GB2372780A

GB2372780A - Gas turbine engine nozzle with noise-reducing tabs

Info

Publication number: GB2372780A
Application number: GB0105352A
Authority: GB
Inventors: Kevin Andrew White; David Sydney Knott; Paul Jonathan Railton Strange
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2001-03-03
Filing date: 2001-03-03
Publication date: 2002-09-04
Also published as: GB0105352D0

Abstract

A gas turbine engine exhaust nozzle arrangement 16, comprises a nozzle wall 15, and a plurality of noise-reducing tabs 20. The tabs 20, extend in a generally downstream direction from the downstream periphery of the nozzle and each tab 20, comprises a fin 27, for the purpose of controlling the vortices shed by the tab. The tabs may be deployable and may be curved. The tabs may also be used in the bypass nozzle of an engine and in conjunction with a lobed mixer.

Description

GAS TURBINE ENGINE EXHAUST NOZZLE

The present invention relates generally to gas turbine engine exhaust nozzles, and in particular to noise reduction and performance improvements to nozzle arrangements used on gas turbine engines suited to aircraft propulsion.

Gas turbine engines are widely used to power aircraft.

As is well known, the engine basically provides propulsive power by generating a high velocity stream of gas which is exhausted rearwards through an exhaust nozzle arrangement.

A single high velocity gas stream is produced by a turbojet gas turbine engine. More commonly nowadays however two streams, a core exhaust and a bypass exhaust, are generated by a ducted fan gas turbine engine or bypass gas turbine engine.

The high velocity gas stream produced by gas turbine engines generates a significant amount of noise, which is referred to as exhaust or jet noise. This noise is generated by the high velocity of the exhaust stream, or streams, and the mixing of the streams with the surrounding atmosphere, and in the case of two streams, as the bypass and core streams mix. The degree of the noise generated is determined by the velocity of the stream and how the streams mix as they exhaust through and away from the exhaust nozzle.

Increasing environmental concerns require that the noise produced by gas turbine engines, and in particular aircraft gas turbine engines, is reduced and there has been considerable work carried out to reduce the noise produced by the mixing of the high velocity gas stream (s). A large number of various exhaust nozzle designs have been used and proposed to control and modify how the high velocity exhaust gas streams mix. With ducted fan gas turbine engines particular attention has been paid to the core stream and the mixing of the core and bypass exhaust

streams. This is because the core stream velocity is considerably greater than the bypass stream and also the surrounding atmosphere and consequently the core exhaust stream generates a significant amount of the exhaust noise. Mixing of the core stream with the bypass stream has also been found to generate a significant proportion of the exhaust noise due to the difference in velocity of the core and bypass streams.

One common current exhaust nozzle design, that is widely used, is a lobed type nozzle which comprises a convoluted lobed core nozzle as known in the art. However, this adds considerable weight, drag, and cost to the installation and nowadays short bypass nozzles are favoured with which the lobed type core nozzles are less effective and are also more detrimental to the engine performance than when used on a long cowl arrangement.

Alternative nozzle designs that are directed to reducing exhaust noise are proposed and described in GB 2,289, 921, UK Application GB 0025727.9, EP0913567, EP0984152 and EP0999358. In all these designs there are number of circumferentially spaced tabs or chevrons, of various specified configurations, sizes, spacing and shapes, and which are disposed to the downstream periphery of a generally circular exhaust nozzle. Such nozzle designs are considerably simpler to manufacture than the conventional lobed designs and are lighter and are generally more aerodynamically efficient. All the designs describe that the tabs or chevrons generate vortices in the exhaust streams. These vortices enhance and control the mixing of the core and bypass streams which it is claimed reduces the exhaust noise. However, in all cases each tab or chevron produces two contra-rotating vortices, one vortex from each of their respective side edges. It is believed that these contra-rotating vortices mix with one another creating their own noise and reducing each

vortices'individual effectiveness for mixing the core and bypass exhaust streams and bypass stream and ambient air.

It is therefore desirable and is an object of the present invention to provide an improved gas turbine engine exhaust nozzle which is quieter than conventional exhaust nozzles and/or which offers improvements generally.

According to a first aspect of the present invention there is provided a gas turbine engine exhaust nozzle arrangement for the flow of exhaust gases therethrough between an upstream end thereof and a downstream end thereof comprising a nozzle wall and a plurality of tabs, the nozzle wall having a downstream periphery, the tabs extend in a generally downstream direction from the downstream periphery wherein each tab comprises a fin.

Preferably the nozzle wall is substantially frusto-conical and the plurality of tabs is circumferentially disposed to the downstream periphery.

Preferably each tab comprises lateral edges, the fin disposed to a lateral edge of each tab. Preferably each tab comprises a high static pressure side and a low static pressure side, the fin generally extending radially from the high static pressure side of each tab.

Preferably the high pressure side is radially inward of the low pressure side and the fin generally extends radially inwardly from a lateral edge of each tab.

Preferably the fin generally radially tapers from the periphery of the nozzle wall to the downstream edge of each tab.

Preferably the fin prevents a vortex forming at and being shed from a lateral edge of each tab. Alternatively the fin substantially reduces the strength of a vortex forming at and being shed from a lateral edge of each tab.

Alternatively the fin is so arranged to control the strength of a vortex being generated and shed from a lateral edge of each tab.

Alternatively a number of the tabs comprise fins disposed to the first edge and a number of the tabs comprise fins disposed to the second edge.

Preferably the radial height of each fin is between 5mm and 50mm.

The present invention will now be described by way of example only with reference to the following figures in which: Figure 1 is a schematic section of a ducted fan gas turbine engine incorporating an exhaust nozzle, which itself comprises noise and/or performance improvement means; Figure 2 is a more detailed schematic perspective view of an exhaust nozzle arrangement of the present invention which comprises noise reducing tabs having fins; Figure 3 is a view looking axially forward toward the rear of the engine and shows the in more detail the tabs having fins ; Figure 4 is a part cutaway schematic view of the core exhaust nozzle of the ducted fan gas turbine engine and exhaust nozzle shown in figures 2 and 3.

With reference to figure 1, which is a schematic section of a ducted fan gas turbine engine incorporating an exhaust nozzle, which itself comprises noise and/or performance improvement means. A ducted fan gas turbine engine 10 comprises, in axial flow series an air intake 5, a propulsive fan 2, a core engine 4 and an exhaust nozzle assembly 16 all disposed about a central engine axis 1.

The core engine 4 comprises, in axial flow series, a series of compressors 6, a combustor 8, and a series of turbines 9. The direction of airflow through the engine 10 in operation is shown by arrow A and the terms upstream and downstream used throughout this description are used with reference to this general flow direction.

Air is drawn in through the air intake 5 and is compressed and accelerated by the fan 2. The air from the

fan 2 is split between a core engine flow and a bypass flow. The core engine flow enters core engine 4, flows through the core engine compressors 6 where it is further compressed, and into the combustor 8 where it is mixed with fuel which is supplied to, and burnt within the combustor 8. Combustion of the fuel mixed with the compressed air from the compressors 6 generates a high energy and velocity gas stream which exits the combustor 8 and flows downstream through the turbines 9. As the high energy gas stream flows through the turbines 9 it rotates turbine rotors extracting energy from the gas stream which is used to drive the fan 2 and compressors 6 via engine shafts 11 which drivingly connect the turbine 9 rotors with the compressors 6 and fan 2. Having flowed through the turbines 9 the high energy gas stream from the combustor 8 still has a significant amount of energy and velocity and it is exhausted, as a core exhaust stream, through the engine exhaust nozzle assembly 16 to provide propulsive thrust. The remainder of the air from, and accelerated by, the fan 2 flows within a bypass duct 7 around the core engine 4. This bypass air flow, which has been accelerated by the fan 2, flows to the exhaust nozzle assembly 16 where it is exhausted, as a bypass exhaust stream to provide further, and in fact the majority of, the useful propulsive thrust.

The velocity of the bypass exhaust stream is significantly lower than that of the core exhaust stream.

Turbulent mixing of the two exhaust streams in the region of, and downstream of, the exhaust nozzle assembly 16, as well as mixing of both streams with the ambient air surrounding and downstream of the exhaust nozzle assembly 16 generates a significant component of the overall noise generated by the engine 10. This noise is known as exhaust or jet noise. Effective mixing and control of the mixing of the two exhaust streams with each other and the ambient air is required in order to reduce the exhaust noise

generated. The mixing and its control is effected by the exhaust nozzle assembly 16.

In the embodiment shown the exhaust nozzle assembly 16 comprises two concentric sections, namely a radially outer bypass exhaust nozzle 12 and an inner core exhaust nozzle 14. The core exhaust nozzle 14 is defined by a generally frusto-conical core nozzle wall 15 (see Figure 2). This defines the outer extent of an annular core exhaust duct 30 through which the core engine flow is exhausted from the core engine 4. The inner extent of the core exhaust duct 30 is defined by an engine plug structure 22.

Referring now to Figure 2 which is a more detailed schematic perspective view of an exhaust nozzle arrangement 16 of the present invention which comprises noise reducing tabs 20 having fins 27. The present invention comprises a plurality of circumferentially spaced tabs 20 extending from the downstream end of the core exhaust nozzle 14 and core nozzle walls 15. As shown the tabs 20 are of a general trapezoidal shape with a first lateral edge 23 of the tabs 20 circumferentially tapering in the rearward or downstream direction. The tabs 20 further comprise a second lateral edge 25, which in this embodiment is substantially parallel to the axis 1. The general trapezoidal shape of the tabs 20 is further defined by a downstream edge 24 and the downstream periphery 29 of the nozzle wall 15.

Disposed to the second edge 25 is a generally inwardly projecting fin 27 which substantially extends along the axial length of the tab 20.

It is the object of the present invention to prevent contra-rotating vortices being generated and shed from a tab 20, which then impinge with one another thereby reducing the effectiveness of each vortex. It is therefore an advantage of providing the tabs 20 with fins 27 to substantially prevent a vortex being generated from the second edge 25. As only one vortex is generated and shed

from the first edge 23, all the vortices shed from all the tabs 20 are rotating in substantially the same rotational direction. Contra-rotating vortices, from one tab 20 or from adjacent tabs 20, have a tendency to merge and in doing so negate one another's rotational momentum and as a consequence reduce each vortex's jet stream mixing ability and noise reducing performance. The present invention thus forms discrete co-rotating vortices where each vortex maintains its integrity for a greater downstream axial length where useful mixing of the exhaust flows occurs and thus the present invention provides improved noise reduction characteristics over the prior art.

Bearing in mind that each vortex causes aerodynamic drag is it a further advantage that the present invention significantly reduces the amount of drag while also improving noise reduction. Although in the embodiment shown, the fin 27 is disposed to the second edge 25 it may equally be disposed to the first edge 23.

Figure 3 is a view looking axially forward toward the downstream end of the engine and shows in more detail the tabs having fins. The same reference numerals refer to the same components for Figures 2 and 3. In operation a core gas stream has a higher velocity and a generally higher static pressure than the bypass which in turn has a higher velocity and a generally higher static pressure gas stream than the ambient. In general tabs 20 generate vortices at their edges as the higher static pressure gas stream spills over the edge of the tab 20 and into the lower static pressure gas stream. The high static pressure gas then mixes with the lower pressure gas along the tapered lateral edge 23 of the tab 20 forming a stream-wise vortex in the process. The fin 27 generally extends in the radially inward direction from the edge 25 of the tab 20. Thus in Figure 3 (and as viewed in relation to the figure) the free or first edge 23 generates an anti-clockwise rotating vortex.

The fin 27, generally extending in the radially inward direction from the edge 25 of the tab 20, prevents the higher static pressure gas spilling over the edge 25 and thus there is a minimal or no vortex generated there. The radial extent of the fin 27 is therefore important as a radially high fin 27 will increase drag and a particularly short radial fin 27 will allow some over-spill of high static pressure gas. It has been estimated for the present embodiment shown in figures 2 and 3 that a suitable radial height range for the fin 27 is 5-50 mm. For the fin 27 to be useful in substantially preventing over-spill it needs to be higher than the depth of the boundary layer.

The maximum height is dictated by the degree of drag the fin 27 produces.

Referring now to Figure 4 which is a part cutaway schematic view of the core exhaust nozzle of the ducted fan gas turbine engine and exhaust nozzle shown in figures 2 and 3. It can be seen that the core nozzle 14 generally converges toward its aft end resulting in the gas stream increasing in velocity, this has the intrinsic effect of reducing the static pressure therethrough. Thus a static pressure PI exists in the duct 30 and downstream toward the aft end of the duct 30 a lower static pressure P2 exists. The static pressures PI and P2 are both generally higher than static pressure P3, which exists in the bypass duct.

Thus the amount of spillage flow around the edge 23 of the tab 20 diminishes when moving along the edge 23 in a downstream direction. Subsequently the shape of the fin 27 may be optimised by tapering its radial extent from the edge 24 of the nozzle 14 to the downstream edge of the tab 20.

Figure 4 also shows in more detail the tab 20 comprising, in this embodiment of the present invention, a high static pressure side 31 which is radially inward of a low static pressure side 32 of the tab. The low static pressure side 32 being so described as it is subject to a

lower static pressure than the high static pressure side 31.

The radial height of the fin 27 will depend on whether it is desired to prevent any vortex from forming or whether to permit a limited vortex forming. It may be beneficial to allow a weak vortex to form at the second lateral edge 25 of the tab 20 so that some additional gas stream mixing will occur over a relatively short axial length (downstream of the periphery 29) before the weak vortex is dissipated.

The vortex shed from the first lateral edge 23 is greater than that from the second edge 25 and therefore penetrates further downstream and enhances the mixing of the gas streams. Thereby, the vortex generated by the first edge is not unduly inhibited by the vortex from the second edge.

Thus it can be appreciated that the fin 27 may be designed to control the vortices generated and shed from the tabs 20 as well as to substantially prevent any vortex from issuing from the edge 25.

A further advantage of the present invention is to provide stiffening for the tab 20. The tab 20 being subject to very high velocity gas streams, particularly the core stream which may have an axial velocity of approximately 400 mus-.

It is not intended that the present invention, described in relation to Figures 2 and 3, only relates to those preferred embodiments. For example, the deployable tabs 20 may be any suitable shape and in particular may also be triangular, and the tabs 20 may be unevenly distributed about the periphery of the nozzle 14.

Similarly, the tabs 20 may be angled both radially inwardly and radially outwardly from the periphery of the nozzle 14.

It is also not intended that the tabs 20 are straight, but may be curved in the plane of the paper on figure 3. Although the preferred embodiment of the present invention is not restricted to a particular angle of the tabs 20 nor their length L.

In certain circumstances it may be preferred for a number of the tabs to comprise fins disposed to the first edge and a number of the tabs to comprise fins disposed to the second edge. For instance, with reference to Figure 3, the tabs 20 on one side of the nozzle 14 have fins 27 disposed on their first edge 23 and on the other side have fins 27 disposed on their second edge 25. The object of this configuration is to provide a symmetrical design and to balance out inherent forces generated by the tabs 20.

The present invention may be also configured with respect to two co-pending applications of the present assignee, which are concerned with having deployable noise reduction means. The term deployable meaning that the noise reduction means, in a first position, may be exposed to the gas stream (s) and be operable as noise reduction means as herein described and in a second position or suitable arrangement may be stowed or not exposed to the gas stream (s) and therefore not operable as a noise reduction means.

Furthermore in yet another embodiment of the present invention a bypass exhaust nozzle using tabs as described above can be used in conjunction with a conventional forced lobed type core exhaust nozzle/mixer.

Although the invention has been described and shown with reference to a short cowl type engine arrangement in which the bypass duct 28 and bypass exhaust nozzle 12 terminate upstream of the core exhaust duct 30 and nozzle 14, the invention may also be applied, in other embodiments, to long cowl type engine arrangements in which the bypass duct 28 and bypass exhaust nozzle 12 terminate downstream of the core exhaust duct 20 and nozzle 14. The invention however is particularly beneficial to short cowl arrangements since with such arrangements conventional noise suppression treatments of the exhaust are not practical in particular where high bypass ratios are also used.

The invention is also not limited to ducted fan gas turbine engines 10 with which in this embodiment it has been described and to which the invention is particularly suited. In other embodiments it can be applied to other gas turbine engine arrangements in which either two exhaust streams, one exhaust stream or any number of exhaust streams are exhausted from the engine though an exhaust nozzle (s).

Claims

1. A gas turbine engine exhaust nozzle arrangement for the flow of exhaust gases therethrough between an upstream end thereof and a downstream end thereof comprising a nozzle wall and a plurality of tabs, the nozzle wall having a downstream periphery, the tabs extend in a generally downstream direction from the downstream periphery wherein each tab comprises a fin.

2. A gas turbine engine exhaust nozzle arrangement as claimed in claim 1 wherein the nozzle wall is substantially frusto-conical and the plurality of tabs is circumferentially disposed to the downstream periphery.

3. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-2 wherein each tab comprises lateral edges, the fin disposed to a lateral edge of each tab.

4. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-3 wherein each tab comprises a high pressure side and a low pressure side, the fin generally extending radially from the high pressure side of each tab.

5. A gas turbine engine exhaust nozzle arrangement as claimed in claim 4 wherein the high pressure side is radially inward of the low pressure side and the fin generally extends radially inwardly from a lateral edge of each tab.

6. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-5 wherein the fin generally radially tapers from the periphery of the nozzle wall to the downstream edge of each tab.

7. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-6 wherein the fin prevents a vortex forming at and being shed from a lateral edge of each tab.

8. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-6 wherein the fin substantially reduces the strength of a vortex forming at and being shed from a lateral edge of each tab.

9. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-6 wherein the fin is so arranged to control the strength of a vortex being generated and shed from a lateral edge of each tab.

10. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-8 wherein a number of the tabs comprise fins disposed to the first edge and a number of the tabs comprise fins disposed to the second edge.

11. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-9 wherein the radial height of each fin is between 5mm and 50mm.

12. A gas turbine engine exhaust nozzle arrangement as hereinbefore described and with reference to figures 1 to 4.

13. A ducted fan gas turbine engine as hereinbefore described and with reference to figures 1 to 4.