WO2017188991A1

WO2017188991A1 - Staged cooling of a turbine component

Info

Publication number: WO2017188991A1
Application number: PCT/US2016/030111
Authority: WO
Inventors: Nicholas F. MARTIN, Jr.; Allister William James
Original assignee: Siemens Aktiengesellschaft
Priority date: 2016-04-29
Filing date: 2016-04-29
Publication date: 2017-11-02

Abstract

A staged cooling system for turbine component (10) includes a near-wall cooling channel (80) defined between an outer wall (12) of the turbine component (10) and a hollow core (30) positioned in an internal cavity (24) of the turbine component (10). The near-wall cooling channel (80) is made up of multiple fluidically separated channels (82, 84) stacked span-wise, including at least a first upstream channel (82) and a second downstream channel (84). The first upstream channel (82) is fluidically separated from a coolant passage (32) through the hollow core (30). The second downstream channel (84) is fluidically connected to the coolant passage (32) of the core (30) via injection openings (34) formed on the core (30). Coolant flowing though the first upstream channel (82) exits the turbine component (10) via exit openings (38) through the outer wall (12). Coolant flowing through the coolant passage (32) of the core (30) is injected into the second downstream channel (84) via said injection openings (34) on the core (30) at a lower temperature than spent coolant in the first upstream channel (82).

Description

STAGED COOLING OF A TURBINE COMPONENT

BACKGROUND 1. Field

[0001] The present invention relates to gas turbine engines. Specific embodiments relate to a staged cooling of a turbine component, such as a turbine airfoil.

2. Description of the Related Art

[0002] In a turbomachine, such as an axial flow gas turbine engine, air is pressurized in a compressor section and then mixed with fuel and burned in a combustor section to generate hot combustion gases. The hot combustion gases are expanded within a turbine section of the engine where energy is extracted to power the compressor section and to produce useful work, such as turning a generator to produce electricity. The hot combustion gases travel through a series of turbine stages within the turbine section. A turbine stage may include a row of stationary airfoils, i.e., vanes, followed by a row of rotating airfoils, i.e., blades, where the turbine blades extract energy from the hot combustion gases for powering the compressor section and providing output power. Since turbine components, such as airfoils (i.e., vanes and blades) are directly exposed to the hot combustion gases, they are typically provided with an internal cooling circuits that utilize a coolant, such as compressor bleed air.

[0003] In any turbine component, achieving a high cooling efficiency based on the rate of heat transfer is a significant design consideration in order to minimize the volume of coolant air diverted from the compressor for cooling.

SUMMARY [0004] Briefly, aspects of the present invention turbine relate staged cooling of a turbine component.

[0005] According to a first aspect of the invention, a turbine component is provided. The turbine component includes an outer wall delimiting an internal cavity and a hollow core positioned in the cavity, wherein a coolant passage is defined through the hollow core. At least one near- wall cooling channel is formed between an outer surface of the core and an inner surface of the outer wall. The near-wall cooling channel is made up of multiple fluidically separated channels stacked along a span of the outer wall. The multiple stacked channels include at least a first upstream channel and a second downstream channel. The first upstream channel is fluidically separated from the coolant passage of the core and the second downstream channel is fluidically connected to the coolant passage of the core via one or more injection openings formed on the core. Coolant flowing though the first upstream channel exits the turbine component via one or more exit openings through the outer wall. Coolant flowing through the coolant passage of the core is injected into the second downstream channel via said injection openings on the core. The coolant thus injected into the second downstream channel is at a lower temperature than spent coolant in the first upstream channel.

[0006] According to a second aspect of the invention, a turbine airfoil is provided. The turbine airfoil comprises an outer wall delimiting an airfoil interior. The outer wall extending span-wise in a radial direction of a turbine engine and formed of a pressure sidewall and a suction sidewall joined at a leading edge and at a trailing edge. At least one insert is positioned in a cavity in the airfoil interior. The insert is hollow, whereby a coolant passage is defined through the insert. The insert extends along a radial extent of the turbine airfoil and comprises an outer surface that is spaced from an inner surface of the outer wall. A near-wall cooling channel is defined between the outer surface of the insert and the inner surface of the outer wall. The near-wall cooling channel is made up of multiple fluidically separated channels stacked along the radial direction, the multiple stacked channels including at least a first upstream channel and a second downstream channel. The first upstream channel is fluidically separated from the coolant passage of the insert and the second downstream channel is fluidically connected to the coolant passage of the insert via one or more injection openings formed on the insert. Coolant flowing though the first upstream channel exits the turbine airfoil via one or more exit openings through the outer wall. Coolant flowing through the coolant passage of the insert is injected into the second downstream channel via said injection openings on the insert. The coolant thus injected into the second downstream channel is at a lower temperature than spent coolant in the first upstream channel.

[0007] According to a third aspect of the invention, a method for forming a turbine component is provided. The method comprises forming an outer wall, wherein the outer wall delimits an internal cavity. The method further comprises forming a near- wall cooling channel for cooling the outer wall. The near-wall cooling channel is formed by inserting an expandable core into the internal cavity in a first insertable configuration, and subsequently expanding the core to a final expanded configuration. In the final expanded configuration, an outer surface of the core is spaced from an inner surface of the outer wall to form said near-wall cooling channel. The near-wall cooling channel is formed of multiple fluidically separated channels stacked along a span of the outer wall, the multiple stacked channels including at least a first upstream channel and a second downstream channel. The first upstream channel is fluidically separated from the coolant passage of the core. The second downstream channel is fluidically connected to the coolant passage of the core via one or more injection openings formed on the core.

BRIEF DESCRIPTION OF THE DRAWINGS

[0008] The invention is shown in more detail by help of figures. The figures show specific configurations and do not limit the scope of the invention.

[0009] FIG 1 is a schematic cross-sectional view looking in a radial direction through a two- wall airfoil with radial internal cooling channels;

[0010] FIG 2 is a perspective view of an airfoil in which exemplary embodiments of the present invention may be incorporated; [0011] FIG 3 is a schematic partial cross-sectional view through the turbine airfoil looking in a radial direction, prior to insertion of the expandable core;

[0012] FIG 4 is a schematic partial cross-sectional view through the turbine airfoil looking in a radial direction, illustrating an exemplary expandable core in an initial non-expanded configuration;

[0013] FIG 5 is a schematic partial cross-sectional view through the turbine airfoil looking in a radial direction, illustrating the exemplary core in a final expanded configuration;

[0014] FIG 6 is a schematic span-wise cross-sectional view along the section VI- VI, according to an example embodiment of the present invention;

[0015] FIG 7 is a schematic cross-sectional view along the section VII-VII in FIG 6, illustrating a staged cooling system according an example embodiment;

[0016] FIG 8 is a schematic cross-sectional view along the section VIII-VIII in FIG 6, illustrating a sharp edged feature according to an example embodiment; and [0017] FIG 9 is a schematic illustration of pre- formed crack arresting features in a core according to an example embodiment.

DETAILED DESCRIPTION

[0018] In the following detailed description, across different embodiments, like reference characters have been used to designate like or corresponding elements for the sake of simplicity.

[0019] In this description, various specific details are set forth in order to provide a thorough understanding of such embodiments. However, those skilled in the art will understand that disclosed embodiments may be practiced without these specific details, that the present invention is not limited to the depicted embodiments, and that the present invention may be practiced in a variety of altemative embodiments. In other instances, methods, procedures, and components, which would be well- understood by one skilled in the art have not been described in detail to avoid unnecessary and burdensome explanation.

[0020] Furthermore, usage of the phrase "in one embodiment" does not necessarily refer to the same embodiment, although it may. It is noted that disclosed embodiments need not be construed as mutually exclusive embodiments, since aspects of such disclosed embodiments may be appropriately combined by one skilled in the art depending on the needs of a given application.

[0021] The terms "comprising", "including", "having", and the like, as used in the present application, are intended to be synonymous unless otherwise indicated. Also, unless otherwise specified, the connector "or", as used herein, implies an inclusive "or", which is to say that the phrase "A or B" implies: A; or B; or both A and B. Lastly, as used herein, the phrases "configured to" or "arranged to" embrace the concept that the feature preceding the phrases "configured to" or "arranged to" is intentionally and specifically designed or made to act or function in a specific way and should not be construed to mean that the feature just has a capability or suitability to act or function in the specified way, unless so indicated.

[0022] As shown in FIG 1, a typical turbine blade or vane may involve a two-wall structure, including an outer wall 12 which is formed of a pressure sidewall 14 and a suction sidewall 16 joined at a leading edge 18 and at a trailing edge 20. Internal cooling cavities 24 may be created by employing partition walls or ribs 22 which connect the pressure and suction sidewalls 14 and 16. The internal cooling cavities 24 may, for example, conduct coolant in alternating radial directions to form one or more serpentine cooling paths, which may be forward and/or aft flowing. In such a cooling scheme, the coolant fills the entire cavity 24, which may result in a greater coolant requirement than is actually needed to cool the component, because it is generally favorable to maintain a minimum coolant flow momentum in order to keep the flow moving in the desired direction.

[0023] A more efficient use of coolant would be possible if the coolant flow could be largely confined to the area very close to the hot outer wall, at the pressure and suction sidewalls. This effect may be referred to as near- wall cooling. For example, in cast turbine airfoils, near-wall cooling has been achieved by forming near-wall cooling channels between the outer wall and metallic inner walls. In such a multi-wall design, the metallic inner wall is typically connected to the outer wall, which thereby leads to high induced thermal stresses due to difference in thermal loading between the hot outer wall and the relatively colder inner walls. Furthermore, in typical near- wall cooling designs, the coolant is made to traverse substantially the entire span of the turbine airfoil, which extends in a radial direction of the turbine engine. This poses a challenge to maintain the coolant temperature at the end of the radial near-wall passage sufficiently low to be able to effectively absorb heat from the outer wall. The present inventors have devised a mechanism for near- wall cooling which eliminates or minimizes the above-mentioned induced thermal stresses that apply to multi-wall designs, while effectively maintaining the heat transfer capacity of the coolant flow by providing a staged near-wall cooling along the span of the airfoil. Aspects of the present disclosure provide an improvement to the near-wall cooling mechanism disclosed in the co-pending International Application No. PCT/US2015/029673 titled "TURBINE AIRFOIL WITH INTERNAL COOLING SYSTEM HAVING COOLING CHANNELS DEFINED IN PART BY AN INNER BLADDER" filed by the present Applicant, which is herein incorporated by reference in entirety. The inventive concepts described in the present disclosure are however not necessarily limited to turbine airfoils, but may be applied to other cooled turbine components, such as combustion system transitions, and impingement cooled ring segments, among others.

[0024] FIG 2 shows a perspective view of a turbine airfoil 10 where embodiments of the present invention may be incorporated. As illustrated, the airfoil 10 is a turbine blade for a gas turbine engine. It should however be noted that aspects of the invention could additionally be incorporated into stationary vanes in a gas turbine engine. The airfoil 10 comprises an outer wall 12 adapted for use, for example, in a high pressure stage of an axial flow gas turbine engine. The outer wall 12 extends span-wise along a radial direction R of the turbine engine, and is formed of a generally concave pressure sidewall 14 and a generally convex suction sidewall 16 joined at a leading edge 18 and at a trailing edge 20. The outer wall 12 delimits a hollow airfoil interior which may comprise one or more internal cooling channels (not shown in FIG 2) that extend along a radial extent of the airfoil 10. As illustrated, the outer wall 12 may be coupled to a root 56 at a platform 58. The root 56 may couple the turbine airfoil 10 to a disc (not shown) of the turbine engine. The outer wall 12 is delimited in the radial direction by a radially outer end face or airfoil tip 52 and a radially inner end face 54 coupled to the platform 58. In an alternate embodiment, in case of a stationary vane, the radially inner end face of the airfoil 10 may be coupled to the inner diameter of the turbine section of the turbine engine and the radially outer end face of the turbine airfoil 10 may be coupled to the outer diameter of the turbine section of the turbine engine. In the illustrated example, the internal cooling channels of the airfoil 10 may receive a coolant, such as air from a compressor section (not shown), via one or more coolant supply passages (not shown) through the root 56. The coolant traverses through the internal cooling channels, and exits the airfoil 10 via exhaust orifices 27, 29 positioned along the leading edge 18 and the trailing edge 20 respectively. Exhaust orifices 38 may also be provided on the outer wall 12 on one or more locations on the pressure sidewall 14 and/or the suction sidewall 16. Although not shown in the drawings, exhaust orifices may further be provided at the airfoil tip 52. In one embodiment, the airfoil 10, including the outer wall 12, the root 56 and the platform 58 is integrally formed by casting, for example from a ceramic casting core. However, other manufacturing techniques may be used, including, for example, additive manufacturing processes such as 3-D printing, among others. [0025] Similar to the configuration shown in FIG 1, the airfoil 10 of the exemplary embodiment comprises multiple radially extending partition ribs 22, formed integrally with the airfoil outer wall 12. The ribs 22 connect the pressure and suction sidewalls 14 and 16, whereby radial internal cavities 24 are defined between adjacent partition ribs 22. FIGS 3-5 show a schematic partial cross-sectional view through the turbine airfoil 10 looking in a radial direction, showing one of the internal cavities 24. In contrast to the configuration of FIG 1, in this case, a hollow core 30 is inserted into one or more of the internal cavities 24. An outer surface 30a of the core 30 forms a gap with an inner surface 12a of the outer wall 12 to define near- wall cooling channels 80 (see FIG 5). FIG 3 illustrates a configuration prior to the core 30 being inserted into the internal cavity 24. In the illustrated embodiment, the core 30 is an expandable insert. FIG 4 shows the core 30 in an initial inserted configuration, while FIG 5 shows the core 30 in a final expanded configuration. In one embodiment, the core 30 may comprise a bladder that is expanded within the internal cavity 24 of the airfoil 10 by a super plastic forming process. An example of such a process is disclosed in the co-pendmg International Application No. PCT US2015/029673, and will not be described in any further detail herein for the sake of brevity. The core 30 may be formed from a material that is different from the material of the outer wall 12. The material forming the core 30 may have a large plastic deformation range such that the material can be super plastically deformed without fracture. An advantage of a super plastic forming process is that it is applicable to very complex internal shapes.

[0026] As shown in FIGS 3-5, the airfoil 10 may comprise a plurality of span-wise extending internal features or standoff ribs 90 along an inner surface 12a of the outer wall 12. In one embodiment, the standoff ribs 90 may be formed integrally with the outer wall 12, for example, by casting. The standoff ribs 90 extend from the inner surface 12a of the outer wall 12 into a non-bonded contact with the core 30 after the core 30 has been expanded to the final configuration, as shown in FIG 5. In the illustrated embodiment, the hollow core 30 forms a continuous non-linear wall having peaks 42 and troughs 44 defining locally minimum and maximum distances, respectively, from the outer wall 12 of the airfoil 10. The area between the outer surface 30a of the core 30 and the inner surface 12a of the outer wall 12 defines a near-wall cooling channel 80 for coolant flow that extends in a span-wise or radial direction (perpendicular to the plane of FIG 5). The standoff ribs 90 may be provided with non-sharp edges, for example configured as rounded tips 92, to prevent fracturing of the core 30. The non-linear shape of the core 30, as illustrated in FIG 5, creates a narrow flow-cross-section of the near-wall cooling channel 80, which may lead to more efficient usage of the coolant and higher coolant velocities in the near- wall cooling channel 80 to enhance heat transfer with the outer wall 12. Furthermore, the standoff ribs 90 may be shaped and configured as cooling features that may positively influence heat transfer in the near-wall cooling channel 80. Since the core 30 is not bonded to the standoff ribs 90, thermal induced stresses in the airfoil 10 due to differential thermal loading between the outer wall 12 and the core 30 may be minimized or eliminated. [0027] It has been observed from trials that in a super plastic forming process, the core would tear or fracture when the core contacts a sharp edge. The precise definition of "sharp" may be a function of a multiplicity of factors, for example, including the thickness of the core material, the expansion rate of the core material and the depth of the near-wall cooling channel, among others. It was also observed from trials that such tears or fractures could be obviated by rounding off the sharp edges and limiting the channel depth between the standoff ribs. [0028] Now, as the coolant flow enters the near-wall cooling channel, it begins to heat up. As a result, the effectiveness of the coolant continues to diminish until the coolant is no longer able to effectively cool the component. This effect may pose a challenge to the component design and often requires a higher coolant flow to adequately cool the span-wise downstream end of the near-wall cooling channel, at the cost of over-cooling the span-wise upstream end of the near- wall cooling channel. This is where the above-mentioned observations from the super plastically formed core development have led the present inventors to develop a unique mechanism, in which new (or relatively under-heated) coolant flow may be injected into a downstream section of the near-wall cooling channel when the coolant flow in an upstream section of the near- wall cooling channel has reached its temperature limit. According to embodiments of the present invention, said "new" coolant is injected into the downstream section of the near- wall cooling channel via injection openings formed by local sharp-edge failures of the super plastically formed core. Embodiments of the present invention thus provide staged cooling flows in a near- wall cooling configuration, whereby the benefits of near-wall cooling are combined with sharp edge failures of the super plastically formed core, to provide "new" coolant flows to the downstream sections of the near-wall cooling channel.

[0029] In the illustrated embodiments, instead of configuring the hollow core 30 to occupy a dead space in the internal cavity 24, the hollow core 30 is configured such that a coolant passage 32 is defined through the hollow core 30. The core 30 may extend along a span-wise direction of the airfoil 10 from an open first end to a closed second end, the first end being in communication with a coolant supply external to the airfoil 10, such as a compressor section of the turbine engine (not shown).

[0030] Referring to FIG 6, the standoff ribs 90 define multiple near-wall channels 80 arranged next to each other along the chordal axis of the airfoil 10. Each of the near-wall cooling channels 80, individually designated as 80a-h, extends in a span- wise or radial direction. At least one of the near-wall cooling channels 80 is made up of multiple fluidically separated channels, stacked span-wise along the outer wall 12. The flow direction of the coolant is indicated by the arrows K. In the present example, as shown in FIGS 6 and 7, two such span-wise stacked channels are shown, namely a first upstream channel 82 and a second downstream channel 84. A flow blocking feature 86 extends from the inner surface 12a of the outer wall 12 and contacts the core 30, so as to fluidically separate the first upstream channel 82 and the second downstream channel 84. Referring to FIG 7, the first upstream channel 82 is fluidically separated from the coolant passage 32 of the core 30 by the wall of the core, while the second downstream channel 84 is fluidically connected to the coolant passage 32 of the core 30 via one or more injection openings 34 formed on the wall of the core 30. In operation, coolant is supplied from a radial end of the airfoil 10, in this case, via the root 56, to the upstream end of the near-wall cooling channel 80, as well as to the coolant passage 32 of the hollow core 30. Spent coolant in the first upstream channel 82 exits the airfoil 10 via one or more of the exit openings 38 on the outer wall 12. The exit openings 38 are configured as film cooling holes and may provide flow metering for the first upstream channel 82. Coolant flowing through the coolant passage 32 of the core 30 is injected into the second downstream channel 84 via the injection openings 34 on the core 30. Since the core 30 is not in contact with the hot gas path, it remains at a significantly lower temperature than the outer wall 12. As a consequence, the coolant flowing through the coolant passage 32 of the core 12 is at a substantially lower temperature than spent coolant in the first upstream channel 82. "Spent" coolant refers to the coolant in the first upstream channel 82 that is heated up by absorbing heat from the outer wall 12, as well as the core 30, as it travels along the span-wise or radial direction. The coolant injected into the second downstream channel 84 via the injection openings 34 is thus relatively "new" or under-heated and has higher heat absorption capacity than the spent coolant in the first upstream channel 82. Coolant may exit the second downstream channel 84, for example, via exit openings 38 on the outer wall or via exit openings (not shown) at the airfoil tip 52.

[0031] In the illustrated embodiment, the injection openings 34 are formed during the super plastic forming process by local fracture of the core 30. The local fracture is caused by sharp-edged features 88 (see FIG 8) located in the internal cavity 24. As shown in FIGS 6 and 7, the sharp edged features 88 are located downstream of the first upstream channel 82 in relation to a span-wise coolant flow direction. In the shown example, the sharp edged features 88 are positioned outboard of the flow blocking feature 86. In one embodiment, internal features such as the flow blocking features 86 and the sharp edged features 88 may be formed integrally with the airfoil 10, for example by casting. In general, the internal features may have generally non- sharp (for example, rounded) edges contacting the core 30, except where a tear or fracture is desired.

[0032] In one embodiment, tear zones thus formed can be configured to provide radial support points for the core 30 in case of a rotating blade. As a further development, the core 30 may be provided with crack arresting features, to limit the tearing and reduce post forming operational crack propagation. In one embodiment, as shown in FIG 9, the core 30 may be manufactured with pre-formed holes 36 prior to being expanded. The pre-formed holes 36 may be precisely positioned in the unexpanded core, so as to be located around the intended tear zone 37 that eventually forms the injection openings 34 by fracturing during the super plastic forming process.

[0033] The above-described embodiments of the present invention provide improved gas turbine efficiency by reducing the net coolant mass flow with optimized radial staging. To this end, the inventive concepts illustrated above may be extended to near-wall configurations having more than two span-wise stacked channels to provide optimized usage of coolant. In this case, adjacently stacked channels may be fluidically separated, for example, by placing respective flow blocking features 86 therebetween, with a sharp edged feature 88 being positioned downstream of the flow blocking feature 86 to form re-injection openings 34. Moreover, as shown in FIG 6, it may not be necessary for every near-wall cooling channel 80a-h to have a flow blocker-reinjection configuration. The number, if any, of flow blocking features 86 in a near-wall cooling channel may vary, for example, based on the coolant flow effect temperature limit. [0034] While specific embodiments have been described in detail, those with ordinary skill in the art will appreciate that various modifications and alternative to those details could be developed in light of the overall teachings of the disclosure. Accordingly, the particular arrangements disclosed are meant to be illustrative only and not limiting as to the scope of the invention, which is to be given the full breadth of the appended claims, and any and all equivalents thereof.

Claims

1. A turbine component (10) comprising:

an outer wall (12) delimiting an internal cavity (24),

a hollow core (30) positioned in the internal cavity (24), wherein a coolant passage (32) is defined through the hollow core (30),

wherein at least one near- wall cooling channel (80) is formed between an outer surface (30a) of the core (30) and an inner surface (12a) of the outer wall (12), the near-wall cooling channel (80) being made up of multiple fluidically separated channels stacked along a span of the outer wall (12), the multiple stacked channels including at least a first upstream channel (82) and a second downstream channel (84),

wherein the first upstream channel (82) is fluidically separated from the coolant passage (32) of the core (30) and the second downstream channel (84) is fluidically connected to the coolant passage (32) of the core (30) via one or more injection openings (34) formed on the core (30),

wherein coolant flowing though the first upstream channel (82) exits the turbine component (10) via one or more exit openings (38) through the outer wall (12), and coolant flowing through the coolant passage (32) of the core (30) is injected into the second downstream channel (84) via said inj ection openings (34) on the core (30) at a lower temperature than spent coolant in the first upstream channel (82).

2. The turbine component (10) according to claim 1 , wherein the core (30) is an expandable insert, which is expandable from an initial insertable configuration to a final expanded configuration,

whereby the near-wall cooling channel (80) is defined between the inner surface (12a) of the outer wall (12) and the outer surface (30a) of the core (30) in the final expanded configuration.

3. The turbine component (10) according to claim 2, wherein the core (30) is expanded to the final expanded configuration by a super plastic forming process.

4. The turbine component (10) according to claim 3, further comprising one or more sharp edged features (88) located in the internal cavity (24),

the sharp edged features (88) being configured to locally fracture the core (30) during the super plastic forming process to form said injection openings (34).

5. The turbine component (10) according to claim 4, wherein the sharp edged features (88) are positioned downstream of the first upstream channel (82) in a span-wise coolant flow direction.

6. The turbine component (10) according to claim 4, wherein the core (30) comprises pre-formed holes (36) prior to being expanded,

the pre-formed holes (36) being positioned so as to be located around a tear zone (37) formed by said fracturing during the super plastic forming process,

the pre-formed holes (36) being configured to arrest post forming crack propagation at the location of the injection openings (34).

7. The turbine component (10) according to claim 1, wherein the core (30) extends along a span-wise direction of the airfoil (10) from an open first end to a closed second end,

the first end being in communication with a coolant supply external to the turbine component (10).

8. The turbine component (10) according to claim 1, further comprising a flow blocking feature (86) positioned between the first upstream channel (82) and the second downstream channel (84),

the flow blocking feature (86) extending from the outer wall (12) to the core (30) so as to fluidically separate the first upstream channel (82) and the second downstream channel (84).

9. The turbine component (10) according to claim 1, further comprising one or more span-wise extending standoff ribs (90),

the standoff ribs (90) extending from the inner surface (12a) of the outer wall (12) into a non-bonded contact with the core (30).

10. The turbine component (10) according to claim 9, wherein the standoff (90) ribs have non-sharp edges (92) that contact the core (30).

1 1. A turbine airfoil (10) comprising:

an outer wall (12) delimiting an airfoil interior, the outer wall extending (12) span-wise in a radial direction (R) of a turbine engine and formed of a pressure sidewall (14) and a suction sidewall (16) joined at a leading edge (18) and at a trailmg edge (20);

at least one insert (30) positioned in a cavity (24) in the airfoil interior, the insert (30) being hollow such that a coolant passage (32) is defined through the insert (30), the insert (30) extending along a radial extent of the turbine airfoil (10) and comprising an outer surface (30a) that is spaced from an inner surface (12a) of the outer wall (12), to define a near- wall cooling channel (80) therebetween,

the near-wall cooling channel (80) being made up of multiple fluidically separated channels stacked along the radial direction, the multiple stacked channels including at least a first upstream channel (82) and a second downstream channel (84),

wherein the first upstream channel (82) is fluidically separated from the coolant passage (32) of the insert (30) and the second downstream channel (84) is fluidically connected to the coolant passage (32) of the insert (30) via one or more injection openings (34) formed on the insert (30),

wherein coolant flowing though the first upstream channel (82) exits the turbine airfoil (10) via one or more exit openings (38) through the outer wall (12), and wherein coolant flowing through the coolant passage (32) of the insert (30) is injected into the second downstream channel (84) via said injection openings (34) on the insert (30) at a lower temperature than spent coolant in the first upstream channel (82).

12. The turbine airfoil (10) according to claim 11, wherein the insert (30) is expandable from an initial msertable configuration to a final expanded

configuration,

whereby the near-wall cooling channel (80) is defined between the inner surface (12a) of the outer wall (12) and the outer surface (30a) of the insert (30) in the final expanded configuration.

13. The turbine airfoil (10) according to claim 12, wherein the insert (30) is expanded to the final expanded configuration by a super plastic forming process.

14. The turbine (10) airfoil according to claim 13, further comprising one or more sharp edged features (88) located in the airfoil interior downstream of the first upstream channel (82),

the sharp edged features (88) being configured to locally fracture the insert (30) during the super plastic forming process to form said injection openings (34).

15. The turbine airfoil (10) according to claim 11, wherein coolant is supplied from a radial end of the airfoil (10) to the first upstream channel (82) as well as the coolant passage (32) of the insert (30).

16. The turbine (10) airfoil according to claim 11, further comprising one or more span-wise extending standoff ribs (90),

the standoff ribs (90) extending from the inner surface (12a) of the outer wall (12) into a non-bonded contact with the insert (30).

17. The turbine airfoil (10) according to claim 16, wherein the insert (30) forms a continuous non-linear wall having peaks (42) and troughs (44) defining locally minimum and maximum distances, respectively, from the outer wall (12) of the airfoil (10).

18. A method for forming a turbine component (10), comprising:

forming an outer wall (12), wherein that the outer wall (12) delimits an internal cavity (24),

forming a near-wall cooling channel (80) for cooling the outer wall (12) by: inserting an expandable core (30) into the internal cavity (24) in a first insertable configuration, and

expanding the core (30) to a final expanded configuration, wherein in the final expanded configuration, an outer surface (30a) of the core (30) is spaced from an inner surface (12a) of the outer wall (12) to form said near-wall cooling channel (80),

wherein the near-wall cooling channel (80) is formed of multiple fluidically separated channels stacked along a span of the outer wall (12), the multiple stacked channels including at least a first upstream channel (82) and a second downstream channel (84),

wherein the first upstream channel (82) is fluidically separated from the coolant passage (32) of the core (30) and the second downstream channel (84) is fluidically connected to the coolant passage (32) of the core (30) via one or more injection openings (34) formed on the core (30).

19. The method according to claim 18, wherein said expanding is effected by a super plastic forming process.

20. The method according to claim 19, comprising forming a plurality of sharp edged features (88) located in the internal cavity (24),

wherein the inj ection openings (34) are formed during the super plastic forming process by local fracture of the core (30) by the sharp-edged features (88), the sharp edged features (88) being located downstream of the first upstream channel (82) in relation to a span-wise coolant flow direction.

21. The method according to claim 18, comprising forming one or more flow blocking features (86) on the inner surface (12a) of the outer wall (12),

wherein each of the flow blocking features (86) extends from the inner surface

(12a) of the outer wall (12) and contacts the core (30) after the core (30) has been expanded to the final expanded configuration, to fluidically separate adjacent channels (82, 84) of said multiple stacked channels of the near-wall cooling channel (80).

22. The method according to claim 18, comprising forming one or more span-wise extending standoff ribs (90) on the inner surface (12a) of the outer wall (12),

the standoff ribs (90) extending from the inner surface (12a) of the outer wall

(12) into a non-bonded contact with the core (30) after the core (30) has been expanded to the final expanded configuration.