GB2381298A

GB2381298A - A turbine blade having a greater thickness to chord ratio

Info

Publication number: GB2381298A
Application number: GB0125739A
Authority: GB
Inventors: Richard Adrian Beverley Mccall
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2001-10-26
Filing date: 2001-10-26
Publication date: 2003-04-30
Also published as: GB0125739D0

Abstract

A turbine blade (30 of fig 2) for a gas turbine engine includes a circumferentially extending shroud (38 of fig 2) at its radially outer tip. At least a tip portion of the blade has a greater thickness (A) to true chord ratio (B) than the normal design value for that particular blade.

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 for each blade extends in a generally circumferential direction from the tip of that blade and the shrouds of adjacent turbine blades together form a generally cylindrical seal around the turbine blades, thus reducing the leakage of hot gases.

Because of the high speed of rotation of the turbine blades, there is a significant force exerted in a radially outward direction on the blades. Particularly because the blades are working at high temperatures (perhaps 9000 to 10000C) the material of the blades tends to creep in a radially outward direction. The shroud on each blade extends circumferentially out from the tip of that blade and therefore the forces on the edge portions or overhang of the shroud are relatively high. There is 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. The greater the overhang, the larger the fillet must be. However, if the fillet is too large it can have a significant adverse effect on the aerodynamic performance of the blade.

According to the invention there is provided a turbine blade for a gas turbine engine, the blade including a circumferentially extending shroud at its radially outer

tip and at least a portion of the turbine blade having a greater thickness to true chord ratio than the optimum value of thickness to true chord ratio for that particular blade.

The turbine blade may include an aerofoil having a radially outer tip portion near the shroud and a radially inner base portion near a root of the blade and at least a tip portion of the aerofoil may have a greater thickness to true chord ratio than the optimum value.

The thickness of the aerofoil may be greater at the tip portion than at the base portion. The thickness may increase gradually from the base portion to the tip portion. The aerofoil may be provided with internal cooling passages which are of greater cross sectional area in the tip portion than in the base portion.

Alternatively, the thickness of the aerofoil may be substantially constant along its length.

According to the invention there is further provided a turbine blade for a gas turbine engine, the blade including a circumferentially extending shroud at its radially outer tip and at least a portion of the turbine blade having a thickness to true chord ratio greater than 0.35.

Preferably the thickness to true chord ratio is between 0.4 and 0.5.

According to the invention there is further provided a high pressure turbine for a gas turbine engine, the turbine including a set of blades according to any of the previous definitions, the shrouds of adjacent blades forming a generally cylindrical seal for hot gases passing through and driving the turbine.

The turbine may include between 55 and 75 blades.

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 diagrammatic cross section through a prior art turbine blade; and Fig. 4 is a diagrammatic cross section through a turbine blade according to the invention.

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 the 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 from the blade in a circumferential direction. The shrouds 38 of adjacent blades together form 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 (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, strengthening fillets 50 are provided for the blade 30, extending between the outer edges 48 of the shroud 38 and the body of the aerofoil at its radially outer end.

Fig. 3 is a cross section through a prior art turbine blade and Fig. 4 a cross section through a turbine blade according to the invention. In each case, the turbine blade includes an aerofoil section 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 and trailing edges 54 and 56 on the high pressure side of the aerofoil. The aerofoil is provided with internal cooling passages 62.

The turbine blade according to the invention, illustrated in Fig. 4, has a greater thickness to axial chord ratio than does the prior art blade. The thickness of a blade is defined as the maximum thickness across the body of the blade as indicated by the arrows A in Figs. 3 and 4. The axial chord is the dimension indicated by the arrows B in Figs. 3 and 4. Increasing the above ratio enables the extent to which the shrouds 38 project circumferentially from the radially outer tip of the aerofoils 34 to be minimised. The inventors have realised that this allows strengthening fillets 50 of a generally conventional size to be used even if larger turbine blades are used.

The thickness to chord ratio of a blade according to the invention is higher than the optimum value from an aerodynamic point of view.

The thickness of the blade may be increased all along

the length of the aerofoil 34. Alternatively, the thickness may be increased only in a radially outer tip region of the blade 30. In this case, the internal cooling passages 40, 62 may be of greater diameter in the radially outer region of the blade than in the root portion. The heat transfer produced by air passing through cooling passages is proportional to the velocity of the air for a fixed pressure drop. Therefore in a smaller diameter passage where air travels faster, the heat transfer is greater. The temperature towards the tip of the blade can be allowed to be greater than towards the root portion, because the radial forces at the tip are lower. Therefore, by making the cooling passages of a greater diameter at the tip portion less cooling takes place in this area and, since the air does not pick up as much heat in this area, it will cool better in the root portion.

There is thus provided a turbine blade which enables fewer, larger blades to be used on a rotor. Blade numbers may be reduced by for example up to 25% without increasing the shroud peak stresses.

A larger blade is not significantly more expensive to manufacture than a smaller blade and therefore cost savings can be made for the whole blade set. Further, because the most difficult parts of a blade to cool are the leading and trailing edges and because there are drag issues to contend with on the trailing edge, the presence of fewer blades can actually improve performance.

Prior to the invention, if a turbine had been designed with fewer blades, it would have been necessary for the shroud 38 of each blade to extend further in the circumferential direction from the blade tip than was the case previously. This increases the tendency of the edges

48 of the shrouds 38 to curl up in the radially outward direction. One way to deal with this problem would have been to increase the size of the fillets 50. However, the fillets would then be required to extend significantly down the turbine blade and would therefore dramatically reduce the work done by that blade. The invention overcomes this problem by increasing the thickness of the blade, thus allowing the strengthening fillets to be a conventional size, whilst still being effective in minimising the curl of the shrouds 38.

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 a gas turbine engine, the blade including a circumferentially extending shroud at its radially outer tip and at least a portion of the turbine blade having a greater thickness to true chord ratio than the optimum value of thickness to true chord ratio for that particular blade.
2. A turbine blade according to claim 1, the turbine blade including an aerofoil having a radially outer tip portion near the shroud and a radially inner base portion near a root of the blade and at least a tip portion of the aerofoil having a greater thickness to true chord ratio than the optimum value.
3. A turbine blade according to claim 2 f wherein the thickness of the aerofoil is greater at the tip portion than at the base portion.
4. A turbine blade according to claim 3, wherein the thickness increases gradually from the base portion to the tip portion.
5. A turbine blade according to claim 4, wherein the aerofoil is provided with internal cooling passages which are of greater cross sectional area in the tip portion than in the base portion.
6. A turbine blade according to claim 1 or claim 2, wherein the thickness of the aerofoil is substantially constant along its length.
7. A turbine blade for a gas turbine engine, the blade including a circumferentially extending shroud at its radially outer tip and at least a portion of the turbine blade having a thickness to true chord ratio greater than 0.35.

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8. A turbine blade according to Claim 7, wherein the thickness to true chord ratio of the portion of the blade is between 0.4 and 0.5.
9. A turbine blade according to Claim 7 or Claim 8, the turbine blade including an aerofoil having a radially outer tip portion near the shroud and a radially inner base portion near a root of the blade and at least a tip portion of the aerofoil having a thickness to true chord ratio greater than 0.35.
10. A high pressure turbine for a gas turbine engine, the turbine including a set of blades according to any preceding claim, the shrouds of adjacent blades forming a generally cylindrical seal for hot gases passing through and driving the turbine.
11. A turbine according to claim 7, wherein the turbine includes between 55 and 75 blades.
12. A turbine blade substantially as herein described with reference to Fig. 4 of the drawings.
13. 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.