EP4088009A1

EP4088009A1 - Control method and unit for controlling the clearance of a high-pressure turbine to reduce the effect of egt overshoot

Info

Publication number: EP4088009A1
Application number: EP21705233.1A
Authority: EP
Inventors: Jérôme Claude George LEMONNIER; Damien Bonneau; Cécile RACINET; Antoine Bruno VAN NOORT
Original assignee: Safran Aircraft Engines SAS
Current assignee: Safran Aircraft Engines SAS
Priority date: 2020-01-08
Filing date: 2021-01-04
Publication date: 2022-11-16
Also published as: CN114945734A; US20230044006A1; FR3105980B1; CN114945734B; WO2021140292A1; FR3105980A1

Abstract

Method for controlling a clearance between blade tips of an aircraft engine turbine rotor and a turbine shroud, which comprises estimating the clearance to be controlled and controlling a valve supplying a flow of air to the turbine shroud depending on the clearance estimated, said method comprising: - detecting (100) a transient acceleration phase on the basis of at least one parameter representative of the engine; - receiving (102) data relating to the aircraft altitude; - determining (104) data representative of the temperature of the rotor during the transient acceleration phase and at a steady speed and calculating a relative difference in temperatures; if the transient acceleration phase is detected and if the relative difference in temperatures is greater than a predetermined minimum temperature difference, -defining a valve opening level and a valve opening time by means of a predefined altitude/relative temperature difference look-up table; and controlling (114) the opening of the valve in order to supply the flow of air to the turbine shroud.

Description

Title of the invention: METHOD AND CONTROL UNIT FOR CONTROL OF THE SET OF A HIGH PRESSURE TURBINE FOR REDUCING THE EFFECT OF

EGT EXCEEDING

Technical area

The present invention relates to the general field of turbomachines for gas turbine aeronautical engines. It is more specifically aimed at controlling the clearance between, on the one hand, the tops of mobile blades of a turbine rotor and, on the other hand, a turbine ring of an external casing surrounding the blades.

Prior art

The clearance existing between the top of the blades of a turbine and the ring which surrounds them is dependent on the differences in dimensional variations between the rotating parts (disc and blades forming the turbine rotor) and the fixed parts (outer casing including the turbine ring that it includes). These dimensional variations are both of thermal origin (related to temperature variations of the blades, disc and 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. On the other hand, during an increase in speed, for example when changing from an idle speed on the ground to a take-off speed in a turbomachine for an aeronautical engine, the centrifugal force exerted on the turbine rotor tends in bringing the blade tops closer to the turbine ring before the turbine ring has had time to expand under the effect of the increase in temperature linked to the increase in speed. There is therefore a risk of contact at this operating point called the pinch point.

It is known practice to use an active control system to control the blade tip set of a turbomachine turbine. A system of this type generally works by directing air taken, for example, at the level of a compressor and / or the blower of the turbine onto the outer surface of the turbine ring. turbomachine. If this air is cool, by being sent to the outer surface of the turbine ring, this has the effect of cooling the latter and thus of limiting its thermal expansion. The game is thus minimized. Conversely, if this air is hot, this promotes thermal expansion of the turbine ring, which increases the clearance and makes it possible, for example, to avoid contact at the aforementioned pinch point.

Such active piloting is controlled by a control unit, for example by the full authority regulation system (or FADEC) of the turbomachine. Typically, the control unit acts on a valve with a regulated position to control the flow rate and / or the temperature of the air directed onto the turbine ring, as a function of a clearance instruction and an estimate of the clearance. of real dawn top.

The turbomachine also has an engine operating limit temperature defined with respect to a combustion gas limit temperature determined downstream of its combustion chamber, more particularly downstream of the high pressure engine turbine. This temperature is commonly referred to as the "Red Line EGT" and is identified as the maximum allowable engine temperature and determined during tests carried out on the ground ("Block Tests") by the manufacturer and then communicated by the latter. In other words, the Red Line EGT is the maximum value declared by the manufacturer, this being certified according to the life cycle of the engine (eg: new or reconditioned engine). Once this limit is reached, the engine is removed for maintenance in order to restore a positive EGT margin. The term EGT margin is used here to mean the difference between the Red Line EGT certified by the manufacturer and a temperature of the combustion gases determined downstream of the engine combustion chamber.

The temperature of the combustion gases downstream of the combustion chamber of the engine is generally maximum during a rapid acceleration phase, taking into account 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 ring which surrounds them increases. Temperatures of the order of 20 to 30K higher than a temperature of the engine in stabilized speed, the stabilized speed being obtained are measured downstream of the combustion chamber, by way of example at the outlet of the high pressure turbine. after a given time interval following the acceleration phase of the engine. The temperature difference between the maximum temperature of the combustion gases determined during an acceleration phase of the turbomachine and the temperature of its stabilized speed determined following this acceleration phase is commonly referred to as "Overshoot EGT" .

During an acceleration phase, optimizing the clearance between the blades of the high pressure turbine rotor and the ring surrounding them can reduce the EGT overshoot, and therefore the maximum temperature of the combustion gases. However, such optimization may present a risk of premature wear of the high pressure turbine. For example, too great a reduction in the EGT overshoot linked to a prolonged reduction in the clearance of the high pressure turbine for a new, hot engine, or already having a minimized clearance of its high pressure turbine, can lead to a pinch point between the blades and the high pressure turbine ring. Thus, limiting an EGT overshoot, the duration of which is of the order of ten minutes, may present a risk of permanent degradation of the high pressure turbine blades, thus impacting the overall performance of the engine and its fuel consumption. fuel.

It therefore appears desirable to reduce or even eliminate the EGT overshoot phenomenon during variations in engine speed, while avoiding the possible risk of degradation of the blades of the high pressure turbine.

With application FR3078362 filed in the name of the applicant, a process is known which makes it possible to reduce this EGT overshoot phenomenon. It uses a threshold temperature T2 (<T1 maximum acceptable temperature) below which the valve bringing the air flow closes. However, a threshold temperature T2 too close to T1 will imply, in order to deal with this phenomenon, a large number of opening / closing of the valve such as to generate significant oscillations in the temperature and a high stress on the control system. The risk of failure of the valve and the entire pilot system will be significantly increased. On the contrary, a threshold temperature T2 that is too low will prevent the oscillation phenomenon but will run the risk of cooling the casing too much, then significantly increasing the risk of contact and wear of the turbine blades. The necessary compromise between these two values, if possible, therefore proves difficult to find. In addition, this method assumes that all EGT overshoots observed are due to high pressure turbine clearance when this is not necessarily the case.

Disclosure of the invention

The object of the present invention is to remedy the aforementioned drawbacks and in particular to provide a method for controlling the valve optimizing the clearance at the top of the turbine blade and making it possible in particular to differentiate the different types of maneuvers and flight conditions that may generate a such EGT overshoot phenomenon (Altitude, idle time, go-around, flight or ground conditions, etc.).

To this end, the invention provides a method of controlling a 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 ring of a casing surrounding said blades of the high pressure turbine, the method comprising the estimation of the clearance to be controlled and the control of a valve delivering an air flow directed towards said ring turbine as a function of the clearance thus estimated, this method being characterized in that it comprises the following steps:

- detection of a transient phase of engine acceleration from at least one parameter representative of the engine;

- receipt of data relating to the altitude of the aircraft;

- determination of data representative of the temperature of said rotor of the high pressure turbine of the engine during the transient phase of acceleration and in steady state and calculation of a relative temperature difference between said temperature data obtained during the transient phase of acceleration and in a stabilized regime;

- if said transient acceleration phase is detected and if said relative temperature difference is greater than a predetermined minimum temperature difference, definition of an opening level and an opening time of the valve by a table of predefined altitude / relative temperature difference correspondence; and

control of the opening of the valve at the opening level and during the opening time defined to deliver said air flow to the turbine ring. Thus, the above method makes it possible to cover all the maneuvers and conditions of use liable to generate an EGT overshoot phenomenon, regardless of the level of wear of the high pressure turbine. The introduction of a dynamic table based on the temperature of the high pressure turbine and the altitude makes it possible to best adapt the level of opening and the duration of opening of the valve during this EGT overshoot phenomenon. .

Preferably, the definition of the opening level and the opening time of the valve from a predefined altitude / relative temperature difference correspondence table is carried out only if said estimated clearance is also greater than a predetermined minimum clearance.

Taking into account in the process a predetermined minimum clearance obtained on engine tests making it possible to prevent any risk of turbine wear, the performance of the turbomachine is therefore improved and the shelf life of an EGT margin is extended. positive, which makes it possible to increase the life of the engine and to postpone its maintenance deposit.

Advantageously, the control of the opening of the valve also includes a delay on its opening defining a period from which the valve is open to the opening level and for the desired opening time following the detection of the transient phase d. 'acceleration.

Preferably, the transient acceleration phase of the engine is detected from the difference between the temperature in steady state Tstab and the temperature in the transient acceleration phase Ttrans.

Advantageously, said at least one parameter representative of the engine is chosen from: the speed of a low pressure turbine of the engine, the speed of the high pressure turbine, a pressure measured at a high pressure compressor, the angular position of an airplane throttle control lever and data representative of the temperature of the gases leaving the combustion chamber of the engine.

Preferably, the data representative of the temperature of the rotor is an estimate of the temperature of a rotor disk of the high pressure turbine on the basis of said at least one parameter representative of the engine. The invention also proposes, according to another aspect, a control unit for controlling a clearance between, on the one hand, the tops of blades of a rotor of a high pressure turbine of an engine. gas turbine airplane and, on the other hand, a turbine ring of a casing surrounding said blades of the high pressure turbine, the control unit comprising means for estimating the clearance to be piloted and control means for 'a valve, the valve being configured to deliver an air flow to said ring of the turbine as a function of the clearance thus estimated, the control unit being characterized in that it comprises:

- detection means configured to detect a transient phase of acceleration of the engine from at least one parameter representative of the engine;

- reception means configured to receive representative data relating to the altitude of the aircraft;

- Calculation means configured to determine data representative of the temperature of said rotor of the high pressure turbine of the engine during the transient acceleration phase and in steady state and to calculate a relative temperature difference between said temperature data determined during the phase transient acceleration and steady state;

- the control means being configured to control an opening of the valve to deliver said air flow to the turbine ring at an opening level and during an opening time defined by an altitude / relative deviation correspondence table of predefined temperatures, if the transient acceleration phase is detected and if said relative temperature difference is greater than a predetermined minimum temperature difference.

Preferably, the predefined altitude / relative temperature difference correspondence table delivers a value of the parameters {X, Y (i), Z (i)} for a given pair of values {Altitude; relative difference (Tstab-Ttrans) / Tstab}, with:

X defining a time from which the valve must open,

Z (i) defining for a given level i the opening level of the valve,

Y (i) defining the opening time of the valve for stage i and the opening level Z (i). Advantageously, the valve is a regulated position valve intended to remain in the closed position in the absence of an electrical power supply and the position of which may be between 0% (full closure), corresponding to a closed valve, and 100% (full open).

The invention also provides, 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 flow directed towards the turbine ring. and wherein the valve is controlled by the control means.

Brief description of the drawings

Other characteristics and advantages of the present invention will emerge from the description given below, with reference to the appended drawings which illustrate an exemplary embodiment thereof without any limiting nature and on which:

[Fig. 1] Figure 1 is a schematic view in longitudinal section of a portion of a gas turbine aircraft engine according to one embodiment of the invention,

[Fig. 2] FIG. 2 is an enlarged view of the engine of FIG. 1 showing in particular the high pressure turbine thereof,

[Fig. 3] Figure 3 is a functional diagram of a control unit of a valve for controlling the set of blade tips in the engine of Figure 1 according to the invention,

[Fig. 4] Figure 4 illustrates the difference between the transient temperature and the stabilized temperature; and

[Fig. 5] FIG. 5 shows the timing logic for controlling the valve according to the invention with the evolution of the engine speed.

Description of the embodiments

FIG. 1 schematically represents a turbojet 10 of the double-flow and double-body type to which the invention applies in particular. Of course, the invention is not limited to this particular type of gas turbine aircraft engine. In a well-known manner, the turbojet 10 of longitudinal axis XX comprises in particular a fan 12 which delivers an air flow in a primary flow stream 14 and in a secondary stream flow stream 16 coaxial with the stream. primary flow. From upstream to downstream in the direction of flow of the gas flow passing through it, the primary flow stream 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 represented more precisely by FIG. 2, the high pressure turbine 24 of the turbojet engine comprises a rotor formed by a disc 28 on which are mounted a plurality of mobile blades 30 arranged in the flow duct of the primary flow 14. The rotor is surrounded by a turbine casing 32 comprising a turbine ring 34 carried by an outer turbine casing 36 via a mounting bracket 37.

Turbine ring 34 may be formed from a plurality of adjacent sectors or segments. On the internal side, it is provided with a layer 34a of abradable material and surrounds the blades 30 of the rotor, leaving a clearance 38 with the tops 30a thereof.

According to the invention, there is provided a system making it possible to control the clearance 38 by modifying, in a controlled manner, the internal diameter of the external turbine casing 36. To this end, a control unit 50 controls the flow rate and / or the flow rate. temperature of the air directed towards the external turbine casing 36. The control unit 50 is for example the full authority control system (or FADEC) of the turbojet 10.

In the example shown, a pilot box 40 is arranged around the outer turbine casing 36. This casing receives fresh air by means of an air duct 42 opening at its upstream end into the air duct. flow of the primary flow at one of the stages of the high pressure compressor 20 (for example by means of a scoop known per se and not shown in the figures). The fresh air circulating in the air duct is discharged onto the outer turbine casing 36 (for example by means of a multi-perforation of the walls of the control box 40) causing cooling thereof and therefore a decrease in its internal diameter. As shown in Figure 1, a valve 44 is arranged in the air duct 42. This valve 44 is controlled by the control unit 50 and is designed to remain in the closed position in the absence of electrical power.

Valve 44 is a position valve continuously regulated between the 0% full closed position (valve closed) and the 100% full open position (fully open valve).

When the valve 44 is fully open (100% position), the fresh air is brought to the outer turbine casing 36, which has the effect of a thermal contraction of the latter and therefore a reduction in the clearance 38. On the contrary, when the valve 44 is fully closed (0% position), the fresh air is not brought to the outer turbine casing 36 which is therefore heated by the primary flow. This has the effect either of a thermal expansion of the casing 36 and an increase in the clearance 38, or at least a controlled limitation (or even a stop) of the expansion of the casing 36. In the intermediate positions, the outer casing of the turbine 36 contracts or expands and clearance 38 increases or decreases, to a lesser extent.

Of course, the invention is not limited to this example. Thus, another example can consist in taking air at the level of two different stages of the compressor and controlling valves 44 to modulate the flow rate of each of these withdrawals in order to adjust the temperature of the mixture to be directed onto the external turbine casing. 36.

We now describe the control of the valve 44 by the control unit 50. According to the invention, the control unit 50 comprises:

- detection means 52 configured to detect a transient phase of acceleration of the turbojet 10 over a predetermined time interval;

- Receiving means 54 configured to receive data relating to the altitude of the aircraft;

- Calculation means 56 configured to determine data representative of the temperature of the rotor of the high pressure turbine 24 of the turbojet 10 during this transient phase of acceleration and in steady state in order to deduce by calculation a relative temperature difference at on the basis of the aforementioned temperature data; control means 58 configured to define an opening level and an opening time of the valve 44 and to control the valve 44 according to a dynamic table predefined as a function of the altitude and of the relative difference between the temperature in steady state and temperature in transient acceleration phase.

The detection means 52, the reception means 54, the calculation means 56 and the control means 58 together form a control module for the valve 44 integrated 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 the programmable logic circuit type) or to a combination of an electronic circuit and a program of computer.

The term “transient acceleration phase of the turbojet 10” is understood here to mean a speed transition linked to an acceleration phase of the turbojet 10 occurring between two stabilized speeds of the latter. The transient acceleration phase that one seeks to detect using the detection means 52 can, by way of example, correspond to a transition between the idle speed on the ground and the stabilized speed (called take-off). , that is to say in the acceleration phase between these two regimes. In another example, the transient acceleration phase can correspond to the acceleration phase between any intermediate speed (eg: half throttle) and the flight speed.

The various steps of the method for controlling the game 38 implemented in the control unit 50 are now described with reference to FIG. 3. It is understood that the order of some of these steps illustrated in these figures is given by way of reference. illustrative, these steps being able in a non-illustrated example to be carried out in parallel.

The first step 100 consists in the detection of this transient phase of acceleration of the turbojet 10 which can be carried out from one or more parameters representative of the turbojet 10.

A representative parameter of the turbojet 10 is for example its rotation speed, but other parameters can also be used such as: the speed of the high pressure turbine 24, the speed of the low pressure turbine 26, the angular position of the control lever. control of the gases of the airplane, a measured or calculated temperature of the combustion gases at the outlet of the combustion chamber 22 or a pressure measured at the high pressure compressor 20. The detection of a transient phase of acceleration of the turbojet 10 is then carried out from a continuous determination of its speed, variations with respect to a setpoint characterizing a variation in speed of the engine. turbojet 10. Thus, if during the predetermined interval the variation in the speed of rotation of the turbojet 10 is greater than or equal to a variation threshold characterizing a transient acceleration phase of the turbojet 10, the detection means 52 detect a transient phase d 'acceleration.

Then in a step 102, or preferably in parallel, the reception means 54 receive data representative of the altitude of the aircraft (also making it possible to define whether the aircraft is in flight or on the ground).

In a following step 104, two data representative of the temperature of the rotor (disc 28 and blades 30) of the high pressure turbine 24 of the turbojet 10 are determined by the calculation means 56. These data representative of the temperature of the rotor are on the one hand, a first temperature Tstab estimated in permanent or stabilized mode and on the other hand a second temperature Ttrans estimated during the transient acceleration phase. The first temperature in steady state Tstab is preferably determined from the engine data detected in step 100 and the second temperature in the transient acceleration phase Ttrans is determined from a formulation as a function of the response time of the representative temperature. rotor disc.

By way of example, the data representative of the first temperature Tstab can be estimated by a polynomial function of the pressures and temperatures measured in the engine and is given by the following formula: i, j being integers;

C0, Ci and Cj represent the coefficients of the polynomial; Pi representing pressures in the engine;

Ti representing temperatures in the engine;

And that relating to the second temperature Ttrans by: with z a function of an engine pressure parameter.

The calculation means 56 then determine by calculation a relative difference between the temperatures Tstab and the temperatures Ttrans, that is to say the ratio (Tstab-Ttrans) / Tstab. This relative temperature difference will make it possible, on the one hand, to confirm the transient acceleration phase (it should be noted that this is an alternative example of detection of this transient phase), the transient temperature exhibiting a delay over the temperature. stabilized, there is necessarily a positive difference (see FIG. 4) between the temperature stabilized after acceleration and the transient temperature at the start of the plateau after acceleration, and on the other hand to detect a high risk of EGT overshoot. Indeed, the greater the gap, the longer the disc 28 will take time to expand and a strong gap in play at the start of a stabilized regime (which is the cause of the EGT overshoot) will take time to expand. to resorb.

The following three steps 106, 108 and 110 are test steps carried out by the control unit 50 to identify from the detection means 52, the reception means 54 and the calculation means 56 the possible occurrence of a EGT overshoot situation for which:

- a transient phase of acceleration of the turbojet engine is detected,

- the relative temperature difference is greater than a predefined minimum difference, and optionally

the play estimated by an on-board play model for the needs of the control system is greater than a predetermined minimum play depending on the operating conditions of the engine.

Reaching the minimum temperature difference threshold at the same time as the detection of a transient phase of acceleration of the turbojet makes it possible to clearly distinguish between Important and rapid accelerations which are the most risky for EGT overshoot.

Likewise, when test 110 is present, reaching the minimum clearance threshold in parallel with the detection of a transient phase of acceleration of the turbojet provides protection against the risk of wear, the specific control of the valve n ' being activated only if the estimated current clearance is greater than the minimum clearance.

After each test step 106, 108, 110, the control unit 50 tries to detect the possible occurrence of one of the above three conditions.

If the occurrence of such a situation is not identified, the control unit 50 deduces the non-occurrence of an EGT overshoot and ensures, in a final step 112, the control of the valve 44 conventionally according to the estimated set 38 set then it is returned to the initial step 100 to detect a possible acceleration phase likely to generate this time an EGT overshoot. This conventional control is based on the comparison of the clearance coming from a clearance computer integrated in the FADEC and a setpoint clearance (which is generally a clearance to be achieved in order to maximize the performance of the turbine or in order to protect it from wear depending on the requirements. Terms of use). This clearance estimate is made as it is instantly known from engine and aircraft data.

Conversely, if the aforementioned situation is detected (positive response to each of the two or three tests of steps 106, 108, 110), the control unit 50 deduces an EGT overshoot situation which it then seeks to minimize. by acting on the clearance 38 of the high pressure turbine 24 by a specific control of the valve 44 in an alternative final step 114. Indeed, in the absence of action, such a situation would risk reducing the EGT margin of the engine and therefore its duration of use before it is deposited for maintenance. This direct action on play 38 is then aimed at keeping a positive EGT margin as long as possible. As before, once the game is controlled and a stabilized speed is reached, it is returned to the initial step 100 to detect a possible new phase of acceleration which could generate an EGT overshoot.

The specific piloting of the valve to reduce the EGT overshoot is illustrated in FIG. 5 and it is carried out by an action on the amplitude and the duration of opening of the valve 44 via the following three parameters: X: This first parameter defines a period from which the valve opens following the detection of an EGT overshoot compared to the detection of the stabilization of the speed reached (delay on opening in seconds),

Z (i): This second parameter defines for a given level i the valve opening level (100% = valve fully open; 50% = valve semi open; 0% = valve fully closed),

Y (i): This third parameter defines the opening time (in seconds) of the valve for stage i and the opening level Z (i).

All these parameters are presented in a predefined dynamic table as a function of the altitude and the relative difference between the stabilized temperature Tstab and the transient temperature Ttrans. For a given pair of values {Altitude; relative deviation (Tstab-Ttrans) / Tstab), there is a value of the parameters {X, Y (i), Z (i)} that is most suitable for reducing the EGT overshoot. The number of steps is only limited by the memory capacities of the control unit 50 and the type of valve used. It is therefore quite possible to integrate more than three levels. If this memory capacity is limited, it is possible to delete the first parameter.

Likewise, the number of levels is determined by the type of valve: if a regulated valve is used, then there is an infinite number of possible levels. If you use an all-or-nothing valve, then there are only two possible levels.

Thus, the control of a valve 44 as described above makes it possible to maintain a positive EGT margin based on the thermal state of the rotor and the speed of rotation of the turbine and to cover all the maneuvers and operating conditions. use likely to generate an EGT overshoot phenomenon. The introduction of a dynamic table based on the relative temperature difference to the high pressure turbine rotor and the altitude allows to better adapt the amplitude and the duration of the valve opening during this phenomenon. Taking into account in the process a minimum clearance established in advance on engine tests also allows precise estimation of the clearance at any time, thus preventing any risk of premature wear of the turbine.

Claims

[Claim 1] A method of controlling a clearance (38) between, on the one hand, the tops (30a) of blades (30) of a rotor of a high pressure turbine (24) of an engine ( 10) of a gas turbine airplane and, on the other hand, a turbine ring (34) of a housing (32) surrounding said vanes (30) of the high pressure turbine (24), the method comprising the estimation of the clearance to be controlled and the control of a valve (44) delivering an air flow directed towards said turbine ring (34) as a function of the clearance thus estimated, this method being characterized in that it comprises the following steps : detection (100) of a transient phase of acceleration of the engine (10) on the basis of at least one parameter representative of the engine (10); reception (102) of data relating to the altitude of the aircraft; determination (104) of data representative of the temperature of said rotor of the high pressure turbine of the engine during the transient phase of acceleration and in steady state and calculation of a relative temperature difference between said temperature data of the transient phase of acceleration and temperature stabilized regime; if said transient acceleration phase is detected and if said relative temperature difference is greater than a predetermined minimum temperature difference, definition of an opening level and an opening time of the valve by a correspondence table preset altitude / relative temperature deviation; control (114) of the opening of the valve (44) at the opening level and during the opening time defined to deliver said air flow to the turbine ring (34).

[Claim 2] The control method according to claim 1, for which the definition of the opening level and of the opening time of the valve from a predefined altitude / relative temperature difference correspondence table is carried out only if said estimated clearance is also greater than a predetermined minimum clearance.

[Claim 3] A method according to claim 1 or claim 2, wherein the control of the opening of the valve also comprises a delay on opening defining a delay from which the valve is open to the opening level and for the desired opening time following detection of the transient acceleration phase.

[Claim 4] A control method according to any one of claims 1 to 3, in which the transient phase of acceleration of the engine (10) is detected from the difference between the temperature in steady state Tstab and the temperature in transient acceleration phase Ttrans.

[Claim 5] A control method according to any one of claims 1 to 4, wherein said at least one parameter representative of the engine is chosen from: the speed of a low pressure turbine of the engine (10), the speed of the engine. high pressure turbine, a pressure measured at a high pressure compressor (20), the angular position of a throttle control lever of the airplane and the data representative of the temperature of the gases leaving the combustion chamber (22) of the motor (10).

[Claim 6] A method according to any one of claims 1 to 5, wherein the data representative of the temperature of the rotor is an estimate of the temperature of a rotor disc (28) of the high pressure turbine (24) from of said at least one parameter representative of the engine.

[Claim 7] Control unit (50) for controlling a clearance (38) between, on the one hand, the tops (30a) of blades (30) of a rotor of a high pressure turbine (24 ) of a gas turbine airplane engine (10) and, on the other hand, a turbine ring (34) of a housing (32) surrounding said vanes (30) of the high pressure turbine (24) , the control unit (50) comprising means for estimating the clearance to be controlled and for controlling (58) a valve (44), the valve (44) being configured to deliver an air flow to said ring of the turbine (34) as a function of the play thus estimated, the control unit (50) being characterized in that it comprises: detection means (52) configured to detect a transient phase of acceleration of the engine (10 ) from at least one parameter representative of the engine (10); reception means (54) configured to receive data relating to the altitude of the aircraft; computing means (56) configured to determine data representative of the temperature of said rotor of the high pressure turbine of the engine (10) during the transient phase of acceleration and in steady state and calculate a relative temperature difference between said temperature data determined during the transient phase of acceleration and in regime stabilized; the control means (58) being configured to control an opening of the valve (44) to deliver said air flow to the turbine ring (34) at an opening level and for an opening time defined by a predefined altitude / relative temperature difference correspondence table, if the transient acceleration phase is detected and if said relative temperature difference is greater than a predetermined minimum temperature difference.

[Claim 8] A control unit according to claim 7, wherein the predefined altitude / relative temperature deviation correspondence table delivers a value of the parameters {X, Y (i), Z (i)} for a given pair of values { Altitude ; relative deviation (Tstab-Ttrans) / Tstab}, with: X defining a time from which the valve must open, Z (i) defining for a given stage i the opening level of the valve, Y (i) defining the valve opening time for stage i and opening level Z (i).

[Claim 9] A control unit according to claim 7 or claim 8, in which the valve (44) is a regulated position valve intended to remain in the closed position in the absence of electrical power and the position of which may be between 0 % (full closing), corresponding to a closed valve, and 100% (full opening).

[Claim 10] A gas turbine aircraft engine (10) comprising a control unit (50) according to any one of claims 7 to 9 and at least one valve (44) for acting on a directed air flow. towards the turbine ring (34) and in which the valve (44) is controlled by the control means (58).