GB2310895A - Turbine shroud assembly - Google Patents

Description

SHROUD ASSEMBLY The invention relates to a gas turbine engine especially the turbine stage thereof, and more especially a power turbine stage. In particular, the invention concerns improvements in the design of a shroud liner or the turbine stage.

The invention is concerned with a gas turbine engine of the kind having a rotary stage encircled by a shroud assembly in which the shroud assembly is suspended from an engine casing or an adjacent static structure and the corresponding rotary assembly is carried by a shaft referenced to the engine casing at a remote location.

The shroud assembly may be a single ring or a plurality of part-circumferential segments. A suitable static structure supporting the shroud assembly may be a turbine nozzle annulus free to grow radially relative to the casing.

Differential thermal growth in the radial direction may result in the blades of the rotary assembly temporarily fouling the shroud assembly thereby tending to inhibit rotation of the rotary assembly. When this happens during powered rotation the rotating part wears a track or groove in the stationary shroud liner, but if it happens when the engine is stopped the engine can temporarily seize. Differential cooling of a stopped hot engine causes differential axial movement of the rotary stage relative to the shroud assembly as a result of which the rotary part may engage the shroud assembly around a circle not normally swept in a rotating engine.

According to one aspect of the invention there is provided a gas turbine engine of the kind having a rotary stage encircled by a shroud liner assembly in which the shroud liner assembly is suspended from a relatively static part of the engine structure and the corresponding rotary assembly is carried by a shaft referenced to the engine casing at a remote location and in which differential thermal growth in the radial direction may result in the shroud assembly carried by the blades of the rotary assembly temporarily fouling the shroud liner assembly thereby tending to inhibit rotation of the rotary assembly characterised in that the shroud liner assembly is formed with a clearance step in a region transitionally swept by the stopped, hot rotary stage.

Gas turbine engines exhibit inherent differential thermal expansion characteristics so that the relative axial alignment of components and radial clearances, irrespective of the effects of centrifugal forces, vary throughout the thermal cycle of an engine. Of particular concern in the present invention is an inevitable hysteresis in the thermal cycle of an engine. Different parts of the engine are heated to different temperatures, at different rates and they cool at different rates. The engine casing may be considered, for the purpose of this description, to be a cylinder which is subject to heating and cooling during a whole engine thermal cycle. Various static components are mounted within this cylinder the major one of which of present concern is the structure which carries the main engine thrust bearing, by means of which all engine shafts are effectively referenced with respect to the engine casing. The thrust bearing is usually forward of the combustor section, and may even divide the compressor, so that the turbine shaft can extend for a considerable distance aft of the thrust bearing. Differential thermal expansion is, therefore, more in evidence the greater the distance from the thrust bearing and is most pronounced in low pressure turbine and power turbine stages. Thus, in the frame of a turbine casing section over a whole engine thermal cycle from cold start, through all phases of operation, back to hot stop and subsequent cooling down a turbine rotor moves substantially both axially and radially.

Thus, in a turbine stage, for example circumferential seal fins of the turbine stage shrouds may make contact with the shroud liner after the engine has been stopped and the engine casing cools more quickly than the internal parts.

The invention will now be described in more detail with reference, by way of example, to a specific embodiment illustrated in the accompanying drawings, in which: Figure 1 shows in diagrammatic form a longitudinal section through a two-spool gas turbine engine with a power turbine, Figure 2 shows a more detailed view of the low pressure turbine and power turbine rotary stages illustrating the power turbine shroud liner arrangement, Figure 3 shows a closer view of the power turbine shroud illustrating the behaviour of the shroud seal fins relative to the shroud linear, and Figure 4 shows a view of a section of the shroud liner of Figure 3 in the direction of arrow A.

Referring in the first instance to Figure 1, there is shown a two-spool gas turbine engine having an integral power turbine stage which is mounted on a third shaft.

Briefly the gas turbine engine comprises an engine outer casing 2 having, at the left hand side of the drawing an air intake 4 and at the opposite side of the drawings a turbine exhaust aperture 6. Air enters the engine intake duct leading to a multi-stage low pressure compressor 8, then passes to a centrifical high-pressure compressor 10 and from vents into a reverse flow combustor generally indicated at 12. Hot gas from the combustor exits through a high pressure turbine stage 14 followed by a low pressure turbine stage 16. The centrifical compressor 10 and high pressure turbine 14 are mounted on a common shaft 18 to form a first, high pressure spool, and the low pressure compressor 8 and low pressure turbine 16 are mounted on a second shaft 20 to form a low pressure spool.

Exhaust gas from low pressure turbine 16 is carried by exhaust duct 22 to a power turbine stage, generally indicated at 24, which consists of two rotary stages 26,28 which are mounted on a third shaft assembly 30 which in this example is mounted concentrically with shafts 18,20 and passes through the engine to terminate in a power take-off hub 32. For example, the hub 32 in a static engine could be connected to drive the rotor of an electric generator, or in a helicopter air frame would be connected through a gearbox to drive the main rotor assembly.

Figure 2 shows the power turbine, comprising rotary stages 26,28 in more detail, together with the adjacent low pressure turbine stage 16. The LP turbine shaft 18 and power turbine shaft 30 are shown concentrically mounted by means of an intershaft bearing 34. Incidental to the invention, but for easy overhaul and repair, the power turbine shaft assembly comprises outer shaft, to which reference 30 is annotated, which is splined to an inner co-axial shaft 36 which carries the forward hub assembly 32 and may be withdrawn from the engine for maintenance without dismantling the gas generator.

The power turbine stages 26,28 comprise respectively conventional arrangements of blades 38,40 circumferentially spaced apart around the periphery of discs 42,44 which are bolted to a flange 46 carried on the right hand end of power stage shaft 30. As is well known in the art blades 38,40 are shrouded blade assemblies. In power turbine stage 26 the shrouds 48 at the radially outer ends of blades 38 abut each other in a circumferential direction to form a shroud ring.

Similarly in power turbine stage 28 all of the blades 40 carry at their radially outer ends shroud segments 50 which abut each other to form a second shroud ring.

In order to minimise gas path leakage the shroud assembly carried by the blades of the rotary assembly includes at least one seal fin which cooperates with a shroud liner assembly which encircles each power turbine stage, such arrangements are well known in the design of conventional turbine stages. So, encircling power turbine stage 26 is an annular array of shroud liner segments 52 which are placed in end to end abutment and mounted inside the engine casing 2. A similar annular array of shroud liner segments 54 encircles the second power turbine stage 28.

In power turbine stage 26 the shroud segments 48 carry twin radially upstanding fins 56 which extend in a circumferential direction. Therefore, in a fully bladed rotary assembly the shroud segments 48 abut to form a substantially continuous shroud ring and the twin fins 56 similarly form substantially continuous rings. To reduce the loss of efficiency through gas leakage around the shroud ring each shroud liner segment 52 carries a lining 58 on its inner surface which closely approaches the annular fins 56 and may, at certain parts of an engine cycle be cut by them. The lining 58 is, therefore, capable of being cut by the fins 56 and may consist of a hollow honeycomb structure, shown in the drawings, a solid abradable layer or a honeycomb structure filled with abradable material.

The second rotary stage 28 is similarly treated, although the shroud ring segments 50 are provided on their radially outer sides with two annular, radially upstanding fin segments 60,62 which are disposed respectively towards the upstream and downstream sides of the stage. Corresponding to these annular fins 60,62 the shroud liner segments 54 carry abradable lining rings 64,66 on their inner surface, but otherwise have the same purpose as the fins 56 and lining segments 58 of the first rotary stage.

For the purposes of explaining the invention all fin seals 56,60,62 and their relationships with their respective linings 58,64,66 may be considered to function in identical manner. Therefore, for the purposes of explaining the construction and operation of the invention reference will be made in Figures 3 and 4 to one fin seal and abradable lining only.

Figure 3 shows the relationship between seal fin 62 and its associated shroud lining 66 during various operational phases of an engine cycle. The initial position of the fin seal is indicated at 62a, this is its cold or build position. It will be understood that there is a build tolerance envelope for the position of the seal and this is indicated by a dashed line across the tip of the fin, and it is to be understood that similar tolerance will apply to other positions of the fin although these will not be indicated nor referred to further. When the engine is started and warms to normal operating temperatures differential thermal expansion occurs between those parts carried by the engine casing 2 and those parts carried by the rotating shafts 18,20,30 which are effectively mechanically referenced to the casing by the thrust bearing 31. Thus, after the engine has been started it is allowed to run for some period of time at an idle speed during which some thermal growth occurs but which is not indicated in the drawing, next the engine is accelerated to take off speed and substantial amounts of heat are released which warms the engine casing faster than the rotating machinery.

Consequently, relative to the shroud lining 66 the seal fin moves forward to position 62b and this is its most forward relative position. Maximum engine speed take off is not held too long and the engine is soon throttled back to a cruise setting and engine temperatures stabilise to a level where the fin has moved in a relative rearwards direction to position 62c. It should be noted that there is a radial difference between positions 62b and 62c due to the greater centrifical forces generated by higher rotational speeds. Engines may be designed with a radial tip clearance such that, at least during the early stages of engine life when the shroud liners are provided with a honeycomb or abradable lining tip rubs may occur at take-off and when high transient forces occur. This allows the seal clearance at cruise condition to be minimised, thus the clearance at cruise becomes independent of the build/manufacturing tolerance. However, it has been found when a hot engine is shut down substantial differential thermal cooling of the engine casing occurs causing it to shrink both radially and axially such that the fin seal takes up position 62d in which the seal tip can engage, or cut into, the shroud lining 66. As a result of this there is a possibility in engines towards an extremity of build tolerance for the power turbine to temporarily seize shortly after shut down. This is not a permanent condition and after a period of time during which the power turbine spool is able to cool sufficiently to match the engine casing and the radial running clearance is restored, but there can remain a period during which the power turbine is not free to rotate.

The solution offered by the present invention is to provide an additional step clearance in the appropriate region of the shroud lining 66 to accommodate the radial and axial differential thermal expansion of the power turbine stage relative to the engine casing. Therefore, in the illustrated embodiment a step 70 is provided towards the trailing edge margin of shroud lining 66 of depth sufficient to accommodate the relative and maximum radial expansion of the turbine rotary stage. The axial length of this step is sufficient to ensure that the locus of the fin tips in transitioning from hot shut down position 62d to cold position 62a does not foul the new step edge of the lining. It will be apparent from the drawing that during operational running of the engine all positions of the seal fins are forward of the cold position and it is only after hot shut down that the fin seal moves aft.

Claims (12)

1A gas turbine engine of the kind having a rotary stage encircled by a shroud liner assembly in which the shroud liner assembly is suspended from a relatively static part of the engine structure and the corresponding rotary assembly is carried by a shaft referenced to the engine casing at a remote location and in which differential thermal growth in the radial direction may result in the shroud assembly carried by the blades of the rotary assembly temporarily fouling the shroud liner assembly thereby tending to inhibit rotation of the rotary assembly characterised in that the shroud liner assembly is formed with a clearance step in a region transitionally swept by the stopped, hot rotary stage.

2 A gas turbine engine as claimed in claim 1 wherein the shroud assembly carried by the blades of the rotary assembly includes at lest one seal fin which extends in a circumferential direction and is radially upstanding to closely approach the shroud liner assembly.

3 A gas turbine engine as claimed in claim 1 or claim

2 wherein the clearance step is formed in an annular margin of the shroud liner assembly.

4 A gas turbine engine as claimed in any preceding claim wherein the clearance step is formed towards the downstream edge of the shroud liner assembly.

5 A gas turbine engine as claimed in any preceding claim wherein the shroud liner assembly comprises a plurality of part-annular shroud liner segments arranged end to end.

6 A gas turbine engine as claimed in any preceding claim wherein the clearance step is formed in an abradable layer carried on the inner annular face of the shroud liner assembly.

7 A gas turbine engine as claimed in clam 6 wherein the abradable layer comprises a honeycomb structure composed of thin metallic walls.

8 A gas turbine engine as claimed in claim 7 wherein the cells of the honeycomb structure are unfilled.

9 A gas turbine engine as claimed in claim 7 wherein the cells of the honeycomb structure are filled with an abradable ceramic material.

10 A gas turbine engine as claimed in any preceding claim wherein the rotary stage comprises a power turbine driven by a the hot turbine exhaust of a gas generator portion of the engine.

11 A gas turbine engine as claimed in claim 10 wherein the power turbine is carried on a shaft arranged concentrically with the engine shafts of the gas generator and referenced by a thrust bearing in the gas generator in respect of axial movement.

12 A gas turbine engine substantially as hereinbefore described with reference to the accompanying drawings.