US20150337687A1

US20150337687A1 - Split cast vane fairing

Info

Publication number: US20150337687A1
Application number: US14/758,377
Authority: US
Inventors: Jonathan Ariel Scott; Matthew Budnick; Conway Chuong
Original assignee: United Technologies Corp
Current assignee: Raytheon Technologies Corp
Priority date: 2012-12-29
Filing date: 2013-12-19
Publication date: 2015-11-26
Also published as: WO2014105621A1

Abstract

A turbine exhaust case (28) comprises a frame (102) and a fairing (118). The frame has inner (106) and outer (104) rings connected by a plurality of radial struts (108). The fairing defines an airflow path within the frame, and has upstream (202) and downstream (204) sections connected together about the radial struts.

Description

BACKGROUND

The present disclosure relates generally to gas turbine engines, and more particularly to a vane fairing for a turbine exhaust case of an industrial gas turbine engine.
A turbine exhaust case is a structural frame that supports engine bearing loads while providing a gas path at or near the aft end of a gas turbine engine. Some aero engines utilize a turbine exhaust case to help mount the gas turbine engine to an aircraft airframe. In industrial applications, a turbine exhaust case is more commonly used to couple gas turbine engines to a power turbine that powers an electrical generator. Industrial turbine exhaust cases may, for instance, be situated between a low pressure engine turbine and a generator power turbine. A turbine exhaust case must bear shaft loads from interior bearings, and must be capable of sustained operation at high temperatures.
Turbine exhaust cases serve two primary purposes: airflow channeling and structural support. Turbine exhaust cases typically comprise structures with inner and outer rings connected by radial struts. The struts and rings often define a core flow path from fore to aft, while simultaneously mechanically supporting shaft bearings situated axially inward of the inner ring. The components of a turbine exhaust case are exposed to very high temperatures along the core flow path. Various approaches and architectures have been employed to handle these high temperatures. Some turbine exhaust case frames utilize high-temperature, high-stress capable materials to both define the core flow path and bear mechanical loads. Other frame architectures separate these two functions, pairing a structural frame for mechanical loads with a high-temperature capable fairing to define the core flow path. In systems with separate structural frames and flow path fairings, the installation of the fairing on the frame without weakening either component presents a technical challenge.

SUMMARY

The present disclosure is directed toward a turbine exhaust case comprising a frame and a fairing. The frame has inner and outer rings connected by a plurality of radial struts. The fairing defines an airflow path within the frame, and has upstream and downstream sections connected together about the radial struts.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a simplified partial cross-sectional view of an embodiment of a gas turbine engine.

FIG. 2 is a cross-sectional view of a turbine exhaust case of the gas turbine engine of FIG. 1.

FIG. 3A is a perspective view of an initial cast of a vane fairing of the turbine exhaust case of FIG. 2.

FIG. 3B is a perspective view of the vane fairing of FIG. 3A, cut apart into upstream and downstream sections

FIG. 3C is a perspective view of the vane fairing of FIG. 3B, with the upstream and downstream sections attached together in an assembled state

FIG. 4 is a method flowchart describing a method for assembling the turbine exhaust case of FIG. 2.

DETAILED DESCRIPTION

FIG. 1 is a simplified partial cross-sectional view of gas turbine engine 10, comprising inlet 12, compressor 14 (with low pressure compressor 16 and high pressure compressor 18), combustor 20, engine turbine 22 (with high pressure turbine 24 and low pressure turbine 26), turbine exhaust case 28, power turbine 30, low pressure shaft 32, high pressure shaft 34, and power shaft 36. Gas turbine engine 10 can, for instance, be an industrial power turbine.
Low pressure shaft 32, high pressure shaft 34, and power shaft 36 are situated along rotational axis A. In the depicted embodiment, low pressure shaft 32 and high pressure shaft 34 are arranged concentrically, while power shaft 36 is disposed axially aft of low pressure shaft 32 and high pressure shaft 34. Low pressure shaft 32 defines a low pressure spool including low pressure compressor 16 and low pressure turbine 26. High pressure shaft 34 analogously defines a high pressure spool including high pressure compressor 18 and high pressure compressor 24. As is well known in the art of gas turbines, airflow F is received at inlet 12, then pressurized by low pressure compressor 16 and high pressure compressor 18. Fuel is injected at combustor 20, where the resulting fuel-air mixture is ignited. Expanding combustion gasses rotate high pressure turbine 24 and low pressure turbine 26, thereby driving high and low pressure compressors 18 and 16 through high pressure shaft 34 and low pressure shaft 32, respectively. Although compressor 14 and engine turbine 22 are depicted as two-spool components with high and low sections on separate shafts, single spool or 3+ spool embodiments of compressor 14 and engine turbine 22 are also possible. Turbine exhaust case 28 carries airflow from low pressure turbine 26 to power turbine 30, where this airflow drives power shaft 36. Power shaft 36 can, for instance, drive an electrical generator, pump, mechanical gearbox, or other accessory (not shown).
In addition to defining an airflow path from low pressure turbine 26 to power turbine 30, turbine exhaust case 28 can support one or more shaft loads. Turbine exhaust case 28 can, for instance, support low pressure shaft 32 via bearing compartments (not shown) disposed to communicate load from low pressure shaft 32 to a structural frame of turbine exhaust case 28.
FIG. 2 is a cross-sectional view of an embodiment of turbine exhaust case 28, illustrating frame 102 (with frame outer ring 104, frame inner ring 106, frame struts 108, low pressure turbine connection 110, and power turbine connection 112), bearing support 114, fasteners 116 a and 116 b, fairing 118 (with fairing outer platform 120, fairing inner platform 122, and fairing vanes 124), forward stiffening flange 126, aft stiffening flange 128, strut heat shield 132, outer heat shield 134, and inner heat shield 136.
As described above with respect to FIG. 1, turbine exhaust case 28 defines at least a portion of an airflow path for core flow F, and carries load radially from bearing support 114 (which in turn connects to bearing components, not shown). These two functions are performed by separate components: frame 102 carries bearing loads, while fairing 118 at least partially defines the flow path of core flow F.
Frame 102 is a relatively thick, rigid support structure formed, for example, of cast steel. Outer ring 104 of frame 102 serves as an attachment point for upstream and downstream components at low pressure turbine connection 110 and power turbine connection 112, respectively. Low pressure turbine connection 110 and power turbine connection 112 can, for instance, include fastener holes for attachment to adjacent low pressure turbine 26 and power turbine 30, respectively. Frame inner ring 106 is mechanically connected to bearing support 114 via fasteners 116 a, which can for instance be bolts, screws, pins or rivets. Frame inner ring 106 communicates bearing load radially from bearing support 114 to frame outer ring 104 via frame struts 108, which extend at angular intervals between frame inner ring 106 and frame outer ring 104. Although only one strut 108 is visible in FIG. 1, turbine exhaust case 28 can include any desired number of struts 108.
Fairing 118 is a high-temperature capable aerodynamic structure at least partially defining the boundaries of core flow F through turbine exhaust case 28. Fairing outer platform 120 generally defines an outer flowpath diameter, while fairing inner platform 122 generally defines an inner flowpath diameter. Fairing vanes 124 surround frame struts 108, and form a plurality of aerodynamic vane bodies. Fairing 118 can, for instance, be formed of a superalloy material such as Inconel or other nickel-based superalloy. Fairing 118 is generally rated for higher temperatures than frame 102, and can be affixed to frame 102 via fasteners 116 b. In the depicted embodiment, fairing 118 is affixed to frame inner ring 106 at the forward inner diameter of fairing 118, although alternative embodiments of turbine exhaust case 28 can secure fairing 118 by other means and/or in other locations. Forward and aft stiffening flanges 126 and 128, respectively, can extend radially outward from the entire circumference of fairing outer platform 120 to provide increased structural rigidity to fairing 118.
Turbine exhaust case 28 includes a plurality of heat shields to protect frame 102 from radiative and convective heating. Strut heat shield 132 is situated between fairing vanes 124 and frame struts 108. Outer heat shield 134 can be situated between fairing outer platform 120 and frame outer ring 104. Inner heat shield 136 can be is situated radially inward of a forward portion of fairing inner platform 122. Like fairing 118, all three heat shields 132, 134, and 136 can be formed of Inconel or a similar nickel-based superalloy. Strut heat shield 132, outer heat shield 134, and inner heat shield 136 act as barriers to heat from fairing 118, which can become very hot during operation of gas turbine 10. Heat shields 132, 134, and 136 thus help to protect frame 102, which can be rated to lower temperatures than fairing 118, from exposure to excessive heat.
FIGS. 3A, 3B, and 3C present successive steps in the formation and installation of fairing 118 with outer platform 120, inner platform 122, and fairing vanes 124. FIGS. 3A, 3B, and 3C illustrate a single angular section 200 of fairing 118. Angular section 200 can be a representative section of a unitary fairing cast as a annular piece, or one of several separately cast angular pieces welded or otherwise joined together to form fairing 118. Angular section 200 includes forward section 202 and aft section 204 separated by sacrificial region 206 situated along cut line CL.
As shown in FIG. 3A, angular section 200 is initially cast in a single piece. As shown in FIG. 3B, angular section 200 is cut along cut line CL to separate forward section 202 from aft section 204, thereby allowing vane section 200 to be situated about frame 102. Fairing section 200 can, for instance, be cut by electric discharge machining (EDM). The cutting process consumes a portion of sacrificial region 206 of angular section 200, leaving reduced forward and aft sacrificial regions 206 a and 206 b attached to forward section 202 and aft section 204, respectively. In alternative embodiments, forward and aft sections 202 and 204 can be separated by other means, such as by mechanically cutting along cut line CL. In some such embodiments, forward and aft sections 202 and 204 can subsequently be finished along cut line CL to smooth the resulting cut. The width of sacrificial region 206 can be scaled as appropriate to the method of cutting selected. Forward and aft sections 202 and 204 are then positioned about strut 108 of frame 102, and welded together along cut line CL as shown in FIG. 3C to form a single unitary piece. This weld consumes the remainder of forward and aft sacrificial sections 206 a and 206 b.
The positioning of cut line CL is selected to distance the resulting eventual weld apart from leading and trailing edges of fairing vane 124 where stresses on fairing 118 from core airflow F are greatest. In the depicted embodiment, cut line CL is situated along a chord slightly upstream of the widest portion of fairing vane 124. This positioning allows easy access from the forward side of fairing 118 to perform a manual weld along cut line CL during installation. In alternative embodiments wherein a manual weld is performed from aft of fairing 118, cut line CL can instead be positioned along a chord downstream of the widest portion of fairing vane 124.
FIG. 4 illustrates the method set out above with respect to FIGS. 3A, 3B, and 3C for installing fairing 118. First, frame 102 is produced in a single piece. (Step S1). Frame 102 can be formed in a single unitary section, or formed from several separately cast pieces joined together. Next, forward and aft sections 202 and 204 of fairing 118 are cast in a single angular piece 200. (Step S2). As discussed above, fairing 118 can be formed from a plurality of separate angular sections 200, or in a single annular piece. Fairing 118 can, for instance, be die cast. Fairing 118 is then cut into forward section 202 and aft section 204, thereby consuming a portion of sacrificial region 206. (Step S3). Forward and aft sections 202 and 204 are assembled about strut 108 of frame 102 (Step S4), and welded back together (Step S5), thereby consuming the remainder of sacrificial region 206.
By splitting each angular section 200 into multiple sections during its installation process, fairing 118 can be installed about the strut of a unitary frame. Splitting fairing 118 into forward and aft sections produces an assembled fairing without weak joints corresponding to weld locations at leading or trailing edges of fairing vanes 124.

Discussion of Possible Embodiments

The following are non-exclusive descriptions of possible embodiments of the present invention.
A turbine exhaust case comprising a frame having inner and outer rings connected by a plurality of radial struts, and a fairing defining an airflow path within the frame. The fairing comprises upstream and downstream sections connected together about the radial struts.
The turbine exhaust case of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
wherein the upstream and downstream sections of the vane are welded together about the radial struts along a cut line.
wherein the cut line defines a radial plane of separation between the upstream and downstream sections.
wherein the fairing comprises radially inner and outer platforms connected by a vane body.
wherein the upstream and downstream sections are connected along a chord near a widest section of the vane body.
wherein the chord is located upstream of the widest section of the vane body.
wherein the fairing is comprised of a plurality of joined-together angular segments further comprising a radiative heat shield disposed between the frame and the fairing.
A method for assembling a, the method comprising casting a fairing with a vane body, an inner platform, and an outer platform in a single piece; cutting the fairing into an upstream section and a downstream section along a along a cut line; assembling the upstream and downstream sections about a turbine exhaust case frame; and attaching the upstream section to the downstream section.
The method of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
wherein the cut line is situated along a chord near the widest section of the vane body wherein the fairing includes a sacrificial region to account for weld shrinkage and material removed while cutting the fairing into upstream and downstream sections.
wherein the vane is cut using electric discharge machining.
wherein attaching the upstream section to the downstream section comprises performing a weld long the chord.
wherein the chord is upstream of the widest section of the vane body.
wherein the chord is downstream of the widest section of the vane body.
wherein the weld is a manual weld.
wherein the manual weld is performed from forward of the fairing.
While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes can be made and equivalents can be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications can 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(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.

Claims

1. A turbine exhaust case comprising:

a frame having inner and outer rings connected by a plurality of radial struts; and

a fairing defining an airflow path within the frame, the fairing comprising upstream and downstream sections connected together about the radial struts.

2. The turbine exhaust case of claim 1, wherein the upstream and downstream sections of the vane are welded together about the radial struts along a cut line.

3. The turbine exhaust case of claim 2, wherein the cut line defines a radial plane of separation between the upstream and downstream sections.

4. The turbine exhaust case of claim 1, wherein the fairing comprises radially inner and outer platforms connected by a vane body.

5. The turbine exhaust case of claim 4, wherein the upstream and downstream sections are connected along a chord near a circumferentially thickest section of the vane body.

6. The turbine exhaust case of claim 5, wherein the chord is located upstream of the circumferentially thickest section of the vane body.

7. The turbine exhaust case of claim 1, wherein the fairing is comprised of a plurality of joined-together angular segments.

8. The turbine exhaust case of claim 1, further comprising a radiative heat shield disposed between the frame and the fairing.

9. A method for assembling a turbine exhaust case, the method comprising;

casting a fairing with a vane body, an inner platform, and an outer platform in a single piece;

cutting the fairing into an upstream section and a downstream section along a cut line;

assembling the upstream and downstream sections about a turbine exhaust case frame; and

attaching the upstream section to the downstream section.

10. The method of claim 9, wherein the cut line is situated at a chord location near the widest section of the vane body.

11. The method of claim 10, wherein the fairing includes a sacrificial region to account for weld shrinkage and material removed while cutting the fairing into upstream and downstream sections.

12. The method of claim 10, wherein the vane is cut using electric discharge machining.

13. The method of claim 10, wherein attaching the upstream section to the downstream section comprises performing a weld long at the chord location.

14. The method of claim 13, wherein the chord location is situated upstream of the widest section of the vane body.

15. The method of claim 13, wherein the chord location is situated downstream of the widest section of the vane body.

16. The method of claim 13, wherein the weld is a manual weld.

17. The method of claim 16, wherein the manual weld is performed from forward of the fairing.