Impeller backface shroud for use with a gas turbine engine

A gas turbine engine comprising:

• a shaft;
• an impeller or axial stage compressor disk affixed to the shaft;
• a turbine affixed to the shaft at a location aft of the impeller;
• and
• an impeller or axial stage compressor disk backface shroud comprising:
• a substantially funnel shaped body having a surface, the substantially funnel shaped body being statically mounted to the gas turbine engine in a position that is substantially coaxial with the impeller or axial stage compressor disk and axially spaced apart therefrom in an aft direction such that the surface and a backface of the impeller or axial stage compressor disk form a cavity, the cavity being configured to guide an airflow portion from the impeller or axial stage compressor disk to the turbine, the airflow portion having a tangential velocity;
• and
• a recessed groove defined in the surface, the recessed groove oriented generally transversely to the tangential velocity of the airflow portion and configured to at least partially interfere with the airflow portion, whereby a static pressure in the cavity is affected.

A method for compensating for an undesirable amount of spool thrust in a gas turbine engine having a shaft, an impeller or axial stage compressor disk affixed to the shaft, a turbine affixed to the shaft at a location aft of the impeller or axial stage compressor disk, and an impeller or axial stage compressor disk backface shroud statically mounted to the gas turbine engine in a position that is substantially coaxial with the impeller and aft thereof such that a surface of the impeller or axial stage compressor disk backface shroud and a backface of the impeller form a cavity configured to guide an airflow portion from the impeller or axial stage compressor disk to the turbine, the airflow portion having a tangential velocity, the method comprising the steps of:

• A. determining a target static pressure;
• B. performing a computational fluid dynamic analysis using a processor to determine a static pressure in the cavity that would result from defining a recessed groove in the surface of the impeller or axial stage compressor disk backface shroud, the recessed groove having a predetermined configuration;
• C. changing the predetermined configuration of the recessed groove if the static pressure in the cavity differs substantially from the target static pressure;
• D. repeating steps B and C until a configuration of the recessed groove that yields a static pressure in the cavity that does not differ substantially from the target static pressure is determined;
• E. manufacturing a second impeller or axial stage compressor disk backface shroud including the recessed groove having the configuration determined at step D;
• and
• F. assembling the second impeller or axial stage compressor disk backface shroud to the gas turbine engine.