U.S. patent application number 10/951618 was filed with the patent office on 2005-08-04 for turbine blade for an aircraft engine and casting mold for its manufacture.
Invention is credited to Blume, Barbara.
Application Number | 20050169762 10/951618 |
Document ID | / |
Family ID | 34178042 |
Filed Date | 2005-08-04 |
United States Patent
Application |
20050169762 |
Kind Code |
A1 |
Blume, Barbara |
August 4, 2005 |
Turbine blade for an aircraft engine and casting mold for its
manufacture
Abstract
A hollow-type turbine blade produced by a casting process has a
cooling air cavity (8) and film cooling ducts (12) originating at
this cooling air cavity (8). Adjacent film cooling ducts with a
large length-diameter ratio are connected by cross-ducts (13) which
are offset relative to each other and arranged vertically to the
film cooling ducts. The casting mold comprises a core corresponding
to a brickwall design of the film cooling ducts and the
cross-ducts, with long core pins being intersupported by cross
pins. The turbine blade, including all film cooling ducts, can be
produced with high quality in a casting process.
Inventors: |
Blume, Barbara; (Berlin,
DE) |
Correspondence
Address: |
Harbin King & Klima
500 Ninth Street SE
Washington
DC
20003
US
|
Family ID: |
34178042 |
Appl. No.: |
10/951618 |
Filed: |
September 29, 2004 |
Current U.S.
Class: |
416/223R |
Current CPC
Class: |
B22C 21/14 20130101;
F01D 5/187 20130101; Y02T 50/676 20130101; Y02T 50/60 20130101;
F05D 2230/21 20130101 |
Class at
Publication: |
416/223.00R |
International
Class: |
B63H 001/14 |
Foreign Application Data
Date |
Code |
Application Number |
Sep 29, 2003 |
DE |
DE 103 46 366.6 |
Claims
What is claimed is:
1. A casting mold for the production of a hollow turbine blade
having at least one radial cavity flown by cooling air and a
multitude of film cooling ducts, radially spaced and arranged one
above the other, extending from an inner surface of the cavity to
an outer surface, the mold comprising an investment surrounding a
core for the production of cavities and ducts which is soluble upon
casting, wherein, the core for the formation of film cooling ducts
having a large length-diameter ratio comprises a multitude of core
pins of an extension corresponding to the duct length which, in a
longitudinal direction of the casting mold, are spaced and arranged
one above the other and in that adjacent core pins are connected to
each other by at least one cross-pin which is arranged generally
normal to the core pins and which intersupports the core pins
during a core making and a casting process.
2. A casting mold in accordance with claim 1, wherein, in the
longitudinal direction of the casting mold, adjacent cross-pins,
that lie above one another, are arranged regularly offset to each
other.
3. A casting mold in accordance with claim 2, wherein a number of
cross-pins arranged between two core pins is variable in dependence
of the length of the core pins.
4. A casting mold in accordance with claim 3, wherein the
cross-pins are related to the core pins in an area of a portion of
the mold forming the trailing edge of the turbine blade.
5. A casting mold in accordance with claim 1, wherein the
cross-pins are related to the core pins in an area of a portion of
the mold forming the trailing edge of the turbine blade.
6. A casting mold in accordance with claim 1, wherein a number of
cross-pins arranged between two core pins is variable in dependence
of the length of the core pins.
7. A cast turbine blade for an aircraft engine having a leading
edge, a trailing edge and at least one internal radial cooling
air-supplied cavity from which a plurality of cast, radially spaced
film cooling ducts extend to an outer surface of the blade to
produce a cooling air film, the turbine blade further comprising a
plurality of cast cross-ducts, each cross-duct interconnecting two
adjacent film cooling ducts, the cross-ducts being positioned
generally normal to the film cooling ducts with adjacent
cross-ducts connected to a film cooling duct being offset from one
another.
8. A turbine blade in accordance with claim 7, wherein a number of
the cross-ducts between adjacent film cooling ducts is set in
dependence of a length of the film cooling ducts.
9. A turbine blade in accordance with claim 8, wherein adjacent
cross-ducts connected to a film cooling duct are offset from one
another along a length of the film cooling duct.
10. A turbine blade in accordance with claim 9, wherein the
cross-ducts are related to the film cooling ducts in an area of the
trailing edge of the turbine blade.
11. A turbine blade in accordance with claim 7, wherein adjacent
cross-ducts connected to a film cooling duct are offset from one
another along a length of the film cooling duct.
12. A turbine blade in accordance with claim 7, wherein the
cross-ducts are related to the film cooling ducts in an area of the
trailing edge of the turbine blade.
13. A method of producing cooling ducts in a turbine blade for an
aircraft engine, the turbine blade having a leading edge, a
trailing edge and at least one radial cooling air-supplied cavity,
the cooling ducts being radially spaced and extending from the
air-supplied cavity to an outer surface of the blade to produce a
cooling air film, comprising: producing a lost core casting mold
for precision casting the turbine blade, forming a core portion
from a material soluble upon casting of the turbine blade, the core
portion having a plurality of core pins of an extension
corresponding to desired lengths of the cooling ducts which, in a
longitudinal direction of the casting mold, are spaced and arranged
one above the other, the core pins forming the cooling ducts in the
turbine blade when the core material is dissolved during casting of
the turbine blade; forming a plurality of core cross-pins, each of
the core cross-pins interconnecting at least two of the core pins
to support the connected core pins until dissolved during the
casting of the turbine blade, the core cross-pins being formed
generally normal to the core pins, the core cross-pins forming
cross-ducts between the interconnected cooling ducts when the core
is dissolved during the casting of the turbine blade; and casting
the turbine blade to remove the core portion and form the turbine
blade.
14. A method in accordance with claim 13, wherein the number of
core cross-pins formed between adjacent core pins is varied in
dependence of the length of the core pins.
15. A method in accordance with claim 14, wherein adjacent core
cross-pins connected to a core pin are offset from one another
along a length of the core pin.
16. A method in accordance with claim 15, wherein the cooling ducts
are formed in an area of the trailing edge of the turbine
blade.
17. A method in accordance with claim 13, wherein adjacent core
cross-pins connected to a core pin are offset from one another
along a length of the core pin.
18. A method in accordance with claim 13, wherein the cooling ducts
are formed in an area of the trailing edge of the turbine blade.
Description
[0001] This application claims priority to German Patent
Application DE10346366.6 filed Sep. 29, 2003, the entirety of which
is incorporated by reference herein.
BACKGROUND OF THE INVENTION
[0002] This invention relates to a turbine blade made in a casting
process with at least one radial cavity flown by cooling air and a
multitude of film cooling ducts, radially spaced and arranged one
above the other, extending from the inner surface of the cavity to
the outer surface and, further, this invention relates to a casting
mold for the manufacture of the turbine blade in a casting process
using a soluble core for the production of the cavities.
[0003] One primary goal in the effort to enhance the performance of
aircraft gas turbine engines is the increase of the temperature of
the turbine gases. Internal cooling of the turbine blades
counteracts the constraints set in this respect by the limited heat
resistance of the available materials. The cooling air, upon
entering central, radial cavities in the interior of the blades, is
routed to the outside via a multitude of minute film cooling ducts
arranged radially spaced from the airfoil bottom to the blade tip
to produce a cooling air film on the outer surface of the blade.
This cooling air film forms a barrier layer between the surface of
the turbine blade and the hot gases impinging onto the blade
surface to cool, in particular, the blade pressure side and here,
especially, the turbine blade trailing edge which, due to its small
thickness, is sensitive to stresses and problematic with regard to
cooling. Turbine blades with a multitude of film cooling ducts in
the area of the trailing edge which, due to the small material
thickness in this blade area, are long and thin and which are
arranged parallelly spaced relative to each other are known from
Patent Specifications EP 0 916 809 A2 or DE 40 03 804 C2, for
example.
[0004] The manufacture of turbine blades with film cooling ducts in
the area of the blade trailing edge is, however, difficult in that
these ducts, upon forming the blades, are to be produced by a
demanding machining method on the basis of electric discharge
processes, namely electro-discharge machining (EDM), and in that
this method is costly and incurs the highest scrap rate in the
entire manufacturing process.
[0005] Currently, the casting of turbine blades together with the
film cooling holes at the trailing edge is not possible or, the
modern casting processes, for example on the basis of virtual
pattern casting or the casting method with water-soluble core, in
which the core residues in the inner of the finished casting are
dissolved out with suitable means, are disadvantageous in that the
respective core sections (thin pins) for the long, thin
trailing-edge film cooling ducts do not withstand the high stresses
occurring in the casting mold making process (baking). In virtual
pattern casting (VPC), the core sections are ceramic pins that are
liable to break during the making of the casting mold or in
consequence of the stresses occurring when the metal cools in the
mold. If the thin ceramic pins forming certain portions of the mold
break before the casting process is fully finished, the film
cooling ducts will not be formed completely, i.e. they will be
fully or partly blocked, thus rendering the blade
unserviceable.
BRIEF SUMMARY OF THE INVENTION
[0006] The present invention, in a broad aspect, provides a casting
mold and a turbine blade for gas turbine engines produced by means
of this mold, enabling the blade to be cost-effectively
manufactured in a casting process, while ensuring high quality and
adequate cooling.
[0007] It is a particular object of the present invention to
provide a solution to the above problems by a casting mold and a
corresponding turbine blade designed in accordance with the
features described herein. Certain features of the present
invention will be apparent from the description below.
[0008] The concept underlying the present invention is to produce
the turbine blade in its entirety, i.e. including the area of the
trailing-edge film cooling ducts with a large length-diameter
ratio, by means of a casting mold using a core soluble upon
performance of the casting process, with very long, thin core pins
being, however, intersupported by cross-pins. Thus, the stresses
occurring during baking and cooling of the core and during and
after the casting process will not damage the long core pins,
allowing the turbine blade, including the film cooling ducts, to be
cost-effectively produced in a casting process.
[0009] The turbine blade according to the present invention
features cross-ducts between the film cooling ducts which
correspond to the cross-pins and which extend vertically to the
film cooling ducts. This arrangement of the cross ducts relative to
the film cooling ducts ensures that the cooling air will only flow
in the direction of the film cooling ducts, i.e. will not be
diverted, thus providing the required cooling air film and the
required cooling effect.
BRIEF DESCRIPTION OF THE DRAWINGS
[0010] The present invention is more fully described in the light
of the accompanying drawings, showing a preferred embodiment. In
the drawings,
[0011] FIG. 1 is a sectional view of a hollow, cast turbine blade
with film cooling ducts provided in the outer wall on the pressure
side, and
[0012] FIG. 2 is a sectional view along line AA in FIG. 1 showing
the long film cooling ducts provided in the area of the trailing
edge of the turbine blade.
DETAILED DESCRIPTION OF THE INVENTION
[0013] The hollow-type turbine blade 1 produced by a precision
casting process with a lost core (soluble upon performance of the
casting process) has a pressure side 2, a suction side 3, a leading
edge 4 and a trailing edge 5. Partitions 6 between the outer walls
7 of the turbine blade 1 define cooling air cavities 8 extending in
the radial direction of the turbine blade which provide the blade
interior with cooling air supplied via openings in the blade root
(not shown). The cooling air enters adjacent cavities 10 in the
blade interior via cooling ducts 9 in the partitions 6 and/or flows
to the outside via film cooling ducts 11 or 12, respectively,
radially spaced in the outer wall 7. The air exiting from the film
cooling ducts 11 or 12, respectively, produces a cooling gas film
flowing along the outer wall to cool, in particular, the pressure
side 2 of the turbine blade 1.
[0014] The film cooling ducts 12 at the trailing edge 5 of the
turbine blade 1 are connected to each other by at least one
cross-duct 13 in the area in which the film cooling ducts 12 exceed
a certain length, i.e. in the area where their length-diameter
ratio is particularly large. The cross-ducts 13 which lie adjacent
to each other in the radial (longitudinal) direction of the turbine
blade 1 are offset relative to each other in brickwall style and
are positioned at angles to the film cooling ducts 12, and in the
embodiment shown, generally normal to the film cooling ducts 12.
This generally normal arrangement of the cross-ducts 13 relative to
the film cooling ducts 12 ensures that the cooling air will pass
through the film cooling ducts 12 completely and actually reach the
required cooling areas at the trailing edge 5 of the turbine blade
1.
[0015] The brickwall design of the film cooling ducts with large
length-diameter ratio is created by the special core structure of
the casting mold (not shown) for the production of the turbine
blade 1, namely in the area of the trailing edge 5. There, the
long, thin core material (core pins) for the production of the film
cooling ducts 12, which is arranged radially spaced and parallel
above one another, is supported by cross-pins which are radially
offset relative to each other and arranged generally normal to said
core material. The representation of the core structure has been
dispensed with in the present embodiment since it corresponds
exactly with the brickwall design of the film cooling ducts
connected by the cross-ducts shown in FIG. 2. This core structure
for the production of the trailing edge 5 prevents the thin core
material provided for the formation of the long film cooling ducts
12 from breaking due to shrinkage stresses resulting from the
cooling of the baked casting mold and the core or due to stresses
resulting from the cooling of the hot metal melt injected into the
casting mold, thus providing for complete formation of the film
cooling ducts 12 upon removal or evacuation of the core material
from the casting, i.e. the turbine blade 1, according to the lost
mold principle and ensuring efficient film cooling at the trailing
edge 5 of the turbine blade 1.
[0016] This core for the casting mold and the resultant form of the
turbine blade in the area of the film cooling ducts 12 at the
trailing edge 5 provides for cost-effective, high-quality
production of the turbine blades, including the film cooling ducts,
by means of a casting process.
LIST OF REFERENCE NUMERALS
[0017] 1 Turbine blade
[0018] 2 Pressure side
[0019] 3 Suction side
[0020] 4 Leading edge
[0021] 5 Trailing edge
[0022] 6 Partition
[0023] 7 Outer wall
[0024] 8 Cooling air cavity
[0025] 9 Cooling ducts of partition 6
[0026] 10 Cavities
[0027] 11 Film cooling ducts of outer wall 7
[0028] 12 Film cooling ducts at trailing edge
[0029] 13 Cross-duct
* * * * *