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1. 再颁专利

USRE45689E1 Swept turbomachinery blade 有权
公开(公告)号：USRE45689E1
公开(公告)日：2015-09-29
申请号：US12785222
申请日：2010-05-21
申请人： David A. Spear , Dennis N. Kantor , Bruce P. Biederman , John A. Orosa
发明人： David A. Spear , Bruce P. Biederman , John A. Orosa
IPC分类号： F01D5/14 , F01D5/16 , F04D21/00 , F04D29/32 , F04D29/38
CPC分类号： F01D5/16 , F01D5/141 , F04D21/00 , F04D29/324 , F04D29/384 , F04D29/386 , F05D2220/327 , F05D2240/302 , F05D2250/70 , F05D2250/711 , F05D2250/712 , F05D2250/713
摘要： A swept turbomachinery blade for use in a cascade of such blades is disclosed. The blade (12) has an airfoil (22) uniquely swept so that an endwall shock (64) of limited radial extent and a passage shock (66) are coincident and a working medium (48) flowing through interblade passages (50) is subjected to a single coincident shock rather than the individual shocks. In one embodiment of the invention the forwardmost extremity of the airfoil defines an inner transition point (40) located at an inner transition radius rt-inner. The sweep angle of the airfoil is nondecreasing with increasing radius from the inner transition radius to an outer transition radius rt-outer, radially inward of the airfoil tip (26), and is nonincreasing with increasing radius between the outer transition radius and the airfoil tip.

2. 再颁专利

USRE43710E1 Swept turbomachinery blade 有权
标题翻译：扫涡轮机叶片
公开(公告)号：USRE43710E1
公开(公告)日：2012-10-02
申请号：US09874931
申请日：2001-06-05
申请人： David A. Spear , Dennis N. Kantor, legal representative , Bruce P. Biederman , John A. Orosa
发明人： David A. Spear , Bruce P. Biederman , John A. Orosa
IPC分类号： F01D5/14
CPC分类号： F01D5/16 , F01D5/141 , F04D21/00 , F04D29/324 , F04D29/384 , F04D29/386 , F05D2220/327 , F05D2240/302 , F05D2250/70 , F05D2250/711 , F05D2250/712 , F05D2250/713
摘要： A swept turbomachinery blade for use in a cascade of such blades is disclosed. The blade (12) has an airfoil (22) uniquely swept so that an endwall shock (64) of limited radial extent and a passage shock (66) are coincident and a working medium (48) flowing through interblade passages (50) is subjected to a single coincident shock rather than the individual shocks. In one embodiment of the invention the forwardmost extremity of the airfoil defines an inner transition point (40) located at an inner transition radius rt-inner. The sweep angle of the airfoil is nondecreasing with increasing radius from the inner transition radius to an outer transition radius rt-outer, radially inward of the airfoil tip (26), and is nonincreasing with increasing radius between the outer transition radius and the airfoil tip.
摘要翻译：公开了一种用于这种叶片级联的扫掠式涡轮机叶片。叶片（12）具有独特地扫过的翼型件（22），使得有限径向范围的端壁冲击（64）和通道冲击（66）重合，并且流经叶片间通道（50）的工作介质（48）以单一的同时震惊而不是个别的冲击。在本发明的一个实施例中，翼型件的最前端限定位于内部过渡半径rt-内部的内部转变点（40）。翼型的扫掠角度随着从内部过渡半径到外部过渡半径rt-半径的增加而减小，这在外部过渡半径（外部过渡半径）外部，在翼型端部（26）的径向内侧，并且随着外部过渡半径和翼型尖端之间的半径的增加而增加。

3. 发明授权

US5642985A Swept turbomachinery blade 失效
标题翻译：扫涡轮机叶片
公开(公告)号：US5642985A
公开(公告)日：1997-07-01
申请号：US559965
申请日：1995-11-17
申请人： David A. Spear , Bruce P. Biederman , John A. Orosa
发明人： David A. Spear , Bruce P. Biederman , John A. Orosa
IPC分类号： F01D5/14 , F01D5/16 , F02K3/06 , F04D21/00 , F04D29/32 , F04D29/38
CPC分类号： F01D5/16 , F01D5/141 , F04D21/00 , F04D29/324 , F04D29/384 , F04D29/386 , F05D2220/327 , F05D2240/302 , F05D2250/70 , F05D2250/711 , F05D2250/712 , F05D2250/713
摘要： A swept turbomachinery blade for use in a cascade of such blades is disclosed. The blade (12) has an airfoil (22) uniquely swept so that an endwall shock (64) of limited radial extent and a passage shock (66) are coincident and a working medium (48) flowing through interblade passages (50) is subjected to a single coincident shock rather than the individual shocks. In one embodiment of the invention the forwardmost extremity of the airfoil defines an inner transition point (40) located at an inner transition radius r.sub.t -inner. The sweep angle of the airfoil is nondecreasing with increasing radius from the inner transition radius to an outer transition radius r.sub.t-outer, radially inward of the airfoil tip (26), and is nonincreasing with increasing radius between the outer transition radius and the airfoil tip.
摘要翻译：公开了一种用于这种叶片级联的扫掠式涡轮机叶片。叶片（12）具有独特地扫过的翼型件（22），使得有限径向范围的端壁冲击（64）和通道冲击（66）重合，并且流经叶片间通道（50）的工作介质（48）以单一的同时震惊而不是个别的冲击。在本发明的一个实施例中，翼型件的最前端限定位于内部过渡半径rt-内部的内部转变点（40）。翼型的扫掠角度随着从内部过渡半径到外部过渡半径rt-半径的增加而减小，这在外部过渡半径（外部过渡半径）外部，在翼型端部（26）的径向内侧，并且随着外部过渡半径和翼型尖端之间的半径的增加而增加。

4. 再颁专利

USRE38040E1 Swept turbomachinery blade 有权
标题翻译：扫涡轮机叶片
公开(公告)号：USRE38040E1
公开(公告)日：2003-03-18
申请号：US09343736
申请日：1999-06-30
申请人： David A. Spear , Bruce P. Biederman , John A. Orosa
发明人： David A. Spear , Bruce P. Biederman , John A. Orosa
IPC分类号： F01D514
CPC分类号： F01D5/16 , F01D5/141 , F04D21/00 , F04D29/324 , F04D29/384 , F04D29/386 , F05D2220/327 , F05D2240/302 , F05D2250/70 , F05D2250/711 , F05D2250/712 , F05D2250/713
摘要： A swept turbomachinery blade for use in a cascade of such blades is disclosed. The blade (12) has an airfoil (22) uniquely swept so that an endwall shock (64) of limited radial extent and a passage shock (66) are coincident and a working medium (48) flowing through interblade passages (50) is subjected to a single coincident shock rather than the individual shocks. In one embodiment of the invention the forwardmost extremity of the airfoil defines an inner transition point (40) located at an inner transition radius rt-inner. The sweep angle of the airfoil is nondecreasing with increasing radius from the inner transition radius to an outer transition radius rt-outer, radially inward of the airfoil tip (26), and is nonincreasing with increasing radius between the outer transition radius and the airfoil tip.
摘要翻译：公开了一种用于这种叶片级联的扫掠式涡轮机叶片。叶片（12）具有独特地扫过的翼型件（22），使得有限径向范围的端壁冲击（64）和通道冲击（66）重合，并且流经叶片间通道（50）的工作介质（48）以单一的同时震惊而不是个别的冲击。在本发明的一个实施例中，翼型件的最前端限定位于内部过渡半径rt-内部的内部转变点（40）。翼型的扫掠角度随着从内部过渡半径到外部过渡半径rt-半径的增加而减小，这在外部过渡半径（外部过渡半径）外部，在翼型端部（26）的径向内侧，并且随着外部过渡半径和翼型尖端之间的半径的增加而增加。

5. 发明授权

US5397215A Flow directing assembly for the compression section of a rotary machine 失效
标题翻译：用于旋转机器的压缩部分的流动引导组件
公开(公告)号：US5397215A
公开(公告)日：1995-03-14
申请号：US076325
申请日：1993-06-14
申请人： David A. Spear , Bruce P. Biederman
发明人： David A. Spear , Bruce P. Biederman
IPC分类号： F01D5/02 , F01D5/14 , F01D5/20 , F01D9/02 , F04D29/32 , F04D29/52 , F04D29/68 , F01D5/22
CPC分类号： F01D5/143 , F01D5/145 , F01D9/02 , F04D29/321 , F04D29/526 , F04D29/681 , Y02T50/673 , Y10S415/914
摘要： A flow directing assembly 24, 24a for a rotary machine 10 is disclosed. Various construction details are developed to decrease aerodynamic losses and promote the ease of design. In one particular embodiment, a concave region 58 of an end wall is employed to create convex wall regions 54, 56 upstream and downstream of the concave region.
摘要翻译：公开了一种用于旋转机器10的流动引导组件24,24a。开发了各种结构细节，以减少空气动力学损失并提高设计的便利性。在一个特定实施例中，端壁的凹入区域58用于在凹入区域的上游和下游产生凸壁区域54,56。

6. 发明授权

US5460488A Shrouded fan blade for a turbine engine 失效
标题翻译：用于涡轮发动机的有风扇叶片
公开(公告)号：US5460488A
公开(公告)日：1995-10-24
申请号：US259705
申请日：1994-06-14
申请人： David A. Spear , Bruce P. Biederman
发明人： David A. Spear , Bruce P. Biederman
IPC分类号： F01D5/22 , F04D21/00 , F04D29/68
CPC分类号： F01D5/22 , F04D21/00 , F04D29/681 , Y02T50/673
摘要： A compressor or fan blade (112) for a gas turbine engine includes a suction surface midspan shroud (122) which is cambered so that the shroud trailing edge (132) is substantially aligned with the direction of the air flow in the vicinity of the shroud trailing edge (132). The cambered portion originates at the chordwise location where a shock wave (60) attached to the leading edge of the neighboring blade crosses the shroud (136). The cambered portion gradually transitions to an uncambered profile at a circumferential location intermediate the suction surface (110) of the blade (112) and the pressure surface shroud of the neighboring blade.
摘要翻译：用于燃气涡轮发动机的压缩机或风扇叶片（112）包括吸入表面中跨护罩（122），其被弯曲成使得护罩后缘（132）基本上与护罩附近的空气流的方向对齐后缘（132）。弧形部分起始于附接到相邻叶片的前缘的冲击波（60）穿过护罩（136）的弦向位置。弧形部分在叶片（112）的吸力表面（110）和相邻叶片的压力表面护罩之间的圆周位置处逐渐转变为无弯曲轮廓。

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