GB2128685A

GB2128685A - Turbine overspeed limiter

Info

Publication number: GB2128685A
Application number: GB08228582A
Authority: GB
Inventors: Neil Milner Evans
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1982-10-06
Filing date: 1982-10-06
Publication date: 1984-05-02

Abstract

The mechanism comprises shrouds 28 each mounted on the outer tip of an aerofoil shaped turbine rotor blade, and static structure which includes a catch means in the form of stops 60 spaced around the periphery of the rotor. Each shroud 28 abuts adjacent shrouds 28 in planes that lie at oblique angles to the direction of rotation of the rotor 14 and has a radially outward extending projection 62 adjacent its leading corner. The projections 62 and the stops 60 are located relative to each other so that if the rotor 14 exceeds a predetermined safe speed and expands radially beyond a predetermined diameter the projections 62 engage the stops 60 and twists the shrouds 28 to open a circumferential gap between them. Simultaneously the aerofoil blades 26 are damaged to diminish their aerodynamic efficiency and decelerate the rotor. <IMAGE>

Description

SPECIFICATION Turbine overspeed limiter for turbomachines This invention relates to a mechanism for preventing a turbine rotor of a gas turbine engine rotating at an unsafe speed.

A primary requisite in the design of gas turbine engines is that a failure of any component of the engine should not jeopardise the safety of the aircraft to which the engine is fitted, no matter how remote the likeliness of such a failure may be.

This invention addresses itself specifically to the problem of a turbine overspeeding as a result of a malfunction of the engine.

During normal running, the compressor and turbine rotors run at a predetermined safe speed.

The aerodynamic forces on the blades of the turbine drive the compressor, and the aerodynamic forces on the compressor oppose the rotation of the turbine rotor. If a shaft connecting the turbine rotor to the compressor were to break, the aerodynamic loads on the turbine rotor accelerate it very rapidly (within a few milliseconds) and there is no opposition or axial constraint provided by the compressor rotor.

Consequently, the turbine rotor can accelerate to a speed at which the disc or drum, retaining the turbine blades bursts. The blades are then released and subjected to an extremely high centrifugal force which can propel them through the engine casings. It is extremely difficult to ensure that in these extreme, and unlikely conditions, all the ejected blades are contained within the engine casings. There is, therefore, a risk that one or more of the blades could damage the aircraft.

Concomitant with a rapid acceleration, the aerodynamic loads on the turbine rotor biases it rearwards with enormous force. In many designs of engine, the axial force is reacted by the thrust bearing which supports the shaft. One method of enabling one to prevent the turbine rotor overspeeding is to design the attachments of the compressor rotor to its driving turbine and to the thrust bearing supporting the shaft, so that if the shaft fails, the turbine rotor is not supported in the thrust bearing but is free to move axially to strike structure that will brake the rotor. However, simply allowing the rotor to run against a fixed structure will have no appreciable effect in slowing the rotor down because the heat generated by friction would melt the rotor and the stator vane structures and provide liquid metal lubrication of the rotor in less time than it takes for the disc to burst.

The breakage of a drive shaft is not the only possible cause of a turbine rotor overspeeding. For example in an engine such as a turbo-propellor engine employing variable pitch propellors, the load on the turbine varies according to the pitch of the propellor blades. During normal running, these variations of loads can be safely accommodated either by controlling the fuel supply to the engine or by varying the geometry of stator vanes of a compressor of the gas generator core of the engine. Even so, it is cutomary in such engines to locate the turbine rotor in thrust bearings so that axial movement of the rotor is limited and constained at all times. Accordingly, if the pitch change mechanism were to fail and the propellor blades were to assume a "feathered" attitude, or any other malfunction were to occur, the turbine rotor could accelerate above a predetermined safe speed.The turbine rotor could then burst releasing the turbine blades. In this case, because the turbine is still located axially by the thrust bearing, it is impossible to use axial movement of the turbine rotor to initiate operation of a mechanism to decelerate the rotor.

The present invention resides in the appreciation that it is possible to design a structure which makes use of the radial expansion of the rotor prior to it bursting to initiate deceleration of the rotor to safe speeds at which the blades are less likely to be ejected through the engine casings.

A separate device which forms no part of this invention is employed to retain the compressor rotor and prevent it travelling axially forwards.

An object of this invention is to provide a mechanism for preventing a turbine rotor of a gas turbine engine exceeding a predetermined speed.

The present invention, as claimed, makes use of the radial expansion of the turbine rotor prior to it bursting to twist the tip shrouds to unlock them, and also damage the aerofoil blades of the turbine rotor to diminish their aerodynamic efficiency and thereby decelerate the rotor.

The invention will now be described, by way of an example, with reference to the accompanying drawings in which, Figure 1 illustrates schematically a gas turbine engine incorporating a mechanism 25 constructed in accordance with the present invention, for preventing a turbine rotor 14 overspeeding.

Figure 2 illustrates schematically a radial cross sectional view through part of the low pressure turbine of the engine of Figure 1.

Figure 3 is an end view of the tip shrouds looking in a radially inwards direction, and, Figure 4 illustrates a radial cross sectional view through part of an alternative design of tip shroud to that shown in Figure 2.

Referring to Figure 1 there is shown a two spool gas turbine engine of the by-pass type. The engine comprises, a low pressure compressor fan 12 driven by a low pressure turbine 14, a multistage axial flow high pressure compressor 1 6 driven by a high pressure turbine 18, a combustion chamber 20 and a jet pipe 22.

The mechanism for preventing the turbine 1 4 exceeding a predetermined safe speed is shown by the reference 25. For convenience only one stage of the turbine 14 has been shown as incorporating the mechanism but it is to be understood that more than one stage of the turbine 14, or one or more stages of the turbine 18, could be provided with a similar mechanism.

Referring specifically to Figures 2 and 3 the turbine 14 is a two stage turbine which employs shrouded turbine blades 26. That is to say each of the blades 26 is provided with a shroud 28 which co-operates with the shrouds 28 of adjacent blades 26 to form a segmented ring surrounding the tips of the blades 26. Located upstream of each stage of the turbine rotor is a nozzle guide vane assembly; only the second stage nozzle guide vane assembly 30 is shown in Figure 2. The NGV assembly 30 comprises a plurality of segments each of which comprises a plurality of stator vanes 32 extending between inner and outer shrouds 34 and 36 respectively.

The stator vane assembly 30 is constrained at its outer downstream end against axial displacement by flanges 38 on the turbine outer casing 40. The radially inner ends of the NGV segments 30 are mounted on an annular fixing 41 that reacts the gas loads exerted on the vanes.

A segmented seal liner 42 extends around the circumference of the turbine rotor shroud ring 28.

Referring specifically to Figure 3 it will be seen that each tip shroud 28 is of trapezoidal shape with two parallel sides 46, 48 lying obliquely to the direction of rotation, (shown by arrow A). Each shroud 28 is shaped and constructed so that it has a leading corner 50 on the upstream side of the rotor, In otherwords, as the rotor rotates, the corner 50 of each shroud 28 leads the remainder of that shroud 28.

Each shroud 28 has two circumferential projections 54, 56 which co-operate with a seal liner 42 to form tip seals to control air leakage.

Each shroud 28 is also provided at an upstream region with a projection 58'that is angled at an oblique angle to the direction of rotation. This projection co-operates with the liner 42 to form a tip seal.

The outer casing 38 is provided with a catch means comprising a plurality of stops 60 spaced around the circumference of the rotor. The stops 60 are angled at an oblique angle to the direction of rotation, A. The stops 60 are spaced radially from the rotor but project radially towards the rotor.

During normal engine running, the tip shroud ring 28 expands and contracts radially due to centrifugal and thermally induced movements of the rotor. The rotor also moves a relatively small amount in the axial direction due to differential thermal expansion and contraction of the rotor and its driving shaft 24 relative to the outer casing 34, and also due to changing thrust loads on the turbine rotor. Nevertheless, generally, the turbine rotor is considered to be located axially by its thrust bearing. The stops 60 are, therefore, positioned axially so that during normal running they do not foul the rotor as it shuffles axially.

It can be shown that in some engines the time taken for the rotor to accelerate from its designed maximum speed to a speed at which it would burst, can be as short as 300 milliseconds. In this time it is possible for the rotor to increase its speed by 100%. At about 20% increase in speed above the designed maximum, the rotor starts to grow plastically in a radial direction beyond the normal growth due to centrifugal and thermally induced growth experienced during normal running. The stops 60 are, therefore, arranged to project inwards to a diameter at which they would not be struck by the rotor during normal running but would only be struck by the projections 58 if the rotor exceeded the designed maximum safe speed.Clearly the amount of overspeed that could be tolerated is not likely to be the same for all engines, but this may be deduced empirically for specific designs of turbine.

The stops 60 are positioned axially relative to the projections 58 so that they are on the downstream side of the projections 58 and overlap the projections 58. The projections 58 and the stops 60 are angled so that radial movement of each stop 58 into engagement with a stop 60, causes that stop 60 to push the leading edge of the shroud segment 28 in a forwards direction.

Continued rotation of the rotor causes the confronting faces of the projections 58 and stops 60 to move towards each other to increase the interference between them. This action causes the shrouds 28 to be twisted and opens up a circumferential gap between adjacent shroud segments 28 (as shown by the dotted lines). It can be shown that only a small amount of rotation of the shrouds, of the order of 3 or 4 degrees, is sufficient to unlock the hoop constraint of the shroud ring and transform the relatively stiff rotor into a floppy assembly of shrouded blades.

The tips of the blades will tend to be knocked rearwards relative to the point of attachments of the aerofoils 26 to their blade root platforms 66.

The individual shroud segments will be pushed in an arc passing under the leading edge of each immediately following shroud segment 28 and the whole bladed assembly will collapse destroying the aerodynamic efficiency of the blades.

Simultaneously, some of the blades will be twisted off by the twisting action of the shrouds.

Debris from the rotor is contained within the turbine casing and ejected rearwards through subsequent stages of the turbine to destroy the aerodynamic efficiency of downstream stages of the turbine. Consequently the rotor ceases to be driven aerodynamically and the rotor decelerates.

The turbine rotor is assembled into the casing 38 by positioning the stops 60 in the helical gaps between the ajuxtaposed projections 58 and carefully rotating the rotor helically to the correct axial position to avoid the projections fouling the stops 60.

Referring now to Figure 4 there is shown an alternative shroud 28 to that shown in Figures 2 and 3. The shroud 28 is again of trapezoidal shape and again contacts adjacent shrouds in planes tha are oblique to the direction of rotation. The projections 62 that in use engage the stops 60 do not constitute part of the tip seal as is the case of the projections 58. A separate circumferential tip seal 64 is provided at the upstream end of each shroud 28. Furthermore, the projections 62 do not extend in a circumferential direction. Instead, they extend in an axial direction and are positioned so as not to foul the liner 42 during radial excursions of the tip shrouds 28 during normal engine running.The stops 60 need not be aligned at an oblique angle to the axis of rotation, instead they extend in an axial direction and are positioned so as not to foul the tip seals. In some cases it may be necessary to omit the middle tip seal 54 and modify the shape of the liner 42 so that it only has two lands on which the seals 56, 64 run.

In a further embodiment of this invention, to cater for radial expansion with or without a significant axial movement of the rotor rearwards when the rotor overspeeds, it may be possible to combine the features of Figures 2 and 4. That is to say to provide each shroud 28 with an angled projection similar two projection 58 of Figure 2 and an axial projection similar to the projection 62 of Figure 4. In this case two sets of stops 60, or a single set of stops 60 capable of engaging either or both projections 58, 62 of each shroud may be provided.

The stops 60 may be of any desired shape that will engage the projections 58 or 62.

Claims

1. A mechanism for preventing a turbine rotor exceeding a predetermined speed wherein the rotor comprises a plurality of circumferentially spaced aerofoil blades each of which is provided at its outer tip with a shroud that co-operates with adjacent shrouds to form a segmented shroud ring around the periphery of the rotor, each shroud being shaped and constructed to abut adjacent shrouds in planes that lie at oblique angles to the direction of rotation of the rotor and each shroud having a leading corner and a radially outwardly extending projection adjacent its leading corner, and the mechanism further comprising static structure which includes a catch means radially spaced from the periphery of the rotor, the catch means being located relative to the projections so that if the turbine exceeds a predetermined speed and expands radially beyond a predetermined diameter, the projections engage the catch means and twists the shrouds to open up a circumferential gap between them and simultaneously damage the aerofoil blades to diminish their aerodynamic efficiency and thereby decelerate the turbine rotor.

2. A mechanism according to claim 1 wherein each shroud, when viewed looking radially inwards, is of a substantially trapezoidal shape with two substantially parallel sides aligned obliquely to the direction of rotation of the rotor.

3. A mechanism according to claim 1 or claim 2 wherein the catch means comprises a plurality of stops circumferentially spaced around the periphery of the rotor, and the stops extend radially inwards to a diameter at which the projections will only engage the stops if the rotor exceeding a predetermined safe speed and expands beyond a predetermined diameter.

4. A mechanism according to claim 3 wherein each projection extends in a circumferential direction at an oblique angle to the direction of rotation of the rotor and each stop also extends at a complementary oblique angle, so as to cause the shrouds to be twisted in a direction that opens up a circumferential gap between adjacent shrouds when the projections engage the stops.

5. A mechanism according to claim 3 wherein the projections extend in an axial direction and the stops also extend in an axial direction.

6. A mechanism according to any one of the preceding claims wherein the leading corner is located on the upstream side of the rotor.

7. A mechanism for preventing a turbine rotor exceeding a predetermined speed substantially as herein described with reference to any one of the accompanying drawings.