Turbine center frame and method

Abstract

Aspects of the present disclosure generally relate to a turbine center frame for a turbine engine, a flow path extending through the turbine center frame. The turbine center frame may include an inner wall radially spaced from an outer wall, wherein the inner wall and the outer wall extend between an inlet and an outlet, and wherein the outlet is downstream of the inlet relative to the flow path. A set of circumferentially spaced airfoils may extend between the inner wall and the outer wall.

Description

Technical Field

[0001] The present disclosure relates generally to frames for turbine engines and, more particularly, to turbine center frames and methods of varying airflow in turbine engines. Background Art

[0002] A turbine engine, particularly a gas or combustion turbine engine, is a rotary engine that extracts energy from the flow of combustion gases through the engine onto a plurality of rotating turbine blades.

[0003] A gas turbine engine may include a compressor for compressing a gas flow and a turbine for driving the compressor using the compressed gas flow. The turbine may be divided into a high-pressure turbine and a low-pressure turbine, wherein a turbine center frame positioned between the high-pressure turbine and the low-pressure turbine serves as a conduit for fluid to flow from the high-pressure turbine to the low-pressure turbine. The turbine center frame may provide for changing the fluid flow between the high-pressure turbine and the low-pressure turbine. Summary of the invention

[0004] Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.

[0005] In one aspect, the present disclosure relates to a turbine center frame for a turbine engine, a flow path extending through the turbine center frame. The turbine center frame includes: an inner wall radially spaced from an outer wall, wherein the inner wall and the outer wall extend between an inlet and an outlet, wherein the outlet is downstream of the inlet relative to the flow path; a set of circumferentially spaced airfoils extending between the inner wall and the outer wall; a first cross-sectional area, the first cross-sectional area is defined between the inner wall and the outer wall along the flow path; a second cross-sectional area, the second cross-sectional area is larger than the first cross-sectional area and is defined between the inner wall and the outer wall along the flow path, downstream of the first cross-sectional area; and a third cross-sectional area, the third cross-sectional area is defined between the inner wall and the outer wall along the flow path, downstream of the second cross-sectional area.

[0006] In another aspect, the present disclosure relates to a turbine engine. The turbine engine comprises: an engine core, the engine core comprising a compressor section, a combustion section, and a turbine section arranged in an axial flow and defining a flow path, wherein the turbine section comprises a high-pressure turbine and a low-pressure turbine; and a turbine center frame, the turbine center frame extending from the high-pressure turbine to the low-pressure turbine, the turbine center frame comprising: an inner wall, the inner wall being radially spaced apart from the outer wall, wherein the inner wall and the outer wall extend between an inlet and an outlet, wherein the outlet is downstream of the inlet relative to the flow path; a set of circumferentially spaced airfoils, the set of circumferentially spaced airfoils extending between the inner wall and the outer wall; a first cross-sectional area, the first cross-sectional area being radially defined between the inner wall and the outer wall; a second cross-sectional area, the second cross-sectional area being larger than the first cross-sectional area and being radially defined between the inner wall and the outer wall, downstream of the first cross-sectional area; and a third cross-sectional area, the third cross-sectional area being radially defined between the inner wall and the outer wall, downstream of the second cross-sectional area.

[0007] In yet another aspect, the present disclosure relates to a method of changing an airflow along a flow path extending through a turbine center frame for a turbine engine. The method includes: directing an airflow along the flow path through a first cross-sectional area of ​​the turbine center frame at a first speed; slowing the airflow to a second speed less than the first speed by flowing the airflow through a second cross-sectional area of ​​the turbine center frame that is larger than the first cross-sectional area; and accelerating the airflow to a third speed greater than the second speed by flowing the airflow through a third cross-sectional area of ​​the turbine center frame.

[0008] These and other features, aspects and advantages of the present invention will be better understood with reference to the following description and appended claims.The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the present invention and, together with the description, serve to explain the principles of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] A full and enabling disclosure of the invention, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification, which refers to the accompanying drawings, in which:

[0010] Figure 1 is a schematic cross-sectional view of a gas turbine engine for an aircraft including an exemplary turbine center frame according to various aspects described herein.

[0011] Figure 2 yes Figure 1 Cross-sectional view of the turbine center frame along line II-II.

[0012] Figure 3 yes Figure 1 Side view of the turbine center frame, showing the inner and outer walls.

[0013] Figure 4 It is shown by Figure 1 Plot of the cross-sectional area of ​​the turbine center frame.

[0014] Figure 5 It is available for Figure 1 A side view of another turbine center frame in a turbine engine showing the inner and outer walls with partial contours. DETAILED DESCRIPTION

[0015] The aspects of the present disclosure described herein are broadly related to turbine frames and their operation and manufacture. For the purpose of illustration, an exemplary environment in which the system can be utilized will be described in the form of a turbine engine having a turbine center frame positioned between a high-pressure turbine and a low-pressure turbine. In a non-limiting example, such a turbine engine can be in the form of a gas turbine engine, a turboprop engine, a turboshaft engine, or a turbofan engine. However, it will be understood that the aspects of the present disclosure described herein are not limited thereto and can have general applicability in other engine systems. For example, the present disclosure can be applied to systems in other engines or vehicles, and can be used to provide benefits in industrial, commercial, and residential applications.

[0016] The turbine center frame (also called the turbine middle frame) is often used in the transition duct between the high-pressure turbine and the low-pressure turbine. Such a frame can provide diffusion or smoothing of the turbulent airflow before leaving the high-pressure turbine and entering the low-pressure turbine. The airflow entering the turbine center frame may produce unstable, three-dimensional and non-uniform turbulent motion. These flow characteristics usually lead to high levels of mixing and poor performance. Such a frame usually includes airfoils to divide, turn or accelerate the airflow. Accelerating the airflow can help reduce this turbulence and can be achieved by redirecting the airflow along the turning airfoils. Reducing turbulence can also be achieved by accelerating the airflow through the frame. For example, the typical minimum acceleration level across the turning vane is usually 20% between the inlet and the outlet. If the airflow acceleration is too small, the level of 3D end wall flow will usually increase, and the turbulence level in the airflow will be high. As a result, engine performance will be reduced.

[0017] As used herein, the term "upstream" refers to a direction opposite to the direction of fluid flow, and the term "downstream" refers to the same direction as the direction of fluid flow. The terms "front" or "forward" refer to in front of something, while "rear" or "rearward" refer to behind something. For example, when used in fluid flow terms, front / forward may refer to upstream and rear / rearward may refer to downstream.

[0018] Additionally, as used herein, the term "radial" or "radially" refers to a direction away from a common center. For example, in the overall context of a turbine engine, radial refers to a direction along a ray extending between a central longitudinal axis of the engine and an outer engine circumference. Additionally, as used herein, the term "group" or a "group" of elements may be any number of elements, including only one element.

[0019] All directional references (e.g., radial, axial, proximal, distal, above, below, upward, downward, left, right, lateral, front, rear, top, bottom, up, down, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, backward, etc.) are only used for identification purposes to help readers understand the present disclosure, and should not be interpreted as limitations on the embodiments, particularly limitations on the position, orientation or use of the aspects of the present disclosure described herein. Connection references (e.g., attachment, connection, fixing, connection, engagement, etc.) should be broadly interpreted, and unless otherwise noted, may include intermediate members between element sets and relative movement between elements. In this way, connection references do not necessarily infer that two elements are directly connected and are in a fixed relationship to each other. Unless the context clearly indicates otherwise, the singular forms "one", "a kind of" and "the" include plural forms.

[0020] As used herein throughout the specification and claims, approximating language is used to modify any quantitative representation that can be permitted to vary without resulting in a change in the basic function to which it is related. Thus, values ​​modified by terms such as "about," "approximately," and "substantially" are not limited to the precise values ​​specified. In at least some cases, approximate language may correspond to the precision of an instrument used to measure a value, or the precision of a method or machine used to construct or manufacture a component and / or system. For example, approximate language may refer to being within 10%.

[0021] The exemplary drawings are for illustrative purposes only, and the dimensions, positions, orders, and relative sizes reflected in the accompanying drawings may vary.

[0022] Figure 1 is a schematic cross-sectional view of a gas turbine engine 10 for an aircraft. The engine 10 has a generally longitudinally extending axis or centerline 12 extending from a front portion 14 to an aft portion 16. The engine 10 includes, in downstream serial flow relationship: a fan section 18 including a fan 20; a compressor section 22 including a supercharger or low pressure (LP) compressor 24 and a high pressure (HP) compressor 26; a combustion section 28 including a combustor 30; a turbine section 32 including an HP turbine 34 and an LP turbine 36; and an exhaust section 38.

[0023] The fan section 18 includes a fan case 40 surrounding the fan 20. The fan 20 includes a plurality of fan blades 42 radially arranged about the centerline 12. The HP compressor 26, the combustor 30, and the HP turbine 34 form a core 44 of the engine 10 that generates combustion gases. The core 44 is surrounded by a core case 46 that can be coupled to the fan case 40.

[0024] An HP shaft or spool 48 disposed coaxially about the centerline 12 of the engine 10 drivingly connects the HP turbine 34 to the HP compressor 26. An LP shaft or spool 50 disposed coaxially about the centerline 12 of the engine 10 within the larger diameter annular HP spool 48 drivingly connects the LP turbine 36 to the LP compressor 24 and the fan 20. The spools 48, 50 are rotatable about the engine centerline and are coupled to a plurality of rotatable elements that may collectively define a rotor 51.

[0025] The LP compressor 24 and the HP compressor 26 each include a plurality of compressor stages 52, 54 in which a set of compressor blades 56, 58 rotate relative to a corresponding set of static compressor vanes 60, 62 to compress or pressurize a fluid flow through the stage. In a single compressor stage 52, 54, the plurality of compressor blades 56, 58 may be arranged in an annular shape and may extend radially outward from a blade platform to a blade tip relative to the centerline 12, while the corresponding static compressor vanes 60, 62 are positioned upstream of and adjacent to the rotating blades 56, 58. Note that Figure 1 The numbers of blades, buckets, and compressor stages shown in FIG. 5 are selected for illustration purposes only, and other numbers are possible.

[0026] The blades 56, 58 for the compressor stage may be mounted to (or integrated into) a disk 61 that is mounted to a respective one of the HP spool 48 and the LP spool 50. The buckets 60, 62 for the compressor stage may be mounted to the core housing 46 in a circumferential arrangement.

[0027] The HP turbine 34 and the LP turbine 36 each include a plurality of turbine stages 64, 66, wherein a set of turbine blades 68, 70 rotate relative to a corresponding set of static turbine vanes 72, 74 (also referred to as nozzles) to extract energy from a fluid flow passing through the stage. In a single turbine stage 64, 66, the plurality of turbine blades 68, 70 may be arranged in an annular shape and may extend radially outward relative to the centerline 12, while the corresponding static turbine vanes 72, 74 are positioned upstream and adjacent to the rotating blades 68, 70. Note that Figure 1 The number of blades, buckets, and turbine stages shown in FIG. 1 are selected for illustration purposes only, and other numbers are possible.

[0028] The blades 68, 70 for the turbine stage may be mounted to a disk 71 that is mounted to a respective one of the HP spool 48 and the LP spool 50. The buckets 72, 74 for the compressor stage may be mounted to the core housing 46 in a circumferential arrangement.

[0029] In addition to the rotor portion, the stationary portion of the engine 10 (such as the static blades 60, 62, 72, 74 in the compressor section 22 and the turbine section 32) are also individually or collectively referred to as stators 63. As such, the stator 63 may refer to the combination of non-rotating elements in the entire engine 10.

[0030] In operation, the airflow leaving the fan section 18 is split so that a portion of the airflow is directed into the LP compressor 24, which then supplies pressurized air to the HP compressor 26, which further pressurizes the air. The pressurized air from the HP compressor 26 is mixed with fuel and ignited in the combustor 30, thereby producing combustion gases. The HP turbine 34 driving the HP compressor 26 extracts some work from these gases. The combustion gases are discharged through the turbine section 32 into the LP turbine 36 defining the main flow 76, which extracts additional work to drive the LP compressor 24, and the exhaust gases are finally discharged from the engine 10 via the exhaust section 38. The drive of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24.

[0031] A portion of the pressurized air stream may be extracted from the compressor section 22 as bleed air 77. The bleed air 77 may be extracted from the pressurized air stream and provided to engine components that require cooling. The temperature of the pressurized air stream entering the combustor 30 is significantly elevated above the bleed air temperature. The bleed air 77 may be used to reduce the temperature of core components downstream of the combustor.

[0032] The remainder of the airflow 78 bypasses the LP compressor 24 and the engine core 44 and exits the engine assembly 10 at the fan exhaust side 84 through the stationary vane row (and more specifically, an exit guide vane assembly 80 including a plurality of airfoil guide vanes 82). More specifically, a circumferential row of radially extending airfoil guide vanes 82 is utilized near the fan section 18 to exert some directional control on the airflow 78.

[0033] Some of the air supplied by the fan 20 may bypass the engine core 44 and be used to cool portions of the engine 10, particularly hot portions, and / or to cool or power other aspects of the aircraft. In the case of a turbine engine, the hot portions of the engine are typically downstream of the combustor 30, particularly downstream of the turbine section 32, with the HP turbine 34 being the hottest portion as it is directly downstream of the combustion section 28. Other sources of cooling fluid may be, but are not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26.

[0034] The turbine center frame 100 may be disposed between the HP turbine 34 and the LP turbine 36. The turbine center frame 100 may be a portion of a transition duct 88 disposed between the HP turbine 34 and the LP turbine 36. In this way, the turbine center frame 100 may be disposed at an inlet 120 ( Figure 2 ) and terminates at an outlet 122 at the front end of the LP turbine 36. The turbine center frame 100 may provide for fluidly coupling the HP turbine 34 to the LP turbine 36. In addition, the turbine center frame 100 may change the flow of fluid discharged from the HP turbine 34, for example by diffusing, smoothing, accelerating, or reducing turbulence. For example, the turbine center frame 100 may turn the flow tangentially or vertically relative to the engine centerline 12. This turning of the flow may reduce the three-dimensional or turbulent airflow from the HP turbine 34 and provide a smoother airflow entering the LP turbine 36, thereby allowing the LP turbine 36 to extract more work from the airflow. In a non-limiting example, the turbine center frame 100 may also serve as a structural member for supporting a through pipe, an auxiliary flow system, or a lubrication system.

[0035] Figure 2 1 is a view of the turbine center frame 100 looking toward the inlet 120 toward the outlet 122. The turbine center frame 100 includes an inner wall 102 radially spaced apart from an outer wall 104. In one non-limiting example, Figure 1 The core shell 46 may at least partially form the outer wall 104, and Figure 1 The stator 63 may at least partially form the inner wall 102. In one example, the inner wall 102 and the outer wall 104 may be axially or circumferentially continuous to form the turbine center frame 100. In another example, the inner wall 102 or the outer wall 104 may be made of a plurality of axial or circumferential wall segments coupled together to form the turbine center frame 100.

[0036] An exemplary rotor element 106 is shown forming a portion of the rotor 51. The rotor element 106 may have any suitable size, shape, or orientation within the engine 10. For example, the rotor element 106 may extend from the LP spool 50 to other rotor 51 components of the engine 10, such as Figure 1LP turbine blades 70( Figure 1 ).

[0037] An annular flow path 108 may be defined between the inner wall 102 and the outer wall 104. In the illustrated view, the flow path 108 defines a flow direction extending into the leaf. In operation, the mainstream flow 76 may move along the flow path 108 through the inlet 120 and out the outlet 122.

[0038] A set 110 of airfoils 112 are circumferentially arranged about the turbine center frame 100 extending between the inner wall 102 and the outer wall 104. Although shown as evenly spaced about the turbine center frame 100, any number of airfoils 112 may be utilized and may also be arranged in any suitable pattern, including varying spacing between adjacent airfoils 112 or between multiple rows or groups of airfoils 112.

[0039] In addition, for the purpose of visual clarity, the airfoil 112 is schematically shown as being radially aligned relative to the engine centerline. It will be understood that the airfoil 112 may be angled or tilted tangentially toward the engine centerline, or positioned within the turbine center frame 100 in any other form or orientation derived from desired aerodynamic design or analysis. Further, the airfoil 112 may include an arcuate or serpentine shape, or any suitable curvature along the body of the airfoil 112. It is contemplated that the shape or orientation of the airfoil 112 may divert a fluid flow (e.g., the mainstream flow 76) passing through the turbine center frame 100 along the airfoil 112.

[0040] The turbine center frame 100 may further include a separator 115. In the example shown, a plurality of separators 115 are provided with one separator 115 positioned between adjacent airfoils 112. The separator 115 may have any suitable shape or geometric profile, including an airfoil profile. In the example shown, the separator 115 extends from the inner wall 102 to the outer wall 104. In addition, a second separator 115A extends from the inner wall 102 to define a gap 117 with the outer wall 104, wherein the separator 115 and the second separator 115A are positioned between adjacent airfoils 112. Additionally or alternatively, the separator may extend from the outer wall 104 to form a gap (not shown) with the inner wall 102.

[0041] In addition, although the separator 115 is shown as having a smaller or thinner profile than the airfoil 112, the display of such a profile is only for visual distinction between the airfoil 112 and the separator 115, and is not intended to limit the geometry of the separator 115. It will be understood that the separator 115 may have the same geometric profile or shape as the airfoil 112, or the separator 115 may have a wider profile, or have a longer or shorter axial length than the airfoil 112. Such a separator 115 can be used to divide or redirect airflow between adjacent airfoils 112 in the group 110. In addition, the orientation, style, shape, curvature or spacing between the airfoil 112 and the separator 115 can vary. Any arrangement, style, orientation or grouping of the separator 115 and the airfoil 112 can be used for the turbine center frame 100.

[0042] It can be located relative to the engine centerline 12 ( Figure 1 ) An exemplary height or spacing distance 116 is measured between the inner wall 102 and the outer wall 104 extending in the radial direction. As used herein, "height" will be considered a synonym for the spacing distance between the inner wall 102 and the outer wall 104. In addition, an exemplary spacing width 117 between the airfoil 112 and the adjacent separator 115 is shown, as shown. It will be understood that the spacing width 117 can be defined between two adjacent airfoils 112 without a separator therebetween, or between two adjacent separators 115, or between any two circumferentially adjacent airfoils or separators within the turbine center frame 100.

[0043] An exemplary annular cross-sectional area 118 may be defined as an annular area taken anywhere between the inlet 120 and the outlet 122 (including the inlet 120 and the outlet 122). In other words, π*(Ro 2 -Ri 2 ) to estimate the annular area. It will be appreciated that the cross-sectional area 118 may vary at different locations along the flow path 108. For example, when moving along the flow path 108 (such as entering or moving out of the Figure 2 116, the cross-sectional area 118 may be increased or decreased. It should be understood that, depending on the flow swirl angle and obstruction, the local cross-sectional area 119 representing a portion of the path along which the mainstream flow 76 actually travels will be different from the annular cross-sectional area 118. The local cross-sectional area 119 may be at least partially formed by the spacing width 117 and the spacing distance 116, and may define an effective flow area. In addition, the local cross-sectional area 119 may be increased or decreased by changing one or both of the spacing distance 116 and the spacing width 117, or by changing the swirl (or flow turning) level within the flow path.

[0044] Now go to Figure 3 , shows a side view of the transition duct 88 and the turbine center frame 100, wherein the main flow 76 moves along the flow path 108. The axial direction 90 is indicated for reference. The axial direction 90 may be parallel to the engine centerline 12 ( Figure 1 ). In the example shown, the inner wall 102 and the outer wall 104 extend axially to the turbine center frame 100 between the inlet 120 and the outlet 122. The airfoil 112 extends from a leading edge 130 to a trailing edge 132. In the example shown, the leading edge 130 is positioned downstream of the inlet 120 and the trailing edge 132 is positioned upstream of the outlet 122 relative to the flow path 108. It is contemplated that the leading edge 130 and the trailing edge 132 may be aligned with, positioned upstream of, or positioned downstream of the respective inlet 120 and outlet 122.

[0045] The splitter 115 also extends from a splitter leading edge 140 to a splitter trailing edge 142. In the example shown, the splitter 115 extends completely from the inner wall 102 to the outer wall 104, wherein both the splitter leading edge 140 and the splitter trailing edge 142 are positioned between the leading edge 130 and the trailing edge 132 of the airfoil 112. It is also contemplated that one or both of the splitter leading edge 140 and the splitter trailing edge 142 may be positioned between the inlet 120 and the outlet 122 based on a function of the axial width of the splitter 115. In a non-limiting example, the splitter trailing edge 142 may be spaced apart from the trailing edge 132 of the airfoil 112 by a distance between 5-20% of the axial width of the splitter 115.

[0046] As shown, a first height 116A, a second height 116B, and a third height 116C can be radially defined between the inner wall 102 and the outer wall 104. A first cross-sectional area 118A, a second cross-sectional area 118B, and a third cross-sectional area 118C, respectively, indicated by dashed lines, can also be defined between the inner wall 102 and the outer wall 104 along the flow path 108. It should be understood that, although only shown as partial cross-sectional areas, the first, second, and third cross-sectional areas 118A, 118B, and 118C are respectively exemplary annular cross-sectional areas 118 as described herein, and can be at least partially formed by the respective first, second, and third heights 116A, 116B, 116C. It should be understood that, Figure 3 The view is perpendicular to Figure 2 The view makes the indication Figure 2 The dashed outline of the exemplary partial cross-sectional area 119 is considered to be "side-up" to form Figure 3 118A, 118B, 118C are schematically illustrated in phantom view of first, second and third cross-sectional areas 118A, 118B, 118C.

[0047] In the example shown, first height 116A is located at inlet 120 of turbine center frame 100. Inlet 120 is shown slightly spaced from first height 116A for visual clarity, and it is understood that inlet 120 may coincide with both first height 116A and first cross-sectional area 118A.

[0048] In addition, second height 116B and second cross-sectional area 118B can be defined at separator leading edge 140. Again, only for visual clarity, second height 116B, second cross-sectional area 118B and separator leading edge 140 are shown as slightly spaced apart. Second cross-sectional area 118B is located downstream of first cross-sectional area 118A relative to flow path 108. Second height 116B can be greater than first height 116A. For example, second height 116B can be between 1.1 and 2.0 times of first height 116A, including between 1.3 and 1.5 times of first height 116A.

[0049] In addition, the third height 116C and the third cross-sectional area 118C may be defined at or near the trailing edge 132 of the airfoil 112. Again, for visual clarity, the third height 116C, the third cross-sectional area 118C, and the outlet 122 are shown slightly spaced apart. The third cross-sectional area 118C is located downstream of the second cross-sectional area 118B relative to the flow path 108. The third height 116C may be between 0.5 and 1.5 times the second height 116B, including between 0.8 and 1.1 times the second height 116B.

[0050] It is further contemplated that other portions of the transition duct 88 may form a height or cross-sectional area consistent with the turbine center frame 100 to change or adjust the mainstream flow 76 along the flow path 108. In the example shown, the fourth height 116D is defined in the transition duct 88 between the core housing 46 and the stator 63 downstream of the third height 116C. The fourth height 116D may be defined at or near the outlet 122 of the turbine center frame 100. The fourth height 116D may be greater than, less than, or equal to the third height 116C. In a non-limiting example, the fourth height 116D may be between 0.8 and 1.5 times the third height 116C, including between 1.0 and 1.3 times the third height 116C. In addition, as shown, the fourth height 116D may at least partially form a fourth cross-sectional area 118D. The fourth cross-sectional area 118D may have any suitable size, including greater than, less than, or equal to the third cross-sectional area 118C. In this way, the mainstream flow 76 may be accelerated, slowed down, or turned downstream of the outlet 122. Furthermore, second cross-sectional area 118B may be larger than first cross-sectional area 118A, and larger than or equal to third cross-sectional area 118C.

[0051] Reference now Figure 4 , a graph 150 is shown schematically showing the flow path 108 ( Figure 3 ) through an exemplary cross-sectional profile of the turbine center frame 100. The axial direction 90 ( Figure 3 ), a cross-sectional area through the turbine center frame 100 and the transition duct 88 is indicated on a vertical axis 154. Axial locations of the first, second, third, and fourth heights 116A, 116B, 116D, 116D are indicated on a horizontal axis 152, and respective first, second, third, and fourth cross-sectional areas 118A, 118B, 118C, 118D are indicated on the vertical axis 154.

[0052] It is contemplated that the cross-sectional area through the turbine center frame 100 may increase or decrease between the heights 116A, 116B, 116C, 116D shown. In the example shown, when moving axially from the second height 116B to the third height 116C, the cross-sectional area through the turbine center frame 100 may decrease and then increase such that the second cross-sectional area 118B and the third cross-sectional area 118C may be approximately the same size, for example within 10%. The cross-sectional area may be increased or decreased between the heights 116A, 116B, 116C, 116D shown. In the example shown, when moving axially from the second height 116B to the third height 116C, the cross-sectional area through the turbine center frame 100 may decrease and then increase such that the second cross-sectional area 118B and the third cross-sectional area 118C may be approximately the same size, for example within 10%. Figure 2 ) to form this increase or decrease in cross-sectional area.

[0053] In one non-limiting example, a ratio of the third cross-sectional area 118C to the second cross-sectional area 118B is less than 1.1 (118C / 118B<1.1).

[0054] In the example shown, the second cross-sectional area 118B can be larger than the first cross-sectional area 118A. In a non-limiting example, the second cross-sectional area 118B can be between 1.0 and 2.0 times the first cross-sectional area 118A, including greater than 1.5 times the first cross-sectional area 118A, wherein the ratio of the second cross-sectional area 118B to the first cross-sectional area 118A is greater than 1.5 (118B / 118A>1.5)

[0055] As shown, the third cross-sectional area 118C can be less than or equal to the second cross-sectional area 118B, but it is also conceivable that the third cross-sectional area is slightly larger than the second cross-sectional area. In a non-limiting example, the third cross-sectional area 118C can be between 0.3 and 1.5 times the second cross-sectional area 118B, including between 0.5 and 1.1 times the second cross-sectional area 118B, wherein the ratio of the third cross-sectional area 118C to the second cross-sectional area 118B is less than 1.1 (118C / 118B<1.1).

[0056] It is further contemplated that the third cross-sectional area 118C may be between 1.0 and 2.0 times the first cross-sectional area 118A, including between 1.0 and 1.8 times the first cross-sectional area 118A. It is further contemplated that the third cross-sectional area 118C may be even greater than 2.0 times the first cross-sectional area 118A. In yet another example, the fourth cross-sectional area 118D within the transition duct 88 may be between 1.0 and 3.0 times the first cross-sectional area 118A in the turbine center frame 100.

[0057] Reference now Figure 5 , another turbine center frame 200 that may be used in turbine engine 10 is shown. Turbine center frame 200 is similar to turbine center frame 100; therefore, like components will be identified with like numerals increased by 100, it being understood that the description of like components of turbine center frame 100 applies to turbine center frame 200 unless otherwise noted.

[0058] The turbine center frame 200 includes an inner wall 202 and an outer wall 204 extending from an inlet 220 to an outlet 222. As described above, the airfoil 112 and the separator 115 extend between the inner wall 202 and the outer wall 204. First, second, third, and fourth heights 216A, 216B, 216C, 216D are defined between the inner wall 202 and the outer wall 204, and corresponding first, second, third, and fourth cross-sectional areas 218A, 218B, 218C, 218D are formed within the turbine center frame 200 and the transition duct 88 as described above.

[0059] As shown, a first height 216A is defined at the leading edge 130 of the airfoil 112, a second height 216B is defined at the separator leading edge 140, a third height 216C is defined at the outlet 222, and a fourth height 216C is defined in the transition duct 88 downstream of the outlet 222. One difference compared to the turbine center frame 100 is that the leading edge 130 of the airfoil 112 is located at the inlet 220 of the turbine center frame 200. Therefore, as shown, the first height 216A is also defined at the inlet 220.

[0060] Another difference is that at least one of the inner wall 202 and the outer wall 204 may include at least one local profile 260. In the example shown, the first local profile 261 is in the form of a convex ridge that narrows the flow path 108 downstream of the separator leading edge 130. The second local profile 262 is in the form of a concave ridge that expands the flow path 108 at the second height 216B. The second local profile 262 also forms a second cross-sectional area 218B, as shown. Multiple local profiles 260 can be set along the inner wall 202 or the outer wall 204 to locally increase or decrease the cross-sectional area through the turbine center frame 200 and adjust the mainstream flow 76 through the transition duct 88. In this way, at least one local profile 260 can form at least one of the first, second or third cross-sectional areas 218A, 218B, 218D.

[0061] It is understood that the velocity of the mainstream flow 76 can increase or decrease in proportion to the size of the cross-sectional area through the turbine center frame 100 or the transition duct 88. For example, the second cross-sectional area 118B, 218B being larger than the first cross-sectional area 118A, 218B can cause the mainstream flow 76 to reduce velocity in the region of the second height 116B, 216B (e.g., at the leading edge of the separator 115). In one example, the second cross-sectional area 118B, 218B can define a maximum cross-sectional area 118M, so that the mainstream flow 76 has a minimum velocity at the maximum cross-sectional area 118M. In addition, the third cross-sectional area 118C, 218C being smaller than the second cross-sectional area 118A, 218A can cause the velocity of the mainstream flow 76 across the separator 115 and toward the outlet 122, 222 to increase. It is contemplated that the velocity of the mainstream flow 76 at the outlet 122, 222 may be between 80% and 125% of the velocity of the mainstream flow 76 at the inlet 120, 220 while still providing the desired smoothing or diffusion of the airflow through the turbine center frame 100, 200. This low overall acceleration or velocity change may be achieved in part due to the slowing of the airflow prior to encountering the separator 115, as opposed to conventional turbine center frames having a stepwise or gradual increase in velocity across the entire frame.

[0062] The aspects described above provide a method of changing an airflow along a flow path extending through a turbine center frame (which passes through a turbine engine), such as a mainstream flow 76 along a flow path 108 through a turbine center frame 100, 200. The method includes directing an airflow through a first cross-sectional area of ​​the turbine center frame, such as a first cross-sectional area 118A, 218A, at a first speed along the flow path. For example, the airflow may be directed through a leading edge 130 of an airfoil 112 within the turbine center frame 100, 200 at a first speed. The method also includes slowing the airflow to a second speed less than the first speed via a second cross-sectional area 118B, 218B that is larger than the first cross-sectional area 118A, 218A. Slowing the airflow may also include slowing the airflow to a minimum speed within the turbine center frame 100, 200 by causing the airflow to flow through a maximum cross-sectional area 118M within the turbine center frame 100. The method further includes accelerating the airflow to a third speed greater than the second speed via a third cross-sectional area 118C, 218C greater than or equal to the first cross-sectional area 118A, 218A. Accelerating the airflow may further include accelerating the airflow to a third speed at an outlet 122, 222 of the turbine center frame 100, 200. In addition, the method may further include splitting the airflow passing through a separator 115 disposed along the flow path 108. It is understood that changing the airflow may include, for example, accelerating, slowing down, or turning the airflow via the geometry or arrangement of the airfoil 112, the separator 115, the inner wall 102, 202, or the outer wall 104, 204.

[0063] The above aspects provide a variety of benefits. For example, changing the cross-sectional area through the turbine center frame can provide a lower total velocity increase across the frame, including a velocity reduction at the outlet compared to the inlet, compared to a conventional turbine center frame. It is known that a total velocity increase of less than 20% can cause engine performance losses due to excessive 3D end wall flow and turbulence in the airflow. At least due to the airflow slowing down at the second height position and accelerating toward the third height position, the turbine center frame described herein can operate without such performance losses even at a lower total airflow acceleration, or even without acceleration or velocity reduction. Compared to a conventional turbine center frame, the greater acceleration across the separator provides improved turbulence reduction and airflow smoothness without the need for a large total acceleration between the inlet and outlet of the turbine center frame.

[0064] In addition, the airflow through the airfoil or separator experiences frictional drag losses, which are known to be proportional to the flow velocity. The lower overall acceleration through the turbine center frame and the lower airflow velocity at the leading edge of the separator can both provide lower frictional drag losses compared to a conventional turbine center frame, thereby improving engine performance. Aspects of the present disclosure provide for achieving low drag losses without being adversely affected by turbulence / three-dimensional flow or undesirable flow interactions from an upstream high-pressure turbine.

[0065] In the undescribed range, different features and structures of various aspects can be used in combination or replace each other as needed. A feature not shown in all examples does not mean that it cannot be shown in this way, but is for the sake of description simplicity. Therefore, regardless of whether new aspects are explicitly described, various features of different aspects can be mixed and matched as needed to form new aspects. All combinations or permutations of the features described herein are covered by the present disclosure.

[0066] This written description uses examples to describe the aspects of the disclosure described herein, including the best mode, and also to enable any person skilled in the art to practice aspects of the disclosure, including making and using any device or system and performing any incorporated method. The patentable scope of aspects of the disclosure is defined by the claims, and may include other examples that occur to one skilled in the art. Such other examples are intended to fall within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims.

[0067] Further aspects of the invention are provided by the subject matter of the following clauses:

[0068] 1. A turbine center frame for a turbine engine, a flow path extending through the turbine center frame, the turbine center frame comprising: an inner wall, the inner wall being radially spaced apart from an outer wall, wherein the inner wall and the outer wall extend between an inlet and an outlet, wherein the outlet is downstream of the inlet relative to the flow path; a set of circumferentially spaced apart airfoils, the set of circumferentially spaced apart airfoils extending between the inner wall and the outer wall; a first cross-sectional area, the first cross-sectional area being defined along the flow path between the inner wall and the outer wall; a second cross-sectional area, the second cross-sectional area being larger than the first cross-sectional area and being defined along the flow path between the inner wall and the outer wall, downstream of the first cross-sectional area; and a third cross-sectional area, the third cross-sectional area being defined along the flow path between the inner wall and the outer wall, downstream of the second cross-sectional area.

[0069] 2. The turbine center frame according to any of the preceding clauses further comprises: a first height, the first height being radially defined between the inner wall and the outer wall, at least partially forming a first cross-sectional area; a second height, the second height being greater than the first height, radially defined between the inner wall and the outer wall, at least partially forming a second cross-sectional area; and a third height, the third height being greater than or equal to the first height, radially defined between the inner wall and the outer wall, at least partially forming a third cross-sectional area.

[0070] 3. A turbine center frame according to any preceding clause, further comprising a separator extending from one of the outer wall or the inner wall and positioned circumferentially between adjacent airfoils in the set of airfoils.

[0071] 4. A turbine center frame according to any preceding clause, wherein the second height is defined at a leading edge of the separator.

[0072] 5. A turbine center frame according to any preceding clause, wherein the second cross-sectional area is at least partially defined at a leading edge of the separator.

[0073] 6. A turbine center frame according to any preceding clause, wherein the splitter comprises a splitter trailing edge located upstream of a trailing edge of an airfoil in the set of airfoils.

[0074] 7. A turbine center frame according to any preceding clause, wherein the first height is defined at a leading edge of an airfoil in the set of airfoils.

[0075] 8. A turbine center frame according to any preceding clause, wherein the leading edge of the airfoil is located at the inlet of the turbine center frame.

[0076] 9. A turbine center frame according to any preceding clause, wherein the third height is located at an outlet of the turbine center frame.

[0077] 10. A turbine center frame according to any preceding clause, wherein an airfoil of the set of airfoils comprises a leading edge located downstream of an inlet of the turbine center frame.

[0078] 11. A turbine center frame according to any preceding clause, wherein an airfoil of the set of airfoils further comprises a trailing edge located upstream of an outlet of the turbine center frame.

[0079] 12. A turbine center frame according to any preceding clause, wherein at least one of the inner wall or the outer wall further comprises at least one local profile forming at least one of the first cross-sectional area, the second cross-sectional area or the third cross-sectional area.

[0080] 13. A turbine centre frame according to any preceding clause, wherein the second cross-sectional area defines a maximum cross-sectional area within the turbine centre frame.

[0081] 14. A turbine engine, comprising: an engine core, the engine core comprising a compressor section, a combustion section and a turbine section arranged in an axial flow and defining a flow path, wherein the turbine section comprises a high-pressure turbine and a low-pressure turbine; and a turbine center frame, the turbine center frame extending from the high-pressure turbine to the low-pressure turbine, the turbine center frame comprising: an inner wall, the inner wall being radially spaced apart from the outer wall, wherein the inner wall and the outer wall extend between an inlet and an outlet, wherein the outlet is downstream of the inlet relative to the flow path; a group of circumferentially spaced airfoils, the group of circumferentially spaced airfoils extending between the inner wall and the outer wall; a first cross-sectional area, the first cross-sectional area being radially defined between the inner wall and the outer wall; a second cross-sectional area, the second cross-sectional area being larger than the first cross-sectional area and being radially defined between the inner wall and the outer wall, downstream of the first cross-sectional area; and a third cross-sectional area, the third cross-sectional area being radially defined between the inner wall and the outer wall, downstream of the second cross-sectional area.

[0082] 15. The turbine engine according to any of the preceding clauses further comprises: a first height, the first height being defined between the inner wall and the outer wall perpendicular to the flow path, at least partially forming a first cross-sectional area; a second height, the second height being greater than the first height, being defined between the inner wall and the outer wall perpendicular to the flow path, at least partially forming a second cross-sectional area; and a third height, the third height being greater than the second height, being defined between the inner wall and the outer wall perpendicular to the flow path, at least partially forming a third cross-sectional area.

[0083] 16. A turbine engine according to any preceding clause, further comprising a separator extending from one of the outer wall or the inner wall and positioned circumferentially between adjacent airfoils in the set of airfoils.

[0084] 17. A turbine engine according to any preceding clause, wherein the second height is defined at a leading edge of the separator and wherein the first height is located at an inlet of the turbine center frame.

[0085] 18. A turbine engine according to any preceding clause, wherein at least one of the inner wall or the outer wall further comprises at least one local contour forming at least one of the first cross-sectional area, the second cross-sectional area or the third cross-sectional area.

[0086] 19. A method for changing an airflow along a flow path extending through a turbine center frame for a turbine engine, the method comprising: directing an airflow along the flow path through a first cross-sectional area of ​​the turbine center frame at a first speed; slowing the airflow to a second speed less than the first speed by causing the airflow to flow through a second cross-sectional area of ​​the turbine center frame that is larger than the first cross-sectional area; and accelerating the airflow to a third speed greater than the second speed by causing the airflow to flow through a third cross-sectional area of ​​the turbine center frame.

[0087] 20. A method according to any preceding clause, further comprising splitting the gas flow through a separator disposed along the flow path.

[0088] 21. The method of any preceding clause, wherein slowing the airflow further comprises causing the airflow to flow along the flow path past a leading edge of an airfoil within a turbine center frame at a second speed.

[0089] 22. The method of any preceding clause, wherein accelerating the airflow further comprises accelerating the airflow to a third speed at an outlet of the turbine center frame.

[0090] 23. A method according to any preceding clause, wherein slowing the airflow to a second speed further comprises slowing the airflow to a minimum speed within the turbine center frame by passing the airflow through a maximum cross-sectional area within the turbine center frame.

Claims

1. A turbine center frame for a turbine engine, through which a flow path extends, characterized in that: The turbine center frame comprises: an inner wall radially spaced from the outer wall, wherein each of the inner wall and the outer wall extends between an inlet and an outlet, wherein the outlet is downstream of the inlet relative to the flow path; a set of circumferentially spaced airfoils extending between the inner wall and the outer wall; a separator extending from one of the outer wall or the inner wall and positioned circumferentially between adjacent airfoils in the set of airfoils; a first cross-sectional area defined along the flow path between the inner wall and the outer wall; a second cross-sectional area, the second cross-sectional area being larger than the first cross-sectional area and being defined along the flow path between the inner wall and the outer wall at a leading edge of the separator; a third cross-sectional area defined along the flow path between the inner wall and the outer wall downstream from the second cross-sectional area; and A fourth cross-sectional area is defined along the flow path between the inner wall and the outer wall downstream from the third cross-sectional area, wherein the fourth cross-sectional area is between 1 and 3 times the first cross-sectional area.

2. The turbine center frame according to claim 1, characterized in that: Further including: a first height defined radially between the inner wall and the outer wall and at least partially defining the first cross-sectional area; a second height, the second height being greater than the first height, radially defined between the inner wall and the outer wall, at least partially forming the second cross-sectional area; as well as A third height, the third height being greater than or equal to the first height, is defined radially between the inner wall and the outer wall, at least partially forming the third cross-sectional area.

3. The turbine center frame according to claim 2, characterized in that: in, The second height is defined at a leading edge of the separator.

4. The turbine center frame according to claim 3, characterized in that: in, The second cross-sectional area is at least partially defined at the leading edge of the separator.

5. The turbine center frame according to claim 2, characterized in that: in, The splitter includes a splitter trailing edge located upstream of a trailing edge of an airfoil in the set of airfoils.

6. The turbine center frame according to claim 2, characterized in that: in, The first height is defined at a leading edge of an airfoil in the set of airfoils.

7. The turbine center frame according to claim 6, characterized in that: in, The leading edge of the airfoil is located at the inlet of the turbine center frame.

8. The turbine center frame according to claim 2, characterized in that: in, The third height is located at the outlet of the turbine center frame.

9. The turbine center frame according to claim 1, characterized in that: in, An airfoil in the set of airfoils includes a leading edge located downstream of the inlet of the turbine center frame.

10. The turbine center frame according to claim 9, characterized in that: in, The airfoil of the set of airfoils further includes a trailing edge located upstream of the outlet of the turbine center frame.

11. The turbine center frame according to claim 1, characterized in that: in, At least one of the inner wall or the outer wall further comprises at least one local contour forming at least one of the first cross-sectional area, the second cross-sectional area, or the third cross-sectional area.

12. The turbine center frame according to claim 1, characterized in that: in, The second cross-sectional area defines a maximum cross-sectional area within the turbine center frame.

13. A turbine engine, characterized in that: include: an engine core comprising a compressor section, a combustion section, and a turbine section arranged in an axial flow and defining a flow path, wherein the turbine section comprises a high pressure turbine and a low pressure turbine; and A turbine center frame, the turbine center frame extending from the high-pressure turbine to the low-pressure turbine, the turbine center frame comprising: an inner wall radially spaced from the outer wall, wherein the inner wall and the outer wall extend between an inlet and an outlet, wherein the outlet is downstream of the inlet relative to the flow path; a set of circumferentially spaced airfoils extending between the inner wall and the outer wall; a separator extending from one of the outer wall or the inner wall and positioned circumferentially between adjacent airfoils in the set of airfoils; a first cross-sectional area defined radially between the inner wall and the outer wall; a second cross-sectional area, the second cross-sectional area being larger than the first cross-sectional area and being defined radially between the inner wall and the outer wall at a leading edge of the separator; a third cross-sectional area defined radially between the inner wall and the outer wall downstream from the second cross-sectional area; and A fourth cross-sectional area is defined along the flow path between the inner wall and the outer wall downstream from the third cross-sectional area, wherein the fourth cross-sectional area is between 1 and 3 times the first cross-sectional area.

14. The turbine engine according to claim 13, characterized in that Further including: a first height, the first height being defined perpendicularly to the flow path between the inner wall and the outer wall and at least partially forming the first cross-sectional area; a second height, the second height being greater than the first height, defined perpendicularly to the flow path between the inner wall and the outer wall, at least partially forming the second cross-sectional area; as well as A third height, which is greater than the second height, is defined perpendicularly to the flow path between the inner wall and the outer wall, at least partially forming the third cross-sectional area.

15. The turbine engine according to claim 14, characterized in that in, The second height is defined at a leading edge of the separator, and wherein the first height is located at the inlet of the turbine center frame.

16. The turbine engine according to claim 13, characterized in that in, At least one of the inner wall or the outer wall further comprises at least one local contour forming at least one of the first cross-sectional area, the second cross-sectional area, or the third cross-sectional area.

17. A method of modifying a flow of gas along a flow path extending through a turbine center frame according to claim 1, characterized in that The method comprises: directing the airflow through the first cross-sectional area of ​​the turbine center frame at a first velocity along the flow path; slowing the airflow to a second speed less than the first speed by passing the airflow through a second cross-sectional area of ​​the turbine center frame that is larger than the first cross-sectional area; and accelerating the airflow to a third speed greater than the second speed by passing the airflow through the third cross-sectional area of ​​the turbine center frame; The third speed is between 80% and 125% of the first speed.

18. The method according to claim 17, characterized in that in, Slowing the airflow further includes flowing the airflow along the flow path past a leading edge of an airfoil within the turbine center frame at the second speed.

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