System and method for the emergency starting
of an aircraft turbomachine
TECHNICAL FIELD
5 The present invention relates to an emergency start system and to
an emergency start method for a turbine engine of an aircraft, and in
particular of a helicopter.
PRIOR ART
In the case of a twin helicopter (FR 2 967 132 and FR 2 967 133),
)
10 critical situations can arise when one of the engines is voluntarily shut
down. Indeed, this mode is recommended to minimise consumption during
search phases and en-route phases. In this context, two exceptional
situations can occur, which then require an emergency restart of the shutdown
engine:
15 - the one active engine stops or slows down considerably following a
malfunction or incident; and
- the flying conditions deteriorate unexpectedly, which requires a return
to twin-engine mode (insufficient flying altitude, for example).
Currently, a turbine engine is typically started by means of an electric
20 starter motor powered by the helicopter's on-board power supply. However,
the performance of this system is incompatible with the demand required
( ,
for an emergency restart. It is possible to adapt the electrical system, but
this requires mass use of expensive and disadvantageous technology
(synchronous machine having permanent magnets, power electronics and
25 dedicated battery packs, etc.).
Typically, a conventional start sequence of an idle engine takes
around 30 seconds, which could be too long depending on the flying
conditions, for example at low altitude in the event of an at least partial
failure of the single active engine. If the idle engine is not restarted in time,
30 I to land with the engine experiencing difficulties.
More generally, the emergency situations that can arise in the
applications considered above need reaction time of around a few seconds,
in order to ensure an emergency start or restart within a sufficient safety
margin.
5 In particular, the present invention provides a simple, effective and
economical solution to this requirement. However, the invention is not
limited to the above application and can be used to ensure the emergency
start of a turbine engine of any type of aircraft or helicopter equipped with
more than two engines, for example a triple engine helicopter.
10 SUMMARY OF THE INVENTION
To this end, the invention proposes an emergency start system for a
turbine engine of an aircraft, characterised in that it comprises at least one
solid-propellant gas generator, an electrically controlled ignition device, a
computer connected to the ignition device, and at least one starter motor
15 comprising a turbine for driving a shaft intended for being coupled to a shaft
of the turbine engine, the gas outlet of the generator being connected to the
inlet of the turbine of the starter motor.
The invention thus proposes a new technology for ensuring the
emergency start of an aircraft turbine engine. This technology uses a solid-
20 propellant gas generator that is relatively compact and can be integrated in
a turbine engine or an aircraft in a simple manner. A solid propellant is an
energetic material containing the oxidiser elements (combustion agent) and
reducing elements (fuel), which allow for high-energy gaseous combustion
products to be generated by combustion (redox reaction). According to the
25 invention, the energetic material is a solid propellant. This propellant is, for
example, a homogenous or composite propellant. .
An emergency start system based on a solid-propellant has a high
power density and high energy density compared with an electric
accumulator, and allows the action to occur much more quickly. This
30 system also boasts complete autonomy with regard to the helicopter's
electric network, in particular if the ignition device is controlled by the
computer of the turbine engine.
When an emergency situation is detected, the computer is designed
to activate the ignition device of the gas generator. The gases produced by
the generator rotate the rotor of the turbine and thus the shaft for driving the
shaft of the turbine engine.
Advantageously, the gas outlet of the generator is connected to the
inlet of the turbine by a distribution valve connected to the computer.
According to another embodiment of the invention, the system
comprises two independent starter motors intended for starting two turbine
engines independently of one another. The gas outlet of the generator is
connected to the inlet of the turbine of each starter motor.
The computer controls the valve so that the gases from the
combustion of the propellant feed the turbine of the starter motor
associated with the turbine engine that requires an emergency start.
The outlet of the turbine of the or each starter motor is preferably
connected to the exhaust. This is advantageous because, if the rotor of the
turbine were locked against rotation, the gases generated by the gas
generator would pass through the rotor to the exhaust, without the risk of
the turbine exploding.
The turbine of the or each starter motor can comprise one single
rotor wheel. The turbine of the starter motor is, for example, designed to
generate an average power of from 40-50 kW over a period of around 3 s. If
higher performance is required, the turbine could consequently be
optimised and would include more than one stage, for example.
The shaft driven by the turbine is preferably connected to a
freewheel designed to transmit a driving torque only when said torque
comes from the starter motor. The rotor of the turbine of the starter motor is
thus not driven by the shaft of the turbine engine during operation, which
ensures an optimum service life of the starter motor. In a variant, the shaft
driven by the turbine can be directly connected to the shaft of the turbine
engine or by transmission means designed to transmit a driving torque
when said torque comes from one shaft or the other of the starter motor
and the turbine engine. The shaft driven by the turbine can be coupled to
the shaft of the turbine engine by means of an accessory gearbox.
The present invention also relates to an aircraft, such as a helicopter,
comprising at least one turbine engine and at least one starter system as
described above.
The aircraft can comprise at least two turbine engines, each turbine
engine being associated with an independent starter system or with an
overall starter system consisting of one starter motor per turbine engine and
a common gas generator.
The present invention also relates to a method for emergency
starting a turbine engine of an aircraft by means of a system as described
above, characterised in that the ignition device is activated by the computer
as soon as an emergency start situation is detected.
DESCRIPTION OF THE DRAWINGS
The invention will be better understood and other details, features
and advantages of the invention will emerge from reading the following
description, given by way of non-limiting example and with reference to the
accompanying drawings, in which:
- Fig. 1 is a very schematic view of a twin helicopter equipped with an
emergency start system according to the invention, and
- Fig. 2 is a schematic view of an emergency start system according to the
invention, together with an axial sectional view of a starter motor of this
system, and
- Fig. 3 and 4 are views similar to Fig. 1 that show alternative embodiments
of the invention.
DETAILED DESCRIPTION
Fig. 1 and 2 show an embodiment of the invention which, in this
case, is applied to a twin helicopter 10, this helicopter comprising two
turbine engines 12 for driving the rotor of the propeller 14 by means of a
main gearbox 16.
Each turbine engine is equipped with a starter motor 18, which is
part of an emergency start system 20 according to the invention (Fig. 2),
the system 20 thus comprising two starter motors 18 in this case.
The system 20 further comprises a solid-propellant gas generator 22,
an electrically controlled device 24 for igniting the solid propellant, a
distribution valve 26 connecting the gas outlet of the generator to the starter
motors 18, and a computer 28 connected to the ignition device 24 and to
the valve so as to control said starter motors.
Here, the gas generator 22 comprises an elongate cylindrical body
that contains one or more loads of solid propellant, which are shaped
according to the desired gas flow law of the generator, said body acting as
a combustion chamber. It should be noted that the desired gas flow law can
be achieved by selecting the shape of the load appropriately andlor by
completely or partly lining some portions of the load.
Once the surface of the propellant load has been ignited, the surface
of the load burns and advances, producing combustion gases under high
pressure, according to the flow law resulting from the shape and the lining
of the load. The gases are ejected at the outlet of the generator. The
combustion temperature inside the generator 22 is typically in the range of
between 1400 K and 2700 K.
The ignition device 24 is controlled electrically by the computer 28
and is designed to activate the combustion of the propellant as soon as a
corresponding signal is output by the computer 28.
The distribution valve 26 is, for example, a proportional three-way
valve and comprises an inlet channel connected to the outlet of the
generator 22 (arrow 30) and two outlet channels connected to the starter
motors 18 (arrows 32), respectively. The valve 26 is such that the inlet
channel can only be connected to one of the outlet channels, such that the
system 20 according to the invention is designed to execute an emergency
restart on only one turbine engine at a time. The inlet channel is brought
into fluid communication with one outlet channel or the other, depending on
a signal output by the computer 28.
The computer 28 is an electronic control unit such as those currently
used in the field of aeronautics. When said computer detects an emergency
situation, it activates the distribution valve 26 and the ignition device 24
such that the starter motor 18 of the turbine engine to be started is fed the
combustion gases generated by the combustion of the solid propellant.
Fig. 2 is a schematic view of one embodiment of the starter motor
18, shown in axial section. Said starter motor substantially comprises a
shaft 34 supporting a rotor wheel 36 of a turbine 38, for example a
supersonic turbine, the shaft 34 being guided in rotation by bearings 40
mounted in a casing 42 of the starter motor. The casing 42 comprises a
radial opening 44, which forms the inlet of the turbine 38 and opens into an
annular cavity 46 for feeding the turbine. Said cavity 46 can have a
constant cross section from upstream to downstream or, on the other hand,
can have a cross section that changes from upstream to downstream.
The combustion gases that enter the cavity 46 are depressurised
and flow through the blades 48 of the wheel 36 (arrows 50), which rotates
the wheel 36 and thus the shaft 34 about its axis (arrow 52). The gases
then escape from the turbine 38 through an exhaust nozzle thereof, and are
ejected to the outside (arrows 50).
Cooling means can be provided upstream of the turbine so as to
reduce the temperature of the combustion gases, for example to 600 K.
Furthermore, a filter 53 can be fitted upstream of the turbine so as to limit
the entry of solid particles into the duct of the turbine.
The shaft 34 is designed to transmit a starting torque to a shaft 54,
for example a high-pressure shaft, of the turbine engine with which the
starter motor 18 is associated. This torque can be transmitted directly, as
shown schematically in dashed lines 56, or by means of transmission
means such as an accessory gearbox 58.
However, it is preferable for the torque to be transmitted between the
shaft 34 and the shaft 54 or the accessory gearbox 58 by means of a
freewheel 60. This freewheel is shown schematically by a diode, since the
function of the freewheel is to transmit a torque if it comes from the shaft
5 34, but not if it comes from the shaft 54. Once the turbine engine has
started, the shaft 34 no longer drives the shaft 54.
In the alternative embodiment shown in Fig. 3, the starter system
only comprises one starter motor 118; the emergency start system 120 is
therefore only associated with one turbine engine 112 and does not
10 comprise a distribution valve, the outlet of the gases from the generator 122
( ) being directly connected to the inlet of the turbine of the starter motor.
In the alternative embodiment shown in Fig. 4, each twin-helicopter
turbine engine 112 is associated with its own starter system 120, which is
thus independent of the other system 120. Each system 120 comprises a
15 generator 122 and a starter motor 118.
In another variant of the invention (not shown), a starter system
could comprise more than one generator for generating gas by combustion
of a solid propellant, for example in the form of a "cluster"; the generators
would be installed in parallel and would be actuated in a manner offset over
20 time in order to adjust the overall flow of hot gases feeding the starter
motor.
CLAIMS
1. Emergency start system (20) for turbine engines (12) of an aircraft
(lo), characterised in that it comprises at least one solid-propellant gas
5 generator (22), an electrically controlled ignition device (24), a computer
(28) connected to the ign~tiond evice, and at least two independent starter
motors (18), each intended for starting one turbine engine, each starter
motor comprising a turbine (38) for driving a shaft (34) intended for being
coupled to a shaft (54) of the corresponding turbine engine, the outlet of
10 the gases from the generator being connected to the inlet (44) of the
turbine of each starter motor by means of the same distribution valve
connected to the computer (28).
2. System (20) according to the preceding claim, characterised in that
the output of the turbine (38) of each starter motor (18) is connected to the
15 exhaust
3. System (20) according to any of the preceding claims, characterised
in that the turbine (38) of each starter motor (18) comprises a single rotor
wheel (36).
4. System (20) according to any of the preceding claims, characterised
20 in that it comprises means for coupling the shaft (34) driven by the turbine
(38) to the shaft (54) of the turbine engine (12), said coupling means
comprising a freewheel (60) designed to transmit a driving torque only
when said torque comes from the starter motor (18).
5. Aircraft comprising at least two turbine engines (12, 112),
25 - characterised in that it comprises an emergency start system (20)
according to any of claims I to 4.
6. Method for emergency starting a turbine engrne (12) of an aircraft by
means of a system (20) according to any of claims 1 to 5, characterised in
that the ignition device (24) is activated by the computer (28) as soon as
30 an emergency start situation is detected.