EP3207224A1

EP3207224A1 - Ignition system for a combustion chamber of a turbo engine

Info

Publication number: EP3207224A1
Application number: EP15791692.5A
Authority: EP
Inventors: Romain Thiriet; Jean-Michel Bazet; Guillaume COTTIN; Camel SERGHINE; Patrick Marconi; Bertrand MOINE; Vincent POUMAREDE
Original assignee: Safran Helicopter Engines SAS
Current assignee: Safran Helicopter Engines SAS
Priority date: 2014-10-13
Filing date: 2015-10-06
Publication date: 2017-08-23
Also published as: CA2963837A1; US20170292491A1; WO2016059319A1; KR20170067770A; JP2017532491A; FR3027059B1; RU2017113350A; FR3027059A1; CN106795777A

Abstract

The invention relates to an ignition system of a combustion chamber (2) of a turbo engine, comprising: a plurality of start-up injectors (21a, 21b, 31a, 31b) capable of injecting fuel into said chamber (2) during a combustion initiation phase; a fuel supply circuit (6) of said start-up injectors, comprising a first sub-circuit, referred to as the primary start-up circuit (20), configured to supply fuel to a part of said plurality of start-up injectors; and a second sub-circuit, referred to as the secondary start-up circuit (30), configured to supply fuel to the remainder of said plurality of start-up injectors.

Description

IGNITION SYSTEM OF A COMBUSTION CHAMBER

FROM A TURBOMOTEUR

1. Technical field of the invention

The invention relates to a system for igniting a combustion chamber of a turbine engine. The invention particularly relates to the ignition system of a combustion chamber of a turbine engine capable of being put in a standby mode and to be reactivated quickly if necessary.

2. Technological background

A twin-engine helicopter or three-engine helicopter present in a known manner a propulsion system comprising two or three turbine engines, each turbine engine comprising a gas generator and a free turbine rotated by the gas generator, and integral with an output shaft. The output shaft of each free turbine is adapted to set in motion a power transmission box, which drives itself the rotor of the helicopter. The gas generator comprises a combustion chamber into which fuel injectors fed by a supply circuit.

It is known that when the helicopter is in a cruising flight (that is to say, when it is operating under normal conditions, during all phases of flight, except for transitional phases of take-off, climb, landing or hovering), the turboshaft engines develop low powers lower than the maximum continuous powers. These low power levels result in a specific consumption (hereinafter Cs), defined as the ratio between the hourly fuel consumption by the turbine engine combustion chamber and the mechanical power supplied by this turbine engine, which is in the order of 30 % Cs of the maximum takeoff power, and therefore overconsumption of fuel in cruising flight.

In addition, the turboshaft engines of a helicopter are designed oversized to be able to keep the helicopter in flight in case of failure of one of the engines. This flight situation arises following the loss of an engine and This translates into the fact that each running engine provides power well beyond its rated power to allow the helicopter to cope with a perilous situation and then to be able to continue its flight.

The turboshaft engines are also oversized to be able to ensure the flight in all the flight range specified by the aircraft manufacturer and in particular the flight at high altitudes and in hot weather. These flight points, very restrictive, especially when the helicopter has a mass close to its maximum take-off weight, are encountered only in certain cases of use.

These oversized turboshaft engines are penalizing in terms of weight and fuel consumption. In order to reduce this consumption during cruising flight, it is envisaged to put at least one of the turboshaft engines in flight. The active engine (s) then operate at higher power levels to provide all the power required and therefore at more favorable Cs levels.

The standby of a turbine engine requires to have a quick reactivation system that allows the turbine engine to quickly get out of the standby state if necessary. This need may for example result from a failure of one of the active engines or unexpected degradation of flight conditions requiring a quick return of full power.

The applicant has therefore sought to optimize the ignition system of a combustion chamber of a turbine engine to be able to quickly reactivate the turbine engine when idle and the flight conditions require a quick return to full power available .

A ignition system of a combustion chamber of a turbine engine of a helicopter comprises, in known manner, starter injectors intended to initiate combustion and main injectors intended to maintain combustion, once this has been initiated. It is known that the main injectors are fueled by a main circuit and the starter injectors are fueled by a starting circuit, separate from the main circuit. A known ignition system makes it possible to initiate combustion by means of the starter injectors associated with at least one spark plug of starting adapted to provide the spark for igniting the mixture of air and fuel present in the combustion chamber. This results in a propagation of the flame of the starter injectors to the main injectors.

When designing a turbine engine ignition system, engineers must make a choice between using a large number of starter injectors, which allows a rapid propagation of the flame to the main injectors, but a slower delivery of fuel to all injectors, and use of a small number of starter injectors, allowing faster fuel delivery to starter injectors, but slower flame spread to injectors key.

The inventors have therefore sought to propose a solution that can both allow rapid propagation of the flame of the starter injectors to the main injectors and rapid filling of the starter injectors.

In other words, the inventors sought to reconcile the two alternatives that are a priori incompatible with each other.

The inventors have also sought to provide an ignition system whose reliability is improved compared to known systems to improve the safety of helicopters equipped with hybrid turbine engines capable of being put in standby mode.

3. Objectives of the invention

The invention aims to provide an ignition system of a combustion chamber of a turbine engine which allows rapid ignition of the combustion chamber compatible with rapid reactivation of the turbine engine.

The invention also aims at providing an ignition system that combines the advantages of a rapid flame spread from the starter injectors to the main injectors and a fast filling of the starter injectors.

The invention also aims to provide an ignition system which has improved reliability compared to the systems of the prior art. The invention also aims to provide a turbine engine equipped with an ignition system according to the invention.

4. Presentation of the invention

To this end, the invention relates to a system for igniting a combustion chamber of a turbine engine of an aircraft comprising:

a plurality of starter injectors opening into said combustion chamber and adapted to inject fuel into said chamber during a combustion initiation phase,

a fuel supply circuit for said starting injectors, called a starting circuit,

- A plurality of main injectors opening into said combustion chamber and adapted to inject fuel into said combustion chamber to maintain combustion, once the combustion initiated by said starter injectors.

The ignition system according to the invention is characterized in that the starting circuit comprises:

a first sub-circuit, called a primary starting circuit, configured to supply fuel to a portion of said plurality of starter injectors, called primary starter injectors,

a second sub-circuit, called a secondary start-up circuit, configured to supply fuel to the other part of said plurality of start injectors, called secondary start injectors.

The ignition system is also characterized in that said primary starting circuit and said secondary starting circuit each comprise a starter electro-valve adapted to be controlled by a control unit so as to allow or prohibit respectively the supply of power. fuel of said primary and secondary starter injectors.

An ignition system according to the invention therefore comprises two separate starting circuits, a primary circuit for supplying primary starter injectors and a secondary circuit intended to feed injectors secondary startup. In addition, each circuit is equipped with an electro-valve controlled by a control unit to allow or prohibit the supply of injectors. An ignition system according to the invention may therefore comprise a large number of starter injectors, without however having the disadvantage of a slow filling of the injectors because these injectors are distributed in two separate supply circuits.

In addition, an ignition system according to the invention has a better reliability than the systems of the prior art by providing two separate starting circuits. Also, in case of failure of an electro-valve of one of the starting circuits, the other circuit can take over and ensure the reactivation of the turbine engine. Such an ignition system is particularly suitable for hybrid turboshaft engines capable of being placed in a standby mode during the flight, with improved reliability to ensure reactivation of the turbine engine when necessary.

Advantageously and according to the invention, the electro-valves are controlled by the control unit according to a sequential or simultaneous procedure, the choice of the procedure depending on the flight conditions of said aircraft.

The flight conditions of the aircraft, for example a helicopter, comprise, for example, the ambient temperature, the ambient pressure, the speed of rotation of the gas turbine engine, and so on. These various parameters are used by the control unit to define the best start-up procedure for the turbine engine to be implemented, taking into account the flight conditions, ie a simultaneous start-up procedure for the two starting circuits, or a sequential procedure of start of both circuits.

Advantageously and according to the invention, said electro-valves are controlled by the control unit so that on the ground, each starting circuit is used alternately at each flight so as to limit a dormancy of a possible breakdown to a single flight.

According to this advantageous variant, the ignition system is configured so that on the ground, the turbine is started alternately with each flight, on a single start circuit. This makes it possible to limit the dormancy of a possible breakdown to a single flight.

Advantageously and according to the invention, each starter injector is associated with an injector fuel supply ramp, said supply ramp of a primary starting injector having a volume less than said supply ramp of a secondary injection nozzle so that it can be filled more quickly with fuel.

According to this advantageous variant, the primary and secondary circuits are different from each other. The primary circuit has injectors with a filling ramp of reduced volume compared to the secondary injectors. Also, the primary injectors can be quickly filled with fuel and quickly initiate combustion of the combustion chamber. The secondary injectors continue the combustion and can, in combination with the primary injectors, ensure the propagation of the flame towards the main injectors, once the combustion initiated.

Advantageously, an ignition system according to the invention comprises a candle adapted to provide a spark for igniting the fuel present in said combustion chamber facing each starter injector.

The presence of a spark plug next to each starter injector - primary and secondary - accelerates the combustion and propagation of the flame to the main injectors.

Advantageously, an ignition system according to the invention comprises two primary starter injectors and two secondary starter injectors.

An ignition system according to the invention, according to one or other of the advantageous variants described, is particularly intended to equip a hybrid turbine engine capable of being put in a standby mode to allow to reactivate it if necessary.

When the helicopter is on the ground, the primary and secondary start circuits are tested independently of each other to verify their integrity. and allow in flight the standby of the hybrid turbine engine.

During a cruise flight of the helicopter, the hybrid turbine engine can be put on standby.

An ignition system according to the invention can also be configured so that on the ground, the turbine is started alternately at each flight, on a single starting circuit. This makes it possible to limit the dormancy of a possible breakdown to a single flight.

If the flight conditions require a normal reactivation of the turbine engine, for example because the helicopter will go from a cruising flight phase to a landing phase, the ignition system according to the invention is used by piloting the engines. two primary and secondary start-up circuits and the different candle supply routes. The control of the primary and secondary circuits can be done simultaneously or sequentially. Normal reactivation of the hybrid turbine engine is a reactivation that occurs within 10 seconds to 1 minute, especially 30 seconds to 1 minute, after the reactivation command.

If the flight conditions require rapid reactivation of the turbine engine, for example because one of the active turbine engines suddenly fails, the ignition system according to the invention is used by successively controlling the primary starting circuit, then the circuit secondary start as soon as the ignition of the chamber is detected. According to another variant, the control of the primary and secondary circuits is simultaneous.

The invention also relates to a turbine engine comprising a combustion chamber, characterized in that it comprises an ignition system according to the invention.

The invention also relates to an aircraft - in particular a helicopter - comprising at least one turbine engine according to the invention.

The invention also relates to an ignition system, a turbine engine and an aircraft, characterized in combination by all or some of the characteristics mentioned above or below. 5. List of figures

Other objects, features and advantages of the invention will become apparent on reading the following description given solely by way of non-limiting example and which refers to the appended FIG. 1, which is a schematic view of an ignition system according to a embodiment of the invention.

6. DETAILED DESCRIPTION OF AN EMBODIMENT OF THE INVENTION In the figure, the scales and the proportions are not respected for purposes of illustration and clarity.

Figure 1 shows schematically an ignition system of a combustion chamber 2 of a turbine engine.

The system comprises starter injectors 21a, 21b, 31a, 31b which open into combustion chamber 2 and which are adapted to inject fuel into chamber 2 during a combustion initiation phase.

The system also includes main injectors 12 which open into the combustion chamber 2 and which are adapted to inject fuel into the chamber 2 with a higher flow rate once the combustion is initiated.

In the figure, the combustion chamber 2 is schematically represented by a rectangle in FIG. 1, for the sake of clarity. In practice, the combustion chamber generally comprises two outer and inner annular walls extending one inside the other and connected by an annular wall of the chamber bottom. The fuel injectors are distributed over the entire circumference of the combustion chamber.

The system also comprises a fuel supply circuit of the main injectors 12, said main circuit 5, and a fuel supply circuit of the injectors 21, 31 of startup, said circuit 6 of startup.

These two circuits are connected to a fuel inlet 7 which is supplied with fuel by a pump configured to draw fuel into a fuel tank, not shown in FIG. 1.

According to the invention, the fuel supply start circuit 6 of the injectors 21, 31 for starting is formed of two sub-circuits, a first subcircuit, said primary circuit 20 of startup, configured to supply fuel injectors 21, said primary start injectors, and a second sub-circuit, said secondary circuit 30 of startup, configured to supply fuel injectors 31 start , called secondary start injectors.

The primary starting circuit 20 further comprises an electro-valve 22 controlled for example by the electronic control computer of the helicopter, better known by the acronym EECU. The secondary start circuit 30 also includes an electro-valve 32 controlled by the EECU. The electro-valve 22 is configured to allow or prohibit the fuel supply of the primary starter injectors. The electro-valve 32 is configured to allow or prohibit the fuel supply of the primary starter injectors 31.

The primary starter injectors 21 have fuel supply ramps of a volume smaller than the volume of the fuel supply ramps of the secondary starter injectors 31. As a result, at the opening of the electro-valves, the primary injectors 21 will be able to quickly start and initiate combustion in the combustion chamber 2. The secondary injectors 31 continue combustion as soon as the corresponding ramps are filled, which, taking into account their larger volume, takes a little longer than for the primary injectors.

Once the active injectors 21, 31 start, the combustion of the combustion chamber is maintained by the actuation of the injectors 12 of the main circuit combined with a propagation of the flame of the injectors 31, 21 of the main 12 injectors. Once the main injectors 12 have taken over the starting injectors 21, 31, the primary and secondary start circuits are purged and the fuel residues are expelled to a manifold via the lines 25, 35. The purge starter injectors after stopping their supply avoids coking (carbonization of the fuel in the pipes) and thus the clogging of the injectors. According to the embodiment of FIG. 1, each starting injector 21a, 21b, 31a, 31b is associated with a spark plug 23a, 23b, 33a, 33b arranged facing the injector. Each spark plug 23a, 23b, 33a, 33b is supplied with electricity from an electric circuit 24, 34 comprising a source of high voltage electrical energy. Each spark plug is configured to produce a spark igniting the air and fuel mixture of the combustion chamber 2.

The presence of a spark plug by starting injector makes it possible to reduce the propagation time of the flame towards the main injectors, and thus ultimately to reduce the starting time of the turbine engine equipped with such a starter system.

The invention is not limited to the embodiment described. In particular, according to other embodiments, the ignition system may comprise more than 4 starter injectors and / or a different number of primary start injectors and secondary start injectors.

Claims

A system for igniting a combustion chamber (2) of a turbine engine of an aircraft comprising:

a plurality of starter injectors (21a, 21b, 31a, 31b) opening into said combustion chamber (2) and adapted to inject fuel into said chamber (2) during a combustion initiation phase,

a fuel supply circuit of said starter injectors (21a, 21b, 31a, 31b), said start circuit (6), a plurality of main injectors (12) opening into said combustion chamber (2) and adapted to injecting fuel into said combustion chamber (2) to maintain combustion, once the combustion initiated by said starter injectors (21a, 21b, 31a, 31b),

characterized in that said starting circuit (6) comprises:

a first sub-circuit, said primary circuit (20) for starting, configured to supply fuel to a portion of said plurality of starter injectors, called primary starter injectors (21a, 21b),

a second sub-circuit, said secondary circuit (30) for starting, configured to supply fuel to the other part of said plurality of starter injectors, said secondary injectors (31a, 31b),

and in that said primary circuit (20) for starting and said secondary circuit (30) for starting each comprise a starter electro-valve (22, 32) adapted to be controlled by a control unit so as to allow or prohibit respectively fueling said primary and secondary starter injectors (21a, 21b, 31a, 31b).

2. Ignition system according to claim 1, characterized in that said electro-valves (22, 32) are controlled by said control unit according to a sequential or simultaneous procedure, the choice of the procedure depending on the flight conditions of said aircraft.

3. Ignition system according to one of claims 1 or 2, characterized in that said electro-valves (22, 32) are controlled by said control unit so that on the ground, each starting circuit is used alternately on each flight so as to limit a dormancy of a possible breakdown to a single flight.

4. System according to one of claims 1 to 3, characterized in that each injector (21a, 21b, 31a, 31b) is associated with a fuel supply ramp of the injector, said supply ramp a primary starter injector having a volume smaller than said supply ramp of a secondary starter injector so that it can be filled more quickly with fuel.

5. System according to one of claims 1 to 4, characterized in that it comprises a spark plug (23a, 23b, 33a, 33b) adapted to provide a spark for igniting the fuel present in said chamber (2) of combustion opposite each starter injector.

6. System according to one of claims 1 to 5, characterized in that it comprises two injectors (21a, 21b) of primary start and two injectors (31a, 31b) secondary start.

7. Turbomotor comprising a combustion chamber, characterized in that it comprises an ignition system of said combustion chamber according to one of claims 1 to 6.

8. Aircraft comprising at least one turbine engine according to claim 7.