Patent ID: 8479492

Claim:
A hybrid slinger combustor system for an aero gas turbine engine powering an aircraft, the combustor system comprising a combustor shell defining a combustion chamber, the combustion chamber having first and second combustion zones; two distinct fuel injector units for respectively spraying fuel into said first and second combustion zones, said two distinct fuel injector units including a rotary fuel slinger for spraying fuel radially outwardly into the first combustion zone through a first opening defined in said combustor shell, and a set of circumferentially spaced-apart fuel nozzles for spraying fuel into the second combustion zone through a set of second openings defined in said combustor shell, the rotary fuel slinger being positioned upstream of the fuel nozzles relative to a flow of combustion gases through said combustion chamber; and a control unit controlling the rate of fuel flow to said rotary fuel slinger and said set of fuel nozzles as a function of the power demand of the gas turbine engine, wherein the control unit is configured so that a major portion of the fuel to be atomized is supplied to said rotary fuel slinger during take-off and climb phases of a flight, whereas the major portion of the fuel is directed to the fuel nozzles at ground idle.