FIELD: turbomachine combustion chambers.
SUBSTANCE: invention relates to a combustion chamber for a turbomachine such as an aircraft turbojet or an aircraft turboprop. The combustion chamber (1) of the turbomachine contains a bottom wall (4) containing at least one hole (5), at least one sleeve (12) installed upstream from the bottom wall (4) and attached to the bottom wall (4), closing the ring (13), limiting the annular groove (11) together with the sleeve (12) and attached to the sleeve (12), at least one air-fuel mixture injection system (6), having an axis (A) and installed in the hole (5) of the bottom wall (4), while the injection system (6) contains an annular collar (10) extending radially relative to the specified axis (A) and installed in the specified groove (11) with a radial clearance, the deflector (14), located downstream from the bottom wall (4), attached to the sleeve (12) and/or to the bottom wall (4) and containing a radially inner part located in the axial direction between the bottom wall (4) and the downstream end of the injection system (6). The injection system (6) contains at least one protruding part configured to enter the hollow part (16) of the deflector (14), or vice versa, in the first angular position of the installation of the injection system (6) relative to the deflector (14), wherein in the second angular position, the said protruding part is configured to come in the axial direction to the stop position against the radial surface or the downstream surface of the deflector (14) or, respectively, the injection system (4) to hold the injection system (6) relative to the deflector (14 ), while the specified protruding part is offset in the angular direction from the hollow part (16) in the specified second position.
EFFECT: invention improves operational reliability.
10 cl, 8 dwg
