Method of repairing a damaged abradable coating

A method of repairing a damaged abradable coating (48) on a surface (46) of a shroud (44) in an assembled gas turbine engine (10) comprises inserting a boroscope (60) through an aperture (52) in the casing (50) of the compressor (26) of the gas turbine engine (10). The boroscope (60) is arranged to carry a conduit (62). The boroscope (60) and hence the conduit (62) are directed to the damaged abradable coating (48) on the surface (46) of the shroud (44). A liquid abradable glue (64) is supplied through the conduit (62) and the liquid abradable glue (64) is directed onto the surface (46) of the shroud (44) in the compressor (26) of the gas turbine engine (10) to repair the damaged abradable coating (48).

The present invention relates to a method of repairing a damaged abradable coating, in particular to a method of repairing a damaged abradable coating on a surface in an assembled engine, particularly a gas turbine engine.

The compressors and turbines of gas turbine engines are provided with abradable coatings at various positions. In particular abradable coatings are provided on the radially inner surfaces of compressor stator component surrounding the compressor rotor blades and abradable coatings are provided on the radially inner surfaces of turbine stator components surrounding turbine rotor blades. Abradable coatings may be provided on other surfaces of other components at other positions.

Currently damaged abradable coatings on components of the gas turbine engine are repaired, or reworked, at overhaul facilities. The repair of the abradable coating involves removing the damaged, or defective, abradable coating before applying a new abradable coating of the same composition/similar composition. The abradable coating is applied by thermal spraying or by plasma spraying. The cost associated with a scheduled overhaul visit, the cost of the abradable coating powder and the spraying time, are relatively small.

However, if an abradable coating is damaged and requires repair at unscheduled overhaul, the costs are more significant. This is due to the requirement to take the gas turbine engine to an overhaul facility and to disassemble the gas turbine engine into its modules, before the damaged abradable coating may be repaired by flame spraying or plasma spraying with a new abradable coating. Even minor damage to an abradable coating may lead to an unscheduled repair, which requires the removal of the compressor module or even the entire gas turbine engine from an aircraft. There are very high costs associated with this type of unscheduled overhaul.

Currently there are no methods of repairing a damaged abradable coating while the gas turbine engine in situ, e.g. while the gas turbine engine is located on an aircraft or on a ship or in an industrial plant.

Accordingly the present invention seeks to provide a novel method of repairing an abradable coating, which reduces, preferably overcomes, the above-mentioned problem.

Accordingly the present invention provides a method of repairing a damaged abradable coating on a surface in an assembled engine, the method comprising the steps of (a) inserting a boroscope through an aperture in a casing of the engine, the boroscope carrying a conduit, (b) directing the boroscope to the damaged abradable coating on the surface, (c) supplying a liquid abradable glue through the conduit, (d) directing the liquid abradable glue onto the surface in the engine to repair the damaged abradable coating.

Preferably the method comprises an additional step of heating the liquid abradable glue such that the liquid abradable glue hardens. Preferably the method comprises running the engine for a predetermined time to harden the abradable glue.

Preferably the liquid abradable glue comprises silica powder, sodium silicate and a dislocator. Preferably the dislocator comprises polyester, graphite or hexagonal-boron nitride.

Preferably the engine comprises a gas turbine engine.

Preferably the surface is a surface of a compressor stator component or a surface of a turbine stator component.

The damaged abradable coating may comprise a plasma sprayed abradable coating or a thermally sprayed abradable coating.

The damaged abradable coating may comprise aluminium, silicon and hexagonal boron nitride clad powder. The damaged abradable coating may comprise 12 wt % silicon, 16 wt % hexagonal boron nitride and the balance aluminium.

The damaged abradable coating may comprise aluminium, silicon and polyester. The damaged abradable coating may comprise 7 wt % silicon, 40 wt % polyester and the balance aluminium.

The damaged abradable coating comprises MCrAlY and bentonite.

A turbofan gas turbine engine10, as shown inFIG. 1, comprises an inlet12, a fan section14, a compressor section16, a combustion section18, a turbine section20and an exhaust22. The fan section14comprises a fan24. The compressor section16comprises an intermediate pressure compressor26and a high-pressure compressor28arranged in flow series. The turbine section20comprises a high-pressure turbine30, an intermediate pressure turbine32and a low-pressure turbine34arranged in flow series. The low pressure turbine34is arranged to drive the fan24, the intermediate pressure turbine32is arranged to drive the intermediate pressure compressor26and the high pressure turbine30is arranged to drive the high pressure compressor24.

The intermediate pressure compressor26, as shown more clearly inFIG. 2, comprises a rotor36carrying a plurality of stages of compressor rotor blades38and a stator40carrying a plurality of stages of compressor stator vanes42. The compressor rotor blades38in each stage are circumferentially spaced and extend generally radially outwardly from the rotor36. The compressor stator vanes42in each stage are circumferentially spaced and extend generally radially inwardly from the stator40. The stator40also comprises a plurality of shrouds44interconnecting the stages of compressor stator vanes42and the shrouds44are positioned radially around a corresponding one of the stages of compressor rotor blades38. The shrouds44have a radially inner surface46and the radially inner surface of each shroud44is provided with an abradable coating48. The stator40of the intermediate pressure compressor26also comprises a casing50and the casing50is provided with one or more apertures52to allow access for boroscopes. In operation of the gas turbine engine10the tips of the compressor rotor blades38pass close to the shrouds44to form a seal and may touch, and wear, the abradable coating48.

The abradable coating48comprises a plasma sprayed abradable coating or a thermally sprayed abradable coating. The abradable coating48may comprise aluminium, silicon and hexagonal boron nitride clad powder, e.g. comprising 12 wt % silicon, 16 wt % hexagonal boron nitride and the balance aluminium, or the abradable coating48may comprise aluminium, silicon and polyester, e.g. comprising 7 wt % silicon, 40 wt % polyester and the balance aluminium. The abradable coating48may comprise MCrAlY and bentonite. M in MCrAlY may be one or more of Ni, Co or Fe.

The high-pressure compressor28, the low-pressure turbine30, the intermediate pressure turbine32and the low-pressure turbine34are also provided with shrouds, which have abradable coatings on their radially inner surfaces.

As mentioned previously, the abradable coatings48on the radially inner surface46of the shrouds44may become damaged during operation of the turbofan gas turbine engine10.

The present invention provides a method of repairing a damaged abradable coating48on the surface46of a shroud44in an assembled gas turbine engine10. The method comprises inserting a boroscopes60through an aperture52in the casing50of the intermediate pressure compressor26of the gas turbine engine10. The boroscope60is also inserted through an aperture56in the radially outer platform54of one of the stator vanes42of the intermediate pressure compressor26of the gas turbine engine10. The boroscope60is arranged to carry a conduit62. The boroscope60and hence the conduit62are directed to the damaged abradable coating48on the surface46of the shroud44. A liquid abradable glue64is supplied from a supply66, e.g. a syringe etc, through the conduit62and the liquid abradable glue64is directed/supplied onto the surface46of the shroud44in the intermediate pressure compressor26of the gas turbine engine10to repair the damaged abradable coating48.

Following the deposition of the liquid abradable glue64, the liquid abradable glue64is heated such that the liquid abradable glue64hardens. The liquid abradable glue64may be heated by running the gas turbine engine10for a predetermined time to harden the liquid abradable glue64. However, other suitable methods of heating the liquid abradable glue64to harden it may be used, for example a microwave heater also directed through the aperture52in the casing50with the boroscope60etc. The liquid abradable glue comprises a dislocator.

The liquid abradable glue64comprises silica powder, sodium silicate and a dislocator. The dislocator may comprise polyester for low temperature use or graphite or hexagonal boron nitride for high temperature use. This liquid abradable glue64comprises in particular a high temperature binary adhesive, Sauereisen 315 (RTM), and a dislocator. Sauereisen 315 (RTM) is a two-part system comprising silica powder and sodium silicate. However, other suitable liquid abradable glues may be used and other suitable dislocators may be used.

Although the present invention has been described with reference to the repair of a damaged abradable coating on a radially inner surface of an intermediate pressure compressor stator shroud it is equally applicable to the repair of the radially inner surfaces of stator shrouds in the high pressure compressor, the high pressure turbine, the intermediate pressure turbine or the low pressure turbine.

Although the present invention has been described with reference to the repair of a damaged abradable coating on an inner surface of a stator shroud it is equally applicable to the repair of abradable coatings on other surfaces of stator or rotor components.

Although the present invention has been described with reference to a turbofan gas turbine engine it is equally applicable to other types of gas turbine engines and is equally applicable to aero gas turbine engines, marine gas turbine engine and industrial gas turbine engines.

Although the present invention has been described with reference to repair of thermally sprayed, or plasma sprayed, abradable coatings it is equally applicable to the repair of cast abradable coatings or other abradable coatings.

The present invention may also be applicable to other types of engine.

The advantage of the present invention is that it allows a damaged abradable coating on a component within an engine to be repaired to extend the life of the abradable coating for a period of time to allow overhaul of the engine to take place at a more convenient time. A further advantage of the present invention is that it allows a damaged abradable coating on a component within an engine to be repaired in situ, e.g. while the gas turbine engine is located on an aircraft, on a ship or in an industrial plant. The present invention allows a Damaged abradable coating on a component within an engine to be repaired without having to remove a module of the engine, or the whole engine, from an aircraft, ship or industrial plant.