Inter-turbine burner in recuperation cycle engine

1. A turbine engine assembly comprising:

a compressor section where a core airflow is compressed;

a combustor section where the compressed core airflow from the compressor section is mixed with fuel and ignited to generate an exhaust gas flow;

a turbine section including at least two turbine stages through which the exhaust gas flow expands to generate a mechanical power output;

an inter-turbine burner disposed between the at least two turbine stages, wherein in the inter-turbine burner an additional amount of fuel is mixed with the exhaust gas flow to generate a reheated gas flow;

a condenser where water is extracted from the reheated gas flow exhausted from the turbine section; and

an evaporator system disposed between a turbine stage of the at least two turbine stages and the condenser, the evaporator system configured for generating a steam flow from at least a portion of water extracted by the condenser for injection into the core airflow, wherein the evaporator system includes a first evaporator disposed between a first turbine stage and a second turbine stage of the at least two turbine stages and a second evaporator disposed between the turbine section and the condenser.

2. The turbine engine assembly as recited in claim 1 , wherein the turbine section includes a first turbine section and at least one other turbine section and the inter-turbine burner is disposed between the first turbine section and the at least one other turbine section.

3. The turbine engine assembly as recited in claim 1 , wherein a portion of the steam flow is communicated to the combustor section and added to the gas flow generated in the combustor section.

4. The turbine section as recited in claim 1 , wherein the first evaporator generates a first steam flow and the second evaporator generates a second steam flow, where both the first steam flow and the second steam flow are communicated to at least one of the combustor and the inter-turbine burner.

5. The turbine engine as recited in claim 1 , wherein the first evaporator receives a flow of steam or hot water from the second evaporator.

6. The turbine engine assembly as recited in claim 1 , wherein a bypass flow through a bypass flow path is communicated to the condenser for cooling the exhaust gas flow.

7. The turbine engine assembly as recited in claim 1 , wherein the inter-turbine burner comprises a different power output capacity than the combustor section.

8. An aircraft propulsion system comprising:

a compressor section where a core airflow is compressed;

a combustor section where the compressed core airflow is mixed with fuel and ignited to generate an exhaust gas flow;

a first turbine section through which the exhaust gas flow expands to generate power;

an inter-turbine burner where the exhaust gas flow from the first turbine section is mixed with fuel and ignited to generate a reheated gas flow;

a second turbine section through which the reheated gas flow expands to generate additional power;

a condenser where water is extracted from the reheated gas flow exhausted from the second turbine section; and

an evaporator system disposed between the inter-turbine burner and the condenser, the evaporator system generating a steam flow by heating the extracted water with heat from the reheated exhaust gas flow, wherein a portion of the steam flow is communicated to the combustor section and added to the gas flow generated in the combustor section, wherein the evaporator system includes a first evaporator separated from a second evaporator, wherein each of the first evaporator and the second evaporator transfer heat from the reheated gas flow into the extracted water.

9. The aircraft propulsion system as recited in claim 8 , wherein the evaporator system communicates a steam flow to each of the combustor section and the inter-turbine burner.

10. The aircraft propulsion system as recited in claim 8 , wherein a mass flow rate of the core flow into the condenser is more than 10% higher than the core airflow within the compressor section.

11. The aircraft propulsion system as recited in claim 10 , wherein the molecular oxygen content of the core flow exiting through a core nozzle is less than 10% oxygen by mass at high engine power.