U.S. patent application number 12/447972 was filed with the patent office on 2010-03-18 for turbine blade.
Invention is credited to Fathi Ahmad.
Application Number | 20100068069 12/447972 |
Document ID | / |
Family ID | 37945313 |
Filed Date | 2010-03-18 |
United States Patent
Application |
20100068069 |
Kind Code |
A1 |
Ahmad; Fathi |
March 18, 2010 |
Turbine Blade
Abstract
A turbine blade, having a plurality of auxiliary cooling
channels which branch off from a main cooling channel, formed
within a blade body, is provided. The plurality of auxiliary
cooling channels open into outlet openings in the leading edge
region of the blade body. A heat shield element is attached to the
blade body in the leading edge region at a predefined spacing,
wherein the heat shield element has a number of outlet channels
which are arranged behind one another in the longitudinal direction
and extend from the main cooling channel to the outer wall face of
the heat shield element.
Inventors: |
Ahmad; Fathi; (Kaarst,
DE) |
Correspondence
Address: |
SIEMENS CORPORATION;INTELLECTUAL PROPERTY DEPARTMENT
170 WOOD AVENUE SOUTH
ISELIN
NJ
08830
US
|
Family ID: |
37945313 |
Appl. No.: |
12/447972 |
Filed: |
September 20, 2007 |
PCT Filed: |
September 20, 2007 |
PCT NO: |
PCT/EP07/59989 |
371 Date: |
April 30, 2009 |
Current U.S.
Class: |
416/97R |
Current CPC
Class: |
F01D 5/186 20130101 |
Class at
Publication: |
416/97.R |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Foreign Application Data
Date |
Code |
Application Number |
Oct 30, 2006 |
EP |
06022622.2 |
Claims
1.-7. (canceled)
8. A turbine blade, comprising: a blade body, provided with a first
outside surface which is an outside surface of the blade body, in a
leading edge region of the blade body; a heat shield element, which
is arranged on the first outside surface at a distance from the
blade body; and a plurality of secondary cooling passages, which
branch off from a main cooling passage formed inside the blade
body, wherein the plurality of secondary cooling passages open into
a plurality of discharge openings in the leading edge region, and
wherein the heat shield element has a plurality of discharge
passages which extend from the main cooling passage to a second
outside surface, the second outside surface is an outside surface
of the heat shield element.
9. The turbine blade as claimed in claim 8, wherein the turbine
blade is used in a gas turbine.
10. The turbine blade as claimed in claim 8, wherein the heat
shield element has a shape that is adapted to a profile of the
turbine blade in the leading edge region.
11. The turbine blade as claimed in claim 10, wherein the heat
shield element is produced from a first material which is more
resistant to temperature than a second material of the blade
body.
12. The turbine blade as claimed in claim 8, wherein the heat
shield element is arranged at the distance of no more than 3 mm
from the blade body.
13. The turbine blade as claimed in claim 8 wherein the heat shield
element is curved whereby a constant distance between an inside
surface of the heat shield element and the first outside surface is
maintained.
14. The turbine blade as claimed in claim 8, wherein the plurality
of secondary cooling passages are arranged essentially
perpendicularly to an inside wall surface of the heat shield
element.
15. The turbine blade as claimed in claim 14, wherein the plurality
of secondary cooling passages, are uniformly distributed behind the
heat shield element.
16. The turbine blade as claimed in claim 8, wherein a plurality of
edge regions of the heat shield element are connected to the blade
body.
17. The turbine blade as claimed in claim 8, wherein the plurality
of discharge passages are arranged one behind the other.
18. A thermal turbomachine, comprising: a plurality of turbine
blades, each turbine blade comprising: a blade body, provided with
a first outside surface which is an outside surface of the blade
body, in a leading edge region of the blade body, a heat shield
element, which is arranged in the first outside surface at a
distance from the blade body, and a plurality of secondary cooling
passages, which branch off from a main cooling passage formed
inside the blade body; wherein the plurality of secondary cooling
passages open into a plurality of discharge openings in the leading
edge region, and wherein the heat shield element has a plurality of
discharge passages which extend from the main cooling passage to a
second outside surface, an outside surface of the heat shield
element.
19. A thermal turbomachine as claimed in claim 18, wherein the
turbomachine is a gas turbine.
20. The thermal turbomachine as claimed in claim 18, wherein the
heat shield element has a shape that is adapted to a profile of the
turbine blade in the leading edge region.
21. The thermal turbomachine as claimed in claim 20, wherein the
heat shield element is produced from a first material which is more
resistant to temperature than a second material of the blade
body.
22. The thermal turbomachine as claimed in claim 18, wherein the
heat shield element is arranged at the distance of no more than 3
mm from the blade body.
23. The turbine blade as claimed in claim 18 wherein the heat
shield element is curved whereby a constant distance between an
inside surface of the heat shield element and the first outside
surface is maintained.
24. The thermal turbomachine as claimed in claim 18, wherein the
plurality of secondary cooling passages are arranged essentially
perpendicularly to an inside wall surface of the heat shield
element.
25. The thermal turbomachine as claimed in claim 24, wherein the
plurality of secondary cooling passages, are uniformly distributed
behind the heat shield element.
26. The thermal turbomachine as claimed in claim 18, wherein a
plurality of edge regions of the heat shield element are connected
to the blade body.
27. The thermal turbomachine as claimed in claim 18, wherein the
plurality of discharge passages are arranged one behind the other.
Description
CROSS REFERENCE TO RELATED APPLICATIONS
[0001] This application is the US National Stage of International
Application No. PCT/EP2007/059989, filed Sep. 20, 2007 and claims
the benefit thereof. The International Application claims the
benefits of European Patent Office application No. 06022622.2 EP
filed Oct. 30, 2006, both of the applications are incorporated by
reference herein in their entirety.
FIELD OF INVENTION
[0002] The invention refers to a turbine blade according to the
claims.
BACKGROUND OF INVENTION
[0003] Turbomachines, especially gas turbines, are used in many
areas for driving generators or driven machines. A gas turbine
customarily has a rotatably mounted rotor which is enclosed by a
fixed casing. The fixed sub-assemblies of a gas turbine are also
collectively referred to as a stator. The energy content of a fuel
in this case is used for producing a rotational movement of the
rotor components. For this purpose, the fuel is combusted in a
combustion chamber, wherein compressed air is supplied by an air
compressor. The operating medium which is produced in the
combustion chamber as a result of the combustion of the fuel, being
under high pressure and at high temperature, is guided in the
process through a turbine unit which is connected downstream to the
combustion chamber, where it is expanded, performing work.
[0004] For producing the rotational movement of the rotor
components, in this case a number of rotor blades, which are
customarily assembled in blade groups or blade rows, are arranged
on these rotor components and drive the rotor components by means
of impulse transmission of the flow medium. For guiding the flow
medium in the turbine unit, stator blade rows, which are connected
to the turbine casing, are moreover customarily arranged between
adjacent rotor blade rows. The turbine blades, especially the
stator blades, in this case customarily have a blade airfoil, which
is extended along a blade axis, for suitable guiding of the
operating medium and upon which a platform, which extends
transversely to the blade axis, can be formed on the end face for
fastening the turbine blade on the respective carrier body.
[0005] With the design of such gas turbines, a particularly high
efficiency is customarily a design aim in addition to the
achievable output. An increase of the efficiency can be achieved in
this case for thermodynamic reasons basically by means of an
increase of the exhaust temperature at which the operating medium
flows from the combustion chamber and into the turbine unit. The
temperatures which are achieved during operation of such a gas
turbine lie at up to 1300.degree. C.
[0006] The components and component parts of the gas turbine which
are exposed to these high temperatures of the operating medium are
therefore subjected to a high thermal stress. In order to therefore
ensure with high reliability a comparatively long service life of
the affected components, the affected components, especially the
rotor blades and/or stator blades of the turbine unit, are cooled.
The turbine blades in this case are customarily provided with
cooling passages, wherein an effective and reliable cooling of the
leading edge of the respective turbine blade, which is thermally
stressed to a particularly high degree, is especially to be
ensured.
SUMMARY OF INVENTION
[0007] As cooling medium, cooling air is customarily used in this
case. This cooling air can be fed to the respective turbine blade
via a number of cooling medium passages which are integrated into
the blade airfoil or the blade profile. From these cooling medium
passages the cooling air flows into discharge passages, which
branch off from these, of the respectively provided regions of the
turbine blades, as a result of which a convective cooling of the
blade interior and of the blade wall is achieved. On the discharge
side, these passages are left open so that after flowing through
the turbine blade the cooling air flows from discharge openings,
which are also referred to as film cooling holes, and form a
cooling film on the surface of the blade airfoil. By means of this
cooling air film, the blade basic body is largely protected on the
surface against a direct and intensive contact with the hot
operating medium which flows past at high velocity.
[0008] In order to enable an especially uniform and effective film
cooling in the leading edge region of the blade airfoil, the
discharge openings in this region are customarily arranged
uniformly along at least two rows which are oriented parallel to
the leading edge. The discharge passages, moreover, as a rule are
oriented at an angle to the longitudinal direction of the turbine
blade, which assists the forming of the protective cooling air film
which flows along the surface.
[0009] Since the leading edge region of the turbine blade is
particularly exposed to a severe thermal stress, the leading edge
of the blade can moreover be provided with a heat shielding
coating. This heat shielding coating expediently consists of a
material which is more resistant to temperature than that of the
blade basic body. Moreover, the heat shielding coating is
characterized by a low coefficient of thermal conductivity, as a
result of which the temperature stress of the base material of the
blade body is reduced. Therefore, the service life of the turbine
blade is increased as a result of such a heat shielding coating in
conjunction with cooling of the leading edge region of the
blade.
[0010] This heat shield, however, has the disadvantage that after a
certain time cracks occur in the heat shielding coating. These
cracks reduce the protection of the blade basic body against the
hot exhaust gas of the gas turbine so that as a consequence of the
increased thermal stress crack development can also occur in the
basic body of the turbine blade. Such cracks in the blade basic
body endanger the operational safety and can lead to the breakdown
of the gas turbine.
[0011] Moreover, a modular turbine blade of the type referred to in
the introduction is known from GB 841 117. The turbine blade
comprises a cast basic body with a blade airfoil upon which a
plurality of cooling air blow-out slots are provided on the leading
edge side, which are covered at a distance by a guard plate which
is fastened on the blade profile on the side. The cooling air which
issues from the slots cools the guarded leading edge in the manner
of an impingement cooling, and subsequent to the impingement
cooling is deflected by the plate in such a way that it can leave
the modular turbine blade in the region of the pressure-side
surface and suction-side surface.
[0012] The invention is therefore based on the object of disclosing
a turbine blade of the aforementioned type which with simple means
ensures an especially high operational safety of the gas turbine,
even when used in high flow temperatures.
[0013] This object is achieved according to the invention by means
of a turbine blade according to the features of the claims.
[0014] The invention in this case is based on the consideration
that particularly with regard to the operational safety and the
economical efficiency of a gas turbine the turbine blades should
have a service life which is as long as possible as a result of a
suitably selected heat shield. At the same time, the fact that
particularly the leading edge of the turbine blade is thermally
severely stressed should especially be taken into consideration.
This leading edge should therefore especially be protected.
[0015] This is consequently achieved by the heat shield element
being attached at a distance on the blade basic body in the leading
edge region, as a result of which a direct contact of the heat
shield element with the blade basic body is avoided. For cooling
the blade basic body, in this case its outside surface in the
leading edge region is provided with a number of secondary cooling
passages, wherein these extend from the main cooling passage to the
outside surface of the blade basic body. These secondary cooling
passages are arranged in a uniformly distributed manner behind the
heat shield element for effective cooling in the leading edge
region of the blade basic body. Therefore, stresses, and cracks
which result from them, can be avoided.
[0016] For cooling the heat shield element, this has a number of
discharge passages which extend from its outside surface in the
direction of the blade basic body. This passage, which is formed
for guiding a cooling flow, additionally also serves as a
connecting element between the heat shield element and the blade
basic body. The discharge passage in this case projects with one
end into a main cooling passage which is formed inside the blade,
wherein the medium which flows in the main cooling passage can flow
for cooling the heat shield element on its outside surface.
[0017] By means of a heat shield system which is formed in such a
way the especially critical region, that is to say the leading edge
region of the turbine blade, is especially effectively protected
against the high temperatures of the operating medium of the
turbine. As a result of cooling the heat shield element and the
blade basic body it is possible to increase the temperature of the
operating medium of the turbine, which flows around the turbine
blade, above the temperature which is possible for the material of
the turbine blade. The cooling is carried out in such a way by a
cooling flow from the main cooling passage being directed in part
through the discharge passages of the heat shield element onto its
outside surface, and in part by a cooling flow from the main
cooling passage flowing via the secondary cooling passages of the
blade basic body through the gap which is formed by the heat shield
element and the blade basic body. By means of cooling medium which
is guided in such a way, a protective film is formed on the outside
surface of the heat shield element. This cooling film prevents a
direct contact of the hot operating medium of the turbine with the
heat shield element, as a result of which the temperature stress of
the outside surface which is exposed to inflow is reduced. The
increase of the temperature of the heat shield element which
therefore occurs does not directly have an effect on the
temperature of the blade basic body in the leading edge region,
however, since the heat shield element is arranged at a distance
from the blade basic body. The heat transfer between the heat
shield element and the blade basic body is significantly reduced,
moreover, by means of the cooling medium which flows between the
inside surface of the heat shield element and the outside surface
of the blade basic body by the heat in the leading edge region
being carried away by means of the internal cooling flow.
[0018] The heat shield element especially preferably has a shape
which is adapted to the profile of the blade basic body in the
leading edge region. Consequently, the effect is achieved of the
turbine blade also having a flow-optimized shape in the leading
edge region after attaching the heat shield element. Moreover, the
shape of the heat shield element which corresponds to the blade
basic body leads to a uniform extent of the gap in the leading edge
region. Consequently, the cooling medium flows with predominantly
constant velocity along the outside surface of the blade basic body
and along the inside surface of the heat shield element, as a
result of which the cooling in the leading edge region of the
turbine blade is carried out especially uniformly. Therefore, no
excessively high stresses, which could lead to crack developments,
occur, particularly in the blade basic body.
[0019] In a further expedient development, the heat shield element
is produced from a material which is more resistant to temperature
in comparison to the blade basic body. Since the heat shield
element is directly exposed to inflow of hot operating medium
during operation of the turbine, this component is particularly
exposed to a high temperature stress. Therefore, the heat shield
element should be produced from an especially temperature-resistant
material in order to particularly ensure the operational safety and
to minimize the downtimes of the turbine.
[0020] In addition to the use of temperature-resistant materials,
for increasing the resistance of the heat shield element this
should be cooled. If in this case the heat shield element is
designed for impingement cooling for an especially effective
cooling, this is achieved by the distance of the heat shield
element from the blade basic body being sufficiently minimized.
[0021] For this purpose, the heat shield element is preferably
arranged at a distance of 1 mm to 3 mm from the blade basic
body.
[0022] A heat shield element which is attached up to this distance
in the leading edge region of the turbine blade particularly
ensures a sufficiently high impingement velocity of the cooling
medium upon the inside surface of the heat shield element, as a
result of which an especially effective cooling by means of
impingement cooling is achieved. Since the static pressure in the
main cooling passage of the blade basic body is predetermined, the
impingement velocity of the cooling flow, in addition, for example,
to the diameter of the secondary cooling passages, is determined in
particular by the distance of the heat shield element from the
blade basic body. A sufficiently high velocity of the cooling
medium directly before impinging on the inside surface of the heat
shield element is necessary since an intimate contact between the
cooling medium and the inside surface of the heat shield element
takes place in this way. By means of such impingement cooling a
significantly more effective carrying away of heat is possible than
is possible, for example, in the case of a film cooling.
[0023] In an especially advantageous development, the secondary
cooling passages are arranged in a manner in which they are
oriented essentially perpendicularly to the inside surface of the
heat shield element. Therefore, the cooling flow from the main
cooling passage impinges perpendicularly upon the inside surface of
the heat shield element, as a result of which a large part of the
kinetic energy of the cooling medium is used for an especially
intimate contact between the particles in the cooling flow and the
inside surface of the heat shield element. As a result, the heat of
the heat shield element in an especially effective manner is
transmitted to the internal cooling flow and carried away.
[0024] In a further variant, the heat shield element is connected
to the blade basic body in the edge regions of the turbine blade.
For foaming the especially effective impingement cooling, the blade
basic body, preferably in its leading edge region, is provided with
a recess. The advantage of this alternative embodiment lies inter
alia in the fact that the original flow-optimized shape of the
turbine blade is maintained.
[0025] The described heat shield element can advantageously be used
at the places of the turbomachine where component parts and
sub-assemblies of the thermal turbomachine are impinged upon with
the hot operating medium. However, the use of the heat shield
element is especially preferable for the protection of the leading
edge region of the turbine blade since the temperature stress of
the blade basic body is especially high in this region. Inter alia,
the downtimes of the gas turbine are minimized by means of such a
heat shield element since the service life is increased because of
the heat shield element.
[0026] As a result of the extremely high thermal stress of the heat
shield element, cracks can occur in the heat shield element, even
when using an especially temperature-resistant material, especially
after a specified operating period of the turbine. In this case,
for example within the scope of the maintenance operations of the
gas turbine, the heat shield elements can be removed in a
relatively simple manner and replaced by new ones. Therefore, the
affected turbine blades do not have to be completely exchanged as
previously in the case of crack development in the leading edge
region of the turbine blade.
[0027] The advantages which are achieved with the invention are
especially that by means of the heat shield element which is
connected upstream to the blade basic body of the turbine blade an
efficient protection of the leading edge region of the turbine
blade against the high temperatures of the operating medium of the
turbines is provided. In particular, such a heat shield system
enables the use of impingement cooling, as a result of which the
heat shield element can be cooled in an especially effective
manner. Furthermore, by means of the heat shield element the
possible occurrence of cracks, which extend from the outside
surface of the heat shield element and spread into the blade basic
body, is prevented. Moreover, the heat shield elements according to
the invention can subsequently be attached on the turbine blades in
a simple manner and with relatively little cost.
BRIEF DESCRIPTION OF THE DRAWINGS
[0028] An exemplary embodiment of the invention is explained in
more detail with reference to a drawing. In the drawing:
[0029] FIG. 1 shows a half-section through a gas turbine,
[0030] FIG. 2 shows a turbine blade, which is provided with a heat
shield, in longitudinal section,
[0031] FIG. 3 shows a heat shield element which is sectioned in the
longitudinal direction,
[0032] FIG. 4 shows a heat shield element in cross section,
[0033] FIG. 5 shows a cross section through a turbine blade which
is provided with a heat shield element,
[0034] FIG. 6 shows in alternative embodiment a turbine blade with
a heat shield element which is integrated into the leading edge
region of the blade basic body.
[0035] Like parts are provided with the same designations in all
the figures.
DETAILED DESCRIPTION OF INVENTION
[0036] The gas turbine 1 according to FIG. 1 has a compressor 2 for
combustion air, a combustion chamber 4, and also a turbine unit 6
for driving the compressor 2 and for driving a generator, which is
not shown, or a driven machine. For this purpose, the turbine unit
6 and the compressor 2 are arranged on a common turbine shaft 8,
which is also referred to as a turbine rotor, to which the
generator or the driven machine is also connected, and which is
rotatably mounted around its center axis 9. The combustion chamber
4, which is designed in the style of an annular combustion chamber,
is equipped with a number of burners 10 for combustion of a liquid
or gaseous fuel.
[0037] The turbine unit 6 has a number of rotatable rotor blades 12
which are connected to the turbine shaft 8. The rotor blades 12 are
arranged on the turbine shaft 8 in the manner of a ring and so form
a number of rotor blade rows. Furthermore, the turbine unit 6
comprises a number of fixed stator blades 14 which are fastened
also in the manner of a ring on an inner casing 16 of the turbine
unit 6, forming stator blade rows. The rotor blades 12 in this case
serve for driving the turbine shaft 8 by means of impulse
transmission by the operating medium M which flows through the
turbine unit 6. The stator blades 14, on the other hand, serve for
flow-guiding of the operating medium M between two rotor blade rows
or rotor blade rings which follow each other in each case as seen
in the flow direction of the operating medium M. A pair consisting
of a ring of stator blades 14 or a stator blade row, and a ring of
rotor blades 12 or a rotor blade row, which follow each other, in
this case is also referred to as a turbine stage.
[0038] Each stator blade 14 has a platform 18 which is arranged as
a wall element for fixing the respective stator blade on the inner
casing 16 of the turbine unit 6. The platform 18, as also the
turbine blade 12, 14, in this case is a comparatively thermally
severely stressed component part. Each rotor blade 12 is fastened
in a similar manner on the turbine shaft 8 via a platform 19 which
is also referred to as a blade root.
[0039] Between the platforms 18, which are arranged at a distance
from each other, of the stator blades 14 of two adjacent stator
blade rows, a guide ring 21 is arranged in each case on the inner
casing 16 of the turbine unit 6. The outer surface of each guide
ring 21 in this case is also exposed to the hot operating medium M
which flows through the turbine unit 6, and in the radial direction
is at a distance from the outer end of the rotor blades 12, which
lie opposite it, by means of a gap. The guide rings 21 which are
arranged between adjacent stator blade rows in this case serve
especially as shroud elements which protect the inner casing 16, or
other installed parts on the casing, against overstressing as a
result of the operating medium M which flows through the turbine
6.
[0040] The combustion chamber 4 in the exemplary embodiment is
designed as a so-called annular combustion chamber, in which a
multiplicity of burners 10, which are arranged around the turbine
shaft 8 in the circumferential direction, open into a common
combustion space. For this purpose, the combustion chamber 4 is
designed in its entirety as an annular structure which is
positioned around the turbine shaft 8.
[0041] For achieving a comparatively high efficiency, the
combustion chamber 4 is designed for a comparatively high
temperature of the operating medium M of about 1000.degree. C. to
1600.degree. C. In order to also enable a comparatively long
service life in the case of these operating parameters which are
unfavorable for the materials, the rotor blades 12, as shown in
FIG. 2, have a heat shield element 22 which is attached in the
leading edge region. Each of the heat shield elements 22 which are
attached to the rotor blades 12 is equipped on the operating medium
side with an especially heat-resistant protective coating, such as
ceramic, or is produced from a high temperature-resistant
material.
[0042] As shown in FIG. 2, the turbine blade 12, 14 is provided
with a number of secondary cooling passages 24 in the leading edge
region. The discharge passages 28, which are also attached in the
leading edge region of the turbine blade 12, 14 and which project
into a main cooling passage 26, serve as fastening elements for the
heat shield element 22 in addition to the guiding of the cooling
medium K. On account of the higher pressure which prevails in the
main cooling passage 26 of the blade basic body 30 compared with
the ambient pressure in the turbine unit 6, the cooling air K,
which is preferably used as cooling medium K, flows via the
secondary cooling passages 24 into the gap which is formed between
the outside surface 32 of the blade basic body 30 and the inside
surface 34 of the heat shield element 22, and also through the
discharge passages 28 of the heat shield element 22, wherein the
cooling air K which flows from the discharge passages 28 forms a
protective film between the operating medium M and the outside
surface 36 of the heat shield element 22. On the other hand, the
cooling air K which escapes from the secondary cooling passages 24
of the blade basic body 30 flows against the inside surface 34 of
the heat shield element 22 and cools this by means of the
impingement cooling effect which occurs as a result.
[0043] FIGS. 3 and 4 show the heat shield element 22 in two
different sectional views in each case, wherein it becomes apparent
from the longitudinal section of the heat shield element 22 which
is shown in FIG. 3 that the discharge passages 28, as seen in the
longitudinal direction of the heat shield element 22, are arranged
one behind the other, and wherein each discharge passage 28 extends
from the outside surface 36 of the heat shield element 22 towards
its inside surface 34. The discharge passages 28 in this case, as
shown in FIG. 4, can be concentrically arranged perpendicularly to
the longitudinal direction of the heat shield element 22.
[0044] As can especially be gathered from the view in FIG. 5, the
heat shield element 22 has a shape which is adapted to the profile
of the blade basic body 30 in the leading edge region.
Consequently, the effect is achieved inter alia of the turbine
blade 12, 14 also having a flow-optimized shape after attaching the
heat shield element 22 on the blade basic body 30. Moreover, a heat
shield element 22 which is curved in such a way results in a
constant distance between the inside surface 34 of the heat shield
element 22 and the outside surface 32 of the blade basic body 30,
as a result of which an especially effective cooling in this region
is made possible. The cooling air K which is required for the
cooling in this case flows from the main cooling passage 26 of the
turbine blade 12, 14, through the secondary cooling passages 24,
and through the discharge passages 28, as a result of which a
cooling film is formed on the outside surface 36 of the heat shield
element 22 on account of the cooling air K which flows from the
discharge passage 28 and on account of the operating medium M which
flows in the turbine unit 6. The cooling of the inside surface 34
of the heat shield element 22 and of the outside surface 32 of the
blade basic body 30 in the leading edge region of the turbine blade
12, 14 is carried out by the discharging of the cooling air K from
the secondary cooling passages 24, wherein the inside surface 34 of
the heat shield element 22 is cooled in an especially effective
manner as a result of the impingement cooling effect which occurs
in the process.
[0045] In order to achieve as far as possible an impingement
cooling on the inside surface 34 of the heat shield element 22 in
each of the regions which are exposed to inflow of the cooling air
K, the secondary cooling passages 24 are preferably arranged in
such a way that the cooling air K which flows from the secondary
cooling passages 24 impinges perpendicularly to the inside surface
34 of the heat shield element 22. The distance of the heat shield
element 22 from the blade basic body 30 in this case is preferably
to be selected so that as a result of a sufficiently high flow
velocity of the cooling medium K when impinging upon the inside
surface 34 of the heat shield element 22 an intimate contact
between the cooling air K and the impingement surface is brought
about, and in this way the impingement effect is established.
[0046] An especially expedient design of the turbine blade 12, 14
with the heat shield element 22 according to the invention is shown
in FIG. 6. In this case, the heat shield element 22 was integrated
into the leading edge region of the blade basic body 30, as a
result of which the original external shape of the turbine blade
12, 14 is advantageously maintained. The aerodynamic design of the
turbomachine is therefore not altered, as a result of which
reduction of the efficiency of the gas turbine, for example on
account of vortex formations on the outer edges when a heat shield
element 22 is attached externally on the blade basic body 30, is
prevented.
[0047] The gap between the heat shield element 22 and the blade
basic body 30 which is required for creating impingement cooling is
consequently achieved in the case of this special embodiment of the
turbine blade 12, 14 by the heat shield element 22 being seated in
a recess 38 which is provided in the blade basic body 30. In this
way, the outside surface of the turbine blade 12, 14 which reaches
into the flow passage of the gas turbine is partially formed by the
outside surface of the heat shield element 22.
[0048] The free ends of the heat shield element 22 according to
FIG. 5 are formed flush on the blade walls which are formed by the
basic body 30 in the case of the design according to FIG. 6 in
order to achieve an offset-free surface of the turbine blade 12,
14. For this, the part of the blade basic body 30 which lies
opposite the heat shield element 22 is set back towards the inside
of the blade so that the edge regions of the heat shield element 22
are connected to the blade body.
* * * * *