Turbulator geometry for a combustion liner

Abstract

A heat transfer mechanism is provided comprising a plurality of turbulators located along a surface of a body, such as a combustion liner. The turbulators have a first side with a first ramp angle, a second side with a second ramp angle, a height, and a base width, where the base width is a function of the height and where the turbulators are spaced an axial distance apart that is a function of the turbulator height.

Description

TECHNICAL FIELD

[0001] This disclosure relates generally to a heat transfer mechanism for use on a surface of a component subjected to elevated temperatures in a gas turbine engine and more specifically to aspects of a turbulator configuration for a combustion system. Especially this disclosure relates to a combustion liner. A combustion liner according to the prior art is known from US 9,511,447 B2.BACKGROUND OF THE DISCLOSURE

[0002] A gas turbine engine typically comprises a multi-stage compressor coupled to a multi-stage turbine via an axial shaft. Air enters the gas turbine engine through the compressor where its temperature and pressure increase as it passes through subsequent stages of the compressor. The compressed air is then directed to one or more combustors where it mixes with a fuel source to create a combustible mixture. This mixture is ignited in the combustors to create a flow of hot combustion gases. These gases are directed into the turbine causing the turbine to rotate, thereby driving the compressor. The output of the gas turbine engine can be mechanical thrust through exhaust from the turbine or shaft power from the rotation of an axial shaft, where the axial shaft can drive a generator to produce electricity.

[0003] The compressor and turbine each comprise a plurality of rotating blades and stationary vanes having an airfoil extending into the flow of compressed air or flow of hot combustion gases. Each blade or vane has a particular set of design criteria which must be met in order to provide the necessary work to the passing flow through the compressor and the turbine.

[0004] Combustion liners frequently contain reactions of fuel and air reaching upwards of 2204 deg. C. To prevent melting and / or erosion of the combustion liner, the combustion liner is typically covered with a protective thermal barrier coating on the surface of the liner in direct contact with the hot combustion gases. The benefit obtained by the thermal barrier coating is a function of the composition and coating thickness, but can reduce combustion liner temperature by approximately 89 deg. C. However, a thermal barrier coating alone is not always enough to protect the combustion liner from the hot combustion gases passing therethrough. Active cooling can be incorporated in the form of cooling holes, where air cooler than the hot combustion gases passes therethrough to cool the wall of the combustion liner. Furthermore, cooling air can pass along an outer surface of the combustion liner in order to cool a backside of the combustion liner.

[0005] An example of backside cooling techniques is shown in FIG. 1 where the combustion liner 10 comprises a series of raised edges or perturbances 12 positioned along a limited portion, such as the upper portion 14, of the combustion liner 10.BRIEF SUMMARY OF THE DISCLOSURE

[0006] The present disclosure discloses an improved heat transfer system and process for actively cooling a heated surface, such as that used in conjunction with a combustion liner having a surface requiring active cooling. The present invention defined in claim 1.

[0007] In the present disclosure, a combustion liner comprises a generally annular body having a first cylindrical portion, a conical portion, and a second cylindrical portion. The combustion liner also comprises an inlet end proximate the first cylindrical portion and an outlet end proximate the second cylindrical portion. A plurality of turbulators are located along an outer surface of the first cylindrical portion and the conical portion, where the turbulators have a first side with a first ramp angle, a second side with a second ramp angle, a height, and a base width extending between the first side and the second side.

[0008] These and other features of the present disclosure can be best understood from the following description and claims.BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

[0009] The present disclosure is described in detail below with reference to the attached drawing figures, wherein: FIG. 1 is an elevation view of a combustion liner for a gas turbine engine. FIG. 2 is an elevation view of a combustion liner in accordance with an embodiment of the disclosure. FIG. 3 is a cross section view of the combustion liner of FIG. 2 in accordance with an embodiment of the present disclosure. FIG. 4 is a detailed cross section view of a portion of the combustion liner of FIG. 3. FIG. 5 is an alternate cross section view of a portion of the combustion liner of FIG. 3. FIG. 6 is a cross section view of a portion of a gas turbine combustor in accordance with an embodiment of the present disclosure. FIG. 7 is a perspective view of a flow sleeve of a gas turbine combustor in accordance with the prior art. FIG. 8 is a detailed perspective view of a portion of the flow sleeve of FIG. 7 in accordance with the prior art. FIG. 9 is a detailed perspective view of a cross fire tube retention system in accordance with the prior art. FIG. 10 is a perspective view of a retention clip in accordance with the prior art. FIG. 11 is a perspective view of a flow sleeve of a gas turbine combustor in accordance with an embodiment of the present disclosure. FIG. 12 is a detailed perspective view of a portion of the flow sleeve of FIG. 11 in accordance with an embodiment of the present disclosure. FIG. 13 is a perspective view of a retention clip in accordance with an embodiment of the present disclosure. FIG. 14 is a detailed perspective view of a cross fire tube retention system in accordance with an embodiment of the present disclosure. DETAILED DESCRIPTION

[0010] The following presents a simplified summary of the disclosure to provide a basic understanding of some aspects thereof. This summary is not an extensive overview of the application. It is not intended to identify critical elements of the disclosure or to delineate the scope of the disclosure. Its sole purpose is to present some concepts of the disclosure in a simplified form as a prelude to the more detailed description that is presented elsewhere herein.

[0011] The present disclosure is intended for use in a gas turbine engine, such as a gas turbine engine used for power generation. As such, the present disclosure is capable of being used in a variety of turbine operating environments, regardless of the manufacturer.

[0012] As those skilled in the art will readily appreciate, a gas turbine engine is circumferentially disposed about an engine centerline, or axial centerline axis. The engine includes a compressor, a combustion section and a turbine with the turbine coupled to the compressor via an engine shaft. As is well known in the art, air compressed in the compressor is mixed with fuel which is burned in the combustion section and expanded in turbine. The air compressed in the compressor is mixed with fuel and the gases are expanded in the turbine. The turbine includes rotors that, in response to the fluid expansion, rotate, thereby driving the compressor. The turbine comprises alternating rows of rotary turbine blades, and static airfoils, often referred to as vanes.

[0013] Various embodiments of the present disclosure are depicted in FIGS. 2- 6. Referring initially to FIG. 2, a combustion liner 200 for use in a gas turbine engine is provided. The combustion liner 200 comprises a generally annular body 202 having a first cylindrical portion 204, a conical portion 206 connected to the first cylindrical portion 204, and a second cylindrical portion 208 connected to the conical portion 206. The combustion liner 200 also has an inlet 210 proximate the first cylindrical portion 204 and an outlet 212 proximate the second cylindrical portion 208.

[0014] In an industrial gas turbine engine, compressed air enters the combustion liner 200 through the inlet 210 where the compressed air mixes with fuel from one or more fuel nozzles, where the one or more fuel nozzles are also positioned adjacent the inlet 210. Proximate the outlet 212 and the second cylindrical portion 208 is a sealing mechanism 214 for sealing the outlet 212 of the combustion liner 200 to an adjacent component, such as a transition duct. The sealing mechanism 214 can be a slotted spring seal comprising of a plurality of sheet metal fingers capable of being compressed when a force, such as that from a mating engine component, is applied to the sealing mechanism 214.

[0015] Referring now to FIGS. 2-5, the combustion liner 200 also comprises a plurality of turbulators 216 positioned along an outer surface 218 of the first cylindrical portion 204 and the conical portion 206. The turbulators 216 are positioned across generally the entire length of the first cylindrical portion 204 and conical portion 206 in order to provide a more effective cooling configuration over the prior art.

[0016] More specific details of the turbulators 216 are shown in FIGS. 3-5. Referring to FIGS. 4 and 5, the plurality of turbulators 216 each have a first side 220 with a first ramp angle α and a second side 222 with a second ramp angle β. The turbulators 216 also have a height 224 extending away from the outer surface 218 and a width 226, where the width 226 is measured from a tangent between each of the first side 220 and second side 222 and the outer surface 218. In the embodiment depicted in FIG. 5, the turbulators 216 comprise a base fillet radius R between the first side 220 and the outer surface 218 and the second side 222 and the outer surface 218 along the first cylindrical portion 204 and the conical portion 206. The exact size of base fillet radius R can be the same or vary as it is not believed to greatly impact heat transfer or pressure loss as air passes over the turbulators 216. The first side 220 and second side 222 are joined together at a tip region 228. In the embodiment shown in FIGS. 4 and 5, the tip region 228 includes a full round radius.

[0017] In general, the plurality of turbulators 216 are axisymmetric. For example, and as depicted in FIGS. 4 and 5, each of the plurality of turbulators 216 has a generally triangular cross section with a plurality of radii at its corners. While the exact size and shape of the plurality of turbulators 216 can vary, the embodiment depicted in FIGS. 3-5 includes a base width 226 that is approximately 1-3 times larger than the height 224. For an embodiment of the disclosure, the height 224 of the turbulator 216 is approximately 0.76 mm while the base width is approximately 2.28 mm wide, or about three times the height 224.

[0018] The first ramp angle α and the second ramp angle β can also vary depending on the preferred cooling design of the turbulators 216 and combustion liner 200. For the embodiment depicted in FIGS. 3-5, the first ramp angle α and the second ramp angle β are approximately 30-45 degrees, as measured from a surface of the first cylindrical portion 204 or the conical portion 206. Depending on the configuration of turbulators 216, the first ramp angle α and the second ramp angle β can be the same or can be different.

[0019] In addition to the specific size and shape of the plurality of turbulators 216, the position of the turbulators 216 can also vary. More specifically, the plurality of turbulators 216 have an axial spacing 230 as measured between centerpoints C of adjacent turbulators 216. For the embodiment depicted in FIGS. 3-5, the axial spacing 230 is approximately 8.6 mm, which, for the height 224 of 0.76 mm is slightly greater than 10 times the height. The axial spacing 230 can be approximately 10-20 times the height 224.

[0020] In an alternate embodiment of the disclosure, a method of providing a heat transfer mechanism is disclosed. The method comprises providing a body having a surface for the heat transfer mechanism and forming the heat transfer mechanism in the surface of the body. The heat transfer mechanism comprises a plurality of turbulators where each turbulator comprises a first side with a first ramp angle and a second side with a second ramp angle, where the first side is connected to the second side at a tip region having a height and a full round tip radius. The plurality of turbulators are spaced apart by an axial distance.

[0021] The plurality of turbulators 216 are provided to enhance the heat transfer along a surface subject to high temperature loads. While the turbulators 216 can be located on an outer surface 218, as shown in FIGS. 3-6, the turbulators 216 can also be incorporated along an inner surface, depending on the heat transfer requirements of the component.

[0022] The heat transfer mechanism can be incorporated into the surface of the body through a variety of means. For example, in an embodiment of the disclosure, the plurality of turbulators can be machined into the surface of the body. Alternatively, the plurality of turbulators can be cast into the surface of the body as part of the body itself. In addition, the plurality of turbulators can be separately fabricated and secured to the surface of the body, such as through a brazing process.

[0023] One such use of the present disclosure is along an external surface of a combustion liner 200, where the combustion liner 200 is positioned within a flow sleeve 240 and a combustor case 242. The combustion liner 200 and the flow sleeve 240 form a passageway 244 located therebetween and through which air passes (indicated by arrows). The air is directed towards a head end 246 of a combustion system and passes over the plurality of turbulators 216 causing the air to come in contact with a greater surface area of the combustion liner 200 operating at an elevated temperature.

[0024] The specific turbulator configuration is determined by maximizing the size of passageway 244 and selecting a height 224 of the turbulator 216 that provides the required level of cooling heat transfer for the airflow and geometry of the passageway 244. The axial spacing 230 is set to minimize pressure loss within the passageway 244 based on the height of the passageway but may be adjusted smaller or larger depending on a streamwise length of the passageway 244.

[0025] The artisan understands that in a typical industrial gas turbine engine, the combustor section comprises a plurality of can-annular combustors. In this configuration, a plurality of individual combustors is arranged about the axis of the gas turbine engine, where each combustor receives a portion of the compressed air from the compressor. However, in order to eliminate the need for ignition sources in each combustor for use at start-up as well as any time a combustor flashes back or when a flame is unintentionally extinguished, the plurality of individual combustors is connected by a plurality of cross fire tubes. In operation, one combustor can be ignited, and the flame will pass through the cross fire tubes to an adjacent combustor, thereby igniting a combustible mixture in an adjacent combustor.

[0026] A cross fire tube arrangement in accordance with the prior art is disclosed in FIGS. 7-10. Referring initially to FIG. 7, a flow sleeve 100 is shown and includes a plurality of openings 102 in the wall of the flow sleeve. The flow sleeve 100 also includes a plurality of lugs 104, which are used for positioning a combustion liner within the flow sleeve 100. Also located within the flow sleeve 100 are a plurality of brackets 106, which are more clearly depicted in FIG. 8. The cross fire tubes 108 are placed through the openings 102 and brackets 106, as shown in FIG. 9.

[0027] Referring now to FIG. 10, the cross fire tubes 108 are secured in the flow sleeve / liner by a clip 110. The clip 110 includes a hook portion 112 which can be used to help install and remove the clip 110 from the flow sleeve 100. As a result of the configuration of the bracket 106 and clip 110, each of these features extend inward and into a flow path between the flow sleeve 100 and combustion liner (not shown), thus interfering and restricting the flow of air passing between the flow sleeve 100 and a combustion liner. In a common configuration, the bracket 106 and clip 110 extend over half an inch (25.4 mm) into the flow path of the passing airflow thus adversely impacting air flow to a combustor and combustion dynamics and emissions.

[0028] Some embodiments of the present disclosure are depicted in FIGS. 11-14. Referring initially to FIG. 11, a flow sleeve 500 for use in a gas turbine combustion system is shown. The flow sleeve 500 comprises a generally annular body 502 having a flange 504 at a forward end 506 of the generally annular body 502. The flow sleeve 500 also includes one or more openings 508 spaced about the generally annular body 502. These one or more openings 508 are preferably two openings, as shown in FIG. 11, and are used for communicating with combustors adjacent to the flow sleeve 500. This communication occurs via a plurality of cross fire tubes which serve as a conduit in which a flame can be passed from one combustor to an adjacent combustor.

[0029] Referring now to FIGS. 11 and 12, the flange 504 of flow sleeve 500 further comprises one or more recessed portions 510. The one or more recessed portions 510 correspond directly to the one or more openings 508 in the generally annular body 502. Thus, for the embodiment of the present disclosure shown in FIG. 11, the flow sleeve 500 includes two recessed portions 510 in the flange 504. Within the recessed portions 510 is a plurality of holes 512 which provide a way of securing a retainer clip, as will be discussed in more detail below. The plurality of holes 512 can be through holes or threaded.

[0030] The flow sleeve 500 also comprises a clip block 514 positioned axially between the one or more recessed portions 510 and the one or more openings 508. The clip block 514, which in one embodiment is welded to the generally annular body 502, is used to secure a retainer clip and cross fire tube in place, as discussed in more detail below. The clip block 514 can be formed of a variety of shapes depending on the specific cross fire tube and retainer clip geometry. For the embodiment depicted in FIGS. 11-14, the clip block 514 has a T-shaped cross section. In an embodiment, the clip block 514 has a first leg 515A and a second leg 515B generally perpendicular thereto. When the clip block 514 is secured (e.g., welded) to the generally annular body 502, a gap 517A may be formed between the annular body 502 and the first leg 515A at one side of the second leg 515B, and a gap 517B may be formed between the annular body 502 and the first leg 515A at an opposing side of the second leg 515B.

[0031] As discussed above, in a can-annular combustor configuration, a combustion liner is located within a flow sleeve. Compressed air from an engine compressor is directed between the combustion liner and flow sleeve in order to cool the combustion liner and direct the air into the combustion liner. As a result, this air is also preheated before entering the combustion liner and undergoes a combustion process to generate hot combustion gases for powering the turbine section. In order to properly locate the combustion liner within the flow sleeve, a plurality of pegs 516 extend radially inward from the generally annular body 502. Mounting tabs extend radially outward from a combustion liner and slide into the slots in the plurality of pegs 516.

[0032] Another feature of the present disclosure is shown in FIG. 13. A retention clip 700 is provided for securing a cross fire tube in a gas turbine combustor. The retention clip 700 comprises a mounting plate 702 having one or more mounting holes 703 located therein and a centerbody 704 extending from the mounting plate 702. As can be seen in FIG. 13, the mounting plate 702 is generally perpendicular to the centerbody 704. The centerbody 704 has a through hole 706 and a slot 708 extending away from the through hole 706. Extending away from the centerbody 704 are two fingers, a first finger 710 and a second finger 712. The first finger 710 is separated from the second finger 712 by an axially extending space 714. The axially extending space 714, the slot 708, and the through hole 706 permit the first and second fingers 710 and 712 to expand in opposing directions in multiple planes, such that the fingers can expand to surround another component positioned in the axially extending space 714.

[0033] Referring still to FIG. 13, another feature of the retention clip 700 is at least one curved portion, or bend, 716 that extends along a portion of the centerbody 704 and / or the first and second fingers 710 and 712. The at least one curved portion 716 shown in FIG. 13 comprises two portions curved in opposing directions. In the embodiment depicted, the at least one curved portion is located along the centerbody 704 between the through hole 706 and the first and second fingers 710 and 712. Furthermore, one curved portion 716 curves in a direction towards the mounting plate 702 while the adjacent curved portion 716 curves away from the mounting plate 702. This set of opposing curve portions creates a spring effect in the clip 700 when the clip 700 is placed against adjacent mating surfaces, such as the annular body 502 of flow sleeve 500.

[0034] The retainer clip can be made from a variety of materials but is preferably made in a flat pattern from a material capable of withstanding the temperatures adjacent the cross fire tubes as well as the adjacent components. Such acceptable materials may include a tool steel as well as Inconel ®< X-750, a nickel-chromium alloy. The retainer clip 700 can be cut from a plate, typically 1.575 mm to 3.175 mm thick. Features such as the through hole 706, slot 708, and axially extending space 714 are cut out of the plate material while in a flat pattern, typically by a laser or wire EDM and then the mounting flange 702 is bent at approximately 90-degree angle relative to the centerbody 704.

[0035] Referring now to FIG. 14, a system 800 for retaining a cross fire tube in a gas turbine engine having multiple combustors is disclosed. The system 800 utilizes the features discussed above with respect to FIGS. 11-13. As such, the terminology used to describe the system 800 and its assembly process will incorporate terms and reference identifiers discussed above.

[0036] The system 800 comprises a flow sleeve 500 having a generally annular body 502, a flange 504 with one or more recessed portions 510 located therein. The generally annular body 502 of the flow sleeve 500 also includes one or more openings 508 as well as a clip block 514 positioned between the one or more recessed portions 510 and the one or more openings 508.

[0037] A tube 802 extends through the one or more openings 508 of the flow sleeve 500. This tube, also known as a cross fire tube may comprise multiple tubes, often in a telescoping arrangement for connecting adjacent combustors. The tube 802 may also include a groove about its outer surface 804 for receiving the retention clip 700. As shown in FIG. 14, the retention clip 700 extends along an inner surface 806 of the generally annular body 502 with the mounting plate 702 engaging the recessed portion 510 in the flange 504. The centerbody 704 of the retention clip 700 extends from the mounting plate 702 and to the first finger 710 and the second finger 712, each of which surround a portion of the tube 802, thus preventing the tube 802 from moving into or out of the one or more openings 508 in the generally annular body 502.

[0038] In operation, once a flow sleeve is installed in adjacent combustor cases, one or more tubes 802, also commonly referred to as cross fire tubes, are passed through the openings 508 in the flow sleeve annular body 502. Then, a combustion liner is installed into the flow sleeve 500. Once the combustion liner is positioned within the flow sleeve 500, the tubes 802 are slid into the corresponding combustion liner. Once the tubes 802 are in the appropriate position through the flow sleeve and into the combustion liner, the retention clip 700 is positioned between the inner surface 806 of the generally annular body 502 and the clip block 514, such that each of the first finger 710 and second finger 712 extends at least partially through one of the gaps 517A and 517B and the fingers 710, 712 expand to surround at least a portion of the tube 802. The retention clip is slid into the flow sleeve 500 until the mounting plate 702 is positioned within the recessed portion 510 of the flange 504. Then, the mounting plate is secured to the flange 504 by placing a plurality of fasteners (not depicted) through mounting holes 703 in the mounting plate 702 and into the holes 512 in the recessed portion 510 of the flange 504.

[0039] Due to the curvatures 716 in the retention clip 700, and as discussed above, the retention clip 700 provides some resistance as it is positioned in place between the inner surface 806 of the generally annular body 502 and the clip block 514. This further aids in preventing accidental removal of the retention clip 700.

[0040] As can be seen from FIG. 14, the clip block 514 and retention clip 700 are positioned closer to the inner surface 806 of the generally annular body 502 than in prior art configurations. More specifically, the retention clip 700 of the present disclosure extends radially into the flow sleeve 500, and thus the airflow between the flow sleeve and combustion liner, by approximately 6.86 mm. The prior art configuration, as depicted in FIGS. 7-11, extends into the flow sleeve more than twice as much, or upwards of 14.2 mm, thereby creating a much larger blockage than the present disclosure. As one skilled in the art will appreciate, a blockage in compressed air can limit the air flow to the combustor, thus adversely impacting combustor emissions and impacting combustion dynamics.

[0041] Although a preferred embodiment of this disclosure has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure. Since many possible embodiments may be made of the disclosure without departing from the scope thereof, it is to be understood that all matter herein set forth or shown in the accompanying drawings is to be interpreted as illustrative and not in a limiting sense.

[0042] From the foregoing, it will be seen that this disclosure is one well adapted to attain all the ends and objects hereinabove set forth together with other advantages which are obvious and which are inherent to the structure.

[0043] It will be understood that certain features and subcombinations are of utility and may be employed without reference to other features and sub combinations.

Claims

1. A combustion liner (200) comprising: A generally annular body (202) having a first cylindrical portion (204), a conical portion (206), and a second cylindrical portion (208); an inlet (210) proximate the first cylindrical portion (204) and an outlet (212) proximate the second cylindrical portion (208); a plurality of turbulators (216) located along an outer surface (218) of the first cylindrical portion (204) and the conical portion (206), the turbulators (216) each having a first side (220) with a first ramp angle (α), a second side (222) with a second ramp angle (β), a height (224), and a base width (226), wherein each of the plurality of turbulators (216) extends circumferentially around the whole periphery of the combustion liner (200) in a direction perpendicular to the axis of the combustion liner (200), wherein the first ramp angle (α) and the second ramp angle (β) are approximately 30 - 45 degrees as measured from a surface (218) of the first cylindrical portion (204) and the conical portion (206), wherein the plurality of turbulators (216) have an axial spacing (230) of approximately 10-20 times the height (224), wherein the plurality of turbulators (216) have a generally triangular cross section, and wherein the base width (226) is approximately 1-3 times the height (224).

2. The combustion liner of claim 1, wherein the plurality of turbulators (216) are axisymmetric.

3. The combustion liner of claim 1 further comprising a sealing mechanism (214) located along an outer surface (218) of the second cylindrical portion (208).

4. The combustion liner of claim 1 further comprising a base fillet radius (R) between the first and second sides (220 and 222) and the outer surface (218) of the first cylindrical portion (204) and the conical portion (206).

5. The combustion liner of claim 1, wherein the plurality of turbulators (216) are integral with the generally annular body (202) and the conical portion (206).

6. The combustion liner of claim 1 further comprising a full round radius at a tip region (228) of the plurality of turbulators (216).

7. Combustor case (242) comprising a flow sleeve (200) and a combustion liner (200) according to any of the previous claims, wherein the combustion liner (200) is positioned within the flow sleeve (240) and the combustor case (242), wherein the combustion liner (200) and the flow sleeve (240) form a passageway (244) located therebetween and through which, in use, air passes.

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