Patent Application
20200325780
The present invention relates to gas turbines, and more particularly to blades or vanes of gas turbines with cooling channels.
Cooling of gas turbine components, such as a turbine blade or a vane is a major challenge and an area of interest in turbine technology. A common technique for cooling a turbine blade/vane, i.e. blade and/or vane, is to have one or more internal passages, referred to as cooling channels or cooling passages, within the blade/vane and to flow a cooling fluid, such as cooling air through the cooling channel. Surfaces of such cooling channel or channels are often lined with turbulators to enhance the heat transfer into the cooling air from the blade/vane internal surfaces forming surfaces of the cooling channel. Often a series of rib turbulators or pin-fin turbulators are arranged along the flow path of the cooling fluid within the cooling channel. The turbulators induce turbulence in the cooling fluid and thereby increase the efficiency of the heat transfer.
The flowing cooling fluid passes over, about and/or around the sequentially arranged rows or members of the turbulators and a rate of heat acceptance by the cooling fluid is increased as the cooling fluid passes over from a turbulator positioned first to downstream turbulators i.e. a turbulator positioned third or fourth in the series, and from there onwards the rate of heat acceptance by the cooling fluid as it passes on to further downstream turbulators, for example a turbulator positioned seventh or eighth in the series, remains substantially constant. Thus the rate of heat acceptance, and thereby an effectiveness of cooling by the cooling fluid passing over a series of sequentially arranged turbulators, increases from the very first turbulator towards the third or fourth turbulator and there it reaches a peak value.
Furthermore, the first turbulator itself is positioned downstream of an inlet of the cooling channel by a distance, primarily due to manufacturing and design constraints. Thus there is a delay in reaching peak value of cooling efficiency between an instance when the cooling fluid enters the inlet of the cooling channel till an instance when the cooling fluid crosses the position of the third or the fourth turbulator of the series of turbulators positioned inside the cooling channel. Thus there is a scope of improvement in a turbomachine component with cooling channels and turbulators by shortening the delay between the instance when the cooling fluid enters the inlet of the cooling passage and when the cooling efficiency reaches the peak value.
Thus an object of the present disclosure is to provide a technique by which in a turbomachine component having an aerofoil and a cooling channel, a cooling effect of a cooling fluid passing through the cooling channel reaches a peak faster and thus enhancing an efficiency of cooling in the turbomachine component.
The above objects are achieved by a turbomachine component according to the present technique. Advantageous embodiments of the present technique are provided in dependent claims. Features of the independent claim can be combined with features of dependent claims, and features of dependent claims can be combined together.
The present technique presents a turbomachine component which has an aerofoil. An example of such turbomachine component is a blade or a vane for a turbomachine or a gas turbine engine. The aerofoil of the turbomachine component includes a suction side wall and a pressure side wall. The side walls, namely the suction side wall and the pressure side wall bordering an aerofoil cavity. The turbomachine component also has at least one cooling channel that extends inside at least a part of the aerofoil cavity. The cooling channel is for flow of a cooling fluid, when present. The cooling channel has an inlet that receives the cooling fluid which then flows through the cooling channel. The cooling channel also has a series of turbulators positioned inside the cooling channel. The cooling fluid flows over and about the turbulators. The turbomachine component includes at least one vortex generating element, hereinafter also referred to as the element. The element is positioned at the inlet of the cooling channel upstream of the turbulators or is positioned adjacent to and upstream of the inlet of the cooling channel. The cooling fluid flows about and contiguous with the element before the cooling fluid reaches the turbulators. The element generates a swirl in the cooling fluid before the cooling fluid reaches the turbulators.
The swirl includes generating a vortex and/or disturbing a stream lined flow and/or generating turbulence. As a result of the swirl generation the cooling effect of the cooling fluid reaches a peak faster than a scenario where the element is not present. In other words, if the cooling fluid were to reach its peak cooling effect after crossing three sequentially arranged turbulators inside the cooling channel, then due to the action of the element, when present according to the present technique, the cooling fluid reaches its peak cooling effect before crossing three sequentially arranged turbulators inside the cooling channel and thus faster. This results in increased efficiency of cooling by the cooling fluid flow.
In an embodiment of the turbomachine, the inlet of the cooling channel is present in the aerofoil cavity, and thus the element is positioned in the aerofoil cavity.
In an embodiment of the turbomachine component, the turbomachine component includes a base wherefrom the aerofoil extends radially. The base is used to fix the turbomachine component to a desired position in the turbomachine. For example the base may in a blade or a vane and may aid in fixing the aerofoil of the blade or the vane onto their respective disks. The base, especially in a blade of a turbomachine, has a circumferentially extending platform and a root section radiating out of the platform in a direction opposite to the aerofoil. The turbomachine component, such as the blade or the vane, may also have a tip, for example a shroud, which is present at an end of the aerofoil that is opposite to the base. The base may have a base cavity, for example a root cavity and/or a platform cavity, and the cooling channel may be supplied with the cooling air through the base cavity and thus the inlet of the cooling channel may be present either in the base, fluidly connected with the base cavity, or may be in the aerofoil but in close proximity of the base. In such a turbomachine component, the element may be positioned within the base, for example in the root cavity or in the platform cavity.
In the turbomachine component, the tip may have a shroud cavity and the cooling channel may be supplied with the cooling air through the shroud cavity and thus the inlet of the cooling channel may be present either in the shroud, connected with the shroud cavity, or may be in the aerofoil but in close proximity of the shroud. In such a turbomachine component, the element may be positioned within the shroud, i.e. the tip of the turbomachine component.
The element may be formed as a protrusion emerging out from a surface at which the vortex generating element is positioned i.e. the protrusion is formed projecting out of the surface but as a part of the surface and not as a separate entity from the surface. The protrusion emerging out from the surface may be formed for example by casting the surface and the protrusion together. Alternatively, the element may be a fixture attached to the surface at which the element is positioned, for example if the element is separately formed and then glued onto the surface.
The element may have a shape selected from a rib shape, split-rib shape, wedge shape, split-wedge shape, pin fin shape, conical shape with straight side, conical frustum shape with straight side, conical shape with curved side, conical frustum shape with curved side, spherical dome shape, tetrahedron shape, tetrahedral frustum shape, pyramidal shape, and pyramidal frustum shape. It may be noted that the element is meant to be representative of one or more individual members. An example of the element having one member is where the element is a rib extending on the parallel to a surface where the turbulators of the cooling channel are present and normal to direction of flow of cooling air when entering the inlet. An example of the element having more than one member is if the element is a split rib which may be visualized as smaller ribs formed by dividing a larger rib perpendicularly to a longitudinal axis of the larger rib and then spacing apart the smaller ribs along the longitudinal axis.
The element, or each member of the element in case where the element has more than one member, may have dimensions relative to the turbulators, for example a height of the element is between 50 percent and 150 percent of a height of the turbulators, especially when the turbulators are rib shaped. In one embodiment a height of the element is greater than the height of the cooling channel, thus physically differentiating the turbulators positioned inside the cooling channel from the element positioned outside the cooling channel. Similarly, the element, or each member of the element in case where the element has more than one member, may have dimensions relative to the cooling channel, for example a height of the element is between 10 percent and 40 percent of a height of the cooling channel at the inlet, especially when the turbulators inside the cooling channel are pin fin shaped. If the cooling channel is of a rectangular or triangular cross section the height at the inlet is the largest height of the cooling channel in the same direction as that measuring the height of the element.
The above mentioned attributes and other features and advantages of the present technique and the manner of attaining them will become more apparent and the present technique itself will be better understood by reference to the following description of embodiments of the present technique taken in conjunction with the accompanying drawings, wherein:
Hereinafter, above-mentioned and other features of the present technique are described in details. Various embodiments are described with reference to the drawing, wherein like reference numerals are used to refer to like elements throughout. In the following description, for the purpose of explanation, numerous specific details are set forth in order to provide a thorough understanding of one or more embodiments. It may be noted that the illustrated embodiments are intended to explain, and not to limit the invention. It may be evident that such embodiments may be practiced without these specific details.
As mentioned hereinabove, in a turbomachine component cooled by the flow of a cooling fluid though a cooling channel having sequentially arranged turbulators, a peak of cooling is reached after the cooling fluid crosses the third or the fourth turbulator in the series, thus the delay. The reason the cooling fluid reaches a peak value of heat acceptance after this delay is that turbulence starts setting into the flow of the cooling fluid only after the cooling fluid encounters the very first turbulator within the cooling passage or channel and an optimum turbulence in the flow of the cooling fluid is reached by further two or three encounters between the flowing cooling fluid and turbulators immediately downstream of the first turbulator.
The basic idea of the present disclosure to shorten or obviate this delay is by introducing a turbulence or swirl in the cooling fluid immediately before the cooling fluid enters the cooling channel through an inlet of the cooling channel or by introducing a turbulence or swirl in the cooling fluid as the cooling fluid enters the cooling channel through the inlet of the cooling channel, i.e. by introducing the swirl in the cooling fluid before the cooling fluid reaches the first turbulator. The introduction of the swirl in the cooling fluid is achieved, according to the present technique, by positioning a vortex generating element either at the inlet of the cooling channel upstream of the turbulators or by positioning the vortex generating element adjacent to and upstream of the inlet of the cooling channel. The positioning of the vortex generating element is such that the cooling fluid has to flow about the vortex generating element before the cooling fluid enters the cooling passage and/or reaches the turbulators. The vortex generating element generates a swirl in the cooling fluid by having a shape that induces swirling of the fluid or generation of vortices in the cooling fluid flow before the cooling fluid reaches the turbulators. The turbulence or swirl is initiated by highly unsteady flow features that are created by the vortex generating device in the cooling air. These features include vortex shedding, fluid shear layers within the passage containing flow recirculations and flow separations and unstable shear layers that do not have a stable location.
In operation of the gas turbine engine 10, air 24, which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16. The burner section 16 comprises a longitudinal axis 35 of the burner, a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28. The combustion chambers 28 and the burners 30 are located inside the burner plenum 26. The compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the burner 30 and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.
This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment. An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine 18.
The turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22. In the present example, two discs 36 each carry an annular array of turbine blades 38. However, the number of blade carrying discs could be different, i.e. only one disc or more than two discs. In addition, guiding vanes 40, which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and turn the flow of working gas onto the turbine blades 38.
The combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotates the shaft 22. The guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.
The turbine section 18 drives the compressor section 14. The compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48. The rotor blade stages 48 comprise a rotor disc supporting an annular array of blades. The compressor section 14 also comprises a casing 50 that surrounds the rotor stages and supports the vane stages 48. The guide vane stages include an annular array of radially extending vanes that are mounted to the casing 50. The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point. Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operational conditions.
The casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14. A radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor which is partly defined by the annular array of blades 48.
The present technique is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine. However, it should be appreciated that the present technique is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications.
The terms upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine unless otherwise stated. The terms forward and rearward refer to the general flow of gas through the engine. The terms axial, radial and circumferential are made with reference to the rotational axis 20 of the engine.
It may be noted that the present technique has been explained in details with respect to an embodiment of a turbine blade, however, it must be appreciated that the present technique is equally applicable and implemented similarly with respect to a turbine vane or any other turbomachine component having an aerofoil and being cooled by a cooling channel with turbulators arranged inside the cooling channel.
The aerofoil 5 includes a suction side wall 2, also called suction side 2, and a pressure side wall 3, also called pressure side 3. The side walls 2 and 3 meet at a trailing edge 92 on one end and a leading edge 91 on another end. The aerofoil 5 has a tip end 93. The aerofoil 5 may be connected to a shroud (not shown) at the tip end 93 of the aerofoil 5. In some other embodiments the aerofoil 5 may be connected to a tip platform (not shown) instead of the shroud. The shroud and the tip platform are commonly referred to a tip (not shown) of the turbomachine component 1. The aerofoil 5 may also include a shroud (not shown) at the tip end 93 of the aerofoil 5. The side walls 2 and 3 of the aerofoil 5 act as boundary for an aerofoil cavity 4.
Referring to
The blade 1 includes at least one vortex generating element 8, hereinafter also referred to as the element 8. The element 8 is positioned at the inlet 66 of the cooling channel 6 upstream of the turbulators 62. The element 8 may also be positioned adjacent to and upstream of the inlet 66 of the cooling channel 6 as depicted in
The element in
Referring to
Alternatively, the cooling channels 6 may be arranged such that one of the cooling channels 6 opens in the aerofoil cavity 4 and the other cooling channel 6 has its inlet 66 at the aerofoil cavity 4, hereinafter also referred to the cavity 4. In such an embodiment, the element 8 for the other cooling channel 6 is positioned inside the cavity 4.
As depicted in
It may be noted that both for
Furthermore, in an exemplary embodiment of the turbomachine component 1, the tip, i.e. for example a shroud (not shown) may have a shroud cavity (not shown) and the cooling channel 6 may be supplied with the cooling air 7 through the shroud cavity and thus the inlet 66 of the cooling channel 6 may be present either in the shroud, connected fluidly with the shroud cavity, or may be in the aerofoil 5 but in close proximity of the shroud. In such embodiments of the turbomachine component 1, the element 8 may be positioned within the shroud, i.e. the tip of the turbomachine component 1.
Referring now to
As shown in
Besides the embodiments for the element 8 depicted in
It may also be noted that the element 8 is meant to be representative of one or more individual members, for example as shown in
The element 8, or each member of the element 8 in case where the element 8 has more than one member, may have dimensions relative to the turbulators 62 that are present within the cooling channel 6, for example a height of the element 8 is between 50 percent and 150 percent of a height of the turbulators 62, especially when the turbulators 62 are rib shaped turbulators 63. Similarly, the element 8, or each member of the element 8 in case where the element has more than one member, may have dimensions relative to the cooling channel 6, for example a height of the element 8 is between 10 percent and 40 percent of a height of the cooling channel 8 at the corresponding inlet 66, especially when the turbulators 62 inside the cooling channel are pin fin shaped 64.
It may be noted that in some cases the same cooling channel 6 may have more than one type of turbulators 62 for example the rib shaped turbulators 63 and the pin fin shaped turbulators 64, as depicted in the exemplary embodiment of
In an embodiment of the element 8, and in reference to
Furthermore, the height H of the element 8 can be up to ⅓ the cooling channel 6 width i.e. the dimension of the cooling channel 6 in the direction of the height of the element 8.
It is necessary for the element 8 to be located upstream of the cooling channel turbulators 63 and effectively in a clear line-of sight of the turbulators 63 such that the vortices or swirl in the cooling fluid 7 generated by the element 8 are effective and impinge on the turbulators 63. No other features such as obstructions or bends in the cooling channel 6 should be between the element 8 and the turbulators 63. The element 8 is located on the same part of the surface of the cooling channel 6 as the turbulators 63 to have the best effect. This is to ensure that the turbulence or vortices from the element 8 interact with the turbulators 63 enhancing their effectiveness.
The element 8 is located in the cooling channel 6 symmetrical about a plane through a central line of the cooling channel 6. The central line of the cooling channel 6 may be curved bearing in mind the general three-dimensional shape of the component and the cooling channel 6.
It will be appreciated by the skilled person that the turbulators 63 have a much greater aspect ratio W to L because they substantially span across the cooling channel and are designed to cause turbulence to interact with the surface of the cooling channel. This is in contrast to the vortex generator elements 8 which are designed, by virtue of their aspect ratios to create turbulence within the stream of cooling fluid before the cooling fluid interacts with the turbulators 63. Furthermore, the pedestals 64 span the entire ‘height’ of the cooling channel because they are design to enhance the surface area for heat transfer purposes.
While the present technique has been described in detail with reference to certain embodiments, it should be appreciated that the present technique is not limited to those precise embodiments. Rather, in view of the present disclosure which describes exemplary modes for practicing the invention, many modifications and variations would present themselves, to those skilled in the art without departing from the scope and spirit of this invention. The scope of the invention is, therefore, indicated by the following claims rather than by the foregoing description. All changes, modifications, and variations coming within the meaning and range of equivalency of the claims are to be considered within their scope.
| Number | Date | Country | Kind |
|---|---|---|---|
| 16168424.6 | May 2016 | EP | regional |
This application is the US National Stage of International Application No. PCT/EP2017/060296 filed Apr. 28, 2017, and claims the benefit thereof. The International Application claims the benefit of European Application No. EP16168424 filed May 4, 2016. All of the applications are incorporated by reference herein in their entirety.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/EP2017/060296 | 4/28/2017 | WO | 00 |
