Full-span forward swept airfoils for gas turbine engines
Rotor of a gas turbine engines having a rotor hub and a plurality of blades extending from the rotor hub, wherein each blade has a full-span forward sweep along a leading edge of the blade that starts at an airfoil root of the blade at the hub and extends to a blade tip, wherein a sweep of a blade is a percentage of a root axial chord length of the respective blade.
1. A rotor of a gas turbine engine comprising:
a rotor hub; and
a plurality of blades extending from the rotor hub, wherein each blade has a full-span forward sweep along a leading edge of the blade that starts at a root leading edge at an airfoil root of the blade at the hub and extends to a tip leading edge at a blade tip, wherein a sweep of a blade is a percentage of a root axial chord length of the respective blade, wherein the leading edge sweep is defined as an axial distance of the tip leading edge from the root leading edge in a direction of the root axial chord length,
wherein each blade has a full-span forward sweep along a trailing edge of the blade that starts at the airfoil root of the blade at the hub and extends to the blade tip, and
wherein the leading edge has a sweep of 14% of the root axial chord length or greater and the trailing edge has a sweep of less than 25% of the root axial chord length.
2. The rotor of claim 1 , wherein the full-span forward sweep along the leading edge has a sweep of between 14% and 25% of the root axial chord length of the blade.
3. The rotor of claim 1 , wherein the trailing edge of each blade has the same sweep as the leading edge.
4. The rotor of claim 1 , wherein the trailing edge of each blade has a sweep that is less than the sweep of the leading edge.
5. The rotor of claim 1 , wherein the leading edge has a sweep of 14% of the root axial chord length and the trailing edge has a sweep of 18% of the root axial chord length.
6. A gas turbine engine comprising:
a plurality of stator portions each having a plurality of vanes; and
a plurality of rotor portions, wherein the stator portions and the rotor portions alternate to form a compressor section of the gas turbine engine, wherein each rotor portion comprises:
a rotor hub; and
a plurality of blades extending from the rotor hub, wherein each blade has a full-span forward sweep along a leading edge of the blade that starts a root leading edge at an airfoil root of the blade at the hub and extends to a tip leading edge at a blade tip, wherein a sweep of a blade is a percentage of a root axial chord length of the respective blade, wherein the leading edge sweep is defined as an axial distance of the tip leading edge from the root leading edge in a direction of the root axial chord length,
wherein each blade has a full-span forward sweep along a trailing edge of the blade that starts at the airfoil root of the blade at the hub and extends to the blade tip, and
wherein the leading edge has a sweep of 14% of the root axial chord length or greater and the trailing edge has a sweep of less than 25% of the root axial chord length.
7. The gas turbine engine of claim 6 , wherein the full-span forward sweep along the leading edge has a sweep of between 14% and 25% of the root axial chord length of the blade.
8. The gas turbine engine of claim 6 , wherein the trailing edge of each blade has the same sweep as the leading edge.
9. The gas turbine engine of claim 6 , wherein the trailing edge of each blade has a sweep that is less than the sweep of the leading edge.
10. The gas turbine engine of claim 6 , wherein the leading edge has a sweep of 14% of the root axial chord length and the trailing edge has a sweep of 18% of the root axial chord length.
11. The gas turbine engine of claim 6 , wherein the full-span forward sweeps of the blades of each rotor portion is the same.
12. The gas turbine engine of claim 6 , wherein the vanes of the stator portions are swept to accommodate the sweep of the blades of the rotor portions.