WO 2012/080669 PCT/FR2011/053000
TURBINE ENGINE BLADE BAVING IMPROVED STACKING LAW
The field of the present invention is that of
thermodynamics and morespeGifically that of the blades
5 for the compressors of turbomachines.
Aeronautical turbomachines are conventionally made up,
from upstream to downstream in the direction in which
the gases flow, of a fan, of one or more compressor
10 stages, for example a low-pressure compressor and a
high-pressure compressor, a combustion chamber, one or
more turbine stages, for example a high-pressure
turbine and a low-pressure turbine, and a gas exhaust
nozzle. The compressor or compressors are produced in
15 the form of a plurality of sets of rotor blading
rotating past a plurality of sets of stator blading
known as guide vanes. The rotor blading is arranged
evenly at the periphery of a disk driven by the rotor
of the turbomachine, and their airfoils extend radially
20 between the rotor disk and a casing enclosing the
airflow path.
Each rotor blade comprises a pressure face over which
the air of the flow path is at a raised pressure with
25 respect to the mean pressure prevailing in the vicinity
of the blade airfoil, and a suction face over which the
air is at a reduced pressure in relation to this mean
pressure. This then causes an air circuit to become
established at the outer tip of the blade, causing air
30 to pass from the pressure face to the suction face
through the clearance there is between the blade and
the casing. In the known way, this circulation of air
develops along the entire length of the chord of the
blade and takes the form of a vortex, referred to as
35 the blade tip clearance vortex, which spreads
downstream of the trailing edge of the blade.
The presence of this vortex disturbs the flow in the
stages further downstream of the compressor and creates
WO 2012/080669 - 2 - PCT/FR2011/053000
losses which are detrimental to the efficiency of the
compressor. It would therefore be desirable to
eliminate this vortex or at the very least, to reduce
the flow rate of air it carries.
5
Attempts have been made to try to control this vortex,
these for example including treatments applied to the
casing surrounding the compressor or the creation of
"trenches", namely cavities hollowed into the casing.
10 One example of such treatments is described in the
applicant's patent application published under the
number FR 2940374. All of these have the disadvantage
of generating additional cost in producing the
turbomachine and of potentially impairing the
15 performance of the compressor in terms of efficiency at
certain operating points.
Patent applications have also been filed in an attempt
to reduce the impact that this vortex has on the
20 efficiency of a compressor or turbine stage, these
including for example applications US 2010/0054946 or
EP 1953341. These applications plan to modify the shape
of the blades by altering the shape given to the
leading edge, i.e. by altering its sweep angle between
25 the root and the tip of the blade along this leading
edge. They do not, with the exception of figure 12 of
the American publication, provide any indication
regarding changes to the line of stacking of the
elemental profiles along the height of the blade.
30
Moreover, document US 6341942 describes undulations
along the height of a compressor blade for the purpose
of increasing the flexural rigidity thereof, without an
increase in its mass. Although it indicates that one
35 undulation may be situated in a position high up on the
blade, it does not specify the position of the point of
inversion of curvature associated therewith, nor a
fortiori does it indicate the position of the lower
WO 2012/080669 - 3 - PCT/FR2011/053000
point of inflection in the case of a double inflection.
Moreover, by highlighting the problem of the
vibrational behavior of the blade, it is not, a priori,
of any benefit to a person skilled in the art wondering
5 how to improve the efficiency of a stage by controlling
the blade tip clearance vortex.
It is an object of the present invention to improve as
far as possible the efficiency of a compressor or
10 turbine stage of a turbomachine by giving the blade a
special shape that reduces the impact of this leakage
flow between the pressure face and the suction face of
the airfoil without any need to modify the compressor
casing.
15
To this end, one subject of the invention is a
turbomachine blade, the airfoil of which extends
radially between a blade root and an airfoil tip,
axially between a leading edge and a trailing edge, and
20 tangentially between a pressure face and a suction
. face, the profile of said blade being made up of a
series of elementary profiles, in the form of vane
sections, stacked on one another along a line known as
the stacking line joining the center of gravity of all
25 of the sections, characterized in that the projection
of said stacking line of the airfoil onto at least one
plane extending radially from the blade root comprises
a double tangential inversion of the direction of its
curvature which inversion is situated in the last 30
30 percent of. the height of
projection being oriented
to the chord of the blade.
the airfoil,
substantially
the plane of
perpendicular
These stacking modifications make it possible, through
. 35 better guidance of the flow, to reduce the blade tip
clearance vortex generated by the airfoil.
WO 2012/080669 - 4 - PCT/FR2011/053000
Indeed, calculations have shown that the beneficial
effect provided by the invention is no longer
maintained if the inversion is positioned lower down
than these last 30 percent. The impact that any
5 undulations further away from the blade tip might have
would be small because of the weak extent to which they
interfere with the blade tip clearance vortex.
As a preference, the two points of tangential inversion
10 are situated in the last 10 percent of the height of
the airfoil.
15
In another particular embodiment, the
comprises an axial inversion, the plane
being oriented substantially parallel to
the blade.
blade further
of pro j ection
the chord of
20
For preference, said projection contains a double
radial inversion.
The invention also relates to a compressor or to a
turbine for a turbomachine comprising at least one
rotor wheel made up of blades as described hereinabove,
and to a turbomachine comprising such a compressor or
25 such a turbine.
The invention will be better understood, and other
obj ects, details, features and advantages thereof will
become more clearly apparent during the course of the
30 detailed explanatory description which will follow of
several embodiments of the invention which are given by
way of purely illustrative and nonlimiting examples
with reference to the attached schematic drawings.
35 In the remainder of the description, the references
axial and tangential are to be understood to be with
reference to the axis of rotation of the turbomachine,
the axial direction coinciding with this axis of
WO 2012/080669 - 5 - PCT/FR2011/053000
rotation and the tangential direction being oriented
along a tangent to the circumference of the
turbomachine. By convention in the remainder of the
description, the direction referred to as axial with
5 reference to a blade is substantially that of a line
parallel to the chord at the tip of the blade, whereas
the direction referred to as tangential corresponds
substantially to a direction perpendicular to the chord
at the tip of the blade.
10
In these drawings:
- figure 1 is a perspective view of two adjacent blades
of a compressor according to the prior art;
figure 2 is a face-on view of a compressor blade
15 according to a first embodiment of the invention;
- figure 3 is a profile view of the blade of figure 2;
- figures 4 and 5 show the changes along the stack of
the profiles applied to a blade , respectively axially
and tangentially, according to the first embodiment,
20 and
- figures 6 and 7 show the changes to the stack of the
profiles applied to a blade , respectively axially and
tangentially, according to a second embodiment,
- figure 8 provides one example of how the efficiency
25 of a compressor stage comprising rotor blades according
to figure 2 or 3 is imp!oved.
Reference is made to figure 1 which shows two blades 1
of a compressor of an aeronautical turbomachine
30 according to the prior art, extending between a leading
edge 2 and a trailing edge 3, with a pressure face 4
and a suction face 5. The two blades depicted are
positioned side by side and guide the main flow 10 -of
the stream of air that is to be compressed. Because of
35 the raised pressure there is on the pressure face 4 of
each blade 1 and because of the reduced pressure there
is on the suction face 5 thereof, a leakage flow 11,
directed from the pressure face toward the suction
WO 2012/080669 - 6 - PCT/FR2011/0S3000
face, is set up at the top end of the blade, at the
clearance between this blade tip and the casing. As
this flow occurs over the entire length of the chord of
the blade 1, it grows into the form of a blade tip
5 clearance vortex 12 which spreads downstream of the
blade, along the axis of the chord thereof, thus
impairing the efficiency of the compressor.
Figures 2 and 3 show the airfoil of a blade 1 according
10 to the invention, face on, viewed from the suction face
side 5, and in profile, viewed from the trailing edge
3, respectively. The blade root, normally situated
toward the bottom of the figure, has not been depicted.
The shape of the airfoil can be defined as a series of
15 elemental profiles, in the form of vane sections, along
which the air that is to be compressed flows, these
profiles being stacked on one another along a line
referred to as the stacking line, starting from the
root and ending at the tip of the airfoil, and
20 connecting the centers of gravity of the various
sections. The shape of the blade can be defined, except
for the changes in the elemental profiles between the
bottom part and the top part of the airfoil, by, on the
one hand, the rotation applied to the elemental profile
25 according to its position along the height of the
airfoil and, on the other hand, the shape given to this
stacking line.
In a blade of the prior art, the curvature of the
30 stacking line changes very little between the root and
the tip of the airfoil of the blade 1; it generally has
a convex shape at its upper part (typically over a
region of between 20 and 100% of the height), which
means that the direction of curvature of the stacking
35 line is maintained. Figure 12 of document
US 2010/0054946 shows an inversion in the curvature of
the blade which is located a short way up the height of
the blade and which, because of this low-down
WO 2012/080669 - 7 - PCT/FR2011/053000
positioning, has no influence on the blade tip
clearance vortex or on the impact this has on the
efficiency of the stage to which the blade belongs.
5 In the case of the invention and, more particularly, in
the case of the blade depicted in figures 2 and 3, this
stacking line has two changes in curvature in the axial
direction and two changes in curvature in the
tangential direction, these changes taking place in the
10 last 10 percent of the height of the airfoil. In the
example depicted, the axial change in stacking is said
to be positive, the first change in curvature
(encountered when progressing from the root toward the
tip) causing the stacking to be offset toward the
15 trailing edge. In another embodiment (depicted in
figure 6), the offset may be negative, the invention
anticipating just one change in curvature which
therefore brings the stacking back toward the trailing
edge from a position that is already highly offset
20 toward the leading edge. As far as the tangential
modification to the stacking is concerned, this is
positive in figures 2 and 3, i. e. the stacking line
deforms in the direction of the suction face in the
first change of curvature; and then comes back toward
25 the pressure face in the second change in curvature.
Reference is now made to figures 4 and 5 which show the
shape of the stacking line of a blade according to the
first embodiment of the invention, in projection along
30 the height of the blade, in planes oriented
substantially axially and tangentially respectively,
which means parallel or perpendicular to the chord of
the blade. In this first embodiment, the change in
curvature of the stacking line occurs only over the
35 last 10 percent thereof (these being measured from a
zero conventionally considered to be at the root of the
blade and measuring toward the tip thereof, the reading
at the tip then corresponding to 100%). Figure 4 shows
the axial deformation of the stacking line of the
airfoil, i. e. the shape its projection has on a plane
oriented radially and parallel to the chord of this
airfoil; it shows this deformation in two
5 configurations, one according to the prior art (the
convex line) and one according to the invention (the
line that has a double change in curvature). Figure 5
shows the same tangential changes to the stacking line,
i.e. the shape of its projection onto a plane oriented
10 radially and perpendicular to the chord of the airfoil.
Whereas the blades of the prior art have, axially, a
convex shape, i. e. have a curvature which remains of
constant sign, the airfoil according to the first
embodiment maintains a curvature identical to that of
15 the prior art over 90% of its height, before turning a
first time towards the trailing edge and then turning
again and returning toward the leading edge; finally,
at the tip of the airfoil , it is practically back at
the same level as the blade of the prior art. There is
20 therefore a double change in the direction of axial
curvature of the airfoil over the last 10 percent of
the height thereof. The same phenomenon is encountered
in the tangential direction, with an airfoil of the
prior art having a curvature of constant sign over the
25 majority of its height, and in any case over its last
80%. The airfoil according to the first embodiment of
the invention has the same shape as the prior art over
the first 90 percent of its height, before differing
from the prior art through the presence of two changes
30 in curvature, a first one which shifts its stacking
line toward the suction face, followed by a second one
which more or less returns this stacking line to the
same position as that of the blade of the prior art.
WO 2012/080669 - 8 - PCT/FR2011/053000
35 In a similar way, figures 6 and 7 show a second
embodiment of the invention, the changes in shape of
the stacking line occurring, in this case, over the
last 30 percent of the height of the stacking line. In
WO 2012/080669 - 9 - PCT/FR2011/053000
this second embodiment, the axial deformation of the
stacking line has just one change in curvature, this
change being more pronounced than in the prior art over
the first 25 of the last 30 percent, reversing in the
5 last 5% and returning toward the stacking line of the
prior art. The tangential deformation for its part has
the same double change in the curvature as in the first
embodiment.
10 In both embodiments, as has been illustrated in figures
5 and 7, it may be seen that the tangential deformation
of the stacking line flattens in a very pronounced
manner as it nears the tip of the blade and that the
tangent to this line at the tip becomes contained in
15 the plane tangential to the cylinder that the chord
describes at the blade tip when the turbomachine is
turning. The same is true of the axial deformation of
this stacking line in the first embodiment (cf. figure
4) •
20
Figure 8 shows the improvement in performance that is
obtained using a three-dimensional Navier-Stokes
calculation on a compressor stage according to the
invention as compared with a stage produced in the
25 conventional way. The two curves show the points
obtained at iso-rotational speed for a stage of the
prior art (the curve at the bottom) and a stage
according to the invention (the top curve). The
abscissa axis represents, in units of 0.5 kg/s per
30 increment, the change in the flow rate of air passing
through the stage and the ordinate axis indicates, at a
scale of 0.1 point per increment, the efficiency
obtained for the various test points. The point
furthest to the left represents the point situated on
35 the surging line for this stage and the point furthest
to the right represents the point situated on the
operating line for the stage, which is, a priori
adopted in the design of the compressor. Between the
WO 2012/080669 - 10 - PCT/FR2011/053000
two, the stage passes through a point referred to as
the point of maximum efficiency, which is the point
aimed for when positioning the operating line for the
whole compressor.
5
The modified blades according to the invention· have
been evaluated from an aerodynamic standpoint using
three-dimensional calculation codes that allow the
Navier-Stokes equations to be solved. The results
10 obtained can be set out as follows, for both
embodiments, the efficiency of a compressor being
defined, on a scale from 0 to 100 points, as the ratio
of the work actually applied to the fluid by a given
increase in pressure, to the ideal work (isentropic
15 conversion) that would need to be supplied for the same
increase in pressure:
- for a blade modified over the last 10 percent of the
height of the airfoil thereof, there is observed, on
average, for each stage, an increase in efficiency of
20 0.15 points, with respect to the prior art, at the
compressor operating line (in a diagram that gives the
ratio of pressure obtained as a function of flow rate).
On the surge line, the improvement achieved is, on
average, around 0.30 points over the prior art. This
25 improvement on the borderline of surge may be converted
into an improvement in efficiency on the operating line
by moving the operating line closer to the surge line
by altering the angles of attack of the blades with
respect to the air stream~ The contribution of the
30 blading according to the invention therefore represents
an appreciable improvement over the bladings of the
prior art.
- the blades modified over the last 30 percent of their
height yield substantially identical results.
35 - by contrast, modifying the blades over a height in
excess of these last 30 percent provides no significant
addi tional improvement. The reason for this might be
the greater distance of the- undulations of the blade
WO 2012/080669 - 11 - PCT/FR2011/053000
from the tip thereof, the influence that these
undulations then have on the blade tip clearance vortex
then becoming negligible.
5 The invention has been described in relation to
compressor blades. Similar improvements may be obtained
on turbine blades which, in the prior art, suffer from
the same problem of controlling blade tip clearance
vortices.
10

WO 2012/080669 - 12 -
CLAIMS
PCT/FR2011/053000
1. A turbomachine blade, the airfoil of which extends
5 radially between a blade root and an airfoil tip,
axially between a leading edge (2) and a trailing edge
(3), and tangentially between a pressure face (4) and a
suction face (5), the profile of said blade being made
up of a series of elementary profiles, in the form of
10 vane sections, stacked on one another along a line
known as the stacking line joining the center of
gravity of all of the sections,
characterized in that the proj ection of said stacking
line of the airfoil onto at least one plane extending
15 radially from the blade root comprises a double
tangential inversion of the direction of its curvature
which inversion is situated in the last 30 percent of
the height of the airfoil, the plane of projection
being oriented substantially perpendicular to the chord
20 of the blade.
2. The blade as claimed in claim 1, in which the two
points of tangential inversion are situated in the last
10 percent of the height of the airfoil.
25
3. The blade as claimed in one of claims 1 and 2,
further comprising an inversion of axial curvature, the
plane of proj ection being oriented substantially
parallel to the chord of the blade.
30
4. The blade as claimed in claim 3, in which said
projection contains a double radial inversion.
5. A turbomachine compressor comprising at least one
35 rotor wheel made up of blades as claimed in one of
claims 1 to 4.
WO 2012/080669 -13 - PCT/FR2011/053000
6. A turbomachine turbine comprising at least one
rotor wheel made up of blades as claimed in one of
claims 1 to 4.
5 7. A
turbine
turbomachine
as claimed
comprising a
in one of
compressor or
claims 5 and
a
6
respectively.
Dated this 13/06/2013 ~~
NEHASRIVASTAVA
OF REMFRY & SAGAR
ATTORNEY FOR THE APPLICANT[S]