Turbine nozzle segment and a turbine nozzle comprising such a turbine nozzle segment

A nozzle segment for a gas turbine engine comprises an outer band having a cooling air inlet, an inner band having a first cooling air outlet, and a nozzle airfoil comprising a cooling flow passage arranged to receive the cooling air as a cooling air stream. A first channel and a second channel are arranged within the cooling flow passage. A deflector is arranged to divide the cooling air stream into a first cooling air stream in the first channel and a second cooling air stream in the second channel, respectively. The deflector deflects the first cooling air stream obliquely to a suction sidewall in the first channel, wherein the first channel is configured to transport the first cooling air stream along the first channel in a swirly flow.

CROSS REFERENCE TO RELATED APPLICATION(S)

This application claims priority to and benefit of E.P. Patent Application No. 20461510.8 filed Feb. 10, 2020, which is incorporated herein in its entirety.

TECHNICAL FIELD

The present disclosure is related to a nozzle segment for a gas turbine engine. More specifically the present disclosure is related in particular to a nozzle segment for a gas turbine engine comprising a high pressure turbine. The present disclosure is also related to a second stage turbine nozzle of a high pressure turbine of a gas turbine engine, especially for a small gas turbine engine. The present disclosure is further related to a turbine nozzle comprising such a turbine nozzle segment. More specifically the present disclosure is related to a high pressure turbine nozzle comprising such a high pressure turbine nozzle segment.

BACKGROUND

In brief, a gas turbine engine comprises a compressor, a combustor and a turbine. The compressor is provided to intake surrounding air into the engine, compress said air and feed compressed air to the combustor. The compressed air is then mixed in the combustor with the fuel to provide an air-fuel mixture. The air-fuel mixture is burned in the combustor, providing hot high-energy combustion gases. The combustion gases are fed further from the combustor to the turbine for extracting energy from the hot high energy combustion gases to power the gas turbine engine. The combustion gases depleted from energy are then discharged to the atmosphere downstream of the turbine.

The turbine in a modern gas turbine engine comprises a high pressure turbine (HPT) and a low pressure turbine (LPT). The HPT is just downstream of the combustor and is adapted to extract energy from the combustion gases for powering the compressor. The LPT is arranged further downstream of the HPT and said LPT is adapted to extract further energy from the combustion gases for producing output work, such as for example powering an upstream fan in a turbofan engine, a turboprop propeller or a helicopter rotor via a gearbox.

A turbine comprises a stationary nozzle having a row of nozzle airfoils for directing the combustion gases into corresponding row of turbine rotor blades extending radially outward from a supporting rotor disc. In most cases a row of nozzle airfoils is provided as a row of nozzle segments. The nozzle airfoils and the corresponding rotor blades cooperate for extracting energy from the combustion gases. The energy is used to rotate a supporting rotor disc. The supporting rotor disk of the turbine is joined via a corresponding shaft with either a compressor rotor or a fan rotor for rotating the corresponding blades thereof by rotating the supporting disk of the turbine.

The above described components of the gas turbine engine define an annular flow path of combustion gases. The annular flow path extends downstream between the corresponding rows of nozzle airfoils and rotor blades of the turbine. Since the combustion gases fed from the combustor are hot, the components defining the flow path have to be cooled during operation of the engine to ensure intended operational life of the turbine. Because the HPT is arranged in the gas turbine engine essentially as the first component of the turbine downstream of the combustor, said HPT receives the hottest combustion gases. The components of the HPT, therefore, are subjected to extreme temperatures. As a consequence, the components of the HPT have to be adequately configured for enhanced cooling to assure intended operational temperature of said components of the HPT. Because of that, there are specific goals necessary to be achieved in terms of cooling, when designing the HPT.

The components of the turbine of the gas turbine engine, especially the components defining the flow path of the combustion gases, particularly nozzles, are cooled with air. The air for cooling the components of the turbine is provided in a modern gas turbine engine by the compressor, the air being bled from said compressor as compressor air. The air provided by the compressor for cooling the components of the turbine is not used in the combustion of the fuel. Because of that, bleeding air from the compressor for cooling the components of the turbine decreases the overall efficiency of the gas turbine engine. Therefore, in order to minimize decrease of the efficiency of the gas turbine engine the volume of the compressor air bled from the compressor for cooling the turbine, especially the components defining a flow path of the combustion gases, needs to be minimized.

The compressor air for cooling the turbine is provided by the compressor under the maximum pressure. Said compressor air has to be channeled through the components of the turbine of the gas turbine engine, respectively, to cool said components to ensure their intended operational temperature. The flow of the cooling air through the components of the turbine to be cooled causes a pressure loss of the compressor air. In other words, the pressure of the cooling air leaving the components to be cooled is lower than the pressure of the air entering said components. The pressure loss, as discussed above, needs to be minimized to ensure a backflow margin. It means that the pressure of cooling air leaving the components of the turbine to be cooled into the flow path of the combustion gases has to be greater that the pressure of said combustion gases in said flow path. Ensuring the backflow margin prevents the combustion gases to flow back into the components of the turbine disturbing cooling of said components of the turbine of the gas turbine engine, decreasing thereby their operational life.

Therefore, upon development of components of a turbine of a gas turbine engine, especially turbine nozzles, specific design goals need to be achieved, the goals being often competing. One goal is to provide components of the turbine having increased operational life. One of attempts for achieving improved operational life is improving cooling of said component, especially cooling the turbine nozzles. The competing goal is to minimize an amount of cooling air bled from a compressor as compressor air in order to minimize decrease of efficiency of the gas turbine engine. The other goal is to provide also adequate backflow margin to prevent backflow of combustion gases into the component of the turbine to be cooled, especially to prevent backflow of said combustion gases into the turbine nozzles. Because of the above the balanced attitude for designing components of the turbine is necessary.

Some efforts to solve the above specified problems of cooling of components of a turbine have been already made.

The document No. U.S. Pat. No. 5,772,397 discloses a gas turbine vane or blade having an internal structure that allows for cooling under diverse pressure ratios. The vane comprises an inlet passage that communicates with an inner cooling cavity positioned between the air passage and the vane's trailing edge. Disposed within the cavity are deflectors, turning members, ribs, and deflecting pins arranged so as to direct the cooling air through the cavity in a manner that minimizes pressure loss.

The document No. U.S. Pat. No. 3,807,892 discloses a cooled guide blade for a gas turbine. The body of said blade is provided with a first coolant flow path at the front of the blade and a second torturous coolant flow path behind the first flow path. Both flow paths terminate at the trailing edge of the blade. The first path is subjected to a low pressure drop and with a narrow cross-section allows the coolant to flow through at high velocity to obtain a rapid heat transfer. The second path is also subjected to a low pressure drop, but with a larger cross-section in the middle part of the blade body and restrictors at the trailing edge, also allows discharge at high velocity.

The document U.S. Pat. No. 6,884,036 discloses a complementary cooled turbine nozzle. The turbine nozzle includes outer and inner bands integrally joined to a vane having a three-pass serpentine flow circuit between opposite pressure and suction sidewalls. The outer band includes an inlet for channeling cooling air into a first channel of the circuit located behind a leading edge of the vane. A first outlet is disposed in the inner band at the bottom of the first channel where it joins the second channel of the circuit. A second outlet is also disposed in the inner band at the bottom of a third channel of the circuit which is disposed in flow communication with the second channel. The first channel behind the leading edge is smooth except for corresponding rows of first and second turbulators spaced laterally apart. The first turbulators bridge the pressure and suction sidewalls directly behind the leading edge, and the second turbulators are disposed behind the suction sidewall.

Nonetheless, the technical solutions, as disclosed in the above documents, are still insufficient in performance, especially in a small gas turbine engine.

Therefore, there is still a demand for a nozzle segment for a turbine nozzle of a gas turbine engine, especially for a nozzle segment for a HPT for a small gas turbine engine, solving or at least diminishing problems of the technical solutions of the state of the art.

BRIEF DESCRIPTION

An aspect of the present disclosure relates to a nozzle segment for a gas turbine engine. The nozzle comprises an outer band comprising a cooling air inlet, an inner band comprising a first cooling air outlet, and at least one nozzle airfoil comprising a pressure sidewall, a suction sidewall, a leading edge and a trailing edge. The pressure sidewall and the suction sidewall are spaced apart and extend in span between the outer band and the inner band and in chord between a leading edge and a trailing edge, the nozzle airfoil comprises a cooling bores arranged within the trailing edge. The nozzle further comprises a cooling flow passage arranged at least within the nozzle airfoil for cooling said nozzle airfoil with cooling air, wherein the cooling flow passage is in a fluid communication with the cooling air inlet for delivering cooling air into the cooling flow passage as a cooling air stream, and a first integral rib protruding from the inner band and arranged spaced apart from the leading edge for defining a first channel and a second channel within the cooling flow passage. The first channel is arranged along the leading edge and is in a fluid communication with the first cooling air outlet for discharging cooling air from the first channel of the cooling flow passage, the second channel is at least in part arranged along the trailing edge and is in a fluid communication with the cooling bores for discharging cooling air from the second channel of the cooling flow passage. A deflector is arranged at the end of the first integral rib and configured for dividing flow of the cooling air stream received through the cooling air inlet into a first cooling air stream and a second cooling air stream and for directing the first cooling air stream into the first channel and for directing the second cooling air stream into the second channel, respectively. The cooling air stream is received through the cooling air inlet is a cooling air jet stream, wherein the first cooling air stream and the second cooling air stream divided by the deflector are a first cooling air jet stream and the second cooling air jet stream, respectively. The deflector is configured to deflect the first cooling air jet stream obliquely to the suction sidewall in the first channel of the cooling flow passage of the nozzle segment, and the first channel is configured to transport the first cooling air jet stream along said first channel in a swirly flow.

In another aspect, the disclosure relates to a nozzle for a gas turbine engine comprising at least one nozzle segment. The nozzle segment comprises an outer band comprising a cooling air inlet, an inner band comprising a first cooling air outlet, and at least one nozzle airfoil comprising a pressure sidewall, a suction sidewall, a leading edge and a trailing edge. The pressure sidewall and the suction sidewall are spaced apart and extend in span between the outer band and the inner band and in chord between a leading edge and a trailing edge, the nozzle airfoil comprises a cooling bores arranged within the trailing edge. The nozzle further comprises a cooling flow passage arranged at least within the nozzle airfoil for cooling said nozzle airfoil with cooling air, wherein the cooling flow passage is in a fluid communication with the cooling air inlet for delivering cooling air into the cooling flow passage as a cooling air stream, and a first integral rib protruding from the inner band and arranged spaced apart from the leading edge for defining a first channel and a second channel within the cooling flow passage. The first channel is arranged along the leading edge and is in a fluid communication with the first cooling air outlet for discharging cooling air from the first channel of the cooling flow passage, the second channel is at least in part arranged along the trailing edge and is in a fluid communication with the cooling bores for discharging cooling air from the second channel of the cooling flow passage. A deflector is arranged at the end of the first integral rib and configured for dividing flow of the cooling air stream received through the cooling air inlet into a first cooling air stream and a second cooling air stream and for directing the first cooling air stream into the first channel and for directing the second cooling air stream into the second channel, respectively. The cooling air stream is received through the cooling air inlet is a cooling air jet stream, wherein the first cooling air stream and the second cooling air stream divided by the deflector are a first cooling air jet stream and the second cooling air jet stream, respectively. The deflector is configured to deflect the first cooling air jet stream obliquely to the suction sidewall in the first channel of the cooling flow passage of the nozzle segment, and the first channel is configured to transport the first cooling air jet stream along said first channel in a swirly flow.

In yet another aspect, the disclosure relates to a gas turbine engine comprising a nozzle having at least one nozzle segment. The nozzle segment comprises an outer band comprising a cooling air inlet, an inner band comprising a first cooling air outlet, and at least one nozzle airfoil comprising a pressure sidewall, a suction sidewall, a leading edge and a trailing edge. The pressure sidewall and the suction sidewall are spaced apart and extend in span between the outer band and the inner band and in chord between a leading edge and a trailing edge, the nozzle airfoil comprises a cooling bores arranged within the trailing edge. The nozzle further comprises a cooling flow passage arranged at least within the nozzle airfoil for cooling said nozzle airfoil with cooling air, wherein the cooling flow passage is in a fluid communication with the cooling air inlet for delivering cooling air into the cooling flow passage as a cooling air stream, and a first integral rib protruding from the inner band and arranged spaced apart from the leading edge for defining a first channel and a second channel within the cooling flow passage. The first channel is arranged along the leading edge and is in a fluid communication with the first cooling air outlet for discharging cooling air from the first channel of the cooling flow passage, the second channel is at least in part arranged along the trailing edge and is in a fluid communication with the cooling bores for discharging cooling air from the second channel of the cooling flow passage. A deflector is arranged at the end of the first integral rib and configured for dividing flow of the cooling air stream received through the cooling air inlet into a first cooling air stream and a second cooling air stream and for directing the first cooling air stream into the first channel and for directing the second cooling air stream into the second channel, respectively. The cooling air stream is received through the cooling air inlet is a cooling air jet stream, wherein the first cooling air stream and the second cooling air stream divided by the deflector are a first cooling air jet stream and the second cooling air jet stream, respectively. The deflector is configured to deflect the first cooling air jet stream obliquely to the suction sidewall in the first channel of the cooling flow passage of the nozzle segment, and the first channel is configured to transport the first cooling air jet stream along said first channel in a swirly flow.

DETAILED DESCRIPTION

Referring now to the figures,FIG. 1shows a second stage nozzle segment1for a HPT for a gas turbine engine. The nozzle segment1comprises an outer band2, an inner band3and at least one nozzle airfoil10, extending between the outer band2and inner band3, as shown inFIG. 1. The outer band2and the inner band3form an upper part and a lower part5of the nozzle segment1, respectively. The outer band2and the inner band3define a flow path for transporting combustion gases from a combustor of a gas turbine engine downstream. The nozzle airfoils10of the HPT are arranged within said flow path of the combustion gases. The nozzle airfoil10comprises a pressure sidewall7and a suction sidewall8extending in span between the inner band3and the outer band2of the nozzle segment1, as shown inFIG. 1. The pressure sidewall7and the suction sidewall8extend also in chord and are joined together at one side forming a leading edge4and at the opposite side forming a trailing edge6of the nozzle airfoil10, as shown inFIGS. 1, 5 and 7. The nozzle airfoil10comprises mid-chord areas5extending between the leading edge4and the trailing edge6on both sides of said nozzle airfoil10, respectively. The nozzle airfoil10is the main working part of the nozzle segment1. The nozzle airfoils10of a HPT provide the required angular component of momentum of combustion gases flowing through the nozzles, in order to produce useful torque on rotor blades being arranged downstream to the nozzles.

The nozzle segment1comprises an internal cooling system for cooling said nozzle segment1. The internal cooling system comprises a cooling flow passage12for transporting cooling air through the nozzle segment1. The cooling flow passage12is comprised within the outer band2, the nozzle airfoil10and the inner band3, as shown inFIG. 3. The internal cooling system comprises further a cooling air inlet9for feeding cooling air into said system. The cooling air inlet9is in a fluid communication with the cooling flow passage12to feed cooling air into said passage12, as shown inFIGS. 2, 3 and 12 to 14. The cooling air inlet9is configured to feed cooling air into the cooling flow passage12as a cooling air jet stream13. The cooling air inlet9is configured to feed cooling air into the cooling flow passage12at a velocity equal or over about 0.1 Mach, preferably at a velocity in a range of about 0.2 to about 0.3 Mach. The cooling air inlet9is arranged in the outer band2, as shown inFIGS. 1 and 2. The cooling air inlet9has a shape suitable to feed cooling air into the cooling flow passage12in a form of the cooling air jet stream13. In one aspect the cooling air inlet9has cylindrical shape. In another aspect, the cooling air inlet9has conical shape, as shown inFIG. 3. The cooling air inlet9comprises the main axis O extending through said inlet9. The cooling air inlet9is configured to feed cooling air into the cooling flow passage12along said main axis O of said inlet9.

The nozzle segment1comprises a first integral rib15protruding from the inner band3and extending inside the nozzle airfoil10towards the outer band2, as shown inFIGS. 3 and 12. The first integral rib15defines within the cooling flow passage12a first channel18. In other words, the first channel18is defined between the first integral rib15, the pressure sidewall7, the suction sidewall8and the leading edge4, and extends along said leading edge4towards the inner band3, as shown inFIGS. 3, 8 and 12. The first channel18is in a fluid communication with the cooling air inlet9for feeding cooling air into said first channel18. The internal cooling system comprises further a first cooling air outlet24. The first cooling air outlet24is arranged in the inner band3of the nozzle segment1in proximity to the leading edge4, as shown inFIG. 3. The first cooling air outlet24is in a fluid communication with the first channel18for discharging cooling air out of said first channel18of the cooling flow passage12and, therefore, out of the cooling flow passage12of the internal cooling system. The first channel18channels a first cooling air jet stream of the cooling air jet stream13for cooling the leading edge4, as well as the pressure sidewall7and the suction sidewall8in a vicinity of said leading edge4of the nozzle airfoil10.

The first integral rib15defines also within the cooling flow passage12a second channel19. In other words, the second channel19is defined between the first integral rib15, the pressure sidewall7, the suction sidewall8and the trailing edge6. The second channel19is in a fluid communication with the cooling air inlet9for feeding cooling air into said second channel19. The second channel19channels a second cooling air jet stream of the cooling air jet stream13for cooling the remaining part of the nozzle airfoil10. The nozzle segment1comprises a second integral rib16arranged next to the first integral rib15in the direction to the trailing edge6. The second integral rib16protrudes from the outer band2and extends inside the nozzle airfoil10towards the inner band3, as shown inFIGS. 3 and 12. The second integral rib16defines within the second channel19of the cooling flow passage12a first cavity19a. In other words, the first cavity19aof the second channel19is defined between the first integral rib15, the pressure sidewall7, the suction sidewall8and the second integral rib16, and extends along said first and second ribs15,16towards the inner band3, as shown inFIGS. 3, 8 and 12. The first cavity19ais in a fluid communication with the cooling air inlet9at the outer band2for feeding cooling air into said first cavity19aof the second channel19. The first cavity19aof the second channel19channels the second cooling air jet stream of the cooling air jet stream13for cooling the mid-chord areas of the nozzle airfoil10located farther to the leading edge4. The nozzle segment1comprises a third integral rib17arranged next to the second integral rib16in the direction to the trailing edge6. Similarly to the first integral rib15, the third integral rib17protrudes from the inner band3and extends inside the nozzle airfoil10towards the outer band2, as shown inFIGS. 3 and 12. The third integral rib17defines within the second channel19of the cooling flow passage12a second cavity19b. In other words, the second cavity19bof the second channel19is defined between the second integral rib16, the pressure sidewall7, the suction sidewall8and the third integral rib17, and extends along said second and third ribs16,17towards the outer band2, as shown inFIGS. 3, 8 and 12. The second cavity19bis in a fluid communication with the first cavity19aat the inner band3for transferring further cooling air into said second cavity19bof the second channel19. The second cavity19bof the second channel19channels the second cooling air jet stream of the cooling air jet stream13for cooling the mid-chord areas5of the nozzle airfoil10located substantially in the middle of the nozzle airfoil10. The third integral rib17defines also inside the second channel19of the cooling flow passage12a third cavity20. In other words, the third cavity of the second channel19is defined between the third integral rib17, the pressure sidewall7, the suction sidewall8and the trailing edge6, and extends along said trailing edge6towards the inner band3, as shown inFIGS. 3, 8 and 12. The third cavity20is in a fluid communication with the second cavity19bat the outer band2for transferring further cooling air into said third cavity20of the second channel19. The third cavity20of the second channel19channels the second cooling air jet stream of the cooling air jet stream13for cooling the mid-chord areas5of the nozzle airfoil10located close to the trailing edge and said trailing edge6as such. The internal cooling system of the nozzle segment1comprises further cooling bores26. The cooling bores26are in a fluid communication with the second channel19for discharging cooling air out of said second channel19of the cooling flow passage12and, therefore, out of the internal cooling system. More specifically, the cooling bores26are in a fluid communication with the third cavity20of the second channel19. The cooling bores26are provided for cooling the trailing edge6with cooling air from the second air flow discharged through said cooling bores26. The cooling bores26are arranged along the trailing edge6. In an aspect the cooling bores26are arranged one next another in a row along the trailing edge6, as shown inFIGS. 3 and 11. Such an arrangement of the cooling bores26provides uniform discharging of cooling air through the trailing edge6enhancing thereby uniform cooling of said trailing edge6. In another aspect, the second air flow is discharged through slots (not shown). The slots, are arranged in the same way, as described above for the cooling bores26. The slots are used to discharge more amount of cooling air for enhanced cooling the trailing edge6of the nozzle airfoil10, if necessary.

The second channel19of the cooling flow passage12, as described above, comprising the first cavity19a, the second cavity19band the third cavity20, is arranged in a meandered form, as shown inFIGS. 3, 13 and 14. Said meander form of the second channel19provides a long path for a flow of cooling air for efficient cooling the mid-chord areas5and the trailing edge6of the nozzle airfoil10.

The nozzle airfoil10comprises a deflector14for dividing the cooling air jet stream13into the first cooling air jet stream and the second cooling air jet stream and for directing said first and second cooling air jet streams into the first channel18and the second channel19of the cooling flow passage12, respectively. The deflector14is arranged on a distal end of the first integral rib15, as shown inFIGS. 3 and 12. The deflector14preferably extends in span between the pressure sidewall7and the suction sidewall8, as shown inFIG. 5. The deflector14is arranged in proximity of the cooling air inlet9. In particular, the deflector14is positioned substantially in front of and at a distance from the cooling air inlet9, as shown inFIGS. 3 and 12. In other words, the deflector14is configured to face towards the cooling air jet stream13entering the cooling flow passage12through the cooling air inlet9. The deflector14comprises a surface inclined downwards the inner band3, as shown inFIG. 3. The surface of the deflector14is also faced towards the suction sidewall8and the leading edge4, as shown inFIGS. 3, 5 and 6. In one aspect, the surface of the deflector14is concave, as shown inFIGS. 3 and 12. In another aspect the surface of the deflector14is flat. In yet another aspect, the surface of the deflector14is convex. The surface of the deflector14can also have a complex geometry, composed of sub-surfaces of the above mentioned geometries. In yet another aspect the surface of the deflector14comprises several inclined sub-surfaces. The surface of the deflector14provides to said deflector14a function of deflecting of the first cooling air jet stream of the cooling air jet stream13.

The deflector14deflects the first cooling air jet stream to hit the leading edge4and impinge the suction sidewall8of the nozzle airfoil10. The first cooling air jet stream is deflected on the deflector14askew with respect to the direction of flow of the cooling air jet stream13entering the nozzle airfoil10through the cooling air inlet9. Avoiding to be bound by any theory, the geometry of the deflection of the first cooling air jet stream is also discussed conventionally below in two dimensions, i.e., in a lateral direction and in a longitudinal direction of the nozzle airfoil10. The lateral direction is understood in the context of the present disclosure as a general direction defined between the leading edge4and the trailing edge6of the nozzle airfoil10of the turbine nozzle segment1. The lateral direction of the nozzle airfoil10may also be understood as a normal to a radial direction of the turbine, comprising such a nozzle airfoil10. The longitudinal direction is understood in the context of the present disclosure as a general direction defined between the outer band2and the inner band3of the turbine nozzle segment1. The longitudinal direction of the nozzle airfoil10may also be understood as a radial direction of the turbine, comprising such a nozzle airfoil10. Considering the lateral direction of the deflection, the first cooling air jet stream is deflected on the deflector14towards the leading edge4and the suction sidewall8in the vicinity of said leading edge4at an angle α, as shown inFIG. 6. The angle α of the lateral component of the deflection of the first cooling air jet stream is in a range from about 5° to about 60°, preferably in a range from about 15° to about 40°, and more preferably is about 30°. The angle α is defined between the tangent to a parting line PL and the line representing the intended direction of flow of the first cooling air jet stream, the starting point of the tangent being defined at most forward edge EE of the surface of the deflector14, as defined in the lateral cross section of the nozzle airfoil10and shown in detail inFIG. 6. The parting line PL is a geometric line extending within the interior of the nozzle airfoil10in the equal distance from both the pressure sidewall7and the suction sidewall8, as shown inFIG. 6. An intersection of the parting line PL with the wall of the nozzle airfoil10around of the leading edge4defines geometrically an intersection point LE. A Specialist in the field is aware how to determine the parting line PL and the related intersection point LE in the nozzle airfoil10. Considering the longitudinal direction of the deflection, the first cooling air jet stream is also deflected on the deflector14towards the leading edge4and the suction sidewall8in the vicinity of said leading edge4at an angle β, as shown inFIG. 4. The angle β of the longitudinal component of the deflection of the first cooling air jet stream is in a range from about 5° to 60°, preferably in a range from about 10° to 50°, more preferably is about 30°. The angle β is defined between the line perpendicular to the main axis O of the cooling air inlet9and the line representing the direction of flow of the first cooling air jet stream, the starting point of the perpendicular line being defined at most forward edge EE of the surface of the deflector14, as defined in the longitudinal cross section of the nozzle airfoil10and shown in detail inFIG. 4. In a consequence the first cooling air jet stream is directed to the leading edge4and the suction sidewall8in the vicinity of said leading edge4at least at the same level as the deflector14. In an aspect the first cooling air jet stream is deflected below the deflector14, as shown inFIG. 14. The deflector14deflects the first cooling air jet stream so that said first cooling air jet stream hits around the area of the turbine nozzle segment1, wherein the outer band2is 35 joined with the nozzle airfoil10. The area of hitting of the first cooling air jet stream, as discussed above, provides efficient flow of the first cooling air jet stream downstream of the first channel18of the cooling flow passage12. In an aspect, the deflector14deflects the first cooling air jet stream so that said first cooling air jet stream hits the area of the turbine nozzle segment1just below, wherein the outer band2is joined with the nozzle airfoil10, as shown inFIG. 14.

The nozzle airfoil10comprises turbulators21arranged within the first channel18of the cooling flow passage12. The turbulators are provided for stripping of a boundary layer along inner walls of said first channel18and for accelerating and swirling cooling air of the first cooling jet stream within said first channel18, increasing thereby further cooling efficiency of the nozzle airfoil10. The number and the orientation of the turbulators21are selected according to the needs. One parameter taken into consideration when determining the number of turbulators21in the first channel18is a spacing S. The spacing S is defined as p/e, wherein p is a distance between two adjacent turbulators21and e is a height of said turbulators21. As a general rule, the number of turbulators21is increased with the spacing S is decreased. Other parameter taken into consideration when determining the number of turbulators21in the first channel18is a blockage B. The blockage B is defined as e/Dh, wherein e is a height of the turbulators21and Dh is a hydraulic diameter of the first channel18for turbulating. And again, as a general rule, if the blockage B is increased the efficiency of cooling increases and the pressure drop of cooling air in the first channel18is increased. For the first channel18of the turbine nozzle segment1according to non-limiting aspects, the spacing S is in a range from about 4% to about 14%, preferably from about 6% to about 12%, whereas the blockage B is in a range from about 20% to about 40%, preferably is about 30%. The number of the turbulators21for the first channel18of the nozzle airfoil10according to a non-limiting aspect is in the range from 2 to 6. In an aspect the first channel18comprises four turbulators21, as shown inFIGS. 3, 8 and 13. The turbulators21are arranged in the first channel18in an inclined pattern, as shown inFIGS. 3, 8 and 13. The turbulators21are inclined to the first integral rib15at an angle in a range from about 40° to about 90°. In one aspect the turbulators21are inclined at an angle about 45°, as shown inFIG. 3. In another aspect, the turbulators21are inclined at an angle about 60°. In yet another aspect, the turbulators21are inclined at an angle about 90°. The turbulators21are arranged in the first channel18in a staggered pattern, as shown in aFIGS. 8 and 13. In other aspect, the turbulators21are arranged in the first channel18in an in-line pattern (not shown in figures). The turbulators21are arranged in the first channel18in such a pattern that said turbulators21at least partially overlap each other. In one aspect adjacent turbulators21are arranged in the overlapping pattern, as shown inFIGS. 8 and 13. In other aspect, every second turbulators21are arranged in the overlapping pattern. The arrangement of the turbulators21in the first channel18increases further turbulence in the first cooling air jet stream, enhancing thereby cooling efficiency.

The first cavity19aand the second cavity19bof the second channel19of the cooling flow passage12are provided with pin fins23for dissipating the second cooling air jet stream in said second channel19for enhancing further cooling efficiency of the nozzle airfoil10. The pin fins23are longitudinal protrusions extending substantially perpendicularly from a curved inner wall11aof the pressure sidewall7and/or a curved inner wall11bof the suction sidewall8, as shown inFIG. 7. In aspects, the pin fins23extend in span between the inner walls11a,11bof the pressure sidewall7and the suction sidewall8, respectively. The pin fins23are arranged in parallel to each other, as shown inFIGS. 5 and 7. In non-limiting aspects, the pin fins23are arranged further in parallel to the first integral rib15, the second integral rib16and third integral rib17, as shown inFIGS. 5 and 7. A cross section of the pin fins23is selected from a group comprising circular, elliptical, rectangular, pentagonal, hexagonal cross section. The pin fins23have substantially circular cross section, as shown inFIGS. 3 and 12. In an aspect, the pin fins23are arranged in a staggered pattern, as shown inFIG. 3. The staggered arrangement of the pin fins23provides uniform dissipation of the second cooling air jet stream in the second channel19and prevents from forming dead zones in said second channel19. In another aspect, the pin fins23are also arranged in an in-line pattern (not shown in figures).

The first cavity19aand the second cavity19bof the second channel19of cooling flow passage12are provided further with half-pin fins23a. The half-pin fins23aare arranged on inner walls of the first integral rib15, the second integral rib16and third integral rib17, as shown inFIGS. 3 and 12. The half-pin fins23aare provided for stripping the second cooling air jet stream of the inner walls of the first integral rib15, the second integral rib16and third integral rib17in the first cavity19aand the second cavity19b, respectively, enhancing yet further cooling efficiency of the nozzle airfoil10. Analogously to the pin fins3023, the half-pin fins23aare longitudinal protrusions extending substantially perpendicularly from a curved inner wall11aof the pressure sidewall7and/or a curved inner wall11bof the suction sidewall8, along the inner walls of the first integral rib15, the second integral rib16and third integral rib17in the first cavity19aand the second cavity19b, respectively. In aspects the half-pin fins23aextend in span between the inner walls11a,11bof the pressure sidewall7and the suction sidewall8, respectively, as shown inFIG. 13. A half cross section of the half-pin fins23ais selected from a group comprising circular, elliptical, rectangular, pentagonal, hexagonal half cross section. The half-pin fins23ahave substantially half circular cross section, as shown inFIGS. 3 and 12. In an aspect, the half-pin fins23aand the pin fins23are arranged in a staggered pattern, as shown inFIG. 3. In another aspect, the pin fins23aare arranged in an inline pattern with the pin fins23(not shown in figures).

The nozzle airfoil10comprises turbulators22arranged within the third cavity20of the second channel19of the flow cooling passage12. Analogously to the turbulators21of the first channel18, the turbulators22of the second channel19are provided for stripping of a boundary layer along inner walls of the third cavity20and for accelerating and swirling cooling air of the second cooling jet stream within the second channel19, increasing thereby further cooling efficiency of the nozzle airfoil10. The number and the orientation of the turbulators21are selected according to the needs. The parameters taken into consideration when determining the number of the turbulators22in the third cavity20of the second channel19are spacing S and blockage B. The spacing S and blockage B are defined in the same way and comply the same relationships, as described in detail above for the turbulators21of the first channel18. For the third cavity20of the second channel19of the turbine nozzle segment1according to non-limiting aspects, the spacing S is in a range from about 4% to about 14%, preferably from about 6% to about 12%, whereas the blockage B is in a range from about 30% to about 40%, preferably is about 30%. The number of the turbulators22for the third cavity20of the second channel19of the nozzle airfoil10according to non-limiting aspects is in the range 8 to 16, more preferably is in a range from10to14. In an aspect, the third cavity20of the second channel19comprises twelve turbulators22, as shown inFIGS. 3, 8 and 13. The turbulators22are arranged in the third cavity20of the second channel19in an inclined pattern, as shown inFIGS. 3, 8 and 13. The turbulators22are inclined to the third integral rib17at an angle in a range from about 40° to about 90°. In one aspect, the turbulators22are inclined at an angle about 60°, as shown inFIG. 3. In another aspect, the turbulators22are inclined at an angle about 45°. In yet another aspect, the turbulators22are inclined at an angle about 90°. The turbulators22are arranged in the third cavity20in a staggered pattern, as shown inFIGS. 10 and 13. In other aspects, the turbulators22are arranged in the third cavity20of the second channel19in an in-line pattern (not shown in figures). The turbulators22are arranged in the third cavity20in such a pattern that said turbulators22at least partially overlap each other. In one aspect, adjacent turbulators22are arranged in the overlapping pattern, as shown inFIG. 10. In other aspect, every second turbulators22are arranged in the overlapping pattern. The arrangement of the turbulators22in the third cavity20of the second channel19increases further turbulence in the second cooling air jet stream, enhancing thereby cooling efficiency of the nozzle airfoil10.

In an aspect, the nozzle airfoil10is provided with a vertical turbulator29. The vertical turbulator29is arranged in the second channel19of the internal cooling system. The vertical turbulator29extends from the third integral rib17towards the outer band2of the nozzle segment1between the second cavity19band the third cavity20of the second channel19, respectively, as shown inFIG. 3. The vertical turbulator29is arranged on the inner wall of the suction sidewall8. The vertical turbulator29is provided for turbulating the second cooling air jet stream within the end of the second cavity19band the beginning of third cavity20of the second channel19, enhancing thereby cooling efficiency of the suction sidewall8within the vertical turbulator29.

In an aspect, the first cooling air outlet24is an ejector for providing purging air to a cavity of an inter-stage disc (not shown in figures) for cooling rotating parts of the turbine. The ejector comprises an outlet at an angle optimized to control circumferential velocity of a cooling air flow in a rotor cavity to decrease friction losses called as “windage losses”.

In an aspect, the nozzle airfoil10comprises a set of outlet holes. The outlet holes are arranged on the leading edge4, the pressure sidewall7and the suction sidewall8. The outlet holes are in a fluid communication with the first channel18and/or the second channel19of the cooling flow passage12, respectively. Cooling air discharged through the outlet holes forms a thin film of the cooling air on the leading edge4, the pressure sidewall7and the suction sidewall8, respectively, enhancing thereby cooling the external surface of the nozzle airfoil10. In an aspect, the outlet holes are arranged on the leading edge4and provide “showerhead cooling”.

Generally, cooling air for cooling the turbine nozzle segment1is provided by the compressor of the gas turbine engine. Cooling air is provided to the outer band2by a feeding system for delivering cooling air, the feeding system being arranged upstream to the turbine nozzle segment1(not shown in figures). The cooling air inlet9is in a fluid communication with the cooling air feeding system. In one aspect, cooling air is provided at the cooling air inlet9, wherein said cooling air inlet9is configured to feed cooling air into the cooling flow passage12at a velocity, as specified above. In another aspect, cooling air is provided already as a cooling air jet stream having a velocity, as specified above. In an aspect, the cooling air feeding system is the system disclosed in the US patent application No. US 2019/0071994 A1, of the present Applicant, the disclosure of which in its entirety is enclosed herewith by way of a reference.

When the turbine nozzle segment1is assembled in the turbine of the gas turbine engine and operating, cooling air is provided to the cooling air inlet9in the outer band2for feeding said cooling air into said turbine nozzle segment1. In one aspect, cooling air is fed to the cooling air inlet9as a cooling air flow and accelerated by said cooling air inlet9to high velocity forming a cooling air jet stream. In another aspect, cooling air is provided to the cooling air inlet9already as a cooling air jet stream having high velocity.

Cooling air is transported through the cooling air inlet9into the cooling flow passage12of the internal cooling system as the cooling air jet stream13at a velocity equal or over about 0.1 Mach, preferably at a velocity in a range of about 0.2 to about 0.3 Mach. The cooling air jet stream13is fed into the cooling flow passage12through the cooling air inlet9substantially along the main axis O. In an aspect, the cooling air jet stream13is fed into the cooling flow passage12along the main axis O of the cooling air inlet9, said main axis O being parallel to the longitudinal direction of the nozzle airfoil10. The cooling air jet stream13is directed on the deflector14. The cooling air jet stream13hits the deflector14and is divided into the first cooling air jet stream and the second cooling air jet stream.

The first cooling air jet stream is directed into the first channel18of the cooling flow passage12of the turbine nozzle segment1. The first cooling air jet stream is deflected on the surface of the deflector14and directed into the first channel18so that said first cooling air jet stream hits the leading edge4and the suction sidewall8in a vicinity of said leading edge4. The first cooling air jet stream is deflected on the deflector14at an angle α and hits the leading edge4and the suction sidewall8in the vicinity of said leading edge4inside the first channel of the cooling flow passage12of the nozzle segment1. As a consequence, the first cooling air jet stream hits the suction side wall8at an acute angle, as considered in a lateral cross section of the nozzle airfoil10and as shown inFIG. 6. The angle α is in a range from about 5° to about 60°, preferably 15° to about 40°, more preferably is about 30°. At the same time the first cooling air jet stream is deflected on the deflector14at an angle β and hits the leading edge4and the suction sidewall8in the vicinity of said leading edge4inside the first channel18of the cooling flow passage12of the nozzle segment1at least at the level of the deflector14. As a consequence, the first cooling air jet stream hits the suction side wall8at an acute angle as well, as considered in a longitudinal cross section of the nozzle airfoil10and as shown inFIG. 4. The angle β is in a range from about 5° to about 60°, preferably 10° to about 50°, more preferably is about 30°. The first cooling air jet stream hits, therefore, the suction sidewall8obliquely, as seen inFIGS. 4 and 6. In an aspect, the first cooling air jet stream hits, therefore, the suction sidewall8of the first cooling air jet stream hits obliquely the suction sidewall8and below the deflector14, as seen inFIGS. 4, 6 and 14. The configuration of impinging of the first cooling air jet stream promotes further flow of said first cooling air jet stream downstream of the first channel18of the cooling flow passage12of the nozzle segment1. The first cooling air jet stream impinging at high velocity the leading edge4and the suction sidewall8is further transported along the first channel18with swirling motion, as shown inFIG. 14. In other words, feeding the first cooling air jet stream into the first channel18, as described above, generates turbulent and swirling flow of cooling air along said first channel18, cooling thereby efficiently hot areas of the nozzle airfoil10, namely hot areas located on the leading edge4, neighboring pressure sidewall7and suction sidewall8. The swirling flow of the first cooling air jet stream provides intense and uniform mixing of cooling air in the first channel18providing efficient cooling. The first cooling air jet stream flowing along the first channel18is also turbulated on the turbulators21arranged in said first channel18. The turbulators21cause stripping the boundary layer of cooling air of the inner walls of the first channel18, enhancing thereby further cooling efficiency within said first channel18.

Cooling air of the first cooling air jet stream exits the first channel18of the cooling flow passage12through the first cooling air outlet24in the inner band3. Cooling air of the first cooling air jet stream is directed to other components of the turbine of the gas turbine engine. In one aspect, the first cooling jet stream is discharged through the first cooling air outlet24being the ejector, as described above, and directed to a rotor cavity as a cooling air flow for decreasing therein windage losses.

The second cooling air jet stream is directed into the second channel19of the cooling flow passage12of the turbine nozzle segment1, as shown inFIG. 14. The second cooling air jet stream enters the first cavity19aof the second channel19first. The second cooling air jet stream dissipates in the first cavity19aon the pin fins23and the half-pin fins23a. Cooling air entering the first cavity19ain the second jet stream and flowing around the pin fins23and the half-pin fins23ais transported along said first cavity19awith local whirls and changes of flow directions of said cooling air in the second cooling air jet stream, providing thereby efficient cooling of the mid-chord areas5of the nozzle airfoil10located farther to the leading edge4. The half-pin fins23acause also stripping the boundary layer of cooling air of the inner walls of the first integral rib15and the second integral rib16, enhancing thereby further cooling efficiency within said first cavity19aof the second channel19.

The second cooling air jet stream exits the first cavity19aand meanders into the second cavity19bof the second channel19of the cooling flow passage12. The second cooling air jet stream dissipates in the second cavity19bon the pin fins23and the half-pin fins23ain the same way as in the first cavity19a. Namely, cooling air entering the second cavity19bin the second cooling air jet stream and flowing around the pin fins23and the half-pin fins23ais transported along said second cavity19bwith local whirls and changes of flow directions of said cooling air in the second cooling air jet stream, providing thereby efficient cooling the mid-chord areas5of the nozzle airfoil10located substantially in the middle of the nozzle airfoil10. Analogously, the half-pin fins23acause also stripping the boundary layer of cooling air of the inner walls of the second integral rib16and the third integral rib17, further enhancing thereby cooling efficiency within said second cavity19bof the second channel19.

The cooling air jet stream exits the second cavity19band meanders farther into the third cavity20of the second channel19of the cooling flow passage12. The second cooling air jet stream enters the third cavity20at a velocity lower than the velocity of said the second cooling air jet stream at the entrance to the first cavity19aof the second channel19. Therefore, formally cooling air of the second cooling air jet stream entering the second channel19of the internal cooling system enters the third cavity20of said second channel19as the second cooling air stream. Cooling air of the second cooling stream flows farther along the third cavity20of the second channel19. The second cooling air stream flowing along the third cavity20of the second channel19turbulates also on the turbulators22arranged in said third cavity20. Analogously to the turbulators21in the first channel18, the turbulators22of the second channel19cause stripping the boundary layer of cooling air of the inner walls of the third cavity20, enhancing thereby cooling efficiency within said third cavity20.

Cooling air of the second cooling air stream exits the third cavity20of the second channel19of the nozzle airfoil10through the cooling bores26arranged in the trailing edge6, cooling thereby efficiently another of the hot areas of said nozzle airfoil10, namely the hot areas located on the trailing edge6, and neighboring pressure sidewall7and suction sidewall8.

The meandering second channel19of the cooling flow passage12of the internal cooling system, as described above, provides even distribution of flowing cooling air within said second channel19, providing thereby efficient cooling of the turbine nozzle segment1, especially the nozzle airfoil10.

In an aspect, comprising the vertical turbulator29, as described above, cooling air traveling from the second cavity19bto the third cavity20of the second channel19turbulates on said vertical turbulator29, enhancing thereby further cooling efficiency of the suction sidewall8around the vertical turbulator29.

In another aspect, comprising set of outlet holes, as described above, cooling air is discharged from the first channel18and/or the second channel19of the cooling flow passage12through said outlet holes and forms a thin film on the external walls of the nozzle airfoil10. The thin film of cooling air flows substantially laminarly on the external walls of the nozzle airfoil10, preventing said external walls from contact with hot combustion gases and providing a showerhead cooling of the nozzle airfoil10.

In one aspect, the turbine nozzle segment1is manufactured as a singlet. The singlet comprises an outer band2, an inner band3and a nozzle airfoil10, as described above and shown in figures. In other aspect, the turbine nozzle segment1is manufactured as a doublet (not shown in figures). The doublet comprises an outer band2, and inner band3and two nozzle airfoils10. The outer band2and inner band3, respectively, are common for the two airfoils10in the doublet. The nozzle airfoils10in the doublet extend radially between the outer band2and the inner band3in a neighboring arrangement. In yet other aspect, the turbine nozzle segment1is manufactured as a triplet (not shown in figures). The triplet comprises an outer band2, and inner band3and three nozzle airfoils10. The outer band2and inner band3, respectively, are common for all the nozzle airfoils10in the triplet. Analogously, the nozzle airfoils10in the triplet extend radially between the outer band2and the inner band3in an neighboring arrangement. The nozzle airfoils10in the doublet and triplet are as described above and shown in figures.

The turbine airfoil10according to an aspect are assembled into a turbine nozzle for a HPT. The turbine nozzle for the HPT comprising the turbine nozzle segments1according to an aspect is shown inFIG. 15. The turbine nozzle according to an aspect is used in the gas turbine engine in the second stage of the HPT.

Providing cooling air into the cooling flow passage12of the turbine nozzle segment1in a form of a cooling air jet stream13and directing divided therefrom the first cooling air jet stream to the areas of the first channel18of said segment, as described above, wherein cooling air of the first cooling air jet stream transported along said the first channel18in a swirled flow, as described above, provides highly efficient cooling of the turbine nozzle, especially highly efficient cooling of the nozzle airfoils10. The turbine nozzle segment1is effectively usable in a turbine nozzle in a contact with the hottest combustion gases exhausted from a combustor. Therefore, the turbine nozzle segment1is especially usable for a turbine nozzle for a HPT. The turbine nozzle segment1according to an aspect is effectively cooled down to the operational temperatures, ensuring thereby intended operational life of a turbine of a gas turbine engine. Because of the improved cooling of the turbine nozzle segment1, less amount of cooling air is needed to maintain the operational temperature of the turbine nozzle, especially the nozzle airfoil10of said turbine nozzle. Therefore, less cooling air for cooling the turbine is necessary to bled from a compressor, reducing thereby loss of efficiency of a gas turbine engine. Because of the improved cooling of the turbine nozzle segment1, a turbine nozzle is successfully downsized and is usable in a small gas turbine engine.

To the extent not already described, the different features and structures of the various aspects can be used in combination with each other as desired. That one feature is not illustrated in all of the aspects is not meant to be construed that it may not be, but is done for brevity of description. Thus, the various features of the different aspects may be mixed and matched as desired to form new aspects, whether or not the new aspects are expressly described. All combinations or permutations of features described herein are covered by this disclosure.

Various characteristics, aspects and advantages of the present disclosure may also be embodied in any permutation of aspects of the disclosure, including but not limited to the following technical solutions as defined in the enumerated aspects:

A nozzle segment for a gas turbine engine comprising: an outer band comprising a cooling air inlet, an inner band comprising a first cooling air outlet, at least one nozzle airfoil comprising a pressure sidewall, a suction sidewall, a leading edge and a trailing edge, wherein the pressure sidewall and the suction sidewall are spaced apart and extend in span between the outer band and the inner band and in chord between a leading edge and a trailing edge, the nozzle airfoil comprises a cooling bores arranged within the trailing edge, a cooling flow passage arranged at least within the nozzle airfoil for cooling said nozzle airfoil with cooling air, wherein the cooling flow passage is in a fluid communication with the cooling air inlet for delivering cooling air into the cooling flow passage as a cooling air stream, a first integral rib protruding from the inner band and arranged spaced apart from the leading edge for defining a first channel and a second channel within the cooling flow passage, the first channel is arranged along the leading edge and is in a fluid communication with the first cooling air outlet for discharging cooling air from the first channel of the cooling flow passage, the second channel is at least in part arranged along the trailing edge and is in a fluid communication with the cooling bores for discharging cooling air from the second channel of the cooling flow passage, a deflector arranged at the end of the first integral rib and configured for dividing flow of the cooling air stream received through the cooling air inlet into a first cooling air stream and a second cooling air stream and for directing the first cooling air stream into the first channel and for directing the second cooling air stream into the second channel, respectively, characterized in that the cooling air stream received through the cooling air inlet is a cooling air jet stream, wherein the first cooling air stream and the second cooling air stream divided by the deflector are a first cooling air jet stream and the second cooling air jet stream, respectively, wherein the deflector is configured to deflect the first cooling air jet stream obliquely to the suction sidewall in the first channel of the cooling flow passage of the nozzle segment, and the first channel is configured to transport the first cooling air jet stream along said first channel in a swirly flow.

The nozzle segment in accordance with the preceding clause, wherein the first cooling air jet stream is deflected by the deflector to impinge the suction sidewall in the first channel at an acute angle laterally and at an acute angle longitudinally, respectively.

The nozzle segment in accordance with any preceding clause, wherein the first cooling air jet stream is deflected by deflector towards the suction sidewall in the first channel at an longitudinal angle (α) and independently at lateral angle (β) with respect to the deflector.

The nozzle segment in accordance with any preceding clause, wherein the angle (α) is in a range from about 5° to about 60°, preferably from about 15° to about 40°, more preferably is about 30°, and independently the angle (β) is in a range from about 5° to about 60°, preferably 10° to about 50°, more preferably is about 30°.

The nozzle segment in accordance with any preceding clause, wherein the cooling air jet stream impinging the deflector has a velocity equal or over about 0.1 Mach, preferably is in a range of about 0.2 to about 0.3 Mach.

The nozzle segment in accordance with any preceding clause, comprising further a second integral rib protruding from the outer band and arranged spaced apart from the first integral rib defining a first cavity in the second channel, the first cavity is in a fluid communication with the cooling air inlet, and the third integral rib protruding from the inner band defining a second cavity and a third cavity, the second cavity is in a fluid communication with the first cavity, the third cavity is in a fluid communication with the second cavity and the third cavity is in a fluid communication with the cooling bores on the trailing edge of the nozzle airfoil.

The nozzle segment in accordance with any preceding clause, wherein the first cavity (19a), the second cavity and the third cavity are arranged within the second channel in a meandric arrangement.

A nozzle for a gas turbine engine comprising at least one nozzle segment according to any preceding clause.

A gas turbine engine comprising a nozzle according to the preceding clause.