Abstract: Gas turbine engine (GTE) airfoils, such as rotor and turbofan blades, having multimodal thickness distributions include an airfoil tip, and an airfoil root opposite the airfoil tip in a spanwise direction. The GTE airfoil has a first, second and third locally-thickened region, with the first locally-thickened region defined at the airfoil root. A maximum thickness of each chord between the airfoil root and the airfoil tip transitions toward the leading edge between the first locally-thickened region and the second locally-thickened region, and the third locally-thickened region extends in the spanwise direction. A chord line that extends through the third locally-thickened region contains a first local thickness maxima and a second local thickness maxima interspersed with at least two local thickness minima, and the first local thickness maxima is defined by the third locally-thickened region and is greater than the second local thickness maxima.

Abstract: A rotor blade for a compressor of a gas turbine engine includes an airfoil. The airfoil has a span that extends from 0% at the root to 100% at the tip and a mean camber line that extends from a leading edge to a trailing edge. The airfoil has a location of local maximum thickness defined as a ratio of a first arc distance along the mean camber line between the leading edge and a position of the local maximum thickness to a total arc distance along the mean camber line from the leading edge to the trailing edge. A value of the ratio increases from the root to a first position value, decreases from the first position value to a second position value and increases from the second position value to the tip. The first position value is at a spanwise location within 20% to 50% of the span.

Abstract: A rotor blade for a compressor of a gas turbine engine includes an airfoil extending from a root to a tip and having a leading edge and a trailing edge. The airfoil has a span that extends from 0% at the root to 100% at the tip and a mean camber line that extends from the leading edge to the trailing edge. The airfoil has a total camber distribution that increases from the root to a maximum value of total camber between 5% of the span and 20% of the span.

Abstract: Gas turbine engine (GTE) airfoils, such as rotor and turbofan blades, having multimodal thickness distributions include an airfoil tip, and an airfoil root opposite the airfoil tip in a spanwise direction. The GTE airfoil has a first, second and third locally-thickened region, with the first locally-thickened region defined at the airfoil root. A maximum thickness of each chord between the airfoil root and the airfoil tip transitions toward the leading edge between the first locally-thickened region and the second locally-thickened region, and the third locally-thickened region extends in the spanwise direction. A chord line that extends through the third locally-thickened region contains a first local thickness maxima and a second local thickness maxima interspersed with at least two local thickness minima, and the first local thickness maxima is defined by the third locally-thickened region and is greater than the second local thickness maxima.

Abstract: Gas turbine engine (GTE) airfoils, such as rotor and turbofan blades, having multimodal thickness distributions are provided. In one embodiment, the GTE airfoil includes an airfoil tip, an airfoil root opposite the airfoil tip in a spanwise direction, and first and second airfoil halves extending between the airfoil tip and the airfoil root. The first airfoil half has a first multimodal thickness distribution, as taken in a cross-section plane extending in the spanwise direction and in a thickness direction substantially perpendicular to the spanwise direction. The first multimodal thickness distribution may be defined by multiple locally-thickened airfoil regions, which are interspersed with multiple locally-thinned airfoil regions. The second airfoil half may or may not have a multimodal thickness distribution.

Abstract: A rotor includes a plurality of airfoils and each airfoil has a span that extends from 0% at a root to 100% at a tip, a chord that extends from 0% at a leading edge to 100% at a trailing edge and a pressure side opposite a suction side. The pressure side of each airfoil has a pressure side surface shape and the suction side of each airfoil has a suction side surface shape based on an operating state of the rotor. Each airfoil has the same suction side surface shape between 10% and 90% of the chord and between 80% and 100% of the span at a first state. At least one first airfoil has a different suction side surface shape between 10% and 90% of the chord and between 80% and 100% of the span than at least one second airfoil at a static state of the rotor.

Abstract: Embodiments of a core-protecting fan module are provided, as are embodiments of a turbofan engine containing such a fan module. In an embodiment, the core-protecting fan module contains a nose member, a fan rotor downstream of the nose member, a full span stator downstream of the fan rotor, and a splitter structure downstream of the fan rotor. The fan rotor includes a plurality of fan blades, which extends from a rotor hub and which is angularly spaced about a rotational axis. Certain fundamental angular relationships are observed between the angles formed by rotational axis, the nose member, the fan rotor, and a leading edge of the splitter structure to reduce contaminant ingestion by the core flow path and to promote moisture shedding to reduce susceptibility to icing within the fan module, while further avoiding or minimizing negative impacts to other structural and functional aspects of the turbofan engine.

Abstract: A rotor blade for a compressor of a gas turbine engine includes an airfoil. The airfoil has a span that extends from 0% at the root to 100% at the tip and a mean camber line that extends from a leading edge to a trailing edge. The airfoil has a location of local maximum thickness defined as a ratio of a first arc distance along the mean camber line between the leading edge and a position of the local maximum thickness to a total arc distance along the mean camber line from the leading edge to the trailing edge. A value of the ratio increases from the root to a first position value, decreases from the first position value to a second position value and increases from the second position value to the tip. The first position value is at a spanwise location within 20% to 50% of the span.

Abstract: A rotor blade for a compressor of a gas turbine engine includes an airfoil extending from a root to a tip and having a leading edge and a trailing edge. The airfoil has a span that extends from 0% at the root to 100% at the tip and a mean camber line that extends from the leading edge to the trailing edge. The airfoil has a total camber distribution that increases from the root to a maximum value of total camber between 5% of the span and 20% of the span.

Abstract: A rotor includes a plurality of airfoils and each airfoil has a span that extends from 0% at a root to 100% at a tip, a chord that extends from 0% at a leading edge to 100% at a trailing edge and a pressure side opposite a suction side. The pressure side of each airfoil has a pressure side surface shape and the suction side of each airfoil has a suction side surface shape based on an operating state of the rotor. Each airfoil has the same suction side surface shape between 10% and 90% of the chord and between 80% and 100% of the span at a first state. At least one first airfoil has a different suction side surface shape between 10% and 90% of the chord and between 80% and 100% of the span than at least one second airfoil at a static state of the rotor.

Abstract: Embodiments of a core-protecting fan module are provided, as are embodiments of a turbofan engine containing such a fan module. In an embodiment, the core-protecting fan module contains a nose member, a fan rotor downstream of the nose member, a full span stator downstream of the fan rotor, and a splitter structure downstream of the fan rotor. The fan rotor includes a plurality of fan blades, which extends from a rotor hub and which is angularly spaced about a rotational axis. Certain fundamental angular relationships are observed between the angles formed by rotational axis, the nose member, the fan rotor, and a leading edge of the splitter structure to reduce contaminant ingestion by the core flow path and to promote moisture shedding to reduce susceptibility to icing within the fan module, while further avoiding or minimizing negative impacts to other structural and functional aspects of the turbofan engine.

Abstract: Embodiments of a forward-swept impeller are provide, as are embodiments of a gas turbine engine containing a forward-swept impeller. In one embodiment, the gas turbine engine includes a shaft and a forward-swept impeller mounted to the shaft. The forward-swept impeller includes, in turn, an inboard impeller section, an outboard impeller section circumscribing the inboard impeller section, and a plurality of hub flow paths extending over the forward-swept impeller from the inboard impeller section to the outboard impeller section. The plurality of hub flow paths each have a flow path exit that is tilted in a forward direction, as taken along a line tangent to the flow path exit.

Abstract: Gas turbine engine (GTE) airfoils, such as rotor and turbofan blades, having multimodal thickness distributions are provided. In one embodiment, the GTE airfoil includes an airfoil tip, an airfoil root opposite the airfoil tip in a spanwise direction, and first and second airfoil halves extending between the airfoil tip and the airfoil root. The first airfoil half has a first multimodal thickness distribution, as taken in a cross-section plane extending in the spanwise direction and in a thickness direction substantially perpendicular to the spanwise direction. The first multimodal thickness distribution may be defined by multiple locally-thickened airfoil regions, which are interspersed with multiple locally-thinned airfoil regions. The second airfoil half may or may not have a multimodal thickness distribution.

Abstract: Embodiments of a forward-swept impeller are provide, as are embodiments of a gas turbine engine containing a forward-swept impeller. In one embodiment, the gas turbine engine includes a shaft and a forward-swept impeller mounted to the shaft. The forward-swept impeller includes, in turn, an inboard impeller section, an outboard impeller section circumscribing the inboard impeller section, and a plurality of hub flow paths extending over the forward-swept impeller from the inboard impeller section to the outboard impeller section. The plurality of hub flow paths each have a flow path exit that is tilted in a forward direction, as taken along a line tangent to the flow path exit.

Abstract: A component for a gas turbine engine includes a rotor disk, a plurality of blade attachments, a plurality of retention flanges, and a plurality of retention tabs. The rotor disk has a plurality of slots formed in an outer surface thereof. Each blade attachment has an attachment region configured to be partially disposed within one of the slots. The plurality of retention flanges extend from the rotor disk outer surface, to thereby at least inhibit axial movement of the blade attachment, when the attachment region is fitted inside its corresponding slot. A retention tab extends from each attachment region and engages one of the retention flanges.

Abstract: A propulsion engine fan incorporates two sets of blades with recessed portions formed along the surface of each blade. One set of blades has a higher resonant frequency than the other set of blades. In an illustrative embodiment of the invention, first and second blade groups each comprise an airfoil portion forming a surface for directing the flow of a working fluid through the engine. Each of the first group of blades has first and second recessed regions formed from the blade surface and has a first resonant frequency. Each of the second group of blades has third and fourth recessed regions formed from the blade surface and has a second resonant frequency less than the first resonant frequency. The first and second group blades are arranged in an alternating pattern along the rotor portion of the fan.

Abstract: A gas turbine engine cooling airflow pickup and delivery system which radially inwardly traverses the torque transmittal path of the compressor section at a relatively large radius location, then extends substantially farther inwardly to a significantly smaller radius location, all with minimal static pressure loss to the cooling airflow. Pressurized, substantially non-swirling air from the compressor stage is supplied to a collector, and from the collector thru pickup holes arranged at an angle of 45 degrees reversely to the direction of rotation of the shaft. The pressurized air is then lead into a vortex spoiler which has a plurality of blades smoothly curved reversely to the direction of rotation of the shaft, the radially outer end of each blade being inclined in matching relation to the pickup holes.