GB2614760A

GB2614760A - Turbine for gas turbine engine

Info

Publication number: GB2614760A
Application number: GB2202928.4A
Authority: GB
Inventors: Bhaumik Soumyik
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2022-01-13
Filing date: 2022-03-03
Publication date: 2023-07-19
Also published as: GB202202928D0

Abstract

Disclosed is a turbine 16 for a gas turbine engine 10 including a stator vane 102 having a vane aerofoil portion 104, having a radially outermost and innermost ends 106, 108 with respect to an axis of rotation X-X’, the stator vane also includes radially inner and outer platforms 112, 118 coupled to the vane aerofoil portion, with the radially outer platforms having trailing edges 114, 120 radially proximal to and downstream of the radial ends of the vane edge, where the outer platform trailing edges are disposed radially either: outwards of a first imaginary line 172, where the first imaginary line 172, extends between the radially outer end of the vane edge and a radially outer end of the subsequent blade edge 148; or inwards of a second imaginary line 174 which extends between the radially inner end of the vane edge and the radially inner end of the subsequent blade edge 146, and the blade includes tip seals which move between contacting and not contacting the casing 162 depending on the running condition of the engine.

Description

TURBINE FOR GAS TURBINE ENGINE

FIELD OF THE DISCLOSURE

The present disclosure relates to a turbine, and in particular to turbines associated 5 with a gas turbine engine The present disclosure further relates to a gas turbine engine for an aircraft.

BACKGROUND

A turbine, such as, a high pressure turbine, an intermediate pressure turbine, or a low pressure turbine, associated with a gas turbine engine includes a rotor having multiple rotor blades. Further, the turbine also includes a stator including a multiple guide vanes. Hot gases flowing along a rotational axis of the turbine from the rotor blades towards the guide vanes may cause the rotor blades to rotate about the rotational axis. Typically, a cavity may be defined between the rotor blades and the guide vanes. A portion of the hot gases flowing through the cavity may be referred to as a cavity flow, whereas a portion of the hot gases flowing along a top surface of the rotor blades may be referred to as a main flow.

During an operation of the gas turbine engine, the cavity flow may move at a relatively slower velocity as compared to the main flow. In some situations, a high circumferential velocity difference between the cavity flow and the main flow may generate a circumferential shear-layer between the cavity flow and the main flow. The circumferential shear-layer may lead to a high circumferential pressure differential at an interface of the cavity flow and the main flow. Such a pressure differential may cause ingestion of the main flow into the cavity. Accordingly, the main flow may interact with the cavity flow. Specifically, the slower moving cavity flow may mix with the faster moving main flow, which may disturb the main flow and may also cause secondary losses in the turbine. The secondary losses may be caused, in part, due to a relatively lower aspect ratio design of the rotor blades.

Further, the ingestion of the main flow into the cavity may also increase a temperature of the rotor blades at a leading edge of a tip shroud of the corresponding rotor blades.

SUMMARY

In a first aspect, there is provided a turbine for a gas turbine engine having a rotational axis and a circumferential direction with respect to the rotational axis. The turbine includes a stator vane. The stator vane includes a vane aerofoil portion including a radially inner vane edge extending at least in the circumferential direction. The vane aerofoil portion also includes a radially outer vane edge disposed radially outwards of the radially inner vane edge with respect to the rotational axis and extending at least in the circumferential direction. The vane aerofoil portion further includes a vane aerofoil trailing edge extending from the radially inner vane edge to the radially outer vane edge. The stator vane also includes a radially outer platform coupled to the vane aerofoil portion. The radially outer platform includes an outer platform trailing edge radially proximal to and downstream of the radially outer vane edge. The outer platform trailing edge extends at least in the circumferential direction and is a most downstream edge of the radially outer platform. The stator vane further includes a radially inner platform coupled to the vane aerofoil portion opposite to the radially outer platform. The radially inner platform includes an inner platform trailing edge radially proximal to and downstream of the radially inner vane edge. The inner platform trailing edge extends at least in the circumferential direction and is a most downstream edge of the radially inner platform. The turbine further includes a rotor blade disposed downstream of the stator vane and configured to receive a fluid flow from the stator vane. The rotor blade includes a blade aerofoil portion including a radially inner blade edge extending at least in the circumferential direction. The blade aerofoil portion further includes a radially outer blade edge disposed radially outwards of the radially inner blade edge with respect to the rotational axis and extending at least in the circumferential direction. The blade aerofoil portion further includes a blade aerofoil leading edge extending from the radially inner blade edge to the radially outer blade edge. The rotor blade also includes a blade tip shroud coupled to the blade aerofoil portion and extending at least radially outwards from the radially outer blade edge of the blade aerofoil portion. The turbine further includes a casing extending at least in the circumferential direction and at least partially enclosing the stator vane and the rotor blade. During a cold condition of the gas turbine engine where at least one of a current engine temperature is lesser than 70% of a maximum rated take-off engine temperature and a current turbine shaft speed is lesser than 80% of a design turbine shaft speed, the blade tip shroud is disposed at a clearance from the casing. Further, during a hot running condition of the gas turbine engine where at least one of the current engine temperature is greater than 70% of the maximum rated take-off engine temperature and the current turbine shaft speed is greater than 80% of the design turbine shaft speed, the blade tip shroud engages the casing. Furthermore, a first imaginary line extends between the radially outer vane edge and the radially outer blade edge. The first imaginary line defines a minimum distance between the radially outer vane edge and the radially outer blade edge.

Moreover, a second imaginary line extends between the radially inner vane edge and the radially inner blade edge. The second imaginary line defines a minimum distance between the radially inner vane edge and the radially inner blade edge. During both the cold idle condition and the hot running condition of the gas turbine engine, the outer platform trailing edge is disposed radially outwards of the first imaginary line with respect to the rotational axis and/or the inner platform trailing edge is disposed radially inwards of the second imaginary line with respect to the rotational axis.

As the outer platform trailing edge may be disposed radially outwards of the first imaginary line with respect to the rotational axis, a main flow of hot gases having a high whirl velocity may impart a whirl velocity to a cavity flow of the hot gases. This phenomenon may reduce a velocity difference between the main flow and the cavity flow, which may in turn reduce secondary losses in the turbine.

Further, as the inner platform trailing edge may be disposed radially inwards of the second imaginary line with respect to the rotational axis, the main flow of hot gases having the high whirl velocity may impart a whirl velocity to the cavity flow of the hot gases. As a result, the velocity difference between the main flow and the cavity flow may reduce, thereby reducing secondary losses in the turbine.

Additionally, an improved design of the stator vane as described herein may also increase an efficiency of the turbine.

In some embodiments, a third imaginary line extends between the radially outer vane edge and the outer platform trailing edge. The third imaginary line defines a minimum distance between the radially outer vane edge and the outer platform trailing edge. The third imaginary line is inclined to the first imaginary line by a first inclination angle.

In some embodiments, the first inclination angle is from about 5 degrees to about 20 degrees. An increase in the first inclination angle may cause the main flow to impart the whirl velocity to the cavity flow.

In some embodiments, a fourth imaginary line extends between the radially inner vane edge and the inner platform trailing edge. The fourth imaginary line defines a minimum distance between the radially inner vane edge and the inner platform trailing edge. The fourth imaginary line is inclined to the second imaginary line by a second inclination angle.

In some embodiments, the second inclination angle is from about 5 degrees to about 20 degrees. An increase in the second inclination angle may cause the main flow to impart the whirl velocity to the cavity flow.

In some embodiments, the radially outer platform includes an outer extending 20 portion extending both downstream and radially outwards at least from the radially outer vane edge. The outer extending portion includes the outer platform trailing edge.

In some embodiments, the outer extending portion is substantially planar.

In some embodiments, the outer extending portion is at least partially curved away from the radially outer vane edge.

In some embodiments, the outer extending portion further extends both 30 downstream and radially outwards from the vane aerofoil portion upstream of the radially outer vane edge.

In some embodiments, the radially inner platform includes an inner extending portion extending both downstream and radially inwards at least from the radially inner vane edge. The inner extending portion includes the inner platform trailing edge.

In some embodiments, the inner extending portion is substantially planar.

In some embodiments, the inner extending portion is at least partially curved away from the radially inner vane edge.

In some embodiments, the inner extending portion further extends both 10 downstream and radially inwards from the vane aerofoil portion upstream of the radially inner vane edge.

In some embodiments, the blade tip shroud includes a tip shroud leading edge radially proximal to and upstream of the radially outer blade edge. The tip shroud leading edge extends at least in the circumferential direction and is a most upstream edge of the blade tip shroud. Furthermore, an improvement in the design of the stator vane as described herein may cause a reduction in a temperature of the blade tip shroud, and more specifically, the tip shroud leading edge, thereby improving an operational lifetime of the rotor blade. Moreover, the reduction in the temperature of the blade tip shroud may also reduce an amount of cooling air that may be otherwise required to cool the blade tip shroud.

In some embodiments, the tip shroud leading edge intersects the first imaginary line.

In some embodiments, the tip shroud leading edge is disposed radially outwards of the first imaginary line with respect to the rotational axis. When the tip shroud leading edge may be disposed radially outwards of the first imaginary line with respect to the rotational axis, the main flow of hot gases having the high whirl velocity may impart the whirl velocity to the cavity flow of the hot gases. As a result, the velocity difference between the main flow and the cavity flow may reduce, thereby reducing secondary losses in the turbine.

In some embodiments, the blade tip shroud further includes a shroud extending portion extending both upstream and radially outwards from the radially outer blade edge towards the radially outer vane edge The shroud extending portion includes the tip shroud leading edge.

In a second aspect, there is provided a gas turbine engine for an aircraft. The gas turbine engine includes the turbine of the first aspect.

As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.

The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with reference to the Figures, in which: Figure 1 is a sectional side view of a gas turbine engine, according to an 25 embodiment of the present disclosure; Figure 2 is a schematic view illustrating a stator vane and a rotor blade associated with a turbine of the gas turbine engine, according to an embodiment of the present disclosure; Figure 3 is a schematic view illustrating a stator vane and the rotor blade 30 associated with the turbine, according to another embodiment of the present disclosure; Figure 4 is a schematic view illustrating a stator vane and the rotor blade associated with the turbine, according to yet another embodiment of the present disclosure; Figure 5 is a schematic view illustrating a stator vane and the rotor blade associated with the turbine, according to an embodiment of the present disclosure; Figure 6 is a schematic view illustrating a stator vane and the rotor blade associated with the turbine, according to another embodiment of the present

disclosure;

Figure 7 is a schematic view illustrating the stator vane and a rotor blade associated with the turbine, according to an embodiment of the present disclosure; Figure 8 is a schematic view illustrating the stator vane and the rotor blade associated with the turbine, according to another embodiment of the present

disclosure;

Figure 9 is a schematic view illustrating the stator vane and the rotor blade associated with the turbine, according to yet another embodiment of the present disclosure; Figure 10 is a schematic view illustrating the stator vane and the rotor blade associated with the turbine, according to an embodiment of the present disclosure; Figure 11 is a plot illustrating a variation in an efficiency of the turbine as per a variation in a first inclination angle defined by the stator vane of the turbine, according to an embodiment of the present disclosure; Figure 12 is a plot illustrating a variation in an ingestion of a fluid flow into a cavity 20 as per the variation in the first inclination angle by the stator vane of the turbine, according to an embodiment of the present disclosure; Figure 13 is an exemplary schematic perspective of a stator vane of a turbine, according to an embodiment of the present disclosure; and Figures 14A to 14D are schematic views illustrating variations in a fluid flow 25 through the turbine of Figure 13 as per the variation in a first inclination angle defined by the stator vane of the turbine, according to an embodiment of the present disclosure.

DETAILED DESCRIPTION

Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art.

Figure 1 shows a sectional view of a gas turbine engine 10 for an aircraft. The gas turbine engine 10 has a rotational axis X-X'. The rotational axis X-X' coincides with a longitudinal centre line 101 of the gas turbine engine 10.

In the following disclosure, the following definitions are adopted. The terms "upstream" and "downstream" are considered to be relative to an air flow through the gas turbine engine 10. The terms "axial" and "axially" are considered to relate to the direction of the rotational axis X-X' of the gas turbine engine 10.

The gas turbine engine 10 includes, in axial flow series, an intake 11, a fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, an intermediate pressure turbine 17, a low-pressure turbine 18 and an engine core exhaust nozzle 19. A nacelle 21 generally surrounds the gas turbine engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.

During operation, air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high-pressure compressor 14 where further compression takes place.

The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate, and low-pressure turbines 16, 17, 18 before being exhausted through the engine core exhaust nozzle 19 to provide additional propulsive thrust. The high, intermediate, and low-pressure turbines 16, 17, 18 respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.

In some embodiments, the gas turbine engine 10 is used in an aircraft. In some embodiments, the gas turbine engine 10 is an ultra-high bypass ratio (UHBPR) engine.

The nacelle 21 further includes an intake lip 31 disposed at an upstream end 32 of the nacelle 21, a fan casing 33 downstream of the intake lip 31, a diffuser 34 disposed between the upstream end 32 and the fan casing 33, and an engine casing 35 downstream of the intake lip 31. The fan 12 is received within the fan casing 33. An engine core 36 of the gas turbine engine 10 including the intermediate pressure compressor 13, the high-pressure compressor 14, the combustion equipment 15, the high-pressure turbine 16, the intermediate pressure turbine 17, the low-pressure turbine 18 and the engine core exhaust nozzle 19 is received within the nacelle 21. Specifically, the engine core 36 is received within the engine casing 35. The nacelle 21 further includes an exhaust 37 disposed at a downstream end 38 of the nacelle 21. The exhaust 37 may be a part of the engine casing 35. The exhaust 37 may at least partly define the engine core exhaust nozzle 19.

The nacelle 21 for the gas turbine engine 10 may be typically designed by manipulating a plurality of nacelle parameters. The nacelle parameters may be dependent on the type of engine the nacelle 21 surrounds, as well as considerations for integration of engine ancillaries, such as a thrust reversal unit (TRU).

Further, the gas turbine engine 10 may operate in a cold idle condition where at least one of a current engine temperature is lesser than 70% of a maximum rated take-off engine temperature and a current turbine shaft speed is lesser than 80% of a design turbine shaft speed. Furthermore, the gas turbine engine 10 may operate in a hot running condition where at least one of the current engine temperature is greater than 70% of the maximum rated take-off engine temperature and the current turbine shaft speed is greater than 80% of the design turbine shaft speed.

Figure 2 illustrates a schematic view of a portion of the turbine 16 of the gas turbine engine 10. Although illustrated as the high-pressure turbine 16, it may be contemplated that the teachings of this disclosure may be applied to the intermediate and/or low-pressure turbines 17, 18, without any limitations. The turbine 16 has the rotational axis X-X' (also shown in Figure 1) and a circumferential direction C with respect to the rotational axis X-X'.

The turbine 16 includes a stator vane 102. The stator vane 102 includes a vane aerofoil portion 104 including a radially inner vane edge 106 extending at least in the circumferential direction C. The vane aerofoil portion 104 also includes a radially outer vane edge 108 disposed radially outwards of the radially inner vane edge 106 with respect to the rotational axis X-X' and extending at least in the circumferential direction C. The vane aerofoil portion 104 further includes a vane aerofoil trailing edge 110 extending from the radially inner vane edge 106 to the radially outer vane edge 108.

The stator vane 102 also includes a radially outer platform 112 coupled to the vane aerofoil portion 104. The radially outer platform 112 includes an outer platform trailing edge 114 radially proximal to and downstream of the radially outer vane edge 108. The outer platform trailing edge 114 extends at least in the circumferential direction C. Further, the outer platform trailing edge 114 is a most downstream edge of the radially outer platform 112.

In some embodiments, the radially outer platform 112 may include an outer extending portion 116 extending both downstream and radially outwards at least from the radially outer vane edge 108. In some embodiments, the outer extending portion 116 may include the outer platform trailing edge 114. In the illustrated embodiment of Figure 2, the outer extending portion 116 may be substantially planar. In other words, the outer extending portion 116 may be generally rectangular in shape.

The stator vane 102 further includes a radially inner platform 118 coupled to the vane aerofoil portion 104 opposite to the radially outer platform 112. The radially inner platform 118 includes an inner platform trailing edge 120 radially proximal to and downstream of the radially inner vane edge 106. The inner platform trailing edge 120 extends at least in the circumferential direction C. Further, the inner platform trailing edge 120 is a most downstream edge of the radially inner platform 118.

In some embodiments, the radially inner platform 118 may further include an inner extending portion 122 extending both downstream and radially inwards at least from the radially inner vane edge 106. In some embodiments, the inner extending portion 122 may include the inner platform trailing edge 120. In the illustrated embodiment of Figure 2, the inner extending portion 122 may be substantially planar. In other words, the inner extending portion 122 may be generally rectangular in shape.

Further, the turbine 16 also includes a rotor blade 142. The rotor blade 142 is disposed downstream of the stator vane 102. Further, the rotor blade 142 is configured to receive a fluid flow from the stator vane 102. The rotor blade 142 includes a blade aerofoil portion 144. The blade aerofoil portion 144 includes a radially inner blade edge 146 extending at least in the circumferential direction C. The blade aerofoil portion 144 also includes a radially outer blade edge 148 disposed radially outwards of the radially inner blade edge 146 with respect to the rotational axis X-X' and extending at least in the circumferential direction C. The blade aerofoil portion 144 further includes a blade aerofoil leading edge 150 extending from the radially inner blade edge 146 to the radially outer blade edge 148.

The rotor blade 142 also includes a blade tip shroud 154 coupled to the blade aerofoil portion 144. The blade tip shroud 154 extends at least radially outwards from the radially outer blade edge 148 of the blade aerofoil portion 144. In some embodiments, the blade tip shroud 154 may include a tip shroud leading edge 156 radially proximal to and upstream of the radially outer blade edge 148. In some embodiments, the tip shroud leading edge 156 may extend at least in the circumferential direction C and is a most upstream edge of the blade tip shroud 154.

In some embodiments, the blade tip shroud 154 may further include a shroud extending portion 158 extending both upstream and radially outwards from the radially outer blade edge 148 towards the radially outer vane edge 108. In some embodiments, the shroud extending portion 158 may include the tip shroud leading edge 156. The shroud extending portion 158 maybe substantially planar. The shroud extending portion 158 may be generally rectangular in shape.

The turbine 16 further includes a casing 160 extending at least in the circumferential direction C and at least partially enclosing the stator vane 102 and the rotor blade 142. In some embodiments, the casing 160 may include a casing tip portion 162 disposed proximate to the radially outer platform 112 of the stator vane 102 and the radially outer blade edge 148 of the rotor blade 142. The casing 160 may define a cavity 152. Specifically, the cavity 152 may be disposed proximate to and radially outwards of each of the vane aerofoil trailing edge 110 and the blade aerofoil leading edge 150. The casing 160 may further include a casing hub portion 164 disposed proximate to the radially inner platform 118 of the stator vane 102 and the radially inner blade edge 146 of the rotor blade 142. The casing hub portion 164 may be further disposed towards the rotational axis X-X'.

Further, during a cold idle condition of the gas turbine engine 10 where at least one of the current engine temperature is lesser than 70% of the maximum rated take-off engine temperature and the current turbine shaft speed is lesser than 80% of the design turbine shaft speed, the blade tip shroud 154 is disposed at a clearance Cl from the casing 160.

Furthermore, during a hot running condition of the gas turbine engine 10 where at least one of the current engine temperature is greater than 70% of the maximum rated take-off engine temperature and the current turbine shaft speed is greater than 80% of the design turbine shaft speed, the blade tip shroud 154 engages the casing 160.

Further, a first imaginary line 172 extends between the radially outer vane edge 108 and the radially outer blade edge 148. The first imaginary line 172 defines a minimum distance D1-1 between the radially outer vane edge 108 and the radially outer blade edge 148. In the illustrated embodiment of Figure 2, during both the cold idle condition and the hot running condition of the gas turbine engine 10, the outer platform trailing edge 114 is disposed radially outwards of the first imaginary line 172 with respect to the rotational axis X-X. In the illustrated embodiment of Figure 2, the shroud extending portion 158 may be substantially parallel to the first imaginary line 172. In the illustrated embodiment of Figure 2; the tip shroud leading edge 156 may intersect the first imaginary line 172. Specifically, the first imaginary line 172 may contain the tip shroud leading edge 156.

Moreover, a second imaginary line 174 extends between the radially inner vane edge 106 and the radially inner blade edge 146. The second imaginary line 174 defines a minimum distance D2-1 between the radially inner vane edge 106 and the radially inner blade edge 146. In the illustrated embodiment of Figure 2, during both the cold idle condition and the hot running condition of the gas turbine engine 10, the inner platform trailing edge 120 is disposed radially inwards of the second imaginary line 174 with respect to the rotational axis X-X'. In the illustrated embodiment of Figure 2, the inner extending portion 122 of the radially inner platform 118 may be substantially parallel to the second imaginary line 174. In the illustrated embodiment of Figure 2, the radially inner vane edge 106 and the inner platform trailing edge 120 may intersect the second imaginary line 174. Specifically, the second imaginary line 174 may contain the radially inner vane edge 106 and the inner platform trailing edge 120.

In some embodiments, a third imaginary line 176 may extend between the radially outer vane edge 108 and the outer platform trailing edge 114. In some embodiments, the third imaginary line 176 may define a minimum distance D3-1 between the radially outer vane edge 108 and the outer platform trailing edge 114. Further, in some embodiments, the third imaginary line 176 may be inclined to the first imaginary line 172 by a first inclination angle 01-1. In some embodiments, the first inclination angle 01-1 may be from about 5 degrees to about 20 degrees. In some embodiments, the first inclination angle 01-1 may be preferably from about 10 degrees to about 14 degrees. Thus, the first imaginary line 172 may not contain the outer platform trailing edge 114 and the first imaginary line 172 may be radially offset from the outer platform trailing edge 114.

Figure 3 illustrates a stator vane 202 and the rotor blade 142 associated with the 5 turbine 16, according to another embodiment of the present disclosure. The stator vane 202 includes a vane aerofoil portion 204 including a radially inner vane edge 206 extending at least in the circumferential direction C. The vane aerofoil portion 204 also includes a radially outer vane edge 208 disposed radially outwards of the radially inner vane edge 206 with respect to the rotational axis X-X' and extending 10 at least in the circumferential direction C. The vane aerofoil portion 204 further includes a vane aerofoil trailing edge 210 extending from the radially inner vane edge 206 to the radially outer vane edge 208.

The stator vane 202 also includes a radially outer platform 212 coupled to the vane aerofoil portion 204. The radially outer platform 212 includes an outer platform trailing edge 214 radially proximal to and downstream of the radially outer vane edge 208. The outer platform trailing edge 214 extends at least in the circumferential direction C. Further, the outer platform trailing edge 214 is a most downstream edge of the radially outer platform 212.

In some embodiments, the radially outer platform 212 may include an outer extending portion 216 extending both downstream and radially outwards at least from the radially outer vane edge 208. In some embodiments, the outer extending portion 216 may include the outer platform trailing edge 214. In the illustrated embodiment of Figure 3, the outer extending portion 216 may be at least partially curved away from the radially outer vane edge 208. Thus, the outer extending portion 216 includes a curved design herein. In other words, the outer extending portion 216 may be arcuate in shape.

The stator vane 202 further includes a radially inner platform 218 coupled to the vane aerofoil portion 204 opposite to the radially outer platform 212. The radially inner platform 218 includes an inner platform trailing edge 220 radially proximal to and downstream of the radially inner vane edge 206. The inner platform trailing edge 220 extends at least in the circumferential direction C. Further, the inner platform trailing edge 220 is a most downstream edge of the radially inner platform 218.

In some embodiments, the radially inner platform 218 may further include an inner 5 extending portion 222 extending both downstream and radially inwards at least from the radially inner vane edge 206. In some embodiments, the inner extending portion 222 may include the inner platform trailing edge 220. In the illustrated embodiment of Figure 3, the inner extending portion 222 may be substantially planar. In other words, the inner extending portion 222 may be generally 10 rectangular in shape.

In some embodiments, the casing 160 may define a cavity 252. Specifically, the cavity 252 may be disposed proximate to and radially outwards of each of the vane aerofoil trailing edge 210 and the blade aerofoil leading edge 150.

Further, a first imaginary line 272 extends between the radially outer vane edge 208 and the radially outer blade edge 148. The first imaginary line 272 defines a minimum distance D1-2 between the radially outer vane edge 208 and the radially outer blade edge 148. In the illustrated embodiment of Figure 3, during both the cold idle condition and the hot running condition of the gas turbine engine 10, the outer platform trailing edge 214 is disposed radially outwards of the first imaginary line 272 with respect to the rotational axis X-X'. Further, in the illustrated embodiment of Figure 3, the shroud extending portion 158 may be substantially parallel to the first imaginary line 272. In the illustrated embodiment of Figure 3, the tip shroud leading edge 156 may intersect the first imaginary line 272. Specifically, the first imaginary line 272 may contain the tip shroud leading edge 156.

Moreover, a second imaginary line 274 extends between the radially inner vane edge 206 and the radially inner blade edge 146. The second imaginary line 274 defines a minimum distance D2-2 between the radially inner vane edge 206 and the radially inner blade edge 146. In the illustrated embodiment of Figure 3, during both the cold idle condition and the hot running condition of the gas turbine engine 10, the inner platform trailing edge 220 is disposed radially inwards of the second imaginary line 274 with respect to the rotational axis X-X'. In the illustrated embodiment of Figure 3, the inner extending portion 222 of the radially inner platform 218 may be substantially parallel to the second imaginary line 274. In the illustrated embodiment of Figure 3, the radially inner vane edge 206 and the inner platform trailing edge 220 may intersect the second imaginary line 274. Specifically, the second imaginary line 274 may contain the radially inner vane edge 206 and the inner platform trailing edge 220.

In some embodiments, a third imaginary line 276 may extend between the radially outer vane edge 208 and the outer platform trailing edge 214. In some embodiments, the third imaginary line 276 may define a minimum distance D3-2 between the radially outer vane edge 208 and the outer platform trailing edge 214. Further, in some embodiments, the third imaginary line 276 may be inclined to the first imaginary line 272 by a first inclination angle 81-2. In some embodiments, the first inclination angle 81-2 may be from about 5 degrees to about 20 degrees. In some embodiments, the first inclination angle 01-2 may be preferably from about 10 degrees to about 14 degrees. Thus, the first imaginary line 272 may not contain the outer platform trailing edge 214 and the first imaginary line 272 may be radially offset from the outer platform trailing edge 214.

Figure 4 illustrates a stator vane 302 and the rotor blade 142 associated with the turbine 16, according to another embodiment of the present disclosure. The stator vane 302 includes a vane aerofoil portion 304 including a radially inner vane edge 306 extending at least in the circumferential direction C. The vane aerofoil portion 304 also includes a radially outer vane edge 308 disposed radially outwards of the radially inner vane edge 306 with respect to the rotational axis X-X' and extending at least in the circumferential direction C. The vane aerofoil portion 304 further includes a vane aerofoil trailing edge 310 extending from the radially inner vane edge 306 to the radially outer vane edge 308.

The stator vane 302 also includes a radially outer platform 312 coupled to the vane aerofoil portion 304. The radially outer platform 312 includes an outer platform trailing edge 314 radially proximal to and downstream of the radially outer vane edge 308. The outer platform trailing edge 314 extends at least in the circumferential direction C. Further, the outer platform trailing edge 314 is a most downstream edge of the radially outer platform 312.

In some embodiments, the radially outer platform 312 may include an outer extending portion 316 extending both downstream and radially outwards at least from the radially outer vane edge 308. In some embodiments, the outer extending portion 316 may include the outer platform trailing edge 314. In the illustrated embodiment of Figure 4, the outer extending portion 316 further extends both downstream and radially outwards from the vane aerofoil portion 304 upstream of the radially outer vane edge 308. In other words, some portion of the outer extending portion 316 may be disposed upstream of the radially outer vane edge 308 and some portion of the outer extending portion 316 may be disposed downstream of the radially outer vane edge 308. In the illustrated embodiment of Figure 4, the outer extending portion 316 may be substantially planar. In other words, the outer extending portion 316 may be generally rectangular in shape.

The stator vane 302 further includes a radially inner platform 318 coupled to the vane aerofoil portion 304 opposite to the radially outer platform 312. The radially inner platform 318 includes an inner platform trailing edge 320 radially proximal to and downstream of the radially inner vane edge 306. The inner platform trailing edge 320 extends at least in the circumferential direction C. Further, the inner platform trailing edge 320 is a most downstream edge of the radially inner platform 318.

In some embodiments, the radially inner platform 318 may further include an inner extending portion 322 extending both downstream and radially inwards at least from the radially inner vane edge 306. In some embodiments, the inner extending portion 322 may include the inner platform trailing edge 320. In the illustrated embodiment of Figure 4, the inner extending portion 322 may be substantially planar. In other words, the inner extending portion 322 may be generally rectangular in shape.

In some embodiments, the casing 160 may define a cavity 352. Specifically, the cavity 352 may be disposed proximate to and radially outwards of each of the vane aerofoil trailing edge 310 and the blade aerofoil leading edge 150.

Further, a first imaginary line 372 extends between the radially outer vane edge 308 and the radially outer blade edge 148. The first imaginary line 372 defines a minimum distance D1-3 between the radially outer vane edge 308 and the radially outer blade edge 148. In the illustrated embodiment of Figure 4, during both the cold idle condition and the hot running condition of the gas turbine engine 10, the outer platform trailing edge 314 is disposed radially outwards of the first imaginary line 372 with respect to the rotational axis X-X'. Further, in the illustrated embodiment of Figure 4, the shroud extending portion 158 may be substantially parallel to the first imaginary line 372. In the illustrated embodiment of Figure 4, the tip shroud leading edge 156 may intersect the first imaginary line 372.

Specifically, the first imaginary line 372 may contain the tip shroud leading edge 156.

Moreover, a second imaginary line 374 extends between the radially inner vane edge 306 and the radially inner blade edge 146. The second imaginary line 374 defines a minimum distance D2-3 between the radially inner vane edge 306 and the radially inner blade edge 146. In the illustrated embodiment of Figure 4, during both the cold idle condition and the hot running condition of the gas turbine engine 10, the inner platform trailing edge 320 is disposed radially inwards of the second imaginary line 374 with respect to the rotational axis X-X'. In the illustrated embodiment of Figure 4, the inner extending portion 322 of the radially inner platform 318 may be substantially parallel to the second imaginary line 374. In the illustrated embodiment of Figure 4, the radially inner vane edge 306 and the inner platform trailing edge 320 may intersect the second imaginary line 374. Specifically, the second imaginary line 374 may contain the radially inner vane edge 306 and the inner platform trailing edge 320.

In some embodiments, a third imaginary line 376 may extend between the radially outer vane edge 308 and the outer platform trailing edge 314. In some embodiments, the third imaginary line 376 may define a minimum distance D3-3 between the radially outer vane edge 308 and the outer platform trailing edge 314. Further, in some embodiments, the third imaginary line 376 may be inclined to the first imaginary line 372 by a first inclination angle 81-3. In some embodiments, the first inclination angle 81-3 may be from about 5 degrees to about 20 degrees.

In some embodiments, the first inclination angle 01-3 may be preferably from about 10 degrees to about 14 degrees. Thus, the first imaginary line 372 may not contain the outer platform trailing edge 314 and the first imaginary line 372 may be radially offset from the outer platform trailing edge 314.

Figure 5 illustrates a stator vane 402 and the rotor blade 142 associated with the turbine 16, according to another embodiment of the present disclosure. The stator vane 402 includes a vane aerofoil portion 404 including a radially inner vane edge 406 extending at least in the circumferential direction C. The vane aerofoil portion 404 also includes a radially outer vane edge 408 disposed radially outwards of the radially inner vane edge 406 with respect to the rotational axis X-X' and extending at least in the circumferential direction C. The vane aerofoil portion 404 further includes a vane aerofoil trailing edge 410 extending from the radially inner vane edge 406 to the radially outer vane edge 408.

The stator vane 402 also includes a radially outer platform 412 coupled to the vane aerofoil portion 404. The radially outer platform 412 includes an outer platform trailing edge 414 radially proximal to and downstream of the radially outer vane edge 408. The outer platform trailing edge 414 extends at least in the circumferential direction C. Further, the outer platform trailing edge 414 is a most downstream edge of the radially outer platform 412.

In some embodiments, the radially outer platform 412 may include an outer extending portion 416 extending both downstream and radially outwards at least from the radially outer vane edge 408. In some embodiments, the outer extending portion 416 may include the outer platform trailing edge 414. In the illustrated embodiment of Figure 5, the outer extending portion 416 may be substantially planar. In other words, the outer extending portion 416 may be generally rectangular in shape.

The stator vane 402 further includes a radially inner platform 418 coupled to the vane aerofoil portion 404 opposite to the radially outer platform 412. The radially inner platform 418 includes an inner platform trailing edge 420 radially proximal to and downstream of the radially inner vane edge 406. The inner platform trailing edge 420 extends at least in the circumferential direction C. Further, the inner platform trailing edge 420 is a most downstream edge of the radially inner platform 418 In some embodiments, the radially inner platform 418 may further include an inner extending portion 422 extending both downstream and radially inwards at least from the radially inner vane edge 406. In some embodiments, the inner extending portion 422 may include the inner platform trailing edge 420. In the illustrated embodiment of Figure 5, the inner extending portion 422 may be substantially planar. In other words, the inner extending portion 422 may be generally rectangular in shape.

In some embodiments, the casing 160 may define a cavity 452. Specifically, the cavity 452 may be disposed proximate to and radially outwards of each of the vane aerofoil trailing edge 410 and the blade aerofoil leading edge 150.

Further, a first imaginary line 472 extends between the radially outer vane edge 408 and the radially outer blade edge 148. The first imaginary line 472 defines a minimum distance D1-4 between the radially outer vane edge 408 and the radially outer blade edge 148. In the illustrated embodiment of Figure 5, during both the cold idle condition and the hot running condition of the gas turbine engine 10, the outer platform trailing edge 414 is disposed radially outwards of the first imaginary line 472 with respect to the rotational axis X-X'. Further, in the illustrated embodiment of Figure 5, the shroud extending portion 158 may be substantially parallel to the first imaginary line 472. In the illustrated embodiment of Figure 5, the tip shroud leading edge 156 may intersect the first imaginary line 472. Specifically, the first imaginary line 472 may contain the tip shroud leading edge 156.

Moreover, a second imaginary line 474 extends between the radially inner vane edge 406 and the radially inner blade edge 146. The second imaginary line 474 defines a minimum distance D2-4 between the radially inner vane edge 406 and the radially inner blade edge 146. In the illustrated embodiment of Figure 5, during 5 both the cold idle condition and the hot running condition of the gas turbine engine 10, the inner platform trailing edge 420 is disposed radially inwards of the second imaginary line 474 with respect to the rotational axis X-X'. In the illustrated embodiment of Figure 5, the radially inner vane edge 406 may intersect the second imaginary line 474. Specifically, the second imaginary line 474 may 10 contain the radially inner vane edge 406.

In some embodiments, a third imaginary line 476 may extend between the radially outer vane edge 408 and the outer platform trailing edge 414. In some embodiments, the third imaginary line 476 may define a minimum distance D3-4 between the radially outer vane edge 408 and the outer platform trailing edge 414. Further, in some embodiments, the third imaginary line 476 may be inclined to the first imaginary line 472 by a first inclination angle 01-4. In some embodiments, the first inclination angle 01-4 may be from about 5 degrees to about 20 degrees. In some embodiments, the first inclination angle 01-4 may be preferably from about 10 degrees to about 14 degrees. Thus, the first imaginary line 472 may not contain the outer platform trailing edge 414 and the first imaginary line 472 may be radially offset from the outer platform trailing edge 414.

In some embodiments, a fourth imaginary line 478 may extend between the radially inner vane edge 406 and the inner platform trailing edge 420. In some embodiments, the fourth imaginary line 478 may define a minimum distance D4 4 between the radially inner vane edge 406 and the inner platform trailing edge 420. In some embodiments, the fourth imaginary line 478 may be inclined to the second imaginary line 474 by a second inclination angle 82-4. In some embodiments, the second inclination angle 02-4 may be from about 5 degrees to about 20 degrees. In some embodiments, the second inclination angle 02-4 may be preferably from about 10 degrees to about 14 degrees. Thus, the fourth imaginary line 478 may not contain the inner platform trailing edge 420 and the second imaginary line 474 may be radially offset from the inner platform trailing edge 420.

Figure 6 illustrates a stator vane 502 and the rotor blade 142 associated with the turbine 16, according to another embodiment of the present disclosure. The stator vane 502 includes a vane aerofoil portion 504 including a radially inner vane edge 5 506 extending at least in the circumferential direction C. The vane aerofoil portion 504 also includes a radially outer vane edge 508 disposed radially outwards of the radially inner vane edge 506 with respect to the rotational axis X-X' and extending at least in the circumferential direction C. The vane aerofoil portion 504 further includes a vane aerofoil trailing edge 510 extending from the radially inner vane 10 edge 506 to the radially outer vane edge 508.

The stator vane 502 also includes a radially outer platform 512 coupled to the vane aerofoil portion 504. The radially outer platform 512 includes an outer platform trailing edge 514 radially proximal to and downstream of the radially outer vane edge 508. The outer platform trailing edge 514 extends at least in the circumferential direction C. Further, the outer platform trailing edge 514 is a most downstream edge of the radially outer platform 512.

In some embodiments, the radially outer platform 512 may include an outer extending portion 516 extending both downstream and radially outwards at least from the radially outer vane edge 508. In some embodiments, the outer extending portion 516 may include the outer platform trailing edge 514. In the illustrated embodiment of Figure 6, the outer extending portion 516 may be substantially planar. In other words, the outer extending portion 516 may be generally rectangular in shape.

The stator vane 502 further includes a radially inner platform 518 coupled to the vane aerofoil portion 504 opposite to the radially outer platform 512. The radially inner platform 518 includes an inner platform trailing edge 520 radially proximal to and downstream of the radially inner vane edge 506. The inner platform trailing edge 520 extends at least in the circumferential direction C. Further, the inner platform trailing edge 520 is a most downstream edge of the radially inner platform 518.

In some embodiments, the radially inner platform 518 may further include an inner extending portion 522 extending both downstream and radially inwards at least from the radially inner vane edge 506. In some embodiments, the inner extending portion 522 may include the inner platform trailing edge 520. In the illustrated embodiment of Figure 6, the inner extending portion 522 may be at least partially curved away from the radially inner vane edge 506. Thus, the inner extending portion 522 includes a curved design herein. In other words, the inner extending portion 522 may be arcuate in shape.

In some embodiments, the casing 160 may define a cavity 552. Specifically, the cavity 552 may be disposed proximate to and radially outwards of each of the vane aerofoil trailing edge 510 and the blade aerofoil leading edge 150.

Further, a first imaginary line 572 extends between the radially outer vane edge 508 and the radially outer blade edge 148. The first imaginary line 572 defines a minimum distance D1-5 between the radially outer vane edge 508 and the radially outer blade edge 148. In the illustrated embodiment of Figure 6, during both the cold idle condition and the hot running condition of the gas turbine engine 10, the outer platform trailing edge 514 is disposed radially outwards of the first imaginary line 572 with respect to the rotational axis X-X'. Further, in the illustrated embodiment of Figure 6, the shroud extending portion 158 may be substantially parallel to the first imaginary line 572. In the illustrated embodiment of Figure 6, the tip shroud leading edge 156 may intersect the first imaginary line 572. Specifically, the first imaginary line 572 may contain the tip shroud leading edge 156.

Moreover, a second imaginary line 574 extends between the radially inner vane edge 506 and the radially inner blade edge 146. The second imaginary line 574 defines a minimum distance D2-5 between the radially inner vane edge 506 and the radially inner blade edge 146. In the illustrated embodiment of Figure 6, during both the cold idle condition and the hot running condition of the gas turbine engine 10, the inner platform trailing edge 520 is disposed radially inwards of the second imaginary line 574 with respect to the rotational axis X-X'. In the illustrated embodiment of Figure 6, the radially inner vane edge 506 may intersect the second imaginary line 574. Specifically, the second imaginary line 574 may contain the radially inner vane edge 506.

In some embodiments, a third imaginary line 576 may extend between the radially 5 outer vane edge 508 and the outer platform trailing edge 514. In some embodiments, the third imaginary line 576 may define a minimum distance D3-5 between the radially outer vane edge 508 and the outer platform trailing edge 514. Further, in some embodiments, the third imaginary line 576 may be inclined to the first imaginary line 572 by a first inclination angle 81-5. In some embodiments, 10 the first inclination angle 81-5 may be from about 5 degrees to about 20 degrees. In some embodiments, the first inclination angle 81-5 may be preferably from about 10 degrees to about 14 degrees. Thus, the first imaginary line 572 may not contain the outer platform trailing edge 514 and the first imaginary line 572 may be radially offset from the outer platform trailing edge 514.

In some embodiments, a fourth imaginary line 578 may extend between the radially inner vane edge 506 and the inner platform trailing edge 520. In some embodiments, the fourth imaginary line 578 may define a minimum distance D4-5 between the radially inner vane edge 506 and the inner platform trailing edge 520.

In some embodiments, the fourth imaginary line 578 may be inclined to the second imaginary line 574 by a second inclination angle 82-5. In some embodiments, the second inclination angle 02-5 may be from about 5 degrees to about 20 degrees. In some embodiments, the second inclination angle 02-5 may be preferably from about 10 degrees to about 14 degrees. Thus, the fourth imaginary line 578 may not contain the inner platform trailing edge 520 and the second imaginary line 574 may be radially offset from the inner platform trailing edge 520.

Figure 7 illustrates a stator vane 602 and the rotor blade 142 associated with the turbine 16, according to another embodiment of the present disclosure. The stator vane 602 includes a vane aerofoil portion 604 including a radially inner vane edge 606 extending at least in the circumferential direction C. The vane aerofoil portion 604 also includes a radially outer vane edge 608 disposed radially outwards of the radially inner vane edge 606 with respect to the rotational axis X-X' and extending at least in the circumferential direction C. The vane aerofoil portion 604 further includes a vane aerofoil trailing edge 610 extending from the radially inner vane edge 606 to the radially outer vane edge 608.

The stator vane 602 also includes a radially outer platform 612 coupled to the vane aerofoil portion 604. The radially outer platform 612 includes an outer platform trailing edge 614 radially proximal to and downstream of the radially outer vane edge 608. The outer platform trailing edge 614 extends at least in the circumferential direction C. Further, the outer platform trailing edge 614 is a most downstream edge of the radially outer platform 612.

In some embodiments, the radially outer platform 612 may include an outer extending portion 616 extending both downstream and radially outwards at least from the radially outer vane edge 608. In some embodiments, the outer extending portion 616 may include the outer platform trailing edge 614. In the illustrated embodiment of Figure 7, the outer extending portion 616 may be substantially planar. In other words, the outer extending portion 616 may be generally rectangular in shape.

The stator vane 602 further includes a radially inner platform 618 coupled to the vane aerofoil portion 604 opposite to the radially outer platform 612. The radially inner platform 618 includes an inner platform trailing edge 620 radially proximal to and downstream of the radially inner vane edge 606. The inner platform trailing edge 620 extends at least in the circumferential direction C. Further, the inner platform trailing edge 620 is a most downstream edge of the radially inner platform 618.

In some embodiments, the radially inner platform 618 may further include an inner extending portion 622 extending both downstream and radially inwards at least from the radially inner vane edge 606. In some embodiments, the inner extending portion 622 may include the inner platform trailing edge 620. In some embodiments, the inner extending portion 624 further extends both downstream and radially inwards from the vane aerofoil portion 604 upstream of the radially inner vane edge 606. In other words, some portion of the inner extending portion 624 may be disposed upstream of the radially inner vane edge 606 and some portion of the inner extending portion 624 may be disposed downstream of the radially inner vane edge 606. In the illustrated embodiment of Figure 7, the inner extending portion 622 may be substantially planar. In other words, the inner extending portion 624 may be generally rectangular in shape.

In some embodiments, the casing 160 may define a cavity 652. Specifically, the cavity 652 may be disposed proximate to and radially outwards of each of the vane aerofoil trailing edge 610 and the blade aerofoil leading edge 150.

Further, a first imaginary line 672 extends between the radially outer vane edge 608 and the radially outer blade edge 148. The first imaginary line 672 defines a minimum distance D1-6 between the radially outer vane edge 608 and the radially outer blade edge 148. In the illustrated embodiment of Figure 7, during both the cold idle condition and the hot running condition of the gas turbine engine 10, the outer platform trailing edge 614 is disposed radially outwards of the first imaginary line 672 with respect to the rotational axis X-X'. Further, in the illustrated embodiment of Figure 7, the shroud extending portion 158 may be substantially parallel to the first imaginary line 672. In the illustrated embodiment of Figure 7, the tip shroud leading edge 156 may intersect the first imaginary line 672.

Specifically, the first imaginary line 672 may contain the tip shroud leading edge 156.

Moreover, a second imaginary line 674 extends between the radially inner vane edge 606 and the radially inner blade edge 146. The second imaginary line 674 defines a minimum distance D2-6 between the radially inner vane edge 606 and the radially inner blade edge 146. In the illustrated embodiment of Figure 7, during both the cold idle condition and the hot running condition of the gas turbine engine 10, the inner platform trailing edge 620 is disposed radially inwards of the second imaginary line 674 with respect to the rotational axis X-X'. Further, in the illustrated embodiment of Figure 7, the radially inner vane edge 606 may intersect the second imaginary line 674. Specifically, the second imaginary line 674 may contain the radially inner vane edge 606.

In some embodiments, a third imaginary line 676 may extend between the radially outer vane edge 608 and the outer platform trailing edge 614. In some embodiments, the third imaginary line 676 may define a minimum distance D3-6 between the radially outer vane edge 608 and the outer platform trailing edge 614. 5 Further, in some embodiments, the third imaginary line 676 may be inclined to the first imaginary line 672 by a first inclination angle 01-6. In some embodiments, the first inclination angle 01-6 may be from about 5 degrees to about 20 degrees. In some embodiments, the first inclination angle 01-6 may be preferably from about 10 degrees to about 14 degrees. Thus, the first imaginary line 672 may not 10 contain the outer platform trailing edge 614 and the first imaginary line 672 may be radially offset from the outer platform trailing edge 614.

In some embodiments, a fourth imaginary line 678 may extend between the radially inner vane edge 606 and the inner platform trailing edge 620. In some embodiments, the fourth imaginary line 678 may define a minimum distance D4-6 between the radially inner vane edge 606 and the inner platform trailing edge 620. In some embodiments, the fourth imaginary line 678 may be inclined to the second imaginary line 674 by a second inclination angle 02-6. In some embodiments, the second inclination angle 02-6 may be from about 5 degrees to about 20 degrees.

In some embodiments, the second inclination angle 02-6 may be preferably from about 10 degrees to about 14 degrees. Thus, the fourth imaginary line 678 may not contain the inner platform trailing edge 620 and the second imaginary line 674 may be radially offset from the inner platform trailing edge 620.

Figure 8 illustrates the stator vane 102 and a rotor blade 742 associated with the turbine 16, according to another embodiment of the present disclosure. The rotor blade 742 is disposed downstream of the stator vane 102. Further, the rotor blade 742 is configured to receive a fluid flow from the stator vane 102. The rotor blade 742 includes a blade aerofoil portion 744. The blade aerofoil portion 744 includes a radially inner blade edge 746 extending at least in the circumferential direction C. The blade aerofoil portion 744 also includes a radially outer blade edge 748 disposed radially outwards of the radially inner blade edge 746 with respect to the rotational axis X-X' and extending at least in the circumferential direction C. The blade aerofoil portion 744 further includes a blade aerofoil leading edge 750 extending from the radially inner blade edge 746 to the radially outer blade edge 748.

The rotor blade 742 also includes a blade tip shroud 754 coupled to the blade 5 aerofoil portion 744. The blade tip shroud 754 extends at least radially outwards from the radially outer blade edge 748 of the blade aerofoil portion 744. In some embodiments, the blade tip shroud 754 may include a tip shroud leading edge 756 radially proximal to and upstream of the radially outer blade edge 748. In some embodiments, the tip shroud leading edge 756 may extend at least in the 10 circumferential direction C and is a most upstream edge of the blade tip shroud 754.

In some embodiments, the blade tip shroud 754 may further includes a shroud extending portion 758 extending both upstream and radially outwards from the radially outer blade edge 748 towards the radially outer vane edge 108. In some embodiments, the shroud extending portion 758 may include the tip shroud leading edge 756. The shroud extending portion 758 may be generally rectangular in shape.

In some embodiments, the casing 160 may define a cavity 752. Specifically, the cavity 752 may be disposed proximate to and radially outwards of each of the vane aerofoil trailing edge 110 and the blade aerofoil leading edge 750.

Further, the first imaginary line 172 defines a minimum distance D-1-7 between the radially outer vane edge 108 and the radially outer blade edge 748. In the illustrated embodiment of Figure 8, the shroud extending portion 758 may extend radially outwards of the first imaginary line 172 with respect to the rotational axis X-X'. Further, in some embodiments, the tip shroud leading edge 756 may be disposed radially outwards of the first imaginary line 172 with respect to the rotational axis X-X'.

Moreover, the second imaginary line 174 defines a minimum distance D2-7 between the radially inner vane edge 106 and the radially inner blade edge 746.

In some embodiments, a fifth imaginary line 778 may extend between the radially outer blade edge 748 and the tip shroud leading edge 756. In some embodiments, the fifth imaginary line 778 may define a minimum distance D5-7 between the radially outer blade edge 748 and the tip shroud leading edge 756. Further, in some embodiments, the fifth imaginary line 778 may be inclined to the first imaginary line 172 by an inclination angle T1-7. Thus, the fifth imaginary line 778 may not contain the tip shroud leading edge 756 and the first imaginary line 172 may be radially offset from the tip shroud leading edge 756.

Figure 9 illustrates the stator vane 202 and the rotor blade 742 associated with the turbine 16, according to another embodiment of the present disclosure. In some embodiments, the casing 160 may define a cavity 852. Specifically, the cavity 852 may be disposed proximate to and radially outwards of each of the vane aerofoil trailing edge 110 and the blade aerofoil leading edge 750.

Further, the first imaginary line 272 defines a minimum distance D1-8 between the radially outer vane edge 208 and the radially outer blade edge 748. In the illustrated embodiment of Figure 9, the shroud extending portion 758 may extend radially outwards of the first imaginary line 272 with respect to the rotational axis X-X'. Further, in some embodiments, the tip shroud leading edge 756 may be disposed radially outwards of the first imaginary line 272 with respect to the rotational axis X-X'. Moreover, the second imaginary line 274 defines a minimum distance D2-8 between the radially inner vane edge 206 and the radially inner blade edge 746.

In some embodiments, the fifth imaginary line 778 may extend between the radially outer blade edge 748 and the tip shroud leading edge 756. In some embodiments, the fifth imaginary line 778 may define the minimum distance D5-7 between the radially outer blade edge 748 and the tip shroud leading edge 756.

Further, in some embodiments, the fifth imaginary line 778 may be inclined to the first imaginary line 272 by the inclination angle rp1-7. Thus, the fifth imaginary line 778 may not contain the tip shroud leading edge 756 and the first imaginary line 172 may be radially offset from the tip shroud leading edge 756 Figure 10 illustrates the stator vane 302 and the rotor blade 742 associated with the turbine 16, according to another embodiment of the present disclosure. In some embodiments, the casing 160 may define a cavity 952. Specifically, the cavity 952 may be disposed proximate to and radially outwards of each of the vane aerofoil trailing edge 110 and the blade aerofoil leading edge 750.

Further, the first imaginary line 372 defines a minimum distance D1-9 between the radially outer vane edge 208 and the radially outer blade edge 748. In the illustrated embodiment of Figure 10, the shroud extending portion 758 may extend radially outwards of the first imaginary line 372 with respect to the rotational axis X-X'. Further, in some embodiments, the tip shroud leading edge 756 may be disposed radially outwards of the first imaginary line 372 with respect to the rotational axis X-X'. Moreover, the second imaginary line 374 defines a minimum distance D2-9 between the radially inner vane edge 206 and the radially inner blade edge 746.

In some embodiments, the fifth imaginary line 778 may extend between the radially outer blade edge 748 and the tip shroud leading edge 756. In some embodiments, the fifth imaginary line 778 may define the minimum distance D5-7 between the radially outer blade edge 748 and the tip shroud leading edge 756. Further, in some embodiments, the fifth imaginary line 778 may be inclined to the first imaginary line 372 by the inclination angle cp1-7. Thus, the fifth imaginary line 778 may not contain the tip shroud leading edge 756 and the first imaginary line 372 may be radially offset from the tip shroud leading edge 756.

Referring to Figures 2 to 10, when the outer platform trailing edge 114, 214, 314, 414, 514, 614 may be disposed radially outwards of the first imaginary line 172, 272, 372, 472, 572, 672 with respect to the rotational axis X-X', a main flow of the hot gases having a high whirl velocity may impart a whirl velocity to a cavity flow of the hot gases. As a result, a velocity difference between the main flow and the cavity flow may reduce, thereby reducing secondary losses in the turbine 16.

Further, when the inner platform trailing edge 420, 520, 620 may be disposed radially inwards of the second imaginary line 474, 574, 674 with respect to the rotational axis X-X', the main flow of hot gases having the high whirl velocity may impart the whirl velocity to the cavity flow of the hot gases. As a result, the velocity difference between the main flow and the cavity flow may reduce, thereby reducing secondary losses in the turbine 16.

Furthermore, an improvement in the design of the stator vane 102, 202, 302, 402, 502, 602 as described herein may cause a reduction in a temperature of the blade tip shroud 154, 754 and more specifically, the tip shroud leading edge 156, 756, thereby improving an operational lifetime of the rotor blade 142, 742. Moreover, the reduction in the temperature of the blade tip shroud 154, 754 may also reduce an amount of cooling air that may be otherwise required to cool the blade tip shroud 154, 754.

Moreover, when the tip shroud leading edge 756 may be disposed radially outwards of the first imaginary line 172, 272, 372 with respect to the rotational axis X-X', the main flow of hot gases having the high whirl velocity may impart the whirl velocity to the cavity flow of the hot gases. As a result, the velocity difference between the main flow and the cavity flow may reduce, thereby reducing secondary losses in the turbine 16.

Figure 11 is a plot 1100 illustrating a variation of a turbine efficiency of the turbine 16 as per a variation in the first inclination angle 01-1 (see Figure 2) associated with the outer extending portion 116 (see Figure 2). Various values for the first inclination angles el-1 are depicted along an X-axis and various values for percentage increase in the turbine efficiency are depicted along a Y-axis. Values X1, X2, ...Xn on the Y-axis depict increase in the turbine efficiency, whereas values Y1, Y2, Yn on the Y-axis depict decrease in the turbine efficiency.

The plot 1100 includes a bar 1102 depicting a turbine efficiency of the turbine 16 when the first inclination angle 01-1 is about -5 degrees. It should be noted that negative values for first inclination angle 01-1 as mentioned in this disclosure may relate to a radially inward extension of the outer extending portion 116 of the turbine 16. The plot 1100 includes a bar 1104 depicting a turbine efficiency of the turbine 16 when the first inclination angle 01-1 is about -2.5 degrees. The plot 1100 includes a bar 1106 depicting a turbine efficiency of the turbine 16 when the first inclination angle 81-1 is about 0 degree. The plot 1100 includes a bar 1108 depicting a turbine efficiency of the turbine 16 when the first inclination angle 911 is about 5 degrees. The plot 1100 includes a bar 1110 depicting a turbine efficiency of the turbine 16 when the first inclination angle 01-1 is about 8 degrees.

As can be observed from the plot 1100, an increase in the first inclination angle 81-1 may increase the turbine efficiency of the turbine 16.

Figure 12 is a plot 1200 illustrating a change in an ingestion of a fluid flow into a cavity (such as, the cavity 152 shown in Figure 2) of the turbine 16 as per a variation in the first inclination angle 01-1 (see Figure 2). Various values for the first inclination angles 91-1 are depicted along an X-axis and various values for percentage change in the ingestion are depicted along a Y-axis. Values X1, X2, 15...Xn on the Y-axis depict increase in the ingestion, whereas values Yl, Y2, ...Yn on the Y-axis depict decrease in the ingestion.

The plot 1200 includes a bar 1202 depicting a change in the ingestion of the fluid flow when the first inclination angle 81-1 is about -5 degrees. The plot 1200 includes a bar 1204 depicting a change in the ingestion of the fluid flow when the first inclination angle 81-1 is about -2.5 degrees. The plot 1200 includes a bar 1206 depicting a change in the ingestion of the fluid flow when the first inclination angle 01-1 is about 0 degree. The plot 1200 includes a bar 1208 depicting a change in the ingestion of the fluid flow when the first inclination angle 81-1 is about 5 degrees. The plot 1200 includes a bar 1210 depicting a change in the ingestion of the fluid flow when the first inclination angle 81-1 is about 8 degrees.

As can be observed from the plot 1200, an increase in the first inclination angle 81-1 may decrease the ingestion of the fluid flow in the turbine 16.

Figure 13 illustrates an exemplary schematic perspective view of a stator vane 1302 of a turbine 1300. The stator vane 1302 includes a vane aerofoil portion 1304. The stator vane 1302 also includes a radially outer platform 1312 coupled to the vane aerofoil portion 1304. Further, Figure 13 illustrates an outer extending portion 1316-1 as defined on conventional turbines. The outer extending portion 1316-1 has a first inclination angle (that may be defined similar to the first inclination angle 81-1 shown in Figure 2) of 0 degree. Moreover, Figure 13 also illustrates an outer extending portion 1316-1 that is similar to the outer extending portion 116 shown in Figure 2. The outer extending portion 1316-2 may have a first inclination angle (that may be defined similar to the first inclination angle 81-'1 shown in Figure 2) of about 8 degrees. The turbine 1300 further defines a cavity 1352 (shown in Figures 14A to 14D) similar to the cavity 152 (see Figure 2).

Figures 14A to 14D illustrate variations in fluid flows 1402, 1404, 1406, 1408 through the turbine 1300 (see Figure 13) as per a variation in the first inclination angle of the outer extending portions 1316-1, 1316-2. The turbine 1300 further includes a shroud extending portion 1358.

Figure 14A illustrates the fluid flow 1402 flowing through the turbine 1300 (see Figure 13). As illustrated in Figure 14B, an increased vortex roll-up of the fluid flow 1402 into the cavity 1352 may be visible when the first inclination angle of the outer extending portion 1316-1 is about -5 degrees.

Figure 14B illustrates the fluid flow 1404 flowing through the turbine 1300 (see Figure 13). As illustrated in Figure 14B, when the first inclination angle of the outer extending portion 1316-1 is about 0 degree, a vortex roll-up of the fluid flow 1404 into the cavity 1352 may be lesser as compared to the vortex roll-up of the fluid flow 1402.

Figure 14C illustrates the fluid flow 1406 flowing through the turbine 1300 (see Figure 13). As illustrated in Figure 14C, when the first inclination angle of the outer extending portion 1316-2 is about 5 degrees, a vortex roll-up of the fluid flow 1406 into the cavity 1352 may be lesser as compared to the vortex roll-up of the fluid flow 1404.

Figure 14D illustrates the fluid flow 1408 flowing through the turbine 1300 (see Figure 13). As illustrated in Figure 14D, when the first inclination angle of the outer extending portion 1316-2 is about 8 degrees, a vortex roll-up of the fluid flow 1406 into the cavity 1352 may be lesser as compared to the vortex roll-up of the fluid flow 1406. Thus, as can be observed from Figures 14A to 14D, the fluid flow 1408 through the turbine 1300 when the first inclination angle is about 8 degrees may be less turbulent compared to the fluid flows 1402, 1404, and 1406.

It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

CLAIMS: 1. A turbine (16) for a gas turbine engine (10) having a rotational axis (X-X') and a circumferential direction (C) with respect to the rotational axis (X-X'), the 5 turbine (16) comprising: a stator vane (102, 202, 302, 402, 502, 602) comprising: a vane aerofoil portion (104, 204, 304, 404, 504, 604) comprising a radially inner vane edge (106, 206, 306, 406, 506, 606) extending at least in the circumferential direction (C), a radially outer vane edge (108, 208, 308, 408, 508, 608) disposed radially outwards of the radially inner vane edge (106, 206, 306, 406, 506, 606) with respect to the rotational axis (X-X') and extending at least in the circumferential direction (C), and a vane aerofoil trailing edge (110, 210, 310, 410, 510, 610) extending from the radially inner vane edge (106, 206, 306, 406, 506, 606) to the radially outer vane edge (108, 208, 308, 408, 508, 608); a radially outer platform (112, 212, 312, 412, 512, 612) coupled to the vane aerofoil portion (104, 204, 304, 404, 504, 604) and comprising an outer platform trailing edge (114, 214, 314, 414, 514, 614) radially proximal to and downstream of the radially outer vane edge (108, 208, 308, 408, 508, 608), wherein the outer platform trailing edge (114, 214, 314, 414, 514, 614) extends at least in the circumferential direction (C) and is a most downstream edge of the radially outer platform (112, 212; 312, 412, 512, 612); and a radially inner platform (118, 218, 318, 418, 518, 618) coupled to the vane aerofoil portion (104, 204, 304, 404, 504, 604) opposite to the radially outer platform (112, 212, 312, 412, 512, 612) and comprising an inner platform trailing edge (120, 220, 320, 420, 520, 620) radially proximal to and downstream of the radially inner vane edge (106, 206, 306, 406, 506, 606), wherein the inner platform trailing edge (120, 220, 320, 420, 520, 620) extends at least in the circumferential direction (C) and is a most downstream edge of the radially inner platform (118, 218, 318, 418, 518, 618); a rotor blade (142, 742) disposed downstream of the stator vane (102, 202, 302, 402, 502, 602) and configured to receive a fluid flow from the stator vane (102, 202, 302, 402, 502, 602), the rotor blade (142, 742) comprising: a blade aerofoil portion (144, 744) comprising a radially inner blade edge (146, 746) extending at least in the circumferential direction (C), a radially outer blade edge (148, 748) disposed radially outwards of the radially inner blade edge (146, 746) with respect to the rotational axis (X-X') and extending at least in the circumferential direction (C), and a blade aerofoil leading edge (150, 750) extending from the radially inner blade edge (146, 746) to the radially outer blade edge (148, 748); and a blade tip shroud (154, 754) coupled to the blade aerofoil portion (144, 744) and extending at least radially outwards from the radially outer blade edge (148, 748) of the blade aerofoil portion (144, 744); and a casing (160) extending at least in the circumferential direction (C) and at 15 least partially enclosing the stator vane (102, 202, 302, 402, 502, 602) and the rotor blade (142, 742); wherein, during a cold idle condition of the gas turbine engine (10) where at least one of a current engine temperature is lesser than 70% of a maximum rated take-off engine temperature and a current turbine shaft speed is lesser than 80% of a design turbine shaft speed, the blade tip shroud (154, 754) is disposed at a clearance (C1) from the casing (160); wherein, during a hot running condition of the gas turbine engine (10) where at least one of the current engine temperature is greater than 70% of the maximum rated take-off engine temperature and the current turbine shaft speed is greater than 80% of the design turbine shaft speed, the blade tip shroud (154, 754) engages the casing (160); wherein a first imaginary line (172, 272, 372, 472, 572, 672) extends between the radially outer vane edge (108, 208, 308, 408, 508, 608) and the radially outer blade edge (148, 748), the first imaginary line (172, 272, 372, 472, 30 572, 672) defining a minimum distance (D1-1, D1-2, D1-3, D1-4, D1-5, D1-6) between the radially outer vane edge (108, 208, 308, 408, 508, 608) and the radially outer blade edge (148, 748); wherein a second imaginary line (174, 274, 374, 474, 574, 674) extends between the radially inner vane edge (106, 206, 306, 406, 506, 606) and the radially inner blade edge (146, 746), the second imaginary line (174, 274, 374, 474, 574, 674) defining a minimum distance (D2-1, D2-2, D2-3, D2-4, 02-5, Dl-6) between the radially inner vane edge (106, 206, 306, 406, 506, 606) and the radially inner blade edge (146, 746); and wherein, during both the cold idle condition and the hot running condition of the gas turbine engine (10): the outer platform trailing edge (114, 214, 314, 414, 514, 614) is disposed radially outwards of the first imaginary line (172, 272, 372, 472, 572, 672) with respect to the rotational axis (X-X'); and/or the inner platform trailing edge (120, 220, 320, 420, 520, 620) is disposed radially inwards of the second imaginary line (174, 274, 574, 674) with respect to the rotational axis (X-X'). 374, 474, 2. The turbine (16) of claim 1, wherein a third imaginary line (176, 276, 376, 476, 576, 676) extends between the radially outer vane edge (108, 208, 308, 408, 508, 608) and the outer platform trailing edge (114, 214, 314, 414, 514, 614), the third imaginary line (176, 276, 376, 476, 576, 676) defining a minimum distance (D3-1, D3-2, D3-3, D3-4, D3-5, D3-6) between the radially outer vane edge (108, 208, 308, 408, 508, 608) and the outer platform trailing edge (114, 214, 314, 414, 514, 614), and wherein the third imaginary line (176, 276, 376, 476, 576, 676) is inclined to the first imaginary line (172, 272, 372, 472, 572, 672) by a first inclination angle (01-1, 01-2, 91-3, 01-4, 01-5, 01-6).3. The turbine (16) of claim 2, wherein the first inclination angle (91-1, 81-2, 81-3, 01-4, 91-5, 81-6) is from about 5 degrees to about 20 degrees.4. The turbine (16) of any one of claims 1 to 3, wherein a fourth imaginary line (478. 578, 678) extends between the radially inner vane edge (106, 206, 306, 406, 506, 606) and the inner platform trailing edge (120, 220, 320, 420, 520, 620), the fourth imaginary line (478. 578, 678) defining a minimum distance (D4-4, D45, D4-6) between the radially inner vane edge (106, 206, 306, 406, 506, 606) and the inner platform trailing edge (120, 220, 320, 420, 520, 620), and wherein the fourth imaginary line (478. 578, 678) is inclined to the second imaginary line (174, 274, 374, 474, 574, 674) by a second inclination angle (92-4, 02-5, 92-6).5. The turbine (16) of claim 4, wherein the second inclination angle (82-4, 92- 5, 82-6) is from about 5 degrees to about 20 degrees.6. The turbine (16) of any one of claims 1 to 5, wherein the radially outer platform (112, 212, 312, 412, 512, 612) comprises an outer extending portion (116, 216, 316, 416, 516, 616) extending both downstream and radially outwards at least from the radially outer vane edge (108, 208, 308, 408, 508, 608), the outer extending portion (116, 216, 316, 416, 516, 616) comprising the outer platform trailing edge (114, 214, 314, 414, 514, 614).7. The turbine (16) of claim 6, wherein the outer extending portion (116, 316, 416, 516, 616) is substantially planar.8. The turbine (16) of claim 6, wherein the outer extending portion (216) is at least partially curved away from the radially outer vane edge (208).9. The turbine (16) of any one of claims 6 to 8, wherein the outer extending portion (316) further extends both downstream and radially outwards from the 20 vane aerofoil portion (304) upstream of the radially outer vane edge (308).10. The turbine (16) of any one of claim 1 to 9, wherein the radially inner platform (118, 218, 318, 418, 518, 618) comprises an inner extending portion (122, 222, 322, 422, 522, 622) extending both downstream and radially inwards at least from the radially inner vane edge (106, 206, 306, 406, 506, 606), the inner extending portion (122, 222, 322, 422, 522, 622) comprising the inner platform trailing edge (120, 220, 320, 420, 520, 620).11. The turbine (16) of claim 10, wherein the inner extending portion (122, 222, 322, 422, 622) is substantially planar.12. The turbine (16) of claim 10, wherein the inner extending portion (522) is at least partially curved away from the radially inner vane edge (506).13. The turbine (16) of any one of claims 10 to 12, wherein the inner extending portion (622) further extends both downstream and radially inwards from the vane aerofoil portion (604) upstream of the radially inner vane edge (606).14. The turbine (16) of any one of claims 1 to 13, wherein the blade tip shroud (154, 754) comprises a tip shroud leading edge (156, 756) radially proximal to and upstream of the radially outer blade edge (148, 748), and wherein the tip shroud leading edge (156, 756) extends at least in the circumferential direction (C) and is a most upstream edge of the blade tip shroud (154, 754).15. The turbine (16) of claim 14, wherein the tip shroud leading edge (156) intersects the first imaginary line (172, 272, 372, 472, 572, 672).16. The turbine (16) of claim 14, wherein the tip shroud leading edge (756) is 15 disposed radially outwards of the first imaginary line (172, 272, 372) with respect to the rotational axis (X-X').17. The turbine (16) of claim 14 wherein the blade tip shroud (154, 754) further comprises a shroud extending portion (158, 758) extending both upstream and radially outwards from the radially outer blade edge (148, 748) towards the radially outer vane edge (108, 208, 308, 408, 508, 608), the shroud extending portion (158, 758) comprising the tip shroud leading edge (156, 756).18. A gas turbine engine (10) for an aircraft, the gas turbine engine (10) 25 comprising the turbine (16) of any one of the preceding claims.