US20120304660A1

US20120304660A1 - Turbomachine combustors having different flow paths

Info

Publication number: US20120304660A1
Application number: US13/153,550
Authority: US
Inventors: Daniel B. Kupratis
Original assignee: United Technologies Corp
Current assignee: Raytheon Technologies Corp
Priority date: 2011-06-06
Filing date: 2011-06-06
Publication date: 2012-12-06
Also published as: EP2532857B1; EP2532857A1

Abstract

An example turbomachine assembly includes a first combustor configured to combust fuel and compressed air and a second combustor configured to combust fuel and compressed air. Flow moves through the first combustor in a first direction and flow moves though the second combustor in a second direction different than the first direction. The first combustor is axially spaced from the second combustor.

Description

FIELD OF INVENTION

This disclosure relates generally to a combustor section of a turbomachine and, more particularly to a combustor section that includes combustors having different flow paths.

BACKGROUND

Turbomachines include multiple sections. A gas turbine engine is a type of turbomachine. Gas turbine engines, and other turbomachines, may include a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section. During operation, air moves into the engine through the fan section. Rotors in the compressor section rotate to compress the air, which is then mixed with fuel and combusted in the combustor section. The products of combustion are expanded to rotatably drive rotors in the turbine section. The rotating turbine section rotatably drives the fan and compressor sections. Flow moves axially through the engine from the fan section to the exhaust section.
High temperatures can damage turbomachine components. Maintaining a desirable overall pressure ratio while avoiding damaging high temperatures is often difficult. For example, incorporating features to increase and maintain the overall pressure ratio can increase the size of the turbomachine.

SUMMARY

An example turbomachine assembly includes a first combustor configured to combust fuel and compressed air and a second combustor configured to combust fuel and compressed air. Flow moves through the first combustor in a first direction and flow moves though the second combustor in a second direction different than the first direction. The first combustor is axially spaced from the second combustor.
An example turbomachine assembly includes a compressor section, a first combustor section that receives a flow of compressed fluid from the compressor section, and a first turbine section that receives a product of combustion from the first combustor section. The example turbomachine assembly further includes a second combustor section that receives a flow of compressed fluid from the first turbine section, and a second turbine section that receives a product of combustion from the second combustor section. Flow moves through the first combustor in a first direction and flow moves though the second combustor in a second direction different than the first direction.
An example method of operating a turbomachine includes moving a mixture of fuel and compressed air through a first combustor section in a first direction. The method also includes moving a mixture of fuel and compressed air through a second combustor section in a second direction different than the first direction.

DESCRIPTION OF THE FIGURES

The various features and advantages of the disclosed examples will become apparent to those skilled in the art from the detailed description. The figures that accompany the detailed description can be briefly described as follows:

FIG. 1 shows a section view of an example gas turbine engine.

FIG. 2 shows a close up view of a portion of a combustion section of the FIG. 1 engine.

FIG. 3 shows a simplified sectional view of the FIG. 1 engine at line 3-3 in FIG. 2.

FIG. 4 shows a simplified sectional view of the FIG. 1 engine at line 4-4 in FIG. 2.

FIG. 5 shows a fuel delivery manifold of the FIG. 1 engine.

DETAILED DESCRIPTION

Referring to FIG. 1, an example turbomachine, such as a gas turbine engine 10, is circumferentially disposed about an axis A. The gas turbine engine 10 includes a fan section 12, a compression section 14, a combustion section 16, a turbine section 18, and an exhaust section 20. Other example turbomachines may include more or fewer sections.
In this example, the fan section 12 includes a first fan 22 and a second fan 24. The example compression section 14 includes a low-pressure compressor 26 and a high-pressure compressor 28. The example combustion section 16 includes a first combustor 30 and a second combustor 32. Also, the example turbine section 18 includes a high-pressure turbine 34, an intermediate-pressure turbine 36, and low-pressure turbine 38.
The example gas turbine engine 10 includes a first spool 40, a second spool 42, and a third spool 44. The first spool 40 rotatably couples the first fan 22 to the low-pressure turbine 38. In this example, a fan drive gear system 46 is used to rotatably link the first fan 22 with the first spool 40. The second spool 42 rotatably couples the second fan 24 and the low-pressure compressor 26 to the intermediate-pressure turbine 36. The third spool 44 rotatably couples the high-pressure compressor 28 to the high-pressure turbine 34.
Except where indicated otherwise, the examples described in this disclosure are not limited to the three-spool gas turbine architecture described, and may be used in other architectures, such as a single-spool axial design, a two-spool axial design, and still other architectures. That is, there are various types of gas turbine engines, and other turbomachines, that can benefit from the examples disclosed herein.
During operation, a flow of air enters the engine 10 through the fan section 12. The flow of air is compressed in the compression section 14. The compressed flow is then mixed with fuel and combusted in the combustion section 16. The products of combustion are then expanded to rotate rotors of the turbine section 18. The rotors rotatably drive the first fan 22, the second fan 24, the low-pressure compressor 26, and the high-pressure compressor 28. Once expanded, flow exits the engine 10 at the exhaust section 20. As can be appreciated, flow generally moves through the engine 10 in a direction F.
Referring to FIG. 2 with continuing reference to FIG. 1, the first combustor 30 includes a nozzle 48 a and a first combustion chamber 50. During operation, a mixture of fuel and compressed air is injected from the nozzle 48 a into the first combustion chamber 50. The nozzle 48 a sprays the mixture of fuel and compressed air into the first combustion chamber 50 in a first direction D₁.
Flow moves through the first combustion chamber 50 in the first direction D₁, which is opposite the general direction of flow F through the engine 10. The flow includes the products of combustion. In some examples, the first combustor 30 is considered a reverse flow combustor because the flow moves through the first combustion chamber 50 in a direction opposite the general direction of flow F through the engine 10.
Once combusted, flow moves from the first combustion chamber 50 through an outlet conduit 52 to drive the high-pressure turbine 34. Flow moves through the outlet conduit 52 in a direction opposite the first direction D₁.
The second combustor 32 includes a nozzle 54 a and a second combustion chamber 56. During operation, a mixture of fuel and compressed air is injected from the nozzle 54 a into the second combustion chamber 56 of the second combustor 32. Notably, the nozzle 54 a sprays the mixture of fuel and compressed air into the second combustion chamber 56 in a second direction D₂, which is the same as the general direction of flow F through the engine 10.
Flow moves from the second combustion chamber 56 through an outlet conduit 58 to drive the intermediate-pressure turbine 36. The flow includes the products of combustion within the second combustor 32. As can be appreciated, flow moves through the first combustion chamber 50 and the second combustion chamber 56 in opposite directions.
In this example, a single fuel manifold 60 is used to control the flow of fuel to both the nozzle 48 a of the first combustor 30 and the nozzle 54 a of the second combustor 32. The single fuel manifold 60 receives fuel from a fuel supply 62. Arranging the nozzle 48 a to be axially adjacent the nozzle 54 a facilitates this communication of fuel from the single fuel manifold 60.
A swirl air conduit 72 swirls and delivers pressurized air from the high-pressure compressor 28 to the combustor 30. Another swirl air conduit 74 swirls and delivers pressurized air from high-pressure compressor 28 and combustor housing to the fuel nozzles 54 a. The air delivered by the conduits 72 and 74 facilitates atomization and cools the nozzles 48 a and 54 a to prevent coking.
Referring now to FIGS. 3-5 with continuing reference to FIG. 2, the example gas turbine engine 10 includes six of the first combustors 30 and six of the second combustors 32. The first combustors 30 and the second combustors 32 are circumferentially distributed about the axis A. The first combustors 30 are axially in front of the second combustors 32 relative to the direction of flow F through the engine 10.
In this example, the single fuel manifold 60 delivers fuel to all twelve of the combustors 30 and 32 within the engine 10. Notably, the fuel manifold 60 is a helix that is centered radially about the axis A, and centered axially relative to the first combustors 30 and the second combustors 32. The example fuel manifold 60 is configured to deliver fuel to the fuel nozzles 54 a-54 f of the second combustors 32 prior to the fuel nozzles 48 a-48 f of the first combustors 30. Notably, the fuel manifold 60 shown in FIG. 5 has been axially stretched for clarity.
The fuel manifold 60 includes a valve 70 that can be closed to block fuel flow to the first combustors 30. When the valve 70 is closed, no fuel is combusted in first combustors 30. When no fuel is combusted in the first combustors 30, the high-pressure turbine 34 windmills, which causes the high-pressure compressor 28 to rotate at an insufficient speed to generate a significant increase in the temperature or pressure of the air. In one example, the valve 70 is closed when the engine 10 (FIG. 1) is propelling an aircraft during takeoff on a hot day. In another example, the valve 70 is opened when the engine 10 is propelling an aircraft during cruise. Opening the valve 70 delivers fuel to the first combustors 30, which drive the high-pressure turbine 34 to maximize the compression pressure of high-pressure compressor 28. The power of the engine 10 is thus controlled in part by controlling the fuel supply 62.
In this example, the products of combustion moving through the outlet conduit 52 drive the third spool 44, and the products of combustion moving through the outlet conduit 58 drive the second spool 42. The example first combustor 30 and the second combustor 32 are thus configured to drive different spools.
The spools 40, 42, and 44 are held by bearing arrangements in a known manner. Sealing of the hot gas path from the shafts 40, 42, and 44 and the bearing arrangements is also executed in a known manner.
In some examples, the nozzles 54 a-54 f introduce air and uncombusted fuel from the outlet conduit 52 into the second combustion chamber 56. In such examples, the nozzle 54 a is able to introduce fuel into the second combustion chamber 56 even if the nozzle 54 a is not connected to the single fuel manifold 60. That is, the first combustor 30 may not combust all of the fuel introduced by the nozzles 48 a-48 f. This fuel is not wasted, because the second combustor 32 is able to combust this fuel.
The example rotors in the turbine section 18 include an array of blades 64 that rotate together with the third shaft 44 about the axis A. The example high-pressure turbine 34 also includes two axial rows of vanes 66 that do not rotate.
In this example, the vanes 66 a in the first row are adjustable vanes. That is, the vanes 66 a can be pivoted, or otherwise actuated, to control flow through the high-pressure turbine 34. Controlling flow through the high-pressure turbine 34 regulates rotation of the high-pressure compressor 28, which is rotatably coupled to the high-pressure turbine 34. In this example, the vanes 66 a are pivoted about a radially extending axis. Other examples include other ways of manipulating the vanes 66 a to control flow through the high-pressure turbine 34.
A person having skill in this art and the benefit of this disclosure would understand how to vary the position of the vanes 66 and 66 a to control flow through the high-pressure turbine 34. United States Published Application No. 20090016871, which is incorporated herein by reference, discloses an example configuration having variable vanes that control flow through a gas turbine engine. Other examples configurations of variable vanes could be used.
Although only the high-pressure turbine 34 is described as having variable vanes 66 a, the intermediate-pressure turbine 36 and the low-pressure turbine 38 could also include variable vanes. That is, any area of the turbine section 18 could include variable vanes to control flow through a desired area of the turbine section 18. Also, the blades 64 in turbine section 18, or other components, could be varied to control flow through a desired area of the turbine section 18.
The example engine 10 may include a clutch 68. The low-pressure compressor 26 and the high-pressure compressor 28 are selectively coupled in rotation together via the clutch 68. In some examples, coupling the low-pressure compressor 26 together with the high-pressure compressor 28 slows the rotational speed of the low-pressure compressor 26 to desirably reduce the pressure ratio of the high-pressure compressor 28.
Features of some of the disclosed examples include a back-to-back arrangement of the first combustor and the second combustor, which decreases the axial length of the engine over other designs. The back-to-back arrangement also enables a delivering fuel to the combustors using a single fuel manifold. Another feature of some of the disclosed examples includes the variable turbines, which can act as chokepoint for flow through the engine.
The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. Thus, the scope of legal protection given to this disclosure can only be determined by studying the following claims.

Claims

1. A turbomachine assembly, comprising:

a first combustor configured to combust fuel and compressed air; and

a second combustor configured to combust fuel and compressed air, wherein flow moves through the first combustor in a first direction and flow moves though the second combustor in a second direction different than the first direction, the first combustor axially spaced from the second combustor.

2. The turbomachine assembly of claim 1, wherein products of combustion from the first combustor are expanded to drive a first spool, and products of combustion from the second combustor are expanded to drive a second spool different than the first spool.

3. The turbomachine assembly of claim 1, including a fuel manifold configured to supply fuel to both the first combustor and the second combustor.

4. The turbomachine assembly of claim 1, wherein the first direction is opposite a general direction of flow through the turbomachine, the general direction of flow extending from a fan section of the turbomachine to an exhaust section of the turbomachine.

5. The turbomachine assembly of claim 1, wherein the first combustor comprises a plurality of reverse flow combustors annularly distributed about an axis of the turbomachine.

6. The turbomachine assembly of claim 1, wherein the turbomachine is a gas turbine engine.

7. The turbomachine assembly of claim 1, wherein a compressor directly provides compressed air to the first combustor, and a turbine section directly provides compressed air to the second combustor.

8. The turbomachine assembly of claim 7, wherein the turbine section includes variable vanes.

9. A turbomachine assembly, comprising:

a compressor section;

a first combustor section that receives a flow of compressed fluid from the compressor section;

a first turbine section that receives a product of combustion from the first combustor section;

a second combustor section that receives a flow of compressed fluid from the first turbine section; and

a second turbine section that receives a product of combustion from the second combustor section, wherein flow moves through the first combustor in a first direction and flow moves though the second combustor in a second direction different than the first direction.

10. The turbomachine assembly of claim 9, wherein the first turbine section is a higher pressure turbine than the second turbine section.

11. The turbomachine assembly of claim 9, wherein the first turbine section drives a first spool, and the second turbine section drives a second spool different than the first spool.

12. The turbomachine assembly of claim 11, wherein the first spool is rotatably coupled to a high-pressure compressor of the compressor section, and the second spool is rotatably coupled to a low-pressure compressor of the compressor section.

13. The turbomachine assembly of claim 12, including a third turbine section that is rotatably coupled to a fan with a third spool.

14. The turbomachine assembly of claim 9, including a common manifold configured to supply fuel to both the first combustor and the second combustor.

15. The turbomachine assembly of claim 9, wherein the first direction is opposite the second direction.

16. A method of operating a turbomachine, comprising:

moving a mixture of fuel and compressed air through a first combustor section in a first direction; and

moving a mixture of fuel and compressed air through a second combustor section in a second direction different than the first direction.

17. The method of claim 16, communicating products of combustion from the first combustor section to a turbine section and communicating compressed air from the turbine section to the second combustor section.

18. The method of claim 16, communicating fuel to the first combustor section and the second combustor section using a common manifold.

19. The method of claim 16, driving a first turbine section with the first combustor section and driving a second turbine section with the second combustor section, the first turbine section and the second turbine section configured to drive different spools.