings on both sides (22). The materials, coating preparation process and requirements of the test pieces are the same as those of the rocket engine nozzle (2). The high-temperature anti-oxidation coating (22) in the middle of the test piece is removed by machining to form an exposed area of ​​a preset width, and a coating transition layer (221) of a preset thickness is retained to form a nozzle substrate (23). The surface roughness of the nozzle substrate (23) is controlled at a preset threshold. The platinum-rhodium thermocouple (3) is fixedly in contact with the nozzle substrate (23), and multiple ceramic protective tubes (4) are configured at the thermocouple measuring end; The test piece was heated by an electric heating test platform under vacuum environment test conditions, and the temperature data of the platinum-rhodium thermocouple (3) was collected. At the same time, the temperature was measured by a dual-color infrared thermometer (1). Module M3: Using the temperature measured by the dual-color infrared thermometer (1) in module M2 as a reference, establish the calibration curve of the platinum-rhodium thermocouple (3); Module M4: The high-temperature anti-oxidation coating (22) on the outer surface of the nozzle throat (21) of the rocket engine nozzle (2) is removed by machining to form an exposed area of ​​a preset width, and a coating transition layer (221) of a preset thickness is retained to form a nozzle substrate (23). The surface roughness of the nozzle substrate (23) is controlled at a preset threshold. The platinum-rhodium thermocouple (3) is fixed in contact with the nozzle substrate (23), and the platinum-rhodium thermocouple (3) is connected to the measurement and control system to collect temperature data; Module M5: Measure the temperature of the nozzle throat (21) of the rocket engine nozzle (2) under standard operating conditions, obtain the temperature Tr of the platinum-rhodium thermocouple (3), and calculate the mapping from the measured value to the true value based on the calibration curve of the platinum-rhodium thermocouple (3) to obtain the temperature Tt of the nozzle throat (21).

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