| 摘要 |
The invention concerns an axially stepped annular combustion chamber, in particular for an aircraft gas turbine, the combustion chamber essentially comprising an independent main combustion chamber (5') and an independent pilot combustion chamber (5). Appropriate design of the inner boundary walls (6a, 6b) of the pilot combustion chamber (5) ensures that the combustion gases thereof enter the main combustion zone (5') substantially in the radial direction. As a result, optimum mixing of the fuel and air is thus ensured in this main combustion zone or main combustion chamber (5'), exhaust gas emissions are minimized and the temperature distribution at the combustion chamber outlet (8) is optimum. The inner boundary wall (6a) can comprise a deflection section (12) or the outer wall section (6b) can be inclined towards the pilot burner longitudinal axis (3a), such that the cross-section of the pilot burner zone (5) decreases in the direction of flow. |
