GB2372779A

GB2372779A - Gas turbine engine nozzle with noise reducing tabs

Info

Publication number: GB2372779A
Application number: GB0105351A
Authority: GB
Inventors: Martyn Richards
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: GB0105351D0

Abstract

A gas turbine engine exhaust nozzle arrangement comprises a nozzle 14, and a plurality of noise-reducing tabs 20, The tabs 20, extend from a downstream periphery of the nozzle 14. Various means are provided for changing the operational condition of the tabs 20, according to whether or not noise reduction is required. The tabs may simply be retractable within the nozzle 14, (figure 4),or projections 34 may be provided on a rotatable (figures 5 and 6) or axially moveable (figure 7) member the projections thereby being movable to and from positions in which they cover/overlap the notches 21, between tabs 20. Movement may be effected by pinion and rack arrangements (figures 8-10) or actuators acting on a pivotal angled arm (figures 11 and 12).

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. 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 due to 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 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.

An alternative nozzle design that is directed to reducing exhaust noise is proposed and described in GB 2,289, 921. In this design, a number of circumferentially spaced notches, of various specified configurations, sizes, spacing and shapes, are provided in the downstream periphery of a generally circular core exhaust nozzle. Such a nozzle design is considerably simpler to manufacture than the conventional lobed designs. This prior proposal describes that the notches 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.

Model testing of nozzles similar to those described in GB 2,289, 921 has shown that significant noise reduction and suppression can be achieved. However the parameters and details of the design proposed in GB 2,289, 921 are not optimal and there is a continual desire to improve the nozzle design further.

A further design, and that of the present Assignee, is proposed in UK Application GB 0025727.9, which claims

priority from UK Application GB 9925193. 6. These applications disclose trapezoidal shaped tabs disposed to the axially rearward exhaust ducts of the bypass and core and which are inclined radially inward to impart vortices to the exhaust streams.

However, the main requirement of reducing exhaust noise is during aircraft take-off and landing. At higher altitudes where the majority of the duration of the flight is, exhaust noise is not a problem. It is therefore not necessary to have noise reduction means operational at higher altitudes especially when one considers the noise reduction means introduces aerodynamic inefficiencies.

Furthermore, gas turbine engine design has moved away from low bypass ratios to high bypass ratios in order to reduce the exhaust gas stream velocities and thereby also reduce the exhaust noise. Although high bypass ratio engines are generally more efficient than low bypass engines, as the relative air speed increases the lower relative velocity exhaust gas stream means that the aircraft must cruise at a lower speed to maintain the necessary thrust from the high bypass engine. It would therefore be an advantage for a high bypass ratio gas turbine engine to provide the advantages of a lower relative velocity exhaust gas stream at take-off and landing and a higher relative velocity exhaust gas stream at cruise.

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 and a downstream end thereof comprising a nozzle and a plurality of tabs, the tabs extend in a generally downstream direction from a

downstream periphery of the nozzle wherein the nozzle comprises an actuation mechanism for moving the tabs between a first deployed position and a second non-deployed position. Preferably the tabs are circumferentially disposed about the nozzle.

Preferably the nozzle comprises a radially inner wall and a radially outer wall, the actuation mechanism being generally disposed between the radially inner and outer walls.

Preferably the actuation mechanism comprises an actuator and a movable member, the actuator working in operative association with the movable member to move the movable member between a first position, where the tabs are deployed, and a second position, where the tabs are not deployed. Preferably the movable member is circumferentially segmented.

Preferably the movable member is any one of a group comprising a sleeve and a cuff. Preferably the movable member comprises the tabs. Alternatively the movable member comprises a plurality of projections.

Alternatively the nozzle wall comprises the tabs furthermore the nozzle wall may comprise the projections.

Preferably the movable member is rotatable relative to a main engine axis. Alternatively the movable member is axially translatable relative to the main engine axis and the movable member may also be rotatable and axially translatable relative to a main engine axis.

Alternatively the inner wall comprises the cuff, the cuff comprising the tabs, the actuator and the cuff working in operative association to axially translate the cuff from a first position to a second position, the first position being axially rearward of the second position and where the tabs extend beyond the rearward edge of the outer wall and into a gas stream. Preferably the cuff comprises an overlap, the overlap configured to provide support for the cuff and a smooth airwash surface between the axially

forward portion of the cuff and the remainder of the inner wall.

Alternatively the inner wall comprises the cuff and the outer wall comprises the projections, the cuff comprising the tabs, the actuator and the cuff working in operative association to rotate the cuff from the first position to the second position, the first position comprising the tabs and the projections being generally circumferentially aligned and in the second position the tabs being generally circumferentially aligned with the spaces between the projections.

Alternatively the actuation mechanism further comprises a sleeve, the sleeve disposed between the inner wall and the outer wall, the sleeve arranged to work in operative association with the actuator.

Preferably the actuator comprises any one of the group comprising an electric motor, a magnetic actuator, a hydraulic actuator, a pneumatic actuator Preferably the actuator comprises an actuation arm, the actuator capable of translating the actuator arm along an actuator axis.

Alternatively the actuator comprises an actuation arm, the actuator capable of rotatably driving the actuator arm about an actuator axis and the actuator arm is disposed to the movable member.

Preferably the actuation mechanism further comprises a pinion, and a rack, the actuator arm at one end connected to and driven by the actuator and the pinion attached to the other end of the actuator arm, the pinion engaging the rack, the rack being disposed to the movable member, so that in operation the actuator moves the movable member between the first position where the tabs are deployed and the second position where the tabs are not deployed.

Furthermore the actuation mechanism further comprises an angled member, rotatably mounted, connected at one end to the actuator arm and at another end to the movable member,

in use the actuator translates the actuator arm thereby rotating the angled member about the rotatable mounting and thereby moving the movable member between the first position where the tabs are deployed and the second position where the tabs are not deployed.

Preferably the sleeve is mounted on a guide track, the guide track itself mounted on the nozzle wall and the guide track is of a dovetail joint arrangement.

Preferably the sleeve extends between the downstream end of the outer wall and the downstream end of the inner wall, thereby providing a smooth nozzle wall for the gas stream. Preferably the sleeve comprises an inflection, the purpose of the inflection being to enable the sleeve to extend between the downstream end of the outer wall and the downstream end of the inner wall, and thereby provide a smooth nozzle wall for the gas stream, in both the first position and the second position.

Preferably the radially inner wall comprises the cuff but alternatively the radially outer wall comprises the cuff.

Preferably the tabs circumferentially taper in the downstream direction and the tabs are radially inwardly angled at an angle of up to 200 relative to the nozzle wall.

Alternatively the tabs are radially outwardy angled at an angle of up to 200 relative to the nozzle wall.

Furthermore the tabs may be circumferentially alternatively radially inwardly angled at an angle of up to 200 relative to the nozzle wall and radially outwardly angled at an angle of up to 200 relative to the nozzle wall.

Preferably the tabs are of a substantially trapezoidal shape alternatively the general shape of the tabs is any one of the group comprising rectangular, square, triangular shape.

Preferably the tabs are circumferentially disposed about the periphery of the nozzle wall to define substantially trapezoidally shaped notches between adjacent tabs. Alternatively the tabs are circumferentially disposed about the periphery of the nozzle wall to define substantially V-shaped notches between adjacent tabs.

Alternatively the edges of the tabs are curved.

Preferably the nozzle tabs are radially inwardly angled at an angle of up to 100 relative to the nozzle wall.

Preferably the exhaust nozzle is a core engine nozzle but may also be a bypass exhaust nozzle. Alternatively the arrangement comprises a core exhaust nozzle and a bypass exhaust nozzle.

Alternatively exhaust nozzle arrangement comprises an outer bypass exhaust nozzle as herein described and an inner core exhaust nozzle of a lobed mixer type.

Preferably the downstream end of the bypass nozzle is further downstream than the downstream periphery of the core exhaust nozzle but alternatively the downstream end of the bypass nozzle is upstream of the downstream periphery of the core exhaust nozzle.

Preferably the arrangement is configured for exhaust noise attenuation, but alternatively the arrangement is configured for reducing the cross sectional area of the exhaust nozzle.

Preferably the tabs, in operation, are deployed substantially prior to the aircraft reaching a predetermined altitude and the tabs, in operation, are not deployed substantially after the aircraft reaches a predetermined altitude.

A further aspect of the present invention is method of operating an aircraft having a gas turbine engine comprising an exhaust nozzle arrangement as claimed in any preceding claim wherein the method comprises the steps of: deploying noise reduction means prior to take-off; not

deploying noise reduction means above a predetermined aircraft altitude and ; deploying the noise reduction means below the predetermined aircraft altitude.

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 the exhaust nozzle of a ducted fan gas turbine engine of UK Application GB 0025727. 9, which claims priority from UK Application GB 9925913.6 ; Figure 3 is a part cutaway schematic view of the core exhaust nozzle of the ducted fan gas turbine engine and exhaust nozzle shown in figures 1 and 2.

Figure 4 shows a schematic of a first embodiment of the present invention comprising deployable nozzle tabs in a second non-deployed position.

Figure 5 shows a schematic of a second embodiment of the present invention comprising deployable nozzle tabs in a second non-deployed position.

Figure 6 shows a schematic of the second embodiment of the present invention comprising deployable nozzle tabs in a first deployed position.

Figure 7 shows a schematic of a third embodiment of the present invention comprising deployable nozzle tabs in a first deployed position.

Figure 8 shows an axial section of the inner core nozzle wall comprising a first actuation mechanism suitable for operative association with the first and third embodiment of the present invention.

Figure 9 is a view on A of Figure 8, showing an axial section of the inner core nozzle wall comprising a variation of the first actuation mechanism suitable for

operative association with the first and third embodiment of the present invention Figure 10 is a view on A of Figure 8, showing the actuation mechanism of the first embodiment of the present invention.

Figure 11 shows an axial section of the inner core nozzle wall comprising a second actuation mechanism suitable for operative association with the second embodiment of the present invention.

Figure 12 which is a view on B of Figure 11, showing the second actuation mechanism of the present invention.

Figure 13 is an axial section through the core nozzle wall showing a third actuation mechanism suitable for operative association with second embodiment of the present invention.

Figure 14 is a view on arrow C of Figure 13, showing the third actuation mechanism of the second embodiment of the present invention.

Figure 15 is an axial section through the core nozzle wall showing a fourth actuation mechanism suitable for operative association with first, second and third embodiments of the present invention.

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 4 flow and a bypass flow. The core engine 4 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 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 large component of the 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 with the ambient air is required in order to reduce 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. 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.

Figure 2 is a more detailed schematic perspective view of the exhaust nozzle of a ducted fan gas turbine engine of UK Application GB 9925913.6. The present invention relates to an exhaust nozzle of a gas turbine engine having deployable noise reduction means. By way of example, and as a preferred embodiment, the present invention is herein described with reference to the fixed noise reduction means of UK Application GB 9925913.6. It is therefore intended that in the first deployed position the present invention has the advantages described in UK Application GB 9925913.6.

Figure 2 shows 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 trapezoidal shape with the sides of the tabs 20 circumferentially tapering towards each other in the downstream direction. The tabs 20 are evenly and circumferentially disposed so that a notch 21 or space is defined by and between adjacent tabs 20. The notches 21 are complimentary to the shape of the tab 20 and accordingly are of a trapezoidal shape on the core nozzle 14, with the notches 21 circumferentially opening out in a downstream direction.

The number of tabs 20, and so notches 21 defined in the core exhaust nozzle 14 and also bypass exhaust nozzle

12 (described below), the width of the notches 21, angle of the notches 21, width of notch 21, angular offset between notches 21, and angular gap between notches 21 are all essentially the same and within the same ranges as described in GB 2,289, 921.

Referring to figure 3, which is a part cutaway schematic view of the core exhaust nozzle 14 of the ducted fan gas turbine engine 10 and exhaust nozzle assembly 16 and showing a preferred embodiment of a deployed arrangement of tabs 20. For simplicity, Figure 3 and the description hereafter relates only to the present invention applied to the core exhaust nozzle 14, however the present invention may equally be applied to the bypass nozzle 12.

The tabs 20 of the core exhaust nozzle 14 are radially inwardly angled so that the tabs 20 impinge into the core duct 30 (relative to an extended line 24, shown in figure 3, of the profile of the core nozzle wall 15 immediately upstream of the tabs 20) and are, in operation, incident on the core exhaust flow which is exhausted through the core exhaust nozzle 14. The angle of incidence P of the tabs 20 is defined relative to an extended line 24 of the profile of the core exhaust nozzle wall 15 immediately upstream of the tabs 20. The profile of the core nozzle wall 15 immediately upstream of the tabs 20 itself is at an angle a (typically between 100 and 200) to the engine axis 1.

During experimentation it has been found that the angle of incidence has a effect on noise suppression. As the angle of incidence is increased up to 200 the noise reductions are improved. However at angles of incidence above 200 there is little further improvement in noise suppression. Furthermore at these higher angles of incidence aerodynamic losses due to the effect the tabs 20 have on the core exhaust flow increase. Therefore preferably the tabs 20 are angled at angles of incidence up to 10 .

The tabs 20 and angling of the tabs 20 reduces the mid and low frequency noise generated by the exhaust and engine 10, typically in the frequency range 50-500 kHz. It does however, in some cases increase the noise generated at higher frequencies. The noise at low and mid frequencies though is the most critical in terms of the perceived noise level and the higher frequency noise is masked by noise generated from elsewhere in the engine 10. Therefore overall the tabs 20 provide a reduction in the perceived exhaust noise generated. The increase in high frequency noise sometimes associated with the angled tabs 20 at higher angles of incidence is a further reason why the tabs 20 are preferably angled at angles of incidence up to 10 .

The tabs 20 induce stream-wise vortices in the exhaust flow through and around the nozzle 14. These vortices are generated and shed from the sides of the tabs 20 and increase the local turbulence levels in a shear layer that develops between the core and bypass exhaust streams downstream of the exhaust nozzle assembly 16. This vorticity and turbulence increases and controls the rate of mixing between the core exhaust stream, bypass exhaust stream, and the ambient air. This reduces the velocities downstream of the exhaust assembly 16, as compared to a conventional nozzle, and so reduces the mid to low frequency noise generated by the exhaust streams. The increased turbulence generated by the tabs 20 in the initial part of the shear layers immediately downstream of the exhaust nozzle assembly 16 causes an increase in the high frequency noise generated. Having tabs 20 angled radially inwards increases the strength of the vortices produced and so improves the reduction in perceived noise. However the angle of incidence of the tabs 20 must not be too large since this can induce flow separation which will generate, rather than reduce the noise as well as adversely affecting aerodynamic performance of the nozzle 14.

The bypass exhaust nozzle 12 is also defined by a generally frusto-conical bypass nozzle wall 17 which is concentric with and disposed radially outwardly of the core exhaust nozzle 14. The bypass nozzle wall 17 defines the outer extent of an annular bypass exhaust duct 28 through which the bypass engine flow is exhausted from the engine 10. The inner extent of the bypass exhaust duct 28 is defined by an outer wall of the core engine 4. The bypass nozzle 12 is similar to the core exhaust nozzle 14 and a plurality of circumferentially spaced tabs 18 extend from the downstream end of the bypass exhaust nozzle 12 and bypass nozzle walls 17. As with the core nozzle 14, the tabs 18 are of a trapezoidal shape with the sides of the tabs 18 circumferentially tapering in the downstream direction. The tabs 18 are evenly circumferentially disposed so that a V-shaped notch 19 or space is defined by and between adjacent tabs 18. The bypass nozzle tabs 18 affect the bypass exhaust flow and noise generated in a similar way to the core exhaust nozzle tabs 20.

The tabs 20 should have a length L sufficient to generate the required stream-wise vortices as described below and GB 2,289, 921 specifies that the tabs 18,20 must have a length L of between 5% to 50% of the nozzle diameter Dc, Db. It has been found however that using long tabs, towards the 50% end of the range given, induces excessive aerodynamic losses which adversely affect the performance particularly when they are angled. Accordingly it has been determined that the core tabs 20 should have a length L of approximately 5%-10% of the core exhaust nozzle diameter Dc, whilst the bypass tabs 18 should have a length L of approximately 2. 5%-5% of the bypass exhaust nozzle diameter Db. The bypass tabs 18 have a smaller percentage length since the bypass provides more of the propulsive thrust of the engine and so any performance loss on the bypass will have a greater effect on the overall engine performance.

In addition although the percentage size is less, since the

bypass is of a greater diameter than the core the actual physical size of the core tabs 20 and bypass tabs 20 are not so different.

In model tests of the exhaust nozzle assembly 16 shown in figure 2 and described above a 5dB reduction in the peak sound pressure level over a conventional plain frusto conical nozzle arrangement has been achieved. It has also been found that the noise reductions provided by using tabs 18 on the bypass exhaust nozzle 12 and by using tabs 20 on the core exhaust nozzle 14 can be cumulative. It will therefore be appreciated that in other embodiments tabs 18,20 can be used on the bypass exhaust nozzle 12 or the core exhaust nozzle 14 alone to give some improved degree of noise suppression. The core exhaust nozzle tabs 20 and the bypass exhaust nozzle tabs 18 can also be angled at different angles of incidence ss.

Whereas in Figure 3 and the description thereof is related to a preferred embodiment it is not intended that the present invention only relates to that preferred embodiment. 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 must be straight, but may be curved in the plane of the paper on figure 3. Although the preferred embodiment the present invention is not restricted to a particular angle of the tabs 20 nor their length L.

The present invention is however, 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 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.

Figure 4 shows a schematic of a first embodiment the present invention comprising axially deployable nozzle tabs 20, in the second, non-deployed or stowed position. It is an object of the present invention to provide either the bypass nozzle 12 or the inner core exhaust nozzle 14 or both with noise reduction means 20. The main requirement of reducing exhaust noise is during aircraft take-off and landing procedures. At higher altitudes where often the majority of the duration of the flight is, exhaust noise is not a problem. It is therefore not necessary to operate noise reduction means 20 at higher altitudes especially when one considers the noise reduction means 20 introduces aerodynamic inefficiencies. Thus, in Figure 4, the dashed lines indicate tabs 20 in a second non-deployed position, where they have been retracted into the nozzle 14 and when they are not required for exhaust noise reduction.

The tabs 20 themselves may be retracted in an axially forward direction or rotated circumferentially and axially forward depending on the actuation mechanism employed.

Suitable actuation mechanisms are described herein with reference to Figure 8,9, 10 which describe a first actuation mechanism and a variation thereof and Figure 15 which shows a fourth actuation mechanism. These actuation mechanisms are merely one way of carrying out the present invention and are not intended to be a limitation thereof.

The tabs 20 shown in Figure 4, when fully deployed in a first position, are substantially configured as shown in Figure 2 and therefore posses the same advantages of the fixed nozzle tabs of UK Application GB 0025727.9. When retracted, the tabs 20 expose the rearward edge 27 of the inner core nozzle 14 and thereby the gas stream may then exit the over the edge 27, which is designed to give the best possible aerodynamic efficiency.

The tabs 20 are attached to a movable member 32 which in this embodiment is a sleeve 29 and is shown by the dashed lines. The sleeve 29 is configured to work in operative association with actuation means as shown in Figures 8,9, 10 and 15. The sleeve 29 may be completely annular, or circumferentially segmented (not shown), each segment having its own actuation means. In a preferred embodiment of the present invention a circumferentially segmented sleeve 29 is advantageous where the tabs 20 are translated axially, as changes in circumferential length may be accommodated by adjacent sleeve 29 segments overlapping when deployed. Alternatively, the segmented sleeve 29 may be configured to have circumferential spaces when they are retracted into the second non-deployed position. In the first deployed position the tabs 20 and sleeve 29 are radially innermost and accommodate the least circumferential length, whereas when retracted the tabs 20 and sleeve 29 are radially outermost and accommodate a greater circumferential length.

Referring now to Figure 5, which shows a schematic of a second embodiment of the present invention comprising deployable nozzle tabs 20 in a non deployed second position. The tabs 20 are fixed to the rearward edge 27 of the inner core nozzle 14 and are therefore not translatable in any manner. A moveable member 32 is disposed to the downstream end of the core nozzle 14 and is substantially annular comprising a substantially annular portion or sleeve 29 and generally downstream extending projections 34, which are themselves disposed to the sleeve. However, much in the same way as in the first embodiment of the present invention the movable member 32 may be circumferentially segmented.

In the second non-deployed position the projections 34 are so arranged and configured to substantially fill or cover the notch 21 or space between the tabs 20. Although the projections 34 are shown radially outward of the tabs

20, which is preferable, the movable member 32, including projections 34, may be disposed radially inward of the tabs 20.

In this first position the projections 34 and tabs 20 co-operate to define a substantially smooth outlet edge 35, which is aerodynamically more efficient than the rearward edge 27 having exposed and deployed tabs 20. This first position is particularly beneficial at high altitude, where the aircraft is at cruise, as the tabs 20 are not exposed and therefore do not generate and shed the noise reducing vortices which inherently invoke unnecessary aerodynamic losses.

A further advantage of this second embodiment of the present invention, when employed particularly on the bypass nozzle 12 is that the cross sectional area of the bypass nozzle 12 is reduced when the noise reduction means 20 are not deployed. This has the effect of increasing the velocity of the exhausted gas stream relative to the ambient air. Therefore for substantially the same engine specific fuel burn the aircraft may travel at a greater velocity or more efficiently for a given velocity. It should be appreciated that the extent of area change may be optimised when considering exhaust noise, engine performance and the operability considerations of the low pressure fan 2.

Figure 6 shows a schematic of the second embodiment of the present invention comprising deployable nozzle tabs 20 in a deployed first position. The moveable member 32 having been rotated about the main engine axis 1 (Figure 1) to the second position, thereby deploying the tabs 20, where the projections 34 are so arranged and configured to circumferentially align with the tabs 20. As previously described hereinbefore, stream-wise vortices are generated and shed from the sides of the tabs 20 which are now exposed to the bypass gas stream. A suitable and second actuation mechanism is shown and described with reference

to Figures 11 and 12, while a further and third actuation mechanism is shown and described with reference to Figures 13 and 14. It is not intended to limit the present invention exclusively to the actuation mechanisms herein described which are merely specific embodiments of a number of ways of carrying out the present invention. Other actuation mechanisms may readily be substituted but which are intended to be within the scope of the present invention.

Figure 7 shows a schematic of a third embodiment of the present invention comprising deployable nozzle tabs 20 in a deployed second position. The figure shows the movable member 32, comprising the projections 34, having been axially translated relative to the main engine axis 1 (Figure 1), from the second non-deployed position as shown in Figure 5, to the first position, thereby deploying the tabs 20. Although it is preferable for the movable member 32 to comprise projections 34 to obstruct the notches 21 or spaces between the tabs 20, a substantially annular downstream periphery to the movable member 32 may be used instead. This may have its own advantages in providing a smoother more aerodynamic outlet edge 35 (Figure 5) for the gas stream exiting the nozzle 14, than the outlet edge 35 created by the projections 34 and tabs 20 when in the second position.

A suitable and first actuating mechanism is shown and described with reference to Figures 8,9, 10 and 15. It is not intended to limit the present invention exclusively to the actuation mechanisms herein described which are merely embodiments of a number many ways of carrying out the present invention. Other actuation mechanisms may readily be substituted but which are intended to be within the scope of the present invention.

Figure 8 shows an axial section of the inner core nozzle wall 14 comprising a first actuation mechanism 39 suitable for operative association with the first and third

embodiment of the present invention. The actuation mechanism 39 comprises an actuator 40 secured to the generally frusto-conical exhaust nozzle wall 15 by mounting means 41. The actuator 40 comprises an actuator axis 43 along which is substantially aligned an actuator arm 48 (Figure 9) having a pinion 45 at its distal end. The pinion 45 engages a rack 44, which forms part of the sleeve 29. The. rack 44 is held in a location track 42 comprising interlocking dovetail root 46 and slot 47 portions. Other types of location and guide means may be utilised without departing from the present invention.

The sleeve 29 is configured as shown in Figure 8 so that it forms an aerodynamic outer profile with the radially outer wall 37 of the inner core exhaust nozzle 14 and the nozzle wall 15. In this embodiment of the present invention it is preferred to axially deploy the tabs 20 at a constant pre-determined angle, as shown in Figure 3, thus the location track 42 is also angled to provide the necessary angle of the tabs 20.

Figure 9 shows an axial section of the inner core nozzle wall 14 comprising a variation of the first actuation mechanism 39 suitable for operative association with the first and third embodiment of the present invention. Reference numerals of Figure 8 are used for like parts in Figure 9. In this variation it is intended to introduce the tabs 20 to the core exhaust gas stream in an arcuate manner. This is advantageous as the effective angle of the tabs 20 and length L, may be readily adjusted and the noise suppression optimised. Again it is also desired to configure the sleeve 29 so that it forms an efficient aerodynamic transition between the radially outer wall 37 and the edge 27. The location track 42 is configured to substantially follow the curvature of the sleeve 29.

Referring to figure 10 which is a view on A of Figure 8, showing the actuation mechanism of the first embodiment

of the present invention. The view is also substantially that of Figure 9. The actuator 40 rotatingly drives the actuator arm 48 and pinion 45 thereby translating the rack 44 and sleeve 29, thus deploying the tabs 20.

The actuator 40 is preferably an electric motor although any actuation means suitable to drive the rack 44 and pinion 45 is intended to be within the scope of the present invention. Similarly the rack 44 and pinion 45 may be substituted for any other suitable means for translating the sleeve 29 and thereby deploying the tabs 20.

Figure 11 shows an axial section of the inner core nozzle wall comprising a second actuation mechanism 49 suitable for operative association with the second embodiment of the present invention. An actuator 50, mounted on the nozzle wall 15, is connected to an angled member 54 at one end via an actuation arm 52. The angled member 54, disposed to the nozzle wall 15 through a rotatable mounting 60, is connected to a rigid section 56 integral with the movable member 32. The movable member 32 also comprises the projections 34. The actuation arm 52 has a central axis 62, which in this embodiment is substantially aligned with the central engine axis 1, along which the arm 52 is translated by the actuator 50.

Referring to figure 12 which is a view on B of Figure 11, showing the second actuation mechanism 49 of the second embodiment of the present invention. The angled member 54 is in a first position, the position in which the projections 34 cover the notches 21 or spaces and thus in Figure 5 the tabs 20 are concealed and thus ineffective for reducing jet noise. When required the actuator 50 axially translates the arm 52, thereby rotating the angled member 54 about the rotatable mounting 60 to the dashed outline shown as 54'. Thus the movable member 32 is rotated and the projections 34 circumferentially align with the tabs 20 as shown on Figure 6. As can be seen in Figure 9 the angled member 54 is slidably connected to both the actuator

arm 52 and the circumferential member 56 via co-operating pins 66 and slots 64 defined in the angled member 54. The slots 66 are designed so as to accommodate the arcuate displacement of the angled member 54.

The actuator 50 is preferably a hydraulic actuator although alternatively a pneumatic, magnetic or electric actuator may be used without departing from the scope of the present invention. A shape memory metal may also be used to provide the actuation means 50. Typical shape memory metal actuators are commonly known in the art and are therefore not discussed in detail herein.

Figure 13 is an axial section through the core nozzle showing a third actuation mechanism 70 suitable for operative association with second embodiment of the present invention. The third actuation mechanism 70 enables the sleeve 29 to be rotated generally about the main engine axis 1 (see Figure 1). The third actuation mechanism 70 comprises an actuator 72 securely mounted to the nozzle wall 15 by conventional means, an actuator arm 82 having a pinion 80 at its distal end. The pinion 80 engages a rack 78 disposed to the sleeve 29, thus rotation of the pinion 80 translates the projections 34 in the circumferential direction generally about the main engine axis 1 (see Figure 1).

The sleeve 29 is secured to the nozzle wall 15 by means 74,76 to allow circumferential displacement and restraint in the axial and radial directions. The securing means 74,76 may be a circumferential track 74 and may further comprise a dovetail fixture 76.

Referring now to Figure 14 which is a view on arrow C of Figure 14, showing the third actuation mechanism 70 of the second embodiment of the present invention. Figure 14 shows the tabs 20 in the second non-deployed position (in accordance with Figure 5), the projections 34 occupying the spaces 21 between the tabs 20. In this second position the projections 34 and tabs 20 provide an aerodynamically

smooth annular nozzle outlet edge 35 for the gas stream. The projections 34, disposed to the axially rearward end of the sleeve 29, may be rotated to axially align with the tabs 20 (in accordance with Figure 6), thereby exposing the tabs 20 to the core exhaust gas stream, thus partially attenuating the jet noise.

Figure 15 is an axial section through the core nozzle showing a fourth actuation mechanism 90 suitable for operative association with second embodiment of the present invention. Whereas the foregoing embodiments show and describe the present invention generally comprising an inner wall 15, a radially outer wall 37 and a sleeve 29, the sleeve 29 being either axially translatable or rotatable or a combination of both, the fourth embodiment does not comprise a sleeve 29. The fourth actuation mechanism 90 comprises the radially inner nozzle wall 15, having a cuff 94, actuator 98 mounted on the wall 15 and actuator arm 92. The cuff 94 comprises an overlap 96 overlapping the wall 15 and a web 97 which co-operates with the actuator arm 92. The cuff 94 further comprises the tabs 20. This fourth actuator mechanism 90 is configured to be associated to the operation of the first and third embodiment of the present invention.

Figure 15 shows the cuff 94 and thus tabs 20 in the second non-deployed position. In operation, activation of the actuator 98 extends the actuator arm 92 generally rearward, with respect to the orientation of the engine, and in turn this axially translates the cuff 94 rearward and beyond the rearward edge of the radially outer wall 37, thereby deploying the tabs 20. When the tabs 20 are deployed the overlap 96 maintains a relatively smooth airwashed surface. In an alternative arrangement of the fourth actuation mechanism the radially outer wall 37 comprises the cuff-portion with operation of this embodiment of the present invention in accordance with the above description.

A further advantage of the present invention is that the degree to which the tabs 18, 20 are extended into the gas streams may be optimised easily during testing and evaluation. Furthermore the tabs 18,20 may be deployed to varying extents during the flight cycle of the host aircraft and thereby attenuate different noise frequencies.

Furthermore in yet another embodiment of the 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. Such an arrangement has also been tested and has shown improved noise suppression over an exhaust assembly which uses a lobed type core nozzle/mixer with a conventional bypass exhaust nozzle.

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 by pass 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 and a downstream end thereof comprising a nozzle and a plurality of tabs, the tabs extend in a generally downstream direction from a downstream periphery of the nozzle wherein the nozzle comprises an actuation mechanism for moving the tabs between a first deployed position and a second non-deployed position.

2. A gas turbine engine exhaust nozzle arrangement as claimed in claim 1 wherein the tabs are circumferentially disposed about the nozzle.

3. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-2 wherein the nozzle comprises a radially inner wall and a radially outer wall, the actuation mechanism being generally disposed between the radially inner and outer walls.

4. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-3 wherein the actuation mechanism comprises an actuator and a movable member, the actuator working in operative association with the movable member to move the movable member between a first position, where the tabs are deployed, and a second position, where the tabs are not deployed.

5. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-4 wherein the movable member is circumferentially segmented.

6. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-5 wherein the movable member is any one of a group comprising a sleeve and a cuff.

7. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-6 wherein the movable member comprises the tabs.

8. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-6 wherein the movable member comprises a plurality of projections.

9. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-6,8 wherein the nozzle wall comprises the tabs.

10. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-7 wherein the nozzle wall comprises the projections.

11. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-10 wherein the movable member is rotatable relative to a main engine axis.

12. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-10 wherein the movable member is axially translatable relative to the main engine axis.

13. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-10 wherein the movable member is rotatable and axially translatable relative to a main engine axis.

14. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-13 wherein the inner wall comprises the cuff, the cuff comprising the tabs, the actuator and the cuff working in operative association to axially translate the cuff from a first position to a second position, the first position being axially rearward of the second position and where the tabs extend beyond the rearward edge of the outer wall and into a gas stream.

15. A gas turbine engine exhaust nozzle arrangement as claimed in claim 14 wherein the cuff comprises an overlap, the overlap configured to provide support for the cuff and a smooth airwash surface between the axially forward portion of the cuff and the remainder of the inner wall.

16. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-15 wherein the inner wall comprises the cuff and the outer wall comprises the

projections, the cuff comprising the tabs, the actuator and the cuff working in operative association to rotate the cuff from the first position to the second position, the first position comprising the tabs and the projections being generally circumferentially aligned and in the second position the tabs being generally circumferentially aligned with the spaces between the projections.

17. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-13 wherein the actuation mechanism further comprises a sleeve, the sleeve disposed between the inner wall and the outer wall, the sleeve arranged to work in operative association with the actuator.

18. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-17 wherein the actuator comprises any one of the group comprising an electric motor, a magnetic actuator, a hydraulic actuator and a pneumatic actuator.

19. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-18 wherein the actuator comprises an actuation arm, the actuator capable of translating the actuator arm along an actuator axis.

20. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-18 wherein the actuator comprises an actuation arm, the actuator capable of rotatably driving the actuator arm about an actuator axis.

21. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-20 wherein the actuation mechanism further comprises an actuator arm, the actuator arm disposed to the movable member.

22. A gas turbine engine exhaust nozzle arrangement as claimed in claim 20 wherein the actuation mechanism further comprises a pinion, and a rack, the actuator arm at one end connected to and driven by the actuator and the pinion attached to the other end of the actuator arm, the pinion engaging the rack, the rack being disposed to the movable

member, so that in operation the actuator moves the movable member between the first position where the tabs are deployed and the second position where the tabs are not deployed.

23. A gas turbine engine exhaust nozzle arrangement as claimed in claim 19 wherein the actuation mechanism further comprises an angled member, rotatably mounted, connected at one end to the actuator arm and at another end to the movable member, in use the actuator translates the actuator arm thereby rotating the angled member about the rotatable mounting and thereby moving the movable member between the first position where the tabs are deployed and the second position where the tabs are not deployed.

24. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-23 wherein the sleeve is mounted on a guide track, the guide track itself mounted on the nozzle wall.

25. A gas turbine engine exhaust nozzle arrangement as claimed in claim 24 wherein the guide track is of a dovetail joint arrangement.

26. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-25 wherein the sleeve extends between the downstream end of the outer wall and the downstream end of the inner wall, thereby providing a smooth nozzle wall for the gas stream.

27. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-26 wherein the sleeve comprises an inflection, the purpose of the inflection being to enable the sleeve to extend between the downstream end of the outer wall and the downstream end of the inner wall, and thereby provide a smooth nozzle wall for the gas stream, in both the first position and the second position.

28. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-27 wherein the radially inner wall comprises the cuff.

29. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-27 wherein the radially outer wall comprises the cuff.

30. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-29 wherein the tabs circumferentially taper in the downstream direction.

31. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-30 wherein the tabs are radially inwardly angled at an angle of up to 20 relative to the nozzle wall.

32. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-30 wherein the tabs are radially outwardy angled at an angle of up to 200 relative to the nozzle wall.

33. A gas turbine engine exhaust nozzle arrangement as claimed in claim 1-30 wherein the tabs are circumferentially alternatively radially inwardly angled at an angle of up to 200 relative to the nozzle wall and radially outwardly angled at an angle of up to 200 relative to the nozzle wall.

34. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-33 wherein the tabs are of a substantially trapezoidal shape.

35. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claim 1-33 wherein the general shape of the tabs is any one of the group comprising rectangular, square, triangular shape.

36. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-33 wherein the tabs are circumferentially disposed about the periphery of the nozzle wall to define substantially trapezoidally shaped notches between adjacent tabs.

37. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-33 wherein the tabs are circumferentially disposed about the periphery of the

nozzle wall to define substantially V-shaped notches between adjacent tabs.

38. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-37 wherein the edges of the tabs are curved.

39. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-38 wherein the nozzle tabs are radially inwardly angled at an angle of up to 10 relative to the nozzle wall.

40. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-39 wherein the exhaust nozzle is a core engine nozzle.

41. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-40 wherein the exhaust nozzle is a bypass exhaust nozzle.

42. A ducted fan gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-41 wherein the arrangement comprises a core exhaust nozzle and a bypass exhaust nozzle.

43. A ducted fan gas turbine engine exhaust nozzle arrangement comprising an outer bypass exhaust nozzle as claimed in any one of claims 1-39, and an inner core exhaust nozzle of a lobed mixer type.

44. A ducted fan gas turbine engine exhaust nozzle arrangement as claimed in claim 43 wherein the downstream end of the bypass nozzle is further downstream than the downstream periphery of the core exhaust nozzle.

45. A ducted fan gas turbine engine exhaust nozzle arrangement as claimed in claim 41 wherein the downstream end of the bypass nozzle is upstream of the downstream periphery of the core exhaust nozzle.

46. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-45 wherein the arrangement is configured for exhaust noise attenuation.

47. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-45 wherein the arrangement

is configured for reducing the cross sectional area of the exhaust nozzle.

48. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-47 wherein the tabs, in operation, are deployed substantially prior to the aircraft reaching a predetermined altitude.

49. A gas turbine engine exhaust nozzle arrangement as claimed in any one of claims 1-48 wherein the tabs, in operation, are not deployed substantially after the aircraft-reaches a predetermined altitude.

50. A gas turbine engine exhaust nozzle arrangement as

hereinbefore described and with reference to figures 1 to 15.

51. A ducted fan gas turbine engine as hereinbefore described and with reference to figures 1 to 15.

52. A method of operating an aircraft having a gas turbine engine comprising an exhaust nozzle arrangement as claimed in any preceding claim wherein the method comprises the steps of: deploying noise reduction means prior to takeoff; not deploying noise reduction means above a predetermined aircraft altitude and; deploying the noise reduction means below the predetermined aircraft altitude.