Cooled cooling air for blade air seal through outer chamber

Information

  • Patent Grant
  • 11773742
  • Patent Number
    11,773,742
  • Date Filed
    Monday, March 2, 2020
    4 years ago
  • Date Issued
    Tuesday, October 3, 2023
    a year ago
Abstract
A gas turbine engine according to an example of the present disclosure include a compressor section, a combustor, and a turbine section. The combustor has a radially outer surface that defines a diffuser chamber radially outwardly of the combustor. The turbine section has a high pressure turbine first stage blade that has an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap for tapping air has been compressed by the compressor and is passed through a heat exchanger. The air downstream of the heat exchanger passes through at least one pipe and into a manifold radially outward of the blade outer air seal, and then passes across the blade outer air seal to cool the blade outer air seal.
Description
BACKGROUND OF THE INVENTION

This application relates to the supply of high pressure cooling air to a blade outer air seal through an outer diameter chamber.


Gas turbine engines are known and typically include a fan delivering air into a bypass duct as propulsion. The fan also delivers air into a compressor where air is compressed and then delivered into a combustor. The air is mixed with fuel and ignited in the combustor. Products of this combustion pass downstream over turbine rotors driving them to rotate. The turbine rotors, in turn, rotate compressor rotors and the fan rotor.


As can be appreciated, many components in the turbine section see very high temperatures. Two such components would be the turbine blades and blade outer air seals. Blade outer air seals typically sit radially outwardly of the blades and maintain clearance to increase the efficient use of the products of combustion.


One type of blade outer air seal is a so-called self-acting clearance control blade outer air seal. In such a blade outer air seal, two components formed of different materials having different coefficients of thermal expansion combine to control the expansion of the blade outer air seals to, in turn, control the clearance with the blade.


Both the blade and the blade outer air seal are provided with cooling air.


Traditionally, a turbine rotated at the same speed as the fan rotor. More recently, it has been proposed to include a gear reduction between a fan drive turbine and the fan rotor. With this change, the pressures and temperatures seen across the turbine sections have increased.


Thus, to drive cooling air into the turbine, the cooling air must be at a higher pressure than in the past. The highest pressure in the gas turbine engine is that downstream of a high pressure compressor. However, this cooling air is also at relatively high temperatures.


Thus, it has been proposed to tap high pressure air from a location downstream of the high pressure compressor and pass it through a heat exchanger prior to being delivered to the turbine section for cooling.


It has also been proposed to tap lower pressure from a more forward station in the compressor. The air is ducted to a separate auxiliary compressor and one or more heat exchangers to increase pressure and decrease the temperature of the air.


SUMMARY OF THE INVENTION

A gas turbine engine according to an example of the present disclosure include a compressor section, a combustor, and a turbine section. The combustor has a radially outer surface that defines a diffuser chamber radially outwardly of the combustor. The turbine section has a high pressure turbine first stage blade that has an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap for tapping air has been compressed by the compressor and is passed through a heat exchanger. The air downstream of the heat exchanger passes through at least one pipe and into a manifold radially outward of the blade outer air seal, and then passes across the blade outer air seal to cool the blade outer air seal.


In a further embodiment of any of the foregoing embodiments, the air downstream of the heat exchanger passes into a mixing chamber and is mixed with higher temperature air from a diffuser chamber outwardly of the combustor, and mixed and passes to cool a first stage blade row in a high pressure turbine.


In a further embodiment of any of the foregoing embodiments, the compressor diffuser chamber has an outer boundary defined by an outer core housing and the pipes are radially outward of the outer core housing.


In a further embodiment of any of the foregoing embodiments, the pipes communicate with the supply of cooled high pressure air at a location upstream of the mixing chamber such that air delivered to the manifold does not include hot air from the diffuser chamber.


In a further embodiment of any of the foregoing embodiments, the manifold is also outwardly of the outer core housing and communicates with passages passing through the outer core housing to the blade outer air seal.


In a further embodiment of any of the foregoing embodiments, the passages passing through the housing extend to a radially outer surface of the blade outer air seal and flow in both upstream and downstream locations around the blade outer air seal.


In a further embodiment of any of the foregoing embodiments, the air flowing upstream of the blade outer air seal is routed through holes in a seal portion of the blade outer air seal to cool adjacent a leading edge of the blade outer air seal and the air passing downstream of the blade outer air seal passing through holes in the seal portion of the blade outer air seal to cool adjacent a trailing edge of the blade outer air seal.


In a further embodiment of any of the foregoing embodiments, the blade outer air seal includes at least two components having different thermal coefficients of expansion to provide clearance control between an outer periphery of the blades and an inner periphery of the seal portion.


In a further embodiment of any of the foregoing embodiments, the mixing chamber is radially outward of a compressor diffuser defined downstream of a downstream most location in a high pressure compressor section and the air from the mixing chamber passes through vanes in the compressor diffuser.


In a further embodiment of any of the foregoing embodiments, the cooling air is tapped from a location downstream of a downstream most location in a high pressure compressor.


In a further embodiment of any of the foregoing embodiments, the air is tapped from a location upstream of a downstream most location in the compressor section.


In a further embodiment of any of the foregoing embodiments, the manifold is also outwardly of the outer core housing and communicates with passages passing through the outer core housing to the blade outer air seal.


In a further embodiment of any of the foregoing embodiments, the compressor diffuser chamber has an outer boundary defined by an outer core housing and the pipes are radially outward of the outer core housing.


In a further embodiment of any of the foregoing embodiments, the pipes communicate with the supply of cooled high pressure air at a location upstream of the mixing chamber such that air delivered to the manifold does not include hot air from the diffuser chamber.


In a further embodiment of any of the foregoing embodiments, the manifold is also outwardly of the outer core housing and communicates with passages passing through the outer core housing to the blade outer air seal.


In a further embodiment of any of the foregoing embodiments, the passages passing through the housing extend to a radially outer surface of the blade outer air seal and flow in both upstream and downstream locations around the blade outer air seal.


In a further embodiment of any of the foregoing embodiments, the air flowing upstream of the blade outer air seal being routed through holes in a seal portion of the blade outer air seal to cool adjacent a leading edge of the blade outer air seal and the air passing downstream of the blade outer air seal passing through holes in the seal portion of the blade outer air seal to cool adjacent a trailing edge of the blade outer air seal.


In a further embodiment of any of the foregoing embodiments, the mixing chamber is radially outward of a compressor diffuser defined downstream of a downstream most location in a high pressure compressor section and the air from the mixing chamber passing through vanes in the compressor diffuser.


In a further embodiment of any of the foregoing embodiments, the blade outer air seal includes at least two components that have different thermal coefficients of expansion to provide clearance control between an outer periphery of the blades and an inner periphery of the seal portion.


In a further embodiment of any of the foregoing embodiments, a valve controls the air passing across the blade outer air seal, and allows control of at least one of an amount, a pressure or a temperature of the air being delivered to the blade outer air seal.


These and other features may be best understood from the following drawings and specification.





BRIEF DESCRIPTION OF THE DRAWINGS


FIG. 1 schematically shows a gas turbine engine.



FIG. 2 shows a cooling system.



FIG. 3 shows details of a blade outer air seal cooling system.





DETAILED DESCRIPTION


FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.


The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.


The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.


The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded through the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.


The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.


A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7°R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second).



FIG. 2 shows a cooling system 100 for cooling turbine components. As shown, a compressor section 101 is provided with a tap 102 for tapping pressurized air.


The tap 102 may be at a location upstream from a downstream most portion of a high pressure compressor, in which case, it is typically provided with a boost compressor to raise its pressure. Alternatively, the air can be tapped from a location 99 where it has been fully compressed by the high pressure compressor.


In either case, pressurized air passes through a heat exchanger 104 where it is cooled, such as by air. In one embodiment, the heat exchanger 104 may be in the bypass duct as described in FIG. 1. From heat exchanger 104, air passes into conduit 106.


From the conduit 106, the air passes into a mixing chamber 108, which may be outward of a compressor diffuser 109. The air passes through vanes in the compressor diffuser 109, such that it is separate from the air downstream of a downstream most compression point 99. The air passes, as shown at 116, to cool a turbine blade 118. In the mixing chamber 108, hot air is shown at 110 mixing with the cool high pressure air from the conduit 106. This air is from a diffuser chamber 112, and is at the pressure downstream of the downstream most point 99. As such, it mixes easily with the air in the mixing chamber such that the air delivered at 116 is not unduly cool.


The chamber 112 is outward of a combustion chamber 114. An outer core housing 113 is positioned outwardly of the chamber 112.


A plurality of pipes 120 (only one of which is shown) tap air from the conduit 106 upstream of the mixing chamber 108. As such, this air is entirely the cooled high pressure cooling air. The air from the plurality of pipes is delivered into a manifold 122 which extends circumferentially over more than 270° about an axis of rotation of the engine. In embodiments, the manifold 122 extends over 360° about the axis of rotation. That air then passes through a plurality of pipes 124 to cool a blade outer air seal 126.


A valve 150 is shown schematically. The valve 150 may be controlled by a control 151 to control the cooling air being sent to the blade outer air seal 126. As an example, the valve may control the amount, pressure or temperature of the air being delivered to the blade outer air seal 126. An optional line 152 may selectively bypass the heat exchanger 104 to allow temperature control, as an example.



FIG. 3 shows details. As shown, air from the pipes 120 enters manifold 122 and then flows through pipe 124 to the blade outer air seal 126. The blade outer air seal 126 is shown to have components 128 and 130 which are formed of materials having distinct coefficients of thermal expansion. These components expand at different rates in response to exposure to heat and provide clearance control for a clearance between an inner portion of a seal 132 and an outer tip of the first stage high pressure turbine blade 118 as known.


As shown, the air flows at 134 upstream of the blade outer air seal components 128 and 130 and through holes 137 to cool a leading edge 136 of the seal portion 132. This cooling air drives, or controls, the expansion of the components and thus the clearance control.


Similarly, the air flows at 138 downstream of the blade outer air seal and through holes 142 to cool a trailing edge 140 of the blade outer air seal.


Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.

Claims
  • 1. A gas turbine engine comprising: a compressor section, a combustor, and a turbine section, said combustor having a radially outer surface defining a diffuser chamber radially outwardly of said combustor;said turbine section including a high pressure turbine first stage blade having an outer tip, and a blade outer air seal positioned radially outwardly of said outer tip;a tap for tapping air having been compressed by said compressor section being passed through a heat exchanger;a first portion of said air downstream of said heat exchanger passing through a plurality of pipes and into a manifold radially outward of said blade outer air seal, and then passing across said blade outer air seal to cool said blade outer air seal;wherein a second portion of said air downstream of said heat exchanger passing into a mixing chamber without passing through said manifold, and is mixed with higher temperature air from a diffuser chamber outwardly of said combustor, and mixed air passed to cool a first stage blade row in a high pressure turbine;wherein said blade outer air seal includes at least two components having different thermal coefficients of expansion to provide clearance control between said outer tip and an inner periphery of said seal portion;wherein a valve controls the air passing to said blade outer air seal, and allows control of at least one of an amount, a pressure or a temperature of the air being delivered to the blade outer air seal;wherein said diffuser chamber has an outer boundary defined by an outer core housing and said pipes are radially outward of said outer core housing;wherein said manifold is also outwardly of said outer core housing and communicates with passages passing through said outer core housing to said blade outer air seal;wherein said air flowing upstream of said blade outer air seal being routed through holes in a seal portion of said blade outer air seal to cool adjacent a leading edge of said blade outer air seal and the air passing downstream of said blade outer air seal passing through holes in said seal portion of said blade outer air seal to cool adjacent a trailing edge of said blade outer air seal; andwherein said mixing chamber is radially outward of a compressor diffuser defined downstream of a downstream most location in a high pressure compressor section and said mixed air from said mixing chamber passing through vanes in said compressor diffuser.
  • 2. A gas turbine engine comprising: a compressor section, a combustor, and a turbine section, said combustor having a radially outer surface defining a diffuser chamber radially outwardly of said combustor;said turbine section including a high pressure turbine first stage blade having an outer tip, and a blade outer air seal positioned radially outwardly of said outer tip;a tap for tapping air having been compressed by said compressor section being passed through a heat exchanger;a first portion of said air downstream of said heat exchanger passing through a plurality of pipes and into a manifold radially outward of said blade outer air seal, and then passing across said blade outer air seal to cool said blade outer air seal;wherein a second portion of said air downstream of said heat exchanger passing into a mixing chamber without passing through said manifold, and is mixed with higher temperature air from a diffuser chamber outwardly of said combustor, and mixed air passed to cool a first stage blade row in a high pressure turbine;wherein said blade outer air seal includes at least two components having different thermal coefficients of expansion to provide clearance control between said outer tip and an inner periphery of said seal portion;wherein a valve controls the air passing to said blade outer air seal, and allows control of at least one of an amount, a pressure or a temperature of the air being delivered to the blade outer air seal; andwherein said mixing chamber is radially outward of a compressor diffuser defined downstream of a downstream most location in a high pressure compressor section and said mixed air from said mixing chamber passing through vanes in said compressor diffuser.
REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No. 15/475,656 filed on Mar. 31, 2017.

US Referenced Citations (131)
Number Name Date Kind
2692476 Schaal et al. Oct 1954 A
3878677 Colvin Apr 1975 A
4254618 Elovic Mar 1981 A
4539945 Bosisio Sep 1985 A
4882902 Reigel et al. Nov 1989 A
5048288 Bessette Sep 1991 A
5056335 Renninger et al. Oct 1991 A
5269135 Vermejan et al. Dec 1993 A
5305616 Coffinberry Apr 1994 A
5392614 Coffinberry Feb 1995 A
5414992 Glickstein May 1995 A
5452573 Glickstein et al. Sep 1995 A
5498126 Pighetti et al. Mar 1996 A
5724806 Horner Mar 1998 A
5758485 Frutschi Jun 1998 A
5867979 Newton et al. Feb 1999 A
5918458 Coffinberry et al. Jul 1999 A
6050079 Durgin et al. Apr 2000 A
6065282 Fukue et al. May 2000 A
6134880 Yoshinaka Oct 2000 A
6430931 Horner Aug 2002 B1
6487863 Chen et al. Dec 2002 B1
6612114 Klingels Sep 2003 B1
6892523 Fetescu et al. May 2005 B2
7237386 Hoffmann et al. Jul 2007 B2
7246484 Giffin, III et al. Jul 2007 B2
7284377 Joshi et al. Oct 2007 B2
7306424 Romanov et al. Dec 2007 B2
7334412 Tiemann Feb 2008 B2
7347637 Kubo et al. Mar 2008 B2
7500365 Suciu et al. Mar 2009 B2
7552591 Bart et al. Jun 2009 B2
7698884 Maguire et al. Apr 2010 B2
7765788 Schwarz Aug 2010 B2
7823389 Seitzer Nov 2010 B2
7882691 Lemmers, Jr. et al. Feb 2011 B2
7886520 Stretton et al. Feb 2011 B2
8015828 Moniz et al. Sep 2011 B2
8037686 Lasker Oct 2011 B2
8087249 Ottaviano et al. Jan 2012 B2
8181443 Rago May 2012 B2
8307662 Turco Nov 2012 B2
8350398 Butt Jan 2013 B2
8397487 Sennoun et al. Mar 2013 B2
8402742 Roberge et al. Mar 2013 B2
8434997 Pinero et al. May 2013 B2
8511967 Suciu et al. Aug 2013 B2
8522529 Martinou et al. Sep 2013 B2
8572982 Tiemann Nov 2013 B2
8602717 Suciu et al. Dec 2013 B2
8621871 McCune et al. Jan 2014 B2
8727703 Laurello et al. May 2014 B2
8776952 Schwarz et al. Jul 2014 B2
8814502 Eleftheriou Aug 2014 B2
8876465 Stretton Nov 2014 B2
8961108 Bergman et al. Feb 2015 B2
9234481 Suciu et al. Jan 2016 B2
9243563 Lo Jan 2016 B2
9255492 Bacic Feb 2016 B2
9297391 Rued et al. Mar 2016 B2
9422063 Diaz Aug 2016 B2
9429072 Diaz et al. Aug 2016 B2
20030046938 Mortzheim et al. Mar 2003 A1
20040088995 Reissig May 2004 A1
20050172612 Yamanaka et al. Aug 2005 A1
20070022735 Henry et al. Feb 2007 A1
20070213917 Bruno et al. Sep 2007 A1
20070245738 Stretton et al. Oct 2007 A1
20080028763 Schwarz et al. Feb 2008 A1
20080230651 Porte Sep 2008 A1
20080253881 Richards Oct 2008 A1
20090007567 Porte et al. Jan 2009 A1
20090090096 Sheridan Apr 2009 A1
20090145102 Roberge et al. Jun 2009 A1
20090196736 Sengar et al. Aug 2009 A1
20090226297 Yanagi et al. Sep 2009 A1
20090272120 Tiemann Nov 2009 A1
20100043396 Coffinberry Feb 2010 A1
20100154434 Kubota et al. Jun 2010 A1
20110036066 Zhang et al. Feb 2011 A1
20110088405 Turco Apr 2011 A1
20110120083 Giffin et al. May 2011 A1
20110247344 Glahn et al. Oct 2011 A1
20120067055 Held Mar 2012 A1
20120102915 Baltas May 2012 A1
20120159961 Krautheim et al. Jun 2012 A1
20120180509 DeFrancesco Jul 2012 A1
20130036747 Fuchs et al. Feb 2013 A1
20130067928 Arias Chao et al. Mar 2013 A1
20130098059 Suciu et al. Apr 2013 A1
20130145744 Lo et al. Jun 2013 A1
20130145774 Duong et al. Jun 2013 A1
20130186102 Lo Jul 2013 A1
20130199156 Ress, Jr. et al. Aug 2013 A1
20130239583 Suciu et al. Sep 2013 A1
20130319002 Sidelkovskiy et al. Dec 2013 A1
20140020506 Duong Jan 2014 A1
20140137417 Silberberg et al. May 2014 A1
20140196469 Finney et al. Jul 2014 A1
20140230444 Hao et al. Aug 2014 A1
20140250898 Mackin et al. Sep 2014 A1
20140260326 Schwarz et al. Sep 2014 A1
20140311157 Laurello et al. Oct 2014 A1
20140341704 Fletcher Nov 2014 A1
20140352315 Diaz Dec 2014 A1
20150114611 Morris et al. Apr 2015 A1
20150285147 Phillips et al. Oct 2015 A1
20150308339 Forcier Oct 2015 A1
20150330236 Beecroft et al. Nov 2015 A1
20150354465 Suciu et al. Dec 2015 A1
20150354822 Suciu et al. Dec 2015 A1
20160010554 Suciu Jan 2016 A1
20160131036 Bintz et al. May 2016 A1
20160131037 Spangler et al. May 2016 A1
20160169118 Duong Jun 2016 A1
20160177830 Guardi Jun 2016 A1
20160215732 Malecki Jul 2016 A1
20160237906 Suciu et al. Aug 2016 A1
20160312797 Suciu et al. Oct 2016 A1
20160341125 Kraft et al. Nov 2016 A1
20160369697 Schwarz et al. Dec 2016 A1
20170009657 Schwarz et al. Jan 2017 A1
20170044980 Duesler et al. Feb 2017 A1
20170044982 Duesler et al. Feb 2017 A1
20170152765 Uechi et al. Jun 2017 A1
20170159568 Sennoun et al. Jun 2017 A1
20170167388 Merry et al. Jun 2017 A1
20170175632 Hanrahan et al. Jun 2017 A1
20170184027 Moniz et al. Jun 2017 A1
20170204736 Varney Jul 2017 A1
20170204787 Duesler et al. Jul 2017 A1
Foreign Referenced Citations (31)
Number Date Country
2852057 Jun 1979 DE
0447886 Sep 1991 EP
0469825 Feb 1992 EP
0608142 Jul 1994 EP
0903484 Mar 1999 EP
1314872 May 2003 EP
1630385 Mar 2006 EP
1923539 May 2008 EP
1944475 Jul 2008 EP
2085599 Aug 2009 EP
2128023 Dec 2009 EP
2362081 Aug 2011 EP
2540991 Jan 2013 EP
2584172 Apr 2013 EP
2604825 Jun 2013 EP
2733322 May 2014 EP
2865981 Apr 2015 EP
2942490 Nov 2015 EP
3085923 Oct 2016 EP
3085924 Oct 2016 EP
3121411 Jan 2017 EP
2851295 Aug 2004 FR
1244340 Aug 1971 GB
2152148 Jul 1985 GB
H1136889 Feb 1999 JP
2003037715 May 2003 WO
2008082335 Jul 2008 WO
2013154631 Oct 2013 WO
2014046713 Mar 2014 WO
2014092777 Jun 2014 WO
2014120125 Aug 2014 WO
Non-Patent Literature Citations (23)
Entry
Dornheim, Michael A., Rolls-Royce Trent 1000 to Drive Boeing 787 Accessories From IP Spool, Aviation Week & Space Technology, Mar. 28, 2005, p. 51, Los Angeles, CA.
U.S. Appl. No. 15/232,101.
U.S. Appl. No. 14/964,984.
U.S. Appl. No. 14/967,446.
U.S. Appl. No. 15/069,197.
U.S. Appl. No. 15/269,014.
U.S. Appl. No. 15/373,072.
European Search Report for European Application No. 16166707.6 dated Sep. 26, 2016.
European Search Report for European Application No. 16166724.1 dated Sep. 26, 2016.
European Search Report for European Patent Application No. 16154635.3 dated Jul. 6, 2016.
European Search Report for European Application No. 16155316.9 completed Jun. 30, 2016.
European Search Report for Application No. 16170021.6 dated Oct. 11, 2016.
European Search Report for Application No. 16174862.9 dated Nov. 7, 2016.
European Search Report for European Application No. 16175531.9 dated Nov. 15, 2016.
European Search Report for European Application No. 16175533.5 dated Nov. 15, 2016.
European Search Report for European Application No. 16175552.5 dated Nov. 17, 2016.
European Search Report for European Application No. 16175760.4 dated Nov. 16, 2016.
European Search Report for Application No. 16178207.3 dated Nov. 21, 2016.
European Search Report for European Application No. 16202876.5 dated Apr. 24, 2017.
European Search Report for European Application No. 16180657.5 dated Dec. 16, 2016.
European Search Report for EP Application No. 17160816.9 dated Jul. 21, 2017.
European Search Report for EP Application No. 18165234.8 dated Aug. 6, 2018.
European Search Report for EP Application No. 20208109.7 dated Mar. 10, 2021.
Related Publications (1)
Number Date Country
20200347741 A1 Nov 2020 US
Continuations (1)
Number Date Country
Parent 15475656 Mar 2017 US
Child 16806168 US