Abstract: A combustor in which fuel is introduced into a combustion zone at a first acute angle with respect to the radially outward direction so as to swirl the fuel about the combustor axis, thereby improving the stability of the combustion process. Combustion air is introduced into the combustion zone at a second acute angle with respect to the radially inward direction so as to swirl the air about the combustor axis in the same circumferential direction that the fuel is swirled, with the first and second acute angles being approximately equal. By swirling the combustion air in a manner that is compatible with the swirling of the fuel, noise associated with the combustion process is reduced.

Abstract: Sleeves (28) are circumferentially spaced from one another about the liner of a combustor body (12) of a dry low NO.sub.x combustor. The sleeves carry dilution air into the dilution zone. Cooling air is supplied a venturi (20) to cool the venturi and the cooling air flows into the reaction volume. The dilution air sleeves penetrate sufficiently to thoroughly mix the dilution air with the core of hot gases of combustion and, by vorticity effects caused by the flow past the sleeves, thoroughly mix the generally annular flow of cooling air from the venturi with the hot gases of combustion. The thorough mixing of both the cooling air and dilution air inhibits or minimizes the formation of cold areas or streaks within the reaction volume such that CO to CO.sub.2 reactions are not quenched, affording reduced CO emissions.

Abstract: A gas turbine engine has a low cost combustor liner cooling system combining the benefits of high internal heat removal with improved film cooling by employing a large number of strategically positioned, laser-drilled cooling passages. Cooling air flows through these specially tailored passages to absorb heat from the liner prior to injection as a protective film on the interior surface. The passages are set in staggered rows on a thickened portion of the liner and have a rough internal heat transfer surface and an exit with a steep injection angle to evenly distribute the cooling film along the interior surface of the liner.

Abstract: An oxidizer intake flow control system is disclosed for a gas turbine engine combustion chamber. The chamber has an oxidizer intake assembly which extends through the wall of the chamber such that an inlet aperture of the oxidizer intake and an exhaust aperture of the oxidizer intake are located on opposite sides of the wall defining the combustion chamber. An oxidizer intake sleeve defines a central passage through which oxidizer may pass into the combustion chamber. The intake passage extends along a central axis and the intake assembly is attached to the wall of the combustion chamber such that it is rotatable about this central axis. The inlet aperture and/or the exhaust aperture is located in a plane extending non-perpendicularly to the central axis. The combustion chamber may have several oxidizer intake assemblies extending through the combustion chamber walls. The intake assemblies may be interconnected such that they may be simultaneously rotated with respect to the combustion chamber wall.

Abstract: A combustion chamber wall structure is formed of a plurality of segments, each segment having an upstream portion and a downstream portion. The upstream portion is attached to an outer surface of an adjacent wall segment at its downstream portion in an overlapping fashion such that the overlapping portions form a cooling air chamber. The main cooling air holes are defined by the upstream portion so as to communicate with the cooling chamber, which has an exit to direct the cooling air along an inner surface of the downstream portion of the wall segment. A primary air hole is defined by the downstream portion of the segment and is located adjacent to the main cooling air holes. This location maximizes the distance between the primary air hole and an upstream portion of the same segment, which, in turn, maximizes the distance along the inner surface of the segment on which the cooling air may flow without being disrupted by air passing through the primary air hole.

Abstract: A burner assembly includes a register including an annular wall defining a chamber, the annular wall including a plurality of coaxial annular bands and a plurality of parallel, axially spaced, circumferentially extending slots which separate the bands and through which air is supplied to the chamber, the annular bands forming a plurality of bluff body elements to the air, the annular bands being formed separate and independent of each other; a plurality of spacer washers inserted between adjacent annular bands to define the slots; a plurality of connecting rods extending through openings in the spacing washers for maintaining the spacing washers in a predetermined relationship and for maintaining the spacing washers between the annular bands; an oil and/or gas gun for supplying a combustible material to the chamber; and a gas pilot for igniting a mixture of the combustible material and the air in the chamber.

Abstract: A gas turbine engine combustor, eg of annular combustor configuration, has air inlets, typically in the dilution zone, which incorporate vortex generators operative to trip the inlet jet flow causing stabilizing vortex flow at the lateral edges of the jet. The vortex generators can comprise an inwardly directed abrupt change of section between forward and rearward parts of the air inlets. The rearward section of the air inlets can have a configuration operative to reinforce the jet stabilizing effect of the vortex flow. The combustor exhibits improved uniformity of jet and crossflow mixing.

Abstract: A plurality of gas diffusion openings (24) direct film cooling air (26) along the surface of sheet (20). Each opening (24) includes a baffle (28) extending into gas stream (10) defining gas chamber (30) with outlet opening (32). Cool air inlet opening (34) meters and directs impingement air against the upstream end (36) of the baffle. cooling air discharge is parallel to the surface to be cooled and the baffle directs the upstream hot air flow.

Abstract: A radial flameholder (18) extends from the augmentor case (22) through an opening (30) in a coaxial liner (20) and into the hot exhaust gas stream (26). A portion of a cool, annular fan air stream (28) enters an upstream facing opening (34) and discharged radially inwardly adjacent the downstream edge (44) of the liner opening (30). The inner flow (42) of cool air protects the liner (20) downstream of the flameholder (18) by both cooling the liner (20) locally and by displacing the combustion reaction (32) attached to the flameholder (18).

Abstract: The expense of fabricating an annular combustor for use with a small diameter gas turbine engine is minimized in a construction including a bell-shaped housing 52, an axially extending sleeve 66 within the housing 52, an annular liner 74 within the housing 52 and about the sleeve 66 and including inner, outer and radial walls 80, 76, 78, respectively, spaced from the sleeve 66 and the housing 52 and all being formed of sheet metal. A circular fuel manifold 108 with angularly spaced fuel dispensing openings 122 is disposed between the housing 52 and the outer wall 76 and a plurality of open ended, elongated tubes 112 extend through the outer wall 76 and are in fluid communication with respective openings 122 in the fuel manifold 108 to introduce air and fuel generally tangentially into the liner 74.

Abstract: A gas turbine combustion chamber with an inlet and outlet has apertures in the combustion chamber wall and air scoops provided in the apertures, the air scoops having an outer tubular member and coaxially spaced inner tubular member. The spaced tubular members have inner cylindrical portions and outwardly extending flanges, with spacers disposed between and secured to the flanges with cooling air passing therebetween and through the annular air flow passage into the combustion chamber. Preferably, a radially outwardly extending arcuate section on the inner tubular member has a radius at least about one-third of the inner diameter of the inner cylindrical portion thereof.

Abstract: An air admission insert particularly adapted for air admission holes in a combustor liner in a gas turbine combustion system comprises a pair of sleeve members, one of which is inserted in the other in axial but eccentric relationship so that the sides of the sleeves contact each other with the intervening space between the sleeves being crescent shaped. The insert is positioned concentrically in a combustor liner air admission hole. Combustion air flows through the insert and into the combustor liner and axially through the crescent space to cool the insert as well as the liner hold perimeter.

Abstract: A combustion chamber for a turbojet engine is disclosed which utilizes double walled construction. Hot walls are attached to adjacent cold walls solely by an orifice member which defines an orifice to supply primary of dilution air to the combustion chamber. The attachment allows relative radial movement of the hot walls with respect to the cold walls so as to accommodate the thermal expansion and contraction of the hot walls. The sleeves forming the walls also cooperate so as to provide an internal cooling film on the interior surface of the hot walls. Convection cooling may also be utilized between the walls to prevent structural damage due to high temperatures.

Abstract: A liner cooling construction for a gas turbine combustor or the like in which twisted or helical tapes in the spaces defined between the liner walls and the partition walls cause the cooling air to swirl so that the efficient cooling of the liner walls can be ensured.

Abstract: A refilmer for placing the film of cool air impeded by combustion/dilution air holes of a combustor for a gas turbine engine includes a grommet with a lip extending axially in the combustion chamber. The outer side wall of the grommet is recessed in the area overlying the lip so that when assembled, it forms a gap with the liner exposing the lip to cool air for forming the film. Tabs on the cool air side prevent the grommet from falling into the combustion chamber. The aperture in the grommet is oriented to direct the combustion/dilution air stream at a relative angle to the flow of combustion products. Indexing surfaces on the grommet fixture locates the grommet to the liner prior to being welded into place.

Abstract: A burner assembly includes a register having a circular front wall and an annular side wall; a plurality of cylindrical bluff body elements circumferentially spaced about the annular side wall in a plurality of axially spaced rows for supplying air to the register; fuel supply assemblies which supply at least one of coal, gas and oil to the register for mixing with the air; and a plurality of bluff body discs, each positioned within a respective bluff body element for increasing the velocity and pressure drop of the air entering the register so as to enhance mixing of the fuel and the air within the register, and to reduce the NOx.

Abstract: A method of fabricating sheet metal panels and the article produced thereby. According to one form, the method of fabrication includes the steps of providing a panel of sheet metal, perforating the panel to provide a plurality of holes, forming the panel into a preselected curve about a longitudinal centerline, forming the leading edge portion of the panel into a front flange, forming a shoulder in the panel centered on the holes and extending perpendicularly from a surface thereof, bending an outer portion of the shoulder into a lip, and bonding the portions of the panel comprising the shoulder and lip.

Abstract: A gas turbine combustion including a combustion wall having a plurality of generally cylindrical passages therein which extend around and are generally parallel to, the chamber axis, each of said passages having a series of air inlets disposed along the passage and tangentially directed thereinto, generally towards the center of the chamber, and means for directing cooling air into the series of air inlets in each passage to set up a forced vortex therein. The series of air inlets may be arranged to set up forced vortices of opposite rotation in adjacent passage. A means for directing cooling air into the tube inlets may comprise an annular shroud which extends around the air inlets and directs airflow into the air inlets.

Abstract: A combustion system comprising a tubular structure having an upstream end and defining within it a longitudinal passage. A tubular wall surrounding the structure and defining with the structure a further longitudinal passage, of annular cross section. The tubular structure having longitudinally extending, circumferentially spaced staves and longitudinally spaced coaxial hoops for the passage of fluid, the hoops being shaped to deflect substantially axially flowing fluid either inwards or outwards through the openings for discharge either inside or outside the assembly respectively.

Abstract: A combustor for a gas turbine is provided, the combustor including a reverse flow type combustion chamber defined by an inner cylindrical wall and an outer wall concentric with the inner wall and connected at one end by an annular dome. The outer wall has three successive sections with the first section being cylindrical and of a predetermined diameter, the second outer wall section being of frusto-conical shape, and the third wall section being of cylindrical shape but of reduced diameter compared to the first section. Injectors are provided in the angled wall of the second outer wall section, and openings are provided annular with the injectors to allow compressor delivered air to enter directly into the combustion chamber surrounding the injector, the injector being directed towards the primary zone.

Abstract: A counter flow combustor for regenerative open cycle turbine systems has an outer casing defining a head cavity at one end, an inner casing mounting in said outer casing, said inner casing defining a combustion chamber. An annular insulated partition and support means is connected at its upper end to the inner wall of the outer casing and at its lower end to the discharge outlet end of the inner casing, said inner casing including a cooling and metering wall assembly about the secondary combustion zone having cooling means communicating at one end with a head cavity extension space and at the other end with the secondary combustion zone for delivering heated combustion air thereto and support means on the insulated partition and support means operatively connected to the cooling and metering wall assembly to hold the same in assembled position in the combustor.

Abstract: The airflow from the compressor of a gas turbine engine into the combustion chamber or chambers principally comprises combustion air which enters the upstream end of the combustion chamber and dilution air which enters the combustion chamber at some point downstream of the combustion chamber inlet, the combustion air being made up of primary and secondary air entering the combustion chamber primary and secondary zones.In order to cope with the control of emissions from the combustion chambers, it is desirable to control the combustion air fuel ratio over the engine operating range by controlling air mass flow to the combustion and dilution zones of the combustion chamber. The invention proposes that this control can be achieved by providing a variable rate diffuser upstream of the primary, secondary and dilution air inlets, the variable rate diffuser comprising vortex generator having an associated variable air bleed.

Abstract: A gas turbine combustion chamber in which the primary air inlets in a peripheral wall of the combustion chamber are defined by open-ended tubes extending inwardly from the peripheral wall by a substantial distance into the combustion chamber and having their inner, that is their outlet, ends facing in an upstream direction within the combustion chamber, whereby each of the tubes will introduce a stream of air with at least a component of motion in the upstream direction along or parallel with the longitudinal axis of the combustion chamber, thereby to effect recirculation of fuel, air and combustion gases within the combustion chamber. Similar air inlet tubes may be provided for introducing secondary and tertiary air into the combustion chamber.

Abstract: The NOx emissions of a stationary gas turbine are reduced by concentrating a NOx suppressant in the reaction zone of the gas turbine combustor by dividing the flow of air to the reaction zone and the dilution zone of the combustor by means of an air flow splitter and by taking advantage of the radially stratified compressor flow. The air flow to the two zones is preferably separated by a common flow shield.

Abstract: A counter flow combustor has a generally elongated hollow cylindrical outer casing or shell with a closure means at the upper end defining a head space and a generally elongated smaller hollow annular inner casing or combustion wall mounted therein in spaced relation to the outer casing to form an annular head space extension between the side walls of the outer casing and the inner casing which is continuous with the head space. The inner casing or combustion wall defines a primary combustion zone, and a secondary combustion zone in communication at one end with the primary combustion zone and at the end remote therefrom with the discharge outlet for delivering combustion gases from the combustor. An inlet assembly is mounted in said head space for mixing fuel and air and for delivering the same in proper ratio for combustion in said primary combustion zone and secondary combustion zone.

Abstract: 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 suppor

Abstract: A combustion chamber of a gas turbine engine has a combustion cowl and transition piece with at least one air gap provided therebetween. The air gap opens into an intermediate chamber having two throttles which are arranged oppositely with respect to one another. The two throttles are provided in a cooling jacket which is divided into two portions by the throttles. The air gap permits the combustion cowl and transition piece to expand freely with respect to each other. The two throttles are arranged so that fuel gas leaving the combustion chamber provides an injector effect to expedite the flow of a coolant such as air through the air gap.

Abstract: A combustion apparatus for a gas turbine engine comprises a combustion chamber having primary and secondary combustion zones, a fuel injector having a series of primary fuel nozzles and a series of secondary fuel nozzles, and primary and secondary fuel and air duct means to direct fuel and air mixtures to the primary and secondary combustion zones respectively.

Abstract: A covered air scoop for a gas turbine combustor comprising a tubular member radially extending through the combustor shell to define an air inlet. A symmetrical cap member is disposed in spaced, overhanging relationship over the inlet. To enter the combustor, the air must flow into the cap member, thereby causing all air to enter the tube in a substantially uniform manner so that the air flow through all similarly constructed air tubes at a common axial location of the combustor is substantially equal at all times.