| 摘要 |
<p>This lnvention is a combustor for a gas turbine engine, particularly for aircraft propulsion, in which a combustion passage (54) is defined by a liner (44) within a housing (38) defining an inlet air plenum chamber (56) for air from a compressor. Normally fuel is injected through nozzles (24) in an annular end wall (46) of the liner and burnt using primary air received through inlets (88); additional air is provided through swirlers (84, 86) and by way of liner cooling passages (70, 72) both of which receive their air from a liner plenum (60) through an inlet (98). The inlet (98) can be progressively opened or closed by actuating means so that the quantity of air in the swirlers an din the cooling passages can be controlled together. For re-ignition at altitude and possibly also for initial starting at ground level, a valve sleeve (96) can close the inlet (98) so that ignition is only in the primary air at (88), and simultaneously an inlet (92) for compressed air at a downstream part of the combustion passage (54) is opened so that that air is not wasted, but is fed into the burning gases for exit at (52) for supply to a turbine.</p> |
