US20160265777A1 - Modified floatwall panel dilution hole cooling - Google Patents
Modified floatwall panel dilution hole cooling Download PDFInfo
- Publication number
- US20160265777A1 US20160265777A1 US14/884,506 US201514884506A US2016265777A1 US 20160265777 A1 US20160265777 A1 US 20160265777A1 US 201514884506 A US201514884506 A US 201514884506A US 2016265777 A1 US2016265777 A1 US 2016265777A1
- Authority
- US
- United States
- Prior art keywords
- panel
- liner
- rail
- dilution hole
- wall
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/06—Arrangement of apertures along the flame tube
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/002—Wall structures
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/60—Support structures; Attaching or mounting means
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/03041—Effusion cooled combustion chamber walls or domes
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- Engine combustors have exhibited hot metal temperatures in liner panels in proximity to combustion/dilution holes. These high temperatures are largely a result of the difficulty in delivering a cooling fluid (e.g., air) to the metal mass (e.g., grommet) that surrounds the hole. Accordingly, a change is needed to the panel and an associating holding structure (e.g., shell) in the region around the dilution holes in order to enhance the availability and reliability of the liner.
- a cooling fluid e.g., air
- the metal mass e.g., grommet
- a liner of an aircraft comprising: a panel including at least one dilution hole, a rail coupled to the panel, and at least one effusion cooling hole located between a wall of the at least one dilution hole and a side of the rail.
- the at least one effusion cooling hole includes a plurality of effusion cooling holes.
- the rail is spaced from the wall by a distance of approximately 6.4 millimeters.
- the rail is spaced from the wall in a radial direction away from the at least one dilution hole.
- the liner further comprises at least a second effusion cooling hole located between a second side of the rail and a perimeter of the panel.
- the liner further comprises a shell coupled to the panel.
- the rail is configured to seal the panel against the shell to prevent leakage.
- the liner is associated with a combustion engine of the aircraft.
- FIG. 1 illustrates an exemplary gas turbine engine.
- FIG. 2 illustrates a portion of a liner incorporating grommets in accordance with the prior art.
- FIG. 3 illustrates the liner of FIG. 2 incorporating a shell in accordance with the prior art.
- FIG. 4 illustrates a first view of a liner incorporating a rail and effusion cooling holes in accordance with aspects of the disclosure.
- FIG. 5 illustrates a second view of the liner of FIG. 4 .
- connections are set forth between elements in the following description and in the drawings (the contents of which are included in this disclosure by way of reference). It is noted that these connections are general and, unless specified otherwise, may be direct or indirect and that this specification is not intended to be limiting in this respect.
- a coupling between two or more entities may refer to a direct connection or an indirect connection.
- An indirect connection may incorporate one or more intervening entities.
- the liner may include a panel coupled to a holding structure/shell.
- the coupling may be facilitated by one or more rails.
- Effusion cooling holes may be incorporated into the panel in order to cool the panel.
- FIG. 1 is a side-sectional illustration of an exemplary gas turbine engine 10 .
- the engine 10 includes a compressor section 12 , a turbine section 14 and one or more engine hot sections.
- the engine hot sections may include, for example, a first engine hot section 16 configured as a combustor section and a second engine hot section 18 configured as an augmentor section.
- the compressor section 12 , the first engine hot section 16 , the turbine section 14 and the second engine hot section 18 may be sequentially aligned along an axial centerline 20 between a forward engine airflow inlet 22 and an aft engine airflow exhaust 24 .
- the liner 200 includes a panel 204 .
- the panel 204 may be referred to as, or correspond to, a float wall panel.
- a float wall panel is a panel that is designed to expand under a heat-load without cracking.
- the panel 204 may take the form of a tiled structure that includes threaded studs 206 that allow the panel 204 to be bolted to a shell 302 (see FIG. 3 ).
- the panel 204 includes grommets 208 . At the center of the grommets 208 are dilution holes 210 that provide a combustion flow to the panel 204 and the associated combustor (e.g., first engine hot section 16 of FIG. 1 ).
- the panel 204 interfaces to the high temperatures associated with the combustor.
- the liner 200 of FIG. 2 is shown in more detail.
- the liner 200 includes the shell 302 at least partially coupled to the panel 204 via a grommet 208 .
- the shell 302 may include impingement holes (not shown) as would be known to one of skill in the art.
- a cooling hole 312 incorporated into the panel 204 .
- a center line associated with the dilution hole 210 is shown via a dashed line in FIG. 3 .
- a size (e.g., a diameter) of the dilution hole 210 may be referenced with respect to an edge/wall 208 a of the grommet 208 .
- the panel 204 /grommet 208 may be subject to distress (e.g., melting and oxidation).
- FIGS. 4-5 illustrate various views of a liner 400 .
- the liner 400 includes a shell 402 (which may be similar to, or correspond to, the shell 302 of FIG. 3 ) coupled to a panel 404 .
- the shell 402 may be at least partially coupled to the panel 404 via one or more rails 418 .
- the rail 418 may seal the panel 404 against the shell 402 to prevent leakage.
- the rail 418 may be cast in connection with the panel 404 , such that the rail 418 may be integral to, and may be made of the same material as, the panel 404 .
- the rail 418 may be spaced (in a radial direction) from an edge/wall 408 a associated with the dilution hole 210 by a distance D.
- the distance D may be approximately 6.4 millimeters.
- the spacing D allows for the insertion of one or more effusion cooling holes 412 adjacent to the wall 408 a, or in a location corresponding to a region of the panel 404 between the wall 408 a and a first side of the rail 418 .
- the holes 412 have a high pressure loss through them.
- Additional cooling holes 422 may be included in the panel 404 on the other/second side of the rail 418 , in between the rail 418 and a perimeter 504 a of the panel 404 .
- Cooling may be provided via effusion holes incorporated in the panel between a rail and an edge/wall of a dilution hole.
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Gasket Seals (AREA)
Abstract
Aspects of the disclosure are directed to a liner of an aircraft. The liner comprises a panel including at least one dilution hole, a rail coupled to the panel, and at least one effusion cooling hole located between a wall of the at least one dilution hole and a side of the rail. The liner may be associated with a combustion engine of the aircraft.
Description
- This application claims priority to U.S. patent application Ser. No. 62/065,312 filed Oct. 17, 2014.
- Engine combustors have exhibited hot metal temperatures in liner panels in proximity to combustion/dilution holes. These high temperatures are largely a result of the difficulty in delivering a cooling fluid (e.g., air) to the metal mass (e.g., grommet) that surrounds the hole. Accordingly, a change is needed to the panel and an associating holding structure (e.g., shell) in the region around the dilution holes in order to enhance the availability and reliability of the liner.
- The following presents a simplified summary in order to provide a basic understanding of some aspects of the disclosure. The summary is not an extensive overview of the disclosure. It is neither intended to identify key or critical elements of the disclosure nor to delineate the scope of the disclosure. The following summary merely presents some concepts of the disclosure in a simplified form as a prelude to the description below.
- Aspects of the disclosure are directed to a liner of an aircraft, comprising: a panel including at least one dilution hole, a rail coupled to the panel, and at least one effusion cooling hole located between a wall of the at least one dilution hole and a side of the rail. In some embodiments, the at least one effusion cooling hole includes a plurality of effusion cooling holes. In some embodiments, the rail is spaced from the wall by a distance of approximately 6.4 millimeters. In some embodiments, the rail is spaced from the wall in a radial direction away from the at least one dilution hole. In some embodiments, the liner further comprises at least a second effusion cooling hole located between a second side of the rail and a perimeter of the panel. In some embodiments, the liner further comprises a shell coupled to the panel. In some embodiments, the rail is configured to seal the panel against the shell to prevent leakage. In some embodiments, the liner is associated with a combustion engine of the aircraft.
- The present disclosure is illustrated by way of example and not limited in the accompanying figures in which like reference numerals indicate similar elements.
-
FIG. 1 illustrates an exemplary gas turbine engine. -
FIG. 2 illustrates a portion of a liner incorporating grommets in accordance with the prior art. -
FIG. 3 illustrates the liner ofFIG. 2 incorporating a shell in accordance with the prior art. -
FIG. 4 illustrates a first view of a liner incorporating a rail and effusion cooling holes in accordance with aspects of the disclosure. -
FIG. 5 illustrates a second view of the liner ofFIG. 4 . - It is noted that various connections are set forth between elements in the following description and in the drawings (the contents of which are included in this disclosure by way of reference). It is noted that these connections are general and, unless specified otherwise, may be direct or indirect and that this specification is not intended to be limiting in this respect. A coupling between two or more entities may refer to a direct connection or an indirect connection. An indirect connection may incorporate one or more intervening entities.
- In accordance with various aspects of the disclosure, apparatuses, systems and methods are described for cooling a liner of an aircraft combustor. The liner may include a panel coupled to a holding structure/shell. The coupling may be facilitated by one or more rails. Effusion cooling holes may be incorporated into the panel in order to cool the panel.
- Aspects of the disclosure may be applied in connection with an aircraft, or portion thereof. For example, aspects of the disclosure may be applied in connection with a gas turbine engine.
FIG. 1 is a side-sectional illustration of an exemplarygas turbine engine 10. Theengine 10 includes acompressor section 12, aturbine section 14 and one or more engine hot sections. The engine hot sections may include, for example, a first enginehot section 16 configured as a combustor section and a second enginehot section 18 configured as an augmentor section. Thecompressor section 12, the first enginehot section 16, theturbine section 14 and the second enginehot section 18 may be sequentially aligned along anaxial centerline 20 between a forwardengine airflow inlet 22 and an aftengine airflow exhaust 24. - Referring to
FIG. 2 , a portion of aliner 200 is shown. Theliner 200 includes apanel 204. Thepanel 204 may be referred to as, or correspond to, a float wall panel. A float wall panel is a panel that is designed to expand under a heat-load without cracking. Thepanel 204 may take the form of a tiled structure that includes threadedstuds 206 that allow thepanel 204 to be bolted to a shell 302 (seeFIG. 3 ). Thepanel 204 includesgrommets 208. At the center of thegrommets 208 aredilution holes 210 that provide a combustion flow to thepanel 204 and the associated combustor (e.g., first enginehot section 16 ofFIG. 1 ). Thepanel 204 interfaces to the high temperatures associated with the combustor. - Referring to
FIG. 3 , theliner 200 ofFIG. 2 is shown in more detail. Theliner 200 includes theshell 302 at least partially coupled to thepanel 204 via agrommet 208. Theshell 302 may include impingement holes (not shown) as would be known to one of skill in the art. Also shown is acooling hole 312 incorporated into thepanel 204. For purposes of reference, a center line associated with thedilution hole 210 is shown via a dashed line inFIG. 3 . A size (e.g., a diameter) of thedilution hole 210 may be referenced with respect to an edge/wall 208 a of thegrommet 208. - As described above, in can be difficult to provide a sufficient cooling flow to the
panel 204 in the proximity of the grommet 208 (via, e.g., the cooling hole 312), due in large to the mass/thickness of thegrommet 208 relative to the remainder of thepanel 204. If an insufficient cooling flow is provided, thepanel 204/grommet 208 may be subject to distress (e.g., melting and oxidation). -
FIGS. 4-5 illustrate various views of aliner 400. Theliner 400 includes a shell 402 (which may be similar to, or correspond to, theshell 302 ofFIG. 3 ) coupled to apanel 404. Theshell 402 may be at least partially coupled to thepanel 404 via one ormore rails 418. For example, therail 418 may seal thepanel 404 against theshell 402 to prevent leakage. Therail 418 may be cast in connection with thepanel 404, such that therail 418 may be integral to, and may be made of the same material as, thepanel 404. - The
rail 418 may be spaced (in a radial direction) from an edge/wall 408 a associated with thedilution hole 210 by a distance D. In some embodiments, the distance D may be approximately 6.4 millimeters. The spacing D allows for the insertion of one or moreeffusion cooling holes 412 adjacent to thewall 408 a, or in a location corresponding to a region of thepanel 404 between thewall 408 a and a first side of therail 418. Theholes 412 have a high pressure loss through them. Additional cooling holes 422 may be included in thepanel 404 on the other/second side of therail 418, in between therail 418 and aperimeter 504 a of thepanel 404. - Technical effects and benefits of this disclosure include a cost-effective design for cooling a panel. Cooling may be provided via effusion holes incorporated in the panel between a rail and an edge/wall of a dilution hole.
- Aspects of the disclosure have been described in terms of illustrative embodiments thereof Numerous other embodiments, modifications, and variations within the scope and spirit of the appended claims will occur to persons of ordinary skill in the art from a review of this disclosure. For example, one of ordinary skill in the art will appreciate that the steps described in conjunction with the illustrative figures may be performed in other than the recited order, and that one or more steps illustrated may be optional in accordance with aspects of the disclosure. One or more features described in connection with a first embodiment may be combined with one or more features of one or more additional embodiments.
Claims (8)
1. A liner of an aircraft, comprising:
a panel including at least one dilution hole;
a rail coupled to the panel; and
at least one effusion cooling hole located between a wall of the at least one dilution hole and a side of the rail.
2. The liner of claim 1 , wherein the at least one effusion cooling hole includes a plurality of effusion cooling holes.
3. The liner of claim 1 , wherein the rail is spaced from the wall by a distance of approximately 6.4 millimeters.
4. The liner of claim 3 , wherein the rail is spaced from the wall in a radial direction away from the at least one dilution hole.
5. The liner of claim 1 , further comprising:
at least a second effusion cooling hole located between a second side of the rail and a perimeter of the panel.
6. The liner of claim 1 , further comprising:
a shell coupled to the panel.
7. The liner of claim 6 , wherein the rail is configured to seal the panel against the shell to prevent leakage.
8. The liner of claim 1 , wherein the liner is associated with a combustion engine of the aircraft.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US14/884,506 US20160265777A1 (en) | 2014-10-17 | 2015-10-15 | Modified floatwall panel dilution hole cooling |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US201462065312P | 2014-10-17 | 2014-10-17 | |
US14/884,506 US20160265777A1 (en) | 2014-10-17 | 2015-10-15 | Modified floatwall panel dilution hole cooling |
Publications (1)
Publication Number | Publication Date |
---|---|
US20160265777A1 true US20160265777A1 (en) | 2016-09-15 |
Family
ID=54364971
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US14/884,506 Abandoned US20160265777A1 (en) | 2014-10-17 | 2015-10-15 | Modified floatwall panel dilution hole cooling |
Country Status (2)
Country | Link |
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US (1) | US20160265777A1 (en) |
EP (1) | EP3009745A1 (en) |
Cited By (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20160209033A1 (en) * | 2015-01-20 | 2016-07-21 | United Technologies Corporation | Combustor dilution hole passive heat transfer control |
EP3392566A1 (en) * | 2017-04-18 | 2018-10-24 | United Technologies Corporation | Combustor panel cooling arrangements |
US10753283B2 (en) | 2017-03-20 | 2020-08-25 | Pratt & Whitney Canada Corp. | Combustor heat shield cooling hole arrangement |
US10816202B2 (en) | 2017-11-28 | 2020-10-27 | General Electric Company | Combustor liner for a gas turbine engine and an associated method thereof |
US11029027B2 (en) * | 2018-10-03 | 2021-06-08 | Raytheon Technologies Corporation | Dilution/effusion hole pattern for thick combustor panels |
Citations (5)
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US20030021324A1 (en) * | 2001-07-24 | 2003-01-30 | Gsi Lumonics, Inc. | Waveguide device with mode control and pump light confinement and method of using same |
US20030213249A1 (en) * | 2002-05-14 | 2003-11-20 | Monica Pacheco-Tougas | Bulkhead panel for use in a combustion chamber of a gas turbine engine |
US20030213250A1 (en) * | 2002-05-16 | 2003-11-20 | Monica Pacheco-Tougas | Heat shield panels for use in a combustor for a gas turbine engine |
US20050086940A1 (en) * | 2003-10-23 | 2005-04-28 | Coughlan Joseph D.Iii | Combustor |
US20110048024A1 (en) * | 2009-08-31 | 2011-03-03 | United Technologies Corporation | Gas turbine combustor with quench wake control |
Family Cites Families (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB9919981D0 (en) * | 1999-08-24 | 1999-10-27 | Rolls Royce Plc | Combustion apparatus |
GB0227842D0 (en) * | 2002-11-29 | 2003-01-08 | Rolls Royce Plc | Sealing Arrangement |
US7748221B2 (en) * | 2006-11-17 | 2010-07-06 | Pratt & Whitney Canada Corp. | Combustor heat shield with variable cooling |
-
2015
- 2015-10-15 US US14/884,506 patent/US20160265777A1/en not_active Abandoned
- 2015-10-16 EP EP15190073.5A patent/EP3009745A1/en not_active Withdrawn
Patent Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20030021324A1 (en) * | 2001-07-24 | 2003-01-30 | Gsi Lumonics, Inc. | Waveguide device with mode control and pump light confinement and method of using same |
US20030213249A1 (en) * | 2002-05-14 | 2003-11-20 | Monica Pacheco-Tougas | Bulkhead panel for use in a combustion chamber of a gas turbine engine |
US20030213250A1 (en) * | 2002-05-16 | 2003-11-20 | Monica Pacheco-Tougas | Heat shield panels for use in a combustor for a gas turbine engine |
US20050086940A1 (en) * | 2003-10-23 | 2005-04-28 | Coughlan Joseph D.Iii | Combustor |
US20110048024A1 (en) * | 2009-08-31 | 2011-03-03 | United Technologies Corporation | Gas turbine combustor with quench wake control |
Cited By (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20160209033A1 (en) * | 2015-01-20 | 2016-07-21 | United Technologies Corporation | Combustor dilution hole passive heat transfer control |
US10132498B2 (en) * | 2015-01-20 | 2018-11-20 | United Technologies Corporation | Thermal barrier coating of a combustor dilution hole |
US10753283B2 (en) | 2017-03-20 | 2020-08-25 | Pratt & Whitney Canada Corp. | Combustor heat shield cooling hole arrangement |
EP3392566A1 (en) * | 2017-04-18 | 2018-10-24 | United Technologies Corporation | Combustor panel cooling arrangements |
US10605169B2 (en) | 2017-04-18 | 2020-03-31 | United Technologies Corporation | Combustor panel cooling arrangements |
US10816202B2 (en) | 2017-11-28 | 2020-10-27 | General Electric Company | Combustor liner for a gas turbine engine and an associated method thereof |
US11415321B2 (en) | 2017-11-28 | 2022-08-16 | General Electric Company | Combustor liner for a gas turbine engine and an associated method thereof |
US11029027B2 (en) * | 2018-10-03 | 2021-06-08 | Raytheon Technologies Corporation | Dilution/effusion hole pattern for thick combustor panels |
Also Published As
Publication number | Publication date |
---|---|
EP3009745A1 (en) | 2016-04-20 |
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Owner name: UNITED TECHNOLOGIES CORPORATION, CONNECTICUT Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:HOKE, JAMES B.;KWOKA, DAVID;REEL/FRAME:036807/0780 Effective date: 20141016 |
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