Background of the invention
The present invention relates to the field of
5 turbomachine blades or vanes, and in particular the field
of turbomachine rotor blades.
The term "turbomachine" is used in the present
context to designate any machine in which energy can be
transferred between a fluid flow and at least one set of
10 blades, such as for example: a compressor, a pump, a
turbine, or indeed a combination of at least two of
these. In the description below, the terms ''upstream"
and "downstream" are defined relative to the normal flow
direction of the fluid through the turbomachine.
15 Such a turbomachine may comprise a plurality of
stages, each stage normally comprising two sets of blades
and vanes, specifically a set of moving blades and a set
of guide vanes. Each set of blades or vanes comprises a
plurality of blades or vanes that are offset from one
20 another in a lateral direction. Typically, such blades
or vanes are arranged radially around a central axis A.
Thus, such a set forms a rotor, when it is a set of
moving blades, or a stator when it is a set of guide
vanes~ The proximal end of each blade or vane relative
25 to the central axis A is normally referred to as its
root, while the distal end is normally referred to as its
tip. The distance between the root.and the tip is
referred to as the ''height''. Between its root and its
tip, a blade or vane is made up of a stack of aerodynamic
30 profiles that are substantially perpendicular to a radial
axis Z. The term "substantially perpendicular" is used
in this context to mean that the plane of each profile
may present an angle close to 90°, e.g. lying in the
range 60°tO 120°, relative to the radial axis Z.
35 The geometrical shape of blades is the subject of
major design efforts in order to optimize the aerodynamic
behavior of blades, thereby increasing the efficiency of
2
the rotary assemblies such as compressors, fans, or
turbines, of which they form a part. Thus, aerodynamic
engineers propose relationships for stacking aerodynamic
profiles that are optimized from the aerodynamic point of
5 view.
Nevertheless, such stacking relationships are not
necessarily optimized, nor even acceptable, from a
mechanical point of view. For example, stacking
relationships that are particularly effective from an
10 aerodynamic point of view have been proposed in which a
major portion of the blade is cantilevered out relative
to the remainder of the blade. Such a large
cantilevered-out mass is then highly sensitive to
centrifugal forces resulting from the rotation of the
15 rotor and leads to significant bending of the high
portion of the airfoil, thereby leading to large
mechanical stresses in the middle of the airfoil with
static stresses that are too high at the "red line"
flight point, i.e. the emergency flight point. Under
20 such circumstances, such blades have only a very small
dynamic margin and, in the event of an impact or in the
event of the rotary assembly surging, they withstand
fatigue poorly.
Conversely, other stacking relationships that are
25 optimized from a mechanical point of view have been
proposed by mechanical engineers, but they have not been
accepted because of their aerodynamic performance being
insufficient.
There therefore exists a real need for a blade that
30 benefits both from good aerodynamic properties and from
good mechanical properties.
Object and summary of the invention
The present description relates to a turbomachine
35 rotor blade comprising a blade root and a blade tip
separated by a blade height, together with at least one
intermediate segment presenting a negative tangential
5
3
slope, and a distal segment situated between the
intermediate segment and the blade head and presenting a
positive tangential slope, when the distal segment
extends over at most 30% of said blade height.
Such a configuration makes it possible to greatly
reduce the mass of the blade that is cantilevered out,
and thus to reduce the bending of the high portion of the
blade, and simultaneously reduce the level of static
stresses, while preserving good aerodynamic properties.
10 Furthermore, the mechanical strength improvements as
obtained in this way make it possible to lighten the
structure of the blade, and in particular to reduce its
thickness, thereby improving its aerodynamic properties.
Also, this stacking relationship leaves design
15 freedom concerning the geometrical shape of the blade in
the axial direction, thereby making it possible to
optimize the blade freely in this direction in order to
optimize its aerodynamic and/or mechanical properties as
a function of specific requirements.
20 In certain embodiments, said distal segment is
directly adjacent to said blade tip.
In certain embodiments, said distal segment is
directly adjacent to said intermediate segment.
In certain embodiments, the junction between said
25 intermediate segment and said distal segment is situated
at a level lying in the range 75% to 80% of the height of
the blade from the blade root.
In certain embodiments, said distal segment extends
over at most 25% of the blade height. This further
30 reduces the cantilevered-out mass.
In certain embodiments, said distal segment extends
over at least 5% of the blade height.
In certain embodiments, said distal segment extends
over at least 15% of the blade height.
35 In certain embodiments, the projection onto a radial
plane of the line passing through the centers of gravity
of each of the tangential sections of the blade is gamma4
shaped. The curve plotting the tangential coordinate,
known as yG, of the center of gravity of each section of
the blade as a function of its position along the blade
thus presents a substantially rectilinear rising segment,
5 imparting increased stiffness to the lower portion of the
blade, and a subhorizontal segment that leaves only a
small weight cantilevered out. Furthermore, the
pronounced curvature at the interface between the
intermediate segment and the distal segment of the blade
10 serves to break the force path between the high portion
and the low portion of the blade, thereby serving to
reduce the magnitude of the stresses generated in the low
portion of the blade as a result of bending in the high
portion of the blade.
15 In certain embodiments, the projection onto a radial
plane of the line passing through the centers of gravity
of each tangential section of the blade possesses, in the
intermediate segment of ·the blade, a segment that is
substantially rectilinear. In other words, the curve
20 plotting the tangential coordinate yG as a function of
position possesses a second derivative that is
substantially zero in this segment. This leads to
greater stiffness in the intermediate segment of the
blade, thus reinforcing its mechanical strength.
25 In certain embodiments, said substantially
rectilinear segment extends over at least 30% of the
blade height, preferably over at least 40% of the blade
height, more preferably over at least 50% of the blade
height.
30 In certain embodiments, said substantially
rectilinear segment extends at least to a level situated
at 50%, preferably 55%, preferably 60%, more preferably
65% of the blade height from the blade root.
In certain embodiments, said substantially
35 rectilinear segment extends at least from a level
situated at at most 30%, preferably at most 20%, of the
blade height from the blade root.
5
In certain embodiments, the difference between the
tangential coordinates of the centers of gravity of the
blade sections situated firstly at the blade tip and
secondly at the interface between the intermediate
5 segment and the distal segment is greater than 150%,
preferably greater than 180%, of the difference between
the tangential coordinates of the centers of gravity of
the blade sections situated firstly at the blade root and
secondly at the interface between the intermediate
10 segment and the distal segment. Such a blade has a large
amount of sweep at the blade tip and benefits from good
aeronautical and mechanical properties, in particular in
terms of static stresses.
In other embodiments, the difference between the
15 tangential coordinates of the centers of gravity of the
blade sections situated firstly at the blade tip and
secondly at the interface between the intermediate
segment and the distal segment lies in the range 100% to
150%, preferably in the range 110% to 140%, of the
20 difference between the tangential coordinates of the
centers of gravity of the blade sections situated firstly
at the blade root and secondly at the interface between
the intermediate segment and the distal segment. Such a
blade presents sweep at the blade tip that is less
25 pronounced than in the preceding situation, thereby
making it easier to integrate in the turbomachine module,
while preserving good aeronautical and mechanical
properties. In particular, this makes it possible to
increase the distance between the tips of the blade of
30 the rotor and of the vane of the stator situated
upstream.
In certain embodiments, the thickness of the blade
is at all points less than 8% of the blade height, and
preferably less than 6%. This reduced thickness provides
35 the blade with good aerodynamic properties.
6
In certain embodiments, the thickness of the distal
segment is less than 5% of the blade height, preferably
less than 3%.
The present description also provides a single-piece
5 blade disk having a plurality of blades in accordance
with any of the above embodiments. Such a single-piece
blade disk presents advantages of robustness and
simplicity. Nevertheless, blades of the present
description may alternatively be individually fastened
10 blades, e.g. having a fir-tree shaped fastener member
under the blade root.
15
The present description also provides a rotor having
a plurality of blades in accordance with any of the above
embodiments.
The present description also provides a turbomachine
having a disk or rotor in accordance with any one of the
above embodiments.
The above-described characteristics and advantages,
and others, appear on reading the following detailed
20 description of embodiments of the proposed blade. The
detailed description refers to the accompanying drawings.
Brief description of the drawings
The invention can be well understood and its
25 advantages appear better on reading the following
detailed description of an embodiment shown as a
nonlimiting example. The description refers to the
accompanying drawings, in which:
· Figure 1 is a diagrammatic longitudinal section of
30 a turbomachine;
Figure 2A:is a diagrammatic perspective view of a
turbomachine rotor;
Figure 2B is a diagrammatic perspective view of a
detail of the Figure 2A rotor;
35 Figure 2C is a cross-section view of one of the
blades of the Figure 2A rotor on a plane IIC-IIC shown in
Figure 2B;
5
7
Figure 3A shows a forwardly-swept rotor blade;
Figure. 3B shows a rearwardly-swept rotor blade;
Figure 3C shows a rotor blade presenting a
negative tangential slope;
· Figure 30 shows a rotor blade presenting a
positive tangential slope;
Figure 4 shows an example blade of the invention;
Figure 5 is a diagram showing the variation,
between the root and the tip, of the tangential slope of
10 a first blade of the invention;
· Figure 6 is a diagram showing the variation,
between the root and the tip, of the tangential slope of
a second blade of the invention;
· Figure 7 is a diagram showing the variation,
15 between the root and the tip, of the tangential slope of
a reference prior art blade;
· Figures SA to BB show the static stress levels of
the first example blade;
· Figures 9A to 9B show the static stress levels of
20 the second example blade;
25
· Figures lOA to lOB show the static stress levels
of a conventional prior art blade.
Detailed description of the invention
Figure 1 shows an illustrative example of a
turbomachine; and more specifically of a bypass axial
turbojet 1. The turbojet 1 shown has a fan 2, a lowpressure
compressor 3, a high-pressure compressor 4, a
combustion chamber 5, a high-pressure turbine 6, and a
30 low-pressure turbine 7. The fan 2 and the low-pressure
compressor 3 are connected to the high-pressure turbine 7
by a first transmission shaft 9, while the high-pressure
compressor 4 and the high-pressure turbine 6 are
connected together by a second transmission shaft 10. In
35 operation, a flow of air compressed by the high- and lowpressure
compressors 3 and 4 feeds combustion in the
combustion chamber 5, and the expansion of the combustion
8
gas drives the high- and low-pressure turbines 6 and 7.
The turbines 6 and 7 thus drive the fan 2 and the
compressors 3 and 4 by means of the shafts 9 and 10. The
air propelled by the fan 2 and the combustion gas leaving
5 the turbojet 1 via a propulsive nozzle (not shown)
downstream from the turbines 6 and 7 together exert
reaction thrust on the turbojet 1, and via the turbojet
on a vehicle such as an airplane (not shown) .
Each compressor 3, 4 and each turbine 6, 7 of the
10 turbojet 1 comprises a plurality of stages, each stage
being made up of a stationary set of vanes or stator, and
a rotary set of blades or rotor. An axial compressor
rotor 11 is shown diagrammatically in Figure 2A. The
rotor 11 has a plurality of blades 12 arranged radially
15 around the axis of rotation A of the rotor 11, which axis
is substantially parallel to the general direction of the
flow of working fluid through the turbojet 1. The blades
12 may be integrated as a single part in the rotor 11,
thereby forming a single-piece blade disk, or else they
20 may be made separately and joined to the rotor by
fastener means that are generally known in the state of
the art, such as fir-tree fasteners.
As shown in greater detail in Figure 2B, each blade
12 presents a three-dimensional reference frame having
25 three orthogonal axes X, Y, and Z. The axis X is
parallel :to the axis of rotation A of the rotor 11, the
axis Y is tangential to the direction of rotation R of
the blade 12 about the axis of rotation A, and the axis Z
is a radial axis in a direction intersecting the axis of
30 rotation A. Each blade 12 compris.es a blade root 13 and
a blade tip 14 that are spaced apart by a blade height h
in the direction of the radial axis Z. Between the blade
root 13 and the blade tip 14, the blade 12 has a stack of
aerodynamic profiles 15 in planes perpendicular to the
35 radial axis Z, forming a leading-edge 16 in the upstream
direction, a trailing edge 17 in the downstream
direction, a suction side 18, and a pressure side 19. In
9
a compressor or fan rotor, the direction of rotation R in
normal operation is such that each blade 12 moves towards
its pressure side 19.
One such profile 15 of the blade 12 is shown in
5 Figure 2C. Each profile 15 presents a chord C between
the leading edge 16 and the trailing edge 17, and a
center of gravity CG defined as the geometrical centroid
of the profile 15. In the field of turbomachine blades
or vanes, the slope of the line passing·through the
10 centers of gravity CG of the successive profiles 15
relat.ive to the radial axis Z is used to define the sweep
and the tangential slope of a blade or vane 12. Thus,
when, on going towards the tip 14, this line 20 slopes at
an angle of inclination -i rn an upstream direction in
15 the XZ plane, as shown in Figure 3A, the blade 12
presents a forward sweep. In contrast, when this line 20
slopes at an angle of inclination i in the downstream
direction in the same plane, as shown in Figure 3B, the
blade 12 presents a backward sweep. In similar manner,
20 the tangential slope is defined by the angle of
inclination of the line 20 relative to the radial axis Z
in the YZ plane. Thus, when, on going towards the tip
14, the line 20 slopes towards the suction side 18 (and
thus in the direction opposite to the direction of
25 rotation R of the rotor) , the blade 12 slopes at a
tangential angle of inclination -j that is negative as
shown in Figure 3C. In contrast, when this angle of
inclination is towards the pressure side 19 (and thus rn
the direction of rotation R of the rotor), the blade 12
30 slopes at a tangential angle of inclination j that is
positive, as shown in Figure 30. Apart from the sweep
and the tangential slope, turbomachine blades or vanes
generally present shapes that are complex, having
profiles 15 in which the angle of attack, the camber, the
35 thickness, and the chord C can also vary along the axis
z.
10
Figure 4 shows a blade or vane 112 in a first
embodiment of the invention that enables this drawback to
be mitigated for forwarding swept blades or vanes. This
blade 112 also has a blade root 113, a.blade tip 114, a
5 leading edge 117, a trailing edge 116, .a pressure side
118, and a suction side 119, and it is also made up of a
stack of aerodynamic profiles 115 over the blade height h
between the blade root 113 and the blade tip 114.
Figure 5 shows the yG relationship of this blade
10 112, i.e. the way the tangential coordinate yG of the
center of gravity CG varies along the radial axis Z. In
Figure 5, the yG abscissa axis is graduated in
millimeters. The angle of inclination of this curve,
i.e. the first derivative of yG, corresponds to the
15 tangential slope of the blade: thus, when the curve goes
towards the left, i.e. yG is negative, .the corresponding
portion of the blade has a negative tangential slope, and
when the curve goes towards the right, i.e. yG is
positive, the corresponding portion of the blade as a
2 0 positive tangential slope. The curvature of this curve,
i.e. the second derivative of yG, corresponds to the
curvature of the blade in the tangential direction.
In this figure, it can be seen that this blade 112
presents an intermediate segment 112a with a negative
25 tangential slope occupying about 70% of the blade height
h up to a dimension corresponding to about 75% of the
blade height h. The blade 112 also presents a distal
segment 112b with a positive tangential slope extending
between the intermediate segment 112a and the blade tip
30 114, and thus occupying about 25% of the blade height h.
It can also be seen that the intermediate segment
112a has a substantially rectilinear segment 122a
extending almost between the dimensions corresponding to
15% and 70% of the blade height h. The distal segment
35 112b also has a substantially rectilinear segment 122b
extending almost from the dimension corresponding to 90%
of the blade height h to the blade tip 114. The curve
11
for yG is thus Gamma-shaped. The curvature of the blade
is then concentrated in a restricted zone 122c of the
blade, mainly between the dimensions corresponding to 70%
and 90% of the blade height h: the curve for yG thus
5· turns through more than 90" in less than 20% of the blade
height h, thereby contributing to decoupling the forces
acting on the intermediate segment 112a and the distal
segment 112b of the blade 112.
It can also be seen that the distal segment 112b
10 extends strong from the positive yG side to reach about
15
4 mm at the blade tip, i.e. practically as much, in
absolute value, as the coordinate reached on the negative
yG side at the interface between the intermediate segment
112a and the distal segment 112b.
Figures 8A and BB are screen captures of software
for calculating mechanical stresses: they show the static
stress levels respectively on the suction side and on the
pressure side of the blade of the first example. For
comparison, Figures lOA and lOB show the static stress
20 levels respectively on the suction side and on the
pressure side of a reference prior art blade for which
the stacking relationship yG, as shown in Figure 8, is Sshaped.
It can thus be seen that the maximum stress level at
25 the suction side of the first example blade is
401 megapascals (MPa), while this maximum level is
542 MPa for the reference blade, i.e. a decrease of 26%.
On the pressure side, the maximum stress level is 368 MPa
for the blade of the first example compared with 457 MPa
30 for the reference blade, i.e. a decrease of 19%.
Figure 6 shows the yG relationship for a second
example blade. In this figure, it. can be seen that the
second blade presents substantially the same Gamma-shape
except that its distal segment 212b does not extend as
3 5 far as in the first example, reaching about 1. 5 mm at the
blade tip, i.e. about 30% in absolute value of the
coordinate reached with negative yG at the interface
12
between the intermediate segment 212a and the distal
segment 212b.
In spite of that, there can be seen the presence of
an intermediate segment 212a of negative tangential slope
5 that occupies about 7 0% of the blade height h up to a
dimension corresponding to about 75% of the blade height
h, and a distal segment 212b with a positive tangential
slope extending between the intermediate segment 212a and
the blade tip, and thus occupying about 25% of the blade
10 height h.
The intermediate segment 212a also has a
substantially rectilinear segment 222a extending almost
between the dimensions corresponding to 15% and to 70% of
the blade height h. The distal segment 212b also has a
15 substantially rectilinear segment 222b extending almost
from the dimension corresponding to 90% of the blade
height h to the blade tip 114. The curvature of the
blade is then concentrated in a restricted zone 222c of
the blade, mainly between the dimensions corresponding to
20 70% and to 90% of the blade height h: the curve for yG
thus turns through more than 90° in less than 20% of the
blade height h, thereby contributing to decoupling the
forces acting on the intermediate segment 112a and the
distal segment 112b of the blade.
25
30
Figures 9A and 98 are once more screen captures of
software for calculating mechanical stresses: they show
the static stress levels respectively on the suction side
and on the pressure side of the blade of the second
example.
It can thus be seen that the maximum stress level at
the suction side of the second example blade is 401 MPa,
while this maximum level is 542 MPa for the reference
blade, i.e. a decrease of 26%. On the pressure side, the
maximum stress level is 331 MPa for the blade of the
35 first example compared with 457 MPa for the reference
blade, i.e. a decrease of 28%.
13
The embodiments described in the present description
are given by way of nonlimiting illustration, and in the
light of this description, a person skilled in the art
can easily modify these embodiments, or envisage others,
5 while remaining within the ambit of the invention.
Furthermore, the various characteristics of these
embodiments may be used singly or they may be combined
with one another. When they are combined, these
characteristics may be combined as described above or in
10 other ways, the invention not being limited to the
specific combinations described in the present
description. In particular, unless specified to the
contrary, any characteristic described with reference any
one embodiment may be applied in analogous manner to any
15 other embodiment.
14

CLAIMS
1. A turbomachine rotor blade comprising a blade root
(113) and a blade tip (114) spaced apart by a blade
height (h), at least one intermediate segment (112a)
5 presenting a negative tangential slope, and a distal
segment (112b) situated between the intermediate segment
(112a) and the blade tip (114) and presenting a positive
tangential slope;
wherein the distal segment (112b) extends over at
10 most 30% of said blade height (h); and
wherein the projection onto a radial plane of the
line passing through the centers of gravity of each
tangential section (115) of the blade (110) possesses, in
the intermediate segment (112a) of the blade, a segment
15 (122a) that is substantially rectilinear and that extends
over at least 40% of the blade height (h).
2. A blade according to any preceding claim, wherein said
distal segment (112b) is directly adjacent to said blade
20 tip (114), and
wherein said distal segment (112b) is directly adjacent
to said intermediate segment (112a).
3. A blade according to either preceding claim, wherein
25 said distal segment (112b) extends over at most 25% of
the blade height (h) .
4. A blade according to any preceding claim, wherein said
distal segment (112b) extends over at least 5% of the
30 blade height (h).
5. A blade according to any one of claims 1 to 4, wherein
said substantially rectilinear segment (112a) extends at
least to a level situated at 50%, preferably 60%, of the
35 blade height (h) from the blade root (113).
5
15
6. A blade according to any one of claims 1 to 5, wherein
said substantially rectilinear segment (122a) extends
from a level situated at at most 30%, preferably at most
20%, of the blade height (h) from the blade root (113).
7. A blade according to any one of claims 1 to 6, wherein
the projection onto a radial plane of the line passing
through the centers of gravity of each tangential section
(115) of the blade (110) possesses, in the distal segment
10 (112b) of the blade, a second segment (122b) that is
substantially rectilinear, and that extends preferably
from a level situated at at most 90% of the blade height
(h) from the blade root (113).
15 8. A blade according to claim 7, wherein the second
substantially rectilinear segment (122b) is
subhorizontal.
9. A blade according to claim 7 or claim 8, wherein the
20 projection onto a radial plane of the line passing
through the centers of gravity of each tangential section
(115) of the blade (110) turns through more than 90° in
less than 20% of the blade height (h) .
25 10. A blade according to any preceding claim, wherein the
difference between the tangential coordinates (yG) of the
centers of gravity of the blade sections situated firstly
at the blade tip and secondly at the interface between
the intermediate segment (212a) and the distal segment
30 (212b) lies in the range 100% to 150%, preferably in the
range 110% to 140%, of the difference between the
tangential coordinates (yG) of the centers of gravity of
the blade sections. situated firstly at the blade root and
secondly at the interface between the intermediate
35 segment (212a) and the distal segment (212b) .
16
11. A blade according to any preceding claim, wherein the
thickness of the blade (112) is at all points less than
8% of the blade height (h), and preferably less than 6%.
5 12. A single-piece blade disk comprising a plurality of
blades (112) according to any preceding claim.
10
13. A rotor having a plurality of blades (112) according
to any one of claims 1 to 11.
14. A turbomachine including a disk according to claim 12
or a rotor according to claim 13.