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Title:
A DUCTING ARRANGEMENT FOR DIRECTING COMBUSTION GAS
Document Type and Number:
WIPO Patent Application WO/2017/023330
Kind Code:
A1
Abstract:
A ducting arrangement (10), including: an annular duct (38) having a plurality of discrete IEPs (18) secured together to form the annular duct that defines an annular chamber (14) configured to deliver an annular flow of combustion gases directly onto turbine blades without a turning vane, wherein the annular duct defines a chamber mid-annulus (98); an IEP intermediate portion (34) for each IEP, each IEP intermediate portion defining a respective intermediate flow path (48) in fluid communication with the annular chamber and defining an intermediate flow path axis (192); and a cone (16) for each IEP intermediate portion, each cone defining a cone flow path (46), the cone flow path defining a cone flow path axis (190). When viewed from upstream toward downstream along a longitudinal axis (134) of the gas turbine engine each cone flow path axis intersects a respective intermediate flow path axis (192).

Inventors:
MORRISON JAY A (US)
CHARRON RICHARD C (US)
KUMAR MANISH (US)
MONTGOMERY MATTHEW D (US)
Application Number:
PCT/US2015/043974
Publication Date:
February 09, 2017
Filing Date:
August 06, 2015
Export Citation:
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Assignee:
SIEMENS AG (DE)
SIEMENS ENERGY INC (US)
International Classes:
F01D9/02
Foreign References:
US8276389B22012-10-02
EP2383518A22011-11-02
US20150114003A12015-04-30
US8276389B22012-10-02
US8230688B22012-07-31
Attorney, Agent or Firm:
MORA, Enrique J. (US)
Download PDF:
Claims:
CLAIMS

The invention claimed is: 1 . A ducting arrangement (10), comprising:

an annular duct (38) comprising a plurality of discrete lEPs (18) secured together to form the annular duct (38), the annular duct (38) defining an annular chamber (14) configured to deliver an annular flow of combustion gases directly onto turbine blades of a gas turbine engine without a turning vane, wherein the annular duct (38) defines a chamber mid-annulus (98) equidistant from an inner diameter (90) and an outer diameter (92) of the annular chamber (14);

an IEP intermediate portion (34) for each IEP (18), a fully bounded portion of each IEP intermediate portion (34) defining a respective intermediate flow path (48) in fluid communication with the annular chamber (14) and defining an intermediate flow path axis (192); and

a cone (16) for each IEP intermediate portion (34), each cone (16) defining a cone flow path (46) in fluid communication with a respective IEP intermediate portion (34), the cone flow path (46) defining a cone flow path axis (190), wherein each cone (16) is configured to establish fluid communication with an outlet of a respective combustor can of a can annular combustion arrangement;

wherein when viewed from upstream toward downstream along a longitudinal axis (134) of the gas turbine engine each cone flow path axis (190) intersects a respective intermediate flow path axis (192). 2. The ducting arrangement (10) of claim 1 , wherein the annular chamber

(14) defines a chamber mid-annulus (98) disposed centrally between an inner diameter (90) and an outer diameter (92) of the annular chamber (14), and wherein when viewed from upstream toward downstream along the longitudinal axis (134) of the gas turbine engine each intermediate flow path axis (192) forms a respective tangent (1 10) with the chamber mid-annulus (98).

3. The ducting arrangement (10) of claim 1 , wherein when viewed from upstream toward downstream along the longitudinal axis (134) of the gas turbine engine each cone flow path axis (190) forms a respective cone radial skew angle (100) of up to ten degrees with a respective intermediate flow path axis (192).

4. The ducting arrangement (10) of claim 1 , wherein when viewed from upstream toward downstream along the longitudinal axis (134) of the gas turbine engine each cone flow path axis (190) forms a respective cone radial skew angle (100) of over two degrees with a respective intermediate flow path axis (192).

5. The ducting arrangement (10) of claim 1 , wherein a flow area of a respective intermediate flow path (48) is less than a flow area of a respective cone flow path (46) at a cone outlet (28).

6. The ducting arrangement (10) of claim 1 , wherein at least one of the cone (16) and the IEP (18) forms a respective accelerating geometry (42) effective to accelerate combustion gases to a speed acceptable for delivery onto the turbine blades.

7. The ducting arrangement (10) of claim 1 , wherein at least one of the cone (16) and the IEP (18) forms a respective throat (44) comprising a constant hydraulic diameter, a constant shape, and a length of at least ten percent of the constant hydraulic diameter.

8. A ducting arrangement (10), comprising:

a cone (16) comprising a cone inlet (26) configured to secure to a combustor outlet of a combustor can of a can annular combustion arrangement, and a cone outlet (28), wherein the cone (16) defines a cone flow path (46) that narrows from the cone inlet (26) to the cone outlet (28);

an lEP (18) comprising an lEP inlet (30) configured to secure to the cone outlet (28), an annular chamber end (36), and an lEP intermediate portion (34), wherein a fully bound portion of the lEP intermediate portion (34) defines an intermediate flow path (48);

wherein the annular chamber end (36) defines a portion of an annular chamber

(14) that defines a annular outlet suitable for fluid communication with a first stage of turbine blades without intervening turning vanes,

wherein the lEP intermediate portion (34) is in fluid communication with the annular chamber (14),

wherein the annular chamber (14) defines a chamber mid-annulus (98) disposed centrally between an inner diameter (90) and an outer diameter (92) of the annular chamber (14), and

wherein a circumferential component (182) of the cone flow path (46) and a circumferential component (180) of the intermediate flow path (48) form a cone radial skew angle (100) with each other.

9. The ducting arrangement (10) of claim 8, wherein the cone radial skew angle (100) is eighty degrees or more. 10. The ducting arrangement (10) of claim 9, wherein the cone radial skew angle (100) is less than eighty-eight degrees.

1 1 . The ducting arrangement (10) of claim 8, wherein the cone (16) defines a circular cross sectional shape for the cone flow path (46) at the cone inlet (26), and wherein the lEP (18) defines a non-circular cross sectional shape for the intermediate flow path (48) in the lEP intermediate portion (34).

12. The ducting arrangement (10) of claim 8, wherein a flow area of the intermediate flow path (48) is less than a flow area of the cone flow path (46) at the cone outlet (28). 13. The ducting arrangement (10) of claim 8, where at least one of the cone

(16) and the IEP (18) forms an accelerating geometry (42) effective to accelerate combustion gases to over Mach 0.6.

14. The ducting arrangement (10) of claim 8, where at least one of the cone (16) and the IEP (18) forms a collimating throat (44).

15. The ducting arrangement (10) of claim 8, further comprising one cone (16) and one IEP (18) for each combustor can of the can annular combustion arrangement. 16. A ducting arrangement (10), comprising:

a cone (16) and an IEP (18) for a combustor can of a gas turbine engine can annular combustion arrangement, wherein the IEP (18) defines part of an annular duct (38) defining an annular chamber (14) configured to deliver an annular flow of combustion gases directly onto turbine blades without a turning vane;

wherein the cone (16) is configured to establish fluid communication with an outlet of the combustor can via a cone flow path (46), and wherein the IEP (18) establishes fluid communication between the cone (16) and the annular chamber (14) via an intermediate flow path (48), and

wherein the annular chamber (14) defines a chamber mid-annulus (98) centrally positioned within the annular chamber (14), wherein the chamber mid-annulus (98) defines a cross sectional shape of an enclosed area (162), wherein a projection of the enclosed area along a longitudinal axis (94) of the annular chamber (14) defines an enclosed volume, and

wherein a cone flow path axis (190) defined by the cone flow path (46) pierces the enclosed volume and also intersects an intermediate flow path axis (192) defined by the intermediate flow path (48).

17. The ducting arrangement (10) of claim 16, wherein an intermediate flow path axis (192) does not pierce the enclosed volume.

18. The ducting arrangement (10) of claim 16, wherein the enclosed volume defines a tangent (1 10) where the cone flow path axis (190) pierces the enclosed volume, and wherein the cone flow path axis (190) forms a cone radial skew angle (100) of up to ten degrees with the tangent (1 10).

19. The ducting arrangement (10) of claim 16, wherein the enclosed volume defines a tangent (1 10) where the cone flow path axis (190) pierces the enclosed volume, and wherein the cone flow path axis (190) forms a cone radial skew angle (100) of more than two degrees with the tangent (1 10).

20. The ducting arrangement (10) of claim 16, wherein at least one of the cone (16) and a respective IEP (18) forms an accelerating geometry (42) effective to accelerate combustion gases to a speed acceptable for delivery onto the turbine blades.

Description:
A DUCTING ARRANGEMENT FOR DIRECTING COMBUSTION GAS

FIELD OF THE INVENTION

The invention related to a ducting arrangement for a can annular gas turbine engine that is free of a first row of turning vanes between the combustor and the first stage of turbine blades.

BACKGROUND OF THE INVENTION

Conventional can annular gas turbine engines include several individual combustor cans that are disposed radially outside of and axially aligned with a rotor shaft. Combustion gases produced in the combustor are guided radially inward and then transitioned to axial movement by a transition duct. Turning vanes receive the combustion gases and accelerate and turn them to a vector appropriate for delivery onto a first stage of turbine blades.

A recent ducting structure dispenses with the turning vanes by creating straight flow paths from reoriented combustors to a common, annular chamber, and then directly onto the first stage of turbine blades. One configuration of such a ducting structure is disclosed in U.S. Patent Number 8,276,389 to Charron et al. and which is incorporated by reference herein in its entirety. Eliminating the turning vanes reduces aerodynamic losses associated with the turning, while the ducting structure itself requires less cooling fluid than the conventional transition ducts. The ducting structure thereby improves engine performance. Another configuration of the ducting structure is disclosed in U.S. Patent Number 8,230,688 to Wilson et al. which is incorporated by reference herein in its entirety.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is explained in the following description in view of the drawings that show:

FIG. 1 shows an exemplary embodiment of a ducting arrangement composed of plural flow directing structures as seen looking fore to aft.

FIG. 2 shows the ducting arrangement of FIG. 1 as seen looking aft to fore. FIG. 3 highlights an individual flow directing structure within the ducting arrangement of FIG. 1 looking from fore to aft.

FIG. 4 shows the flow directing structure of FIG. 3 looking radially inward along line A-A of FIG. 3.

FIG. 5 schematically represents the geometry of the ducting arrangement shown in FIG. 3, also showing streamlines within a flow of combustion gases.

FIG. 6 schematically represents the geometry of the ducting arrangement shown along line D-D of FIG. 5, also showing the streamlines of FIG. 5.

FIG. 7 schematically represents the flow lines of FIG. 5.

FIG. 8 schematically represents the flow lines of FIG. 6.

FIG. 9A schematically represents the flow lines of FIG. 8 along line E-E of FIG. 8. FIG. 9B is a close-up of FIG. 9A.

FIG. 10 schematically represents the flow lines of FIG. 9A along line F-F of FIG.

9A.

FIG. 1 1 is a graph showing predicted improvement in flow angle when compared to the prior art.

FIG. 12 is a graph showing predicted improvement in the mass flow when compared to the prior art.

FIG. 13 depicts the canted cone and associated combustor superimposed on an uncanted cone and its associated combustor.

DETAILED DESCRIPTION OF THE INVENTION

The present inventors have recognized that some ducting structures for reoriented combustors have been designed with a high degree of interest in optimizing aerodynamics where the discrete flows merge into a singular flow. For example, in some configurations initially large and round flow paths must blend smoothly into a single, annular chamber, and so the ducting arrangement has been designed to optimize the flow interactions at the merge. However, the inventors have also recognized that a ducting arrangement designed to optimally blend the flows at the merge may create a blended flow that is not optimal for the downstream environment. The inventors propose a ducting arrangement herein that may sacrifice optimal blending where the flows merge in order to deliver a blended flow that is better suited for the downstream environment. The losses incurred where the flows merge are offset by gains realized downstream, resulting in a net gain.

FIGS. 1 -2 show the innovative ducting arrangement 1 0 that includes a flow path 12 for each combustor can (not shown). The flow path 12 is configured to deliver combustion gases formed in the combustor to a first stage of turbine blades (not shown) without intervening turning vanes. Each flow path 12 includes a cone 16 and an integrated exit piece (IEP) 18, where the cone 16 is canted with respect to the IEP 18. Each cone 16 has a cone inlet 26 having a circular cross section and configured to receive the combustion gases from a combustor outlet. The cone 16 narrows toward a cone outlet 28 that is associated with an IEP inlet 30 in fluid communication with each other. In the configuration shown the cone outlet 28 and the IEP inlet 30 may both be characterized by a circular cross section. In such a configuration any morphing of the cross sectional shape of a flow area (at locations other than at the interface of the cone 16 and the IEP 18) may then occur downstream of this cone to IEP interface 32.

Alternately, the morphing of the cross sectional shape may occur at or upstream of the cone to IEP interface 32.

The IEP 18 includes an IEP intermediate portion 34 and an annular chamber end 36. Combustion gases enter the I EP intermediate portion 34 discrete from combustion gases from other combustors and fully bounded (bounded on all sides) by physical walls. Within the IEP intermediate portion 34 those combustion gases transition to being only partially bounded by physical walls. Combustion gases in the annular chamber end 36 are not fully bounded and are merged with combustion gases from other combustors to form a unified flow of combustion gases in the annular chamber 14. The annular chamber 14 is defined by an annular duct 38 that is formed when all of the annular chamber ends 36 are assembled together as shown.

Together each cone 16 and respective IEP 18 define the flow path 12

therethrough. Each flow path 12 starts at the cone inlet 26 and terminates at on outlet of the annular duct 38. At the cone inlet 26 each flow path 12 is discrete from the other flow path 12. In each I EP 18 physical barriers between the adjacent flow path 12 disappear until the flow path 12 are no longer distinct, at which point the flow paths form a single, annular flow defined by the annular chamber 14 inside the annular duct 38.

The cone 16 and the IEP 18 reduce the flow area between the cone inlet 26 and the annular chamber 14, and this reduction forms an accelerating geometry 42 effective to increase a speed of the combustion gases to a speed appropriate for delivery directly onto a first stage of turbine blades (not shown). In an exemplary embodiment the constriction in area is approximately 6:1 . However, this ratio may vary in this configuration or in other configurations. For example, in another combustion system or different basket counts, the ratio might be higher or lower. A range of ratios expected includes 5:1 to 10:1 . The orientation of the cones 16 and the lEPs 18 together with the speed imparted by the accelerating geometries 42 provide an annular flow in the annular chamber 14 that is suitable for delivery directly onto the first stage of turbine blades without intervening turning vanes.

Optionally included in each flow path 12 may be a throat 44 configured to collimate the flow of combustion gases. This may be accomplished by having a perimeter of a constant size and shape for a certain distance. For irregular cross sectional shapes, the size may be a hydraulic diameter, and the length may be at least ten percent of the diameter or hydraulic diameter.

The cone 16 includes a cone flow path 46. An IEP 18 secured to the cone 16 includes an intermediate flow path 48, and an annular chamber end flow path 50. The cone may fully bound the combustion gases and the cone flow path 46 defines a cone flow path trajectory 52. The IEP 18 may fully bound the combustion gases at an IEP intermediate portion upstream end 54 and may partially bound the combustion gases at an IEP intermediate portion downstream end 56. The fully bounded IEP intermediate portion upstream end 54 may define an intermediate flow path trajectory 58. The annular chamber end flow path 50 may partially bound the combustion gases and may define an annular chamber end flow path trajectory 60. In the exemplary embodiment shown the annular chamber end 36 defines the annular chamber end flow path 50 of an adjacent flow of combustion gases. Accordingly, in this exemplary embodiment, the flow of combustion gases through the selected IEP intermediate portion 34 flows into the annular chamber end flow path 50 of a downstream adjacent IEP 18. Consequently, in the exemplary embodiment shown, two flows of combustion gases enter and given IEP 18: the first is a flow of combustion gases from the cone 16 attached to the IEP 18 and into the IEP intermediate portion 34; and the second is a flow from an adjacent I EP 18 and into the annular chamber end flow path 50. Within the IEP 18 a barrier between the flows disappears as the flow from the cone 16 attached to the IEP 18 transitions from being fully bounded to being partially bounded. It is in this manner that discrete flows enter the respective lEPs 18 while one flow exits the annular chamber 14. The embodiment shown is not meant to be limiting, and various IEP configurations may define a portion of one or more adjacent flows, or may define a portion of part or none of an adjacent flow.

Within the cone flow path 46 the cone flow path trajectory 52 may be defined as a line that connects centroids of each cross section of the cone flow path 46. In the cone flow path 46 this "centroidal" cone flow path trajectory 52 may be straight. Within the IEP intermediate portion 34 the intermediate flow path trajectory 58 may be defined as a line that connects centroids of each section in the fully bounded IEP intermediate portion upstream end 54 of the intermediate flow path 48. In the intermediate flow path 48 this "centroidal" intermediate flow path trajectory 58 may be straight. Here it can be seen that the cone flow path 46 is angled with respect to the intermediate flow path 48 so that the cone flow path trajectory 52 and the intermediate flow path trajectory 58 are not aligned with each other, but instead they are radially angled (with respect to a longitudinal axis of the gas turbine engine) with respect to each other. Therefore, the cone flow path trajectory 52 and the intermediate flow path trajectory 58 form a non- straight flow path centroid 80.

FIG. 3 highlights an individual flow directing 70 structure within the ducting arrangement 1 0 of FIG. 1 looking from fore to aft. A flow directing structure 70 includes one cone 16 and the IEP 18 to which the cone 16 is secured. Also visible in FIG. 3 is a downstream adjacent IEP 72. FIG. 4 shows the flow directing structure 70 of FIG. 3, looking radially inward along line A-A of FIG. 3. From this it can be seen that combustion gases exit a given combustor (not shown) and flow along a respective flow path 12 through the associated cone 16, then through the associated IEP 18, and then into annular chamber end flow path 50 of the downstream adjacent IEP 72. The combustion gases exit the annular chamber 14 at an annular chamber outlet 74 that may be an outlet plane 76. The IEP 18 and the downstream adjacent IEP meet at an IEP interface 78.

In the shown exemplary embodiment, the cone flow path trajectory 52, the intermediate flow path trajectory 58, and the annular chamber end flow path trajectory 60 are different from each other as can be seen in FIG. 3. In FIG. 4, looking from radially outside inward, these differences (relative angles) are not visible.

Consequently, the cone flow path trajectory 52, the intermediate flow path trajectory 58, and the annular chamber end flow path trajectory 60 appear to be the same line, though they diverge when looked at from the different perspective of FIG. 3. In the FIG. 4 view it can be seen that the non-straight flow path centroid 80 forms an intersection angle 82 with the outlet plane 76. A complement of the intersection angle 82 is an exit angle 84. While the angle between the cone flow path trajectory 52 and the intermediate flow path trajectory 58 are not visible in FIG. 4 in this exemplary embodiment, it is understood that an angle between the cone flow path trajectory 52 and the intermediate flow path trajectory 58 may exist in this view in addition to the angle that is visible in FIG. 3.

For further simplicity the remaining explanation relies on schematic

representations of FIGS. 3-4. Accordingly, FIGS. 5 schematically represents geometry of the flow directing structure of FIG. 3 in the same view of FIG. 3 showing streamlines of a flow of combustion gases therein. FIG. 6 schematically represents geometry of the flow directing structure of FIG. 4 along the same view of FIG. 4 showing the streamlines of FIG. 5.

Visible in FIG. 5 are an inner diameter 90, an outer diameter 92, an annular chamber longitudinal axis 94, a reference radial 96, and a chamber mid-annulus 98 of the annular chamber 14. In exemplary embodiments where the inner diameter 90 and the outer diameter 92 are circular, the chamber mid-annulus 98 is circular. In exemplary embodiments where the inner diameter 90 and/or the outer diameter 92 are not perfectly circular, (e.g. the diameters' surfaces include contours etc.), the chamber mid-annulus 98 may have an irregular shape that locally follows the contours of the inner diameter 90 and the outer diameter 92. The chamber mid-annulus 98 may have an axial length as long as an axial length of the annular chamber 14. Should the annular chamber outlet 74 be considered an axial end of the annular chamber 14, the chamber mid-annulus 98 may be an annulus connecting centroids of cross sections of the annular chamber 14. The chamber mid annulus 98 may be circular. Alternately, the chamber mid annulus 98 may not be perfectly circular throughout its entire

circumference, but may have local variations in shape due to local curvatures in the inner diameter 90 and/or the outer diameter 92.

In this exemplary embodiment a circumferential component 180 of the intermediate flow path trajectory 58 is collocated with a tangent 1 10 of the chamber mid-annulus 98 taken where the intermediate flow path trajectory 58 is at a same radial distance 1 60 from the annular chamber longitudinal axis 94 as the chamber mid- annulus 98 (e.g. where they intersect). Thus, a circumferential component 182 of the cone flow path trajectory 52 forms a positive, non-zero cone radial skew angle 100 with the tangent 1 10 (as well as with the circumferential component 182 of the intermediate flow path trajectory 58), while the circumferential component 180 of the intermediate flow path 48 is collocated with the tangent 1 10 where the intermediate flow path 48 and the tangent 1 10 intersect. (E.g. the circumferential component 180 of the intermediate flow path 48 overlies the tangent 1 10 in FIG. 5.) The circumferential component 180 of the intermediate flow path 48 intersects the tangent 1 10 at an intersection point 86 having an axially, radially, and circumferentially identifiable position.

Stated another way, the cone flow path 46 defines a cone flow path axis 190 that forms a cone radial skew angle 100 with the tangent 1 10 of the chamber mid-annulus 98. An intermediate flow path axis 192 is defined by the intermediate flow path trajectory 58. The tangent 1 10 is formed where the intermediate flow path axis 192 intersects the chamber mid-annulus 98. Stated in yet another way, when viewed looking from fore to aft along the annular chamber longitudinal axis 94, the cone flow path axis 190 intersects the intermediate flow path 48 (and the tangent 1 10) to form the cone radial skew angle 100. Since the annular chamber longitudinal axis 94 and the gas turbine engine longitudinal axis 134 are the same, it can also be stated that when viewed looking from fore to aft along the gas turbine engine longitudinal axis 134 the cone flow path axis 190 intersects the intermediate flow path 48 (and the tangent 1 10) to form the cone radial skew angle 100. When the cone radial skew angle 100 is positive and non-zero, it necessarily follows that the cone flow path axis 1 90 must pass radially inward of the chamber mid- annulus 98 as is shown in FIG. 1 1 . A complement of the cone radial skew angle 100 is a reference radial angle 158. Accordingly, a cone radial skew angle 100 of, for example, ten degrees, would provide a reference radial angle 158 of eighty degrees.

The cone radial skew angle 100 may be any positive angle desired. In various exemplary embodiments the skew angle may be ten degrees or less (e.g. a reference radial angle of eighty degrees or more), seven degrees or less (e.g. a reference radial angle of eighty-three degrees or more), five degrees or less (e.g. a reference radial angle of eighty-five degrees or more), and may be over two degrees (e.g. a reference radial angle of less than eighty-eight degrees). A skew angle of over two degrees overcomes any machining tolerances etc. that might be found in the prior art seeking a zero degree cone radial skew angle 100. In an embodiment, a range of cone radial skew angles 100 may be three to seven degrees.

FIG. 6 shows the non-straight flow path centroid 80, the cone flow path axis 190, and the intermediate flow path axis 192 from the perspective of FIG. 4. The cone radial skew angle 100 is not visible in this view because the skew is out of the page.

With the geometry established, discussion can turn to the flow of combustion gases. The cone flow path axis 190 is, by definition, essentially a center of the flow of combustion gases flowing in the cone 16. Thus, it represents a middle streamline 122 of a central slice 124 of the flow of combustion gases flowing through the middle of the cone 16 and then the associated IEP 18. An upstream end 126 of the slice 124 is parallel to the reference radial 96. The upstream end 126 of the slice 124 includes the middle streamline 122 (in the middle), a tip streamline 128 at a radially outside end of the slice 124, and a hub streamline 1 30 at a radially inside end (a hub end) of the slice 124. The slice 124 remains planar inside the cone 1 6 until reaching the IEP interface 32. For sake of explanation, the slice will remain planar in the IEP 18 until the middle streamline 122 intersects the reference radial 96. In other words, until the middle streamline 122 intersects the reference radial 96, at any given location the three streamlines form a straight line that is parallel to a reference radial 96. From that point forward each streamline is locally influenced and their paths may differ from each other enroute to the first stage of turbine blades. In the exemplary embodiment described herein, the tip streamline 128 and the hub streamline 130 orbit the middle streamline 122 as the streamlines move helically downstream enroute to the first stage of turbine blades. The orbit appears to be clockwise when looking downstream at the streamlines from upstream of the streamlines. Each streamline can be thought of as a path taken by a respective molecule in the flow of combustion gases.

FIGS. 7-10 are simplified schematics that show the divergence of the streamlines after they pass through the reference radial 96. FIG. 7 shows the streamlines of FIG. 5. In addition, in FIG. 7 the chamber mid-annulus 98 may be seen as defining a cross sectional shape of an enclosed area 162 identified in FIG. 13 with hash marks. An extension of the enclosed area 162 along the annular chamber longitudinal axis 94 (i.e. in and out of the page) defines an enclosed volume 164. Since the cone flow path axis 190 passes radially inward of the chamber mid-annulus, it necessarily pierces the enclosed volume 164. The axial location of where the cone flow path axis 190 pierces the enclosed volume can be selected as desired. The tangent 1 10 against which the cone radial skew angle 100 is measured is taken axially and circumferentially at the location where the intermediate flow path axis 192 reaches/contacts the chamber mid- annulus 98.

FIG. 8 shows the streamlines of FIG. 6 as they travel circumferentially and axially toward the first stage of turbine blades. The inventors have recognized that after passing through the reference radial 96, the tip streamline 128 begins to overturn while the hub streamline 1 30 begins to underturn with respect the respective intersection angle 82 (and necessarily the respective exit angle 84) at the reference radial 96. As used herein, overturning means that there is more circumferential travel per unit of axial travel as there was at the reference radial 96. Conversely, as used herein, underturning means that there is less circumferential travel per unit of axial travel as there was at the reference radial 96. Axial travel is shown in FIG. 8 as axial direction 132 along the annular chamber longitudinal axis 94, which is common with a gas turbine engine longitudinal axis 1 34. FIG. 8 shows this underturning and overturning schematically during a hypothetical ninety degree turn. Ninety degrees is only used for illustrative purposes and in reality may be more or fewer degrees. Such overturning and underturning in the flow of combustion gases may not provide the best results when the flow interacts with turbine blades which may be optimized for a more uniform radial distribution of the flow of combustion gases (where a uniform radial distribution means no overturning or underturning throughout a flow as the flow reaches the turbine blades).

For illustrative purposes the middle streamline 122 is considered to maintain a consistent intersection angle 82 (and respective exit angle 84) as it travels in the axial direction 132. In contrast, the tip streamline 128 begins to overturn as it travels in the axial direction 132 along the gas turbine engine longitudinal axis 134. Accordingly, for a given amount of axial travel the tip streamline 128 travels more circumferentially than the middle streamline 122. The hub streamline begins to underturn as it travels in the axial direction 132. Accordingly, for a given axial amount of axial travel the hub streamline 130 travels less circumferentially than the middle streamline 122.

This may seem counter-intuitive, given than the tip streamline 128 is at a greater radius from the annular chamber longitudinal axis 94, and hence the tip streamline has a greater arc-length to travel for the given axial length. However, this is a highly complex fluid environment having multiple factors influencing the paths the streamlines take, including inertia that tends to move the combustion gases radially outward, friction, discrete gas flows that are uniting with adjacent flows at angle to each other, pressures etc.

The differences in circumferential travel can be seen in FIG. 9A, which shows the streamlines of FIG. 8 as seen from line B-B. In FIG. 9A, the underturning hub streamline 130 cover less distance in the circumferential direction 138 than does the middle streamline 122 for the same axial distance 136. The overturning tip streamline covers more distance in the circumferential direction 138 than does the middle streamline 122 for the same axial distance 136. A circumferential distance 152 therefore exists between the hub streamline 130 and the tip streamline 128 where they are at the same axial distance 136. While there is still a circumferential distance 152, it is smaller than a prior art circumferential distance 1 52' that occurs when the cone flow path axis 190 and the intermediate flow path axis 192 form a straight flow path axis having a radial skew angle of zero (e.g. when the cone flow path axis 190 and the intermediate flow path axis 192 form a straight flow path axis that overlies the tangent 1 10 when viewed from the perspective shown in FIG. 5).

. FIG. 9B is a close up of FIG. 9A. A middle streamline intersection angle 144 and a middle streamline exit angle 146 are the same as the intersection angle 82 and the exit angle 84 at the reference radial 96 in FIG. 8. The hub streamline intersection angle 140 is increased and the hub streamline exit angle 142 is decreased when compared to the intersection angle 82 and the exit angle 84 at the reference radial 96 in FIG. 8. Conversely, the tip streamline intersection angle 148 is decreased and the tip streamline exit angle 1 50 is increased when compared to the intersection angle 82 and the exit angle 84 at the reference radial 96 in FIG. 8. However, these latter two respective angles are closer to the intersection angle 82 and the exit angle 84 at the reference radial 96 than occurs in the prior art where the cone flow path axis 190 and the intermediate flow path axis 192 form a straight flow path axis having a radial skew angle of zero (e.g. when the cone flow path axis 1 90 and the intermediate flow path axis 192 form a straight flow path axis that overlies the tangent 1 10 when viewed from the perspective shown in FIG. 5)..

It is understood that the tip and hub streamline angles shown in FIG. 9B would be most accurate if taken at the annular chamber longitudinal axis 94. However, these streamlines from this slice 124 are not so located due to the underturning and overturning. Tip and hub streamlines from adjacent combustion gas molecules may be so located, but these are not shown for sake of clarity. Consequently, for illustrative purposes these streamline angles are used in hopes of conveying the necessary information.

FIG. 10 shows the streamlines of FIG. 9A along line C-C of FIG. 9A. The circumferential distance 152 can be seen. Further, it can be seen that the slice 124 is no longer radially oriented, but it is tilted relative to a local radial 154 through the middle streamline 122. While there is still a circumferential distance 152, it is smaller than a prior art circumferential distance 1 52' that occurs when the cone flow path axis 190 and the intermediate flow path axis 192 form a straight flow path axis having a radial skew angle of zero (e.g. when the cone flow path axis 1 90 and the intermediate flow path axis 192 form a straight flow path axis that overlies the tangent 1 10 when viewed from the perspective shown in FIG. 5).

The closer the slice 124 is to being parallel with the local radial 154, the more uniform the radial distribution of the flow of combustion gases becomes. Accordingly, the geometry disclosed herein provides for a more uniform radial distribution of the combustion gases than does the prior art geometry. A more uniform radial distribution provides numerous advantages in terms of how the combustion gases interact with the first stage of turbine blades etc., and this results in an improvement in engine efficiency.

Also visible is a distance 156 between the middle streamline 122 and the inner diameter 90 of the annular chamber 14. This schematically represents that the non-zero cone radial skew angle 100 may be directing the streamlines on respective trajectories that "cut the corner", where the corner is the inner diameter 90 of the annular chamber 14. In other words, the non-zero cone radial skew angle 100 directs the flow of combustion gases slightly radially inward toward the inner diameter 90 of the annular chamber 14 (e.g. toward the hub). This helps to counter the inertia and other factors that tend to move the combustion gases radially outward and otherwise influence the flow of combustion gases. It is understood that some aerodynamic loss may occur as a result of directing the flow of combustion gases this way, (e.g. turning losses), but the losses incurred are more than offset by the benefit gained downstream when a more uniformly radially distributed flow is realized. Accordingly, the net effect is still an improvement in engine efficiency.

FIG. 1 1 is a graph showing prior art predicted exit angles 168, prior art measured exit angles 170, and the predicted exit angles 172 along a radial span at a given axial location. The prior art predicted exit angles 168 and the prior art measured exit angles 170 are fairly consistent with each other, lending confidence to the predicted exit angles 172. The predicted exit angles 172 show a much more uniform exit angle 84

throughout most of the radial span when compared to the prior art predicted exit angles 168, prior art measured exit angles 170.

FIG. 12 is a graph showing the prior art mass flow 174 and the predicted mass flow 176 along a radial span at a given axial location. The predicted mass flow 176 shows a much more uniform mass flow throughout most of the radial span when compared to the prior art mass flow 174.

In addition to the aerodynamic benefits realized at the downstream location, radially canting the cone 16 with respect to the lEP 18 allows for a shorter cone 16 and/or lEP intermediate portion 34. FIG. 13 shows a canted cone 200 and a canted combustor 202 superimposed on an uncanted cone 204 as well as its uncanted combustor 206. The uncanted cone 204 and the uncanted combustor 206 are aligned with the intermediate flow path axis 192 as occurs in the prior art where the cone flow path axis 190 and the intermediate flow path axis 192 form a straight flow path axis.

The cone inlet 26 is characterized by a cone inlet diameter 210. Discussion herein focuses on shortening of the cone 16. The lEP intermediate portion 34 may or may not be shortened as well. Any shortening of the cone 16 is limited by the cone inlet diameter 210. If a length of the cone 16 is reduced, the cone inlet diameter 210 eventually reaches an interference point 220 where the cone 16 contacts an adjacent lEP 212. Here the uncanted cone 204 is in a theoretically shortest uncanted position 222. When wall thicknesses, flanges, and attachment features are considered, a realistically shortest uncanted position 224 is more practical. This leaves an uncanted minimum cone length 226 from the cone inlet 26 to the cone to lEP interface 32.

When the cone 16 is canted as disclosed above, the cone flow path axis 190 rotates clockwise in this view, pivoting at, for example, a rotation point 228 where the cone flow path axis 190 of the uncanted cone 204 intersects the cone to lEP interface 32 (e.g. a centroid of the lEP inlet). While this rotation point 228 is used in this exemplary embodiment, any point in the cone to lEP interface 32 (or in the lEP inlet 30) could be used as the rotation point, as could any other point deemed appropriate, and the interfacing geometry of the cone 16 and/or of the lEP 18 can be modified as necessary to accommodate the reoriented canted cone 200.

The canted cone 200 is thereby moved away from the adjacent lEP 212 (in an upward direction in this view). This changes a location of the interference point 220 where the cone 16 contacts the adjacent l EP 212. Here the canted cone 200 is in a theoretically shortest canted position 230. When wall thicknesses, flanges, and attachment features are considered, a realistically shortest canted position 232 is more practical. This leaves a canted minimum cone length 234 from the cone inlet 26 to the cone to IEP interface 32. The canted minimum cone length 234 is shorter than the uncanted minimum cone length 226, and the difference is made possible because of the cant of the canted cone 200.

Shortening the cone 16 any amount reduces the amount of cooling air used to cool the cone 16. Accordingly, the additional shortening that canting allows simply reduces further the amount of cooling air used, thereby increasing engine efficiency.

A shorter cone 16 also pulls the combustor canted combustor 202 radially inward with respect to the uncanted combustor 206. Since the combustors dictate the size of a combustion section outer casing (of the gas turbine engine) and associated

components, moving the combustors inward means that the combustion section outer casing can be reduced in diameter, and this saves in materials costs. The combustion section outer casing contains compressed air at high pressures. Since the force seen by the combustion section outer casing is associated with a square of its diameter, any reduction in the diameter yields a substantial reduction in force. The reduced strength requirements reduces further the amount of material needed to make the combustion section outer casing, further saving material costs.

In the exemplary embodiment of FIG. 13 the cone 16 is rotated clockwise within the plane of the page, described above as radially outward. Alternately, or in addition, the cone 16 may be rotated axial ly, into the page and away from the adjacent IEP 212. This increases an axial distance (i.e. along the gas turbine engine longitudinal axis 134) between the cone 16 and the adjacent IEP 212 much the same as canting the cone 16 radially increases a radial distance between the cone 16 and the adjacent IEP 212.

Canting the cone 16 in both the radial direction and the axial direction creates a compound canted cone (not shown). There is a physical and/or aerodynamic limit to the amount of radial canting that can be done. Likewise, there is a physical and/or aerodynamic limit to the amount of axial canting that can be done. Therefore, a compound canted cone taking advantage of the radial and the axial cant may have a greater distance between it and the adjacent I EP 212 than either the radially canted cone 200 or an axially cant cone alone could. To maximize the distance the compound cant would move the uncanted cone 204 in a compound cant direction perpendicular to an interference point. However, as shown above, the location of the interference point 220 changes as the amount of the cant changes. Similarly, the location of the interference point 220 changes as the direction of the cant changes. As the location of the interference point 220 changes, a direction perpendicular to the interference point 220 may also change due to the three- dimensionality and/or local contour changes in the cone 16 and/or the adjacent IEP 212 etc. Thus, the orientation of the compound cant determines the location of the interference point, and the location of the interference point determines the best orientation of the compound cant, making the two interdependent. Therefore, selection of the optimum compound cant may require consideration of several factors to maximize the distance between the compound canted cone and the adjacent IEP 212, and thereby maximize the cone length reduction.

In light of the foregoing it has been shown that the inventors have identified a source of inefficiency in the gas turbine engine and have provided a counter-intuitive solution that involves reducing aerodynamic efficiency at the ducting structure in order to provide an aerodynamic improvement of greater value at another location.

Consequently, this represents an improvement in the art.

While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.