WO2017023331A1

WO2017023331A1 - Ducting arrangement for directing combustion gas

Info

Publication number: WO2017023331A1
Application number: PCT/US2015/043983
Authority: WO
Inventors: Jay A. Morrison; Richard C. Charron; Manish Kumar; Gary D. Snyder
Original assignee: Siemens Aktiengesellschaft; Siemens Energy, Inc.
Priority date: 2015-08-06
Filing date: 2015-08-06
Publication date: 2017-02-09

Abstract

A ducting arrangement (10), including: an annular duct (38) having a plurality of discrete integrated exit pieces (IEP 18) secured together to form the annular duct and an annular chamber (14) configured to deliver an annular flow of combustion gases directly onto turbine blades, 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) and defining a respective straight flow path axis (80); and a cone (16) for each IEP intermediate portion, each cone defining a cone flow path (46) and sharing the respective straight flow path axis. When viewed from upstream toward downstream along a longitudinal axis (134) of the gas turbine engine the straight flow path axis intersects the chamber mid-annulus.

Description

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. 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, and canted with respect to an annular chamber 14 (as opposed to delivering the flow tangential to the common chamber 14). Each flow path 12 includes a cone 16 and an integrated exit piece (IEP) 1 8. 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 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 bound 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. Further, the cone flow path trajectory 52 and the intermediate flow path trajectory 58 may form a single, straight flow path axis 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.

The straight flow path axis 80 in FIG. 3 can be seen diverging from the annular chamber end flow path trajectory 60 where the annular chamber end flow path trajectory 60 begins to curve. In FIG. 4, looking from radially outside inward, the curve of the annular chamber end flow path trajectory 60 is not visible. Consequently, the annular chamber end flow path trajectory 60 and the straight flow path axis 80 appear to be the same line, though they diverge when looked at from a different perspective. Here it can be seen that the straight flow path axis 80 forms an intersection angle 82 with the outlet plane 76. A complement of the intersection angle 82 is an exit angle 84.

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 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 mid annulus 98 may be circular. Alternately, the 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.

Also visible is the straight flow path axis 80 of FIGS. 3-4. It can be seen that the straight flow path axis 80 forms positive radial skew angle 100 with the chamber mid- annulus 98 where the two intersect. As defined herein, the radial skew angle 100 is an angle formed between: 1 ) a tangent 1 10 of the chamber mid-annulus 98 formed where the straight flow path axis 80 intersects the chamber mid-annulus 98, and 2) a circumferential component 1 12 of the straight flow path axis 80 (seen in FIG. 6). An axial component 1 14 of the straight flow path axis 80 (visible in FIG. 6) is not visible in FIG. 5. Accordingly, in FIG. 5, looking from upstream toward downstream along the gas annular chamber longitudinal axis 94, the radial skew angle appears to be formed between the straight flow path axis 80 and the tangent 1 10 where the two intersect. The straight flow path axis 80 intersects the tangent 1 10 at an intersection point 86 having an axially, radially, and circumferentially identifiable position.

Regarding the positive skew angle 100, and stated another way, the

circumferential component 1 12 of the straight flow path axis 80 (visible in FIG. 6) forms a radial skew angle 100 with the tangent 1 10 of the chamber mid-annulus 98 where the straight flow path axis 80 is at a same radial distance 160 from the chamber mid- annulus 98 (e.g. where they intersect). Since the radial skew angle 100 is non-zero and positive, it necessarily follows that the straight flow path axis 80 must pass radially inward of the chamber mid-annulus 98 as is shown in FIG. 5. A complement of the radial skew angle 100 is a reference radial angle 158. Accordingly, a radial skew angle 100 of, for example, ten degrees, would provide a reference radial angle 158 of eighty degrees.

The 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 radial skew angle 100. In an embodiment, a range of skew angles 100 may be three to seven degrees.

With the geometry established, discussion can turn to the flow of combustion gases. The straight flow path axis 80 is, by definition, a center of the flow of combustion gases flowing in the cone 16 and the fully bounded IEP intermediate portion upstream end 54. 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 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 130 at a radially inside end (a hub end) of the slice 124. The slice 124 remains planar 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 straight flow path axis 80 passes radially inward of the chamber mid-annulus, it necessarily pierces the enclosed volume 164. The axial location of where the straight flow path axis 80 pierces the enclosed volume can be selected as desired. The tangent 1 10 against which the radial skew angle 1 00 is measured is taken axially and circumferentially at the location where the straight flow path axis pierces the enclosed volume 164.

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 radial skew angle 100 is zero.

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 when the skew angle 100 is zero as in the prior art.

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 the prior art circumferential distance 152' that occurs when the radial skew angle 100 is zero. 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 tip streamline 128 and the inner diameter 90 of the annular chamber 14. This schematically represents that the non-zero 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 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 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.

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 a respective straight flow path axis (80); and

a cone (16) for each IEP intermediate portion (34), each cone (16) defining a cone flow path (48) in fluid communication with a respective IEP intermediate portion (34) and sharing the respective straight flow path axis (80), 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 the straight flow path axis (80) intersects the chamber mid-annulus (98).

2. The ducting arrangement (10) of claim 1 , wherein the chamber mid- annulus (98) defines a respective tangent (1 10) where the respective straight flow path axis (80) passes radially inward of the chamber mid-annulus (98), and wherein when viewed from upstream toward downstream along the longitudinal axis (134) of the gas turbine engine the respective straight flow path axis (80) forms a respective radial skew angle (100) of up to ten degrees with the respective tangent (1 10).

3. The ducting arrangement (10) of claim 1 , wherein the chamber mid- annulus (98) defines a respective tangent (1 10) where the respective straight flow path axis (80) passes radially inward of the chamber mid-annulus (98), and wherein when viewed from upstream toward downstream along the longitudinal axis (134) of the gas turbine engine the respective straight flow path axis (80) forms a respective radial skew angle (100) of over two degrees with the respective tangent (1 10).

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

5. The ducting arrangement (10) of claim 1 or 4, wherein a flow area of a respective intermediate flow path (48) is less than a flow area of a respective cone flow path (48) at a cone outlet (28).

6. The ducting arrangement (10) of claim 1 or 4, 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 (48) that narrows from the cone inlet (26) to the cone outlet (28);

an IEP (18) comprising an IEP inlet (30) configured to secure to the cone outlet (28), an annular chamber end (36), and an IEP intermediate portion (34), wherein a fully bound portion of the IEP 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 IEP 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 the cone flow path (48) and the intermediate flow path (48) share a straight flow path axis (80), wherein a reference radial (96) of the annular chamber (14) is defined where the straight flow path axis (80) and the chamber mid-annulus (98) intersect each other, and wherein the straight flow path axis (80) is radially angled with respect to the reference radial (96).

9. The ducting arrangement (10) of claim 8, wherein the straight flow path axis (80) forms a reference radial angle (158) of eighty degrees or more with the reference radial (96).

10. The ducting arrangement (10) of claim 9, wherein the straight flow path axis (80) forms a reference radial angle (158) of less than eighty-eight degrees with the reference radial (96).

1 1 . The ducting arrangement (10) of claim 8, wherein the cone (16) defines a circular cross sectional shape for the cone flow path (48) 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 or 1 1 , wherein a flow area of the intermediate flow path (48) is less than a flow area of the cone flow path (48) at the cone outlet (28).

13. The ducting arrangement (10) of claim 8 or 1 1 , where at least one of the cone (16) and the lEP (18) forms an accelerating geometry (42) effective to accelerate combustion gases to over Mach 0.6.

14. The ducting arrangement (10) of claim 8 or 1 1 , where at least one of the cone (16) and the lEP (18) forms a collimating throat (44).

15. The ducting arrangement (10) of any of claims 8-14, further comprising one cone (16) and one lEP (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 a cone outlet (28) of the combustor can, and wherein the IEP (18) establishes fluid

communication between the cone (16) and the annular chamber (14), 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 (162) along a longitudinal axis (94) of the annular chamber (14) defines an enclosed volume, and wherein a straight axis (80) of a flow path through the cone (16) and the IEP (18) pierces the enclosed volume.

17. The ducting arrangement (10) of claim 16, wherein the enclosed volume defines a tangent (1 10) where the straight flow path axis (80) pierces the enclosed volume, and wherein the straight flow path axis (80) forms a radial skew angle (100) of up to ten degrees with the tangent (1 10).

18. The ducting arrangement (10) of claim 16, wherein the enclosed volume defines a tangent (1 10) where the straight flow path axis (80) pierces the enclosed volume, and wherein the straight flow path axis (80) forms a radial skew angle (100) of more than two degrees with the tangent (1 10).

19. 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.

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