COMBUSTION CHAMBER ASSEMBLY OF A GAS TURBINE AND AVIATION GAS TURBINE

摘要

The present invention relates to a combustion chamber of a gas turbine comprisingan annular combustion chamber (15) with an inner ring wall (7) and an outer ring wall 8),a combustion chamber head (3) with a multiplicity of fuel nozzles (6),a first admixed air row (Z1) having a multiplicity of first admixing air holes (10) designed as passage openings, which are arranged in the inner ring wall (7) and / or the outer ring wall (8),a second admixing air row (Z2) having a multiplicity of second admixing air holes (11) designed as passage openings, which are arranged in the inner ring wall (7) and / or the outer ring wall (8),wherein the first admixing air holes (10) have first inner centers (10a) and first outer centers (10c) and the second admixing air holes (11) have second inner centers (11a) and second outer centers (11c), the first and second inner Center points (10a, 11a) respectively on a side facing the combustion chamber (15) side of the first and second admixing air holes (10, 11), and the first and second outer centers (10c, 11c) on a side facing away from the combustion chamber (15) side the first and second admixing air holes (10, 11) are located.where the equation L = D2 / D1 * (D2-D1) / C 2 is fulfilled,wherein L is a distance between the first and second inner centers (10a, 11a) and / or the first and second outer centers (10c, 11c) of the first and / or second admixing air holes (10, 11),wherein D1 is a first flow diameter of the first admixing air holes (10) at an inlet side and / or an outlet side to the combustion chamber (15) and D2 is a second flow diameter of the second admixing air holes (11) at the inlet side and / or the outlet side to the combustion chamber (15) is andwhere C is a mean flow coefficient of the first and second admixing holes (10, 11).