AIRCRAFT PROPULSION SYSTEM WITH LAUNCH CONTROL DEVICE

Description

Technical field of the invention

[0001] The present invention relates to a system and method for propelling an aircraft, as well as a computer program for implementing the method. Technological background

[0002] Document FR 2 995 345 A1 describes an aircraft gas turbine comprising a combustion chamber, a compressor shaft for supplying compressed air to the combustion chamber and a starter connected to the shaft so as to provide it with sufficient starting torque to drive it in rotation. The document also describes the starting phases necessary for starting the gas turbine which include a so-called pre-acceleration phase during which the starter is activated to drive the shaft in rotation followed by a so-called stabilization phase during which the rotational speed of the shaft is maintained at a level. This stabilization phase corresponds to a time delay during which the various ignition equipment of the combustion chamber are activated to be ready for use. These phases have the particularity of being carried out when the compressor shaft is rotating at low speed, that is to say with a low rotational speed.Once the shaft rotation speed has stabilized, an attempt is made to ignite the combustion chamber using the activated ignition equipment. If the ignition is successful, the shaft is accelerated until it reaches a high rpm, i.e. high rotational speeds.

[0003] The prior art also includes documents WO 2013 / 007912 A1 and EP 2 006 496 A1 which describe methods for starting an aircraft gas turbine.

[0004] The invention proposes an improvement to these technologies, which in particular makes it possible to complete the starting strategy of a gas turbine by proposing a propulsion system improving the time and chances of starting. Summary of the invention

[0005] An aircraft propulsion system is therefore proposed comprising: a gas turbine comprising: a combustion chamber, a device for igniting the combustion chamber, a compressor for supplying air to the combustion chamber, the compressor being provided with a compressor shaft, and an electric starter for driving the compressor shaft, the gas turbine being designed so that the combustion chamber can be ignited by the ignition device within a first ignition range of rotational speeds of the compressor shaft; a start control device designed to: while the combustion chamber is extinguished, control the electric starter to accelerate the compressor shaft, and with the compressor shaft accelerated, control the ignition device to attempt ignition of the combustion chamber; characterized in that, the gas turbine being designed so that the combustion chamber can further be ignited by the ignition device in a second ignition range of rotational speeds of the compressor shaft, higher than the first ignition range, but not between these two ignition ranges, the control device is designed, upon acceleration of the compressor shaft, to accelerate the latter to the second ignition range to attempt ignition of the combustion chamber in the latter.

[0006] Thus, thanks to the invention, the starting time can be reduced, by eliminating a stabilization phase, or timing of the rotation speed. Indeed, during the acceleration phase up to high speed, with the combustion chamber off, the various ignition equipment such as the ignition device or one or more fuel injectors can be activated so as to be already ready for use when the second ignition range is reached. Also thanks to the invention, the reliability and reproducibility of the start is improved, operability problems such as, for example, the risk of overtemperatures, being eliminated.

[0007] The system according to the invention may comprise one or more of the following characteristics: the second ignition range has a lower limit at least equal to 40% of a nominal rotational speed of the compressor shaft; the first ignition range has an upper limit at most equal to 30% of the nominal rotational speed of the compressor shaft; the control device is designed, in the event of failure of the ignition attempt, to: decelerate the compressor shaft, until reaching the first ignition range; and with the rotational speed in the first range, control the ignition device to attempt ignition of the combustion chamber; the control device is designed to control the electric starter and the ignition device in response to an extinction of the combustion chamber and to the detection that the rotational speed falls below the second ignition range;the control device is configured to determine a lower limit of the second ignition range from at least one atmospheric environment parameter of the gas turbine.;

[0008] The invention also relates to an aircraft comprising a propulsion system according to the invention.

[0009] The invention also relates to a method for starting a gas turbine of an aircraft, comprising: while a combustion chamber of the gas turbine is extinguished, controlling an electric starter of the gas turbine to accelerate a compressor shaft driving a compressor supplying air to the combustion chamber; and with the compressor shaft accelerated, controlling an ignition device of the gas turbine to attempt ignition of the combustion chamber; characterized in that the gas turbine is designed so that the combustion chamber can be ignited by the ignition device in a first range ignition and in a second ignition range of rotational speeds of the compressor shaft, the second ignition range being greater than the first ignition range, but not between these two ignition ranges, the compressor shaft is accelerated to the second ignition range to attempt ignition of the combustion chamber in the latter.

[0010] The method according to the invention may comprise the following characteristic: a lower limit of the second ignition range is previously determined by testing on the gas turbine or by simulation of the gas turbine.

[0011] The invention also relates to a computer program downloadable from a communication network and / or recorded on a computer-readable medium, characterized in that it comprises instructions for executing a startup method according to the invention, when the computer program is executed in a control device according to the invention. Brief description of the figures

[0012] The invention will be better understood with the aid of the following description, given solely by way of example and with reference to the appended drawings in which: [ Fig. 1 ] there figure 1 is a very schematic view of an aircraft propulsion system with its start control device, according to the invention, [ Fig. 2 ] there figure 2 is a block diagram of an example of a method for determining ignition thresholds for rotational speed according to the invention, [ Fig. 3 ] there figure 3is a block diagram of a first example of a starting method according to the invention, [ Fig. 4 ] there figure 4 is a block diagram of a second example of a start-up method according to the invention, in the case of a prior shutdown of the gas turbine, [ Fig. 5 ] there Figure 5 is a diagram representing the evolution of an aerodynamic load parameter as a function of a rotation speed of a compressor shaft of the propulsion system according to the invention, [ Fig. 6 ] there figure 6 is a diagram representing the start-up phases of a gas turbine of an aircraft equipped with the system according to the invention, and [ Fig. 7 ] there figure 7 is a diagram representing the start-up phases of a gas turbine of an aircraft equipped with the system according to the invention, in the case of a prior shutdown of the gas turbine. Detailed description of the invention

[0013] There figure 1illustrates an example of a propulsion system 100 of an aircraft according to the invention comprising a gas turbine 102 and a control device 106 for starting the gas turbine 102.

[0014] The gas turbine 102 comprises a gas generator 104 having a combustion chamber 110 and a fuel injector 116 in the combustion chamber 110. In the latter, pressurized air is intended to be introduced to mix with the fuel in order to carry out combustion. The gas generator 104 further comprises an ignition device 108 of the combustion chamber 110 designed to initiate combustion.

[0015] In order to raise the pressure of the air intended to be introduced into the combustion chamber 110, the gas generator 104 further comprises a compressor 113 designed to supply pressurized air to the combustion chamber 110.

[0016] To drive the compressor 113, the gas generator 104 further comprises a turbine 111 and a mechanical shaft, called the compressor shaft 112, connecting the compressor 113 to the turbine 111. The turbine 111 is designed to be driven in rotation by the gases resulting from the combustion in the combustion chamber 110 and thus to drive the compressor shaft 112 in rotation at a rotational speed N GG . The compressor 113 is thus driven by the compressor shaft 112.

[0017] The propulsion system 100 comprises, for example, a sensor 115 which makes it possible, for example, to measure this rotation speed N GG .

[0018] The surplus thermal energy generated at the outlet of the gas generator 104 can be recovered by a member 117 of the propulsion system 100. This member 117 can be a nozzle for generating thrust, a free or linked turbine for generating mechanical energy to drive a rotor or a propeller or even an electric generator.

[0019] At start-up, the thermal energy at the outlet of the combustion chamber 110 is insufficient to drive the turbine 111 with the compressor shaft 112. To overcome this, the gas turbine 102 further comprises an electric machine 114 (for example reversible) which can be controlled in motor mode for driving the compressor shaft 112, the electric machine 114 then having a starter function. In the following, the electric machine 114 is called the electric starter 114.

[0020] The electric starter 114 is connected to the compressor shaft 112 and provides torque to the compressor shaft 112 to drive it in rotation. The assistance provided by the electric starter 114 also makes it possible to overcome all of the resistive torques of the rotating parts.

[0021] When reversible, the electric machine 114 can also have an alternator mode in which it can draw torque from the compressor shaft 112 and thus become an electric generator.

[0022] To electrically power the electric starter 114, the gas turbine 102 comprises, for example, an electrical storage device 118 connected to the electric starter 114. The electrical storage device 118, for example one or more batteries, is designed to deliver high powers in order to allow the electric starter 114 to drive the compressor shaft 112 in rotation. The storage device 118 can be recharged by the electric starter 114 when the latter is in alternator mode.

[0023] The storage device 118 can be configured to deliver a power that can reach between 1 and 15% of a total power delivered by the gas turbine 102 when the latter is at 100% of its nominal speed (in standard atmospheric conditions on the ground), and advantageously between 5 and 11% of this total power.

[0024] The propulsion system 100 further comprises a measurement system 120 configured to measure one or more atmospheric environmental parameters of the gas turbine 102 and to provide the measurements to the controller 106.

[0025] The control device 106 is for example designed to receive the measurements from the measuring system 120 and from the sensor 115 of the rotation speed of the compressor shaft 112.

[0026] In the example described, the control device 106 comprises a computer system comprising a data processing unit 122 (such as a microprocessor) and a main memory 124 accessible by the processing unit 122. This computer system further comprises a computer program 128 containing instructions for the processing unit 122. This computer program 128 is for example intended to be loaded into the main memory 124, all at once or in part, so that the processing unit 122 executes its instructions.

[0027] Prior to its loading into the main memory 124, the computer program 128 is for example recorded on a medium 126 suitable for this purpose (such as a local mass memory like a hard disk) and / or downloaded via a network interface of the computer system.

[0028] For example, this computer system is an on-board calculator of the ECU type (from the English “Electronic Control Unit”).

[0029] The instructions of the computer program 128 are designed so that, when executed by the data processing unit 122, the control device 106 performs the steps which will be described later.

[0030] Alternatively, all or part of these steps can be carried out by electronic circuits, for example micro-wired, not involving a computer program.

[0031] The control device 106 comprises a memory, such as the memory 124, in which data are recorded defining in particular thresholds n1, n2 for the rotation speed N GG of the compressor shaft 112. These data will be described in more detail later.

[0032] There figure 2illustrates an example of a method 200, according to the invention, for determining the thresholds n1, n2 for the rotation speed N GG of the compressor shaft 112. This method 200 is for example implemented during the design of the gas turbine 102.

[0033] Indeed, the capacity of the combustion chamber 110 to burn the introduced fuel is expressed by an aerodynamic load Ω which depends on the atmospheric environment in which the gas turbine 102 is supposed to operate (this atmospheric environment being characterized for example by an external pressure P0 and / or an external temperature T0 at the gas turbine 102) and the rotation speed of the compressor shaft 112. The aerodynamic load Ω thus makes it possible to evaluate the performance of the combustion chamber 110.

[0034] The aerodynamic load curve Ω depends on the architecture of the gas turbine 102 considered, which defines, among other things, the residence time of the gases in the combustion chamber 110 as well as the speed of the combustion reactions. In other words, the aerodynamic load curve Ω is specific to each gas turbine model.

[0035] For example, the aerodynamic load Ω considered is equivalent to the combustion chamber loading given in the book Gas Turbine Performance, Walsh, PP and Fletcher, P., Blackwell Science, Oxford (1998). Thus, the aerodynamic load Ω is for example given by the following equation: Ω = W 32 VOL . P 32 1.8 .10 0.00145 . T 32 where VOL is the volume of the combustion chamber 110, P 32 and T 32 are respectively the inlet pressure and temperature of the gases into the combustion chamber 110, and W 32 is the inlet mass flow rate of the gases into the combustion chamber 110.

[0036] The volume VOL of the combustion chamber 110 is a design parameter of the gas turbine 102, while the pressure P 32 , the temperature T 32 and the flow rate W 32 depend on the design of the gas turbine 102, but also on its atmospheric environment and the rotation speed N GG of the compressor shaft 112.

[0037] In the example described, the ISA standard (International Standard Atmosphere) is used to define the atmospheric environment. This ISA standard associates a predefined temperature and pressure with each altitude. Thus, in the example described, the atmospheric environment is completely defined by the altitude H. Thus, the aerodynamic load Ω is a function of the altitude H and the rotational speed N GG of the compressor shaft 112.

[0038] Alternatively, the aerodynamic load Ω is further defined as a function of a difference between the temperature ISA at this altitude H and the measured outside temperature T0. Thus, in this alternative, the aerodynamic load Ω is a function of the altitude H, the difference from the temperature ISA at this altitude H and the rotation speed N GG of the compressor shaft 112.

[0039] There Figure 5shows two curves A, B of aerodynamic load Ω of the combustion chamber 110 as a function of the rotation speed N GG of the compressor shaft 112, for respectively two atmospheric environments defined by the ISA standard: the first when the altitude H is zero, i.e. at sea level (SLS, from the English for “Sea Level Static”), and the second when the altitude H is higher than sea level, for example at an altitude of 6,000 m. The rotation speed N GG is expressed as a percentage of a nominal speed of the gas generator 104. This nominal speed is generally a high-speed speed of the compressor shaft 112 taken as a reference.

[0040] As can be seen on the Figure 5, the aerodynamic load curves A, B are "bell-shaped", that is to say that, for each one, the aerodynamic load Ω is zero when the rotation speed N GG is zero, then increases with the rotation speed N GG , reaches a maximum, then decreases with the rotation speed N GG to tend towards zero when the rotation speed N GG is high, that is to say approaches the nominal speed.

[0041] It has been determined that the combustion chamber 110 can only be ignited when the aerodynamic load Ω is below a certain ignition limit L. This limit depends on the aerodynamic load and the air flow rate W 32 entering the combustion chamber 110, the air flow rate being able to be brought back to the rotational speed N GG of the compressor shaft 112.

[0042] Due to their bell shape, each of the curves A and B passes above the ignition limit L at a first threshold n1, then passes below it at a second threshold n2. Between these two thresholds n1, n2, the curve remains above the ignition limit L, which indicates an impossibility of igniting the combustion chamber 110. On the other hand, the combustion chamber 110 can be ignited for a rotation speed N GG of the compressor shaft 112 lower than the threshold n1 or higher than the threshold n2.

[0043] Thus, the gas turbine 102 is designed to allow ignition of the combustion chamber 110 over each of the ranges P1 = (approximately 0%, n1) and P3 = (n2, 100%). On the contrary, the gas turbine 102 is designed such that ignition is not possible over the range P2 (n1, n2). The threshold n1 marks an upper limit of the first ignition range P1 = (approximately 0%, n1), while the threshold n2 marks a lower limit of the second ignition range P3 = (n2, 100%).

[0044] There Figure 5 also shows that these thresholds n1, n2 can vary depending on the atmospheric environment of the gas turbine 102. Comparing curve A with curve B shows that the threshold n1 decreases with altitude H so that n1(A) > n1(B) and that the threshold n2 increases with altitude so that n2(A) < n2(B). In other words, the range P2 = (n1, n2) over which ignition of the combustion chamber 110 is not possible is wider at altitude than at sea level.

[0045] Furthermore, it was determined that, for a deviation from the ISA temperature at the altitude H considered, the range P2 (n1, n2) over which ignition is not possible is wider with a temperature lower than the ISA temperature at this altitude.

[0046] Threshold n1 is generally at most equal to 30% of the nominal speed, while threshold n2 is generally at least equal to 40% of the nominal speed, and more particularly generally at least equal to 50% of the nominal speed.

[0047] Thus, at a step 202, for at least one atmospheric environment in which the gas turbine 102 is supposed to operate, the thresholds n1 and n2 are determined, for example by numerical simulation of the gas turbine 102, or by testing the latter.

[0048] For example, for each atmospheric environment, an increasingly high air flow rate is sent by a dedicated device into the combustion chamber 110 (the compressor is not used and can be detached from the combustion chamber 110), while attempts are made to ignite the combustion chamber 110. This makes it possible to know the air flow rate ranges where ignition is possible. However, the relationship between the rotation speed N GG of the compressor shaft 112 which will make it possible to obtain each air flow rate value is known, which makes it possible to deduce the ranges P1, P2, P3 and therefore the thresholds n1, n2 from the air flow rate ranges.

[0049] Alternatively, at this step 202, the aerodynamic load curve Ω is determined, for example by simulation or testing as previously, and compared to the predetermined ignition limit L, to determine the thresholds n1, n2.

[0050] Preferably, thresholds n1 and n2 are determined for several atmospheric environments.

[0051] In a step 204, the thresholds n1 and n2 are recorded in the memory 124 of the control device 106, in association with the atmospheric environment for which they were determined.

[0052] For example, a table is recorded in the control device 106, this table indicating for each of several altitudes H, the associated thresholds n1 and n2: [Table 1] External atmospheric conditions first threshold n1 second threshold n2 ISA, altitude H = 0 n1(A) n2(A) ISA, altitude H = a value greater than 0 n1(B) n2(B)

[0053] In the case where the difference between the ISA temperature and the measured outside temperature T0 is taken into account, the table can indicate the thresholds n1 and n2 for several altitudes H and each time for several differences from the ISA temperature at the altitude H considered.

[0054] In this way, the control device 106 knows, for several altitudes H and, where appropriate for several deviations from the temperature ISA, the rotation speed ranges N GG over which the combustion chamber 110 is capable or incapable of igniting.

[0055] Alternatively, the table could be replaced by a function linking the thresholds n1, n2 to the atmospheric environment parameter(s) (altitude H or altitude H and the deviation from the ISA temperature in the example described).

[0056] THE figures 3 And 6 illustrate a first example of a method 300 for starting the gas turbine 102. In this example, the gas turbine 102 is initially assumed to be stopped, with the combustion chamber 110 extinguished.

[0057] In a step 302, the control device 106 obtains the thresholds n1 and n2. For example, the measurement system 120 measures the atmospheric environment parameter(s) of the gas turbine 102. In the example described, the altitude H is measured, as well as, where appropriate, the outside temperature T0 to know the deviation from the temperature ISA at the altitude H. The control device 106 then determines the thresholds n1, n2 from the measured environmental parameter(s). In the example described, the thresholds n1 and n2 associated with the measured outside atmospheric conditions are retrieved from the recorded table.

[0058] At step 304 (time t0 on the figure 6 ), the control device 106 activates the electric starter 114 so that the latter provides a driving torque to the compressor shaft 112 and accelerates the latter, for example with constant acceleration.

[0059] At a step 306, the control device 106 controls the injector 116 to inject the fuel into the combustion chamber 110. This step can for example be carried out from the moment when 90% of the value of the threshold n2 is reached. Simultaneously, the control device 106 can control the ignition device 108 to activate so that it is ready to operate as soon as the threshold n2 is crossed.

[0060] At step 308 (time t1 on the figure 6), the control device 106 detects that the rotational speed N GG of the compressor shaft 112 crosses the threshold n2 and, in response, controls the ignition device 108 to attempt to ignite the combustion chamber 110. Several ignition attempts can be made. During this time, the rotational speed N GG of the compressor shaft 112 increases up to a maximum rotational speed n (1)< at which the electric starter 114 alone can drive the compressor shaft 112. This will subsequently be called the maximum assistance speed n (1)< .

[0061] In the event of successful ignition, the gas generator 104 starts at a step 310 and the turbine 111 drives the compressor shaft 112 so that its rotation speed N GG increases until it reaches a rotation speed nc called the end of start setpoint (dotted curve portion on the figure 6 ).

[0062] After starting the gas generator 104, assistance by the electric starter 114 is no longer necessary. The control device 106 can then deactivate the electric starter 114 to stop providing torque to the compressor shaft 112 or, alternatively, continue to assist the compressor shaft 112 until a so-called cut-off speed is reached from which the turbine 111 alone ensures the rotation of the compressor shaft 112, the cut-off speed being lower than the end-of-start setpoint speed.

[0063] In the event of failure of ignition of the combustion chamber 110 in step 308, for example if the control device 106 detects that the rotation speed N GG remains for a certain predefined time at the maximum assistance speed n (1) < without ignition of the combustion chamber 110, the control device 106 preferably implements a second chance start at low speed, in the first ignition range P1. The chances of starting are thus improved.

[0064] For this, at a step 312 (time t2 on the figure 6), the control device 106 controls the electric starter 114 to stop driving the compressor shaft 112, so that the rotational speed N GG decreases to the first firing range P1. For example, the control device 106 controls the electric starter 114 in generator mode to take torque from the compressor shaft 112 and thus decelerate the latter. The use of the electric starter 114 allows rapid deceleration. Alternatively, the deceleration can be free, that is to say, the control device 106 can deactivate the electric starter 114 so that the latter does not act on the compressor shaft 112.

[0065] At step 314 (time t3 on the figure 6 ), the control device 106 detects that the rotation speed N GG crosses the threshold n1 and, in response, controls the ignition device 108 to attempt to ignite the combustion chamber 110 in the first ignition range P1.

[0066] At step 316 (time t4 on the figure 6 ), the control device 106 controls the electric starter 114 to stabilize the rotation speed N GG of the compressor shaft 112 at a speed n (2) < included in the first ignition range P1 in order to allow the continuation of ignition attempts if the ignition has not succeeded in step 314 or to maintain the flame in the combustion chamber 110 in the event of successful ignition in step 314.

[0067] At step 318 (time t5 on the figure 6), while the combustion chamber 110 is ignited, the control device 106 can continue to control the electric starter 114 to continue assisting the compressor shaft 112 until it reaches the cut-off speed from which the turbine 111 alone ensures the rotation of the compressor shaft 112. The control device 106 deactivates, for example, the electric starter 114 in response to the detection of this cut-off speed. The turbine 111 then continues to accelerate the compressor shaft until the set speed nc at the end of start-up.

[0068] In reference to the figures 4 And 7 , a second example of a method 400 for starting the gas turbine 102 will now be described. In this second example, the start follows an undesired flameout of the combustion chamber 110 at high speed.

[0069] In such a case, at the time of switching off, the rotational speed N GG of the compressor shaft 112 is generally high and is located for example (as illustrated in the figure 7 ) in the second ignition range P3, above threshold n2.

[0070] Thus, at a step 402 (time t'1 on the figure 7 ), an extinction of the combustion chamber 110 occurs and the rotation speed N GG of the compressor shaft 112 begins to decrease.

[0071] This deceleration is for example free, that is to say that the electric starter 114 is inactive.

[0072] The method 400 then comprises the step 302 described previously for obtaining the threshold n2. For example, the threshold n2 is obtained from the measured environmental parameter(s), and more precisely from the altitude H or from the altitude H and the difference from the temperature ISA in the example described.

[0073] In an optional step 404, the control device 106 controls the ignition device 108 to attempt to relight the combustion chamber 110 at least once before reaching the threshold n2.

[0074] At step 406 (time t'2 on the figure 7 ), the threshold n2 is crossed and the combustion chamber 110 is no longer able to ignite. The control device 106 detects this crossing and then activates in response the electric starter 114 to supply torque to the compressor shaft 112. This makes it possible to limit the drop in the rotation speed N GG of the compressor shaft 112, then to increase this rotation speed N GG to reach the second ignition range P3.

[0075] At step 408 (time t'3 on the figure 7), the rotation speed N GG of the compressor shaft 112 crosses the threshold n2, which the control device 106 detects to control, in response, the ignition device 108 to attempt to ignite the combustion chamber 110 as quickly as possible after crossing the threshold n2.

[0076] At step 410 (time t'4 on the figure 7 ), following ignition of the combustion chamber 110, the control device 106 detects that the gas generator 104 is autonomous. The assistance of the electric starter 114 is then no longer necessary and the control device 106 then deactivates the electric starter 114.

[0077] Such a start-up allows rapid re-ignition of the combustion chamber 110 after it has been extinguished. This saves time for restarting the gas turbine 102, compared to the case where it would be necessary to reduce the rotation speed N GG of the compressor shaft 112 to the first ignition range P1 to carry out the restart there.

[0078] In some embodiments, the control device 106 may previously check the amount of energy available in the electrical storage device 118, for example during a scheduled shutdown of the gas turbine 102. In the case where the available energy is insufficient to allow a high-speed start (i.e. in the second ignition range P3) as described previously, the control device 106 may command the recharging of the electrical storage device 118, for example from an onboard generator (not shown) designed to generate electrical energy on board an aircraft, before the gas turbine 102 is shut down. Thus, the electrical storage device 118 is capable of powering the electric starter 114 for a future start (normal or following an unexpected shutdown).

[0079] It is clear that an aircraft propulsion system as described above makes it possible to reduce the start-up time by eliminating a phase of stabilization of the rotation speed N GG to allow the activation of different ignition equipment.

[0080] The propulsion system described above also has the advantage of improving the chances of starting the gas turbine by allowing ignition of the combustor over a low speed range after a failed ignition attempt over a high speed range.

[0081] Another advantage is that it allows for rapid re-ignition of the combustion chamber after it has been extinguished when the rotational speed is still close to or within the high speed range. This can be understood to save time when restarting the gas turbine.

[0082] Another advantage is that it improves the reliability and reproducibility of starting. Indeed, operability issues such as the risk of overtemperatures, which can impact the lifespan of the hot parts of the engine, are eliminated. The risks of overtemperatures can be linked, among other things, to an accumulation of fuel in the combustion chamber before successful ignition. In the context of the invention, the air flow is high when ignition is attempted, which allows rapid evacuation of fuel from the combustion chamber where it does not have time to accumulate.

[0083] It will also be noted that the invention is not limited to the embodiments described above. It will indeed appear to the person skilled in the art that various modifications can be made to the embodiments described above, in light of the teaching which has just been disclosed to him.

[0084] For example, the thresholds used could be constant, and thus not depend on atmospheric environmental parameters of the gas turbine. Thus, the control device would use in particular a single value of threshold n1 and a single value of threshold n2. For example, these values ​​would be taken in the worst case, for example at a maximum altitude that the aircraft is designed to reach. Thus, these constant thresholds could be used in all situations, in particular both for a start from a standstill and for a restart following an untimely flameout of the combustion chamber in full flight.

Claims

1. A propulsion system (100) for an aircraft comprising: - a gas turbine (102) comprising : a combustion chamber (110), an ignition device (108) for igniting the combustion chamber (110), a compressor (113) for supplying air to the combustion chamber (110), the compressor (113) being equipped with a compressor shaft (112), and an electric starter (114) for driving the compressor shaft (112), the gas turbine (102) being designed so that the combustion chamber (110) can be ignited by the ignition device (108) in a first ignition range (P1) of rotational speeds of the compressor shaft (112); - a start-up control device (106) designed to: with the combustion chamber (110) switched off, control the electric starter (114) to accelerate the compressor shaft (112), and with the compressor shaft (112) accelerated, control the ignition device (108) to attempt an ignition of the combustion chamber (110); characterised in that, since the gas turbine (102) is designed in such a way that the combustion chamber (110) can additionally be ignited by the ignition device (108) in a second ignition range (P3) of rotational speeds of the compressor shaft (112) which is greater than the first ignition range (P1), but not between these two ignition ranges (P1, P3), the control device (106) is designed, when the compressor shaft (112) is accelerated, to accelerate the latter up to the second ignition range (P3) in order to attempt to ignite the combustion chamber (110) in the latter.

2. The propulsion system (100) according to claim 1, wherein the second ignition range (P3) has a lower bound (n2) at least equal to 40% of a nominal rotational speed of the compressor shaft (112).

3. The propulsion system (100) as claimed in claim 2, wherein the first ignition range (P1) has an upper bound (n1) of at most 30% of the nominal rotational speed of the compressor shaft (112).

4. The propulsion system (100) according to any one of claims 1 to 3, wherein the control device (106) is designed, in the event of failure of the ignition attempt, to: - decelerate the compressor shaft (112) until it reaches the first ignition range (P1); and - with the rotational speed (NGG) in the first range (P1), control the ignition device (108) to attempt an ignition of the combustion chamber (110).

5. The propulsion system (100) according to any one of claims 1 to 4, wherein the control device (106) is arranged to control the electric starter (114) and the ignition device (108) in response to an extinction of the combustion chamber (110) and to the detection that the rotational speed (NGG) falls below the second ignition range (P3).

6. The propulsion system (100) according to any one of claims 1 to 5, wherein the control device (106) is configured to determine a lower bound (n2) of the second ignition range (P3) on the basis of at least one measured parameter of the atmospheric environment of the gas turbine (102).

7. An aircraft comprising a propulsion system (100) according to any one of claims 1 to 6.

8. A method (300; 400) for starting up a gas turbine (102) of an aircraft, comprising: - while a combustion chamber (110) of the gas turbine (102) is switched off, controlling (304; 406) an electric starter (114) of the gas turbine to accelerate a compressor shaft (112) driving a compressor (113) supplying air to the combustion chamber (110); and - with the compressor shaft (112) accelerated, controlling (308; 408) an ignition device (108) of the gas turbine (102) to attempt an ignition of the combustion chamber (110); characterised in that, the gas turbine (102) being designed so that the combustion chamber (110) can be ignited by the ignition device (108) in a first ignition range (P1) and in a second ignition range (P3) of rotational speeds of the compressor shaft (112), the second ignition range (P3) being greater than the first ignition range (P1), but not between these two ignition ranges (P1, P3), the compressor shaft (112) is accelerated to the second ignition range (P3) in order to attempt to ignite the combustion chamber (110) in the latter.

9. The method (300; 400) according to claim 8, wherein a lower bound (n2) of the second ignition range (P3) is previously determined by testing the gas turbine (102) or by simulating the gas turbine (102).

10. A computer program (128) downloadable from a communications network and / or recorded on a computer-readable medium, characterised in that it comprises instructions for executing a start-up method (300, 400) according to claim 9, when the computer program (128) is executed in a control device (106) according to claim 1.