| 摘要 |
The present invention relates to an axial compressor for aircraft engines that comprises at least two stages arranged in series. The compressor is fitted with a system for deactivating a portion of the stages, i.e. one or more of the last and/or front stages. This system is used for instantly modifying the compression stage in the compressor without modifying the number of revolutions of the rotor (1) thereof. The stage deactivation system consists in connection couplers of driving devices or in air bleed-off devices (5). The dimensions, the number and the arrangement of the bleed-off windows (5) are determined according to the optimal deactivation conditions of the stages. The free circulation of air between the opened bleed-off windows (5) and the annular chamber (6) depends on the compliance to the equation C1u = C2u in the impeller (2). The annular chamber (6) together with the compressor outlet are connected to the inlet of the combustion chamber (9) when the last stages are deactivated. When the front stages are deactivated, the annular chamber (6) surrounding the same is connected to the outlet into the atmosphere or to the compressor inlet. The opening of the bleed-off windows (5) corresponds to the deactivation of the stages. This invention can be used in both subsonic and supersonic engines. |
