Combustion chamber comprising additional injection devices opening up directly into corner recirculation zones, turbomachine comprising such a chamber and fuel supply method for such a chamber

A combustion chamber for an aircraft turbomachine, comprising:

• an annular chamber end wall and two coaxial annular walls connected to each other by said annular chamber end wall and centred on a longitudinal axis of the combustion chamber;
• and
• an annular row of injection systems mounted in the annular chamber end wall, each injection system comprising at least one air inlet swirler and a main fuel injection nozzle configured to output a fuel stream centred on an injection axis and comprising a central recirculation zone and a corner recirculation zone extending as an annulus around the central recirculation zone;
• wherein:
• the combustion chamber also comprises a plurality of additional fuel injection devices mounted in the chamber end wall and configured to inject fuel directly into the corresponding corner recirculation zones produced by the injection systems;
• each additional injection device comprises at least one secondary fuel injection nozzle connected to a fuel supply source, and installed to pass through the chamber end wall at a spacing from the corresponding injection system;
• and
• said at least one secondary injection nozzle of each additional injection device is fitted with a valve configured to go into an open state when a fuel pressure corresponds to an operating speed of the turbomachine less than or equal to the idle speed, and to go into a closed state when a fuel pressure corresponds to an operating speed of the turbomachine greater than the idle speed.