The present invention relates generally to cooling of gas turbine engine components and more specifically to cooling of adjoining mate faces in cooled gas turbine engine components, such as shrouds and platforms.
Gas turbine engines operate by passing a volume of high energy gases through a plurality of stages of vanes and blades, each having an airfoil, in order to drive turbines to produce rotational shaft power. The shaft power is used to drive a compressor to provide compressed air to a combustion process to generate the high energy gases. Additionally, the shaft power is used to drive a generator for producing electricity, or to drive a fan for producing high momentum gases for producing thrust. In order to produce gases having sufficient energy to drive the compressor, generator and fan, it is necessary to combust the fuel at elevated temperatures and to compress the air to elevated pressures, which also increases its temperature. Thus, the vanes and blades are subjected to extremely high temperatures, often times exceeding the melting point of the alloys comprising the airfoils. High pressure turbine blades are subject to particularly high temperatures.
In order to maintain gas turbine engine turbine blades at temperatures below their melting point, it is necessary to, among other things, cool the blades with a supply of relatively cooler air, typically bled from the compressor. The cooling air is directed into the blade to provide convective cooling internally and film cooling externally. For example, cooling air is passed into interior cooling channels of the airfoil to remove heat from the alloy, and subsequently discharged through cooling holes to pass over the outer surface of the airfoil to prevent the hot gases from contacting the vane or blade directly. Various cooling air channels and hole patterns have been developed to ensure sufficient cooling of various portions of the turbine blade.
A typical turbine blade is connected at its inner diameter ends to a rotor, which is connected to a shaft that rotates within the engine as the blades interact with the gas flow. The rotor typically comprises a disk having a plurality of axial retention slots that receive mating root portions of the blades to prevent radial dislodgment. Blades typically also include integral inner diameter platforms that prevent the high temperature gases from escaping through the radial retention slots. It is desirable to further provide targeted cooling to the platforms to cool the surfaces between adjacent platforms. There is a continuing need to improve cooling of turbine blade platforms to increase the temperature to which the blade can be exposed, thereby increasing the overall efficiency of the gas turbine engine.
The present invention is directed toward a gas turbine engine component, such as a shroud, platform or blade outer air seal. The gas turbine engine component comprises a shroud, a U-channel, an internal cooling air passage and a U-channel cooling hole. The shroud comprises a forward face, an aft face, a first side face and a second side face. The U-channel is disposed in the aft face of the shroud. A gas path surface connects the forward face, aft face, first side face and second side face. A cooled surface connects the forward face, aft face, first side face and second side face opposite the gas path face. The internal cooling air passage extends through the shroud. The U-channel cooling hole extends into the first side face of the shroud adjacent the U-channel to intersect the internal cooling passage.
Inlet air A enters engine 10 and it is divided into streams of primary air AP and secondary air AS after it passes through fan 12. Fan 12 is rotated by low pressure turbine 22 through shaft 24 to accelerate secondary air AS (also known as bypass air) through exit guide vanes 26, thereby producing a major portion of the thrust output of engine 10. Shaft 24 is supported within engine 10 at ball bearing 25A, roller bearing 25B and roller bearing 25C. Low Pressure Compressor (LPC) 14 is also driven by shaft 24. Primary air AP (also known as gas path air) is directed first into LPC 14 and then into high pressure compressor (HPC) 16. LPC 14 and HPC 16 work together to incrementally step-up the pressure of primary air AP. HPC 16 is rotated by HPT 20 through shaft 28 to provide compressed air to combustor section 18, which includes inlet guide vanes 29. Shaft 28 is supported within engine 10 at ball bearing 25D and roller bearing 25E. The compressed air is delivered to combustors 18A and 18B, along with fuel through injectors 30A and 30B, such that a combustion process can be carried out to produce the high energy gases necessary to turn turbines 20 and 22, as is known in the art. Primary air AP continues through gas turbine engine 10 whereby it is typically passed through an exhaust nozzle to further produce thrust.
HPT 20 and LPT 22 each include a circumferential array of blades extending radially from discs 31A and 31B connected to shafts 28 and 24, respectively. Similarly, HPT 20 and LPT 22 each include a circumferential array of vanes extending radially from HPT case 23D and LPT case 23E, respectively. Specifically, HPT 20 includes blades 32A and 32B and vanes 34A. Blades 32A and 32B include internal channels or passages into which compressed cooling air AC air from, for example, HPC 16 is directed to provide cooling relative to the hot combustion gasses. Blade 32A of the present invention includes a platform having mate face cooling holes for cooling a trailing edge U-channel. Although described with reference to blade 32A, the cooling holes of the present invention may be used in other gas turbine engine components having a U-channel, such as turbine vanes, shrouds and blade outer air seals.
Airfoil 36 and airfoil 40 extend from their respective inner diameter supports toward engine case 23D, across gas path 52. Hot combustion gases of primary air AP are generated within combustor 18 (
The cooling air AC directed into blade 32A is passed into airfoil 40 to cool exterior surfaces of airfoil 40, which includes film cooling holes as is known in the art. In the present invention, a portion of cooling air AC is directed to side faces of platform 42 that abut or adjoin mating faces of adjacent platforms. This cooling air provides direct impingement cooling of the platform mate faces, but also provides film and impingement cooling to U-channel 56, as is discussed with reference to
Turbine blade 32A is positioned in gas path 52 such that a flow of primary air AP flows across airfoil 40 and over gas path surface 82 of platform 42. Cooling air AC travels underneath platform 42 against inner surface 84, and through blade 32A within passages 78A and 78B. In one embodiment, second flange 94 comprises an angel wing seal that cooperates with a seal fin of an adjacent vane. A fin of stator vane 34 (
The labyrinth seal formed by U-channel 56 prevents the ingestion of primary air AP into cavity 54 (
Cooling air for U-channel cooling hole 70 is provided from passage 78B. Cooling air exiting U-channel cooling hole 70 directly impacts a platform 42 of an adjacent turbine blade, thereby providing direct impingement cooling. Cooling hole 70 is positioned so that the cooling air impinges on portions of platform 42 forming U-channel 56. Specifically, U-channel cooling hole 70 is positioned at the juncture, or apex, of first flange 92, second flange 94 and base 96, beneath trailing edge 64 of airfoil 40. Thus, from hole 70, the cooling hole can disperse along mate face 80. Furthermore, the cooling air fills the gap between adjacent platforms 42 with a shroud of cooling air that shrouds over the top of U-channel 56. Thus, a film of cooling air forms an air dam that blocks ingestion of primary air AP into U-channel 56. Additionally, the cooling air ultimately curls around base 96 to enter into U-channel 56 to further dilute any primary air AP that may have entered therein.
Cooling air from U-channel cooling hole 70 is supplemented with cooling air from forward, augmenting cooling holes 72 and 74. Cooling air for cooling holes 72 and 74 is provided from passage 78A. Cooling air from holes 72 and 74 directly impacts a platform 42 of an adjacent turbine blade, thereby providing direct impingement cooling. Cooling air from holes 72 and 74 also fortifies cooling air from hole 70 such that a stronger, more forceful combined flow of cooling air is formed to more effectively block primary air AP. Furthermore, the combined flow is cooler and better able to dilute primary air that has entered U-channel 56.
As indicated in
The U-channel cooling hole scheme of the present invention has been described with respect to a platform of a turbine blade, but may also be used in other gas turbine engine components such as turbine vanes, compressor blades, compressor vanes, shrouds and blade outer air seals. For example, cooling holes 70, 72 and 74 may be positioned in mate faces of platform 38B of vane 29, or in blade outer air seal (BOAS) 100 (
While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.
The following are non-exclusive descriptions of possible embodiments of the present invention.
A turbine blade comprises: an airfoil, a platform surrounding a base of the airfoil, a U-channel disposed in an aft face of the platform, a root extending from the platform opposite the airfoil, an internal cooling passage extending through the turbine blade, and a U-channel cooling hole extending from the internal cooling passage to a mate face of the platform upstream of the U-channel.
The turbine blade of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
the airfoil comprises: a leading edge, a trailing edge, a pressure side extending between the leading edge and the trailing edge with a predominantly concave curvature, a suction side extending between the leading edge and the trailing edge with a predominantly convex curvature, and a span extending radially from an inner diameter base to a outer diameter tip, wherein the U-channel cooling hole extends into a pressure side mate face of the platform;
the U-channel cooling hole is positioned radially inward of a trailing edge of the airfoil;
the platform comprises: the aft face, a forward face opposite the aft face, an upper surface defining an end wall from which the airfoil extends, a lower surface opposite the upper surface and from which the root extends, a first side face, and a second side face comprising the mate face into which the U-channel cooling hole extends;
the U-channel comprises: a first flange comprising: a first proximate end extending from the platform, and a first distal end opposite the first proximate end; a base extending radially inward from the first proximate end; and a second flange comprising: a second proximate end extending from the base, and a second distal end opposite the second proximate end;
the second flange comprises an angel wing seal and is longer than the first flange;
the base is arcuate;
the U-channel cooling hole is positioned at an apex between the base, the first flange and the second flange;
the internal cooling channel passage comprises: forward and aft channels extending through the airfoil, wherein the U-channel cooling hole extends to the aft channel;
the internal cooling channel further comprises: first and second feed channels extending through the root and joining to the forward channel, and third and fourth feed channels extending through the root and joining to the aft channel;
a pair of forward cooling holes extending into the side face of the platform upstream of the U-channel cooling hole;
the U-channel cooling hole extends straight between an inlet and an outlet; and
the U-channel cooling hole extends from the internal cooling passage to the side face of the platform with a downstream vector component.
A method for cooling a U-channel in a gas turbine engine shroud comprises: flowing cooling air through an internal cooling passage of the turbine engine shroud; directing a portion of the cooling air through a U-channel cooling hole extending from the internal cooling passage to a mate face of the gas turbine engine shroud upstream of the U-channel; and passing the portion of the cooling air into the U-channel.
The method of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features and/or additional steps:
the step of forming an air dam above the U-channel with the portion of the cooling air to prevent hot combustion gas from entering the U-channel;
the step of augmenting the portion of the cooling air passing through the U-channel cooling hole with additional cooling air from an additional cooling hole extending from the internal cooling passage to the mate face upstream of the U-channel cooling hole; and
the step of forming a layer of film cooling air along the mate face with the portion of the cooling air.
A gas turbine engine component comprises: a shroud comprising a forward face, an aft face, a first side face and a second side face; a U-channel disposed in the aft face of the shroud; a gas path surface connecting the forward face, aft face, first side face and second side face; a cooled surface connecting the forward face, aft face, first side face and second side face opposite the gas path face; an internal cooling air passage extending through the shroud; and a U-channel cooling hole extending into the first side face of the shroud adjacent the U-channel to intersect the internal cooling passage.
The gas turbine engine component of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
a first flange comprising: a first proximate end extending from the aft face of the platform, and a first distal end opposite the first proximate end; a base extending radially inward from the first proximate end; and a second flange comprising: a second proximate end extending from the base, and a second distal end opposite the second proximate end;
a pair of forward cooling holes extending into the first side face of the shroud upstream of the U-channel cooling hole; and
an airfoil extending radially outward from the gas path surface, the airfoil having a leading edge, a trailing edge, a pressure side, a suction side, an outer diameter end and an inner diameter end, and a root extending radially inward from the cooled surface.