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A method for making a gas turbine engine ceramic matrix composite airfoil is disclosed. The method includes fabricating an airfoil preform that has a slotted forward end and a continuous trailing end. The slotted forward end of the airfoil preform is coupled to an airfoil core insert. A ceramic matr

A method for making a gas turbine engine ceramic matrix composite airfoil is disclosed. The method includes fabricating an airfoil preform that has a slotted forward end and a continuous trailing end. The slotted forward end of the airfoil preform is coupled to an airfoil core insert. A ceramic matrix composite covering is applied to cover the slots of the airfoil perform. The continuous trailing end of the airfoil preform is removed to expose the slots. A gas turbine engine airfoil is also disclosed.

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1. A method comprising: fabricating an airfoil preform having a slotted forward end and a continuous trailing end;coupling the slotted forward end of the airfoil preform to an airfoil core insert;applying a ceramic matrix composite covering to cover the slots of the airfoil preform; andremoving the

1. A method comprising: fabricating an airfoil preform having a slotted forward end and a continuous trailing end;coupling the slotted forward end of the airfoil preform to an airfoil core insert;applying a ceramic matrix composite covering to cover the slots of the airfoil preform; andremoving the continuous trailing end of the airfoil preform to expose the slots;wherein:the fabricating step comprises providing a cooling passage preform having a spanwise extending trailing end portion and a plurality of flow dividing members projecting from the spanwise extending trailing end portion that define cooling passages therebetween;the coupling step comprises coupling a cooling delivery core to a forward end of the cooling passage preform to close a forward end of the cooling passages;the applying step comprises covering the cooling passages with a ceramic matrix material; andthe removing step comprises trimming the spanwise extending trailing end portion of the cooling passage preform to expose the cooling passages. 2. The method of claim 1 in which the airfoil preform comprises a monolithic ceramic. 3. The method of claim 1 in which the slots are substantially perpendicular to the continuous trailing end of the airfoil preform. 4. The method of claim 1 in which the slots extend through the thickness of the airfoil preform. 5. The method of claim 1 in which the slots are formed by machining material from the airfoil preform. 6. The method of claim 1 in which the coupling comprises capturing the slotted forward end of the airfoil preform in a spanwise groove in the trailing end of the airfoil core insert. 7. The method of claim 1 in which adjacent ones of the plurality of flow dividing members define cooling passages therebetween that have a substantially linear configuration in the chordwise direction. 8. The method of claim 1 in which the cooling passages are substantially the same size. 9. The method of claim 1 in which the cooling passages are equally spaced apart in the spanwise direction. 10. The method of claim 1 in which the cooling passages have a substantially uniform width. 11. A method comprising: fabricating an airfoil preform as a cooling passage preform having a spanwise extending trailing end portion and a plurality of flow dividing members projecting from the spanwise extending trailing end portion towards a forward end and defining cooling passages therebetween;coupling a cooling delivery core to the forward end of the cooling passage preform to close a forward end of the cooling passages;covering the cooling passages with a ceramic matrix material; andtrimming the spanwise extending trailing end portion of the cooling passage preform to expose the cooling passages. 12. The method of claim 11, wherein the airfoil preform comprises a monolithic ceramic. 13. The method of claim 11, wherein the cooling passages are substantially perpendicular to the trailing end portion of the airfoil preform. 14. The method of claim 11, wherein the cooling passages extend as slots through the thickness of the airfoil preform. 15. The method of claim 11, wherein the cooling passages are formed by machining material from the airfoil preform. 16. The method of claim 11, wherein the coupling comprises capturing the forward end of the airfoil preform in a spanwise groove of the cooling delivery core. 17. The method of claim 11, wherein the plurality of flow dividing members are substantially linear in the chordwise direction. 18. The method of claim 11, wherein the cooling passages are substantially the same size. 19. The method of claim 11, wherein the cooling passages are equally spaced apart. 20. The method of claim 11, wherein the cooling delivery core is formed as at least a portion of an airfoil core insert.