GB2361035A - Gas turbine engine vane with noise attenuation features - Google Patents

Abstract

A hollow aerofoil vane (21) for a gas turbine engine contains a honeycomb material (30) that abuts the internal surface of the wall (24) defining the concave pressure surface (25) of the vane (21). A series of apertures (33) in the wall (24) provide communication between the exterior of the wall (24) and the chambers (31) of the honeycomb material (30). The vane (21), in combination with similar vanes, provides attenuation of axially and tangentially propagating noise.

Description

2361035 1 GAS TURBINE ENGINE NOISE SUPPRESSION This invention relates to

the suppression of noise in gas turbine engines.

There is a constant demand to reduce the noise produced by gas turbine engines. This is particularly so in the case of gas turbine engines that are used for aircraft propulsion. Airport regulations and aircraft noise certification requirements which dictate the maximum levels lo of noise that aircraft are permitted to produce are extremely demanding and likely to become more so in the future.

The significant sources of noise in a typical gas turbine engine are its fan, compressor, turbine and exhaust jet. Noise from these sources obey different laws and mechanisms of generation. However, they all tend to produce increased noise levels with correspondingly increased airflow velocity. Exhaust jet noise varies by a larger factor than compressor or turbine noise. Consequently, noise reduction efforts have tended to concentrate on the reduction of exhaust jet velocity since this has a stronger influence on noise reduction than equivalent reductions in compressor and turbine airflow velocities.

As gas turbine engines have become more sophisticated, the noise generated by compressor and turbine components has become an increasing problem. There is a trend towards the use of higher by-pass ratios in gas turbine engines in which a higher proportion of the total gas flow through the engine passes through its by-pass duct rather than through its core engine. One result of this is that the diameters of the fans of such engines are becoming larger. Because of the need to limit the tip velocities of such fans (for noise reasons) the rotational speeds of the fans and the low pressure turbines that power them are becoming slower. This lower rotational speed results in lower frequencies of interaction between the aerofoil blade rows in the turbine, 2 so causing the generated noise to move into the audible range. The perceived noise generated by the engine thus increases.

It is the interaction of pressure fields and turbulent wakes from rotating aerofoil blades and stationary aerofoil vanes that generates noise in both compressors and turbines. The noise can be defined as being of two distinct types: discrete tone (single frequency) and broadband (a wide range of frequencies). Discrete tones are produced by jo the regular passage of blade wakes over downstream stages, causing a series of tones and harmonics from each stage. The wake intensity is largely dependent upon the distance between the rows of blades and vanes. If the distance is very short, there is an intense pressure field interaction which results in the generation of a strong tone. With a high by-pass ratio engine, the low pressure fan blade wakes passing over downstream vanes produces such tones, but they are of a lower intensity due to lower velocities and larger blade/vane separations. Broadband noise is produced by the reaction of each blade to the passage of air over its suirface, even with a smooth airstream. Turbulence in the airstream passing over the blades increases the intensity of the broadband noise and can also induce tones.

It is known to provide noise 0 suppression in gas 25 turbine engines by lining the duct walls of the engine with a sound absorbent material. Typically such material comprises a double skin structure in which one of the skins is perforate and the skins are maintained in spaced apart relationship by a honeycomb material. While such materials are effective in bringing about reductions in noise, there are drawbacks to their use. In particular, the perforate material which necessarily confronts the airflow through the duct has a frictional effect upon that airflow which in turn has a slight but significant detrimental effect upon engine efficiency. Additionally, the use of such sound a_sorbent material lining on duct walls is not particularly 3 effective in absorbing noise which propagates in generally axial directions through a gas turbine engine. Moreover, there are locations within a gas turbine engine in which is not possible to utilise duct wall sound absorbent material.

It is an object of the present invention to provide means for bringing about a reduction in the noise emitted by a gas turbine engine and in particular the noise associated with the aerofoil members in such an engine.

According to the present invention, a hollow aerofoil 10 member for use in a gas turbine engine includes a leading edge and a trailing edge, contains a structure defining a plurality of enclosed chambers, and is provided with a first wall defining a convex suction surface and a second wall defining a concave pressure surface, said first and second walls being joined to define said leading and trailing edges, at least part of said second wall defining said concave suction surface being perforate to provide communication between the exterior of said concave suction surface and said enclosed chambers.

Such aerofoil members are effective in the attenuation of generally axially and tangentially propagating noise within the engine and noise generated through the interaction between axially adjacent aerofoil members.

Said chambers are preferably of generally constant depth.

Said enclosed chambers may be defined by a honeycomb material.

Said perforate nature of said second wall defining said pressure surface is preferably facilitated by the provision of a plurality of apertures that have been drilled in said second wall.

Said honeycomb material is preferably supported by support structure provided within said member.

Said perforate part of said second wall may constitute between 5 and 15% of the total surface area of that part of said second wall member adjacent said enclosed chambers.

4 Preferably, the area of said second wall member that is perforate does not extend any further than approximately 80% of the axial chord of said aerofoil member measured from its leading edge.

Said aerofoil may be in the form of a vane for said gas turbine engine.

Alternatively, said aerofoil may be in form of a blade for said gas turbine engine.

Said vane may be in the f orm of an outlet guide vane lo for location downstream of the fan of a ducted fan gas turbine engine.

Said vane may be in the form of a stator vane for location in the low pressure turbine of said engine.

Said vane may be in the form of an outlet guide vane is for the low pressure turbine of said engine.

Said perforate portion of said second wall may extend over between 5 and 95% of the total span-wise extent of said member.

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

Figure 1 is a schematic sectioned side view of a ducted fan gas turbine engine that includes hollow aerofoil members in accordance with the present invention.

0 Figure 2 is a sectioned view of a hollow aerofoil member in accordance with the present invention.

Figure 3 is a partially broken away view of part of the hollow aerofoil member shown in Figure 2.

Figure 4 is a sectioned view of an alternative embodiment of a hollow aerofoil member in accordance with the present invention.

With reference to Figure 1, a ducted fan gas turbine engine generally indicated at 10 is of conventional configuration comprising a core engine 11 that drives a propulsive fan 12. The fan 12 is contained within an annular casing 13 that is mounted from the core engine 11 by means of a plurality of radially extend vanes 14.

r ing outlet guide The core engine 11 comprises, in axial flow series, an intermediate pressure compressor 15, a high pressure compressor 16, combustion equipment 17, a high pressure turbine 18, an intermediate pressure turbine 19 and a low pressure turbine 20. An annular array of radially exte.nding outlet guide vanes 8 is provided downstream of the low pressure turbine 20.

Air drawn in by the fan 12 is divided into two flows. The first flow passes over the outlet guide vanes 14 before being exhausted to atmosphere to provide propulsive thrust. The second flow passes into the core engine 11 where it is compressed by the intermediate and high pressure compressors 15 and 16. The resultant compressed air is then directed into the combustion equipment 17 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 18, 19 and 20, before being exhausted to atmosphere at the rear of the engine 10 to provide additional propulsive thrust.

Thq high, intermediate and low pressure turbines 18, 19 and 20 are respectively drivingly interconnected with the high and intermediate pressure compressors 16 and 15, and the fan 12.

In the low pressure turbine 20, an annular array of aerofoil stator vanes 21 is interposed, in the conventional manner, between two annular arrays of aerofoil rotor blades. The aerofoil stator vanes 21 are hollow as can be seen from the cross-sectional view of one such vane that can be seen in Figure 2.

Each aerofoil stator vane 21 comprises a first wall 22 that defines a convex suction surface 23 and a second wall 24 that defines a concave pressure surface 25. The first and second walls 22 and 23 are joined at each of their 6 is chordal extents to define the leading and trailing edges 26 and 27 respectively of the stator vane 21.

The interior 28 of each vane 21 contains a support member 29 which, in combination with the second wall 23, is of approximately rectangular cross-sectional shape. The support member 29 extends throughout the majority of the radial extent of the vane 21 and serves to support a honeycomb material 30 which similarly extends throughout the majority of the radial extent of the vane 21. The lo actual span-wise extent of honeycomb material 30 is dependent upon the particular noise problems in its immediate vicinity. However, generally speaking, the spanwise extent of the honeycomb material 30 is within the range 5-95% of the total aerofoil span.

The honeycomb material 30 is of generally conventional configuration, as can be seen from Figure 3, comprising a network of chambers 31, each of which extends from the rear wall 32 of the support member 29 to the to the second wall 24 defining the concave pressure surface. The rear wall 32 of the support member 29 is so configured as to be a constant distance from the second wall 24, so ensuring that the honeycomb chambers 31 are all enclosed and generally of the same depth.

The second wall 24 defining the concave pressure surface 25 is partially perforate by virtue of a plurality of apertures 33 provided by drilling over part of its overall extent. The apertures 33, which in this particular embodiment of the present invention constitute approximately 7% of that portion the surface area of the second wall 24 adjacent the honeycomb material 30, are in a regular pattern and serve to provide communication between the chambers 31 and the exterior of the second wall 24. Although in this particular case the apertures 33 constitute approximately 7% of the surface area of that portion of the second wall 24 adjacent the honeycomb material 30, they could be anywhere within the range 5-15% 7 of the surface area depending upon the particular characteristics of the engine 10.

It will be seen, therefore, that the chambers 31, by virtue of their communication with the exterior of the second wall 24 via the apertures 33, constitute a sound absorption structure. The depths of the chambers 31 are so arranged that they are of a value consistent with the effective absorption of noise in their immediate vicinity. Thus sound waves enter the chambers 31 through the lo apertures 33 whereupon they are dissipated within the chambers 31, heat.

The apertures 33 may be formed by any convenient drilling process. Alternatively, if the vane 21 is produced by casting, it may be convenient to define the apertures 33 during the casting process.

Although in the present embodiment, the chambers 31 are defined by the honeycomb material 30, it will be appreciated that other means could be employed to define those chambers 31.

The vane 21, together with the accompanying vanes in its annular array, effectively span the annular gas passage through the low pressure turbine 20. Consequently, a large proportion of the pressure turbine directions will their energy ultimately being converted to noise propagating through the low 20 in generally axial and tangential impinge those portions of the vanes 21 provided with the apertures 33. That noise will, in the main, have been produced by the interaction between the blades and vanes in the engine 10 and the airflows over them. The noise will thus be attenuated, thereby in turn reducing the overall noise emitted by the gas turbine engine 10.

The sound absorption material conventionally used on duct walls in gas turbine engines is effective, to a certain extent, in absorbing tangentially propagating noise. However, aerofoil members in accordance with the 8 present invention ensure that there is more efficient absorption of that tangentially propagating noise as well absorption of the generally axially propagating noise.

Moreover, the use of aerofoil members in accordance with the present invention ensures that there is effective attenuation of generally axially and tangentially propagating noise in those parts of a gas turbine engine in which the use of duct wall sound absorption material is not possible.

It is inevitable that the modification of any surface on an aerofoil member will affect the airflow over that surface. Thus the provision of a large number of apertures 33 in the surface of the aerofoil vane 21 will have a detrimental effect on the airflow over that vane 21.

Specifically, the apertures 33 disturb the boundary layer flow over the aerofoil member 21 and this would normally be expected to add to the aerodynamic losses suffered by the vane 21. However, in the case of the present invention, the apertures 33 are present only in the concave suction surface 25 of the aerofoil vane 21. This is a region in which air flow velocities are low so that aerodynamic losses are correspondingly low. Indeed, the boundary layer of the concave pressure surface 25 normally only contributes less than 20% of the total aerodynamic loss of the vane 21. Additionally, air flowing over the concave pressure surface 25 is subsequently subject to strong acceleration which acts to dampen disturbances induced in the boundary layer. This being so, the aerodynamic penalties resulting from the provision of the apertures 33 are negligible. In order to ensure that this remains the case, the apertures 33 are located on the concave pressure surface 25 such that they are only positioned in low air velocity areas. Typically, in order to achieve this, the apertures 33 are so located that they do not lie beyond 80% of the axial chord from the vane leading edge 26.

9 During assembly, the honeycomb material 30 is anchored within the support member 29 by brazing. Thereupon, the support member 29, together with its associated honeycomb material 30, is located and retained within the hollow vane s 21 by brazing so that the apertures 33 in the second wall 24 are aligned with the enclosed chambers 31 defined by the honeycomb material 30.

It will be appreciated, however, that the aerofoil vane 21 could be manufactured in other ways. Thus, it could lo be cast with support member 29 as an integral part thereof but without the second wall 24 defining the concave pressure surface 25. The honeycomb material 30 could then be brazed in place within the support member 29 and a separate perforate second concave pressure surface wall 24 brazed in place on top of that honeycomb material 30.

Although the present invention has been described with reference to a stator vane 21 positioned in the low pressure turbine 20, it could also be applied to other aerofoil members, whether they are stator vanes or indeed rotary blades.

A further example of an aerofoil member in accordance with the present invention is shown in Figure 4. Figure 4 depicts a cross-sectional view of one of the hollow outlet guide vanes 14 that are located downstream of the fan blades 12. Although of the same general configuration as the stator vanes 21, the outlet guide vanes 14 are larger. Nevertheless, components common to both embodiments are provided with the same reference numbers.

The larger size of the outlet guide vanes 14 together with its slightly different cross-sectional configuration and the particular acoustic environment downstream of the fan 12, dictates that the honeycomb material 30 occupies a slightly smaller proportion of the interior of the outlet guide vane 14 than is the case with the low pressure turbine stator vane 21. This being so, the honeycomb support member 34 is of slightly different configuration to that of the low pressure turbine vane 21 in order to ensure effective location of the honeycomb material 30 within the outlet guide vane 14. However, in other respects, the honeycomb material 30 functions in the same manner as described earlier, serving to absorb axially and tangentially noise propagating downstream of the fan 12.

The turbine outlet guide vanes 8 are generally of the same general configuration as the fan outlet guide vanes 14 and serve to provide further attenuation of noise generated lo within the low pressure turbine 20.

It will be seen, therefore that the present invention provides means for ensuring that generally axially and tangentially propagating noise within a gas turbine engine is attenuated.

11

Claims (14)

1. A hollow aerofoil member for use in a gas turbine engine including a leading edge and a trailing edge, said aerofoil member containing a structure defining a plurality of enclosed chambers, and is provided with a first wall defining a convex suction surface and a second wall defining a concave pressure surface, said first and second walls. being joined to define said leading and trailing edges, at least part of said second wall defining said concave suction surface being perforate to provide communication between the exterior of said concave suction surface and said enclosed chambers.

2. A hollow aerofoil member for a gas turbine engine as 15 claimed in claim 1 wherein said chambers are of generally constant depth.

3. A hollow aerofoil member for a gas turbine engine as claimed in claim 1 or claim 2 wherein said enclosed chambers are defined by a honeycomb material.

4. A hollow aerofoil member for a gas turbine engine as claimed in any one preceding claim wherein the perforate nature of said second wall defining said pressure surface is facilitated by the provision of a plurality of apertures that have been drilled in said second wall.

5. A hollow aerofoil member for a gas turbine engine as claimed in claim 2 wherein said honeycomb material is supported by a support structure provided within said member.

6. A hollow aerofoil member for a gas turbine engine as 30 claimed in any one preceding claim wherein said perforate part of said second wall constitutes between 5 and 15% of the total surface area of that part of said second wall member adjacent said enclosed chambers.

7. A hollow aerofoil member for a gas turbine engine as claimed in any one preceding claim wherein the area of said second wall member that is perforate does not extend any 12 further than approximately 80% of the axial chord of said aerofoil member measured from its leading edge.

8. A hollow aerofoil member for a gas turbine engine as claimed in any one preceding claim wherein said aerofoil member is in the form of a stator vane.

9. A hollow aerofoil member as claimed in any one of claims 1-7 wherein said aerofoil member is in the form of a rotor blade.

10. A hollow aerofoil member for a gas turbine engine as claimed in claim 8 wherein said member is in the form of an outlet guide vane for location downstream of the fan of a ducted fan gas turbine engine.

11. A hollow aerofoil member for a gas turbine engine as claimed in claim 8 wherein said member is in the form of a stator vane for location in the low pressure turbine of said engine.

12. A hollow aerofoil member as claimed in claim 8 wherein said member is in the form of a outlet guide vane for the low pressure turbine of said engine.

13. A hollow aerofoil member as claimed in any one preceding claim wherein said perforate portion of said second wall extends span-wise over between 5 and 95% of the total span-wise extent of said member.

14. A hollow aerofoil vane for a gas turbine engine substantially as hereinbefore described with reference to, and as shown in the accompanying drawings.

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