Geared turbofan engine with high compressor exit temperature

1. A gas turbine engine comprising:

a fan including a plurality of fan blades rotatable about an axis;

a plurality of shafts that include at least a first shaft rotating at a first speed and a second shaft rotating at a second speed that is faster than the first speed;

a compressor section including at least a low pressure compressor associated with the first shaft and a high pressure compressor associated with the second shaft;

a combustor downstream of the compressor section;

a turbine section downstream of the combustor, wherein the turbine section comprises a low pressure turbine that drives the low pressure compressor via the first shaft and a high pressure turbine that drives the high pressure compressor via the second shaft;

a geared architecture driven by the low pressure turbine and coupling the first shaft to the fan, wherein the geared architecture is positioned upstream of the low pressure compressor; and

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, 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 an integrally bladed rotor such that there are no rim cavities or associated honeycomb seals, and wherein the first stage is located upstream of the second stage.

2. The gas turbine engine as set forth in claim 1 , wherein components of the high pressure compressor are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.

3. The gas turbine engine as set forth in claim 2 , wherein the average exit temperature is between 1100° F. and up to 1450° F.

4. The gas turbine engine as set forth in claim 1 , including an inducer comprising a third compression section that is upstream of the low pressure compressor.

5. The gas turbine engine as set forth in claim 4 , wherein the inducer comprises a blade that can be a slotted blade or a blisk coupled to a hub that is coupled to rotate with the fan.

6. The gas turbine engine as set forth in claim 4 , wherein the inducer is configured to rotate at a higher speed through an additional output of the geared architecture.

7. 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.

8. The gas turbine engine as set forth in claim 1 , wherein the integrally bladed rotor comprise a second disk with associated second blades that are machined out of a common material such that no dovetails, slots, or other fastening features are required to fix the second blades to the second disk.

9. The gas turbine engine as set forth in claim 8 , including an inducer comprising a third compression section that is upstream of the low pressure compressor.

10. The gas turbine engine as set forth in claim 1 , wherein the plurality of fan blades comprises no more than twenty-six fan blades.

11. The gas turbine engine as set forth in claim 10 , wherein the plurality of fan blades comprises no more than twenty fan blades.

12. The gas turbine engine as set forth in claim 10 , wherein the low pressure turbine includes no more than six turbine rotors.

13. The gas turbine engine as set forth in claim 12 , wherein the low pressure turbine includes three turbine rotors.

14. The gas turbine engine as set forth in claim 12 , wherein a ratio between a number of fan blades and a number of turbine rotors for the low pressure turbine is between 3.3 and 8.6.

15. The gas turbine engine as set forth in claim 1 , wherein the first stage is located at a front of the high pressure compressor and the second stage is located at a rear of the high pressure compressor.

16. The gas turbine engine as set forth in claim 1 , wherein the compressor section drives air along a core flow gaspath for compression and communication into the combustor and then expansion through the turbine section, and wherein the integrally formed blisks have associated rims that are flush with the gaspath.