CPC F04D 29/324 (2013.01) [F01D 5/14 (2013.01); F01D 5/141 (2013.01); F01D 9/02 (2013.01); F02C 3/04 (2013.01); F02C 7/36 (2013.01); F02K 3/06 (2013.01); F04D 29/384 (2013.01); F04D 29/544 (2013.01); F05D 2220/32 (2013.01); F05D 2240/35 (2013.01); F05D 2250/71 (2013.01); F05D 2260/4031 (2013.01); Y02T 50/60 (2013.01)] | 20 Claims |
1. A gas turbine engine comprising:
a combustor section arranged between a compressor section and a turbine section, wherein the compressor section includes at least a low pressure compressor and a high pressure compressor, the high pressure compressor arranged upstream of the combustor section;
a fan section, wherein the low pressure compressor is downstream from the fan section;
an airfoil arranged in the low pressure compressor and including pressure and suction sides extending in a radial direction from a 0% span position to a 100% span position, wherein the airfoil has an axial stacking offset defined as a distance between a center of gravity for a particular span section and a reference point, the reference point corresponding to an axial center of one of a root of the airfoil or a rotor bore, a positive axial stacking offset being on an axially aft side of the reference point and a negative axial stacking offset being on an axially forward side of the reference point; and
wherein the axial stacking offset at a 10% span position is positive and greater than or equal to an axial stacking offset at the 100% span position, and a maximum positive axial stacking offset value of the axial stacking offset across the 0% span position to the 100% span position of the airfoil is greater than an absolute value of any negative axial stacking offset at the 100% span position.
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