A cooling assembly for a gas turbine system includes a turbine nozzle having at least one channel comprising a channel inlet configured to receive a cooling flow from a cooling source, wherein the at least one channel directs the cooling flow through the turbine nozzle in a radial direction at a first pressure to a channel outlet. Also included is an exit cavity for fluidly connecting the channel outlet to a region of a turbine component, wherein the region of the turbine component is at a second pressure, wherein the first pressure is greater than the second pressure.

An aerodynamic element of a turbomachine including a body extending according to a radial main direction and a radial end located at a radial end of the body. The aerodynamic element further includes a heat exchanger between an internal fluid of the turbomachine and an airflow flowing around the body of the aerodynamic element. The heat exchanger includes a heat pipe in which a working fluid circulates and as well as an evaporation portion in which the working fluid exchanges heat with the internal fluid and a condensation portion in which the working fluid exchanges heat with the airflow.

An aerodynamic element of a turbomachine including a body extending according to a radial main direction and a radial end located at a radial end of the body. The aerodynamic element further includes a heat exchanger between an internal fluid of the turbomachine and an airflow flowing around the body of the aerodynamic element. The heat exchanger includes a heat pipe in which a working fluid circulates and as well as an evaporation portion in which the working fluid exchanges heat with the internal fluid and a condensation portion in which the working fluid exchanges heat with the airflow.