Patent Publication Number: US-11639668-B2

Title: Method and control unit for controlling the play of a high-pressure turbine

Description:
BACKGROUND OF THE INVENTION 
     The present invention relates to the general field of turbomachines for gas turbine aeronautical engines. It more precisely concerns the control of the clearance between, on the one hand, the moving blade tips of a turbine rotor and, on the other hand, a turbine shroud of an outer casing surrounding the blades. 
     The clearance existing between the blade tips of a turbine and the shroud that surrounds them is dependent on the differences in dimensional variation between the rotating parts (disc and blades forming the turbine rotor) and the fixed parts (outer casing including the turbine shroud it comprises). These dimensional variations are both of thermal origin (related to the temperature variations of the blades, the disc and the casing) and of mechanical origin (in particular related to the effect of the centrifugal force exerted on the turbine rotor). 
     To increase the performance of a turbine, it is desirable to minimize the clearance as much as possible. Additionally, when there is an increase in rating, for example when passing from a ground idle rating to a take-off rating in a turbomachine for an aeronautical engine, the centrifugal force exerted on the turbine rotor tends to bring the blade tips closer to the turbine shroud before the turbine shroud has had time to expand from the effect of the temperature increase related to the increase in rating. There is therefore a risk of contact at this operating point known as the pinch point. 
     It is known to employ an active control system to control the clearance of the blade tips of a turbomachine turbine. A system of this type generally operates by directing air bled off, for example at the level of a compressor and/or the turbomachine fan, onto the outer surface of the turbine shroud. Cool air sent onto the outer surface of the turbine shroud has the effect of cooling the latter and thus limiting its thermal expansion. The clearance is therefore minimized. Conversely, hot air promotes the thermal expansion of the turbine shroud, which increases the clearance and makes it possible for example to avoid contact at the aforementioned pinch point. 
     An active control of this kind is operated by a control unit, for example by the full authority regulation system (or FADEC) of the turbomachine. Typically, the control unit acts on a controlled-position valve to control the flow rate and/or temperature of the air directed onto the turbine shroud, as a function of a clearance setpoint and an estimate of the actual blade tip clearance. 
     The turbomachine also has an operating limit temperature. The operating limit temperature of the engine is defined with respect to a limit temperature of the combustion gas determined downstream of its combustion chamber, for example deduced from at least one measurement made within the high-pressure or low-pressure turbine of the engine. This temperature is commonly referred to as the “Red Line EGT”. The Red Line EGT is identified during tests carried out on the ground (Block Tests) by the manufacturer, then communicated thereby. In other words, the Red Line EGT is the maximum value declared by the manufacturer, this value being certified according to the engine lifecycle (e.g. new or reconditioned engine). Once this limit is reached the engine is sent off for maintenance in order to restore a positive EGT margin. The term “EGT margin” is understood to mean the difference between the Red Line EGT certified by the manufacturer and a combustion gas temperature determined downstream of the combustion chamber of the engine. 
     The combustion gas temperature downstream of the combustion chamber of the engine is generally at a maximum during a phase of rapid acceleration, given the thermal response of the engine. Typically, approximately 60 seconds after an acceleration phase, the clearance between the blades of the rotor of the high-pressure turbine and the shroud surrounding them increases. The increase in this clearance manifests as an increase in the combustion gas temperature. Downstream of the combustion chamber, by way of example at the outlet of the high-pressure turbine, temperatures are measured in the order of 20 to 30K greater than a temperature of the engine in stabilized rating, the stabilized rating being obtained after a given time interval following the acceleration phase of the engine. 
     The temperature difference between the maximum combustion gas temperature determined during a phase of acceleration of the turbomachine and the temperature of its stabilized regime determined after this acceleration phase is currently referred to as the “Overshoot”. 
     In practice, the more the engine ages, the more the maximum combustion gas temperature increases. The maximum combustion gas temperature therefore tends to approach the operating limit temperature of the engine (Red Line EGT) as the latter ages. This temperature degradation is generally justified, at least in part, by a degradation of the high-pressure turbine manifesting as an increase in its clearance. 
     In this context, taking into account the aging of the engine, it would be beneficial to keep a positive EGT margin for as long as possible in order to postpone sending the engine off for maintenance. 
     During an acceleration phase, the optimization of the clearance between the blades of the rotor of the high-pressure turbine and the shroud surrounding them can make it possible to reduce the Overshoot, and therefore the maximum combustion gas temperature. However, such an optimization can pose a risk of premature wear to the high-pressure turbine. By way of example, too great a reduction of the Overshoot related to a prolonged reduction of the clearance of the high-pressure turbine for a new, hot engine, or an engine that already has minimized clearance of its high-pressure turbine, can result in a pinch point between the blades and the shroud of the high-pressure turbine. Thus, the limitation of an Overshoot during a phase/transient state of the engine can pose a risk of permanent degradation of the blades of the high-pressure turbine, thus affecting the overall performance of the engine and its fuel consumption. 
     It would therefore be desirable to minimize the temperature Overshoot of the high-pressure turbine during a variation in the engine rating, while eliminating any risk of degradation of the blades of the high-pressure turbine. 
     SUBJECT AND SUMMARY OF THE INVENTION 
     The aim of the present invention is to remedy the aforementioned drawbacks. 
     For this purpose, the invention proposes a method for controlling the clearance between, on the one hand, the blade tips of a rotor of a high-pressure turbine of a gas turbine aircraft engine and, on the other hand, a turbine shroud of a casing surrounding said blades of the high-pressure turbine, the method comprising the controlling of a valve delivering a stream of directed air to said turbine shroud, this method being characterized in that it comprises the following steps:
         the detection of a transient acceleration phase of the engine on the basis of at least one parameter representative of the engine;   the receiving of an item of data representative of the gas temperature at the outlet of the combustion chamber of the engine;   a valve opening command, to deliver said air stream to the turbine shroud or to increase the flow rate of said delivered air stream, if the transient acceleration phase is detected and if the gas temperature at the outlet of the combustion chamber of the engine is greater than a first temperature threshold corresponding to a degraded clearance characteristic of an aged engine, the first temperature threshold being less than an operating limit temperature of the engine.       

     Advantageously, the method above makes it possible to adapt the control of clearance during an acceleration phase of the engine, while taking into account the residual margin existing between the operating limit temperature of the engine and the combustion gas temperature at the outlet of the combustion chamber of the engine. As explained previously, as the engine ages, the maximum combustion gas temperature of the engine increases and tends to approach the operating limit temperature of the engine (Red Line EGT). In other words, the EGT margin tends to decrease when the engine ages. The taking into account of the separation between the operating limit of the engine and the combustion gas temperature of the engine, via the first temperature threshold, therefore makes it possible to take into account the aging of the engine. Thus, the clearance setpoint of the high-pressure turbine is adapted as a function of the aging of the engine. Subsequently, the adaptation of this clearance setpoint itself influences the variation in the combustion gas temperature at the outlet of the combustion chamber of the engine, thus making it possible to reduce the Overshoot. The clearance of the high-pressure turbine as well as the Overshoot are therefore regulated in a closed loop and adaptively as a function of the aging of the engine. This method is applicable throughout the engine lifecycle. Typically an aged engine has greater clearance in its high-pressure turbine than a new engine. As a function of the aging of the engine, the method described above then makes it possible to minimize the clearance of its high-pressure turbine, via control of the valve, without risking damage to the turbine blades. The performance of the turbomachine is thus optimized throughout its lifecycle. This therefore extends the time over which a positive EGT margin is kept for the engine, which makes it possible to increase the life of the engine and postpone its being sent off for maintenance. 
     Preferably, in this method a higher percentage of valve opening is commanded if the combustion gas temperature temporarily exceeds the first temperature threshold. 
     In an exemplary embodiment of this method, said at least one parameter representative of the engine is the engine rating and the detection of a transient acceleration phase of the engine comprises the continuous determination of the engine rating and the determination of a variation in the engine rating for a predetermined time interval, the transient acceleration phase of the engine being detected during said predetermined time interval if the variation in the engine rating is greater than or equal to a variation threshold characterizing a transient acceleration phase of the engine. 
     In an exemplary embodiment, said at least one parameter representative of the engine is chosen from among: the rating of a low-pressure turbine of the engine, the rating of the high-pressure turbine, the angular position of an aircraft throttle lever and the item of data representative of the gas temperature at the outlet of the combustion chamber of the engine. 
     In an exemplary embodiment of this method, the valve is a valve of on-off type configured to switch between an open state and a closed state, the method further comprising, following the opening of the valve, a command to close the valve when the gas temperature at the outlet of the combustion chamber of the engine is less than a second temperature threshold, the second temperature threshold being less than the first temperature threshold. 
     In another exemplary embodiment of this method, the valve is a controlled-position valve, the method comprising a command to gradually open the valve as a function of a predefined control law taking into account a separation between the gas temperature at the outlet of the combustion chamber of the engine and the first temperature threshold. 
     In an exemplary embodiment of this method, the item of data representative of the gas temperature at the outlet of the combustion chamber is a temperature measurement taken at the level of the high-pressure turbine. 
     The invention also proposes, according to another aspect, a control unit for controlling the clearance between, on the one hand, a number of blade tips of a rotor of a high-pressure turbine of a gas turbine aircraft engine, and, on the other hand, a turbine shroud of a casing surrounding said blades of the high-pressure turbine, the control unit comprising means for controlling a valve, the valve being configured to deliver a stream of air to said shroud of the turbine, the control unit being characterized in that it comprises:
         detection means configured to detect a transient acceleration phase of the engine on the basis of at least one parameter representative of the engine;   receiving means configured to receive an item of data representative of the gas temperature at the outlet of the combustion chamber of the engine;   the control means being configured to command the opening of the valve to deliver said air stream to the turbine shroud, or to control an increase in the flow rate of said stream of delivered air, if the transient acceleration phase is detected and if the gas temperature at the outlet of the combustion chamber of the engine is greater than a first temperature threshold corresponding to a degraded clearance characteristic of an aged engine, the first temperature threshold being less than an operating limit temperature of the engine.       

     Preferably, the control means are furthermore configured to command a greater percentage of opening of the valve if the combustion gas temperature temporarily exceeds the first temperature threshold. 
     Advantageously, to judge the state of aging of the engine, the control unit counts a trigger number to trigger the additional valve opening command. 
     In an exemplary embodiment, in this control unit, said at least one parameter representative of the engine is the engine rating and the detection means are configured to:
         continuously determine the engine rating;   determine a variation in the engine rating for a predetermined time interval;   detect the transient acceleration phase of the engine during said predetermined time interval if the variation in the engine rating is greater than or equal to a variation threshold characterizing a transient acceleration phase of the engine.       

     In an exemplary embodiment, in this control unit, the valve is a valve of on-off type configured to switch between an open state and a closed state, the control means being configured to command, following the opening of the valve, the closing of the valve when the gas temperature at the outlet of the combustion chamber of the engine is less than a second temperature threshold, the second temperature threshold being less than the first temperature threshold. 
     In another exemplary embodiment, in this control unit, the valve is a controlled-position valve, the control means being configured to command the gradual opening of the valve as a function of a predefined control law taking into account a separation between the gas temperature at the outlet of the combustion chamber of the engine and the first temperature threshold. 
     The invention also proposes, according to another aspect, a gas turbine aircraft engine comprising the control unit summarized above and at least one valve for acting on an air stream directed toward the turbine shroud and wherein the valve is controlled by the control means. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Other features and advantages of the invention will become apparent from the following description of particular embodiments of the invention, given by way of non-limiting example, with reference to the appended drawings, wherein: 
         FIG.  1    is a schematic and longitudinal section view of a part of a gas turbine aircraft engine according to an embodiment of the invention; 
         FIG.  2    is a magnified view of the engine of  FIG.  1    in particular showing the high-pressure turbine of the engine; 
         FIG.  3    is a functional diagram of a module for controlling a valve making it possible to control the blade tip clearance in the engine of  FIG.  1    according to a first embodiment; 
         FIG.  4    is a functional diagram of a module for controlling a valve making it possible to control the blade tip clearance in the engine of  FIG.  1    according to a second embodiment. 
     
    
    
     DETAILED DESCRIPTION OF EMBODIMENTS 
       FIG.  1    schematically represents a jet engine  10  of double-flow, twin-spool type to which the invention in particular applies. Of course, the invention is not limited to this particular type of gas turbine aircraft engine. 
     In a well-known manner, the jet engine  10  of longitudinal axis X-X particularly comprises a fan  12  which delivers a stream of air in a primary stream flow duct  14  and in a secondary stream flow duct  16  coaxial with the primary stream duct. From upstream to downstream in the direction of flow of the gas stream passing through it, the primary stream flow duct  14  comprises a low-pressure compressor  18 , a high-pressure compressor  20 , a combustion chamber  22 , a high-pressure turbine  24  and a low-pressure turbine  26 . 
     As shown more precisely by  FIG.  2   , the high-pressure turbine  24  of the jet engine comprises a rotor formed by a disc  28  on which are mounted a plurality of blades  30  disposed in the primary stream flow duct  14 . The rotor is surrounded by a turbine casing  32  comprising a turbine shroud  34  carried by an outer turbine casing  36  by way of attachment spacers  37 . 
     The turbine shroud  34  can be formed by a plurality of adjacent sections or segments. On the inner side, it is provided with a layer  34   a  of abradable material and surrounds the blades  30  of the rotor, leaving a clearance  38  between itself and the tips  30   a  of the blades. 
     In accordance with the invention, provision is made for a system making it possible to control the clearance  38  by modifying, in a controlled manner, the inner diameter of the outer turbine casing  36 . For this purpose, a control unit  50  controls the flow rate and/or the temperature of the air directed toward the outer turbine casing  36 . The control unit  50  is for example the full authority regulation system (or FADEC) of the jet engine  10 . 
     In the example shown, a control box  40  is disposed around the outer turbine casing  36 . This box receives cool air by means of an air conduit  42  opening at its upstream end into the flow duct of the primary stream at one of the stages of the high-pressure compressor  20  (for example by means of a scoop known perse and not shown in the figures). The cool air circulating in the air conduct is discharged onto the outer turbine casing  36  (for example using multiple perforations on the walls of the control box  40 ) causing it to cool and its inside diameter to thus be reduced. 
     As shown in  FIG.  1   , a valve  44  is disposed in the air conduit  42 . This valve  44  is controlled by the control unit  50 . 
     In a first exemplary embodiment, the valve  44  can be an on-off valve able to switch between an open state and a closed state. The use of such a valve is advantageous, particularly in terms of cost, bulk, reliability and power necessary for control. 
     It will be understood that by controlling the valve  44  to act, on the one hand, on the opening frequency and on the other hand, on the cyclic opening/closing ratio of the valve, it is possible to obtain a variation in the average flow rate of the air directed toward the casing. Different architectures of on-off valve are well-known to those skilled in the art and will therefore not be described here. Preferably, an electrically controlled valve would be chosen control which would remain in the closed position in the absence of an electrical power supply (thus guaranteeing that the valve remains closed in the event of a control fault). 
     In a second exemplary embodiment, the valve  44  can be a controlled-position valve. The position of the valve  44  can be between 0%, corresponding to a closed valve, and 100%, corresponding to an open valve. When the valve  44  is open (position at 100%), the cool air is conveyed toward the outer turbine casing  36 , which results in the thermal contraction of the latter and therefore a reduction in the clearance  38 . When, on the contrary, the valve  44  is closed (position at 0%), the cool air is not conveyed toward the outer turbine casing  36  which is therefore heated by the primary stream. This results either in the thermal expansion of the casing  1  and an increase in the clearance  38 , or at least the controlled limitation (or stopping) of the expansion of the casing  1  and the control of the clearance  38 . In the intermediate positions, the outer turbine casing  36  contracts or expands and the clearance  38  increases or decreases, to a lesser extent. As will be seen later, control of the clearance  38  is used in such a way as to keep a positive EGT margin, thus making it possible to extend the lifetime of the jet engine  10 . 
     Of course, the invention is not limited to these two examples. Thus, another example can consist in bleeding off air at two different stages of the compressor and controlling valves  44  to modulate the flow rate of each of these bleed-offs to regulate the temperature of the mixture to be directed onto the outer turbine casing  36 . 
     We will now describe the controlling of the valve  44  by the control unit  50 . 
     In accordance with the invention, the control unit  50  comprises:
         detection means  51  configured to detect a transient acceleration phase of the jet engine  10  over a predetermined time interval;   receiving means  52  configured to receive at least one item of data representative of the temperature of the combustion gases coming from the combustion chamber  22  of the jet engine  10 ;   control means  53  configured to control the valve  44 .       

     The detection means  51 , the receiving means  52  and the control means  53  together form a module for controlling the valve  44  incorporated into the control unit  50 . This control module corresponds for example to a computer program executed by the control unit  50 , to an electronic circuit of the control unit  50  (for example of programmable logic circuit type) or to a combination of an electronic circuit and a computer program. 
     The term “transient acceleration phase of the jet engine  10 ” is understood to mean a transition in rating related to an acceleration phase of the jet engine  10  occurring between two stabilized ratings of it. The transitional acceleration phase that one is seeking to detect using the detection means  51  can by way of example correspond to a transition between the ground idle rating and the stabilized flight rating, i.e. to the phase of acceleration between these two ratings. In another example, the transient acceleration phase can correspond to the phase of acceleration between any intermediate rating (e.g. half-throttle) and the flight rating. 
     The detection, where applicable, of a transient acceleration phase of the jet engine  10  can be done on the basis of one or more parameters representative of the jet engine  10 . 
     A parameter representative of the jet engine  10  is by way of example its rotation rating. The detection of a transient acceleration phase of the jet engine  10  is then done on the basis of a continuous determination of its rating. The detection of the variation in the rating of the jet engine  10  by the detection means  51  then makes it possible to identify a transient acceleration phase of the jet engine  10  over a predefined period, for example chosen between 1 second and 5 minutes. During this predetermined time interval, the detection means  51  can identify a transient acceleration phase by observing the variations in rating of the jet engine  10 . These variations are then compared to a setpoint characterizing a variation in rating of the jet engine  10 . Thus, if during the predetermined time interval the variation in the rotation rating of the jet engine  10  is greater than or equal to a variation threshold characterizing a transient acceleration phase of the jet engine  10 , the detection means  51  detect a transient acceleration phase. 
     In other examples, the determination of the rating of the jet engine  10 , as well as the detection of a transient acceleration phase of the jet engine  10  can be done on the basis of any parameter(s) representative of the engine. 
     By way of example, the determination of the rotation rating of the jet engine  10  as well as the detection of a transient acceleration phase thereof can be done on the basis of one or more of the following parameters: the rating of the high-pressure turbine  24 , the rating of the low-pressure turbine  26 , the angular position of the aircraft throttle lever, a measured or computed combustion gas temperature at the outlet of the combustion chamber  22 . 
     In parallel, the receiving means  52  receive at least one item of data representative of the combustion gas temperature at the outlet of the combustion chamber  22  of the jet engine  10 . The item of data representative of the combustion gas is by way of example a temperature measurement taken somewhere between the outlet of the combustion chamber  22  of the jet engine and the aircraft nozzle, for example at any point of the high-pressure turbine  24  or of the low-pressure turbine  26 . The receiving means  52  then obtain the temperature of the combustion gas in a known manner, directly on the basis of the representative item of data or indirectly by computation on the basis thereof. By way of example, the item of data representative of the gas temperature at the outlet of the combustion chamber  22  is a temperature measurement taken at the level of the high-pressure turbine  24 , i.e. taken in or at the outlet of the latter, allowing the receiving means  52  to access the gas temperature at the outlet of the combustion chamber  22 . 
     The configuration of the control means  53  depends on the type of valve  44  implemented as will be described in  FIGS.  3  and  4   . These figures respectively illustrate the method for controlling the valve  44 , of on-off and regulated position type respectively. 
     The steps  301 ,  401  and  302 ,  402  are similar in these figures. These steps correspond to a step  301 ,  401  of detecting a variation in the rating of the jet engine  10  by the detection means  51 , and to a step  302 ,  402  of receiving at least one item of data representative of the gas temperature at the outlet of the combustion chamber  22  of the engine by the receiving means  52 . It is understood that the order of the steps illustrated in these figures is given by way of illustration, these steps being able to be done in parallel in a non-illustrated example. 
     The control unit  50  is configured to identify from the detection means  51  and receiving means  52  any occurrence of a situation for which:
         a transient acceleration phase of the jet engine  10  is detected, and   the temperature of the combustion gas at the outlet of the combustion chamber ( 22 ) of the engine ( 10 ) is greater than a first temperature threshold T1.       

     The first temperature threshold T1 is chosen beforehand to be less than the Red Line EGT characterizing the operating limit temperature of the jet engine  10 , such as to keep a positive EGT margin (difference between the Red Line EGT and the combustion gas temperature) if the combustion gas temperature of the jet engine  10  reaches the temperature threshold T1. The temperature threshold T1 is by way of example defined to be lower by 1 to 10° C. than the Red Line EGT. This temperature threshold T1 thus constitutes a protection threshold of the Red Line EGT, the reaching of this threshold parallel to a detection of a transient acceleration phase of the jet engine  10  then manifesting as an Overshoot situation for an aged engine or an engine exhibiting degraded performance. 
     Moreover, the temperature threshold T1 is chosen with regard to the state of health of the jet engine  10 , the temperature value T1 only being meant to be reached by the combustion gas for an aged engine, for example exhibiting a degraded clearance  38 . Specifically, as explained previously, the more an engine ages, the more the maximum temperature of its combustion gas increases and tends to approach the Red Line EGT. Conversely, a jet engine which is new or just out of maintenance is not subject to the risk of the gas temperature at the outlet of the combustion chamber approaching the temperature T1, still less the Red Line EGT. The identification by the control unit  50  of a situation for which a transient acceleration phase of the jet engine  10  is detected and for which the combustion gas temperature is greater than the temperature threshold T1 can therefore only occur for an engine that is aged and/or exhibiting degraded performance. 
     After each step  301 ,  302 ,  401 ,  402  the control unit  50  attempts to detect (steps  303 ,  403 ) any occurrence of the aforementioned situation. The step  303  can, by way of example, be carried out by the control means  53  or by other dedicated detection means. 
     If the occurrence of such a situation is not identified, the control unit  50  deduces the non-occurrence of an Overshoot of the combustion gas temperature at the outlet of the combustion chamber  22  which might run the risk of approaching the Red Line EGT. The steps  301 ,  302 ,  401 ,  402  are then executed again. 
     Conversely, if the aforementioned situation is detected, the control unit  50  deduces a situation of Overshoot of the combustion gas temperature that potentially runs the risk of approaching the Red Line EGT. The control unit  50  then seeks to minimize the Overshoot by optimizing the clearance  38  of the high-pressure turbine  24 . Specifically, in the absence of optimization of the clearance  38 , an Overshoot situation for an aged or degraded engine would run the risk of reducing its EGT margin and therefore its lifetime before it is sent off for maintenance. The optimization of the clearance  38  then has the aim of keeping a positive EGT margin for as long as possible. 
     When the valve  44  is of on-off type ( FIG.  3   ) the control means  53  are then configured to command an opening (step  304 ) of the valve  44  such as to deliver a stream of air to the turbine shroud  34  and thus reduce the clearance  38  of the high-pressure turbine  24 . The reduction of the clearance  38  makes it possible to optimize the performance of the high-pressure turbine  24 , causing a reduction in the combustion gas temperature at the outlet of the combustion chamber  22 . The combustion gas temperature is then periodically compared (step  305 ) to a second temperature threshold T2 chosen as equal to or less than the first temperature threshold T1 to avoid oscillation effects. As long as the combustion gas temperature remains greater than the second temperature threshold T2, the valve  44  is kept open. When the combustion gas temperature is detected as less than the second temperature threshold T2, the control means  53  command (step  306 ) the closing of the valve  44 . 
     When the valve  44  is of regulated position type, the control means  53  are configured to control (step  404 ) the percentage of opening of the valve  44  as a function of the separation between the current combustion gas temperature and the first temperature threshold T1. In other words, the opening of the valve  44  is done gradually as a function of a control law previously stored in the control means  53 , this law taking into account the separation between the combustion gas temperature at the outlet of the combustion chamber  22  and the first temperature threshold T1. The control means  53  are by way of example configured to command a greater percentage of opening of the valve  44  (resulting from an over-setpoint value) and therefore an increase in the stream of air delivered to the turbine shroud  34 , if the combustion gas temperature temporarily exceeds the first temperature threshold T1. Thus, the clearance  38  of the high-pressure turbine  24  is once again optimized, subsequently causing the reduction of the combustion gas and therefore of the Overshoot. In other words, when the temperature threshold T1 is reached, a closing clearance over-setpoint value incurring an additional valve opening (of up to 200%) with respect to an open valve position (at 100%) is triggered. 
     Thus, the controlling of a valve  44  of on-off type or with regulated position as described above makes it possible to keep a positive EGT margin while reducing the combustion gas temperature. 
     The embodiments described above have the following advantages. The controlling of the clearance  38  of the high-pressure turbine  24  during an acceleration phase of the engine  10  takes into account the residual margin existing between the Red Line EGT and the combustion gas temperature at the outlet of the combustion chamber  22 . The taking into account of this margin is made possible by the comparison of the combustion gas temperature with the first temperature threshold T1, chosen with respect to the Red Line EGT as protection threshold. 
     As explained in the introduction, as the high-pressure turbine  24  ages, the maximum combustion gas temperature tends to gradually approach the Red Line EGT. The taking into account of the separation between the Red Line EGT and the combustion gas temperature, via the temperature T1, therefore makes it possible to take into account the aging of the engine  10  of the jet engine. The exceeding of the temperature T1 by the combustion gas in particular indicates the aging or degradation of the performance of the jet engine  10  requiring a reduction of its Overshoot in order to limit any risk of approaching the Red Line EGT. 
     The setpoint of the clearance  38  of the high-pressure turbine  24  is then adapted by the control means  53  as a function of the aging of the engine. The adapting of this clearance setpoint itself influences the variation of the combustion gas temperature of the combustion chamber  22  and makes it possible to reduce the Overshoot in the temperature of the reactor  10 . 
     In the same way, the trigger number of the over-setpoint value giving rise to a greater percentage of opening of the valve can be counted and stored in the control unit in order to be made use of later in maintenance to judge the state of aging of the engine. 
     The clearance  38  of the high-pressure turbine  24  as well as the Overshoot are therefore regulated in a closed loop and adaptively as a function of the aging of the engine, and this occurs throughout the lifecycle of the jet engine  10 . Typically the high-pressure turbine  24  of an aged engine has more significant clearance than a new engine. The method described above therefore makes it possible to minimize the clearance  38  of the high-pressure turbine  24  as a function of the aging of the jet engine  10 , via the controlling of the valve  44 , without risking damage to the blades of the turbine. The performance of the jet engine  10  is therefore optimized throughout its lifecycle. The EGT margin is in particular kept positive for as long as possible, extending the lifetime of the jet engine  10  before it is sent off for any maintenance.