Patent Grant
7467922
The invention relates to a turbine blade or vane for a gas turbine, having a blade or vane root, which is successively adjoined by a platform region with a transversely running platform and then a blade or vane profile which is curved in the longitudinal direction, having at least one cavity which is open on the root side, through which a coolant can flow and which extends through the blade or vane root and the platform region into the blade or vane profile. The invention also relates to the use of a turbine blade or vane of this type.
EP 1 355 041 A2 has disclosed a turbine blade or vane of this type. The cast turbine blade has a cavity which extends from the blade root through the platform into the blade profile. The cross section of the cavity is substantially constant along its extent. The cavity is surrounded by an inner wall and has a cross section which is enlarged only in the region of the platform, by virtue of the inner wall being set back in the region of the platform. The material thickness in the transition region between blade profile and platform projecting transversely to it consequently remains constant, so that the transition between them can be cooled more successfully.
The invention presented is directed toward a turbine blade or vane for a gas turbine, comprising a blade or vane root that is successively adjoined by a platform region with a transversely extending platform and then a blade profile that is curved in the longitudinal direction a platform surface that is provided at the platform and exposed to hot gas; and at least one cavity that is open on the root side, through which a coolant can flow and which extends through the blade or vane root and at least into the platform region and is surrounded by an inner wall, the contour of which, extending in the platform region, is set back with respect to the contour running in the blade or vane root so as to form a recess which widens the cavity, wherein the recess that widens the cavity extends into the region below the platform surface so as to form an at least partially hollow platform and in that there is at least one means for diverting the coolant into the partial cavity.
Moreover,
While the gas turbine is operating, mechanical centrifugal force loads and thermal stresses occur at the turbine blade between the relatively cold, thin-walled blade profile and the often hotter platform. The high stresses in the platform and in the transition region limit the fatigue service life of the turbine blade as a whole. Moreover, particularly in the case of turbine blades with a high diverting action and accordingly a strong curvature the fatigue service life is further reduced by the platforms which in sections overhang on one side. The wide platform overhangs are difficult to cool, and high thermal stresses, which also restrict the fatigue service life, may be formed there in particular.
Moreover, the difficulty in cooling the platform is on the one hand that of guiding the cooling air into the platform and on the other hand that of establishing as uniform as possible a dissipation of heat in order to lengthen the fatigue service life, while at the same time taking account of the need to make economical use of cooling air.
Therefore, it is an object of the invention to provide a turbine blade or vane for a gas turbine in which the fatigue service life is lengthened while at the same time cooling air is saved. A further object of the invention is to provide the use of a turbine blade or vane of this type.
The object relating to the turbine blade or vane is achieved by a turbine blade or vane of the generic type which is designed with the features of the claims.
The invention is based on the discovery that the platform can be cooled in a particularly simple way if the recess which widens the cavity projects into the region below the platform surface, so as to form an at least partially hollow platform, and at least one means for diverting the coolant into the partial cavity is provided.
The platforms which are of hollow design can be produced by the use of suitable cores when casting the turbine blade or vane. On account of the recess projecting into the platform, therefore, transitions between blade profile and platform which, as seen in cross section, have a constant material thickness, are possible. In particular as a result of this, it is possible to reduce the thermal stresses in the transition region and in the platforms, which has a beneficial effect on the service life of the turbine blade or vane. The invention therefore institutes a step which is a significant advance on the quoted prior art.
To enable coolant to flow into the recess, there is at least one means for diverting the coolant into the partial cavity. Without a means of this type, cooling air which flows in on the root side would simply flow through the turbine blade or vane in the radial direction. Only standing swirls or what are known as dead water regions, in which a small proportion of the cooling air would be recirculated, would be formed in the recesses running transversely with respect to the radial direction. The use of these means forces the coolant which flows in at the root side to be diverted in the direction of the recess, so that as a result coolant flows around the rear side of the platform surface. This leads to extremely effective convective cooling of the transition and of the platform.
Advantageous configurations are given in the subclaims.
Open platform cooling can be achieved if at least one outlet opening, through which the coolant can flow out of the partial cavity, is provided in the partial cavity as means for guiding the coolant. The outlet opening is provided in the vicinity of the platform edge, so that coolant can flow into the recess and can flow out again on the opposite side. It is advantageous for the outlet opening to open out into the platform surface. This allows film cooling of the platform as well as convective cooling, in order to effectively protect particularly hot regions of the platform from hot gas.
If, on the other hand, the outlet opening opens out into an end side of the platform, it is advantageously possible to block a gap which is formed by the end-side longitudinal edges of platforms of adjacent gas turbine blades or vanes from the penetration of hot gas.
In a further advantageous configuration of the invention, a pin which is located in the cavity and extends from the blade or vane root into the platform region is provided as means for guiding the coolant. This pin divides the cavity into two supply passages which run close to the surface. Accordingly, coolant which flows therein is guided relatively close to the inner wall of the passage for the purpose of cooling the turbine blade or vane.
The configuration in which the pin, in the platform region, has a widening, which diverts the coolant, which can flow along the pin, in the direction of the partial cavity, is particularly effective. The widening which extends in the transverse direction causes the coolant which flows in radially through the supply passages to be diverted in the transverse direction into the hollow platform.
In a further advantageous configuration of the invention, at least one guiding element, which is L-shaped in cross section, extends from the blade or vane root toward the platform region as means for guiding the coolant, so as to form supply passages, the limbs of which guiding element, at the end located in the platform region, at least partially project into the hollow partial cavity. This allows the coolant which flows into the supply passages to be diverted particularly effectively into the partial cavity, since the L-shaped guiding element runs parallel to the inner wall which delimits the cavity and the partial cavity. On account of the L-shaped guiding element, the coolant which is diverted into the partial cavity is guided to the platform edge, where it can then flow radially outward and then back inward around the free end of the limb of the L-shaped guiding element. On account of the flow conditions which are present in the turbine blade or vane, the coolant then flows onward in the direction of the blade or vane profile and during this period cools the transition region between blade profile and platform extremely effectively.
On account of the uniform platform cooling and the uniform cooling of the transition, the fatigue service life of the turbine blade or vane can be effectively lengthened in this configuration.
In a variant of the invention, at least one guiding element extends from the blade or vane root toward the platform region as means for guiding the coolant, until it merges into an inner wall, delimiting the cavity, of the blade or vane profile.
The abovementioned cooling concepts can be used particularly effectively in a turbine blade or vane in which the blade or vane root runs in the longitudinal direction of the blade profile, and the platform has two platform longitudinal edges bent parallel and running in the longitudinal direction, and in which the respective blade or vane root surface facing the suction-side and pressure-side profile walls is convexly and concavely curved in a corresponding way to the associated platform longitudinal edge. In a turbine blade or vane of this type with a curved blade or vane root and a curved platform, a pressure-side platform and a suction-side platform, each having an approximately constant platform width along the main blade or vane part, automatically result along the longitudinal direction. Constant platform widths of this type are heated more uniformly and accordingly can be combined particularly successfully with the cooling concepts according to the invention.
Cooling concepts of this nature can be used to advantageous effect even if the suction-side and/or pressure-side platform overhang are designed as platform stubs with a relatively short platform width.
It is preferable for the turbine blade or vane to be cast and to have a blade or vane root which, when seen in cross section, is in dovetail, hammer or fir tree shape.
The object relating to a use of the turbine blade or vane is achieved by the features of claim 12. It is proposed that the turbine blade or vane as claimed in one of claims 1-11 be used in a preferably stationary gas turbine.
The invention is explained with reference to figures, in which:
The blade profile 56 has a pressure-side, concavely curved profile wall 62 and a suction-side, convexly curved profile wall 64, which extend from a leading edge 66 of the blade profile 56 to a trailing edge 68. When the gas turbine 1 is operating, the hot gas 11 flows around the turbine blade 50, along the profile walls 62, 64, from the leading edge 66 toward the trailing edge 68.
In a corresponding way to the curvature of the blade profile 56, the platform 54 is curved along the axial direction A, the longitudinal edges 55 of the platform 54 do not run in a straight line, but rather on an arc. Accordingly, the platform longitudinal edge 54 arranged at the pressure-side profile wall 62 is curved concavely and the platform longitudinal edge arranged at the suction-side profile wall 64 is curved convexly. The platform 54 has a platform transverse edge 53, which runs transversely at the end side, in the region of the leading edge 66 and in the region of the trailing edge 68.
As can be seen from the perspective illustration presented in
The blade root surface 72 is to be understood as meaning that surface of the blade root 52 which runs in the axial direction A. The end-side blade root surfaces are excluded from this term.
The platform 54 has a platform overhang 75 projecting transversely with respect to the radial direction, i.e. in the transverse direction. The width of the platform overhang 75 is determined by the distance from suction-side profile wall 64 or pressure-side profile wall 62 to the respectively immediately adjacent platform longitudinal edge 55.
On account of the curved shape of the blade root 52, it is possible to realize platform overhangs 75 which, along the axial direction A, have an approximately constant platform width B on the suction side and on the pressure side, in a particularly successful way. On account of the constant platform width B, the platform can be cooled particularly uniformly, as described below.
In accordance with the cross-sectional illustrations presented in
When the gas turbine 1 is operating, the cavity 58 has a coolant 60, preferably cooling air, flowing through it. For the coolant 60 to be supplied, the cavity 58 in the blade root 52 is open on the root side. Based on the installation position in the gas turbine 1, the turbine blade 50, in the region of the platform 54, has a recess 63 which runs transversely with respect to the radial direction R and extends sufficiently deep into the platform 54 for it to lie opposite the surface 61 of the platform 54 as a partial cavity 51 therein.
The recess 63 extends over at least 30% of the width B of the platform overhang 75. On account of the pocket-shaped recess 63 extending relatively deep into the platform 54 compared to the prior art, it is possible not only to realize extremely efficient cooling of the transition region 48 of blade profile 36 and platform 54 running transversely to it, but also to realize efficient internal, convective cooling of the platform 54 and/or of the platform overhang 75.
To divert the coolant 60, which flows in on the root side, in the direction of the recesses 63 and into the hollow platform 54, there is, as shown in
The configuration of the outlet openings 73 shown in
In the configuration shown in
In a further variant of the invention, as shown in
In another configuration according to the invention,
After two coolant streams 60a, 60c which flow into the supply passages 96a, 96c on the root side have been passed into the recesses 63 to cool the platform 54, these coolant streams are combined in the blade profile 56, where the coolant 60 can be used to cool the blade profile 56 using a conventional cooling method, such as for example impingement cooling, convective cooling, film cooling or effusion cooling.
The two guiding elements 92 divide the cavity 58 into three supply passages 96a, 96b and 96c on the blade root side. The coolant 60 which flows in via the supply passages 96a, 96c convectively cools the platforms 54 of the turbine blade 50 according to the invention, since the guiding elements 92 force the coolant 60 to be diverted into the recesses 63. By contrast, the coolant 60 which flows in via the supply passage 96b can flow into the blade profile 56 without being used by the blade root 52 and the platform region, and can be used in the blade profile 56 to cool for the first time the latter.
Consequently, these solutions allow coolant 60 to be passed in targeted fashion into the recesses 63 and/or the partial cavity 51, so as to form closed platform cooling, which leads to particularly efficient cooling of the platform 54 and of the transition region 48 or the transition radius. Moreover, on account of the approximately constant platform width B along the axial direction A, particularly uniform cooling of the transition is possible.
The turbine blades 50 proposed in
A final variant of a turbine blade 50 according to the invention is shown in cross section in
To illustrate the geometry shown,
The supply passage 96b is arranged centrally on the leading side and passes coolant 60 into the hollow blade profile 56. The supply passages 96a and 96c are provided adjacent to it on the pressure side and the suction side. In the blade root 52, the supply passages 96a, 96c initially run substantially in the radial direction, and in the region of the platform 54 they bend in the transverse direction and then in the axial direction A, so that they form the hollow platforms 54. Consequently, the coolant 60 is supplied in the root-side end of the turbine blade 50.
The supply passages 96a, 96c merge into cooling passages 57a, 57c which run in the axial direction A along and approximately parallel to the curved platform longitudinal edges 55 by virtue of guiding elements 92, starting from the blade root 52, extending in the direction of the platform region and merging into the inner wall 59, delimiting the cavity 58, of the blade profile 56.
The turbine blades 50 shown preferably have the blade root 52 and platform 54 designed with a curvature in the axial direction of the gas turbine, so that there are no asymmetric overhangs of platforms 54 formed. On account of the associated more uniform platform width (platform overhang along the axial direction), all the novel cooling concepts are particularly simple and particularly efficient in use.
Overall, the invention provides novel cooling concepts for gas turbine blades as running blades and vanes as guiding blades which have platforms which can be cooled particularly efficiently and uniformly. On account of the more uniform cooling, the fatigue service life of the turbine blade is lengthened. The platforms which are of hollow design can be internally cooled convectively either by means of suitable pins or guiding elements and/or by the provision of bores for producing a discharge of cooling air. The excellent coolability of the platforms also allows particularly efficient use of TBC coatings (thermal barrier coating). Moreover, it is possible to save cooling air compared to the platform cooling concepts which have been known hitherto and this cooling air can then be burnt in the gas turbine, increasing the efficiency of the latter.
This application claims the benefit of U.S. Provisional Patent Application Ser. No. 60/702,313, filed Jul. 25, 2005.
| Number | Name | Date | Kind |
|---|---|---|---|
| 5820343 | Kraft et al. | Oct 1998 | A |
| 6402471 | Demers et al. | Jun 2002 | B1 |
| Number | Date | Country |
|---|---|---|
| 1 355 041 | Oct 2003 | EP |
| Number | Date | Country | |
|---|---|---|---|
| 20070020100 A1 | Jan 2007 | US |
| Number | Date | Country | |
|---|---|---|---|
| 60702313 | Jul 2005 | US |
