GB2384275A

GB2384275A - Cooling of blades for turbines

Info

Publication number: GB2384275A
Application number: GB0123214A
Authority: GB
Inventors: Peter Antony Evans
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2001-09-27
Filing date: 2001-09-27
Publication date: 2003-07-23
Also published as: GB0123214D0

Abstract

A turbine blade 30 comprises an aerofoil 34 with a hollow interior 43, a fir tree root (32, Fig 2) and a shroud 38 at the radially outer tip of the aerofoil 34. The shroud 38 merges with the aerofoil 34 by a corresponding strengthening fillet 50 such that the join between the shroud 38 and the aerofoil 34 helps prevent deformation of the shroud 38 by creep. Cooling orifices 66 and 68 extend from the hollow interior 43 through the shroud 38 to the exterior to further improve the structural integrity of the blade by direct cooling. Orifices 66 extend through the fillets 50 while orifice 68 is located centrally to the shroud 38.

Description

Turbine Blades The invention relates to turbine blades for gas turbine engines and particularly to high pressure turbine blades.

It is known to manufacture high pressure turbine blades for gas turbine engines with shrouds at their radially outer tips. The shroud of each blade extends transversely from the tip of the blade aerofoil, and the shrouds of adjacent turbine blades together form a generally cylindrical seal around the periphery of the turbine blades, thus reducing the leakage of hot gases.

Because of the high speeds of rotation of the turbine blades, there are significant forces exerted in a radially outward direction on the blades. Particularly because the blades are working at high temperatures (perhaps 900 to 1000 C) the material of the blades tends to creep in a radially outward direction. The shroud on each blade extends axially and circumferentially out from the aerofoil tip and therefore the forces. on the edge portions or overhang of the shroud are relatively high, particularly at locations furthest from the body of the blade aerofoil.

There is therefore a tendency for the shroud overhang to curl outwardly in the radial direction. To minimise this problem a strengthening fillet is incorporated into the blade design, the fillet bridging an area between the overhang of the shroud and the body of the turbine blade aerofoil.

Although the presence of the fillet reduces the stresses at the intersection of the shroud and the aerofoil, the levels of stress and the high temperatures can nevertheless cause the shrouds to creep as discussed above, and eventually to rupture where the maximum creep strain has been exceeded. This is often in the shroud fillet area.

According to the invention there is provided a turbine

blade for mounting on a rotor of a gas turbine engine to extend radially therefrom, the blade including an elongate aerofoil and a shroud extending transversely from a radially outer tip of the aerofoil and merging into the aerofoil in a fillet region, wherein the blade includes an elongate orifice for receiving cooling air, the orifice passing through the fillet region.

Where the terms"radial","axial"and "circumferential"are used in relation to the blade, they refer to the orientation of the blade when mounted on a rotor of a gas turbine engine, for rotation thereon. Thus, the radial direction is along the length of the aerofoil, the circumferential direction is transverse to the radial direction, in the direction of rotation of the aerofoil, and the axial direction is along the axis of the gas turbine engine, perpendicular to the circumferential direction.

Preferably the orifice extends from a hollow interior of the blade to the exterior of the blade.

The shroud may extend transversely from the aerofoil around an entire perimeter of the aerofoil. The extent to which the shroud extends from the aerofoil may vary around the perimeter of the aerofoil. Preferably the fillet region extends around the entire perimeter of the aerofoil, where it joins the shroud.

A plurality of the said cooling orifices may be provided, preferably located in the fillet region adjacent to the parts of the shroud which have the highest levels of uncooled creep strain. These may be the parts which extend furthest from the aerofoil. The cooling orifices may each be generally parallel to one another and may be oriented in a plane substantially perpendicular to the axial direction of the aerofoil. Alternatively the orifices may each be oriented substantially perpendicularly to the perimeter of the aerofoil.

The or each orifice may slope radially outwardly and

may be oriented at an angle of between 250 and 700 to the radial direction of the aerofoil. The diameter of the or each orifice may be between 0.3mm and 2.5mm.

The turbine blade may further include an orifice located in a central portion of the shroud. This orifice may extend generally radially from a central region of a radially outer tip of the aerofoil to an exterior of the blade. The diameter of this orifice is preferably greater than the diameters of the aforesaid orifices.

According to the invention, there is further provided a gas turbine engine including a turbine blade as defined in any of the preceding six paragraphs.

An embodiment of the invention will be described for the purpose of illustration only with reference to the accompanying drawings in which: Fig. 1 is a diagrammatic sectional view of a ducted fan gas turbine engine; Fig. 2 is a diagrammatic sectional view of a prior art wall cooled high pressure turbine blade; Fig. 3 is a highly diagrammatic top view of a turbine blade according to the invention, viewed from the radial direction; and Fig. 4 is a highly diagrammatic sectional view of a tip portion of the turbine blade of Fig. 3, sectioned along the line A-A.

With reference to Fig. 1 a ducted fan gas turbine engine generally indicated at 10 comprises, in axial flow series, an air intake 12, a propulsive fan 14, an intermediate pressure compressor 16, a high pressure compressor 18, combustion equipment 20, a high pressure turbine 22, an intermediate pressure turbine 24, a low pressure turbine 26 and an exhaust nozzle 28.

The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by T : he fan 14 to produce two air flows, a first air flow into the intermediate pressure compressor 16 and a second

airflow which provides propulsive thrust. The intermediate pressure compressor 16 compresses the air flow directed into it before delivering the air to the high pressure compressor 18 where further compression takes place.

The compressed air exhausted from the high pressure compressor 18 is directed into the combustion equipment 20 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through and thereby drive the high, intermediate and low pressure turbines 22,24 and 26 before being exhausted through the nozzle 28 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 22,24 and 26 respectively drive the high and intermediate pressure compressors 16 and 18 and the fan 14 by suitable interconnecting shafts.

Fig. 2 illustrates a wall-cooled high pressure turbine blade 30 suitable for the high pressure turbine 22 of the gas turbine engine 10. The blade 30 includes a fir tree root 32 by means of which the blade is mounted for rotation on a rotor (not illustrated) and an aerofoil 34 over which hot gases flow, causing the blade 30 to rotate. At a radially inner end of the aerofoil 34, blade platforms 36 extend from the blade in the circumferential direction. The blade platforms of adjacent blades together form a radially inner barrier for hot gases.

At a radially outer end of the blade 30 a turbine shroud 38 extends transversely of the aerofoil 34. The shrouds 38 of adjacent blades together form a generally cylindrical seal around the periphery of the blades. This creates a radially outer barrier for the hot gases.

The blade illustrated in Fig. 2 is wall-cooled, including cooling passages 40 within its walls 42, and a central space 43. Film cooling is also provided by orifices 44 in the wall of the blade. Air passing through these orifices as indicated by the arrows 46 cools the outer surface of the blade and protects it from the hot

gases.

The high speed of rotation of the blade exerts a radial load on the blade, in the radially outward direction. Particularly because of the high temperatures (typically around 9000 to 10000C) at which the blade operates, the radial load tends to cause creep of the blade. Such creep is a particular problem for outer edges 48 of the shrouds 38. In time, these outer edges 48 tend to curl upwardly as viewed in Fig. 2, i. e. in the radially outward direction.

In order to minimise the above upward curl, a strengthening fillet 50 is provided in the region where the aerofoil merges into the shroud.

Figs. 3 and 4 illustrate a turbine blade 30 according to the invention. Referring particularly to Fig. 3, the blade 30 includes an aerofoil 34 having a leading edge 54 and a trailing edge 56. A convex suction surface 58 extends between the leading and trailing edges 54 and 56 on a low pressure side of the aerofoil and a concave pressure surface 60 extends. between the leading the trailing edges 54 and 56 on the high pressure side of the aerofoil. The suction surface 58 and pressure surface 60 together define a perimeter of the aerofoil. The broken line in Fig. 4 illustrates the position of the perimeter of the aerofoil, below the region where the aerofoil starts to merge into the shroud. The shroud 38 extends transversely of the aerofoil all around its perimeter, so as to overhang the aerofoil. However, it may be seen that in some areas the overhang is greater than in others.

Referring to Fig. 4, a strengthening fillet 50 extends around the perimeter of the aerofoil, where it merges into the shroud 38. The fillet 50 provides additional material in the region where the aerofoil 34 joins the shroud 38.

The fillet 50 includes a smoothly curved outer surface 64 defining an edge of this additional material and providing a smooth join between the aerofoil 34 and the shroud 38.

The fillet 50 strengthens the join between the shroud and the aerofoil and avoids any sharp junctions therebetween, which may cause weakness.

Nevertheless, it has been realised by the inventors that the fillet region is still vulnerable to creep and in the worst case eventually to cracking, particularly in the regions where the shroud overhang is greatest, such as in a central region of the concave pressure surface 60 of the aerofoil. In order to minimise this problem, a turbine blade according to the invention includes elongate cooling orifices 66 provided within the strengthening fillet 50.

Each orifice 66 extends from the central space 43 within the aerofoil, through the fillet 50, to a radially outer exterior of the aerofoil. In the illustrated embodiment, the orifices 66 lie at an angle of about 450 to the radial direction of the aerofoil. This allows the orifices to be as close to the curved outer surface 64 of the strengthening fillet 50 as possible whilst allowing maximum cooling of the fillet. However, the orientation of the orifices may vary depending upon the shape of the fillet 50.

The orifices 66 are typically between 0. 3mm and 2. 5mm in diameter, depending upon the size and shape of the blade 30, in particular the fillet 50.

Referring to Fig. 3, a plurality of orifices may be provided, generally in the regions of the aerofoil where the shroud overhang is greatest. The orifices may be parallel to one another, as indicated by the set of orifices 66a in Fig. 3. The orifices 66a lie in parallel planes each of which is perpendicular to the axial direction of the engine. Having the orifices parallel to one another is the easiest option from a manufacturing point of view. However, the orifices may alternatively be oriented at respectively different angles and may not be parallel to the axial direction of the engine, as illustrated by the orifices 66b in Fig. 4. For example, it

might be most efficient from a cooling point of view if the orifices lie generally perpendicular to the perimeter of the aerofoil. The orifices 66 are illustrated in Fig. 3 with full lines for clarity, although they would not in reality be visible.

In use, cooling air flows from the central space 43 in the aerofoil 34, though the orifices 66, to the exterior of the blade 30, as indicated by the arrows in Fig. 4.

The orifices 66 allow direct cooling of the fillet 50 and minimise the likelihood of creep occurring within the fillet region.

Referring to Fig. 4, an additional cooling orifice 68 extends from a radially outer part of the central space 43 to the exterior of the blade. The diameter of this orifice 68 is greater than the diameter of the orifices 66 in the illustrated example. This orifice 68 allows any dirt particles to be thrown out of the blade without clogging the orifices 66. Because the entrances 70 to the orifices 66 are located radially inwardly of the tip of the blade, there is not any particular tendency for dust particles to enter these orifices. However, if they do so, they are likely to be swept through the orifices rather than clogging them because the angle of the orifices does not force the air through any sharp corners where dust is likely to accumulate.

There is thus provided a turbine blade for a gas turbine engine in which the highly stressed fillet region is cooled directly. The holes may be produced by laser drilling or EDM (electro-discharge machining) approaching from the shroud outer surface. The use of orifices which cool the fillet region directly enables turbine blades to be made with thin shrouds, because the need to provide cooling holes within the shrouds is minimised or avoided. However, the invention may be used with blades having shrouds incorporating cooling holes.

Various modifications may be made to the above

described embodiment without departing from the scope of the invention. In particular, the number and orientation of the orifices 66 may vary depending upon the precise shape of the fillet 50 and the aerofoil. It may be easier from a manufacturing point of view if the orifices 66 are all parallel and such orifices may extend generally in a plane perpendicular to the axial direction of the engine, as indicated by the orifices 66a in Fig. 4. Alternatively, the orifices may extend, for example, generally perpendicularly to the perimeter or the aerofoil, as indicated by the orifices 66b in Fig. 4. Although the invention has been described as applied to a triple spool engine, it may be accommodated in an engine with any number of spools. Also, although the invention has been described with reference to a wall cooled blade, it may be accommodated with any different turbine aerofoil cooling system.

Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon.

Claims

Claims 1. A turbine blade for mounting on a rotor of a gas turbine engine to extend radially therefrom, the blade including an elongate aerofoil and a shroud extending transversely from a radially outer tip of the aerofoil and merging into the aerofoil in a fillet region, wherein the blade includes an elongate orifice for receiving cooling air, the orifice passing through the fillet region.
2. A blade according to Claim 1, wherein the orifice extends from a hollow interior of the blade to the exterior of the blade.
3. A blade according to Claim 1 or Claim 2, wherein the shroud extends transversely from the aerofoil around an entire perimeter of the aerofoil.
4. A blade according to Claim 3 wherein the extent to which the shroud extends from the aerofoil varies around the perimeter of the aerofoil.
5. A blade according to Claim 4 wherein a plurality of the said cooling orifices are provided, located in the fillet region adjacent to the parts of the shroud which have the highest level of uncooled creep strain.
6. A blade according to any preceding claim wherein the cooling orifices are parallel to one another and are oriented in a plane substantially perpendicular to the axial direction of the aerofoil.
7. A blade according to any of Claims 3 to 5 wherein the orifices are oriented substantially perpendicularly to the perimeter of the aerofoil.
8. A blade according to any preceding claim wherein the or each orifice slopes radially outwardly and is oriented at an angle of between 250 and 700 to the radial direction of the aerofoil.
9. A blade according to any preceding claim wherein the diameter of the or each orifice is between 0.3mm and 2.5mm.

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10. A blade according to any preceding claim, the blade including a further orifice located in a central portion of the shroud and extending generally radially from a central region of a radially outer tip of the aerofoil to an exterior of the blade.
11. A blade according to Claim 10 wherein the diameter of the further orifice is greater than the diameters of the aforesaid orifices.
12. A gas turbine engine including a turbine blade according to any preceding claim.
13. A turbine blade substantially as herein described with reference to Figs. 3 and 4 of the drawings.
14. Any novel subject matter or combination including novel subject matter disclosed herein, whether or not within the scope of or relating to the same invention as any of the preceding claims.