A GAS TURBINE BLADE WITH TIP SECTIONS OFFSET TQWARDS THE
PRESSURE SIDE AND WITH COOLING CHANNELS
The field of the present invention relates to hollow
blades, in particular gas turbine blades, and more
particularly to the moving blades of turbine engines,
specifically the moving blades of a high pressure
turbine.
In known manner, a blade comprises in particular an
airfoil extending in a longitudinal direction, a root,
and a tip opposite from the root. For a moving turbine
blade, the blade is fastened to the disk of a turbine
rotor by means of its root. The blade tip is situated
facing the inside face of the stationary annular casing
surrounding the turbine. The longitudinal direction of
the airfoil corresponds to the radial direction of the
rotor or of the engine, with this being relative to the
axis of rotation of the rotor.
The airfoil may be subdivided into airfoil sections
that are stacked in a stacking direction that is radial
relative to the axis of rotation of the rotor disk. The
blade sections thus build up an aifoil surface that is
subjected directly to the gas passing through the
turbine. From upstream to downstream in the fluid flow
direction, this aifoil surface extends between a leading
edge and a trailing edge, these edges being connected
together by a pressure side face and a suction side face,
also referred to as the pressure side and the suction
side.
The turbine having such moving blades has a flow of
gas passing therethrough. The aerodynamic surfaces of
its blades are used for transforming a maximum amount of
the kinetic energy taken from the flow of gas into
mechanical energy that is transmitted to the rotary shaft
of the turbine rotor.
However, like any obstacle present in a gas flow,
the airfoil of the blade generates kinetic energy losses
that need to be minimized. In particular, it is known
that a non-negligible portion of these losses (in the
range 20% to 30% of total losses) can be attributed to
the presence of functional radial clearance between the
tip of each blade and the inside surface of the casing
5 surrounding the turbine. This radial clearance allows a
flow of gas to leak from the pressure side of the blade
(zone where pressure is higher) towards the suction side
(zone where pressure is lower). This leakage flow
represents a flow of gas that does no work and that does
10 not contribute to expansion in the turbine. Furthermore,
it also gives rise to turbulence at the tip of the blade
(known as the tip vortex), which turbulence generates
high levels of kinetic energy losses.
In order to solve that problem, it is known to
15 modify the stacking of the sections of the blade at the
level of the blade tip, in order to offset the stacking
towards the pressure side face, this offset preferably
taking place progressively, being more pronounced for
sections that are closer to the free end of the tip.
2 0 Blades of this type are referred to as blades with
an "advanced blade top" or as blades with a "tip section
off set".
Furthermore, turbine blades, and in particular the
moving blades of a high pressure turbine, are subjected
25 to high temperature levels by the external gas coming
from the combustion chamber. These temperature levels
exceed the temperatures that can be accepted by the
material from which the blade is made, thus requiring the
blades to be cooled. Recently-designed engines have
30 ever-increasing temperature levels for the purpose of
improving overall performance, and these temperatures
make it necessary to install innovative cooling systems
for the high pressure turbine blades in order to ensure
that these parts have a lifetime that is acceptable.
35 The hottest location in a moving blade is its tip,
so cooling systems seek firstly to cool the top of the
blade.
A w-i de variety of techniques have already been
proposed for cooling blade tips, and mention may be made
in particular to those described in EP 1 505 258,
FR 2 891 003, and EP 1 726 783.
5 Consequently, it can be understood that the
particular configuration that arises when using the "tip
section offset" technique disturbs the performance and
the effectiveness of conventional cooling systems in the
tip zone of the blade.
10 Unfortunately, the top of a blade is always the
hottest location of a moving blade, so it is essential
for the "tip section offset" technique to be capable of
coexisting with a cooling system that remains effective
in order to conserve a lifetime for the part in this zone ,
15 that is sufficient when subjected to high temperature
conditions upstream.
It is found that those solutions are not compatible
with the "tip section offset" technique.
An object of the present invention is thus to
20 propose a blade structure that makes it possible to
conserve high effectiveness of the cooling system at the
top of a blade, even when the blade has an advanced top
of the "tip section offset" type.
To this end, the present invention relates to a
25 hollow blade having an airfoil extending along a
longitudinal direction, a root, and a tip, an internal
cooling passage inside the airfoil, a cavity (or
"bathtub") situated in the tip, being open towards the
free end of the blade and defined by an end wall and a
30 rim, said rim extending between the leading edge and the
trailing edge and comprising a suction side rim along the
suction side and a pressure side rim along the pressure
side, and cooling channels connecting said internal
cooling passage with the pressure side, said cooling
35 channels sloping relative to the pressure side, the stack
of airfoil sections of the blade at the level of the rim
of the blade tip presenting an offset towards the
pressure side, this offset increasing on approaching the
free end of the tip of the blade.
This hollow blade is characterized in that the
pressure side wall of the airfoil presents a projecting
5 portion with more than half of its length extending along
a longitudinal portion of the internal cooling passage,
and with an outside face that slopes relative to the
remainder of the pressure side of the airfoil, and
presenting a terminal face at its end facing towards the
10 cavity, the end wall being connected to the pressure side
wall at the location of said end of the projecting
portion and said cooling channels being arranged in said
projecting portion in such a manner as to open out in the
terminal face of said projecting portion, whereby the
15 distance -d between the axes of the cooling channels and
the outer limit A of the free end of the pressure side
rim is greater than or equal to a non-zero minimum value
dl. This value dl thus corresponds to a threshold value
that is predetermined depending on the type of blade and
20 on the operating conditions that apply to drilling the
channels.
Overall, by means of the solution of the present
invention, the position of the pressure side wall portion
that includes the cooling channels is offset towards the
25 pressure side so as to enable drilling tools to access
the appropriate location, while not degrading the
performance of the cooling, and possibly even while
improving it.
This solution also presents the additional advantage
30 of making it possible to further improve the cooling of
the pressure side wall portion carrying the cooling
channels by means of thermal pumping so as to obtain
better film cooling of the pressure side rim of the
cavity (or bathtub).
35 The present invention also provides a turbine engine
rotor, a turbine engine turbine, and a turbine engine
including at least one blade as defined in the present
specification.
Other advantages and characteristics of the
invention appear on reading the following description
5 made by way of example and with reference to the
accompanying drawings, in which:
Figure 1 is a perspective view of a conventional
hollow rotor blade for a gas turbine;
Figure 2 is a perspective view on a larger scale
10 of the free end of the Figure 1 blade;
Figure 3 is a view analogous to the view of
Figure 2, but partially in longitudinal section after the
trailing edge of the blade has been removed;
Figure 4 is a fragmentary longitudinal section
15 view on line IV-IV of Figure 3;
Figures 5 to 7 are views similar to the view of
Figure 4, for blades incorporating the "tip section
offset" technique;
Figures 8 and 9 show the solution of the present
20 invention; and
Figures 10 and 11 are views similar to the view of
Figure 8 for first and second variant embodiments.
In the present application, unless specified to the
contrary, upstream and downstream are defined relative to
25 the normal flow direction of gas through the turbine
engine (from upstream to downstream). Furthermore, the
term "axis of the engine" is used to designate the axis
X-X' of radial symmetry of the engine. The axial
direction corresponds to the direction of the axis of the
30 engine, and a radial direction is a direction
perpendicular to said axis and intersecting it.
Likewise, an axial plane is a plane containing the axis
of the engine, and a radial plane is a plane
perpendicular to said axis and intersecting it. The
35 transverse (or circumferential) direction is a direction
perpendicular to the axis of the engine and not
intersecting it. Unless specified to the contrary, the
adjectives axial, radial, and transverse (and the adverbs
axially, radially, and transversely) are used relative to
the above-specified axial, radial, and transverse
directions. Finally, unless specified to the contrary,
5 the adjectives inner and outer are used relative to the
radial direction such that an inner (i.e. radially inner)
portion or face of an element is closer to the axis of
the engine than is an outer (i.e. radially outer) portion
or face of the same element.
10 Figure 1 is a perspective view of an example of a
conventional hollow rotor blade 10 for a gas turbine.
Cooling air (not shown) flows inside the blade from the
bottom of the root 12 of the blade, along the airfoil 13,
in a longitudinal direction R-R' of the blade 13 (the
15 vertical direction in the figure and the radial direction
relative to the axis of rotation X-X' of the rotor),
towards the tip 14 of the blade (at the top in Figure I),
and this cooling air then escapes via an outlet to join
the main gas stream.
20 In particular, this cooling air flows in an internal
cooling passage situated inside the blade and terminating
at the tip 14 of the blade in through holes 15.
The body of the blade is profiled so as to define a
pressure side wall 16 (to the left in all of the figures)
25 and a suction side wall 18 (to the right in all of the
figures) .
The pressure side wall 16 is generally concave in
shape and it is the first wall encountered by the hot gas
stream, i.e. its outside face facing upstream is on the
30 gas pressure side and is referred to as the "pressure
side face" or more simply as the "pressure side" 16a.
The suction side wall 18 is convex and encounters
the hot gas stream subsequently, i.e. it is on the gas
suction side along its outer face that faces downstream
35 and referred to as the "suction side face" or more simply
as the "suction side" 18a.
The pressure and suction side walls 16 and 18 meet
at a leading edge 20 and at a trailing edge 22 that
extend radially between the tip 14 of the blade and the
top of the root 12 of the blade.
5 As can be seen from the enlarged views of Figures 2
to 4, at the tip 14 of the blade, the internal cooling
passage 24 is defined by the inside face 26a of an end
wall 26 that extends over the entire tip 14 of the blade
between the pressure side wall 16 and the suction side
10 wall 18, and thus from the leading edge 20 to the
trailing edge 22.
At the tip 14 of the blade, the pressure and suction
side walls 16 and 18 form a rim 28 of a cavity 30 that is
open facing away from the internal cooling passage 24,
15 i.e. radially outwards (upwards in all of the figures).
More precisely, the rim 28 is constituted by a pressure
side rim 281 beside the pressure side wall 16 and a
suction side rim 282 beside the suction side wall 18.
As can be seen in the figures, this open cavity 30
20 is thus defined laterally by the inner face of the rim 28
and in its low portion by the outer face 26b of the end
wall 26.
The rim 28 thus forms a thin wall along the profile
of the blade that protects the Free end of the tip 14 of
25 the blade 10 from making contact with the corresponding
inner annular surface of the turbine casing 50 (see
Figure 4) .
As can be seen more clearly in the section view of
Figure 4, which shows the prior art cooling technology
30 involving holes under the bathtub, sloping cooling
channels 32 pass through the pressure side wall 16 in
order to connect the internal cooling passage 24 to the
outside face of the pressure side wall 16, i.e. the
pressure side 16a.
3 5 These cooling channels 32 slope so as to open out
towards the top 28a of the rim in order to cool it by
means of a jet of air that goes towards the top 28a of
the rim 2' along the pressure side wall 16.
The effectiveness of the cooling that results from
these cooling channels 32 is governed mainly by two
5 geometrical parameters of these cooling channels 32 (see
Figure 4) :
the total radial extent D of the cooling channels
32 between the two radii R1 and R2 (respectively the
height of the inlet opening 32b and the height of the
10 outlet opening 32a of the cooling channels 32 in the
pressure side 16); the greater this radial extent D, the
more the phenomenon of cooling by thermal pumping applies
to a large portion of the blade along the axis R-R'; and
the height of the outlet openings 32a of the
15 cooling channels 32 in the pressure side 16 specified by
the radius R2 referred to as the "outlet" radius; the
greater this radius R2, the more effective the external
film of cooling air all the way to the top of the
bathtub, i.e. the top 28a of the pressure side rim 281.
20 Finally, the industrial feasibility of making
cooling channels 32 (which are generally made by electron
discharge machining (EDM)), requires an angle a between
the axis of the cooling channel 32 and the outside face
281a of the pressure side rim 281 that is sufficient to
25 leave enough clearance to allow the EDM nozzle to pass.
It can be seen that if the geometrical configuration
of the cooling channel 32 in Figure 4 is used unchanged
for a blade 10' that also includes a "tip section offset"
(Figure 5), then the clearance of the axis of the cooling
30 channel 32 (angle a) is no longer sufficient. Under such
circumstances, the axis of the cooling channel 32
interferes with the pressure side rim 281', either by
being too close to it or by intersecting it as shown in
Figure 5. It is thus no longer possible to make the
35 cooling channel 32 by drilling.
In Figure 5, the blade 10' with a "tip section
offset" is given the same reference signs as those used
for the blade in Figures 1 to 4, together with a prime
symbol (" "') for portions that are modified.
Specifically, the differences relate solely to the shape
of the rim 28' that is no longer parallel to the
5 longitudinal direction R-R' of the blade lo', i.e. to the
radial direction.
The sections S of the airfoil are considered as
corresponding to the outline of the airfoil in sections
on section planes that are orthogonal to the longitudinal
10 direction R-R' of the blade, i.e. the radial direction.
For the blade 10, all of the airfoil sections S are
stacked in a stacking direction parallel to the
longitudinal direction R-R' of the blade, i.e. the radial
direction, the sections being superposed on one another
15 (see Figure 4).
For the blade 10' in Figure 5, the airfoil sections
S of the airfoil portion including the internal cooling
passage 24 and the end wall 26 are likewise stacked in
the radial direction of the blade; nevertheless, the
20 airfoil sections S1, S2, S3, and S4 of the rim 28' (i.e.
the tip sections) are stacked so that their stacking is
offset towards the pressure side 16a, w'ith this taking
place progressively and increasingly for sections closer
to the top 28a' (in the order S1, S2, 53, and 54 in
25 Figure 5).
"A" designates the outer limit of the free end of
the pressure side rim 281', with this being referred to
below as the end A of the pressure side rim 281'.
Furthermore, the rim 28' shown also has an
30 enlargement 283' in the pressure side rim 281' at the
location of the outer limit A of the free end of said
pressure side rim 281', i.e. at the location of the
margin of the pressure side at the top 28a1.
This enlargement 283' is present in some of the
35 stacked sections (53 and S4) of Figure 5 and leads to the
end A having a pointed shape in section, with the axis of
the cooling channel 32 intersecting this pointed shape.
This pointed shape, which appears during the machining of
the blade 10, should be considered as being optional and
not essential.
In order to mitigate this problem and to make a tip
5 section offset compatible with holes under the bathtub,
it is natural to modify the shape of the bathtub and thus
to degrade its thermal efficiency:
a first solution, as shown in Figure 6, has
cooling channels 32' that are easily drilled, by reducing
10 theheight of the outlet radius R2 to the value R2'
without modifying the total radial extent D (the height
of the cooling channel inlet radius R1 is lowered to the
value R1'); under such circumstances, by reducing the
radius R2 and lowering the position of the outlets from
15 the cooling channels, it is no longer possible to obtain
satisfactory cooling of the blade tip formed by the rim
28'; and
a second solution, as shown in Figure 7, has
cooling channels 32" that are easy to drill, and consists
20 in reducing the total radial extent D to a value D"
without changing the height of the outlet radius R2;
under such circumstances, by increasing the radius R1 to
a value Rl", it is possible to obtain satisfactory
cooling of the blade tip formed by the rim 28', but the
25 phenomenon of thermal cooling by pumping is no longer
sufficient, since it is effective over only a small
portion of the blade along the axis R-R'.
In order to mitigate those drawbacks, the present
invention proposes the solution presented in Figures 8 to
30 11 and described below.
The blade 110 has a rim 28' provided with a tip
section offset as described above with reference to
Figure 5.
The pressure side wall 16 is modified in its
35 intermediate portion that is adjacent to the pressure
side rim 281', in that this intermediate portion forms a
protrusion towards the pressure side 16a.
More precisely, the intermediate portion is a
projecting portion 161 such that, in this projecting
portion, the pressure side 16a is no longer directed in
the longitudinal direction R-R', i.e. the radial
5 direction, but slopes so as to depart progressively
further from the suction side 18a on approaching the rim
28' in the longitudinal direction R-R'.
More than half the length of this projecting portion
161 extends along a longitudinal portion of the internal
10 cooling passage 24 (specifically the radially outermost
portion in the assembled engine).
By offsetting the pressure side wall 16 in this way
where the hole is drilled, it is possible to conserve the
radii R2 and R1 of Figure 4 and to move the axis of the
15 cooling channels 132 at the end A of the pressure side
rim 281' far enough away to allow drilling to be
undertaken.
This projecting portion 161 extends over the full
height of the cooling channels 132 between the radii R2
20 and R1 (where R2>R1) and is visible on the pressure side
16a in the form of an outside face or pressure side face
161a, a terminal face 161b facing towards the rim 28',
and an internal face 161c facing towards the internal
cooling passage 24.
25 The pressure side face 161a of the projecting
portion 161 slopes progressively away from the radial
direction R-R' on approaching the terminal face 161b.
The angle of inclination P formed between the pressure
side face 161a of the projecting portion 161 and the
30 longitudinal direction R-R', i.e. the radial direction,
preferably lies in the range 10" to GO0, more preferably
in the range 20" to 50°, and advantayeously in the range
25" to 35", in particular being close to 30°.
Furthermore, the angle of inclination a of the
35 cooling channels 132 relative to the loniitudinal
direction R-R', i.e. the radial direction, lies in the
range 10" to 60°, preferably in the range 20" to 50°, and
advantageously in the range 25" to 35O, specifically
being close to 30°.
With this configuration, a non-zero minimum distance
dl is available on measuring the difference -d between the
5 parallel to the longitudinal direction R-R' passing
through the end A of the pressure side rim 281' and the
end B or outer edge of the projecting portion 161 as
situated between the pressure side face 161a and the
terminal face 161b. In other words, the end B is set
10 back relative to the end A.
Preferably, said minimum value dl is greater than or
equal to 1 millimeter (rnm), or indeed 2 mm, and depends
on the material used for performing the drilling of the
cooling channels 132.
15 In characteristic manner, said cooling channels 132
are arranged in the projecting portion 161 so as to open
out into the terminal face 161b of said projecting
portion 161.
In this way, a stream of cooling air F1 is obtained
20 (see Figure 8) that is pushed back by the external flow
of hot gas passing from the pressure side 16a towards the
suction side 18a via the clearance that exists between
the top of the blade and the corresponding inner annular
surface of the turbine casing 50 as a result of the
25 positive pressure gradient between the pressure side 16a
and the suction side 18a.
This configuration generates a stream F2 in a
recirculation zone (corner zone) that ensures effective
mixing between the cooling gas stream F1 and the external
30 hot gas, regardless of the position of the outlet
openings of the cooling channels 132 in the terminal face
161b of said projecting portion 161.
Thus, the use of a projecting portion 161 of the
invention makes it possible to further improve the
35 effectiveness of the cooling generated by the air coming
from the cooling channels 132.
In a preferred geometrical arrangement shown in
Figures 8 to 11 the distance A (see Figure 9) between the
end B of the terminal face 161b of the projecting portion
161 and the remainder of the pressure side wall 16 is not
5 less than the difference between firstly the offset E
measured between the end A of the pressure side rim 281'
and the remainder of the pressure side wall 16, and
secondly said distance -d between the axes of the cooling
channels 132 and the end A of the pressure side rim 281';
10 this distance A corresponds to the axial extent of the
terminal face 161b of said projecting portion 161. In
other words:
A 2 E - d .
In order to avoid increasing the weight of the
15 structure, the thickness -e of the pressure side wall 16
of the airfoil of the blade 110 is substantially constant
both in the projecting portion 161 and in the remainder
of the pressure side wall 16, and is also substantially
equal to the thickness of the wall in the zone 161d of
20 the projecting portion 161 (see Figure 9) connected to
the end wall level with and in front of the base of the
pressure side rim 281'.
It should be observed that the wall thicknesses are
considered along a direction orthogonal to the outside
25 face of the zone under consideration.
This characteristic is shown in Figure 9, where this
thickness - e can be seen: below the projecting portion
161; at locations in the projecting portion 161 along the
cooling channels 132; and in the zone 161d situated
30 between the terminal face 161b and the internal cooling
passage, and connecting the projecting portion 161 to the
end wall 26.
In order to avoid penalizing the mechanical
robustness of the blade root 12, it is necessary to avoid
35 thickening the pressure side wall 16 at the location of
the projecting portion 161. For this purpose, the rear
face of the pressure side wall is cut away in the
location of the projecting portion 161. Specifically,
the zone to be removed behind the projecting portion 161
compared with the conventional profile for the pressure
side wall 16 and represented by lines PI and P2 in
Figure 8 corresponds to the shaded.zone referenced C in
Figure 9.
Advantageously, this design in accordance with the
invention with a projecting portion 161 that does not
involve increasing wall thickness can be obtained with a
minimum of modification to existing tooling; for casting,
the already existing core box is dug into for a volume
equivalent to the extruded surface C (across the entire
width of the pressure side) so as to produce cores having
the inside profile of the cavity suitable for obtaining
the projecting portion 161, and this volume is dug away
from the wax mold forming the outer envelope of the
blade.
In this configuration, the outside face 161a and the
inside face 161c of the projecting portion 161 are
mutually parallel.
The terminal face 161b of the projecting portion 161
is preferably plane.
In Figures 8 and 9, the terminal face 161b of the
projecting portion 161 is horizontal; it is directed
orthogonally to the longitudinal direction R-R' of the
blade at the location where the cooling channels 132 open
out into said terminal face 161b.
In the example shown, the entire terminal face 161b
of the projecting portion 161 extends orthogonally to the
longitudinal direction R-R' of the blade.
In a first variant shorin in Figure 10, a chamfer is
used at the terminal face 161b, so that the terminal face
161b of the projecting portion 161 is inclined so as to
form a non-zero obtuse angle y l with the longitudinal
direction R-R' of the blade at the location where the
cooling channels 132 open out into said terminal face
161b. In this arrangement, an acute angle y2 is formed
between the terminal face-I6lb of the projecting portion
161 and the horizontal direction parallel to the rotary
axis X-X' of the rotor and orthogonal to the longitudinal
direction R-R' of the blade. This angle y2 preferably
5 lies in the range 10" to 60°, more preferably in the
range 20° to 50°, and advantageously in the range 25" to
35", and in particular it is close to 30°.
In this way, the axis of the cooling channels 132 is
orthogonal to the terminal face 161b of the projecting
10 portion 161 at the location where the cooling channels
132 open out into said terminal face 161b. The advantage
of this variant is that the shape of the outlet openings
of the cooling channels 132 in the terminal face 161b is
round, in contrast to the more oval shape when the
15 terminal face 161b is horizontal, thus making it possible
to obtain better control over the outlet section of the
cooling channels 132, and thus over the flow rate of
cooling air.
In Figures 8 to 10, the end wall 26 extends
20 orthogonally to the longitudinal direction R-R' of the
blade, which corresponds to a conventional configuration.
Furthermore, in Figures 8 to 'lo, the terminal face
161b of the projecting portion 161 is arranged at the
height of the outlet radius R2 that is less than the
25 radius R3 corresponding to the outside face 26b of the
end wall 26 (see Figures 8 and 9) that faces towards the
cavity 30. Thus, R2R3, then
the bottom of the bathtub would not be impacted by the
30 cooling coming from the cooling channel 32).
Also, in these Figures 8 to 10, the terminal face
161b of the projecting portion 161 is located at the
height of the outlet radius R2 that is greater than the
radius R4 corresponding to the inside face 26a of the end
35 wall 26 (see Figures 8 and 9) that faces towards the
internal cooling passage 24. This situation with R2>R4
makes it possible to guarantee that the blade 110 is
properly cooled above the zone that is not thermally __
covered by the cooling generated by the cavity 30.
Consequently, having R2R4 represents the
best thermal compromise that can be found.
5 In the second variant of Figure 11, a bathtub is
used having a sloping bottom wall with the end wall 126
sloping to form an angle 61 that is not a right angle and
that is not zero relative to the longitudinal direction
R-R' of the blade.
10 More precisely, the top face of said end wall 126 in
the location adjacent to the pressure side rim 281' forms
an acute angle 61 that preferably lies in the range 45O
to 8 g 0 , more preferably in the range 50" to 65', and
advantageously in the range 55" to 65", specifically
15 being close to 60°, which corresponds to an acute angle
62 between the top face of said end wall 126 and the
horizontal direction parallel to the axis of rotation
X-X' of the rotor and orthogonal to the longitudinal
direction R-R' of the blade.
20
CLAIMS
1. A hollow blade (110) having an airfoil (13) extending
along a longitudinal direction (R-R'), a root (12), and a
tip (14), an internal cooling passage (24) inside the
5 airfoil, a cavity (30) situated in the tip, being open
towards the free end (14) of the blade (110) and defined
by an end wall (26, 126) and a rim (28' ), said rim (28')
extending between the leading edge (20) and the trailing
edge (22) and comprising a suction side rim (282') along
10 the suction side (18a) and a pressure side rim (281')
along the pressure side (16a), and cooling channels (132)
connecting said internal cooling passage (24) with the
pressure side (16), said cooling channels (32) sloping
relative to the pressure side (16a), the stack of airfoil
15 sections (S, S2, S3, 54) of the blade at the level of the
rim (28') of the blade tip presenting an offset towards
the pressure side (16a), this offset increasing on
approaching the free end of the tip (14) of the blade
(110), the blade being characterized in that the pressure
20 side wall (16) of the airfoil presents a projecting
portion (161) with more than half of its length extending
along a longitudinal portion of the internal cooling
passage (24), and with the outside face (161a) that
slopes relative to the remainder of the pressure side
25 (16a) of the airfoil, and presenting a terminal face
(161b) at its end facing towards the cavity (30), the end
wall (26) being connected to the pressure side wall (16)
at the location of said end of the projecting portion
(161) and said cooling channels (132) being arranged in
30 said projecting portion (161) in such a manner as to open
out in the terminal face (161b) of said projecting
portion (161), whereby the distance d between the - axes of
the cooling channels (132) and the outer limit A of the
free end of the pressure side rim (281') is greater than
35 or equal to a non-zero minimum value dl.
2. A blade accord-ing to any one of the preceding claims,
characterized in that said minimum value dl is greater
than or equal to 1 mm.
5 3. A blade (110) according to either preceding claim,
characterized in that the distance (A) between the end
(B) of the terminal face (161b) of the projecting portion
(161) and the remainder of the pressure side wall (16) is
not less than the difference between the offset (E)
10 measured between the end (A) of the pressure side rim
(281') and the remainder of the pressure side wall (16)
and said distance (-d ) between the axes of the cooling
channels (132) and the end (A) of the pressure side rim
(281') .
15
4. A blade (110) according to any preceding claim,
characterized in that the thickness (-e ) of the pressure
side wall (16) of the airfoil is substantially constant
in the projecting portion (161) and in the remainder of
20 the pressure side wall (16).
5. A blade (110) according to any preceding claim,
characterized in that the outside face (161a) and the
inside face (161~)o f the projecting portion (161) are
25 mutually parallel.
6. A blade (110) according to any preceding claim,
characterized in that the terminal face (161b) of the
projecting portion (161) is plane.
30
7. A blade (110) according to claim 6, characterized in
that the terminal face (161b) of the projecting portion
(161) slopes to form a non-zero obtuse angle yl relative
to the longitudinal direction (R-R') of the blade at the
35 location where the cooling channels (132) open out into
said terminal face (161b) .
. . ' 8. A blade (110) according to the preceding claim,
characterized in that the axes of the cooling channels
(132) are orthogonal to the terminal face (161b) of the
projecting portion (161) at the location where the
5 cooling channels (132) open out into said terminal Pace
(161b) .
9. A blade (110) according to any preceding claim,
characterized in that said end wall (26) is arranged
10 ~Fthogonally relative to the longitudinal direction of
the blade.
10. A blade (110) according to any one of claims 1 to 8,
characterized in that said end wall (126.) extends along a
15 slope so as to form a non-zero angle (61) other than a
right angle relative to the longitudinal direction (R-R')
of the blade (110).
11. A turbine engine rotor including at least one-blade
(110) according to any one of claims 1 to 10.
12. A turbine engine turbine including at least one blade
(IlQ) according to any one of claims 1 to 10.
13. A turbine engine including at least one blade (110)
according to any one of claims 1 tolo.