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1. 发明授权

US08992172B2 Turbo engine 有权
标题翻译：涡轮发动机
公开(公告)号：US08992172B2
公开(公告)日：2015-03-31
申请号：US13131040
申请日：2009-12-02
申请人： Martin Hoeger , Franz Malzacher , Marc Nagel
发明人： Martin Hoeger , Franz Malzacher , Marc Nagel
IPC分类号： F01D9/00 , F01D5/14 , F01D9/04
CPC分类号： F01D5/143 , F01D9/04 , F05D2240/301 , F05D2250/70 , Y02T50/673
摘要： A turbo engine, particularly a gas turbine aircraft engine, has compressor components, turbine components, and at least one combustion chamber. At least one support rib is in flow channel between two turbine components, connected one behind the other. Each support rib diverts a flow through the flow channel. A preferably cylindrical guide element runs within each support rib. Each support rib has a suction side with a greater thickness toward a radially inner flow channel wall as well as toward a radially outer flow channel wall, when viewed in the radial direction. Each support rib has a pressure side with a greater thickness toward a radially inner flow channel wall as well as toward a radially outer flow channel wall, when viewed in the radial direction. The front edge and the rear edge of each support rib are inclined in the meridian direction.
摘要翻译：涡轮发动机，特别是燃气涡轮机飞机发动机，具有压缩机部件，涡轮机部件和至少一个燃烧室。至少一个支撑肋在两个涡轮机部件之间的流动通道中，一个在另一个之后连接。每个支撑肋使流过该流动通道的流动。优选圆柱形导向元件在每个支撑肋内延伸。当从径向观察时，每个支撑肋具有朝向径向内部流动通道壁以及朝向径向外部流动通道壁的较大厚度的吸力侧。当从径向观察时，每个支撑肋具有朝向径向内部流动通道壁以及朝向径向外部流动通道壁的较大厚度的压力侧。每个支撑肋的前边缘和后边缘在子午线方向上倾斜。

2. 发明申请

US20140064951A1 ROOT BOW GEOMETRY FOR AIRFOIL SHAPED VANE 审中-公开
标题翻译：用于飞机形状的VANE的ROOT BOW几何
公开(公告)号：US20140064951A1
公开(公告)日：2014-03-06
申请号：US13603634
申请日：2012-09-05
申请人： Renee J. Jurek , Thomas J. Praisner , Martin Hoeger
发明人： Renee J. Jurek , Thomas J. Praisner , Martin Hoeger
IPC分类号： F01D9/02 , F01D9/04
CPC分类号： F01D9/041 , F01D5/141 , F01D5/145
摘要： A vane assembly used for controlling a turning gas flow includes multiple vanes, each of which is bowed toward a pressure side of the vane at the root of the vane.
摘要翻译：用于控制转向气流的叶片组件包括多个叶片，每个叶片在叶片的根部处朝向叶片的压力侧弯曲。

3. 发明申请

US20110225979A1 TURBO ENGINE 有权
标题翻译：涡轮发动机
公开(公告)号：US20110225979A1
公开(公告)日：2011-09-22
申请号：US13131040
申请日：2009-12-02
申请人： Martin Hoeger , Franz Malzacher , Marc Nagel
发明人： Martin Hoeger , Franz Malzacher , Marc Nagel
IPC分类号： F02C7/20
CPC分类号： F01D5/143 , F01D9/04 , F05D2240/301 , F05D2250/70 , Y02T50/673
摘要： The invention relates to a turbo engine, in particular a gas turbine aircraft engine, having a plurality of compressor components, at least one combustion chamber and a plurality of turbine components, wherein at least one support rib (36) is positioned in a flow channel (35) between two turbine components (32, 34) connected one behind the other, wherein the support rib (36) or each support rib (36) has a suction side, a pressure side, a front edge (37) and a rear edge (38), wherein the support rib or each support rib diverts a flow that flows through flow channel (35), and wherein a preferably cylindrical guide element runs in an inside space of the support rib or of each support rib. According to the invention, the turbo engine comprises at least the following features: a) the suction side (39) of the support rib (36) or of each support rib (36) is contoured in such a way that, viewed in the radial direction, a thickness of the respective support rib (36) is enlarged or increases in the direction onto a radially inner boundary wall (42) of flow channel (35) as well as in the direction onto a radially outer boundary wall (43) of flow channel (35); b) the pressure side (40) of support rib (36) or of each support rib (36) is contoured in such a way that, viewed in the radial direction, the thickness of the respective support rib (36) is enlarged or increases at least directly in the region of the radially inner boundary wall (42) of flow channel (35) as well as directly in the region of the radially outer boundary wall (43) of flow channel (35); c) the front edge (37) and the rear edge (38) of support rib (36) or of each support rib (36) are inclined in the meridian direction.
摘要翻译：本发明涉及一种具有多个压缩机部件，至少一个燃烧室和多个涡轮机部件的涡轮发动机，特别是燃气涡轮机飞机发动机，其中至少一个支撑肋（36）定位在流动通道（36）和所述支撑肋（36）或每个支撑肋（36）具有吸入侧，压力侧，前边缘（37）和后部（36），所述两个涡轮机部件（32,34）边缘（38），其中所述支撑肋或每个支撑肋转向流过流动通道（35）的流动，并且其中优选圆柱形引导元件在所述支撑肋或每个支撑肋的内部空间中延伸。根据本发明，涡轮发动机至少包括以下特征：a）支撑肋（36）或每个支撑肋（36）的吸力侧（39）的轮廓是这样的，即从径向方向上，相应的支撑肋（36）的厚度在沿着流动通道（35）的径向内侧边界壁（42）的方向上增大或增大，并且朝向径向外边界壁（43）的方向流通道（35）; b）支撑肋（36）或每个支撑肋（36）的压力侧（40）的轮廓是这样的，即从径向方向看，相应的支撑肋（36）的厚度被扩大或增大至少直接在流动通道（35）的径向内部边界壁（42）的区域中，以及直接位于流动通道（35）的径向外边界壁（43）的区域中; c）支撑肋（36）或每个支撑肋（36）的前边缘（37）和后边缘（38）沿子午线方向倾斜。

4. 发明授权

US6017186A Rotary turbomachine having a transonic compressor stage 失效
标题翻译：旋转涡轮机具有跨音速压缩机级
公开(公告)号：US6017186A
公开(公告)日：2000-01-25
申请号：US984606
申请日：1997-12-03
申请人： Martin Hoeger , Uwe Schmidt-Eisenlohr
发明人： Martin Hoeger , Uwe Schmidt-Eisenlohr
IPC分类号： F01D5/14 , F04D21/00 , F04D29/32 , F04D29/54 , F04D29/68 , F01D1/10 , B64C11/00 , F01D1/02
CPC分类号： F04D29/321 , F01D5/143 , F04D21/00 , F04D29/544 , F04D29/681 , Y02T50/673 , Y10S415/914
摘要： A rotary turbomachine, and particularly a turbojet engine, has a compressor stage that is especially adapted for operation in the transonic region. The compressor stage includes a rotor (1) having a hub (2) and a plurality of compressor blades (3) extending radially therefrom, and a stator having a plurality of compressor blades extending radially between a stator hub and a housing. The hub (2) of the rotor (1), and/or the hub of the stator, and/or the housing of the stator, have a circumferential surface contour that is not continuously rotationally symmetrical. Namely, a concave contour (K) is provided in the circumferential surface of the hub near the base of each blade (3) on the pressure side (PS) thereof, while the circumferential surface on the suction side (SS) of the base of each blade has a contour that is linear, slightly convex, convex/concave, or slightly concave to a lesser degree than the concave contour (K) on the pressure side of each blade. The hub contour slopes and transitions smoothly radially outwardly from the concave contour (K) on the pressure side to the flatter or convex contour on the suction side. In this manner, fluid flow velocities on both the pressure side and the suction side of each blade are reduced, especially under transonic operating conditions, and compression shock losses can be minimized.
摘要翻译：旋转涡轮机，特别是涡轮喷气发动机，具有特别适于在跨音速区域中操作的压缩机级。压缩机级包括具有轮毂（2）的转子（1）和从其径向延伸的多个压缩机叶片（3），以及具有在定子毂和壳体之间径向延伸的多个压缩机叶片的定子。转子（1）的毂（2）和/或定子的毂和/或定子的壳体具有不是连续旋转对称的圆周表面轮廓。也就是说，在轮毂的周向表面上设置有凹形轮廓（K），该圆周表面靠近其每个叶片（3）的压力侧（PS）的基部，而基部的吸力侧（SS）每个叶片具有与每个叶片的压力侧上的凹轮廓（K）更小的线性，稍凸起，凸/凹或稍微凹入的轮廓。轮毂轮廓从压力侧的凹形轮廓（K）平滑地向外平滑地向外倾斜到吸力侧的平坦或凸形轮廓。以这种方式，减小了每个叶片的压力侧和吸入侧的流体流动速度，特别是在跨音速操作条件下，可以使压缩冲击损失最小化。

5. 发明授权

US08573941B2 Tandem blade design 有权
标题翻译：串联刀片设计
公开(公告)号：US08573941B2
公开(公告)日：2013-11-05
申请号：US13145753
申请日：2010-03-15
申请人： Martin Hoeger
发明人： Martin Hoeger
IPC分类号： F01D9/04 , F01D5/14
CPC分类号： F01D9/041 , F01D5/142 , F01D5/146 , F04D21/00 , F04D29/544 , F05D2240/12 , Y02T50/671 , Y02T50/673
摘要： A tandem blade design for an axial turbomachine, comprising a front blade and a rear blade disposed with an offset thereto in the circumferential direction and in the axial direction. The rear blade is profiled and positioned with respect to the front blade such that it raises the speed level at the trailing edge of the front blade in a predetermined working range in interaction with the front blade.
摘要翻译：一种用于轴向涡轮机的串联叶片设计，包括在沿圆周方向和轴向方向上偏置的前叶片和后叶片。后叶片相对于前刀片成型并定位，使得其在与前刀片相互作用的预定工作范围内将前刀片的后缘升高速度水平。

6. 发明授权

US09382806B2 Intermediate housing of a gas turbine having an outer bounding wall having a contour that changes in the circumferential direction upstream of a supporting rib to reduce secondary flow losses 有权
标题翻译：具有外围壁的燃气涡轮机的中间壳体具有在支撑肋的上游沿圆周方向变化的轮廓，以减少二次流动损失
公开(公告)号：US09382806B2
公开(公告)日：2016-07-05
申请号：US13699202
申请日：2012-01-16
申请人： Martin Hoeger , Inga Mahle , Jochen Gier
发明人： Martin Hoeger , Inga Mahle , Jochen Gier
IPC分类号： F01D9/02 , F01D25/14 , F01D25/24 , F01D5/14
CPC分类号： F01D9/02 , F01D5/143 , F01D25/14 , F01D25/24 , F05D2240/11 , F05D2270/17
摘要： Intermediate housing (14), in particular of turbines (11, 13) of a gas turbine engine, having a radially inner bounding wall (23) and having a radially outer bounding wall (24, 24′), having a crossflow channel (33), which is formed by the bounding walls (23, 24, 24′) and within which at least one supporting rib (15) is positioned that has a leading edge (16), a trailing edge (17), as well as side walls (18) extending between the leading edge (16) and the trailing edge (17) that direct a gas flow traversing the crossflow channel (33); the radially outer bounding wall (24) having a contour that changes in the circumferential direction at least in one section upstream of the supporting rib (15).
摘要翻译：中间壳体（14），特别是燃气涡轮发动机的涡轮机（11,13），具有径向内部界限壁（23）并且具有径向外侧边界壁（24,24'），具有横流通道），其由边界壁（23,24,24'）形成，并且其中至少一个支撑肋（15）被定位成具有前缘（16），后缘（17）以及侧面在前缘（16）和后缘（17）之间延伸的引导穿过横流通道（33）的气流的壁（18）; 所述径向外边界壁（24）具有至少在所述支撑肋（15）上游的一个部分中在圆周方向上变化的轮廓。

7. 发明申请

US20110318172A1 TANDEM BLADE DESIGN 有权
标题翻译： TANDEM刀片设计
公开(公告)号：US20110318172A1
公开(公告)日：2011-12-29
申请号：US13145753
申请日：2010-03-15
申请人： Martin Hoeger
发明人： Martin Hoeger
IPC分类号： F01D5/12 , F01D1/02
CPC分类号： F01D9/041 , F01D5/142 , F01D5/146 , F04D21/00 , F04D29/544 , F05D2240/12 , Y02T50/671 , Y02T50/673
摘要： The present invention relates to a tandem blade design for an axial turbomachine, comprising a front blade and a rear blade disposed with an offset thereto in the circumferential direction and in the axial direction. The rear blade is profiled and positioned with respect to the front blade such that it raises the speed level at the trailing edge of the front blade in a predetermined working range in interaction with the front blade.
摘要翻译：本发明涉及一种用于轴向涡轮机的串联叶片设计，其包括在沿圆周方向和轴向偏置的前叶片和后叶片上。后叶片相对于前刀片成型并定位，使得其在与前刀片相互作用的预定工作范围内将前刀片的后缘升高速度水平。

8. 发明授权

US07789631B2 Compressor of a gas turbine and gas turbine 有权
标题翻译：燃气轮机和燃气轮机的压缩机
公开(公告)号：US07789631B2
公开(公告)日：2010-09-07
申请号：US10591996
申请日：2005-03-03
申请人： Martin Hoeger
发明人： Martin Hoeger
IPC分类号： F01D5/14
CPC分类号： F04D21/00 , F04D29/324
摘要： A compressor, particularly a high-pressure compressor, of a gas turbine, particularly of an aircraft engine, includes at least one rotor and a number of blades (11, 12), which are assigned to the or to each rotor and which rotate together with the respective rotor. Each blade (11, 12) is delimited, in essence, by a flow entry edge or leading edge (16), a flow exit edge or trailing edge (17), and by a blade surface (20), which extends between the leading edge (16) and the trailing edge (17) while forming a suction side (18) and a pressure side. The leading edges (16) of the blades (11, 12) are slanted at a sweep angle that changes with the height of the respective blade (11, 12) in such a manner that the leading edges (16) comprise, in a radially external area (23) of the same, at least one forward sweep angle, a backward sweep angle or zero-sweep angle following in a radially external manner, and a forward sweep angle following, in a radially external manner, the backward sweep angle or the zero-sweep angle.
摘要翻译：特别是飞机发动机的燃气轮机的压缩机，特别是高压压缩机包括至少一个转子和多个叶片（11,12），其被分配给每个转子或转动到每个转子上与相应的转子。每个叶片（11,12）本质上由流入边缘或前缘（16），流出口边缘或后缘（17）以及叶片表面（20）限定，叶片表面（20）在引导边缘（16）和后缘（17），同时形成吸力侧（18）和压力侧。叶片（11,12）的前缘（16）以以相应叶片（11,12）的高度改变的扫掠角倾斜，使得前缘（16）以径向外部区域（23），以径向外部方式跟随的至少一个正向扫掠角度，反向扫掠角度或零点扫掠角度以及沿径向外部方式的向后扫掠角度的后向扫掠角度零扫角。

9. 发明申请

US20060051200A1 Flow structure for a gas turbine 有权
公开(公告)号：US20060051200A1
公开(公告)日：2006-03-09
申请号：US11220745
申请日：2005-09-06
申请人： Martin Hoeger
发明人： Martin Hoeger
IPC分类号： F04D29/42
CPC分类号： F02C7/20 , F01D5/145 , F01D9/023 , F04D29/545 , F05D2250/70 , Y02T50/673 , Y02T50/675
摘要： A flow structure is for a gas turbine, e.g., for an aircraft engine, in a transition channel between two compressors or in a transition channel between two turbines or in a transition channel of a turbine outlet housing downstream from a low-pressure turbine, having support ribs positioned in the transition channel spaced apart from one another around the circumference of the transition channel. A channel wall which delimits the transition channel radially on the inside and/or a channel wall which delimits the transition channel radially on the outside may be drawn inwardly into the transition channel in the area of the outflow edges of the support ribs.

10. 发明授权

US07553129B2 Flow structure for a gas turbine 有权
标题翻译：燃气轮机的流动结构
公开(公告)号：US07553129B2
公开(公告)日：2009-06-30
申请号：US11190447
申请日：2005-07-27
申请人： Martin Hoeger , Franz Malzacher
发明人： Martin Hoeger , Franz Malzacher
IPC分类号： F01D9/02
CPC分类号： F01D9/041 , F01D5/143 , F01D9/02 , F04D29/547 , Y02T50/673
摘要： A flow structure of a gas turbine, in particular for an aircraft engine, in a transitional channel between two compressors or in a transitional channel between two turbines or in a transitional channel of a turbine outlet housing downstream from a low-pressure turbine is disclosed. Supporting ribs are positioned in the transitional channel and spaced a distance apart in the circumferential direction of the transitional channel. At least one guide vane and/or guide rib is positioned between two supporting ribs spaced a distance apart from one another. The flow outlet edge of the guide rib or each guide rib runs upstream from the flow outlet edges of the supporting ribs.
摘要翻译：公开了一种燃气涡轮机的流动结构，特别是用于飞行器发动机的两个压缩机之间的过渡通道，或在两个涡轮机之间的过渡通道中或在低压涡轮机下游的涡轮出口壳体的过渡通道中。支撑肋定位在过渡通道中，并且在过渡通道的圆周方向上间隔一段距离。至少一个引导叶片和/或引导肋定位在彼此间隔开一段距离的两个支撑肋之间。引导肋或每个引导肋的流出口边缘从支撑肋的流出口边缘的上游延伸。

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