U.S. patent application number 11/366747 was filed with the patent office on 2010-07-08 for engine noise.
This patent application is currently assigned to Rolls-Royce plc. Invention is credited to Kevin M. Britchford, Rory D. Stieger, Paul J. R. Strange.
Application Number | 20100170261 11/366747 |
Document ID | / |
Family ID | 34509061 |
Filed Date | 2010-07-08 |
United States Patent
Application |
20100170261 |
Kind Code |
A1 |
Stieger; Rory D. ; et
al. |
July 8, 2010 |
ENGINE NOISE
Abstract
In order to provide noise suppression, bumps or undulations are
provided on a nozzle surface in order to vary the
convergent-divergent ratio between that surface and an opposed
nozzle surface. By such an approach, a circumferential variation in
the shock cell pattern is created and the flow is deflected so as
to enhance turbulent mixing thereby suppressing noise.
Inventors: |
Stieger; Rory D.; (Derby,
GB) ; Britchford; Kevin M.; (Loughborough, GB)
; Strange; Paul J. R.; (Duffield, GB) |
Correspondence
Address: |
OLIFF & BERRIDGE, PLC
P.O. BOX 320850
ALEXANDRIA
VA
22320-4850
US
|
Assignee: |
Rolls-Royce plc
London
GB
|
Family ID: |
34509061 |
Appl. No.: |
11/366747 |
Filed: |
March 3, 2006 |
Current U.S.
Class: |
60/771 |
Current CPC
Class: |
Y02T 50/60 20130101;
F02K 1/46 20130101; F02K 1/827 20130101; B64D 33/06 20130101; Y02T
50/671 20130101; F05D 2260/96 20130101; F02K 1/48 20130101; F05D
2240/127 20130101 |
Class at
Publication: |
60/771 |
International
Class: |
F02K 1/06 20060101
F02K001/06 |
Foreign Application Data
Date |
Code |
Application Number |
Mar 15, 2005 |
GB |
0505246.9 |
Claims
1. A nozzle for a gas turbine engine, the nozzle comprising a
nozzle surface defining a plurality of undulations to vary in a
circumferential direction, perpendicular to a jet flow direction,
an available cross sectional area across the nozzle between the
nozzle surface and an opposed surface of the nozzle over a desired
convergent-divergent ratio range for noise control of the jet
passing through the nozzle in use, wherein the undulations have
amplitude in the range 0.1-2.0% of the nozzle exit diameter to
attenuate shock cell noise.
2. A nozzle as claimed in claim 1, the undulations also providing
variation in an angle of flow of the jet passing through the
nozzle.
3. A nozzle as claimed in claim 1 wherein the variation in cross
sectional area is adjusted to provide stimulation in mixing of a
shear layer of the jet for relative noise reduction in comparison
with that without mixing of the shear layer of the jet.
4. A nozzle as claimed in claim 1 wherein the undulations alter a
repeat cycle or provide variation of intensity of shock cells
generated by the jet.
5. A nozzle as claimed in claim 1 wherein the undulations comprise
bumps formed in the nozzle surface.
6. A nozzle as claimed in claim 1 wherein the undulations are
sinusoidal in the circumferential direction.
7. A nozzle as claimed in claim 1 wherein the undulations each
comprise maximum amplitude, the maximum amplitude situated a
distance between 2% and 15% of the nozzle diameter along the nozzle
surface upstream from the nozzle exit plane.
8. A nozzle as claimed in claim 7 wherein the maximum amplitude
situated a distance equivalent to 6% of the nozzle diameter
upstream of nozzle exit plane.
9. A nozzle as claimed in claim 1 wherein the undulations each
comprise a maximum amplitude, the maximum amplitude situated within
a distance equivalent to +/-2% of the nozzle diameter along the
nozzle surface from the nozzle throat plane.
10. A nozzle as claimed in claim 1 wherein an approximate
cross-section shape of the undulations are from a group comprising
triangular, trapezoidal, part-circular, and
asymmetric-sinusoidal.
11. A nozzle as claimed in claim 1 wherein the undulations have an
aerodynamically smooth gradual spline in an axial direction of jet
flow in use.
12. A nozzle as claimed in claim 11 wherein the aerodynamically
smooth spline is between radii at three fixed axial locations along
the nozzle surface from the nozzle exit plane, one radii within a
distance upstream equivalent to 20% of the nozzle diameter, one
radii at a point of maximum undulation amplitude within an upstream
distance equivalent to 15% of the nozzle diameter and one radii
within 10% of the nozzle exit plane diameter.
13. A nozzle as claimed in claim 1 wherein the convergent-divergent
ratio is in the range of 1 to 1.01.
14. (canceled)
15. A nozzle as claimed in claim 1 wherein the nozzle comprises a
bypass nozzle of a gas turbine engine with the undulations on an
inner surface of an outer wall or outer surface of an inner wall of
the bypass nozzle.
16. A nozzle as claimed in claim 1 wherein the nozzle is a core
nozzle of a gas turbine engine with undulations on an inner surface
of an outer wall or outer surface of an inner wall of a cone
nozzle.
17. A nozzle as claimed in claim 1 wherein the undulations are
symmetrically regularly circumferentially distributed about the
nozzle surface.
18. A nozzle as claimed in claim 1 wherein the undulations are
asymmetrically or irregularly circumferentially distributed about
the nozzle surface.
19. A nozzle as claimed in claim 1 wherein the undulations are at
different axial positions relative to an exit plane of the
nozzle.
20. A nozzle as claimed in claim 1 wherein the undulations have
groups of differing amplitudes circumferentially or axially in the
nozzle surface.
21. A nozzle as claimed in claim 17 wherein the undulations may
have different groups of axial length or width relative to each
other.
22. A nozzle as claimed in claim 1 wherein the nozzle surface has
an edge with serrations or tabs.
23. A nozzle as claimed in claim 22 wherein the undulations are
arranged reciprocally with the serrations for additional variation
in convergent-divergent ratio range.
24. A nozzle as claimed in claim 22 wherein the tabs are deployable
for noise reduction.
25. A nozzle as claimed in claim 22 wherein alternate tabs are
deployable for noise reduction.
26. A nozzle as claimed in claim 24 wherein undulations are formed
on one or more of the tabs.
27. A nozzle as claimed in claim 1 wherein the undulations are
transformable between a deployed position and a non-deployed
position, the non-deployed position being less aerodynamically
obtrusive than the deployed position.
28. A nozzle as claimed in claim 27 wherein the undulations are
transformable to a second deployed position, between the deployed
and non-deployed positions.
29. A nozzle as claimed in claim 27 wherein the undulations
comprise a shape memory material element.
30. A nozzle as claimed in claim 29 wherein the shape memory
material element comprises two layers of shape memory material,
each layer having different switch temperatures and capable of
deploying in a first shape and a second shape, the second shape
having a greater amplitude that the first shape.
31. A nozzle as claimed in claim 29 wherein the shape memory
material element comprises two layers, one layer of shape memory
material and the other layer of resilient material to provide a
spring force to the element.
32. A nozzle as claimed in claim 1 wherein the undulations are
integral with the nozzle.
33. A nozzle as claimed in claim 1 wherein the undulations are
created by attached elements individually or as part of an assembly
secured to the nozzle surface.
34. A nozzle as claimed in claim 1 wherein the undulations are
variable in terms of amplitude or position or distribution in the
nozzle surface.
35. A nozzle as claimed in claim 27 wherein such variation is by
use of inflatable features with the nozzle surface or deployable
mechanical portions of the nozzle surface.
36. A nozzle as claimed in claim 1 wherein the number of
undulations is in the range of one to forty two undulations
distributed about the nozzle surface.
37. A gas turbine engine incorporating a nozzle as claimed in claim
1.
38. A nozzle as claimed in claim 1 wherein the convergent-divergent
ratio is 1.008.
39. A nozzle as claimed in claim 1 wherein the number of
undulations distributed about the nozzle surface is twenty.
Description
[0001] The present invention relates to gas turbine engine noise
and more particularly to jet noise under cruise conditions.
[0002] The general stages of gas turbine operation are known. In
particular, it will be understood that there is a downstream jet
created as the various gas streams are forced out of the engine in
order to create propulsion. Inherently such jet flows create noise
as the jet shear layer breaks down. This shear layer breakdown
along with other factors such as the presence of shock waves
produces noise.
[0003] Clearly, noise is a detrimental factor with respect to gas
turbine engine operation. Thus, there is a continuing objective to
reduce engine noise in all phases of engine operation including
whilst an engine is propelling an aircraft through the air at
altitude and under cruise conditions.
[0004] At cruise conditions the nozzle of a jet engine is not
perfectly expanded. As a result a shock structure occurs in the
jet. This shock structure is strongest near the nozzle exit but
extends several diameters downstream of the nozzle in a repeat but
progressively fading shock cell pattern.
[0005] Shock cell noise is generated as the turbulence of the jet
shear layer passes through and interacts with the shock structure
of the jet (see Harper-Bourne, M. and Fisher, M. J., 1973, "The
Noise from Shock Waves in Supersonic Jets", Proceedings (No. 131)
of the AGARD Conference on
[0006] Noise Mechanisms, Brussels, Belgium). If one considers a
cross section through half of a typical high bypass ratio civil
engine nozzle system, with the bypass jet imperfectly expanded,
there is a shock structure set up in the bypass stream. The shear
layer between the flight stream and the bypass stream becomes
turbulent as it develops and the turbulence that results convects
through the shock structure generating noise. The region in which
the shock cell noise is generated may be several nozzle diameters
downstream of the nozzle exit plane.
[0007] Shock cell noise may be reduced by previous serrations at
the nozzle exit that enhance the mixing of the shear layer so that
the turbulence intensity is lower in regions where the turbulence
interacts with the shock structure.
[0008] Noise suppression by serrations has been demonstrated but
this has typically been for environmental noise at take off or
landing conditions.
[0009] Generally, serrated nozzles consist of flaps or tabs added
to or cut out of a nozzle so as to generate circumferential flow
non-uniformities. The circumferential flow non-uniformities enhance
mixing of the jet thereby breaking up coherent structures leading
to lower noise.
[0010] In order for serrations to reduce the noise of a nozzle they
need to disturb the nozzle flow. This typically requires the
serrations to deflect the flow by having some incidence or
insertion to the flow. This results in increased drag and an
associated loss of performance. The performance loss and noise
reduction mechanism are inherently linked for serrations.
[0011] The increased surface area of serrations also increases the
drag. Increased surface area also increases overall weight.
[0012] In accordance with the present invention there is provided a
nozzle for a gas turbine engine, the nozzle comprising a nozzle
surface including a plurality of undulations to vary available
cross sectional area across the nozzle between the nozzle surface
and an opposed surface of the nozzle over a desired
convergent-divergent ratio range for noise control of a jet passing
through the nozzle in use.
[0013] Additionally, the undulations also provide variation in the
angle of flow of the jet passing through the nozzle.
[0014] Generally, the variation in cross sectional area is adjusted
to provide stimulation in mixing of a shear layer of the jet for
relative noise reduction in comparison with that without mixing of
the shear layer of the jet. Additionally, the undulations alter the
repeat cycle and/or provide variation of intensity of shock cells
generated by the jet.
[0015] Generally, the undulations comprise bumps formed in the
nozzle surface.
[0016] Possibly, the undulations are sinusoidal in a
circumferential direction or in a plane perpendicular to a jet flow
direction in use.
[0017] Preferably, the undulations each comprise maximum amplitude,
the maximum amplitude situated a distance between 2% and 15% of the
nozzle diameter along the nozzle surface upstream from the nozzle
exit plane.
[0018] Preferably, the maximum amplitude situated a distance
equivalent to 6% of the nozzle diameter upstream of nozzle exit
plane.
[0019] Alternatively, the undulations each comprise a maximum
amplitude, the maximum amplitude situated within a distance
equivalent to +/-2% of the nozzle diameter along the nozzle surface
from the nozzle throat plane.
[0020] Preferably, the approximate cross-section shape of the
undulations are from the group comprising triangular, trapezoidal,
part-circular, sinusoidal and asymmetric-sinusoidal.
[0021] Alternatively, the undulations have an aerodynamically
smooth gradual spline in an axial direction of jet flow in use.
Typically, the aerodynamically smooth spline is between radii at
three fixed axial locations along the nozzle surface from the
nozzle exit plane, one radii within a distance upstream equivalent
to 20% of the nozzle diameter, one radii at a point of maximum
undulation amplitude within an upstream distance equivalent to 15%
of the nozzle diameter and one radii within 10% of the nozzle exit
plane diameter.
[0022] Generally, the convergent-divergent ratio is in the range of
1 to 1.01 preferably 1.008.
[0023] Typically, the undulations have amplitude 39 in the range
0.1-2.0% of the nozzle exit diameter.
[0024] Typically, the nozzle comprises a bypass nozzle of a gas
turbine engine with the undulations on an inner surface of the
outer wall and/or outer surface of an inner wall of the bypass
nozzle.
[0025] Possibly, the nozzle is a core nozzle of a gas turbine
engine with undulations on an inner surface of the outer wall
and/or outer surface of the inner wall of the cone nozzle.
[0026] Possibly, the undulations are symmetrically regularly
circumferentially distributed about the nozzle surface.
Alternatively, the undulations are asymmetrically and/or
irregularly circumferentially distributed about the nozzle surface.
Further, the undulations may be at different axial positions
relative to an exit plane of the nozzle. Possibly, the undulations
have groups of differing amplitudes circumferentially and/or
axially in the nozzle surface. Additionally, the undulations may
have different groups of axial length and/or width relative to each
other.
[0027] Possibly, the nozzle surface has an edge with serrations or
tabs. Additionally, the undulations may be arranged reciprocally
with the serrations for additional variation in
convergent-divergent ratio range.
[0028] Alternatively, the tabs are deployable for noise reduction.
Possibly, alternate tabs are deployable for noise reduction and the
undulations are formed on any one or more of the tabs. Preferably,
the undulations are transformable between a deployed position and a
non-deployed position, the non-deployed position being less
aerodynamically obtrusive than the deployed position. Further, the
undulations may be transformable to a second deployed position,
between the deployed and non-deployed positions.
[0029] Possibly, the undulations comprise a shape memory material
element. Alternatively, the shape memory material element comprises
two layers of SMM material, each layer having different switch
temperatures and capable of deploying in a first shape and a second
shape, the second shape having a greater amplitude that the first
shape. Alternatively, the shape memory material element comprises
two layers, one layer of SMM material and the other layer of
resilient material to provide a spring force to the element.
[0030] Preferably, the undulations are integral with the nozzle.
Alternatively, the undulations are created by attached elements
individually or as part of an assembly secured to the nozzle
surface.
[0031] Alternatively, the undulations are variable in terms of
amplitude and/or position and/or distribution in the nozzle
surface.
[0032] Alternatively, such variation is by use of inflatable
features with the nozzle surface or deployable mechanical portions
of the nozzle surface.
[0033] Preferably, the number of undulations is in the range one to
forty-two and preferably twenty undulations distributed about the
nozzle surface.
[0034] Embodiments of the present invention will now be described
by way of example and with reference to the accompanying drawings
in which;
[0035] FIG. 1 is a schematic half cross-section of a typical
turbine engine nozzle;
[0036] FIG. 2 is a schematic perspective view of a prior art
turbine engine nozzle incorporating serrations;
[0037] FIG. 3 is a graphic core representation illustrating
available flow cross-sectional area against axial distance for a
convergent nozzle and a convergent-divergent nozzle as illustrated
in FIG. 1;
[0038] FIG. 4 is a schematic perspective view of a first embodiment
of a nozzle incorporating undulations in accordance with the
present invention;
[0039] FIG. 5 is a schematic enlarged view of a portion of a nozzle
surface in accordance with the present invention;
[0040] FIG. 5a is a section A-A through the nozzle shown in FIG.
5;
[0041] FIG. 5b is a section A-A through the nozzle shown in FIG. 5
showing alternative embodiments to FIG. 5a;
[0042] FIG. 6 is a schematic cross-section showing variation in
cross-sectional area between a nozzle surface and an opposed
surface of a nozzle in accordance with the present invention;
[0043] FIG. 7 is a graphic representation of variation in available
flow area across a nozzle against axial distance in accordance with
the present invention;
[0044] FIG. 8 provides graphical illustrations with respect to
variation in convergent-divergent ratio and available flow area
against azimuthal angle (circumferential direction) for an
undulation in accordance with the present invention;
[0045] FIG. 9 provides a graphic illustration of variation in
overall SPL against jet Mach number;
[0046] FIG. 10 provides a schematic illustration of a first
alternative undulation distribution in accordance with the present
invention;
[0047] FIG. 11 provides a schematic perspective view of a second
alternative undulation distribution in accordance with the present
invention;
[0048] FIG. 12 provides a schematic perspective view of a third
alternative undulation distribution in accordance with the present
invention;
[0049] FIG. 13 provides a schematic perspective view of a fourth
alternative embodiment of an undulation distribution in accordance
with the present invention;
[0050] FIGS. 13a and b provide schematic perspective view of
alternative embodiments of the fourth embodiment shown in FIG. 13
and in accordance with the present invention;
[0051] FIG. 14 provides a schematic perspective view of a fifth
alternative undulation distribution in accordance with the present
invention;
[0052] FIG. 15 provides a schematic perspective view of a sixth
alternative embodiment of an undulation distribution in accordance
with the present invention;
[0053] FIG. 16 provides a schematic perspective view of a seventh
alternative embodiment of an undulation distribution in accordance
with the present invention;
[0054] FIG. 17 provides a schematic perspective view of an eighth
alternative embodiment of an undulation distribution in accordance
with the present invention;
[0055] FIG. 18 provides a schematic perspective view of a ninth
alternative embodiment of an undulation distribution in accordance
with the present invention; and
[0056] FIG. 19 provides a schematic part cross-section illustrating
location of undulations in accordance with the present
invention.
[0057] FIG. 20 provides a schematic perspective view of a tenth
alternative embodiment of an undulation distribution combined with
a nozzle having deployable tabs in accordance with the present
invention.
[0058] FIG. 21 is a view in the direction of arrow A on FIG.
20.
[0059] FIGS. 22 and 22A is a part section A-A through an
alternative embodiment than the nozzle shown in FIG. 5.
[0060] FIG. 1 shows a half cross section through a typical high
bypass ratio civil aero turbine engine exhaust nozzle 1. The bypass
stream, that is to say the outer stream that passes only through
the fan has a nozzle exit area 2. The area 2 is thus available for
flow to pass through and exit a nozzle 3. In the nozzle 3 there is
a nozzle throat area 4 which is the minimum area for the flow to
pass through at any point in the bypass nozzle 1. This limits the
possible mass flow rate so that there may be choking of the flow
with expansion after the throat area 4.
[0061] The throat area 4 and exit area 2 may be of different
magnitudes and may occur at axially separated positions. A
convergent nozzle is one in which a flow area 5 is continually
decreasing in a direction of flow A (or axial direction) and
therefore one in which the exit area 2 is the minimum area and thus
also the throat area 4. A convergent-divergent nozzle (FIG. 1) is
one in which the throat area 4 is upstream of the exit area 2 so
that the flow area 2 decreases in the direction A of flow until it
reaches a minimum point at the throat area 4 and then increases to
the exit area 2.
[0062] The respective variations in the available flow area is
depicted graphically against axial distance in the flow direction A
in FIG. 3. Thus, as can be seen, with a convergent nozzle the
relationship is given by line 10, whilst with a
convergent-divergent nozzle the relationship is given by line 11.
The convergent-divergent ratio of a nozzle is therefore given by
the ratio of the exit area 2 in comparison with the throat area 4
(FIG. 1). The convergent nozzle has a convergent-divergent ratio of
unity (1) whilst, in general, a convergent-divergent nozzle has a
convergent-divergent ratio greater than 1. With a typical engine
used for aircraft, the nozzle configurations have a
convergent-divergent ratio in the range 1.00 to 1.02 and for the
Applicant's commercial production engines typically a range between
1.00 and 1.01.
[0063] FIG. 2 is included to provide an illustration of a typical
prior serrated nozzle in order to provide noise suppression. As can
be seen, the nozzle 20 has serrations 21 which can take the form of
flaps or tabs added to or cut from the nozzle in order to generate
circumferential flow non-uniformities, which as indicated above
break up coherent structures in the jet flow in order to give rise
to noise suppression. In effect, the serrations 21 deflect the flow
so as to enhance turbulent mixing thereby suppressing noise.
However, as indicated above, serrations can add significantly to
cost, weight and drag upon the engine reducing efficiencies.
[0064] The present nozzle provides a circumferentially varying
convergent-divergent nozzle by incorporating a number of
undulations or bumps into at least one nozzle surface. Typically,
twenty sinusoidal and evenly spaced bumps are machined into an
inner surface of the outer wall of a bypass nozzle such that the
radius varies through the pitch of the sinusoidal bumps. As will be
described later, a number of varying alternative embodiments will
provide undulations and bumps in differing patterns and
distributions in accordance with particular operational
requirements. With regard to the first embodiment described, as
indicated sinusoidal oscillations in the form of bumps in the
circumferential direction about the nozzle will be provided.
[0065] Referring again to FIG. 1 and also seen in FIG. 6, the bumps
40 comprise an upstream surface 37, a point of maximum amplitude 38
and a downstream surface 39. The undulations or bumps 40 will
generally have a smooth spline in the axial direction with radii at
three fixed axial locations along the nozzle surface. Defining the
upstream surface 37 is a first fixed axial radii defined at a
position up to 20%, but in a preferred example 10% of the nozzle
exit plane diameter 2, upstream of the nozzle throat position 4. A
second fixed radius is provided between the upstream and downstream
surfaces 37, 39 and defines the maximum bump amplitude 38. The
second fixed radius is positioned approximately 6% of the nozzle
exit plane diameter upstream of the nozzle exit position, but in
other embodiments may be positioned between 2% and 15% of the exit
plane diameter upstream of the nozzle exit position. A third fixed
radius preferably is positioned at the nozzle exit plane itself,
but may be positioned up to 10% of the nozzle exit plane diameter
upstream of the exit plane.
[0066] As indicated above in a first embodiment of a nozzle
undulations or bumps will be provided circumferentially in a
regular distribution pattern. FIGS. 4 and 5 illustrate the first
embodiment of the invention with FIG. 4 providing a schematic
perspective illustration of a nozzle whilst FIG. 5 provides an
enlarged view of a section of the nozzle bypass inner wall
surface.
[0067] Undulations or bumps 40 are regularly circumferentially
distributed about an inner surface 41 of an outer wall 42 of a
bypass nozzle of an engine 43. Thus, as can be seen in both FIGS. 4
and 5, the effect of the undulations or bumps 40 is to provide a
convergent-divergent nozzle form along the axial length of the
undulation or bumps 40 with generally areas between the undulations
or bumps 40 being flatter and therefore creating a convergent
nozzle format. In such circumstances, as described previously,
noise suppression is achieved through mixing of the shear layer so
that the turbulence intensity is lower in the regions where the
turbulence interacts with the shock structure from the nozzle.
Furthermore, the shock cell repeat pattern will vary across the
undulations again leading to noise suppression.
[0068] Referring to FIG. 5a, which shows the first embodiment in
more detail, generally the undulations 40 are sinusoidal and have
amplitude 39 in the range 0.1-3.0% of the nozzle exit diameter, but
preferably in the range between 0.3% and 1.5%. Typically, the
undulations have maximum amplitude point 38 positioned within
.+-.2% nozzle exit diameter of the nozzle throat position 4. The
circumferential extent 26 of the undulations 40 is defined by a
length equivalent to between 1.degree. and 45.degree. and the
angular spacing 27 between maximum amplitude points 38 is between
2.degree. and 90.degree. , i.e. the total number of bumps is
between 180 and 4, however, a preferred number of maximum amplitude
points is between 12 and 45.
[0069] Referring to FIG. 5b, alternate shapes of bumps 40a-e are
shown. These alternately shaped bumps 40a-e may be either attached
by suitable means to an existing smooth nozzle surface or may be
machined into the nozzle wall. Where the bumps are machined into
the nozzle wall they may either be proud of the wall surface 23 (as
shown) or machined in as for the sinusoidal wave form having peaks
38 and troughs 49 defined by the respective inverse bump shape.
Here the undulations 40 are preferably shaped in cross-section as
shown by bump 40a, which comprises maximum amplitude 38 defined by
a radius 24 and blend radii 25 smoothing the shape into a
circumferential profile 23 of the nozzle wall 42. The maximum
amplitude 38 is in this case the radial height above the existing
or original wall surface 23.
[0070] FIG. 5b also shows alternative shapes the undulations or
bumps may take without departing from the scope of the invention.
Bump 40b is generally trapezoidal in cross-section; bump 40c is
triangular; bump 40d is defined by a constant radius
(part-circular) and bump 40e is an asymmetric version of bump 40a
and similarly defined by three radii but radius 25' is greater than
25''.
[0071] Each bump 40a-e, in their respective array of bumps, are
angularly spaced apart a corresponding distance 27 peak-to-peak
(38) dependent on the number required around the nozzle's
circumference.
[0072] For a rectangular nozzle or other non-circular nozzles the
spacing of the bumps peak or maximum amplitude point 38 is the
total length of side divided by the number of bumps.
[0073] In such circumstances, for a scale model tested by the
Applicant, the nozzle exit diameter was 58 millimetres, the first
radius is at approximately 11.2 millimetres, the second is at a
position 3.6 millimetres upstream of the nozzle exit plane and the
final radius at the exit plane itself. In such a situation, the
amplitude of the undulations or bumps is in the order of 0.8
millimetres (1.38% of the nozzle exit diameter) with an axial
position of maximum amplitude as indicated at 3.6 millimetres
upstream of the nozzle exit plane (6% of the nozzle diameter). The
scale model comprised 20 bumps 40 (FIG. 5a) evenly spaced at
18.degree. intervals. A 2 db noise reduction was achieved over a
similar nozzle without bumps.
[0074] In such circumstances, for a production gas turbine engine
of the Applicant's, the nozzle diameter is 1450 millimetres, having
a bump amplitude in the order of 4.5 millimetres (0.31% of the
nozzle exit diameter) with the first radius at approximately 280
millimetres upstream of the nozzle exit plane, the second radius at
approximately 90 millimetres upstream of the nozzle exit plane and
the third radius at the nozzle exit plane. This nozzle comprised 20
sinusoidal bumps 40a (FIG. 5a) evenly spaced at 18.degree.
intervals.
[0075] However, for other applications and depending on specific
engine operating circumstances different distributions, amplitudes
and axial lengths may be used within the ranges indicated
throughout this specification and depending on particular noise
reduction requirements.
[0076] FIGS. 6 and 7 illustrate respectively a schematic
cross-section through one of the undulations or bumps 40 in the
inner nozzle surface 41 of an outer wall of a bypass nozzle 42 with
respect to an opposed surface 44 of the nozzle (FIG. 6) and in FIG.
7 a graphic representation illustrating differences in the
available nozzle flow area relative to axial distance along the
bump 40. Thus, as can be seen, the available flow area 45 between
the nozzle surface 41 and the opposed surface 44 is varied in the
circumferential direction through the bump and the spaces in the
area 41 between the bumps 40. This is illustrated in FIG. 7 through
the representative relationship lines 46 showing the variation in
available flow area 45 with axial distance in the direction 46 at
different circumferential positions across one bump.
[0077] The above circumferential variation in available area is
further illustrated across, that is to say circumferentially
around, the nozzle in FIG. 8 where variations in the
convergent-divergent ratio as well as the available flow area 45
are shown relative to the azimuthal angle across the undulation or
bump 40. As can be seen, with a sinusoidal undulation or bump 40, a
similar sinusoidal relationship is provided in the graphic
representations depicted in FIG. 8. In FIG. 8a, the
convergent-divergent ratio is depicted against circumferential
angle across an undulation or bump without a flat space section
between undulations such that there is a continuous sinusoidal
variation from one bump to the next such that the
convergent-divergent ratio oscillates sinusoidally around an
average value 47, but it will be understood where there is an
undulation of bump formed with relatively flat spaces either side
then a half sinusoidal relationship will be provided in terms of
the variation in convergent-divergent in ratio as the bump or
undulation amplitude moves into and out of the nozzle across the
circumferential width of the bump or undulation. Similarly, the
available flow area 46 will vary sinusoidally across the
circumferential width of the bump or undulation and relative to a
maximum exit area 48 defined at the exit plane of the nozzle.
[0078] Referring to FIG. 9, overall shock cell noise 50 is
dependent on Mach number 51 of jet. A fully expanded nozzle has
minimum noise (52) with over expanded and under expanded nozzles
having greater noise (Tam, C. K. W. and Tanna, H. K., Journal of
Sound and Vibration, 1982, 81(3), 337-358) shown in FIG. 9.
[0079] For a given pressure ratio there exists a
convergent-divergent ratio to give a fully (perfectly) expanded jet
and this will have the minimum noise as there will be no shock
structure established. A fully expanded nozzle does not necessarily
meet all operational requirements and so it is impractical for a
fixed geometry nozzle to achieve a fully expanded jet at cruise
conditions. For information solid line 53 shows the effect of a
convergent-divergent nozzle whilst broken line 54 shows a simple
convergent nozzle.
[0080] If some sectors of a nozzle operate at perfectly expanded
conditions then no shock structure will be formed in those sectors
and the mechanism for shock cell noise generation will disappear
locally. The closer a sector of the nozzle is to being perfectly
expanded the weaker the shock structure and the less shock cell
noise will be generated. The undulations provide a range of
available flow areas to increase the likelihood of a perfect or
near perfect expansion for noise suppression.
[0081] For an imperfectly expanded supersonic jet from a nozzle of
fixed geometry, the angle of the flow relative to the axis of the
jet in the region just behind the nozzle exit is a function of the
nozzle pressure ratio. This is a result of the flow emerging from
the nozzle expanding to match the conditions outside of the nozzle.
Moreover, the mass flow of the jet is fixed by the area of the
nozzle throat. The final flow area of the jet (outside of the
nozzle) is dependent on the mass flow and the freestream
conditions. The freestream conditions are very nearly
circumferentially uniform and so the flow area of the jet is
proportional to the throat area these being linked by the mass
flow. A circumferential variation in the throat area thus leads to
a circumferential variation in the final flow area of the jet. This
mimics the effects of serrations and produces a circumferentially
non-uniform flow field.
[0082] The circumferential variation in convergent-divergent ratio
as a result of a circumferential variation in throat area thus
produces a circumferentially non-uniform flow field downstream of
the nozzle exit. This enhances mixing of the shear layer reducing
the extent of turbulent flow. The interaction of the turbulence and
the shock cell structure responsible for the shock cell noise is
thus further reduced as the turbulence is reduced.
[0083] In cases where the circumferential variation of
convergent-divergent ratio is achieved with a circumferentially
constant throat area (i.e. circular throat and sinusoidal variation
of exit area) the circumferential non-uniformity in flow would be
reduced but the circumferential variation in shock strength would
persist and this would still reduce shock cell noise.
[0084] By contrast to serrations, circumferential variation of
convergent-divergent ratio avoids the performance degradation due
to tabs inserted into flow with incidence as this increases the
drag on the serration. Serrations and tabs also have increased
surface area exposed to the flow and this increases drag. The
length (perimeter) of the trailing edge of the nozzle is a minimum
for a circular nozzle in a plane perpendicular to the engine axis.
The application of serrations or tabs increases the length of the
nozzle trailing edge and thus increases the base drag.
[0085] The mixing achieved by varying the throat area
circumferentially with undulations is as a result of manipulating
the shock waves rather than deflecting the flow. Manipulating the
shock waves to change flow directions is a near lossless process
unlike deflecting the flow.
[0086] Serrations necessarily add weight to the design. The
circular planar nozzle exit permitted by this invention is the
minimum weight design. Mechanical challenge of tab and associated
stress concentration are avoided. However, serration tabs suppress
shock cell noise by enhanced mixing of shear layer.
[0087] A number of alternative embodiments to the regular
sinusoidal or other shaped undulations presented circumferentially
about the nozzle can be provided in accordance with the invention.
Thus, the undulations or bumps may be provided on an inner surface
of the outer wall of a bypass nozzle as described above, or
alternatively the bumps or undulations can be provided on the outer
surface of an inner wall of the bypass nozzle or bump undulations
provided on the inner surface of the outer wall of the core nozzle
or bumps and undulations provided on the outer surface of the inner
wall of the core nozzle or combinations of these configurations. In
the specific embodiment described above, it will be appreciated
that there is a circular nozzle exit with sinusoidal variation in
available throat flow area, but alternatively there could be a
circular throat area with variation in the exit area by corrugating
the nozzle exit area edge to create undulating correlations.
Furthermore, there may be variation in the available throat flow
area and variation in the exit area in such a way that leads to
undulations that enhance shear layer turbulence and mixing as
described above for noise reduction. Additionally, although
described with regular spacing of the sinusoidal bumps or
undulations in the embodiment described above, it will also be
understood that there may be a range of different bump or
undulation distributions as described below in a number of
alternative embodiments.
[0088] The particular combination of bump or undulation position in
relation to distribution as well as exit plane area will depend
upon particular engine design requirements.
[0089] FIG. 10 illustrates a first alternative embodiment of a
nozzle in which bumps 140 are arranged with a regular distribution
about an inner nozzle surface 141 of an outer wall of a bypass
nozzle. The bumps have undisturbed regions 143 between them in
which the nozzle therefore acts as a simple convergent nozzle in
these parts with the bumps 140 providing the convergent-divergent
variation in available flow throat area as required for noise
suppression. It would be appreciated that an opposed surface (not
shown) of the nozzle may include itself undulations or bumps which
may directly oppose the bumps 140 or interleave with those bumps
140 such that these bumps in the opposed surface directly oppose
the undisturbed regions 143.
[0090] FIG. 11 illustrates a second alternative embodiment of a
nozzle. Thus, bumps 240 are arranged with irregular spacing in an
inner nozzle surface 241 of an outer wall 242 of a bypass nozzle.
By such irregular spacing of the bumps 240, it is possible that
there is further disturbance with respect to circumferential modes
for shear layer turbulence or there may be variation in the noise
suppression level at certain directions of the nozzle in comparison
with others dependent upon operational requirements.
[0091] FIG. 12 illustrates a third alternative embodiment of the
present nozzle. Thus, bumps 340 are located in groups or
individually in an inner nozzle surface 341. Thus in a similar
fashion to that with regard to the second alternative embodiment
depicted in FIG. 11, there is circumferential variation in the
distribution of the bumps 340a compared to bumps 340b in order to
again disturb the circumferential modes and vary the noise
suppression level at different directions of the nozzle.
[0092] FIG. 13 illustrates a fourth alternative embodiment of the
present nozzle in which undulations or bumps are provided at
different axial positions as compared to circumferential conditions
in previous embodiments. Thus, bumps 440 are provided in an inner
nozzle surface 441 of an outer wall 442 of a bypass nozzle. It will
be noted that undulations or bumps 440a are essentially based at an
exit plane edge 443 of the nozzle, whilst bumps or undulations 440b
are slightly displaced from that exit edge 443, whilst the
undulations or bumps 440c are even further displaced from the edge
443. Such an arrangement will provide a variation in the
convergent-divergent ratio over a broader axial length of the
nozzle and therefore provide different operational performance
compared to previous embodiments.
[0093] For each of the bumps 440a,b,c,d their first radius, which
defines their upstream surface, are located at respectively 0%, 5%,
2.5% and 7.5% of the nozzle exit diameter, upstream of the nozzle
throat. However, it should be appreciated that each of the bumps
440a,b,c,d are located and sized within the ranges defined
hereinbefore.
[0094] FIGS. 13a and 13b show alternative embodiments to that shown
in FIG. 13, where a number of adjacent bumps (440a,b,c,d in FIG.
13) are merged circumferentially to form one or more larger and
therefore more complex undulations 443, 444. For each of these
embodiments the bumps 443, 444 are generally sinusoidal or part
sinusoidal in a circumferential sense and comprise their first
radii axially located between 0% and 7.5% of the nozzle exit
diameter, upstream of the nozzle throat. The second and third
cross-section radii are accordingly located relative to the first
radius at any given axial cross-section through each bump 443, 444.
In these embodiments the maximum amplitude is constant (i.e. it
forms a ridge of maximum amplitude) except at the circumferential
extents, where each bump blends out to the nozzle wall. It should
be appreciated that in other embodiments the maximum amplitude may
be varied circumferentially along the bumps 443, 444.
[0095] FIG. 14 illustrates a fifth alternative embodiment of the
present nozzle in which bumps 540 are provided which have different
amplitudes at different circumferential and axial locations in an
inner nozzle surface 541 of an outer wall 542. In such
circumstances, the different amplitudes for the bumps or
undulations 540 may provide differing levels of noise reduction in
different directions of the nozzle, and through changing the
available flow area throat could improve mixing to allow further
noise control. As can be seen, bump or undulations 540b have a
greater amplitude than bumps or undulations 540a, 540c and have a
greater axial length whilst undulations 540b, 540c respectively
have different circumferential widths and axial lengths compared to
each other and undulation 540a.
[0096] FIG. 15 illustrates a sixth alternative embodiment of a
nozzle in which bumps or undulations 640 are presented in an inner
nozzle surface 641 of an outer wall 642 of a bypass nozzle. Thus,
the bumps or undulations 640 are grouped in regions with other
undisturbed regions between them such that differing levels of
noise suppression will be provided and therefore quieter areas
achieved relative to normal nozzle operation. Such regionalisation
of the bumps or undulations will provide a similar effect to a
regular spacing and the bumps or undulations described with regard
to the third alternative embodiment (FIG. 12) above.
[0097] FIG. 16 illustrates a seventh alternative embodiment of a
nozzle. Thus, undulations or bumps 740 are presented in a serrated
outer wall 742 of a bypass nozzle. The bumps or undulations 740 are
again presented on an inner nozzle surface 741 of the wall 742. As
described previously, the bumps or undulations could be provided in
an inner or outer wall of a nozzle, but FIG. 16 only illustrates
provision of the bumps or undulations 740 in the inner surface 741
of the outer wall 742. In such circumstances, the effects of
serrations 743 are enhanced by this circumferential change in the
convergent-divergent ratio created by the bumps and undulations
740. Such an arrangement may provide enhanced noise suppression,
although as described previously, provision of serrations 743 may
add to drag and other factors with respect to engine operation.
[0098] FIG. 17 provides a further eighth alternative embodiment of
the present nozzle. Thus, as with the seventh alternative
embodiment depicted in FIG. 16, bumps or undulations 840 are
provided upon an inner nozzle surface 841 of an outer wall 842 of a
bypass nozzle. In comparison with the embodiment depicted in FIG.
16, the bumps or undulations 840 circumferentially change the
convergent-divergent, ratio, but do not extend into tab or
serration portions 843, but are upon a fixed portion of the nozzle
prior to such serrations 43. Again, such an approach will provide
an alternative for particular operational requirements in terms of
noise suppression and shear layer turbulence.
[0099] FIG. 18 illustrates a further ninth alternative embodiment
of a nozzle. Thus, in comparison with the embodiment depicted in
FIG. 17, bumps or undulations 940 are again provided in an inner
nozzle surface 941 of an outer wall 942 of a bypass nozzle.
However, the bumps or undulations 940 still remain prior to
serrations 942 in the exit plane of the nozzle and in comparison
with the embodiment depicted in FIG. 17, these bumps 940 are out of
phase with the serrations 943 in order to provide a further
enhancement or variation in noise suppression performance dependent
upon operational requirements.
[0100] It will be understood that noise and therefore noise
suppression requirements will vary dependent upon an engine's
operational state. In such circumstances, bumps or undulations in
accordance with the present invention may be variable dependent
upon operational conditions or desired requirements. In such
circumstances, the bumps or undulations may have a shaped memory
alloy type function and therefore vary according to temperature or
other requirements in terms of amplitude and shape for variation in
the turbulence created in the shear layer for noise suppression.
Where possible, the bumps or undulations may be arranged to be
electively deployable through use of inflation or deflectable
mechanical panels or otherwise in order to change their amplitude,
both in terms of inward deflection as well as axial length and
circumferential spacing for operational requirements.
[0101] FIG. 19 provides a schematic illustration of an engine
nozzle arrangement in which the respective nozzle surfaces are
illustrated. Thus, a bypass nozzle is provided by an outer wall
1002 and an inner wall 1003 such that surfaces 1001 and 1004 may
incorporate bumps or undulations in accordance with the present
invention in order to vary the convergent and divergent ratio
across the nozzle surfaces 1001, 1004 in accordance with the
present invention. Similarly, an outer wall 1005 and an inner wall
1006 present opposing surfaces 1007, 1008 of a core nozzle. These
nozzle surfaces 1007, 1008 may also incorporate undulations or
bumps in accordance with the present invention in order to vary the
convergent-divergent ratio across the core nozzle. In such
circumstances, additional noise suppression may be provided by
creating turbulence in the shear layer between jets for noise
suppression as described above.
[0102] It will be understood that undulations particularly in a
core nozzle will be subject to erosion at high temperatures, thus
provision may be made for replacement of undulations as securable
elements or assembly to a nozzle surface.
[0103] Alternative embodiments and modifications of the present
invention will be understood by those skilled in the art. Thus, for
example, rather than providing smooth splines for undulations or
bumps as described above, more angular bumps or undulations may be
provided. For example, a triangular cross-section bump, defining an
apex at its point of maximum amplitude, may be used. Furthermore,
there may be axial cycling in the bump or undulation amplitude
axially or circumferentially if required in order to create mini
turbulence in the jet flow for noise suppression.
[0104] In FIG. 20, the nozzle 42 comprises an arrangement of
deployable noise reducing tabs 80, 82 which are described in U.S.
Pat. No. 6,813,877, the teachings of which are incorporated herein.
Briefly, circumferentially alternate tabs 80 are rigidly fixed in a
`deployed` position as shown in the figure, where they interact
with the gas streams to enhance mixing out the noise creating shear
layer between gas streams. The deployable tabs 82 comprise shape
memory material and are moveable between a deployed position as
shown in FIG. 20 and a non-deployed position, where they are
aligned and abutting tabs 80. During take-off and climb the tabs 82
are deployed, angled radially outwardly, for noise reduction
purposes and the exit area of the nozzle 12 is enlarged. This
enlargement reduces the velocity of the gas stream issuing from the
bypass duct 12 and which intrinsically reduces exhaust noise. At
aircraft cruise the tabs 82 are in the non-deployed position, where
adjacent tabs' edges 88, 90 are in sealing engagement with one
another, and the exit area is therefore reduced. This reduction
increases the velocity of the exhaust gas stream and improves
engine efficiency.
[0105] As the bumps 40 are designed primarily for reducing aircraft
cabin noise at cruise, the exhaust exit plane 36 in this case is
defined by a downstream edge 86 of the tabs 80, 82 when in their
non-deployed position. The bumps 40 are still positioned within the
range of positions specified hereinbefore and may therefore be
situated on one or more of the tabs' radially inner surface,
depending on the axial length of the tab and the
convergent-divergent ratio.
[0106] FIG. 21 show a bump 40 situated on each of the tabs 80, 82
around the nozzle 42 circumference. The shape and configuration of
the bumps 40 are as hereinbefore described.
[0107] In further embodiments of the present invention shown in
FIGS. 22, 22A and 23, the bumps 40 are deployable and preferably
comprise shape memory material as a means for actuating the bumps
between a deployed position 40' and a non-deployed position 40''.
Shape memory material (SMM) is well known in the industry and is
not discussed further except that its operation is similar to that
for the deployable nozzle tabs as disclosed by the present
Applicant in U.S. Pat. No. 6,813,877, the teachings of the use of
shape memory material are incorporated herein. The main advantage
of having deployable bumps is to reduce aerodynamic drag when they
are not required.
[0108] In the non-deployed position 40'' the gas stream through the
nozzle 42 is not disturbed by any bump 40 as would otherwise be the
case and described hereinbefore. In the deployed position 40',
particularly used at aircraft cruise, the bumps 40 interact with
the gas stream and reduce exhaust noise as herein described.
[0109] In each figure the bumps 40 are formed from a shape memory
material element 60 which is prestressed to a particular shape and
changes shape, at a predetermined temperature, between the deployed
and non-deployed positions. In FIG. 22 securing means 61 attaches a
continuous ring of SMM defining bump elements 60. However,
individual SMM elements may equally be used and aatched to the
nozzle wall by the securing means 61. The securing means 61 may be
a nut and bolt, weld, screw or other capturing member. The dashed
lines define the non-deployed position 40'' of the SMM element 60.
In the left hand part of the figure, the nozzle wall 42 defines a
bump 62 having amplitude between the maximum amplitude 38 and the
otherwise `original` nozzle wall profile indicated by the dashed
line 63. This arrangement is advantageous in that there are two
bump amplitudes which are help to attenuate cabin noise at two
different engine operating points.
[0110] In FIG. 22A shows a further embodiment of the SMM element
60, where there are two layers of SMM material 64, 65 which have
different switch temperatures. At a first temperature element 64
switches and the bump obtains a first shape 40''' and at a second
temperature element 65 obtains a second shape, the second shape
having a greater amplitude that the first shape.
[0111] In an alternative embodiment of FIG. 22A, the layer 64 is a
spring element, which comprises titanium or other suitable
resilient material, such that the spring element provides a force
to retain or return the bump in the non-deployed position or
perhaps in the deployed position. The element 60 is arranged such
that the change in modulus of the SMM element 65 is capable of
bending the element 60 into the desired shape.
[0112] In the embodiments shown all the downstream surfaces of the
bumps blend out at or just upstream of the final exit plane 36.
Thus the exit plane itself is a smooth and in these cases circular
shape. However, it is possible that the downstream surface is
intersected by the nozzle exit plane particularly where the
convergent-divergent is 1.00 or very close thereto.
* * * * *