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

A technology of liquid rockets and calculation methods, applied in design optimization/simulation, special data processing applications, etc., can solve the problems of incomparability and the inability to obtain the wall temperature of the thrust chamber, and achieve the effect of improving the probability of success and facilitating structural parameters.

Pending Publication Date: 2020-11-17
江苏深蓝航天有限公司
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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

Method used

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

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Embodiment Construction

[0033] The present disclosure will be further described in detail below with reference to the drawings and embodiments. It can be understood that the specific implementation manners described here are only used to explain relevant content, rather than to limit the present disclosure. It should also be noted that, for ease of description, only parts related to the present disclosure are shown in the drawings.

[0034] It should be noted that, in the case of no conflict, the implementation modes and the features in the implementation modes in the present disclosure can be combined with each other. The present disclosure will be described in detail below with reference to the drawings and embodiments.

[0035] see figure 2 A simplified schematic of the thrust chamber heat transfer structure shown, and image 3 A schematic diagram of the enlarged structure is shown. The side wall of the thrust chamber extends in a cylindrical shape around the axis of the thrust chamber, and g...

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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), ...

Claims

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Application Information

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IPC IPC(8): G06F30/20
CPCG06F30/20
Inventor 不公告发明人
Owner 江苏深蓝航天有限公司
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