GB2299378A - Cooling compressor guide vanes - Google Patents

Abstract

A guide vane stator ring for a compressor outlet of a gas turbine bypass engine comprises an annular array of vanes 12 spaced apart circumferentially. In order to provide additional cooling for the solid portions of the leading edges of the vanes each has a longitudinally extending passageway 14 which is supplied from a common source of cooling air. This cooling air is first cooled by passage through a heat exchanger 20 located in the bypass duct.

Description

STRUCTURAL GUIDE VANE The invention relates to a structural guide vane or the like member. In particular it concerns providing such a member with leading edge cooling.

It is an ever present requirement of gas turbine aeroengine design to keep engine weight as low as safely possible commensurate with providing the static structure with sufficient strength and stiffness. It follows that the load bearing capacity of any conveniently situated part of the static structure of the engine may be utilised if possible. One such development particularly intended to improve the axial and torsional rigidity of the structural casing surrounding the combustion chamber, especially an annular chamber, uses a plurality of generally radially disposed, structural vanes crossing the diffuser passage leading from the high pressure compressor outlet to the combustion chamber outer volume in order to support loads arising from the turbine entry nozzle guide vanes and the pressure differential across the combustor casing walls.

With this kind of arrangement in which structural vanes cross the gas passage leading from the high pressure (HP) compressor load carrying members are exposed to the HP gas temperature and are exposed to temperature fluctuations which arise directly from engine speed changes. Moreover the speed of the airflow is such that the leading edges of these vanes are subject to frictional heating by the compressor outlet airstream.

The purpose of the diffuser passages is to slow the airstream, and to recover pressure head, sufficiently to support combustion in the combustor chamber. However, the restricted lengths of the diffuser for practical purposes is such that the air retains sufficient speed at the vane leading edge to result in significant frictional heating. This drawback inevitably results in the vanes becoming subject to thermal stress cycling and over a period of time or number of cycles this results in crack propagation within the vane material. It is an object of the present invention to extend the useful life of these vanes by mitigating the effects of the thermal cycling.

According to the present invention there is provided a structural guide vane for a gas turbine engine having in the region of its leading edge a cooling air passage.

The invention and how it may be carried into practice will now be described, by way of example only, with reference to the accompanying drawings, in which: Figure 1 shows an axial section through a twin-spool by-pass turbojet engine, Figure 2a shows a more detailed section view of an existing combustor diffuser passage of the engine of Figure 1, and Figure 2b is a view on section line AA of Figure 2a.

Referring now to the drawings in more detail and to Figure 1 in the first instance, there is shown a sectioned view along the axis of an engine illustrating a twin-spool by-pass turbo jet of conventional, well-known type having a low pressure compressor 2, driven through a first shaft 3 by a low pressure turbine 4. The output of compressor 2 is divided into two flows which pass down a by-pass duct 5 and into a high pressure compressor 6 which is driven via a second shaft 7 by a high pressure turbine 8. Between the high pressure compressor 6 and its associated turbine 8 is located an annular combustor 9 which receives air from the compressor 6 via an annular diffuser passage 10.Within the passage 10 there are a multiplicity of radially extending structural vanes 12 spaced apart circumferentially and forming an integral load-bearing part of a structure generally indicated at 11 extending between the HP turbine and HP compressor and surrounding the combustor. A typical diffuser design would have about 22 vanes spaced apart equidistant from each other around the diffuser passage.

In accordance with the present invention the structural vanes 12 are formed with a hollow interior cavity 13 and an internal cooling passage 14 in the vane leading edge extending the full radial height of the vane near to the leading edge of the vane.

Referring now to the more detailed view of Figures 2a, shown sectioned at Figure 2b, the HP compressor outlet flow is indicated by an arrow labelled "HP AIR". This flow which normally has a tangential velocity component is straightened by an annular array of outlet guide vanes (OGVs) 13 spaced apart around the annular diffuser passage 10. Towards the downstream end of this passage 10 is a further array of generally straight, vanes 12 which extend across the passage between inner and outer walls of the diffuser duct 15 and 16 respectively. The particular example referred to is for a rotary axial flow compressor but, it will be understood, the invention may be applied also to an engine having a centrifugal compressor. Generally, however, the compressor has an annular outlet within which a multiplicity of outlet guide vanes 19 are spaced apart circumferentially.Thus it follows also that the inner and outer walls 15,16 are concentric about the longitudinal, rotational axis of the engine.

It is axiomatic in the design of aeroengines to reduce weight whenever possible while simultaneously maintaining or improving the inherent stiffness of the static structure and its ability to absorb or resist bending and torsional loads. In order to carry loads around the combustion chamber and between bearing housings and the engine casing, it is now common to have a plurality of structural ie load-carrying vanes 12 crossing the diffuser passage in order to support loads arising from the turbine entry nozzle guide vanes and pressure differences across the combustor inner casing walls.

These vanes 12 are disposed to extend radially across the passage and are spaced apart circumferentially around the annular diffuser passage 10.

The leading edges of these diffuser vanes 12 are heated or cooled by HP compressor delivery air. The rate at which the vanes are heated or cooled is dependent upon the rapidity of engine acceleration of deceleration and can be at a significantly faster rate than adjacent part of the diffuser. The resulting local differences in thermal expansion of the vanes can generate substantial stresses in the vane material which, over a period of time, leads to crack initiation and propagation and eventually component failure. The present invention is intended to overcome these drawbacks by providing means to reduce the rate of heat transfer in region of the vanes most susceptible to high thermally generated stress. As a result the vanes, and the structure of which they form part, are endowed with substantially enhanced crack resistance and life usage.

Air is extracted from the diffuser exit region 17 through a bleed aperture 18 and passed through a heat exchanger, generally indicated at 20, before being returned to pass through the hollow interior of the diffuser vanes 12 and thence to cool other parts of the engine such as the shaft 7 and turbine rotors 4,8. It is known in connection with heat exchangers in gas turbine engines to use by-pass air, fuel or oil as the cooling medium. In the particular example the chosen medium is by-pass air, thus the heat exchanger 20 is mounted within the engine by-pass duct 5 and on the outer surface of the core engine casing 21. In Figure 2a heat exchanger 20 is illustrated schematically but it will be understood that other types of heat exchanger construction may be employed.HP air bled through aperture 18 is conveyed by an array of tubes 22 to make a series of passes through a heat exchanger matrix comprising the tubes 22 and a series of baffles 24 which are disposed in the by-pass airstream. Heat is thus transferred from the air carried by tubes 22 according to the temperature differential between the temperature of the HP air and the by-pass air.

Cooled air exits the heat exchanger 20 through exit aperture 26, which may be one of series of such apertures in an annular or part-annular manifold, into an annular plenum chamber volume 28 created by bulkhead walls 30,32 which extend in a circumferential direction between the diffuser outer wall 16 and the core engine casing 21.

These bulkhead walls 30,32 form complete annulii around the diffuser wall 16 in the vicinity of the leading and trailing edges of the vanes 12. The cooled air flow flows from plenum chamber 28 through hollow vanes 12, a proportion of the air passing through passages 14 close to the leading edge of the vanes. Typically the passages 14 are parallel sided ie circular cross-section holes of approximately 1 mm diameter. However the shape of the holes may be adapted to suit cooling requirements, for example the holes may be elliptical in cross-section or have a different cross-section. Towards either or both ends of the hole its shape may be flared in order to reduce turbulence for example.By carefully matching the cooling air velocity, passage shape and proximity to the leading edge it is possible to reduce the cyclic stress range experienced by the structural vanes thereby usefully increasing the component life.

After passage through the vanes the cooling air passes into an annular cavity radially inboard of an inner casing of the combustion chamber 9. From there the air is ducted via preswirl nozzles etc to the disc or discs of the HP turbine stage 8 and the turbine blades, some air is also bled to cool the shafts 3,7. Spent cooling air is eventually allowed to exhaust through effusion cooling cools and through leakage paths into the high velocity, low pressure hot gas exhaust stream in well known manner. The pressure differential between this low pressure and the high pressure achieved at the HP compressor exit diffuser is sufficient to drive the cooling air through the cooling system including the by-pass heat exchanger.

Claims (6)

1A guide vane arrangement for a compressor outlet diffuser of a gas turbine engine includes at least one guide vane having a solid leading edge region along the length of the vane through which is formed a longitudinally extending cooling air passage which is in communication with a source of cooling air.

2 A guide vane arrangement as claimed in claim 1 wherein the cooling air passage is in communication with a source of cooled, cooling air.

3 A guide vane arrangement as claimed in claim 2 wherein the source of cooled, cooling air comprises a heat exchanger connected to receive a supply of air from a high pressure compressor outlet bleed.

4 A guide vane arrangement as claimed in claim 3 wherein the heat exchanger is mounted in an engine air by-pass duct.

5 A guide vane as claimed in either claim 3 or claim 4 wherein air from the heat exchanger is collected in a plenum chamber surrounding the diffuser vanes and the vane leading edge passages are in open communication with said chamber.

6 A guide vane substantially as hereinbefore described with reference to the accompanying drawings.