TURBINE ROTOR BLADE OF A GAS TURBINE AND METHOD FOR COOLING A BLADE TIP OF A TURBINE ROTOR BLADE OF A GAS TURBINE

Abstract

The present invention relates to a turbine rotor blade of a gas turbine with a blade tip, on which means for duct-type guidance of cooling air extending from a front suction-side area of the blade tip to a rear area of the blade tip are provided, and to a method for cooling a blade tip of a turbine rotor blade of a gas turbine, where air from a hot gas flow is guided from a front suction-side area of a blade tip to a rear area of the blade tip through a duct-type guidance.

Description

This application claims priority to German Patent Application DE102013224998.5 filed Dec. 5. 2013, the entirety of which is incorporated by reference herein.

This invention relates to a turbine rotor blade of a gas turbine with a blade tip. Furthermore, this invention relates to a method for cooling such a blade tip of a turbine rotor blade.

It is known from the state of the art that a leakage mass flow caused by the pressure difference from a blade pressure side to a blade suction side arises at a radial gap between a turbine rotor and a casing. Attempts are therefore being made to design the blade tip of the turbine rotor such that the leakage mass flow is reduced. Another objective is to reduce the negative effect of the blade tip leakage vortex caused by the leakage mass flow on the turbine aerodynamics.

To improve the flow over the blade tips of the turbine rotor, circumferential sealing edges (squealers) are used. Designs are also known, where overhangs at the blade tip (winglets) are provided. The circumferential sealing edges can contribute to an improvement in the aerodynamics. The overhangs on the suction side and/or on the pressure side can reduce the leakage mass flow and also improve the aerodynamics around the blade tip.

It is furthermore necessary in turbine rotor blades, particularly under operating conditions with high thermal loads due to the hot gas flow, to achieve high durability and a long service life for the blade tips. To that end, it is known to cool the turbine rotor blades internally by convection using compressor air and to discharge this air for the purpose of film cooling on the blade outer surfaces or blade tips. Besides the blade leading edge, the thermally most highly loaded area of a blade tip is often the rear pressure-side area. It is known to provide on the blade tip pressure-side holes or recesses through which blade-internal film cooling air is passed out near the blade tip. In addition, cooling air can be passed out through openings on the blade tip (dust holes), in order to prevent dirt accumulations in the internal cooling air passages and to achieve additional cooling.

The measures known from the state of the art entail a number of disadvantages:

When film cooling air is used, it is necessary to consume a large quantity of “expensive” film cooling air in order to protect the blade tip from the high temperatures of the hot gas. The film cooling air must be taken from the compressor of the gas turbine, thereby reducing the efficiency of the thermodynamic cyclic process inside the gas-turbine engine.

Convective internal cooling of the blade tip by internal cooling air passages is relatively ineffective in the area of the blade tip. This is partly due to the fact that only insufficiently high internal heat transfer coefficients can be generated directly underneath the blade tip. In addition, conventional casting techniques, using which the turbine rotor blades are manufactured, require relatively high minimum wall thicknesses, so that the temperature in the outer wall area is relatively close to the hot gas temperature.

Cooling air that must be forcibly discharged at the blade tip in order to prevent dirt accumulations in the internal blade ducts contributes only to a limited extent to the cooling of the blade tip, since it cannot reach the thermally most highly loaded points of the blade tip.

The object underlying the present invention is to provide a method for cooling a blade tip of a turbine rotor blade of a gas turbine as well as a turbine rotor blade suitable for carrying out the method, which, while being simply designed guarantees effective cooling.

It is a particular object of the present invention to provide solution to the above problematics by a combination of the features of the independent Claims. Further advantageous embodiments of the invention become apparent from the sub-claims.

In accordance with the invention, it is thus provided in respect of the turbine rotor blade that means for duct-type guidance of cooling air extend from its front suction-side area to a rear area. The invention thus provides for the thermally highly loaded rear area of the turbine rotor blade tip to be supplied at least partly with passive air, i.e. with relatively cold air from the hot gas flow. It is self-evident that the hot gas temperature must be below the maximum temperature sustainable by the blade material. There is an area with relatively low hot gas temperature on the front suction-side blade tip. In particular in the case of additional discharge of cooling air at the turbine casing upstream of the rotor blades, the hot gas temperature in this area can be considerably below the highest permissible metal temperature.

In accordance with the invention, the relatively “cold” hot gas flow is thus used for cooling the blade tip by being routed to the thermally most highly loaded pressure-side rear area of the blade tip. This is achieved by the means provided in accordance with the invention for duct-type guidance of this cooling air. In a particularly favourable embodiment of the invention, it is provided that a cover forming a flow duct is attached at the blade tip. This creates a duct or a cavity on the blade tip, through which the cooling air can be passed. This duct or cavity, forming the means in accordance with the invention for duct-type guidance of cooling air, preferably has an inlet opening on the suction-side blade leading edge. It is however also possible to introduce cooling air from the blade interior. In the area of the blade trailing edge too, an opening is provided through which the cooling air flows out. The pressure difference over the blade row ensures here a flow of relatively cold hot gas through the duct or cavity.

The means provided in accordance with the invention can extend over the entire length or only over part of the length of the blade tip.

The cooling air can exit the duct or cavity in accordance with the invention at the blade tip, depending on the required counter-pressure level, in the area of the pressure-side or the suction-side rear blade tip. The discharge at the pressure side has the further advantage, besides lower aerodynamic losses, that the cooling air discharged there is in turn sucked into the blade tip gap, so that the external blade tip surface too is supplied with relatively cold air.

In a preferred embodiment of the invention, a protective cover as mentioned is fitted onto a circumferential sealing edge of the blade tip (winglet, squealer) and fastened there. The result is an effective, flat and contourless blade tip geometry with a duct or cavity underneath. In an even more advantageous embodiment of the invention, it is provided that the cover is suitably shaped or contoured such that the blade tip can retain the contour of the circumferential sealing edge.

It is particularly advantageous when the wall thicknesses of the cover are designed relatively thin, so that its maximum metal temperature can be kept as low as possible.

The protective cover and/or the cavity on the blade tip can be but do/does not necessarily have to be, designed up to the blade leading edge. Since the pressure level in the front blade tip area is relatively constant and only drops steeply towards the blade trailing edge, it can be advantageous to design the protective cover and hence the duct or cavity only starting from a middle position of the blade tip. With this embodiment, the dust holes can then be near the cover and hence close to the inflow area of the duct or cavity, in order to promote the aspiration of the cold dust hole air into the duct or cavity. It is thus possible to use the dust hole air, otherwise not readily usable for cooling, to cool the pressure-side blade tip close to the trailing edge.

In a further embodiment of the invention, it is possible to design the protective cover at the blade leading edge closed (without opening). Hence only one opening of the cavity or of the duct is provided close to the blade trailing edge. With this embodiment of the invention, the cavity or duct can be flooded completely with cold blade-internal air. This embodiment is suitable in particular for very high hot gas temperatures, when blade-external air is no longer usable for cooling the blade tip.

For guiding the flow inside the blade tip cavity or in the flow duct provided at the blade tip, additional webs or supports can be used. By means of these webs, the air can be guided to the thermally most highly loaded areas. The webs or supports are furthermore used as fastening surfaces for the protective cover and thereby contribute to the mechanical stability of the blade tip. Furthermore, the webs or supports can increase the internal heat transfer in the cavity or duct, so that the cooling effect can be further improved.

The following advantages result in accordance with the invention, as already partially explained above:

With the invention, the quantity of film cooling air required for cooling the rotor blade tip can be considerably reduced.

Due to the more effective cooling of the blade tip and the reduction of the blade tip temperatures that this entails, the wear on the blade tip can be reduced and hence the service life of the turbine blade extended.

Due to the reduced wear on the blade tip during operation, the decrease in turbine efficiency over the period of operation can be reduced.

In accordance with the invention the operating costs of the gas turbine are reduced.

The present invention is described in the following in light of the accompanying drawing showing exemplary embodiments. In the drawing,

FIG. 1 shows a schematic representation of a gas-turbine engine in accordance with the present invention,

FIG. 2 shows a simplified sectional view of a blade tip designed in accordance with the present invention,

FIG. 3 shows a view, by analogy with FIG. 2, of a further exemplary embodiment of the present invention,

FIGS. 4 shows a further exemplary embodiment, by analogy with FIGS. 2 and 3 without lateral blade tip overhang, and

FIGS. 5 to 12 show simplified perspective representations of exemplary embodiments in accordance with the present invention.

The gas-turbine engine 10 in accordance with FIG. 1 is a generally represented example of a turbomachine where the invention can be used. The engine 10 is of conventional design and includes in the flow direction, one behind the other, an air inlet 11, a fan 12 rotating inside a casing, an intermediate-pressure compressor 13, a high-pressure compressor 14, a combustion chamber 15, a high-pressure turbine 16, an intermediate-pressure turbine 17 and a low-pressure turbine 18 as well as an exhaust nozzle 19, all of which being arranged about a central engine axis 1.

The intermediate-pressure compressor 13 and the high-pressure compressor 14 each include several stages, of which each has an arrangement extending in the circumferential direction of fixed and stationary guide vanes 20, generally referred to as stator vanes and projecting radially inwards from the engine casing 21 in an annular flow duct through the compressors 13, 14. The compressors furthermore have an arrangement of compressor rotor blades 22 which project radially outwards from a rotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine 16 or the intermediate-pressure turbine 17, respectively.

The turbine sections 16, 17, 18 have similar stages, including an arrangement of fixed stator vanes 23 projecting radially inwards from the casing 21 into the annular flow duct through the turbines 16, 17, 18, and a subsequent arrangement of turbine rotor blades 24 projecting outwards from a rotatable hub 27. The compressor drum or compressor disk 26 and the blades 22 arranged thereon, as well as the turbine rotor hub 27 and the turbine rotor blades 24 arranged thereon rotate about the engine axis 1 during operation.

FIG. 2 shows a simplified sectional view of a blade tip 29 of a turbine rotor blade 24. The reference numeral 35 shows the suction side, while the reference numeral 36 indicates the pressure side. Front-side sealing edges 33 are provided on the blade tip 29. Furthermore, the blade tip 29 can have lateral overhangs (winglets) 34. A flow duct 37 is provided on the front side of the blade tip 29 and closed by a protective cover 38. The protective cover 38 is designed profiled, so that any contouring of the blade tip 29 can be retained.

FIG. 3 shows a view, by analogy with FIG. 2, where additional webs or supports 39 are provided which support the protective cover 38. The webs or supports 39 furthermore enable several flow ducts 37 to be formed, or the airflow through the flow duct 37 to be optimized.

FIG. 4 shows a view, by analogy with FIGS. 2 and 3, where in the exemplary embodiment of FIG. 4 the blade tip has no lateral overhang (winglet overhang).

FIGS. 5 to 12 show differing exemplary embodiments of the invention, where the perspective view is schematic and where the protective cover 38 is only shown in simplified form in order to make clear the flow through the flow duct 37.

In all exemplary embodiments of FIGS. 5 to 12, the front suction-side area of the blade tip 29 has the reference numeral 30, while the rear area has the reference numeral 31. A blade trailing edge 32 is designed in the usual way.

FIG. 5 shows an exemplary embodiment in which the protective cover 38 is provided on the entire front-side area of the blade tip 29. The flow necessary for cooling the blade tip is introduced centrally at the blade leading edge, while the outflow through a suction-side opening takes place in the rear area of the blade tip.

FIG. 6 shows a design variant of the exemplary embodiment of FIG. 5, where the protective cover 38 extends over only part of the total length of the blade tip 29.

Complementing the design variant of FIG. 5, a centric support 39 is provided in the exemplary embodiment of FIG. 7, which divides the flow duct 37.

In the exemplary embodiment of FIG. 8, the outflow is provided, in a variation of the exemplary embodiment of FIG. 7 on the pressure side of the rear area 31 of the blade tip 29.

The exemplary embodiment of FIG. 9 represents a variant of the exemplary embodiments of FIGS. 7 and 8 and has a centric outlet opening in the area of the blade trailing edge 32.

In the exemplary embodiment of FIG. 10, the protective cover extends completely over the front area of the blade tip 29 and has an outlet opening only on the suction side of the rear area 31. The cooling air is supplied from the blade interior via ducts.

The exemplary embodiment of FIG. 11 also shows a protective cover closed in the front area, by analogy with FIG. 10, where a central web 39 divides the flow duct 37.

In the exemplary embodiment of FIG. 12, it is provided, in a variation from the exemplary embodiments of FIGS. 10 and 11, that individual supports acting as turbulators are arranged inside the flow duct 37.

The exemplary embodiments of FIGS. 10 to 12 each show the supply of the flow through cooling air holes (dust holes) 41.

List of Reference Numerals

• 1 Engine axis
• 10 Gas-turbine engine/core engine
• 11 Air inlet
• 12 Fan
• 13 Intermediate-pressure compressor (compressor)
• 14 High-pressure compressor
• 15 Combustion chamber
• 16 High-pressure turbine
• 17 Intermediate-pressure turbine
• 18 Low-pressure turbine
• 19 Exhaust nozzle
• 20 Guide vanes
• 21 Engine casing
• 22 Compressor rotor blades
• 23 Stator vanes
• 24 Turbine rotor blades
• 25 —
• 26 Compressor drum or disk
• 27 Turbine rotor hub
• 28 Exhaust cone
• 29 Blade tip
• 30 Front suction-side area of blade tip 29
• 31 Rear area of blade tip 29
• 32 Blade trailing edge
• 33 Sealing edge
• 34 Overhang
• 35 Suction side
• 36 Pressure side
• 37 Flow duct
• 38 Cover/protective cover
• 39 Web/support
• 40 Casing
• 41 Cooling air hole

Claims

1. Turbine rotor blade of a gas turbine with a blade tip, on which means for duct-type guidance of cooling air extending from a front suction-side area of the blade tip to a rear area of the blade tip are provided.

2. Turbine rotor blade in accordance with claim 1, wherein the means are provided at least over part of the length of the blade tip.

3. Turbine rotor blade in accordance with claim 1, wherein the means for the inlet of cooling air from the hot gas flow are provided in the front suction-side area of the blade tip.

4. Turbine rotor blade in accordance with claim 1, wherein the means for the outlet of cooling air are provided in the pressure-side area of the rear blade tip or in the suction-side area of the rear blade tip or in the area of the blade trailing edge.

5. Turbine rotor blade in accordance with claim 1, wherein a sealing edge and/or an overhang are/is provided at the blade tip.

6. Turbine rotor blade in accordance with claim 1, wherein the means are designed in the form of a cover arranged at the blade tip and forming a flow duct.

7. Turbine rotor blade in accordance with claim 1, wherein the means are designed for the introduction of air exiting at least one air duct extending inside the turbine rotor blade.

8. Turbine rotor blade in accordance with claim 1, wherein at least one web or support is provided at the blade tip for guiding the cooling air in the flow duct.

9. Turbine rotor blade in accordance with claim 6, wherein the cover is arranged on the web.

10. Method for cooling a blade tip of a turbine rotor blade of a gas turbine, where air from a hot gas flow is guided from a front suction-side area of a blade tip to a rear area of the blade tip through a duct-type guidance.