Geared turbofan engine with high compressor exit temperature
A gas turbine engine includes a fan with a plurality of fan blades rotatable about an axis, and a compressor section that includes at least first and second compressor sections. An average exit temperature of the compressor section is between about 1000° F. and about 1500° F. The engine also includes a combustor that is in fluid communication with the compressor section, and a turbine section that is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.
1. 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, and 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., and wherein the fan is configured to operate at a redline speed of at least 3200 rpm at the average exit temperature;
a combustor in fluid communication with the compressor section;
a turbine section in fluid communication with the combustor, 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;
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 high and low pressure compressors.
2. The gas turbine engine according to claim 1 wherein the average exit temperature is between 1100° F. and 1450° F.
3. The gas turbine engine according to claim 1 wherein the fan drives air along a bypass flow path in a bypass duct defined between a fan nacelle and a core nacelle, and wherein a bypass ratio is greater than ten.
4. The gas turbine engine according to claim 1 wherein the geared architecture has a gear ratio that is greater than 2.4.
5. The gas turbine engine according to claim 1 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.
6. The gas turbine engine according to claim 1 wherein the inducer is configured to rotate at a speed common with that of the fan.
7. The gas turbine engine as set forth in claim 1 , wherein the geared architecture is positioned intermediate the fan and a compressor rotor driven by the low pressure turbine.
8. The gas turbine engine according to claim 1 wherein the first compressor section is configured to rotate at a redline speed of at least 10,000 rpm and the second compressor section is configured to rotate at a redline speed of at least 22,000 rpm at the average exit temperature.
9. The gas turbine engine as set forth in claim 1 wherein the average exit temperature is defined at Sea Level, end of takeoff power and at a rated thrust for the gas turbine engine.
10. 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, and 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.;
the high pressure compressor including a plurality of stages with each stage comprising a disk with a plurality of stages includes at least a first stage having a first blade and disk configuration and 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 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;
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 high and low pressure compressors.
11. The gas turbine engine according to claim 10 wherein the first stage is positioned forward of the second stage.
12. 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, and 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.;
a combustor in fluid communication with the compressor section;
a turbine section in fluid communication with the combustor, 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;
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 high and low pressure compressors, and wherein the inducer is configured to rotate at a higher speed than the fan through an additional output of the geared architecture.
13. 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, and 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.;
a combustor in fluid communication with the compressor section;
a turbine section in fluid communication with the combustor, 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;
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 high and low pressure compressors; and
an intermediate pressure turbine section that drives a compressor rotor.
14. 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, and 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.;
a combustor in fluid communication with the compressor section;
a turbine section in fluid communication with the combustor, 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;
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 high and low pressure compressors, and wherein the geared architecture is positioned upstream of the low pressure turbine and downstream of a compressor rotor driven by the low pressure turbine.
15. 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, and 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.;
a combustor in fluid communication with the compressor section;
a turbine section in fluid communication with the combustor, 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;
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 high and low pressure compressors, and wherein the turbine section includes a fan drive turbine that is coupled to drive the geared architecture, an intermediate pressure turbine configured to drive a first compressor rotor, and a turbine rotor configured to drive a second compressor rotor.