Abstract: A gas turbine engine centrifugal compressor axial forward thrust apparatus bleeds impeller tip aft bleed flow from between an annular centrifugal compressor impeller of a high pressure rotor and a diffuser directly downstream of the impeller. The apparatus then uses the aft bleed flow to pressurize an annular cavity extending radially between an inner combustor casing and the rotor and extending axially between forward and aft thrust balance seals. Forward and aft thrust balance lands are in sealing engagement with the forward and aft thrust balance seals on the high pressure rotor respectively. An annular stator plenum in fluid communication with the annular cavity is bounded in part by a forward end wall having conical diffusion holes therethrough which may be axially or circumferentially or axially and circumferentially angled. The forward thrust balance seal may be on an aft conical arm of the impeller.

Abstract: A gas turbine engine turbine cooling system includes an impeller and a diffuser directly downstream of the impeller, a bleed for bleeding clean cooling air from downstream of the diffuser, and one or more channels in fluid communication with the bleed. Each of the channels having a generally radially extending section followed by a generally axially aftwardly extending section terminating at an annular cooling air plenum connected to accelerators. The radially and axially aftwardly extending sections may be connected by a bend section of the cooling air channel and the axially aftwardly extending section may be angled radially inwardly going from the bend section to the cooling air plenum. Each of the cooling channels includes an inner wall formed by a forward end wall extending radially outwardly from an inner combustor casing, an annular cover covering a radially inner portion of the forward end wall, and the inner combustor casing.

Abstract: A gas turbine engine centrifugal compressor axial forward thrust apparatus bleeds impeller tip aft bleed flow from between an annular centrifugal compressor impeller of a high pressure rotor and a diffuser directly downstream of the impeller. The apparatus then uses the aft bleed flow to pressurize an annular cavity extending radially between an inner combustor casing and the rotor and extending axially between forward and aft thrust balance seals. Forward and aft thrust balance lands are in sealing engagement with the forward and aft thrust balance seals on the high pressure rotor respectively. An annular stator plenum in fluid communication with the annular cavity is bounded in part by a forward end wall having conical diffusion holes therethrough which may be axially or circumferentially or axially and circumferentially angled. The forward thrust balance seal may be on an aft conical arm of the impeller.

Abstract: A gas turbine engine turbine cooling system includes an impeller and a diffuser directly downstream of the impeller, a bleed for bleeding clean cooling air from downstream of the diffuser, and one or more channels in fluid communication with the bleed. Each of the channels having a generally radially extending section followed by a generally axially aftwardly extending section terminating at an annular cooling air plenum connected to accelerators. The radially and axially aftwardly extending sections may be connected by a bend section of the cooling air channel and the axially aftwardly extending section may be angled radially inwardly going from the bend section to the cooling air plenum. Each of the cooling channels includes an inner wall formed by a forward end wall extending radially outwardly from an inner combustor casing, an annular cover covering a radially inner portion of the forward end wall, and the inner combustor casing.

Abstract: A method facilitates manufacturing a combustor for a gas turbine engine. The combustor includes an inner and an outer liner, a dome, and flanges. The method comprises assembling the inner and outer liner such that each includes a series of liner panels, positioning a spatter shield against a downstream side of the inner and outer liners, wherein the spatter shield comprises molybdenum, positioning a mounting flange against the downstream side of the inner and outer liners, such that the spatter shield is positioned at least partially between the mounting flange and the inner and outer liners, and coupling the mounting flange to the inner and outer liners.

Abstract: A method facilitates manufacturing a combustor for a gas turbine engine. The combustor includes an inner and an outer liner, a dome, and flanges. The method comprises assembling the inner and outer liner such that each includes a series of liner panels, positioning a spatter shield against a downstream side of the inner and outer liners, wherein the spatter shield comprises molybdenum, positioning a mounting flange against the downstream side of the inner and outer liners, such that the spatter shield is positioned at least partially between the mounting flange and the inner and outer liners, and coupling the mounting flange to the inner and outer liners.

Abstract: A method facilitates manufacturing a combustor for a gas turbine engine. The combustor includes an inner and an outer liner, a dome, and flanges. The method comprises assembling the inner and outer liner such that each includes a series of liner panels, positioning a spatter shield against a downstream side of the inner and outer liners, wherein the spatter shield comprises molybdenum, positioning a mounting flange against the downstream side of the inner and outer liners, such that the spatter shield is positioned at least partially between the mounting flange and the inner and outer liners, and coupling the mounting flange to the inner and outer liners.