1. TECHNICAL FIELD OF THE INVENTION

The invention relates to a propulsion system of an aircraft. In particular, the invention relates to a propulsion system of an aircraft, comprising at least two turboshaft engines connected to a main gearbox, typically in a twin-engine or multi-engine helicopter.

2. BACKGROUND

The prior art comprises in particular the documents US-A1-2016/003144, FR-A1-2 962 488, U.S. Pat. No. 4,644,744, US-A1-2003/094001 and U.S. Pat. No. 4,197,699.

In a known manner, twin-engine or multi-engine helicopters have a propulsion system comprising two, three or more turboshaft engines, each turboshaft engine comprising a gas generator and a free turbine driven in rotation by the gas generator and attached to an output shaft. The output shaft of each free turbine is adapted to set in motion a power gearbox (hereinafter referred to as BTP), which itself drives the rotor of the helicopter equipped with variable pitch blades. The gas generator comprises at least one compressor and at least one turbine connected by a shaft, and a combustor receiving air compressed by the compressor and burning an air/fuel mixture to form a gas transmitted to the gas generator turbine and the free turbine.

At the outlet of the free turbine, an exhaust nozzle allows the evacuation of the outlet gases, called exhaust gas. These exhaust gases are at high temperature. It was thus envisaged to recover part of the thermal energy from these exhaust gases, in particular thanks to a heat exchanger. In particular, a heat exchanger can be used to use the thermal energy of the exhaust gases to heat the air entering the combustion chamber at the outlet of the compressor. A turboshaft engine equipped with such a heat exchanger is called a recuperative cycle turboshaft engine. Such a turboshaft engine is therefore more energy efficient.

However, the exchangers currently used in the turboshaft engines of today's propulsion systems are subject to thermal cycles due to rapid (transient) power variations, which cause strong thermal and mechanical stresses. In particular, the thermal and mechanical stresses due to these transients require adequate sizing and mechanical strength of the exchanger to limit its degradation. The exchanger is thus penalized because it is less efficient, heavy and expensive, and the turboshaft engine is therefore less competitive than a turboshaft engine without an exchanger.

The inventors therefore sought a way to use an exchanger in a propulsion system so that it would be competitive with existing propulsion systems.

3. OBJECTIVES OF THE INVENTION

The invention is intended to overcome at least some of the disadvantages of known propulsion systems.

In particular, the invention aims at providing, in at least one embodiment of the invention, a propulsion system allowing to recover part of the thermal energy of the exhaust gases.

The invention also aims at providing, in at least one embodiment, a propulsion system that does not require an oversized exchanger to resist the stresses generated by the transients of turboshaft engines. Thus, the invention aims to provide, in at least one embodiment, a propulsion system in which the thermal and mechanical stresses generated by the transients of a turboshaft engine in the heat exchanger associated with it are limited.

The invention also aims at providing, in at least one embodiment of the invention, a high-performance propulsion system.

The invention also aims to provide, in at least one embodiment, a propulsion system capable of managing transient phases with at least the same efficiency as a conventional propulsion system.

The invention also aims to provide, in at least one embodiment, a propulsion system operating in the event of failure of one of the turboshaft engines.

4. STATEMENT OF THE INVENTION

For this purpose, the invention relates to a propulsion system of an aircraft, comprising a main gearbox and at least two turboshaft engines connected to the main gearbox, respectively a first turboshaft engine and a second turboshaft engine, each turboshaft engine comprising a free turbine, wherein the first turboshaft engine comprises a heat exchanger configured to recover part of the thermal energy from the exhaust gases exiting the free turbine, and the propulsion system comprises at least one computer, configured to control the two turboshaft engines and to limit the acceleration and deceleration of the first turboshaft engine when neither of the turboshaft engines is broken down, so as to limit power transients at the heat exchanger.

A propulsion system according to the invention thus makes it possible to take advantage of the presence of at least two turboshaft engines so that one of them (the first turboshaft engine) is equipped with a heat exchanger, in return for limiting its acceleration and deceleration so as not to cause significant power transients at the heat exchanger, and so that the other turboshaft engine (the second turboshaft engine) meets the remaining needs while operating normally, without limitation of acceleration/deceleration. If the second turboshaft engine is equipped with a heat exchanger, this one will undergo the degradations evoked previously, thus the second turboshaft engine is preferably not equipped with a heat exchanger.

Thus, the first turboshaft engine can be used at an almost constant power level to supply part of the propulsive energy required by the aircraft, for example between 50% and 100% of the propulsive energy in cruise mode, preferably around 70% of the propulsive energy in cruise mode. The second turboshaft engine will provide the necessary power, either to achieve the propulsive energy at cruise power or to meet additional punctual energy requirements (helicopter maneuvering, load lifting, etc.).

In the event of a failure of the first turboshaft engine, the second turboshaft engine will take over and provide all the power needed to extend the flight or enable an emergency landing. Because of the failure, the deceleration of the first turboshaft engine is no longer controlled by the computer.

In the event of a failure of the second turboshaft engine, the computer stops limiting the acceleration and deceleration of the first turboshaft engine so that the first turboshaft engine can provide all the power required, even in the event of additional punctual power requirements, in order to guarantee the integrity of the aircraft and the safety of its occupants (by avoiding an accident). This limitation is to the detriment of the exchanger, which can be degraded more quickly in this critical case.

Turboshaft engines are of the type described above comprising a gas generator and a free turbine. Preferably, the exchanger recovers the exhaust gases at the free turbine outlet, and heats the gas leaving the compressor before it enters the combustion chamber.

The propulsion system can comprise more than two turboshaft engines, as long as it comprises at least one turboshaft engine with an exchanger with limited acceleration/deceleration and at least a second “conventional” turboshaft engine without an exchanger. The other turboshaft engines are either with an exchanger like the first turboshaft engine or without an exchanger like the second turboshaft engine.

Advantageously and according to the invention, the acceleration and deceleration limits imposed by the computer on the first turboshaft engine, when none of the turboshaft engines has failed, can be predefined as a function of the heat exchanger so as to limit the thermal and mechanical stresses in the latter.

Advantageously and according to the invention, the acceleration and deceleration limits can be predefined as a function of the dimensioning and mechanical strength of the heat exchanger.

Advantageously and according to the invention, the acceleration and deceleration limits predefined as a function of the heat exchanger can be lower than the physical acceleration and deceleration limits of the first turboshaft engine, and the heat exchanger is unable to withstand without degradation the power transients corresponding to said physical limits.

The physical acceleration and deceleration limits of the turboshaft engine are physical limits of operability of the turboshaft engine during transients. In a manner known per se, these physical limits translate into acceleration and deceleration stops which are imposed by the computer. Beyond these physical acceleration and deceleration limits, the turboshaft engine is no longer operable. The operability of the turboshaft engine is its ability to manage the transients, and in particular its ability to accelerate and decelerate the turboshaft engine while managing the acceleration and deceleration limits.

The computer of the propulsion system according to the invention can use, in a manner known per se, acceleration and deceleration stops to prevent pumping or shutdown of a turboshaft engine. Classically, curves including a stabilized operating line and a pumping line characterize the relationship between a compression ratio and a reduced mass flow rate of gases in a compressor of a turboshaft engine. On a graph representing this relationship, the stabilized operating line is located below the pumping line, and each line forms an increasing curve.

During transients, the acceleration line of the turboshaft engine and the deceleration line of the turboshaft engine deviate above and below the stabilized operating line, respectively. In order not to risk pumping during rapid acceleration of the turboshaft engine, the acceleration line must remain between the stabilised operating line and the pumping line, leaving a certain margin below the pumping line. This margin is due to the fact that the ordinate position of the pumping line, on the graph linking the compression ratio to the reduced mass flow rate, can vary in particular with the ageing or the deterioration of the engine, and moreover this position is not necessarily the same for two identical engines due to manufacturing distortions or dispersions. An acceleration stop dN/dt is generally defined as corresponding to a line, usually a straight line, maintaining a margin greater than or equal to a predefined minimum margin with respect to the pumping line. Thus, during transients, the acceleration line can reach the acceleration stop, but does not encroach on the pumping margin.

Similarly, on the graph linking the compression ratio to the reduced mass flow rate, the deceleration line of the turboshaft engine must remain between the stabilised operating line and a shutdown line of the combustion chamber of the turboshaft engine. A deceleration stop is usually defined so that during rapid deceleration of the turboshaft engine, the deceleration line remains sufficiently above the shutdown line.

The computer can therefore impose acceleration and deceleration stops on the first turboshaft engine that correspond to the physical acceleration and deceleration limits beyond which the turboshaft engine may pump or shutdown. Advantageously and according to the invention, the acceleration and deceleration limits provided to limit the power transients at the heat exchanger are lower than said acceleration and deceleration stops.

Advantageously and according to the invention, the propulsion system comprises an electric motor connected to the second turboshaft engine and configured to assist the second turboshaft engine during an acceleration, starting and/or standby exit phase.

According to this aspect of the invention, the electric motor allows the second turboshaft engine to be assisted during transient phases. In particular, several assistances are possible, the motor can be configured to provide one or more of them:acceleration/deceleration assistance: since the second turboshaft engine is the only one to provide transients and sudden variations in energy demand, the electric motor can allow the second turboshaft engine to accelerate/decelerate more than it is capable of doing alone;starting or standby exit assistance: in situations where the first turboshaft engine can provide the propulsive power alone, the second turboshaft engine can be switched off or in standby, according to several possible configurations. The electric motor can thus, in the event of an increase in the energy required by the aircraft or in the event of a failure of the first turboshaft engine, quickly restart or exit of standby the second turboshaft engine. In particular, if the first turboshaft engine fails, the second turboshaft engine must be able to supply power quickly enough so as not to endanger the aircraft and its occupants. Conventional starting of the second turboshaft engine without assistance may not be fast enough for these critical situations. For example, a power-assisted startup or standby exit may be performed in one-third of the time without assistance (e.g. 10 seconds for a power-assisted startup/standby exit versus 30 seconds for a conventional startup/standby exit).

The different shutdown or standby modes of the second turboshaft engine are for example:a complete shutdown of the combustion chamber and a shutdown of the rotating shafts. In this mode, fuel consumption is zero, but starting may take a considerable amount of time, which is not compatible with a possible need for a quick start following the sudden loss of the other turboshaft engine;a complete shutdown of the combustion chamber and a switch to “turning” mode, i.e. the shafts remain in rotation. In this mode, the electric motor drives the gas generator at a speed of about 10% of the nominal operating speed, allowing significant gains in emergency re-ignition time with zero fuel consumption;a partial extinction of the combustion chamber and a setting in “super-idling” mode. In this mode, the gas generator is driven partly by the combustion of gas and fuel in the combustion chamber and partly by the electric motor. It consumes little fuel but allows very fast reactivation.

The first mode is considered a shutdown and the other two modes are considered standby, according to two different levels. Preferably, the second mode is implemented in the invention during the standby of the second turbo-shaft engine.

Turboshaft engines are of the type described above comprising a gas generator and a free turbine. The electric motor added to the second turboshaft engine drives the shaft of the gas generator.

Advantageously and according to the invention, each turboshaft engine comprises:a gas generator comprising a compressor and a turbine connected by a shaft, and a combustion chamber receiving air compressed by the compressor and burning an air/fuel mixture to form a gas transmitted to the turbine of the gas generator,the free turbine driven in rotation by the gas generator and integral with an output shaft, said output shaft being connected to the main gearbox,

and the heat exchanger is configured to heat the air compressed by the compressor of the first turboshaft engine via the recovered thermal energy portion.

Advantageously and according to the invention, each turboshaft engine comprises an exhaust nozzle, the heat exchanger being configured to recover part of the thermal energy from the exhaust gases at the exhaust nozzle of the first turboshaft engine.

Advantageously and according to the invention, the heat exchanger is configured to recover part of the thermal energy from the exhaust gases of both turboshaft engines.

The invention also relates to a propulsion system of an aircraft, comprising a main gearbox and N turboshaft engines connected to the main gearbox, N being an integer greater than or equal to 3, each turboshaft engine comprising a free turbine,

characterized in that at least one turboshaft engine among the N turboshaft engines, referred to as the first turboshaft engine, comprises a heat exchanger configured to recover part of the thermal energy from the exhaust gases exiting the free turbine,

and in that the propulsion system comprises at least one computer, configured to control the N turboshaft engines and to limit the acceleration and deceleration of said first turboshaft engine when none of the N turboshaft engines is broken down, so as to limit power transients at the heat exchanger.

Advantageously and according to the invention, the propulsion system comprises an electric motor connected to at least one other turboshaft engine, referred to as the second turboshaft engine, and configured to assist said second turboshaft engine during an acceleration, starting and/or standby exit phase.

Advantageously and according to the invention, the heat exchanger is configured to recover part of the thermal energy from the exhaust gases of each turboshaft engine.

The invention also relates to a method for managing a propulsion system according to the invention, characterized in that it comprises the following steps:a step of controlling the energy supplied by the first turboshaft engine and the second turboshaft engine so as to ensure the propulsive energy requirements,during a cruise flight of the aircraft, a shutdown or standby step of the second turboshaft engine, so that the first turboshaft engine provides all the propulsive energy,a restart or standby exit step for the second turboshaft engine if it is shut down or in standby mode and the energy supplied by the first turboshaft engine is no longer sufficient to meet the propulsive energy requirements.

A method according to the invention therefore allows the propulsion system to be used in cruise flight in “single engine operative” (also called SEO) with the first turboshaft engine. In the event of a failure or additional power requirement, the second turboshaft engine is back in operation (restarted or standby exit).

Preferably, the restart or standby exit step comprises an assistance step with an electric motor connected to the second turboshaft engine.

Assistance via an electric motor ensures that the time required for the second engine to be operational is limited, especially in an emergency.

The invention also concerns a propulsion system and a management method characterized in combination by all or part of the characteristics mentioned above or below.

