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US 12,540,577 B2Granted
Aircraft propulsion system with start-up control device
AssigneesSafran, Safran Helicopter Engines
InventorsAntoine Pierre Emmanuel Ferrand, Fabien Mercier-Calvairac
Life of this patent
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Abstract
A propulsion system includes a gas turbine designed so that a combustion chamber can be ignited in a first ignition range of rotational speeds of a compressor shaft. The system further includes a control device designed to control an electric starter to accelerate the compressor shaft and, when the compressor shaft is accelerated, to control an attempt to ignite the combustion chamber. The gas turbine is designed so that the combustion chamber can be ignited in a second ignition range which is higher than the first ignition range, but not between these two ignition ranges, and the ignition attempt is carried out in the second ignition range.
Claims12
A propulsion system for an aircraft comprising:
- a gas turbine comprising:
- a combustion chamber, an ignition device configured to ignite the combustion chamber, a compressor configured to supply air to the combustion chamber, the compressor being equipped with a compressor shaft, and an electric starter configured to drive the compressor shaft, the gas turbine being configured so that the combustion chamber is ignited by the ignition device in a first ignition range (P1) of rotational speeds of the compressor shaft;
- a start-up control device designed to:
- with the combustion chamber switched off, control the electric starter to accelerate the compressor shaft, and with the compressor shaft accelerated, control the ignition device to attempt an ignition of the combustion chamber;
- wherein the combustion chamber is further configured to be ignited by the ignition device in a second ignition range (P3) of rotational speeds of the compressor shaft, which is greater than the first ignition range (P1), but not in a third rotational speed range (P2) between the first ignition range (P1) and the second ignition range (P3), the control device being designed, when the compressor shaft is accelerated, to accelerate the latter up to the second ignition range (P3) in order to attempt to ignite the combustion chamber in the latter.
2 more elements
A method for starting up a gas turbine of an aircraft, the gas turbine comprising:
- a combustion chamber, an ignition device configured to ignite the combustion chamber, a compressor configured to supply air to the combustion chamber, the compressor being equipped with a compressor shaft, and an electric starter configured to drive the compressor shaft, the qasr turbine being configured so that the combustion chamber is ignited by the ignition device in a first ignition range (P1) of rotational speeds of the compressor shaft;
- a start-up control device designed to:
- with the combustion chamber switched off, control the electric starter to accelerate the compressor shaft, and with the compressor shaft accelerated, control the ignition device to attempt an ignition of the combustion chamber, wherein the combustion chamber is further configured to be ignited by the ignition device in a second ignition range (P3) of rotational speeds of the compressor shaft, which is greater than the first ignition range (P1), but not in a third rotational speed range (P2) between the first ignition range (P1) and the second ignition range (P3), the control device being designed, when the compressor shaft is accelerated, to accelerate the latter up to the second ignition range (P3) in order to attempt to ignite the combustion chamber in the latter, the method comprising:
- while the combustion chamber of the gas turbine is switched off, controlling the electric starter of the gas turbine to accelerate the compressor shaft to drive the compressor configured to supply the air to the combustion chamber;
- and with the compressor shaft accelerated, controlling the ignition device of the gas turbine to attempt an ignition of the combustion chamber, and accelerating the compressor shaft to the second ignition range (P3) in order to attempt to ignite the combustion chamber.
2 more elements
A propulsion system for an aircraft comprising:
- a gas turbine comprising:
- a combustion chamber, an ignition device configured to ignite the combustion chamber, a compressor configured to supply air to the combustion chamber, the compressor being equipped with a compressor shaft, and an electric starter configured to drive the compressor shaft, the gas turbine being configured so that the combustion chamber is ignited by the ignition device in a first ignition range (P1) of rotational speeds of the compressor shaft;
- a start-up control device designed to:
- with the combustion chamber switched off, control the electric starter to accelerate the compressor shaft, and with the compressor shaft accelerated, control the ignition device to attempt an ignition of the combustion chamber;
- wherein the combustion chamber is further configured to be ignited by the ignition device in a second ignition range (P3) of rotational speeds of the compressor shaft, which is greater than the first ignition range (P1), but not between the ignition ranges (P1, P3), the second ignition range (P3) having a lower bound (n2) equal to at least 40% of a nominal rotational speed of the compressor shaft, the control device being designed, when the compressor shaft is accelerated, to accelerate the latter up to the second ignition range (P3) in order to attempt to ignite the combustion chamber in the latter.
2 more elements
Dependencies are inferred from the claim text.