Abstract: A compressor rotor airfoil in a gas turbine engine is presented. Opposed pressure and suction sides are joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. A leading edge dihedral angle is defined at a point on the leading edge between a tangent to the airfoil and a vertical. The leading edge dihedral angle has a span-wise distribution. The distribution has at least one inflection point. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade is also presented.

Abstract: An airfoil for a gas turbine engine comprises at least three successive sections defined between a root and a tip. A first decrease of maximum thickness-to-chord ratios defined by a difference between a maximum thickness-to-chord ratio at the root and a maximum thickness-to-chord ratio at a first spanwise position, a second decrease of the maximum thickness-to-chord ratios defined by a difference between the maximum thickness-to-chord ratio at the first spanwise position and a maximum thickness-to-chord ratio at a second spanwise position, a third decrease of the maximum thickness-to-chord ratios defined by a difference between the maximum thickness-to-chord ratio at the second spanwise position and a maximum thickness-to-chord ratio at the tip, the second decrease being greater than the corresponding first and third decreases.

Abstract: A rotary airfoil in a gas turbine engine is provided. The airfoil includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. An axial component of a center of gravity of a cross-section taken chordally toward the tip of the airfoil being upstream relative to an axial component of a center of gravity of a cross-section taken chordally toward the root of the airfoil. A method for forming such blade is also presented.

Abstract: A method for repairing a blade in a gas turbine engine comprises the steps of: isolating the damage on the airfoil of the blade; forming a cut back in the shape of elongated “D” shaped recess with a pair of fillets, a depth and a longitudinal axis of the “D” shaped recess having a length along the leading or trailing edge of the airfoil; and the fillets having a respective radius.

Abstract: An airfoil for a gas turbine engine comprises at least three successive sections defined between a root and a tip. A first decrease of maximum thickness-to-chord ratios defined by a difference between a maximum thickness-to-chord ratio at the root and a maximum thickness-to-chord ratio at a first spanwise position, a second decrease of the maximum thickness-to-chord ratios defined by a difference between the maximum thickness-to-chord ratio at the first spanwise position and a maximum thickness-to-chord ratio at a second spanwise position, a third decrease of the maximum thickness-to-chord ratios defined by a difference between the maximum thickness-to-chord ratio at the second spanwise position and a maximum thickness-to-chord ratio at the tip, the second decrease being greater than the corresponding first and third decreases.

Abstract: A compressor rotor airfoil in a gas turbine engine is presented. Opposed pressure and suction sides are joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. A leading edge dihedral angle is defined at a point on the leading edge between a tangent to the airfoil and a vertical. The leading edge dihedral angle has a span-wise distribution. The distribution has at least one inflection point. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade is also presented.

Abstract: An airfoil in a gas turbine engine includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend spanwise from a root to a tip of the airfoil. The airfoil has a spanwise distribution of maximum thicknesses of chordwise cross-sections of the airfoil. The spanwise distribution of maximum thicknesses decreases from the root to the tip. In one aspect, the spanwise distribution is stepped between a first portion extending from the root and a second portion extending to the tip.

Abstract: A compressor rotor airfoil in a gas turbine engine is presented. Opposed pressure and suction sides are joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. A leading edge dihedral angle is defined at a point on the leading edge between a tangent to the airfoil and a vertical. The leading edge dihedral angle has a span-wise distribution. The distribution has at least one inflection point. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade is also presented.

Abstract: An impeller for increasing the pressure of a fluid circulating in an annular fluid path, the impeller comprising: a plurality of centrifugal compressor vanes circumferentially interspaced around the axis of the annular fluid path, the plurality of compressor vanes extending from an axially-oriented inlet to a radially-oriented outlet, and each having an inner edge and a free edge, the free edge of the plurality of compressor vanes coinciding with an outer limit of the annular fluid path, and a hub having a solid-of-revolution shape centered around an axis, the hub having an outer hub surface forming an inner limit to the annular fluid path and to which the inner edge of the plurality of centrifugal vanes is secured, the outer hub surface having a portion which leans forward, forming an axial recess therein.

Abstract: An airfoil in a gas turbine engine includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend spanwise from a root to a tip of the airfoil. The airfoil has a spanwise distribution of maximum thicknesses of chordwise cross-sections of the airfoil. The spanwise distribution of maximum thicknesses decreases from the root to the tip. In one aspect, the spanwise distribution is stepped between a first portion extending from the root and a second portion extending to the tip.

Abstract: A rotary airfoil in a gas turbine engine is provided. The airfoil includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. An axial component of a center of gravity of a cross-section taken chordally toward the tip of the airfoil being upstream relative to an axial component of a center of gravity of a cross-section taken chordally toward the root of the airfoil. A method for forming such blade is also presented.

Abstract: A compressor rotor airfoil in a gas turbine engine is presented. Opposed pressure and suction sides are joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. A leading edge dihedral angle is defined at a point on the leading edge between a tangent to the airfoil and a vertical. The leading edge dihedral angle has a span-wise distribution. The distribution has at least one inflection point. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade is also presented.

Abstract: The bolt has, in sequence along a bolt axis: a threaded portion, a thread run-out portion, a shank, and a head, the threaded portion having a thread with a given root radius and a given depth, the thread run-out portion connected to the shank via a thread run-out fillet having a thread run-out fillet radius, the thread run-out fillet radius being between two and six times the thread root radius.

Abstract: The bolt has, in sequence along a bolt axis: a threaded portion, a thread run-out portion, a shank, and a head, the threaded portion having a thread with a given root radius and a given depth, the thread run-out portion connected to the shank via a thread run-out fillet having a thread run-out fillet radius, the thread run-out fillet radius being between two and six times the thread root radius.

Abstract: A method for repairing a blade in a gas turbine engine comprises the steps of: isolating the damage on the airfoil of the blade; forming a cut back in the shape of elongated “D” shaped recess with a pair of fillets, a depth and a longitudinal axis of the “D” shaped recess having a length along the leading or trailing edge of the airfoil; and the fillets having a respective radius.

Abstract: A spigot arrangement for split impeller (inducer and exducer) includes a recess of the exducer and means for reducing exducer blade root stresses and localized contact stresses between inducer and exducer.

Abstract: A spigot arrangement for split impeller (inducer and exducer) includes a recess of the exducer and means for reducing exducer blade root stresses and localized contact stresses between inducer and exducer.