U.S. patent application number 13/497302 was filed with the patent office on 2012-09-13 for airfoil and corresponding guide vane, blade, gas turbine and turbomachine.
Invention is credited to Andrew Shepherd, Philip Twell.
Application Number | 20120230818 13/497302 |
Document ID | / |
Family ID | 41409226 |
Filed Date | 2012-09-13 |
United States Patent
Application |
20120230818 |
Kind Code |
A1 |
Shepherd; Andrew ; et
al. |
September 13, 2012 |
Airfoil and corresponding guide vane, blade, gas turbine and
turbomachine
Abstract
An airfoil suitable for being arranged in a compressor section
of the gas turbine includes a platform, a pressure surface and a
suction surface, and a tip region. The tip region is located
between the pressure and the suction surface opposite to the
platform in respect of the pressure and the suction surfaces, being
a distant end of the airfoil in direction along the first direction
perpendicular to the platform. The tip region includes a first
surface with a first distance to the platform along the first
direction, and a second surface with a second distance to the
platform along the first direction. The first distance is greater
than the second distance and the first surface and the second
surface together build a profiled surface. The profiled surface is
a ribbed surface. The material of the first surface is deformable
to bend away from the rotation direction of the airfoil.
Inventors: |
Shepherd; Andrew; (Branston,
GB) ; Twell; Philip; (Lincoln, GB) |
Family ID: |
41409226 |
Appl. No.: |
13/497302 |
Filed: |
August 9, 2010 |
PCT Filed: |
August 9, 2010 |
PCT NO: |
PCT/EP10/61553 |
371 Date: |
March 21, 2012 |
Current U.S.
Class: |
415/208.1 ;
416/236R |
Current CPC
Class: |
F01D 5/20 20130101; F01D
11/127 20130101; F16J 15/44 20130101; F16J 15/4472 20130101 |
Class at
Publication: |
415/208.1 ;
416/236.R |
International
Class: |
F01D 5/20 20060101
F01D005/20; F01D 1/04 20060101 F01D001/04 |
Foreign Application Data
Date |
Code |
Application Number |
Sep 30, 2009 |
EP |
09012405.8 |
Claims
1-10. (canceled)
11. An airfoil of a gas turbine, the airfoil suitable for being
arranged in a compressor section of the gas turbine, the airfoil
comprising: a platform; a pressure surface and a suction surface,
extending substantially perpendicular to the platform along a first
direction, the first direction being defined as a direction
substantially perpendicular to the platform and progressing with a
distance to the platform; a tip region between the pressure and the
suction surface opposite to the platform, the tip region being a
distant end of the airfoil in direction along the first direction
and being configured such that the tip region comprises: a first
surface with a first surface expanse less than a cross sectional
plane of the airfoil perpendicular to the first direction and with
a first distance to the platform along the first direction, a
second surface with a second surface expanse and with a second
distance to the platform along the first direction, wherein the
first distance is greater than the second distance, and wherein the
first surface and the second surface together build a profiled
surface, the profiled surface being a ribbed surface, and wherein
the material of the first surface is deformable such that the
material of the first surface bends away from a direction of
rotation of the airfoil during operation.
12. The airfoil according to claim 11, wherein the first surface
includes a plurality of elevated areas of the profiled surface and
the second surface includes a plurality of recessed areas of the
profiled surface.
13. The airfoil according to claim 12, wherein the plurality of
elevated areas of the profiled surface and the plurality of
recessed areas of the profiled surface are arranged
alternating.
14. The airfoil according to claim 11, the elevated areas and the
recessed areas are arranged as ribs between the suction surface and
the pressure surface.
15. The airfoil according to claim 11, the elevated areas and the
recessed areas are arranged as ribs substantially parallel to the
suction surface or to the pressure surface.
16. The airfoil according to claim 11, the elevated areas and the
recessed areas are arranged in a grid pattern.
17. The airfoil according to claim 16, wherein the grid pattern has
a shape of a first plurality of ribs intersected with a second
plurality of ribs, the first plurality of ribs being diagonal in
direction of a trailing edge of the airfoil and the second
plurality of ribs being diagonal in direction of a leading edge of
the airfoil.
18. The airfoil according to claim 16, wherein the grid pattern
having a shape of an inverse profile of a first plurality of ribs
intersected with a second plurality of ribs, the first plurality of
ribs being diagonal in direction of a trailing edge of the airfoil
and the second plurality of ribs being diagonal in direction of a
leading edge of the airfoil.
19. The airfoil according to claim 11, wherein the tip region is
coated with an abrasive coating for abrading an opposing
surface.
20. The airfoil according to claim 11, wherein the tip region is
coated with an abradable coating.
21. A component of a gas turbine, comprising an airfoil according
to claim 11.
22. The component according to claim 21, wherein the component is a
variable guide vane.
23. The component according to claim 21, wherein the component is a
blade.
24. A gas turbine, comprising: at least one of a compressor section
or a turbine section, the compressor section or the turbine section
comprising at least one airfoil configured according to claim 11.
Description
FIELD OF THE INVENTION
[0001] The invention relates to an airfoil, a guide vane, and a
blade, particularly of a compressor or a turbine section of a
turbomachine. Furthermore the invention relates to a turbomachine,
especially a gas turbine, comprising at least one airfoil.
BACKGROUND OF THE INVENTION
[0002] A conventional gas turbine engine comprises a compressor
which is used for pressurising ambient air. The compressed air is
mixed with fuel and burned in a combustor. Energy is extracted from
the hot combustion gas by passing it through a turbine section of
the gas turbine engine. The rotatable sections of an engine
typically comprise annular arrays of compressor or turbine rotor
blades, the blades normally being intersected with annular arrays
of static aerodynamic guide vanes, also called stator vanes. Each
disc of guide vanes and each disc of rotor blades are arranged
axially alternately. One disc of rotor blades and one disc of guide
vanes are referred to as a stage. The guide vanes ensure the gas
impinges on the rotor blades at the correct angle.
[0003] Optionally, the first stages of a compressor may be equipped
with variable guide vanes that are pivotable to adapt the stagger
angle during different operating modes of the compressor.
[0004] A rotor comprises a plurality of rotor blades arranged at
even intervals in a circumferential direction, each of which is
obliquely positioned to both the front direction and the rotating
direction so as to compress air downstream by rotation thereof. In
each blade, a first face--e.g. directed forward--is to suck air and
therefore referred to as a suction side or suction surface and a
second face--e.g. directed aftward--is to compress air and referred
to as a pressure side or pressure surface within this application.
To realise the compression of the air, the suction side is
typically made convex and the pressure side is made concave.
[0005] For a good performance of a turbine it is important to
establish proper clearance between the rotating blades and a
stationary component which encloses the rotating blades. This
clearance may be defined as the distance between these mentioned
elements after their assembly but prior to initial operation of the
turbine (cold clearance measurement).
[0006] It is a goal to have as little leakage from the path of the
air within the compressor. Hardly any air should be able to leave
the path of the air because this would lead to an unwanted drop of
air pressure. Besides, the complete air stream should be compressed
in the downstream direction. Reverse air flow should be prevented.
Thus, all gaps between rotor and stator should be limited to a
minimum so that all air that is present in the path of air will be
compressed and guided in the downstream direction. This also
includes gaps at the tip of the rotor blades between the blades and
a surrounding compressor component or gaps at the tip of the stator
vanes between the stator vane and the inner rotating shaft.
[0007] To minimise these gaps, a distal end of each blade--i.e. a
tip section of the blade--may be in frictional contact with an
inner face of a case of the compressor and may be coated with a
hard coating to abrade the inner face of the case of the
compressor. Because of the abrasiveness of the hard coating the
distal end itself is protected from deterioration by frictional
contact. The opposite member preferentially wears in comparison
between the tip section and the opposite member. Such a
configuration allows a perfect match of the tip of the blade and
the opposing compressor casing, even with manufacturing or assembly
tolerances.
[0008] According to EP 1 930 547 A2 a hard coating is applied to a
specific face of the tip of a blade. This may lead to an improved
fatigue lifetime of the blade. On the other hand the use of special
coatings may be costly.
[0009] In U.S. Pat. No. 4,671,735 A1 a rotor of an axial-flow
compressor is discussed with means for sealing the rotor blade tips
relative to a casing wall provided with a coating is to be capable
of abrasion but at the same time also low in wear.
[0010] In U.S. Pat. No. 4,295,786 A1 a gas path seal suitable for
use with a turbine engine or compressor is provided. A shroud
wearable or abradable by the abrasion of the rotor blades of the
turbine or compressor shrouds the rotor blades. With this approach
close tolerance between the tips of the blades and the surrounding
shroud or housing which seals one side of the blades from the other
can be reach even though the clearance dimensions are dynamic, i.e.
the clearance dimensions increase or decrease with temperature and
with mechanical and aerodynamic forces and may increase or decrease
faster than the rotor.
[0011] U.S. Pat. No. 4,526,509 A1 discloses a further type of seal
that is in rubbing contact with a tip of a rotor blade, but not
being abradable but resilient.
SUMMARY OF THE INVENTION
[0012] The present invention seeks to mitigate drawbacks of the
prior art. Specifically it is an object of the invention to provide
an airfoil or a component comprising such an airfoil that leakage
over the tip of the airfoil is reduced, by also considering
material costs and/or durability of the airfoil.
[0013] This objective is achieved by the independent claims. The
dependent claims describe advantageous developments and
modifications of the invention.
[0014] In accordance with the invention there is provided an
airfoil of a turbomachine, particularly a gas turbine, the airfoil
suitable for being arranged in a compressor section of the gas
turbine, the airfoil being in particular a vane or a blade. The gas
turbine may be an axial flow turbine and the compressor section an
axial flow compressor. Furthermore, the invention may also apply to
blades or vanes within a turbine section. The airfoil comprises a
platform, a pressure surface, a suction surface, and a tip region.
The pressure surface and a suction surface both extend
substantially perpendicular to the platform along a first
direction, the first direction being defined as a direction
substantially perpendicular to the platform and progressing with a
distance to the platform. The tip region is located between the
pressure and the suction surface opposite to the platform in
respect of the pressure and the suction surfaces. The tip region is
defined as a distant end of the airfoil in direction along the
first direction and is configured such that the tip region
comprises a first and a second surface. The first surface has a
first surface expanse less than a cross sectional plane of the
airfoil perpendicular to the first direction and has a first
distance to the platform along the first direction. The second
surface has a second surface expanse and a second distance to the
platform along the first direction. Additionally the first distance
is greater than the second distance. Furthermore the first surface
and the second surface together build a profiled surface, the
profiled surface being a ribbed surface. Material of the first
surface is deformable such that the material of the first surface
will bend away from a direction of rotation of the airfoil during
operation.
[0015] With "tip region" the part of the airfoil is meant, when
being assembled within the gas turbine, that is directed to an
opposing surface of a component of the gas turbine--e.g. a
compressor casing or a rotor.
[0016] Tip region "opposite to the platform" means opposite in
respect of the pressure and the suction surfaces, thus that the
pressure and suction surfaces both terminate in one direction at
the platform and in the other direction at the tip region.
"Opposite" means, if the platform defines a radial inward end of
the pressure and the suction surfaces once assembled within a
component of a gas turbine engine, the tip region is located at a
radial outward end.
[0017] The first surface expanse is compared to the cross sectional
plane of the airfoil perpendicular to the first direction, i.e.
substantially parallel to the platform or parallel to the spread of
the tip section. Typically a cross sectional plane of a blade
differs in size based on the position where the measurement is
taken. Particularly a smallest possible cross sectional plane, an
average cross sectional plane, a largest cross section, or a cross
section of a plane close to the tip region could be taken for the
to be compared cross sectional plane.
[0018] The first and second distance may be average values, if the
distance of the first surface (or the second surface) in relation
to the platform is different based on at which actual position
within the first surface (or the second surface) the measurement is
taken.
[0019] The inventive airfoil may be advantageous--especially when
also considering the subject matter of the dependent claims--,
because this form of the airfoil provides only a small contact
surface--the first surface--to an opposing surface, once assembled.
This may lead to the possibility to have a deformable tip section.
Thus a reduced operating clearance between the tip region and the
opposing surface may be provided which leads to reduced leakage and
eventually to an improved compressor efficiency.
[0020] According to the invention material of the first surface is
deformable such that the material of the first surface will bend
away from a direction of rotation of the airfoil during
operation.
[0021] Deformable means that during operation of the turbomachine,
an airfoil that is rotating circumferentially about an axis of the
turbomachine--an airfoil of a rotor blade--, the tip region will
deform in the sense that the tip region or more specifically
elevated regions of the tip region, i.e. regions of the first
surface, is bent in circumferential direction opposite to the
direction of rotation. For an airfoil that does not rotate about
the axis of the turbomachine but about a radially oriented rotation
line for varying a pitch of the airfoil--an airfoil of a variable
guide vane--, the tip region--i.e. regions of the first
surface--will bend away from the direction of rotation, i.e.
opposite to the direction of rotation.
[0022] With pitch an angle of attack with a main fluid flow is
meant in respect of the direction of the flow of the main
fluid.
[0023] For both types of airfoils, the first surface of the tip
region is composed of a deformable material such that the material
of the first region will bend away from the direction of
rotation--pitch rotation or circumferential rotation--of the
airfoil during operation.
[0024] Additionally the tip region may also show a radial
deformation, like a slight compression of the tip region in radial
direction.
[0025] The first surface is not required to be continuous and may
be a plurality of separated partial surfaces together building the
overall first surface. The same applies to the second surface.
Actually, in a preferred embodiment the first surface may be a
plurality of elevated areas of the profiled surface and the second
surface being a plurality of recessed areas of the profiled
surface. Additionally the plurality of elevated areas of the
profiled surface and the plurality of recessed areas of the
profiled surface may be arranged alternating. For example the tip
region may have a plurality of alternating ribs and slots, the
surface of the ribs defining the first surface and the surface of
the slots defining the second surface. The transition from one rib
to an adjacent slot may be via a step down from the first distance
to the second distance. All ribs may belong to the first surface
and may be located in one common plane with the first distance. All
slots may belong to the second surface and may be located in one
further common plane with the second distance.
[0026] In yet another embodiment the elevated areas and/or the
recessed areas may be arranged as ribs between the suction surface
and the pressure surface. The ribs may be substantially straight.
Furthermore the elevated areas and/or the recessed areas may be
arranged as ribs substantially parallel to the suction surface or
to the pressure surface or may be straight from the leading edge
and the trailing edge of the airfoil independently of the
orientation of the suction or pressure surface. As a further option
the elevated areas and/or the recessed areas may be arranged as a
grid pattern.
[0027] The grid pattern may have a shape of a first plurality of
ribs intersected with a second plurality of ribs. Alternatively,
the grid pattern may have a shape of an inverse profile of a first
plurality of ribs intersected with a second plurality of ribs. In
both cases the first plurality of ribs may be diagonal in direction
of a trailing edge of the airfoil and the second plurality of ribs
may be diagonal in direction of a leading edge of the airfoil.
Thus, in case of a plurality of ribs intersected with a second
plurality of ribs, the first surface may substantially be formed of
a plurality of exalted diamonds surrounded by recesses forming the
second surface. Thus, in case of an inverse profile of a first
plurality of ribs intersected with a second plurality of ribs, the
second surface may be formed of a plurality of diamond-shaped
cavities, surrounded by walls with a top surface being the first
surface.
[0028] The invention is particularly advantageous, as the tip
region is composed of a deformable material. Therefore the cold
clearance of the airfoil to its opposing mating part can be
smaller. If contact is made with the adjoining part during
operation, the tip region of the airfoil is allowed to deform. Thus
no precise manufacturing or configuration of blades is required.
Tolerances are allowed so that manufacturing or installation costs
can be reduced. Also no expensive abradable or abrasive coating is
required. If no contact of the tip region with the opposing part is
present during operation due to tolerances and operating
conditions, the operating gap and the leakage will be reduced. With
reduced clearance and reduced leakage the overall compressor
efficiency will be improved.
[0029] In yet another embodiment the tip region--particularly the
first surface and possibly only the first surface--may either be
coated with an abrasive coating for abrading an opposing surface or
may be coated with an abradable coating, so that the tip region
gets abraded. Such a coating may be used as an alternative to the
deformable tip or may even be combined with the deformable tip.
This is possible by using only less abradable or abrasive coatings
on the tip region, because only the raised first surface may be
coated. This again may reduce costs.
[0030] The invention is also directed to a blade and/or a guide
vane, particularly a variable guide vane of a gas turbine,
comprising an airfoil as explained. The blade and/or vane may be
located in a compressor or a turbine section of the gas
turbine.
[0031] Besides, the invention is also directed to a gas turbine
comprising at least one of a compressor section or a turbine
section, the compressor section or the turbine section comprising
at least one airfoil configured as explained.
[0032] Beside a gas turbine, the invention is directed to all kinds
of turbomachinery equipment, like turbines, compressors, pumps that
comprise rotating parts.
[0033] It has to be noted that embodiments of the invention have
been described with reference to different subject matters. In
particular, some embodiments have been described with reference to
apparatus type claims whereas other embodiments have been described
with reference to method type claims. However, a person skilled in
the art will gather from the above and the following description
that, unless other notified, in addition to any combination of
features belonging to one type of subject matter also any
combination between features relating to different subject matters,
in particular between features of the apparatus type claims and
features of the method type claims is considered as to be disclosed
with this application.
[0034] The aspects defined above and further aspects of the present
invention are apparent from the examples of embodiment to be
described hereinafter and are explained with reference to the
examples of embodiment.
BRIEF DESCRIPTION OF THE DRAWINGS
[0035] Embodiments of the invention will now be described, by way
of example only, with reference to the accompanying drawings, of
which:
[0036] FIG. 1 shows a gas turbine in a longitudinal partial
section;
[0037] FIG. 2 shows a cross-sectional view of a section of a flow
duct of a compressor;
[0038] FIG. 3 is a perspective view of a prior art turbine
blade;
[0039] FIG. 4 shows several profiles of the tip region of a blade,
according to the invention in a perspective view.
[0040] The illustration in the drawing is schematical. It is noted
that for similar or identical elements in different figures, the
same reference signs will be used.
[0041] Some of the features and especially the advantages will be
explained for an assembled gas turbine, but obviously the features
can be applied also to the single components of the gas turbine but
may show the advantages only once assembled and during operation.
But when explained by means of a gas turbine during operation none
of the details should be limited to a gas turbine while in
operation.
[0042] In the following most of the features are explained for an
axial compressor of a stationary gas turbine. Most of the features
can also be applied to blades and vanes of a compressor of such a
gas turbine but will also apply to a turbine section.
DETAILED DESCRIPTION OF THE INVENTION
[0043] FIG. 1 shows a gas turbine 1 in a longitudinal partial
section. In the interior, it has a rotor 3 which is rotatably
mounted about a rotation axis 2 and is also referred to as turbine
rotor or rotor shaft. Following one another along the rotor 3 are
an intake casing 4, a compressor 5, torus-like annular combustion
chambers 6 (only one can be seen in the figure) each having a
burner 7, a turbine unit 8 and an exhaust-gas casing 9. As an
alternative to the annular combustion chambers 6, also separate can
combustion chambers may be present, each with one or more burners,
of which there may be a plurality of them arranged around the axis
of the gas turbine engine.
[0044] Provided in the compressor 5 is an annular compressor duct
10 as a flow path for air which narrows in cross section in the
direction of the annular combustion chamber 6, i.e. downstream of a
to be compressed fluid. Arranged at the downstream end of the
compressor 5 is a diffuser 11, which is fluidically connected to
the annular combustion chambers 6. Each of the annular combustion
chambers 6 form a combustion space 12 for a mixture of liquid
and/or gaseous fuel and compressed air from the compressor 5. A
hot-gas duct 13 arranged in the turbine unit 8 is fluidically
connected to the combustion spaces 12, the exhaust-gas casing 9
being arranged downstream of the hot-gas duct 13.
[0045] Sections with blades and vanes are arranged in the
compressor duct 10 and in the hot-gas duct 13. In each case a
turnable blade section 17 of blades 16 alternately follows a vane
section 15 of non-rotatable guide vanes 14. These fixed guide vanes
14 are in this case connected to one or more guide vane carriers
18, whereas the moving blades 16 are fastened to the rotor 3 by
means of a disc 19.
[0046] The turbine unit 8 has a conically widening hot-gas duct 13,
the outer guide surface 21 of which widens concentrically in the
direction of flow of the working fluid 20. The inner guide surface
22, on the other hand, is oriented essentially parallel to the
rotation axis 2 of the rotor 3. At their free ends, the moving
blades 16 each have an edge as a tip region 29 (see FIG.
2)--possibly grazing edges--, which form a radial gap 23 with the
outer guide surfaces 21 opposite them.
[0047] The same applies to the compressor 5, in which edges are
present at a tip region 29 of a blade 16 and a radial gap 23 is
present with the outer opposing guide surfaces 21.
[0048] During operation of the gas turbine 1, air L is drawn in
from the compressor 5 through the intake casing 4 and is compressed
in the compressor duct 10. The air L provided at the burner-side
end of the compressor 5 is directed through the diffuser 11 to the
burners 7 and is mixed there with a fuel. The mixture is then
burned, with a working fluid 20 being formed in the combustion
space 12. The working fluid 20 flows from there into the hot-gas
duct 13. At the moving blades 16 arranged in the turbine unit 8,
the working fluid expands in an impulse-transmitting manner, so
that the rotor 3 is driven.
[0049] An inlet-side compressor bearing 32 may serve, in addition
to the axial and radial mounting, as an adjusting device for a
displacement of the rotor 3. The rotor 3, in the steady state, may
be displaced, to the left in FIG. 1, from an initial position into
a steady operating position against the direction of flow of the
working fluid 20. As a result, the radial gap 23 formed in the
turbine unit 8 by moving blades 16 and the outer guide surface 21
is reduced. This leads to a reduction in the flow losses in the
turbine unit 8 and therefore to an increase in the efficiency of
the gas turbine 1.
[0050] Besides, the radial gap 23 may be different for each of the
blades 16 due to tolerances.
[0051] A section of the annular compressor duct 10 of the
compressor 5 with two sections 17 of blades 16 and with a
guide-vane section 15 arranged in between is shown in FIG. 2. The
annular compressor duct 10 is in this case designed as a flow duct
for air L as the flow medium.
[0052] In FIG. 2, the guide vane 14 is fastened to an external
wall--e.g. the compressor casing--, whereas the blades 16 are
connected via a disc to the rotor 3. At its fixed end, each blade
16 has a respective platform 25, the surfaces of which define the
compressor duct 10 on the radial inside. Likewise, each guide vane
14, at its fixed end, has a platform 25, which defines the
compressor duct 10 on the radial outside. Extending from the
platform 25 of the blade 16 (or of the guide vane 14) into the
compressor duct 10 is a rotatable airfoil 27 (or respectively a
fixed guide airfoil 28) which compress the air L during operation
of the compressor 5. The free ends of the airfoils 27, 28 opposite
to the platform-side ends form a tip region 29 of the airfoils 27,
28 and are opposite respective guide rings 30, with the radial gap
23 being formed.
[0053] According to FIG. 2, as viewed in the axial direction, the
radial gap 23 is in each case oriented substantially parallel to
the rotation axis 2 in one section. On the other hand, the
platforms 25 arranged in the section may be each inclined relative
to the rotation axis 2 of the rotor 3, so that the flow duct 24
narrows as viewed in the axial direction. A cylindrical contour of
the flow duct 10 is obtained in the regions of the radially
opposite fixed and rotating components, which as viewed in the
axial direction lie in sections and in the radial direction lie
inside and respectively outside the guide profiles and moving
profiles, respectively. In the axial direction, therefore, both the
outer guide surface 21 and the inner guide surface 22 alternately
run cylindrically and in such a way as to be inclined relative to
the rotation axis 2 of the rotor 3, the cylindrical guide surface
21, 22 in each case being opposite an inclined guide surface 21, 22
as viewed in the radial direction of the rotor 3. This is one
possible option of orientation, but alternative orientations are
possible.
[0054] Referring to FIG. 3, an exemplary turbine blade 16 for a gas
turbine engine is illustrated. The blade 16 comprises an airfoil
27. The airfoil 27 has an outer wall 41 comprising a pressure
sidewall or pressure surface 42 and a suction sidewall or suction
surface 43. The pressure and suction surfaces 42, 43 are joined
together along an upstream leading edge 44 and a downstream
trailing edge 45, where the leading and trailing edges 44, 45 are
spaced axially or chordally from each other with respect to a
chordal direction C. The airfoil 27 extends radially along a
longitudinal direction of the blade 16--a radial direction of the
compressor 4--, from a radially inner airfoil platform 25 to a
radially outer tip region 29.
[0055] The tip region 29 includes a blade tip surface 46 having an
airfoil shape, and pressure tip side 51 and suction tip side 50
which are joined together at spaced apart leading tip edge 55 and
trailing tip edge 56 of the tip region 29. The pressure and suction
tip sides 51, 50 form the radial ends of the pressure and suction
surfaces 42, 43, respectively, of the airfoil 27.
[0056] Optionally cooling holes 52 may be provided along the
trailing edge of the airfoil 27, as indicated in the FIG. 3.
[0057] According to FIG. 3, no specific profile for the tip region
29 is shown. Simply a flat surface is depicted.
[0058] For easier reference a first direction D is defined in FIG.
3. This direction D should also apply to the following figures and
defines a direction substantially perpendicular to the platform 25
and progressing with a distance to the platform 25. To be more
precise, the first direction is perpendicular to inner guide
surface 22 of the platform 25.
[0059] In FIG. 4, several profiles of the tip region 29 are
depicted schematically. The actual airfoil form of the tip region
29 as shown in FIG. 3 is disregarded in FIG. 4 to keep the focus on
the profile itself.
[0060] In FIG. 4 an airfoil 27 of the compressor blade 16 of an
axial flow gas turbine is depicted. The airfoil 27 comprises the
elements as presented according to FIG. 3, e.g. the platform 25
(not shown in FIG. 4), the pressure surface 42 and the suction
surface 43 (the latter is hidden in FIG. 4), and a tip region 29.
As before, the pressure surface 42 and the suction surface 43
extend substantially perpendicular to the platform 25 along a first
direction D. D is not shown, but is defined as a direction
substantially perpendicular to the platform 25 and progressing with
a distance to the platform 25 as it can be seen in FIG. 3. A first
distance D1 and a second distance D2, which will be discussed
later, are marked along the direction D.
[0061] The tip region 29 is located between the pressure and the
suction surface 42, 43 opposite to the platform 25 at a distant end
of the airfoil 27 in direction along the first direction D.
[0062] The tip region 29 comprises a first surface 70 and a second
surface 71. The first surface 70 has a first distance D1 to the
platform 25 along the first direction D, and the second surface 71
has a second distance D2 to the platform 25 along the first
direction D. The first distance D1 may be taken from a most distant
or from an intermediate location of the first surface 70, if the
first surface 70 has different distances to the platform 25
depending upon the location on the first surface 70. The second
distance D2 may be taken from a least distant or from an
intermediate location of the second surface 71, if the second
surface 71 has different distances to the platform 25 depending
upon the location on the second surface 71.
[0063] According to the invention the first surface 70 and the
second surface 71 together build a profiled surface, particularly a
ribbed surface. The first surface 70 is comprised of a plurality of
elevated partial surfaces being the elevated part of the ribbed
surface, i.e. the ribs themselves. The second surface 71 is
comprised of a plurality of recessed partial surfaces forming the
recesses of the ribbed surface, i.e. grooves between the ribs.
[0064] The first surface 70 has a first surface expanse which is
the overall surface expanse of all elevated partial surfaces. The
second surface 71 has a second surface expanse which is the overall
surface expanse of all recessed partial surfaces.
[0065] According to the invention, the first surface expanse 70 is
less than a cross sectional plane of the airfoil 27 perpendicular
to the first direction D, i.e. parallel to the platform 25. This
means that the first surface expanse 70 has a lesser expanse than
compared to an airfoil with a flat surface at the tip region 29, as
it can be seen for example in FIG. 3.
[0066] With cross sectional plane of the airfoil 27 any possible
plane is meant that could be defined parallel to the platform 25,
e.g. a plane close to the tip region 29, a plane close to the
platform 25, or any plane in between, especially a plane with the
smallest possible cross section.
[0067] According to the invention, the first distance D1 is greater
than the second distance D2.
[0068] With this configuration and once the blades are assembled
and during operation of the gas turbine, the total surface of the
tip region 29 that may be in contact with an opposing surface of
the compressor casing, may be only the first surface 70, which has
a first surface expanse less than a flat tip region as it is shown
in FIG. 3. This is particularly advantageous if the tip region
29--especially the material of the first surfaces 70--is composed
of a deformable material. In this case the ribs can adjust to the
opposing contour by deformation during operation if contact is made
to the opposing material.
[0069] FIG. 4A specifies a first embodiment, in which the ribs
extend in parallel between the leading tip edge 55 and the trailing
tip edge 56. The ribs, defined by alternating first surfaces 70 and
second surfaces 71, may be completely straight, or may follow the
curved form of the airfoil along the pressure surface 42 and/or the
suction surface 43. Alternatively the ribs may be arranged
perpendicular to the direction of rotation of the airfoil 27 or may
be arranged at a specific angle in respect of the direction of
rotation.
[0070] In a second embodiment according to FIG. 4B, again parallel
ribs of alternating first surfaces 70 and second surfaces 71 are
shown. The direction of the ribs is substantially perpendicular to
at least one area of the suction surface 43 and/or pressure surface
42. In FIG. 4B each partial surface of the first surface 70 is
substantially rectangular. Alternatively first surfaces 70 may be
in form of a trapezium if the grooves between the ribs are not
perfectly parallel.
[0071] Not shown in the figures, the embodiments of FIG. 4A and
FIG. 4B may be combined to gain a grid pattern with a set of
parallel grooves extending between the leading tip edge 55 and the
trailing tip edge 56 and a set of parallel grooves extending
between the suction surface 43 and the pressure surface 42.
[0072] Further variants of grid patterns are shown in FIGS. 4C and
4D. In both variants a first plurality of ribs or grooves are
arranged diagonally in direction of the trailing tip edge 56 of the
airfoil 27 and the second plurality of ribs or grooves are arranged
diagonally in direction of the leading tip edge 55 of the airfoil
27, the ribs and/or grooves intersecting each other. According to
the figures, the intersection occurs with a right angle, but
different angles are possible.
[0073] According to FIG. 4C, the grid pattern has a shape of a
first plurality of ribs intersected with a second plurality of
ribs. Thus the first surface 70 results in a plurality of rhombus
like surfaces. The second surface 71 results in a grid of grooves
between the rhombus like surfaces of the first surface 70.
[0074] According to FIG. 4D, the grid pattern has a shape of an
inverse profile of the previously discussed pattern of a first
plurality of ribs intersected with a second plurality of ribs. This
results in a plurality rhombus like cavities as the second surface
71. The first surface 70 surrounds these cavities like meshed walls
that get intersected.
[0075] The given profiles may be advantageous because only a
fraction of the tip region 29 may be in contact with an opposing
surface like the outer guide surface 21. Particularly, only the
first surface 70 may be in contact.
[0076] Thus, all of these embodiments have the advantage that for
the elevated regions underneath the first surface 70 a flexible
material can be used that is deformable under pressure. So the
operating clearance between the tip region 29 and an opposing
surface may be reduced.
[0077] Besides, it is particularly advantageous if--additionally or
alternatively to the deformable material--the tip region 29 is
designed to abrade. In this case, and according to the invention,
it is sufficient to only coat the first surface 70 with an
abradable or abrasive coating, because this is the only surface
which is in contact with the opposing surface (which itself may be
coated with the "opposite" coating, i.e. with abrasive coating if
the first surface 70 comprises an abradable coating and with
abradable coating if the first surface 70 comprises an abrasive
coating). By this the amount of coating material may be reduced.
Furthermore the force on the blades to abrade is reduced.
[0078] According to the embodiments, the radial gap 23--which may
be affected by manufacturing or assembly tolerances--during
operation may be reduced so that also leakage of air will be
reduced. This again improves the efficiency of the compressor.
[0079] As a further positive side effect, the invention reduces the
need to grind the fully assembled compressors, particularly the
need to grind the rotor blades to size, because the tip region of
the blade will adapt to the opposing surface simply during
operation of the gas turbine.
[0080] Even though the embodiments show a compressor section as an
example, the same principles also apply to airfoils in the turbine
section of a gas turbine engine. Besides, the principles can be
applied to blades but also to guide vanes.
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