A gas turbine engine including blade tip clearance control, and method

Abstract

Disclosed herein is a gas turbine including a turbine casing and a turbine rotor arranged for rotation in the turbine casing. The gas turbine further includes a stator shroud concentric with the rotor and positioned radially outward from the blade tips. The stator shroud comprises a plurality of shroud segments. A clearance control assembly, is provided, which includes a plurality of actuators that are functionally coupled to the shroud segments. Each actuator comprises: a bimetallic element, adapted to displace an actuator rod in a radial direction from the first radial position to a second radial position, through deformation of the bimetallic element caused by a temperature variation generated by a heater. The actuator rods are connected to the shroud segments and control a radial displacement thereof.

Description

A GAS TURBINE ENGINE INCLUDING BLADE TIP CLEARANCE CONTROL,AND METHODDESCRIPTIONTECHNICAL FIELD

[0001] The present disclosure is directed generally to gas turbine engines, and more particularly to gas turbine engines for industrial applications, such as mechanical drive or electric power generation. More specifically, the present disclosure relates to improvements for controlling the clearance between the rotor blade tips and the stator of in the turbine section of gas turbine engines.BACKGROUND ART

[0002] Typically, gas turbine engines include a compressor section, a combustor section and a turbine section. The compressor section compresses an inflowing stream of air. The compressed air is mixed with a fuel, usually a gaseous or a liquid fuel, in the combustor section and the air-fuel mixture is ignited to generate a hot and pressurized combustion gas. The combustion gas is expanded in the turbine section to convert gas enthalpy into mechanical power. A portion of the mechanical power generated by the turbine section is used to drive the compressor section, and the remaining mechanical power is available on an output shaft of the gas turbine engine and used to drive a load, such as an electric generator, or a dynamic compressor, for instance.

[0003] The turbine section includes a casing which houses a rotor that is supported for rotation within the casing. Rotor blades, also known as buckets, extend radially outwards from the rotor in an expansion flow path where the combustion gas flows and expands. One or more annular rows of rotor blades are usually mounted on the rotor. Upstream of each annular row of rotor blades is arranged an annular row of stator blades, to properly direct the expanding combustion gas toward the rotor blades.

[0004] A stationary annular structure, usually referred to as stator shroud, is supported in the casing and extends around each annular array of rotor blades. A spacing or gap is provided between the tips of the rotor blades and an inner surface of the respective stator shroud, to avoid rubbing between the stationary stator shroud and therotating rotor blades when the turbine is in use. The gap is referred to in the art as “blade tip clearance”, and should be minimized to reduce combustion gas leakages through the gap, as the combustion gas which escapes through the gap expands without producing useful mechanical power.

[0005] Losses due to leakages through blade tip clearance cannot be eliminated, since during transient conditions, e.g. during start-up of the gas turbine engine, or sometimes during operation under partial load conditions, the rotary parts, including the rotor with the respective rotor blades, and the stationary parts, including the casing, the stator blades and the stator shroud, expand at different rates. Usually, at start up the rotary parts expand faster than the stationary parts. The blade tip clearance must be large enough to ensure that during start-up transient, for example, the blade tips do not rub against the stator shroud. The minimum dimension of the blade tip clearance is achieved during the transient operation of the gas turbine engine, when the rotor expands faster than the stator. Such minimum blade tip clearance is usually referred to as the “pinch point”. When the gas turbine engine reaches the steady state condition, the blade tip clearance becomes larger than the pinch point. The difference between the clearance at pinch point and the clearance at steady state is usually referred to as the ’’closure function”.

[0006] Figs. 1, 2 and 3 illustrate how the blade tip clearance changes during transient operation of a gas turbine engine, specifically during a start-up transient. Fig.1 illustrates a partial sectional view of the expansion flow path FP across one annular row of rotor blades, according to a plane containing the rotation axis of the turbine. More specifically, Fig.1 illustrates a portion of a turbine casing 1, which houses a stator shroud including a shroud support 3 and shroud segments 4. Each shroud segment 4 has a radially inner surface 4.1 facing the rotor and specifically the tips 5.1 of rotor blades 5. Fig.1 illustrates a situation in which the turbine engine is shut down and cold, i.e. both the casing 1 and the rotor are at ambient temperature. The blade tip clearance is indicated Cl.

[0007] Fig.2 illustrates a condition of the turbomachine during transient start-up, when the blade tip clearance is at a minimum, due to fast rise in rotor speed which in turn causes fast rise in rotor centrifugal displacement and fast rise in thermal expansionof the rotor blades. The thermal expansion of the rotor components is faster than those of the stator components, which contributes to a reduction of the clearance. The blade tip clearance during this start-up transient is labeled C2. This is the minimum clearance achieved during transient expansion of the rotor and therefore C2 represents the pinch point clearance.

[0008] When the temperature of the outer casing 1 and of the stator shroud 4 and relevant supporting structure 3 increases, the blade tip clearance increases as well until a steady state clearance C2 is achieved, as shown in Fig.3. Due to the different rate at which the rotor parts and the stator parts expand, the steady state clearance C3 must be larger than the pinch point clearance C2. The difference (C3-C2) is the closure function.

[0009] Fig.4 shows a diagram where time is plotted on the horizontal axis and a dimension depending upon the blade tip clearance is plotted on the vertical axis. More specifically, the radial closure is plotted on the vertical axis. Point A is the starting point of the radial closure, when the turbomachine is cold and before start-up. Thus, point A represents the condition at ambient temperature and with the turbine engine in non-operating conditions. Point B represents the pinch point and C is the point at which the steady state clearance C3 is achieved. CF is the closure function.

[0010] Similar issues arise during normal shut down and hot restart of the gas turbine engine, as depicted in the diagram of Fig.5, which again reports time on the horizontal axis and the blade tip clearance on the vertical axis. At instant tO the gas turbine engine is shut down. The blade tip clearance increases sharply, which is due to several factors. One factor is the abrupt reduction of the centrifugal force, with a consequent radial inward displacement of the rotor components. Moreover, since no combustion occurs in the combustor, air at ambient temperature flows in the flowpath, which leads to a fast reduction of the thermal expansion of the rotor.

[0011] At instant tl the gas turbine engine is restarted and the sudden expansion of the rotor and faster thermal growth of rotor blades and relevant shrouds causes the blade tip clearance to reduce reaching the restart pinch point clearance shown at RSP. Gradual expansion of the outer casing brings the blade tip clearance slowly back to the steady state clearance at time t2.

[0012] As can be understood from the above description, the steady state clearance C3 cannot be reduced at will and can be substantially larger than the minimum clearance (pinch point clearance C2) needed to avoid rubbing between the blade tips 5.1 and the stator shroud 4. The steady state clearance C3 causes non-negligible losses and which negatively affect the overall efficiency of the gas turbine engine.

[0013] An object of embodiments disclosed herein is to alleviate the above-discussed drawbacks of the current art gas turbine engines.SUMMARY

[0014] According to one aspect, the present disclosure concerns a gas turbine, specifically a gas turbine engine for industrial applications, comprising a turbine casing and a turbine rotor arranged for rotation in the turbine casing about a rotation axis. The rotor comprises one or more annular rows of rotor blades. The gas turbine further includes a stator shroud concentric with the rotor and positioned radially outward from the blade tips. The stator shroud comprises a plurality of shroud segments supported by the turbine casing and aligned circumferentially around the annular row of rotor blades. Each shroud segment includes an inner surface forming a portion of a hot gas flow path within the turbine.

[0015] According to embodiments disclosed herein, the gas turbine further includes a clearance control assembly, adapted to control the clearance between the inner surfaces of the shroud segments and the tips of the rotor blades. The clearance control assembly comprises, in turn, a plurality of actuators that are functionally coupled to the shroud segments. Each actuator comprises: an actuator rod extending generally radially outwards from the respective shroud segment, a resilient feature adapted to apply a resilient force on the actuator rod and to bias the actuator rod towards a first radial position; a bimetallic element, adapted to displace the actuator rod in a radial direction from the first radial position to a second radial position, against the resilient force applied by the resilient feature, through deformation of the bimetallic element caused by a temperature variation; and a heater adapted to heat the bimetallic element.

[0016] The bimetallic element comprises a bimetallic washer, mounted coaxially with and around the actuator rod. The washer is held between a stationary stop, orretainer, forming an abutment, which is integral with the turbine casing, and a movable stop, or retainer, forming a further abutment,, integral with the actuator rod, such that a change in shape of the washer caused by a change in temperature causes a radial displacement of the actuator rod between the first radial position and a second radial position.

[0017] In some embodiments, the heater can be an induction heater. The use of an induction heater is particularly beneficial, since control thereof is easy and efficient, and the heat transfer from the heater to the bimetallic element can be particularly fast and easy to control. Other heaters are not ruled out. For instance, a heater using a hot fluid, such as a hot gas, can be used instead of (or in combination with) an induction heater.

[0018] Further embodiments and advantageous features of the gas turbine according to the present disclosure are set out in the dependent claims and are described below with reference to the attached drawings.

[0019] According to a further aspect, disclosed herein is a method for blade clearance control in a gas turbine. The method comprises the following steps: through the clearance control assembly, maintaining inner surfaces of shroud segments surrounding rotor blades of the gas turbine at a first distance from the rotation axis of the gas turbine during a transient operating condition of the turbine; and through the clearance control assembly, moving the inner surfaces of the shroud segments at a second distance from the rotation axis when the gas turbine is operating at steady state condition, the second distance being smaller than the first distance.

[0020] Further embodiments and advantageous features of the method according to the present disclosure are set out in the dependent claims and are described below with reference to the attached drawings.BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Reference is now made briefly to the accompanying drawings, in which:Figs.l to 5, already described, illustrate the current art;Fig.6 is a sectional view of a gas turbine engine, which may include blade clearance control features of the present disclosure;Figs. 7, 8, 9 and 10 show a first embodiment of a blade tip clearance control assembly and how the blade tip clearance evolves over time during a gas turbine startup transient;Figs. 11, 12, 13, 14 and 15 show a second embodiment of a clearance control assembly for blade tip clearance control and how the blade tip clearance evolves over time during a gas turbine start-up transient;Figs. 16 and 17 illustrate a diagram which shows the blade tip clearance as a function of time during start-up and hot restart transients;Figs 18 and 19 illustrates a further embodiment of a clearance control assembly for blade tip clearance control according to the present disclosure, in two operating conditions;Figs. 20 and 21 illustrate a modified embodiment of a clearance control assembly for blade tip clearance control according to the disclosure, intended to enhance the amount of actuation displacement;Figs. 20 and 21 illustrate a further modified embodiment of a clearance control assembly for blade tip clearance control according to the disclosure intended to enhance the amount of actuation force; andFigs. 22 and 23 illustrate a further embodiment of a clearance control assembly according to the present disclosure.DETAILED DESCRIPTION

[0022] Fig.6 illustrates a sectional view of an aeroderivative gas turbine engine 2 for industrial applications, e.g. for mechanical drive or electric power applications. The gas turbine engine shown in Fig.6 is only an example; those skilled in the art will understand that the gas turbine engine can take several configurations. Embodiments of the arrangement for blade tip clearance control disclosed herein can be used in a wide number of different gas turbine engines, whenever the need to control and reduce the blade tip clearance arises.

[0023] The gas turbine engine 2 of Fig.6 comprises a compressor section 12, a combustor section 14 and a turbine section 16. The casing 1 of the gas turbine engine 1 houses a rotor 17, which is supported in the casing 1 for rotation about a rotation axis A-A. The rotor 17 comprises a plurality of annular rows of rotor blades 5. Annularrows of stator vanes 21 are interspersed with the annular rows of rotor blades 5.

[0024] At least one or some annular rows of rotor blades 5 can be provided with an assembly for controlling the blade tip clearance. A first embodiment of the assembly for controlling the blade tip clearance and the operation thereof is shown in Figs. 7 to 10. These figures show, similarly to Figs. 1-3, a sectional view along a plan containing the rotation axis A-A, of the expansion flow path FP along which hot and pressurized combustion gas generated in the combustor section 14 expands generating mechanical power by interacting with the rotor blades 5. The same reference numbers used in Figs.1-3 are used in Figs. 7 to 10 to designate the same or corresponding components.

[0025] Thus, with reference to Figs. 7 to 10, the turbine casing 1 houses a stator shroud including a shroud support 3 and shroud segments 4. Each shroud segment 4 has a radially inner surface 4.1 facing the rotor and specifically the tips 5.1 of rotor blades 5.

[0026] Fig.7 illustrates a situation in which the gas turbine engine 2 is shut down and cold, i.e. both the casing 1 and the rotor are at ambient temperature. The blade tip clearance is indicated Cl. Reference number 30 designates a blade tip clearance control assembly, which for the sake of brevity is herein also referred to simply as “clearance control assembly”.

[0027] It is to be understood that for at least one, some or all annular rows of rotor blades 5 a clearance control assembly can be provided, which includes a plurality of actuators, as described in detail below, for the various shroud segments combined with the annular row of rotor blades 5. In the following description reference is made to a single actuator of the clearance control assembly, but it is to be understood that a plurality of actuators forming part of the clearance control assembly are provided around one, some or all of the annular rows of rotor blades 5, to simultaneously and synchronously control all the shroud segments belonging to the same expansion stage of the gas turbine engine.

[0028] Turning now to Fig. 7, the clearance control assembly 30 comprises a plurality of actuators 33 functionally coupled to the shroud segments 4. In some embodiments, the clearance control assembly 30 can comprise one actuator 33 for each shroudsegment 4. In other embodiments, one actuator 33 can control a plurality of sequentially arranged shroud segments 4.

[0029] In some embodiments, each actuator comprises 33 comprises an actuator rod 35 which extends generally radially outwards from the respective shroud segment 4. In the embodiment of Figs. 7 to 10 each actuator rod 35 comprises a distal end 35.1 constrained to the respective shroud segment 4, and a proximal end extending outside the casing 1 of the gas turbine engine 2.

[0030] The actuator 33 further comprises a resilient feature adapted to apply a resilient force on the actuator rod 35 and to bias the actuator rod 35 towards a first radial position. In the embodiment of Figs. 7 to 10 the resilient feature comprises a helical compression spring 37 coaxial to the actuator rod 35. The helical compression spring 37 is retained between an outer surface of the casing 1 and a Seeger ring 39 or another retainer or stop connected to the actuator rod 35 and forming an abutment along the actuation rod 35. The helical compression spring 37 and the retainer 39 are contained in a housing 43 which is fixed to the outer surface of the casing 1. The helical compression spring 37 therefore biases the actuator rod 35 radially outwardly, i.e. away of the rotor blades 5.

[0031] In addition to the helical compression spring 37, the housing 43 further contains a bimetallic element 41 that, in the embodiment shown in Figs. 7 to 10, is in the form of a washer.

[0032] In this embodiment, the bimetallic washer 41 has a central hole, through which the actuator rod 35 extends, such that the bimetallic washer 41 is adapted to slide with respect to the actuator rod 35. The bimetallic washer 41 is retained between the retainer 39 and a ceiling wall 43.1 of the housing 43. The retainer 39 forms a movable stop, integral with the actuator rod 35, i.e. integrally movable therewith in the radial direction. The ceiling wall 43.1 forms a stationary stop, integral with the casing of the turbine.

[0033] The bimetallic washer 41 is adapted to displace the actuator rod in a radial direction from the first radial position to a second radial position, against the resilient force applied by the helical compression spring 37 when the bimetallic washer 41undergoes a deformation caused by a temperature variation. For this purpose, the actuator 33 further includes a heater 47, adapted to heat the bimetallic washer 41. In the exemplary embodiment of Figs 7 to 10 the heater is an induction heater and includes a heating spool which can be embedded in the ceiling 43.1 of the housing 43.

[0034] The distal end 35.2 of the actuator rod 35 extends through the ceiling 43.1 of the housing 43 and is provided with a second Seeger ring or another retainer, or stop, 49 positioned outside the housing 43, and specifically on the exterior of the ceiling 43.1 of the housing 43.

[0035] The operation of the actuator 33 can be best understood by comparing Figs.7 and 10. In Fig.7 the bimetallic washer 41 has a planar shape. The upper and lower surfaces of the bimetallic washer 41 are flat. In Fig.10 the bimetallic washer 41 has an arcuate shape in the form of a spherical cap. The upper surface of the bimetallic washer 41 is concave and the lower surface of the bimetallic washer 41 is convex. The change of shape from planar to arcuate is obtained by heating the bimetallic washer 41. From Fig.7 to Fig. 10 the central portion of the bimetallic washer 41 is moved longitudinally along the actuator rod in radial direction towards the rotation axis A-A of the turbine. The deformation of the bimetallic washer 41 has caused a displacement of the actuator rod 35 against the biasing force of the helical compression spring 37 and a consequent displacement of the shroud segment 4 towards the tips 5.1 of the rotor blades 5.

[0036] The operation of the clearance control assembly 30 described so far will now be explained with reference to the sequence of Figs. 7 to 10.

[0037] In Fig. 7 the gas turbine engine is non-operating and at ambient temperature the blade tip clearance is shown at Cl and is such that the pinch point clearance, i.e. the minimum value of the blade tip clearance, is reached during the start-up transient, when the rotor expands at a higher rate than the stator. Fig.8 shows the situation when the pinch point clearance C2 is reached. In both conditions the actuators 33 are inactive, the bimetallic washers 41 are flat and the shroud segments are maintained at the maximum distance from the rotation axis A-A.

[0038] Fig. 9 shows the steady state condition, when the rotor and the stator of the turbine have reached their final, steady state temperature and therefore their final radialdimension. The blade tip clearance is C3, which is less than the clearance Cl in cold condition, but larger than the pinch point clearance C2. The actuators 33 are still inactive.

[0039] When the steady state condition is reached, the heater 47 can be activated to deform the bimetallic washer 41 from the planar shape of Figs. 7, 8, 9 to the arcuate shape of Fig.10. As a result of the displacement of the actuator rod 35 caused by heating the bimetallic washer 41 of each actuator 33, the shroud segments 4 approach the blade tips 5.1. The bimetallic washers 41 of each actuator 33 and the temperature to which the heater 41 brings the bimetallic washers 41 can be selected such that in the activated condition of Fig.10 and under steady state operating conditions of the turbine the blade tip clearance C4 is equal to the pinch point clearance C2.

[0040] The heater 47 remains on and maintains the bimetallic washer 41 in the curved shape during the whole period of operation of the gas turbine engine 2, such that the turbine operates with a pinch point clearance C2 and thus the leakage of expanding combustion gas along the blade tip clearance is minimized.

[0041] The clearance control assembly 30 is such that if any of the induction heaters 47 of the clearance control assembly 30 fails, the corresponding shroud segment 4 is displaced radially outwardly by the helical compression spring 37 and the blade tip clearance is increased. There is no risk of the blade tips rubbing against the shroud segment. The only consequence of the heater failure is a reduced efficiency of the gas turbine engine.

[0042] Figs. 11 to 15 illustrate a different embodiment of the clearance control assembly 30. The same reference numbers used in Figs. 7 to 10 are used in Figs. 11 to 15 to designate the same or equivalent parts, elements, or components.

[0043] The differences between Figs. 7-10 and 11-15 are mainly the following. The bimetallic washer 41 is positioned outside the housing 43, on the outer surface of the ceiling 43.1 thereof which forms a stationary stop, integral with the turbine casing. The first Seeger ring or other retainer 39, integral with the actuator rod 35 (i.e. movable therewith in the radial direction), is positioned under the helical compression spring 37, i.e., between the helical compression spring and the outer surface of the casing 1.Thus, the helical compression spring is retained between the ceiling 43.1 of the housing 43 and biases the actuator rod 35 of the relevant actuator 33 radially inwardly, i.e. towards the rotor blades 5.

[0044] A second Seeger ring or other retainer 49, forming a movable stop integral with the actuator rod 35 (i.e. movable therewith in the radial direction), is positioned outside the housing 43, near the distal end 35.2 of the actuator rod 35. However, in contrast to the embodiment of Figs. 7 to 10, the bimetallic washer 41 is positioned between the outer surface of the ceiling 43.1 and the second retainer or stop 49.

[0045] Furthermore, in contrast to the previous Figs. 7 to 10, in the embodiment of Figs. 11 to 15 the bimetallic washer 41 is mounted in a reversed condition. When the heater 47 is activated, the bimetallic washer 41 deforms with a convex surface facing radially outwardly, i.e. opposite the shroud segment 4, and with a concave surface facing radially inwardly, i.e. towards the shroud segment 4 and the rotation axis A-A of the turbine.

[0046] The sequence of Figs. 11 to 15 illustrates the operation of the control assembly 30 in this embodiment. In Fig.11 the gas turbine engine is in a cold state. The blade tip clearance is shown at Cl. As will become apparent from the following description, the blade tip clearance Cl is selected such that under steady state operating conditions the clearance reduces to the pinch point clearance C2 with the actuators 33 inactive.

[0047] Before start-up or upon start-up, the heater 47 is activated and the bimetallic washer 41 is deformed in the curved state, as shown in Fig. 12 through heating. The blade tip clearance increases from Cl (Fig.11) to Clx (Fig.12).

[0048] During the start-up transient the rotor blades 5 expands thermally at a higher rate than the stator. This causes a pinch point clearance C2 to be reached (see Fig.13), which is the minimum clearance allowable to prevent rubbing between the stator shroud and the rotor blade tips 5.1. As the stator continues to expand the blade tip clearance increases again. This is shown in Fig.14, which illustrates the situation at the beginning of the steady state condition, where the blade tip clearance is C3, greater than the pinch point clearance C2. In this situation the actuators 33 are still on, i.e. the heaters 47 maintain the curved shape of the bimetallic washers 41.

[0049] Once the steady state condition is fully reached, the heaters 47 are turned off, the bimetallic washers 41 return to the planar shape and the actuator rod 35 is displaced radially inwardly under the biasing force of the helical compression spring 37. This causes the blade tip clearance to reduce to C4, equal to C2, i.e. the pinch point clearance. The steady state operating condition is shown in Fig.15.

[0050] The embodiment of Figs. 11 to 15 has, over the embodiment of Figs. 7-10, the following advantage: During steady state operation the heaters 47 are inactive, which results in a substantial energy saving, specifically in industrial applications, where, differently from e.g. aircraft applications the gas turbines typically operate at steady state condition for most of their life span.

[0051] The effect in terms of blade tip clearance is the same in both embodiments. The difference is that in the embodiment of Figs. 7 to 10 the heaters 47 are activated during steady state operation of the gas turbine engine. Conversely, in the embodiment of Figs. 11 to 15 the heaters 47 are activated for a much shorter time, during the startup transient until the steady state operation condition is reached.

[0052] Fig. 16 illustrates a diagram where time is plotted on the horizontal axis and the radial closure is plotted on the vertical axis. The diagram shows two curves. A first curve LI is identical to the curve plotted in the diagram of Fig.4 and represents the radial closure during the start-up transient in a gas turbine engine according to the state of the art. As mentioned above, point A is the starting point, when the gas turbine engine is at ambient temperature, before start-up. After ignition of the gas turbine engine, the rotor expands faster than the stator and the radial closure reaches point B, wherefrom the radial clearance increases again to point C, where the steady state condition is reached. The blade tip clearance at point C is larger than the pinch point clearance (point B).

[0053] The second curve, labeled L2, depicts the radial closure in a gas turbine engine provided with a clearance control assembly clearance according to the present disclosure. The same curve L2 is valid for the embodiment of Figs 7-10 and for the embodiment of Figs. 11-15. Starting at point C the condition of the clearance control assembly is switched and the blade tip clearance is reduced, such that the radial closure reaches the same value as in point B, i.e. the blade tip clearance reaches the pinch pointclearance at steady state operating condition.

[0054] As described above, the two embodiments differ in that according to the embodiment of Figs. 7-10 the actuators 33 of the clearance control assembly are activated upon reaching of the steady state condition, i.e. at time tss, and remain in the active condition from that instant onwards. During the time interval tO (start-up) and tss the heaters and thus the clearance control assembly are inactive.

[0055] The opposite happens in the embodiment of Figs. 11-15, where the heaters 47 of the clearance control assembly are activated at or before tO and remain active till time tss, when the steady state condition is achieved. After tss the heaters 47 are turned off.

[0056] The different behavior of the clearance control assembly in the two embodiments is plotted under the diagram of Fig. 16, where the “ON” and “OFF" condition of the heaters for the first embodiment I (Figs. 7-10) and for the second embodiment II are shown.

[0057] Fig. 17 illustrates the same diagram of Fig.16 but during a normal shutdown and hot restart, i.e. in a situation where the gas turbine engine 2 is shut down for a short period of time, e.g. 1-3 hours, and restarted before reaching ambient temperature. The diagram of Fig. 17 is therefore similar to Fig.5 described above. Similar to Fig. 16, in Fig. 17 two curves are plotted, which are again labeled LI and L2. Curve LI is identical to the curve plotted in Fig.5 and refers to a gas turbine engine according to the state of the art.

[0058] At time instant tO the gas turbine engine 2 is shut down. The radial closure decreases abruptly, i.e., the blade tip clearance increases. Since the starting point at time tO is different, the peak value of the radial closure reached upon shutdown in the two cases (curve LI and curve L2 is different). From the peak value, the radial closure increases gradually (as the natural cooling down of the turbine casing 1 is faster than the cooling down of the rotor 17) till restart of the gas turbine engine 2 at time tl. Following hot restart (i.e., restart before the entire turbine cools down to ambient temperature), the pinch-point clearance is reached at point B and the steady state clearance is reached again at point C. From point C, the two curves LI and L2 diverge. In a gasturbine engine of the prior art, the radial closure remains at the same value reached in point C. By using the clearance control assembly of the disclosure, at time t2 the clearance control assembly is switched to reduce the blade tip clearance to the pinch point value.

[0059] As in the case of start-up, also in the case of a hot restart the two embodiments (Figs. 7-10 and Figs 11-15) operate in different ways.

[0060] In the first embodiment (line I at the bottom of Fig.17), the heaters 47 are switched on till tl, when the hot restart begins. During hot restart the heaters 47 remain inactive and the bimetallic washers 41 are in the non-deformed condition. At time t2, i.e at steady state condition of the turbine, the heaters are activated again to bring the bimetallic washers 41 back in the thermally deformed condition and reduce the clearance to reach the pinch-point clearance.

[0061] In the second embodiment (line II at the bottom of Fig.17), the heaters remain turned-off till tl, when the hot restart begins. During hot restart the heaters 47 are active and heat the bimetallic washers 41 to bring them in the thermally deformed condition and maintain such condition till time t2, i.e, at steady state condition of the turbine. At t2 the heaters are de-activated again to bring the bimetallic washers 41 back in the un-deformed condition and reduce the clearance to the pinch-point clearance.

[0062] A further embodiment of the clearance control assembly is shown in Figs. 18 and 19. The same or equivalent elements, parts or components as shown in Figs. 7-10 are identified by the same reference numbers.

[0063] The main difference between the first embodiment of Figs. 7-10 and the third embodiment of Figs. 18 and 19 is the following. In Figs. 18 and 19 the helical compression spring 37 is arranged outside the housing 43. More specifically, the helical compression spring 37 is retained between the outer surface of the ceiling 43.1 of the housing and a Seeger ring or other retainer 40 mounted on the actuator rod 35, near the distal end 35.2 thereof. The bimetallic washer 41 is slidingly mounted on the actuator rod 35, as in the previous embodiments, and is positioned between the inner surface of the ceiling 43.1 and a Seeger ring or other retainer 39. A third Seeger ring or other retainer 49 is positioned between the retainer 39 and the outer surface of thecasing 1 of the gas turbine engine 2.

[0064] When the heater 47 is inactive, as shown in Fig.18, the bimetallic washer 41 is in the flat shape and the actuator rod 35 is maintained in a position distanced from the blade tips 5.1 by the helical compression spring 37 which biases the actuator rod pushing it radially outwardly. The retainer 39 defines the axial position of the actuator rod 35.

[0065] When the heater 47 is activated, the heat generated by the heater 47 thermally deforms the bimetallic washer 41 in the curved shape as shown in Fig. 19. The deformation of the bimetallic washer 41 pushes the actuator rod 35 radially inwardly with a movement parallel to the axis of the actuator rod 35. The pushing force is transmitted by the bimetallic washer 41 to the actuator rod 35 through the retainer 39. The displacement of the actuator rod 35 is defined by the retainer 49, which defines the radially innermost position of the actuator rod 35 shown in Fig. 19.

[0066] The operation of the clearance control assembly in the embodiment of Figs.18 and 19 is the same as the operation of the embodiment shown in Figs 7-10.

[0067] Modified embodiments of the actuator 33 are shown in Figs.20 and 21. In Fig.20 the bimetallic element, still labeled 41, comprises two bimetallic washers 41 A, 41B. The two bimetallic washers are mounted in such a way that the thermal deformation thereof causes the bimetallic washer 41 A to deform with a concavity facing the rotation axis and the bimetallic washer 4 IB to deform with a convexity facing the rotation axis A-A of the gas turbine engine 1. In this way, the displacement of the actuator rod 35 is double the displacement which can be obtained with a single bimetallic washer 41 as described in the previous embodiments.

[0068] A further modified embodiment is shown in Figs. 22 and 23. The difference with respect to the embodiment of Figs 20 and 21 is that the washers 41 A, 41B are mounted such that in the thermally deformed condition, the curved washers have both a convexity oriented radially inwardly, i.e. towards the rotation axis A-A. The displacement of the actuator rod 35 which is obtained in this embodiment is the same as in Figs. 7-10, but the thrust applied by the bimetallic washers 41 A, 4 IB is double.

[0069] A variety of different arrangements of bimetallic washers or other bimetallicelements in various configurations can be foreseen for all three embodiments described above and shown in Figs. 7-10, 11-15 and 18-19, respectively.

[0070] Exemplary embodiments have been disclosed above and illustrated in the accompanying drawings. It will be understood by those skilled in the art that various changes, omissions and additions may be made to that which is specifically disclosed herein without departing from the scope of the invention as defined in the following claims.

Claims

CLAIMS1. A gas turbine comprising:a turbine casing; a turbine rotor arranged for rotation in the turbine casing about a rotation axis; said rotor comprising at least one annular row of rotor blades, each rotor blade having an airfoil extending from a radially inner blade base to a radially outer blade tip; a stator shroud concentric with the rotor and positioned radially outward from the blade tips; wherein the stator shroud comprises a plurality of shroud segments supported by the turbine casing and aligned circumferentially around the annular row of rotor blades; wherein each shroud segment includes an inner surface forming a portion of a hot gas flow path within the turbine; anda clearance control assembly, adapted to control the clearance between the inner surfaces of the shroud segments and the tips of the rotor blades;wherein the clearance control assembly comprises a plurality of actuators functionally coupled to said shroud segments; and wherein each actuator comprises: - an actuator rod extending generally radially outwards from the respective shroud segment;- a resilient feature adapted to apply a resilient force on the actuator rod and to bias the actuator rod towards a first radial position.- a bimetallic element, adapted to displace the actuator rod in a radial direction from the first radial position to a second radial position, against the resilient force applied by the resilient feature, through deformation of the bimetallic element caused by a temperature variation; and- a heater adapted to heat the bimetallic element;wherein the bimetallic element comprises a bimetallic washer, mounted coaxially with and around the actuator rod; and wherein the washer is held between a stationary stop, integral with the turbine casing, and a movable stop, integral with the actuator rod, such that a change in shape of the washer caused by a change in temperature causes a radial displacement of the actuator rod between the first radial position and a second radial position.

2. The gas turbine of claim 1, wherein the heater comprises an induction heater.

3. The gas turbine of claim 1 or 2, wherein the second radial position is closer to the rotation axis than the first radial position.

4. The gas turbine of claim 1 or 2, wherein the first radial position is closer to the rotation axis than the second radial position.

5. The gas turbine of claim 3 or 4, wherein the bimetallic element is configured and arranged such that heating of the bimetallic element causes a deformation thereof which displaces the actuator rod from the first radial position toward the second radial position against the resilient force applied by the resilient feature on the actuator rod.

6. The gas turbine of claim 3, wherein the clearance control assembly is adapted to maintain the bimetallic components in a heated and thermally deformed configuration when the gas turbine is operating at steady state condition, and to maintain the bimetallic components in an un-heated and thermally non-deformed configuration during transient operating conditions at cold start-up or hot restart of the gas turbine.

7. The gas turbine of claim 4, wherein the clearance control assembly is adapted to maintain the bimetallic components in an un-heated and thermally nondeformed configuration when the gas turbine is operating at steady state condition, and to maintain the bimetallic components in a heated and thermally deformed configuration during transient operating conditions at cold start-up or hot restart of the gas turbine.

8. The gas turbine of any one of the preceding claims, wherein each actuator rod extends through the turbine casing and comprises a distal portion located outside the turbine casing; and wherein the bimetallic element and the heater are located outside the turbine casing.

9. The gas turbine of any preceding claim, wherein: a distal portion of the actuator rod extends through a ceiling of a housing positioned outside the turbine casing; the ceiling forms the stationary stop; the bimetallic element and the resilient feature are positioned in the housing; and the movable stop is located in the housing between the bimetallic washer and the resilient feature.

10. The gas turbine of an one of claims 1 to 8, wherein; a distal portion of the actuator rod extends through a ceiling of a housing positioned outside the turbine casing; the ceiling forms the stationary stop;; the movable stop is positioned outside the ceiling; the bimetallic element is located outside the ceiling, between the movable stop and an outer surface of the ceiling; and the resilient feature is positioned in the housing between an inner surface of the ceiling and a further movable stop integral with the actuator rod.

11. The gas turbine of any one of claims 1 to 8, wherein: a distal portion of the actuator rod extends through a ceiling of a housing positioned outside the turbine casing; the ceiling forms the stationary stop; the bimetallic element and the movable stop are located in the housing, the bimetallic element being positioned between the ceiling and the movable stop; the resilient feature is positioned outside the housing and retained between an outer surface of the ceiling and a second movable stop integral with the actuator rod; and wherein a third movable stop integral with the actuator rod is located in the housing, between the movable stop and the turbine casing.

12. The gas turbine of any preceding claim, wherein the heater is contained in the housing.

13. The gas turbine of any one of the preceding claims, wherein the resilient feature comprises a helical spring coaxial with the actuator rod.

14. A method for blade clearance control in a gas turbine according to any one of the preceding claims, the method comprising the following steps:- through the clearance control assembly, maintaining the inner surfaces of the shroud segments at a first distance from the rotation axis during a transient operating condition of the turbine; and- through the clearance control assembly, moving the inner surfaces of the shroud segments at a second distance from the rotation axis when the gas turbine is operating at steady state condition, the second distance being smaller than the first distance.

15. The method of claim 14, wherein the step of maintaining the inner surfaces of the shroud segments at the second distance from the rotation axis during-19-the steady state condition comprises the step of heating a plurality of bimetallic elements and maintaining the bimetallic elements in a deformed configuration, in which the bimetallic elements maintain the shroud segments in the second position against a biasing force applied to the shroud segments by a plurality of biasing features; wherein during the transient operating condition the biasing features maintain the shroud segments in the first position, while the bimetallic elements are in a non-deformed configuration.

16. The method of claim 14, wherein the step of maintaining the inner surface of the shroud segments at the second distance from the rotation axis during the steady state condition comprises the step of biasing the shroud segments towards the second position through a plurality of biasing features; and wherein during the transient operating condition the shroud segments are displaced in the first position against the biasing features by heating the bimetallic elements and maintaining the bimetallic elements in a deformed configuration.

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