FIELD: engines and pumps.
SUBSTANCE: proposed aircraft turbojet combustion chamber comprises walls, bodies of revolution, arranged one inside the other and coupled via combustion chamber annular bottom wall. Inner and outer walls have forged primary air inlets and thinning air inlets that have edges extending inside combustion chamber. Combustion chamber comprises means of relaxation or reduction of mechanical strain in edges or nearby them for, at least, a part of said inlets. Said means comprise one, two or three slots per one inlet made in said edge or nearby it. Every said slot is connected, at least, by one its end, with orifice designed to stop crack propagation. Walls have microperforations to allow cooling airflow, inclined inward relative to normal to wall outer surface. Said slots and orifices are arranged in wall in parallel to adjacent microperforations to allow combustion chamber cooling via air circulation though said orifices.
EFFECT: decrease blade root temperature, longer life.
15 cl, 8 dwg
