Abstract: A fan section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan rotor having fan blades, and a plurality of fan exit guide vanes positioned downstream of the fan rotor. The fan rotor is configured to be driven through a gear reduction. A ratio of a number of fan exit guide vanes to a number of fan blades is defined. The fan exit guide vanes are provided with optimized sweep and optimized lean.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a compressor section driven by the turbine section, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.

Abstract: A fan section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan rotor having fan blades, and a plurality of fan exit guide vanes positioned downstream of the fan rotor. The fan rotor is configured to be driven through a gear reduction. A ratio of a number of fan exit guide vanes to a number of fan blades is defined. The fan exit guide vanes are provided with optimized sweep and optimized lean.

Abstract: Disclosed is an acoustic liner for a gas turbine engine, having: a face sheet; a first plurality of resonator chamber cavities, distributed non-uniformly against the face sheet, the first plurality of resonator chamber cavities defining a first aggregated volume that is configured to target noise attenuation at a first frequency, the first plurality of resonator chamber cavities include first and second resonator chamber cavities fluidly coupled to each other, which have different configurations with respect to each other; and a second plurality of resonator chamber cavities, distributed non-uniformly against the face sheet, the second plurality of resonator chamber cavities define a second aggregated volume that differs from the first aggregated volume and is configured to target noise attenuation at a second frequency, the second plurality of resonator chamber cavities include third and fourth resonator chamber cavities fluidly coupled to each other, which have different configurations with respect to each

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a compressor section driven by the turbine section, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.

Abstract: A fan section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan rotor having fan blades, and a plurality of fan exit guide vanes positioned downstream of the fan rotor. The fan rotor is configured to be driven through a gear reduction. A ratio of a number of fan exit guide vanes to a number of fan blades is defined. The fan exit guide vanes are provided with optimized sweep and optimized lean.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.

Abstract: A gas turbine engine has a fan section including a fan, a compressor section including a low pressure compressor and a high pressure compressor, a turbine section including a low pressure turbine and a high pressure turbine, and a gear reduction. The low pressure compressor and the low pressure turbine have a number of blades in each of at least one of a plurality of blade rows. The blades are rotatable at least some of the time at a rotational speed in operation. The number of blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor turbine: (number of blades×rotational speed)/60s?5500 Hz, and the rotational speed is in revolutions per minute.

Abstract: A gas turbine engine has a fan section including a fan, a compressor section including a low pressure compressor and a high pressure compressor, a turbine section including a low pressure turbine and a high pressure turbine, and a gear reduction. The low pressure compressor and the low pressure turbine have a number of blades in each of at least one of a plurality of blade rows. The blades are rotatable at least some of the time at a rotational speed in operation. The number of blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor turbine: 5500?(number of blades×rotational speed)/60?10000, the rotational speed being in revolutions per minute.

Abstract: A fan section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan rotor having fan blades, and a plurality of fan exit guide vanes positioned downstream of the fan rotor. The fan rotor is configured to be driven through a gear reduction. A ratio of a number of fan exit guide vanes to a number of fan blades is defined. The fan exit guide vanes are provided with optimized sweep and optimized lean.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a compressor section driven by the turbine section, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.

Abstract: An exemplary section of a gas turbine engine according to this disclosure includes, among other things, a first array of airfoils including a first number of airfoils, and a second array of airfoils downstream of the first array of airfoils. The second array includes a second number of airfoils. The second number of airfoils is at least 1.19 times the first number of airfoils thereby providing a predetermined modal.

Abstract: An exemplary section of a gas turbine engine according to this disclosure includes, among other things, a first array of airfoils including a first number of airfoils, and a second array of airfoils downstream of the first array of airfoils. The second array includes a second number of airfoils. The second number of airfoils is at least 1.19 times the first number of airfoils thereby providing a predetermined modal.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan, a compressor section having a low pressure compressor and a high pressure compressor, a combustor section, and a turbine section having a low pressure turbine, the low pressure turbine for driving the low pressure compressor and the fan; a gear reduction effecting a reduction in the speed of the fan relative to a speed of the low pressure turbine and the low pressure compressor; and at least one stage of the compressor section having a ratio of vanes to blades that is greater than or equal to 1.8. The corrected tip speed of the blades is greater than or equal to 480 ft/sec at an approach speed.

Abstract: A disclosed fan section of a gas turbine engine includes a fan rotor having a plurality of fan blades and a duct defining a passageway aft of the fan rotor. A fan exit guide vane is disposed within the duct downstream of the fan blades. The fan exit guide vane includes a plurality of exit guide vanes positioned downstream of the fan rotor with at least two of the plurality of exit guide vanes including different aft geometries for guiding airflow through the passage to reduce pressure distortions at the fan blades.

Abstract: In accordance with one aspect of the disclosure, a rotor for a gas turbine engine is disclosed. The rotor may include a rotor disk and a plurality of blade extending radially outward from the rotor disk. At least one of the blades may have a physical nonuniformity. The blades may be distributed about the rotor disk based on any physical nonuniformities of the blades to generate at least one decay-resistant harmonic.

Abstract: A fan section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan rotor having fan blades, and a plurality of fan exit guide vanes positioned downstream of the fan rotor. The fan rotor is configured to be driven through a gear reduction. A ratio of a number of fan exit guide vanes to a number of fan blades is defined. The fan exit guide vanes are provided with optimized sweep and optimized lean.

Abstract: A gas turbine engine impact liner is disclosed. The impact liner may include a base sheet, a plurality of stanchions extending from the base sheet, and a plurality of supports, each being operatively associated with one of the plurality of stanchions.