GB2394751A

GB2394751A - Anti creep turbine blade with internal cavity

Info

Publication number: GB2394751A
Application number: GB0225592A
Authority: GB
Inventors: Roderick Miles Townes
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2002-11-02
Filing date: 2002-11-02
Publication date: 2004-05-05
Also published as: GB0225592D0

Abstract

An anti-creep uncooled turbine blade 16a comprises a solid root portion 24 and a an aerofoil portion 17, the aerofoil having an internal cavity or void 20 within it, extending between the tip 22 and the root 24. This void reduces the centrifugal forces during operation in a gas turbine engine, and thus reduces creep in the blade. Also disclosed is a method for producing the blade, involving the steps of positioning a soluble core within a mould in a spaced relationship with said mould, and filling the space with wax. When solid, the wax is then covered by a flit, before being melted away. Molten metal can then be poured into the shell, and when cooled, the shell is removed, and the core dissolved and ejected through an aperture 28. A cap can then be placed over the aperture to create a sealed volume within the blade.

Description

239475 1

ANTI-CREEP TURBINE BLADE

The present invention relates to a turbine blade, and has particular efficacy when used in a stage or stages of 5 turbine blades in a gas turbine engine.

Whilst it is common practice to provide cooling air to the first high pressure stage of turbine blades in a gas turbine engine, intermediate and low pressure stages of turbine blades are not normally cooled. The drop in 10 temperature of gas that has passed the first high pressure stage of blades and its immediately following stage of high pressure guide vanes is such as to render the steps unnecessary. However, a drawback is experienced by way of residual gas temperature combined with centrifugal forces 15 generated during operational rotation of an associated turbine disk, causing the aerofoil portions to creep, ie to increase their respective lengths. This in turn causes blade tip rub on surrounding structure with consequent increase in tip clearances and reduction in performance 20 efficiency.

It is known to cast blades around a relatively light ceramic core, respective ends of which protrude through the tip and root of the blade. The core is retained in the blade during its lifetime and is shaped so as to take up 25 more space near the blade tip than at the blade root. This provides lightness at the blade end near its tip and strength at the root, enabling better absorbance of stresses exerted by centrifugal forces generated during operation of a stage of the blades in a gas turbine engine.

30 The present invention seeks to provide an improved uncooled turbine blade.

According to the present invention, an uncoated turbine blade comprises an aerofoil portion and a solid root portion by which said turbine blade is locatable on

turbine disk for operation in a gas turbine engine, wherein said aerofoil portion contains an internal space volume that extends from adjacent its tip extremity to a point of termination short of said root portion.

5 The invention will now be described by way of example and with reference to the accompanying drawings in which: Figure 1 is a diagrammatic part cross-sectional view of a gas turbine engine including a stage of turbine blades in accordance with the present invention.

10 Figure 2 is an enlarged view of a turbine blade in the stage of turbine blades in Figure 1 in accordance with the present invention.

Figure 3 is a view in the direction of arrow 3 in Figure 2.

15 Referring to Figure 1, a gas turbine engine 10 includes a compressor 12, combustion equipment 14, three stages of turbine blades 16, 16a and 16b and an exhaust nozzle 18. Each blade in the second (intermediate) stage of blades 16a has an internal space volume 20 formed in the 20 radially outer portion of its aerofoil 17 with respect to the axis of rotation of the stage of blades.

Referring now to Figure 2, space volume 20 is provided so as to reduce the weight of the outer portion of blade 16a and thereby reduce the stresses exerted on the aerofoil 25 17 which stresses are generated by centrifugal forces dur ng rotation of the stage of blades in an operating associated engine and cause the material of the aerofoil 17 to creep, ie to elongate.

In the example depicted in the accompanying drawings 30 aerofoil 17 is an extension of a solid root 24. Space volume 20 extends from near the tip of aerofoil 17 to a pc-,n. of termination 24 at about mid length thereof.

However, the form and extent of space volume 20 as depicted sho.:ld not be regarded as limitive. The aim is provide in

combination a space volume 20 to lighten aerofoil 17 and solid root 24, by which blade 16a is attachable to an associated disk 26 in known manner and which is sufficient bulk to absorb the stresses mentioned hereinbefore. In all 5 cases the form and extent of the space volumes will depend upon the stresses involved and the aerofoil form, particularly it thickness. This latter point is exemplified in Figure 3 wherein space volume 20 is seen to leave this wall sections in aerofoil 17.

10 The person skilled in the art, having read this specification, will appreciate that any stage of uncooled

turbine blades, including a high pressure stage the operating temperature of which is low enough, can incorporate the features of the present invention.

15 Manufacture of a turbine blade in accordance with the present invention is achieved by a known process in cooled turbine blade manufacture, ie forming a soluble core of appropriate shape, positioning it within a could in spaced relationship with adjacent wall surfaces of the mould, and 20 filling the space with wax. When the wax has solidified, it is removed from the mould and dipped in a wet silicon frit, and when that has dried, the wax is melted and run from the resulting shell. Molten metal is then poured into the regained space. Upon solidification of the metal, the 25 shell is broken away leaving the cast blade containing the solid core, which is then dissolved and removed, in the present example, via an aperture 28 in the tip of the aerofoil 17.

If appropriate aperture 28 may be sealed with a brazed 30 cap (not shown) after the core (not shown) has been removed.

Claims

1. An uncoated turbine blade comprising an aerofoil portion and a solid root portion by which said turbine blade is locatable on a turbine disk for operation in a gas 5 turbine engine, wherein said aerofoil portion contains an internal space volume that extends from adjacent its tip extremity to a point of termination short of said root portion.

2. An uncooled turbine blade as claimed in claim 1 10 wherein said internal space volume is totally enclosed.

3. A stage of turbine blades as claimed in claim 1 or claim 2 when assembled on a turbine disk.

4. A gas turbine engine including a stage of turbine blades as claimed in claim 3.

15

5. An uncoated turbine blade substantially as hereinbefore described with reference to, and as shown in, the accompanying drawings.

6. A method of manufacturing an uncoated turbine blade as claimed in claim 1 or claim 2 comprising the steps of 20 positioning a soluble core in a moult in spaced relationship with the interior surfaces thereof, filling said space with wax, covering said wax with a frit, hardening said frit, melting said wax out of said frit, and then dissolving said soluble core and emptying it from said 25 recovered space via an aperture in the tip of the aerofoil of said turbine blade so as to form a space volume therein.

7. A method of manufacturing an uncooled turbine blade as claimed in claim 6 including a final step of covering said aper Ore with a cap after draining said dissolved core 30 therefrom, so as to totally enclose said space volume.