WO2007102807A1

WO2007102807A1 - Angled flow annular combustor for turbine engine

Info

Publication number: WO2007102807A1
Application number: PCT/US2006/007898
Authority: WO
Inventors: Steven W. Burd; Albert K. Cheung
Original assignee: United Technologies Corporation
Priority date: 2006-03-06
Filing date: 2006-03-06
Publication date: 2007-09-13

Abstract

An annular combustor for a turbine engine includes a plurality of nozzles (32) that are configured to provide increased efficiency and to permit the use of simplified turbine vanes (36). The plurality of nozzles are angled such that the flow into the combustion chamber has an angular component relative to an axis of the combustor.

Description

ANGLED FLOW ANNULAR COMBUSTOR FOR TURBINE ENGINE

BACKGROUND OF THE INVENTION

The present invention relates to a turbine engine, and more particularly to a combustor for a tip turbine engine.

An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan and a low pressure compressor, a middle core engine, and an aft low pressure turbine, all located along a common longitudinal axis. A high pressure compressor and a high pressure turbine of the core engine are interconnected by a high spool shaft. The high pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in a combustor, where it is ignited to form a high energy gas stream. The gas stream flows axially aft to rotatably drive the high pressure turbine, which rotatably drives the high pressure compressor via the high spool shaft. The gas stream leaving the high pressure turbine is expanded through the low pressure turbine, which rotatably drives the bypass fan and low pressure compressor via a low spool shaft.

Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship results in a relatively complicated elongated engine structure of considerable length relative to the engine diameter.

This elongated shape may complicate or prevent packaging of the engine into particular applications.

A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines locate an axial compressor forward of a bypass fan which includes hollow fan blades that receive airflow from the axial compressor therethrough such that the hollow fan blades operate as a centrifugal compressor. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor, where it is ignited to form a high energy gas stream which drives the turbine that is integrated onto the tips of the hollow bypass fan blades for rotation therewith as generally disclosed in U.S. Patent Application Publication Nos.: 20030192303; 20030192304; and 20040025490. The tip turbine engine provides a thrust-to-weight ratio equivalent to or greater than conventional turbofan engines of the same class, but within a package of significantly shorter length.

SUMMARY OF THE INVENTION

The present invention provides a turbine engine annular combustor that includes a plurality of fuel nozzles that are configured to provide increased efficiency and to permit the use of simplified turbine vanes. The plurality of nozzles are angled such that the flow into the combustion chamber has an angular component relative to an axis of the combustor. The angled fuel flow is in a direction that is not parallel to the engine centerline and not normal to a plane perpendicular to the engine centerline.

The angled flow allows for simplification of the combustor design. Since the flow exiting the front of the combustor exits with significant momentum, the fuel-air mixtures distribute transversely or circumferentially more effectively than in a conventional design. The same pattern factor as found with a conventional combustor design can be achieved with fewer fuel injectors and/or air swirlers or be improved with the same or less nozzle and/or air swirler count. Additionally, residence time of the fuel in the combustor is increased by the angled flow. This permits the combustor to be axially shorter. The angled flow also enhanced shearing of the flow or a modified recirculation zone that can promote enhanced mixing for reduced emissions or improved combustion efficiency.

Angled flow permits simplification of the turbine vane. Since the flow received by the turbine locally and spatially has a mean velocity component or sense that is substantially transverse in its direction, the amount of turning required by the turbine vane is reduced. As a result, the turbine vane itself can be simplified is shape/geometry, length and complexity. This simplification can also provide for reduced vane stage losses or count, netting other system performance benefits. Under special circumstances, the turning provided by the combustor may preclude the need for a vane stage.

BRIEF DESCRIPTION OF THE DRAWINGS

Other advantages of the present invention can be understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:

Figure 1 is a partial sectional perspective view of a tip turbine engine according to the present invention.

Figure 2 is a longitudinal sectional view of the tip turbine engine of Figure 1 taken along an engine centerline.

Figure 3 is an enlarged view of the diffuser, combustor and turbine area of Figure 2. Figure 4 is an interior perspective partial view of the combustor of Figure 3.

Figure 5 is a plan view of the combustor and the turbine vanes of Figure 3.

Figure 6 is an interior perspective partial view of an alternative combustor.

Figure 7 is a plan view of the combustor of Figure 6 in front of the turbine vanes of Figure 2.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Figure 1 illustrates a general perspective partial sectional view of a tip turbine engine type gas turbine engine 10. Although the invention is shown as used in a tip turbine engine, it could also be used in conventional gas turbine engines. The engine 10 includes an outer nacelle 12, a rotationally fixed static outer support structure 14 and a rotationally fixed static inner support structure 16. A plurality of fan inlet guide vanes 18 are mounted between the static outer support structure 14 and the static inner support structure 16. Each inlet guide vane preferably includes a variable trailing edge 18 A. A nosecone 20 is preferably located along the engine centerline A to improve airflow into an axial compressor 22, which is mounted about the engine centerline A behind the nosecone 20.

A fan-turbine rotor assembly 24 is mounted for rotation about the engine centerline A aft of the axial compressor 22. The fan-turbine rotor assembly 24 includes a plurality of hollow fan blades 28 to provide internal, centrifugal compression of the compressed airflow from the axial compressor 22 for distribution to an annular combustor 30 located within the rotationally fixed static outer support structure 14. A turbine 32 includes a plurality of tip turbine blades 34 (two stages shown) which rotatably drive the hollow fan blades 28 relative a plurality of tip turbine vanes 36 which extend radially inwardly from the rotationally fixed static outer support structure 14. The annular combustor 30 is disposed axially forward of the turbine 32 and communicates with the turbine 32. Referring to Figure 2, the rotationally fixed static inner support structure 16 includes a splitter 40, a static inner support housing 42 and a static outer support housing 44 located coaxial to said engine centerline A.

The axial compressor 22 includes the axial compressor rotor 46, from which a plurality of compressor blades 52 extend radially outwardly, and a fixed compressor case 50. A plurality of compressor vanes 54 extend radially inwardly from the compressor case 50 between stages of the compressor blades 52. The compressor blades 52 and compressor vanes 54 are arranged circumferentially about the axial compressor rotor 46 in stages (three stages of compressor blades 52 and compressor vanes 54 are shown in this example). The axial compressor rotor 46 is mounted for rotation upon the static inner support housing 42 through a forward bearing assembly 68 and an aft bearing assembly 62.

The fan-turbine rotor assembly 24 includes a fan hub 64 that supports a plurality of the hollow fan blades 28. Each fan blade 28 includes an inducer section 66, a hollow fan blade section 72 and a diffuser section 74. The inducer section 66 receives airflow from the axial compressor 22 generally parallel to the engine centerline A and turns the airflow from an axial airflow direction toward a radial airflow direction. The airflow is radially communicated through a core airflow passage 80 within the fan blade section 72 which acts as a compressor chamber where the airflow is centrifugally compressed. From the core airflow passage 80, the airflow is diffused and turned once again by the diffuser section 74 toward the annular combustor 30.

Generally the airflow through the core airflow passage 80 is core airflow directed by the diffuser section 74 axially forward toward the combustor 30.

Minimal amounts of airflow may be directed radially outwardly from the diffuser section 74 through the tip turbine blades 34 (paths not shown) to cool the tip turbine blades 34. This cooling airflow is then discharged through radially outer ends of the tip turbine blades 34 and then into the combustor 30. However, at least substantially all of the airflow is core airflow directed by the diffuser section 74 toward the combustor 30. As used herein, "core airflow" is airflow that flows to the combustor 30.

A plurality of fuel injectors 82, or "nozzles," (one shown) supply fuel to the combustor 30. Fuel is delivered to the fuel injectors 82 from a fuel manifold or ring

84 extending circumferentially about the engine centerline A. As will be explained in more detail subsequently, the fuel injectors 82 are all canted relative to the engine centerline A.

A gearbox assembly 90 aft of the fan-turbine rotor assembly 24 provides a speed increase between the fan-turbine rotor assembly 24 and the axial compressor 22.

The annular combustor 30, fuel injector 82 and fuel manifold or ring 84 are shown in greater detail in Figure 3. The annular combustor 30 includes an annular combustion chamber 112 defined between an annular inner combustion chamber wall 114 and annular outer combustion chamber wall 116 and having the engine centerline A (Figure 2) as its axis. A bulkhead 118 at a forward end of the combustion chamber 112 has mounted thereto a fuel injector 82, which directs fuel into the combustion chamber 112 along its fuel injector axis. The fuel injector 82 may be coupled and mated with a swirler or air guide 85 (in the embodiment shown) to provide air to mix with the fuel from the injector. The bulkhead 118 includes a plurality of altematingly angled panels 120, 122. The fuel injector 82 and air guide 85 are mounted over an opening through the angled panel 122. The angled panel 122 extends at an acute angle relative to the engine centerline A, and the plane defined by the panel 122 extends at an acute angle relative to a plane perpendicular to the engine centerline A. This positions the fuel injector 82 at an acute angle relative to the engine centerline A and relative to the plane perpendicular to the engine centerline A. In the embodiment shown, each fuel injector axis is contained in plane tangent to a cylinder having the engine centerline A as its axis. It may also be desirable to change a pitch of the fuel injector axis toward or away from the engine centerline A.

The annular inner and outer combustion chamber walls 114, 116 and the bulkhead 118 are perforated (not shown in Figure 3) to permit core airflow into the combustion chamber 112. An annular diffuser case 128 substantially encloses the annular inner combustion chamber wall 114, outer combustion chamber wall 116 and the bulkhead 118.

An interior perspective partial view of the combustor 30 is shown in Figure 4, wherein the bulkhead 118, annular inner combustion chamber wall 114 and fuel injector 82 and associated air guide are shown. The angled panels 120, 122 shown in Figure 4 alternate about the engine centerline A (Figure 2) to form the bulkhead 118. A fuel injector 82 is mounted on the outside of each angled panel 122 over an opening through the angled panel 122. Figure 5 is a plan view of a plurality of fuel injectors 82 mounted to the bulkhead 118 in the manner described above. In Figure 5 it can be seen that each fuel injector 82 includes an air guide 85 (or passage or swirler) and a concentric fuel nozzle tip 83. The air guide 85 supplies core airflow into the combustor 30 at the angle of the fuel injector 82 and potential rotation or swirl. The fuel nozzle tip 83 receives fuel from the fuel manifold or ring 84 (Figure 3) and supplies the fuel to the combustor 30 at the angle of the fuel injector 82.

As shown, the fuel injectors 82 are all directed toward the vanes 36 at an acute angle relative to the engine centeiiine A and at an acute angle relative to a plane P perpendicular to the engine centerline A. This provides the flow within the annular combustor 30 with an angular, transverse or circumferential component. As is apparent from Figure 5, the residence time of the fuel in the combustor 30 is increased for a given combustor 30 axial length. As a result, the axial length of the combustor 30 can be reduced. The angled fuel injectors 82 provide enhanced shearing of the flow or a modified recirculation zone that can promote enhanced mixing for reduced emissions or improved combustion efficiency. Providing angled flow locally or in prescribed spatial patterns relative to the turbine vanes also allows for simplification of the vanes 36, since less turning of the flow is required by the vane 36. As a result, the losses from the vanes can be reduced and the number of vanes 36 can be reduced. In certain configurations, the initial stage of vanes 36 may even be eliminated. The hole patterns in the annular inner combustion chamber wall 114 may be positioned such that the dilution flow into the combustion chamber 112 promotes the angled flow (or alternatively, attenuate the angled flow if it achieves a desired mixing). Figures 6 and 7 illustrate an alternative combustor 30a that could be used in the turbine engine of Figures 1-3. The combustor 30a includes a generally annular bulkhead 118a through which a plurality of fuel injectors 82a (each having air guides 85a and fuel nozzle tips 83a) are mounted. The bulkhead 118a provides a generally annular inner surface through which the fuel injectors 82a extend at an acute angle relative to the engine centerline A and at an acute angle relative to a plane P perpendicular to the engine centerline A. The fuel injector 82a includes a shield or housing 86a protruding into the annular combustion chamber 112a at an acute angle relative to the bulkhead 118a. The fuel is provided at the angle indicated in Figure 7, similar to Figures 4-5. In operation, referring to Figure 2, air enters the axial compressor 22, where it is compressed by the three stages of the compressor blades 52 and compressor vanes 54. The compressed air from the axial compressor 22 enters the inducer section 66 in a direction generally parallel to the engine centerline A, and is then turned by the inducer section 66 radially outwardly through the core airflow passage 80 of the hollow fan blades 28. The airflow is further compressed centrifugally in the hollow fan blades 28 by rotation of the hollow fan blades 28. From the core airflow passage 80, the airflow is turned and diffused axially forward in the engine 10 by diffuser section 74 into the annular combustor 30. The compressed core airflow from the hollow fan blades 28 then flows radially outwardly and through the annular inner and outer combustion chamber walls 114, 116 and the bulkhead 118 to the combustion chamber 112.

The fuel is injected into the annular combustor 30 where it is mixed with the core airflow and ignited to form a high-energy gas stream. Because the fuel is injected at an angle, the residence time of the fuel in the combustion chamber 112 is increased, improving fuel burn and/or decreasing the required length of the combustor 30.

The high-energy gas stream expands through the turbine vanes 36 and the tip turbine blades 34. The high-energy gas stream rotatably drives the plurality of tip turbine blades 34 mounted about the outer periphery of the fan-turbine rotor assembly 24 to drive the fan-turbine rotor assembly 24, which in turn drives the axial compressor 22 via the gearbox assembly 90.

The fan-turbine rotor assembly 24 discharges fan bypass air axially aft to merge with the core airflow from the turbine 32 in an exhaust case 106. A plurality of exit guide vanes 108 are located between the static outer support housing 44 and the rotationally fixed static outer support structure 14 to guide the combined airflow out of the engine 10 and provide forward thrust. An exhaust mixer 110 mixes the airflow from the tip turbine blades 34 with the bypass airflow through the fan blades In accordance with the provisions of the patent statutes and jurisprudence, exemplary configurations described above are considered to represent a preferred embodiment of the invention. However, it should be noted that the invention can be practiced otherwise than as specifically illustrated and described without departing from its spirit or scope. For example, although the invention is shown as used in a tip turbine engine, the present invention would be beneficial in most or all conventional gas turbine engines.

Claims

1. A combustor for a turbine engine comprising: a wall at least partially defining an annular combustion chamber about a combustor axis; and at least one fuel injector directed along an injector axis, the injector axis disposed at an angle not normal to a plane perpendicular to the combustor axis.

2. The combustor of claim 1 wherein the at least one fuel injector includes a plurality of fuel injectors spaced circumferentially about the combustor axis.

3. The combustor of claim 2 wherein each injector axis of the plurality of fuel injectors is not normal to the plane perpendicular to the combustor axis.

4. The combustor of claim 2 wherein each injector axis is generally contained within a different plane tangential to a common cylinder defined about the combustor axis.

5. The combustor of claim 1 wherein the wall includes a bulkhead to which the at least one fuel injector is mounted.

6. The combustor of claim 5 wherein the bulkhead includes an opening over which the at least one fuel injector is mounted.

7. The combustor of claim 5 wherein the bulkhead includes an opening through which the at least one fuel injector extends.

8. The combustor of claim 5 wherein the bulkhead includes a plurality of angled planar portions circumferentially spaced about the combustor axis, the at least one fuel injector mounted to at least one of the plurality of angled planar portions.

9. The combustor of claim 1 wherein each at least one fuel injector includes a fuel nozzle.

10. The combustor of claim 9 wherein each at least one fuel injector includes an air guide for supplying core airflow to the combustor.

11. A turbine engine including the combustor of claim 1 and further including a plurality of turbine blades aft of the combustor, the turbine engine having an engine centerline generally coinciding with the combustor axis.

12. The turbine engine of claim 11 further including a fan rotatable about the engine centerline, the plurality of turbine blades operatively coupled to an outer portion of the fan.

13. The turbine engine of claim 12 further including a diffuser case defining a core airflow path between the diffuser case and the wall, the core airflow path having an inlet through the wall.

14. A combustor comprising: a bulkhead circumferential disposed about a combustor axis, the bulkhead at least partially defining an annular combustion chamber; and the bulkhead including a plurality of circumferentially-spaced injector openings each defining a plane not parallel to a plane perpendicular to the combustor axis.

15. The combustor of claim 14 wherein the combustion chamber includes an outlet that is axially spaced from the injector openings.

16. The combustor of claim 14 further including an annular outlet from the combustion chamber.

17. The combustor of claim 14 wherein the bulkhead includes a plurality of angled panel portions, the injector openings formed through the plurality of angled panel portions.

18. A method for operating a turbine engine combustor including the steps of: a) supplying compressed airflow into an annular combustion chamber having a combustor axis; b) supplying fuel to the annular combustion chamber; and c) providing the supply of fuel with an angular component in the combustion chamber.

19. The method of claim 18 wherein said step c) further includes the step of directing the supply of fuel not parallel to an axis of the annular combustion chamber about which the annular combustion chamber is annular.

20. The method of claim 18 wherein said step c) further includes the step of directing the supply of fuel not normal to an axis of the annular combustion chamber about which the annular combustion chamber is annular.