Gas turbine engine with improved high pressure compressor leakage path guide structure

1 . A gas turbine engine having an axial centerline, comprising:

a compressor section having a high pressure compressor (HPC), wherein the HPC includes an HPC aft hub;

a combustor section having a combustor;

a turbine section;

wherein a core gas path extends through the compressor section, the combustor section, and the turbine section;

a tangential on board injector (TOBI) system that extends circumferentially around the engine axial centerline, the TOBI system having a plurality of nozzles and a TOBI inner radial flange; and

an HPC leakage guide structure (LGS) having a LGS forward end and a LGS aft end, wherein the LGS forward end is disposed and configured to receive a leakage flow from the HPC and the LGS aft end is engaged with the TOBI inner radial flange;

wherein the HPC leakage guide structure and the HPC aft hub are separated from one another and define an HPC aft hub cavity therebetween, wherein the HPC aft hub cavity extends between the LGS forward end and the LGS aft end,

wherein the HPC aft hub cavity monotonically increases in a direction from the LGS forward end to the LGS aft end,

wherein the HPC aft hub cavity monotonically increasing from the LGS forward end of the HPC leakage guide structure to a position where the LGS aft end and the TOBI inner radial flange engage, such that a radial height of the HPC aft hub cavity at the LGS forward end is less than a radial height of the HPC aft hub cavity at the position where the LGS aft end and the TOBI inner radial flange engage; and

wherein the LGS aft end axially overlaps and is radially biased against the TOBI inner radial flange.

2 . The gas turbine engine of claim 1 , wherein the LGS aft end and the TOBI inner radial flange are configured for mechanical engagement with one another.

3 . The gas turbine engine of claim 2 , wherein the mechanical engagement of the LGS aft end and the TOBI inner radial flange is a mating configuration.

4 . The gas turbine engine of claim 3 , wherein the mating configuration is a splined configuration.

5 . The gas turbine engine of claim 1 , wherein the HPC leakage guide structure has a thickness that extends between two surfaces and the HPC leakage guide structure is a solid panel.

6 . The gas turbine engine of claim 1 , wherein the HPC leakage guide structure has a thickness that extends between two surfaces and the HPC leakage guide structure includes a plurality of apertures extending through the thickness.

7 . The gas turbine engine of claim 1 , further comprising: an inner diffuser case disposed radially inside of and spaced apart from the combustor, wherein the inner diffuser case and the combustor define a diffuser inner diameter cavity therebetween; and

wherein the inner diffuser case is disposed radially outside of and spaced apart from the HPC leakage guide structure, wherein the inner diffuser case and the HPC leakage guide structure define a diffuser secondary cavity therebetween.

8 . The gas turbine engine of claim 7 , wherein the HPC leakage guide structure is connected to the inner diffuser case.

9 . The gas turbine engine of claim 7 , wherein the HPC leakage guide structure includes an LGS mounting flange disposed adjacent the LGS forward end and the LGS mounting flange is connected to the inner diffuser case.

10 . The gas turbine engine of claim 7 , wherein the HPC leakage guide structure has a thickness that extends between two surfaces and the HPC leakage guide structure is a solid panel.

11 . The gas turbine engine of claim 7 , wherein the HPC leakage guide structure has a thickness that extends between two surfaces and the HPC leakage guide structure includes a plurality of apertures extending through the thickness thereby providing a fluid passage between the HPC aft hub cavity and the diffuser secondary cavity.

12 . A gas turbine engine having an axial centerline, comprising:

a compressor section having a high pressure compressor (HPC), wherein the HPC includes an HPC aft hub;

a combustor section having a combustor;

a turbine section;

wherein a core gas path extends through the compressor section, the combustor section, and the turbine section;

a tangential on board injector (TOBI) system that extends circumferentially around the engine axial centerline, the TOBI system having a plurality of nozzles and a TOBI inner radial flange; and

an HPC leakage guide structure (LGS) extending between a LGS forward end and a LGS aft end, wherein the LGS forward end engages the HPC aft hub, the LGS forward end is disposed and configured to receive a leakage flow from the HPC, and the LGS aft end is engaged with the TOBI inner radial flange;

wherein the HPC leakage guide structure and the HPC aft hub are separated from one another and define an HPC aft hub cavity therebetween,

wherein the HPC aft hub cavity extends between the LGS forward end and the LGS aft end, wherein the HPC aft hub cavity monotonically increases from the LGS forward end to a distal end of the TOBI inner radial flange which engages the LGS aft end, such that a radial height of the HPC aft hub cavity at the LGS forward end is less than a radial height of the HPC aft hub cavity at the distal end of the TOBI inner radial flange.

13 . The gas turbine engine of claim 12 , wherein the LGS aft end axially overlaps and is radially biased against the TOBI inner radial flange.