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公开(公告)号：US20160245305A1
公开(公告)日：2016-08-25
申请号：US15046881
申请日：2016-02-18
CPC分类号： F04D29/522 , F01D5/143 , F02C3/30 , F02C7/1435 , F04D19/00 , F04D29/321 , F04D29/542 , F04D29/563 , F05D2220/32 , F05D2230/50 , F05D2230/51
摘要： An object is to reduce the compressor flow rate in comparison with the reference model while maintaining a compression ratio equivalent to that in the reference model. Annulus areas required of a compressor 38 of a derivative gas turbine 200 are determined based on a compressor flow rate and a compression ratio required of the compressor 38 of the derivative gas turbine 200. Under the condition that the annulus area of each stage of the compressor 38 becomes equal to the determined annulus area, an inner radius increment and an outer radius decrement of an initial stage 36a are determined, the inner radius increment and the outer radius decrement of each of intermediate stages 36b-36e are determined so that the inner radius increment is not more than the inner radius increment of the previous stage and the outer radius decrement is not less than the outer radius decrement of the previous stage, and the inner radius increment and the outer radius decrement of a final stage 36f are determined so that the outer radius decrement is not less than the inner radius increment. The compressor 38 is designed by updating design data of components of the reference compressor 15 that deviated from the specifications due to the determination of the inner radius increment and the outer radius decrement so that the updated design data fulfill the specifications in each of the stages 36a-36f.
摘要翻译： 目的是与参考模型相比降低压缩机流量，同时保持与参考模型中的压缩比相当的压缩比。 衍生燃气轮机200的压缩机38所需的环形区域基于压缩机流量和衍生燃气轮机200的压缩机38所需的压缩比来确定。在压缩机的各级的环空区域 38变得等于所确定的环面积，初始阶段36a的内半径增量和外半径减量被确定，每个中间级36b-36e的内半径增量和外半径减量被确定为使得内半径 增量不大于前一级的内半径增量，外半径减量不小于前一级的外半径减量，并且确定最终级36f的内半径增量和外半径减量，使得 外半径减量不小于内半径增量。 压缩机38是通过更新参考压缩机15的部件的设计数据来设计的，该设计数据由于确定了内半径增量和外半径减量而偏离规格，使得更新后的设计数据满足每个级36a中的规格 -36f。
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公开(公告)号：US20180066662A1
公开(公告)日：2018-03-08
申请号：US15559164
申请日：2016-04-11
摘要： In a gas turbine manufacturing method for manufacturing a derivative gas turbine having a different cycle from a reference gas turbine including a reference compressor, a compressor of the derivative gas turbine is designed to add at least one additional stage further on an upstream side than a last stage of the reference compressor and on a downstream side of a bleed slit of a bleed chamber of the reference compressor, the compressor of the derivative gas turbine is manufactured on the basis of the design, and the derivative gas turbine is manufactured. Consequently, it is possible to manufacture a gas turbine that can secure a surge margin of a compressor with respect to fluctuation in the composition of fuel.
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公开(公告)号：US20170248020A1
公开(公告)日：2017-08-31
申请号：US15417277
申请日：2017-01-27
摘要： A turbine blade includes cooling passages formed inside the blade and extending in a blade height direction, blade surfaces on a suction side and a pressure side being covered with thermal barrier coating, a design point on a suction side being set on the blade surface on the suction side of each blade section perpendicular to the blade height direction within a range from a position on a back side of and including a throat position, and to a position in front of and not including a tailing end of a final cooling passage. Thickness distribution of the thermal barrier coating on the suction side of each blade section is configured such that a thickness of the thermal barrier coating is uniform from a blade leading edge to the design point and gradually reduces from the design point toward the back side up to the blade trailing edge.
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