U.S. patent application number 10/383554 was filed with the patent office on 2004-01-15 for variable area nozzle.
Invention is credited to Whurr, John R..
Application Number | 20040006969 10/383554 |
Document ID | / |
Family ID | 9932729 |
Filed Date | 2004-01-15 |
United States Patent
Application |
20040006969 |
Kind Code |
A1 |
Whurr, John R. |
January 15, 2004 |
Variable area nozzle
Abstract
An exhaust nozzle assembly (18) for a gas turbine engine (10),
the assembly (18) comprises a main axis (20), an inner nozzle (19),
an outer nozzle (17), a translatable centre-body (44) and a means
for translating the centre-body (49) between a forward position and
a rearward position. The centre-body (44) is disposed radially
inwardly of the inner nozzle (19) thereby partly defining an inner
duct (35) for a core engine gas flow (28), an outer duct (17) for a
bypass gas flow (30) is defined by the inner nozzle (19) and the
outer nozzle (17). The outer nozzle (17) extends downstream of the
inner nozzle (19) and with the centre-body (44) defines a final
mixing duct (36) having a final exhaust exit area (38). When the
centre-body (44) is in the forward position the final exhaust exit
area (38) is at a maximum area, and when the centre-body (44) is in
a rearward position the final exhaust exit area (38) is at a
minimum area.
Inventors: |
Whurr, John R.; (Derby,
GB) |
Correspondence
Address: |
MANELLI DENISON & SELTER
2000 M STREET NW SUITE 700
WASHINGTON
DC
20036-3307
US
|
Family ID: |
9932729 |
Appl. No.: |
10/383554 |
Filed: |
March 10, 2003 |
Current U.S.
Class: |
60/262 ;
60/771 |
Current CPC
Class: |
F02K 1/08 20130101 |
Class at
Publication: |
60/262 ;
60/771 |
International
Class: |
F02K 003/04 |
Foreign Application Data
Date |
Code |
Application Number |
Mar 12, 2002 |
GB |
0205701.6 |
Claims
1. An exhaust nozzle assembly for a gas turbine engine, the
assembly comprises a main axis, an inner nozzle, an outer nozzle, a
translatable centre-body and a means for translating the
centre-body between a forward position and a rearward position; the
centre-body is disposed radially inwardly of the inner nozzle
thereby partly defining an inner duct for a core engine gas flow,
an outer duct for a bypass gas flow is defined by the inner nozzle
and the outer nozzle, the outer nozzle extends downstream of the
inner nozzle and with the centre-body defines a final mixing duct
having a final exhaust exit area; wherein when the centre-body is
in the forward position the final exhaust exit area is at a maximum
area, and when the centre-body is in a rearward position the final
exhaust exit area is at a minimum area.
2. An exhaust nozzle assembly as claimed in claim 1 wherein the
inner nozzle defines a core exhaust exit area, and when the
centre-body is in the forward position the core exhaust exit area
is at a maximum area, and when the centre-body is in a rearward
position the core exhaust exit area is at a minimum area.
3. An exhaust nozzle assembly as claimed in claim 1 wherein the
inner nozzle defines a core exhaust exit area, and when the
centre-body is in the forward position the core exhaust exit area
is at a minimum area, and when the centre-body is in a rearward
position the core exhaust exit area is at a maximum area.
4. An exhaust nozzle assembly as claimed in claim 1 wherein when
the centre-body is in the rearward position the centre-body and the
outer nozzle define, in axial flow series, a converging portion and
a diverging portion, the mixed exhaust flow then exits the final
exhaust exit area which is at the minimum area.
5. An exhaust nozzle assembly as claimed in claim 1 wherein the
centre-body comprises, in a downstream direction, a parallel
portion, a waist, a diverging section, a maximum diameter portion
and a taper; the parallel portion is substantially parallel and a
part of which is slidably engaged with a static core engine
structure.
6. An exhaust nozzle assembly as claimed in claim 5 wherein when
the centre-body is in the forward position the waist is generally
axially aligned with the core exhaust exit area thereby providing
the maximum core exhaust exit areas; when the centre-body is in the
rearward position the parallel portion is axially aligned with core
exhaust exit area thereby minimising the exit area.
7. An exhaust nozzle assembly as claimed in claim 1 wherein the
centre-body comprises, in a downstream direction, a parallel
portion, a diverging section, a maximum diameter portion and a
taper; the parallel section is substantially parallel and a part of
which is slidably engaged with a static core engine structure.
8. An exhaust nozzle assembly as claimed in claim 7 wherein when
the centre-body is in the forward position or the rearward position
the diverging section is generally axially aligned with core
exhaust exit area thereby providing a constant core exhaust exit
area.
9. An exhaust nozzle assembly as claimed in claim 7 wherein when
the centre-body is in the forward position the waist is axially
aligned with the core exhaust exit area thereby providing the
maximum core exhaust exit areas; when the centre-body is in the
rearward position the parallel portion is axially aligned with core
exhaust exit area and thereby minimising the core exhaust exit
area.
10. An exhaust nozzle assembly as claimed in claim 5 wherein when
the centre-body is in the forward position the downstream end of
the taper is generally axially aligned with the final exhaust exit
area thereby providing the maximum final exit area; when the
centre-body is in the rearward position the maximum diameter
portion is generally axially aligned with the final exhaust exit
area thereby minimising the final exit area.
11. An exhaust nozzle assembly as claimed in claim 5 wherein a
means for translating the centre-body between a forward position
and a rearward position comprises at least one actuator mounted
within and to the static core engine structure, the actuator having
a piston which is attached to a mounting on the translating
centre-body.
12. An exhaust nozzle assembly as claimed in claim 1 wherein when
the centre-body is in the rearward position the engine is in
high-speed mode and when the centre-body is in the forward position
the engine is in low-speed mode.
13. An exhaust nozzle assembly as claimed in claim 1 wherein when
the engine operates between low-speed mode and high-speed mode the
centre-body is positioned between the forward position and the
rearward position.
14. A gas turbine engine comprising an exhaust nozzle assembly as
claimed in claim 1.
Description
[0001] The present invention relates to a variable exhaust area
nozzle for a gas turbine engine.
[0002] An aircraft is required to perform many different
operational modes in a single flight cycle. Such modes comprise
taxiing, take-off, climb, cruise, holding, decent and landing to
which the engines of the aircraft must also comply with related
performance requirements.
[0003] Conventional gas turbine engines, particularly turbofans,
operate at relatively low exhaust nozzle pressure ratios, where it
is sufficient to use a convergent nozzle to develop thrust, as the
pressure loss due to external expansion of the working exhaust
gases is relatively low. With increasing flight speed the nozzle
pressure ratio increases, and the expansion of the gases needs to
be controlled to reduce losses, normally using a
convergent-divergent nozzle. In the ideal setting at cruise, the
nozzle exit pressure is reduced to ambient by the divergent part of
the nozzle. If the nozzle comprises fixed convergent-divergent
geometry, optimised for cruise, the losses at lower nozzle pressure
ratios would increase due to either over expansion of the exhaust
gases or exhaust flow separation. Thus it is preferable for the
area ratio of the convergent-divergent nozzle to be varied with
flight speed and nozzle pressure ratio.
[0004] It is well known to vary the area of an exhaust nozzle; one
such application is the Olympus.TM. engines on Concorde.TM.
aircraft as well as the propulsion systems on most of the World's
military fighter aircraft. However, these are generally complex
multi-variable devices, typically consisting of systems of
overlapping petals and comprising numerous moving components and
multiple actuation systems.
[0005] U.S. Pat. No. 4,527,388 granted to The Garrett Corp.,
discloses a turbofan engine comprising a core engine and a fan
bypass passage, both discharging streams of pressurised gases
through a mixing section and into an exhaust nozzle. The exhaust
nozzle defines a throat through which the combined gas streams are
accelerated and discharged therethrough. The exhaust nozzle also
comprises a single axially moveable and rearwardly tapering
centre-body, which is moveable relative to the remainder of the
exhaust nozzle to simultaneously vary both the throat area and one
of the respective fluid flow areas through which the core engine
and fan bypass passage fluid streams flow. When the engine produces
high power the centre-body is translated rearward, for instance at
takeoff, such that the areas of the exhaust nozzle and core are
both minimum. This increases the velocity of the gas stream and
intrinsically disadvantageously increases the amounts of exhaust or
jet noise. Furthermore, at a lower engine power, for instance
relatively slow cruise; the centre-body is in its forward-most
position where the mixing section is generally divergent in area.
This cruise position leads to a reduced gas stream velocity
relative to the ambient air and disadvantageously provides a less
thrust-capable engine.
[0006] Therefore it is an object of the present invention to
provide a means for varying the areas of the final exhaust nozzle
and the core exhaust exit such that; the final exhaust nozzle area
is increased at take-off to maximise airflow and minimise jet
velocity; the core exhaust exit area is adjustable during take-off
to optimise the jet velocity profile to further reduce jet
noise.
[0007] A further object of the present invention is to provide a
means for varying the aerodynamic profile of the exhaust nozzle
such that; at high nozzle pressure ratios, for instance at cruise,
the nozzle comprises a convergent-divergent profile; at lower
nozzle pressure ratios, for instance aircraft take-off, decent or
hold, the nozzle comprises a predominantly convergent profile.
[0008] Accordingly the present invention seeks to provide an
exhaust nozzle assembly for a gas turbine engine, the assembly
comprises a main axis, an inner nozzle, an outer nozzle, a
translatable centre-body and a means for translating the
centre-body between a forward position and a rearward position; the
centre-body is disposed radially inwardly of the inner nozzle
thereby partly defining an inner duct for a core engine gas flow,
an outer duct for a bypass gas flow is defined by the inner nozzle
and the outer nozzle, the outer nozzle extends downstream of the
inner nozzle and with the centre-body defines a final mixing duct
having a final exhaust exit area; wherein when the centre-body is
in the forward position the final exhaust exit area is at a maximum
area, and when the centre-body is in a rearward position the final
exhaust exit area is at a minimum area.
[0009] Preferably, the inner nozzle defines a core exhaust exit
area, and when the centre-body is in the forward position the core
exhaust exit area is at a maximum area, and when the centre-body is
in a rearward position the core exhaust exit area is at a minimum
area.
[0010] Alternatively, the inner nozzle defines a core exhaust exit
area, and when the centre-body is in the forward position the core
exhaust exit area is at a minimum area, and when the centre-body is
in a rearward position the core exhaust exit area is at a maximum
area.
[0011] Preferably, when the centre-body is in the rearward position
the centre-body and the outer nozzle define, in axial flow series,
a converging portion and a diverging portion, the mixed exhaust
flow then exits the final exhaust exit area which is at the minimum
area.
[0012] Preferably, the centre-body comprises, in a downstream
direction, a parallel portion, a waist, a diverging section, a
maximum diameter portion and a taper; the parallel portion is
substantially parallel and a part of which is slidably engaged with
a static core engine structure.
[0013] Preferably, when the centre-body is in the forward position
the waist is generally axially aligned with the core exhaust exit
area thereby providing the maximum core exhaust exit areas; when
the centre-body is in the rearward position the parallel portion is
axially aligned with core exhaust exit area thereby minimising the
exit area.
[0014] Alternatively, the centre-body comprises, in a downstream
direction, a parallel portion, a diverging section, a maximum
diameter portion and a taper; the parallel section is substantially
parallel and a part of which is slidably engaged with a static core
engine structure. Preferably, when this centre-body is in the
forward position or the rearward position the diverging section is
generally axially aligned with core exhaust exit area thereby
providing a constant core exhaust exit area. Alternatively, when
the centre-body is in the forward position the waist is axially
aligned with the core exhaust exit area thereby providing the
maximum core exhaust exit areas; when the centre-body is in the
rearward position the parallel portion is axially aligned with core
exhaust exit area and thereby minimising the core exhaust exit
area.
[0015] Preferably, when the centre-body is in the forward position
the downstream end of the taper is generally axially aligned with
the final exhaust exit area thereby providing the maximum final
exit area; when the centre-body is in the rearward position the
maximum diameter portion is generally axially aligned with the
final exhaust exit area thereby minimising the final exit area.
[0016] Preferably, a means for translating the centre-body between
a forward position and a rearward position comprises at least one
actuator mounted within and to the static core engine structure,
the actuator having a piston that is attached to a mounting on the
translating centre-body.
[0017] Preferably, when the centre-body is in the rearward position
the engine is in high-speed mode and when the centre-body is in the
forward position the engine is in low-speed mode.
[0018] Preferably, when the engine operates between low-speed mode
and high-speed mode the centre-body is positioned between the
forward position and the rearward position.
[0019] Preferably, a gas turbine engine comprising an exhaust
nozzle assembly as claimed in any one of the preceding
paragraphs.
[0020] The present invention will be more fully described by way of
example with reference to the accompanying drawings in which:
[0021] FIG. 1A is a schematic section of a ducted fan gas turbine
engine incorporating an exhaust nozzle assembly, which itself
comprises a translating centre-body in accordance with the present
invention;
[0022] FIG. 1B is a part schematic section of the exhaust nozzle
assembly of FIG. 1A;
[0023] FIG. 2 is a schematic section of a ducted fan gas turbine
engine incorporating an exhaust nozzle assembly, which itself
comprises a translating centre-body in accordance with the present
invention;
[0024] FIG. 2B is a part schematic section of the exhaust nozzle
assembly of FIG. 2A;
[0025] FIG. 3A is a part schematic section of the exhaust nozzle
assembly comprising a second embodiment of the translating
centre-body;
[0026] FIG. 3A is a part schematic section of the exhaust nozzle
assembly comprising a second embodiment of the translating
centre-body.
[0027] FIG. 4 is a graph showing velocity profiles for a final
exhaust exit.
[0028] With reference to FIGS. 1 and 2 where reference numerals of
one figure denote like elements of the other figure, a ducted fan
gas turbine engine 10 comprises, in axial flow series an air intake
12, a propulsive fan 14, a core engine 16 and an exhaust nozzle
assembly 18 all disposed about a main engine axis 20. The exhaust
nozzle assembly 18 comprises an outer nozzle 17 and a radially
inner nozzle 19. The core engine 16 comprises, in axial flow
series, a series of compressors 22, a combustor 24, and a series of
turbines 26. 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 12 and is compressed and accelerated by the fan 14. The air
from the fan 14 is split between a core engine flow 28 and a bypass
flow 30. The core engine flow 28 enters core engine 16, flows
through the core engine compressors 22 where it is further
compressed, and into the combustor 24 where it is mixed with fuel,
which is supplied to, and burnt within the combustor 24. Combustion
of the fuel with the compressed air from the compressors 22
generates a high energy and velocity gas stream, which exits the
combustor 24 and flows downstream through the turbines 26. As the
high energy gas stream flows through the turbines 26 it rotates
turbine rotors extracting energy from the gas stream which is used
to drive the fan 14 and compressors 22 via engine shafts 32 which
drivingly connect the turbines rotors with the compressors 22 and
fan 14. Having flowed through the turbines 26 the high energy gas
stream from the combustor 24 still has a significant amount of
energy and velocity and it is exhausted, as a core exhaust stream
28, through a core exhaust exit area 42 of the engine exhaust
nozzle assembly 18 into a mixing area 36. The remainder of the air
from, and accelerated by, the fan 14 flows within a bypass duct 34
around the core engine 16. This bypass airflow 30, which has been
accelerated by the fan 14, flows through a bypass exit area 40 of
the exhaust nozzle assembly 18 where it is mixed with the core
engine flow 28 in the mixing area 36 and exhausted through a final
exhaust area 38 to provide propulsive thrust. The bypass flow 30
provides the majority of the useful propulsive thrust of the engine
10.
[0029] A conventional flight cycle for a commercial aircraft
comprises take-off, climb, cruise, descent and landing modes. If
the aircraft is required to wait for a landing slot, then the
flight cycle may also comprise a holding mode. At take-off the
engine is at full power and at near full power while the aircraft
climbs to cruise altitude. During the cruise phase the engine
operates at typically 70-80% of its full power. While the aircraft
is descending the engines are reduced to 10-20% of full power.
[0030] For commercial aircraft, the engines are designed to be at
their most efficient at cruise mode as the greatest proportion of
fuel is used during this phase. Thus, for the engine to be
efficient at other modes, it is desirable to have a variable area
nozzle providing the ability to control the working line and
operating points of the engine's propulsive fan 14. Importantly,
with the final exhaust area 38 at a maximum, the engine 10 is then
able to operate at its maximum flow capacity at takeoff, reducing
jet velocity, and consequently jet noise, for any given fan size
and thrust level requirement. Variation of the core engine nozzle
area 42 enables the core engine performance to be matched and
adjusted in accordance with the exhaust jet velocity for the
different flight cycle modes. This would enable the jet velocity
profile to be optimised to achieve minimum specific fuel
consumption at cruise and minimum jet noise at takeoff. In
particular, it is the variation of the pressure ratios (the
pressure difference of a gas stream and ambient) and gas stream
velocities exiting the final nozzle at the different flight cycle
conditions that the variable area exhaust nozzle assembly 18 of the
present invention is suited to.
[0031] In accordance with the present invention, a translating
centre-body 44 is incorporated as part of the exhaust nozzle
assembly 18. The centre-body 44 is disposed radially inwardly of
the inner or core nozzle 19 thereby partly defining an inner or
core duct 35 for the flow of core engine gas 28 therethrough. The
inner nozzle 19 and the outer nozzle 17 define the outer or bypass
duct 34 for the flow of bypass gas 30 therethrough. The outer
nozzle 17 extends downstream of the inner nozzle 19 and with the
centre-body 44 defines a final mixing duct/area 36. Both the core
gas flow 28 and the bypass gas flow 30 discharge into the final
mixing duct 36. The centre-body 44 is axially translatable and
provides modification of the final exhaust exit area 38 between
takeoff and cruise flight conditions.
[0032] FIG. 1 shows the centre-body 44 in a rearward position
suitable for cruise flight of an associated aircraft. In this
position the final exhaust exit area 38 comprises a
convergent-divergent profile that is beneficial for high nozzle
pressure ratios, i.e. where there is a relatively large difference
in exhaust gas and ambient pressures. FIG. 2 shows the centre-body
44 in a forward position suitable for take-off or low air-speed of
an associated aircraft. In this position the final exhaust exit
area 38 comprises a convergent profile for relatively low nozzle
pressure ratios. Although FIG. 2 shows the sectional lines of the
outer nozzle 17 and centre-body 44 diverging it should be
appreciated that the cross-sectional area converges towards the
final exhaust exit area 38.
[0033] Referring to FIG. 3B, which shows the nozzle assembly in
more detail, when the centre-body 44 is in the rearward position
the centre-body 44 and the outer nozzle 17 define, in axial flow
series, a converging portion 72 and a diverging portion 74. The
transition between the converging portion 72 and diverging portion
74 is generally axially aligned with the maximum diameter portion
58 of the centre-body 44. Although divergent, the final exit area
38 is in fact at a minimum exit area for this mode of operation.
The divergent portion 74 is required, as the pressure ratio across
the exhaust nozzle is such that a convergent nozzle alone cannot
achieve complete expansion of the gas stream. The divergent portion
74 is provided to allow full expansion of the exhausted gas stream
therefore achieving an increase in thrust over a convergent only
exhaust nozzle.
[0034] The invention is particularly advantageous for application
to high-speed aircraft where the engine/nacelle diameter is more
critical, and where a greater variation in nozzle pressure ratio
between takeoff and cruise will exist than for a conventional
sub-sonic aircraft. The present invention will allow the diameter
of the propulsive fan 14 to be minimised for given takeoff thrust
and noise requirements, and will facilitate the achievement of good
nozzle performance at both takeoff and cruise.
[0035] The centre-body 44 is slideably mounted to a static core
engine structure 46 and is translated axially by a means for
translating the centre-body 49. In this embodiment an annular array
of actuators 48 are provided. Each actuator 48 is mounted to the
static core engine structure 46 with its piston 50 connected to a
mounting 52 disposed to the centre-body 44. The actuator 48 is a
hydraulic ram but alternatively it may be an electric screw jack or
any other means suitable for translating the centre-body 44 axially
between its rearward and forward positions.
[0036] FIGS. 2A and 2B show the shape of the centre-body 44
comprising, in axially rearward series, a parallel portion 54, a
waist 56, a maximum diameter portion 58 and a taper 60. The
parallel portion 54 is substantially parallel and a part of which
is disposed radially inwardly of the static core engine structure
46. Alternatively, the parallel portion 54 may be disposed radially
outwardly of the static core engine structure 46. Although not
shown an airtight sealing means may be provided between the
parallel portion 54 and the static core engine structure 46. The
waist 56 is of a smaller diameter than the parallel portion 54.
[0037] In the axially forward position, e.g. for take-off, the
centre-body 44 allows both the core exhaust exit area 42 and the
final exhaust exit area 38 to be at their maximum flow areas. It
should be appreciated that the mixing duct 36 is substantially
convergent towards the final exhaust exit area 38. In this
position, the waist 56 is axially aligned with the core exhaust
exit area 42 and the downstream end of the taper 60 is axially
aligned with the final exhaust exit area 38. For cruise operation,
the centre-body 44 is moved axially rearward partially filling both
the core exhaust exit area 42 and the final exhaust exit area 38
thereby reducing their effective areas. It should be appreciated
that these axial positions of the centre-body 44 are preferential
for the two mode of operation, however, for modes between the
cruise and takeoff it is intended that the position of the
centre-body 33 is altered to be beneficial to the efficient
operation of the core engine 16 and propulsive fan 14.
[0038] Referring to FIGS. 1A, 1B, 2A and 2B, in summary when the
centre-body 44 is in the forward position the final exhaust exit
area 38 is at a maximum area, and when the centre-body 44 is in a
rearward position the final exhaust exit area 38 is at a minimum
area. It is preferable that when the centre-body 44 is in the
forward position the core exhaust exit area 42 is at a maximum
area, and when the centre-body 44 is in a rearward position the
core exhaust exit area 42 is at a minimum area.
[0039] FIGS. 3A and 3B show an alternative second embodiment of the
centre-body 44. The shape of the centre-body 44 comprises, in
axially rearward series, a parallel portion 54, a diverging section
55, a maximum diameter portion 58 and a taper 60. This alternative
embodiment of the centre-body 44 differs from the preceding design
in that there is no waist 56 and the profile of the centre-body 44
comprises the parallel portion 54 blending into the diverging
portion 55 to the maximum diameter portion 58. Thus in the axially
forward position (as shown in FIG. 3A), e.g. for take-off, the
centre-body 44 allows the core exhaust exit area 42 to be at a
minimum flow area and the final exhaust exit area 38 to be at its
maximum flow area. In this position, part of the diverging portion
55 is axially aligned with the core exhaust exit area 42 and the
downstream end of the taper 60 is axially aligned with the final
exhaust exit area 38. Although preferred this is not essential and
depending on a particular engine application the end of the taper
60 may be axially forward or rearward of the final exhaust exit
area 38.
[0040] For cruise operation (as shown in FIG. 3B), the centre-body
44 is moved axially rearward increasing the core exhaust exit area
42 and reducing the final exhaust exit area 38. In the cruise or
rearward position the mixing duct 36 comprises a similar convergent
and divergent portion as shown and described with reference to FIG.
1B.
[0041] Both embodiments of the centre-body 44 are beneficial in
lowering the working line of the fan as the pressure ratio is
further reduced for the propulsive fan 14. Furthermore, either
arrangement is particularly advantageous in reducing exhaust noise
at take-off.
[0042] FIG. 4 indicates how the final exhaust exit gas stream
velocity profile is advantageously modified for the second
embodiment. The x-axis represents gas velocity 62 and the y-axis
represents radial height from the central engine axis 20. A typical
exit velocity profile for fixed final exhaust nozzle geometry is
shown by line 64. The radially outer portion of the exhausted gas
flow forms a shear layer region 70 as the exhaust gases mix with
the ambient air. It is this shear layer region that generates the
exhaust noise. Exhaust noise is related to the velocities of the
exhausted gas stream and ambient and generally speaking the greater
the velocity difference the greater the noise. It is therefore
advantageous to reduce the velocity of the radially outer portion
of the exhausted gas flow thereby reducing exhaust noise
particularly during take-off. During take-off the centre-body 44 is
in its forward most position and the final exhaust exit is a
maximum and the core exit area is minimum. This configuration
results in the exit velocity profile of line 68. Here the core gas
stream velocity is relatively greater, at the centre of the
combined gas flow, and the bypass gas stream velocity is relatively
less than that of a fixed geometry nozzle. It should be appreciated
that the shear layer region 70 is subject to significantly reduced
exit gas stream velocity thereby reducing exhaust noise.
[0043] Alternatively, in some applications, it may be more
advantageous to produce a more uniform velocity profile, as
indicated by line 66. This would be achieved by the first
embodiment of the present invention. As exhaust noise is related to
the gas stream velocity, the first embodiment is beneficial as it
produces a more constant velocity profile without a significant
velocity peak.
[0044] Whilst endeavouring in the foregoing specification to draw
attention to those features of the invention believed to be of
particular importance it should be understood that the Applicant
claims protection in respect of any patentable feature or
combination of features hereinbefore referred to and/or shown in
the drawings whether or not particular emphasis has been placed
thereon.
* * * * *