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Thickness optimization design method for thermal barrier coatings of turbine blade

A technology of turbine blades and thermal barrier coatings, applied in the field of optimization design of the thickness of turbine blades thermal barrier coatings, can solve the problems of inability to optimize the design of thickness distribution, lack of evaluation of the pros and cons of TBCs thickness design, etc., and achieve simple and efficient TBCs thickness optimization Design, reduced analysis calculations, effects of low stress levels

Active Publication Date: 2017-05-10
XI AN JIAOTONG UNIV
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Problems solved by technology

However, the current existing technology can only qualitatively give the rough distribution of the thickness of TBCs of turbine blades, and cannot realize the accurate optimal design of thickness distribution, and lacks the evaluation of the advantages and disadvantages of TBCs thickness design

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  • Thickness optimization design method for thermal barrier coatings of turbine blade
  • Thickness optimization design method for thermal barrier coatings of turbine blade
  • Thickness optimization design method for thermal barrier coatings of turbine blade

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Embodiment Construction

[0048] The present invention will be further described in detail below in conjunction with specific embodiments, which are explanations of the present invention rather than limitations.

[0049] The optimized design process of the present invention is as figure 1 As shown, in order to better understand the technical scheme of the present invention, the above optimization design method is applied to the TBCs thickness design of the gas turbine blade, the turbine blade is as figure 2 shown.

[0050] The specific process of this embodiment includes the following steps:

[0051] Step 1: Determine the thickness of each layer of TBCs.

[0052] The TBCs include a ceramic layer, an adhesive layer and a thermally grown oxide layer. The thickness of the selected ceramic layer is k×100 μm, where k is the number of analysis times, that is, the number of repetitions from step 1 to step 6, k=1,2,…,10. In this embodiment, the selected thickness of the adhesive layer is 150 μm, and the s...

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Abstract

The invention discloses a thickness optimization design method for thermal barrier coatings (TBCs) of a turbine blade, and belongs to the technical field of surface coating protection. Representative nodes uniformly distributed on the turbine blade are selected, the states of local regions are reflected by temperatures and stress results of representative node positions, and TBCs thickness analysis of a complex blade is equivalent to thickness optimization design of a finite number of representative node positions, so that the analysis calculation amount is reduced; a mathematic formula is established for reflecting design objectives of high heat insulation performance, low stress level and low preparation cost, calculation is performed by introducing a multi-objective optimization algorithm to obtain optimal ceramic layer thickness of each representative node position, and a total objective function value is taken as a TBCs thickness optimization design and evaluation parameter of the blade, so that the advantages and disadvantages of a TBCs thickness distribution scheme can be quantitatively evaluated and the shortcoming that an existing method only can perform qualitative evaluation is overcome; and the method can ensure the service safety of the coatings and improve the usage efficiency of the coatings.

Description

technical field [0001] The invention belongs to the technical field of surface coating protection, and in particular relates to an optimal design method for the thickness of a thermal barrier coating of a turbine blade. Background technique [0002] Thermal barrier coatings (TBCs) are advanced ceramic-metal multilayer material systems widely used in modern turbine engines. By coating TBCs with low thermal conductivity on the surface of high-temperature hot-end parts such as engine combustion chambers and turbine blades, on the one hand, the surface temperature of the metal parts can be reduced or the gas inlet temperature of the engine turbine blades can be further increased, and on the other hand, the metal parts can be protected from Corrosion and oxidation of high-temperature gas, so as to achieve the purpose of prolonging the service life of hot-end components and improving engine efficiency. Taking the turbine blade of a gas turbine as an example, the combined effect o...

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Application Information

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Patent Type & Authority Applications(China)
IPC IPC(8): G06F17/50
CPCG06F2111/06G06F2111/04G06F30/17G06F30/23
Inventor 李彪王铁军范学领李定骏江鹏
Owner XI AN JIAOTONG UNIV
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