Components for gas turbine engines

1. A component for a gas turbine engine, the component comprising:

an airfoil body having a leading edge, a trailing edge, a pressure side, and a suction side, wherein the airfoil defines at least a leading edge cavity located proximate the leading edge and defined between the leading edge and a separator rib of the airfoil body in an axial direction and between the pressure side and the suction side in a circumferential direction;

a counter-flow insert member installed within the leading edge cavity, the insert member having one or more metering flow apertures along a pressure side of the insert member and along a suction side of the insert member and proximate an aft end of the insert member and one or more leading edge impingement apertures at a forward end of the insert member, wherein the insert member is in contact with an supported by the separator rib at the aft end of the insert member; and

a plurality of axially extending and parallel ribs extending along an exterior surface of the insert member from the aft end of the insert member to the forward end of the insert member;

wherein a plurality of segregated and axial extending flow channel passages are defined between the plurality of axial extending ribs and between the exterior surface of the insert member and an interior surface of the airfoil body and extending from the aft end of the insert member to the forward end of the insert member in an axial direction, and wherein the plurality of axial extending ribs are configured to contact the interior surface of the airfoil body to position the insert member within the leading edge cavity and to define the plurality of axial extending flow channel passages, and

wherein air flow flows from an internal insert cavity of the insert member, through the one or more metering flow apertures at the aft end of the insert member, into the plurality of axial extending flow channel passages, and flows forward through the plurality of axial extending flow channel passages toward the leading edge of the airfoil body, and

at least one film cooling aperture formed on the leading edge of the airfoil body to fluidly connect the leading edge cavity to an exterior of the airfoil body, wherein the at least one film cooling aperture causes a pressure differential to draw the airflow through the metering flow apertures at the aft end of the insert member toward the leading edge of the airfoil, and wherein the flow through the one or more metering flow apertures and the flow through the one or more impingement apertures combines at the leading edge with the air from the one or more impingement apertures resupplying the air from the one or more metering flow apertures and then expelled through the at least one film cooling aperture,

wherein the plurality of axial extending flow channel passages do not include fluid connection to the interior of the insert member at locations between the one or more metering flow apertures and the one or more leading edge impingement apertures.

2. The component of claim 1 , further comprising at least one heat transfer augmentation feature formed on the exterior surface of the insert member and within at least one axial extending flow channel passage of the plurality of axial extending flow channel passages.

3. The component of claim 1 , wherein each axial extending rib has a variable radial height in a direction from the aft end to the forward end.

4. The component of claim 1 , wherein the airfoil body is formed of ceramic matrix composites.

5. The component of claim 1 , wherein the insert member is a baffle inserted into the airfoil body.

6. The component of claim 1 , wherein the insert member is a structural spar inserted into the airfoil body.

7. The component of claim 6 , wherein the insert member is integrally formed with a platform.

8. The component of claim 1 , wherein the insert member is formed from sheet metal.

9. The component of claim 1 , wherein the ribs extend from an aft end of the ribs at the aft end of the insert to a forward end of the ribs at the forward end of the insert.

10. A gas turbine engine comprising:

an airfoil having an airfoil body defining a leading edge, a trailing edge, a pressure side, and a suction side, wherein the airfoil defines at least a leading edge cavity located proximate the leading edge and defined between the leading edge and a separator rib of the airfoil body in an axial direction and between the pressure side and the suction side in a circumferential direction;

a counter-flow insert member installed within the leading edge cavity, the insert member having one or more metering flow apertures along a pressure side of the insert member and along a suction side of the insert member and proximate an aft end of the insert member and one or more leading edge impingement apertures at a forward end of the insert member, wherein the insert member is in contact with an supported by the separator rib at the aft end of the insert member; and

a plurality of axially extending and parallel ribs extending along an exterior surface of the insert member from the aft end of the insert member to the forward end of the insert member;

wherein a plurality of segregated and axial extending flow channel passages are defined between the plurality of axial extending ribs and between the exterior surface of the insert member and an interior surface of the airfoil body and extending from the aft end of the insert member to the forward end of the insert member in an axial direction, and wherein the plurality of axial extending ribs are configured to contact the interior surface of the airfoil body to position the insert member within the leading edge cavity and to define the plurality of axial extending flow channel passages, and

wherein air flow flows from an internal insert cavity of the insert member, through the one or more metering flow apertures at the aft end of the insert member, into the plurality of axial extending flow channel passages and flows forward through the plurality of axial extending flow channel passages toward the leading edge of the airfoil body, and

at least one film cooling aperture formed on the leading edge of the airfoil body to fluidly connect the leading edge cavity to an exterior of the airfoil body, wherein the at least one film cooling aperture causes a pressure differential to draw the airflow through the metering flow apertures at the aft end of the insert member toward the leading edge of the airfoil, and wherein the flow through the one or more metering flow apertures and the flow through the one or more impingement apertures combines at the leading edge with the air from the one or more impingement apertures resupplying the air from the one or more metering flow apertures and then expelled through the at least one film cooling aperture,

wherein the plurality of axial extending flow channel passages do not include fluid connection to the interior of the insert member at locations between the one or more metering flow apertures and the one or more leading edge impingement apertures.

11. The gas turbine engine of claim 10 , further comprising at least one heat transfer augmentation feature formed on the exterior surface of the insert member and within at least one axial extending flow channel passage of the plurality of axial extending flow channel passages.

12. The gas turbine engine of claim 10 , wherein each axial extending rib has a variable radial height in a direction from the aft end to the forward end.

13. The gas turbine engine of claim 10 , wherein the airfoil body is formed of ceramic matrix composites.

14. The gas turbine engine of claim 10 , wherein the insert member is a baffle inserted into the airfoil body.

15. The component of claim 10 , wherein the insert member is a structural spar inserted into the airfoil body.

16. The gas turbine engine of claim 15 , wherein the insert member is integrally formed with a platform.