Gm Global Technolgoy Operations, Inc.

Samuel B Reider

Avrum S Herman

Thomas E Scott

Arthur H LeFebvre

Jerry G Tomlinson

Cecil H Sharpe

Henry M Mar

A jet propulsion engine with a fan bypass duct includes a duct burner with a plurality of flame stabilizers therein each mounted to inner case and outer case members through spherical bearings. Each of the stabilizers consists of two blade members having integral arms thereon actuated by fore and aft motion of an external actuating ring to assume an expanded position to increase duct turbulence for mixing air flow therethrough with a fuel supply and into a retracted position against each other to reduce pressure drop under nonafterburning operation. Each of the flame stabilizer blades has a platform that controls communication between a hot air source and a duct for improving fuel vaporization during afterburner operation thereby to increase afterburning limits; the platforms close communication between the hot air source and the duct during nonafterburning operation when flame stabilization is not required.

Low loss duct burner

In a gas turbine engine having an annular combustor and an axial flow compressor discharging compressed air through a discharge annulus upstream of said combustor, the diffuser being supported on the engine independently of the combustor between the discharge annulus and a dome on the combustor and including inner and outer walls operative to direct first and second portions of compressor discharge to radially inner and outer air plenums at the combustor and a remaining third portion of compressor discharge flow to a combustor dome feed chamber ahead of the dome, an improvement in the form of radially oriented pins on the diffuser and guides on the combustor dome slidably received on the pins, the pins and guides cooperating in maintaining a predetermined positional relationship between the diffuser and the combustor during relative thermal growth of the latter whereby at normal operating temperature the diffuser outer wall seals against the combustor while a flow balancing slot is developed between the diffuser inner wall and the combustor.

Diffuser for gas turbine engine

An air blast nozzle includes a ring of air swirler vanes therein for producing a first turbulent flow of air therethrough and a turbulence generator located within the air flow through said swirler vanes to produce an intensification of the turbulent flow of the air swirl produced by the swirler vanes. The intensified swirled air is immediately directed into mixed flow relationship with fuel from said air blast nozzle to maintain increased turbulence intensity at the fuel from said nozzle without attenuation of the air energy level prior to burning of the air/fuel mixture so as to improve combustor efficiency by evaporation rate control due to decreased fuel droplet size attributable to the turbulence intensification.

High intensity air blast fuel nozzle

A gas turbine engine combustor has a porous wall structure for directing coolant from a compressed air plenum into a combustion chamber to cool its inner surface; an outer lamina has a first hole pattern therein and an inner lamina diffusion bonded to the outer lamina has a second hole pattern offset from the first hole pattern to direct air from the air plenum in cross-flow relationship across internal walls to cool the outer and inner lamina and the outer lamina has a plurality of relief notches formed part way through its depth which are responsive to thermal gradients across the porous wall structure to break the outer lamina into a plurality of separated shingles thereby to prevent excessive stress formation in the porous wall structure; the inner lamina including flow control means to maintain a regulated flow of coolant from an inlet air plenum to the combustion chamber irrespective of the number of relief notches broken during operation of the gas turbine engine.

Shingled laminated porous material

A porous, laminated metal fabrication includes first and second porous walls, each having laminae therein with a free edge portion across at least one end thereof; each of the first and second walls including an outer lamina with a first preformed hole pattern therein, each of the first and second walls further including an inner lamina having a second preformed hole pattern therein combined to form a tortuous air flow path through the first and second walls for cooling the metal therein; and each of the lamina in the laminae including a solid metal annulus therein of uniform density for welding; the annulus being located between the hole patterns and the free edge portions to define a weldable region between the first and second walls having an axial width limited to the axial width of the solid metal annulus whereby air flow through the first and second walls will flow freely throughout the full extent of all the performed hole patterns therein so as to maintain full coolant flow from exteriorly of the porous laminated fabrication to an inner surface thereon.

Porous laminated material

A combustor assembly for a gas turbine engine includes a tubular, multi-layered porous metal wall with pores and cavities for distribution of compressor discharge air into a combustion chamber and wherein a rigid combustor support ring connects one end of the porous wall to fixed support components of the gas turbine engine; the combustor assembly further including a porous metal transition member for joining the tubular porous metal wall to a downstream nozzle plate; the transition member including free formed side walls and top and bottom walls without substantial reduction in permeability due to forming and each of the transition walls further including a sharp radius bend at the side edge thereof defining a compressed porous metal section for connection to an adjacent sharp radius bend by means of a longitudinal seam weld along the length thereof to maximize the flow of secondary cooling air into the hot combustion gases passing from the outlet of the combustion chamber of the apparatus.

Porous laminated combustor structure

A combustor apparatus for use in gas turbine engines includes an annular dome of porous laminated material joined by a solid metal joint ring to inner and outer annular porous wall segments also constructed of porous laminated material; and wherein the solid metal joint ring includes a plurality of small diameter offset effusion holes for directing coolant through the solid metal ring at a rate to cool it without entrapment of unburned fuel against the combustor walls; the combustor further including a transition mount including a solid metal support ring welded to the aft end of a porous laminated metal combustor wall, the mount including a seal ring retained by a keeper and a thin conical, axially extending beam portion that is flexible to serve to accommodate differential thermal expansion of the combustor relative to a grounded support component and wherein the solid metal support also includes a plurality of effusion cooling holes located immediately outboard of a weld bead between the solid metal support and the porous laminated combustor wall for cooling a downstream nozzle.

Combustor liner joints

An airblast atomizer with air swirl vanes has an outlet configured to produce a fuel/air mixture pattern from the fuel nozzle to match an annular combustion chamber formed between an annular inner combustor wall and an annular combustor wall having inlet ends thereof connected to an annular dome for supporting each of a plurality of the fuel nozzles; each of the nozzles including a pair of spaced, large radius slots in a cylindrical outlet wall of the nozzle operative to produce a vortex pattern that will spread the mixed fuel/air pattern from the nozzle to define the major axis of an elliptically formed fuel/air pattern from the nozzle and wherein the remainder of the airblast atomizer nozzle has a reduced radius to conform the fuel/air mixture to the diameter of an outlet swirl chamber from the nozzle to produce the minor axis of an elliptically configured fuel/air spray pattern and wherein the major axis is formed as a chordal line of the combustion chamber to conform the fuel/air mixture pattern from each of the fuel nozzles to the annular shape of the combustor chamber thereby to reduce the number of airblast atomizer fuel nozzles required to completely cover a predetermined diameter of combustion chamber annulus.

Elliptical airblast nozzle

An air blast fuel supply system for a gas turbine engine comprises a floating swirler separated from the fuel injector and means for radially supporting both the swirler and fuel injector for free radial movement with respect to a combustor dome; a fuel atomization lip on the floating swirler is located in spaced overlying relationship to a tangential fuel director to form an annular fuel film at the outlet of the fuel injector and an outer annular air flow directing lip on the floating swirler directs inlet air flow against the fuel film as it leaves the atomization lip. The fuel injector includes a nozzle tube that slips to permit free axial movement of said fuel injector with resepct to the dome and wherein the tangential fuel director maintains the annular fuel film throughout axially shifted positions of said nozzle tube. This allows the fuel nozzle to be inserted through a small opening in the engine case while maintaining the integrated relationship with the swirler attached to the combustor. The fuel atomization lip has an outlet edge thereon and an outer annular air flow directing lip has outlet edge thereon maintained at a constantly fixed dimensional relationship therebetween throughout axial shifted positions of the nozzle tube whereby the fuel break-up point for atomization of fuel and air remains the same with respect to the combustor during engine operation.

Integrated, replaceable combustor swirler and fuel injector

A combustor assembly for a gas turbine engine includes a tubular, multi-layered porous metal wall with pores therethrough for distribution of compressor discharge air into a combustion chamber and wherein a rigid combustor support ring is connected to one end of the wall to receive an inlet diffuser member including an ovate inlet and a circular outlet connected to the support ring for axially directing primary air flow into the combustion chamber; and the diffuser member including a flow divider on one outer surface thereof including means for fixedly securing the inlet diffuser member with respect to a wall of a gas turbine engine and wherein coacting means are provided between the inlet diffuser member and the rigid combustor support ring of the combustor to radially connect it in place at one end thereof and wherein further coacting means are provided between the inner engine wall and a portion of the porous sleeve for axially indexing of the combustor assembly.

Combustor support structure

An industrial gas turbine engine includes an inclined annular combustor made up of a plurality of support segments each including inner and outer walls of trapezoidally configured planar configuration extents and including side flanges thereon interconnected by means of air cooled connector bolt assemblies to form a continuous annular combustion chamber therebetween and wherein an air fuel mixing chamber is formed at one end of the support segments including means for directing and mixing fuel within a plenum and a perforated header plate for directing streams of air and fuel mixture into the combustion chamber; each of the outer and inner walls of each of the support segments having a ribbed lattice with tracks slidably supporting porous laminated replaceable panels and including pores therein for distributing combustion air into the combustion chamber while cooling the inner surface of each of the panels by transpiration cooling thereof.

Segmented annular combustor

A short length afterburner assembly for a jet propulsion engine having a fan bypass includes cold and hot air cross-over passages and a plurality of flame stabilization swirler vanes associated with a balanced load controller for positioning the vanes parallel to hot gas stream flow from a jet engine core when the afterburner is off and in an inclined position to such gas stream flow when fuel is injected therein during afterburner operation thereby to produce flame spread within the afterburner core by a combination of translatory and swirling motions; and wherein atomized fuel for afterburner combustion is injected into hot gases ducted through hot air cross-over passages from the jet engine core to produce premix and prevaporization of fuel upstream of fixed flameholders and wherein cold fan bypass air cross-over passages have movable turbulator grids positioned during afterburner operation for mixing cold air flow with the bypassed hot core gas at the fixed flameholders during afterburner operation and wherein the turbulator grids are positioned parallel to gas flow when the afterburner is not in operation by means of the balanced load controller to produce a balanced variable geometry mechanical load to that on the flame stabilization swirler vanes.

Afterburner flow mixing means in turbofan jet engine

Combustion apparatus for a gas turbine engine includes two separate combustion zones; one having a first air flow thereto defined by a plurality of holes in combustion apparatus liner; the second zone has air flow thereto through a second plurality of holes in the combustion apparatus liner. Parallel fuel injectors supply fuel into the respective zones and include a wide angle fuel flow nozzle that supplies fuel to the pilot zone of combustion and a narrow cone fuel injector that directs fuel through the pilot zone without combustion and into a downstream main combustion region where the fuel is then burned with air flow though the second plurality of holes to produce a staged combustion of air fuel in the combustion apparatus to achieve high combustion efficiency over a broad ange of fuel/air ratios as well as more uniform combustor outlet temperatures and low smoke output.

Growing community of inventors

Indianapolis, IN, United States of America

Samuel B Reider