Liquid rocket thrust chamber test run inner wall temperature calculation method and calculation equipment

Abstract

The invention provides a liquid rocket thrust chamber test run inner wall temperature calculation method and calculation equipment. The calculation method comprises the steps that distribution data ofan outer wall temperature measurement value of a thrust chamber along the axis of the thrust chamber are acquired; the thrust chamber is divided into n temperature sections in the axis direction of the thrust chamber, wherein n is larger than or equal to 2; the outer wall temperature measurement value i of the ith temperature section is greater than or equal to 1 and less than or equal to n; thevalue of i is gradually increased in the direction from the upstream of a cooling flow path to the downstream of the cooling flow path; according to a heat transfer function relationship, calculationis conducted to obtain the heat transfer amount of the inner wall of the ith temperature section when the calculated value of the outer wall temperature of the ith temperature section is equal to themeasured value of the outer wall temperature of the ith temperature section, and calculation is conducted to obtain the cooling outlet temperature Tci and the inner wall temperature of the ith temperature section corresponding to the heat transfer amount of the inner wall of the ith temperature section. According to the calculation method disclosed by the invention, the thrust chamber inner wall temperature can be obtained according to a test run experiment result.

Description

technical field [0001] The present disclosure relates to the technical field of liquid rocket thrust chamber test run, and in particular to a calculation method and computing equipment for the inner wall temperature of liquid rocket thrust chamber test run. Background technique [0002] In the thrust chamber of a liquid rocket engine, the propellant components are burned at high temperature (3000-4000K) and high pressure (5-20MPa or higher), and the high-temperature and high-pressure gas flow rate is high (throat flow rate is as high as 1000-1500m / s) , the heat flux passing through the wall of the thrust chamber is high (10~160MW / m 2 ), therefore, thermal protection is an important consideration in thrust chamber design. In order to ensure the thermal strength and structural stability of the thrust chamber structure, thermal protection measures need to be taken, which can usually be divided into external cooling (regenerative cooling, discharge cooling, radiation cooling), ...

AI Technical Summary

Problems solved by technology

However, only the temperature of the outer wall can be obtained from the experimental results, and the temperature of the inner wall of the thrust cannot be obtained from the existing test runs

Therefore, it is impossible to compare the test results with theoretical calculations, so that there is no quantitative data reference for the next thrust chamber thermal protection design

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