IGNITION SYSTEM FOR A COMBUSTION CHAMBER OF A
TURBOSHAFT ENGINE
1. Technical field of the invention
The invention relates to a system for igniting a combustion chamber of a
turboshaft engine. The invention relates in particular to a system for igniting a
combustion chamber of a turboshaft engine which is capable of being put into a
standby mode and of being quickly reactivated if needed.
2. Technological background
As is known, a twin-engine or three-engine helicopter has a propulsion
system comprising two or three turboshaft engines, each turboshaft engine
comprising a gas generator and a free turbine which is rotated by the gas
generator and is rigidly connected to an output shaft. The output shaft of each free
turbine is suitable for inducing the movement of a power transmission unit, which
15 itself drives the rotor of the helicopter. The gas generator comprises a combustion
chamber into which injectors for fuel supplied by a supply circuit lead.
It is known that, when the helicopter is in a cruise flight situation (i.e.
when it is progressing in normal conditions, during all flight phases apart from
transitional phases of take-off, ascent, landing or hovering flight), the turboshaft
20 engines develop low power levels, below their maximum continuous output.
These low power levels give rise to a specific consumption (hereinafter SC),
defined as the ratio between the hourly consumption of fuel by the combustion
chamber of the turboshaft engine and the mechanical power supplied by this
turboshaft engine, of greater than approximately 30% of the SC of the maximum
25 take-off power, and they therefore give rise to overconsumption of fuel in cruising
flight.
Furthermore, the turbo shaft engines of a helicopter are designed so as to be
oversized in order to be able to keep the helicopter in flight in the event of failure
of one of the engines. This flight situation occurs following the loss of an engine
30 and results in each operating engine supplying a power level much beyond its
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nominal power to allow the helicopter to deal with a hazardous situation, and then
to be able to continue its flight.
The turboshaft engines are also oversized so as to be able to ensure flight
over the entire flight range specified by the aircraft manufacturer, and in particular
5 flight at high altitudes and during hot weather. These flight points, which are
highly demanding, in particular when the helicopter has a weight close to its
maximum take-off weight, are encountered only in certain circumstances of use.
These oversized turboshaft engines are disadvantageous in terms of weight
and fuel consumption. In order to reduce this consumption in cruising flight, it is
1 0 envisaged to put at least one of the turbo shaft engines on standby in flight. The
active engine or engines then operate at higher power levels in order to provide all
the necessary power, and therefore at more favourable SC levels.
Putting a turboshaft engine on standby requires the provision of a rapid
reactivation system which makes it possible to quickly take the turboshaft engine
15 out of the standby state if needed. This need may arise, for example, when one of
the active engines fails or if the flight conditions deteriorate unexpectedly,
meaning that the total power is required once again.
The applicant has therefore sought to optimise the system for igniting a
combustion chamber of a turboshaft engine so as to be able in particular to
20 quickly reactivate the turboshaft engine when it is on standby and when the flight
conditions mean that the total available power is required once again.
As is known, a system for igniting a combustion chamber of a turboshaft
engine of a helicopter comprises start7up injectors intended for initiating
combustion and main injectors intended for maintaining the combustion once it
25 has been initiated. It is known that main injectors are supplied with fuel by a main
circuit and the start-up injectors are supplied with fuel by a start-up circuit, which
is separate from the main circuit. A known ignition system makes it possible to
initiate combustion by means of start-up injectors associated with at least one
start -up spark plug suitable for providing the spark for setting alight the mixture
30 of air and fuel in the combustion chamber. The flame then spreads from the startup
injectors towards the main injectors.
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When designing an ignition system for a turboshaft engine, engineers have
to choose between using a large number of start-up injectors, which allows the
flame to spread rapidly towards the main injectors but means that it takes longer
for the fuel to be conveyed to all of the injectors, and using a small number of
5 start-up injectors, which allows fuel to be conveyed to the start-up injectors more
quickly but means that it takes longer for the flame to spread towards the main
injectors.
The inventors have therefore sought to propose a solution which makes it
possible for the flame to spread rapidly from the start-up injectors towards the
10 main injectors, while at the same time allowing the start-up injectors to be quickly
filled with fuel.
In other words, the inventors have sought to reconcile the two alternatives
which are, in principle, incompatible.
The inventors have also sought to provide an ignition system having
15 improved reliability compared with known systems, in order to improve the safety
of helicopters provided with hybrid turboshaft engines capable of being put into
standby mode.
3. Aims of the invention
The invention aims to provide a system for igniting a combustion chamber
20 of a turboshaft engine and makes it possible to quickly ignite the combustion
chamber, while allowing the turboshaft engine to be reactivated quickly.
The invention also aims to provide an ignition system which combines the
advantages of the flame spreading rapidly from the start-up injectors towards the
main injectors and of the start-up injectors being filled up quickly.
25 The invention also aims to provide an ignition system which has improved
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reliability by comparison with systems from the prior art.
The invention also aims to provide a turboshaft engine provided with an
ignition system according to the invention.
4. Disclosure of the invention
In order to achieve this, the invention relates to a system for igniting a
combustion chamber of an aircraft turbo shaft engine, comprising:
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a plurality of start-up injectors which lead into said combustion
chamber and are suitable for injecting fuel into said chamber during
a combustion-initiating phase,
a circuit for supplying fuel to said start-up injectors, referred to as
the start-up circuit,
a plurality of main injectors which lead into said combustion
chamber and are suitable for injecting fuel into said combustion
chamber so as to maintain the combustion once said combustion has
been initiated by said start-up injectors.
The ignition system according to the invention is characterised in that the
start-up circuit comprises:
a first sub-circuit, referred to as the primary start-up circuit,
designed to supply fuel to some of said plurality of start -up
injectors, referred to as the primary start-up injectors,
a second sub-circuit, referred to as the secondary start-up circuit,
designed to supply fuel to the other start-up injectors of said
plurality, referred to as the secondary start-up injectors.
The ignition system is also characterised in that said primary start-up
circuit and said secondary start-up circuit each comprise a solenoid start-up valve
20 suitable for being controlled by a control unit so as to allow or prevent the supply
of fuel to said primary and secondary start-up injectors, respectively.
An ignition system according to the invention therefore comprises two
separate start-up circuits, namely one primary circuit intended for supplying fuel
to primary start-up injectors and one secondary circuit intended for supplying fuel
25 to secondary start-up injectors. Furthermore, each circuit is provided with a
solenoid valve controlled by a control unit for allowing or preventing the supply
of fuel to the injectors. An ignition system according to the invention may
therefore comprise a large number of start-up injectors, and yet without having the
disadvantage of it taking a long time to fill up the injectors, since said injectors
30 are distributed across two separate supply circuits.
Furthermore, an ignition system according to the invention is more reliable
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than the systems from the prior art as a result of being provided with two separate
start-up circuits. Moreover, if a solenoid valve of one of the start-up circuits fails,
the other circuit can take over and ensure that the turboshaft engine is reactivated.
An ignition system of this kind is therefore particularly suitable for hybrid
5 turboshaft engines capable of being put into a standby mode during flight, on
account of having improved reliability which makes it possible to guarantee that
the turboshaft engine is reactivated if needed.
Advantageously and according to the iqvention, the solenoid valves are
controlled by the control unit using a sequential or simultaneous procedure, the
10 procedure being selected according to the flight conditions of said aircraft.
The flight conditions of the aircraft, for example a helicopter, include for
example the ambient temperature, ambient pressure, rotational speed of the gas
generator of the turbo shaft engine, etc. These different parameters are used by the
control unit to defme which procedure is the best to implement in order to start up
15 the turboshaft engine, taking account of the flight conditions, from either a
simultaneous start-up procedure for the two start-up circuits or a sequential startup
procedure for the two circuits.
Advantageously and according to the invention, said solenoid valves are
controlled by the control unit such that, on the ground, each start-up circuit is used
20 alternately for each flight so as to limit dormancy of a possible failure to a single
flight.
According to this advantageous variant, the ignition system is designed
such that, on the ground, the turbine is started alternately for each flight in a single
start-up circuit. This makes it possible to limit the dormancy of a possible failure
25 to a single flight.
Advantageously and according to the invention, each start-up injector is
associated with a rail for supplying fuel to said injector, said supply rail of a
primary start-up injector having a lower volume than said supply rail of a
secondary start-up injector so as to be able to be filled up with fuel more quickly.
30 According to this advantageous variant, the primary and secondary circuits
are different from one another. The primary circuit has injectors having a filling
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rail of a reduced volume by comparison with the secondary injectors. Therefore,
the primary injectors can be quickly filled up with fuel and can quickly initiate
combustion in the combustion chamber. The secondary injectors continue the
combustion and can, in combination with the primary injectors, ensure that the
5 flame spreads towards the main injectors once the combustion has been initiated.
Advantageously, an ignition system according to the invention comprises
one spark plug opposite each start-up injector, which spark plug is suitable for
providing a spark for setting alight the fuel in said combustion chamber.
A spark plug being opposite each start-up injector, i.e. both primary and
10 secondary start-up injectors, makes it possible to speed up the combustion and the
spreading of the flame towards the main injectors.
Advantageously, an ignition system according to the invention comprises
two primary start-up injectors and two secondary start-up injectors.
An ignition system according to the invention, according to one or the
15 other advantageous variants described, is particularly intended for being fitted in a
hybrid turboshaft engine capable of being put into a standby mode, so as to be
able to reactivate said engine if needed.
When the helicopter is on the ground, the primary and secondary start-up
circuits are tested independently of one another so as to check the integrity thereof
20 and allow the hybrid turboshaft engine to be put on standby during flight.
When the helicopter is in cruise flight, the hybrid turboshafl engine can
therefore be put on standby.
An ignition system according to the invention can also be designed such
that, on the ground, the turbine is started alternately for each flight in a single
25 start-up circuit. This makes it possible to limit dormancy of a possible failure to a
single flight.
If the flight conditions require the turboshaft engine to be reactivated in the
normal manner, for example because the helicopter is going to transition from a
cruise flight phase to a landing phase, the ignition system according to the
30 invention is used by controlling the two start-up circuits, namely the primary startup
circuit and the secondary start-up circuit, and the different power supply paths
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of the spark plugs. The pnmary and secondary circuits can be controlled
simultaneously or sequentially. Normal reactivation of the hybrid turboshaft
engine is reactivation which occurs 1 0 seconds to 1 minute, in particular
30 seconds to 1 ruinute, after the reactivation command.
5 If the flight conditions require the turboshaft engine to be reactivated
quickly, for example because one of the active turboshaft engines suddenly fails,
the ignition system according to the invention is used by consecutively controlling
the primary start-up circuit and then the secondary start-up circuit once it has been
detected that the chamber is ignited. According to another variant, the primary
l 0 and secondary circuits are controlled simultaneously.
The invention also relates to a turboshaft engine comprising a combustion
chamber, characterised in that said engine comprises an ignition system according
to the invention.
The invention also relates to an aircraft, in particular a helicopter,
15 comprising at least one turboshaft engine according to the invention.
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The invention also relates to an ignition system, to a turboshaft engine and
to an aircraft, characterised in combination by all or some of the features
mentioned above or below.
5. List of figures
Other aims, features and advantages of the invention will emerge from
reading the following description, which is given purely by way of non-limiting
example and relates to the accompanying Fig. 1, which is a schematic view of an
ignition system according to an embodiment of the invention.
6. Detailed description of an embodiment of the invention
In the figure, the scales and proportions are not respected for the sake of
illustration and clarity.
Fig. 1 is a schematic view of a system for igniting a combustion chamber 2
of a turboshaft engine.
The system comprises start-up injectors 21a, 21b, 31a, 31b which lead into
the combustion chamber 2 and are suitable for injecting fuel into the chamber 2
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during a combustion-initiating phase ..
The system also comprises mam injectors 12 which lead into the
combustion chamber 2 and are suitable for injecting fuel into the chamber 2 at a
higher flow rate once combustion has been initiated.
5 The combustion chamber 2 is shown schematically by a rectangle in Fig. 1
for the sake of clarity. In practice, the combustion chamber generally comprises
two annular walls, namely an outer wall and an inner wall, which extend one
inside the other and are connected by an annular. bottom wall of the chamber. The
fuel injectors are distributed over the entire circumference of the combustion
10 chamber.
The system also comprises a circuit for supplying fuel to the main injectors
12, referred to as the main circuit 5, and a circuit for supplying fuel to the start-up
injectors 21, 31, referred to as the start-up circuit 6.
These two circuits are connected to a fuel inlet 7 which is supplied with
15 fuel by a pump designed to withdraw fuel from a fuel reservoir (not shown in Fig.
1).
According to the invention, the start-up circuit 6 for supplying fuel to the
start-up injectors 21, 31 is formed of two sub-circuits, namely a first sub-circuit,
referred to as the primary start -up circuit 20, which is designed to supply fuel to
20 the injectors 21, referred to as the primary start-up injectors, and a second subcircuit,
referred to as the secondary start-up circuit 30, which is designed to
supply fuel to the start-up injectors 31, referred to as the secondary start-up
injectors.
The pnmary start-up circuit 20 also compnses a solenoid valve 22
25 controlled for example by the engine electronic control unit (better known by the
acronym EECU) of the helicopter. The secondary start-up circuit 30 also
comprises a solenoid valve 32 controlled by the EECU. The solenoid valve 22 is
designed to allow or prevent the supply of fuel to the primary start-up injectors
21. The solenoid valve 32 is designed to allow or prevent the supply of fuel to the
3 0 primary start-up injectors 31.
The primary start-up injectors 21 have fuel supply rails that have a volume
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that is smaller than the volume of the rails for supplying fuel to the secondary
start -up injectors 31. This means that, when the solenoid valves are open, the
primary injectors 21 are quickly activated and initiate combustion in the
combustion chamber 2. The secondary injectors 31 continue the combustion once
5 the corresponding rails are filled, and this process takes slightly longer for said
secondary injectors than for the primary injectors owing to said secondary
injectors having a larger volume.
Once the start-up injectors 21, 31 are active, the combustion in the
combustion chamber is maintained by the activation of the injectors 12 of the
10 main circuit combined with the spreading of the flame from the start-up injectors
31, 21 to the main injectors 12. Once the main injectors 12 have taken over from
the start-up injectors 21, 31, the primary and secondary start-up circuits are bled
and the fuel residue is discharged to a collector via channels 25, 35. Bleeding the
start-up injectors after they have stopped supplying fuel makes it possible to avoid
15 coking (carbonisation of the fuel in the pipes) and therefore prevents the injectors
from becoming clogged.
According to the embodiment of Fig. 1, each start-up injector 21a, 2lb,
31a, 31 b is associated with a spark plug 23a, 23b, 33a, 33b arranged opposite the
injector. Each spark plug 23a, 23b, 33a, 33b is supplied with electricity from an
20 electrical circuit 24, 34 comprising a high-voltage electrical power source. Each
spark plug is designed to produce a spark that sets alight the mixture of air and
fuel in the combustion chamber 2.
There being one spark plug per start-up injector makes it possible to
reduce the time taken for the flame to spread towards the main injectors, and
25 therefore to ultimately reduce the start-up time of the turboshaft engine provided
with an ignition system of this kind.
The invention is not limited to the described embodiment. In particular,
according to other embodiments, the ignition system may comprise more than
four start-up injectors and/or a different number of primary start-up injectors and
30 secondary start-up injectors.

CLAIMS
1. System for igniting a combustion chamber (2) of an aircraft turboshaft engine,
comprising:
a plurality of start-up injectors (21a, 21b, 31a, 31 b) which lead into
said combustion chamber (2) and are suitable for injecting fuel into
said chamber (2) during a combustion-initiating phase,
a circuit for supplying fuel to said start-up injectors (21a, 21b, 31a,
31 b), referred to as the start-up circuit ( 6),
a plurality of main injectors (12) which lead into said combustion
chamber (2) and are suitable for injecting fuel into said combustion
chamber (2) so as to maintain the combustion once said
combustion has been initiated by said start-up injectors (2la, 2lb,
31a, 3lb),
characterised in that said start-up circuit (6) comprises:
a first sub-circuit, ·referred to as the primary start-up circuit (20),
designed to supply fuel to some of said plurality of start -up
injectors, referred to as the primary start-up injectors (2la, 2lb),
a second sub-circuit, referred to as the secondary start-up circuit
(30), designed to supply fuel to the other start-up injectors of said
plurality, referred to as the secondary start-up injectors (31a, 31b),
and in that said primary start-up circuit (20) and said secondary start-up circuit
(30) each comprise a solenoid start-up valve (22, 32) suitable for being controlled
25 by a control unit so as to allow or prevent the supply of fuel to said primary and
secondary start-up injectors (2la, 2lb, 31a, 31b), respectively.
2. Ignition system according to claim 1, characterised in that said solenoid valves
(22, 32) are controlled by said control unit using a sequential or simultaneous
procedure, the procedure being selected according to the flight conditions of
30 said aircraft.
3. Ignition system according to either claim 1 or claim 2, characterised in that
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said solenoid valves (22, 32) are controlled by said control unit such that, on
the ground, each start-up circuit is used alternately for each flight so as to limit
dormancy of a possible failure to a single flight.
4. System according to any of claims 1 to 3, characterised in that each start-up
5 injector (2la, 2lb, 3la, 3lb) is associated with a rail for supplying fuel to said
injector, said supply rail of a primary start-up injector having a lower volume
than said supply rail of a secondary start-up injector so as to be able to be
filled up with fuel more quickly.
5. System according to any of claims 1 to 4, characterised in tbat it comprises
10 one spark plug (23a, 23b, 33a, 33b) opposite each start-up injector, which
spark plug is suitable for providing a spark for setting alight the fuel in said
combustion chamber (2).
6. System according to any of claims 1 to 5, characterised in that it comprises
two primary start-up injectors (2la, 2lb) and two secondary start-up injectors
15 (3la, 3lb).
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7. Turboshaft engine comprising a combustion chamber, characterised in that
said engine comprises a system for igniting said combustion chamber
according to any of claims 1 to 6.
8. Aircraft comprising at least one turboshaft engine according to claim 7.