CA2568692A1 - Vane platform tangential injection - Google Patents
Vane platform tangential injection Download PDFInfo
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- CA2568692A1 CA2568692A1 CA 2568692 CA2568692A CA2568692A1 CA 2568692 A1 CA2568692 A1 CA 2568692A1 CA 2568692 CA2568692 CA 2568692 CA 2568692 A CA2568692 A CA 2568692A CA 2568692 A1 CA2568692 A1 CA 2568692A1
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- gas path
- platform
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Abstract
An improved vane platform flow arrangement comprising a plurality of holes defined in at least one vane platform downstream edge at an angle substantially parallel to the airfoil trailing edge angle, the holes adapted to eject a cooling flow into the annular gas path generally in the direction of the gas path flow.
Description
VANE PLATFORM TANGENTIAL INJECTION
TECHNICAL FIELD
The invention relates generally to a gas turbine engine and, more particularly, to an improved vane platform cooling flow arrangement.
BACKGROUND OF THE ART
Cooling air that is reintroduced into the hot gas path of a gas turbine engine is known to have a significant impact on turbine efficiency. This is particularly the case in a turbine stage where the total pressure loss caused by the cooling air mixing with the hot combustion gases is significant. Therefore, to reduce inefficiencies caused by colliding air flows, it is desirable for the cooling air flow to merge with the hot gas path flow in an optimal aerodynamic manner. Such improved fluid flow control is advantageous in improving turbine performance.
Accordingly, there is a need for controlling the cooling air entering the hot gas path of gas turbine engines.
SUMMARY OF THE INVENTION
It is therefore an object of this invention to provide an improved vane platform arrangement for a gas turbine engine which addresses the above-mentioned issues.
In one aspect, the present invention provides a cooled stationary vane, the cooled stationary vane comprising a vane platform having a downstream edge, an airfoil extending radially from said vane platform and having a trailing edge, and at least one coolant exit hole extending through said downstream edge of said vane platform, the at least one coolant exit hole extending along a direction substantially parallel to a trailing edge portion of the airfoil in a plane normal to the airfoil.
In another aspect, the present invention provides a vane assembly for a gas turbine engine, comprising a first and a second platform having an upstream edge and a downstream edge and defining an annular gas path therebetween, a plurality of airfoils extending radially between the first and second platforms across the annular gas path and having leading and trailing edges, the trailing edges provided at an angle from an axial direction of the annular gas path, and at least one hole defined in at least one of the platform downstream edges at an angle substantially parallel to the trailing edge angle, the hole adapted to eject a cooling flow into the annular gas path in the direction of the gas path flow.
In another aspect, the present invention provides a stator vane assembly in a turbine stage of a gas turbine engine, the vane assembly comprising an inner and an outer shroud having an upstream edge and a downstream edge and defining an annular gas path therebetween, a plurality of airfoils extending radially between the inner and outer shrouds across the annular gas path and having leading and trailing edges, the trailing edges provided at an angle from an axial direction of the annular gas path, and a plurality of holes defined along the downstream edge of the inner shroud at an angle substantially parallel to the trailing edge angle, wherein a cooling flow exiting the holes entrains a purge flow entering the annular gas path adjacent the downstream edge of the inner shroud and redirects the purge flow to enter the annular gas path in a direction substantially tangential to a direction of the gas path flow.
In another aspect, the present invention provides a stator vane for a gas turbine engine comprising: an upper and a lower platform having an upstream edge and a downstream edge and defining a gas path therebetween; an airfoil extending between the upper and lower platforms having a leading edge, a trailing edge, a pressure side, a suction side, and adapted to extend transverse to an axial direction of the gas path, the trailing edge of the airfoil defined at an acute angle with the axial direction; and a plurality of holes defined on at least one of the platform downstream edges at an angle substantially parallel to the trailing edge angle, the holes providing a fluid flow path ejecting a platform cooling air flow into the gas path in a direction of the gas path flow.
In a still further general aspect of the present invention, there is provided a method of redirecting a radially flowing purge flow in a direction of a gas path flow in a gas turbine engine comprising: ejecting a cooling flow from a stator vane platform at an angle substantially parallel to a trailing edge of an airfoil of the stator vane; and entraining the purge flow from the radial direction to a direction tangential to the gas path flow.
TECHNICAL FIELD
The invention relates generally to a gas turbine engine and, more particularly, to an improved vane platform cooling flow arrangement.
BACKGROUND OF THE ART
Cooling air that is reintroduced into the hot gas path of a gas turbine engine is known to have a significant impact on turbine efficiency. This is particularly the case in a turbine stage where the total pressure loss caused by the cooling air mixing with the hot combustion gases is significant. Therefore, to reduce inefficiencies caused by colliding air flows, it is desirable for the cooling air flow to merge with the hot gas path flow in an optimal aerodynamic manner. Such improved fluid flow control is advantageous in improving turbine performance.
Accordingly, there is a need for controlling the cooling air entering the hot gas path of gas turbine engines.
SUMMARY OF THE INVENTION
It is therefore an object of this invention to provide an improved vane platform arrangement for a gas turbine engine which addresses the above-mentioned issues.
In one aspect, the present invention provides a cooled stationary vane, the cooled stationary vane comprising a vane platform having a downstream edge, an airfoil extending radially from said vane platform and having a trailing edge, and at least one coolant exit hole extending through said downstream edge of said vane platform, the at least one coolant exit hole extending along a direction substantially parallel to a trailing edge portion of the airfoil in a plane normal to the airfoil.
In another aspect, the present invention provides a vane assembly for a gas turbine engine, comprising a first and a second platform having an upstream edge and a downstream edge and defining an annular gas path therebetween, a plurality of airfoils extending radially between the first and second platforms across the annular gas path and having leading and trailing edges, the trailing edges provided at an angle from an axial direction of the annular gas path, and at least one hole defined in at least one of the platform downstream edges at an angle substantially parallel to the trailing edge angle, the hole adapted to eject a cooling flow into the annular gas path in the direction of the gas path flow.
In another aspect, the present invention provides a stator vane assembly in a turbine stage of a gas turbine engine, the vane assembly comprising an inner and an outer shroud having an upstream edge and a downstream edge and defining an annular gas path therebetween, a plurality of airfoils extending radially between the inner and outer shrouds across the annular gas path and having leading and trailing edges, the trailing edges provided at an angle from an axial direction of the annular gas path, and a plurality of holes defined along the downstream edge of the inner shroud at an angle substantially parallel to the trailing edge angle, wherein a cooling flow exiting the holes entrains a purge flow entering the annular gas path adjacent the downstream edge of the inner shroud and redirects the purge flow to enter the annular gas path in a direction substantially tangential to a direction of the gas path flow.
In another aspect, the present invention provides a stator vane for a gas turbine engine comprising: an upper and a lower platform having an upstream edge and a downstream edge and defining a gas path therebetween; an airfoil extending between the upper and lower platforms having a leading edge, a trailing edge, a pressure side, a suction side, and adapted to extend transverse to an axial direction of the gas path, the trailing edge of the airfoil defined at an acute angle with the axial direction; and a plurality of holes defined on at least one of the platform downstream edges at an angle substantially parallel to the trailing edge angle, the holes providing a fluid flow path ejecting a platform cooling air flow into the gas path in a direction of the gas path flow.
In a still further general aspect of the present invention, there is provided a method of redirecting a radially flowing purge flow in a direction of a gas path flow in a gas turbine engine comprising: ejecting a cooling flow from a stator vane platform at an angle substantially parallel to a trailing edge of an airfoil of the stator vane; and entraining the purge flow from the radial direction to a direction tangential to the gas path flow.
Further details of these and other aspects of the present invention will be apparent from the detailed description and figures included below.
DESCRIPTION OF THE DRAWINGS
Reference is now made to the accompanying figures depicting aspects of the present invention, in which:
Figure 1 is a schematic side view of a gas turbine engine, in partial cross-section;
Figure 2 is a cross-sectional view of a vane assembly and a rotor assembly of a turbine stage of the gas turbine engine of Fig. 1;
Figure 3 is a perspective view of a vane of the vane assembly of Fig. 2, in accordance with a preferred embodiment of the present invention; and Figure 4 is a transverse sectional view taken along cross-section line 4-4 of Fig. 3, showing an airfoil extending from a platform of the vane, the airfoil having a trailing edge and the platform defining a plurality of angled cooling holes aligned with the trailing edge.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
Figure 1 illustrates a gas turbine engine 10 of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan 12 through which ambient air is propelled, a multistage compressor 14 for pressurizing the air, a combustor 16 in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section 18 for extracting energy from the combustion gases.
The turbine section 18 may comprise one or more turbine stages, in this case two are shown including a first, or high pressure (HP), turbine stage 20, which includes a turbine rotor assembly 22 and a turbine vane assembly 24 in accordance with the present invention.
Referring now to Figure 2, the turbine vane assembly 24 comprises generally a first or inner platform 26 and a second or outer platform 28 and a plurality of circumferentially spaced-apart airfoils 30 extending radially therebetween. The inner platform 26 or shroud and the outer platform 28 or shroud form an annular stator vane ring defining an annular gas path 32 therebetween, with the airfoils 30 extending across the annular gas path 32. More particularly, the inner vane platform 26 sealingly engages the inner combustion chamber wall 34 and the outer vane platform 28 sealingly engages the outer combustion chamber wall 36, thereby defining therebetween the annular path 32 for directing the stream of hot combustion gases exiting the combustion chamber outlet therethrough (not actually shown - inventor should provide complete Figure). Throughout this description, the axial, radial and circumferential directions are defined respectively with respect to the central axis, radius and circumference of the annular stator vane ring.
It should be noted that the turbine vane assembly 24 can be manufactured as a one-piece annular stator vane ring, wherein the inner platform 26, the airfoils 30 and the outer platform 28 are all integrally formed, or may be provided as a plurality of individual vane segments 40 arranged and interlocked in an annular array to form the annular stator vane ring.
Now referring concurrently to Figures 2 to 4, the vane segment 40 each comprise one airfoil 30 integrally formed with, and radially extending between inner and outer platform segments 26 28 in a conventional fashion. The inner and outer platforms 26 and 28 have upstream and downstream edges 42 and 44. The airfoil has a leading edge 46, a trailing edge 48, a pressure side 50 and a suction side 52.
The airfoil 30 is adapted to extend transverse to an axial direction of the annular gas path 32 identified by arrows 54 in Figure 2 and 4. The trailing edge 48 of the airfoil is defined at an acute angle a (Fig. 4) with the axial direction 54. In one embodiment, angle a is between 65-75 degrees with respect to the axial direction 54.
Naturally, other embodiments exist.
25 Still referring to Figure 2, the turbine rotor assembly 22 is shown including a disc 56 drivingly mounted to the engine shaft (not shown) linking the turbine section 18 to the compressor 14. The disc 56 has a forwardly mounted coverplate 58 and carries at its periphery a plurality of circumferentially distributed blades 60 that extend radially outwardly into the annular path 32, one of which is 30 shown in Figure 2. Similar to the vane segments 40, the blades 60 have a leading edge and a trailing edge 62 and 64 and extend from a blade platform 66. The blade platform 66 has an upstream edge 68 and a downstream edge 70. Notably, each blade 60 may extend from an individual blade platform 66.
Thus, the combustion gases of the gas path flow enter the turbine section 18 in the generally axial downstream direction 54 and are redirected at by the trailing edges 46 of the vane airfoils 30 of an upstream vane assembly 24 at an optimum angle a toward the leading edges 62 of the rotating turbine blades 60 of a downstream turbine rotor assembly 22.
Since, the vanes segments 40 and turbine blades 60 are exposed to the hot combustion gases discharging from the combustor 16, it is imperative that these components are cooled. Typically, cooling is accomplished by bleeding air from the compressor 14 and directing it to the turbine section 18 such that the combustor 16 is bypassed. Notably, any other source of coolant may also be used.
A controlled amount of fluid from the cooling air is permitted to re-enter the annular gas path 32 via a labyrinth leakage path identified by arrows 72.
The leakage path 72 is defined between the forward vane assembly 24 and the rotor assembly 22. The fluid flowing through the labyrinth leakage path 72 will be referred to as the purge flow herein. More particularly, the purge flow progresses through the leakage path 72 until introduced into the annular hot gas path 32 such that it comes into contact with parts of the vane assembly 24, the forward surface of the coverplate 58, the rotor disc 56, the blade platform 66 and the blades 60. The purge flow flows through the labyrinth leakage path 72 to purge hot combustion gases that may have migrated into the area between the vane and rotor assemblies 24 and 22, which are detrimental to the cooling system. Thus, the purge flow creates a seal that prevents the entry of the combustion gases from the annular path 32 into the leakage path 72.
A secondary function of the purge flow is to moderate the temperature of the components defining the leakage path 72.
The purge flow is introduced into the annular gas path 32 by passing through a rearward open nozzle 74 defined by the downstream edge 44 of the vane inner platform 26 and the upstream edge of the blade platform 66.
Furthermore, the cooling air from the cooling air source discharged from the compressor is used to cool the vane assembly 24; however other suitable cooling air sources can also be used. Particularly, the inner and outer platforms 26 and 28 that form the radial boundaries of the annular gas path 32 and are subjected to the temperatures of the hot gases passing therethrough are internally cooled by a flow of cooling air. The cooling air flow can be redirected back into the gas path flow through a plurality of holes 76 defined along the downstream edges 44 of the inner and outer platforms 26 and 28. More particularly, the holes 76 are defined at an angle (3 (Fig. 4) substantially parallel to the trailing edge angle a. The holes 76 are adapted to eject a cooling air flow into the annular gas path 32 in the direction of the gas path flow. Thus, the holes 76 provide a fluid flow path for ejecting the platform cooling air flow into the gas path tangential to the gas path flow.
It is to be understood that the term "substantially parallel" encompasses a reasonable range of angles 0 that would be known by a person skilled in the art. In this exemplary embodiment, the holes 76 are defined at angle 0 ranging between degrees greater than or less than the trailing edge angle a of the airfoil 30.
As best seen in Figure 3, the holes 76 are defined on a vertical face 78 of the downstream edges 44 of the inner and outer platforms 26 and 28. The holes are equidistantly spaced along the vertical face 78. Still further, the holes 76 can be jets that increase the velocity of the cooling air flow exiting the inner and outer platforms 26 and 28 and entering the annular gas path 32. The holes 76 acts as jets that tangentially discharge the cooling air flow back into the gas path flow of the turbine engine 10.
Furthermore, in the case of the plurality of holes 76 defined on the downstream edge 44 of the inner platform 26, the cooling air flow being tangentially injected into the gas path flow entrains the purge flow flowing radially and directs it in a general direction of the trailing edge angle a. Therefore, the vane platform cooling air can be reused by discharging it tangentially into the gas path flow to create high momentum jets which entrain the low momentum purge flow and redirect it in the tangential direction of the gas path flow. The purge flow then mixes with the cooling air flow being discharged which consequently causes an increase in the purge flow momentum that better matches the high momentum of the gas path flow and with a relative direction that substantially matches that of the gas path flow.
DESCRIPTION OF THE DRAWINGS
Reference is now made to the accompanying figures depicting aspects of the present invention, in which:
Figure 1 is a schematic side view of a gas turbine engine, in partial cross-section;
Figure 2 is a cross-sectional view of a vane assembly and a rotor assembly of a turbine stage of the gas turbine engine of Fig. 1;
Figure 3 is a perspective view of a vane of the vane assembly of Fig. 2, in accordance with a preferred embodiment of the present invention; and Figure 4 is a transverse sectional view taken along cross-section line 4-4 of Fig. 3, showing an airfoil extending from a platform of the vane, the airfoil having a trailing edge and the platform defining a plurality of angled cooling holes aligned with the trailing edge.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
Figure 1 illustrates a gas turbine engine 10 of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan 12 through which ambient air is propelled, a multistage compressor 14 for pressurizing the air, a combustor 16 in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section 18 for extracting energy from the combustion gases.
The turbine section 18 may comprise one or more turbine stages, in this case two are shown including a first, or high pressure (HP), turbine stage 20, which includes a turbine rotor assembly 22 and a turbine vane assembly 24 in accordance with the present invention.
Referring now to Figure 2, the turbine vane assembly 24 comprises generally a first or inner platform 26 and a second or outer platform 28 and a plurality of circumferentially spaced-apart airfoils 30 extending radially therebetween. The inner platform 26 or shroud and the outer platform 28 or shroud form an annular stator vane ring defining an annular gas path 32 therebetween, with the airfoils 30 extending across the annular gas path 32. More particularly, the inner vane platform 26 sealingly engages the inner combustion chamber wall 34 and the outer vane platform 28 sealingly engages the outer combustion chamber wall 36, thereby defining therebetween the annular path 32 for directing the stream of hot combustion gases exiting the combustion chamber outlet therethrough (not actually shown - inventor should provide complete Figure). Throughout this description, the axial, radial and circumferential directions are defined respectively with respect to the central axis, radius and circumference of the annular stator vane ring.
It should be noted that the turbine vane assembly 24 can be manufactured as a one-piece annular stator vane ring, wherein the inner platform 26, the airfoils 30 and the outer platform 28 are all integrally formed, or may be provided as a plurality of individual vane segments 40 arranged and interlocked in an annular array to form the annular stator vane ring.
Now referring concurrently to Figures 2 to 4, the vane segment 40 each comprise one airfoil 30 integrally formed with, and radially extending between inner and outer platform segments 26 28 in a conventional fashion. The inner and outer platforms 26 and 28 have upstream and downstream edges 42 and 44. The airfoil has a leading edge 46, a trailing edge 48, a pressure side 50 and a suction side 52.
The airfoil 30 is adapted to extend transverse to an axial direction of the annular gas path 32 identified by arrows 54 in Figure 2 and 4. The trailing edge 48 of the airfoil is defined at an acute angle a (Fig. 4) with the axial direction 54. In one embodiment, angle a is between 65-75 degrees with respect to the axial direction 54.
Naturally, other embodiments exist.
25 Still referring to Figure 2, the turbine rotor assembly 22 is shown including a disc 56 drivingly mounted to the engine shaft (not shown) linking the turbine section 18 to the compressor 14. The disc 56 has a forwardly mounted coverplate 58 and carries at its periphery a plurality of circumferentially distributed blades 60 that extend radially outwardly into the annular path 32, one of which is 30 shown in Figure 2. Similar to the vane segments 40, the blades 60 have a leading edge and a trailing edge 62 and 64 and extend from a blade platform 66. The blade platform 66 has an upstream edge 68 and a downstream edge 70. Notably, each blade 60 may extend from an individual blade platform 66.
Thus, the combustion gases of the gas path flow enter the turbine section 18 in the generally axial downstream direction 54 and are redirected at by the trailing edges 46 of the vane airfoils 30 of an upstream vane assembly 24 at an optimum angle a toward the leading edges 62 of the rotating turbine blades 60 of a downstream turbine rotor assembly 22.
Since, the vanes segments 40 and turbine blades 60 are exposed to the hot combustion gases discharging from the combustor 16, it is imperative that these components are cooled. Typically, cooling is accomplished by bleeding air from the compressor 14 and directing it to the turbine section 18 such that the combustor 16 is bypassed. Notably, any other source of coolant may also be used.
A controlled amount of fluid from the cooling air is permitted to re-enter the annular gas path 32 via a labyrinth leakage path identified by arrows 72.
The leakage path 72 is defined between the forward vane assembly 24 and the rotor assembly 22. The fluid flowing through the labyrinth leakage path 72 will be referred to as the purge flow herein. More particularly, the purge flow progresses through the leakage path 72 until introduced into the annular hot gas path 32 such that it comes into contact with parts of the vane assembly 24, the forward surface of the coverplate 58, the rotor disc 56, the blade platform 66 and the blades 60. The purge flow flows through the labyrinth leakage path 72 to purge hot combustion gases that may have migrated into the area between the vane and rotor assemblies 24 and 22, which are detrimental to the cooling system. Thus, the purge flow creates a seal that prevents the entry of the combustion gases from the annular path 32 into the leakage path 72.
A secondary function of the purge flow is to moderate the temperature of the components defining the leakage path 72.
The purge flow is introduced into the annular gas path 32 by passing through a rearward open nozzle 74 defined by the downstream edge 44 of the vane inner platform 26 and the upstream edge of the blade platform 66.
Furthermore, the cooling air from the cooling air source discharged from the compressor is used to cool the vane assembly 24; however other suitable cooling air sources can also be used. Particularly, the inner and outer platforms 26 and 28 that form the radial boundaries of the annular gas path 32 and are subjected to the temperatures of the hot gases passing therethrough are internally cooled by a flow of cooling air. The cooling air flow can be redirected back into the gas path flow through a plurality of holes 76 defined along the downstream edges 44 of the inner and outer platforms 26 and 28. More particularly, the holes 76 are defined at an angle (3 (Fig. 4) substantially parallel to the trailing edge angle a. The holes 76 are adapted to eject a cooling air flow into the annular gas path 32 in the direction of the gas path flow. Thus, the holes 76 provide a fluid flow path for ejecting the platform cooling air flow into the gas path tangential to the gas path flow.
It is to be understood that the term "substantially parallel" encompasses a reasonable range of angles 0 that would be known by a person skilled in the art. In this exemplary embodiment, the holes 76 are defined at angle 0 ranging between degrees greater than or less than the trailing edge angle a of the airfoil 30.
As best seen in Figure 3, the holes 76 are defined on a vertical face 78 of the downstream edges 44 of the inner and outer platforms 26 and 28. The holes are equidistantly spaced along the vertical face 78. Still further, the holes 76 can be jets that increase the velocity of the cooling air flow exiting the inner and outer platforms 26 and 28 and entering the annular gas path 32. The holes 76 acts as jets that tangentially discharge the cooling air flow back into the gas path flow of the turbine engine 10.
Furthermore, in the case of the plurality of holes 76 defined on the downstream edge 44 of the inner platform 26, the cooling air flow being tangentially injected into the gas path flow entrains the purge flow flowing radially and directs it in a general direction of the trailing edge angle a. Therefore, the vane platform cooling air can be reused by discharging it tangentially into the gas path flow to create high momentum jets which entrain the low momentum purge flow and redirect it in the tangential direction of the gas path flow. The purge flow then mixes with the cooling air flow being discharged which consequently causes an increase in the purge flow momentum that better matches the high momentum of the gas path flow and with a relative direction that substantially matches that of the gas path flow.
Advantageously, imparting a velocity to the purge flow that is tangential to the gas path flow significantly reduces the impact on turbine efficiency.
As a result of the holes 76 being substantially aligned with the trailing edges 48 of the airfoils 30, the cooling air flow and the purge flow merge with the hot gas path flow in a more optimal aerodynamic manner thereby reducing inefficiencies caused by colliding air flows. Such improved fluid flow control is advantageous in improving turbine performance. Moreover, the improved cooling air flow and gas path flow mixing is accomplished using a relatively simple and cost effective vane platform cooling arrangement.
The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without department from the scope of the invention disclosed. Still other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.
As a result of the holes 76 being substantially aligned with the trailing edges 48 of the airfoils 30, the cooling air flow and the purge flow merge with the hot gas path flow in a more optimal aerodynamic manner thereby reducing inefficiencies caused by colliding air flows. Such improved fluid flow control is advantageous in improving turbine performance. Moreover, the improved cooling air flow and gas path flow mixing is accomplished using a relatively simple and cost effective vane platform cooling arrangement.
The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without department from the scope of the invention disclosed. Still other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.
Claims (17)
1. A cooled stationary vane, the cooled stationary vane comprising a vane platform having a downstream edge, an airfoil extending radially from said vane platform and having a trailing edge, and at least one coolant exit hole extending through said downstream edge of said vane platform, the at least one coolant exit hole extending along a direction substantially parallel to a trailing edge portion of the airfoil in a plane normal to the airfoil.
2. The cooled stationary vane as defined in claim 1, wherein the at least one coolant exit hole is defined at an angle ranging between about 5 degrees greater than or less than the trailing edge portion.
3. The cooled stationary vane as defined in claim 1, wherein the at least one coolant exit hole is a jet that increases the velocity of the coolant exiting the platform.
4. A vane assembly for a gas turbine engine, comprising a first and a second platform having an upstream edge and a downstream edge and defining an annular gas path therebetween, a plurality of airfoils extending radially between the first and second platforms across the annular gas path and having leading and trailing edges, the trailing edges provided at an angle from an axial direction of the annular gas path, and at least one hole defined in at least one of the platform downstream edges at an angle substantially parallel to the trailing edge angle, the hole adapted to eject a cooling flow into the annular gas path in the direction of the gas path flow.
5. The vane assembly as defined in claim 4, wherein the at least one hole is defined at an angle within a range of about 5 degrees greater than or less than the trailing edge angle.
6. The vane assembly defined in claim 4, wherein the at least one hole includes a plurality of laterally spaced-apart holes.
7. The vane assembly defined in claim 6, wherein the holes are equidistantly spaced along a vertical face of both of the inner and outer platform downstream edges.
8. The vane assembly defined in claim 6, wherein the holes are jets that increase the velocity of the cooling flow exiting the platform and entering the annular gas path.
9. A stator vane assembly in a turbine stage of a gas turbine engine, the vane assembly comprising an inner and an outer shroud having an upstream edge and a downstream edge and defining an annular gas path therebetween, a plurality of airfoils extending radially between the inner and outer shrouds across the annular gas path and having leading and trailing edges, the trailing edges provided at an angle from an axial direction of the annular gas path, and a plurality of holes defined along the downstream edge of the inner shroud at an angle substantially parallel to the trailing edge angle, wherein a cooling flow exiting the holes entrains a purge flow entering the annular gas path adjacent the downstream edge of the inner shroud and redirects the purge flow to enter the annular gas path in a direction substantially tangential to a direction of the gas path flow.
10. The stator vane assembly as defined in claim 9, wherein the holes are defined at an angle within a range of about 5 degrees greater than or less than the trailing edge angle.
11. The stator vane assembly as defined in claim 9, wherein the holes are equidistantly spaced along a vertical face of the downstream edge of the inner shroud.
12. The stator vane assembly as defined in claim 9, wherein the holes are jets that increase the velocity of the cooling air flow exiting the inner shroud and entering the annular gas path.
13. A stator vane for a gas turbine engine comprising:
an upper and a lower platform having an upstream edge and a downstream edge and defining a gas path therebetween;
an airfoil extending between the upper and lower platforms having a leading edge, a trailing edge, a pressure side, a suction side, and adapted to extend transverse to an axial direction of the gas path, the trailing edge of the airfoil defined at an acute angle with the axial direction; and a plurality of holes defined on at least one of the platform downstream edges at an angle substantially parallel to the trailing edge angle, the holes providing a fluid flow path ejecting a platform cooling air flow into the gas path in a direction of the gas path flow.
an upper and a lower platform having an upstream edge and a downstream edge and defining a gas path therebetween;
an airfoil extending between the upper and lower platforms having a leading edge, a trailing edge, a pressure side, a suction side, and adapted to extend transverse to an axial direction of the gas path, the trailing edge of the airfoil defined at an acute angle with the axial direction; and a plurality of holes defined on at least one of the platform downstream edges at an angle substantially parallel to the trailing edge angle, the holes providing a fluid flow path ejecting a platform cooling air flow into the gas path in a direction of the gas path flow.
14. The stator vane as defined in claim 13, wherein the holes are defined at an angle ranging between about 5 degrees greater than or less than the trailing edge.
15. The stator vane as defined in claim 13, wherein the holes are jets that increase the velocity of the cooling air flow exiting the platform and entering the gas path.
16. A method of redirecting a radially flowing purge flow in a direction of a gas path flow in a gas turbine engine comprising:
ejecting a cooling flow from a stator vane platform at an angle substantially parallel to a trailing edge of an airfoil of the stator vane; and entraining the purge flow from the radial direction to a direction tangential to the gas path flow.
ejecting a cooling flow from a stator vane platform at an angle substantially parallel to a trailing edge of an airfoil of the stator vane; and entraining the purge flow from the radial direction to a direction tangential to the gas path flow.
17. A method as defined in claim 16, wherein the cooling flow is ejected from the stator vane platform at an angle ranging between about 5 degrees greater than or less than the direction defined by the trailing edge in a plane normal to the airfoil.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US28641805A | 2005-11-25 | 2005-11-25 | |
| US11/286,418 | 2005-11-25 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA2568692A1 true CA2568692A1 (en) | 2007-05-25 |
Family
ID=38066754
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA 2568692 Abandoned CA2568692A1 (en) | 2005-11-25 | 2006-11-23 | Vane platform tangential injection |
Country Status (1)
| Country | Link |
|---|---|
| CA (1) | CA2568692A1 (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9074484B2 (en) | 2010-09-30 | 2015-07-07 | Rolls-Royce Plc | Cooled rotor blade |
-
2006
- 2006-11-23 CA CA 2568692 patent/CA2568692A1/en not_active Abandoned
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9074484B2 (en) | 2010-09-30 | 2015-07-07 | Rolls-Royce Plc | Cooled rotor blade |
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