GAS TURBINE DRIVE WITH COOLED TURBINE BLADES WITH COVER PLATE WITH FLOW INDUCTOR

Description

FIELD OF THE INVENTION

[0001] This invention relates generally to a flow inducer assembly for cooling turbine blades of a gas turbine engine, in particular, the last stage turbine blades of the gas turbine engine, using ambient air.DESCRIPTION OF RELATED ART

[0002] An industrial gas turbine engine typically includes a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, a turbine section for producing mechanical power, and a generator for converting the mechanical power to an electrical power. The turbine section includes a plurality of turbine blades that are attached on a rotor disk. The turbine blades are arranged in rows axially spaced apart along the rotor disk and circumferentially attached to a periphery of the rotor disk. The turbine blades are driven by the ignited hot gas from the combustor and are cooled using a coolant, such as a cooling fluid, through cooling passages in the turbine blades.

[0003] Typically, cooling fluid may be supplied by bleeding compressor air. However, bleeding air from the compressor may reduce turbine engine efficiency. Due to high operation pressures of the first, second and third stage turbine blades, bleeding compressor air may be required for cooling the first, second and third stage turbine blades. The last stage turbine blades operate under the lowest pressure, ambient air may be used for cooling the last stage turbine blades. In order to sufficiently cool the last stage turbine blades to achieve required boundary conditions, an efficient flow inducer system is needed to bring sufficient amount of the ambient air into cooling passages of the last stage turbine blade. There is a need to provide an easy and simple system to capture sufficient amount of ambient air into the cooling passages of the last stage turbine blade for sufficiently cooling the last stage turbine blades. US 2006 / 120855 A1 discloses a gas turbine engine with turbine blades having roots received in blade mounting sections of disk grooves of a rotor disk, each disk groove further comprising a disk cavity. Seal plates are attached to upstream side of the rotor disk, wherein each seal plate comprises an upper seal plate portion covering the blade root and a lower seal plate portion covering the disk cavity. Flow inducer assemblies are integrated to each seal plate and are configured to function as a paddle due to rotation of the rotor disk during operation of the gas turbine engine to induce a cooling fluid into the disk cavity and cause it to enter inside the turbine blade from the blade root for cooling the turbine blade. US 3,137,478 A1 describes a cover plate assembly for sealing spaces between turbine buckets, wherein a rim of a turbine wheel is provided with an axial projection and an undercut groove, which corresponds to an axial projection and an undercut groove on a cover plate base. EP 3 121 372 A1 describes a cooled turbine runner for a gas turbine, in particular an aircraft engine, with turbine blades that are radially arranged at the circumferential surface of a rotor disk, wherein respectively a turbine blade is inserted with a profiled blade root into a correspondingly profiled disk finger groove at the circumferential surface of the rotor disk. A cooling device with at least one cooling air supply channel is provided, which extends at least substantially axially and at least over a part of the axial length of the blade root, and which has at least one inlet with an inlet opening at an inflow side of the blade root, wherein the inlet is embodied with a projection.SUMMARY OF THE INVENTION

[0004] Briefly described, aspects of the present invention relate to a gas turbine engine as defined in independent claim 1.

[0005] According to an aspect, a gas turbine engine is presented. The gas turbine engine comprises a rotor disk comprising a plurality of circumferentially distributed disk grooves. Each disk groove comprises a blade mounting section and a disk cavity. The gas turbine engine comprises a plurality of turbine blades. Each turbine blade comprises a platform and a blade root that extends radially downward from the platform and is inserted into the blade mounting section of a respective disk groove. The gas turbine engine comprises a plurality of seal plates attached with respect to an axial flow direction to an aft side circumference of the rotor disk. Each seal plate comprises a radially extending upper seal plate wall and a radially extending lower seal plate wall. The upper seal plate wall is configured to cover the blade root of a respective turbine blade. The gas turbine engine comprises a plurality of flow inducer assemblies. Each flow inducer assembly is integrated to each seal plate at a side facing away from the rotor disk. Each flow inducer assembly aligns with a respective disk cavity. The disk cavity is formed between the respective blade root and a radial bottom of the respective disk groove. The lower seal plate wall of each flow inducer assembly comprises an aperture that is configured to align with the respective disk cavity. Each flow inducer assembly comprises a curved plate that is integrated to the lower seal plate wall and axially extends out from the lower seal plate wall. Each curved plate is attached radially along the aperture at a downstream side with respect to a rotation direction of the rotor disk and has a similar curvature with the aperture. Each flow inducer assembly is configured to function as a paddle due to rotation of the rotor disk and each seal plate therewith during operation of the gas turbine engine to drive a cooling fluid into the respective disk cavity and cause it to enter inside of the respective turbine blade from blade root for cooling the turbine blade.

[0006] Various aspects and embodiments of the application as described above and hereinafter may not only be used in the combinations explicitly described, but also in other combinations. Modifications will occur to the skilled person upon reading and understanding of the description.BRIEF DESCRIPTION OF THE DRAWINGS

[0007] Exemplary embodiments of the application are explained in further detail with respect to the accompanying drawings. In the drawings. FIG. 1 illustrates a schematic perspective view of a portion of a gas turbine engine showing the last stage, in which embodiments of the present invention may be incorporated; FIGs. 2 to 3 illustrate schematic perspective views of flow inducer assemblies according to various embodiments of the present invention;

[0008] To facilitate understanding, identical reference numerals have been used, where possible, to designate identical elements that are common to the figures.DETAILED DESCRIPTION OF THE INVENTION

[0009] A detailed description related to aspects of the present invention is described hereafter with respect to the accompanying figures.

[0010] FIG. 1 illustrates a schematic perspective view of a portion of a gas turbine engine 100 showing the last stage looking in an aft side with respect to an axial flow direction. The gas turbine engine 100 includes a flow inducer assembly 300 according to embodiments of the present invention. As illustrated in FIG. 1, the gas turbine engine 100 includes a last stage rotor disk 120 and a plurality of last stage turbine blades 140 that are attached along an outer circumference of the rotor disk 120. A plurality of seal plates 200 are attached to the aft side circumference of the last stage rotor disk 120. The seal plate 200 may prevent hot gas coming into the aft side of the rotor disk 120. The seal plates 200 are secured to the rotor disk 120. The rotor disk 120 may rotate in a direction as indicated by the arrow R during operation of the gas turbine engine 100, which rotates the turbine blades 140 and the seal plates 200 therewith in the same direction R. For clarity purpose, one turbine blade 140 and one seal plate 200 are removed from the rotor disk 120.

[0011] With reference to FIG. 1, the rotor disk 120 includes a plurality of disk grooves 122. Each disk groove 122 includes a blade mounting section 124 and a disk cavity 126. Each turbine blade 140 includes a platform 142 and a blade root 144 that extends radially downward from the platform 142. Each turbine blade 140 is attached to the rotor disk 120 by inserting the blade root 144 into the blade mounting section 124 of the rotor disk groove 122. The disk cavity 126 is formed between the blade root 144 and bottom of the disk groove 122. Each seal plate 200 includes an upper seal plate wall 220 and a lower seal plate wall 240. A seal arm 230 may extend axially outward between the upper seal plate wall 220 and the lower seal plate wall 240. The upper seal plate wall 220 covers the blade root 144. A flow inducer assembly 300 is attached to the lower seal plate wall 240. The flow inducer assembly 300 aligns with the disk cavity 126 of the disk groove 122.

[0012] During engine operation, rotation of the last stage turbine blades 140 creates pumping force to drive cooling fluid into the disk cavity 126 of the disk groove 120 as indicated by the cooling flow arrow 130 due to centrifugal force. The cooling fluid enters inside of the turbine blade 140 from the blade root 144 for cooling the turbine blade 140 and exits through openings in the turbine blade 140 to gas path of the gas turbine engine 100. The cooling fluid may be ambient air. According to embodiments of the present invention, the flow inducer assembly 300 arranged on the seal plate 200 provides further driving force to induce ambient air entering the disk cavity 126 for sufficiently cooling the last stage turbine blade 140. The flow inducer assembly 300 and the seal plate 200 may be manufactured as an integrated single piece.

[0013] FIGs. 2 to 3 illustrate schematic perspective views of a seal plate 200 having an integrated flow inducer assembly 300 according to various embodiments of the present invention.

[0014] FIG. 2 illustrates a schematic perspective view of a seal plate 200 having an integrated flow inducer assembly 300 according to an embodiment of the present invention. As shown in FIG. 2, the seal plate 200 includes an upper seal plate wall 220 and a lower seal plate wall 240. A seal arm 230 extends axially outward between the upper seal plate wall 220 and the lower seal plate wall 240. The seal plate 200 may have a hook 202 displaced at a side of the upper seal plate wall 220 facing to the rotor disk 120. The hook 202 may have a U-shape that attaches to the rotor disk 120. The seal plate 200 may have a protrusion 204 protruded from a side of the lower seal plate wall 240 facing to the rotor disk 120. The protrusion 204 may have a dovetail shape that attaches to the rotor disk 120. The hook 202 and the protrusion 204 secure the seal plate 200 to the rotor disk 120. The seal plate 200 has an aperture 242 axially penetrating through the lower seal plate wall 240. The aperture 242 may be located at the lower seal plate wall 240 with a radial distance below the seal arm 230. The aperture 242 aligns with the disk cavity 126 of the disk groove 122 after assembly. The aperture 242 may generally have a similar shape with the disk cavity 126.

[0015] According to an exemplary embodiment as illustrated in FIG. 2, a flow inducer assembly 300 is integrated to the seal plate 200 at a side facing away from the rotor disk 120 extending outward in an axial direction. The flow inducer assembly 300 includes a curved plate 310 attached radially along the aperture 242 at a downstream side with respect to the rotation direction R of the rotor disk 120 as shown in FIG. 1. The curved plate 310 may be blended with the aperture 242 in a tangential direction of the aperture 242. The curved plate 310 has a similar curvature with the aperture 242. During operation of the gas turbine engine 100, rotation of the rotor disk 120 and the seal plate 200 therewith makes the curved plate 310 of the flow inducer assembly 300 functioned as a paddle that further induces cooling air 130, such as ambient air from outside of the gas turbine engine 100, in addition to centrifugal force caused by rotation of the turbine blades 140, into the aperture 242 and the disk cavity 126 and causes it to enter inside of the turbine blades 140 from the blade roots 144 for cooling the turbine blades 140. The curved plate 310 may have a scoop shape.

[0016] Dimensions of the flow inducer assembly 300 may be designed to achieve cooling requirement for sufficiently cooling the turbine blades 140. Dimensions of the flow inducer assembly may include a radial height of the curved plate 310, an axial length of the curved plate 310, etc. A radial height of the curved plate 310 may be less than, or equal to, or greater than a radial height of the aperture 242. For illustration purpose, FIG. 2 and FIG. 3 show the curved plates 310 having different radial heights. According to an exemplary embodiment as illustrated in FIG. 2, a radial height of the curved plate 310 is equal to a radial height of the aperture 242. As illustrated in FIG. 2, the curved plate 310 is attached along the aperture 242 at the downstream side starting from the lowest point of the aperture 242 and ending at the highest point of the aperture 242.

[0017] According to another exemplary embodiment as illustrated in FIG. 3, a radial height of the curved plate 310 is greater than a radial height of the aperture 242. As illustrated in FIG. 3, the curved plate 310 is attached along the aperture 242 at the downstream side starting from the lowest point of the aperture 242 and ending at the seal arm 230. Such embodiment may also improve mechanical properties of the flow inducer assembly 300, such as increasing mechanical strength, reducing vibration, etc. It is understood that the curved plate 310 may be attached along the aperture 242 at the downstream starting at a radial point that is below the lowest point of the aperture 242, or above the lowest point of the aperture 242. It is also understood that the curved plate 310 may be attached along the aperture 242 at the downstream side ending at a radial point that is below the highest point of the aperture 242, or between the highest point of the aperture 242 and the seal arm 230.

[0018] An axial length of the curved plate 310 may change along a radial direction. According to exemplary embodiments as illustrated in FIG. 2 and FIG. 3, the axial length of the curved plate 310 may be shorter in the lower portion and longer in the upper portion. For example, the maximum axial length of the curved plate 310 from the lower seal plate wall 240 may be located at the upper portion of the curved plate 310 that is near a region of the top of the curved plate 310.

[0019] According to an aspect, the proposed flow inducer assembly 300 may enable using ambient air as cooling fluid 130 for sufficiently cooling the last stage of turbine blades 140 of a gas turbine engine 100. During operation of the gas turbine engine 100, rotation of the rotor disk 120 and the seal plate 200 therewith makes the flow inducer assembly 300 function as a paddle that drives sufficient amount of ambient air from outside of the gas turbine engine 100 as the cooling air 130 into disk cavities 126 of rotor disk 120 and enters insides of the turbine blades 140 from the blade roots 144 for cooling the turbine blades 140. The proposed flow inducer assembly 300 eliminates bleeding compressor air for cooling the last stage of turbine blades 140, which increases turbine engine efficiency.

[0020] According to an aspect, the proposed flow inducer assembly 300 may be manufactured as an integrated piece of the seal plate 200. The seal plate 200 and the integrated flow inducer assembly 300 provide a lightweight design for preventing hot gas coming into the rotor disk 120 and simultaneously driving enough ambient air for sufficiently cooling the last stage of turbine blades 140 to achieve required boundary condition. The seal plate 200 and the integrated flow inducer assembly 300 provide sufficient cooling of the last stage of the turbine blades 140 with minimal cost.

[0021] Although various embodiments that incorporate the teachings of the present invention have been shown and described in detail herein, those skilled in the art can readily devise many other varied embodiments that still incorporate these teachings. The invention is not limited in its application to the exemplary embodiment details of construction and the arrangement of components set forth in the description or illustrated in the drawings. The invention is capable of other embodiments and of being practiced or of being carried out in various ways. Also, it is to be understood that the phraseology and terminology used herein is for the purpose of description and should not be regarded as limiting. The use of "including," "comprising," or "having" and variations thereof herein is meant to encompass the items listed thereafter and equivalents thereof as well as additional items. Unless specified or limited otherwise, the terms "mounted," "connected," "supported," and "coupled" and variations thereof are used broadly and encompass direct and indirect mountings, connections, supports, and couplings. Further, "connected" and "coupled" are not restricted to physical or mechanical connections or couplings.Reference List:

[0022] 100:Gas Turbine Engine 120:Rotor Disk 122:Disk Groove 124:Blade Mounting Section 126:Disk Cavity 128:Disk Slot 130:Cooling Flow 140:Turbine Blade 142:Blade Platform 144:Blade Root 200:Seal Plate 202:Seal Plate Hook 204:Seal Plate Protrusion 220:Upper Seal Plate Wall 230:Seal Arm 240:Lower Seal Plate Wall 242:Aperture on Lower Seal Plate Wall 244:Seal Plate Root 246:Locking Plate 300:Flow Inducer Assembly 310:Curved Plate having Scoop Shape 320:Floor Plate 330:Inner Side Wall 340:Outer Side Wall 342:Vertical Wall 350:Cooling Fluid Inlet

Claims

1. A gas turbine engine (100) comprising: a rotor disk (120) comprising a plurality of circumferentially distributed disk grooves (122), wherein each disk groove (122) comprises a blade mounting section (124) and a disk cavity (126); a plurality of turbine blades (140), wherein each turbine blade (140) comprises a platform (142) and a blade root (144) that extends radially downward from the platform (142) and is inserted into the blade mounting section (124) of a respective disk groove (122); a plurality of seal plates (200) attached with respect to an axial flow direction to an aft side circumference of the rotor disk (120), wherein each seal plate (200) comprises a radially extending upper seal plate wall (220) and a radially extending lower seal plate wall (240), wherein the upper seal plate wall (220) is configured to cover the blade root (144) of a respective turbine blade (140); and a plurality of flow inducer assemblies (300), wherein each flow inducer assembly (300) is integrated to each seal plate (200) at a side facing away from the rotor disk (120), wherein each flow inducer assembly (300) aligns with a respective disk cavity (126), wherein the disk cavity (126) is formed between the respective blade root (144) and a radial bottom of the respective disk groove (122), wherein the lower seal plate wall (240) of each flow inducer assembly (300) comprises an aperture (242) that is configured to align with the respective disk cavity (126), wherein each flow inducer assembly (300) comprises a curved plate (310) that is integrated to the lower seal plate wall (240) and axially extends out from the lower seal plate wall (240), wherein each curved plate (310) is attached radially along the respective aperture (242) at a downstream side with respect to a rotation direction of the rotor disk (120) and has a similar curvature with the aperture (242), and wherein each flow inducer assembly (300) is configured to function as a paddle due to rotation of the rotor disk (120) and each seal plate (200) therewith during operation of the gas turbine engine (100) to induce a cooling fluid (130) into the respective disk cavity (126) and cause it to enter inside of the respective turbine blade (140) from its blade root (144) for cooling the turbine blade (140).

2. The gas turbine engine (100) as claimed in claim 1, wherein each curved plate (310) comprises a scoop shape.

3. The gas turbine engine (100) as claimed in claim 1, wherein the cooling fluid (130) comprises ambient air.