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1. 发明授权

US06988872B2 Turbine moving blade and gas turbine 有权
标题翻译：涡轮动叶片和燃气轮机
公开(公告)号：US06988872B2
公开(公告)日：2006-01-24
申请号：US10351479
申请日：2003-01-27
申请人： Friedrich Soechting , Charles Ellis , Yasuoki Tomita , Shunsuke Torii
发明人： Friedrich Soechting , Charles Ellis , Yasuoki Tomita , Shunsuke Torii
IPC分类号： B63H1/14
CPC分类号： F01D5/186 , F01D5/187 , F05D2260/22141 , Y02T50/676
摘要： The object of the present invention is to provide a turbine moving blade that has high heat resistance and can be used for a long period of time by improving the cooling efficiency of the turbine moving blade, and to improve both the thermal efficiency and operating efficiency of a gas turbine through the use of this turbine moving blade. In order to achieve the object, the present invention provide a turbine moving blade arranged in a combustion gas flow path in which a plurality of blow-out openings for blowing out a cooling medium are formed in its outer surface, wherein among a plurality of cooling medium trailing edge blow-out openings arranged from the vicinity of a blade base to the vicinity of a blade tip along a blade trailing edge of turbine moving blade, the opening area of a blade tip trailing edge blow-out opening located in the vicinity of blade tip is set to be larger than the opening area of the other trailing edge blow-out openings.
摘要翻译：本发明的目的是提供一种具有高耐热性并可通过提高涡轮机动叶片的冷却效率而长时间使用的涡轮机动叶片，并提高热效率和操作效率通过使用该涡轮机动叶片的燃气轮机。为了实现该目的，本发明提供了一种设置在燃烧气体流路中的涡轮机动叶片，其中在其外表面形成有用于吹出冷却介质的多个吹出口，其中在多个冷却从叶片底部附近到涡轮机动叶片的叶片后缘的叶片尖端附近布置的中间后缘吹出开口，位于涡轮机动叶片附近的叶片尖端后缘吹出开口的开口面积叶片尖端被设定为大于其他后缘吹出口的开口面积。

2. 发明授权

US06994514B2 Turbine blade and gas turbine 有权
标题翻译：涡轮叶片和燃气轮机
公开(公告)号：US06994514B2
公开(公告)日：2006-02-07
申请号：US10299711
申请日：2002-11-20
申请人： Friedrich Soechting , Yasuoki Tomita , Shunsuke Torii
发明人： Friedrich Soechting , Yasuoki Tomita , Shunsuke Torii
IPC分类号： F01D5/18
CPC分类号： F01D5/187 , F01D5/20 , F01D11/122 , F05D2250/292 , F05D2260/201
摘要： Holes 38 and 39 have upstream opening portions 38b and 39b and downstream opening portions 38a and 39a which have a larger cross-sectional area than upstream opening portions 38b and 39b, and are formed at top portion TP of each moving blade. Holes 38 and 39 have tapered shapes T1 and T2 or step portions, and preferably, downstream opening portions 38a and 39a are eccentrically formed toward the moving direction. When tip squealer 37 is formed, hole 38 is formed so that its opening portion is provided at the side surface of tip squealer 37. Without covering the holes for cooling which are formed at the top portion of the turbine blade due to rubbing or the like, the turbine blade is accurately cooled and stably driven.
摘要翻译：孔38和39具有上游开口部分38b和39b以及具有比上游开口部分38b和39b大的横截面积的下游开口部分38a和39a，并且形成在每个移动的顶部TP 刀。孔38和39具有锥形T 1和T 2或台阶部分，并且优选地，下游开口部分38a和39a朝向移动方向偏心地形成。当形成尖端尖叫器37时，形成孔38，使得其开口部设置在尖端尖叫器37的侧表面。不会由于摩擦等而覆盖在涡轮机叶片的顶部形成的用于冷却的孔，所以涡轮叶片被精确地冷却并稳定地驱动。

3. 发明授权

US06860108B2 Gas turbine tail tube seal and gas turbine using the same 有权
标题翻译：燃气轮机尾管密封和燃气轮机使用相同
公开(公告)号：US06860108B2
公开(公告)日：2005-03-01
申请号：US10347745
申请日：2003-01-22
申请人： Friedrich Soechting , Yasuoki Tomita , Shunsuke Torii
发明人： Friedrich Soechting , Yasuoki Tomita , Shunsuke Torii
IPC分类号： F02C7/28 , F01D9/02 , F01D9/04 , F01D11/00 , F02C1/00 , F02C7/00 , F02C7/18 , F02G3/00 , F23R3/42
CPC分类号： F02C7/28 , F01D9/023 , F01D11/005
摘要： In order to provide a gas turbine tail tube seal which can prevent an outer shroud and an inner shroud in a first row stationary blade from being damaged by heat and from being worn, and to provide a gas turbine having the above-mentioned gas turbine tail tube seal, in the gas turbine tail tube seal, the width of a combustion gas flow channel in the gas turbine tail tube seal is smaller than the width of a combustion gas flow channel made between an outer surface of the inner shroud and an inner surface of the outer shroud, and the combustion gas flow channel is provided with inclined surfaces having a gradually enlarging cross section toward the first row stationary blade in at least a downstream end section.
摘要翻译：为了提供一种燃气轮机尾管密封件，其能够防止第一排固定叶片中的外护罩和内护罩受到热量的破坏和磨损，并提供具有上述燃气轮机尾部的燃气轮机管密封件，在燃气轮机尾管密封件中，燃气轮机尾管密封件中的燃烧气体流动通道的宽度小于在内护罩的外表面和内表面之间形成的燃烧气体流动通道的宽度并且燃烧气体流动通道在至少下游端部分中设置有朝向第一排固定叶片逐渐扩大的横截面的倾斜表面。

4. 发明授权

US06983659B2 Turbine blade creep life evaluating method, turbine blade creep elongation strain measuring apparatus, and turbine blade 有权
标题翻译：涡轮叶片蠕变寿命评估方法，涡轮叶片蠕变伸长应变测量装置和涡轮叶片
公开(公告)号：US06983659B2
公开(公告)日：2006-01-10
申请号：US10635494
申请日：2003-08-07
申请人： Friedrich Soechting , Charles Ellis , Yasuoki Tomita , Yugo Tokunaga
发明人： Friedrich Soechting , Charles Ellis , Yasuoki Tomita , Yugo Tokunaga
IPC分类号： G01M15/00
CPC分类号： G01B5/205 , F01D21/003 , F05D2260/80 , F05D2270/11 , G01B5/30 , G01M13/00
摘要： In order to provide a turbine blade creep life evaluating method for quantitively evaluating the life of a turbine blade, the turbine blade is determined to be within its allowable life if the creep elongation strain in the longitudinal direction of the turbine blade is less than 0.5% of an initial length, and is determined to exceed its allowable life if the creep elongation strain is 0.5% or more than the initial length. A turbine blade creep elongation strain measuring apparatus 20 comprises a first fixed end 21, a second fixed end 22, and a dial gauge 24. A dimension in the longitudinal direction is stamped on the surface of a turbine blade.
摘要翻译：为了提供用于定量评估涡轮叶片的寿命的涡轮机叶片蠕变寿命评估方法，如果涡轮叶片的纵向方向上的蠕变伸长应变小于0.5％，则涡轮叶片被确定在其允许寿命内如果蠕变伸长应变为初始长度的0.5％或更大，则确定其超过其允许寿命。涡轮叶片蠕变伸长应变测量装置20包括第一固定端21，第二固定端22和千分表24。纵向上的尺寸被压印在涡轮叶片的表面上。

5. 发明授权

US07552647B2 Turbine blade fatigue life evaluating method 有权
标题翻译：涡轮叶片疲劳寿命评估方法
公开(公告)号：US07552647B2
公开(公告)日：2009-06-30
申请号：US11246137
申请日：2005-10-11
申请人： Friedrich Soechting , Charles Ellis , Yasuoki Tomita , Yugo Tokunaga
发明人： Friedrich Soechting , Charles Ellis , Yasuoki Tomita , Yugo Tokunaga
IPC分类号： G01M15/00
CPC分类号： G01B5/205 , F01D21/003 , F05D2260/80 , F05D2270/11 , G01B5/30 , G01M13/00
摘要： In order to provide a turbine blade fatigue life evaluating method for quantitively evaluating the fatigue life of a turbine blade, the turbine blade is determined to be within its fatigue life if the creep elongation strain in the longitudinal direction of the turbine blade is less than 0.5% of an initial length, and is determined to exceed its fatigue life if the creep elongation strain is 0.5% or more than the initial length. A turbine blade creep elongation strain measuring apparatus 20 comprises a first fixed end 21, a second fixed end 22, and a dial gauge 24. A dimension in the longitudinal direction is stamped on the surface of a turbine blade.
摘要翻译：为了提供一种用于定量评估涡轮机叶片的疲劳寿命的涡轮叶片疲劳寿命评估方法，如果涡轮叶片的纵向上的蠕变伸长应变小于0.5，则涡轮叶片被确定在其疲劳寿命内如果蠕变伸长应变为初始长度的0.5％或更大，则确定其超过其疲劳寿命。涡轮叶片蠕变伸长应变测量装置20包括第一固定端21，第二固定端22和千分表24.在涡轮叶片的表面上冲压沿纵向的尺寸。

6. 发明申请

US20060201257A1 Turbine blade fatigue life evaluating method 有权
标题翻译：涡轮叶片疲劳寿命评估方法
公开(公告)号：US20060201257A1
公开(公告)日：2006-09-14
申请号：US11246137
申请日：2005-10-11
申请人： Friedrich Soechting , Charles Ellis , Yasuoki Tomita , Yugo Tokunaga
发明人： Friedrich Soechting , Charles Ellis , Yasuoki Tomita , Yugo Tokunaga
IPC分类号： F01L1/46
CPC分类号： G01B5/205 , F01D21/003 , F05D2260/80 , F05D2270/11 , G01B5/30 , G01M13/00
摘要： In order to provide a turbine blade fatigue life evaluating method for quantitively evaluating the fatigue life of a turbine blade, the turbine blade is determined to be within its fatigue life if the creep elongation strain in the longitudinal direction of the turbine blade is less than 0.5% of an initial length, and is determined to exceed its fatigue life if the creep elongation strain is 0.5% or more than the initial length. A turbine blade creep elongation strain measuring apparatus 20 comprises a first fixed end 21, a second fixed end 22, and a dial gauge 24. A dimension in the longitudinal direction is stamped on the surface of a turbine blade.
摘要翻译：为了提供一种用于定量评估涡轮机叶片的疲劳寿命的涡轮叶片疲劳寿命评估方法，如果涡轮叶片的纵向上的蠕变伸长应变小于0.5，则涡轮叶片被确定在其疲劳寿命内如果蠕变伸长应变为初始长度的0.5％或更大，则确定其超过其疲劳寿命。涡轮叶片蠕变伸长应变测量装置20包括第一固定端21，第二固定端22和千分表24。纵向上的尺寸被压印在涡轮叶片的表面上。

7. 发明授权

US06761529B2 Cooling structure of stationary blade, and gas turbine 有权
公开(公告)号：US06761529B2
公开(公告)日：2004-07-13
申请号：US10202654
申请日：2002-07-25
申请人： Friedrich Soechting , Satoshi Hada , Masamitsu Kuwabara , Junichiro Masada , Yasuoki Tomita
发明人： Friedrich Soechting , Satoshi Hada , Masamitsu Kuwabara , Junichiro Masada , Yasuoki Tomita
IPC分类号： F01D906
CPC分类号： F01D5/187 , F01D9/041 , F05D2240/81 , F05D2260/201
摘要： A collision plate having plural small holes is provided at an interval from a bottom surface of an inner shroud to form a chamber and guides cooling air from the small holes into the chamber. A leading edge flow path is provided at a leading edge side along a width direction and introduces the cooling air. A side flow path is provided along both sides of the inner shroud and guides the cooling air to a trailing edge side. A header is formed along the width direction near the trailing edge and guides the cooling air from the side flow path. Plural trailing edge flow paths are formed at the trailing edge side at intervals along a width direction, in which one end of each flow path is connected to the header and the other end is open at the trailing edge, and the cooling air in the header is ejected from the trailing edge.

8. 发明授权

US06887039B2 Stationary blade in gas turbine and gas turbine comprising the same 有权
标题翻译：包括燃气轮机和燃气轮机的固定叶片
公开(公告)号：US06887039B2
公开(公告)日：2005-05-03
申请号：US10191448
申请日：2002-07-10
申请人： Friedrich Soechting , Satoshi Hada , Masamitsu Kuwabara , Junichiro Masada , Yasuoki Tomita
发明人： Friedrich Soechting , Satoshi Hada , Masamitsu Kuwabara , Junichiro Masada , Yasuoki Tomita
IPC分类号： F01D9/02 , F01D11/02 , F02C7/28 , F01D11/00
CPC分类号： F01D11/02 , F02C7/28
摘要： A stationary blade of a gas turbine, which can reduce thermal stress produced at a portion in the vicinity of a rear edge of an inner shroud of the stationary blade. The stationary blade is positioned adjacent to at least one of moving-blade disks in an axial direction of the gas turbine. A concave portion is provided in the inner shroud in a manner such that the concave portion is formed in the vicinity of a rear edge of the inner shroud and on an inner-peripheral face of the inner shroud, where cooling air passes along the inner-peripheral face which faces a rotation shaft of the moving-blade disks; and a protruding portion which protrudes towards the rotation shaft is formed at the rear edge of the inner shroud.
摘要翻译：燃气轮机的固定叶片，其可以降低在固定叶片的内护罩的后边缘附近的部分产生的热应力。固定叶片沿着燃气轮机的轴向方向定位成与活动叶片盘中的至少一个相邻。在内护罩中设置有凹部，使得凹部形成在内护罩的后缘附近和内护罩的内周面附近，在内罩的内周面上，面向移动刀片盘的旋转轴的周边面; 并且在内护罩的后边缘处形成有朝向旋转轴突出的突出部。

9. 发明授权

US07033138B2 Ring segment of gas turbine 有权
标题翻译：燃气轮机的环段
公开(公告)号：US07033138B2
公开(公告)日：2006-04-25
申请号：US10235825
申请日：2002-09-06
申请人： Yasuoki Tomita , Osamu Isumi , Shinichi Inoue , Friedrich Soechting , Vincent Laurello , Hiroshige Matsuoka
发明人： Yasuoki Tomita , Osamu Isumi , Shinichi Inoue , Friedrich Soechting , Vincent Laurello , Hiroshige Matsuoka
IPC分类号： F01D25/08
CPC分类号： F01D25/12 , F01D11/005 , F01D11/08 , F01D25/246 , F05D2240/55 , F05D2260/201
摘要： An object of the present invention is to provide a ring segment of a gas turbine in which the temperature is maintained low, damage due to high temperature oxidation is prevented, and high temperature deformation is prevented. In order to achieve the object, the present invention provides a ring segment of a gas turbine which comprises a blade ring, a main shaft and moving blades comprising a plurality of individual units which define an annular form by being arranged around the peripheral direction of the main shaft, and disposed so that its inner peripheral surface is maintained at a constant distance from the tips of the moving blades, wherein grooves which extend along the axial direction of the main shaft of the turbine are formed upon of the individual units so as mutually to confront one another; a seal plate which is inserted into each mutually confronting pair of the grooves so as to connect together the adjacent pair of individual units; and contact surfaces which are formed at positions more radially inward than the seal plates, which extend in the axial direction and the peripheral direction and which mutually contact one another.
摘要翻译：本发明的目的是提供一种燃气轮机的环段，其中温度保持较低，防止了由于高温氧化引起的损坏，并且防止了高温变形。为了实现该目的，本发明提供了一种燃气轮机的环形段，其包括叶片环，主轴和包括多个独立单元的活动叶片，所述多个单独的单元通过围绕主轴，并且配置成使其内周面与动叶片的前端保持恒定的距离，其中沿涡轮主轴的轴向延伸的槽形成在各个单元上，以便相互相互对抗密封板，其插入每个相互面对的一对槽中，以将相邻的一对单独的单元连接在一起; 以及形成在比轴向和周向延伸并彼此相互接触的密封板更径向向内的位置处的接触表面。

10. 发明授权

US06742991B2 Turbine blade and gas turbine 有权
标题翻译：涡轮叶片和燃气轮机
公开(公告)号：US06742991B2
公开(公告)日：2004-06-01
申请号：US10192676
申请日：2002-07-11
申请人： Friedrich Soechting , Satoshi Hada , Masamitsu Kuwabara , Junichiro Masada , Yasuoki Tomita
发明人： Friedrich Soechting , Satoshi Hada , Masamitsu Kuwabara , Junichiro Masada , Yasuoki Tomita
IPC分类号： F01D518
CPC分类号： F01D5/188 , F01D5/186 , F01D5/189 , F05D2250/70 , F05D2260/202
摘要： A turbine blade applicable to a gas turbine has a turbine blade body having film cooling holes, the interior space of which is partitioned into two cavities by a rib. Hollow inserts each having impingement holes are respectively arranged in the cavities to form cooling spaces therebetween. Communication is ensured between the cavities by bypass hole and slits, so that the impingement cooling is interrupted with respect to the prescribed side having a good heat transmission in the turbine blade body. A partition wall is further arranged between the rib and the insert arranged in the trailing-edge side, thus providing a separation between the cooling spaces respectively arranged in the rear side and front side. Thus, it is possible to noticeably reduce the amount of cooling air in the turbine blade body; and it is possible to reduce temperature differences entirely over the turbine blade body as small as possible.
摘要翻译：适用于燃气轮机的涡轮叶片具有涡轮叶片体，其具有膜冷却孔，其内部空间通过肋分隔成两个空腔。各个具有冲击孔的中空插入件分别设置在空腔中，以在它们之间形成冷却空间。通过旁通孔和狭缝确保空腔之间的通信，使得冲击冷却相对于在涡轮叶片体中具有良好的热传递的规定侧被中断。分隔壁还布置在布置在后缘侧的肋和插入件之间，从而在分别布置在后侧和前侧的冷却空间之间分开。因此，可以显着地减少涡轮叶片体内的冷却空气量; 并且可以尽可能小地将涡轮叶片体上的温差完全降低。

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