Abstract: A combustion chamber for a turbine engine, in particular for an aircraft turbojet engine or turboprop engine. The combustion chamber includes a radially outer annular shroud, a radially inner annular shroud coaxial with the radially outer shroud, and an end wall connecting the radially outer shroud and the radially inner shroud. The combustion chamber further includes a first annular sealing member coaxial with said radially inner and outer shrouds. The first annular sealing member is radially interposed between the end wall and the radially outer shroud.

Abstract: The provision of air passage holes through a wall of a gas turbomachine combustion chamber. Multi-perforations are virtually positioned and distributed, even in a first safety zone without air passage openings. Multi-perforations with a virtual inlet or outlet in this first security zone are virtually removed. According to certain criteria, at least some of said removed multi-perforations are then virtually reintegrated, and, from then on a perimeter passing through the virtual inlets and outlets of all the multi-perforations present is defined, in the direction of a primary or dilution hole to be installed, a modified safety zone is defined, then, respecting around said hole and with the freedom to reposition it within this limit, the shape of this hole is redefined.

Abstract: An annular combustion chamber of a turbomachine is described. The combustion chamber has an axis of revolution and is delimited by coaxial internal and external annular walls joined upstream by a bottom of chamber substantially transverse to the walls. In some embodiments, the chamber includes at least one annular deflector placed in the chamber and substantially parallel to the bottom of chamber. The bottom of chamber may have openings adapted to be traversed by air for cooling the deflector. In some embodiments, the bottom of chamber and the deflector include mounting openings for mounting an annular row of injection devices for injecting a mixture of air and fuel into the chamber. At least a portion of the air for cooling the deflector is conveyed into the chamber through holes in the injection devices.

Abstract: The provision of air passage holes through a wall of a gas turbomachine combustion chamber. Multi-perforations are virtually positioned and distributed, even in a first safety zone without air passage openings. Multi-perforations with a virtual inlet or outlet in this first security zone are virtually removed. According to certain criteria, at least some of said removed multi-perforations are then virtually reintegrated, and, from then on a perimeter passing through the virtual inlets and outlets of all the multi-perforations present is defined, in the direction of a primary or dilution hole to be installed, a modified safety zone is defined, then, respecting around said hole and with the freedom to reposition it within this limit, the shape of this hole is redefined.

Abstract: An annular turbine engine combustion chamber wall including air admission orifices to create zones of steep temperature gradient, and cooling orifices to enable the air flowing on the cold side to penetrate to the hot side in order to form a film of cooling air along the annular wall, the annular wall being further includes, in the zones of steep temperature gradient, multi-perforation holes having respective bends of an angle ? greater than 90°, the angle ? being measured between an inlet axis Ae and an outlet axis As of the multi-perforation hole, the outlet axis of the multi-perforation hole being inclined at an angle ?3 relative to the normal N to the annular wall through which the multi-perforation holes with bends are formed, in a “gyration” direction that is at most perpendicular to the axial flow direction D of the combustion gas.

Abstract: An annular combustion chamber with an axis of revolution of a turbine engine delimited by coaxial inner and outer annular walls joined upstream by a substantially transverse bottom of the chamber, the chamber further comprising at least one annular deflector placed in the chamber and substantially parallel to the bottom of the chamber the bottom of the chamber having orifices designed to be passed through by the impact cooling air of the deflector and coming from upstream. The deflector is attached to the inner and outer walls in a sealed manner, and the cooling air of the deflector is discharged from the chamber through exhaust holes formed in the inner and outer walls.

Abstract: An annular combustion chamber of a turbomachine is described. The combustion chamber has an axis of revolution and is delimited by coaxial internal and external annular walls joined upstream by a bottom of chamber substantially transverse to the walls. In some embodiments, the chamber includes at least one annular deflector placed in the chamber and substantially parallel to the bottom of chamber. The bottom of chamber may have openings adapted to be traversed by air for cooling the deflector. In some embodiments, the bottom of chamber and the deflector include mounting openings for mounting an annular row of injection devices for injecting a mixture of air and fuel into the chamber. At least a portion of the air for cooling the deflector is conveyed into the chamber through holes in the injection devices.

Abstract: A combustion chamber for a turbine engine, in particular for an aircraft turbojet engine or turboprop engine. The combustion chamber includes a radially outer annular shroud, a radially inner annular shroud coaxial with the radially outer shroud, and an end wall connecting the radially outer shroud and the radially inner shroud. The combustion chamber further includes a first annular sealing member coaxial with said radially inner and outer shrouds. The first annular sealing member is radially interposed between the end wall and the radially outer shroud.

Abstract: An annular combustion chamber with an axis of revolution of a turbine engine delimited by coaxial inner and outer annular walls joined upstream by a substantially transverse bottom of the chamber, the chamber further comprising at least one annular deflector placed in the chamber and substantially parallel to the bottom of the chamber the bottom of the chamber having orifices designed to be passed through by the impact cooling air of the deflector and coming from upstream. The deflector is attached to the inner and outer walls in a sealed manner, and the cooling air of the deflector is discharged from the chamber through exhaust holes formed in the inner and outer walls.

Abstract: An annular turbine engine combustion chamber wall including air admission orifices to create zones of steep temperature gradient, and cooling orifices to enable the air flowing on the cold side to penetrate to the hot side in order to form a film of cooling air along the annular wall, the annular wall being further includes, in the zones of steep temperature gradient, multi-perforation holes having respective bends of an angle ? greater than 90°, the angle ? being measured between an inlet axis Ae and an outlet axis As of the multi-perforation hole, the outlet axis of the multi-perforation hole being inclined at an angle ?3 relative to the normal N to the annular wall through which the multi-perforation holes with bends are formed, in a “gyration” direction that is at most perpendicular to the axial flow direction D of the combustion gas.

Abstract: A combustion chamber for an aviation turbine engine, the combustion chamber being annular about a longitudinal axis and including an outer side wall, an inner side wall, and an annular chamber end wall connecting one end of the outer side wall to one end of the inner side wall. The outer side wall includes, distributed along its circumference, spark plugs, primary holes, and dilution holes situated downstream from the primary holes in the direction of the longitudinal axis. The primary holes situated in each of the adjacent zones adjacent to one of the spark plugs present a configuration that is different from the configuration of the primary holes situated outside the zones, such that the supply of air in the adjacent zones is different from the supply of air outside the zones.

Abstract: A diffuser for a turbine engine that includes two annular webs extending inside one another and connected between them by substantially radial vanes, wherein the downstream peripheral edge of at least one of the webs includes indentions evenly distributed about the longitudinal axis of the diffuser.

Abstract: A baffle for the bottom of a combustion chamber of a gas turbine engine is disclosed. The baffle includes a flat portion of wall with an opening for an injector of the combustion chamber to pass through, two longitudinal edges for the assembly to two adjacent baffles and two transverse edges. At least one of the edges includes a joint cover arranging a housing along the edge for an edge of an adjacent baffle so as to seal the junction between the two edges. A combustion chamber incorporating the baffles is also disclosed.

Abstract: A baffle for the bottom of a combustion chamber of a gas turbine engine is disclosed. The baffle includes a flat portion of wall with an opening for an injector of the combustion chamber to pass through, two longitudinal edges for the assembly to two adjacent baffles and two transverse edges. At least one of the edges includes a joint cover arranging a housing along the edge for an edge of an adjacent baffle so as to seal the junction between the two edges. A combustion chamber incorporating the baffles is also disclosed.

Abstract: An annular combustion chamber for a gas turbine engine includes radially inner and outer walls connected together by a chamber end wall including openings, each of which receives a fuel injection system. Heat protection deflectors are fastened to the chamber end wall. Holes are formed through the chamber end wall to pass cooling air onto an upstream face of each deflector. The inner or outer edge of a deflector presents a sealing rim engaging the respective inner or outer wall of the chamber.

Abstract: An annular combustion chamber for a turbomachine is disclosed. The chamber includes an inner wall, an outer wall, and a chamber end wall disposed between the inner and outer walls in the upstream region of the chamber. The chamber end wall presents an outer fastener rim and/or an inner fastener rim, and the outer and/or inner wall presents a respective upstream fastener rim. The chamber end wall and the outer and/or inner wall are fastened together via their fastener rims. Cooling channels are made between the fastener rims and open out to the inside of the combustion chamber. Advantageously, a spacer is placed between the fastener rims, and the channels are formed in the thickness or in the sides of the spacer.

Abstract: A rotationally symmetrical wall for an aircraft turbomachine combustion chamber, including primary holes, dilution holes positioned downstream from the primary holes, and a plug installation aperture upstream from the primary holes, wherein each first primary hole, considered as one moves away circumferentially from the median axial plane (D) in either direction is positioned at a circumferential position between the circumferential positions of the first dilution hole and of the second dilution hole considered as one moves away circumferentially from the median axial plane (D), wherein the other primary holes are distributed equidistantly, at a predefined interval P, and the distance between the two primary holes is greater than the interval P separating two of the other adjacent primary holes.

Abstract: A combustion chamber for an aviation turbine engine, the combustion chamber being annular about a longitudinal axis and including an outer side wall, an inner side wall, and an annular chamber end wall connecting one end of the outer side wall to one end of the inner side wall. The outer side wall includes, distributed along its circumference, spark plugs, primary holes, and dilution holes situated downstream from the primary holes in the direction of the longitudinal axis. The primary holes situated in each of the adjacent zones adjacent to one of the spark plugs present a configuration that is different from the configuration of the primary holes situated outside the zones, such that the supply of air in the adjacent zones is different from the supply of air outside the zones.

Abstract: A turbomachine combustion chamber including coaxial walls forming surfaces of revolution that include inlet orifices for admitting primary air and dilution air into the chamber, the orifices in each wall being substantially in alignment with one another along the longitudinal axis of the chamber and forming a single annular row of orifices.

Abstract: A turbomachine comprising an annular combustion chamber (40) with injection orifices (44); a centrifugal compressor (10) having a centrifugal downstream stage and an annular diffuser (20) comprising: a radially-oriented upstream portion (21)with diffusion passages (22) connected to the outlet of the compressor and an elbow-shaped intermediate portion (24)and a downstream portion (25) comprising a series of flow-straightening vanes (26) and inclined relative to the axis (A) of the turbomachine. In the section plane containing the axis (A) of the turbomachine and passing via the center (C) of one of said injection orifices (44), the curvilinear abscissa distance along a flow line (L) between the middle (0) of the flow path at the trailing edges of the flow-straightening vanes (26) and said center (C) is greater than or equal to three times the height (h) of said flow path.