Turbocooled vane of a gas turbine engine
The present invention discloses a novel apparatus and methods for providing a flow of cooling air to one or more turbine nozzles or turbine blade outer air seals. The flow of cooling air is provided by an external source and regulated in order to improve turbine nozzle and air seal cooling efficiency and component life.
1. A method of cooling a turbine nozzle in a gas turbine engine, the gas turbine engine having a compressor section, a compressor discharge plenum, a combustor section, and a turbine section fluidly connected to each other, the method comprising:
compressing a portion of a cooling air for the turbine nozzle through a separate process external to the gas turbine engine to form a supply of compressed air having a pressure above a pressure of air in the compressor discharge plenum;
directing the supply of compressed air to a leading edge of the turbine nozzle; and,
directing a portion of the supply of compressed air from the leading edge to a portion of the turbine nozzle aft of the leading edge;
wherein, the supply of compressed air is used to cool a trailing edge of the turbine nozzle after back side cooling the leading edge of the turbine nozzle.
2. The method according to claim 1 , wherein the supply of compressed air is used to cool a mid-body portion of the turbine nozzle after cooling the leading edge of the turbine nozzle.
3. The method of claim 1 , wherein the separate process external to the gas turbine engine comprises a fueled engine coupled to one or more compressors.
4. A method of cooling a plurality of turbine nozzles in a gas turbine engine having a compressor section, a compressor discharge plenum, a combustor section, and a turbine section all fluidly connected to each other, the method comprising:
compressing a portion of a cooling air to form compressed air for the plurality of turbine nozzles, where the cooling air is compressed in a separate process external to the gas turbine engine; and,
controlling distribution of the compressed air circumferentially to a stage of the plurality of turbine nozzles so as to vary the compressed air provided to each turbine nozzle within the stage;
wherein distribution of the compressed air is based on a predetermined control parameter.
5. The method of claim 4 , wherein air pressure is the predetermined control parameter.
6. The method of claim 4 , wherein temperature is the predetermined control parameter.
7. The method of claim 4 , wherein flow rate is the predetermined control parameter.
8. The method of claim 4 , wherein the predetermined control parameter is a combination of air pressure, temperature and flow rate.
9. A method of cooling a plurality of circumferentially disposed static components in a gas turbine engine, the gas turbine engine having a compressor section, a compressor discharge plenum, a combustor section, and a turbine section fluidly connected to each other, the method comprising:
compressing a portion of cooling air to form compressed air for the plurality of circumferentially disposed static components of the turbine section in a separate process external to the gas turbine engine; and,
controlling distribution of the compressed air circumferentially around the plurality of circumferentially disposed static components based on a control parameter such that a first component of the plurality of circumferentially disposed static components is cooled differently relative to a second component of the plurality of circumferentially disposed static components.
10. The method of claim 9 , wherein the control parameter is air pressure.
11. The method of claim 9 , wherein the control parameter is temperature.
12. The method of claim 9 , wherein the control parameter is flow rate.
13. The method of claim 9 , wherein the control parameter is any combination of air pressure, temperature, and flow rate.