Abstract: A gas turbine engine assembly includes a fan. A diameter of the fan has a dimension D. The fan has a pressure ratio of greater than 1.20 and less than 1.45. A leading edge on an inlet portion of a nacelle is within a first reference plane oriented at an oblique angle. A forward most portion on the fan blade leading edges is in a second reference plane. A length of the inlet portion has a dimension L different at a plurality of locations on the inlet portion. A geared architecture has a gear reduction ratio of greater than 2.3, a bypass ratio is greater than 10, and a low pressure turbine includes a pressure ratio greater than 5:1. A dimensional relationship of UD is between 0.25 and 0.45. The leading edge on the inlet portion is further from the second reference plane near the top of the assembly.

Abstract: A compressor section for a gas turbine engine according to an example of the present disclosure includes, among other things, a low pressure compressor including a plurality of rotor blades arranged about an axis, a high pressure compressor, and a core flowpath passing through the low pressure compressor. The core flowpath at the low pressure compressor defines an inner diameter and an outer diameter relative to the axis. The outer diameter has a slope angle relative to the axis.

Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is between 0.20 and 0.40.

Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is no more than 0.45.

Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan, a nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is no more than 0.45.

Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is no more than 0.45.

Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan that has a plurality of fan blades. A diameter of the fan has a dimension D that is based on a dimension of the fan blades. Each fan blade has a leading edge. An inlet portion is situated forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and about 0.45.

Abstract: A nose cone assembly for a gas turbine engine and method of circulating air in a gas turbine engine are disclosed. The nose cone assembly includes a nose cone having an aperture communicating air to an interior space of the nose cone and a discharge member communicating the air out of the nose cone.

Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan that has a plurality of fan blades. A diameter of the fan has a dimension D that is based on a dimension of the fan blades. Each fan blade has a leading edge. An inlet portion is situated forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and 0.45.

Abstract: A nose cone assembly for a gas turbine engine and method of circulating air in a gas turbine engine are disclosed. The nose cone assembly includes a nose cone having an aperture communicating air to an interior space of the nose cone and a discharge member communicating the air out of the nose cone.

Abstract: A propulsor for a gas turbine engine includes, among other things, a case including a duct disposed along an axis to define a flow path. A rotor includes a row of propulsor blades extending in a generally radial direction outwardly from a hub, the hub rotatable about the axis such that the propulsor blades deliver airflow into the flow path. A row of guide vanes are situated in the flow path. At least two of the guide vanes extend in the generally radial direction between inner and outer surfaces of the duct, extends in a chordwise direction between a first leading edge and a first trailing edge to define a vane chord dimension (VCD) at a first span position of the corresponding guide vane, and defines a vane circumferential pitch (VCP) at the first span position of the corresponding guide vane and an adjacent one of the guide vanes. The row of guide vanes has a vane solidity (VR) defined as VCD/VCP, the vane solidity (VR) being equal to or less than 1.43.

Abstract: A compressor section for a gas turbine engine according to an example of the present disclosure includes, among other things, a low pressure compressor including a plurality of rotor blades arranged about an axis, a high pressure compressor, and a core flowpath passing through the low pressure compressor. The core flowpath at the low pressure compressor defines an inner diameter and an outer diameter relative to the axis. The outer diameter has a slope angle relative to the axis.

Abstract: A gas turbine engine assembly according to an example of the present disclosure includes, among other things, a fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, a geared architecture configured to drive the fan, a turbine section configured to drive the geared architecture, a compressor section including a first compressor and a second compressor, and an inlet portion forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and about 0.45.

Abstract: A gas turbine engine includes a spool, a turbine coupled to drive the spool and a propulsor that is coupled to be driven by the turbine through the spool. A gear assembly is coupled between the propulsor and the spool such that rotation of the spool drives the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. The row includes no more than 16 of the propulsor blades.

Abstract: A gas turbine engine includes a spool, a turbine coupled to drive the spool, a propulsor coupled to be rotated about an axis by the turbine through the spool, and a gear assembly coupled between the propulsor and the spool such that rotation of the turbine drives the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extend from the hub. Each of the propulsor blades has a span between a root at the hub and a tip, and a chord between a leading edge and a trailing edge. The chord forms a stagger angle ? with the axis, and the stagger angle ? is less than 15° at a position along the propulsor blade that is within an inboard 20% of the span.

Abstract: A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flowpath. A shaft provides a rotational axis. A hub is operatively supported by the shaft. A rotor is connected to the hub and supports a compressor section. The compressor section is arranged in a core flow path axially between the inlet case flow path and the intermediate case flow path. The core flowpath has an inner diameter and an outer diameter. At least a portion of inner diameter has an increasing slope angle relative to the rotational axis. A bearing is mounted to the hub and supports the shaft relative to one of the intermediate case and the inlet case.

Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan that has a plurality of fan blades. A diameter of the fan has a dimension D that is based on a dimension of the fan blades. Each fan blade has a leading edge. An inlet portion is situated forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and about 0.45.

Abstract: A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flowpath. A shaft provides a rotational axis. A hub is operatively supported by the shaft. A rotor is connected to the hub and supports a compressor section. The compressor section is arranged in a core flow path axially between the inlet case flow path and the intermediate case flow path. The core flowpath has an inner diameter and an outer diameter. At least a portion of inner diameter has an increasing slope angle relative to the rotational axis. A bearing is mounted to the hub and supports the shaft relative to one of the intermediate case and the inlet case.

Abstract: A core flowpath through a low pressure compressor section of a gas turbine engine includes an outer diameter, which has a slope angle relative to an axis defined by the core flowpath. The slope angle is a slope angle that is operable to prevent flow separation of a fluid passing through the core flowpath.

Abstract: A method of protecting a gas turbine engine according an exemplary aspect of the present disclosure includes, among other things, the steps of determining at least one flight condition of an aircraft and comparing the at least one flight condition to a programmed condition. The method further includes the steps of moving a plurality of inlet vanes of a low pressure compressor from a first position to a second position if the step of comparing the at least one flight condition to the programmed flight condition determines the programmed flight condition are met and deflecting any foreign objects with the plurality of inlet vanes.