Abstract: A gas turbine engine assembly includes a combustor configured to combust an air-fuel mixture to produce combustion gases in a first direction; a transition liner coupled to the combustor and adapted to receive the combustion gases from the combustor and to redirect the combustion gases in a second direction; and a turbine coupled to the transition liner and adapted to receive the combustion gases from the transition liner. The transition liner has a plurality of effusion holes that include a first group that extend at least partially in a tangential direction.

Abstract: A combustor module for a gas turbine engine is provided that includes a first annular liner assembly extending along a longitudinal axis of the engine. The first annular liner assembly includes a first annular support shell and a plurality of first heat shield panels coupled to the first annular support shell. The first heat shield panels form a segmented ring defining a plurality of first axial seams therebetween. The combustor module further includes a bulkhead coupled to the first annular liner assembly. The bulkhead provides a plurality of fuel nozzles for passing a first mass flow comprising fuel and air. The combustor module further includes a second annular liner assembly coupled to the bulkhead. The second annular liner assembly is in spaced-apart generally coaxial relationship from the first annular liner assembly by a channel height H. The second annular liner assembly includes an air admittance hole having a mean diameter D extending along a hole axis.

Abstract: An annular combustor for a gas turbine has a combustion chamber having an interior volume that, in longitudinal section, includes a forward volume, an intermediate volume and an aft volume. The forward volume represents from about 30% to about 40% of the combustor interior volume, the intermediate volume represents from about 10% to about 20% of the combustor interior volume, and the aft volume represents from about 40% to about 60% of the combustor interior volume.

Abstract: A combustion system includes a combustor having a forward annular liner having a first plurality of effusion holes, and an aft annular liner having a second plurality of effusion holes and forming a combustion chamber with the forward annular liner. The first plurality of effusion holes and the second plurality of effusion holes are adapted to receive compressed air from a compressor and contribute to a single toroidal recirculation air flow pattern in the combustion chamber. The combustion system further includes a rotary fuel slinger further adapted to receive a flow of fuel from a fuel source and to centrifuge the received fuel into the combustion chamber; and an igniter extending at least partially into the combustion chamber to ignite the fuel and compressed air in the combustion chamber, to thereby generate combusted gas.

Abstract: The invention relates to a heat shield element for lining a combustion chamber wall, to a combustion chamber and to a gas turbine. The heat shield element comprises a hot side that can be exposed to a hot medium, a wall side opposite said hot side, and a peripheral side adjoining the hot side and the wall side and having a peripheral side surface. Relief slots are introduced into the material in an area of the heat shield element that is susceptible to material cracks induced by thermal stress, thereby limiting crack propagation.

Abstract: A turbomachine combustion chamber has primary and dilution air inlet orifices formed by die stamping to have edges that project into the inside of the combustion chamber, and stress relaxation and/or reduction means in the edges or in the vicinity of the edges of said orifices, said means comprising, for each orifice, one, two, or three slots formed in the edge or around a fraction of the edge of said orifice.

Abstract: A combustor includes an outer liner having a first row and a second row of circumferentially distributed air admission holes penetrating therethrough. The combustor further includes an inner liner circumscribed by the outer liner and having a first row and a second row of circumferentially distributed air admission holes penetrating therethrough. The combustor further includes a plurality of fuel injectors extending into the combustion chamber and configured to deliver an air-fuel mixture to the combustion chamber. Each of the plurality of fuel injectors are associated with two air admission holes in the first row of the outer liner, two air admission holes in the first row of the inner liner, four air admission holes in the second row of the outer liner, and two air admission holes of the second row of the inner liner.

Abstract: A crossfire tube assembly is configured for connecting adjacent combustion cans in a gas turbine, and includes a first tube segment having a first end and an opposite female end. A second tube segment has a first end and an opposite male end fitted concentrically within the female end with an overlap region between the female and male ends. Each of the first ends of the tube segments is configured for securing to a liner of a respective combustion can. Oppositely oriented first and second impingement sleeves extend from the female end of the first tube segment to the respective first ends of the tube segments, Combustion cooling air is directed through metering holes in the impingement sleeves and flows axially along concentric cavities defined between the impingement sleeves and the first and second tube segments. The combustion cooling air vents from the cavities into an axial combustion air flow stream between the respective combustion can liners and sleeves.

Abstract: The present invention provides near-surface cooled airfoils that can be made with near-surface cooling passages that are completely free of any leachable or otherwise sacrificial material in the recessed portion of the outer surface of the core. The turbine airfoil comprises a metallic core or substrate having an outer surface and one or a plurality of recessed portions of the outer surface; an intermediate metallic skin or foil having a back surface and a top surface, the back surface of the intermediate skin being bonded to the outer surface of the core such that the recessed portion(s) is sufficiently enclosed so as to form at least one or more near-surface cooling passages or pathways; and at least one or more metallic coatings of a high temperature-resistant metallic material deposited on a top surface of the intermediate skin.

Abstract: A gas turbine engine has a combustor module including an annular combustor having a liner assembly that defines an annular combustion chamber and includes a circumferential row of a plurality of relatively large combustion dilution air admission holes and a circumferential row of a plurality of smaller quench air admission holes disposed downstream with respect to the flow of combustion gas products. The plurality of quench air admission holes are arranged with respect to the plurality of relatively large dilution air admission holes disposed upstream thereof such that there is associated with each dilution air admission hole a first quench air admission hole and a second quench air hole, the first quench air hole being offset laterally in a first lateral direction and the second quench air hole being offset laterally in a second lateral direction opposite to the first direction.

Abstract: A method and apparatus to facilitate reducing NOx emissions in turbine engines is provided. The method includes providing an annular shell including a plurality of circumferentially extending panels. The plurality of circumferentially extending panels includes a first panel positioned at an upstream end of the shell and a second panel positioned downstream from, and adjacent to, the first panel. The method also includes forming a plurality of primary dilution holes in the first panel and forming a plurality of secondary dilution holes in the second panel. The dilution holes are configured to discharge dilution air into the shell.

Abstract: An annular combustion chamber of a turbomachine is provided. The combustion chamber includes an end wall provided at an upstream end of the chamber and side walls extending longitudinally from the end wall to an orifice for discharging a stream of combustion gases provided at a downstream end of the chamber. The side walls includes at least one row of openings for the intake of air for diluting the stream of combustion gases. At least one dilution opening has an upstream edge which projects toward the inside of the chamber and a downstream edge which projects toward the outside of the chamber and is asymmetric to the upstream edge with respect to a plane extending transversely to the wall. An aperture of the opening having an axis oriented in an oblique direction with respect to the wall. This direction being oriented toward the inside and toward the downstream end of the chamber.

Abstract: Embodiments of a transition (400) of the present invention comprise a cooling channel (20, 30, 410) defined in part by an outer wall (23) and an inner wall (24). The cooling channel (20, 30, 410) also comprises lateral side walls (33, 34). Two or more subdomains (AA, BB, A, B1, B2, C1, C2, D) of impingement jets (25, 35) are provided through the outer wall (23), and one or more metering outlets (26, 36, 37, 38) is provided through the inner wall (24), all communicating with a respective channel (20, 30, 410). The impingement jets of each respective subdomain are designed to supply cooling fluid to one of the one or more metering outlets.

Abstract: A cooling arrangement for a surface of a wall in a gas turbine engine, the wall having a plurality of effusion holes each with an outlet onto the surface for supplying an effusion flow to the surface and an inlet, the inlets of the effusion holes being arranged at the peripheries of groups tessellated on an opposing surface of the wall, each inlet being located on the peripheries of three groups. The arrangement comprises a second wall spaced apart from the opposing surface having impingement orifices each for directing a flow of air in use to a respective impingement location on the opposing surface, each group having a centrally positioned impingement location.

Abstract: A gas turbine combustion chamber wall has an outer wall skin and an inner wall skin, with the outer wall skin (9) and the inner wall skin (10) being arranged essentially parallel to each other and spaced apart from each other by a gap (14). The outer wall skin (9) is provided with inlet openings (8) for the supply of cooling air. The inner wall skin (10) is provided with dampening openings (17), whose center axes are perpendicular to the inner wall skin (10), and with cooling openings (18), whose center axes are inclined at a certain angle to the inner wall skin (10).

Abstract: A combustion chamber head of a gas turbine has a confinement enclosing a dampening volume (207) and including a combustion chamber-opposite confinement (206) and a combustion chamber-side confinement (210). The combustion chamber-side confinement (210) is provided as perforated wall (210). In the edge area of the combustion chamber-side confinement (210), cooling air can be routed onto the combustion chamber-side confinement (210) via recesses (203) in the confinement (206). This cooling air, which flows along the combustion chamber-side confinement (210), crosses the cooling air flow through the perforated wall (210) in the combustion chamber (101) without mixing with the latter, as both are separated by walls.

Abstract: The present application provides a combustor for a gas turbine engine. The combustor may include a combustor can with a number of nozzles therein and a flow conditioner positioned around the combustor can. The flow conditioner may include a number of apertures therein.

Abstract: A combustor having liners with a plurality of angled effusion holes defined therethrough at a first angle with respect to a surface of the liners and at a second angle with respect to a corresponding radial plane. A density of the effusion holes defined in a primary section receiving the fuel nozzles is at least equal to a density of the effusion holes defined in a secondary downstream section.

Abstract: Apparatus for adaptive cooling comprising a first component having at least one aperture extending therethrough with a sacrificial component positioned within the at least one aperture. The first component is operable at a maximum duty temperature and the sacrificial component has a melting or sublimation point below the maximum duty temperature of the first component. The sacrificial component defines an effective aperture the size of which may be increased if, in use, the sacrificial component is subjected to a temperature between the melting or sublimation point of the sacrificial component and the maximum duty temperature of the first component.

Abstract: A combustor liner (231) for a gas turbine engine combustor (200) comprises an inner wall (232), an outer wall (238), a flow channel (244) formed there between, and an end-capping ring (246). The end-capping ring (246) is sealingly attached to the downstream end of the inner wall (232). In operation air passes within the end-capping ring (246), into the flow channel (244), and through holes (250) disposed in the inner wall (232). In some embodiments, an end-capping ring variation, a flow-diverting ring (357) comprises a plurality of holes (360) that, during gas turbine engine operation, may additionally dispense a flow of cooling air. One or more surfaces may be coated with a thermal barrier coating (237) to provide additional protection from thermal damage.

Abstract: The invention relates to a heat shield block, particularly for lining a combustion chamber wall, with a hot side that can be subjected to the action of hot medium and with a wall side situated opposite the hot side. A core area, which has a core material, extends inside the heat shield block from the hot side to the wall side. The core area is surrounded by an edge area with an edge material whose heat conductivity is lower than that of the core material. This targeted thermal insulation in the edge area provided in the form of a material bond between the core material and the edge material renders the heat shield block particularly unsusceptible to the formation and growth of cracks in the core area on the hot side. The invention also relates to a combustion chamber provided with heat shield blocks of the aforementioned type, and to a gas turbine provided with a combustion chamber comprising such a heat shield block.

Abstract: A combustor may include an outer liner having a first group of air admission holes and defining a plurality of outer liner regions. The combustor may further include an inner liner having a second group of air admission holes and defining a plurality of inner liner regions. The first group of air admission holes within a respective outer liner region may include a first plunged air admission hole approximately axially aligned with the respective fuel injector, a second plunged air admission hole approximately on the outer boundary line between the respective outer liner region and a first adjacent outer liner region, the first air admission hole being downstream of the second air admission hole, and a third plunged air admission hole approximately on the outer boundary line between the respective outer liner region and a second adjacent outer liner region.

Abstract: A combustor may include an outer liner having a first row and a second row of circumferentially distributed air admission holes. The second row of the outer liner may be downstream of the first row of the outer liner, and the air admission holes of the first row of the outer liner may be larger than the air admission holes of the second row of the outer liner. An inner liner is circumscribed by the outer liner and has a third and fourth row of circumferentially distributed air admission holes. The air admission holes of the third row of the inner liner may be larger than the air admission holes of the fourth row of the inner liner. The inner and outer liners may form a combustion chamber, and at least a portion of the air admission holes of the first, second, third, or fourth rows may be plunged.

Abstract: A burner (27) of a gas turbine engine (10) includes a cylindrical basket (60) comprising an air flow reversal region (86). The flow reversal region ends at an air inlet plane (84) of the basket. The burner also includes a flow conditioner (90) disposed in the flow reversal region transecting an air flow (80) flowing non-uniformly through the flow reversal region, the flow conditioner being effective to mitigate variation of the air flow entering the basket across the inlet plane.

Abstract: A method facilitates assembling a gas turbine engine including a combustor assembly and a nozzle assembly. The method comprises providing a transition piece including a first end, a second end, and a body extending therebetween, where the body includes an inner surface, an opposite outer surface, coupling the first end of the transition piece to the combustor assembly, and coupling the second end of the transition piece to the nozzle assembly such that a turbulator extending helically over the outer surface of the transition piece extends from the transition piece first end to the transition piece second end to facilitate inducing turbulence to cooling air supplied to the combustor assembly.

Abstract: A cooling arrangement for cooling a single-piece, combined combustor liner/transition piece substantially enclosed within a surrounding flow sleeve, with a cooling annulus radially between the flow sleeve and the single-piece combined combustor liner/transition piece, the cooling arrangement including a first plurality of impingement cooling holes in the flow sleeve, the plurality of impingement cooling holes having first diameters and arranged to direct cooling air onto designated areas of the single-piece, combined combustor liner/transition piece; and a second plurality of effusion cooling holes in the single-piece, combined combustor liner/transition piece having second diameters smaller than the first diameters, and located to cool by effusion other areas of the single-piece, combined combustor liner/transition piece.

Abstract: A gas turbine combustor includes a liner having a forward end and an aft end; a flow sleeve surrounding the liner, the flow sleeve also having forward and aft ends, the aft end of the flow sleeve supporting an annular ring formed with a plurality of cooling bores and extending through the flow sleeve, at least some of the plurality of cooling bores formed at an acute angle relative to a longitudinal axis of the liner.

Abstract: An annular wall for the combustion chamber of a turbomachine has a cold side and a hot side, said wall being provided with a plurality of primary holes and a plurality of dilution holes distributed in circumferential rows, together with a plurality of cooling orifices that are distributed in a plurality of circumferential rows that are spaced apart axially from one another, the number of cooling orifices being identical in each row thereof, and the wall further including bores disposed immediately downstream from the primary holes and from the dilution holes and distributed in circumferential rows, the bores in any one row presenting a substantially identical diameter, being spaced apart at a pitch that is constant, and presenting intrinsic characteristics that are different from the intrinsic characteristics of the cooling orifices of the adjacent rows.

Abstract: A combustor for a turbine engine includes an outer liner having a first group of air admission holes and defining a plurality of outer liner regions. The combustor further includes an inner liner circumscribed by the outer liner and forming a combustion chamber therebetween, the inner liner having a second group of air admission holes and defining a plurality of inner liner regions. The combustor further includes a plurality of fuel injectors extending into the combustion chamber and configured to deliver an air-fuel mixture to the combustion chamber, each of the plurality of fuel injectors being associated with one of the outer liner regions and one of the inner liner regions. The first group within a respective outer liner region includes air admission holes that circumferentially alternate between approximately a first size and approximately a second size, the first size being different than the second size.

Abstract: This invention relates to a solid oxide fuel cell system comprising at least one longitudinally extending tubular solid oxide fuel cell and a longitudinally extending heater mounted in thermal proximity to the fuel cell to provide heat to the fuel cell during start up and during operation as needed. The heater and fuel cell can be encased within a tubular thermal casing; the inside of the casing defines a first reactant chamber for containing a first reactant, such as oxidant. The fuel cell comprises a ceramic solid state electrolyte layer and inner and outer electrode layers concentrically arranged around and sandwiching the electrolyte layer. The outer electrode layer is fluidly communicable with the first reactant, and the inner electrode layer is fluidly isolated from the first reactant and fluidly communicable with a second reactant, such as fuel.

Abstract: A combustor for a gas turbine engine is provided, the combustor having an outer shell with an outer surface exposed to cooling air and an inner surface, and at least one floatwall panel attached to the inner surface of the outer shell and having a trailing edge. At least one dilution hole is in the floatwall panel near the trailing edge and in communication with the outer surface of the outer shell, and at least one local air impingement hole is in the outer shell downstream of each at least one dilution hole, that directs the cooling air towards the trailing edge of the floatwall panel.

Abstract: A combustor includes a first liner; and a second liner forming a combustion chamber with the first liner. The combustion chamber is configured to receive an air-fuel mixture for combustion therein, and the first liner is a first dual wall liner having a first hot wall facing the combustion chamber and a first cold wall. The first cold wall has a first cold wall orifice and the first hot wall has a first hot wall orifice. A first insert is mounted in the first hot wall orifice and is configured to receive a first air jet of pressured air from the first cold wall orifice, and guide the first jet through the first hot wall orifice and into the combustion chamber.

Abstract: A dual wall structure for a combustor of a gas turbine engine including an inner liner and an outer liner coupled to a combustor dome and defining a combustion chamber there between. Each of the inner liner and the outer liner include an outer wall and an inner wall. Each of the outer walls includes a plurality of impingement holes formed therein for allowing a coolant to flow therethrough. Each of the inner walls is coupled to the outer wall and a combustor dome and includes a plurality of heat shield panels. Each of the plurality of heat shield panels extends a longitudinal length of the combustor chamber and includes a plurality of side rails, an aft rail and a forward flange that when coupled to the outer wall defines a single cavity there between. A plurality of cavities being formed by the plurality of heat shield panels.

Abstract: A gas turbine engine combustor liner with at least one of the inner and outer liners that is effusion cooled and has a row of groups of circumferentially spaced apart dilution holes defined therethrough. Each group is located within a respective zone of the combustor liner defined by an overlap of adjacent conical sections corresponding to the sprays of adjacent fuel nozzles.

Abstract: Disclosed are exemplary structures with adaptive cooling and methods of adaptively cooling structures. A liner is affixed to a support and the liner deflects away from the support when exposed to a localized hot spot in a hot fluid stream. The liner deflection creates a chamber between the liner and support, allowing cooling air to impinge against the liner, thus mitigating the effects of the hot spot. By providing impingement cooling only where needed, the amount of air needed for cooling is reduced.

Abstract: An apparatus and method of providing a gas turbine combustor having increased combustion stability and reducing pressure drop across a gas turbine combustor is disclosed. A plurality of vanes is fixed to a flow sleeve radially between the flow sleeve and a combustion liner. The plurality of vanes serve to direct a flow of air entering the region between the flow sleeve and combustion liner in a substantially axial direction, such that components of tangential velocity are removed thereby providing a more uniform flow of air the combustion chamber and reducing the amount of pressure lost due attempting to straighten the airflow by pressure drop alone.

Abstract: A gas turbine engine reverse-flow sheet metal combustor for an aero gas turbine engine has a louverless single-piece liner design which includes a plurality of effusion patches, containing multiple rows of effusion cooling holes, bounded by cooled effusionless bands to aid in repairability of the combustor. The cooling of the bands improves the durability of the combustor and provides a method of repair for the combustor.

Abstract: A combustor liner includes a liner having an upstream end and a downstream end having a longitudinal axis extending therethrough, and a plurality of cooling holes formed in the liner, the cooling holes are arranged along the longitudinal axis into a plurality of circumferentially extending rows that are variably spaced apart along the longitudinal axis.

Abstract: A combustor includes an inner liner; an outer liner circumscribing the inner liner and forming a combustion chamber with the inner liner; a combustor dome coupled to the inner and outer liners; and a plurality of heat shields coupled to combustor dome. Each of the heat shields includes a heat shield plate defined by a first edge facing the inner liner and a second edge facing the outer liner; and a plurality of baffles extending from the heat shield plate. Each of the plurality of baffles includes two ribs and a connection portion connecting the two ribs to form a closed portion and an opposite open portion. The open portion of each of the plurality of baffles faces the first edge or the second edge.

Abstract: A combustor for use with a gas turbine. The combustor may include liner, an impingement sleeve, and with the liner and the impingement sleeve defining an airflow channel. The impingement sleeve may include a number of contoured holes therethrough.

Abstract: The invention relates to an open-cooled component for a gas turbine having an outer wall which is subjected to hot gas and which at least partly defines a first cavity for a first medium and in which through-openings are arranged, which through-openings open into the cavity on the one side and into the hot-gas space on the other side, and having at least one second cavity for admixing a second medium, this second cavity being fluidically connected to the through-openings. In order to specify a component for a gas turbine with which flashback and spontaneous ignition during feeding of fuel into the cooling air can be reduced, it is proposed that the second cavity be formed by supply passages which are provided in the outer wall and are connected via transverse passages (4) to the through-openings designed as through-bores, so that the two media cannot be mixed until inside the through-bores.

Abstract: A gas turbine transition piece includes a duct body having a forward end and an aft end, the duct body defining an enclosure for confining a flow of combustion products from a combustor to a turbine first stage nozzle. A plurality of dilution holes are formed in the duct body, located at selected X, Y, Z coordinates measured from a zero reference point at a center of an exit plane of the transition piece.

Abstract: The low emission combustor includes a combustor housing defining a combustion chamber having a plurality of combustion zones. A liner sleeve is disposed in the combustion housing with a gap formed between the liner sleeve and the combustor housing. A secondary nozzle is disposed along a centerline of the combustion chamber and configured to inject a first fluid comprising air, at least one diluent, fuel, or combinations thereof to a downstream side of a first combustion zone among the plurality of combustion zones. A plurality of primary fuel nozzles is disposed proximate to an upstream side of the combustion chamber and located around the secondary nozzle and configured to inject a second fluid comprising air and fuel to an upstream side of the first combustion zone. The combustor also includes a plurality of tertiary coanda nozzles. Each tertiary coanda nozzle is coupled to a respective dilution hole.

Abstract: A combustor assembly includes an inner and an outer liner defining a combustion chamber. The inner and outer liner includes a plurality of cooling holes that are spaced a specified distance apart. The cooling holes include a specified inclination angle and circumferential angle. A first group of cooling holes is spaced apart according to a uniform geometric pattern and density. A second group disposed between the first group and some structural feature within the liner assembly is disposed at a non-uniform pattern and a hole density equal to the density of the first group of cooling holes. The non-uniform cooling hole arrangement increases cooling flow effectiveness to accommodate local disturbances and thermal properties.

Abstract: A method for generating combustion products within a gas turbine engine includes directing an internal air/fuel mixture towards a stagnation point in close proximity to an inner surface of a porous wall defining a combustion chamber. The internal air/fuel mixture is ignited to generate combustion products including a pilot flame. A quantity of air is externally mixed with a quantity of fuel to produce an external air/fuel mixture. The external air/fuel mixture is directed through the porous wall and into the combustion chamber such that the external air/fuel mixture is ignited by the pilot flame. A direction of flow of the combustion products is reversed at the stagnation point.

Abstract: A combustor assembly includes an inner and outer liner defining a combustion chamber. The inner and outer liners include a plurality of cooling holes spaced a specified distance apart. The cooling holes include first, second and third groups. The first group of cooling holes is the most densely spaced, followed by the second group and then the third group. The first group of cooling holes begin upstream of a leading edge of a large opening and terminates downstream of the leading edge. The increased density of cooling holes adjacent the large openings provide increased cooling airflow in areas where cooling may be affected by local disturbances in cooling airflow.

Abstract: An axial-flow pyrospin combustor comprises inner and outer combustor liners and a plurality of pyrospin effusion holes. The inner liner is coaxially mounted inside the outer liner, about a central combustor axis. The pyrospin effusion holes are formed in at least one of the outer combustor liner and the inner combustor liner. Each of the pyrospin effusion holes has a down angle and a back angle, which control a global swirl flow about the central axis, and promote film cooling without detachment.

Abstract: A combustor for a turbine engine includes an outer liner and an inner liner circumscribed by the outer liner to form a combustion chamber therewith in which an air and fuel mixture is combusted to form streamlines of combustion gases. A first plurality of effusion cooling holes is formed in at least one of the inner or outer liners, with the first plurality of effusion cooling holes being oriented as a function of the streamlines.

Abstract: A combustor wall of a gas turbine engine is provided with an acoustic damper component. The component has a first metering passage, a first damping chamber, a first damping passage, a second damping chamber and a second damping passage. Air flows through the damper to be ejected into the combustion chamber from the second damping passage at a selected velocity and volumetric flow. The flow being sufficient to damp instabilities from the combustion process.

Abstract: A combustor includes two or more sets of effusion cooling holes that extend through the inner and outer liners. Each set of effusion cooling holes includes one or more initial rows of effusion cooling holes, one or more final rows of effusion cooling holes disposed downstream of the one or more initial rows, and a plurality of interposed rows of effusion cooling holes disposed between the initial and final rows. Each effusion cooling hole is disposed at a tangential angle relative to an axial line. The tangential angle of the effusion cooling holes in each set of effusion cooling holes gradually transitions from a substantially transverse tangential angle in each initial row to a substantially axial tangential angle in each final row.