Title:
FACETED DOME ASSEMBLIES FOR GAS TURBINE ENGINE COMBUSTORS
Kind Code:
A1


Abstract:
A dome assembly for a gas turbine engine includes a plurality of faceted segments and a plurality of openings. The plurality of faceted segments are coupled together to form a dome structure that is configured to be disposed between an inner liner and an outer liner that circumscribes the inner liner. The plurality of openings are each opening formed within a respective faceted segment, and are configured to at least partially house an atomizer therein. Each faceted segment is at least substantially flat.



Inventors:
Pardington, Ronald B. (Gilbert, AZ, US)
Kettinger, Jonathan N. (Litchfield Park, AZ, US)
Application Number:
11/867195
Publication Date:
04/09/2009
Filing Date:
10/04/2007
Assignee:
HONEYWELL INTERNATIONAL, INC. (Morristown, NJ, US)
Primary Class:
International Classes:
F23R3/42
View Patent Images:
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Primary Examiner:
KIM, TAE JUN
Attorney, Agent or Firm:
HONEYWELL/LKGLOBAL (Patent Services 115 Tabor Road P.O.Box 249, MORRIS PLAINS, NJ, 07950, US)
Claims:
What is claimed is:

1. A dome assembly for a gas turbine engine, the dome assembly comprising: a plurality of faceted segments coupled together to form a dome structure that is configured to be disposed between an inner liner and an outer liner that circumscribes the inner liner; and a plurality of openings, each opening formed within a respective faceted segment and configured to at least partially house an atomizer therein; wherein each faceted segment is at least substantially flat.

2. The dome assembly of claim 1, wherein: each faceted segment is configured to have a heat shield disposed proximate thereto; and a gap is formed between the faceted segment and at least a portion of the heat shield disposed proximate thereto, the gap being at least substantially uniform in size.

3. The dome assembly of claim 2, wherein each faceted segment further comprises a blocking mechanism configured to minimize rotation of the heat shield disposed proximate thereto.

4. The dome assembly of claim 3, wherein each heat shield has a plurality of mating slots, and the blocking mechanism for each faceted segment comprises three clocking tabs, each clocking tab configured to be disposed against one of the plurality of mating slots of a respective heat shield.

5. A combustor for a gas turbine engine, comprising: an inner liner; an outer liner circumscribing the inner liner; and a dome assembly coupled to and disposed between the inner liner and the outer liner, the dome assembly comprising a plurality of faceted segments, each faceted segment being at least substantially flat.

6. The combustor of claim 5, wherein each faceted segment includes an opening, the opening configured to allow an atomizer to be placed at least partially within the opening.

7. The combustor of claim 5, wherein the dome assembly further comprises: a first wall coupling the plurality of faceted segments to the inner liner; and a second wall coupling the plurality of faceted segments to the outer liner.

8. The combustor of claim 5, further comprising: a heat shield unit disposed proximate one of the faceted segments.

9. The combustor of claim 8, wherein the heat shield unit comprises a plurality of heat shields, each heat shield disposed proximate a respective faceted segment.

10. The combustor of claim 9, wherein a gap is formed between the dome assembly and a portion of the heat shield unit, the gap being at least substantially uniform in size.

11. The combustor of claim 9, wherein each faceted segment further comprises a blocking mechanism configured to minimize rotation of the heat shield disposed proximate thereto.

12. The combustor of claim 11, wherein each heat shield has a plurality of mating slots, and the blocking mechanism for each faceted segment comprises three clocking tabs, each clocking tab configured to be disposed against one of the plurality of mating slots of a respective heat shield.

13. The combustor of claim 5, wherein the dome assembly further comprises: a first wall coupling the plurality of faceted segments to the inner liner; and a second wall coupling the plurality of faceted segments to the outer liner.

14. A gas turbine engine, comprising: a compressor having an inlet and an outlet and operable to supply compressed air; a combustor coupled to receive at least a portion of the compressed air from the compressor and operable to supply combusted air, the combustor comprising: an inner liner; an outer liner circumscribing the inner liner; and a dome assembly coupled to and disposed between the inner liner and the outer liner, the dome assembly comprising a plurality of faceted segments, each faceted segment being at least substantially flat; and a turbine coupled to receive the combusted air from the combustor.

15. The gas turbine engine of claim 14, wherein each faceted segment includes an opening, the opening configured to allow an atomizer to be placed at least partially within the opening.

16. The gas turbine engine of claim 14, further comprising: a heat shield unit disposed proximate one of the faceted segments.

17. The gas turbine engine of claim 16, wherein the heat shield unit comprises a plurality of heat shields, each heat shield disposed proximate a respective faceted segment.

18. The gas turbine engine of claim 16, wherein a gap is formed between the dome assembly and a portion of the heat shield unit, the gap being at least substantially uniform in size.

19. The gas turbine engine of claim 17, wherein each faceted segment further comprises a blocking mechanism configured to minimize rotation of the heat shield disposed proximate thereto.

20. The gas turbine engine of claim 19, wherein each heat shield has a plurality of mating slots, and the blocking mechanism for each faceted segment comprises three clocking tabs, each clocking tab configured to be disposed against one of the plurality of mating slots of a respective heat shield.

Description:

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This invention was made with Government support under contract number F33615-03-D-2355-D006 awarded by the United States Air Force. The Government has certain rights in this invention.

TECHNICAL FIELD

The present invention generally relates to gas turbine engines, and more particularly relates to combustors for gas turbine engines.

BACKGROUND OF THE INVENTION

A gas turbine engine may be used to power various types of vehicles and systems. A particular type of gas turbine engine that may be used to power aircraft is a turbofan gas turbine engine. A turbofan gas turbine engine may include, for example, five major sections: a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section.

The fan section typically is positioned at the front or “inlet” section of the engine and includes a fan that induces air from the surrounding environment into the engine and accelerates a fraction of this air toward the compressor section. The remaining fraction of air induced into the fan section is accelerated into and through a bypass plenum and out the exhaust section. The compressor section raises the pressure of the air it receives from the fan section to a relatively high level.

The compressed air from the compressor section then enters the combustor section where a ring of fuel nozzles injects a steady stream of fuel. The injected fuel is ignited by a burner, which significantly increases the energy of the compressed air. The high-energy compressed air from the combustor section then flows into and through the turbine section, causing radially mounted turbine blades to rotate and generate energy.

Due to the high temperatures involved, combustor sections of gas turbine engines typically include a plurality of heat shields, which are often located near a protective dome of the combustor section. Generally, there are gaps formed between the heat shields and the protective dome, for example to allow for the passage of cooling air therethrough and/or to allow for possible thermal expansion. While such heat shields are generally flat, the protective dome disposed nearby is generally a single structure forming an annular ring. As a result, the gaps are typically not uniform throughout the protective dome. This may cause uneven heat distribution and resulting “hot spots”, and/or wear on the combustor and/or components thereof.

Accordingly, it is desirable to provide an improved dome assembly for a gas turbine engine that potentially results in more uniform gaps between the dome assembly and heat shields disposed proximate thereto, and/or that results in a more even heat distribution or in reduced wear on the combustor or components thereof. It is further desirable to provide a combustor for a gas turbine engine with an improved dome assembly that potentially results in more uniform gaps between the dome assembly and heat shields disposed proximate thereto, and/or that results in a more even heat distribution or in reduced wear on the combustor or components thereof. It is further desirable to provide a gas turbine engine with a combustor having an improved dome assembly that potentially results in more uniform gaps between the dome assembly and heat shields disposed proximate thereto, and/or that results in a more even heat distribution or in reduced wear on the combustor or components thereof. Furthermore, other desirable features and characteristics of the present invention will be apparent from the subsequent detailed description and the appended claims, taken in conjunction with the accompanying drawings and the foregoing technical field and background.

SUMMARY OF THE INVENTION

In accordance with an exemplary embodiment of the present invention, a dome assembly for a gas turbine engine is provided. The dome assembly comprises a plurality of faceted segments and a plurality of openings. The plurality of faceted segments are coupled together to form a dome structure that is configured to be disposed between an inner liner and an outer liner that circumscribes the inner liner. The plurality of openings are each formed within a respective faceted segment, and are configured to at least partially house an atomizer therein. Each faceted segment is at least substantially flat.

In accordance with another exemplary embodiment of the present invention, a combustor for a gas turbine engine is provided. The combustor comprises an inner liner, an outer liner, and a dome assembly. The outer liner circumscribes the inner liner. The dome assembly is coupled to and disposed between the inner liner and the outer liner. The dome assembly comprises a plurality of faceted segments. Each faceted segment is at least substantially flat.

In accordance with a further exemplary embodiment of the present invention, a turbine engine is provided. The turbine engine comprises a compressor, a combustor, and a turbine. The compressor has an inlet and an outlet, and is operable to supply compressed air. The combustor is coupled to receive at least a portion of the compressed air from the compressor, and is operable to supply combusted air. The combustor comprises an inner liner, an outer liner, and a dome assembly. The outer liner circumscribes the inner liner. The dome assembly is coupled to and disposed between the inner liner and the outer liner. The dome assembly comprises a plurality of faceted segments. Each faceted segment is at least substantially flat. The turbine is coupled to receive the combusted air from the combustor.

Other independent features and advantages of the preferred apparatus and methods will become apparent from the following detailed description, taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a simplified cross-sectional side view of an upper portion of an annular multi-spool turbofan gas turbine jet engine in accordance with an exemplary embodiment of the present invention;

FIG. 2 is an end view of a portion of a combustor section of a gas turbine engine, such as the multi-spool turbofan gas turbine jet engine of FIG. 1, in accordance with an exemplary embodiment of the present invention;

FIG. 3 is a close-up, plan view of a portion of the combustor section of FIG. 2; and

FIG. 4 is a simplified cross-sectional view of a portion of the combustor section of FIG. 2, shown with a heat shield disposed near the dome assembly, in accordance with an exemplary embodiment of the present invention.

DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT

Before proceeding with the detailed description, it is to be appreciated that the described embodiment is not limited to use in conjunction with a particular type of turbine engine. Thus, although the present embodiment is, for convenience of explanation, depicted and described as being implemented in a multi-spool turbofan gas turbine jet engine, it will be appreciated that it can be implemented in various other types of turbines engines, and in various other systems and environments.

An exemplary embodiment of an upper portion of an annular multi-spool turbofan gas turbine jet engine 100 is depicted in FIG. 1. The engine 100 includes an intake section 102, a compressor section 104, a combustor section 106, a turbine section 108, and an exhaust section 1 10. The intake section 102 includes a fan 112, which is mounted in a fan case 114. The fan 112 draws air into the intake section 102 and accelerates it. A fraction of the accelerated air exhausted from the fan 112 is directed through a bypass section 116 disposed between the fan case 114 and an engine cowl 118, and provides a forward thrust. The remaining fraction of air exhausted from the fan 112 is directed into the compressor section 104.

In the embodiment of FIG. 1, the compressor section 104 includes two compressors, namely an intermediate pressure compressor 120 and a high pressure compressor 122. However, in other embodiments the number of compressors in the compressor section 104 may vary. In the depicted embodiment, the intermediate pressure compressor 120 raises the pressure of the air directed into it from the fan 112 and directs the compressed air into the high pressure compressor 122. The high pressure compressor 122 compresses the air still further, and directs a majority of the high pressure air into the combustor section 106. In addition, a fraction of the compressed air bypasses the combustor section 106 and is used to cool, among other components, turbine blades in the turbine section 108. In the combustor section 106, which includes an annular combustor 124, the high pressure air is mixed with fuel and combusted. The high-temperature combusted air is then directed into the turbine section 108.

In the embodiment of FIG. 1, the turbine section 108 includes three turbines disposed in axial flow series, namely, a high pressure turbine 126, an intermediate pressure turbine 128, and a low pressure turbine 130. However, it will be appreciated that the number of turbines, and/or the configurations thereof, may vary, as may the number and/or configurations of various other components of the exemplary engine 100. For example, in one alternate embodiment, the compressor section 104 includes a single compressor, and the turbine section 108 includes two turbines. It will be similarly appreciated that the number of compressors in the compressor section 104, the number of turbines in the turbine section 108, or both, may vary in other embodiments.

In the embodiment depicted in FIG. 1, the high-temperature combusted air from the combustor section 106 expands through each turbine, causing it to rotate. The air is then exhausted through a propulsion nozzle 132 disposed in the exhaust section 110, providing additional forward thrust. As the turbines rotate, each drives equipment in the engine 100 via concentrically disposed shafts or spools. Specifically, the high pressure turbine 126 drives the high pressure compressor 122 via a high pressure spool 134, the intermediate pressure turbine 128 drives the intermediate pressure compressor 120 via an intermediate pressure spool 136, and the low pressure turbine 130 drives the fan 112 via a low pressure spool 138.

FIG. 2 is an end view of a portion of a combustor section 200 of a gas turbine engine, such as the multi-spool turbofan gas turbine jet engine 100 of FIG. 1, that includes a dome assembly 202, shown along with a portion of an inner liner 204 and an outer liner 206 of the combustor section 200, in accordance with an exemplary embodiment of the present invention. As shown in FIG. 2, the outer liner 206 circumscribes the inner liner 204. The dome assembly 202 is coupled to and disposed between the inner liner 204 and the outer liner 206.

As depicted in FIG. 2, the dome assembly 202 comprises a plurality of faceted segments 208, a first wall 209, and a second wall 211. The plurality of faceted segments 208 are coupled together to form a dome structure. Each faceted segment 208 is at least substantially flat. Each faceted segment 208 includes an opening 210 to allow an atomizer (not depicted in FIG. 2) to be disposed at least partially within, to provide access for the atomizer to a combustion chamber (not depicted in FIG. 2) of the combustor section 200. As will be described in greater detail below in connection with FIGS. 2 and 3, each faceted segment 208 is also configured to have a heat shield disposed nearby, and each faceted segment 208 includes one or more blocking mechanisms 212 configured to at least minimize rotation of the heat shield, in a preferred embodiment.

In the depicted embodiment, the first wall 209 couples each faceted segment 208 to the inner liner 204. Likewise, the second wall 211 couples each faceted segment 208 to the outer liner 206. For example, in one preferred embodiment, the first wall 209 connects each faceted segment 208 to the inner liner 204, and the second wall 211 connects each faceted segment 208 to the outer liner 206. Also in a preferred embodiment, the first and second walls 209, 211 are at least partially conical in shape. In addition, in one preferred embodiment, the faceted segments 208 are welded together, and are also each welded to the first and second walls 209, 211, which in turn are welded to the inner and outer liners 204, 206.

FIG. 3 is a close-up, top view of a portion of the combustor section 200 of FIG. 2, showing a portion of the dome assembly 202, including some of the faceted segments 208 thereof and a portion of the first and second walls 209, 211, along with a portion of the inner liner 204 and outer liner 206 of the combustor section 200, in accordance with an exemplary embodiment of the present invention. In a preferred embodiment, each faceted segment 208 has a respective heat shield 314 disposed nearby. In FIG. 3, each heat shield 314 is depicted as being disposed at least substantially underneath a respective faceted segment 208. In a preferred embodiment, each heat shield 314 comprises a full coverage heat shield, as is commonly referred to in the industry, and is at least substantially flat in shape.

As shown in FIG. 3 as well as in FIG. 2, the blocking mechanisms 212 help to minimize, and preferably at least substantially prevent, rotation of the heat shields 314 proximate the respective faceted segments 208, in the depicted embodiment. In this embodiment, each faceted segment 208 includes three blocking mechanisms 212, and each blocking mechanism 212 comprises a clocking tab. Each clocking tab is configured to be disposed against a respective slot (not depicted in FIGS. 2 and 3) of a respective heat shield 314. In various other embodiments, one or more pins and/or other devices may be formed as part of, and/or may be disposed against, one of the heat shields 314 and/or a respective faceted segment 208, in order to minimize rotation of the heat shield 314.

Also as shown in FIGS. 2 and 3, in a preferred embodiment, each of the individual, flat, faceted segments 208 are joined and connected together to form an annular ring for the dome assembly 202. In the depicted embodiment, there are twenty-two faceted segments 208 in the dome assembly 202 (depicted in FIG. 2), corresponding with twenty-two atomizers for the combustor section 200. The number of faceted segments 208 and/or atomizers may vary. However, in a preferred embodiment, the number of faceted segments 208 in the dome assembly 202 is equal to the number of atomizers in the combustor section 200. Also in a preferred embodiment, and as will be described in greater detail below in connection with FIG. 4, each heat shield 314 is at least partially mounted adjacent to a respective faceted segment 208, with a small and uniform gap existing between each heat shield 314 and a respective faceted segment 208.

FIG. 4 is a simplified cross-sectional view of a portion of the combustor section 200 of FIG. 2, shown with a representative heat shield 314 disposed near the dome assembly 202, and also shown along with a portion of the inner and outer liners 204, 206, in accordance with an exemplary embodiment of the present invention. As shown in FIG. 4, in a preferred embodiment, the heat shield 314 is preferably mounted at least in part adjacent to a respective faceted segment 208 of the dome assembly 202 via a mounting piece 416. Also in this preferred embodiment, the heat shield 314 is mounted such that a gap 418 is formed between the heat shield 314 and the respective faceted segment 208. The gap 418 allows for passage of cooling air therethrough, and also allows for thermal growth of the respective heat shield 314 and/or faceted segment 208.

Preferably, the gap 418 is uniform across the faceted segment 208. Additionally, preferably the various gaps 418 between the various heat shields 314 and their respective faceted segments 208 are uniform not only across each faceted segment 208, but are also uniform with respect to one another, for example being of a common size and shape across the entire dome assembly 202. The uniform nature of the gaps 418 is made possible due to the structure of the dome assembly 202 as described above in connection with the exemplary embodiments of FIGS. 2 and 3, and in particular due to the use of flat faceted segments 208. The uniform nature of the gaps 418 potentially results in a more even heat distribution in the combustor section 200, a reduction or elimination in “hot spots”, and reduced wear on the combustor section 200 and various components thereof

Accordingly, an improved dome assembly 202 has been provided for a gas turbine engine that potentially results in more uniform gaps 418 between the dome assembly 202 and heat shields 314 disposed proximate thereto, and that potentially results in a more even heat distribution, and in reduced wear on the combustor and components thereof. In addition, a combustor section 200 for a gas turbine engine has been provided that includes such an improved dome assembly 202. Further, a gas turbine engine 100 has been provided that includes a combustor section 200 having such an improved dome assembly 202.

While at least one exemplary embodiment has been presented in the foregoing detailed description of the invention, it should be appreciated that a vast number of variations exist. It should also be appreciated that the exemplary embodiment or exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration of the invention in any way. Rather, the foregoing detailed description will provide those skilled in the art with a convenient road map for implementing an exemplary embodiment of the invention, it being understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope of the invention as set forth in the appended claims and their legal equivalents.