GAS TURBINE ENGINE CASE WITH HEATING LAYER AND METHOD

Abstract

An assembly of a rotor and case for a gas turbine engine comprises a rotor having a plurality of circumferential blades each with a blade tip. A case comprises a structural layer forming an annular body receiving therein the rotor, with a tip clearance being defined between the blade tips and the annular body, and a heating layer connected to the at least one structural layer. The heating layer has a resistivity to heat the structural layer when an electric current passes through the heating layer. An electric power source supplies current to the heating layer, the heating layer configured to in use change the diameter of the annular body to adjust the tip clearance. A method for adjusting a tip clearance between blade tips of a rotor and inner surface of a case in a gas turbine engine is also provided.

Description

TECHNICAL FIELD

The application relates generally to gas turbine engines and, more particularly, to a method and means to control the clearance between compressor or turbine blades and the corresponding compressor/engine case.

BACKGROUND OF THE ART

Compressor blade tip clearances are conventionally built larger than optimal to avoid blade rub in some flight conditions, such as takeoff and slam acceleration. The clearances are larger than optimal to compensate for a faster blade growth rate relative to the case growth rate because of a faster blade heat ramp rate and blade growth from centrifugal force. Compressor blade growth rate may also be faster than the case growth rate because the blades heat up faster due to their lower thermal mass, larger surface area to volume ratio, thinner wall thickness, and/or exposure to heat from all sides rather than from the gas side only as for the compressor case or engine case.

SUMMARY

In one aspect, there is provided an assembly of a rotor and case for a gas turbine engine comprising: a rotor having a plurality of circumferential blades, each of the blade having a blade tip; a case comprising at least one structural layer forming an annular body receiving therein the rotor, with a tip clearance being defined between the blade tips and an inner surface of the annular body, and a heating layer connected to the at least one structural layer, the heating layer having a resistivity to heat the structural layer when an electric current passes through the heating layer; and an electric power source connected to the heating layer and supplying current to the heating layer, the heating layer configured to in use change the diameter of the annular body to adjust the tip clearance between the rotor and the case.

In a second aspect, there is provided a method for adjusting a tip clearance between blade tips of a rotor and inner surface of a case in a gas turbine engine, comprising: applying a heating layer on the case; determining that gas turbine engine conditions require an increase in diameter of the case for tip clearance; and supplying electric current to the heating layer on the case to increase the diameter.

Further details of these and other aspects of the present invention will be apparent from the detailed description and figures included below.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures, in which:

FIG. 1 is a schematic cross-sectional view of a turbofan gas turbine engine; and

FIG. 2 is a schematic sectional view of a clearance between a blade and a case, with an electrically-powered heating layer in accordance with the present disclosure.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

FIG. 1 illustrates a turbofan gas turbine engine 10 of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan 12 through which ambient air is propelled within an engine case 13, a multistage compressor 14 within a compressor case 15 for pressurizing the air, a combustor 16 in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section 18 for extracting energy from the combustion gases.

Referring to FIG. 2, a blade is shown enlarged at 20. The blade 20 is one of multiple blades of a rotor, whether it be part of the fan 12 or the compressor 14, The blade 20 is illustrated as having a blade tip 21. The blade tip 21 is shown as having a straight profile. It is understood that the blade tip 21 may have any appropriate profile as alternatives to the straight profile of

A clearance A is defined between the blade tip 21 and an inner surface 30 of the case 13/15. The case 13/15 is illustrated as consisting of one structural layer (e.g., sheet metal) and has an outer surface 31. However, the case 13/15 may have additional structural layers, typically radially outward of the outer surface 31.

A heating layer is generally shown at 32 and is against the outer surface 31. The heating layer 32 is of the type heating up when an electrical current is passed through it. In accordance with an embodiment, the heating layer 32 is an electric heater coating that is applied (e.g., sprayed, painted, printed, etc.) to the outer surface 31. One suitable type of coating used as heating layer 32 is a product provided by Datec Coating Corporation. In accordance with another embodiment, the heating layer 32 consists of heating foil that is applied against the outer surface 31, In both embodiments, the installation of the heating layer 32 is readily effected when compared to existing prior art system using coils, etc. When heater coating is used, applying the coating may be similar to applying a coat of paint. Moreover, the heating layer 32 may be on either side of the case 13/15 where possible, of sandwiched between layers of the case 13/15, etc. In both cases, the heating layer 32 must be made of a material having a minimum resistivity to generate sufficient heat from the electric current circulated therein.

The heating layer 32 is therefore wired to an electric power source 40, for instance by way of wires 41 (i.e., leads, lead wires) on opposite sides of the layer 32. Any appropriate type of arrangements may be used to allow a current supply through the heating layer 32 from the electric power source 40. It is also considered to have multiple circuits for instance in parallel to heat up the heating layer 32 in segments.

According to the present disclosure, there is provided a smaller cold tip clearance between blade tip and the case then in conventional arrangements, such that the engine as the smallest tip clearance at cruise condition. The case is electrically heated for the case growth rate to better match the blade growth rate, in some flight conditions. For instance, at takeoff or at slam acceleration where blade rub could occur, the case is heated up by circulating current from the electric power source 40 in the heating layer 32, thereby increasing the case growth rate. The electric heating from the electric power source 40 may be switched off at cruise condition. In other words, the case 13/15 is heated up when large tip clearance is needed to avoid rubbing, and the electric heating is switched off when the small clearance is needed for better performance. In an embodiment, the heating layer 32 is sandwiched between different layers of the case 13/15.

The electric power source 40 may therefore comprise a controller or like processing unit to determine the flight conditions in which the gas turbine engine is operated. For instance, the electric power source 40 may have its controller connected to aircraft command center or to an engine control unit to receive this information. Alternatively, the electric power source 40 may have its controller connected to various probes to determine the flight conditions, such as temperature probes, manometers, etc. The electric power source 40 may therefore monitor the flight conditions to calculate an actual tip clearance. The electric power source 40 may then actively adjust the size of the case 13/15 by determining the amount of current required to reach the desired case size, as a function of the known resistivity of the heating layer 32.

The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. For example, multiple heating layers (e.g., of different types) may be provided on the case to increase heating. Various wiring arrangements may be used to supply electric current to the heating layer 32. The heating layer 32 may cover a portion or most of the case 13/15 opposite the rotor. Still other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.

Claims

1. An assembly of a rotor and case for a gas turbine engine comprising:

a rotor having a plurality of circumferential blades, each of the blade having a blade tip;

a case comprising at least one structural layer forming an annular body receiving therein the rotor, with a tip clearance being defined between the blade tips and an inner surface of the annular body, and a heating layer connected to the at least one structural layer, the heating layer having a resistivity to heat the structural layer when an electric current passes through the heating layer; and

an electric power source connected to the heating layer and supplying current to the heating layer, the heating layer configured to in use change the diameter of the annular body to adjust the tip clearance between the rotor and the case.

2. The assembly according to claim 1, wherein the heating layer is a heater coating applied to one of the surfaces of the structural layer.

3. The assembly according to claim 1, wherein the heating layer is a heating foil applied to one of the surfaces of structural layer.

4. The assembly according to claim 1, wherein the heating layer is against an outer surface of structural layer.

5. The assembly according to claim 1, wherein the case is a compressor case and the rotor is a compressor rotor.

6. The assembly according to claim 1, wherein the case is a turbine case and the rotor is a turbine rotor.

7. A method for adjusting a tip clearance between blade tips of a rotor and inner surface of a case in a gas turbine engine, comprising:

applying a heating layer on the case;

determining that gas turbine engine conditions require an increase in diameter of the case for tip clearance; and

supplying electric current to the heating layer on the case to increase the diameter.

8. The method according to claim 7, wherein determining that gas turbine engine conditions require an increase in diameter comprises determining that the gas turbine engine is in takeoff or slam acceleration.

9. The method according to claim 7, further comprising stopping a supply of electric current when gas turbine engine conditions no longer require an increase in diameter of the case.

10. The method according to claim 9, wherein stopping the supply of electric current occurs during cruise conditions of the gas turbine engine.

11. The method according to claim 7, wherein applying a heating layer on the case comprises applying a heater coating to one of the surfaces of the case.

12. The method according to claim 7, wherein applying a heating layer on the case comprises applying a heating foil to one of the surfaces of the case.