Abstract: A compressor airfoil of a turbine engine having a geared architecture includes pressure and suction sides extending in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between leading edge sweep angle and span position defined by a curve in which the leading edge sweep angle is positive at 0% span and crosses to a negative leading edge sweep angle at a span position less than 80% span. A negative sweep angle is in the forward direction, and a positive sweep angle is in the rearward direction.

Abstract: An airfoil for a turbine engine includes pressure and suction sides extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between an axial leading edge location and a span position that is at least a third order polynomial with a generally U-shaped curve having an initial negative slope followed by a positive slope. The positive slope leans aftward and the negative slope leans forward. The curve has a critical point in the range of 30-50% span position at which the curve changes from the negative slope to the positive slope. The curve is generally linear from 55% span to 75% span and has a positive slope that increases at a rate of about 0.0875 inch (2.22 mm) per 1% span, +/?0.04 inch (1.01 mm) per 1% span.

Abstract: An airfoil for a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a tangential leading edge location and a span position that corresponds to a curve that is at least a third order polynomial with a generally S-shaped curve that has an initial negative slope followed by a positive slope and then a second negative slope. The positive slope leans toward the suction side and the negative slopes lean toward the pressure side.

Abstract: In one exemplary embodiment, an airfoil for a turbine engine includes pressure and suction sides extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil geometry corresponds to tangential leading and trailing edge curves and a tangential stacking offset curve. The airfoil extends from a root. A zero tangential reference point corresponds to tangential center of the root. YLE corresponds to a tangential distance from a leading edge to the reference point at a given span position. YTE corresponds to a tangential distance from a trailing edge to the reference point at a given span position. Yd corresponds to a tangential stacking offset at a given span position. (YLE?Yd)/(Yd?YTE) at 40% span position is about 1.5 and (YLE?Yd)/(Yd?YTE) at 20% span position is about 2.

Abstract: An airfoil of a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a tangential stacking offset and a span position that is at least a third order polynomial curve that includes at least one positive and negative slope. The positive slope leans toward the suction side and the negative slope leans toward the pressure side. An initial slope starts at the 0% span position is either zero or positive. The first critical point is less than 15% span.

Abstract: A fan section for a gas turbine engine according to an example of the present disclosure includes, among other things, a rotor hub defining an axis, and an array of airfoils circumferentially spaced about the rotor hub. Each of the airfoils include pressure and suction sides between a leading edge and a trailing edge and extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip, facing pressure and suction sides of adjacent airfoils defining a channel in a chordwise direction having a width between the facing pressure and suction sides at a given span position of the adjacent airfoils. The width at each pressure side location along the channel is defined as a minimum distance to a location along the suction side, the width converging and diverging along the channel for at least some span positions.

Abstract: An airfoil for a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a stacking offset and a span position that includes at least one positive and negative slope. The positive slope leans aftward and the negative slope leans forward relative to an engine axis. The positive slope crosses an initial axial stacking offset corresponding to the 0% span position at a zero-crossing position. A first axial stacking offset X1 is provided from the zero-crossing position to a negative-most value on the curve. A second axial stacking offset X2 is provided from the zero-crossing position to a positive-most value on the curve. A ratio of the second to first axial stacking offset X2/X1 is less than 2.0.

Abstract: An airfoil of a turbine engine includes pressure and suction sides and extends in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between a camber angle and span position that defines a curve with the camber angle having a positive slope from 0% span to 100% span.

Abstract: In one exemplary embodiment, an airfoil for a turbine engine includes an airfoil that has pressure and suction sides and extends in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a curve that corresponds to a relationship between a leading edge dihedral and a span position. The leading edge dihedral has a portion of the curve with a change in dihedral in the range of 90% to 100% span position of greater than 10°. A positive dihedral corresponds to suction side-leaning. A negative dihedral corresponds to pressure side-leaning.

Abstract: An airfoil for a turbine engine includes an airfoil that has pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a curve that corresponds to a relationship between a trailing edge sweep angle and a span position. The trailing edge sweep angle is in a range of 10° to 20° in a range of 40-70% span position. The trailing edge sweep angle is positive from 0% span to at least 95% span. The airfoil has a relationship between a leading edge dihedral and a span position. The leading edge dihedral is negative from the 0% span position to the 100% span position. A positive dihedral corresponds to suction side-leaning. A negative dihedral corresponds to pressure side-leaning. The trailing edge sweep angle is positive from 0%-95% span. A positive-most trailing edge sweep angle is within the range of 10° to 20° in the range of 40-70% span position. The airfoil is a fan blade for the turbine engine.

Abstract: A gas turbine engine includes a bypass flow passage that has an inlet and defines a bypass ratio in a range of approximately 8.5 to 13.5. A fan is arranged at the inlet. A first turbine is coupled with a first shaft such that rotation of the first turbine will drive the fan. A first compressor is coupled with the first shaft and includes three stages, and a second turbine is coupled with a second shaft and includes two stages. The fan includes a row of 16 (N) fan blades that has a solidity value (R) that is from 1.0 to 1.2 and a ratio of N/R that is from 13.3 to 16.0.

Abstract: A gas turbine engine includes a gear assembly, a bypass flow passage, and a core flow passage. The bypass flow passage includes an inlet. A fan is arranged at the inlet of the bypass flow passage. A first shaft and a second shaft are mounted for rotation about an engine central longitudinal axis. A first turbine is coupled with the first shaft such that rotation of the first turbine is configured to drive the fan, through the first shaft and gear assembly, at a lower speed than the first shaft. The fan includes a hub and a row of fan blades that extend from the hub. The row includes 12 (N) of the fan blades, a solidity value (R) that is from 1.0 to 1.2, and a ratio of N/R that is from 10.0 to 12.0.

Abstract: An airfoil for a turbine engine includes an airfoil having pressure and suction sides extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a curve corresponding to a relationship between a trailing edge sweep angle and a span position. The trailing edge sweep angle is in a range of 0° to 10° in a range of 10-20% span position. The trailing edge sweep angle is in a range of the trailing edge sweep angle is positive from 0% span to at least 95% span.

Abstract: An airfoil for a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a total chord length and a span position and corresponds to a curve that has an increasing total chord length from the 0% span position to a first peak. The first peak occurs in the range of 45-65% span position, and the curve either remains constant or has a decreasing total chord length from the first peak toward the 100% span position. The total chord length is at the 0% span position in the range of 8.2-10.5 inches (20.8-26.7 cm). The curve is at least a third order polynomial and has an initial positive slope.

Abstract: A gas turbine engine includes a gear assembly and a bypass flow passage that has an inlet. A fan is arranged within the bypass flow passage. A first turbine is coupled with a first shaft, which is coupled with the fan. A second turbine is coupled with a second shaft. The fan includes a hub and a row of fan blades that extend from the hub. The row includes 12 to 14 (N) of the fan blades, a solidity value (R) at tips of the fan blades that is greater than 0.85 and less than 1.1, and a ratio of N/R that is from 9 to 20.

Abstract: A gas turbine engine component is described. The gas turbine engine component includes an inner diameter edge, an outer diameter edge, a trailing edge and a leading edge. The leading edge has a positive (aft) aerodynamic sweep across substantially an entire span of the leading edge. The gas turbine engine component has a camber angle greater than 50 degrees across substantially an entire span of the component. The gas turbine engine component may have asymmetrical tangential stacking of the component in the radial direction.

Abstract: A gas turbine engine includes a bypass flow passage and a core flow passage. The bypass flow passage defines a bypass ratio in a range of approximately 8.5 to 13.5. A fan is located upstream of the bypass flow passage. The bypass flow passage includes an inlet and an outlet that define a design fan pressure ratio of approximately 1.3 to 1.55. A first, inner shaft and a second, outer shaft are concentric. A first turbine is coupled with the first shaft, and the first shaft is coupled with the fan. The fan includes a hub and a row of fan blades that extend from the hub. The row includes a number of the fan blades, the number (N) being 18, a solidity value (R) at tips of the fan blades that is from 1.0 to 1.1, and a ratio of N/R that is from 16.4 to 18.0.

Abstract: An airfoil for a turbine engine includes an airfoil having pressure and suction sides extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a curve corresponding to a relationship between a trailing edge sweep angle and a span position. The trailing edge sweep angle is in a range of 10° to 20° in a range of 40-70% span position. The trailing edge sweep angle is positive from 0% span to at least 95% span.

Abstract: A gas turbine engine includes a gear assembly and a bypass flow passage that includes an inlet and an outlet that define a design pressure ratio between 1.3 and 1.55. A fan is arranged at the inlet. A first turbine is coupled with a first shaft such that rotation of the first turbine will drive the fan, through the first shaft and the gear assembly, at a lower speed than the first shaft. The fan includes a row of fan blades. The row includes 12-16 (N) fan blades, a solidity value (R) that is from 1.0 to 1.3, and a ratio of N/R that is from 10.0 to 16.

Abstract: A fan exit guide vane assembly for a gas turbine engine includes a plurality of guide vanes having a pressure side and a suction side extending between a leading edge and a trailing edge. The vanes further include a span extending between a root and tip with a stagger angle defined as an angle between a longitudinal axis parallel to an engine axis of rotation and a line connecting the leading edge and the trailing edge that is less than about 15°.