WO2006087383A1

WO2006087383A1 - Variable geometry control of gas turbine engine aircraft

Info

Publication number: WO2006087383A1
Application number: PCT/EP2006/060079
Authority: WO
Inventors: Jean-Marie Brocart; Régis DELDALLE; Philippe Galozio; Michel Martini; Alain Varizat; Maurice Vernochet
Original assignee: Hispano Suiza
Priority date: 2005-02-17
Filing date: 2006-02-17
Publication date: 2006-08-24
Also published as: FR2882097A1; FR2882097B1

Abstract

The invention concerns variable geometry control of gas turbine engine aircraft, including in particular variable setting stator blade setting angles, compressor discharge valve positions and turbine blade top clearances, by means of electromechanical actuators (AEM, AEM') associated with electronic control circuits (SC, SC') using an electric energy source, at least part of the electronic control circuits (SC, SC') capable of being locally implanted at the corresponding electromechanical actuators.

Description

Control of variable geometries of a gas turbine engine.

BACKGROUND OF THE INVENTION The invention relates to the control of variable geometries of a gas turbine engine.

By "variable geometries" is meant here the dimensions and / or positions of motor members and / or positions or speeds of equipment members, other than the motor rotary members, which may be modified according to detected events or engine speeds. Examples of "variable geometries" are variable valve stator blade timing angles of compressor rectifiers, compressor discharge valve positions, turbine blade tip sets, and fuel pump speeds. Traditionally, these variations in dimensions, shapes or positions are achieved by hydraulic actuation systems using the fuel as the actuating fluid and controlled by the electronic control module of the engine control system or ECU (for Electronic Control Unit) . The fuel used for the actuation systems is taken off-line at the high-pressure outlet of a pump of the general fuel supply system of the engine. Heating and filtration of the fuel thus taken are generally necessary to avoid a risk of icing at certain actuating systems and pollution of the operating systems by possible debris of various origins conveyed by the fuel from the pump. In addition, the risks of fuel leakage exist at the level of the actuating systems and their supply circuits.

Object and summary of the invention

The aim of the invention is to modify the control of variable geometries of a gas turbine engine engine in order to improve the engine configuration and the detection of faults and to limit the causes of fuel leakage of the engine and proposes for this purpose a device in which these variable geometries are controlled by electromechanical or electric actuators using a power source electromechanical or electrical actuators being associated with electronic control circuits at least a portion of which are locally located at the corresponding electromechanical or electrical actuators. The use of fuel as a hydraulic fluid for the actuation of variable geometries such as stator vane timing angles, positions of gas discharge valves and turbine blade tip sets is no longer necessary. switch from a fuel heating system to avoid icing. In addition, for the choice or design of the main pump of the general fuel supply circuit of the engine, the constraint of having to deliver for any flight regime a fuel at a pressure and / or a flow rate enabling it to be used as a fluid. actuation is lightened. Moreover, by implanting at least a portion of the electronic control circuits of the electromechanical actuators at the level thereof, the ECU is simplified and its reliability is improved. In addition, the nature of the links between the ECU and the actuators can be changed from analogue to digital and it is possible to carry on these links not only control information but also control and diagnostic information of the actuators and equipment. associates. The electronic circuits can then be installed in the vicinity of the corresponding electromechanical actuators or be integrated therewith.

According to one embodiment, at least one and preferably several of the variable geometries are associated with at least one electromagnetic or electrical actuator with redundant controls, preferably a redundant actuator by nature or two redundant actuators. The supply circuits of the actuator or actuators are preferably doubled as redundancy, and are advantageously electrically isolated from each other to prevent the propagation of failure from one circuit to another.

Advantageously, for the electromechanical actuators power supply of variable geometries, use is made of a circuit having at least a first continuous or alternative electrical energy distribution bus to first equipment of the motor requiring an electric power below a given power threshold. and minus a second electric power distribution bus continuous or alternative to other electrical equipment of the engine, the first bus and the second bus being connected to a third bus itself connected to a source of electrical energy. The first electric power distribution bus can be used to power electromechanical actuators for transient compressor discharge valves, fuel circuit control and for clearance adjustments to turbine rotor blade tips.

The second electric power distribution bus may be used to supply electromechanical actuators for discharge valves and for compressor stator blade timing members. The second bus can also power an electric fuel pump motor.

The first bus can be connected to the third bus via at least one voltage converter or transformer. The third bus can be connected to the electrical distribution network of the aircraft in order to take on the latter the electrical power required for the electrical equipment of the engine. Alternatively, the third bus is connected to an electric generator dedicated to the motor and driven by it.

According to an alternative embodiment, turbine headset clearance adjustment systems comprise electric heating devices fed directly by the third bus.

Brief description of the drawing

The invention will be better understood on reading the description given below by way of indication but without limitation with reference to the accompanying drawings, in which:

FIG. 1 schematically illustrates an embodiment of a device for powering equipment of a gas turbine engine engine including a control device according to the invention,

FIG. 2 is a more detailed view of the variable geometry control device of FIG. 1, and

FIG. 3 illustrates an alternative embodiment of the power supply device of FIG. 1. Detailed description of an embodiment of the invention

In FIG. 1, the references 10 and 20 designate the respective perimeters of an airplane and a gas turbine engine fitted to this aircraft.

The engine 20 comprises several "variable geometries", namely:

the transient discharge valve positions of the TBV ("Transient Bleed Vanes") compressor which are controlled during particular phases of flight, in particular during take-off,

- Sets of turbine rotor blade tips (clearance between the top of rotating blades and the turbine casing), LPTACC low pressure turbine (Low Pressure Turbine Active Clearance Control) and HPTACC high pressure turbine (" High Pressure Turbine Active Clearance Control "),

the positions of compressor discharge valves VBV ("Variable Bleed Vanes") which are controlled during the entire duration of the flight, and

VSV ("Variable Stator Vane") stator vane clamping members of compressor rectifier stages.

These variable geometries are preferably all electrically controlled by means of equipment with electromechanical or electrical actuators associated with electronic control or servocontrol circuits.

In the illustrated embodiment, the electrical energy necessary for the operation of the electrical equipment of the engine is taken from the electrical distribution network of the aircraft via supply lines 12, 12 '.

The electrical energy required for the electrical distribution network of the aircraft is provided by at least one electric generator 21 driven by the engine 20. Advantageously, an electric machine is used which can function as an electric starter and then as a generator driven by the turbine of the aircraft. engine, a machine commonly referred to as S / G ("Starter / Generator"). As a redundancy, a similar generator driven by another engine of the aircraft also provides electrical power to the electrical distribution network of the aircraft, in parallel with the generator or generators 21. The electrical energy supplied is converted into the network of electrical distribution of the aircraft in an AC or DC voltage

(typically 115 or 230 Vac / 400 Hz or variable frequency, or 270 Vdc, approximately).

The supply line 12 is connected to a bus 22 which is part of the power supply circuit of the engine and receives the voltage of the aircraft's onboard network.

In the example illustrated, the voltage of the aircraft's electrical system is an AC voltage and the DC voltage distribution bus 24 is connected to the bus 22 via a circuit breaker 26 and a DC converter circuit. voltage 28. The converter circuit 28 converts the AC voltage supplied by the electrical distribution network of the aircraft into a DC voltage of lower amplitude, typically a voltage of about 28 Vdc. It is advantageous to use a secure converter 28 providing protection against short-circuits to maintain the power supply of the bus 24 in the event of a brownout of the received AC voltage. A second AC voltage distribution bus 30 is connected to the bus 22 via a circuit breaker 32.

Alternatively, the bus 24 may be an AC voltage distribution bus connected to the bus 22 via a circuit breaker and a transformer transforming the AC voltage supplied by the distribution network of the aircraft into an AC voltage. of lower amplitude.

Still alternatively, the bus 30 may be a DC voltage distribution bus connected to the bus 22 by means of a circuit breaker and a voltage converter or transformer-rectifier transforming the AC voltage supplied by the distribution network. the airplane in a continuous voltage.

In the case where the voltage of the aircraft electrical system is a DC voltage, the bus 24 is a DC voltage distribution bus connected to the bus 22 via a circuit breaker and a step-down converter. while the bus 30 is a DC voltage distribution bus connected to the bus 22 by a circuit breaker.

The bus 24 is used to provide electrical energy necessary for the operation or actuation of motor equipment requiring a relatively low power, typically less than 100 W. This equipment includes those associated with variable geometries TBV, HPTACC and LTPACC. The TBV transient discharge valves can be controlled by electromechanical actuators in the form of cylinders or electric motors. The sets of turbine rotor blade tips can be controlled by electromechanical actuators acting on valve adjusters to vary a flow of air projected on the turbine casing at the blades to adjust the geometry of the crankcase by action on its temperature.

Other electrical equipment can be powered by the bus 24, such as:

- a redundant full authority automatic engine control system or ECU ("Electronic Control Unit"), shown in the figure by two identical circuits (for redundancy) designated by "1/2 ECU", - a flow control valve Fuel Flow (FFCV)

Control Value ") of a fuel flow control circuit supplied to the engine,

- an over-speed protection valve of the OSV engine ("Over SpeedValve") of the fuel flow control circuit; - a health management system and the use of HUMS engine components ("Health and Usage Management System ") providing useful information for fault diagnosis and maintenance of engine components,

valves of a fuel flow control system supplied to injectors of the combustion chamber of the engine such as a TAPS ("Twin Annular Pre-Swirl Combustor") system.

The bus 30 is used to provide electrical power to the engine equipment requiring higher power. This equipment includes those associated with VSV and VBV variable geometry. The pitch angles of the variable stator stator vanes of the compressor stator can be controlled by means of electromechanical actuators, which can be in the form of jacks or electric motors associated with a control ring connected to the valves by articulated levers. VBV variable relief valve positions can be controlled by actuators electromechanical for example in the form of cylinders or electric motors.

The bus 30 may also be used to provide electrical power to an electric pump drive motor of a general fuel supply system of the engine, such as a gear pump GP ("Gear Pump").

Of course, the lists of equipment given above are not exhaustive.

As a redundancy, a bus 24 'similar to the bus 24 is connected by a circuit breaker 26' and a voltage converter 28 '(or transformer) to a bus 22' similar to the bus 22 and receiving the voltage of the vehicle's electrical system. plane by line 12 '. Similarly, a bus 30 'anologue bus 30 is connected to the bus 22' by a circuit breaker 32 '(and optionally a voltage converter or transformer-rectifier). The buses 24 and 24 'feed the same equipment in parallel, as well as the buses 30, 30'. The buses 24, 24 ', as well as the buses 30, 30', are electrically isolated from one another, thus avoiding the propagation of breakdown, the insulation being obtained in particular by the segregation of the lines 12 and 12 'of the network of the aircraft. The operation of some equipment simply requires a power supply. This is the case in the illustrated example of the equipment 1/2 ECU and HUMS powered in parallel by the buses 24, 24 '.

The operation of one or more other equipment simply requires the power supply of an excitation circuit. In the illustrated example, this is the case of the spark plug of the "IGNITION" circuit which is connected to an electronic excitation circuit powered in parallel by the buses 24, 24 '.

The operation of the GP electric pump requires a redundant motor and redundant motor control circuit.

The operation of one or more of the remaining equipment is controlled by an electromechanical actuator comprising drive means such as cylinder, motor or electric coil. This is the case in particular for equipment associated with variable geometries TBV, HPTACC, LPTACC, VSV, VBV and GP as already indicated. When operational safety requires it, the electromechanical actuator is doubled as redundancy. This may be the case in particular for the equipment associated with the variable geometries TBV, VSV and VBV which are shown in FIG. 2 with their redundant electromechanical actuators AEM and AEM ', as well as the pump GP with redundant electric drive motors M, M '. Each actuator is supplied in parallel by the buses 24, 24 'or the buses 30, 30'. In other cases, a single electromechanical actuator AEM may be provided, for example for the HPTACC and LPTACC equipment, this actuator being here supplied in parallel by the buses 24, 24 '. The devices with adjustable position or speed can be furthermore associated with servo-control circuits making it possible to maintain their position or real speed detected by sensor according to a position or speed of reference. This may be the case, for example, with VSV, VBV, HPTACC, LPTACC and GP equipment whose electromechanical actuators AEM, M 'and optionally AEM' M 'are controlled by respective electronic servocontrol circuits SC and SC.

At least some of them, SC and SC electronic circuits are located locally, in the vicinity or at the associated equipment, or are integrated therein. The circuits SC, SC are supplied in parallel by the buses 24, 24 'or 30, 30' and are connected to the equipment 1/2 ECU by links (not shown) to receive control information or setpoint information provided by that of the two 1/2 ECU equipment that is active. It will be noted that an electronic circuit of a device powered by an AC bus 30 (or 30 ') may receive its power supply from a DC bus 24 (or 24').

However, it is not excluded that the functions of one or more electronic circuits SC, SC are implanted in equipment 1/2 ECU by providing appropriate links between them and the equipment, motors or actuators concerned.

In the embodiment of FIGS. 1 and 2, it has been considered that the functions LPTACC and HPTACC are provided by air flow control coming to impact on turbine ring sectors to control dimensional variations by action on the temperature. ring areas. These functions may alternatively be provided, in a manner known per se, by electric heating systems associated with the ring sectors. This variant is illustrated in FIG. 3, which differs from FIG. 1 in that the LTPACC and HTPACC systems are supplied directly in parallel by the buses 22, 22 'with the interposition of circuit breakers 23, 23' and 25, 25 '. Switching circuits (not shown) are associated with the LTPACC and HTPACC systems and are controlled by the Vi ECUs to control the power supply via the buses 22, 22 'or the interruption of this power supply. In the embodiment illustrated in FIG. 1, the electrical energy required for the electrical equipment of the motor is taken from the electrical distribution network of the aircraft. This does not have a significant drawback when the power available on the electrical distribution network of the aircraft is important to cope with the increasing need for electrical energy for the equipment of the aircraft, the power required for the engine does not representing a small part of this power.

As a variant, it is however possible to directly supply the bus 22 (and the bus 22 ') from a dedicated redundant generator specific to the motor and driven by the latter, preferably a redundant generator.

Claims

1. Device for controlling variable geometries of a gas turbine engine, said variable geometries notably comprising variable valve timing angles of the stators, positions of compressor discharge valves and games at the vertices of turbine blades, wherein said variable geometries are controlled by electromechanical or electrical actuators using a source of electrical energy, the electromechanical or electrical actuators being associated with electronic control circuits at least a part of which are locally installed at level of the corresponding electromechanical or electrical actuators.

2. Device according to claim 1, wherein at least one of the variable geometries is associated with at least one electromechanical or electrical actuator with redundant controls, and the power supply circuit for actuating this variable geometry comprises two buses. redundant electrical power distribution, each bus being connected in parallel with this actuator.

3. Device according to claim 2, wherein the redundant buses are electrically isolated from each other.

4. Device according to claim 1, comprising a power supply circuit of the electromechanical actuators of variable geometries, which circuit comprises at least a first continuous or alternative electrical energy distribution bus to the first equipment of the engine requiring a lower electrical power. at a given power threshold and at least a second continuous or alternative electric power distribution bus to other electrical equipment of the engine, the first and the second bus being connected to a third bus which is itself connected to a source of power. 'electric energy.

5. Device according to claim 4, wherein the first bus feeds actuators for compressor transient discharge valves and for clearance adjustments to the turbine rotor blade tips.

6. Device according to any one of claims 3 and 4, wherein the second bus feeds actuators for valves of discharge and for compressor stator vane timing members.

7. Device according to any one of claims 4 to 6, wherein the second bus feeds an electric motor fuel pump.

8. Device according to any one of claims 4 to 7, wherein the first bus is connected to the third bus via at least one voltage converter or a transformer.

9. Device according to claim 4, wherein gaming systems at the top of turbine blades comprise electric heaters powered directly by the third bus.

10. Device according to any one of claims 3 to 9, wherein the second bus is connected to an electrical distribution network of the aircraft.

11. Device according to any one of claims 3 to 9, wherein the third bus is connected to an electric generator dedicated to the motor and driven by it.