U.S. patent application number 11/984976 was filed with the patent office on 2008-06-12 for transition duct for a gas turbine engine.
This patent application is currently assigned to ROLLS-ROYCE PLC. Invention is credited to Matthew J. Green, Neil W. Harvey, Howard P. Hodson, Robert J. Miller, Edward M. J. Naylor, Cecilia Ortiz-Duenas.
Application Number | 20080138197 11/984976 |
Document ID | / |
Family ID | 37671893 |
Filed Date | 2008-06-12 |
United States Patent
Application |
20080138197 |
Kind Code |
A1 |
Green; Matthew J. ; et
al. |
June 12, 2008 |
Transition duct for a gas turbine engine
Abstract
A transition duct for a gas turbine engine typically contains
one or more non-turning fairings. These fairings disturb the air
flow through the duct, leading to flow separation and aerodynamic
inefficiency. The invention proposes non-axisymmetric perturbations
in the duct end walls to minimise these undesirable aerodynamic
effects. This permits the construction of shorter and lighter
transition ducts, with more pronounced curvature.
Inventors: |
Green; Matthew J.; (Derby,
GB) ; Harvey; Neil W.; (Derby, GB) ; Naylor;
Edward M. J.; (Market Deeping, GB) ; Ortiz-Duenas;
Cecilia; (Minneapolis, MN) ; Hodson; Howard P.;
(Godmanchester, GB) ; Miller; Robert J.;
(Cambridge, GB) |
Correspondence
Address: |
OLIFF & BERRIDGE, PLC
P.O. BOX 320850
ALEXANDRIA
VA
22320-4850
US
|
Assignee: |
ROLLS-ROYCE PLC
London
GB
|
Family ID: |
37671893 |
Appl. No.: |
11/984976 |
Filed: |
November 26, 2007 |
Current U.S.
Class: |
415/182.1 |
Current CPC
Class: |
Y02T 50/60 20130101;
F01D 9/023 20130101; F01D 5/143 20130101; F05D 2240/126 20130101;
Y02T 50/673 20130101; Y02T 50/671 20130101; F05D 2250/70 20130101;
F01D 5/145 20130101 |
Class at
Publication: |
415/182.1 |
International
Class: |
F01D 25/24 20060101
F01D025/24 |
Foreign Application Data
Date |
Code |
Application Number |
Dec 5, 2006 |
GB |
0624294.5 |
Claims
1. A transition duct for a gas turbine engine, the duct comprising
circumferentially-extending radially inner and outer end walls
defining A an annular flow path for a gas flow, the duct having
different radii from the engine centreline at its upstream and
downstream ends, the duct further comprising a radially extending
fairing, in which at least one of the end walls has a profile
defined by a plurality of non-axisymmetric perturbations at a
succession of axial positions so as to reduce or prevent separation
of the end wall boundary layer flow near to the fairing in use.
2. A transition duct as in claim 1, in which the perturbations are
applied to the radially inner end wall.
3. A transition duct as in claim 1 in which the fairing has a
camber line and the perturbations are symmetrical about the camber
line.
4. A transition duct as in claim 3, in which the fairing is
symmetrical about the camber line.
5. A transition duct as in claim 1, in which at least some of the
perturbations alter the local radius of curvature of the end
wall.
6. A transition duct as in claim 5, in which the local convex
radius of curvature of the end wall, viewed in the circumferential
direction, is increased in the region of the fairing leading edge
and the local concave radius of curvature of the end wall, viewed
in the circumferential direction, is increased in the region of the
fairing trailing edge.
7. A transition duct as in claim 5, in which the circumferential
extent of the perturbations is between 5% and 100% of the fairing
pitch.
8. A transition duct as in claim 5, in which the circumferential
extent of the perturbations is substantially 50% of fairing
pitch.
9. A transition duct as in claim 5, in which the maximum amplitude
of the perturbations is between 2% and 15% of chord.
10. A transition duct as in claim 5, in which the maximum amplitude
of the perturbations is substantially 7% of chord.
11. A transition duct as in claim 1, in which at least some of the
perturbations locally alter the cross-sectional area of the duct
close to the fairing surface.
12. A transition duct as in claim 11, in which the cross-sectional
area of the duct is reduced in the region of the fairing leading
edge and increased in the region of the fairing trailing edge.
13. A transition duct as in claim 11, in which the circumferential
extent of the perturbations is between 5% and 100% of the fairing
pitch.
14. A transition duct as in claim 11, in which the circumferential
extent of the perturbations is substantially 50% of fairing
pitch.
15. A transition duct as in claim 11, in which the maximum
amplitude of the perturbations is between 2% and 15% of chord.
16. A transition duct as in claim 11, in which the maximum
amplitude of the perturbations is substantially 7% of chord.
17. A transition duct as in claim 1, in which at least one spline
fit is applied through circumferentially corresponding points at
different axial locations to blend the perturbations smoothly into
the axisymmetric end wall profile.
18. A gas turbine engine incorporating a transition duct according
to claim 1.
19.-20. (canceled)
Description
[0001] This invention relates to transition ducts in gas turbine
engines.
[0002] In a multi-spool jet engine the low-pressure (LP) system has
a lower rotational speed and larger radius than the high-pressure
(HP) core system. Hence, intermediate "S" shaped transition ducts
are needed to connect the compressor or turbine of the high-radius
LP system with the corresponding compressor or turbine of the
low-radius HP system. These intermediate ducts often carry loads,
support bearings and have thick structural struts passing through
them, making them large, heavy and expensive structures of
considerable complexity. Improving these ducts can lead to
significant benefits both in the weight and in the performance of
the engine.
[0003] In particular, if the curvature of the duct is made more
pronounced, the required change of radius can be accomplished with
a shorter duct and the whole engine can be made shorter and
lighter. However, more pronounced curvature exacerbates the
undesirable aerodynamic effects in the duct. This may worsen the
aerodynamic performance of a downstream compressor or turbine, and
may cause instability in either an upstream or a downstream
compressor. These effects limit the design of current transition
ducts.
[0004] It is therefore an object of the invention to provide a
transition duct for a gas turbine engine that is shorter and
lighter than known ducts, without introducing aerodynamic
problems.
[0005] According to the invention, there is provided a transition
duct for a gas turbine engine as claimed in claim 1.
[0006] The invention will now be described, by way of example, with
reference to the accompanying drawings in which
[0007] FIG. 1 is a perspective view of part of a hub end wall of a
known transition duct containing a fairing, showing contours of
static pressures near to the wall;
[0008] FIG. 2 is a sectional side view of the duct of FIG. 1,
showing contours of velocity close to the fairing surface;
[0009] FIG. 3 is a schematic plan view of a fairing in a transition
duct, showing the changes to the hub end wall in a first embodiment
of the invention;
[0010] FIG. 4 shows the total pressure contours in axial planes
close to the fairing leading edge (FIG. 4(a)) and trailing edge
(FIG. 4(b)) for the embodiment of FIG. 3;
[0011] FIG. 5 is a perspective view of part of a hub end wall of a
transition duct containing a fairing, showing contours of static
pressures near to the wall in a known duct (FIG. 5(a)) and in a
duct according to the invention (FIG. 5(b));
[0012] FIGS. 6(a) and 6(b) are sectional side views of the ducts of
FIGS. 5(a) and 5(b), showing contours of velocity close to the
fairing surface;
[0013] FIG. 7 shows contours of total pressure in the axial plane
downstream of a fairing in a transition duct, without (FIG. 7(a))
and with (FIG. 7(b)) non-axisymmetric end wall profiling;
[0014] FIG. 8 is a graph of static pressure against fairing surface
perimeter for the boundary layer air flow in the embodiment of FIG.
9; and
[0015] FIG. 9 is a schematic plan view of a fairing in a transition
duct, showing the changes to the hub end wall in a second
embodiment of the invention.
[0016] A typical duct used between compressors today has an area
ratio (A.sub.in/A.sub.out) of 1.0-2.5 and an aspect ratio
(.DELTA.R/length) of 0.3-0.6. The ducts used in current engines may
in fact achieve only minimal diffusion. This is a very conservative
choice in order to be absolutely sure that there is no risk of
separations in the duct. When optimizing a compression system a
conservative duct leads to significant compromises in both upstream
and downstream compressors. Rear stages of the upstream compressor
have to be designed at a non-optimum radius, and the design flow
coefficient of the downstream compressor is limited by the exit
Mach number of the upstream compressor and the choice of a
conservative area ratio through the duct. There are inevitable
cost, weight and fuel-burn penalties associated with such
compromises, and these will be present for the complete life cycle
of the product. If it were possible to use more pronounced
curvatures in duct designs and design tools, much better
optimization of the whole compression system would be possible.
[0017] Transition ducts between compressors should ideally diffuse
the flow through them in order to minimise the flow velocity into
the downstream compressor. That is to say, the cross-sectional area
of the duct should ideally increase in the flow direction.
[0018] Where structural struts are present they will typically have
a form of aerodynamic fairing around them to present an acceptably
smooth surface to the gas flow. Typically the fairing is
symmetrical in shape around the camber line of the strut, and both
its surfaces are largely convex in plan view.
[0019] In some instances the strut and fairing will not be
separate. Rather, the strut itself will be aerodynamically
shaped.
[0020] Throughout this specification, the term "fairing" will be
used for such an aerodynamic fairing, or for a fairing-shaped
strut.
[0021] These fairings do not turn the air flow through the duct,
and so there is no aerodynamic lift on them. These fairings should,
therefore, not be confused with vanes, or static turning aerofoils,
in a compressor or a turbine of a gas turbine engine.
[0022] In the case that the inlet flow is axially aligned, the
fairing will also be so aligned. However, if the inlet flow is
swirling (has a tangential velocity component) then the fairings
will be staggered to be aligned to this flow.
[0023] In the case of swirling inlet flow and the duct changing
radius (which typically it will) then the flow angle will change
through the duct as a) the axial velocity changes (with area
change) and b) the tangential velocity changes (with radius,
following the conservation of angular momentum). If the flow angle
through the duct does change then the stagger angle of the camber
line will be modified accordingly. The fairing stagger angle will
follow the flow--but it will not be turning it.
[0024] Although fairings improve the gas flow past the struts,
their presence will exacerbate the diffusion of the flow locally
and may give rise to flow separations, resulting in aerodynamic
losses in the duct and poor inlet conditions to the downstream
compressor. The latter may cause a loss of efficiency of the
downstream compressor and/or make it more prone to surge.
[0025] FIGS. 1 and 2 illustrate the problems with a diffusing duct
with a (thick) fairing 12 in it. In this case the flow into the
duct has no tangential component.
[0026] FIG. 1 shows a plan view of contours of static pressure on
the hub (radially inner) end wall 14. The circumferential direction
is shown by the arrow 16, and the direction of the inlet gas flow
by the arrow 18. The lowest values of static pressure are seen over
the portion 20 of the end wall that has convex curvature, and it
can be seen that low static pressure persists in the region 22 near
to the fairing, even as the static pressure further away from the
fairing (and at the same axial position) increases. It is these
regions of low static pressure that give rise to the flow
separations. Further downstream, the static pressure rises in the
region 24 of the end wall that has concave curvature, and as the
cross-sectional area of the duct increases.
[0027] FIG. 2 shows, in sectional side view, contours of flow
velocity close to the fairing 12 of FIG. 1. The direction of the
inlet gas flow is shown by the arrow 18, and the radially outward
direction by the arrow 26. The inlet 23 to the transition duct is
therefore at a greater radius than the outlet 25. The velocities
shown are just outside the aerofoil boundary layer, but the effects
of the end wall boundary layer can be seen. Note also that the
static pressure and fluid velocity are in an inverse
relationship.
[0028] In FIG. 2 the curvature of the hub end wall 14 can be
clearly seen. In the front part 28 of the transition duct the
curvature is convex. This accelerates the flow, raising velocities
and reducing the static pressure. This reduction in static pressure
can be seen in FIG. 1 (region 20). Further downstream, in the
rearward part 30 of the transition duct, the hub end wall curvature
is concave. This will have the effect of diffusing the flow.
Because the duct is reducing in radius, from inlet to exit, it will
be understood that this arrangement, with first convex and then
concave curvature on the hub end wall, will generally be
unavoidable. The flow field on the hub end wall is a particular
problem. The early convex curvature raises the local flow velocity
above the value that would be expected from the area change alone,
and it may in fact be significantly higher than the inlet value.
From this local peak in velocity the flow must then diffuse to the
exit of the duct. Thus the maximum (local) diffusion may be much
higher than that obtained from simply considering the ratio of exit
to inlet flow velocities.
[0029] These effects would arise even for the duct alone. However,
the presence of the fairing 12 exacerbates them. The region of
lower static pressure due to the flow accelerating over the convex
end wall curvature 28 extends further downstream near to the
fairing 12 than it does further away from it (region 22 in FIG. 1).
This is due both to the blockage caused by the fairing 12 and to
the convex curvature of its two sides, which continues to
accelerate the flow compared with what it would have been in the
duct alone. In FIG. 2, region 34 shows the increased flow velocity
arising from the blockage and convex surface curvature of the
fairing 12. The result is that the diffusion gradient experienced
by the flow in the later part of the fairing 12/end wall 14 corner
region is increased.
[0030] The aerodynamic conditions may be so adverse that flow
separation occurs. This is shown in FIG. 2 where a significant
region of reverse flow 32 occurs towards the downstream end of the
fairing 12 surface near the hub. This significantly increases the
losses in the duct. Note that this region of reverse flow typically
does not extend across the whole circumferential width of the
passage, but is confined to the region near both surfaces of the
fairing 12.
[0031] It should also be understood that the location of the
maximum thickness of the fairing 12 is typically in its forward
(upstream) part. This will usually correspond to the location of
the maximum thickness of the strut within it. By having the maximum
thickness in the forward part of the strut the diffusion gradient
along the surface of the fairing is reduced and the losses in the
boundary layer (away from the end walls) are minimised.
[0032] It will be clear that if the aim of reducing duct length,
and/or increasing the radius change for a given length, is to be
achieved the radii of curvature of the different parts of the end
walls must increase. The local acceleration and diffusion of the
flow will also increase, making flow separation more likely or
adding to the severity of any separation that has already
occurred.
[0033] Three-dimensional shaping of vanes in gas turbine engines is
a well-known technique, used to improve aerodynamic efficiency and
suppress end wall flow separation. Examples of such shaping are the
application of lean (often "compound lean"), sweep and dihedral.
However, in the case of the transition duct described here, the
strut will typically be a major load-bearing part of the gas
turbine engine and the loads must be carried in a largely radial
manner through it, so it is not possible to apply any significant
three-dimensional shaping to such components.
[0034] Boundary layer bleed, from the strut and/or the end walls,
is another known means of preventing separation in strongly
diffusing flows. However, this is rarely if ever applied to
practical gas turbine engines, because usually the benefit in
aerodynamic efficiency will be lost due to the cost and weight of
the bleed system. In addition the action of bleeding off the flow
and either dumping it overboard or re-introducing it back into the
gas flow in some other part of the engine will typically incur as
much extra loss elsewhere as was gained in the duct.
[0035] FIG. 3 illustrates a first embodiment of the invention, and
shows schematically profiling of the hub end wall around a fairing
12 in a transition duct.
[0036] Typically this arrangement will be found aerodynamically
connecting two compressors in a multi-shaft gas turbine, the duct
and compressors being arranged axisymmetrically around the
centre-line of the gas turbine.
[0037] Typically the strut, and the fairing around it, will be
aligned substantially radially. They may be leant in an axial sense
but they will not usually be leant tangentially.
[0038] Typically in such an arrangement the first compressor will
be at a higher radius than the second and rotating more slowly than
it. Thus this transition duct will typically have a reducing radius
through it.
[0039] In FIG. 3, the direction of the inlet gas flow is shown by
the arrow 18. At two axial locations, the end wall has been
profiled to define perturbations 36, 38. Both perturbations have a
circumferential extent of about 1/3 of the pitch between
circumferentially-adjacent fairings 12, at their respective axial
locations. Each perturbation 36, 38 will generally take the form of
a protruding "blister" or a recessed "hollow" in the axisymmetric
annulus profile. The perturbations 36, 38 will be described in more
detail in due course.
[0040] FIG. 4 illustrates the total pressure contours in axial
planes close to the fairing 12 leading edge 46 (FIG. 4(a)) and
trailing edge 48 (FIG. 4(b)). The positions of the perturbations
36, 38 are also shown in FIGS. 4(a) and 4(b) respectively.
[0041] FIG. 5 illustrates the effect of the hub end wall profiling
of FIG. 3, showing contours of static pressures near to the wall in
a known duct (FIG. 5(a)) and in a duct according to the invention
with hub end wall profiling (FIG. 5(b)).
[0042] FIG. 6 shows contours of velocity close to the vane fairing
surface in a known duct (FIG. 6(a)) and in a duct according to the
invention with hub end wall profiling (FIG. 6(b)).
[0043] Referring to FIG. 3, in the front part of the passage, near
the location of fairing 12 maximum thickness, the perturbation 36
is symmetrical about the fairing camber line 40. Noting that this
is a hub end wall, the radius (from the engine centreline) is
reduced at the camber line. The perturbation 36 has its lowest
point at the camber line, and therefore defines a "hollow" in the
end wall. Away from the camber line, the amplitude of the
perturbation reduces steadily until the end wall radius (from the
engine centreline) is the same as that of the axisymmetric annulus.
This occurs at the position indicated by the notional line 37,
which defines the circumferential extent of the perturbation. The
amplitude of the perturbation also reduces in an axially upstream
direction from its maximum, to blend into the axisymmetric annulus
at the appropriate point on the notional line 37.
[0044] Near the leading edge of the fairing (perhaps slightly
upstream of it) this increases the local (convex) radius of
curvature of the end wall (as viewed in the circumferential
direction). This locally lowers the static pressure and raises the
free stream velocity. Referring to FIG. 5 it can be seen that the
region of lowest static pressure is extended further
circumferentially in the leading edge region (compare region 520 in
FIG. 5(b) with region 20 in FIG. 5(a)) and the high pressure at the
leading edge stagnation point is reduced (compare regions 552 and
52). However, downstream of this, the (convex) radius of curvature
is reduced over most of the front part of the passage adjacent to
the fairing thus reducing the acceleration of the flow in this
region. This locally raises the static pressure and lowers the free
stream velocity.
[0045] In the rear part of the passage the perturbation 38 has the
opposite arrangement to the perturbation 36. Its highest point
(i.e. the point of maximum radius from the engine centreline) is at
the camber line 40, and it therefore defines a "blister" in the end
wall. Away from the camber line, the amplitude of the perturbation
reduces steadily until the end wall radius (from the engine
centreline) is the same as that of the axisymmetric annulus. This
occurs at the position indicated by the notional line 39, which
defines the circumferential extent of the perturbation. The
amplitude of the perturbation also reduces in an axially downstream
direction from its maximum, to blend into the axisymmetric annulus
at the appropriate point on the notional line 39.
[0046] This increases the (concave) radius of curvature (as viewed
in the circumferential direction) of the end wall adjacent to the
fairing over most of the rear part of the passage, thus generally
reducing the diffusion of the flow in this region. Downstream of
the trailing edge, the perturbed annulus line is faired back into
the axisymmetric annulus 42. Locally the (concave) radius of
curvature is increased thus increasing the diffusion.
[0047] The perturbations will have a circumferential extent, on
each side of the fairing, typically about 50% of the fairing pitch
at each location. The range of useful values will lie between 5%
and 100% of the pitch. The circumferential extent of a perturbation
is defined as described in the discussion of FIG. 3 above,
particularly in terms of the notional lines shown as 37 and 39 in
that drawing. For example, if two adjacent perturbations each had a
circumferential extent of 50% of pitch, then one or both of their
notional lines 37, 39 would be coincident.
[0048] In the case where the fairings are uniformly spaced
circumferentially and all have to have the same end wall profiling,
then the perturbations will be symmetrical about the fairing camber
lines and also about the mid-pitch lines. Thus the maximum
circumferential extent could only be 50% of pitch in this case.
[0049] However, it may be the case that the fairing arrangement is
not periodically uniform. It might happen, for example, that only
every alternate fairing has the profiling applied. In this case the
perturbation around one fairing could extend beyond the mid-pitch
position, so that its circumferential extent would be greater than
50%, and it may extend as far as the next fairing, in which case
its circumferential extent would be 100%.
[0050] The perturbations at different axial locations may have
different maximum amplitudes. Typically the maximum amplitude will
be 7% of the chord, but may lie in the range 2% to 15% of chord
depending on the details of the flow conditions. (Higher speed
flows require smaller amplitudes to achieve the same aerodynamic
effects as lower speed flows).
[0051] The effect of these perturbations is that the static field
on the hub end wall is more uniform and the corner separation is
reduced (seen in FIG. 6 as the reduced extent of the region of low
or reversed flow velocity in the downstream part of the
duct--compare region 632 in FIG. 6(b) with region 32 in FIG. 6(a)
or FIG. 2). The static pressure experienced by the hub boundary
layer close to the fairing 12 is now closer to that experienced by
the hub boundary layer at mid-pitch (that is to say, the maximum
diffusion gradient it experiences is reduced, and is closer to that
at mid-pitch).
[0052] A smooth end wall shape is obtained by applying a spline in
the streamwise direction through circumferentially corresponding
points at these axial locations, to blend the perturbations
smoothly into the axisymmetric end wall shape so as to present a
smooth surface to the gas flow.
[0053] FIG. 7 shows contours of total pressure in an axial plane at
the exit from the duct, for a known duct (FIG. 7(a)) and for a duct
according to the invention with non-axisymmetric perturbations
(FIG. 7(b)). The areas of greatest loss are shown by the regions 72
in FIG. 7(a) and by the region 74 in FIG. 7(b), from which it can
be seen that the extent and depth of the losses have been
significantly reduced. Reductions in loss of 20% may be achieved,
with a corresponding improvement in the aerodynamic efficiency.
[0054] As noted already, the presence of a fairing in an annular
duct changes the pressure field on the duct wall. Around the
upstream part of the fairing the flow is accelerated and around the
downstream part the flow is diffused. This extra diffusion on the
duct wall can cause the boundary layer to separate close to the
strut surface. By locally altering the area of the duct close to
the fairing surface the effect of this acceleration and diffusion
can be reduced.
[0055] In regions where the fairing is accelerating the flow the
end walls must be opened out. Increasing the passage area would, on
its own, act to decelerate the flow.
[0056] Where the fairing is decelerating the flow the end walls
must be contracted. Decreasing the passage area in this way would,
on its own, act to accelerate the flow.
[0057] Careful design of the area ruling should enable the flow
near the fairing to experience the same diffusion (from inlet to
exit) that the mid-passage region experiences. This embodiment is
illustrated in FIGS. 8 and 9.
[0058] FIG. 8 illustrates the typical static pressure distribution
along one surface of the fairing where it intersects with the
axisymmetric end wall--in the particular case where there is no net
diffusion through the duct. Static pressure, P, is plotted against
fairing surface perimeter, S. The static pressure rises (82) near
the leading edge as the inlet flow tends to stagnate here.
Downstream of this the static pressure falls (84) as the flow is
accelerated by the blockage of the fairing. After the plane of
maximum thickness the fairing thins, the flow diffuses and the
static pressure rises again (86).
[0059] To minimise this pressure variation a second embodiment of
the invention is proposed, in which the end wall profiling is as
shown in FIG. 9. The amplitudes of the perturbations, and their
axial locations, are defined such that they compensate for the
reduction in the cross-sectional area of the passage due to the
blockage of the fairing. This will be known as "non-axisymmetric
end wall area ruling". In this case there will be a perturbation 92
in the leading edge region such that the radius of the end wall
with respect to the engine centreline is increased at the
intersection with the camber line 40. This will locally reduce the
flow area, raising flow velocities and reducing the static
pressure. This will compensate for the increase in static pressure
where the flow stagnates at the leading edge. The maximum amplitude
of the perturbation may be at the leading edge or upstream of
it.
[0060] Downstream of this there will be another perturbation 94
such that the radius of the end wall with respect to the engine
centreline is decreased at the intersection with the camber line.
This will locally increase the flow area adjacent to the fairing,
lowering flow velocities and raising the static pressure. This will
compensate for the decrease in static pressure due to the blockage
of the fairing in the flow. Typically the maximum amplitude 96 of
the perturbation will be axially located in the plane of maximum
vane thickness. The chord-wise extents and amplitudes of the
perturbations will typically lie within the same ranges as
before.
[0061] As in the first embodiment, a smooth shape is obtained for
the end wall by applying a spline fit in the streamwise direction
through circumferentially corresponding points at different axial
locations, to blend the perturbations smoothly into the
axisymmetric end wall shape.
[0062] The end wall profiling is defined by perturbations at, at
least, two axial locations.
[0063] The end wall profiling may be applied to either (radially
inner or radially outer) end wall of the transition duct. Typically
for a transition duct with reducing radius it will be most
effective if applied to the hub end wall.
[0064] In practice the optimum design will incorporate a
combination of controlling local end wall curvature and non
axisymmetric end wall area ruling.
[0065] To achieve the required change in curvature the annulus must
locally be raised or lowered, which will alter the flow area.
Conversely, the area ruling cannot be implemented without changing
the local surface curvature.
[0066] Ultimately the designer must finalise the choice of
perturbation amplitudes and locations to minimise the effect of the
presence of the fairing on the hub end wall static pressure
distribution.
[0067] An optimum aerodynamic design for the annulus wall may be
obtained by the designing the duct without a fairing present to
achieve maximum diffusion along the annulus wall while avoiding
(2-D) boundary layer separation. The fairing is then replaced and
non axisymmetric end wall profiling applied to restore the end wall
static pressure field to be as close as possible to what it was in
the absence of the fairing.
[0068] Where the duct wall static pressure field is made more
uniform, the circumferential variation in static pressure
experienced by upstream and/or downstream blade rows will also be
reduced. This may be of benefit in reducing the aero-mechanical
forcing on these blade rows and improve the component lives by
reducing the unsteady stresses induced in them.
[0069] It may be the case that the fairings are not arranged
uniformly in the circumferential direction. In this case the
dimensions of the perturbations may be based on the chord and pitch
local to each fairing.
[0070] Where the interaction of the static pressure field of the
fairing with adjacent blade rows is a major issue, i.e. the
aero-mechanical forcing significantly reduces these component
lives, then it is a known solution to adopt an arrangement of
"Bragg struts". In this case small vanes or fairings, sometimes in
a helical spiral arrangement, are added between the larger fairings
around the struts. In this case, non axisymmetric end wall
profiling may be applied to the larger and the smaller fairings but
in each case the perturbations will be based on the chord and pitch
local to each.
* * * * *