RU2773081C2 - Method for determination of ignition of gas turbine engine - Google Patents

Abstract

FIELD: engine building.

SUBSTANCE: invention relates to a combustion chamber of a gas turbine engine. A method for the determination of the ignition of a combustion chamber of a gas turbine engine contains stages, at which: first measurement (MT1) of an exhaust gas temperature behind combustion chamber (3) is obtained (E11) before attempting to ignite mentioned combustion chamber (3), temperature threshold (ST) is obtained (E12), secondary determination criterion (CS) is obtained (E13), obtained temperature threshold (ST) is updated (E14) depending on obtained secondary determination criterion (CS), second measurement (MT2) of an exhaust gas temperature is obtained (E15) after attempting to ignite combustion chamber (3), updated temperature threshold (ST) and the difference between first (MT1) and second (MT2) measurements of an exhaust gas temperature are compared (E16), and the ignition state of combustion chamber (3) is determined (E17), corresponding to a successful or unsuccessful ignition attempt, depending on the result of comparison stage (E16).

EFFECT: invention makes it possible to detect the ignition of a combustion chamber of a gas turbine engine in case of a restart of the gas turbine engine on the ground or in flight.

10 cl, 6 dwg

Description

The field of technology to which the invention belongs

The invention relates to a method for determining the ignition of a gas turbine engine.

More specifically, the invention relates to a method for determining ignition of a combustion chamber of a gas turbine engine based on an exhaust gas temperature value. In particular, the invention relates to real-time determination of the ignition state of a combustion chamber corresponding to a successful or unsuccessful ignition attempt.

State of the art

A known gas turbine engine, such as an aircraft turbojet, typically includes a compressor, a combustion chamber, and a turbine arranged in this order from inlet to outlet with respect to the direction of air flow in the gas turbine engine. Air pre-compressed by the compressor enters the combustion chamber, then fuel is injected into the combustion chamber and the resulting mixture is ignited, which, when passing through the turbine, expands and is discarded, for example, to create the thrust necessary for the movement of the aircraft. Part of the energy received in the combustion chamber is additionally taken by the turbine to drive the compressor.

Determination of ignition in the combustion chamber is an essential criterion. In fact, it determines the actions taken to successfully start or restart a turbojet engine in flight.

Once combustion has started in the combustion chamber, a stream of hot air, referred to as "exhaust gases", passes through the gas turbine engine, causing a rapid rise in temperature downstream of the combustion chamber, especially at the turbine and its casing. Therefore, one or more temperature sensors are placed at this location to measure the change in temperature (EGT) of the exhaust gas.

A known method for determining ignition in a combustion chamber is to monitor the EGT measurements reported by these sensors. In addition, a number of parallel control logics may be implemented, for example by a gas turbine engine computer.

Regardless of currently known control logic, a gas turbine engine will be considered "fired" if the EGT has increased to a certain level since the start of fuel injection. To do this, the value of the EGT is first stored, for example after initialization of the startup sequence. After the fuel is injected and the spark plug is discharged (ie, after reaching the ignition speed), the subsequent measured EGT values are compared with the stored value. If a predetermined threshold is reached, it is determined that combustion occurs in the combustion chamber. In this case, the ignition control laws change in favor of the control laws relating to the acceleration or revolutions of the gas turbine engine until the idle speed is reached. On the other hand, if the EGT values do not exceed this threshold after a predetermined amount of time, then the pilot is given an alarm indicating that no combustion is occurring in the combustion chamber.

However, it is possible that the EGT value has already increased prior to launch. This phenomenon occurs, for example, when an ignition attempt has been made for the first time and then aborted, or after a rotating stall due to aerodynamic instability of the compressor. In this case, when the combustion chamber has been ignited, the EGT values are already too high and do not change as in the nominal case. Thus, it was not possible to reach the expected EGT increase threshold in time, and the start was incorrectly aborted when the engine actually ignited.

Therefore, there is a need to improve the method for detecting ignition in the combustion chamber, in particular in the case where the exhaust gas temperature is higher than usual at the time of fuel injection and spark plug discharge.

In addition, document FR3044703A1 is known, which describes a method for determining when ignition occurs in the combustion chamber of a gas turbine engine during a successful ignition attempt. Thus, the method makes it possible to determine the duration of the ignition, which can be used as an indicator of the deterioration of the performance of the entire system for starting a gas turbine engine in the event of, in particular, a confirmed deviation from the reference duration. The logic of the method is applied to data recorded during the flight, and which is usually processed after the flight, so that the state of health ("health state") of the gas turbine engine can be monitored. The use of the real-time method remains possible, but requires, in particular, significant computing resources.

Disclosure of the essence of the invention

One of the objectives of the invention is to detect the ignition of the combustion chamber of a gas turbine engine in the event of a restart of the gas turbine engine on the ground or in flight.

Another object of the invention is to detect the ignition of the combustion chamber of a gas turbine engine after a rotating separation of air at the outlet of a compressor.

Yet another object of the invention is to detect the ignition of the combustion chamber of a gas turbine engine in the event of a malfunction of the exhaust gas temperature sensor.

Another object of the invention is to detect ignition of the chamber in the combustion chamber of a gas turbine engine in the event of an early fuel cut.

Another object of the invention is to detect ignition of the combustion chamber of a gas turbine engine in the event of reinitialization of one or more computer channels during start-up or in flight.

The invention, in particular, proposes a method for determining the ignition of the combustion chamber of a gas turbine engine, comprising the following steps, in which:

- a first measurement of the temperature of the exhaust gas behind the combustion chamber is obtained before attempting to ignite said combustion chamber,

- get the temperature threshold,

- receive a secondary criterion for determining,

- updating the received temperature threshold depending on the received secondary determination criterion,

- obtain a second measurement of the exhaust gas temperature after an attempt to ignite the combustion chamber,

comparing the updated temperature threshold and the difference between the first and second exhaust gas temperature measurements, and

determining the ignition state of the combustion chamber corresponding to a successful or unsuccessful ignition attempt, depending on the result of the comparison step.

In this type of method, the step of updating the temperature detection threshold advantageously makes it possible to take into account the operating state of the gas turbine engine, which affects the starting of the combustion chamber. Thus, the change in exhaust gas temperature is not an absolute determination criterion, but becomes dependent on the general state of the gas turbine engine. Depending on secondary criteria related to the elements that surround the combustion chamber, such as, for example, engine speed or compressor outlet pressure, the change in exhaust gas temperature differs from its nominal change, even if the ignition state is the same. With this method, the ignition sensitive alarm given to the pilot is no longer garbled, which reduces misidentifications of ignition failure or re-ignition on the ground or in flight.

The method according to the invention may also contain the following features, taken alone or in combination:

- the step of obtaining the first measurement of the exhaust gas temperature is carried out at the end of the initiation phase of the turbine engine startup sequence, when the exhaust gas temperature is at its minimum,

- updating the temperature threshold consists in decreasing the value of said threshold if the secondary criterion is met and maintaining the value of said threshold otherwise,

- the secondary criterion is information relating to the change in pressure at the outlet of the high pressure compressor of the gas turbine engine, the secondary criterion being met if a sharp increase in said pressure is detected,

the secondary criterion is information related to the rotational speed of the gas turbine engine, wherein the secondary criterion is met if the engine speed is within a certain range for a predetermined period of time,

- it also contains the following steps:

○ get the first engine speed threshold,

○ get the engine speed measurement,

○ compare the measurement with the first engine speed threshold, and

○ determine the ignition state of the combustion chamber depending on the result of the comparison step,

- it also contains the following steps, in which:

○ get the engine speed acceleration threshold,

○ get the measurement time interval,

○ take a plurality of successive measurements of the acceleration of the engine speed during the measurement time interval obtained in the previous step, and

○ compare a set of consecutive measurements with the obtained acceleration threshold, and

○ determine the ignition state of the combustion chamber depending on the result of the comparison step, and

- it also contains the following steps, in which:

○ get the second engine speed threshold,

○ receive information related to the control state of the gas turbine engine,

○ get the engine speed measurement,

○ compare the measurement with the second engine speed threshold, and

○ determining the ignition state of the combustion chamber depending on the result of the comparison step and information related to the control state.

The invention also relates to a gas turbine engine combustion chamber ignition detection system, the system comprising an exhaust gas temperature sensor, the system also comprising a computer capable of implementing the method described above.

Finally, the invention is applicable to a gas turbine engine comprising the system described above.

Brief description of the drawings

Other features, objects and advantages of the present invention will become apparent upon reading the following detailed description with reference to the accompanying drawings, given by way of non-limiting examples, and in which:

in fig. 1 is a schematic representation of a known gas turbine engine comprising a detection system according to an embodiment of the invention,

in fig. 2a-2d schematically show various embodiments of the determination method according to the invention, and

in fig. 3 shows examples of changes in engine speed, exhaust gas temperature, and high pressure compressor outlet pressure during start-up of a gas turbine engine.

Implementation of the invention

Next, an exemplary implementation of the method according to the invention will be described with reference to the drawings.

With reference to FIG. 1, a conventional gas turbine engine 1, such as an aircraft turbojet, comprises a combustion chamber 3 in which combustion is necessary to start or restart the gas turbine engine 1 on the ground or in flight. The combustion chamber 3 is located between the compressor stage 2, called the high pressure compressor, and the turbine stage 4, called the high pressure turbine, connected to each other by a rotating shaft 5, called the high pressure stage shaft. Each of the stages of the compressor 2 and the turbine 4 may also in a known manner comprise several high pressure stages, respectively 21, 23 and 41, 43, connected to each other by the same rotating shaft 5 to form the high pressure casing of a two-shaft gas turbine engine. During operation, the air is compressed in the compressor section 2, then it passes into the combustion chamber 3, into which fuel is injected through nozzles 31 made for this purpose. Then, the fuel-air mixture is ignited by means of spark plugs 33. The products of this combustion, called exhaust gases, then expand while passing through the turbine section 4. In the following, the air passing through the turbine section 4 after the combustion chamber will be referred to as "exhaust gas", whether this air has been pre-ignited in combustion chamber 3 (nominal operation) or not (beginning of the start-up sequence). Therefore, the exhaust gas temperature is directly related to the temperature of the turbine stage 4 and indicates the ignition state of the combustion chamber 3 .

Typically, the start-up sequence of the gas turbine engine 1 comprises a first phase during which the speed of the rotating shaft 5 increases during a period of time between the receipt of a command to start the gas turbine engine 1, for example, from the pilot, and the moment when fuel injection begins into the combustion chamber 3 of the gas turbine engine 1. During this first phase, the rotating shaft 5 is driven independently of the operation of the turbine section 4, for example by means of a starter motor (not shown). This first phase can be qualified as the pre-injection phase.

In addition, the starting sequence comprises a second phase following the first phase, which ends when the rotating shaft is no longer driven independently of the operation of the turbine stage 4, for example by disengaging the starter clutch. This second phase begins with the ignition of the air-fuel mixture in the combustion chamber 3 and can be qualified as the post-ignition phase.

Further, the initiation phase of the start sequence preferably corresponds to the first phase described above, called the pre-injection phase, which continues until the moment when ignition of the air-fuel mixture begins.

When starting on the ground, the gas turbine engine 1 is first driven by a starter (not shown) to compress enough air to ensure successful ignition of the combustion chamber 3. During re-start in flight, if the speed of the gas turbine engine is high enough (a phenomenon called "autorotation" ), the combustion chamber 3 can be directly ignited. Otherwise, you must use the starter, as when starting on the ground. During start-up, as well as during restart, the combustion chamber 3 is ignited by fuel injection and discharge of spark plugs 33 .

In the first embodiment, with reference to FIG. 2a, the determination method E comprises a first step E11 in which a first measurement MT1 of the exhaust gas temperature, called the initial measurement, is obtained before attempting to ignite the combustion chamber 3. This initial measurement MT1 serves as a reference for controlling the ignition of the combustion chamber 3 . For this reason, the initial MT1 measurement is generally retained for later comparison purposes, as will be described in more detail.

Preferably, the step of obtaining E11 of the initial measurement MT1 is carried out at the end of the initiation phase of the gas turbine engine 1 startup sequence. It is in fact important to keep the minimum value of the exhaust gas temperature reached after the start of the gas turbine engine 1 startup or restart sequence. Thus, if said sequence is initiated, then actuation of the rotating shaft 5 by a starter or autorotation generates air circulation in the gas turbine engine 1, which is thus ventilated and gradually cooled, as shown in FIG. 3. Therefore, the temperature of the exhaust gas drops during the initial startup sequence of the gas turbine engine 1, so that when ignition of the combustion chamber 3 is performed, the temperature of the exhaust gas is at a minimum. This allows the use of an initial reference measurement, which is an accurate picture of the temperature of the combustion chamber 3 at the time of its ignition.

During the second step E12, the temperature threshold ST is received and also, according to the preferred embodiment, stored. This temperature threshold ST provides a reference point for the increase in exhaust gas temperature at which the combustion chamber 3 is considered to be ignited. This temperature threshold ST depends on the gas turbine engine 1, its wear condition, as well as its operating conditions. Advantageously, this temperature threshold ST is 35 K.

During the third step E13, the secondary determination criterion CS is received and also, according to the preferred embodiment, stored. The secondary determination criterion CS makes it possible to determine the state of the gas turbine engine 1 which affects the temperature of the exhaust gas. Depending on the fulfillment of the criterion, the secondary determination criterion CS allows the temperature threshold value ST obtained in the previous step E12 to be corrected to improve the ignition determination of the combustion chamber 3 based on the change in exhaust gas temperature.

During the fourth step E14, the temperature threshold ST is updated depending on the secondary determination criterion CS obtained in the previous step E13. Preferably, the update consists in decreasing the value of said threshold ST if the secondary criterion CS is met and maintaining the value of said threshold ST otherwise. In fact, if the secondary criterion CS is met, then the gas turbine engine 1 is potentially in operation, so that the change in exhaust gas temperature does not correspond to the nominal change, for example, it has already risen regardless of the ignition of the combustion chamber 3. This could be due to the first canceled start or due to a stall phenomenon occurring downstream of the compressor section 2. Thus, lowering the value of the threshold ST allows the ignition of the combustion chamber 3 to be detected more quickly, or even to ensure that the ignition of the combustion chamber 3 is detected. Likewise, updating the value of the temperature threshold ST depends on the gas turbine engine 1, its wear condition, as well as its operating conditions. Advantageously, if the secondary criterion CS is met, then the value of the updated temperature threshold ST is 15 K.

During the fifth step E15, a second measurement MT2 of the exhaust gas temperature after an attempt to ignite the combustion chamber 3 is obtained and, according to a preferred embodiment, stored. Preferably, several second exhaust gas temperature measurements MT2 are obtained in succession.

During the sixth step E16, the updated temperature threshold ST is compared with the difference between the second MT2 and the first MT1 exhaust gas temperature measurements obtained in the previous steps. Advantageously, the sixth step E16 is repeated sequentially for a plurality of second measurements MT2.

In the last step E17, the ignition state of the combustion chamber is determined depending on the result of the previous comparison step E16. More specifically, if the difference between the second MT2 and the first MT1 exhaust gas temperature measurements is greater than the updated temperature threshold ST, then the combustion chamber 3 is deemed to be ignited. Otherwise, it is considered that ignition has not occurred. Therefore, step E17 makes it possible to judge whether the ignition attempt of the combustion chamber 3 of the gas turbine engine 1 succeeded or failed. If necessary, an alarm can be given to the pilot if an ignition failure condition is determined after several ignition attempts.

The steps of the first embodiment of the determination method E may be implemented in an order different from that described above. Specifically, the step E12 of obtaining the temperature threshold ST may be performed independently of the steps E11 and E15 of performing the first temperature measurement MT1 and the second MT2. Thus, the temperature threshold ST can only be stored once during the first start of the gas turbine engine 1, and then it can be systematically used in each implementation of the determination method E. Similarly, obtaining E13 and updating E14 of the secondary determination criterion CS may be performed before obtaining E11 of the first temperature measurement MT1.

In ignition determination method E, various secondary determination criteria CS may be used.

The first secondary determination criterion CS1 is information regarding the pressure change at the outlet of the high pressure stage 23 farthest from the compressor section 2 of the gas turbine engine 1 with respect to the air circulation direction. In fact, as shown in FIG. 3, at the time of ignition of the combustion chamber 3, a sudden increase in pressure occurs, also referred to as a pressure surge. Thus, this first secondary criterion CS1 will be met if said pressure surge is determined. To do this, successive pressure measurements are obtained at the outlet of the high pressure stage 23 of the compressor 2. Advantageously, these measurements are separated by approximately one hundred milliseconds, for example 200 ms. If the deviation between successive measured values is greater than a certain predetermined pressure threshold, then a pressure jump is detected and the obtained secondary criterion CS1 is confirmed. Preferably, the pressure surge threshold is obtained during engine testing and is, for example, up to 4% of the value of the initial pressure measurement. In all cases, the first secondary criterion CS1 is used only when the gas turbine engine 1 is being started in order to limit its impact.

The second secondary criterion CS2 is information related to the rotational speed of the gas turbine engine 1, i. e. to the speed of the rotating shaft 5, for example, the high pressure shaft 5.

This second criterion CS2 is advantageously used in the case where the speed of the turbine engine 1 stagnates at a level lower than the idle threshold SR but higher than the maximum ventilation threshold SMV. In fact, when the engine speed crosses the maximum ventilation threshold SMV, for example, 7000 rpm, it is considered that ignition of the combustion chamber 3 has occurred. In effect, this SMV threshold is defined as the maximum threshold that can be crossed with the torque provided by the starter alone. Without ignition torque, this SMV threshold cannot be crossed, and therefore its crossing indicates that combustion chamber 3 has been ignited. This SMV threshold can be determined experimentally through airflow motion tests carried out on the gas turbine engine 1.

In order to fulfill this second criterion CS2, the operation state of the gas turbine engine 1 corresponding to the stagnation of the engine speed is first confirmed. To do this, monitor the change in engine speed for a predetermined duration T1 time, typically within 20 seconds. This duration T1 corresponds to the maximum time threshold for which the engine speed can be maintained above the maximum ventilation threshold SMV without starting. This duration T1 also depends on the type of gas turbine engine 1 and on its degree of wear. Therefore, the second secondary criterion CS2 is met if the engine speed is between the maximum ventilation threshold SMV and the idle threshold SR for a time corresponding to the maximum time threshold T1 described above.

Preferably, the second secondary criterion CS2 is considered to be met if, in addition, the following conditions are met for a time corresponding to the maximum time threshold T1 described above:

- the fuel supply valve in the combustion chamber 3 is open, and

- the starter is off.

The condition relating to the opening of the supply valve makes it possible to safely detect false ignition when fuel is not supplied to the combustion chamber 3 . This additional test condition allows the temperature threshold ST not to be updated in the event that the supply valve has been closed, i. e. when the launch was canceled or the flight ended. Thus, the determination method logic is reset to zero for the next re-ignition and/or restart.

In the second embodiment, with reference to FIG. 2b, the determination method E comprises a first step E21 in which the first engine speed threshold SRM1 is obtained, for example, the maximum ventilation threshold SMV described above is 7000 rpm.

During the second step E22, a measurement of the engine rotation speed MRM is obtained and, in accordance with the preferred embodiment, stored.

During the third step E23, the obtained MRM measurement is compared with the first engine speed threshold SRM1.

During the fourth step E24, the combustion conditions of the combustion chamber 3 are determined depending on the result of the comparison step E23. In this particular case, if the received measurement MRM is greater than the first engine speed threshold SMR1, then ignition of the combustion chamber 3 is deemed to have been carried out. In fact, as mentioned earlier, the first engine speed threshold SMR1 is advantageously chosen to indicate limits beyond which the gas turbine engine 1 cannot operate without the energy produced in the ignited combustion chamber 3. If the obtained measurement MRM is smaller than the first engine speed threshold SMR1, the combustion chamber 3 is considered to be extinguished. If desired, an alarm may be issued to the pilot if a failed ignition condition is detected after several ignition attempts.

The steps of the second embodiment of the determination method E may be implemented in an order different from that described above. Specifically, the step E21 of obtaining the first engine speed threshold SRM1 may be implemented independently of the step E22 of obtaining the engine speed measurement MRM. Thus, the first engine speed threshold can only be stored once during the first start of the gas turbine engine 1, and then it can be systematically used in each implementation of the determination method E.

With reference to FIG. 2c, in the third embodiment of the determination method E supplementing the second embodiment described above, step E31 of obtaining the threshold SA corresponding to the engine speed acceleration is performed. The acceleration threshold SA corresponds to an increase in the engine speed such that the engine needs to be driven by the combustion carried out in the combustion chamber 3 . This acceleration threshold SA depends on the type of turbine engine 1 and on its level of wear, and corresponds, for example, to an acceleration of the engine speed of 50 revolutions per minute per second.

Then, the measurement time interval E32 IT is also obtained and, according to the preferred embodiment, stored. This IT time interval is, for example, one second.

Thereafter, several successive measurements of the engine speed acceleration MA during the time interval IT obtained in the previous step E32 are obtained E33, and then compared with the acceleration threshold SA obtained in the first step E31.

Finally, the ignition state of the combustion chamber 3 is determined E34 depending on the results of the comparison step E33. In this case, if all acceleration measurements MA are greater than or equal to the received threshold SA, then the combustion chamber 3 is considered to have been ignited. Otherwise, the combustion chamber 3 is considered to be extinguished and an alarm may be given to the pilot if an extinguished condition is determined after several ignition attempts.

This embodiment of the detection method E is advantageously implemented in order to detect the re-ignition of the combustion chamber 3 when the gas turbine engine 1 is operating at a speed above the maximum ventilation threshold SMV and below the idle threshold SR.

The steps of the third embodiment of the determination method may be implemented in an order different from that described above. Specifically, the step E31, E32 of obtaining the acceleration threshold SA and the time interval IT can be performed independently of the step E33 of obtaining a plurality of engine speed acceleration measurements MA. Thus, the acceleration thresholds SA and the time interval IT can only be stored once during the first start of the gas turbine engine 1, and then they can be used systematically in each implementation of the determination method E.

With reference to FIG. 1, the turbojet 1 comprises one (or more) computer(s) 7 configured to control the operation of the turbojet 1. This type of computer 7 typically contains two channels that perform the same operations in parallel based on the same same received data. In general, one of the two channels is subordinate to the other. The expressions active channel and passive channel are used. The channels of the computer 7 receive measurements from one (or more) exhaust gas temperature sensor(s) 6 located behind the combustion chamber 3. Thus, the sensor 6 or this type of probe measures the temperature of the exhaust gas. Therefore, the gas turbine ignition detection system 8 advantageously comprises a temperature sensor 6 and a computer 7 of this type.

Moreover, the computer 7 is configured to perform ignition detection method E in accordance with one or more of the previously described embodiments. To this end, the computer 7 also receives other measurements from other sensors, such as, for example, measurements relating to the engine speed or the pressure downstream of the compressor stage 2, eg the high pressure stage 23.

In the event of a failure of the passive link, computer 7 may automatically be reinitialized and therefore lose all calculations currently in progress. As a result, the system 8 must be able to reinitialize itself into the ignition state of the combustion chamber 3. To do this, the passive channel may first restore the ignition state by communicating with the active channel.

Alternatively, with reference to FIG. 2d, the fourth embodiment of the method E for determining the ignition of the combustion chamber 3 is carried out by means of the passive channel of the computer 7.

During the first step E41, the second engine speed threshold SRM2 is received and, according to one preferred embodiment, stored. This second engine speed threshold SRM2 corresponds, for example, to the idle speed threshold SR described previously.

During the second step E42 also receive information regarding the state of control of the gas turbine engine, such as pilot commands related to the power requirement.

During the third step E43, the engine speed measurement MRM is received and, in accordance with one preferred embodiment, stored.

During the fourth step E44, the MRM measurement taken in the previous step is compared with the second engine speed threshold SRM2.

During the fifth step E45, the combustion conditions of the combustion chamber 3 are determined depending on the result of the comparison step E44 and information related to the control state. In this particular case, if the engine speed is greater than or equal to its idle speed SR (i.e., gas turbine engine 1 is in stable operation above idle), and there are corresponding pilot commands, then the passive channel concludes, that the combustion chamber 3 is ignited and controlled by the active channel.

The steps of the fourth embodiment of the determination method E may be implemented in an order different from that described above. Specifically, in the acquisition step E21, the second engine speed threshold SRM2 can only be stored once during the first start of the gas turbine engine 1, and then it can be systematically used in each implementation of the determination method E. Similarly, the steps E42, E43 of obtaining information regarding the control state and the measurement MRM of the engine speed can be interchanged.

Advantageously, the various embodiments described above are implemented in parallel via the channels of the computer 7 independently of each other. Thus, the first control logic that detects ignition of the combustion chamber 3 can send corresponding information to the pilot.

That is, if the gas turbine engine 1 is started when the exhaust gas temperature sensor 6 is defective, the logic according to the second embodiment of the determination method E will still detect ignition of the combustion chamber 3 .

Similarly, if there is an untimely reinitialization of one of the channels of the computer 7, then the corresponding logic circuit of the fourth embodiment of the determination method E will still allow the ignition of the combustion chamber 3 to be detected.

Similarly, if the engine speed stagnates after a fuel cut during startup of the gas turbine engine 1, then the determination logic implemented will depend on the level of the stagnation frequency (ground start or in-flight restart).

If the engine speed is less than the maximum ventilation threshold SMV, then the first embodiment of the determination method E will detect the ignition of the combustion chamber 3 by means of the first secondary criterion CS1, which is a pressure surge.

If the engine speed is transient, then the second embodiment of the detection method E will detect ignition of the combustion chamber 3 using the maximum ventilation threshold SMV as the first engine speed threshold SRM1.

If the engine speed is greater than the maximum ventilation threshold SMV, then the first embodiment of the determination method E will detect ignition of the combustion chamber 3 by means of the second secondary criteria CS2 representing the engine speed.

Claims (31)

1. Method (E) for determining the ignition of the combustion chamber (3) of a gas turbine engine (1), comprising the steps of: receive (E11) the first measurement (MT1) of the temperature of the exhaust gas behind the combustion chamber (3) before trying to ignite the said combustion chamber (3), get (E12) temperature threshold (ST), receive (E13) secondary criterion (CS) determination, updating (E14) the received temperature threshold (ST) depending on the received secondary determination criterion (CS), receive (E15) the second measurement (MT2) of the exhaust gas temperature after an attempt to ignite the combustion chamber (3), comparing (E16) the updated (ST) temperature threshold and the difference between the first (MT1) and second (MT2) exhaust gas temperature measurements and determining (E17) the ignition state of the combustion chamber (3) corresponding to a successful or unsuccessful ignition attempt, depending on the result of the comparison step (E16). 2. Method (E) according to claim 1, wherein the step (E11) of obtaining the first measurement (MT1) of the exhaust gas temperature is carried out at the end of the initiation phase of the startup sequence of the gas turbine engine (1) when the exhaust gas temperature is at a minimum. 3. Method (E) according to one of paragraphs. 1 or 2, wherein the update (E14) of the temperature threshold (ST) consists in decreasing the value of said threshold (ST) if the secondary criterion (CS) is met and maintaining the value of said threshold (ST) otherwise. 4. Method (E) according to claim 3, wherein the secondary criterion (CS) is information related to the pressure change at the outlet of the high pressure compressor (23) of the gas turbine engine (1), wherein the secondary criterion (CS) is met if a sharp increase in the mentioned pressure was revealed. 5. The method (E) according to claim 3, wherein the secondary criterion (CS) is information related to the engine speed of the gas turbine engine (1), wherein the secondary criterion (CS) is met if the engine speed is within a certain range of for a predetermined period (T1) of time. 6. Method (E) according to one of paragraphs. 1-5, further comprising the following steps, wherein: receive (E21) the first threshold (SRM1) engine speed, receive (E22) measurement (MRM) of the engine speed, compare (E23) the measurement (MRM) with the first threshold (SRM1) of the engine speed and determining (E24) the ignition state of the combustion chamber (3) depending on the result of the comparison step (E23). 7. The method (E) according to claim 6, further comprising the following steps, in which: get (E31) threshold (SA) acceleration of the engine speed, receive (E32) interval (IT) measurement time, receiving (E33) a plurality of successive measurements (MA) of engine speed acceleration during the measurement time interval (IT) obtained in the previous step (E32), comparing (E34) the plurality of successive measurements (MA) with the obtained acceleration threshold (SA) and determine (E35) the state of ignition of the combustion chamber depending on the result of step (E34) comparison. 8. Method (E) according to one of paragraphs. 1-7, further comprising the following steps, wherein: - receive (E41) the second threshold (SRM2) of the engine speed, - receive (E42) information related to the control state of the gas turbine engine (1), - receive (E43) measurement (MRM) of the engine speed, - compare (E44) the measurement (MRM) with the second threshold (SRM2) of the engine speed and determining (E45) the ignition state of the combustion chamber (3) depending on the result of the comparison step (E44) and information related to the control state. 9. The system (8) for determining the ignition of the combustion chamber (3) of the gas turbine engine (1), containing the exhaust gas temperature sensor (6) and the computer (7), configured to implement the method (E) according to one of paragraphs. 1-8. 10. Gas turbine engine containing the system according to claim 9.

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