Pneumatic-thermal collaborative optimization method for scallop hole air film cooling structure of turbine blade

A technology of film cooling and turbine blades, applied in special data processing applications, instruments, electrical digital data processing, etc., can solve the problems of high cost and time-consuming, etc., achieve less resource consumption, strong global approximation ability, and reduce all time-consuming and cost-effective effects

Active Publication Date: 2017-09-22
NANJING UNIV OF AERONAUTICS & ASTRONAUTICS
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  • Abstract
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Problems solved by technology

The optimization of the traditional film hole cooling structure needs to rely on a large number of experiments and numerical simulations to summarize the rules, which not only takes a lot of time, but also costs relatively high

Method used

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  • Pneumatic-thermal collaborative optimization method for scallop hole air film cooling structure of turbine blade
  • Pneumatic-thermal collaborative optimization method for scallop hole air film cooling structure of turbine blade
  • Pneumatic-thermal collaborative optimization method for scallop hole air film cooling structure of turbine blade

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Embodiment Construction

[0061] The present invention will be further explained below in conjunction with the accompanying drawings.

[0062] figure 1 It is a flow chart of the aerodynamic-thermal synergistic optimization method for a fan-shaped hole film cooling structure of a turbine blade of the present invention, and will be referred to below figure 1 This method will be described.

[0063] Step 1, based on figure 2 The fan-shaped air film cooling geometric model of the turbine blade, the mainstream inlet velocity is 140m / s, the main flow inlet temperature is 540K, the fan-shaped air film hole cold air inlet velocity is 80.4m / s, and the air film hole cold air inlet temperature is 310K.

[0064] It is stipulated that the fan hole spacing P is 4.5D, the fan hole height h is 2.5D, and the fan hole forward expansion length is l 1 is D, the lateral extension length of the fan-shaped hole l 2 is 2D, the length of the cylindrical part of the fan-shaped hole l 3 is (h / sinα-3D). Select turbine blade...

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Abstract

The invention discloses a pneumatic-thermal collaborative optimization method for a scallop hole air film cooling structure of a turbine blade. According to the method, a pneumatic-thermal feature agent model of the scallop hole air film cooling structure of the turbine blade is established based on a radial basis neural network, and a particle swarm optimization algorithm is introduced to realize pneumatic-thermal collaborative optimization of the scallop hole air film cooling structure of the turbine blade. The method overcomes the defect that a traditional air film cooling structure optimization method needs to depend on a large quantity of samples. The method has a nonlinear prediction ability, high prediction precision, a strong memorizing ability, high robustness and a good global approximation ability. Moreover, weights of pneumatic performance and thermal performance can be adjusted according to actual demands to design an optimal scallop hole air film cooling structure with both the pneumatic performance and the thermal performance.

Description

technical field [0001] The invention relates to an aerodynamic-thermal synergistic optimization method for a fan-shaped hole film cooling structure of a turbine blade, and belongs to the technical field of enhanced cooling. Background technique [0002] During the technical development of aviation gas turbine engines, the pressure ratio of the compressor and the gas temperature at the inlet of the turbine show a trend of increasing. According to the future development trend of advanced aviation gas turbine engine technology, the thrust-to-weight ratio of the new generation engine will reach about 15, and the gas temperature at the turbine inlet will reach 2200K-2300K. The increase in the gas temperature at the turbine inlet greatly intensifies the thermal load on the hot-end components such as the turbine blades, the combustion chamber flame tube, and the exhaust nozzle; at the same time, the increase in the compression ratio also leads to a decrease in the quality of coolin...

Claims

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Application Information

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Patent Type & Authority Applications(China)
IPC IPC(8): G06F17/50G06N3/00
CPCG06F30/17G06N3/006
Inventor 黄莺张靖周王春华
Owner NANJING UNIV OF AERONAUTICS & ASTRONAUTICS
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