US20050169762A1

US20050169762A1 - Turbine blade for an aircraft engine and casting mold for its manufacture

Info

Publication number: US20050169762A1
Application number: US10/951,618
Authority: US
Inventors: Barbara Blume
Original assignee: Rolls Royce Deutschland Ltd and Co KG
Current assignee: Rolls Royce Deutschland Ltd and Co KG
Priority date: 2003-09-29
Filing date: 2004-09-29
Publication date: 2005-08-04
Also published as: DE10346366A1; EP1518619A1

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.

Description

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

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.
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.
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.
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

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.
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.
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.
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

The present invention is more fully described in the light of the accompanying drawings, showing a preferred embodiment. In the drawings,
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
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

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.
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.
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.
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

1 Turbine blade
2 Pressure side
3 Suction side
4 Leading edge
5 Trailing edge
6 Partition
7 Outer wall
8 Cooling air cavity
9 Cooling ducts of partition 6
10 Cavities
11 Film cooling ducts of outer wall 7
12 Film cooling ducts at trailing edge
13 Cross-duct

Claims

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.