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: Bleed valve assemblies in a gas turbine engine are disclosed herein. A bleed valve assembly in a low pressure compressor may include a bleed valve and a resonator chamber. A manifold may allow passage of air from a bleed duct into the resonator chamber. The resonator chamber may alter resonation properties of the bleed duct in order to prevent damage to components in the low pressure compressor.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan and a turbine section having a fan drive turbine rotor, and a compressor rotor. A gear reduction effects a reduction in a speed of the fan relative to an input speed from the fan drive turbine rotor. The compressor rotor has a number of compressor blades in at least half of a plurality of blade rows of the compressor rotor. The blades are configured to operate at least some of the time at a rotational speed. The number of compressor blades in the at least half of the blade rows and the rotational speed is such that the following formula holds true for each row of the at least half of the blade rows of the compressor rotor: (the number of blades×the rotational speed)/60 s?about 5500 Hz.

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 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: 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 gas turbine engine has a propulsor including a fan and a liner positioned upstream of the fan. The liner has a backing plate, a cellular structure with cells extending from the backing plate, and a perforated sheet with a depth defined as a distance between the perforated sheet and the backing sheet. The depth is selected to achieve a desired ratio of the depth relative to a gap?. A depth to gap ratio is substantially in a range of 0.035 to 0.08. A method is also disclosed.

Abstract: A gas turbine engine has a fan and a turbine having a fan drive turbine rotor. The fan drive turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive turbine rotor that drives the compressor rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor 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)/60 s?5500 Hz, and the rotational speed is in revolutions per minute. A method of designing a gas turbine engine and a compressor module are also disclosed.

Abstract: A gas turbine engine has a fan and a turbine having a fan drive turbine rotor. The fan drive turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive turbine rotor that drives the compressor rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor 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)/60 s?5500 Hz, and the rotational speed is in revolutions per minute. A method of designing a gas turbine engine and a compressor module are also disclosed.

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 comprises a fan and a turbine section having a first turbine rotor. The first turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from a fan drive turbine rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor 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: (the number of blades×the rotational speed)/(60 seconds/minute)?5500 Hz; and the rotational speed being in revolutions per minute. A compressor module and a method of designing a gas turbine engine are also disclosed.

Abstract: A gas turbine engine has a fan and a turbine having a fan drive turbine rotor. The fan drive turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive turbine rotor that drives the compressor rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor 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)/60 s?5500 Hz, and the rotational speed is in revolutions per minute. A method of designing a gas turbine engine and a compressor module are also disclosed.

Abstract: A gas turbine engine has a fan, a compressor section having a low pressure portion and a high pressure portion, a combustor section, and a turbine having a low pressure portion. The low pressure turbine portion drives the low pressure compressor portion and the fan. A gear reduction effects a reduction in the speed of the fan relative to a speed of the low pressure turbine and the low pressure compressor portion. At least one of the low pressure turbine portion and low pressure compressor portion has a number of blades in each of a plurality of rows. The blades operate at least some of the time at a rotational speed. The number of blades and the rotational speed are such that the following formula holds true for at least one of the blade rows of the at least one of the low pressure turbine portion and/or the low pressure compressor sections: (number of blades×rotational speed)/60?5500. The rotational speed is an approach speed in revolutions per minute.

Abstract: A gas turbine engine has a fan and a turbine having a fan drive turbine rotor. The fan drive turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive turbine rotor that drives the compressor rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor 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)/60 s?5500 Hz, and the rotational speed is in revolutions per minute. A method of designing a gas turbine engine and a compressor module are also disclosed.

Abstract: A method of manufacturing a compressor section includes the steps of defining a compressor section having a number of blades, and having one or more stator sections, each with numbers of vanes. Each stator section has at least two sections wherein the spacing between the vanes in a first of the sections is not equal to a spacing between the vanes in a second of the sections. The number of blades, and the number of vanes where all of the sections are selected to achieve acoustic cutoff.

Abstract: A fan section for a gas turbine engine has a fan rotor with a plurality of fan blades. A plurality of exit guide vanes are positioned to be downstream of the fan rotor. The fan rotor is driven through a gear reduction relative to a turbine section. The exit guide vanes are desired to address resultant sound from interaction of wakes from the fan blades across exit guide vanes. A gas turbine engine incorporating a fan section is also disclosed.

Abstract: A method of manufacturing a compressor section includes the steps of defining a compressor section having a number of blades, and having one or more stator sections, each with numbers of vanes. Each stator section has at least two sections wherein the spacing between the vanes in a first of the sections is not equal to a spacing between the vanes in a second of the sections. The number of blades, and the number of vanes where all of the sections are selected to achieve acoustic cutoff.