Abstract: A combustion chamber (18) for aircraft turbomachine, comprising an annular chamber end wall (40), an annular row of injection systems (42) mounted in the annular chamber end wall (40), each injection system comprising at least one air inlet swirler (56, 58) and a main fuel injection nozzle (54) configured to output a fuel stream centred on an injection axis (44) and comprising a central recirculation zone (62) and a corner recirculation zone (64) extending as an annulus around the central recirculation zone (62) chamber, and a plurality of additional fuel injection devices (70) mounted in the chamber end wall (40) and configured to inject fuel (71) directly into the corresponding corner recirculation zones (64) produced by the corresponding injection systems (42) at an operating speed less than or equal to the idling speed.