有限元方法分析冷却气膜孔对涡轮叶片TBCs温度场和应力场的影响
发布时间：2018-08-25 18:31
【摘要】：热障涂层因其较低的导热率具有良好的隔热效果,能够有效提高涡轮叶片的工作温度,改善发动机的工作效率,提高飞机的推重比,对于提高国力具有举足轻重的作用,广泛应用于航空发动机涡轮叶片。因热失配而产生的残余应力是热障涂层不可避免的,也是导致涂层失效的关键因素。本文以考虑冷却气膜孔的实际热障涂层涡轮叶片作为研究对象,采用流固耦合的方法,分别使用FLUENT软件和ABAQUS软件建立流体域计算模型和固体域计算模型。主要研究内容如下:(1)建立带冷却气膜孔涡轮叶片热障涂层计算模型和流体域计算模型。围绕几何模型的构建和网格划分提出了详细的方法和介绍,并赋予材料参数,设置分析步和加载。流体域采用有限体积法解流体动力学方程,涡轮叶片采用有限单元法解固体热应力方程,通过第三方软件MPCCI实现流体计算域和固体计算域的联合模拟仿真。(2)考虑单列冷却气膜孔的热障涂层涡轮叶片温度场和应力场的结果分析。流固耦合仿真计算得到单列冷却气膜孔的热障涂层涡轮叶片的高温稳态温度场,分析发现热障涂层具有较好的隔热效果,平均隔热达到了100K左右,压力面和吸力面涂层的隔热效果要优于前缘和后缘;冷却气膜覆盖的区域由于气膜冷却和热障涂层的共同作用具有最好的隔热效果,并且冷却气膜的冷却效果占据了主导作用,涂层的隔热效果相较于其他区域并不明显。并基于共轭温度场计算得到热障涂层冷却至室温后的残余应力。研究结果表明在气膜孔最右端的涂层表面是最有可能由于水平残余应力11?而导致剥落破坏的危险区域;在气膜孔最左端的涂层界面处则是由法向残余应力22?产生界面裂纹的最可能出现区域。(3)对不含冷却气膜孔的热障涂层涡轮叶片模型与带冷却气膜孔的热障涂层涡轮叶片模型进行比较分析。对于不含冷却气膜孔的叶片模型,其在温度较高的前缘和尾缘位置涂层的隔热效果要优于温度相对较低的压力面和吸力面;热障涂层中陶瓷层的热应力高于过渡层,并且在陶瓷层和过渡层中热应力都是在前缘和尾缘两侧的叶根部位最大,热障涂层剥落失效易发生在这些位置。而对于含冷却气膜孔的叶片模型,其压力面和吸力面涂层的隔热效果要优于前缘和后缘;其热障涂层中也是陶瓷层的热应力高于过渡层,但在气膜孔处出现应力集中,陶瓷层和过渡层中热应力在气膜孔处最大,故热障涂层剥落失效易发生气膜孔处。总之,本文采用流固耦合分析方法实现了实际带冷却气膜孔的热障涂层涡轮叶片高温稳态温度场和残余应力的有限元模拟,分析了冷却气膜孔对涡轮叶片热障涂层温度场和应力场的影响,为真实叶片涂层的失效预测提供了一些依据。
[Abstract]:Thermal barrier coating has good thermal insulation effect because of its low thermal conductivity. It can effectively improve the working temperature of turbine blade, improve the efficiency of engine, and increase the ratio of propulsion to weight of aircraft, which plays an important role in improving the national strength. Widely used in aero-engine turbine blades. The residual stress caused by thermal mismatch is inevitable and the key factor leading to the failure of thermal barrier coating. In this paper, the actual thermal barrier coated turbine blades with cooling film holes are taken as the research object. The fluid-solid coupling method is used to establish the fluid domain calculation model and the solid domain calculation model by using FLUENT software and ABAQUS software, respectively. The main contents are as follows: (1) the thermal barrier coating calculation model and fluid domain calculation model of turbine blade with cooling film hole are established. This paper presents a detailed method and introduction about the construction of geometric model and mesh generation, and gives the material parameters, sets the analysis steps and loads. The hydrodynamic equation is solved by finite volume method in fluid domain, and the solid thermal stress equation is solved by finite element method for turbine blade. The third party software MPCCI is used to realize the joint simulation of fluid and solid fields. (2) the results of temperature field and stress field of heat-barrier coated turbine blade considering single-row cooling film holes are analyzed. The high temperature steady state temperature field of the turbine blade with single row cooling film hole is obtained by fluid-solid coupling simulation. It is found that the thermal barrier coating has a good thermal insulation effect, and the average thermal insulation is about 100K. The thermal insulation effect of the pressure surface and suction surface coating is better than that of the front and rear edge, the area covered by the cooling film has the best thermal insulation effect due to the co-action of the film cooling and the thermal barrier coating, and the cooling effect of the cooling film occupies the leading role. The thermal insulation of the coating is not obvious compared with other regions. The residual stress of thermal barrier coating after cooling to room temperature was calculated based on conjugate temperature field. The results show that the coating surface at the right end of the film hole is most likely due to the horizontal residual stress of 11? At the far left end of the film hole, the coating interface is from normal to residual stress 22? (3) the model of thermal barrier coating turbine blade without cooling film hole is compared with that of thermal barrier coating turbine blade model with cooling film hole. For the blade model without cooling film hole, the thermal insulation effect of the coating on the front edge and the tail edge at higher temperature is better than that on the pressure surface and suction surface with relatively low temperature, and the thermal stress of the ceramic layer in the thermal barrier coating is higher than that in the transition layer. The thermal stress in ceramic layer and transition layer is the largest at the two sides of the front edge and the tail edge, and the flaking failure of the thermal barrier coating is easy to occur in these positions. For the blade model with cooling film hole, the heat insulation effect of pressure surface and suction surface coating is better than that of leading edge and rear edge, and the thermal stress of ceramic coating is higher than that of transition layer, but the stress concentration appears in the gas film hole. The thermal stress in ceramic layer and transition layer is the largest at the gas film hole, so the flaking failure of thermal barrier coating is easy to happen at the film hole. In a word, the fluid-solid coupling analysis method is used to simulate the high temperature steady-state temperature field and residual stress of the thermal barrier coated turbine blade with cooling film hole. The effect of cooling film hole on temperature field and stress field of turbine blade thermal barrier coating is analyzed, which provides some basis for the failure prediction of real blade coating.
【学位授予单位】：湘潭大学
【学位级别】：硕士
【学位授予年份】：2017
【分类号】：V232.4;TG174.4
本文编号：2203741
[Abstract]:Thermal barrier coating has good thermal insulation effect because of its low thermal conductivity. It can effectively improve the working temperature of turbine blade, improve the efficiency of engine, and increase the ratio of propulsion to weight of aircraft, which plays an important role in improving the national strength. Widely used in aero-engine turbine blades. The residual stress caused by thermal mismatch is inevitable and the key factor leading to the failure of thermal barrier coating. In this paper, the actual thermal barrier coated turbine blades with cooling film holes are taken as the research object. The fluid-solid coupling method is used to establish the fluid domain calculation model and the solid domain calculation model by using FLUENT software and ABAQUS software, respectively. The main contents are as follows: (1) the thermal barrier coating calculation model and fluid domain calculation model of turbine blade with cooling film hole are established. This paper presents a detailed method and introduction about the construction of geometric model and mesh generation, and gives the material parameters, sets the analysis steps and loads. The hydrodynamic equation is solved by finite volume method in fluid domain, and the solid thermal stress equation is solved by finite element method for turbine blade. The third party software MPCCI is used to realize the joint simulation of fluid and solid fields. (2) the results of temperature field and stress field of heat-barrier coated turbine blade considering single-row cooling film holes are analyzed. The high temperature steady state temperature field of the turbine blade with single row cooling film hole is obtained by fluid-solid coupling simulation. It is found that the thermal barrier coating has a good thermal insulation effect, and the average thermal insulation is about 100K. The thermal insulation effect of the pressure surface and suction surface coating is better than that of the front and rear edge, the area covered by the cooling film has the best thermal insulation effect due to the co-action of the film cooling and the thermal barrier coating, and the cooling effect of the cooling film occupies the leading role. The thermal insulation of the coating is not obvious compared with other regions. The residual stress of thermal barrier coating after cooling to room temperature was calculated based on conjugate temperature field. The results show that the coating surface at the right end of the film hole is most likely due to the horizontal residual stress of 11? At the far left end of the film hole, the coating interface is from normal to residual stress 22? (3) the model of thermal barrier coating turbine blade without cooling film hole is compared with that of thermal barrier coating turbine blade model with cooling film hole. For the blade model without cooling film hole, the thermal insulation effect of the coating on the front edge and the tail edge at higher temperature is better than that on the pressure surface and suction surface with relatively low temperature, and the thermal stress of the ceramic layer in the thermal barrier coating is higher than that in the transition layer. The thermal stress in ceramic layer and transition layer is the largest at the two sides of the front edge and the tail edge, and the flaking failure of the thermal barrier coating is easy to occur in these positions. For the blade model with cooling film hole, the heat insulation effect of pressure surface and suction surface coating is better than that of leading edge and rear edge, and the thermal stress of ceramic coating is higher than that of transition layer, but the stress concentration appears in the gas film hole. The thermal stress in ceramic layer and transition layer is the largest at the gas film hole, so the flaking failure of thermal barrier coating is easy to happen at the film hole. In a word, the fluid-solid coupling analysis method is used to simulate the high temperature steady-state temperature field and residual stress of the thermal barrier coated turbine blade with cooling film hole. The effect of cooling film hole on temperature field and stress field of turbine blade thermal barrier coating is analyzed, which provides some basis for the failure prediction of real blade coating.
【学位授予单位】：湘潭大学
【学位级别】：硕士
【学位授予年份】：2017
【分类号】：V232.4;TG174.4
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