There are described a method and system for operating a gas turbine engine. The method comprises governing the engine with a primary governing logic to adjust a fuel flow to the engine and governing the engine with at least one complementary governing logic concurrently to the primary governing logic, the at least one complementary governing logic adjusting a position of at least one variable geometry mechanism (VGM) of the engine to prevent a rotational speed of the engine from operating in one or more restricted speed range.

There are described a method and system for operating a gas turbine engine. The method comprises governing the engine with a primary governing logic to adjust a fuel flow to the engine and governing the engine with at least one complementary governing logic concurrently to the primary governing logic, the at least one complementary governing logic adjusting a position of at least one variable geometry mechanism (VGM) of the engine to prevent a rotational speed of the engine from operating in one or more restricted speed range.

A method of operating a multi-engine aircraft includes operating a first engine of the aircraft to provide motive power; and operating a second engine of the aircraft in a standby mode to provide substantially no motive power, a rotor of the second engine having an idle rotational speed in the standby mode, and in the standby mode, sequentially executing cycles, a given cycle of the cycles including: opening a set of variable guide vanes upstream a compressor section of the second engine, spiking a rate of a fuel flow to a combustor of the second engine, the spiking and the opening timed to increase a rotational speed of the rotor of the second engine to an upper threshold above the idle rotational speed of the rotor, and in response to the rotational speed reaching the upper threshold, at least substantially closing the set of variable guide vanes.

A method of operating a multi-engine aircraft includes operating a first engine of the aircraft to provide motive power; and operating a second engine of the aircraft in a standby mode to provide substantially no motive power, a rotor of the second engine having an idle rotational speed in the standby mode, and in the standby mode, sequentially executing cycles, a given cycle of the cycles including: opening a set of variable guide vanes upstream a compressor section of the second engine, spiking a rate of a fuel flow to a combustor of the second engine, the spiking and the opening timed to increase a rotational speed of the rotor of the second engine to an upper threshold above the idle rotational speed of the rotor, and in response to the rotational speed reaching the upper threshold, at least substantially closing the set of variable guide vanes.

A gas turbine engine that includes a stage of guide vanes, a compressor downstream from the stage of guide vanes, and a combustor downstream from the compressor. The combustor includes a primary combustion zone and a secondary combustion zone downstream from the primary combustion zone. The primary combustion zone includes an exit configured to channel combustion gases towards the secondary combustion zone. A controller is communicatively coupled with the stage of guide vanes, the controller configured to monitor a temperature at the exit of the primary combustion zone, and selectively open and close the guide vanes to maintain the temperature within a predefined temperature range.

A gas turbine engine that includes a stage of guide vanes, a compressor downstream from the stage of guide vanes, and a combustor downstream from the compressor. The combustor includes a primary combustion zone and a secondary combustion zone downstream from the primary combustion zone. The primary combustion zone includes an exit configured to channel combustion gases towards the secondary combustion zone. A controller is communicatively coupled with the stage of guide vanes, the controller configured to monitor a temperature at the exit of the primary combustion zone, and selectively open and close the guide vanes to maintain the temperature within a predefined temperature range.

This combustion adjustment method applied to the combustion control of a combustor executes a first step for selecting a combustion parameter for a load of a gas turbine and increasing or decreasing a command value of the combustion parameter from an original location. When the command value reaches a target tolerance upper limit value or a target tolerance lower limit value, the first step is terminated. Furthermore, a second step for decreasing or increasing the command value in a direction opposite to that of the first step is executed. When the command value reaches the target tolerance upper limit value or the target tolerance lower limit value, the present invention includes a combustion tolerance confirmation step for confirming a combustion tolerance range of the combustion parameter at which the second step is terminated.

This combustion adjustment method applied to the combustion control of a combustor executes a first step for selecting a combustion parameter for a load of a gas turbine and increasing or decreasing a command value of the combustion parameter from an original location. When the command value reaches a target tolerance upper limit value or a target tolerance lower limit value, the first step is terminated. Furthermore, a second step for decreasing or increasing the command value in a direction opposite to that of the first step is executed. When the command value reaches the target tolerance upper limit value or the target tolerance lower limit value, the present invention includes a combustion tolerance confirmation step for confirming a combustion tolerance range of the combustion parameter at which the second step is terminated.

Methods and systems for operating a gas turbine engine having variable geometry mechanisms are described. The method comprises detecting a failure event associated with the gas turbine engine, identifying a location of the failure event, selecting an aerodynamic load modulation schedule for the variable geometry mechanisms of the gas turbine engine as a function of the location of the failure event, and applying the aerodynamic load modulation schedule as selected to the variable geometry mechanisms during the failure event.

Methods and systems for operating a gas turbine engine having variable geometry mechanisms are described. The method comprises detecting a failure event associated with the gas turbine engine, identifying a location of the failure event, selecting an aerodynamic load modulation schedule for the variable geometry mechanisms of the gas turbine engine as a function of the location of the failure event, and applying the aerodynamic load modulation schedule as selected to the variable geometry mechanisms during the failure event.