GB2225388A

GB2225388A - Rotor blade tip clearance setting in gas turbine engines

Info

Publication number: GB2225388A
Application number: GB8921959A
Authority: GB
Inventors: Lance Peter Bell
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1988-10-01
Filing date: 1989-09-28
Publication date: 1990-05-30
Anticipated expiration: 2009-09-28
Also published as: GB8921959D0; GB8823094D0; GB2225388B

Abstract

A turbine rotor stage has at least one blade 24 tipped with abrasive material which extends to a diameter greater than the diameter of the remaining blades 22 and which is capable of cutting a running groove in an internal ceramic liner 34 on the turbine casing 32. A method of setting a clearance gap G comprises overspeeding the rotor assembly by a predetermined RPM and at a temperature such that the tips of the cutting blades 24 cut into the liner and thereby set a clearance gap for the remaining blades. The remaining blades 22 may be ground to a uniform height when the rotor assembly is assembled without the slightly longer cutting blades. The abrasive tips may comprise crystals of boron nitride carried in a non-oxidisable matrix. <IMAGE>

Description

IMPROVEMENTS IN TIP CLEARANCE SETTING IN GAS TURBINE ENGINES This inventions relates to tip clearance setting in gas turbine engines. In particular the invention concerns setting the tip clearances in the turbine stage of the engine in such a way as to minimise the progressive loss of efficiency during engine life.

The efficiency of a turbine stage having unshrouded blades is greatly affected by the size of the tip clearance gap which, of course, determines the amount of tip leakage which takes place. Manufacturing variations and degradation during service as well as cyclic variations all contribute to a loss of efficiency although cyclic gap variations arising due to speed and thermal transients are only temporary.

GB 2,153,447 describes a compressor assembly in which one or only a few blades are provided with a hard facing on the radially outer end which, when contact occurs between the blades tips and a casing liner, do the rubbing work. An abradable lining is applied to the inside of the compressor casing and the hard facing applied to a thickened shroud-like end region of a blade adapted (with regard to wear) to suit the abradable coating. However, excessive wear and undesirable rubbing should be avoided especially when the casing grows out-of-round to produce pinch-points or to take up an oval shape. This may happen in transient operating states, such as start, acceleration, deceleration and shut-down which give rise to irregular thermal conditions and an even or unmatched expansion or contraction and irregular mechanical conditions.

The present invention has for one of its objects to provide one or more blades in a rotor assembly having a sacrificial cutting tip which is deliberately intended to cut into an abradable casing lining during the early stages of engine running.

Accordingly, the present invention provides a turbine rotor stage for a gas turbine engine comprising a turbine casing having an intewnal- liner composed of abradable refractory material and a rotor- assembly having a multiplicity of circumferentially spaced shroudless turbine blades one or several of which each has an overall height slightly greater than the other blades and has a tip capable of cutting the turbine casing liner.

The invention will now be described by way of example only with reference to the accompanying drawings, in which: Figure 1 is a front elevation of a rotor assembly including three cutting blades, Figure 2 is a detail view illustrating the height difference of a cutting blade, and Figure 3 is a development of figure 2 illustrating a portion of the casing and liner after the establishment of the blade tip clearance.

Referring now to the drawings, figure 1 shows in schematic form a turbine rotor assembly in which a plurality of turbine blades are attached to the periphery of a rotor hub or disc 20. The blades are of two types which differ only in that blades 22 of the first type are conventional untipped and unshrouded blades, whereas blades 24 of the second type are tipped in a manner to be described in more detail below.

The blades 22, 24 are fixed to the periphery of the hub 20, for example, by means of a 'fir-tree' root fixing as indicated at 25 or by some other suitable attachment means. There is at least one blade 24 of the tipped type but, preferably there are several, typically three, which are evenly spaced apart from each other around the hub in place of the same number of untipped blades.

The tipped blades 24 are provided with a tip or coating 26 of abrasive material. In order to provide the cutting action mentioned above the abrasively tipped blades 24 have an overall height slightly greater than the untipped blades 22. With reference to figure 2 the untipped blades 22 describe a pitch circle 28 of slightly smaller diameter d than the diameter D of the corresponding pitch circle 30 described by the tipped blades 24. Preferably the total depth H of the abrasive cutting tip is greater, by an amount S, than the difference P between the height of the two types of blade 22, 24 or the radii of their respective pitch circles.

Figure 3 illustrates a detail view of the turbine blades 22, 24, similar to that shown in figure 2, in conjunction with a fragment of turbine casing denoted by reference 32. The inner surface 34 of the casing 32 is provided with an abradable layer or coating 36. This layer or coating essentially comprises ceramic material the density and hardness of which has been selected or otherwise adjusted with respect to the characteristics of the abrasive tips of the blades 24 to be abradable.

As initially deposited the abradable layer 38 extends radially inwardly as far as the surface 38. The further intermediate surface 40 lies at the depth within the layer 36 to which the abrasive tips 26 of blades 24 remove ceramic material.

The abrasive tips 26 referred to above preferably comprises crystals of hard cutting material embedded in a relatively softer but non-oxidising matrix material.

The cutting material must be necessity for the purpose it is to perform to be extremely hard and good results have been achieved using boron nitride in cubic form.

The boron nitride crystals are embedded in a layer of ceramic refractory material such as CoCrALY or MCrALY.

The crystals are not fully embedded in the matrix material but stand a little way proud of the surface, a few thousandths of an inch is sufficient, to form an abrasive cutting surface on the outermost endface, of the blade. Preferably the layer of matrix material at the blade tip is deposited to a blade height substantially equal to the blade height of the untipped blades. It is found in practice that after an initial period of operation a proportion of the abrasive cutting material is lost either by physical removal or deterioration as a result of oxidation. The matrix material itself does not oxidise and it is important that the tip layer has sufficient depth to avoid oxidisable blade material being exposed as a result of loss of abrasive material.

When the turbine section is first assembled being both cold and stationary there is a substantial clearance or gap between the end faces of all of the turbine blades and the radially inner surface of the ceramic casing liner. That is, there is no contact between the blades and the casing. This state of affairs is intended to prevail over a range of speeds and temperatures until the engine reaches a predetermined speed of rotation and/or temperature. At this speed the longer abrasively tipped blades make contact with the ceramic casing liner. Overrotation of the turbine rotor will result in the tipped blades cutting a progressively deeper groove or track in the ceramic layer, certain other combinations of speed and temperature also influence the cutting depth of the groove.

The turbine casing is not perfectly round, of course, nor always concentric with the axis of rotation of the rotor and a certain amount of ovality can be brought about by mechanical and thermal stresses. Therefore cutting occurs first at pinch-points on the casing circumference and will be to a greater depth in some regions than in others.

In order to optimise the performance of the rotor assembly it is important that the lengths 1 of each of the non abrasively tipped blades are substantially equal to each other within very close tolerances. This matching may be achieved by selective assembly where the blades are selected for length. Alternatively the blades 22 may be assembled into the disc or rotor 20, that is without the abrasively tipped blades and tip-ground to the desired radius. The tipped blades are then assembled subsequently.

Referring to figure 3 the total height of the abrasively tipped blades from disc circumference to cutting surface is given as L. The difference between height L and the lower height 1 of the non-cutting blades determines the total running clearance G of the turbine rotor measured to the bottom of the groove. However, the groove has a depth g and viewed in an axial direction the turbine blades appear to present a tip clearance gap of only (G-g).

By utilising the above mentioned method of producing tip clearances, it is possible to produce a track or groove which not only forms the rotor path but which is round and concentric with the rotor assembly. This together with the accurate length matching of the non abrasively tipped blades can improve efficiency by a significant amount. The required running clearance G for the blades may be created by overspeeding the engine to a predetermined RPM thereby allowing the outer diameter D of the abrasively tipped blades 24 to expand radially and cut the required clearance into the abradable coating 36.

Figure 1 shows for example three abrasively tipped blades 24 equidistantly spaced around the periphery of the disc 20 to aid balancing thereof. However, the invention is not limited to a particular number of tipped blades. If just one abrasively tipped blade 24 is employed then a simple counter balance weight may be applied to the diametrically opposite side of the disc 20 to achieve dynamic and static balancing. There are advantages in having only a relatively small number of tipped blades since when the abrasive material wears or oxidises and is lost the tip clearance increases only for those blades. The inherent loss of efficiency is minimised by limiting the increase in tip clearance to only the affected blades.

Claims

1. A turbine rotor stage for a gas turbine engine comprises a turbine casing having an internal liner composed of abradable refractory material and a rotor assembly having a multiplicity of circumferentially spaced shroudless turbine blades one or several of which each has an overall height slightly greater than the other blades and has a tip capable of cutting the turbine casing liner.

2. A turbine rotor as claimed in claim 1 wherein the capability of the tips of the longer blades to cut the casing lining is provided by particles of hard material upstanding from the said tip.

3. A turbine rotor as claimed in claim 2 wherein the hard cutting particles at the blade tip are composed of material which progressively oxidises and disappears in the operating environment of the turbine.

4. A turbine rotor as claimed in claim 3 wherein the particles of hard material comprises crystals of a hard substance such as boron nitrite.

5. A turbine rotor as claimed in any one of claims 2 to 4 wherein the hard particles at the blade tip are carried in a relatively softer non-oxidisable matrix material which remains.

6. A turbine rotor as claimed in claim 5 wherein the said matrix material comprises a coating of ceramic material such as CoCrALY or MCrALY.

7. A turbine rotor as claimed in any preceding claim wherein there are at least three of the longer cutting tipped blades equally spaced circumferentially around the rotor assembly.

8. A turbine rotor as claimed in any preceding claim wherein the height of the tipped blades excluding the protrusion of the cutting particles is slightly less than the height of the other blades.

9. A method of setting tip clearance in a turbine rotor stage of any preceding claim comprises the steps of: assembling the rotor assembly including the tipped and untipped blades; installing the rotor assembly within the internally lined casing, and rotating the rotor assembly at a-speed and temperature selected to produce sufficient centrifugal and thermal expansion to cause the tipped blades to cut a groove in the casing liner thereby setting a tip clearance between the other blades and the casing.

10. A method as claimed in claim 9 comprises a further step of: initially assembling the rotor assembly without the cutting tipped blades and tip grinding the remaining blades to a height slightly less than the overall height of the tipped blades.

11. A turbine rotor stage substantially as described with reference to the accompanying drawings.

12. A method of setting the tip clearance of a turbine rotor stage substantially as described with reference to the accompanying drawings.