Abstract: A multi-axial pivoting liner within the combustion system of a turbine engine allows the system to work with minimum thermal interference, especially during system operation at transient conditions, by allowing the liner to pivot and slide about its centerline and relative to the turbine scroll. The pivoting liner has the ability to control and minimize air leakage from part to part, for example, from the liner to the turbine scroll, during various operating conditions. Additionally, the liner provides for easy assembly with no flow path steps. Finally, the pivoting liner tolerates thermal and mechanical stresses and minimizes thermal wear.

Abstract: A gas turbine combustor in which a part or all of the wall of the combustor disposed within an intake chamber is formed as an acoustic energy absorbing member that can absorb the acoustic energy of a combustion variation generated within the combustor. The acoustic energy absorbing member is constructed of a thin corrugated plate in a circumferential direction, a high-temperature-proof perforated material, or a back plate disposed at the outside of a perforated plate in a radial direction with a distance from the perforated plate. It is also possible to provide a covering member at the outside of the acoustic energy absorbing member in a radial direction, for covering the acoustic energy absorbing member with a distance from the acoustic energy absorbing member. It is preferable that the acoustic energy-absorbing member and/or the covering member are reinforced with a frame that extends in a circumferential direction and/or a longitudinal direction.

Abstract: Aspects of the present invention relate to hollow elongated turbine engine components, such as transition ducts, having a cooling system. The component can have first and second ends, and can be defined by an outer peripheral wall. The cooling system extends longitudinally within the wall between the first and second ends. While configured to cool the entire component, the cooling system is particularly configured to reduce thermal gradients that can occur in regions proximate to the first and second ends of the component caused by higher loads in those regions. The cooling system can include at least one pair of cooling channels. Each channel pair includes a first channel and a second channel. The first and second channels can be arranged such that a coolant supplied to each channel can exchange heat with itself in at least the region substantially proximate to one of the ends.

Abstract: A gas-turbine engine combustor having a plurality of venturi mixers, each connected to an air supply path that passes air compressed by a compressor and to a supply source of gaseous fuel, which mix the air and the gaseous fuel to produce an air-fuel mixture and supply the air-fuel mixture to a combustion chamber for generating premixed combustion or diffusive combustion such that produced combustion gas is supplied to a turbine through a turbine nozzle to rotates the turbine that outputs its rotation through an output shaft, while driving the compressor by the rotation.

Abstract: The invention relates to a gas turbine (3) with reheat in the turbine section (11), whereby a fuel injection device (37) is provided for the injection of fuel (33) in to a combustion air flow (61) in the direction of flow before a cooling air flow (65) branching. By means of the early mixing of fuel (63) in the combustion air flow (61) and concomitant additional use of the cooling air flow (65) as combustion air, the specific output and the efficiency of the gas turbine (3) are improved.

Abstract: A gas turbine combustor includes a side wall, for defining a combustion volume, having upstream and downstream ends, a pilot nozzle, disposed adjacent the upstream end of the side wall, for discharging a pilot fuel to form a diffusion flame in the combustion volume, and a plurality of main nozzles, provided around the pilot nozzles, for discharging a fuel-air mixture to form premixed flames in the combustion volume. Film air is supplied into the combustion volume downstream of the main nozzles along the inner surface of the side wall to reduce the fuel-air ratio in a region adjacent the inner surface of the side wall and to restrain a combustion-driven oscillation in the combustion volume.

Abstract: In order to provide a method of air guidance and a design for a gas turbine combustion chamber which permits improved cooling and reduced pressure losses and which avoids the overcooling phenomena by specific reaction to hot spots, a gas turbine combustion chamber includes a casing and at least one hollow element. The at least one hollow element is in connection with air guidance regions of the combustion chamber via at least two openings. As such, a tangential through-flow exists with a subsequent axial flow to the burner.

Abstract: A gas turbine combustor, in particular for an aircraft engine, has at least one chamber, through which combustion gas flows in use, and which is defined by a lateral wall with channeling for feeding cooling air into the chamber and cooling the lateral wall; part of the channeling is defined by at least one double wall for guiding a relative stream of cooling air into the chamber in a tangential direction with respect to the lateral wall.

Abstract: A connector segment for connecting a combustor liner and a transition piece in a gas turbine has a substantially cylindrical shape and is of double-walled construction including inner and outer walls and a plurality of cooling channels extending axially along the segment, between the inner and outer walls. The cooling channels are defined in part by radially inner and outer surfaces, wherein at least one of the radially inner and outer surfaces is formed with an array of concavities.

Abstract: An internal cylinder which has the opening section from which combustion gas is blown out is inserted into the tail cylinder which becomes a flow path of the combustion gas, and is engaged with the tail cylinder, while securing a clearance between the edge of the opening section and the internal wall surface of the tail cylinder. Air compressed with a compressor is jetted from the clearance along the internal wall surface of the tail cylinder, and a film of cooling air is formed on the internal wall surface of the tail cylinder.

Abstract: The invention relates to a combustion-chamber arrangement for gas turbines, having an annular combustion chamber connected to at least one burner and leading into a turbine space, an annular-combustion-chamber wall which defines the annular combustion chamber being of double-shell construction, and an inner shell of the annular-combustion-chamber wall being composed of lining elements releasably arranged on an outer shell of the annular-combustion-chamber wall, and a gap space through which a cooling medium can flow being formed between the inner shell and the outer shell of the annular-combustion-chamber wall.

Abstract: A combustor wall louver for ducting a flow of compressed air through an inlet opening in the combustor wall from a source of compressed air outside the combustor where the louver is a circumferentially extending member, mounted to an interior surface of the combustor wall and covering the inlet opening with outlet openings fed by a channel in flow communication between each outlet opening and the inlet opening. Preferably, the circumferential member is made of arcuate segments of cast metal removably mounted to the interior surface of the combustor wall with threaded studs.

Abstract: A turbine includes a sealing element with a receiving area for sealing the guide blade vanes which are adjacent to each other in the peripheral direction of the turbine. The foot plates of the guide blade vanes extend into the receiving area. The edge area of the foot plates does not have to be reinforced compared to a conventional seal, which enables the entire foot plate to be cooled homogeneously. A closed cooling system can therefore be used for cooling, especially with steam.

Abstract: A turbine installation, especially a gas turbine installation, includes foot plates of the guide blades of adjacent turbine stages being interconnected with a clip-type sealing element on their rear sides facing away from the gas area. This provides a simple seal between adjacent foot plates which is effective regardless of the thermal expansion of the foot plates. The clip-type sealing element is also suitable for sealing the tiles of a combustor of the turbine installation together.

Abstract: A combustor liner for use in a gas turbine engine includes an annular panel section having a cooling nugget formed on one end thereof. An annular cooling lip is formed on the cooling nugget so as to define a cooling slot. The cooling lip includes a hot side, a cold side and a trailing edge surface. The hot side defines a compound taper for minimizing the thickness of the trailing edge surface. In one embodiment, the compound taper includes a first tapered surface disposed on a forward portion of the cooling lip and a second tapered surface disposed on an aft portion of the cooling lip, wherein the second tapered surface has a greater taper than the first tapered surface.

Abstract: A combustor liner has an annular shell which includes a first portion and a second portion. The first portion is provided with slot film cooling and the second portion is provided with multi-hole film cooling. The multi-hole cooling portion can be located either forward or aft of the slot film cooling portion, depending on the nature of the combustor that the liner is to be used in. In one possible embodiment, the liner includes a first annular panel, a second annular panel section joined at its forward end to the aft end of the first panel section, and a third annular panel section being joined at its forward end to the aft end of the second panel section. At least one of the panel sections has multi-hole film cooling and at least one other of the panel sections has slot film cooling.

Abstract: A combustor liner is provided having first and second annular bands which define an overlapping circumferential joint area, wherein a weld is disposed in the joint area encompassing substantially all of the axial length of the joint area. A method for producing such a combustor liner is also provided.

Abstract: An effusion cooled transition duct for transferring hot gases from a combustor to a turbine is disclosed. The transition duct includes a panel assembly with a generally cylindrical inlet end and a generally rectangular exit end with an increased first and second radius of curvature, a generally cylindrical inlet flange, and a generally rectangular end frame. Cooling of the transition duct is accomplished by a plurality of holes angled towards the end frame of the transition duct and drilled at an acute angle relative to the outer wall of the transition duct. The combination of the increase in radii of curvature of the panel assembly with the effusion cooling holes reduces component stresses and increases component life. An alternate embodiment of the present invention is shown which discloses shaped angled holes for improving the film cooling effectiveness of effusion holes on a transition duct while reducing film blow off.

Abstract: An effusion cooled transition duct for transferring hot gases from a combustor to a turbine is disclosed. The transition duct includes a panel assembly with a generally cylindrical inlet end and a generally rectangular exit end with an increased first and second radius of curvature, a generally cylindrical inlet sleeve, and a generally rectangular end frame. Cooling of the transition duct is accomplished by a plurality of holes angled towards the end frame of the transition duct and drilled at an acute angle relative to the outer wall of the transition duct. Effusion cooling geometry, including coverage area, hole size, and surface angle will be optimized in the transition duct to tailor the temperature levels and gradients in order to minimize thermally induced stresses. The combination of the increase in radii of curvature of the panel assembly with the effusion cooling holes reduces component stresses and increases component life.

Abstract: A cooling structure of a combustor tail tube is capable of avoiding the formation of cracks in the tail tube by lessening thermal stress and preventing thermal deformation, thus extending the service life. A multiplicity of cooling jackets (5) extend in the longitudinal direction of the tail tube of a gas turbine combustor along the entire circumference of the tail tube wall. The passage sectional area of the cooling jackets (5) is varied depending on the metal temperature of parts of the tail tube (4). For example, the passage sectional area of the cooling jackets (5) formed at the rotor side wall and the mutually opposite side walls of the adjacent tail tube is larger than the passage sectional area of the cooling jackets (5) formed at the casing side wall.

Abstract: In a twin wall combustion chamber for a gas turbine engine, the outer wall has impingement holes so that compressed air surrounding the chamber can pass through the holes to impinge on the inner wall, and the inner wall has effusion holes whereby air can effuse into the combustion chamber. The number of effusion holes is greater than the number of impingement holes, the effusion-holes preferably being arranged in groups of seven, comprising six holes equi-spaced around a central seventh hole, each group having an impingement hole in a fixed positional relationship to the central hole, preferably downstream of it.

Abstract: A combustor liner for use in a gas turbine engine includes a first annular panel section having a forward end, an aft end and a first cooling nugget located at the forward end thereof, and a second annular panel section having a forward end, an aft end and a second cooling nugget located at the forward end thereof. The second panel section is joined at its forward end to the aft end of the first panel section. A first row of cooling holes is located in the first cooling nugget, and a second row of cooling holes is located in the second cooling nugget. A group of dilution holes is located in the first panel section. The dilution holes are located at the aft end of the first panel section, immediately upstream of the second cooling nugget. Furthermore, each one of the dilution holes defines an aftmost edge, and all of the aftmost edges are axially aligned, even if the dilution holes have different hole diameters.

Abstract: The present invention is directed to a combustor/turbine successive dual cooling arrangement in which the combustor has a one-piece hot combustor wall and front and rear cold combustor walls, and cooling air is forced under pressure through perforations in the cold combustor walls and impinges on the annular front and rear sections of the hot combustor wall for the backside cooling of the hot combustor wall. The exhaust combustor backside cooling air is directed to gain access to the hot end of the engine, that is, the turbine section, to cool the turbine components. The combustor/turbine successive dual cooling arrangement according to the present invention enables all the air typically used to cool the hot end of the engine downstream of the combustor, to be used as combustor backside cooling as well, to significantly reduce the amount of air needed for combustor and turbine cooling.

Abstract: A transition piece assembly for a gas turbine includes a transition duct having one end adapted for connection to a gas turbine combustor and an opposite end adapted for connection to a first turbine stage, and a pair of side panels. The assembly also includes an impingement sleeve surrounding the transition duct and establishing a cooling path therebetween. The impingement sleeve is formed with a plurality of rows of cooling holes therein; and a plurality of flow catcher devices on an external surface of the impingement sleeve, each flow catcher device at least partially surrounding one of the cooling holes.

Abstract: A wall structure for a gas turbine combustor arranged to have a general direction of fluid flow therethrough includes inner and outer walls defining a space therebetween. The inner wall is made up of a plurality of tiles having axial edges aligned generally with the direction of fluid flow, a gap being defined between axial edges of adjacent tiles. Orifices are provided within the axial edges to direct leakage air passing through the gaps to give the leakage air a flow component in the general direction of fluid flow through the combustor.

Abstract: A method of fabricating an annular liner for a combustion chamber of a gas turbine engine. The combustion chamber has a dome including a fuel nozzle for delivering fuel to the combustion chamber. A metal piece is formed into an annular section constituting at least a portion of the liner and having at least one seam extending generally axially with respect to an axial centerline of the liner. An upstream end of the liner is machined to have a registration feature for connecting the liner to the combustion chamber dome in at least one predetermined orientation with respect to the dome. Further, the method includes ensuring the registration feature and the seam are located relative to each other on the liner so the seam is at a preselected position with respect to the combustion chamber dome when the liner is connected to the combustion chamber dome in the predetermined orientation.

Abstract: An improved dual-stage, dual-mode turbine combustor capable of reducing nitric oxide (NOx) emissions is disclosed. This can-annular combustor utilizes multiple, single wall sheet metal combustor liners, generally annular in shape, and each liner having multiple hole film cooling means, which includes at least one pattern of small, closely spaced film cooling holes angled sharply in the downstream direction and various circumferential angles for improved liner cooling and improved fuel/air mixing within the liner, resulting in lower NOx emissions.

Abstract: A wall element (29A, 29B) for a combustor (20) of a gas turbine engine (10). The wall element (29A, 29B) defines an axis. In use, the axis is arranged generally parallel to the principal axis of the engine (10). In one aspect, the length of the wall element (29B) along the axis is at least substantially 20% of the length of the wall element (29B) transverse to the axis. In another aspect, the wall element (29A, 29B) has a first pair of opposite edges extending transverse to the axis and a second pair of opposite edges (48, 50) extending transverse to the first pair, at least one of the second pair of edges (48, 50) being angled relative to the axis of the wall element (29A, 29B).

Abstract: A combustor of a gas turbine includes a combustion tube acting as a combustion chamber. In the combustion tube, a cylindrical heat insulating member is disposed and forms a wall of the combustion chamber. Numerous cooling air passages are formed in the heat insulating member and extending in the axial direction between cooling air inlet disposed at one end of the heat insulating member and cooling air outlet disposed at the other end of the heat insulating member. The heat insulating member is attached to the inner surface of the combustion tube at its cooling air inlet side end and a seal ring is disposed around the outer circumference of the heat insulating member at a portion between the cooling air inlet side end and the cooling air outlet side end. The seal ring is made of heat resistant material and prevents cooling air flowing through the annular space between the inner surface of the combustion tube and the outer surface of the heat insulating member.

Abstract: A liner for a gas turbine engine combustor, including an upstream section, a downstream section oriented at an angle to the upstream liner section, and a cooling nugget joining the upstream and downstream liner sections for providing mechanical stiffness at a junction of the upstream and downstream liner sections. The cooling nugget further includes a first portion connected to the upstream liner section, a second portion connected to the downstream liner section, and a third portion joining the first and second cooling nugget portions at a first end, the third cooling nugget portion extending radially from the first end, wherein the cooling nugget is in flow communication with a cool air supply and is configured to provide a starter film of cooling air along respective surfaces of the upstream liner section and the downstream liner section.

Abstract: Cooling structure of gas turbine combustor in which cooling medium flows through grooves in wall is improved so that adjustment of flow velocity, pressure loss and heat transfer rate of cooling medium flow in the wall becomes possible and cooling effect thereof is enhanced. Wall of combustor tail tube is made in double structure in which outer plate (1) and inner plate (4) are jointed together being lapped one on another. The outer plate (1) has air inlet hole (3) and groove (2) formed therein. The groove (2) is closed by jointing of the inner plate (4) to the outer plate (1). The inner plate (4) has air outlet hole (5) formed therein. The groove (2) communicates with the air inlet hole (3) and the air outlet hole (5). Cross sectional shape of the groove (2) is changed two-dimensionally or three-dimensionally such that width enlarges toward the hole (5) from the hole (3) or depth is constant or changed in tapered form.

Abstract: In a twin wall combustion chamber for a gas turbine engine, the outer wall has impingement holes so that compressed air surrounding the chamber can pass through the holes to impinge on the inner wall, and the inner wall has effusion holes whereby air can effuse into the combustion chamber. The number of effusion holes is greater than the number of impingement holes, the effusion holes preferably being arranged in groups of seven, comprising six holes equi-spaced around a central seventh hole, each group having an impingement hole in a fixed positional relationship to the central hole, preferably downstream of it.

Abstract: Gas turbine combustor tail tube cooling structure is improved so that temperature distribution along tail tube axial direction is made gentle and large thermal stress is prevented from occurring resulting in enhancement of safety and reliability of the combustor tail tube. Supply jacket (16) is disposed between upstream side discharge jacket (14) and downstream side discharge jacket (15). The upstream side discharge jacket (14) is disposed apart from tail tube inlet (2) at position close to by-pass passage (10) so as to form non-cooled zone (A) in the vicinity of the tail tube inlet (2). On the downstream side thereof, the supply jacket (16) is disposed at position in tail tube (1) contracting portion and near the downstream side discharge jacket (15) so as to shorten length of cooling pipes there.

Abstract: This invention concerns the use of pressurized steam as the cooling medium for a gas turbine combustor. It is distinguished by the following. Steam supply manifolds or ports for the cooling steam are provided on the gas inlet and outlet sides of the combustion chamber in the gas turbine combustor. A steam exhaust manifold or port for the cooling steam is provided between the gas inlet and outlet sides, in approximately the center of the chamber. Cooling channels for the steam are created in the external wall panel of the chamber between the steam supply manifolds and the exhaust manifold. The steam supplied into the wall panel through the steam supply manifolds on the gas inlet and outlet sides of the chamber is exhausted to the exterior via the steam exhaust manifold in the center of the chamber. This design allows pressurized steam with a high thermal capacity to be used to effectively cool the wall panels of the combustor, which are exposed to extremely hot combustion gases.

Abstract: This invention concerns the use of pressurized steam as the cooling medium in the combustor wall of a gas turbine combustor. It is distinguished by the following. The combustor wall is configured by 1) a plurality of cooling channels for cooling steam, sealed by an exterior wall panel and a heat-resistant plate which are assembled by soldering or some other method; 2) a supply manifold for supplying the cooling steam into the cooling channels, which is provided on one end of the cooling channels; and 3) a recovery manifold for recovering the cooling steam from the cooling channels, which is provided on another end of the cooling channels. This arrangement can form strong enough steam-channels that do not allow any leakage of the high pressure steam from the cooling system.

Abstract: The invention relates to the control of air flows for cooling the axial walls of a high-temperature combustion chamber in which the walls are provided with through holes forming multiple perforations, and involves arranging the multiple perforations so that the cooling airflow permeability of the walls in the downstream zone of the chamber decreases in the downstream direction in order to compensate for the effects of the increase in pressure drop due to variation of the gas flow velocity. The invention is particularly applicable to sharply convergent combustion chambers such as twin-head combustion chambers.

Abstract: A low emissions combustor can 18 for an aircraft gas turbine engine includes a louvered combustor liner 24 with three arrays of dilution holes 52, 54, 56. The first hole array comprises twelve equally sized, equiangularly distributed holes that penetrates the liner about midway along its axial length. The second hole array comprises twelve equally sized holes, smaller than the holes of the first array, that penetrate the liner a predetermined distance aft of the first hole array. The holes of the second array are equiangularly distributed and each second hole is circumferentially aligned with a first hole. A third hole array penetrates the liner a predefined distance aft of the first array. The third holes are nonuniformly sized and nonequiangularly distributed to regulate the spatial temperature profile of combustion gases exiting the combustor can. The quantity, size, distribution and location of the holes mitigates undesirable exhaust emissions without affecting the performance or durability of the engine.

Abstract: A heat shield configuration, particularly for protecting supporting structures and structural parts of gas turbine plants against a hot fluid. The heat shield has an inner lining which consists of a heat-resistant material and which is composed of plate-shaped heat-shield elements resistant to high temperatures and are disposed next to one another with gaps in-between. The heat-shield elements cover the entire area of the structural part. The heat shield is anchored so as to be thermally movable to the supporting structures and parts by bolts. The heat-shield element is formed of an erosion-proof and corrosion-proof material. An insulating brick formed of a refractory ceramic is disposed in each case between each heat-shield element and the supporting structure or parts.

Abstract: A one piece cold combustor wall for lining the reverse flow hot combustor wall of a gas turbine engine, disposed at a distance from the outer surface of the hot combustor wall. Improved cooling of the hot combustor wall results from the addition of impingement cooling air injected through orifices in the cold combustion wall directed at the hot combustion wall together with film cooling by air conducted between the hot and cold combustor walls. The cold combustor wall is perforated with a pattern of air impingement inlet orifices through the cold combustor wall conducting compressed air from the outer surface of the cold combustor wall in compressed air jets directed at the outer surface of the hot combustor wall. The provision of a cold combustor wall also improves conventional air film cooling by adding impingement cooling and reusing the air after impingement to form a contained air film between the hot and cold walls.

Abstract: A combustion chamber is disclosed for a turbojet engine in which the combustion chamber has a forward intake, a rear exit and a perforated casing extending between the intake and exit. The combustion chamber also has at least two rows of insulating tiles arrayed on the casing in forward and rear rows with both rows extending circumferentially around the inner surface of the casing bounding the combustion chamber. Each rear tile has a forward edge portion with a sloping wall and each forward tile has a tapered rear edge portion overlapping the forward edge portion with a clearance therebetween such that the tampered rear edge portion and the sloping wall bound a slot opening into the combustion chamber and slanting towards the rear of the combustion chamber to facilitate the formation of an air cooling film on inner surfaces of the tiles.

Abstract: A modular cooling panel for cooling a turbine member is provided. The cooling panel comprises a first panel having a relative width, relative length, upper surface and lower surface. The upper surface defines at least one corrugated portion traversing along a portion of the relative width of the upper surface. The corrugated portion defines a cooling flow channel through which a cooling fluid can travel to cool the turbine member. The cooling flow channel has at least one inlet opening for enabling the cooling fluid to enter into the cooling flow channel. The lower portion surface of the first panel is adapted to be coupled in fluid communication with the turbine member.

Abstract: Provided are a gas turbine multi-hole film-cooled combustor liner, which can be manufactured with improved accuracy in shape and position of cooling holes in a very short time, ensuring desirable buckling strength and satisfactory cooling performance. A planar flat member is curved to form a liner shell of a cylindrical shape, the cylindrical member is welded in the longitudinal direction thereof to form a cylindrical shell. Then a wavy configuration and a corrugated configuration are formed on the cylindrical shell by a hydro-bulging method, cooling holes are formed through the liner shell by laser drilling at or near wave crest portions of the wavy configuration of the liner shell, and inner rings are attached to the liner shell by resistance spot welding or vacuum brazing.

Abstract: A combustion chamber, particularly for a turbomachine, has at least one generally axially extending wall provided with a plurality of through holes defining a multi-hole cooling system for the wall. The wall also has a plurality of dilution holes which affect locally the flow of burnt gases in the chamber, and the cooling through holes are oriented according to the local flow of burnt gases in the chamber. In particular the wall is subdivided into first zones, downstream of the dilution holes, in which the through holes are oriented substantially in countercurrent to the overall flow direction of the burnt gases in the chamber, second zones and third zones which are disposed on opposite sides of the first zones and in which the orientation of the through holes is inclined both axially and circumferentially, and a fourth zone covering the remainder of the wall and in which the through holes are inclined axially in the overall flow direction of the burnt gases.

Abstract: The exhaust in a gas turbine engine is lined with stacks of hollow cooling blocks made of a composite material. Air is pumped into each stack and escapes into the exhaust flow from a slot between the blocks in each stack. In addition, the blocks contain passages so that some of the airflow escapes to a space between the stacks. The blocks are attached using special fasteners that does not create mechanical stresses on the composite material.

Abstract: A combustor liner includes an inner layer for facing combustion gases, and an opposite outer layer for facing a cooling fluid. The outer layer has a greater coefficient of thermal conductivity than the inner layer for reducing temperature gradients in the liner. In a preferred embodiment, the outer layer significantly reduces temperature gradients in the liner which are caused by the varying cooling ability of impingement cooling air jets for more uniformly cooling the combustor liner and reducing the maximum temperature thereof.

Abstract: A compact gas turbine engine having improved fuel efficiency, lower emissions, better starting capability, and longer operational life is provided at low cost through the use of a combustor incorporating tangential fuel injection in conjunction with certain strategically placed and sized oxidant inlet jets. The oxidant inlet jets include deflector jets which deflect a generally circumferentially spiralling flow of gases within the combustor in a manner causing fuel and oxidant to recirculate for a longer period of time within a primary combustion zone, to thereby improve combustion efficiency and reduce undesirable emissions. Apparatus and methods for regulating injection of the fuel and oxidant in or primary, secondary, and dilution zones of the combustor in a manner providing improvements in fuel efficiency, high altitude starting and other advantages are also defined.

Abstract: A wall is disclosed for a gas turbine engine combustion chamber having a generally annular configuration around a central axis wherein the wall has an inner surface bounding at least a portion of the combustion chamber and a plurality of rows of cooling openings, each cooling opening having an elongated, generally rectangular configuration with a length l and a width e such that l is greater than e, the cooling openings in each row being equidistantly spaced apart a distance d in which d is not greater than the length l. The cooling openings of a given row are circumferentially displaced from cooling openings of an adjacent row and, since the distance d is not greater than the length l of the cooling openings, the opening areas overlap in adjacent rows to provide an even distribution of the cooling film within the combustion chamber.

Abstract: A wall adapted for use in a gas turbine engine between a first and a hotter second fluid includes a first side over which is flowable the first fluid, and an opposite second side over which is flowable the second fluid. An elongate slot extends inwardly from the second side and is disposed in flow communication with a plurality of longitudinally spaced apart holes extending inwardly from the first side. The holes are disposed at a compound angle relative to the second side for discharging the first fluid obliquely into the slot and at a shallow discharge angle from the slot along the second side. The holes are also disposed in converging pairs to impinge together in the slot the first fluid channeled therethrough. In a preferred embodiment, the slot has an aft surface including a plurality of longitudinally spaced apart grooves extending from the holes to the wall second side.

Abstract: A jet engine component, such as an aircraft gas turbine engine rotor blade or a scramjet engine fuel injector. The component has a wall portion including a first surface exposable to a cooler, higher static pressure fluid and a second surface exposable to a hotter, lower static pressure gas flow flowing across the second surface. The component further includes a generally straight film coolant passageway having an inlet on the first surface and an outlet on the second surface. The second surface has a seamless groove which is open substantially entirely along its longitudinal dimension extending from the outlet along the lower static pressure gas flow for improved film cooling of the second surface.

Abstract: A wall adapted for use in a gas turbine engine between a first and a hotter second fluid includes a first side over which is flowable the first fluid, and an opposite second side over which is flowable the second fluid. An elongate slot extends inwardly from the second side and is disposed in flow communication with a plurality of longitudinally spaced apart holes extending inwardly from the first side. The holes are disposed at a compound angle relative to the second side for discharging the first fluid obliquely into the slot and at a shallow discharge angle from the slot along the second side. In a preferred embodiment, the slot has an aft surface including a plurality of longitudinally spaced apart grooves extending from the holes to the wall second side.