The present invention contemplates an apparatus for forming a combustion mixture in a gas turbine engine. In one form, the apparatus includes a diffuser having at least two flowpath structures spaced apart to define a flowpath for directing fluid flow. At least one of the flowpath structures includes a flowpath surface and a plurality of edges disposed along a trailing end portion of the flowpath surface. The edges extend perpendicularly from the flowpath surface and are arranged generally parallel to the fluid flow to define a plurality of corners. A fluid vortex is generated as the fluid flow rolls over each of the corners. A spray ring is disposed along a trailing edge of one of the flowpath structures and is integrally attached to the diffuser. The spray ring includes a plurality of fuel delivery apertures adapted to spray fuel into respective ones of the fluid vortices to form the combustion mixture.

Apparatus for forming a combustion mixture in a gas turbine engine

Warping of the rear turbine shroud (54) and the resulting cracking of turbine nozzle vanes (24) may be eliminated in a radial inflow turbine having an annular combustor (30) with a radially outer wall (32) opening on a compressed air plenum (48) by locating a plurality of tubes (70) on the radially outer wall (32) in fluid communication with the plenum (48) and extending the tubes (70) inwardly so that the radially inner ends (76) are adjacent the radially outer edge (56) of the turbine shroud (54). The tubes (70) thus direct sweeping streams of cooling air along the rear turbine shroud (54) to cool the same and prevent warpage.

Jet air cooled turbine shroud for improved swirl cooling and mixing

A curved diffuser ( 210 ) in a gas turbine engine ( 201 ) directs a primary portion of air flow from a compressor ( 202 ) through a curved discharge opening ( 213 ) into a plenum ( 220 ). The curved diffuser ( 210 ) also comprises ports ( 217 ) through which a secondary portion of air passes into confined space ( 225 ) that is defined in part by a pressure boundary element that may be comprised of at least one plate ( 222 ) or at least one conduit ( 306 ). The at least one plate ( 222 ) and the at least one conduit ( 306 ) respectively comprise apertures ( 246, 312 ) through which pass the secondary portion of air to provide impingement-type cooling to transitions ( 230, 320 ). In various embodiments the velocity of the air between adjacent transitions ( 230, 320 ) may flow at relatively uniform velocity along the longitudinal distance of the respective transitions ( 230, 320 ).

Gas turbine engine curved diffuser with partial impingement cooling apparatus for transitions

Improved combustor volume in turbine engines without an increased engine envelope is accomplished in an engine (10) having a radial compressor (12), an axial turbine (16), and a radial combustor (20) in which a generally radial flow path extends between a compressed air inlet (30) and a turbine nozzle (22) for gases of combustion produced by combusting fuel introduced through a fuel injector (34) with air introduced through the compressed air inlet (30).

Axial flow gas turbine engine combustor

A combustor assembly is provided that includes a forward annular liner, an aft annular liner, and a stator. The forward annular liner has an inner annular edge and an outer annular edge. The aft annular liner is spaced apart from the forward annular liner to form a combustion chamber. The aft annular liner has an inner annular edge and an outer annular edge. The stator has a forward end and an aft end. The stator forward end includes an inner annular edge and an outer annular edge. The stator forward end inner annular edge is coupled to the aft annular liner outer annular edge via a first bayonet mount system, and the stator forward end outer annular edge is coupled to the forward annular liner outer annular edge via a second bayonet mount system.

Dual bayonet engagement and method of assembling a combustor liner in a gas turbine engine

A turbine engine construction comprising a support housing, a compressor, combustor and turbine expander contained within the support housing. The combustor includes a combustor housing defining an inner combustion chamber and an outer chamber between the combustor housing and the support housing. Compressed air from the compressor outlet is communicated to the combustion chamber into which fuel is injected and ignited so that the resulting combustion products expand through a turbine nozzle and into the turbine expander. The novelty of the present invention resides in the construction of a diffuser assembly integrated with both the turbine nozzle and combustor. The diffuser assembly comprises a plurality of diffuser tubes, each tube having one end open to and in registration with the compressed air outlet of the compressor. The other ends of the tubes are arranged in at least two discharge planes for the combustor to improve and increase induced mixing and swirl in the combustion chamber. Preferably, at least one set of diffuser tubes discharges compressed air within the mixing zone of the combustion chamber and at a predetermined axial angle with respect to the turbine engine axis to provide a more uniform temperature into and through the turbine nozzle. The angular discharge from this set of diffuser tubes enables the turbine nozzle to be unwrapped from the conventional axial direction thus reducing both the axial length of the combustor and also the aerodynamic loading on the turbine nozzle.

Dennis E. Barbeau

Papers

Turbine engine construction