GB2366600A - Cooling arrangement for trailing edge of aerofoil - Google Patents

Abstract

An aerofoil member, such as a gas turbine nozzle guide vane or turbine blade, comprises a suction surface that extends beyond a pressure surface at a trailing edge, to define an overhang portion 40. Cooling holes/slots 42 are provided to convey cooling air from a cavity within the aerofoil, the holes/slots being angled so as to direct the cooling air onto the overhang portion.

Description

2366600 GAS TURBINE ENGINE SYSTEM This invention relates to a gas turbine

engine. More particularly this invention is concerned with the cooling 5 of turbine nozzle guide vanes or turbine blades of suchlan engine.

An important consideration at the design stage of a gas turbine engine is the need to ensure that certain parts of the engine do not absorb heat to an extent that is to detrimental to their safe operation. One area of the engine where this. consideration is of particular importance is its turbine.

The efficiency of a gas turbine engine is dependent on high turbine entry temperatures which are limited by the 15 turbine blade and nozzle guide vane materials. Continuous cooling of these components allows their environmental operating temperatures to exceed the material's melting point without affecting the blade and vane integrity.

There have been numerous previous methods of turbine 20 vane and turbine blade cooling. The use of internal cooling, external film cooling and holes or passageways providing impingement cooling are now common in the design of both turbines and combustors.

The shape of a nozzle guide vane or a turbine blade 25 can also substantially affect the efficiency of the turbine. The hot combustion gases flowing over the surface of a nozzle guide vane forms a boundary layer around both the pressure and suction sides of the vane. Ideally these flows should meet at the trailing edge of the vane 30 providing pressure recovery, the only losses then being friction losses. In practice, however, the boundary layers lose energy and the flow separates at the trailing edge causing drag and trailing edge losses in addition to the friction losses. It is. therefore desirable to manufacture the vane or turbine blade with its trailing edge as thin as possible.

A nozzle guide vane is usually hollow and its trailing edge is commonly provided with an 'overhang' portion, that 5 being the portion where the suction surface extends beyond the pressure surface at the trailing edge. Cooling air is expelled from 'letterbox' slots formed between the suction and pressure sides in the trailing edge, however the overhang portion remains poorly cooled.

10 It has previously been proposed to provide such letter box slots with ribs extending across the slot width in an attempt to control the exhaust cooling flow and permit the use of a thinner overhang. However such ribs have been found to cause overhang cracking and form wakes in the 15 cooling flow which cause mixing with the mainstream air flow.

The mainstream air mixes with the pressure side cooling air and the heat transfer coefficient is not maximised hence providing inefficient cooling of the 20 overhang. The overhang is therefore particularly susceptible to overheating and cracking.

There is a need, therefore, to improve the cooling of trailing edge portions of turbine nozzle guide vanes or blades and/or to try and avoid mixing of the cooling air 25 with the mainstream air thus alleviating the aforementioned problems.

According to the present invention there is provided an aerofoil member comprising a pressure surface, a suction surface and a trailing edge portion, said suction surface 30 being arranged to extend beyond said pressure side at the trailing edge portion of said aerofoil member, said aerofoil member also comprising at least one cavity for receiving cooling air and at least one aperture formed between the pressure and suction surfaces of the aerofoil member at its trailing edge portion f or exhausting said cooling air, wherein said at least one aperture is angled toward the portion of the suction surf ace extending beyond the pressure surface so as to direct cooling air onto said 5 extending portion and thereby provide cooling thereof.

Advantageously the use of angled cooling holes to direct air onto the portion of the suction side extending beyond the pressure side enables this portion to be manufactured more thinly as it is more efficiently cooled.

10 Additionally the cooling air spreads over the "extending' or 'overhanging' portion of the suction side thus reducing the mixing of this cooling air flow with the mainstream airflow hence increasing aerodynamic efficiency.

Preferably the aerofoil member comprises a nozzle 15 guide vane or blade suitable f or use in a gas turbine engine.

Preferably the aerofoil member comprises a series of angled cooling apertures at its trailing edge portion.

Figure 1 is a schematic sectioned view of a ducted gas 20 turbine engine which incorporates a number of aerofoil members in accordance with the present invention.

Figure 2 is a view of a nozzle guide vane and turbine arrangement in accordance with the present invention.

Figure 3 is a section view of a nozzle guide vane of 25 figure 2.

Figure 4 is an enlarged sectioned view of the trailing edge of a nozzle guide vane in accordance with the present invention.

With reference to figure 1, a ducted gas turbine 30 engine shown at 10 is of a generally conventional configuration. It comprises in axial flow series a fan 11, intermediate pressure compressor 12, high pressure compressor 13, combustion equipment 14, high, intermediate and low pressure turbines 15, 16 and 17 respectively and an exhaust nozzle 18. Air is accelerated by the fan 11 to produce two flows of air, the larger of which is exhausted from the engine 10 to provide propulsive thrust. The smaller flow of air is directed into the intermediate 5 pressure compressor 12 where it is compressed and then directed into the high pressure compressor 13 where further compression takes place. The compressed air is then mixed with the fuel in the combustion equipment 14 and the mixture combusted. The resultant combustion products then 10 expand through the high, intermediate and low pressure turbines 15, 16 and 17 respectively before being exhausted to atmosphere through the exhaust nozzle 18 to provide additional propulsive thrust.

Now referring to figure 2 in which a portion of the 15 high pressure turbine 15 for the gas turbine engine 10 is shown more clearly. The high pressure turbine 15 includes an annular array of similar radially extending air cooled aerofoil turbine blades 20 located upstream of an annular array aerofoil nozzle guide vanes 22. The remainder of the 20 turbine is provided with several more axially extending alternate annular arrays of nozzle guide vanes and turbine blades, however these are not shown in figure 2 for reasons of clarity.

The nozzle guide vanes 22 each comprise an aerofoil 25 portion 24 with the passage between adjacent vanes forming a convergent duct 26. The vanes 22 are located in a casing that contains the turbine 15 in a manner that allows for expansion of the hot air from the combustion chamber 14. The nozzle guide vanes 22 are hollow and cooled by the 30 passage of compressor delivery air through them to reduce the effects of high thermal stresses and gas loads. Arrows A indicate the flow of that cooling air. Cooling holes 28 provide both film cooling and impingement cooling of the nozzle guide vanes 22 and turbine blades 20.

In operation hot gases flow through the annular gas passage 30, These hot gases act upon the aerofoil portions of the turbine blades 20 to provide rotation of a turbine disc (not shown) upon which the blades 20 are mounted. The 5 gases are extremely hot and internal cooling of the vanes 22 and the blades 20 is necessary. Both the vanes 22 and the blades 20 are hollow in order to achieve this and in the case of vanes 22 cooling air derived from the high pressure compressor 13 is directed into their radially 10 outer extents through apertures 32 formed within their radially outer platforms 34. The air then flows through the vanes 22 to exhaust therefrom through a large number of cooling holes 28 provided in the aerofoil portion 24 into the gas stream flowing through the annular gas passage 30.

15 Each nozzle guide vane aerofoil 24 comprises a pressure side 24a and a suction side 24b and these portions meet at the trailing edge 36 of the nozzle guide vane 22.

Some of the cooling air within the nozzle guide vane exhaust through a 'letterboxl exhaust portion formed at the 20 trailing edge 36 of the nozzle guide vane 22. Figures 3 and 4 show more clearly the 'letterbox' exhaust 38. This exhaust portion 38 is known generally as a 'letterbox slot' due to its shape and configuration. An overhang portion 40 is formed where the pressure side 24a of the aerofoil 24 25 meets the suction side 24b such that the suction side 24a extends beyond the pressure side 24b.

In the present embodiment of the invention, the Iletterbox exhaust' comprises a series of holes or slots 42 formed within the portion of vane material adjoining the 30 pressure and suction sides 24a, 24b at the trailing edge.

These holes exhaust cooling air, directed from the hollow portions 44 of the vane 22, along the length of the trailing edge 36 of the vane 22.Although such holes are usually drilled or cast any suitable manufacturing technique may be used.

The holes 42 are angled so as to direct air onto the overhang portion 40 and thus provide impingement cooling 5 thereof. The overhang portion 40 also comprises a curved, angled or stepped portion 46 which also assists in directing the cooling air onto the overhang portion 40.

In order to direct the cooling air efficiently onto the overhang portion the apertures 42 may be angled such 10 that their axes intersect the overhang portion 40.

The cooling flow advantageously spreads to substantially cover the overhang portion 40 without mixing with the mainstream flow to an extent where the cooling affect is substantially compromised. The angled slots 42 15 also allow the overhang portion to be manufactured more thinly by improving its cooling. This improves the aerodynamic efficiency of the aerofoil.

Although the above described embodiment of the present invention is directed to a turbine nozzle guide vane it is 20 to be appreciated that the invention is suitable for any aerofoil member requiring cooling, for example a turbine blade.

Claims (1)

1. An aerofoil member comprising a pressure surface, a suction surf ace and a trailing edge portion, said suction 5 surface being arranged o extend beyond said pressure surface at the trailing edge portion of said aerofoil member, said aerofoil member also comprising at least one cavity for receiving cooling air and at least one aperture formed between the pressure and suction surfaces of the 10 aerofoil member at its trailing edge portion for exhausting said cooling air, wherein said at least aperture is angled toward the portion of the suction surface extending beyond the pressure surface so as to direct cooling air onto said extending portion and thereby provide cooling thereof.

15 2. An aerofoil member as claimed in claim 1 wherein said aerofoil member is a turbine nozzle guide vane suitable for a gas turbine engine.

3. An aerofoil member as claimed in claim 1 or claim 2 wherein said aperture comprises an angled cooling hole.

20 4. An aerofoil member as claimed in claim 1 or claim 2 wherein said aperture comprises an angled cooling siot.

5. An aerofoil member as claimed in any one of the preceding.claims wherein the trailing edge of said aerofoil member is provided with a series of angled cooling holes 25 positioned along its length.

6. An aerofoil member as claimed in any one of the preceding claims wherein the angle of said aperture is such that the central axis of said aperture intersects the suction surface portion which extends beyond the pressure 30 surface portion.

8. An aerofoil member substantially as herein described with reference to the accompanying drawings.