TURBINE ENGINE RAPID REACTIVATION METHOD AND SYSTEM
Technical field and prior art
The invention relates to a method and a system for
5 rapidly reactivating a turbine engine.
10
The field of application of the invention is more
particularly that of controlling the starting of gas
turbine propulsion aeroengines such as helicopter
turboshaft engines or turboprops for fixed-wing aircraft.
In conventional manner, an aircraft turboshaft
engine comprises a combustion chamber, a compressor shaft
having a compressor wheel mounted thereon to feed said
combustion chamber with compressed air, and at least one
starter or generator-starter connected to said shaft in
15 order to deliver sufficient starting torque thereto for
driving it in rotation.
In order to start the turboshaft engine, the starter
begins by accelerating the compressor shaft during a
first starting stage in which the fuel circuit upstream
20 from the starting injectors is put under pressure and
flushed. Thereafter, in a second stage of starting, fuel
injection is initiated prior to igniting said fuel in the
combustion chamber of the engine. Finally, in a third
stage of starting, at a predefined speed of rotation, the
25 action of the starter is stopped and the engine can
continue to accelerate as a result of the combustion of
said fuel.
In order to enable the fuel to be ignited, the air
delivered by the compressor wheel to the combustion
30 chamber must satisfy certain pressure and speed
conditions at the fuel injectors, so as to guarantee a
precise fuel/air ratio and so as to avoid blowing out the
flame. However, since the volume of air delivered by the
compressor wheel to the combustion chamber is
35 proportional to the speed of rotation ·of the compressor
shaft, the speed of rotation of the gas generator shaft
must therefore lie within a range of speeds known as the
2
ignition window, and this must continue for a length of
time that is sufficient to enable ignition to take place
correctly.
Traditionally, the turboshaft engines of nearly all
5 light or medium helicopters, and even some engines of
heavy helicopters and numerous turboprops of light fixedwing
airplanes, are started using a direct current (DC)
starter or a generator-starter that is powered at
10
28 volts (V) DC.
The invention applies more particularly to
helicopters having at least two turboshaft engines. Each
engine is designed so as to be oversized and capable on
its own of keeping the helicopter in flight in the event
of the other engine failing. Such oversized engines
15 operate under partial load for most of the time, with the
power needed for keeping the helicopter in cruising
flight being relatively low. Such operation is therefore
penalizing in terms of fuel consumption. That is why, in
order to reduce fuel consumption under cruising
20 conditions, it is possible to stop one of the engines.
The active engine then operates at a higher power rating
and thus with a more favorable level of specific
consumpt~on (Cs) .
Regulator systems for optimizing specific
25 consumption are described in particular in Documents
FR 2 967 132 Al and FR 2 967 133 Al. In those documents,
on a twin-engined helicopter in low-cost cruising flight,
i.e. in a stage of flight characterized by a relatively
low power demand on each engine, typically of the order
30 of 50% to 60% of its maximum continuous power (MCP), and
leading to very high specific consumption, one of the two
engines is put on standby (combustion chamber ignited or
extinguished and with the compressor turning), such that
the other engine operates at a high rating and therefore
35 benefits from a level of specific fuel consumption that
is much lower. Nevertheless, under such circumstances it
is appropriate, for safety reasons, to be able to
3
reactivate an engine in standby mode on board a twinengined
helicopter quickly in a manner that is simple and
reliable.
By way of example, proposals have already been made
5 in Document EP 2 264 297 A for a turboshaft engine fitted
with a 28 V generator-starter that is coupled to the gas
generator, to assist starting by means of a booster
system made up of a bank of supercapacitors, which bank
is connected in parallel with the 28 V battery of the
10 helicopter. Nevertheless, that system presents drawbacks
insofar as the voltage level used is constant, and not
adapted to rapid reactivation, and the electrical machine
constituted by the generator-starter cannot deliver the
power needed for the rapid starting function throughout
15 the entire transient stage. Furthermore, the known
architecture that is proposed seeks only to start a gas
generator that is stationary.
Document EP 2 581 586 A describes a system for
starting helicopter engines, which uses electrical energy
20 sources based on traditional helicopter battery
technology for normal starting and which additionally
uses electrical energy storage systems of the electric
double layer capacitor (EDLC) supercapacitor type. The
nominal voltage is that of the on-board network, i.e.
25 28 VDC. As a result, the starter is used in its nominal
mode and the starting performance is obtained by varying
the electrical characteristics of the source: during
normal starting, the traditional helicopter battery is
connected to the starter, while during rapid starting,
30 the second device having the supercapacitor type storage
systems is connected to the starter. As a result, under
all circumstances the voltage level used remains that of
the traditional helicopter batteries (e.g. 28 V), and at
that voltage the electrical machine constituted by the
35 starter cannot deliver the power needed for the rapid
starting function (emergency reactivation) throughout its
transient stage.
4
It would therefore be desirable to have a system
that enables the igniting and starting turboshaft engines
to be made more robust, including when they are in
standby mode, however if that were done in conventional
5 manner, it would require a DC-DC converter that is
imposing, since it would need to be dimensioned for
currents that are very high, possibly exceeding one
thousand amps.
10 Definition and object of the invention
The invention seeks to remedy the above-mentioned
drawbacks and in particular to make it possible on a
twin-engined helicopter to perform a function of
emergency restarting (rapid reactivation) of one of the
15 engines from a standby mode.
20
The invention seeks more particularly to provide an
electrical architecture for a turboshaft engine starter
system that constitutes an electrical hybridizing device
satisfying in particular the following objects:
· being capable of performing the conventional
functions of a starter, i.e. enabling the turbine engine
to be started normally and to perform dry motoring;
· being capable of delivering the necessary
performance for providing emergency reactivation
25 functions, given that the electrical characteristics
(voltage, impedance) of the generator elements of the onboard
network of the aircraft - battery, generator or
alternator - are designed to be capable of performing
normal starting of the turbine engines, but are generally
30 not sufficient for delivering the level of current needed
for obtaining a burst of starting torque that is
considerably higher, as is necessary of emergency
reactivation;
· minimizing electrical constraints and impacts on
35 the helicopter's on-board network: generation,
distribution, battery, ... ,
5
· being as lightweight and as well optimized as
possible; and
· being capable, where necessary, of being
incorporated easily in existing turbines and thus of
5 being compatible with "conventional" brush
generator/starters.
To solve the above-mentioned problems, there is
provided an aircraft having a turbine engine with a rapid
reactivation system, the system comprising an electrical
10 machine that is DC powered from an on-board electrical
power supply network included in said aircraft, and the
aircraft being characterized in that it further comprises
a switch interposed between the on-board electrical power
supply network and the electrical machine, said switch
15 being open to isolate the electrical machine from the onboard
electrical power supply network when emergency
reactivation is selected, an additional set comprising a
plurality of N electrical energy storage elements, and a
control unit adapted for controlling a device for
20 discharging the electrical energy storage elements, the
device for discharging the electrical energy storage
elements being incorporated in the aircraft and being
adapted to enable a series circuit comprising at least
some of the N electrical energy storage elements to be
25 connected in parallel with the on-board electrical power
supply network, the voltage across the terminals of the
electrical machine being configured by sequentially
switching the number of the N storage elements so as to
accompany the increase in the back-electromotive force
30 from the electrical machine progressively as the speed of
the gas generator associated with the turbine engine
increases in such a manner that, while the rapid
reactivation system is in operation, the electrical
machine is powered by a voltage level above the level of
35 its nominal characteristics.
Advantageously, storage elements have source
impedance that is lower and power density that is higher
6
than the source impedance and the power density of the
on-board electrical power supply network, so as to be
compatible with the high torque levels and thus the high
current levels that are required for emergency
5 reactivation of the turbine engine.
10
The storage elements may be of the electric doublelayer
capacitor (EDLC) supercapacitor type.
The storage elements may also be of the hybrid
lithium capacitor (LIC) type.
In a particular embodiment of the invention, the
control unit is associated with a device for charging and
balancing and with the switch in order to control the
charging of the storage elements from the electrical
machine operating as an electricity generator, outside
15 periods of rapid reactivation.
In another particular embodiment of the invention,
the control unit is associated with a device for charging
and balancing to control the charging of the storage
elements from the on-board electrical power supply
20 network, outside periods of rapid reactivation.
The starting system of the invention is
advantageously applied to a turbine engine of a twinengined
helicopter.
The invention also provides a rapid reactivation
25 method for rapidly reactivating an aircraft turbine
engine including an electrical machine that is DC powered
from an on-board electrical power supply network included
in the aircraft, the method being characterized in that
it comprises the steps consisting in selectively
30 interrupting the electrical connection between said onboard
electrical power supply network and said electrical
machine using a switch that is in an open position in
order to isolate the electrical machine from the on-board
electrical power supply network when emergency
35 reactivation is selected, and is using a control unit and
a device for discharging storage elements to enable a
series circuit comprising at least some of the N
7
electrical energy storage elements to be connected in
parallel with the on-board electrical power supply
network, the voltage across the terminals of the
electrical machine being configured by sequentially
5 switching the number of the N storage elements so as to
accompany the increase in the back-electromotive force
from the electrical machine progressively as the speed of
the gas generator associated with the turbine engine
increases in such a manner that, while the rapid
10 reactivation system is in operation, the electrical
machine is powered by a voltage level above the level of
its nominal characteristics.
The on-board electrical power supply network may
include an alternator or an electricity generator, or it
15 may be connected to a ground power unit (GPU) (when the
aircraft is on the ground) or indeed it may be connected
to a storage battery, e.g. a 28 V battery.
In a particular implementation, the starting method
of the invention further includes a step of controlling
20 the charging of the storage elements by a charging and
balancing device from the on-board electrical power
supply network, outside periods of rapid reactivation.
In another particular implementation, the starting
method of the invention further comprises a step of
25 controlling the charging of the storage elements by means
of a charging and balancing device and the switch from
the electrical machine operating as an electricity
generator, outside periods of rapid reactivation.
The invention applies most particularly to systems
30 for starting turboshaft engines of aircraft, and in
particular of helicopters.
Brief description of the drawings
Other characteristics and advantages of the
35 invention appear from the following description of
particular embodiments given as examples and with
reference to the accompanying drawings, in which:
8
· Figure 1 is a diagrammatic overall view of an
embodiment of a device in accordance with the invention
for rapidly reactivating a turbine engine;
· Figures 2A to 2C show switch commands for
5 discharging electrical energy storage elements in the
context of a device in accordance with the invention for
rapidly reactivating a turbine engine;
· Figure 3 is a diagram showing an example of how
the speed or the back-emf of an electrical machine
10 controlled in accordance with the invention varies as a
function of time;
· Figure 4 is a diagram showing how the current in
an electrical machine controlled in accordance with the
invention varies as a function of time;
15 · Figure 5 is a diagram showing how the voltage
applied to an electrical machine controlled in accordance
with the invention varies as a function of time;
· Figure 6 shows a first embodiment of a device for
balancing and charging electrical energy storage cells
20 that is suitable for being used in the device in
accordance with the invention for rapidly reactivating a
turbine engine: and
· Figure 7 shows a second embodiment of a device for
balancing and charging electrical energy storage cells
25 that is suitable for being used in the device in
accordance with the invention for rapidly reactivating a
turbine engine.
30
Detailed description of preferred embodiments
Figure 1 is a diagram showing the general
configuration of a device of the invention.
The emergency restarting system, i.e. the system for
rapidly reactivating a turbine engine on standby,
comprises an on-board electrical power supply network 10
35 that includes, amongst other things: a storage battery 13
that may be a single battery or a group of batteries and
that may be constituted by the conventional power supply
9
of an on-board network of an aircraft, e.g. at a voltage
of 28 V, however the invention is not limited to such a
value.
The on-board electrical power supply network 10 may
5 also be associated with an alternator or an electricity
generator 11, or it may be connected to a ground power
unit (GPU) 12 (when the aircraft is on the ground) in
addition to being able to be connected to a storage
battery 13, e.g. at 28 V.
10 An electrical machine 60 may be constituted by a
simple DC starter or by a generator-starter (GS) capable
of operating not only in motor mode, but also in
generator mode once the stage of starting has terminated,
e.g. for the purpose of powering the on-board network 10.
15 In the description below, the term "starter" is used to
cover both a device that is a starter only and a device
that is a generator-starter, unless specified to the
contrary.
Figure 1 does not show the main elements of the
20 turbine engine, which are conventional, and may comprise
a gas generator, itself comprising a compressor, a
combustion chamber, and a high pressure turbine, together
with a free turbine and starting accessories. Likewise,
Figure 1 does not show a sensor for sensing the speed of
25 rotation of the starter, nor a sensor for sensing the
speed of rotation of the compressor shaft of the engine.
Nevertheless, in diagrammatic manner, a line 61 is shown
for transmitting to the control unit 20 information about
the speed or the back-emf of the electrical machine 60
30 that may be constituted by a starter or by a generatorstarter.
The starting system of the invention includes a
control unit 20. Figure 1 does not show the various
sensors for measuring the operation of the engine, such
35 as temperature sensors serving in particular to measure
the operating state in the combustion chamber.
10
The control unit 20, which may be associated with
the conventional electronic computer 21 of the engine,
also known as an engine electronic control unit (EECU),
or which may be directly incorporated therein, serves to
5 manage the measurements provided by the sensors and to
control the starting system via the module for managing
the on-board network of an aircraft. The control unit 20
is adapted to receive a normal starting command (line 22)
or an emergency reactivation command (line 23).
10 The starting device of the invention also includes a
switch 50 (K0 ) between the on-board network 10 and the
electrical machine 60, a set 30 of N electrical energy
storage elements 30a, ... , 30n (capacitors of capacitance
Cl' C2 , ••• , CN), and a discharge device 40 for
15 discharging the storage elements and for receiving
commands K"' K2, ••• , Ki, .•• , K0 via lines lOa, 20b,
20i, .... , 20n from the control unit 20.
The control unit 20 controls the switch 50 via a
line 51 (command K0 ) and also controls a set of other
• • • I
20 switches 40a, 40b, ... , 40n of the device 40 for the
storage elements (commands K1 , K2, ••• , KN) enabling a
series circuit comprising some or all of the N electrical
energy storage elements 30a, ... , 30n to be connected in
parallel with the on-board electrical power supply
25 network 10 in order to deliver the energy needed for
emergency restarting, which constitutes rapid
reactivation of the turbine previously put on standby.
Diodes 4la, 4lb, ... , 4ln are connected in series with
the switches 40a, 40b, ... , 40n.
30 Because of the additional subassemblies 20, 30, 40,
50, and 70.! the system of the invention enables the
voltage source that is to power the electrical machine 60
to be configured in such a manner that when the rapid
reactivation system is in "emergency reactivation"
35 operation, a voltage level is applied to the terminals of
the electrical machine 60 that is higher than the level
of the nominal characteristics of that electrical machine
11
60 in order to be able, transiently, to increase the
mechanical torque it delivers, while applying the voltage
gradually so as to limit the current drawn at the
beginning of emergency reactivation and so as to
5 accompany the increase in the back-emf from the starter
progressively as the speed of the gas generator rises.
The storage elements 30a, ... , 30n have source
impedance that is smaller and power density that is
larger than the source impedance and the power density of
10 the storage elements of the on-board electrical power
supply network 10, and they are therefore suitable for
delivering high starting current during the short
duration of emergency reactivation.
The additional storage elements 30a, ... , 30n may in
15 particular be constituted by (EDLC) supercapacitors or by
hybrid lithium ion capacitors (LICs).
The operation of the starting system of the
invention is described below in greater detail.
When normal starting is selected, e.g. when the
20 aircraft is on the ground and the engine is initially
stationary, the EECU 21 sends the normal starting command
over the line 22 to the control unit 20, which closes the
switch 50 by means of the control line 51. The starter
60 is then powered directly by the on-board network 10
25 and it applies a starting torque to the gas generator of
30
the engine. In known manner, the voltage level and the
impedance of the generator elements of the on-board
network 10 are suitable for delivering the moderate
current needed for normal starting of the engine. The
same procedure is also used in flight for normal
reactivation of an engine that has previously been put
standby, when restarting does not present any urgent
nature.
When emergency starting is selected, while the
35 aircraft is in flight and the engine is initially in
standby mode, the EECU 21 sends the emergency
on
12
reactivation command via the line 23 to the control unit
20 which performs the following functions:
· opening the switch 50 so as to isolate the
electrical machine 60 from the on-board network 10; and
5 · configuring the voltage across the terminals of
the electrical machine 60 by sequentially switching the
number of storage elements 30a, ... , 30n needed within
the assembly 30 for the purposes firstly of managing the
currents delivered to the electrical machine 60 so as to
10 obtain mechanical torque that is considerably greater
than the normal starting torque, and secondly to
accompany the increase in the back-emf from the
electrical machine 60 progressively as the speed of the
gas generator increases.
15 The combinations in which the contactors 40a, 40b,
20
... , 40n are closed may differ as a function of the
nature of the storage elements 30a, 30b, ... , 30n and
also as a function of the characteristics of the
electrical machine 60.
Figures 2A to 2C and 3 to 5 show an example of
timing for controlling the switches 40a to 40n (commands
K1 to KN) for limiting current and thus torque in the
machine during the initial stages of emergency starting
while continuing to increase its speed at the end of
25 starting, where the electrical machine 60 is subjected to
an operating voltage that is higher than its nominal
operating point.
More particularly, Figure 2A shows a signal 101
corresponding to a command K1 that closes the contactor
30 40a between an initial instant T1 and a final instant T •.
Figure 28 shows a signal 102 corresponding to a
command K2 that closes the contactor 40b between an
initial instant T2 later than the initial instant T1 , and
a final instant T •.
35 Figure 2C shows a signal 109 corresponding to a
command KN that closes the last contactor 40n between an
13
initial instant TN later than any of the initial instants
Tu T2 , ••• , and a final instant TF.
From Figures 2A, 2B, and 2C, it can be understood
that staggered control of the contactors 40a, 40b, ... ,
5 40n makes it possible to connect the storage elements
30a, 30b, ... , 30n successively in series so as to apply
the sum of the resulting voltages to the armature winding
of the electrical machine 60.
Figure 3 shows how the speed or the back-emf of the
10 electrical machine 60 controlled in accordance with the
sequence of Figures 2A to 2C varies as a function of
time. The curve has a first segment 111 that varies
between instants T1 and T2 between a threshold S1 (equal
to zero) and a threshold S2 , followed by successive
15 segments 112, ... , varying between the instants T2 to TN
between thresholds S2 to SN, and finally a last segment
119 that varies between the instant TN and a final instant
TF between the threshold SN and a final threshold SF
presenting a maximum value.
20 Figure 4 shows how the current flowing through the
electrical machine 60 controlled in accordance with the
sequence of Figures 2A to 2C varies as a function of
time. At each initial instant that a command K1 , K2 , • • • r
KN is applied at successive instants T1 , T2 , ••• , TN' the
25 current reaches a maximum value I~x and then decreases
(segments 121, 122, ... , 129) and returns to zero at the
final instant TF. Suitable contactor commands K1 , K2 ,
... , KN serve to limit the maximum current absorbed by the
electrical machine 60, and thus its torque, to a level
30 that is acceptable for the mechanical train driving the
gas generator, while nevertheless maintaining a high mean
level of torque throughout the emergency activation
sequence.
Figure 5 shows how the voltage applied to the
35 electrical machine 60 controlled in accordance with the
sequence of Figures 2A to 2C varies as a function of
time. At each initial instant that a command K1 , K2 , • • • I
14
KN is appl.ied at successive instants Tu T2 , ••• , TN, the
voltage applied to the electrical machine 60 increases
from a value VelD' Ve2D, ... , VeND equal to the value
corresponding to the initial charge of the storage
5 elements 30a, 30b, ... , 30n, and then decreases (segments
131, 132, ... , 139) during discharging of the storage
elements as connected in parallel in this way with the
electrical machine 60, while nevertheless remaining at a
value that increases progressively until it reaches a
10 value U~x at the initial instant of the command KN for
closing the last contactor 40n. Appropriate contactor
commands K1 , K2 , ... , KN thus enable the level of the
voltage applied to the electrical machine 60 to be
adapted gradually as the speed, and thus the back-emf, of
15 the machine increases, thereby enabling a large mean
starting torque
to high speeds.
to be maintained on the gas
It should be observed that
generator up
although the
maximum voltage level U~x that is applied to the
electrical machine 60 at the end of emergency activation
20 exceeds the nominal voltage of the on-board network, it
nevertheless remains compatible with the strength of the
insulation and with the commutator of the electrical
machine, for exceptional use.
When operating under normal conditions, i.e. other
25 than during a period of emergency activation, the control
unit 20 and the device 70 for charging and balancing the
cells also has the function of charging and keeping
charged the storage elements 30a, ... , 30n of the set 30
of additional storage elements, and in general manner of
30 monitoring these storage elements 30a to 30n.
The storage elements 30a, ... , 30n of the additional
set 30 may be charged from the power supply network 10 of
the helicopter, or in a variant from the electrical
machine 60 operating as an electricity generator.
35 Figures 6 and 7 show two examples of the operation
of the device 70 for balancing and charging cells
constituted by the storage elements 30a, 30b, ... , 30n.
15
As shown in Figure 1, the device 70 for balancing and
charging the cells may be powered from the on-board
network 10 by a connection 72 and may be controlled
selectively by the control unit 20 (connection 71) during
5 periods lying outside emergency reactivation. The
balancing and charging devices described in Figures 6 and
7 are given respective references 170 and 270.
The Figure 6 balancing and charging device is of
"flyback" structure and comprises at its input a filter
10 unit 171 followed by a capacitor 172 connected in
parallel with a circuit having a primary winding 174 of a
transformer and an electronic control member 173. A set
of N secondary windings 175a, 175b, ... , 175n of the
transformer are connected via rectifier diodes 176a,
15 176b, ... , 176n to a set of N electrical energy storage
elements 30a, 30b, ... , 30n so as to deliver voltages
VelD' Ve2D, ••• , VeND to the terminals of the various
electrical energy storage elements having capacitances C1 ,
Cz, ... , eN.
20 The balancing and charging device of Figure 7 is of
"forward" structure and comprises at its input a filter
unit 271, followed by a capacitor 272 connected in
parallel with a circuit comprising a set of electronic
components for powering a primary winding 274 of a
25 transformer. The electronic components may comprise an H
bridge made up of four diodes 281 to 284 and four
electronic control members 273, 277, 278, 279. A set of
N secondary transformer windings 275a, 275b, ... , 275n
are connected via rectifier diodes 276a, 276b, ... , 276n
30 to feed a set of N electrical energy storage elements
30a, 30b, ... , 30n in such a manner as to deliver
35
voltages VelD' Ve2D, ••• , VeND to the terminals of the
various electrical energy storage elements having
capacitances cl' c2'
The embodiment
. ' eN.
of Figure 6, in which the primary
winding 174 of the transformer is considered as being a
current source, enables energy to be transferred, with
16
energy being stored in the primary winding. It is thus
possible to control the energy that is transferred to the
transformer secondary windings 175a, 175b, ... , 175n.
This solution optimizes the weight of the control
5 electronics to the detriment of the weight of the
inductive element.
The embodiment of Figure 7, in which the transformer
primary 274 is considered as being a voltage source,
serves to optimize the size of the inductive element, to
10 the detriment of the control electronics. The system of
the invention is suitable for achieving emergency
restarting in flight of a turboshaft engine in a few
seconds.
It should be observed that advantageously, the
15 storage elements 30a, 30b, ... , 30n may be recharged on
the ground, with the engines idling during a stage of
preparing the aircraft prior to takeoff, so that the
corresponding electrical energy which is taken off may be
spread over a duration that is relatively long (a few
20 tens of seconds to a few minutes) without any negative
operational impact, which thus makes it possible firstly
to avoid oversizing the generator elements of the onboard
network, and secondly to reduce the power for which
the recharging devices as described in Figures 6 and 7
25 are designed, thereby in particular enabling their weight
and volume to be limited.
The additional equipment of the system of the
invention is very simple to put into place and it is very
compact. Thus, the additional set 30 of storage
30 elements, the control unit 20, the balancing device 70,
and the discharge device 40 can be incorporated directly
in the engine compartment of the engine.
The invention is also suitable for being used on
helicopters that are already in operation, given that the
35 modifications that need to be made to existing circuits
can be implemented simply.
17
The invention thus proposes using practical
technical means for implementing, on board a twin-engined
helicopter, a function of emergency restarting (rapid
reactivation) from a standby mode. In the invention, the
5 electric starter 60 of a turbine is thus used outside its
nominal operating range in order to handle the call for
mechanical power that is needed for emergency restarting
in flight.

CLAIMS
1. An aircraft having a turbine engine with a rapid
reactivation system, the system comprising an electrical
machine (60) that is DC powered from an on-board
5 electrical power supply network (10) included in said
aircraft, and the aircraft being characterized in that it
further comprises a switch (50) interposed between the
on-board electrical power supply network (10) and the
electrical machine (60), said switch (50) being open to
10 isolate the electrical machine (60) from the on-board
electrical power supply network (10) when emergency
reactivation is selected, an additional set (30)
comprising a plurality of N electrical energy storage
elements (30a, ... , 30n), and a control unit (20) adapted
15 for controlling a device (40) for discharging the
electrical energy storage elements (30a, ... , 30n), the
device (40) for discharging the electrical energy storage
elements (30a, ... , 30n) being incorporated in the
aircraft and being adapted to enable a series circuit
20 comprising at least some of the N electrical energy
storage elements (30a, ... , 30n) to be connected in
parallel with the on-board electrical power supply
network (10), the voltage across the terminals of the
electrical machine (60) being configured by sequentially
25 switching the number of theN storage elements (30a, ... ,
30n) so as to accompany the increase in the backelectromotive
force from the electrical machine (60)
progressively as the speed of the gas generator
associated with the turbine engine increases in such a
30 manner, that while the rapid reactivation system is in
operation, the electrical machine (60) is powered by a
voltage level above the level of its nominal
characteristics.
35 2. An aircraft according to
that storage elements (30a,
claim 1, characterized in
... , 30n) have source
impedance that is lower and power density that is higher
19
than the source impedance and the power density of the
on-board electrical power supply network (10).
3. An aircraft according to claim 1 or claim 2,
5 characterized in that the storage elements (30a, ... ,
30n) are of the electric double-layer capacitor (EDLC)
supercapacitor type.
4. An aircraft according to claim 1 or claim 2,
10 characterized in that the storage elements (30a, ... ,
30n) are of the hybrid lithium capacitor (LIC) type.
5. An aircraft according to any one of claims 1 to 4,
characterized in that the control unit (20) is associated
15 with a device (70) for charging and balancing and with
the switch (50) in order to control the charging of the
20
storage
machine
elements (30a,
(60) operating
... ' 30n) from the electrical
as an electricity generator,
outside periods of rapid reactivation.
6. An aircraft according to any one of claims 1 to 4,
characterized in that the control unit (20) is associated
with a device (70) for charging and balancing to control
the charging of the storage elements (30a, ... , 30n) from
25 the on-board electrical power supply network (10),
outside periods of rapid reactivation.
7. An aircraft according to any one of claims 1 to 6,
characterized in that it is constituted by a twin-engined
30 helicopter.
8. A rapid reactivation method for rapidly reactivating
an aircraft turbine engine including an electrical
machine (60) that is DC powered from an on-board
35 electrical power supply network (10) included in said
aircraft, the method being characterized in that it
comprises the steps consisting in selectively
20
interrupting the electrical connection between said on.
board electrical power supply network (10) and said
electrical machine (60) using a switch (50) that is in an
open position in order to isolate the electrical machine
5 (60) from the on-board electrical power supply network
(10) when emergency reactivation is selected, and is
using a control unit (20) and a device (40) for
discharging storage elements to enable a series circuit
comprising at least some of the N electrical energy
10 storage elements (30a, ... , 30n) to be connected in
parallel with the on-board electrical power supply
network (10), the voltage across the terminals of the
electrical machine (60) being configured by sequentially
switching the number of theN storage elements (30a, ... ,
15 30n) so as to accompany the increase in the backelectromotive
force from the electrical machine (60)
progressively as the speed of the gas generator
associated with the turbine engine increases in such a
manner that, while the rapid reactivation system is in
20 operation, the electrical machine (60) is powered by a
voltage level above the level of its nominal
characteristics.
9. A rapid reactivation method according to claim 8,
25 characterized in that it further includes a step of
controlling the charging of the storage elements (30a,
... , 30n) by a charging and balancing device (70) from
the on-board electrical power supply network (10),
outside periods of rapid reactivation.
30
10. A rapid reactivation method according to claim 8,
characterized in that it further comprises a step of
controlling the charging of the storage elements (30a,
... , 30n) by means of a charging and balancing device
35 (70) and the switch (50) from the electrical machine (60)
operating as an electricity generator, outside periods of
rapid reactivation ..