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A method of assembling a gas turbine engine includes providing a gas turbine engine including a compressor, a combustor downstream from the compressor, and a turbine coupled to the compressor, and coupling a seal assembly aft of the compressor such that air discharged through the seal assembly facil

A method of assembling a gas turbine engine includes providing a gas turbine engine including a compressor, a combustor downstream from the compressor, and a turbine coupled to the compressor, and coupling a seal assembly aft of the compressor such that air discharged through the seal assembly facilitates reducing the cavity pressure, and therefore reducing the axial forces induced to an aft side of the compressor.

대표청구항 ▼

What is claimed is: 1. A method of assembling a gas turbine engine comprising: coupling a combustor downstream from a compressor; coupling a turbine to the compressor, wherein a recess is formed in a turbine mid-seal support structure of said turbine; and orienting the turbine relative to the compr

What is claimed is: 1. A method of assembling a gas turbine engine comprising: coupling a combustor downstream from a compressor; coupling a turbine to the compressor, wherein a recess is formed in a turbine mid-seal support structure of said turbine; and orienting the turbine relative to the compressor such that a cavity is defined between the compressor and the turbine and such that a seal assembly extending from an aft side of the compressor is positioned such that when air is discharged through the seal assembly, the seal assembly facilitates reducing the operating pressure within the cavity, and facilitates reducing axial forces induced to the aft side of the compressor. 2. The method in accordance with claim 1 wherein the compressor comprises a high-pressure centrifugal compressor. 3. The method in accordance with claim 1 wherein orienting the turbine relative to the compressor further comprises inserting a tab extending from the high-pressure compressor at least partially into the recess to form a two-stage labyrinth seal assembly. 4. The method in accordance with claim 3 further comprising positioning the tab such that the tab and recess define a first orifice stage and a second orifice that is downstream from the first orifice stage. 5. The method in accordance with claim 4 wherein the first orifice has a first diameter and the second orifice has a second diameter that is different than the first diameter. 6. The method in accordance with claim 5 wherein the first and second diameters are sized to regulate the quantity of airflow channeled through the seal assembly. 7. The method in accordance with claim 1 wherein orienting the turbine relative to the compressor further comprises positioning a labyrinth seal assembly extending from the axially aft side of the compressor such that air discharged through the seal assembly is utilized to create a pressure drop aft of the compressor and to facilitate reducing the forward axial load of the compressor. 8. A seal assembly for a gas turbine engine including a combustor downstream from a centrifugal compressor, a turbine coupled to the compressor, and a turbine mid-seal support structure, said seal assembly comprising: a recess formed in the turbine mid-seal support structure; and a labyrinth seal assembly extending aft of the centrifugal compressor such that air discharged through said seal assembly facilitates reducing the operating pressure within a cavity formed between said compressor and said turbine, and reduces axial forces induced to an aft side of the centrifugal compressor. 9. The seal assembly in accordance with claim 8 wherein said seal assembly is coupled proximate to a radially outer edge of the centrifugal compressor. 10. The seal assembly in accordance with claim 8 further comprising a tab extending from the compressor at least partially into said recess to form a two-stage labyrinth seal assembly. 11. The seal assembly in accordance with claim 10 wherein said two-stage labyrinth seal assembly comprises a first orifice stage and a second orifice that is downstream from the first orifice stage. 12. The seal assembly in accordance with claim 11 wherein said first orifice has a first diameter and said second orifice has a second diameter that is different than said first diameter. 13. The seal assembly in accordance with claim 12 wherein said first and second diameters are sized to regulate the quantity of airflow channeled through said seal assembly. 14. The seal assembly in accordance with claim 8 wherein said labyrinth seal assembly is positioned on the axially aft side of the compressor such that air discharged through the seal assembly creates a pressure drop aft of the compressor and facilitates reducing the forward axial load of the compressor. 15. A gas turbine engine assembly comprising: a centrifugal compressor; a combustor downstream from said centrifugal compressor; a turbine coupled to said compressor, such that a cavity is defined between the compressor and the turbine; a turbine mid-seal support structure, such that a recess is formed therein; and a seal assembly extending aft of said centrifugal compressor such that air discharged through said seal assembly facilitates reducing the operating pressure within said cavity, and reduces axial forces induced to an aft side of said centrifugal compressor. 16. The gas turbine engine assembly in accordance with claim 15 wherein said seal assembly is positioned proximate to a radially outer edge of the centrifugal compressor. 17. The gas turbine engine assembly in accordance with claim 15 wherein said seal assembly further comprises a tab extending from said compressor at least partially into said recess to form a two-stage labyrinth seal assembly. 18. The gas turbine engine assembly in accordance with claim 16 wherein said seal assembly comprises a first orifice stage and a second orifice that is downstream from the first orifice stage. 19. The gas turbine engine assembly in accordance with claim 18 wherein said first orifice has a first diameter and said second orifice has a second diameter that is different than said first diameter. 20. The gas turbine engine assembly in accordance with claim 19 wherein said first and second diameters are sized to regulate the quantity of airflow channeled through said seal assembly.