COMBUSTOR LINER SUPPORT AND SEAL ASSEMBLY

Abstract

A method is disclosed herein for reducing binding between a combustor liner of a gas turbine engine and a free-standing ring disposed about the combustor liner. The method comprises the step of disposing a rolling assembly to roll between the combustor liner and the free-standing ring during relative radial displacement.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a divisional application of U.S. patent application Ser. No. 11/755,793, for a COMBUSTOR LINER SUPPORT AND SEAL ASSEMBLY, filed on May 31, 2007, which is hereby incorporated by reference in its entirety.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

The U.S. Government has a paid-up license in this invention and the right in limited circumstances to require the patent owner to license others on reasonable terms as provided for by the terms of N00019-04-C-0093 awarded by the Department of Defense.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates generally to gas turbine engines. More particularly, the present invention relates to a structure for supporting a combustor liner in a gas turbine engine.

2. Description of Related Prior Art

U.S. Pat. No. 6,347,508 sets forth a combustor liner support and seal assembly. Applying the reference numerals used in the '508 patent, an outer combustor liner 28b is supported at its aft end 71 with an inner mounting ring 60 having a plurality of projections or lugs 74. An outer ring 62 includes first and second flanges 92, 94 disposed on opposite sides of the lugs 74. Pins 100 are welded in apertures 104 defined in the flange 94 and extend into slots 76 defined in the lugs 74. The cooperation between the pins 100 and the slots 76 allows the ring 62 and the lugs 74 to move radially with respect to another. Relative radial movement can be desirable because different thermal coefficients of expansion between the combustor liner 28b/mounting ring 60 and the outer ring 62 can lead to undesirable higher thermal gradients and stresses within the liner if the two parts are fixed to one another.

SUMMARY OF THE INVENTION

In summary, the invention a method for reducing binding between a combustor liner of a gas turbine engine and a free-standing ring disposed about the combustor liner. The method comprises the step of disposing a rolling assembly to roll between the combustor liner and the free-standing ring during relative radial displacement.

BRIEF DESCRIPTION OF THE DRAWINGS

Advantages of the present invention will be readily appreciated as the same becomes better understood by reference to the detailed description below when considered in connection with the accompanying drawings, wherein:

FIG. 1 is a simplified schematic view of a gas turbine engine according to the exemplary embodiment of the invention;

FIG. 2 is a cross-sectional view of a rolling assembly according to the exemplary embodiment of the invention;

FIG. 3 is the cross-sectional view of FIG. 2 shown in perspective view;

FIG. 4 is a perspective view looking aft of a plurality of projections according to the exemplary embodiment of the invention and a portion of a free-standing ring according to the exemplary embodiment of the invention;

FIG. 5 is a close-up of a portion of the perspective view of FIG. 4;

FIG. 6 is a perspective view of a portion of an annular seal according to the exemplary embodiment of the invention;

FIG. 7 is a perspective view of a cross-section through the combustor liner support and seal assembly; and

FIG. 8 is a perspective view looking forward of the plurality of projections and a portion a free-standing ring.

DETAILED DESCRIPTION OF THE EXEMPLARY EMBODIMENT

The present invention sets forth several improvements to the combustor liner support and seal assembly set forth in U.S. Pat. No. 6,347,508. Therefore, the '508 patent is hereby incorporated by reference in its entirety.

FIG. 1 is a schematic representation of a gas turbine engine 10. The gas turbine engine 10 extends along a longitudinal axis 12. As used herein, forms of the terms “radial” and “circumference” as applied to some structure refer to the relationship between the structure and the axis 12. The gas turbine engine 10 has a generally annular configuration, however other configurations can be practiced in alternative embodiments of the present invention. The exemplary gas turbine engine 10 includes a fan section 14, a compressor section 16, a combustor section 18, and a turbine section 20 that are integrated to produce an aircraft flight propulsion engine. This particular type of gas turbine engine is generally referred to as a turbo-fan. An alternate form of a gas turbine engine that can be practiced with the invention includes a compressor, a combustor, and a turbine integrated to produce an aircraft flight propulsion engine without a fan section. It should be understood that the term aircraft is generic, including without limitation helicopters, airplanes, missiles, space devices and other substantially similar devices. It is also noted that numerous configurations of turbine engines can practiced with the invention. For example, multiple compressor and turbine sections can be incorporated, with intercoolers connected between the compressor stages. Also, reheat combustion chambers can be added between the turbine stages. All of the various configurations of gas turbine engines described above and/or known in the art can be practiced with the invention. It is also noted that the present invention can be practiced in operating environments other than aircraft propulsion, such as industrial applications including but not limited to pumping sets for gas and oil transmission lines, electricity generation, and naval propulsion.

The compressor section 16 includes a rotor 22 having a plurality of compressor blades 24. The rotor 22 is fixed to a rotatable shaft 26. A plurality of compressor vanes 28 are positioned adjacent to the compressor blades 24 to direct the flow of air through compressor section 16. The combustor section 18 includes an inner combustor liner 30 and an outer combustor liner 32. The liners 30, 32 cooperate with one another to define the inner and outer boundaries of an annular combustion chamber 34. The outer combustor liner 32 is concentrically mounted relative to an outer casing or housing 36 to define an annular fluid passage 38 that surrounds the chamber 34. Also, the inner combustor liner 30 is concentrically mounted relative to the shaft 26 to define an annular fluid passage 40 surrounded by the chamber 34. Fuel is introduced into combustion chamber 34 via a plurality of fuel nozzles (not shown). The inner and outer liners 30, 32 are each formed of materials that are capable of withstanding high temperature environments. Materials such as metallic superalloys and inter-metallic materials, and structures such as Lamilloy®, are contemplated as being within the scope of embodiments of the invention.

The turbine section 20 includes a plurality of turbine blades 42, each coupled to a rotor disk 44. The rotor disk 44 is fixed to the shaft 26. A plurality of turbine vanes 46 are positioned adjacent to the turbine blades 42 to direct the flow of the hot gaseous fluid stream through the turbine section 20. A turbine nozzle 61, sometimes referred to as inlet guide vanes 46, is positioned downstream of the combustor section 18 to direct the hot gaseous fluid stream exiting the combustion chamber 34 toward the turbine blades 42. In the exemplary embodiment of the invention, the gaseous fluid comprises combustion gases.

In operation, the turbine section 20 provides rotational power to one or more shafts 26 to drive the fan section 14 and the compressor section 16, respectively. The fan section 14 includes a fan 48. Air enters the gas turbine engine 10 in the direction indicated by arrows 50, 52 and passes through the fan section 14. The air stream is then divided and fed into both the compressor section 16 and a bypass duct 54. The compressed air exiting compressor section 16 is routed into both the combustion chamber 34 and also the annular fluid passages 38, 40. The compressed air enters the combustion chamber 34 at a forward end 56 of the combustor section 18 and is intermixed with fuel, to becoming a combustible air/fuel mixture. The air/fuel mixture is ignited and burned in the combustor section 18, generating a hot gaseous fluid stream. The hot gaseous fluid stream exits an aft end 58 of the combustor section 18 and is fed into the turbine section 20 to provide the energy applied to power the gas turbine engine 10. During normal operation of the gas turbine engine 10, the air flowing through passages 38, 40 is at a higher pressure than the hot gaseous air stream flowing through combustion chamber 34 and is also at a lower temperature.

Two operational considerations relating to the turbine nozzle flow from the arrangement described above. First, the combustor liners 30, 32 move relative to the turbine section 20 and nozzle 61 and it is thus desirable to seal fluid passageways 38, 40 from the turbine blades 42 and turbine vanes 46. Second, the pressure differential between the fluid streams moving around the outside of the combustion chamber 34 in the passage 38 and the hot gaseous fluid stream moving inside the combustion chamber 34 results in a buckling load on the combustor liner 32 and it is therefore desirable to support the floating ends of the combustor liner 32 against inward deflection.

Referring now to FIGS. 2 and 3, a combustor liner support and seal assembly 60 is positioned between the aft end 58 of the combustor liner 32 and the turbine nozzle 61. The aft end 58 of liner 32 is spaced apart from the turbine nozzle, defining a passageway. The assembly 60 closes and seals this passageway while allowing the aft end 58 to expand and contract radially. The assembly 60 also supports the liner 32 and helps prevent buckling. The assembly 60 includes an inner mounting ring 62 with a forward end 64 fixed to the aft end 58 of the combustor liner 32 about the entire periphery of the outer combustor liner 32. A plurality of projections 66 extend from the combustor liner 32 by way of the mounting ring 62. The projections 66 are spaced from one another circumferentially about the axis 12. The assembly 60 also includes a free-standing ring 68 disposed about the combustor liner 32 and positioned adjacent to the plurality of projection 66 along the axis 12. The free-standing ring 68 floats relative to the combustor liner 32 in a plane normal to the axis 12. The plurality of projections 66 engage a corresponding circumferentially-facing portion of the free-standing ring 68 and circumferentially support the combustor liner 32 while allowing relative radial displacement between said combustor liner 32 and said free-standing ring 68.

The exemplary combustor liner 32 is formed of a metallic material and has a thermal mass less than the thermal mass of the ring 68. Also, the combustor liner 32 has a coefficient of thermal expansion that is equal to the coefficient of thermal expansion of the ring 68 and the ring 68 has a higher moment of inertia than the liner 32 and inner mounting ring 60. As a result, during operation the combustion liner 32 and the ring 62 will radially expand and contract together in response to the thermal cycle operation of the gas turbine engine 10 and the ring 68 will radially expand and contract at a slower rate. To compensate for this variation in radial expansion and contraction, relative radial displacement between the ring 62 and the ring 68 is permitted. The radial movement of the ring 60 and the liner 32 reduces undesirable hoop stresses from developing within the liner 32 which might otherwise result in low cycle fatigue (LCF) and the eventual failure of the liner 32.

The exemplary embodiment of the invention includes a rolling assembly 70 operably disposed between the free-standing ring 68 and the plurality of projections 66 to reduce binding during the relative radial displacement. In operation of the exemplary embodiment, for example, the pins 74 are substantially fixed against movement along the axis 12 by the ring 68 but are disposed to rotate relative to the ring 68. When the ring 68 and projections 66 move radially relative to one another, the pins 74 can roll along the side surfaces of the slots 72. The pins 74 may rotate one or more full turns during relative radial movement between the ring 68 and projections 66, or may rotate less than a full turn. The pins 74 may rotate back and forth during operation, or pivot, over only a portion of the slot 72 or may rotate along the entire length of the slot 72.

In the exemplary embodiment of the invention, the rolling assembly 70 includes a plurality of slots 72 and a plurality of pins 74. The plurality of slots 72 are defined in one of the free-standing ring 68 and the plurality of projections 66. In the exemplary embodiment, the slots 72 extend in the projections 66. In alternative embodiments of the invention, the slots 72 can extend in the ring 68. In such embodiments, the pins 74 would move radially relative to the ring 68. The slots 72 extend in a radial direction with respect to the axis 12. The plurality of pins 74 are each rotatably engaged with at least one of the free-standing ring 68 and the plurality of projections 66 and received in one of the plurality of slots 72 to thereby engage the combustor liner 32, through the mounting ring 62, and the free-standing ring 68 together for the relative radial displacement.

Any particular pin 74 could move in one of the slots 72 without contacting the surfaces that define the slot 72. This event is not likely, but could occur and is contemplated as an embodiment of the invention. Even in such a situation, the pin 74 would still be operable to rotate within the slot 72 even if rotation does not in fact occur.

In the exemplary embodiment of the invention, the pins 74 are rotatably engaged with both the projections 66 and the ring 68. The exemplary ring 68 includes a first radially-extending flange 76 disposed on a first side of the plurality of projections 66 along the axis 12 and a second radially-extending flange 78 disposed on a second side of the plurality of projections 66 along the axis 12 opposite the first side. Each of the plurality of pins 74 includes a first end 80 rotatably engaged with the first radially-extending flange 76 and a second end 82 rotatably engaged with the second radially-extending flange 78. The end 80 is received in an aperture 84 defined in the flange 76. The end 80 and aperture 84 are sized such that a close-tolerance, loose-fit is defined between the end 80 and the aperture 84. The end 82 is received in an aperture 86 defined in the flange 78. The cooperative relationship between the end 82 and the aperture 86 is the same as the cooperative relationship between the end 80 and the aperture 84. Thus, during relative radial movement between the liner 32 and the ring 68, the ends 80, 82 can substantially freely rotate within the apertures 84, 86, respectively.

The pins 74 also include a center portion 88 between the first and second ends 80, 82. The center portion 88 is rotatably disposed in one of the plurality of slots 72. Thus, during relative radial movement between the liner 32 and the ring 68, the center portion 88 can substantially freely rotate within the slot 72. The slots 72 define a width extending circumferentially with respect to the axis 12. The pins 74 are sized relative to the width of the slots 72 to substantially prevent circumferential movement between the plurality of projections 66 and the free-standing ring 68.

In the exemplary embodiment of the invention, the center portion 88 of at least some of the plurality of pins 74 have a greater diameter than the first and second ends 80, 82 such that a first shoulder 90 is defined between the center portion 88 and the first end 80 and a second shoulder 92 is defined between the center portion 88 and the second end 82. The first and second shoulders 90, 92 ensure that a minimum distance is maintained between the first and second radially-extending flanges 76, 78 to prevent binding between the flanges 76, 78 and the projections 66.

The exemplary embodiment of the invention also includes a limiting structure operably disposed between the free-standing ring 68 and the plurality of projections 66 to limit the relative radial displacement. The slots 72 extend radially-outward from a first closed end 94 to a second closed end 96 spaced radially outward of the first closed end 94. The pins 74 are limited in movement by the first and second ends 94, 96 of the slot 72, thereby limiting the relative radial displacement between the liner 32 and the ring 68. Limiting radial relative movement through the closed slots 72 ensures that the pins 74 will not pass out of the slots 72 and also prevents deformation of the liner 32 and ring 62, caused by a thermal growth differential, beyond a range deemed acceptable. The length of slot 72 between the closed ends 94, 96 can be selected so that the outer closed end 96 is not engaged by a pin 74 during expected thermal growth so as not to reduce low-cycle fatigue life. The liner 32 could buckle if a predetermined amount of expansion is prevented. The length of the slot 72 defines the predetermined amount of design travel.

Referring now to FIGS. 6-8, the exemplary embodiment of the invention also includes a self-supporting annular seal 98 disposed between the free-standing ring 68 and the plurality of projections 66. The annular seal 98 has a variable thickness 99 about the axis 12. The thickness 99 is defined along the axis 12. Relatively wider portions of the annular seal 98 are disposed between adjacent projections 66 about the axis 12. Relatively narrower portions of the annular seal 98 are aligned with the plurality of projections 66 about the axis 12.

The exemplary seal 98 includes a first panel 100 and second panel 102 that are segmented about the axis 12. FIG. 6 shows one of the first panels 100 and one of the second panels 102. The seal 98 includes a plurality of the panels 100 that are placed in adjoining, side-by-side relationship to one another to define a substantially continuous sealing surface extending circumferentially around the axis 12. Also, the seal 98 includes a plurality of the panels 102 that are placed in adjoining, side-by-side relationship to one another to define a substantially continuous supporting ring extending circumferentially around the axis 12. The plurality of panels 100 are disposed back-to-back with the plurality of second panels 102.

Each of the panels 100 extends between two adjacent projections 66 to seal the space between the projections 66. Since the pressure of the fluid stream moving through the passageway 38 is greater than the pressure of the fluid stream moving out of the combustion chamber 34, each panel 100 is subjected to pressure tending to push the panel 100 toward the flange 78. This pressure would tend to cause the panel 100 to buckle between the projections 66 and create a leak path. The panels 102 provide support to the panels 100 to prevent this buckling.

Each of the panels 102 includes a flat portion 104 and a protrusion 106 extending from the flat portion 104 through the space between the projections 66. The protrusions 106 define the wider portions of the annular seal 98 about the axis 12, thus resulting in the exemplary seal 98 having a variable width. The first panel 100 contacts the first radially-extending flange 76 and the protrusion 106 extends through the space between two adjacent projections 66 to contact the second radially-extending flange 78 and thereby support the first panel 100 against buckling.

Referring again to FIGS. 2 and 3, each of the panels 100, 102 are maintained to cover the space between adjacent projections 66. However, the panels 100, 102 can move with the projections 66 in response to expansion and contraction resulting from temperature changes. Each of the plurality of projections 66 extends from a base 108 adjacent to the combustor liner 32. As the liner 32 expands, the projections 66 move radially outward. Thus, the projections 66 move circumferentially apart from one another as the liner 32 expands.

The panels 100, 102 move with the projections 66 during expansion of the liner 32. Referring to FIGS. 2, 5 and 6, the first panel 100 includes an aperture 118 aligned with one of the slots 72 and defining a circumferential width 120. One of the pins 74 is received in the aperture 118. Each panel 100 can move with the pin 74 that is received in its aperture 118. When the liner 32 expands, the panels 100, which are disposed in adjoining, side-by-side relationship to one another, can circumferentially separate from one another.

Each of the second panels 102 includes an aperture 114 aligned with one of the slots 72. One of the pins 74 extends through the aperture 114 and prevents relative circumferential displacement between the second panel 102 and the plurality of projections 66. When the liner 32 expands, the panels 102, which are disposed in adjoining, side-by-side relationship to one another, can circumferentially separate from one another.

As best shown in FIG. 4, the first and second panels 100, 102 can be staggered to enhance sealing. The aperture 118 of a first panel 100 can be aligned with a first one of the slots 72 and the aperture 114 of the immediately adjacent second panel 102 can be aligned with a second one of the slots 72 different than the first one. When the liner 32 expands, a panel 100 and a panel 102 disposed back-to-back can move relative to one another, sliding circumferentially away from one another. The back-to-back panels 100 and 102 do not slide fully apart and the gap between adjacent projections 66 remains closed.

Each of the plurality of projections 66 extends from the base 108 to a distal end 110 and includes a flange 112 projecting from the distal end 110. The flange 112 extends parallel to the axis 12 and substantially prevents radial movement of the panels 100, 102 outward relative to the projection 66. The panels 100, 102 define heights in the radial direction that are substantially equal to a distance between the base 108 and the flange 112. The second panels 102 include a lip 116 received at the base 108. The lip 116 receives a radially-innermost edge of the first panel 100. The first panel 100 defines a height in the radial direction that is substantially equal to a distance between the lip 116 and the flange 112 to maintain the first panel 100 in a fixed radial position relative to the projections 66.

While the invention has been described with reference to an exemplary embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims.

Claims

1. A method for reducing binding between a combustor liner of a gas turbine engine and a free-standing ring disposed about the combustor liner comprising the of:

disposing a rolling assembly to roll between the combustor liner and the free-standing ring during relative radial displacement while substantially preventing relative circumferential movement.

2. The method of claim 1 further comprising the step of:

limiting the extent of relative radial displacement between the combustor liner and the free-standing ring to a predetermined design amount.

3. The method of claim 1 further comprising the step of:

disposing a seal of variable width between the combustor liner and the free-standing ring, the variable width varying in at least one of a radial direction and a circumferential direction relative to an axis along which the gas turbine engine extends.

4. The method of claim 1 wherein the rolling assembly rolls radially along a length of a surface defined by one of the combustor liner and the free-standing ring during the relative radial displacement.

5. The method of claim 4 wherein the rolling assembly substantially prevents relative circumferential movement between the combustor liner and the free-standing ring while permitting the relative radial displacement between the combustor liner and the free-standing ring in a radial direction.

6. The method of claim 1 wherein the rolling assembly is rotatably engaged with both the combustor liner and the free-standing ring.

7. The method of claim 1 wherein the rolling assembly includes a plurality of pins engaged between the combustor liner and the free-standing ring.

8. The method of claim 7 wherein the plurality of pins are each received in one of a plurality of slots extending in a substantially radial direction relative to an axis along which the gas turbine engine extends to permit the relative radial displacement between the combustor liner and the free-standing ring.

9. The method of claim 7 wherein the plurality of pins are rotatably engaged with both the combustor liner and the free-standing ring.

10. A method for reducing binding between a combustor liner of a gas turbine engine and a free-standing ring disposed about the combustor liner comprising the step of:

disposing a rolling assembly to roll between the combustor liner and the free-standing ring during relative radial displacement, wherein the rolling assembly substantially prevents relative circumferential movement between the combustor liner and the free-standing ring while permitting the relative radial displacement between the combustor liner and the free-standing ring in a radial direction.

11. A method for reducing binding between a combustor liner of a gas turbine engine and a free-standing ring disposed about the combustor liner comprising the step of:

disposing a rolling assembly to roll between the combustor liner and the free-standing ring during relative radial displacement, wherein the rolling assembly comprises:

a plurality of slots defined in one of the combustor liner and the free-standing ring and extending in a substantially radial direction with respect to an axis along which the gas turbine engine extends; and

a plurality of pins each rotatably engaged with the other of the combustor liner and the free-standing ring and each received in one of the plurality of slots to thereby engage the combustor liner and the free-standing ring together to permit the relative radial displacement between the combustor liner and the free-standing ring.

12. The method of claim 1 wherein the free-standing ring is supported solely by the combustor liner.

13. The method of claim 12 wherein the free-standing ring floats relative to the combustion liner in a radial direction.

14. The method of claim 13 wherein the free-standing ring provides circumferential support to the combustion liner in a circumferential direction.

15. The method of claim 1 wherein the free-standing ring floats relative to the combustion liner in a radial direction and provides circumferential support to the combustion liner in a circumferential direction.

16. A method for reducing binding between a combustor liner of a gas turbine engine and a free-standing ring disposed about the combustor liner comprising the step of:

disposing a rolling assembly to roll between the combustor liner and the free-standing ring during relative radial displacement;

limiting the extent of relative radial displacement between the combustor liner and the free-standing ring to a predetermined design amount, wherein the limiting of the relative radial displacement between the combustor liner and the free-standing ring to the predetermined design amount comprises positioning a projection extending from one of the combustor liner and the free-standing ring in a radially-extending slot defined by the other of the combustor liner and the free-standing ring.

17. The method of claim 16 wherein the radially-extending slot includes opposite first and second closed ends to provide limits to the extent of the relative radial displacement.