WEAR MONITOR FOR A GAS TURBINE ENGINE

Abstract

An abradable liner for a gas turbine engine compressor stage, comprising: an abradable body having a radially inner surface over which an aerofoil tip passes in use; a plurality abradable depth indicators extending into the body from the surface; each depth indicator may be filled with a visual indicator. The depth indicator can be observed by a boroscope to provide a simple mechanical visual indication which can be used to provide quick and reliable information as to the extent and pattern of wear in an attrition liner.

Description

TECHNICAL FIELD OF INVENTION

This invention relates to an abradable liner for a compressor of a gas turbine engine. In particular, the invention relates to a wear indication system incorporated into the abradable liner.

BACKGROUND OF INVENTION

FIG. 1 shows a typical three shaft gas turbine engine 10. The gas turbine engine 10 includes an air intake 1, a fan 2 having rotating blades, a bypass duct 18 and an engine core 12. The engine core 12 includes an intermediate pressure compressor 3, a high pressure compressor 4, a combustor 5, a turbine arrangement comprising a high pressure turbine 6 an intermediate pressure turbine 7, a low pressure turbine 8 and an exhaust nozzle 9. Air entering the intake 1 is accelerated by the fan 2 and directed into two air flows. The first air flow passes into the engine core 12, and the second air flows along the bypass 18 to provide propulsive thrust.

The engine core air flow travels through the intermediate 3 and high 4 pressure compressors in turn. The compressed air exhausted from the high pressure compressor 4 is mixed with fuel from an injector 14 and burnt in the combustor 5. The hot gas expands through and drives the high 6, intermediate 7 and low 8 pressure turbines before being exhausted through the nozzle 9 and adding to the propulsive thrust created by the first air flow. The high 6, intermediate 7 and low 8 pressure turbines respectively drive the high 4 and intermediate 3 pressure compressors and the fan 2 via respective shafts.

It is well known that to maintain an efficient gas turbine engine the gap between compressor blade tips and the engine casing is closely controlled to minimise the leakage of compressed air over the blade tips and back upstream. To this end, the engine casings often include an attrition or abradable liner which provides a close fitting seal with the blade tips. The abradable liner is initially installed so as to be in contact with the compressor blade tips such that, the liner is scored by the rotating compressor (or fan as the case may be) during the first few rotations which removes enough material to allow a close fitting free rotation of the blades.

However, during normal engine use the radial position of the rotating blade tips move due to, for example, centrifugal forces, thermal expansion and vibration, and also during harsh operating conditions such as heavy landings or sharp manoeuvres. This can cause in-service damage to the abradable liner, which, in severe cases, can erode large arcuate sections which then require replacement. Replacement of the liners is expensive both in terms of overhaul cost and the associated loss of service of the engine.

The compressor casing may vary in its symmetry during operation from circular to elliptical or other non-round shape and it is difficult to accurately predict the actual axisymmetric and asymmetric rubs in a compressor. The rubs in an abradable liner are typically measured after running and striping a compressor. This information is used to adjust the geometry of the blades and/or casing liner to ensure the optimum sealing.

It will be appreciated that there is significant effort and cost required to strip the compressor to measure the level of rub and the full strip of the compressor may be some time after the initial entry into service of the engine. It is therefore difficult to frequently check the level of rub.

The present invention seeks to provide a solution to help monitor and control abradable liner damage.

STATEMENTS OF INVENTION

According to a first aspect there is provided an abradable liner for a gas turbine engine compressor stage, the liner extending about an axis and comprising: an abradable body having a radially inner surface over which an aerofoil tip passes in use; a plurality depth indicators extending into the body from the surface; wherein, the depth indicators are arranged in two or more groups each group being circumferentially spaced on the surface from another group; wherein each group has a first subgroup of depth indicators that are circumferentially offset from other depth indicators in the first subgroup, the depth indicators also being axially offset from other depth indicators in the first subgroup; each group further having a second subgroup of depth indicators circumferentially offset from the first subgroup, the depth indicators in the second subgroup being circumferentially aligned with other depth indicators in the second subgroup.

The depth indicators may be visible using a borescope or other inspection equipment allowing an easy determination of whether the depth indicator is still present in the abradable body or whether material has been removed from the abradable body that is at least equal to the depth of the depth indicator.

Preferably at least one depth indicator extends into the body a greater distance than at least one other depth indicator. This may be used to provide a scale from which the amount of material removed can be more easily calculated

Each depth indicator may be a blind aperture and may present a circular or elliptical cross-section to the surface. The maximum length of the longest axis of the cross-section may be 1 cm but more preferably may be less than 0.5 cm.

The relatively small size of the apertures presented to the surface ensures that aerodynamic effects on the compressor are limited so as to be almost negligible. This allows for depth indicators to be provided on the first engines entering development or service to confirm and check operation but not provided on subsequent engines and performance of the engines to be the same regardless of whether the holes are present or not.

Each depth indicator may be filled with a visual indicator. The visual indicator further reduces the aerodynamic effects of the depth indicators on the engine performance. The visual indicator may be coloured or fluorescent. The colouring may be used to help the operator of the inspection device confirm the portion of the liner that he is viewing.

The depth indicators may be arranged in two or more groups with the depth indicators of each group being filled with coloured material of a different colour to at least one of the other groups.

The different colours can allow easier inspection as it can easily be determined which group is being inspected.

The liner may extends about an axis with the surface defining an inner wall of a tube, wherein the groups extend in at least one array extending circumferentially along the inner wall.

Two or more arrays may be provided with each array being axially spaced from another of the arrays.

The liner may extend about an axis with the surface defining an inner wall of a tube, wherein the depth indicators are arranged in two or more groups each group being circumferentially spaced on the surface.

The liner may be used in a gas turbine engine compressor unit having a row of blades which rotate about an axis, and a liner extending about the axis; the liner may have an abradable body with a radially inner surface over which an blade tip passes in use; a plurality of abradable depth indicators extending into the body from the surface; wherein, the depth indicators are arranged in two or more groups each group being circumferentially spaced on the surface from another group; each group having a plurality of axially spaced depth indicators aligned to respectively correspond with a leading edge, a trailing edge and a mid-chord of the blade tip.

The axially spaced depth indicators may be circumferentially offset from each other.

Each group may have a first subgroup of depth indicators that are circumferentially offset from each depth indicator in the first subgroup and one depth indicator is axially aligned to correspond with the leading edge, a second depth indicator is axially aligned to correspond with a trailing edge and third depth indicator is axially aligned to correspond with a mid-chord of the aerofoil tip; each group further having a second subgroup of depth indicators circumferentially offset from the first subgroup, the depth indicators in the second subgroup being circumferentially aligned with each other.

Each depth indicator in the second subgroup may extend into the body the same distance as the one depth indicators in the second subgroup.

Each depth indicator in the first subgroup may extend into the body to a different distance as the other depth indicators in the first subgroup.

Each depth indicator may be a hole.

Each depth indicator may taper as it extends into the abradable liner. This allows the cross-sectional size of the hole to be measured and, provided the taper angle of the hole is known and the original cross-sectional area presented to the surface it is possible to calculate the amount of the liner removed.

Where coloured material is provided it may be provided in two or more layers of different colour. There may be a distinct boundary between the layers.

DESCRIPTION OF DRAWINGS

FIG. 1 shows a typical three shaft gas turbine engine;

FIG. 2 shows a cross section of a ring compressor case with an attrition liner in accordance with the present invention;

FIG. 3 shows a section of the compressor casing and depth indicators before wear;

FIG. 4 shows a section of the compressor casing and depth indicators after wear;

FIG. 5 is a plan view showing an arrangement of the depth indicators of FIG. 2;

FIG. 6 shows alternative, conical, depth indicators.

DETAILED DESCRIPTION OF INVENTION

FIG. 2 shows a cross section of a compressor casing 110 having an attrition liner 112 which incorporates an annulus of abradable material. In use, the liner 112 is located around the compressor blade 2 so as to provide a sealing function.

The casing of the embodiment shown is a ring compressor casing although this application is equally applicable to split casings, or casings used in turbine or fan applications.

Split casings are usually formed as two semicircles joined together at an attachment flange whilst the ring casings are provided by a single hoop of material. A separate case may be provided for one or more stages of the compressor and are joined together to form a casing that extends the length of the compressor. The attrition liner is sprayed onto the compressor casing or rotor path

The liner includes an abradable portion 116 which is designed to be contacted and abraded by the compressor blade in use. The abradable portion may be any suitable type known in the industry such as polyester based material.

Within the liner there is provided one or more depth indicators (55-57) that provide a visual indication of the wear of the abradable liner around the circumference of the liner. The relative axial and radial position of the blade tips and casing varies throughout the flight cycle and understanding the relative movement around the full circumference assists in providing more efficient components by helping to ensure that the optimum blade length can be provided to deliver the smallest running clearances between the blade tip and liner and also that the liner shape could be adjusted to correct for flight asymmetries.

Each depth indicator is provided by a hole or that is provided within the liner at a known circumferential and axial position and extending from the surface of the liner into the liner body. The holes and grooves can be arranged in arrays extending about the circumference with each array having depth indicators of identical, or different depths, as required.

In one embodiment as shown in FIG. 3 three depth indicators 55, 56, 57 are provided with each indicator being of a different depth. The depth indicators extend into the liner from the abradable surface 50. In a pristine case, where the engine is yet to be used, each depth indicator may be observed visually from a borescope inserted into the engine at a routine, or scheduled, inspection. The operator need not conduct time consuming measurement activities to determine that each depth indicator is observed.

After time, the liner may be eroded to a new liner surface 50 depicted in FIG. 4. The erosion may be to an extent greater than the depth of one or more of the depth indicators which means that one or more of the depth indicators is not visible. Upon inspection of the liner of FIG. 4 by a borescope it is observable that depth indicator 55 is not visible but indicators 56 and 57 can still be observed. Without extensive measurement and calculation it is possible for the inspector to determine that the liner has been worn to a distance between the depth of 55′ and the depth of 56′.

Inspection may be facilitated by filling each depth indicator with a material that is easily identifiable against the background of the liner. The material may be identifiable by colour or by the presence of a marker e.g. a fluorescent compound. Where colour is used the depth indicators of a particular depth may each have the same colour which is different to the depth indicators of another depth. The colours or pattern of colours indicate the depth to which the liner has been removed.

Due to the simplicity of inspection the abrasion of the liner can be regularly reviewed; even up to the frequency of after every flight. This level of assessment finds particular advantage when testing the engine as the effect of different manoeuvres or flight schedules on tip rub can be assessed.

It is possible to form new depth indicators in a pre-used abradable liner in order to further monitor the rub on the abradable liner over a greater use period. The relative axial and radial displacements between rotor and casing can be observed over a longer use period that could help to determine the optimum scheduling for replacement of the abradable liners and/or blades.

The filled depth indicators may be circumferential slots and/or drilled holes arranged at a plurality of different circumferential locations. Where holes are used they may be 1 cm in diameter, or less. A diameter of 4-5 mm has been found to be a suitable size.

Where holes are used of this size, or below, the aerodynamic effect on the blades passing over the surface of the liner is negligible even if they are not filled.

In one embodiment depth indicators are provided at four circumferential locations around the liner and they may be offset from each other to limit the danger of a weakened liner caused by too many depth indicators in a given locale.

The depth indicators may be arranged in a group of five as shown in FIG. 5 with two of the holes 58a, 58b aligned circumferentially to help to understand if there is any asymmetry in the angle of the blade tip 60 relative to the liner as it passes over the liner. Greater wear in one or more of the depth indicators 58a, 58b is indicative that the blade tip is not parallel to the liner surface 50.

In a further embodiment as shown in FIG. 6 the depth indicators 62 have a conical profile such that the depth that the liner is abraded can be determined by measuring the diameter of the hole at the surface.

Each hole or groove can be filled with a coloured material in order to provide a substantially smooth surface free of the substantial discontinuities that would be exhibited if the holes or grooves were not filled. The coloured material may be provided as one colour per hole, or arranged in layers such that there are multiple colours per hole. Where layers are provided the boundaries may be distinct or blended.

Where the holes are arranged in groups, all the holes of a given depth may be filled with the same colour material to help identify which hole is part of which group.

The invention provides a simple mechanical visual indication which can be used to provide quick and reliable information as to the extent and pattern of wear in an attrition liner. This can be used by maintenance staff to determine when an engine requires an overhaul and allows for efficient scheduling.

The arrangements allow an early indication of rubs around the circumference of the compressor casing. This allows the geometry of the compressor to be modified as soon as an engine is run—rather than waiting for a full strip and measurement. The forms of depth indicators: conical, cylindrical or grooved may be used independently or together as deemed appropriate to achieve the best depth determination. Where grooves are used they should be kept small to keep the aerodynamic effects to a minimum.

Similarly, the embodiments may find application in determining rub depth and location on any appropriate liner in pump, fan or turbine applications.

Claims

1. An abradable liner for a gas turbine engine compressor stage, the liner extending about an axis and comprising:

an abradable body having a radially inner surface over which an aerofoil tip passes in use;

a plurality depth indicators extending into the body from the surface;

wherein, the depth indicators are arranged in two or more groups each group being circumferentially spaced on the surface from another group; wherein each group has a first subgroup of depth indicators that are circumferentially offset from other depth indicators in the first subgroup, the depth indicators also being axially offset from other depth indicators in the first subgroup;

each group further having a second subgroup of depth indicators circumferentially offset from the first subgroup, the depth indicators in the second subgroup being circumferentially aligned with other depth indicators in the second subgroup.

2. An abradable liner according to claim 1, wherein at least one depth indicator extends into the body a greater distance than at least one other depth indicator.

3. An abradable liner according to claim 1, wherein each depth indicator is a blind aperture.

4. An abradable liner according to claim 3, wherein each blind aperture presents a circular or elliptical cross-section to the surface.

5. An abradable liner according to claim 4, wherein the maximum length of the longest axis of the cross-section is between 4 to 5 mm

6. An abradable liner according to claim 3, wherein each depth indicator tapers as it extends into the abradable liner.

7. An abradable liner according to claim 1, wherein each depth indicator is filled with a visual indicator of a coloured material.

8. An abradable liner according to claim 7 wherein the coloured material is provided as two or more discrete layers.

9. A gas turbine engine compressor unit having a row of blades which rotate about an axis, and a liner extending about the axis;

the liner having an abradable body with a radially inner surface over which an blade tip passes in use;

a plurality of abradable depth indicators extending into the body from the surface;

wherein, the depth indicators are arranged in two or more groups each group being circumferentially spaced on the surface from another group;

each group having a plurality of axially spaced depth indicators aligned to respectively correspond with a leading edge, a trailing edge and a mid-chord of the blade tip.

10. A gas turbine engine compressor unit according to claim 9, wherein the groups extend in at least one array extending circumferentially along the inner wall.

11. A gas turbine engine compressor unit according to claim 10, wherein two or more arrays are provided with each array being axially spaced from another of the arrays.

12. A gas turbine engine compressor unit according to claim 9, wherein each group has a first subgroup of depth indicators that are circumferentially offset from each depth indicator in the first subgroup and one depth indicator is axially aligned to correspond with the leading edge, a second depth indicator is axially aligned to correspond with a trailing edge and third depth indicator is axially aligned to correspond with a mid-chord of the aerofoil tip; each group further having a second subgroup of depth indicators circumferentially offset from the first subgroup, the depth indicators in the second subgroup being circumferentially aligned with each other.

13. A gas turbine engine compressor unit according to claim 9, wherein each depth indicator in the second subgroup extend into the body the same distance as the one depth indicators in the second subgroup.

14. A gas turbine engine compressor unit according to claim 9, wherein each depth indicator in the first subgroup extend into the body to a different distance as the other depth indicators in the first subgroup.

15. A gas turbine engine compressor unit according to claim 9, wherein each depth indicator is filled with a visual indicator of a coloured material.

16. An abradable liner for a gas turbine engine compressor stage, comprising: an abradable body having a radially inner surface over which an aerofoil tip passes in use; and

a plurality abradable depth indicators extending into the body from the surface; wherein each abradable depth indicator comprises an aperture filled with a coloured visual indicator.

17. An abradable liner according to claim 16, wherein each blind aperture presents a circular or elliptical cross-section to the surface.

18. An abradable liner according to claim 17, wherein the maximum length of the longest axis of the cross-section is between 4 to 5 mm

19. An abradable liner according to claim 16, wherein each depth indicator tapers as it extends into the abradable liner.