Airfoil profile for a blade in a turbine engine
An apparatus and method for an engine component having a working airflow separated into a cooling airflow and a combustion airflow, the engine component comprising a plurality of circumferentially spaced airfoils rotatable about a centerline defining an axial direction, each airfoil comprising an outer wall bounding an interior and defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a chord-wise direction and extending between a root and a tip along a span-wise direction to define a span length; wherein each of the airfoils have an unguided turning angle to stagger angle ratio defined with respect to a percentage of the span length.
1. An airfoil for a turbine engine, the airfoil comprising:
an outer wall bounding an interior, defining a pressure side and a suction side, extending between a leading edge and a trailing edge, and extending between a root and a tip to define a span length;
wherein the airfoil has an unguided turning angle to stagger angle ratio (UGT:S) defined with respect to a percentage of the span length (SL) bounded by the relationship:
UGT:S=− 0.48 SL+ 0.66±0.1.
2. The airfoil of claim 1 wherein the lower bound and the upper bound of the relationship are separated by a range of ±0.07.
3. The airfoil of claim 1 wherein the lower bound and the upper bound of the relationship are separated by a range of ±0.03.
4. The airfoil of claim 1 further comprising an inflection point along the pressure side.
5. The airfoil of claim 4 wherein at least a portion of the leading edge includes an enlarged portion terminating at the inflection point.
6. The airfoil of claim 5 further comprising cooling passages in the interior.
7. The airfoil of claim 1 wherein the tip is unshrouded.
8. The airfoil of claim 1 wherein the airfoil is a blade.
9. An engine component for a turbine engine having a working airflow separated into a cooling airflow and a combustion airflow, the engine component comprising:
a plurality of circumferentially spaced airfoils rotatable about a centerline defining an axial direction, each airfoil comprising:
an outer wall bounding an interior, defining a pressure side and a suction side, extending between a leading edge and a trailing edge along a chord-wise direction to define a chord line, and extending between a root and a tip along a span-wise direction to define a span length;
wherein each airfoil has an unguided turning angle to stagger angle ratio (UGT:S) defined with respect to a percentage of the span length (SL) bounded by the relationship:
UGT:S=− 0.48 SL+ 0.66±0.1.
10. The engine component of claim 9 wherein the plurality of airfoils have a solidity of between 0.4 and 0.8.
11. The engine component of claim 9 wherein the plurality of airfoils comprises consecutive first and second airfoils spaced from each other to define a throat having a width defined as a minimum distance between the trailing edge of the first airfoil and a point along the suction side of the second airfoil.
12. The engine component of claim 11 wherein the unguided turning angle is a measured angle between a first line tangent to the suction side of the second airfoil at the point and a second line tangent to the trailing edge of the second airfoil along the suction side.
13. The engine component of claim 9 wherein the stagger angle is a measured angle between the chord line and an axial line intersecting the trailing edge and extending parallel to the centerline.
14. The engine component of claim 9 wherein the lower bound and the upper bound of the relationship are separated by a range of ±0.07.
15. The engine component of claim 9 wherein the lower bound and the upper bound of the relationship are separated by a range of ±0.03.
16. The engine component of claim 9 wherein at least one of the plurality of airfoils includes an inflection point along the pressure side.
17. The engine component of claim 16 wherein at least a portion of the leading edge includes an enlarged portion terminating at the inflection point.
18. The engine component of claim 9 wherein the tip is unshrouded.
19. A method of forming an engine component, the method comprising:
forming a turbine disk rotatable about a centerline;
forming a plurality of blades, each blade comprising a wall having a first side and a second side, each side extending between a root and a tip along a radial direction to define a span length; and
forming each of the plurality of blades with an unguided turning to stagger ratio (UGT:S) defined with respect to a percentage of the span length (SL) bounded by the relationship:
UGT:S=− 0.48 SL+ 0.66±0.07.
20. The method of claim 19 further comprising mounting the plurality of blades to the turbine disk with a solidity of between 0.4 and 0.8.