阅读说明：本技术 用于检测涡轮发动机的点火的方法 (Method for detecting ignition of a turbine engine ) 是由 阿德尔·塞德里克·阿比尔·维德曼 尼古拉斯·莱拉德 赛德瑞克·德杰拉希 于 2018-12-13 设计创作，主要内容包括：本发明涉及用于检测涡轮发动机燃烧室的点火的方法(E),该方法(E)包括以下步骤：在尝试对所述燃烧室进行点火之前,接收(E11)燃烧室下游的第一废气温度测量值；接收(E12)温度阈值；接收(E13)次级检测标准；根据接收到的次级检测标准来更新(E14)接收到的温度阈值；在尝试对燃烧室进行点火之后,接收(E15)第二废气温度测量值；将更新后的温度阈值与第一废气温度测量值和第二废气温度测量值之间的差值进行比较(E16)；以及确定(E17)燃烧室的点火状态。(The invention relates to a method (E) for detecting the ignition of a combustion chamber of a turbine engine, comprising the following steps: receiving (E11) a first exhaust gas temperature measurement downstream of the combustion chamber before attempting to ignite the combustion chamber; receiving (E12) a temperature threshold; receiving (E13) secondary detection criteria; updating (E14) the received temperature threshold according to the received secondary detection criterion; receiving (E15) a second exhaust gas temperature measurement after attempting to ignite the combustion chamber; comparing (E16) the updated temperature threshold to a difference between the first exhaust temperature measurement and the second exhaust temperature measurement; and determining (E17) the ignition state of the combustion chamber.)

1. Method (E) for detecting the ignition of a combustion chamber (3) of a turbine engine (1), said method (E) comprising the steps of:

receiving (E11) a first exhaust gas temperature measurement (MT1) downstream of the combustion chamber (3) before attempting to ignite the combustion chamber (3),

receiving (E12) a temperature threshold (ST),

receiving (E13) a secondary detection Criterion (CS),

updating (E14) the received temperature threshold (ST) according to the received secondary detection Criterion (CS),

-receiving (E15) a second exhaust gas temperature measurement (MT2) after attempting to ignite the combustion chamber (3),

comparing (E16) the updated temperature threshold (ST) with the difference between the first exhaust gas temperature measurement (MT1) and the second exhaust gas temperature measurement (MT2), and

-determining (E17) the ignition status of the combustion chamber (3) as a function of the result of the comparison step (E16), said ignition status corresponding to the success or failure of an ignition attempt.

2. A method (E) according to claim 1, wherein the step (E11) of receiving a first exhaust gas temperature measurement (MT1) is carried out at the end of a start-up phase of a start-up procedure of the turbine engine (1) when the exhaust gas temperature is minimal.

3. Method (E) according to claim 1 or 2, wherein the updating (E14) of the temperature threshold (ST) comprises: -reducing the value of Said Threshold (ST) if the secondary Criterion (CS) is verified, otherwise retaining the value of Said Threshold (ST).

4. A method (E) according to claim 3, wherein said secondary Criterion (CS) is information relating to the evolution of a high-pressure compressor output pressure (23) of said turbine engine (1), said secondary Criterion (CS) being verified if a sudden increase in said pressure is detected.

5. A method (E) according to claim 3, wherein the secondary Criterion (CS) is information relating to the engine speed of the turbine engine (1), said secondary Criterion (CS) being verified if the engine speed is within a determined interval for a predetermined duration (T1).

6. Method (E) according to one of claims 1 to 5, further comprising the steps of:

receiving (E21) a first engine speed threshold (SRM1),

receiving (E22) a measured value (MRM) of the engine speed,

comparing (E23) the measured value (MRM) with the first engine speed threshold (SRM1), and

-determining (E24) the ignition state of the combustion chamber (3) as a function of the result of the comparison step (E23).

7. The method (E) according to claim 6, further comprising the steps of:

receiving (E31) an acceleration threshold (SA) of the engine speed,

receiving (E32) a measurement time Interval (IT),

receiving (E33) during the reception of the measurement time Interval (IT) in the preceding step (E32)

A plurality of successive Measurements (MA) of the acceleration of the engine speed, and

comparing (E34) a plurality of consecutive measured values (MA) with a received acceleration threshold value (SA), and

-determining (E35) the ignition state of the combustion chamber (3) as a function of the result of the comparison step (E34).

8. Method (E) according to one of claims 1 to 7, further comprising the steps of:

receiving (E41) a second engine speed threshold (SRM2),

receiving (E42) information relating to a control state of the turbine engine (1),

receiving (E43) a measured value (MRM) of the engine speed,

comparing (E44) the measured value (MRM) with the second engine speed threshold (SRM2), and

-determining (E45) the ignition state of the combustion chamber (3) as a function of the result of the comparison step (E44) and the information relating to the control state.

9. System (8) for detecting ignition of a combustion chamber (3) of the turbine engine (1), the system (8) comprising an exhaust gas temperature sensor (6), the system (8) further comprising a computer (7) configured to implement a method (E) according to one of claims 1 to 8.

10. Turbine engine comprising a system (8) according to claim 9.

Technical Field

The invention relates to a method for detecting ignition of a turbine engine.

More specifically, the present disclosure relates to a method of detecting ignition of a combustion chamber of a turbine engine based on exhaust gas temperature. In particular, the present invention relates to determining, in real time, an ignition status of a combustion chamber, the ignition status corresponding to success or failure of an attempted ignition.

Background

Known turbine engines, such as aircraft turbojet engines, typically include a compressor, a combustor, and a turbine, arranged in that order from upstream to downstream with respect to the direction of air flow in the turbine engine. The combustor takes in air previously compressed by the compressor, injects fuel into the combustor, and combusts the mixture, which is then expanded through a turbine and ultimately discharged, thereby generating thrust required for movement of the aircraft, for example. A portion of the energy generated within the combustion chamber is further extracted by the turbine to drive the compressor into rotation.

Detection of ignition in the combustion chamber is a basic criterion. In fact, this detection defines the measures to be taken for a successful start or for restarting the turbojet in flight.

After ignition in the combustion chamber, a flow of hot air, called "exhaust gas", flows through the turbine engine, which causes a rapid increase in the temperature downstream of the combustion chamber, in particular in the turbine and its casing. It is therefore known to place one or more Temperature sensors at this location in order to measure the evolution of the Exhaust Gas Temperature (EGT).

Known methods of detecting ignition in the combustion chamber include monitoring EGT measurements provided by these sensors. Furthermore, a plurality of control logics may be implemented in parallel, for example by a computer of the turbine engine.

Regardless of the presently known control logic, the turbine engine will be considered "fired" if the EGT has increased to a certain level since the start of fuel injection. For this purpose, the EGT value is first stored, for example, when a start-up procedure is initiated. After fuel injection and spark plug discharge (i.e., once ignition speed is reached), successive EGT measurements are compared to stored values. If a predetermined increase threshold is reached, ignition is detected in the combustion chamber. In this case, the ignition control law is changed to a control law that favours the acceleration or the cranking of the turbine engine until the idle speed is reached. On the other hand, if the EGT value does not exceed the threshold after a predetermined length of time, an alert is reported to the pilot indicating that the combustion chamber is not firing.

However, the EGT may have increased prior to start-up. This phenomenon occurs, for example, when a first ignition attempt has been carried out and then interrupted or after a rotating stall associated with an aerodynamic instability of the compressor. In this case, the EGT value is already too high when ignition of the combustion chamber has taken place and does not evolve in the same way as in the normal case. Thus, the expected EGT increase threshold cannot be reached in time, and in fact at the time of engine ignition, the start has been erroneously interrupted.

Therefore, there is a need for an improved method of detecting ignition in a combustion chamber, particularly where the exhaust gas temperature is above normal at the time of fuel injection and spark plug discharge.

Furthermore, publication FR3044703a1 is known, which relates to a method for determining the instant at which ignition occurs in a turbine engine combustion chamber during a successful ignition attempt. The method thus enables the determination of the duration of the ignition, which can be used as an indicator of the degradation of the overall starting system of the turbine engine, in particular if a deviation from a reference duration is determined. The logic of the method is applied to the data recorded during the flight mission and this data is generally processed after the mission to enable monitoring of the health state (health) of the turbine engine. It is still possible to implement real-time applications of the method, but especially a large amount of computational resources is required.

Disclosure of Invention

It is an object of the present invention to detect ignition of a turbine engine combustor in the event of restarting the turbine engine on the ground or in flight.

It is another object of the present invention to detect ignition of a turbine engine combustor after air rotating stall at the output of the compressor.

It is another object of the present invention to detect ignition of a turbine engine combustor in the event of a failure of an exhaust gas temperature probe.

It is another object of the present invention to detect ignition of a turbine engine combustor in the event of a premature interruption of fuel supply.

It is another object of the present invention to detect ignition of a turbine engine combustor during startup or flight with one or more channels of the computer reinitialized.

The invention proposes, in particular, a method for detecting the ignition of a combustion chamber of a turbine engine, comprising the steps of:

-receiving a first exhaust gas temperature measurement downstream of a combustion chamber before attempting to ignite said combustion chamber,

-receiving a temperature threshold value,

-receiving a secondary detection criterion,

-updating the received temperature threshold in accordance with the received secondary detection criterion,

-receiving a second exhaust gas temperature measurement after attempting to ignite the combustion chamber,

-comparing the updated temperature threshold with the difference between the first exhaust gas measurement and the second exhaust gas measurement, and

-determining, as a function of the result of the comparison step, an ignition status of the combustion chamber, which ignition status corresponds to the success or failure of the ignition attempt.

In such a method, the step of updating the temperature detection threshold advantageously enables to take into account the operating conditions of the turbine engine which affect the starting of the combustion chamber. In this way, the evolution of the exhaust gas temperature is not an absolute detection criterion, but rather with respect to the overall state of the turbine engine. Depending on secondary criteria (e.g. engine speed or compressor output pressure) related to the elements surrounding the combustion chamber, the evolution of the exhaust gas temperature differs from its nominal evolution even if the ignition regime is the same. Due to this type of method, the ignition-related warning reported to the pilot is no longer biased, thereby reducing the number of false identifications of ignition or re-ignition faults on the ground or in flight.

The method according to the invention may further comprise the following features taken alone or in combination:

-performing the step of receiving a first exhaust gas temperature measurement at the end of a start-up phase of a start-up procedure of the turbine engine when the exhaust gas temperature is minimal,

-the updating of the temperature threshold comprises: reducing the value of the threshold value if the secondary criterion is verified, otherwise retaining the value of the threshold value,

the secondary criterion is information relating to the evolution of the high-pressure compressor output pressure of the turbine engine, which secondary criterion is verified if a sudden increase in said pressure is detected,

the secondary criterion is information relating to the engine speed of the turbine engine, the secondary criterion being verified if the engine speed is within a determined interval for a predetermined duration;

the method according to the invention further comprises the steps of:

o receiving a first engine speed threshold value,

o receiving the engine speed measurement and,

o comparing the measurement with a first engine speed threshold, an

o determining an ignition state of the combustion chamber based on the result of the comparing step,

the method according to the invention further comprises the steps of:

o receives an acceleration threshold value for the engine speed,

o receiving a measurement time interval of the time interval,

o receiving a plurality of successive measurements of the acceleration of the engine speed during the measurement time interval received in the preceding step, an

o comparing a plurality of successive measured values with the received acceleration threshold, and

o determining the ignition state of the combustion chamber on the basis of the result of the comparison step, an

The method according to the invention further comprises the steps of:

o receiving a second engine speed threshold value,

o receiving information relating to the control state of the turbine engine,

o receiving the engine speed measurement and,

o comparing the measured value with a second engine speed threshold value, an

And o determining the ignition state of the combustion chamber based on the result of the comparing step and the information on the control state.

The invention further relates to a system for detecting ignition of a turbine engine combustor, the system comprising an exhaust gas temperature sensor, the system further comprising a computer configured to implement the method as described above.

Finally, the invention is applicable to a turbine engine comprising a system as described previously.

Drawings

Other characteristics, objects and advantages of the invention will appear upon reading the following detailed description and upon reference to the drawings, given by way of non-limiting example, in which:

FIG. 1 schematically shows a known turbine engine comprising an exemplary embodiment of a detection system according to the invention,

figures 2a to 2d schematically show different embodiments of the detection method according to the invention, and

fig. 3 illustrates the variation of the engine speed, the exhaust gas temperature and the high-pressure compressor output pressure during the starting of the turbine engine.

Detailed Description

Exemplary embodiments of the method according to the present invention will now be described with reference to the accompanying drawings.

With reference to fig. 1, a known turbine engine 1, for example an aircraft turbojet, comprises a combustion chamber 3, the ignition of which combustion chamber 3 is essential for starting or restarting the turbine engine 1 on the ground or in flight. The combustion chamber 3 is located between a compressor 2 section, called high-pressure compressor, and a turbine 4 section, called high-pressure turbine, which compressor 2 section and turbine 4 section are connected to each other by a rotating shaft 5, called high-pressure shaft. In a known manner, each of the compressor 2 section and the turbine 4 section may also comprise a plurality of high-pressure stages, respectively 21, 23 and 41, 43, respectively, which are mutually connected by the same rotary shaft 5, respectively, to form the high-pressure body of the two-body turbine engine. In operation, air is compressed by the compressor section 2 and then circulated in the combustion chamber 3, and fuel in the combustion chamber 3 is injected by means of an injector 31, which is configured for this purpose. The fuel-air mixture is then ignited by the action of the spark plug 33. The products of this combustion, referred to as exhaust gases, are then expanded through the turbine 4 section. Hereafter, the air circulating through the section of the turbine 4 downstream of the combustion chamber will be referred to as "exhaust gas" whether it has been previously ignited (at nominal operation) in the combustion chamber 3 or not ignited (at the start of the start-up sequence) in the combustion chamber 3. The exhaust gas temperature is therefore directly linked to the temperature of the turbine 4 section and indicates the ignition state of the combustion chamber 3.

Generally, the starting procedure of the turbine engine 1 comprises a first phase during which the rotation speed of the rotating shaft 5 is increased in a time range from the reception of a command to start the turbine engine 1, for example issued by a pilot, to the start of the injection of fuel into the combustion chamber 3 of the turbine engine 1. During this first phase, the drive of the rotating shaft 5 is effected independently of the action of the turbine 4 section, for example by means of a starter (not shown). This first phase may then be referred to as a pre-injection phase.

Furthermore, the starting sequence comprises a second phase following the first phase, which ends when the rotating shaft is no longer driven independently of the action of the turbine 4 section, for example by disengaging the starter clutch. This second phase begins with ignition of the fuel-air mixture in combustion chamber 3 and may be referred to as a post-ignition phase.

The start phase of the starting procedure then preferably corresponds to the first phase described above (referred to as the pre-injection phase), which lasts until the instant at which ignition of the fuel-air mixture begins.

When starting on the ground, the turbine engine 1 is first driven by a starter (not shown) in order to compress enough air to enable successful ignition of the combustion chamber 3. During the in-flight restart, ignition of the combustion chamber 3 can be carried out directly if the engine speed of the turbine engine is sufficiently high (a phenomenon known as "spinning"). Otherwise, it is necessary to rely on the starter as during start on the ground. As during the restart, during the start, ignition of the combustion chamber 3 is carried out by injection of fuel and discharge of the ignition plug 33.

In a first embodiment, with reference to fig. 2a, the detection method E comprises a first step E11 of receiving a first exhaust gas temperature measurement MT1 (referred to as initial measurement) before attempting to ignite the combustion chamber 3. This initial measurement MT1 is used as a reference for controlling the ignition of the combustion chamber 3. This is why the initial measurement value MT1 is typically stored for later comparison, as will be described more precisely.

Advantageously, the step of receiving the E11 initial measurement MT1 is carried out at the end of the starting phase of the starting procedure of the turbine engine 1. In fact, it is important to store the minimum value of the exhaust gas temperature reached after the start or restart procedure of starting the turbine engine 1. Therefore, when the routine is started, the driving of the rotary shaft 5 by the spinning starter causes air to circulate through the turbine engine 1, thereby ventilating and gradually cooling the turbine engine, as shown in fig. 3. Therefore, the exhaust gas temperature falls during the initial start-up procedure of the turbine engine 1, so that the exhaust gas temperature is minimized when the ignition of the combustion chamber 3 is carried out. This enables to benefit from an initial reference measurement which is an accurate map of the temperature of the combustion chamber 3 at its ignition.

During a second step E12, the temperature threshold ST is received and also stored according to an advantageous embodiment. This temperature threshold ST provides a reference for the rise in the exhaust gas temperature at which the combustion chamber 3 is considered to be ignited. The temperature threshold ST depends on the turbine engine 1, the state of wear of the turbine engine and the operating conditions of the turbine engine. Advantageously, the temperature threshold ST is 35K.

During a third step E13, according to an advantageous embodiment, the secondary detection criterion CS is received and stored. The secondary detection criterion CS enables the state of the turbine engine 1 to be determined which has an effect on the exhaust gas temperature. Thus, according to the verification criterion, the secondary detection criterion CS enables the value of the temperature threshold ST received in the preceding step E12 to be corrected in order to improve the detection of the ignition of the combustion chamber 3 on the basis of the evolution of the exhaust gas temperature.

During a fourth step E14, the temperature threshold ST is updated according to the secondary detection criterion CS received in the previous step E13. Advantageously, the updating comprises: -if the secondary criterion CS is verified, reducing the value of said threshold ST; otherwise, the value of the threshold ST is retained. In fact, when the secondary criterion CS is verified, the turbine engine 1 is potentially in an operating state such that the exhaust gas temperature does not follow the nominal evolution, for example the exhaust gas temperature has increased independently of the ignition of the combustion chamber 3. This may be due to the first start-up being aborted or due to a separation phenomenon occurring downstream of the compressor section 2. Therefore, the reduction in the value of the threshold ST enables the ignition of the combustion chamber 3 to be detected more quickly, or even ensures the detection of the ignition of the combustion chamber 3. Also, the updating of the value of the temperature threshold ST depends on the turbine engine 1, the state of wear of the turbine engine and the operating conditions of the turbine engine. Advantageously, when the secondary criterion CS is verified, the updated temperature threshold ST has a value of 15K.

During a fifth step E15, a second temperature measurement MT2 of the exhaust gases is received after an attempt is made to ignite the combustion chamber 3 and, according to an advantageous embodiment, is stored. Advantageously, a plurality of second temperature measurements MT2 of the exhaust gases is received continuously.

During a sixth step E16, the updated temperature threshold ST is compared with the difference between the second measurement MT2 and the first measurement MT1 of the exhaust gas temperature received in the previous step. Advantageously, the sixth step E16 is repeated consecutively with a plurality of second measurements MT 2.

During the final step E17, the ignition status of the combustion chamber is determined according to the results of the previous comparison step E16. More precisely, the combustion chamber 3 is considered to be ignited if the difference between the second temperature measurement MT2 and the first temperature measurement MT1 of the exhaust gases is greater than the updated temperature threshold ST. Otherwise, the combustion chamber is deemed to have extinguished. Thus, the state E17 enables the success or failure of an attempted ignition of the combustion chamber 3 of the turbine engine 1 to be inferred. If necessary, if a misfire condition is determined after a number of ignition attempts, an alert may be reported to the pilot.

The steps of the first embodiment of the detection method E may be carried out in a different order than that previously described. In particular, the step E12 of receiving the temperature threshold ST may be implemented independently of the steps of receiving the E11 first temperature measurement MT1 and of receiving the E15 second measurement MT 2. Thus, the temperature threshold ST can only be stored once during the first start of the turbine engine 1 and then used systematically each time the detection method E is implemented. Likewise, the receiving E13 and updating E14 of the secondary detection criterion CS may be carried out before the receiving E11 of the first temperature measurement MT 1.

Different secondary detection criteria CS may be used in the ignition detection method E.

The first secondary standard CS1 is information relating to the evolution of the pressure at the output of the high-pressure stage 23 furthest downstream of the compressor 2 section of the turbine engine 1 with respect to the air flow direction. In fact, as shown in fig. 3, during the ignition of the combustion chamber 3, a sudden increase of this pressure occurs, which is also referred to as pressure jump. Thus, if the pressure jump is detected, the first secondary criterion CS1 will be verified. For this purpose, continuous pressure measurements are received at the output of the high-pressure stage 23 of the compressor 2 stage. Advantageously, these measurements are spaced apart by approximately several hundred milliseconds, for example 200 milliseconds. If the deviation between successive measured values is greater than some predetermined pressure threshold, a pressure jump is detected and the received secondary criterion CS1 is verified. Advantageously, the pressure jump threshold is derived from engine tests and is for example equal to 4% of the initial pressure measurement taken. In any case, the first secondary criterion CS1 is only used when the turbine engine 1 is being started, in order to limit the effect of this first secondary criterion.

The second secondary criterion CS2 is information relating to the rotational speed of the turbine engine 1, that is, the rotational speed of the rotating shaft 5 (e.g., the high-pressure shaft 5).

This second criterion CS2 is advantageously used in the event of the turbine engine 1 stalling at a speed below its idle speed threshold SR but greater than the maximum ventilation threshold SMV. In practice, when the engine speed crosses the maximum ventilation threshold SMV (for example 7000 revolutions per minute), the combustion chamber 3 is considered to be ignited. In practice, this threshold SMV is determined as a maximum threshold which can only be crossed when torque is supplied by the starter. In the absence of torque produced by combustion, the threshold SMV cannot be crossed, so crossing of the threshold indicates that the combustion chamber 3 is ignited. The threshold value SMV may be determined experimentally by means of a ventilation test on the turbine engine 1.

In order to verify this second criterion CS2, an operating state of the turbine engine 1 corresponding to a standstill of the engine speed is first determined. For this purpose, the evolution of the engine speed is monitored over a predetermined duration T1, typically 20 seconds. The duration T1 corresponds to a maximum time threshold during which the engine speed may be maintained above the maximum ventilation threshold SMV without starting. The duration T1 also depends on the type of turbine engine 1 and the degree of wear of the turbine engine 1. Thus, if during the duration corresponding to the previously described maximum time threshold T1, the engine speed is within the following interval: this interval extends between the maximum ventilation threshold SMV for the engine speed and the idle threshold SR for the engine speed, the second secondary criterion CS2 is verified.

Advantageously, the second secondary criterion CS2 is considered verified if the following condition is also satisfied during the duration corresponding to the maximum time threshold T1 described previously:

the fuel supply valve in the combustion chamber 3 is opened, and

-the starter has been switched off.

The conditions relating to the opening of the supply valve are such that there is no risk of detecting a false ignition when no fuel is supplied to the combustion chamber 3. This additional validity condition makes it impossible to update the temperature threshold ST in the case where the supply valve has been closed, i.e. when the start is cancelled or the task is terminated. Thus, the logic of the detection method is reset to zero for subsequent re-ignition and/or restart.

In a second embodiment, referring to fig. 2b, detection method E comprises a first step E21 of receiving a first engine speed threshold SRM1 (e.g. the previously described maximum ventilation threshold SMV) of 7000 revolutions per minute.

During a second step E22, according to an advantageous embodiment, an engine speed measurement MRM is received and stored.

During a third step E23, the received measurement MRM is compared with a first engine speed threshold SRM 1.

During a fourth step E24, the ignition conditions of the combustion chamber 3 are determined according to the result of the comparison step E23. In this particular case, if the received measured value MRM is greater than the first engine speed threshold SMR1, combustion chamber 3 is considered to be ignited. Indeed, as previously mentioned, the first engine speed threshold SMR1 is advantageously selected to indicate the following limits: beyond this limit, the turbine engine 1 cannot operate without energy originating from the ignited combustion chamber 3. If the received measurement MRM is less than the first threshold engine speed SMR1, the combustion chamber is deemed to be misfired. If necessary, if a misfire condition is determined after a number of firing attempts, an alert may be reported to the pilot.

The steps of the second embodiment of the detection method E may be carried out in a different order than that previously described. In particular, step E21 of receiving the first engine speed threshold SRM1 may be implemented independently of step E22 of receiving the engine speed measurement MRM. Therefore, this first engine speed threshold value can only be stored once at the first start of the turbine engine 1 and then used systematically each time the detection method E is implemented.

Referring to fig. 2c, in a third embodiment of the detection method E, complementary to the previously described second embodiment, a step E31 of receiving the engine speed acceleration threshold SA is implemented. This acceleration threshold SA corresponds to an increase in the engine speed, so that the engine must be driven by combustion performed in the combustion chamber 3. The acceleration threshold SA depends on the type of turbine engine 1 and on the degree of wear of the turbine engine and corresponds to an acceleration of the engine speed of, for example, 50 revolutions per minute per second.

Then, E32 a measurement time interval IT is also received and, according to an advantageous embodiment, stored. This time interval IT is, for example, one second.

Thereafter, during the time interval IT received in the preceding step E32, a measurement MA of the acceleration of a plurality of successive engine speeds is received E33, which is then compared with the acceleration threshold SA received during the first step E31.

Finally, the ignition condition E34 of the combustion chamber 3 is determined as a function of the result of the comparison step E33. In this case, the combustion chamber 3 is considered to be ignited if all the acceleration measurements MA are greater than or equal to the received threshold SA. Otherwise, if a flameout condition is determined after a number of ignition attempts, the combustion chamber 3 is considered to have flameout and an alert may be reported to the pilot.

Such an embodiment of the detection method E is advantageously implemented to detect the re-ignition of the combustion chamber 3 when the turbine engine 1 is operating above its maximum ventilation threshold SMV and below its idle threshold SR.

The steps of the third embodiment of the detection method may be carried out in a different order than that previously described. In particular, step E31 of receiving the acceleration threshold SA and step E32 of receiving the time interval IT may be carried out independently of step E33 of receiving the measured values MA of the acceleration of the engine speeds. Thus, the acceleration threshold SA and the time interval threshold IT can only be stored once at the first start of the turbine engine 1 and then used systematically each time the detection method E is implemented.

With reference to fig. 1, the turbojet 1 comprises one (or more) computers 7 configured to control the operation of the turbojet 1. This type of computer 7 typically comprises two channels that perform the same operations in parallel based on the same data received. Typically, one of the two channels is subordinate to the other. The terms active channel and passive channel are used. The channel of the computer 7 receives measurements originating from one (or more) exhaust gas temperature sensors 6 located downstream of the combustion chamber 3. This type of sensor 6 or probe therefore measures the exhaust gas temperature. The turbine engine ignition detection system 8 therefore advantageously comprises a temperature sensor 6 and a computer 7 of this type.

Furthermore, the computer 7 is configured to implement an ignition detection method E according to one of the previously described embodiments. To this end, the computer 7 also receives other measurements originating from other sensors, for example measurements relating to the engine speed or to the pressure downstream of the section of the compressor 2 (for example the high-pressure stage 23).

In case of a failure of the passive channel, the computer 7 can be automatically reinitialized, thus losing all ongoing calculations. As a result, the system 8 must be able to reinitialize itself to the ignition conditions of the combustion chamber 3. To this end, the passive channel may first restore the ignition condition by communicating with the active channel.

Alternatively, with reference to fig. 2d, a fourth embodiment of the ignition detection method E of the combustion chamber 3 is implemented by a passive channel of the computer 7.

During a first step E41, a second engine speed threshold SRM2 is received and, according to an advantageous embodiment, stored. The second engine speed threshold SRM2 corresponds, for example, to the idle speed threshold SR previously described.

During a second step E42, information relating to the control state of the turbine engine is also received, for example pilot commands relating to the power demand.

During a third step E43, a measured value MRM of the engine speed is received and, according to an advantageous embodiment, stored.

During a fourth step E44, the measured value MRM received in the previous step is compared with a second engine speed threshold SRM 2.

During a fifth step E45, the ignition conditions of the combustion chamber 3 are determined as a function of the result of the comparison step E44 and the information relating to the control state. In this particular case, if the engine speed is greater than or equal to its idle speed SR (i.e. the turbine engine 1 is in a steady state of operation above its idle speed) and the pilot commands are consistent, the passive channel will conclude that the combustion chamber 3 is ignited and is controlled by the active channel.

The steps of the fourth embodiment of the detection method E may be carried out in a different order than that previously described. In particular, step E21 of receiving the second engine speed threshold SRM2 may be implemented independently of step E22 of receiving an engine speed measurement. Therefore, the second engine speed threshold value SRM2 can only be stored once when the turbine engine 1 is started for the first time and then used systematically each time the detection method E is implemented. Likewise, the steps E42 of receiving information relating to the control state and E43 of receiving the engine speed measurement MRM may be interchanged.

Advantageously, the different embodiments described previously are implemented in parallel, independently of each other, by the channels of the computer 7. In this way, the first control logic, which detects the ignition of the combustion chamber 3, is able to send corresponding information to the pilot.

Thus, if the turbine engine 1 is started when the exhaust gas temperature probe 6 fails, the logic corresponding to the second embodiment of the detection method E will still be able to detect the ignition of the combustion chamber 3.

In the same way, if one or more channels of the computer 7 are reinitialized in a premature manner, the corresponding logic of the fourth embodiment of the detection method E will still be able to detect the ignition of the combustion chamber 3.

Also, if the engine speed is stalled after the fuel supply is shut off during the start of the turbine engine 1, the determination logic implemented will depend on the level of the stalled speed (start on the ground or restart in flight).

The first embodiment of the detection method E will allow the detection of the ignition of the combustion chamber 3 by means of the first pressure jump secondary criterion CS1 if the engine speed is less than the maximum ventilation threshold SMV.

The second embodiment of the detection method E will allow to detect the ignition of the combustion chamber 3 by using the maximum ventilation threshold SMV as the first engine speed threshold SRM1, if the engine speed is transient.

The first embodiment of the detection method E will allow the ignition of the combustion chamber 3 to be detected by means of the second engine speed secondary criterion CS2 if the engine speed is greater than the maximum ventilation threshold SMV.