|Publication number||US3759038 A|
|Publication date||Sep 18, 1973|
|Filing date||Dec 9, 1971|
|Priority date||Dec 9, 1971|
|Also published as||CA961654A, CA961654A1, DE2258719A1|
|Publication number||US 3759038 A, US 3759038A, US-A-3759038, US3759038 A, US3759038A|
|Inventors||A Scalzo, Laurin L Mc|
|Original Assignee||Westinghouse Electric Corp|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (5), Referenced by (109), Classifications (9)|
|External Links: USPTO, USPTO Assignment, Espacenet|
United States Patent 1191 Scalzo et a1.
1451 Sept. 18, 1973  SELF ALIGNING COMBUSTOR AND 3,609,968 10/1971 Mierley et a1 60/3932 TRANSITION STRUCTURE FOR A GAS 2,592,060 4/1952 Oulianoff 60/3932 TURBINE 2,774,618 12/1956 Alderson 285/302 X 2,494,659 1/1950 Huyton 60/3932 UX  Inventors: Augustine J. Scalzo, Philadelphia;
Leroy McLaurm Springfield Primary ExaminerCarlton R. Croyle both of Assistant ExaminerRobert E. Garrett  Assignee: Westinghouse Electric Corporation, y Stratum et Pittsburgh, Pa.
 Filed: Dec. 9, 1971  ABSTRACT P1 0 33 Gas turbine combustion apparatus in which the combustor is anchored at its upstream end and the transi- 52 us. c1. 60/3932 415/134 member is anchmd its dwmream 51 1111. c1 F02C 7/20 rold 9/04 kW/Stream end cmbustr and the uPStream  Field of Search 60/3932 3937- and member being disPSed mutual 285/164 165 302 319 415/134 engagement and alignment and freely slidable relation during thermal expansion by a clevis support structure  References Cited and sealed at their engaging portions by a slotted twin UNITED STATES PATENTS cuff 2,748,567 6/1956 Dougherty 60/39.37 ux 9 Claims, 8 Drawing Figures 54 r f' f 44b 1 441 52 55 55 47 45 446 44C i9 520 Q 1 I O O o 0 E i I 530 46 A 4 48 49 52b g; 20 i I A 68 k 22 76 H H H n H H 42 1 e Q 4o Patented Sept. 18, 1973 3,759,038
4 ShcetsSheet I FIG. I
Patented Sept. 18, 1973 3,759,038
4 Sheets-Sheet 4 1 SELF ALIGNING COMBUSTOR AND TRANSITION STRUCTURE FOR A GAS TURBINE BACKGROUND OF THE INVENTION Present high efficiency gas turbines require ever increasing operating temperature capabilities in order to minimize cost of operation and to extract the greatest amount of useful power from the fuel consumed. In such turbines the supporting and sealing structure of the combustion apparatus must be capable of accommodating higher temperatures without over-stressing associated components. Present apparatus of this type either does not permit the required thermal displacements without undesirable thermal stress or utilizes support walls that rob the components of the beneficial cooling effects of the pressurized compressor air dis charge.
This invention provides gas turbine combustion apparatus that permits combustor and transition member thermal displacement with a minimum of elastic and frictional restraint. In addition, the connection between the combustor and the associated transition member permits axial displacement of both the combustor and the transition member and limited radial displacement of the combustor at the connection, while restraining the transition member against radial displacement.
SUMMARY OF THE INVENTION Briefly, the present invention relates to gas turbine combustion apparatus of the type in which a plurality of combustors of the canister type are disposed in a plenum chamber and arranged in an annular array about the rotational axis of the rotor. Each combustor is anchored at its upstream end to the upstream wall of the outer casing and is telescopically connected at its downstream end by a slotted twin layer cuff spring to the upstream end of its associated transition member.
The transition member is supported by a Y-shaped yoke having its leg anchored to the compressor diffuser casing and its diverging arms straddling the transition member. Each of the arms carries a clevis member that slidably supports an associated clevis guide that is, in turn, attached to the transition member. The yoke restrains the transition member against movement in a plane normal to'the longitudinal axis of the combustor but permits axial movement. I
The cuff spring provides an annular seal between the transition member and the combustor that permits transverse as well as axial movement of the combustor to occur relative to the transition member. The cuff spring has two layers of springs to restrict leakage and to dampen vibrations, in operation, and is further formed with a back up ring to limit spring stress, and
spring catches to prevent pieces of the spring from entering the transition member in the event of breakage. The-transition mouth, i.e., the downstream end of the transition member, is locked and sealed to theturbine blade ring by an arrangement that prevents leakage therepast but permits some sliding relative thereto.
BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a longitudinal sectional view of a portion of an axial flow gas turbine showing fuel combustion apparatus having the invention incorporated therein;
FIG. 2 is an enlarged cross-sectional view taken on line II II of FIG. 1 and showing three of the combustor transition members with their associated support structure;
FIG. 3 is a cross-sectional view, similar to FIG. 2, but taken on a'still larger scale and showing only one of the combustor transition members;
FIG. 4 is a fragmentary view taken on line lVIV in FIG. 3 and showing one of the clevis assemblies;
FIG. 5 is an enlarged fragmentary sectional view showing the mouth or downstream end portion of one of the transition members and its locking structure;
FIG. 6 is a fragmentary perspective view of the mouth portions of two adjacent transition members;
FIG. 7 is a diametric axial sectional view of one of the cuff springs, and
FIG. 8 is an enlarged fragmentary sectional view showing the cuff spring in assembled and sealing relation with the combustor and transition member.
DESCRIPTION OF THE PREFERRED EMBODIMENT Referring to the drawings in detail, in FIG. 1 there is shown a radial sectional view of the central portion of a gas turbine power plant 10 having the invention incorporated therein. The gas turbine 10 is of the wellknown axial flow type comprising a multi-stage axial flow compressor portion 12, an axial flow turbine portion 14 and combustion apparatus 16.
The turbine 14 includes at least one motive fluid expansion stage including an annular row of stationary blades 18 disposed in a blade ring structure 20 and preceding an annular row of rotatable blades 21 carried by a turbine rotor 22.
The compressor 12 includes a plurality of air compressing stages, each stage comprising an annular row of rotatable blades 24 carried by a rotor 25 and preceding an annular row of stationary blades 26.
The compressor rotor 25 is drivenly connected to the turbine rotor 22 by a torque tube 28.
The combustion apparatus 16 includes a plurality of combustors 30 and associated transition members 31 arranged in an annular array concentric with the rotor aggregate (25, 22, 28). The combustors 30 are disposed in a plenum chamber 32 whose outer periphery is defined by outer casing or housing structure including a central tubular casing portion 33 and upstream and downstream casing portions 34 and 35. The inner periphery is defined generally by a portion of the com- .pressor casing structure 36 and 37, and by a tubular fairing structure, 39, 40 and 41 extending from the compressor casing structure 36 to the turbine blade ring 20. The fairing structure 39, 40, 41 encompasses the torque tube 28 and is anchored at its upstream end to the compressor casing 36 by the last row of stationary blades 26 and at its downstream end to the blade ring structure 20 by annular flange structure 42.
The fairing structure 39 also cooperates with the compressor casing structure 37 to form a diffuser structure having a passageway 43 diverging in the direction of flow of the compressed air from the compressor into the plenum chamber 32, as indicated by the arrow A.
A plurality of combustors 30 and associated transition members 31 are disposed in an equally spaced annular array, as indicated in FIG. 2; however, since they may all be substantially identical, only one will be described.
The combustor 30, as best shown in FIG. I, is of the stepped liner construction employing a plurality of cylindrical liners 44a-44f of graduated diameter disposed in axially spaced telescopic engagement with each other and defining a fuel combustion chamber 45 into which fuel from a suitable supply (not shown) is admit-- ted by a suitable fuel injection device 46 and ignited by an igniter 47.
To support the combustion of fuel, pressurized pri-- mary air is admitted from the plenum chamber 32 into the combustion chamber 45 by a plurality of apertures 48 in the liners 44a and 44b. In addition, secondary air is admitted through a plurality of apertures 49 in the last liner 44fto dilute the hot gaseous products of combustion to a temperature that the hot components of the turbine 14 can safely withstand.
The downstream end portion 51 of the last combustion liner 44f is disposed in internal telescopic engagement with the tubular upstream end portion 52 of the transition member 31. The end portion 52 is divided into two mating semi-cylindrical portions 52a and 52b bolted together at diametrically opposed mating flanges 53a and 53b (only one pair shown) by suitable bolts.
As best seen in FIGS. 1 and 8, the liner end portion 51 is of smaller diameter than the transition end portion 52 and is loosely received therein. Referring to FIG. 8, cuff spring 55 of annular shape comprising an inner segmented spring member 56 and an outer segmented spring member 57 is attached at one end 58 to the liner portion 51 and has its other end in spring biased relation with the inner wall surface of the transition end portion 52. The segments 57a of the outer spring member 57 are in peripherally staggered relation with the segments 56a of the inner spring members to provide a peripheral seal between the transition member and the combustor during operation.
A back-up ring member 59 having a radial, peripherally upstanding flange portion 60 is attached to the liner end portion 51 adjacent the free end portion of the cuff spring and positioned in radially spaced relation therewith during normal operation, but is so proportioned as to abut the inner spring 56 to limit deflection of the springs 56 and 57 during assembly, thereby to protect the cuff spring against excessive spring stress.
The outer spring segments 57a are provided with radially inwardly bent free end portions 61 extending beyond the flange 60 of the back-up ring (with regard to the fixed fulcrum point 58), so that should one of the segments 56a or 57a break away from the cuff spring, the broken segment will be prevented from entering the transition member 31 and being entrained in the motive gas flow. More specifically, should an outer segment 57a break away, movement of the bent end portion 61 to the right is arrested by abutment with the flange60. Should an inner segment 56a break away it is inherently held captive by the outer segments 574. Thus, broken segments 56a and/or 57a can cause no damage to the turbine blades.
The combustor and transition structure 30, 31 is supported intermediate its ends by a support structure 63. The support structure comprises a Y-shaped yoke member 64 having a pair of spaced divergent arms 65 ing the cuff spring 55 and the back-up ring member 59.
while the anchoring portion 68 extends across the diffusing passage 42 and is attached to the fairing member 41.
Referring to FIG. 3 the arms 65 of the yoke embrace a quadrant portion of the transition member within their bight 69 and are slidably attached to the transition member by a pair ofvclevis structures 70, one at each arm 65. The clevis structures are substantially identical and comprise a U-shaped male member 71 attached to the transition member 31 and a U-shaped female member 72 attached to the associated arm 65 at a right angle to the male member 71. The female member 72 has an open-ended groove 73 slidably embracing the central portion 74 of the male clevis member. As best seen in FIG. 4, the clevis permits left-to-right movement of the transition member, i.e., movement parallel to the transition members central longitudinal axis. Also, as best seen in FIG. 3, the clevis prevents or at least restricts movement in any direction in a plane normal to the central axis T of the transition member 31.
The transition member 31, as well known in the art, changes in cross-sectional shape from circular at its upstream or inlet end 52 to conform to the circular shape of the combustor 30, to arcuate at its downstream end or mouth 76. More particularly, the transition mouth 76, as best seen in FIG. 6 wherein two neighboring transition member mouths are shown in provided with a pair of circumferentially spaced radial walls 77 and 78 and a pair of radially spaced inner and outer arcuate walls 79 and 80.
The radial wall 77 is provided with three parallel ribs 77a, while the radial wall 78 is provided with two parallel ribs 780. The neighboring transition mouths 76 are disposed in close lateral relation with each other and the ribs 770 and 780 on adjacent walls 77 and 78 are spaced in complementary fashion to form an axial interlock therebetween. 1
The inner and outer arcuate walls 79 and 80 are provided with grooved members 79a and 80a (see FIGS. 5 and 6) which cooperate with mating flanged mem bers 82 and 83 retained by the fairing member flange 42 and the blade ring structure 20, respectively.
The transition mouth 76 is retained in assembled relation with the blade ring by a bracket 85 joined to the arcuate wall 80 by welding or the like and bolted to the blade ring 20 by suitable bolts 86.
It will now be apparent that the combustion apparatus 16 is provided with combustors 30 and associated transition members 31 that in operation may expand and elongate thermally with respect to each other and with respect to their anchoring means with a minimum of thermal stress, warpage and/or damage.
More particularly, the clevis structures 70 permit the transition members 31 to freely move in axial direction, i.e., towards the end casing wall 34 and/or towards the blade ring structure 20. The transition member mouths 76 are anchored to the blade ring structure 20 in the manner described above to permit radial and circumferentialexpansion while restricting blow by" of hot motive gases. g
In addition, the transition members 31 and the combustors 30 are supported at their free ends in a sturdy and reliable manner by the yoke support structure 63, while the combustors are maintained in sealing, yet freely movable, relation with their associated transition members 31 by the novel sealing arrangement compris- We claim:
1. A gas turbine comprising casing structure defining a plenum chamber,
a first motive expansion stage including a blade ring having a row of stationary blades and a rotor having a row of rotatable blades,
at least one fuel combustor disposed within said plenum chamber with its upstream end anchored to said casing structure, and effective to generate hot motive gases,
a transition member connected at its upstream end to the downstream end of said combustor and communicating at its downstream end with said first stage, and effective to transmit said gases to said first stage,
said transition member having its upstream end portion telescopically engaging the downstream end portion of said combustor in a manner to permit relative axial movement therebetween,
said transition member having its downstream end connected to said blade ring in a manner to restrain axial movement,
a support structure anchored to said casing structure and slidably supporting said transition member in a manner permitting movement parallel to its axis but restraining movement transverse thereto, and
an annular cuff spring interposed between the telescoping end portions of said transition member and said combustor,
said cuff spring comprising an inner segmented spring member and an outer segmented spring member, with said segments in peripherally staggered relation with each other to provide a peripheral seal between said transition member and said combustor during operation.
2. The structure recited in claim 1 in which one of the segmented spring members extends axially beyond the other and has its free end portions extending radially.
3. The structure recited in claim 1 in which the cuff spring is connected at one end to one of telescoping portions and has its other end in slidable abutment with the other of said telescoping portions, and
one of the telescoping portions having an annular radially extending stop portion in spaced relation with the cuff spring but engagable by the cuff spring during deflection thereof.
4. The structure recited in claim 1 in which the support structure comprises a yoke member having a pair of spaced arms, and
the transition member is disposed in the bight of said arms.
5. The structure recited in claim 4 in which the support structure further includes a leg portion extending radially inwardly,
the casing structure includes a tubular inner casing portion defining the plenum chamber in part, and
said leg portion is attached to said inner casing portion.
6. In a gas turbine power plant comprising an air compressor portion, an axial flow turbine portion, and combustion apparatus for generating hot motive combustion gases for motivating said turbine,
said turbine having a blade ring carrying a row of stationary blades and a rotor carrying a row of rotatable blades,
inner and outer casing structure defining an annular plenum chamber,
said combustion apparatus including an annular array of fuel combustors disposed in said plenum chamber with their upstream end portions attached to said outer casing,
each of said combustors having a transition member connected in axially slidable relation at its upstream end to the downstream end of its associated combustor and connected at its downstream end to said blade ring, whereby to conduct the hot motive gases from said combustor to said turbine,
a support structure carried by said inner casing and slidably supporting said transition member in a manner permitting movement parallel to its axis, but restraining movement transverse thereto,
said inner casing being concentric with the outer casing and having an annular portion for directing the pressurized air from the compressor to the plenum chamber,
said support structure comprising a yoke member having a pair of spaced arms and a leg member, said arms receiving the transition member within their bight, and
said leg member extending radially inward and being anchored to said annular portion.
7. The structure recited in claim 6 wherein the support structure further includes a clevis carried by. each of the arms,
said clevis having male and female members slidably associated with each other,
one of said clevis members being attached to the transition member,
the other of said clevis members being attached to the associated arm.
8. The structure recited in claim 6 in which each transition member has an arcuate mouth portion with circumferentially spaced radial walls at least partly defining an outlet for the motive gases, and
said radial walls have radial ribs so arranged that the mouth portions of neighboring transition members are keyed to each other.
9. The structure recited in claim 8 in which the mouth portion is further provided with radially spaced inner and outer arcuate walls cooperating with the radial walls to define the gas outlet, and
said arcuate walls having channel members for securing the mouth portion to the blade ring.
a: it a:
|Cited Patent||Filing date||Publication date||Applicant||Title|
|US2494659 *||Feb 6, 1946||Jan 17, 1950||Lucas Ltd Joseph||Pipe joint|
|US2592060 *||Mar 3, 1947||Apr 8, 1952||Rolls Royce||Mounting of combustion chambers in jet-propulsion and gas-turbine power-units|
|US2748567 *||Oct 13, 1949||Jun 5, 1956||Gen Motors Corp||Gas turbine combustion chamber with telescoping casing and liner sections|
|US2774618 *||Sep 13, 1954||Dec 18, 1956||Alderson Winston T||Combination ball and slip joint for pipes|
|US3609968 *||Apr 29, 1970||Oct 5, 1971||Westinghouse Electric Corp||Self-adjusting seal structure|
|Citing Patent||Filing date||Publication date||Applicant||Title|
|US3824030 *||Jul 30, 1973||Jul 16, 1974||Curtiss Wright Corp||Diaphragm and labyrinth seal assembly for gas turbines|
|US4009569 *||Jul 21, 1975||Mar 1, 1977||United Technologies Corporation||Diffuser-burner casing for a gas turbine engine|
|US4016718 *||Jul 21, 1975||Apr 12, 1977||United Technologies Corporation||Gas turbine engine having an improved transition duct support|
|US4413470 *||Mar 5, 1981||Nov 8, 1983||Electric Power Research Institute, Inc.||Catalytic combustion system for a stationary combustion turbine having a transition duct mounted catalytic element|
|US4422288 *||Mar 2, 1981||Dec 27, 1983||General Electric Company||Aft mounting system for combustion transition duct members|
|US4790137 *||Jul 17, 1987||Dec 13, 1988||The United States Of America As Represented By The Secretary Of The Air Force||Aircraft engine outer duct mounting device|
|US4921401 *||Feb 23, 1989||May 1, 1990||United Technologies Corporation||Casting for a rotary machine|
|US5265412 *||Jul 28, 1992||Nov 30, 1993||General Electric Company||Self-accommodating brush seal for gas turbine combustor|
|US5394687 *||Dec 3, 1993||Mar 7, 1995||The United States Of America As Represented By The Department Of Energy||Gas turbine vane cooling system|
|US5400586 *||Oct 25, 1993||Mar 28, 1995||General Electric Co.||Self-accommodating brush seal for gas turbine combustor|
|US5414999 *||Nov 5, 1993||May 16, 1995||General Electric Company||Integral aft frame mount for a gas turbine combustor transition piece|
|US5474306 *||Nov 16, 1994||Dec 12, 1995||General Electric Co.||Woven seal and hybrid cloth-brush seals for turbine applications|
|US5749218 *||Sep 28, 1995||May 12, 1998||General Electric Co.||Wear reduction kit for gas turbine combustors|
|US5749584 *||Sep 25, 1996||May 12, 1998||General Electric Company||Combined brush seal and labyrinth seal segment for rotary machines|
|US5761898 *||Aug 1, 1996||Jun 9, 1998||General Electric Co.||Transition piece external frame support|
|US5951250 *||Mar 5, 1998||Sep 14, 1999||Mitsubishi Heavy Industries, Ltd.||Turbine cooling apparatus|
|US5983641 *||Apr 30, 1997||Nov 16, 1999||Mitsubishi Heavy Industries, Ltd.||Tail pipe of gas turbine combustor and gas turbine combustor having the same tail pipe|
|US5987879 *||Jan 14, 1997||Nov 23, 1999||Mitsubishi Jukogyo Kabushiki Kaisha||Spring seal device for combustor|
|US6006523 *||Apr 30, 1997||Dec 28, 1999||Mitsubishi Heavy Industries, Ltd.||Gas turbine combustor with angled tube section|
|US6010132 *||May 9, 1995||Jan 4, 2000||General Electric Co.||Hybrid labyrinth and cloth-brush seals for turbine applications|
|US6027121 *||Oct 23, 1997||Feb 22, 2000||General Electric Co.||Combined brush/labyrinth seal for rotary machines|
|US6042119 *||May 14, 1997||Mar 28, 2000||General Electric Co.||Woven seals and hybrid cloth-brush seals for turbine applications|
|US6045134 *||Feb 4, 1998||Apr 4, 2000||General Electric Co.||Combined labyrinth and brush seals for rotary machines|
|US6105967 *||Aug 6, 1999||Aug 22, 2000||General Electric Co.||Combined labyrinth and brush seals for rotary machines|
|US6131910 *||May 13, 1997||Oct 17, 2000||General Electric Co.||Brush seals and combined labyrinth and brush seals for rotary machines|
|US6139018 *||Mar 25, 1998||Oct 31, 2000||General Electric Co.||Positive pressure-actuated brush seal|
|US6168162||Aug 5, 1998||Jan 2, 2001||General Electric Co.||Self-centering brush seal|
|US6173958||May 19, 1998||Jan 16, 2001||General Electric Co.||Hybrid labyrinth and cloth-brush seals for turbine applications|
|US6250640||Aug 17, 1998||Jun 26, 2001||General Electric Co.||Brush seals for steam turbine applications|
|US6257586||Apr 15, 1997||Jul 10, 2001||General Electric Co.||Combined brush seal and labyrinth seal segment for rotary machines|
|US6290232||Nov 16, 1999||Sep 18, 2001||General Electric Co.||Rub-tolerant brush seal for turbine rotors and methods of installation|
|US6331006||Jan 25, 2000||Dec 18, 2001||General Electric Company||Brush seal mounting in supporting groove using flat spring with bifurcated end|
|US6345494 *||Sep 20, 2000||Feb 12, 2002||Siemens Westinghouse Power Corporation||Side seal for combustor transitions|
|US6435513||Feb 26, 2001||Aug 20, 2002||General Electric Company||Combined brush seal and labyrinth seal segment for rotary machines|
|US6869082||Jun 12, 2003||Mar 22, 2005||Siemens Westinghouse Power Corporation||Turbine spring clip seal|
|US7093837||Sep 26, 2002||Aug 22, 2006||Siemens Westinghouse Power Corporation||Turbine spring clip seal|
|US7229249||Aug 27, 2004||Jun 12, 2007||Pratt & Whitney Canada Corp.||Lightweight annular interturbine duct|
|US7246995||Dec 10, 2004||Jul 24, 2007||Siemens Power Generation, Inc.||Seal usable between a transition and a turbine vane assembly in a turbine engine|
|US7377116||Apr 28, 2005||May 27, 2008||Siemens Power Generation, Inc.||Gas turbine combustor barrier structures for spring clips|
|US7421842||Jul 18, 2005||Sep 9, 2008||Siemens Power Generation, Inc.||Turbine spring clip seal|
|US7596954 *||Jul 9, 2004||Oct 6, 2009||United Technologies Corporation||Blade clearance control|
|US7631501 *||Dec 15, 2009||Alstom Technology Ltd||Profiled sealing body with spring section|
|US7909569||Jun 9, 2005||Mar 22, 2011||Pratt & Whitney Canada Corp.||Turbine support case and method of manufacturing|
|US7909570||Mar 22, 2011||Pratt & Whitney Canada Corp.||Interturbine duct with integrated baffle and seal|
|US7930891 *||May 10, 2007||Apr 26, 2011||Florida Turbine Technologies, Inc.||Transition duct with integral guide vanes|
|US8038389||Jan 4, 2006||Oct 18, 2011||General Electric Company||Method and apparatus for assembling turbine nozzle assembly|
|US8051663||Nov 9, 2007||Nov 8, 2011||United Technologies Corp.||Gas turbine engine systems involving cooling of combustion section liners|
|US8065881||Nov 29, 2011||Siemens Energy, Inc.||Transition with a linear flow path with exhaust mouths for use in a gas turbine engine|
|US8091365||Jan 10, 2012||Siemens Energy, Inc.||Canted outlet for transition in a gas turbine engine|
|US8142148||Mar 27, 2012||Snecma||Diffuser-nozzle assembly for a turbomachine|
|US8281602 *||Oct 9, 2012||General Electric Company||Circumferentially self expanding combustor support for a turbine engine|
|US8307656||Nov 13, 2012||United Technologies Corp.||Gas turbine engine systems involving cooling of combustion section liners|
|US8322146 *||Aug 22, 2008||Dec 4, 2012||Alstom Technology Ltd||Transition duct assembly|
|US8322976 *||Dec 4, 2012||General Electric Company||High temperature seal for a turbine engine|
|US8403634||Mar 26, 2013||General Electric Company||Seal assembly for use with turbine nozzles|
|US8418474 *||Sep 29, 2008||Apr 16, 2013||Alstom Technology Ltd.||Altering a natural frequency of a gas turbine transition duct|
|US8429919||Apr 30, 2013||General Electric Company||Expansion hula seals|
|US8448444 *||Feb 18, 2011||May 28, 2013||General Electric Company||Method and apparatus for mounting transition piece in combustor|
|US8499565||Oct 1, 2009||Aug 6, 2013||Siemens Energy, Inc.||Axial diffusor for a turbine engine|
|US8662828 *||May 25, 2009||Mar 4, 2014||Snecma||High pressure turbine of a turbomachine with improved assembly of the mobile blade radial clearance control box|
|US8769963 *||Jan 30, 2007||Jul 8, 2014||Siemens Energy, Inc.||Low leakage spring clip/ring combinations for gas turbine engine|
|US8888445||Aug 19, 2011||Nov 18, 2014||General Electric Company||Turbomachine seal assembly|
|US8978388 *||Jun 3, 2011||Mar 17, 2015||General Electric Company||Load member for transition duct in turbine system|
|US9328664||Nov 8, 2013||May 3, 2016||Siemens Energy, Inc.||Transition support system for combustion transition ducts for turbine engines|
|US20040251639 *||Jun 12, 2003||Dec 16, 2004||Siemens Westinghouse Power Corporation||Turbine spring clip seal|
|US20050235647 *||Mar 1, 2005||Oct 27, 2005||Alstom Technology Ltd.||Sealing body|
|US20060005529 *||Jul 9, 2004||Jan 12, 2006||Penda Allan R||Blade clearance control|
|US20060045730 *||Aug 27, 2004||Mar 2, 2006||Pratt & Whitney Canada Corp.||Lightweight annular interturbine duct|
|US20060127219 *||Dec 10, 2004||Jun 15, 2006||Siemens Westinghouse Power Corporation||Seal usable between a transition and a turbine vane assembly in a turbine engine|
|US20060242964 *||Apr 28, 2005||Nov 2, 2006||Siemens Westinghouse Power Corp.||Gas turbine combustor barrier structures for spring clips|
|US20060277922 *||Jun 9, 2005||Dec 14, 2006||Pratt & Whitney Canada Corp.||Turbine support case and method of manufacturing|
|US20070012043 *||Jul 18, 2005||Jan 18, 2007||Siemens Westinghouse Power Corporation||Turbine spring clip seal|
|US20070154305 *||Jan 4, 2006||Jul 5, 2007||General Electric Company||Method and apparatus for assembling turbine nozzle assembly|
|US20070214792 *||Mar 17, 2006||Sep 20, 2007||Siemens Power Generation, Inc.||Axial diffusor for a turbine engine|
|US20080050229 *||Aug 25, 2006||Feb 28, 2008||Pratt & Whitney Canada Corp.||Interturbine duct with integrated baffle and seal|
|US20080179837 *||Jan 30, 2007||Jul 31, 2008||Siemens Power Generation, Inc.||Low leakage spring clip/ring combinations for gas turbine engine|
|US20090120096 *||Nov 9, 2007||May 14, 2009||United Technologies Corp.||Gas Turbine Engine Systems Involving Cooling of Combustion Section Liners|
|US20090145137 *||Aug 22, 2008||Jun 11, 2009||Alstom Technologies, Ltd., Llc||Transition duct assembly|
|US20090188258 *||Sep 29, 2008||Jul 30, 2009||Alstom Technologies Ltd. Llc||Altering a natural frequency of a gas turbine transition duct|
|US20090212504 *||Apr 7, 2008||Aug 27, 2009||General Electric Company||High temperature seal for a turbine engine|
|US20090214333 *||Jan 29, 2009||Aug 27, 2009||Snecma||Diffuser-nozzle assembly for a turbomachine|
|US20100037617 *||Aug 12, 2008||Feb 18, 2010||Richard Charron||Transition with a linear flow path with exhaust mouths for use in a gas turbine engine|
|US20100037619 *||Feb 18, 2010||Richard Charron||Canted outlet for transition in a gas turbine engine|
|US20100058768 *||Oct 1, 2009||Mar 11, 2010||Robert Bland||Axial diffusor for a turbine engine|
|US20100300116 *||May 28, 2009||Dec 2, 2010||General Electric Company||Expansion Hula Seals|
|US20110061397 *||Mar 17, 2011||General Electric Company||Circumferentially self expanding combustor support for a turbine engine|
|US20110076135 *||May 25, 2009||Mar 31, 2011||Snecma||High pressure turbine of a turbomachine with improved assembly of the mobile blade radial clearance control box|
|US20120036857 *||Aug 10, 2010||Feb 16, 2012||General Electric Company||Combustion liner stop blocks having insertable wear features and related methods|
|US20120180500 *||Jan 13, 2011||Jul 19, 2012||General Electric Company||System for damping vibration in a gas turbine engine|
|US20120210729 *||Feb 18, 2011||Aug 23, 2012||General Electric Company||Method and apparatus for mounting transition piece in combustor|
|US20120304653 *||Dec 6, 2012||General Electric Company||Load member for transition duct in turbine system|
|US20150047358 *||Aug 14, 2013||Feb 19, 2015||General Electric Company||Inner barrel member with integrated diffuser for a gas turbomachine|
|US20150184856 *||Sep 26, 2013||Jul 2, 2015||Peter John Stuttaford||Thermally free liner retention mechanism|
|CN102022754A *||Sep 10, 2010||Apr 20, 2011||通用电气公司||Circumferentially self expanding combustor support for a turbine engine|
|CN102022754B||Sep 10, 2010||Dec 4, 2013||通用电气公司||Circumferentially self expanding combustor support for a turbine engine|
|CN102119268B *||Feb 18, 2009||Dec 3, 2014||西门子能源公司||Transition with a linear flow path with exhaust mouths for use in a gas turbine engine|
|CN102171413B||Feb 17, 2009||Mar 12, 2014||西门子能源公司||Canted outlet for transition in gas turbine engine|
|CN102588503A *||Jan 13, 2012||Jul 18, 2012||通用电气公司||System for damping vibration in a gas turbine engine|
|CN102808659A *||Jun 4, 2012||Dec 5, 2012||通用电气公司||Load member for transition duct in turbine system|
|CN102808659B *||Jun 4, 2012||Feb 10, 2016||通用电气公司||用于涡轮系统中过渡管道的载荷部件|
|EP1403584A1 *||Sep 3, 2003||Mar 31, 2004||Siemens Westinghouse Power Corporation||Turbine spring clip seal|
|EP2096266A1 *||Dec 11, 2008||Sep 2, 2009||Snecma||Nozzle-synchronising ring assembly for a turbomachine|
|WO1998016764A1 *||Jul 29, 1997||Apr 23, 1998||Siemens Westinghouse Power Corporation||Brush seal for gas turbine combustor-transition interface|
|WO2006118655A1 *||Mar 7, 2006||Nov 9, 2006||Siemens Power Generation, Inc.||Gas turbine combustor barrier structures for spring clips|
|WO2010019174A2 *||Feb 17, 2009||Feb 18, 2010||Siemens Energy, Inc.||Canted outlet for transition in a gas turbine engine|
|WO2010019174A3 *||Feb 17, 2009||Sep 10, 2010||Siemens Energy, Inc.||Canted outlet for transition in a gas turbine engine|
|WO2010019177A2 *||Feb 18, 2009||Feb 18, 2010||Siemens Energy, Inc.||Transition with a linear flow path with exhaust mouths for use in a gas turbine engine|
|WO2010019177A3 *||Feb 18, 2009||Sep 2, 2010||Siemens Energy, Inc.||Transition with a linear flow path with exhaust mouths for use in a gas turbine engine|
|WO2015069453A1 *||Oct 22, 2014||May 14, 2015||Siemens Energy, Inc.||Adjustable support system for a combustor transition duct|
|U.S. Classification||60/800, 415/117, 415/134|
|International Classification||F02C7/20, F01D9/02, F23R3/60|
|Cooperative Classification||F05D2230/642, F01D9/023|