GB2127105A

GB2127105A - Improvements in cooled gas turbine engine aerofoils

Info

Publication number: GB2127105A
Application number: GB08226357A
Authority: GB
Inventors: Neil Milner Evans; Peter James Mccloskey
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1982-09-16
Filing date: 1982-09-16
Publication date: 1984-04-04
Also published as: JPS59136502A; FR2533262A1; DE3333018A1; GB2127105B

Abstract

The aerofoil is provided with an external lengthwise leading edge slot intersecting the chordwise cooling air passages 12 whereby the passage outlets 13 communicate with the slot and are less likely to be blocked by debris. <IMAGE>

Description

SPECIFICATION improvements in cooled gas turbine engine aerofoils The present invention relates to cooled gas turbine engine aerofoils e.g. rotor blades and statorvanes.

One particular form of cooling applied to the leading edge of a cooled turbine blade is film cooling, in which a plurality of small diameter cooling air passages are spaced along the leading edge of the blade and extend chordwise into the interior of the blade to communicate with a common cooling air supply duct which extends longitudinally of the blade. Air flowing from the outlets of the passages at the leading edge flows rearwardly over the two flanks of the blade thus providing a cooling film on the blade surfaces.

A problem which arises with this type of system is that debris ingested by the engine compressor or from compressor abradable coatings and which is heated in the combustion chamber, impinges on the leading edges of the turbine blades in a molten or softened state and sticks to the blades. This eventually causes blockage of the outlets of the cooling air passages, depriving the leading edge area of the blade of its cooling film and causing the blade to become overheated with a resulting reduction in useful blade life.

Various attempts have been made to overcome this problem including enlarging the passages or providing a bell-mouthed outlet for each passage.

Each have shown a marginal improvement due to the increased area of cooling air flow, but in the first case the overall cooling efficiency of the air passing through the enlarged holes was reduced, and in the second case, although the initial outlet area was greater, the rate of build up of debris was greater.

Thus neither of these two modifications proved acceptable.

The object of the present invention is to provide means for reducing the blockage of the cooling air passages adjacent the leading edge of a cooled gas turbine engine aerofoil without detriment to the overall efficiency of cooling of the blade.

This object is achieved by the invention as claimed in the appended claims by the provision of one or more slots extending longitudinally of the aerofoil adjacent the leading edge and intersecting at least some of the cooling air passages whereby the outlets of the intersected passages communicate with the slot.

The width and depth of the slot may vary over a range of values which depend on the angle at which the debris-containing air approaches the aerofoil, and the available pressure of the cooling air supplied to the cooling air passages.

The invention will now be more particularly described, by way of example only with reference to the accompanying drawings in which: Figure 1 is a cross-sectional elevation of a turbine rotor blade made in accordance with the invention, Figure2 is an enlarged cross-section on the line ll-ll of Figure 1, Figure 3 is a further eniarged view of the leading edge only of the blade of Figure 2, and Figure 4 is a view similar to that of Figure 3 of an unmodified blade.

Figures 5 and 6 are views looking towards the leading edges of two alternative slotted blades.

Referring now to the drawings, the turbine rotor blade of Figure 1 has an aerofoil portion 2 and a root portion 4, with a conventional shroud 6 and platform 8. Within the interior of the blade are a plurality of longitudinally extending cooling air supply ducts, A, B and C, which may take any desired shape, and which may be arranged in any desired pattern, except that the duct A is formed adjacent the leading edge 10 of the blade.

The cooling air supply ducts A, B and C extend through both the aerofoil and root portions of the blade to supply cooling air to the blade from a gas turbine engine (not shown) of which the blade forms a part.

From the leading edge cooling air supply duct A, the cooling air is directed to the exterior of the blade at the leading edge, through a plurality of cooling air passages 12 which extend chordwise of the blade.

The cooling air passages 12 are spaced longitudinally of the blade. A slot 14 is cut into the leading edge of the blade to intersect the cooling air passages 12 so that the outlets 13 of the passages open out into the base of the slot.

The build up of debris is most severe near the mid-span and tip of the blade so that it may not be necessary for the slot to run the whole length of the blade. Thus, that part of the leading edge close to the root of the blade may remain unslotted.

By adjusting the width and depth of the slots to the prevailing gas flow velocities and directions over the leading edge of the blade, we have found that the build up of debris in the passages, although not prevented, is very much reduced. The debris builds up as indicated at 16 on the blade surface and in the slot, and it can be seen that although there is blockage of the slot the passages 12 themselves are only marginally affected. It can be seen from Figure 4 that an unmodified blade tested simultaneously side by side with the modified blade suffered significant debris build up in the exit of the passages (show at 20).

Thus the provision of the slot not only significantly reduces the build up of debris in the cooling air passages, but also has the advantage that the cooling air passages may be maintained at the small diameter required to maintain the optimum pressure and velocity of the cooling air flow while using the minimum quantity thereof for most efficient cooling of the leading edge area of the blade. This is because, the provision of the slot, which preferably has a width equal to the diameter of the cooling air passages, increases the exit area of the total cooling airflow, allowing sideways expansion of the cooling air.A still further advantage which occurs in this preferred case is that the increased exit area of the cooling air flow reduces the velocity and increases the static pressure of the cooling air flow leaving the slot, which aids in establishing and maintaining the cooling air films over the flanks of the blades. The slot should be as narrow as possible to minimise adverse aerodynamic disturbances, but a minimum width would be that which gave the same totai outlet area for the flow as the sum of the outlet areas of the passages.

The depth of the slot has to be selected taking into account the positions of the passages, which need not necessarily be at the stagnation point on the leading edge of the blade, but could be slightly to one side of this point, and the approach vector of the oncoming gases.

With the slots downstream of the leading edge and in the wake of the flow over the leading edge, the slots may have a depth of no more than one half of the diameter of the cooling air passages. With the slots at the stagnation point of the leading edge, they may have a depth of two to five times the passage diameter. However, to avoid reducing the length "i" of the cooling air passages which would reduce the effectiveness of the cooling air in cooling the leading edge region of the blade, the slots should not be deeperthan is necessary to minimise the blockage of the passages in operation.

In one example of the invention the cooling air passages were 0.012 ins dia, the slot was 0.014 ins wide, by 0.028 ins deep and extended halfway along the leading edge of the blade from the tip.

Compared to an unmodified blade, after the same length of operation, the blockage in the passages of the slotted blade was less than 20% of the blockage of the unmodified blade, and the maximum temperature detected on the leading edge of the blade was less than that of the unmodified blade.

Figure 5 shows one embodiment of a slotted blade in which the slot is narrower than the width of the passages. In this embodiment the passages 12 extend to the leading edge of the blade so that the slot interconnects them through the spaces therebetween.

Figure 6 shows an alternative embodiment in which the passages 12 terminate short of the leading edge of the blade and the slot runs across the ends of the passages.

Although the invention is described with reference to a rotor blade the principle of the invention is applicable to any cooled aerofoil blades or stator vane in which the system of film cooling from leading edge cooling passages is used.

The cooling air passages have been shown having circular cross-section in which case the diameter of the passages equates to their width. However, non-circular passages may be used.

Also the slot need not be continuous, but may interconnect groups of passages leaving a barrier at each end to guard against centrifugal pumping of air along the slot.

Claims

1. A cooled gas turbine engine aerofoil having a longitudinally extending internal cooling air supply duct and a plurality of cooling air passages which communicate with the duct and extend chordwise of the aerofoil terminating in outlets in an exterior surface of the aerofoil adjacent the leading edge thereof, and characterised by one or more slots extending longitudinally of the aerofoil adjacent the leading edge and intersecting at least some of the cooling air passages whereby the outlets of the intersected passages communicate with the slot.

2. A cooled aerofoil as claimed in claim 1 and in which the width of the slot substantially equal to the width of the cooling air passages.

3. A cooled aerofoil as claimed in claim 1 or claim 2 and in which the depth of the slot lies in the range of between one half and five times the width of the cooling air passages.

4. A cooled aerofoil as claimed in any preceding claim and in which the slot extends from the tip region of the aerofoil at least to the mid-span region thereof.

5. A cooled aerofoil according to any preceding claim and in which the aerofoil is a turbine rotor blade.

6. A cooled aerofoil according to any one of claims 1 to 4 and in which the aerofoil is a turbine stator vane.

7. A cooled gas turbine engine aerofoil substantially as hereinbefore more particularly described with reference to the accompanying drawings.