GB2061396A

GB2061396A - Turbine blade tip clearance control

Info

Publication number: GB2061396A
Application number: GB7936915A
Authority: GB
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1979-10-24
Filing date: 1979-10-24
Publication date: 1981-05-13
Also published as: FR2467979B1; DE3038603A1; DE3038603C2; JPS5666408A; FR2467979A1; GB2061396B

Abstract

The clearance between the tips of turbine rotor blades 22 and the turbine outer casing, is controlled to match closely the thermal expansions and contractions of the turbine rotor during the whole engine cycle. A segmented liner 48, 49 is spaced from the blade tips and the liner segments are supported upstream of the rotor by the nozzle guide vanes 27, and downstream of the rotor either by guide vanes 39 of a subsequent turbine stage or by other structure. The radial displacement of the linear segments is controlled by miniature static annular members 44, 45, the thermal response of which matches that of the turbine rotor. <IMAGE>

Description

SPECIFICATION Control of clearance between turbine blade tips and casings This invention relates to gas turbine engines in which the clearance between the turbine blade tips and the engine casing or other fixed structure is required to be controlled.

On engine start up and on accelerations of the engine, the thermal growth of the turbine rotor assembly lags behind that of the turbine casing.

During steady-state running the growth of the turbine casing matches more closely that of the rotor assembly and on decelerations or-engine shut-downs the casing contracts more rapidly than the rotor assemblies.

If no means are provided to match the rates of thermal expansion and contraction of the casing and the rotor assembly, the clearance over the blade tips must be large enough to avoid rubbing under transient conditions of relative growth, but then during steady state running the clearances are too great and this gives rise to poor engine performance, loss of power and high specific fuel consumption.

Many proposals have been put forward to control the size of the clearances, these have ranged from blowing cold air from the compressor onto the turbine casing or other structure that cooperates with the blade tips to form a seal to control the gap, to redirecting hot combustion gases from the flow path through the turbine for the purposes of heating and thus delaying the contraction of the casing or other structure during engine deceleration.

In both of the above mentioned proposals, a valve and control device is necessary and the engine efficiency is compromised. By using cold air the compressor work is wasted; by using hot air the turbine efficiency is lowered. Furthermore, neither of these solutions provide adequate control of the clearances between the turbine blades and the casing at all times, that is to say, on acceleration, steady state running and on rundown.

An object of the present invention is to provide a method of controlling the clearance between the turbine blades and the turbine outer casing, or other structure with which the blades co-operate to form a seal, which is closely matched to the expansion and contractions of the turbine rotor assembly during all conditions of operation of the engine.

A further object is to achieve this by deflecting the outer casing, or other structure with which the turbine blades co-operate in response to expansion and contractions of the rotor assembly.

A further object is to divorce the seal surface co-operating with the blade tip from distortions of the engine casing due to structural loads and thermal patterns to which the said casing may be subjected.

According to the present invention a turbine for a turbo-machine, the turbine comprising a bladed rotor assembly, a segmented annular inlet guide vane assembly upstream of the rotor assembly, the segments of which are mounted for radial movement relative to an outer casing, a segmented hollow cylindrical liner spaced radially from the tips of the rotor blades, each segment of the liner being supported at a first end by a segment of the guide vane assembly and being supported at a second end by further radially movable structure located downstream of the rotor assembly, and first and second annular members the thermal response of each of which is designed to match the radial growth and contractions of the rotor assembly which occur in use, the first annular member being located concentrically with, and operable on, the segments of the guide vane assembly to move them and hence the first ends of the segments of the liner, radially when the first annular member undergoes thermal expansion or contractions in radial directions, and the second annular member being operable on the said further structure to move the further structure, and hence the second ends of the segments of the liner, radially in unison and parallel with the first ends of the segments of the liner when the second annular member undergoes thermal expansion or conctractions in radial directions.

The said further structure downstream of the rotor assembly may comprise a second segmented annular guide vane assembly such as an inlet guide vane assembly for a subsequent stage of the turbine or second turbine. The guide vane segments may be mounted so that their radially outer extremities can slide radially against radially extending abutments carried by the outer casing. In this case, the abutments provide an axial restraint against axial gas loads on the guide vane segments. Alternatively, each guide vane segment may pivotally be mounted with respect to the outer casing so that the annular member that operates on the respective guide vane segments imparts a turning moment about the pivot of each respective segment to move the liner segments in a radial direction.

An embodiment of the present invention will now be described by way of an example, with reference to the accompanying drawings in which: Figure 1 shows schematically a gas turbine aero-engine incorporating a turbine constructed according to the present invention; and Figure 2 illustrates in greater detail part of the high pressure turbine of the engine of Figure 1 constructed in accordance with the present invention.

Referring to Figure 1 there is shown a triple spool gas turbine engine 10 of the bypass type.

The engine 10 comprises, in flow series, a low pressure compressor 11 mounted in a bypass duct 12 a multi-stage intermediate pressure axial flow compressor 1 3 a high pressure axial flow compressor 14, a combustion chamber 15, a high pressure turbine 16, an intermediate pressure turbine 17, a low pressure turbine 18 and a jet pipe 19.

The high pressure turbine 1 6 is shown in greater detail in Figure 2, and comprises a turbine rotor assembly comprising an annular turbine disc 20 with a relatively massive central hub 21 and a plurality of equispaced turbine blades 22 spaced around the rim of the disc. The blades 22 are retained in the disc 20 by conventional fir-treeroot fixings. The turbine disc 20 is provided with flanges 23 by which it is secured to its respective shaft 24 to drive the HP compressor 14.

The inlet guide vane assembly 25 of the HP turbine comprises a plurality of segments 26 each comprising a plurality of spaced guide vanes 27 mounted between a fixed platform 28 and a tip shroud 9. The platforms 28 of the segments 26 are provided on their underside with two spaced flanges 29 which project radially inwards. Holes are provided in the flanges 29 to accommodate one or more pins 30 per each segment 26.

The guide vane segments 26 are retained at their innermost ends by the pins 30 which locate in elongate slots 31 in a radial flange 32 carried by the fiame tube structure 33 of the combustion chamber of the engine.

The segments 26 are retained at their outermost ends, adjacent the leading edges of the guide vanes 27, by means of hooks 34 which engage an abutment surface 35 provided on the outer casing 36 of the turbine. Axial gas loads on the guide vanes 27 are reacted by the casing 36 through the hooks 34, rotation of the segments about hooks 34 being resisted by pins 30 and annular member 44.

Downstream of the HP turbine disc 20 is located the inlet guide vane assembly 37 of the intermediate pressure turbine (the rotor of which is not shown in Figure 2). This guide vane assembly 37 is of similar construction to the guide vane assembly 26 of the HP turbine in that it comprises a plurality of segments 38 each comprising a plurality of guide vanes 39 mounted between fixed platforms 40 and tip shrouds 41.

The segments 38 guide vane assembly 37 are retained at their outermost ends adjacent the trailing edge of the guide vanes 39 by means of a flange 42 which engages an abutment surface 43 on the turbine outer casing 36. Axial gas loads on the guide vanes 39 are reacted by the outer casing 36 through the flanges 42 and abutment surfaces 43.

Both the inlet guide vane assembly 25 of the HP turbine and the inlet guide vane assembly 37 of the IP turbine are provided with an annular member 44, 45 the thermal response of each of which is designed to match that of the HP turbine rotor assembly.

In the case of the HP inlet guide vane assembly 25 the annular member 44 is secured at its periphery to all of the segments 26 by means of the pins 30 at a location beneath the platform 28 but adjacent the trailing edges of the guide vanes 27. Radial expansions and contractions of the annular member 44 push the segments 26 radially outwards and impart a turning moment to the segments about hook 34. Similarly, radial expansions and contractions of the annular member 45 push the segments 38 radially putwards and inwards, and impart a turning moment to the segments about abutment 42.

Each segment 26 of the HP guide vane assembly 25 is provided with a recess 47 at the trailing edge of the tip shroud 9 into which is located the front ends of a segmented hollow cylindrical liner 48 and an air impingement ring 49. The rear ends of the segments of the liner 48 and ring 49 locate in a recess 50 in the front edge of the tip shroud of the segments 38 of the IP guide vane assembly 37. A seal plate 51, which permits radial movements of the liner segments 48 and the impingement ring 49, is provided between the ring 49 and a seal face 52 on the turbine outer casing 36.

The impingement ring 49 may be alternately slotted from each end to provide sufficient flexibility to enable it to follow the movements of the guide vane tip shrouds 9 and 41, and 1 5 provided with a plurality of holes through which cooling air, from around the outside of the inner combustion chamber casing can flow to cool the segmented liner 48.

The annular members 44, 45 are each constructed as dummy miniature turbine discs and have a central hub 53 even though they are static.

The hub 53 provides a slugging mass to control the thermal response of the annular members 44, 45 and the members 44, 45 are made of materials which together with the shape of the members ensure that the thermally induced radial expansions and contractions of the members 44,45 match that of the turbine rotor assembly.

To control the thermal response of the annular members 44, 45 a cover 54, 55 is provided around the outside of the members 44, 45 to shield them. In the case of the member 44, the cover 54 is bolted to the structure 56 of the engine and holds pins 30 to permit rotation of segments 26 about hook 34, some axial movement must be allowed at flanges 29, for example by allowing structure 56 to slide axially.

The holes in the fixed structure 56 through which the pins 30 pass are elongate to allow for radial movements of the member 44, pins 30 and guide vane segments 26. The cover 55 is effectively carried by the IP guide vane assembly 37 via the pins and a radially slidable spigot 58. The holes in the cover 55 through which the pins 46 pass are elongate to allow the member 45, pins 46 and the segments 38 of the guide vane assembly 37 to move radially relative to the cover 55 and its supporting structure 57. The cover 55 and its supporting structure 57 constitute the static part of a labyrinth seal for the turbine rotor assembly.

In operation, the members 44 and 45 operate on their respective guide vane assemblies 25 or 37 to move the segmented liner 48 radially to maintain a suitable tip clearance between the turbine blades 22 and the liner. The front and rear ends of the segments of the liner 48 are moved radially in unison and because the thermal response of the members 44 and 45 is matched to that of the rotor assembly the gap between the blades 22 and the liner can be controlled more accurately than hitherto possible.

The movement of the guide vane assemblies 25 and 37 is achieved by pivoting the guide vane segments about the points of contact between them and the outer casing. That is to say, by rocking guide vane assemblies 25 and 37 respectively anti-clockwise and clockwise as viewed in Figure 2 about an axis normal to the plane of the drawing to increase the gap between the blades 22 and the liner 48 and vice-versa to reduce the gap.

In an alternative construction the attachments of the segments 26 and 38 to the annular members 44 and 45 are arranged to resist rocking of the segments (for example, members 44 and 45 may be duplicated, one of each pair being attached to the front and rear respectively of the segment inner platforms 28, 40). In this construction the abutments of the segments are arranged to slide radially.

In an alternative construction the rear end of the segments of the liner 48 may be carried by a supporting structure which is not an inlet guide vane assembly but which is capable of moving in the same manner as the guide vane assembly 37.

In this case, the member 45 would be secured to this supporting structure to operate on it in the same manner as described above. This alternative construction may be suitable for a single stage turbine or a turbine which is not followed by a further turbine stage.

Claims

1. A turbine for a turbo machine, the turbine comprising a bladed rotor assembly, a segmented annular inlet guide vane assembly upstream of the rotor assembly, the segments of which are mounted for radial movement relative to an outer casing, a segmented hollow cylindrical liner spaced radially from the tips of the rotor blades, each segment of the liner being supported at a first end by a segment of the guide vane assembly and being supported at a second end by further radially movable structure located downstream of the rotor assembly, and first and second annular members the thermal response of each of which is designed to match the radial growth and contractions of the rotor assembly which occur in use, the first annular member being located concentrically with, and operable on, the segments of the guide vane assembly to move them and hence the first ends of the segments of the liner, radially when the first annular member undergoes thermal expansion or contractions in radial directions, and the second annular member being operable on the said further structure to move the further structure, and hence the second ends of the segments of the liner, radially in unison and parallel with the first ends of the segments of the liner when the second annular member undergoes thermal expansion or contractions in radial directions.

2. A turbine according to claim 1 wherein the said further structure downstream of the rotor assembly comprises a second segmented annular guide vane assembly.

3. A turbine according to claim 2 wherein the second guide vane assembly constitutes an inlet guide vane assembly for a subsequent stage of the turbine.

4. A turbine according to claim 2 wherein the second guide vane assembly constitutes an inlet guide vane assembly of a second turbine.

5. A turbine assembly according to any one of the preceding claims wherein the guide vane segments are mounted so that their radially outer extremities can slide radially against radially extending abutments carried by the outer casing, which abutments provide an axial restraint against axial gas loads on the guide vane segments.

6. A turbine assembly according to any one of claims 1 to 5 wherein each guide vane segment is pivotally mounted with respect to the outer casing and the annular member that operates on the respective guide vane segments imparts a turning moment about the pivot of each respective segment to move the liner segments in a radial direction.

7. A turbine according to any one of the preceding claims wherein each guide vane segment comprises a radially inner platform and one or more guide vanes mounted on the platform, and the respective annular member is positioned to operate on the radially inboard side of the platform.

New claim filed on 7/10/80 New claim:

8. A turbine substantially as herein described with reference to the accompanying drawing.