Source: https://patents.google.com/patent/DE60316513T2/en
Timestamp: 2020-02-17 12:52:02
Document Index: 480589745

Matched Legal Cases: ['art 7', 'art 22', 'art 35', 'art 37', 'art 35', 'art 37']

DE60316513T2 - Aircraft gas turbine engine with Kontrollitschaufel for opposing low-pressure turbine - Google Patents
Aircraft gas turbine engine with Kontrollitschaufel for opposing low-pressure turbine
DE60316513T2
DE60316513T2 DE2003616513 DE60316513T DE60316513T2 DE 60316513 T2 DE60316513 T2 DE 60316513T2 DE 2003616513 DE2003616513 DE 2003616513 DE 60316513 T DE60316513 T DE 60316513T DE 60316513 T2 DE60316513 T2 DE 60316513T2
DE2003616513
DE60316513D1 (en
2002-07-30 Priority to US10/208,303 priority Critical patent/US6684626B1/en
2002-07-30 Priority to US208303 priority
2007-11-08 Publication of DE60316513D1 publication Critical patent/DE60316513D1/en
2008-06-26 Publication of DE60316513T2 publication Critical patent/DE60316513T2/en
The The invention relates to counter-rotating aircraft gas turbine engines with opposing ones blowers, those of opposite Low-pressure turbine rotors are driven, and in particular on paddled machines to uneven energy distributions among the opposing ones To effect low-pressure turbine rotors.
One A gas turbine engine of the twin-jet jet engine type usually includes a front one blowers and a front booster compressor, a middle core engine and a rear low pressure power turbine. The core engine includes a high pressure compressor, a combustion chamber and a high pressure turbine, the fluidically serially are connected. The high pressure compressor and the high pressure turbine of the core are interconnected by a high pressure shaft. High-pressure compressor, turbine and shaft essentially form the High pressure rotor. The high pressure compressor is rotatably driven, to enter the core engine To compress air to a relatively high pressure. This high pressure air is then mixed with the fuel in the combustion chamber and ignited to to form a high-energy gas stream. The gas stream flows after behind and through the high pressure turbine, taking her and the high pressure shaft rotationally drives, and wherein the high-pressure shaft turn the compressor rotating drives.
The effluent gas stream from the high pressure turbine is expanded by a second or low pressure turbine. The low-pressure turbine drives the blower and the booster compressor in rotation by means of a low-pressure shaft, with all the parts just mentioned forming the low-pressure rotor. The low pressure shaft extends through the high pressure rotor. Some low-pressure turbines have been designed with counter-rotating turbines that drive counter-rotating fans and booster or low-pressure compressors. The U.S. Patents No. 4,860,537 . 5,307,622 and 4,790,133 disclose counter-rotating turbines driving counter-rotating fans and booster or low-pressure compressors. Most of the generated thrust is generated by the blower.
progressive Gas turbine engines have opposing front ones and backwinds, and there are opposing ones Booster developed. It is desirable an opposite one Designing a machine that provides a top performance. It was found that peak performance can be achieved when the front fan with a higher one Blower pressure ratio and a higher speed works as the rear fan. This can in a material non-compliance in terms of performance and Speed of the counter-rotating rotors result. The opposite Low pressure turbine is needed with the front and rear blaster the speed of the respective fan with to supply the required energy. A conventional counter-rotating turbine will then work with peak efficiency when the energy is below equally distributed between two waves and when the speeds are the same and are opposite. In such a case, the speed and performance ratios of the two rotors and turbines is substantially equal to 1. It is highly desirable to have a gas turbine engine to dispose of that with opposing Low-pressure turbines are provided, the different speed and performance ratios such as speed and power ratios of 1.2 or more, to maximize the efficiency of the fan to reach.
Of the according to the present invention comprises Aircraft gas turbine a high-pressure rotor with a high-pressure turbine, which is drivingly connected by a high pressure shaft with a high pressure compressor and is rotatable about an engine centerline. A low pressure turbine with a low pressure flow path is located behind the high-pressure rotor. The low pressure turbine includes opposing low pressure internal shaft rotors and outer shaft rotors with low pressure inner shafts and outer shafts that at least partially rotatable and coaxial with and radially inward of the high pressure rotor are arranged. The low pressure internal shaft rotor has first Low-pressure turbine blade rows extending through the low-pressure turbine flowpath extend and drivingly connected to a first fan blade row at the low pressure inner shaft are. The low-pressure outer-shaft rotor has second Low-pressure turbine blade rows extending through the low-pressure turbine flowpath extend and drivingly connected to a second fan blade row at the low pressure outer shaft are. The first and second fan blade rows are arranged in a bypass channel, the radially outward from limited to a fan case becomes. The first low pressure turbine blade rows include at least a first interlocking turbine blade row between at least a second adjacent pair of the second low pressure turbine blade rows is arranged. The second low-pressure turbine blade rows included at least one second intermeshing turbine blade row, between at least a first adjacent pair of the first Low-pressure turbine blade rows is arranged. The low pressure turbine includes a plurality of rows of non-rotatable low pressure blades. Each of the rows of non-rotatable low pressure blades extends through the low-pressure turbine flow path, not between each one intermeshing adjacent pair of first and second low pressure turbine blade rows without an interlocking row of turbine blades interposed therebetween.
In the exemplary embodiment At least one booster is drivingly connected to the invention the low pressure inner shafts and outer shafts and axially between the first fan blade row and the high-pressure rotor. A low-pressure turbine nozzle is axially forward, upstream arranged to and next to the first low-pressure turbine blade rows.
He can different versions of the low-pressure turbine are used. A forefront of the second low-pressure turbine blade rows can mesh with a farthest pair of first low-pressure turbine blade rows. The Low-pressure turbine may have an odd number of either the first Low pressure turbine blade rows or second low pressure turbine blade rows and an even number of another of the first low pressure turbine blade rows or the second low-pressure turbine blade rows have. The low pressure turbine may be an odd number of the first low-pressure turbine blade rows and an even number of the second low-pressure turbine blade rows exhibit. The low-pressure turbine can be three of the first and four the second low-pressure turbine blade rows have. The two foremost rows of second low-pressure turbine blade rows can mesh with three farthest back rows of first low-pressure turbine blade rows.
The above aspects and other characteristics of the invention in the following description in conjunction with the following Drawings explains:
1 Figure 4 is a longitudinal section of a front portion of an exemplary embodiment of an aircraft gas turbine engine having a counter-rotating low pressure turbine with vanes.
2 is a longitudinal section of a rear part of the engine.
3 is an enlarged view of the in 1 illustrated opposing low-pressure turbine.
4 is a longitudinal section of a rear portion of the engine with an alternative low pressure reciprocating turbine for the in 1 shown engine.
5 is an enlarged view of the counter-rotating low-pressure turbine in the rear part of the in 4 shown engine.
In 1 becomes the front part 7 an exemplary aircraft gas turbine 10 represented, which is about a central engine axis 8th is circumscribed, and a fan section 12 comprising an inlet air flow of ambient air 14 receives. The drive 10 has a housing structure 32 with a front or fan case 34 passing through the engine case 45 with a turbine center housing 60 and a turbine rear housing 155 connected is. The engine is mounted on an aircraft or mounted externally on an aircraft, for example by means of a pylon (not shown) extending down from an aircraft wing.
The fan section 12 is with the opposite wind players 4 and 6 provided with the first and second fan blade rows 13 and 15 and, in the exemplary embodiment, also includes a booster 16 on. The booster 16 is axially behind the opposing first and second fan blade rows 13 and 15 arranged and surrounded by a splinter coat 17 with a leading edge splitter 9 , The boosters are usually arranged axially between a first fan blade row and a core engine and may be disposed between opposing first and second fan blade rows. An annular, radially inwardly located channel wall 29 borders the booster 16 radially inward. On the fan section 12 followed by a high pressure compressor (HPC) 18 who also in 2 is shown. 2 is a schematic representation of a rear part 22 of the engine 10 ,
Downstream of the HPC 18 there is a combustion chamber 20 fuel with that of the HPC 18 pressurized air 14 mixed to produce combustion gases downstream of a high pressure turbine (HPT) 24 and an opposing low-pressure turbine (LPT) 26 flow, after which the combustion gases from the engine 10 be left out. A high pressure shaft 27 connects the HPT 24 with the HPC 18 to essentially a first or high pressure rotor 33 to build. The high pressure compressor 18 , the combustion chamber 20 and the high-pressure turbine 24 become together as the core engine 25 indicates that, for the purpose of this patent, the high pressure shaft 27 contains. The core engine 25 may be configured modularly such that it can be replaced as a single unit independently of the other parts of the gas turbine.
Relegated to 1 : A bypass channel 21 becomes radially outward from a fan case 11 and partly from the splinter cover 17 limited. The first and second fan blade rows 13 and 15 are in a bypass channel 21 arranged radially outward from a fan case 11 is limited. The splinter coat 17 and the leading edge splitter 9 divide that from the second fan shovel row 15 outflowing fan airflow 23 into one in the booster 16 directed first fan airflow section and one around the booster 16 directed around the second fan air flow part, wherein the second part in the bypass channel 21 from where he leads the fan section 12 through a horn exit 30 leaves and provides thrust for the engine. The one from the booster 16 pressurized booster air 31 Leaves the booster and gets through an inlet channel splitter 39 in the first booster air part 35 and the second booster air part 37 divided up. The inlet duct splitter 39 directs the first booster air part 35 into a core engine intake 19 leading to the high pressure compressor 18 of the core engine 25 leads. The inlet duct splitter 39 directs the second booster air part 37 around the core engine 25 around in the bypass channel 21 from where he then the horn section 12 through the horn exit 30 leaves.
Related to the 2 and 3 : The low-pressure turbine 26 includes a low pressure turbine flowpath 28 , The low pressure turbine 26 includes the counter-rotating low pressure inner shaft rotors and outer shaft rotors 200 and 202 with the low pressure inner shafts and outer shafts 130 and 140 at least partially rotatable coaxially with the high-pressure rotor 33 and are disposed radially inside thereof. The low pressure internal shaft rotor 200 is equipped with first low-pressure turbine blade rows 148 extending through the low pressure flow path 28 extend and through the low pressure inner shaft 130 drivingly connected to a first fan blade row 13 are.
The low-pressure outer-shaft rotor 202 has second low pressure turbine blade rows 138 up, passing through the low pressure flow path 28 extend and through the low pressure outer shaft 140 drivingly connected to a second fan blade row 15 are. In the in the 2 and 3 In the exemplary embodiment illustrated, there are four rows each of the first and second low pressure turbine blade rows 148 and 138 , The booster 16 is drivingly connected to one of the low pressure inner and outer shafts 130 and 140 ,
As in the 2 and 3 shown, contain the first low-pressure turbine blade rows 148 at least one first intermeshing turbine blade row 58 that is between at least one second adjacent pair 214 the second low-pressure turbine blade rows 138 is arranged. The second low-pressure turbine blade rows 138 include at least one second intermeshing turbine blade row 62 that is between at least one first adjacent pair 212 the first low-pressure turbine blade rows 138 is arranged. A series of non-rotatable low pressure vanes 210 extends through the low pressure flowpath 28 between each non-interlocking adjacent pair 218 the first and second low pressure turbine blade rows 148 and 138 that have no intermeshing turbine blade row in between.
In the 3 illustrated special embodiment of the low-pressure turbine 26 has a series of non-rotatable low pressure vanes 210 on, the axially between the backmost pair 52 the first low-pressure turbine blade rows 148 are arranged. Each of the rows of non-rotatable low pressure vanes 210 is located between each non-intermeshing adjacent pair 218 from the first and second low pressure turbine blade rows 148 and 138 that have no intermeshing turbine blade row in between. There may be two or more first intermeshing turbine blade rows 58 between the second adjacent pairs 214 the second low-pressure turbine blade rows 138 and two or more second intermeshing turbine blade rows 62 can be between the first adjacent pairs 212 the first low-pressure turbine blade rows 148 be arranged. A turbine nozzle 220 is axially forward, upstream of and adjacent to the second low pressure turbine blade rows 138 arranged. The first interlocking turbine blade row 58 serves as a foremost row of the first low-pressure turbine blade rows 148 that with a backmost couple 214 the second low-pressure turbine blade rows 138 interlocked.
In 3 becomes a low-pressure outer-shaft rotor 202 in which three of the four second low pressure turbine blade rows 138 on second low-pressure turbine disks 238 are mounted. It becomes a low pressure internal shaft rotor 200 in which all first low pressure turbine blade rows 148 on first low-pressure turbine disks 248 are mounted. The last row of second low-pressure turbine blade rows 138 is the interlocking turbine blade row 62 , The interlocking turbine blade row 62 stands from an outer ring-shaped drum extension 70 low pressure internal shaft and outer shaft rotors 200 and 202 inside. 3 generally shows an embodiment of the invention in which one of the counter-rotating low pressure internal shaft and outer shaft rotors 200 and 202 the rotatable drum extension 70 of which the blades are one of the first and second intermeshing turbine blade rows 58 . 62 the first and second low pressure turbine blade rows 148 . 138 stand radially inwards.
In the 4 and 5 becomes an alternative embodiment of the low pressure turbine 26 shown in which a first and second row 102 and 104 the second low-pressure turbine blade rows 138 on a radially outward second low pressure turbine drum 100 are mounted. The second turbine drum 100 is a part of the low pressure internal shaft rotor 200 , The third row 106 the second low-pressure turbine blade rows 138 is a part of the rotating housing 108 , which is the radially outward located second turbine drum 100 supported and rotatably supported by the middle housing 60 and the rear turbine housing 155 , A back or fourth row 110 the second low-pressure turbine blade rows 138 is located on a with the rotating housing 108 connected turbine disk 112 the last stage. The first and second rows 102 and 104 the second low-pressure turbine blade rows 138 grab with first and second adjacent pairs 120 and 122 the first low-pressure turbine blade rows 148 each other. The first low-pressure turbine blade rows 148 are on first low-pressure turbine disks 238 assembled. A series of non-rotatable low pressure vanes 210 extends through the low pressure turbine flowpath 28 between a last or last pair 216 the second low-pressure turbine blade rows 138 , In the drawing are the non-rotatable low pressure vanes 210 shown as being between the third row 106 the second low-pressure turbine blade rows 138 on the rotating housing 108 and the last or fourth row 110 on the turbine disk 112 arranged in the last stage.
Alternatively, the first and second low pressure turbine blade rows 148 and 138 on radially inwardly and radially outwardly located first and second low pressure turbine drums. There may also be more first and second low pressure turbine blade rows 148 and 138 to be present as in the 4 and 5 and more than two rows of the first and second low pressure turbine blade rows may be shown 148 and 138 with more than two adjacent pairs of the first and second low pressure turbine blade rows 148 and 138 mesh.
A gas turbine engine component comprising: a low pressure turbine ( 26 ) with a low-pressure flow path, an inverse low-pressure inside and a low-pressure outer shaft rotor ( 200 . 202 ) with a low pressure inner or a low pressure outer shaft ( 130 . 140 ) coaxial with the high pressure rotor ( 33 ) and are disposed radially inside thereof, wherein the low pressure internal shaft rotor ( 200 ) first low-pressure turbine blade rows ( 148 ) extending through the low pressure flowpath ( 28 ), the low-pressure outer-shaft rotor ( 202 ) second low pressure turbine blade rows ( 138 ) passing through the low pressure flow path ( 28 ), the first low pressure turbine blade rows ( 148 ) at least one first intermeshing turbine blade row ( 58 ) between at least one second adjacent pair ( 214 ) of the second low-pressure turbine blade rows ( 138 ), the second low-pressure turbine blade rows ( 138 ) at least one second intermeshing turbine blade row ( 62 ) between at least one first pair ( 212 ) of the first low-pressure turbine blade rows ( 148 ) and characterized by at least one row of non-rotating low-pressure vanes ( 210 ) extending through the low pressure flowpath ( 28 ) extending between at least one first non-intermeshing adjacent pair ( 218 ) of the first and second low-pressure turbine blade rows ( 148 . 138 ) having no intermeshing turbine blade row therebetween.
Component according to claim 1, further comprising rows of non-rotatable low-pressure guide vanes ( 210 ) and non-interlocking adjacent pairs ( 218 ) from the first and second turbine blade rows ( 148 . 138 ) having therebetween no intermeshing turbine blade row, characterized in that each of the rows of non-rotating low pressure guide vanes ( 210 ) between one of the non-intermeshing adjacent pairs ( 218 ) from the first and second low pressure turbine blade rows ( 148 . 138 ) is arranged.
Component according to claim 2, characterized in that a turbine nozzle ( 220 ) axially in front, above and adjacent to the second low pressure turbine blade rows ( 138 ) is arranged.
The component of claim 2, further comprising the first low pressure turbine blade rows (10). 148 ) a front row ( 58 ), which belongs to the last pair ( 214 ) of the second low-pressure turbine blade rows ( 138 ) interlocks.
Component according to Claim 2, in which the second low-pressure turbine blade rows ( 138 ) two foremost rows ( 58 ) with the three backmost rows of the first low-pressure turbine blade rows ( 148 ) mesh.
Component according to claim 2, further comprising one of the counter-rotating low pressure inner and low pressure outer shaft rotors ( 200 . 202 ), including a rotating drum extension ( 70 ), from which the blades of one of the first and second intermeshing turbine blade rows ( 58 . 62 ) of the first and second low pressure turbine blades rows ( 148 . 138 ) stand radially inwards.
Gas turbine engine component ( 10 ), comprising: a high-pressure rotor ( 33 ) including a high-pressure turbine ( 24 ), which via a high-pressure shaft ( 27 ) with a high pressure compressor ( 18 ) and about an engine center axis ( 8th ), and a gas turbine engine component according to any one of the preceding claims, which via the low-pressure inner shaft ( 130 ) with a first fan blade row ( 13 ) and via the low pressure outer shaft ( 140 ) with a second fan blade row ( 15 ) is at least one auxiliary power unit ( 16 ) with the low pressure inner shaft or the low pressure outer shaft ( 130 . 140 ) is drivingly connected, and axially between the first Bläserschaufelreihe ( 13 ) and the high-pressure rotor ( 33 ), characterized in that the first and second fan blade rows ( 13 . 15 ) within a sheath current housing ( 21 ) are arranged radially outwardly and by a fan case ( 11 ).
Component according to claim 7, further comprising a plurality of rows of non-rotatable low-pressure guide vanes ( 210 ), characterized in that each of these rows of non-rotatable low-pressure guide vanes ( 210 ) between each adjacent pair ( 218 ) from first and second low pressure turbine blade rows ( 148 . 138 ) and has no intermeshing turbine blade row therebetween.
Component according to claim 8, characterized in that a turbine nozzle ( 220 ) axially upstream, upstream of and adjacent to the second low pressure turbine blade rows (138).
A component according to claim 8, wherein the first low pressure turbine blade rows ( 148 ) also a foremost row ( 59 ), which with the rearmost pair ( 214 ) of the second low-pressure turbine blade rows ( 138 ) interlocks.
DE2003616513 2002-07-30 2003-07-29 Aircraft gas turbine engine with Kontrollitschaufel for opposing low-pressure turbine Active DE60316513T2 (en)
US10/208,303 US6684626B1 (en) 2002-07-30 2002-07-30 Aircraft gas turbine engine with control vanes for counter rotating low pressure turbines
US208303 2002-07-30
DE60316513D1 DE60316513D1 (en) 2007-11-08
DE60316513T2 true DE60316513T2 (en) 2008-06-26
ID=30115205
DE2003616513 Active DE60316513T2 (en) 2002-07-30 2003-07-29 Aircraft gas turbine engine with Kontrollitschaufel for opposing low-pressure turbine
US (1) US6684626B1 (en)
EP (1) EP1387060B1 (en)
JP (1) JP4346375B2 (en)
CN (1) CN1309944C (en)
CA (1) CA2435360C (en)
DE (1) DE60316513T2 (en)
US6935837B2 (en) * 2003-02-27 2005-08-30 General Electric Company Methods and apparatus for assembling gas turbine engines
WO2006113900A2 (en) 2005-04-20 2006-10-26 Helicor, Inc. Methods and devices for relieving stress
US7360988B2 (en) * 2005-12-08 2008-04-22 General Electric Company Methods and apparatus for assembling turbine engines
US7921634B2 (en) * 2006-10-31 2011-04-12 General Electric Company Turbofan engine assembly and method of assembling same
RU2460905C2 (en) * 2010-07-29 2012-09-10 Открытое акционерное общество "Национальный институт авиационных технологий" (ОАО НИАТ) Axial-flow fan or compressor impeller and fan of bypass fanjet incorporating said impeller
CN104093720B (en) 2012-01-26 2017-04-12 H.隆德贝克有限公司 PDE9 inhibitors with imidazo triazinone backbone
RU2646987C2 (en) * 2013-12-10 2018-03-13 Виктор Михайлович Морозов Centrifuge-axial fan "sherdor"
FR3014945B1 (en) * 2013-12-16 2019-03-15 Snecma Exhaust case having a turbine floor for turbomachine
CN107810187A (en) 2015-07-07 2018-03-16 H.隆德贝克有限公司 For treating the PDE9 inhibitor with imidazotriazinones skeleton and Imidazopyrazines ketone skeleton of peripheral diseases
CN107208552A (en) * 2015-09-09 2017-09-26 苏犁 Multiple shaft sleeves transmission bidirectional rotary runner fan formula turbine and bushing wheel fan formula compressor
GB2155110A (en) * 1984-03-02 1985-09-18 Gen Electric High bypass ratio counter-rotating turbofan engine
GB2194593B (en) * 1986-08-29 1991-05-15 Gen Electric High bypass ratio, counter rotating gearless front fan engine
DE4122008C2 (en) 1991-07-03 1993-04-22 Mtu Muenchen Gmbh
2002-07-30 US US10/208,303 patent/US6684626B1/en not_active Expired - Fee Related
2003-07-17 CA CA 2435360 patent/CA2435360C/en not_active Expired - Fee Related
2003-07-29 EP EP03254728A patent/EP1387060B1/en not_active Expired - Fee Related
2003-07-29 DE DE2003616513 patent/DE60316513T2/en active Active
2003-07-30 JP JP2003282331A patent/JP4346375B2/en not_active Expired - Fee Related
2003-07-30 CN CNB031551858A patent/CN1309944C/en not_active IP Right Cessation
US6684626B1 (en) 2004-02-03
CN1487180A (en) 2004-04-07
EP1387060A3 (en) 2005-10-12
DE60316513D1 (en) 2007-11-08
CA2435360C (en) 2010-11-09
JP2004060661A (en) 2004-02-26
EP1387060A2 (en) 2004-02-04
EP1387060B1 (en) 2007-09-26
US20040020186A1 (en) 2004-02-05
CN1309944C (en) 2007-04-11
CA2435360A1 (en) 2004-01-30
JP4346375B2 (en) 2009-10-21