CN-120291988-B - Rocket engine spray pipe temperature measurement method and system

CN120291988BCN 120291988 BCN120291988 BCN 120291988BCN-120291988-B

Abstract

The invention provides a temperature measuring method and a temperature measuring system for a rocket engine nozzle, which comprise the following steps of S1, setting standard temperature Te under a non-shielding working condition, S2, preparing a test piece, forming a matrix, enabling a platinum-rhodium thermocouple to be in fixed contact with the matrix, configuring a plurality of ceramic protection tubes at a thermocouple measuring end, collecting temperature data of the platinum-rhodium thermocouple, measuring temperature by using a bicolor infrared thermometer, S3, establishing a platinum-rhodium thermocouple calibration curve by taking the temperature measured by the bicolor infrared thermometer in S2 as a reference, S4, removing the outer surface of a throat of the rocket engine nozzle through machining to form the matrix, enabling the platinum-rhodium thermocouple to be in fixed contact with the matrix, enabling the thermocouple to be connected with a measurement and control system to collect temperature data, and S5, obtaining the throat temperature Tt according to interpolation calculation of the platinum-rhodium thermocouple calibration curve. The invention solves the problem that the throat wall temperature of the key parameter of the spray pipe can not be measured by an infrared thermometer.

Inventors

YAO FENG
CHEN RUIDA
ZHAO TING
LIU CHANGGUO
TIAN ZENG
LU WENJIE
WU LINGFENG

Assignees

上海空间推进研究所

Dates

Publication Date: 20260512
Application Date: 20250410

Claims (10)

1. The temperature measurement method of the rocket engine nozzle is characterized by comprising the following steps of: Step S1, under the non-shielding working condition, measuring the temperature of a nozzle throat (21) of a rocket engine nozzle (2) under the standard working condition by adopting a bicolor infrared thermometer (1), and setting the temperature as a standard temperature Te; S2, preparing a test piece, and preparing a high-temperature oxidation-resistant coating (22) on both sides, wherein the material and the coating preparation process and the requirements of the test piece are consistent with those of a rocket engine nozzle (2); the high-temperature oxidation-resistant coating (22) at the middle position of the test piece is removed through turning to form a naked area with a preset width, a coating transition layer (221) with a preset thickness is reserved to form a spray pipe matrix (23), and the surface roughness of the spray pipe matrix (23) is controlled to be a preset threshold value; a platinum rhodium thermocouple (3) is fixedly contacted with a spray pipe substrate (23), and a plurality of ceramic protection pipes (4) are arranged at the thermocouple measuring end; heating a test piece through an electric heating test platform in a vacuum environment test state, collecting temperature data of a platinum-rhodium thermocouple (3), and measuring the temperature by a bicolor infrared thermometer (1); S3, establishing a calibration curve of the platinum-rhodium thermocouple (3) by taking the temperature measured by the bicolor infrared thermometer (1) in the step S2 as a reference; s4, turning to remove a high-temperature oxidation-resistant coating (22) on the outer surface of a nozzle throat (21) of a rocket engine nozzle (2) to form a bare area with a preset width, reserving a coating transition layer (221) with a preset thickness to form a nozzle substrate (23), and controlling the surface roughness of the nozzle substrate (23) to be a preset threshold value; Fixedly contacting a platinum-rhodium thermocouple (3) with a spray pipe matrix (23), and connecting the platinum-rhodium thermocouple (3) into a measurement and control system to acquire temperature data; And S5, measuring the temperature of a nozzle throat (21) of the rocket engine nozzle (2) under the standard working condition, obtaining the temperature Tr of the platinum-rhodium thermocouple (3), and carrying out interpolation calculation according to a calibration curve of the platinum-rhodium thermocouple (3) to realize the mapping from the measured value to the true value so as to obtain the temperature Tt of the nozzle throat (21).
2. A rocket engine nozzle temperature measurement method according to claim 1, wherein in step S2, test pieces having dimensions of 10mm x 70mm x 1.5mm are prepared; Forming a bare area with the width of 3-6 mm, reserving a coating transition layer (221) with the thickness of 10-20 mu m, and controlling the surface roughness of the spray pipe substrate (23) to be not more than 3.2 mu m; The test temperature was adjusted to cover standard temperature values Te and Te.+ -. 50° C, te.+ -. 100° C, te.+ -. 150° C, te.+ -. 200 ℃ temperature values.
3. The rocket engine nozzle temperature measurement method according to claim 1, wherein in the step S4, a bare area with a width of 3-6 mm is formed, a coating transition layer (221) with a thickness of 10-20 μm is reserved, and the surface roughness of the nozzle substrate (23) is controlled to be not more than 3.2 μm.
4. Rocket engine nozzle temperature measurement method according to claim 1, characterized in that the platinum rhodium thermocouple (3) is an S-type outcrop thermocouple or a B-type outcrop thermocouple.
5. Rocket engine nozzle temperature measurement method according to claim 1, characterized in that the temperature resistant range of the platinum rhodium thermocouple (3) is not lower than 1400 ℃.
6. Rocket engine nozzle temperature measurement method according to claim 1, characterized in that the nozzle base (23) of the rocket engine nozzle (2) is made of Nb521 niobium tungsten alloy.
7. A rocket engine nozzle temperature measurement method according to claim 1, wherein the high temperature oxidation resistant coating (22) is a silicon chromium titanium hafnium silicide coating of grade 056.
8. Rocket engine nozzle temperature measurement method according to claim 1, characterized in that the material of the ceramic protection tube (4) is alumina or silicon carbide.
9. The rocket engine nozzle temperature measurement method according to claim 1, wherein the method for fixedly contacting the platinum-rhodium thermocouple (3) with the nozzle base body (23) is that firstly, a titanium sheet is used for pressing the platinum-rhodium thermocouple (3), and then, a pulse microbeam plasma arc welding spot is adopted for connecting the titanium sheet (6) with the nozzle base body (23).
10. A rocket engine nozzle temperature measurement system, comprising the following modules: the module M1 is used for measuring the temperature of a nozzle throat (21) of a rocket engine nozzle (2) under standard working conditions by adopting a bicolor infrared thermometer (1) under the non-shielding working conditions, and setting the temperature as the standard temperature Te; preparing a test piece, and preparing a high-temperature oxidation-resistant coating (22) on two sides, wherein the material and the coating preparation process and the requirements of the test piece are consistent with those of a rocket engine nozzle (2); the high-temperature oxidation-resistant coating (22) at the middle position of the test piece is removed through turning to form a naked area with a preset width, a coating transition layer (221) with a preset thickness is reserved to form a spray pipe matrix (23), and the surface roughness of the spray pipe matrix (23) is controlled to be a preset threshold value; a platinum rhodium thermocouple (3) is fixedly contacted with a spray pipe substrate (23), and a plurality of ceramic protection pipes (4) are arranged at the thermocouple measuring end; heating a test piece through an electric heating test platform in a vacuum environment test state, collecting temperature data of a platinum-rhodium thermocouple (3), and measuring the temperature by a bicolor infrared thermometer (1); the module M3 is used for establishing a calibration curve of the platinum-rhodium thermocouple (3) by taking the temperature measured by the bicolor infrared thermometer (1) in the module M2 as a reference; The module M4 is used for removing a high-temperature oxidation-resistant coating (22) on the outer surface of a nozzle throat (21) of the rocket engine nozzle (2) through turning to form a naked area with a preset width, reserving a coating transition layer (221) with a preset thickness to form a nozzle matrix (23), and controlling the surface roughness of the nozzle matrix (23) to be a preset threshold value; Fixedly contacting a platinum-rhodium thermocouple (3) with a spray pipe matrix (23), and connecting the platinum-rhodium thermocouple (3) into a measurement and control system to acquire temperature data; And the module M5 is used for measuring the temperature of the nozzle throat (21) of the rocket engine nozzle (2) under the standard working condition, obtaining the temperature Tr of the platinum-rhodium thermocouple (3), carrying out interpolation calculation according to the calibration curve of the platinum-rhodium thermocouple (3), and realizing the mapping from the measured value to the true value to obtain the temperature Tt of the nozzle throat (21).

Description

Rocket engine spray pipe temperature measurement method and system Technical Field The invention relates to the technical field of testing of liquid rocket engines of spacecrafts, in particular to a rocket engine spray pipe temperature measuring method and system, in particular to a rocket engine high-temperature spray pipe accurate temperature measuring method under a space-limited environment, which is particularly suitable for high-altitude simulated thermal test verification of rocket engines arranged in a concave cabin or a sleeve. Background The temperature of the outer wall of the throat part of the spray pipe of the space attitude and orbit control liquid rocket engine for the spacecraft is up to 1400-1550 ℃, the base material of the spray pipe is niobium alloy, high-temperature oxidation-resistant silicide coatings are prepared on the inner surface and the outer surface, and in the prior art, non-contact measurement is mainly carried out by depending on a single-color or double-color infrared thermometer. When the spacecraft assembly layout is compact, the engine spray pipe is required to be embedded into the concave cabin, and the throat and the area above the throat are shielded. In addition, reusable spacecraft must be of telescopic construction to accommodate aerodynamic loads and thermal environments during the round trip, with the rocket engine high temperature nozzle hidden within the sleeve while blocking the thermal impact of the nozzle on the surrounding heat sensitive components. In order to ensure flight reliability, the on-orbit vacuum environment and the installation environment of the spacecraft must be simulated on the ground, and test run check verification is carried out on the engine. However, the infrared light path is blocked by the mechanical structure of the concave cabin or the sleeve, the accurate temperature of the outer wall of the throat cannot be obtained, the temperature of the outer wall of the engine spray pipe is high, the silicide coating coated on the outer surface is poor in conductivity, the temperature cannot be measured by a direct spot welding thermocouple, the conventional K-type thermocouple (the highest temperature of 1200 ℃ and the T-type thermocouple (the highest temperature of 400 ℃) is insufficient in temperature resistance, and the measuring error caused by heat conduction of a coating material cannot be eliminated by the conventional contact type measuring means. The patent document with the publication number of CN119572375A discloses an automatic temperature measuring device after test of a solid rocket engine and a using method, the automatic temperature measuring device comprises a mechanical arm and a temperature measuring device, the temperature measuring device is connected with the mechanical arm through a flange plate and comprises a temperature measuring rod, a temperature measuring rod heat insulation layer and a thermocouple, wherein the temperature measuring rod is provided with a wiring groove and a plurality of through holes, the wiring groove is used for arranging a thermocouple wire harness, the through holes are used for fixing the thermocouple, and the temperature measuring rod heat insulation layer is coated on the outer side of the temperature measuring rod. However, the patent document still has the defect that measurement errors caused by heat conduction of the coating material cannot be eliminated. Disclosure of Invention Aiming at the defects in the prior art, the invention aims to provide a rocket engine nozzle temperature measurement method and system. The invention provides a rocket engine nozzle temperature measurement method, which comprises the following steps: Step S1, under the non-shielding working condition, measuring the nozzle throat temperature of a rocket engine nozzle under the standard working condition by adopting a bicolor infrared thermometer, and setting the nozzle throat temperature as the standard temperature Te; S2, preparing a test piece, and preparing a high-temperature oxidation-resistant coating on two sides, wherein the material and the coating preparation process and the requirements of the test piece are consistent with those of a rocket engine spray pipe; The high-temperature oxidation-resistant coating at the middle position of the test piece is removed through turning to form a naked area with a preset width, a coating transition layer with a preset thickness is reserved to form a spray pipe matrix, and the surface roughness of the spray pipe matrix is controlled at a preset threshold value; Fixedly contacting a platinum-rhodium thermocouple with a spray pipe matrix, wherein a plurality of ceramic protection pipes are arranged at a thermocouple measuring end; Heating a test piece through an electric heating test platform in a vacuum environment test state, collecting temperature data of a platinum-rhodium thermocouple, and measuring the temperature by a bicolor infrared thermometer; S3, establ