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公开(公告)号：EP1831519A1
公开(公告)日：2007-09-12
申请号：EP04822081.8
申请日：2004-12-01
CPC分类号： F02C3/14 , F01D5/022 , F02C3/067 , F02C3/073 , F02K3/068 , F02K3/072 , F05D2260/40311 , Y02T50/673
摘要： A tip turbine engine (10) provides first and second turbines (32) rotatably driven by a combustor (30) generating a high-energy gas stream. The first turbine (32) is mounted at an outer periphery of a first fan (24a) , such that the first fan is rotatably driven by the first turbine (32a) . The second turbine (32b) is mounted at an outer periphery of a second fan (24b) , and is rotatably driven by the high-energy gas stream. In one embodiment, the first turbine (32a) rotatably drives a plurality of stages of first compressor blades (54) in an axial compressor (22) in a first rotational direction, while the second turbine (32b) rotatably drives a plurality of stages of second compressor blades (52) in the axial compressor (22) in a second rotational direction opposite the first. By rotatably driving alternating stages of compressor blades in opposite directions, the efficiency of the axial compressor (22) is increased and/or the number of stages of compressor blades can be reduced. Other variations are described in additional embodiments.
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12.发明公开FAN BLADE WITH INTEGRAL DIFFUSER SECTION AND TIP TURBINE BLADE SECTION FOR A TIP TURBINE ENGINE 审中-公开随着对顶尖涡轮机发动机积分扰流板部分和顶部叶片部分风叶
公开(公告)号：EP1819907A2
公开(公告)日：2007-08-22
申请号：EP04822063.6
申请日：2004-12-01
IPC分类号： F01D5/30
CPC分类号： F01D5/3007 , F01D1/023 , F01D1/06 , F01D5/022 , F01D5/147 , F01D5/30 , Y02T50/672
摘要： A fan-turbine rotor assembly for a tip turbine engine includes an outer periphery scalloped by a multitude of elongated openings which define an inducer receipt section to receive an inducer section and a hollow fan blade section. Each fan blade section includes a turbine section which extends from a diffuser section.
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公开(公告)号：EP1147291B1
公开(公告)日：2007-08-22
申请号：EP99937125.5
申请日：1999-02-26
CPC分类号： F02K3/075 , F01D5/022 , F01D5/145 , F04D27/023 , F04D29/542 , F04D29/682 , Y02T50/673 , Y10S415/914
摘要： A compression system endwall bleed system having a plurality of bleed slots (41) formed in a mechanical housing (38) downstream of a rotating shrouded rotor (35). In one embodiment, the plurality of bleed slots (41) bleed off a separated tip boundary layer to relieve back pressure associated with this blockage. A sealing structure (47, 48) downstream of the shrouded rotor (35) is utilized to minimize working fluid leakage ahead of the plurality of bleed slots (41).
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公开(公告)号：EP1302640A3
公开(公告)日：2006-08-09
申请号：EP02257185.5
申请日：2002-10-16
CPC分类号： B64D41/00 , B64D27/02 , B64D2033/0213 , F01D5/022 , F02C3/073 , F02C3/13 , F02C7/32 , F02K3/075 , F05D2220/50
摘要： The present invention relates to a variable cycle boost propulsor system 10 for use on an aircraft. The variable cycle boost propulsor system 10 includes an engine 14, a turbine 30, a fan 34 connected to the turbine 30, and a valve system 46 for delivering the fluid output from the engine 14 to the turbine 30 for driving the turbine 30 and the fan 34 and thereby generating additional thrust for the aircraft. The engine is also used to provide power to one or more systems 27 onboard the aircraft.
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15.发明公开
公开(公告)号：EP1365154A3
公开(公告)日：2004-07-07
申请号：EP03253260.8
申请日：2003-05-23
发明人： Prentice, Ian Francis , Busbey, Bruce Clark , Bond, Donald Ray , Crall, David William , Glynn, Christopher Charles
CPC分类号： F01D5/022 , F02C3/067 , F02K3/072 , F04D19/024 , F04D25/04
摘要： A counterrotatable booster compressor assembly (36) for a gas turbine engine having a counterrotatable fan section with a first fan blade row (18) connected to a first drive shaft and a second fan blade row (22) axially spaced from the first fan blade row (18) and connected to a second drive shaft (48). The counterrotatable booster compressor assembly (36) includes a first compressor blade row connected to the first drive shaft, a plurality of fan shaft extensions (100) connected to the second drive shaft (48) for driving the second fan blade row (22), and at least one compressor blade (124) integral with each fan shaft extension (100) so as to form a second compressor blade row interdigitated with the first compressor blade row. The counterrotatable booster compressor (36) further includes a first platform member (116) integral with each fan shaft extension (100) at a first location (118) so as to form a portion of an inner flowpath for the counterrotatable booster compressor (36) and a second platform member (120) integral with each fan shaft extension (100) at a second location (122) so as to form a portion of an outer flowpath for the counterrotatable booster compressor (36), where each compressor blade (124) of the second compressor blade row is positioned between the first and second platform members (116,120).
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16.发明公开Counter-rotatable booster compressor assembly for a gas turbine engine 有权GegenläufigeNiederdruckverdichteranlagefüreine Gasturbine
公开(公告)号：EP1367250A1
公开(公告)日：2003-12-03
申请号：EP03251104.0
申请日：2003-02-25
摘要： A counter-rotatable booster compressor assembly (36) for a gas turbine engine (10) has a counter-rotatable fan section (12) with a first fan blade row (18) connected to a first drive shaft (46) and a second fan blade row (22) axially spaced from the first fan blade row (18) and connected to a second drive shaft (48). The counter-rotatable booster compressor assembly (36) includes a first compressor blade row (38) connected to the first drive shaft (46) and a second compressor blade row (40) interdigitated with the first compressor blade row (38) and connected to the second drive shaft (48). A portion of each fan blade (22) of the second fan blade row (22) extends through a flowpath (80) of the counter-rotatable booster compressor (36) so as to function as a compressor blade (104) in the second compressor blade row (40). The counter-rotatable booster compressor (36) further includes a first platform member (94) integral with each fan blade (22) of the second fan blade row (22) at a first location so as to form an inner flowpath for the counter-rotatable booster compressor (36) and a second platform member (96) integral with each fan blade (22) of the second fan blade row (22) at a second location so as to form an outer flowpath for the counter-rotatable booster compressor (36).
摘要翻译： 一种用于燃气涡轮发动机（10）的可反向旋转的增压压缩机组件（36）具有可逆转的风扇部分（12），第一风扇叶片排（18）连接到第一驱动轴（46）和第二风扇 叶片排（22）与第一风扇叶片排（18）轴向间隔开并连接到第二驱动轴（48）。 反向可旋转的增压压缩机组件（36）包括连接到第一驱动轴（46）的第一压缩机叶片排（38）和与第一压缩机叶片排（38）相互指向的第二压缩机叶片排（40），并连接到 第二驱动轴（48）。 第二风扇叶片排（22）的每个风扇叶片（22）的一部分延伸穿过可反向旋转的增压压缩机（36）的流路（80），以便起第二压缩机中的压缩机叶片（104）的作用 刀片排（40）。 可反向旋转的增压压缩机（36）还包括在第一位置与第二风扇叶片排（22）的每个风扇叶片（22）成一体的第一平台构件（94），以形成用于反向旋转的增压压缩机 可旋转增压压缩机（36）和与第二风扇叶片排（22）的每个风扇叶片（22）成一体的第二平台构件（96），以形成用于可逆转的增压压缩机的外部流路 36）。
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公开(公告)号：EP0021459A1
公开(公告)日：1981-01-07
申请号：EP80200318.6
申请日：1980-04-09
CPC分类号： F04D25/045 , F01D5/022 , F02C3/073
摘要： A turbinedriven turbocompresser is a multi-stage machine composed of a number of uncoupled concentrical work-wheels which within a common housing (1) between two stationary guiding-blade wheels (2 and 3) are alternately counterrotating and which are equipped with blade rings that in the row of wheels form corresponding flow passages for respectively both expansion (8-9) and compression (10-11). Flow directions for compression and expansion are opposite, stationary guiding-blade wheels as well as rotating work-wheels are constructed in the form of disks. The rotating wheels (4) and (6) are identical. Because of the low angular speeds of the rotating wheeldisks grease-filled spiral-groove bearings as well as gaslubricated bearings can be applied.
摘要翻译： 涡轮混合涡轮压缩机是由多个非耦合的同心圆工作轮构成的多级机，它们在两个固定的引导叶轮（2和3）之间的公共壳体（1）内交替反向旋转，并且配备有叶片环， 在行轮中形成相应的流动通道，用于分别膨胀（8-9）和压缩（10-11）。 压缩和膨胀的流动方向相反，固定的引导叶片轮以及旋转的工作轮以盘的形式构成。 旋转轮（4）和（6）相同。 由于旋转砂轮的角速度较低，因此可以应用润滑脂填充的螺旋槽轴承以及气润轴承。
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18.发明公开
公开(公告)号：EP3406847A1
公开(公告)日：2018-11-28
申请号：EP17173117.7
申请日：2017-05-26
发明人： KEEGAN, Colm
CPC分类号： F01D5/066 , F01D5/022 , F05D2250/73 , F05D2260/941
摘要： A gas turbine engine rotor disc is presented. The rotor disc includes a hub, a web, a blade retention arrangement, a rotational axis, a first axial side and a second axial side. The hub has a central bore around the rotational axis. The web is integrally formed with and extends radially outwards from the hub to the blade retention arrangement. The blade retention arrangement has a centre of mass. A radial plane perpendicular to the rotational axis passes through the centre of mass. The first axial side is for engaging a tension bolt. The radial plane intersects the hub defining a first axial side portion towards the first axial side and a second axial side portion towards the second axial side. The second axial side portion has an axial extent which is between 10% and 30% greater than an axial extent of the first axial side portion. Corresponding gas turbine rotor disc assembly and gas turbine engine are also provided.
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公开(公告)号：EP3170971B1
公开(公告)日：2018-07-18
申请号：EP16199395.1
申请日：2016-11-17
发明人： MARIANO, Thomas A. , ARNOLD, Jason
CPC分类号： F01D25/246 , F01D5/022 , F01D5/026 , F01D5/06 , F01D5/066 , F05D2230/60 , F05D2260/36 , Y02T50/671
摘要： Disclosed is a coupling system comprising joints (102) configured to friction lock under compressive pressure via tapered teeth (104) for joining rotating components, which may be rotating components of a gas turbine, or other engine. The coupling system provides the alignment orientation for the joined components, and eliminates the need for extraneous fasteners, bolts, interference fits and/or keying arrangements. The coupling system further enables material optimization by allowing for use of different materials throughout different engine sections, depending on the operating parameters.
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公开(公告)号：EP2492484A3
公开(公告)日：2018-05-09
申请号：EP12152601.6
申请日：2012-01-26
申请人： Rolls-Royce plc
CPC分类号： F01D5/022 , B64C11/48 , B64D29/00 , B64D2027/005 , F01D5/143 , F01D5/145 , F02C3/067 , F02K3/072 , F05D2220/324 , F05D2220/325 , F05D2240/14 , F05D2250/70 , Y02T50/66 , Y02T50/671 , Y02T50/673
摘要： The present disclosure relates to a propfan engine (100) comprising: one or more rotor stages (110, 120) comprising a plurality of rotors; and an outer wall (130) comprising an outer profile (132), at least a portion of the outer profile defining a substantially circular cross-section, wherein the diameter of the substantially circular cross-section increases in the direction of flow over the outer wall and downstream of a leading edge (112, 122) of the rotors, and the diameter increases at substantially all points defining the circumference of the substantially circular cross-section. The present disclosure further relates to propfan engine (100) comprising: one or more rotor stages (110, 120) comprising a plurality of rotors; and an outer wall (130) comprising an outer profile (132), the outer profile defining a cross-section, wherein the cross-section of the outer profile comprises a maximum diameter (134, 136) at a point upstream of a leading edge (112, 122) of the rotors, the diameter reducing between the maximum diameter and the leading edge of the rotors, and wherein the outer profile comprises a point of inflection (135, 137) between the maximum diameter and the leading edge of the rotors.
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