GB2363167A

GB2363167A - Compressor/fan blade tip treatment bars in a gas turbine engine

Info

Publication number: GB2363167A
Application number: GB0013772A
Authority: GB
Inventors: Peter Allford; Mark John Houle
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2000-06-06
Filing date: 2000-06-06
Publication date: 2001-12-12
Anticipated expiration: 2020-06-06
Also published as: US20020000086A1; GB2363167B; US6409470B2; GB0013772D0

Abstract

A gas turbine engine casing includes a tip treatment ring 20, surrounding the blade tips of a compressor or fan for the purpose of improving blade stall/surge characteristics. The ring comprises annular end supports 18 which support tip treatment bars 16. Each tip treatment bar 16, is mounted within the respective end supports 18 by means of a damping element 22, (eg. made from an elastomer) which isolates the bar 16, from the end supports 18. The element 22, may be formed in situ, and damps vibrations induced in the tip treatment bars 16 so inhibiting the initiation of high cycle fatigue cracking. The bars may be hollow and both bars and supports may be made form a carbon fibre/bismaleimide composite material.

Description

2363167 TIP TREATMENT BARS IN A GAS TURBINE ENGINE This invention relates

to tip treatment bars of a rotor casing for a gas turbine engine.

W094/20759 discloses an anti-stall tip treatment 5 means in a gas turbine engine, in which an annular cavity is provided adjacent the blade tips of a compressor rotor. The cavity communicates with the gas flow path through the compressor past a series of bars extending across the mouth of the cavity.

10 Such tip treatments are applicable to both fans and compressors of gas turbine engines. and their purpose is to improve the blade stall characteristics or surge characteristics of the compressor.

The passage of the blade tips past the bars is creates vibrations in the bars which, over time, can result in high cycle fatigue failure of the bars. This failure is caused by vibration resonance between the tip treatment bars and the natural engine order modes.

According to one aspect of the present invention, 20 there is provided a gas turbine engine casing comprising tip treatment bars extending between annular end supports, each tip treatment bar being supported at each end by the end supports and being isolated, at at least one end, from the respective end support by 25 damping means.

In a preferred embodiment in accordance with the present invention, both ends of each tip treatment bar are isolated by damping means from the respective end support. The damping means may comprise a damping 30 material, and preferably a damping material having a high degree of damping at higher frequencies (i.e.

frequencies in excess of 1000 Hz).

The damping may be a polymer, and preferably an elastomeric polymer. Silicone elastomers may be used, 35 for example the silicone elastomer available under the name SILASTIC J.

Preferably, the tip treatment bars or the end supports, or both, are bonded to the damping material, for example by means of a silicone adhesive. A suitable adhesive is that available under the name 5 SILCOSET 152.

The damping material may comprise a moulded component which is assembled, after manufacture, with the tip treatment bars and the end supports.

Alternatively, the damping material may be moulded in 10 situ as the tip treatment bars are fitted to the end supports.

The bars may be solid, but they may alternatively be hollow, or provided with pockets, to lighten the structure. The bars may be made from any suitable is material,' for example from alloys commonly used in the aircraft industry. In a preferred embodiment, the tip treatment bars and, preferably, the end supports are made from a composite material such as a carbon fibre/bismaleimide composite.

20 For a better understanding of the present invention, and to show more clearly how it may be carried into effect, reference will now be made, by way of example, to the accompanying drawings, in which:

Figure 1 is a partial axial sectional view of a 25 fan stage in a gas turbine engine; Figure 2 is a view of tip treatment bars suitable for use in the engine of Figure 1; and Figure 3 is a sectional view of a single tip treatment bar.

30 Figure 1 represents a fan casing 2 of a gas turbine engine. A fan, represented by a single blade 4, is mounted for rotation in the casing 2. Guide vanes 6 and 8 are provided upstream and downstream, respectively, of the fan 4. The casing 2 includes a 35 circumferentially extending chamber 10, which communicates with the main gas flow through the fan (represented by an arrow 12) through an array of slots 14 defined between tip treatment.bars 16 disposed around the casing. The function of the chamber 10 in delaying the onset of stalling of the blades 4 is 5 disclosed in International Patent Publication W094/20759.

The tip treatment bars 16 are supported by annular end supports 18 to provide a tip treatment ring 20 which is fitted within the casing 2 and extends around 10 the fan 4. In the embodiments of Figures 1 to 3, the end supports 18 and the bars16 are separate components made from a carbon fibre/bismaleimide composite material, which enables the tip treatment ring to be light in weight while being capable of withstanding the is relatively high temperatures (in excess of 2000C) encountered in operation. However, in other embodiments the end supports may be integral with adjacent parts of the casing 2.

Vibration is induced in the bars 16 in operation 20 of the engine at a frequency determined by the passage of the blades 4. This vibration in a solid construction can lead to fatigue failure of the bars 16. The vibrating bars 16 deflect in a generally circumferential direction as indicated diagrammatically 25 in Figure 2 by an arrow 21, and consequently fatigue failure tends to be initiated by cracking at the slot ends.

The bars 16 are formed separately from the end support 18. Damping means is provided in the form of 30 damping elements or boots 22 of damping material. The boots 22 isolate each end of each tip treatment bar 16 from the respective end support 18. For this purpose, each end support 18 has an array of openings 24 which have generally the same shape as the cross-section of 35 the tip treatment bars 16, but is substantially larger.

The space between each tip treatment bar 16 and the wall of the opening 24 is filled by the damping material of the boot 22. As shown in Figure 3, each boot 22 has a flange 26 which extends for a short distance over the face of each end support 18 facing 5 away from the other end support 18. This flange 26 serves to locate the boot 22 against the end support 18.

Each boot 22 may be formed as a separate component before assembly with its respective tip treatment bar 10 16 and the end supports 18. Alternatively, the boots may be formed by moulding the damping material in situ between the tip treatment bar 16 and the end support 18, in a potting process. The boots 22 are bonded to the respective bars 16 and end supports 18 by means of 15 a-suitable adhesive, such as a silicone adhesive as is available under the name SILCOSET 152. The damping material itself is a silicone elastomer, such as the material available under the name SILASTIC J.

As shown in Figure 3, the tip treatment bar 16 is 20 hollow along its length and includes a lateral, radially inwardly directed, extension 28 along most of its length.

In operation, the vibration induced in the bars 16 is effectively damped by the damping material of the boots 22. Thus, the amplitude of vibration is reduced, so inhibiting the initiation and propagation of high cycle fatigue cracking.

Additionally, the construction shown in the drawings enables the use of relatively frangible tip 30 treatment bars 16. This frangibility is assisted by making the bars 16 hollow, as shown in Figure 3, or relatively thin, and of a suitable frangible material.

The use of frangible tip treatment bars allows a released portion of aerofoil to pass more easily into 35 the cavity, thus minimising consequential damage to further blade stages or to the casing,2. The blade -5 or fragment may be ejected past the tip treatment bars into the chamber 10, or may be held by the bars themselves, so preventing it from reaching the later compressor or turbine stages.

Claims

1. A gas turbine engine casing comprising tip treatment bars extending between annular end supports, each tip treatment bar being supported at each end by 5 the end supports and being isolated, at at least one end, from the respective end support by damping means.

2. A gas turbine engine casing as claimed in claim 1, in which each tip treatment bar is isolated, at both ends, from the respective end supports.

10

3. A gas turbine engine casing as claimed in claim 1 or 2, in which the damping means comprises a damping element situated between each tip treatment bar and the respective end support.

4. A gas turbine engine casing as claimed in claim 3, in which the tip treatment bars are disposed within openings in the end supports, each damping element being provided between a respective one of the tip treatment bars and the wall of the respective opening.

20

5. A gas turbine engine casing as claimed in claim 3 or 4, in which the damping elements are formed from an elastomer.

6. A gas turbine engine casing as claimed in claim 5, in which the elastomer is a silicone 25 elastomer.

7. A gas turbine engine casing as claimed in any one of claims 3 to 6, in which an adhesive is provided between each tip treatment bar and the respective damping element.

30

8. A gas turbine engine casing as claimed in any one of claims 3 to 7, in which an adhesive is provided between the respective damping element and each end support.

9. A gas turbine engine casing as claimed in 35 claim 7 or 8, in which the adhesive is a silicone adhesive.

10. A gas turbine engine casing as claimed in any one of claims 3 to 9, in which each damping element comprises a moulded component between each tip treatment bar and the respective end support.

5

11. A gas turbine engine casing as claimed in any one of claims 3 to 9, in which each damping element is formed in situ between each tip treatment bar and the respective end support.

12. A gas turbine engine casing as claimed in any 10 one of the preceding claims, in which each tip treatment bar is hollow.

13. A gas turbine engine casing as claimed in any one of the preceding claims, in which each tip treatment bar includes a pocket.

15

14. A gas turbine engine casing as claimed in any one of the preceding claims, in which each tip treatment bar is made from a carbon fibre/bismaleimide composite material.

15. A gas turbine engine casing as claimed in any 20 one of the preceding claims, in which each end support is made from a carbon fibre/bismaleimide composite material.

16. A gas turbine engine casing substantially as described herein with reference to, and as shown in, the accompanying drawings.

17. A gas turbine engine including an engine casing in accordance with any one of the preceding claims.