有限元方法分析冷卻氣膜孔對(duì)渦輪葉片TBCs溫度場(chǎng)和應(yīng)力場(chǎng)的影響
發(fā)布時(shí)間：2018-08-25 18:31
【摘要】：熱障涂層因其較低的導(dǎo)熱率具有良好的隔熱效果,能夠有效提高渦輪葉片的工作溫度,改善發(fā)動(dòng)機(jī)的工作效率,提高飛機(jī)的推重比,對(duì)于提高國力具有舉足輕重的作用,廣泛應(yīng)用于航空發(fā)動(dòng)機(jī)渦輪葉片。因熱失配而產(chǎn)生的殘余應(yīng)力是熱障涂層不可避免的,也是導(dǎo)致涂層失效的關(guān)鍵因素。本文以考慮冷卻氣膜孔的實(shí)際熱障涂層渦輪葉片作為研究對(duì)象,采用流固耦合的方法,分別使用FLUENT軟件和ABAQUS軟件建立流體域計(jì)算模型和固體域計(jì)算模型。主要研究?jī)?nèi)容如下:(1)建立帶冷卻氣膜孔渦輪葉片熱障涂層計(jì)算模型和流體域計(jì)算模型。圍繞幾何模型的構(gòu)建和網(wǎng)格劃分提出了詳細(xì)的方法和介紹,并賦予材料參數(shù),設(shè)置分析步和加載。流體域采用有限體積法解流體動(dòng)力學(xué)方程,渦輪葉片采用有限單元法解固體熱應(yīng)力方程,通過第三方軟件MPCCI實(shí)現(xiàn)流體計(jì)算域和固體計(jì)算域的聯(lián)合模擬仿真。(2)考慮單列冷卻氣膜孔的熱障涂層渦輪葉片溫度場(chǎng)和應(yīng)力場(chǎng)的結(jié)果分析。流固耦合仿真計(jì)算得到單列冷卻氣膜孔的熱障涂層渦輪葉片的高溫穩(wěn)態(tài)溫度場(chǎng),分析發(fā)現(xiàn)熱障涂層具有較好的隔熱效果,平均隔熱達(dá)到了100K左右,壓力面和吸力面涂層的隔熱效果要優(yōu)于前緣和后緣;冷卻氣膜覆蓋的區(qū)域由于氣膜冷卻和熱障涂層的共同作用具有最好的隔熱效果,并且冷卻氣膜的冷卻效果占據(jù)了主導(dǎo)作用,涂層的隔熱效果相較于其他區(qū)域并不明顯。并基于共軛溫度場(chǎng)計(jì)算得到熱障涂層冷卻至室溫后的殘余應(yīng)力。研究結(jié)果表明在氣膜孔最右端的涂層表面是最有可能由于水平殘余應(yīng)力11?而導(dǎo)致剝落破壞的危險(xiǎn)區(qū)域;在氣膜孔最左端的涂層界面處則是由法向殘余應(yīng)力22?產(chǎn)生界面裂紋的最可能出現(xiàn)區(qū)域。(3)對(duì)不含冷卻氣膜孔的熱障涂層渦輪葉片模型與帶冷卻氣膜孔的熱障涂層渦輪葉片模型進(jìn)行比較分析。對(duì)于不含冷卻氣膜孔的葉片模型,其在溫度較高的前緣和尾緣位置涂層的隔熱效果要優(yōu)于溫度相對(duì)較低的壓力面和吸力面;熱障涂層中陶瓷層的熱應(yīng)力高于過渡層,并且在陶瓷層和過渡層中熱應(yīng)力都是在前緣和尾緣兩側(cè)的葉根部位最大,熱障涂層剝落失效易發(fā)生在這些位置。而對(duì)于含冷卻氣膜孔的葉片模型,其壓力面和吸力面涂層的隔熱效果要優(yōu)于前緣和后緣;其熱障涂層中也是陶瓷層的熱應(yīng)力高于過渡層,但在氣膜孔處出現(xiàn)應(yīng)力集中,陶瓷層和過渡層中熱應(yīng)力在氣膜孔處最大,故熱障涂層剝落失效易發(fā)生氣膜孔處�？傊�,本文采用流固耦合分析方法實(shí)現(xiàn)了實(shí)際帶冷卻氣膜孔的熱障涂層渦輪葉片高溫穩(wěn)態(tài)溫度場(chǎng)和殘余應(yīng)力的有限元模擬,分析了冷卻氣膜孔對(duì)渦輪葉片熱障涂層溫度場(chǎng)和應(yīng)力場(chǎng)的影響,為真實(shí)葉片涂層的失效預(yù)測(cè)提供了一些依據(jù)。
[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.
【學(xué)位授予單位】：湘潭大學(xué)
【學(xué)位級(jí)別】：碩士
【學(xué)位授予年份】：2017
【分類號(hào)】：V232.4;TG174.4
本文編號(hào)：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.
【學(xué)位授予單位】：湘潭大學(xué)
【學(xué)位級(jí)別】：碩士
【學(xué)位授予年份】：2017
【分類號(hào)】：V232.4;TG174.4
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