GB2117269A

GB2117269A - Thermal barrier coating

Info

Publication number: GB2117269A
Application number: GB08207109A
Authority: GB
Inventors: Kenneth George Page
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1982-03-11
Filing date: 1982-03-11
Publication date: 1983-10-12
Also published as: GB2117269B

Abstract

In order to reduce or eliminate the tendency of thermal barrier coatings to spall, the coating is applied to a component such as a turbine blade in the form of a large number of discrete lamina. The strain forces between the coating and the blade which cause the spalling are reduced due to the large number of small continuous areas covered by the coating.

Description

SPECIFICATION Thermal barrier coating This invention relates to thermal barrier coatings for gas turbine engine components, such as blades and vanes, combustor casings and flame tubes which operate at elevated temperatures.

The benefits of using thermal barrier coatings are well-known, these being a reduction in the quantity of cooling air required to maintain a component at a certain temperature, and protection against overheating. However these barrier coatings do tend to suffer from spalling in the severe operating conditions which occur in gas turbine engines. This problem arises from the temperature gradients across the component and coating, and the differences between the thermal conductivities and expansion coefficients of the coating and the component.

Various methods of more closely matching these characteristics of the component and the barrier coating have been proposed. These include the use of bond coatings applied to the component before the application of the barrier coating, and by gradually grading the barrier coating with the elements of the bond coating so that only the outermost layers of the barrier coating consist of 100% barrier coating material.

In general, these methods have met with limited success only.

Given that there will always be differences between the component and the coating in terms of thermal conductivity and expansion coefficient, the problem can be seen as the difference between the strain forces in the component and the coating, over a sufficiently large continuous area to cause the coating to become detached from the component. The strain forces which cause this spalling reduce with the size of the continuous area covered by the coating and the present invention proposes that the thermal barrier coating is attached to the component in the form of a relatively large number of separate and distinct zones.

Accordingly, the present invention provides a gas turbine engine component which in use is subjected to elevated temperatures, having a thermal barrier attached to at least a part of the surface area of the component, the thermal barrier coating comprising a plurality of discrete Iaminae.

The laminae can be in the shape of regular or irregular polygons or quasi-polygons. In one arrangement, the laminae can be hexagonal and evenly spaced from one another and the coating material can be magnesium zirconate or a stabilised zirconia.

The coating can be supplied to the component by the use of a mask which defines the shape and distribution of each lamina, in which the mask is laid over the component to which a bond coat may or may not have been applied. The barrier coating is then flame or plasma sprayed onto the component, and the mask removed to leave the discrete lamina attached to the component. In another method, the mask can be attached directly to the component by electro-plating a cell structure in a relatively soft metal onto the component and flame or plasma spraying the coating material onto the component. The cell structure can then be removed by electrochemical machining.

The present invention will now be more particularly described with reference to the accompanying drawing in which, Fig. 1 shows a gas turbine engine blade coated with a thermal barrier coating, according to the present invention, Fig. 2 is a view, to an enlarged scale, or part of the coating shown in fig. 1, and Fig. 3 is a view of part of a mask for use in applying a thermal barrier coating to an engine component.

Referring to figs., a gas turbine engine blade 10 includes an aerofoil portion 1 2 which is protected by a thermal barrier coating 14. The coating 14 is in the form of a plurality of discrete laminae 16, in this case each being hexagonal in shape, although any regular, irregular or quasi-polygonal form can be used. The laminae relatively small as compared with the surface area of the aerofoil portion 12, e.g. each lamina can be approxlmately 0.050" to 0.100" across the flats with a gap of approximately 0.005" between adjacent lamina, although these dimensions can be varied to suit the component being coated and the operating conditions. Similarly the thickness of the lamina can be varied to suit, but typically can be of the order 0.01 5".

The coating can be applied to the blade by first making a mask 1 8 which defines the shape and distribution of the lamina. The mask can be made by photo-etching a thin spring steel sheet. The mask is attached to the blade and the coating material is flame or plasma sprayed onto the blade. The metal mask is then removed to leave a pattern of discrete lamina of the coating material on the blade. Preferably a bond coat is applied to the blade by plasma spraying prior to the application of the thermal barrier coating, such a bond coating typically consisting of one or more elements from the group comprising Ni, Cr, Co, Al and Y.

In another method, the mask can be created by electro-plating the mask in a relatively soft metal directly onto the blade, and then applying the bond coat and the barrier coating as described above. The mask can then be removed by electrochemical machining.

In a further method, a bond coat and a subsequent barrier coating are both applied to the blade, and portions of the bond and barrier coating are removed by laser beam to define discrete laminae comprising the bond coating and the barrier coating.

By dividing the total area of the thermal barrier coating into a large number of discrete small areas, the strain forces acting over each small area is much reduced as compared with the strain forces acting over the initial larger area, and the tendency for the barrier coating to spall is much reduced.

The discrete Iaminae may vary in size and distribution on the component and selected areas only of the component can be coated. In particular, the invention is not restricted to the method of application of the barrier coating.

The barrier coating can be one of a number of known coatings, e.g. magnesium zirconate or a calcia or ythrium stabilised zirconia.

The gas turbine engine components which can be usefully coated in accordance with the invention include turbine blades, nozzle guide vanes, combustor casings and flame tubes.

Claims

1. A gas turbine engine component which, in use is subjected to elevated temperatures having a thermal barrier coating attached to at least a part of the surface area of the component, the thermal barrier coating comprising a plurality of discrete laminae.

2. A component as claimed in claim 1 in which each lamina is a regular polygon in shape.

3. A component as claimed in claim 1 in which each lamina is a hexagon, each hexagonal lamina being evenly spaced from adjacent laminae.

4. A component as claimed in claim 3 in which each lamina has a thickness not exceeding 0.025" and the gap between adjacent lamina does not exceed 0.010.

5. A component as claimed in any one of the preceding claims in which the thermal barrier coating is selected from the group comprising magnesium zirconate and a stabilised zirconia.

6. A component as claimed in any one of the preceding claims in which the component is a blade or vane, a flame tube, or a combustion casing.

7. A component as claimed in any one of the preceding claims in which the thermal barrier coating is applied to the component by attaching a metal mask over the component, the metal mask defining the shape and distribution of each discrete lamina, flame or plasma spraying an even thickness of the thermal barrier coating material onto the component, and removing the mask to leave a pattern of discrete laminae of thermal barrier material adhering to the component.

8. A component as claimed in any one of the preceding claims 1 to 6 in which the thermal barrier coating is applied to the component by electro-plating a cell structure in a relatively soft metal onto the component, plasma or flame spraying an even thickness of the thermal barrier coating material onto the component, and removing the metal cell structure by electrochemical machining.

9. A component as claimed in claim 7 or claim 8 in which a bond coating is applied to the component prior to the attachment of the metal mask, the bond coating comprising one or more elements selected from the group comprising Co, Ni, Cr, Al and Y.

10. A component as claimed in claim 9 in which the bond coating includes a proportion of the thermal barrier coating material.

11. A component as claimed in any one of the preceding claims 1 to 6 in which the thermal barrier coating is applied to the component by first applying a bond coating and a subsequent barrier coating to the component, removing portions of the bond and barrier coatings to define discrete laminae of the bond and barrier coatings.

12. A gas turbine engine component including a thermal barrier coating as herein described, and with reference to the accompanying drawing.