Geared turbofan engine with high compressor exit temperature

A gas turbine engine comprising:

• a fan including a plurality of fan blades rotatable about an axis;
• a compressor section including at least a first compressor section and a second compressor section, wherein the first compressor section comprises a low pressure compressor and the second compressor section comprises a high pressure compressor, and wherein components of the second compressor section are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.;
• wherein the high pressure compressor includes a plurality of stages with each stage comprising a disk with a plurality of blades extending radially outwardly from a rim of the disk, and wherein the plurality of stages includes at least a first stage having a first blade and disk configuration and a second stage having a second blade and disk configuration that is different than the first blade and disk configuration, and wherein the first blade and disk configuration comprises a plurality of slots to receive the plurality of blades and including a plurality of rim cavities for honeycomb seals, and wherein the second blade and disk configuration comprises integrally formed blades such that there are no rim cavities or associated honeycomb seals;
• a combustor in fluid communication with the compressor section;
• a turbine section in fluid communication with the combustor;
• and
• a geared architecture driven by the turbine section for rotating the fan about the axis, and wherein the turbine section comprises a low pressure turbine that drives the low pressure compressor via a first shaft and a high pressure turbine that drives the high pressure compressor via a second shaft, and wherein the geared architecture couples the first shaft to the fan.

A gas turbine engine comprising:

• a fan including a plurality of fan blades rotatable about an axis;
• a compressor section including at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.;
• a combustor in fluid communication with the compressor section;
• a turbine section in fluid communication with the combustor, wherein the turbine section comprises three separate turbine sections including a fan drive turbine, an intermediate pressure turbine, and a secondary pressure turbine;
• the intermediate pressure turbine driving a first compressor rotor of the first compressor section;
• the secondary pressure turbine driving a second compressor rotor of the second compressor section;
• and
• a geared architecture driven by the fan drive turbine for rotating the fan about the axis.

A gas turbine engine comprising:

• a fan including a plurality of fan blades rotatable about an axis;
• a compressor section including at least a first compressor section rotating at a first speed and a second compressor section rotating at a second speed greater than the first speed, wherein components of the second compressor section are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.;
• the second compressor section including a plurality of stages with each stage comprising a disk with a plurality of blades extending radially outwardly from a rim of the disk, and wherein the plurality of stages includes at least a first stage having a first blade and disk configuration and a second stage having a second blade and disk configuration that is different than the first blade and disk configuration, and wherein the first blade and disk configuration comprises a plurality of slots to receive the plurality of blades and including a plurality of rim cavities for honeycomb seals, and wherein the second blade and disk configuration comprises integrally formed blades such that there are no rim cavities or associated honeycomb seals;
• a combustor in fluid communication with the compressor section;
• a turbine section in fluid communication with the combustor, wherein the turbine section comprises a first turbine section that drives the first compressor section via a first shaft at the first speed and a second turbine section that drives the second compressor section via a second shaft at the second speed;
• and
• a geared architecture driven by the turbine section for rotating the fan about the axis, and wherein the geared architecture couples the first shaft to the fan.

A gas turbine engine comprising:

• a fan including a plurality of fan blades rotatable about an axis;
• a compressor section including at least a first compressor section rotating at a first speed and a second compressor section rotating at a second speed greater than the first speed, wherein components of the second compressor section are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.;
• the second compressor section including a plurality of stages with each stage comprising a disk with a plurality of blades extending radially outwardly from a rim of the disk, and wherein the plurality of stages includes at least a first stage having a first blade and disk configuration and a second stage having a second blade and disk configuration that is different than the first blade and disk configuration;
• a combustor in fluid communication with the compressor section;
• a turbine section in fluid communication with the combustor, wherein the turbine section comprises a first turbine section that drives the first compressor section via a first shaft at the first speed and a second turbine section that drives the second compressor section via a second shaft at the second speed;
• a geared architecture driven by the turbine section for rotating the fan about the axis, and wherein the geared architecture couples the first shaft to the fan;
• and
• an inducer forming an additional compression section positioned in front of the first and second compressor sections.

A gas turbine engine comprising:

• a fan including a plurality of fan blades rotatable about an axis;
• a compressor section including at least a first compressor section rotating at a first speed and a second compressor section rotating at a second speed greater than the first speed, wherein components of the second compressor section are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.;
• a combustor in fluid communication with the compressor section;
• a turbine section in fluid communication with the combustor;
• a geared architecture driven by the turbine section for rotating the fan about the axis;
• an inducer forming an additional compression section positioned in front of the first and second compressor sections, wherein the inducer is configured to rotate at a higher speed than the fan through an additional output of the geared architecture.