Combustor liner for gas turbine engine
View Patent ↗A combustor liner ( 230 ) for a gas turbine engine combustor ( 200 ) comprises an inner wall ( 232 ), an outer wall ( 238 ), a cooling air flow channel ( 244 ) formed there between, and a flow control ring ( 246 ). The flow control ring ( 246 ) is sealingly attached to the downstream ends of the inner wall ( 232 ) and the outer wall ( 238 ), and comprises a plurality of holes ( 250 ) that, during gas turbine engine operation, may regulate a flow of cooling air that passes through the cooling air flow channel ( 244 ). One or more surfaces may be coated with a thermal barrier coating ( 237 ) to provide additional protection from thermal damage.
1. A can combustor for a gas turbine engine comprising:
an intake, an outlet, and at least one swirler assembly disposed there between;
a combustion liner comprising an inner wall that partly defines a combustion zone, wherein the combustion zone and the inner wall share a common longitudinal axis and the common longitudinal axis is disposed in the combustion zone, and an outer wall disposed radially outward of the inner wall with respect to the common longitudinal axis, the inner wall and outer wall defining there between a flow channel for passage of a cooling air flow; and
a flow control ring directly connected to downstream ends of the inner and outer walls of the combustor proximate the outlet, wherein the flow control ring comprises a plurality of holes in fluid communication with the flow channel and the combustion zone effective to deliver a plurality of discrete flows of cooling fluid into the combustion zone, and wherein during operation the plurality of holes is effective to control the cooling air flow.
2. The combustor of claim 1 , additionally comprising a number of effusion holes through the inner wall.
3. The combustor of claim 1 , additionally comprising a thermal barrier coating on a portion of an inner surface of the inner wall.
4. The combustor of claim 3 , wherein the portion is a major portion of the inner surface.
5. The combustor of claim 1 additionally comprising a protective barrier covering an upstream opening to the flow channel, connecting to at least one of the inner wall and the outer wall, and comprising a plurality of holes for passage of cooling air into the flow channel.
6. The combustor of claim 1 , wherein the flow control ring supports by rigid attachment thereto a spring clip assembly extending radially outward.
7. The combustor of claim 6 , wherein the outer wall supports by rigid attachment thereto a cylindrical barrier structure formed to limit inward movement of the spring clip assembly and to restrict passage of spring clip fragments.
8. A gas turbine engine comprising the combustor of claim 1 .
9. In a can annular gas turbine engine combustor having a combustion liner comprising an inner wall partly defining a combustion zone, wherein the combustion zone and the inner wall share a common longitudinal axis and the common longitudinal axis is disposed in the combustion zone, and an outer walls disposed radially outward of the inner wall with respect to the common longitudinal axis, the inner wall and outer wall defining a flow channel therebetween, the improvement comprising a flow control regulator connected directly to downstream ends of inner and outer walls of the combustor, wherein the flow control regulator further comprises a plurality of holes in fluid communication with the flow channel and the combustion zone that are effective to deliver a plurality of discrete flows of cooling fluid into the combustion zone.
10. A combustor liner assembly for a gas turbine engine can combustor comprising an outer wall disposed radially outward of an inner wall, wherein the inner wall partly defines a combustion zone, wherein the combustion zone and the inner wall share a common longitudinal axis and the common longitudinal axis is disposed in the combustion zone, each said wall comprising an inlet end and an outlet end, a channel between the inner wall and the outer wall, a flow control regulator welded to downstream ends of the inner and outer walls of the combustor proximate the outlet ends and comprising a plurality of holes in fluid communication with the flow channel that are effective to deliver a plurality of discrete flows of cooling fluid into the combustion zone.
11. A gas turbine engine combustor comprising the combustor liner assembly of claim 10 .
12. A gas turbine engine comprising the combustor of claim 11 .
13. A can annular gas turbine combustion engine comprising a plurality of combustor cans disposed therein, each said combustor can comprising:
an intake, an outlet, and at least one swirler assembly disposed there between;
a combustion liner comprising an inner wall that partly defines a combustion zone, wherein the combustion zone and the inner wall share a common longitudinal axis and the common longitudinal axis is disposed in the combustion zone and an outer wall disposed radially outward of the inner wall with respect to the common longitudinal axis, the inner wall and outer wall defining there between a flow channel for passage of a cooling air flow; and
a flow control ring connected directly to downstream ends of the inner and outer walls and comprising a plurality of holes in fluid communication with the flow channel and the combustion zone effective to deliver a plurality of discrete flows of cooling fluid into the combustion zone, wherein collectively said plurality of holes are effective to provide a uniformly controlled cooling among each respective combustor liner wall.