Abstract: A method for assembling an articulated fan front frame for a gas turbine engine facilitates improving engine performance. The method includes forming a strut including a pair of sidewalls connected at a leading edge and a trailing edge, forming a flap including a first sidewall and a second sidewall connected at a leading edge and a trailing edge, and extending in radial span between a root endwall and a tip endwall, wherein each endwall extends between the first and second sidewalls, and wherein at least one of the root endwall and the tip endwall is contoured in a radial direction extending between the flap leading and trailing edges, and pivotally coupling the flap downstream from the strut such that a gap is defined between the flap and the strut.

Abstract: A method for assembling an articulated fan front frame for a gas turbine engine facilitates improving engine performance. The method includes forming a strut including a pair of sidewalls connected at a leading edge and a trailing edge, forming a flap including a first sidewall and a second sidewall connected at a leading edge and a trailing edge, and extending in radial span between a root endwall and a tip endwall, wherein each endwall extends between the first and second sidewalls, and wherein at least one of the root endwall and the tip endwall is contoured in a radial direction extending between the flap leading and trailing edges, and pivotally coupling the flap downstream from the strut such that a gap is defined between the flap and the strut.

Abstract: A blade is comprised of a plurality of a family of airfoil sections, each section being characterized by a large leading edge radii, a camber having an apogee at approximately 45% of the airfoil section chord length and relatively blunt trailing edge. The airfoil sections taper in thickness ratio (h/b) from about 20% near the blade root to about 2% near the blade tip. The maximum thickness for each airfoil section increases along the chord length from about 39% of the airfoil section for a 20% airfoil section to about 42% of the airfoil section chord length of a 2% airfoil section. As such, the novel blade of the invention provides superior aerodynamic performance to the NACA Series 16 airfoil in the Mach 0.50 to 0.70 range with less susceptibility to foreign object damage and handling problems.