GB2237071A

GB2237071A - Compressor assembly

Info

Publication number: GB2237071A
Application number: GB8923526A
Authority: GB
Inventors: John Edmond Hatfield
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1989-10-18
Filing date: 1989-10-18
Publication date: 1991-04-24
Also published as: JPH03151600A; DE4031620A1; GB8923526D0

Abstract

A gas turbine engine compressor assembly comprises a centrifugal fluid flow impeller 12 and a vaned diffuser 13 located downstream of the impeller 12. The diffuser vanes 26 each have a leading edge 28 which is so configured that efficiency losses arising from the interaction between the diffuser vanes 26 and the compressor 12 are minimised. The leading edge configuration is related to the local stream energies of fluid flow exhausted from the impeller. <IMAGE>

Description

0 r 4 1 COMPRESSOR ASSEMBLY This invention relates to a compressor

assembly and in particular to a compressor assembly which includes a vaned diffuser.

It is well known in gas turbine and related fields to utilise a device known as a diffuser in order to reduce the velocity of a fluid flow and thereby provide a corresponding increase in its pressure. Diffusers typically consist of a duct which progressively increase6 in crosssectional area in the direction of fluid flow or alternatively fixed vanes which, with other structure, define passages of increasing crosssectional area. It is the latter type of diffuser with which the present invention is concerned.

In a gas turbine engine. a diffuser is frequently located downstream of an air compressor in order to increase the pressure of the air exhausted from the compressor prior to that air being directed into combustion apparatus. Generally speaking, the reduction in velocity of the air flow entering the combustion apparatus and the increase in its pressure provides improved combustion efficiency as well as ensuring that the air compressor operates at its optimum effectiveness.

If the compressor is of the centrifugal flow type, the diffuser typically comprises an annular array of radially extending vanes surrounding the air flow outlet of the compressor rotor. In order to ensure that the diffuser is as effective as possible in performing its task of decreasing the air flow velocity and increasing its pressure, the leading edges of the vanes are desirably positioned as closely as possible to the compressor air f low outlet. However, this can bring about undesirable fatigue failure problems in the compressor rotor. Essentially the forward wakes caused by the interaction of the compressor rotor and the diffuser vanes cause 1 2 excitation of the radially outermost extents of the compressor rotor blades which in turn can lead to their failure through fatigue.

The problem of fatigue failure can be overcome by spacing apart the diffuser vanes and compressor rotor by a distance which is so large as to ensure that the compressor rotor/diffuser vanes interaction is sufficiently small as to ensure that compressor rotor fatigue failure does not occur. However this can bring about in turn a reduction in the effectiveness of the diffuser.

It is an object of the present invention to provide a vaned diffuser which may be positioned sufficiently close to a compressor air flow outlet as to achieve an effective degree of air diffusion while at the same time minimising the extent of forward wake interaction which could have a deleterious effect upon the fatigue life of the compressor rotor.

According to the present invention, a compressor assembly suitable for a gas turbine engine comprises a fluid flow compressor hairing an axis of rotation and carrying a plurality of rotor blades, and a vaned diffuser located downstream of the fluid flow outlet of said compressor, said diffuser comprising a plurality of stator vanes and associated structure which cooperate to define a plurality of diffusion passages, each of said vanes having a_ leading edge in flow-wise spaced apart relationship with the outlet of said compressor, each of said leading edges being so configured that the distance between each position on said leading edge and said compressor fluid flow outlet is proportional to the local stream energy of the fluid flow operationally exhausted from said compressor and is also sufficiently large as to minimise interaction between static fluid "wakes associated with said diffuser vanes and passing rotor blades on said compressor.

1 3 The invention will now be described, by way of example, with reference to the accompanying drawings in which:

Figure 1 is a sectioned side view of a gas turbine engine incorporating a compressor assembly in accordance with the present invention.

Figure 2 is a sectioned side view of a portion of the compressor assembly of the gas turbine engine shown in Figure 1.

Figure 3 is a view on section line A-A of Figure 1.

With reference to Figure 1, a gas turbine engine generally indicated at 10 comprises a radial flow air intake 11 which directs air to a centrifugal flow compressor 12. Air compressed by the centrifugal compressor 12 is directed through a diffuser 13 which is located radially outwardly thereof where its velocity is reduced and pressure increased. From the diffuser 13 the air passes through a duct 14 and into a heat exchanger 15 where its temperature is raised by being placed in heat exchange relationship with the hot exhaust efflux from the engine 10.

The heated air from-the heat exchanger 15 is directed through a further duct 16 into an annular reverse flow combustion chamber 17. There the air is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through a first axial flow turbine 18, which drives the compressor 12, and a second axial flow turbine 19, which is a power turbine driving a power output shaft 20. The power output shaft 20 extends through appropriate apertures provided in the compressor 12 and the first turbine 18 to emerge at the upstream end of the engine 10.

The hot combustion products exhausted from the power turbine 19 are directed radially outwards through the heat exchange 15 where, as previously described, they are placed in heat exchange relationship with the air flow 1 4 from the diffuser 13. The now cooled combustion products are finally exhausted to atmosphere through a radial exhaust outlet 21.

The present invention is particularly concerned with the compressor assembly of the engine 10,, that is the assembly comprising the centrifugal compressor 12 and the diffuser 13 located downstream of it. The construction of the compressor assembly can be seen more easily if reference is now made to Figures 2 and 3.

Referring to Figures 2 and 3, the centrifugal compressor 12 comprises a rotor 21 of conventional configuration which carries a plurality of blades 22, one of which can be seen in Figure 2. In operation air enters the centrifugal compressor 12 in an axial direction at its inlet end 23 and is exhausted in a radial direction from its outlet end 24.

Air exhausted from the outlet end 24 of the centrifugal compressor 12 is directed into the diffuser 13 which, as previously stated, serves to reduce the velocity of the air and increase its pressure. The diffuser 13 is a static annular structure and comprises two axially spaced apart walls 25 which are interconnected by a plurality of diffuser vanes 26, one of which can be seen in Figure 2. Each of the diffuser vanes 26 extends in a generally radially outward direction so that circumferentially adjacent vanes 26 define generally radially extending. passages 27 which increase in cross-sectional area in the direction of air flow. thereby providing diffusion of that air flow.

As air is exhausted from the compressor 12 it follows a logarithmic spiral path and so frictional losses occur prior to the air entering the diffuser 13. It is important therefore that the diffuser vanes 26 are as close as possible to the outlet end 24 of the compressor 12 in order to minimise these losses. However, as air flows over the stator vanes 26, static wakes are p r established in the region of the generally axially extending leading edges 28 of the vanes 26. These wakes extend in a generally upstream direction with respect to the direction of air f low so that there is a danger that they will be intercepted by the passing blades 22 on the compressor rotor 21. If this occurs to any major extent, high frequency fatigue failure of the blades 22 can occur. In order to minimise this ef fect, the stator vanes 26 leading edges are spaced apart from the rotor blades 22 by a distance which ensures that such wake interaction is minimised.

Unfortunately the velocity distribution of air exhausted from the outlet 24 of the compressor 12 is not uniform as a result of boundary layer effects. Essentially the closer the air flow is to the boundaries. of the compressor 12, the lower is its velocity. This means that air exiting from the axial extents of the outlet 24 of the compressor 12 has lower velocities than air exiting from the region intermediate those axial extents. A reduction in air flow velocity at the outlet 24 of the compressor 12 affects the angle at which the air flow exits the compressor 12. Thus referring to Figure 3, as the velocity of the air f low decreases, so does the angle at which the air flow exits the compressor 12. It will be seen therefore that as the angle decreases, the distance the air flows before contacting the leading edges 28 of the diffuser vanes 26 increases correspondingly. This leads in turn to efficiency losses.

In order to counter such efficiency losses, the leading edge 28 of each diffuser stator vane 26 is curved so that the axial extents of each leading edge 28 are closer to outlet 24 of the compressor 12 than the region intermediate those axial extents. The curvature of each leading edge 28 is so arranged that the distance between each position on the leading edge 28 and the outlet 24 of the compressor 12 is proportional to the local stream 0 6 energy of the air flow exhausted from the compressor outlet 24. Since the local stream energy of the air flow is directly related to its velocity, then those regions of each diffuser vane leading edge 28 which are at the axial extents of the diffuser 13, where air flow velocities are lower, are closer to the compressor outlet 24 than the remaining leading edge 28 regions.

A further factor which must be taken into account in the positioning of the diffuser vane leading edges 28 is wake interaction. Thus as previously stated the diffuser vane leading edges 28 must be so positioned that the interaction between the static wakes associated therewith and the compressor rotor blades 22 is minimised. This is achieved by locating the furthermost portion of each leading edge 28 from the compressor outlet 24 at the distance B from the outlet 24; B being the optimum distance for ensuring that wake interaction is minimized while diffuser 13 efficiency is not excessively compromised. A result of this is of course that certain portions of the diffuser vane leading edges will be less than the distance B from the compressor outlet 24. However this is acceptable since the air flow stream energy, and hence its velocity, is lower in these regions and so wakes formed on those leading edge 28 regions will be smaller and therefore less prone to interaction with the compressor rotor blades 22.

The present invention therefore provides a compressor assembly in which rotor blade fatigue failure through wake interaction is minimised while at the same time efficiency losses through excessive diffuser vane/compressor outlet spacing are avoided.

Although the present invention has been described with reference to a compressor assembly which comprises a centrifugal flow compressor 12, it will be appreciated that it is also applicable to axial flow compressors.

3 1 7

Claims

Claims: -

I. A compressor ass'embly suitable for a gas turbine engine comprising a fluid flow compressor having an axis of rotation and carrying a plurality of rotor blades, and a vaned diffuser located downstream of the fluid flow outlet of said compressor, said diffuser comprising a plurality of stator vanes and associated structure which cooperate to define a plurality of diffusion passages, each of said vanes having a leading edge in flow-wise spaced apart relationship with the outlet of said compressor rotor, each of said leading edges being so configured that the distance between each position on said leading edge and said compressor fluid flow outlet is proportional to the local stream energy of the fluid flow operationally exhausted from said compressor rotor and is also sufficiently large as to minimize interaction between static fluid wakes associated with said diffuser vanes and passing rotor blades on said compressor.
2. A compressor assembly as claimed in claim I wherein said fluid flow compressor is of the centrifugal flow type.
3. A compressor assembly as claimed in claim 2 wherein each of said leading edges of said diffuser vanes is generally axially extending.
4. A compressor assembly as claimed in claim 2 or claim 3 wherein said diffuser is located radially outwardly of the outlet of said compressor.
5. A compressor assembly as claimed in any one preceding claim wherein said associated structure of said diffuser comprises two axially spaced apart walls which are interconnected by said diffuser vanes.
6. A compressor assembly substantially as hereinbefore described with reference to and as shown in the accompanying drawings.
7. A gas turbine engine provided with a compressor assembly as claimed in any onepreceding claim.

Published 1991 atThe Patent Ofricc. State House. 65/71 High Holborn, londonWCIR4TP. Further copies may be obtained from SI ales Branch..Unit 6. Nine Mile Pbint. Cwmfelinfach. Cross Keys. Newport NP1 7HZ. Printed by Multiplex techniques ltd. St Mary Cray. Kent.