METHOD AND CONTROL UNIT FOR CONTROLLING THE GAP OF A HIGH-PRESSURE TURBINE TO REDUCE THE IMPACT OF ICING

Abstract

METHOD AND CONTROL UNIT FOR CONTROLLING THE CLEARANCE OF A HIGH-PRESSURE TURBINE TO REDUCE THE IMPACT OF ICING. Method for controlling the clearance between the blade tips of an aircraft engine turbine rotor and a turbine ring, comprising estimating a nominal clearance to be controlled and controlling a valve delivering an airflow directed to the turbine ring as a function of the nominal clearance thus estimated, method wherein, to prevent the formation of ice that could affect the aircraft engine, the estimated nominal clearance is increased if the on / off result of a predetermined control logic indicates that icing conditions are present and until conditions for returning to normal operation are present, the predetermined control logic being a function of the following parameters: ambient temperature, altitude, temperature T12, and rotational speed of a low-pressure shaft. Figure for abstract: Fig. 3.

Description

Title of the invention: METHOD AND CONTROL UNIT FOR CONTROLLING THE GAP OF A HIGH-PRESSURE TURBINE TO REDUCE THE IMPACT OF ICING technical field

[0001] The present invention relates to the general field of turbomachinery for aeronautical gas turbine engines. More specifically, it aims at controlling the clearance between, on the one hand, the movable blade tips of a turbine rotor and, on the other hand, a turbine ring of an external casing surrounding the blades. Previous technique

[0002] The clearance between the tips of turbine blades and the surrounding ring depends on the differences in dimensional variations between the rotating parts (disk and blades forming the turbine rotor) and the stationary parts (outer casing, including the turbine ring it comprises). These dimensional variations are both thermal in origin (related to temperature variations of the blades, the disc, and the casing) and mechanical in origin (particularly related to the effect of the centrifugal force acting on the turbine rotor).

[0003] To increase turbine performance, it is desirable to minimize backlash as much as possible. On the other hand, during an increase in engine speed, for example when transitioning from idle speed on the ground to takeoff speed in a turbomachine for an aircraft engine, the centrifugal force acting on the turbine rotor tends to bring the blade tips closer to the turbine ring before the turbine ring has had time to expand due to the temperature increase associated with the increased engine speed. There is therefore a risk of contact at this operating point, known as the pinch point.

[0004] It is known to use an active control system to control the blade tip clearance of a turbomachine turbine. Such a system generally operates by directing air, drawn for example from a compressor and / or the turbomachine's fan, onto the outer surface of the turbine ring. Fresh air directed onto the outer surface of the turbine ring cools it and thus limits its thermal expansion. The clearance is therefore minimized. Conversely, hot air promotes the thermal expansion of the turbine ring, which increases the clearance and, for example, prevents contact at the aforementioned pinch point.

[0005] Such active control is controlled by a control unit, for example by the full authority control system (or FADEC) of the turbomachine. Typically, the control unit acts on a position-controlled valve to control the flow rate and / or temperature of the air directed onto the turbine ring, according to a setpoint and an estimate of the actual blade tip clearance implemented in the FADEC and based on the thermodynamic conditions of the turbomachine and the flight conditions of the aircraft.

[0006] The turbomachine is also subject to icing conditions, which generate imbalances related to the centrifugal force of the ice accumulating on the fan. Therefore, the engine dynamics may exhibit unwanted high-pressure / low-pressure coupled vibration modes within a certain engine operating range.

[0007] Due to these imbalances related to icing, these undesired modes can generate a significant loss of clearance between stator and rotor which can lead to contacts between the blades and the turbine ring which can generate wear on the blade tips and a decrease in the efficiency of the high-pressure turbine and consequently a decrease in performance and therefore an increase in the fuel consumption of the turbomachine.

[0008] It therefore appears desirable to reduce or even eliminate the impact of icing under certain engine operating conditions, while eliminating the possible risk of degradation of the high-pressure turbine blades. Description of the invention

[0009] The present invention aims to remedy the aforementioned drawbacks and in particular to propose a method of controlling the valve reducing in operation the clearance at the top of the turbine blade and allowing to limit as much as possible the impact of clearance consumption related to imbalances from icing.

[0010] To this end, 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 ring of a casing surrounding said high-pressure turbine blades, the method comprising estimating the clearance to be controlled and controlling a valve delivering an airflow directed towards said turbine ring as a function of the clearance thus estimated, this method being characterized in that, to prevent the formation of ice likely to affect said aircraft engine, said estimation of the nominal clearance is increased if the on / off result of a predetermined control logic indicates that ice formation conditions are met and as long as conditions for a return to normal operation are not present, said predetermined control logic being a function of the following parameters: an ambient temperature, an altitude,a temperature T12 and a low shaft rotation speed, pressure.

[0011] Thus, the above process makes it possible to limit the impact of the consumption of play linked to the imbalance from icing conditions likely to appear in a risk zone defined by the piloting logic thus predetermined.

[0012] Preferably, said predetermined piloting logic follows this logic: Tamb > X°C AND Altitude < Y km AND T12 < Z°C AND XI rpm < LP speed < X2 rpm, where "Tamb" is the ambient temperature, "Altitude" is the aircraft altitude, "T12" is the engine inlet temperature, and "LP speed" is the low-pressure shaft rotational speed, and X, Y, Z, XI, X2 are predetermined values ​​defining the icing conditions likely to affect said aircraft engine. By way of example, it may be considered that: -70 < X ​​< 0°C; 0 < Y < 20 km; Z < 0°C; XI = 0 rpm and 4000 < X2 < 6500 rpm.

[0013] Advantageously, the said conditions for returning to normal operation follow the logic below: BP regime > X3 rpm OR T12 > XX °C with X3 a speed for example 100 rpm higher than X2 allowing the ice to be centrifuged and XX a temperature higher than 0°C allowing the ice to be melted.

[0014] The invention also proposes, according to another aspect, a control unit 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 ring of a casing surrounding said high-pressure turbine blades, the control unit comprising means for estimating the clearance to be controlled and means for controlling a valve, the valve being configured to deliver an airflow to said turbine ring according to the clearance thus estimated, the control unit being characterized in that, to prevent the formation of ice likely to affect said aircraft engine, it further comprises: acquisition means configured to acquire data relating to the engine and the aircraft, namely an ambient temperature "Tamb", an aircraft altitude "Altitude",an engine inlet temperature "T12" and a low-pressure shaft rotation speed "BP regime" and logic calculation means configured to apply a predetermined control logic to the data thus collected and deliver an on / off result indicating whether or not the conditions for ice formation are met, and control means configured to increase the estimated nominal clearance if the on / off result of the predetermined control logic indicates that the conditions for ice formation are met and as long as conditions for returning to normal operation are not present.

[0015] Preferably, the control unit includes a game calculator integrated into the engine control system.

[0016] 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 act on an airflow directed towards the turbine ring and in which the valve is controlled by the control means. Brief description of the drawings

[0017] Other features and advantages of the present invention will become apparent from the description given below, with reference to the accompanying drawings which illustrate an example of an embodiment without being limiting in any way and on which:

[0018] [Fig-1] [Fig.1] is a schematic and longitudinal sectional view of part of a aircraft engine with gas turbine according to an embodiment of the invention,

[0019] [Fig.2] [Fig.2] is an enlarged view of the engine of [Fig.1] showing in particular its high-pressure turbine, and

[0020] [Fig.3] [Fig.3] is a functional diagram of a control unit for a valve allowing control of the vane tip clearance in the motor of [Fig.1] according to the invention. Description of the implementation methods

[0021] The basic principle of the invention is based on the definition of a specific logic enabling the detection of whether the turbomachine is subjected to conditions which are risky with regard to the accumulation (accretion) of ice in an operating range of the turbomachine defined as risky with regard to coupled high-pressure-low-pressure vibration modes.

[0022] Figure 1 schematically represents a turbojet engine 10 of the twin-spool, twin-flow type to which the invention applies in particular. Of course, the invention is not limited to this particular type of gas turbine aircraft engine.

[0023] As is well known, the turbojet 10 with longitudinal axis XX includes, in particular, a fan 12 which delivers an airflow into a primary flow duct 14 and into a secondary flow duct 16 coaxial with the primary flow duct. From upstream to downstream in the direction of the gas flow passing through it, the primary flow duct 14 includes 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.

[0024] As more precisely represented by [Fig.2], the high-pressure turbine 24 of the turbojet comprises a rotor formed of a disc 28 on which are mounted a plurality of movable blades 30 arranged in the flow channel of the primary flow 14. The rotor is surrounded by a turbine casing 32 comprising a turbine ring 34 carried by an external turbine casing 36 by means of a mounting support 37.

[0025] The turbine ring 34 can be formed from a plurality of adjacent sectors or segments. On its inner side, it is provided with a layer 34a of abradable material and surrounds the rotor blades 30, maintaining a clearance with the tips 30a thereof. 38 whose control is provided by modifying, in a controlled manner, the internal diameter of the external turbine casing 36. For this purpose, a control unit 50 controls the flow rate and / or 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 incorporating a clearance computer delivering at each instant an estimate of the clearance to be controlled.

[0026] In the illustrated example, a pilot housing 40 is arranged around the outer turbine casing 36. This housing receives fresh air by means of an air duct 42 opening at its upstream end into the primary flow path 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 multiple perforations in the walls of the pilot housing 40), causing it to cool and thus reducing the internal diameter of the outer turbine casing 36.

[0027] As shown in [Fig. 1], a valve 44 is arranged in the air duct 42. This valve 44 is controlled by the control unit 50 and is intended to remain in the closed position in the absence of power supply.

[0028] The valve 44 is a continuously regulated position valve between the fully closed position at 0% (valve closed) and the fully open position at 100% (valve fully open).

[0029] When the valve 44 is fully open (100% position), fresh air is supplied to the outer turbine casing 36, resulting in thermal contraction of the latter and thus a reduction in clearance 38. Conversely, when the valve 44 is fully closed (0% position), fresh air is not supplied to the outer turbine casing 36, which is therefore heated by the primary flow. This results either in thermal expansion of the casing 36 and an increase in clearance 38, or at least a controlled limitation (or even a complete halt) of the expansion of the casing 36. In intermediate positions, the outer turbine casing 36 contracts or expands, and clearance 38 increases or decreases, to a lesser extent.

[0030] Of course, the invention is not limited to this example. Thus, another example may consist of drawing air from two different stages of the compressor and controlling valves 44 to modulate the flow rate of each of these two draws in order to regulate the temperature of the mixture to be directed onto the external turbine casing 36.

[0031] According to the invention, the control unit 50 comprises acquisition means 52 configured to acquire data relating to the engine and the aircraft, namely the ambient temperature "Tamb", the aircraft altitude "Altitude", the engine inlet temperature "T12" and the low-pressure shaft rotation speed "BP regime" and logic calculation means 54 configured to apply to the data thus collected a predetermined control logic and deliver a binary calculation result, and finally control means 56 configured to define a level and duration of opening of the valve 44 according to the on or off calculation result of the predetermined control logic.

[0032] The acquisition means 52, the logic calculation means 54 and the control means 56 correspond 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 type programmable logic circuit) or to a combination of an electronic circuit and a computer program.

[0033] The different steps of the game control process 38 implemented in the control unit 50 are now described with reference to [Fig.3].

[0034] The first stage 100 is a normal operating stage of the turbomachine where the clearance setpoint estimated by a clearance model of the clearance computer onboard in the FADEC for the needs of the control system is the nominal setpoint, the valve 44 delivering under the control of the FADEC an air flow appropriate to the operating conditions of the engine, in particular the engine speed considered.

[0035] In a second step 102, it is tested whether the conditions for ice formation are met, i.e. whether the turbomachine is subjected to conditions that are risky with regard to ice accretion in an operating range of the turbomachine defined as risky with regard to coupled high-pressure-low-pressure vibration modes, the criteria for defining these risky conditions conducive to ice accretion being as follows:

[0036] - Tamb > X°C AND - Altitude < Y km AND T12 < Z°C AND - XI rpm < BP speed < X2 rpm

[0037] The parameter “Tamb” is the ambient temperature measured in °C outside the turbomachine;

[0038] The parameter "Altitude" is the altitude of the aircraft in thousands of meters;

[0039] The parameter “T12” represents the temperature measured at the inlet of the motor the aircraft (upstream of the turbomachine fan);

[0040] The parameter "BP regime" represents the rotational speed in revolutions per minute of the low pressure shaft;

[0041] X, Y, Z and XI, X2 being the values ​​of the ambient temperature, the altitude, the temperature T12 and the operating range of the turbomachine engine speed considered to be at risk of ice formation and dependent on the architecture and dynamics of the engine, and which a person skilled in the art will be able to determine without making evidence of inventive activity. For example, the following ranges can be retained for these different values: -70 < X ​​< 0°C; 0 < Y < 20km; Z=< 0°C; XI = 0 rpm and 4000 < X2 < 6500 rpm.

[0042] If these criteria are met (yes answer to the test in step 102), the setpoint governed by the control logic will, in a further step 104, be increased relative to the nominal setpoint defined in step 100 by a predetermined value to minimize the risk of wear on the high-pressure turbine blade tips. This will cause a decrease, or even a complete interruption, in the fresh air intake. Alternatively, in the case of hot air intake, this will result in an increase in the heat input.

[0043] This step continues until conditions for a return to normal operation are present and determined by the test of the final step 106 in which the risk zone is considered to have passed once the following criteria are met:

[0044] - BP regime > X3 rpm OR T12>XX°C

[0045] The rotation speed X3 corresponds to a speed allowing the ice to be centrifuged following an acceleration of the pilot which, depending on the architecture and dynamics of the engine, is typically 100 rpm higher than the rotation speed X2.

[0046] The temperature XX corresponds to a temperature above 0°C which allows the ice to melt.

[0047] Thus, the method of the invention makes it possible to take into account the risk of imbalances related to icing and therefore the risk of high blade tip wear of the high-pressure turbine, while limiting the fresh air intake flow rate thanks to an increased clearance setpoint (via the valve control logic) in the event of icing conditions. Wear at the blade tips, as well as fuel consumption, irreversible degradation of the high-pressure turbine's efficiency, and the performance / operability of the turbomachine, are thus limited.

Claims

Demands

1. A method for controlling the clearance between, on the one hand, the tips (30a) of blades (30) of a rotor of a high-pressure turbine (24) of a gas turbine aircraft engine (10) and, on the other hand, a turbine ring (34) of a casing (32) surrounding said blades (30) of the high-pressure turbine (24), the method comprising estimating by a control unit (50) a nominal clearance (38) to be controlled and controlling a valve (44) delivering an airflow directed towards said turbine ring (34) as a function of the nominal clearance thus estimated, this method being characterized in that, in order to prevent the formation of ice on a fan of said aircraft engine from affecting the clearance consumption at the blade tips related to the resulting imbalances, said estimation of the nominal clearance is increased to limit the impact of the clearance consumption,if the all-or-none result of a predetermined control logic indicates that ice formation conditions are met, and as long as conditions for a return to normal operation are not present, said predetermined control logic being a function of the following parameters: an ambient temperature, an altitude, a temperature T12 and a rotational speed of a low-pressure shaft.

2. A piloting method according to claim 1, wherein said predetermined piloting logic responds to the following logic: Tamb > X°C AND Altitude < Y km AND T12 < Z°C AND XI rpm < BP speed < X2 rpm with "Tamb" the ambient temperature, "Altitude" the altitude of the aircraft, "T12" the temperature at the engine inlet and "BP speed" the rotational speed of the low pressure shaft, and X, Y, Z, XI, X2 predetermined values ​​defining the ice formation conditions on the fan of said aircraft engine.

3. A piloting method according to claim 2, wherein: -70 < X ​​< 0°C; 0 < Y < 20km; Z=< 0°C; XI = 0 rpm and 4000 < X2 < 6500 rpm.

4. Method according to claim 1 or claim 2, wherein said conditions for returning to normal operation follow the logic below: BP regime > X3 rpm OR T12 > XX °C with X3 a speed enabling the ice to be centrifuged and XX a temperature enabling the ice to be melted.

5. A piloting method according to claim 4, wherein X3 = X2 +100 rpm and XX > 0°C.

6. Control unit (50) for game control between, on the one hand, tip vanes (30a) of blades (30) of a rotor of a high-pressure turbine (24) of a gas turbine aircraft engine (10) and, on the other hand, a turbine ring (34) of a casing (32) surrounding said blades (30) of the high-pressure turbine (24), the control unit (50) comprising means for estimating a nominal clearance (38) to be controlled and for controlling (58) a valve (44) configured to deliver an airflow to said turbine ring (34) as a function of the nominal clearance thus estimated, the control unit (50) being characterized in that, to prevent the formation of ice on a fan of said aircraft engine from affecting the blade tip clearance consumption related to the resulting imbalances, it further comprises: acquisition means (52) configured to acquire data relating to the engine and the aircraft, namely a temperature ambient temperature "Tamb", aircraft altitude "Altitude",an engine inlet temperature "T12" and a low-pressure shaft rotation speed "BP regime" and logic calculation means (54) configured to apply a predetermined control logic to the data thus collected and deliver an all-or-nothing result indicating whether or not the conditions for ice formation are met, and control means (56) configured to increase the estimated nominal clearance and thus limit the impact of clearance consumption, if the all-or-nothing result of the predetermined control logic indicates that the conditions for ice formation are met and as long as conditions for returning to normal operation are not present.

7. Control unit according to claim 6, wherein said predetermined control logic follows the following logic: Tamb > X°C AND Altitude < Ykm AND T12 < Z°C AND Xlrpmcrégime BP <X2rpm avec X, Y, Z, XI, X2 des valeurs prédéterminées définissant les conditions de formation de givre sur la soufflante dudit moteur d’avion.

8. Control unit according to claim 6 or claim 7, wherein said conditions for returning to normal operation follow the logic below: BP regime > X3 rpm OR T12 > 0°C with X3 a speed enabling the ice to be centrifuged.

9. Control unit according to any one of claims 6 to 8, comprising a play calculator integrated into the engine control system (FADEC 50).

10. A gas turbine aircraft engine (10) comprising a control unit (50) according to any one of claims 6 to 9 and at least one valve (44) controlled by said control unit to act selectively on an airflow directed towards the turbine ring (34).

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