WO2007122368A1

WO2007122368A1 - Variable area exhaust nozzle

Info

Publication number: WO2007122368A1
Application number: PCT/GB2007/001100
Authority: WO
Inventors: John Stanley Richardson
Original assignee: Short Brothers Plc
Priority date: 2006-04-25
Filing date: 2007-03-27
Publication date: 2007-11-01
Also published as: GB0608093D0

Abstract

A gas turbine engine is described, said gas turbine engine comprising: a by-pass fan, a fuel-burning gas turbine core structure having an exhaust nozzle for hot exhaust gas, said core structure being at least partially surrounded by a by-pass airstream cowling structure, said core structure and said by-pass airstream cowling structure defining a by-pass airstream duct therebetween, said duct having a by-pass airstream exhaust nozzle adjacent a downstream end thereof wherein an area of the exhaust nozzle for the hot exhaust gas and/or the by-pass airstream is selectively variable, said exhaust nozzle comprising at least first and second generally annular ring members, said ring members each having circumferential discontinuities at a trailing edge thereof, said ring members being mutually displaceable so as to bring said discontinuities into and out of register with each other so as to increase or decrease said nozzle area.

Description

VARIABLE AREA EXHAUST NOZZLE

The present invention relates to gas turbine engines having a by-pass fan.

There is a continuing need to be able to reduce the noise generated by gas turbine engines, particularly when used as powerplants of aeroplanes, especially civil aeroplanes. This need is most felt when the aeroplane is at low altitude and at high power such as, for example, at the take-off and climb-out phase. There is continual pressure from national and international legislation to reduce noise in absolute terms and particularly in the vicinity of airports in urban areas.

One way of reducing noise generated by gas turbine engines is to increase the ratio of incoming air directed through the by-pass fan duct to that passing through the hot turbine core exhaust stream. Furthermore, the noise generated by the fan and by-pass airstream can be reduced by increasing the area of the fan nozzle itself, this being due to the reduction in velocity of the by-pass airstream. This latter point also applies to the core exhaust.

It has also been demonstrated that reductions in noise may be achieved by improving the mixing of the air issuing from the by-pass fan nozzle exhaust with the surrounding air freestream and/or the hot exhaust from the turbine core and the by-pass fan airstream.

Whilst most effort has been directed to noise reduction of engines used for civil aviation purposes, engines used for military purposes could also benefit from noise reduction measures.

Whilst the above measures have been shown to reduce noise there is also an accompanying performance penalty in utilising some or all of these measures when the aeroplane is at altitude and in the cruise mode, for example. In general terms the penalty is manifested in reduced fuel efficiency which itself is environmentally undesirable.

Desirably, therefore, there is a need to be able to reduce noise generated by a gas turbine engine when the aeroplane is at low altitude and has a high power requirement and to optimise fuel efficiency at other times of the flight envelope when noise reduction is not such a high priority.

According to the present invention there is provided a gas turbine engine, said gas turbine engine comprising: a by-pass fan, a fuel-burning gas turbine core structure having an exhaust nozzle for hot exhaust gas, said core structure being at least partially surrounded by a by-pass airstream cowling structure, said core structure and said bypass airstream cowling structure defining a by-pass airstream duct therebetween, said duct having a by-pass airstream exhaust nozzle adjacent a downstream end thereof wherein an area of the exhaust nozzle for at least one selected from the group consisting of the hot exhaust gas and the by-pass airstream is selectively variable, said exhaust nozzle comprising at least first and second generally annular ring members, said ring members each having circumferential discontinuities at a trailing edge thereof, said ring members being mutually displaceable so as to bring said discontinuities into and out of register with each other so as to increase or decrease said nozzle area.

In the present invention, either the exhaust nozzle for the by-pass fan airstream or the core exhaust nozzle for the hot combustion gases or both nozzles may be provided with a variable area exhaust nozzle according to the present invention. In principle, the structure and operation of both nozzles will be similar but differences in scale and in materials employed will clearly be directed to the greatly differing temperatures experienced in the two exhaust streams.

The core structure of the gas turbine engine may be essentially conventional up to the exhaust nozzle and include the usual compressor means, combustor means, turbine means and an outer cowl structure surrounding the core to define the inner surface of the by-pass airstream duct of the engine.

The first and second ring members may be formed separately and adapted to fit the structure of the engine, however, in one embodiment of the present invention, one of the ring members, the first for example, may comprise an integral part of a cowling structure which surrounds the by-pass airstream. The ring member may form an integral part of the outer cowling structure and may be located at the trailing edge thereof adjacent the exhaust nozzle. The trailing edge of the ring member may be circumferentially discontinuous in the sense that the trailing edge does not form a plain circular ring but is discontinuous when viewed with respect to a plane normal to the turbine axis and passing through the trailing. edge of the exhaust nozzle. In this embodiment, the second ring member may comprise a generally circumferential ring located either radially within or radially outside the first member and able to be displaced circumferentially relative thereto. The trailing edge of the second ring member also has a series of circumferentially directed discontinuities at least generally corresponding to those of the first ring member.

In the above embodiment of the present invention, the second ring member may be rotated relative to the first ring member which may form part of the by-pass airstream duct cowling so as to bring the discontinuities of the first and second ring members into and out of register with each other.

Suitable actuating means such as hydraulic, pneumatic or electrical actuators may be employed to effect displacement of the ring members relative to each other.

In the above embodiment the ring members have impliedly been described as being continuous 360^° rings which they may be but are not necessarily so. The present invention envisages ring members which comprise a plurality of arcs or ring segments to cater for the circumstance where the by-pass exhaust duct nozzle is not a continuous circle but may have portions about the circumference thereof having different curvatures or obstructions, for example. For example, an exhaust nozzle may comprise three "lobes" when viewed end on and, in which case, the second ring member may comprise three separate segmental portions having co-operating curvatures and each displaceable so as to bring the discontinuities at their trailing edges into and out of register.

Even where the by-pass exhaust duct may be essentially circular in form it may nevertheless be advantageous to provide the displaceable ring member in segmental form. For example, it may be advantageous to provide the displaceable ring member in a plurality of segments which may be selectively independently displaceable or displaceable in groups relative to a stationary ring member in order to provide a degree of thrust vectoring for aeroplane .aerodynamic trimming purposes, for example. This advantage may, of course, also be applied to the case described above where the exhaust duct may be non-circular. In one embodiment of the present invention the discontinuities at the ring member trailing edges may comprise a formation of so-called "chevrons", i.e. a formation having a saw-tooth appearance or a series of adjacent triangles. Whilst the dimensions and hence areas of the chevrons may be substantially equal so that when the chevrons of the first and second ring members are in register, i.e. overlie each other, they form a substantially "mirror image" of each other. However, this need not be the case and the chevrons of the displaceable ring member may have chevrons of smaller area, for example, than those on the other ring member. The relative areas of the discontinuities will, however, be dependent upon the performance characteristics of the gas turbine engine in question and the performance characteristics which it is desired to achieve in, for example, the cruise mode in order to minimise any performance penalty.

It is believed that the chevron form of nozzle trailing edge has particular benefits in that it assists in promoting more efficient mixing of the by-pass fan exhaust stream and the surrounding air freestream or between the core exhaust and the surrounding by-pass fan airstream which is known to help in noise reduction.

In an alternative embodiment of the present invention, the discontinuities may comprise generally rectangular shapes.

Whilst chevron and rectangular shapes for discontinuities have been exemplified above, it will be clear to those people skilled in the gas turbine art that other shapes may alternatively be employed or that combinations of shapes may be utilised in one installation if desired for a particular performance benefit.

Whilst in general, the maximum exhaust nozzle area will be when the discontinuities overlie each other to their maximum extent and the minimum exhaust nozzle area will be when the discontinuities are at their maximum relative displacement, it is permissible for the exhaust nozzle to operate at some intermediate value between maximum and minimum area depending upon performance requirements.

In a particularly advantageous embodiment of the present invention it is envisaged that the variable area by-pass fan exhaust duct of the present invention may be incorporated into a translating cowl arrangement currently used in a cold-stream thrust reversal system for an aeroplane without compromising the safety of the known thrust reversal system. The variable area exhaust nozzle of the present invention as applied to the core exhaust nozzle for hot combustion gases may be of essentially the same structure as that applied to the by-pass fan exhaust nozzle. However, due to the greatly increased temperature of the core exhaust, suitable heat resisting metal alloys will need to be employed in the construction. The core exhaust stream will be surrounded by the bypass fan airstream. Any noise reduction due to the present invention being applied to the core exhaust nozzle may be due to improvements in mixing of the gases between the hot core exhaust and the surrounding by-pass fan airstream due to the discontinuities present in the trailing edge shape of the variable area nozzle and also due to reduction in gas velocity by providing a larger nozzle area. However, a further benefit of a variable area core exhaust nozzle is the possibility of optimising engine operation.

The discontinuities of the variable area exhaust nozzles described herein may be inclined in a generally radially inwardly direction towards the turbine axis. Consequently, when viewed in cross section, normal to the turbine axis, the variable area exhaust nozzle may have a generally frusto-conical appearance. However, the effect of variable area may also or alternatively be generated by discontinuities generally parallel to the engine axis provided that the cowl structure around the core and/or by-pass duct is contoured such that the flow has a partially radially outwardly directed element of velocity in the vicinity of the nozzle. Thus, discontinuities need to be inclined relative to the exhaust flow direction. In some practical applications of gas turbine engines according to the present invention there may be a combination of discontinuity deflection and cowl contouring to achieve the desired variable area nozzle effect.

A further advantage of the present invention is that the loads on the displaceable parts of the variable area exhaust nozzle of the present invention are mainly frictional due to the translation direction being circumferential.

In order that the present invention may be more fully understood, examples will now be described by way of illustration only with reference to the following drawings, of which:

Figure 1 shows a perspective view of the aft end of a by-pass fan equipped gas turbine engine to show the relationship between the by-pass airstream duct exhaust nozzle and also the hot gas core exhaust nozzle; Figures 2A and 2B show perspective views indicating the general principle of the present invention of a first embodiment of a variable area exhaust duct nozzle in maximum and minimum area configurations, respectively;

Figures 3A and 3B show a similar view to Figure 2 but of a portion of the exhaust duct with some schematic actuating structure in maximum and minimum area configurations, respectively;

Figures 4A and 4B show a schematic cross section of a second embodiment showing an alternative form of discontinuity in the maximum and minimum nozzle area configurations, respectively;

Figures 5A and 5B show a partial cross section showing schematically the mixing of the by-pass fan duct exhaust stream with the surrounding air freestream with the duct nozzle in maximum area and minimum area configurations, respectively; and

Figures 6A and 6B which show partial cross sections of an embodiment of a variable area by-pass fan exhaust duct nozzle according to the present invention incorporated into a translating cowl, cold stream thrust reversal system in both forward and reverse thrust modes, respectively.

Referring now to the drawings and where the same or similar features are denoted by common reference numerals.

Figure 1 shows a perspective view of the aft end of a by-pass fan equipped gas turbine engine. At its aft end the engine comprises an outer nacelle 10 and the rear end of the engine core 11. The by-pass fan airstream exhaust nozzle 12 is an annular gap surrounding the core exhaust portion 28 and the core exhaust nozzle 29 lies axially beyond the by-pass fan exhaust nozzle 12.

Figure 2 shows a schematic perspective view of the exhaust duct nozzle end of an engine outer nacelle 10. Only the exhaust nozzle 12 is shown with Figure 2A showing the nozzle in the maximum area (take off for example) configuration and Figure 2B in the minimum area (cruise mode for example) configuration. The nozzle comprises a first, stationary ring member 14 which in this embodiment is part of the nacelle outer cowl structure 16 and which first ring member has a trailing edge comprising a discontinuous B2007/001100

7 annular ring of a plurality of chevron shapes 18. The nozzle also comprises a second ring member 20 lying radially inwardly of the first ring member 14 and able to rotate circumferentially relative to the first ring member 14. The second ring member 20 also has a trailing edge comprising a discontinuous annular ring of a plurality of chevron shapes 22 (best seen in Fig. 2B). The second ring is mechanically attached to the structure of which the first ring member is a part in a manner allowing the second ring member to rotate relative thereto (to be described below). In Figure 2A the chevrons 18, 22 overlie each other to allow maximum nozzle area with the intervening gap areas 24 between the chevrons open to mixing of the airflow from the by-pass fan exhaust and the surrounding freestream air. In Figure 2B the second ring member 20 has been rotated by a amount sufficient to bring the chevrons 22 thereof into position to minimise the area of the intervening gaps 24 and so produce the minimum by-pass fan exhaust nozzle area. In Figure 2A the exhaust of the engine core is shown schematically at 28 (the core exhaust is shown truncated by means of a wavy line since the core exhaust extends appreciably beyond the by-pass fan exhaust duct 12 in a rearwardly direction) and the exhaust nozzle 12 of the by-pass fan duct is the annulus surrounding the core exhaust 28.

Figure 3 shows similar views to those of Figure 2 but with some schematic operating structure included. Figures 3A and 3B also show the by-pass fan exhaust duct nozzle in maximum and minimum area configurations, respectively. In this embodiment the second ring member 20 is depicted as a segment 30 having a plurality of chevrons 22. There is a plurality of the segments 30 extending around the exhaust duct nozzle 12 of sufficient number to encircle the duct 12. The segments 30 are attached to the first ring member 14 by "T"-headed bolts 32 and the segments are constrained to slide circumferentially by virtue of slots 34 which are a sliding fit on the shanks of the "T"- headed bolts. The segments each have an axially (relative to the engine axis) extending tongue 36. A hydraulic actuator 38 is provided to move each segment 30 in the circumferential direction, the actuator being connected pivotally 40 at the tongue 36 and anchored at its distal end (not shown) to suitable structure (not shown). The actuator is covered by a nacelle cowl member (not shown in Fig. 3 but see Fig. 5 or 6) for aerodynamic purposes. Figure 3A shows the actuator extended with the chevrons 18, 22 overlying each other in the maximum area nozzle configuration. In Figure 3B the actuator 38 is retracted moving the segment 30 and "T"-headed bolts 32 to the opposite ends of the slots 34 which corresponds to the position of minimum by-pass duct nozzle area. The actuator may be controlled to position the segment at any position between the ends of the slots 34 to provide a nozzle area intermediate maximum and minimum.

By means of a suitable control system (not shown) the segments may be moved independently to provide, for example, a degree of thrust vectoring by the exhaust nozzle 12.

The arrangement shown in Figure 3 may be employed on an exhaust duct nozzle which is circular or on an exhaust duct which comprises arcs of differing radii of curvature.

Figure 4 shows a similar view to that shown in Figure 2 except for the fact that the shape of the discontinuities is rectangular 50, 52 rather than chevron-shaped. As in Figure 2 there is a first ring member 14 having rectangular-shaped discontinuities or projections 50 and, a second ring member 20 having corresponding discontinuities/ projections 52. Although the discontinuities 50, 52 are shown to be quite wide circumferentially, this is merely for the purposes of illustration and their circumferential extent will be determined on the basis of noise attenuation and fuel efficiency calculations and tests.

A possibility may be that the projections 50 on the stationary first ring member may be quite wide circumferentially leaving relatively narrower gaps 54 therebetween and consequently narrower projections 52 on the displaceable ring member 20. Again, such considerations will be determined according to the performance requirements of the particular engine installation under consideration.

As will be observed with this embodiment of the present invention, the actual trailing edge of the by-pass duct exhaust nozzle ring when in the minimum area configuration is circular thus, this may have performance benefits for the engine in the cruise mode, for example.

The exhaust duct nozzle described with reference to Figures 2, 3 and 4 may be adapted and used as a variable area exhaust nozzle for the hot core exhaust nozzle of the fuel- burning turbine core.

Figure 5 shows a partial cross section through a portion of a by-pass fan equipped gas turbine engine in the vicinity of the by-pass airstream exhaust duct and nozzle 12. The reference numerals are as used in Figures 1 to 3. As may be seen from the cross section in Figure 5 the by-pass fan airstream duct is an axially directed annular passage 60. Figure 5A shows the by-pass duct nozzle 12 having maximum area, i.e. the discontinuities or chevrons 18, 22, for example, are in the positions shown in Figure 3A or 2A. In this position there is both reduced velocity of the actual by-pass airstream 62 and there is maximum mixing of the by-pass airstream 62 with the surrounding air freestream 64, both factors tending to reduce overall noise generation. Figure 5B shows the by-pass exhaust duct in minimum area configuration where the discontinuities or chevrons 18, 22 are as shown in Figures 2B and 3B, for example. In this configuration, by-pass exhaust airstream velocity is at a maximum and mixing of the airstreams 62, 64 is at a minimum.

Figure 6 shows similar cross sections to those of Figure 5 but with additional explanation relating to the installation of a variable area by-pass fan nozzle into a gas turbine engine having a known cold-stream thrust reversal system. Figure 6A shows the engine in normal forward thrust with the variable area by-pass duct nozzle 12 in minimum area configuration. All of the by-pass fan airstream 62 is passing through the nozzle 12. The mechanism for the variable area nozzle 12 is incorporated into a translating cowl portion 70 at the rear of the engine nacelle 10. Below the cowl portion 70 is a thrust reversing cascade 72. Gas flow to the cascade 72 is prevented by a blocker door 74 which forms part of an inner cowl portion 76 which is arranged to move with the translating cowl portion 70. The blocker door 74 is pivotally connected 76 to a leading edge portion 78 of the inner cowl portion 76 and supported by a strut 80 which is pivotally connected to both the blocker door 82 and to the core structure 84. When the aeroplane lands and begins to brake, reverse thrust is selected which causes the translating cowl portion 70, inner cowl portion 76 and variable area nozzle structure to move rearwardly relative to the other engine features as indicated in Figure 6B, the movement being effected, for example, by a known screw-jack mechanism (not shown).The effect of the cowl 70, 76 being moved rearwardly is to cause the blocker door 74 to articulate into a vertical position and to prevent gas flow through the duct 60, the by-pass airstream now being deflected and forced to flow through the cascade 72 which directs the air flow 62 in a forwardly direction to help slow the aeroplane. When the aeroplane has slowed sufficiently, the reverse thrust arrangement is retracted to adopt the geometry shown in Figure 6A.. It should be noted that the above described trust reversal system is, apart from the incorporation of the variable area by-pass exhaust nozzle according to the present invention, an example of a known and conventional thrust reversal system. Throughout the description and claims of this specification, the words "comprise" and "contain" and variations of the words, for example "comprising" and "comprises", means "including but not limited to", and is not intended to (and does not) exclude other moieties, additives, components, integers or steps.

Throughout the description and claims of this specification, the singular encompasses the plural unless the context otherwise requires. In particular, where the indefinite article is used, the specification is to be understood as contemplating plurality as well as singularity, unless the context requires otherwise.

Features, integers, characteristics or groups described in conjunction with a particular aspect, embodiment or example of the invention are to be understood to be applicable to any other aspect, embodiment or example described herein unless incompatible therewith.

Claims

1. A gas turbine engine, said gas turbine engine comprising: a by-pass fan, a fuel-burning gas turbine core structure having an exhaust nozzle for hot exhaust gas, said core structure being at least partially surrounded by a by-: pass airstream cowling structure, said core structure and said by-pass airstream cowling structure defining a by-pass airstream duct therebetween, said duct having a by-pass airstream exhaust nozzle adjacent a downstream end thereof wherein an area of the exhaust nozzle for at least one selected from the group consisting of the hot exhaust gas and the by-pass airstream is selectively variable, said exhaust nozzle comprising at least first and second generally annular ring members, said ring members each having circumferential discontinuities at a trailing edge thereof, said ring members being mutually displaceable so as to bring said discontinuities into and out of register with each other so as to increase or decrease said nozzle area.

2. A gas turbine engine according to claim 1 wherein one of the first or second ring members is an integral part of a cowling structure and is stationary relative to the other ring member.

3. A gas turbine engine according to either claim 1 or claim 2 wherein circumferential discontinuities are in the form of chevrons.

4. A gas turbine engine according to any one preceding claim wherein one of the first or the second ring members is a continuous annular ring.

5. A gas turbine engine according to any one of preceding claims 1 to 3 wherein one of the ring members comprises a plurality of arcs or ring segments.

6. A gas turbine engine according to claim 5 wherein said arcs or segments may be selectively moved independently relative to the other ring member.

7. A gas turbine engine according to any one of claims 1 , 2 or 4 to 6 wherein said discontinuities are in the form of rectangles.

8. A gas turbine engine according to any one preceding claim wherein said variable area nozzle is incorporated into a cold stream thrust reversal system.

9. A gas turbine engine according to any one preceding claim wherein said first and second ring members are displaced in a circumferential direction relative to each other.

10. A gas turbine engine substantially as hereinbefore described with reference to the accompanying description and Figures 2 and 3; or Figure 4; or Figure

5; or Figure 6 of the drawings.