Control system for cooling onboard equipment of a vertical takeoff and landing aircraft

Abstract

A vertical takeoff and landing aircraft. The aircraft includes an air inlet (22) connected to a cooling component (14). The aircraft further includes a horizontal drive propeller (16) arranged to generate airflow in the air inlet (22). The aircraft also includes a control system arranged to receive a trajectory command as input and output an actuation command for a set of actuators. At least some of the actuators are capable of generating predominantly vertical motion. During predominantly vertical flight, the control system is further arranged to receive a cooling requirement as input and to output a propeller rotation command (16). During predominantly vertical flight, the control system is further arranged to determine the actuation command based on the propeller rotation command (16). [Fig. 3]

Description

Title of the invention: Control system for cooling onboard equipment of a vertical takeoff and landing aircraft

[0001] The invention relates to the field of vertical take-off and landing aircraft, also known as VTOL in the art (from the English equivalent "Vertical Take-Off and Landing"). More particularly, the invention relates to a control system for cooling onboard equipment of a VTOL aircraft when the latter is in the vertical flight phase.

[0002] A VTOL aircraft generally comprises a propulsion system with a plurality of rotors, each rotor configured to rotate about an axis of rotation relative to a respective stator. The rotating rotors are capable of jointly producing a predominantly vertical motion of the aircraft, particularly during takeoff and landing. A VTOL aircraft can take off from and land on a small ground infrastructure. This makes its use particularly suitable in highly constrained environments, such as cities, and reduces the land footprint required for its operation.

[0003] An aircraft, particularly a VTOL aircraft, includes numerous onboard components that require cooling during flight. For example, these components may include engines, such as turbogenerators, batteries, or avionics equipment.

[0004] Typically, the aircraft has one or more air inlets connected to the equipment to be cooled. The equipment is cooled by an airflow from these air inlets. For example, to cool the engine driving a propeller, a prior art solution consists of having an air inlet connected to the engine downstream of the propeller, so that the airflow generated by the rotation of the propeller cools the engine.

[0005] During takeoff and landing, a non-VTOL aircraft benefits from a forward speed sufficient to generate, in its air intakes, an airflow with a flow rate adequate to cool the onboard equipment. Typically, this forward speed is greater than or equal to 8 m / s.

[0006] Conversely, a VTOL aircraft during takeoff and landing has a low or even zero forward speed. Typically, this forward speed is less than 8 m / s. Under these conditions, the airflow generated in the air intakes is insufficient to effectively cool the onboard equipment.

[0007] One prior art solution involves placing one or more fans near the equipment to be cooled. These fans, through their rotation, ensure an airflow at the necessary flow rate, even when the VTOL aircraft is in a predominantly vertical flight phase. However, this solution is not satisfactory because the presence of the fans adds mass and bulk to the aircraft. Furthermore, the cooling circuit incorporating these fans may be oversized for high-speed flight phases, during which the airflow around the aircraft is high. This oversizing can result in additional drag. In addition, the fans represent a potential source of further failure for the aircraft.

[0008] The invention improves the situation. To this end, it provides a vertical takeoff and landing aircraft. The aircraft includes an air intake connected to a cooling system. The aircraft further includes a horizontal drive propeller arranged to generate airflow in the air intake. The aircraft also includes a control system arranged to receive a trajectory command as input and output an actuation command for a set of actuators. At least some of the actuators are capable of generating a predominantly vertical motion. During the predominantly vertical flight phase, the control system is further arranged to receive a cooling input and output a propeller rotation command. During the predominantly vertical flight phase, the control system is further arranged to determine the actuation command based on the propeller rotation command.

[0009] The invention makes it possible to efficiently cool onboard equipment, even when the aircraft is in a predominantly vertical flight phase. This eliminates the need for fans, or at least reduces their size and power. The reliability of the onboard equipment cooling is thus improved.

[0010] In addition, the control system makes it possible to adapt the rotation of the propeller to the cooling requirements of the on-board equipment while compensating for the effects of this rotation on the trajectory of the aircraft.

[0011] Optional features of the invention, complementary or alternative, are stated below: - the control system is arranged to determine the actuation control according to a PID type servo function; - the actuator assembly includes a vertical drive rotor and the actuation control includes a rotation control of this vertical drive rotor; - the actuator assembly includes a control surface and the actuation control includes a deflection control of this control surface; - the control surface includes an aileron surface, a rudder surface or a elevator surface; - the control system is further arranged to determine the actuation command according to an angle of incidence or a sideslip angle of the aircraft; - the actuation control includes changing the inclination of a longitudinal axis of the aircraft relative to a horizontal plane; - the aircraft exhibits, in a predominantly vertical flight phase, a forward speed strictly less than 8 m / s.

[0012] The invention also relates to a method for piloting a vertical takeoff and landing aircraft during its predominantly vertical flight phase. The aircraft comprises a control system arranged to receive as input a trajectory command and a cooling requirement from equipment to be cooled. The control system is further arranged to output a rotation command for a horizontal drive propeller, so as to generate airflow in an air intake connected to the equipment to be cooled. The control system is further arranged to determine an actuation command for a set of actuators based on the trajectory command and the propeller rotation command. The control system is also arranged to output the actuation command to the set of actuators.

[0013] Other features and advantages of the invention are described in detail in the following description, made with reference to the accompanying drawings, in which: - [Fig.1] is a schematic diagram of the operation of a VTOL aircraft control system according to the invention; - [Fig.2] represents an aircraft according to an embodiment of the invention, in side view; - [Fig.3] is analogous to [Fig.2].

[0014] The drawings and the description below contain, essentially, elements of a definite nature. They may therefore not only serve to better understand the present invention, but also contribute to its definition, if necessary.

[0015] Reference is made to figures 1 to 3.

[0016] Figure 1 shows a schematic view of the operation of a piloting device 10 for a vertical takeoff and landing aircraft, or aircraft 1, according to the invention. Figures 2 and 3 show an embodiment of aircraft 1.

[0017] The aircraft 1 comprises a fuselage 3, which generally extends along a first axis, called the "longitudinal axis" of the aircraft 1. The aircraft 1 is further characterized by a second axis, or lateral axis, perpendicular to its longitudinal axis. This lateral axis intersects the longitudinal axis at a point corresponding to the center of gravity of the aircraft 1. The aircraft 1 is also characterized by a third axis, perpendicular to its longitudinal and lateral axes and passing through the center of gravity of the aircraft 1. This third axis is called the "vertical axis" of the aircraft 1 in the technique, although this axis does not always extend in the vertical direction. The center of gravity, the longitudinal axis, the lateral axis, and the vertical axis of the aircraft 1 together form a frame of reference attached to the aircraft 1, which is referred to as the "aircraft frame of reference" in the following description.

[0018] Aircraft 1 is generally symmetrical, along a plane comprising its longitudinal axis and its vertical axis. This plane is referred to as the "plane of symmetry" of aircraft 1 in the following description. The "transverse plane" of aircraft 1 is the plane comprising the longitudinal axis and the lateral axis of aircraft 1. The "pitch angle" of aircraft 1 is the angle formed between the transverse plane of aircraft 1 and a horizontal plane.

[0019] The "angle of incidence" of aircraft 1 is defined as the angle formed between the longitudinal axis of aircraft 1 and a projection of the relative wind in the plane of symmetry of aircraft 1.

[0020] The axis resulting from a rotation of the longitudinal axis of aircraft 1 in the plane of symmetry is called "axis Al", this rotation having as its center the center of gravity of aircraft 1 and as its angle value the angle of incidence of aircraft 1. The angle formed between this axis Al and the direction of the relative wind is called "angle of sideslip" of aircraft 1.

[0021] The fuselage 3 of the aircraft 1 has a forward portion, shaped as a nose 5, and a rear portion, shaped as a tail 7, which are mutually opposed. In the embodiment shown in Figures 2 and 3, the aircraft 1 comprises a first wing, or front wing 11, attached to the fuselage 3 between the nose 5 and the tail 7, and a second wing, or rear wing 13, attached to the fuselage 3 behind the front wing 11. In addition, the aircraft 1 has an empennage 9, disposed on the tail 7 of the fuselage 3. The empennage 9 has a portion called the "vertical plane", which generally extends along the plane of symmetry of the aircraft 1, and a portion called the "horizontal plane", which generally extends along the transverse plane of the aircraft 1.

[0022] The "speed of forward movement" is defined as the speed of aircraft 1 along its longitudinal axis. The speed of forward movement is determined with reference to a frame of reference independent of aircraft 1, which is referred to as the "terrestrial frame of reference" in the following description. The terrestrial frame of reference includes, for example, a vertical direction and a horizontal direction.

[0023] A "primarily vertical flight phase" is defined as a flight phase during which the aircraft 1 has a low or even zero forward speed. Typically, this speed is strictly less than 8 m / s. The primarily vertical flight phase may, for example, correspond to a takeoff phase, a landing phase, or a hovering phase of the aircraft 1.

[0024] A flight phase during which the aircraft 1 exhibits a high forward speed is called the "forward flight phase". Typically, this value is greater than or equal to 8 m / s.

[0025] The aircraft 1 includes a set of actuators 18, at least some of which participate in creating the essentially vertical movement of the aircraft 1 during the essentially vertical flight phase.

[0026] The actuator assembly 18 can thus comprise one or more vertical drive rotors, which are referred to as rotors in the following description. The rotating rotors are capable of jointly producing an essentially vertical motion of the aircraft, particularly for takeoff and landing phases. Each rotor is configured to rotate about an axis of rotation relative to a respective stator. Generally, the axes of rotation of the rotors are directed substantially parallel to the vertical axis of the aircraft 1. In one embodiment of the invention, the direction of one or more of these axes can be modified relative to the direction of the vertical axis of the aircraft 1. Each rotor further has a plurality of blades, which may have variable pitch.In the embodiment shown in Figures 2 and 3, the actuator assembly 18 comprises eight rotors, distributed over the front wing 11 and the rear wing 13.

[0027] Furthermore, the actuator assembly 18 may include one or more control surfaces. These control surfaces are movable elements attached to the fairing of the aircraft 1. In particular, the control surfaces are attached to the rotating fairing of the aircraft 1. The control surfaces contribute to controlling the stability of the aircraft 1 in roll (rotational movement of the aircraft 1 about its longitudinal axis), in pitch (rotational movement of the aircraft 1 about its lateral axis), and in yaw (rotational movement of the aircraft 1 about its vertical axis). The control surfaces may include aileron surfaces provided on the wings of the aircraft 1, which contribute to controlling the stability of the aircraft 1 in roll.The control surfaces may further include rudder surfaces provided on the vertical portion of the tail assembly 9, which contribute to yaw stability control of the aircraft 1. The control surfaces may also include elevator surfaces provided on the horizontal portion of the tail assembly 9, which contribute to pitch stability control of the aircraft 1.

[0028] Aircraft 1 further comprises a horizontal drive propeller, or propeller 16. The propeller 16 is configured to rotate about an axis of rotation relative to the fuselage 3 of the aircraft 1. This axis is directed substantially parallel to the longitudinal axis of the aircraft 1. The propeller further has a plurality of blades, which may have variable pitch. When the propeller 16 is rotating, it generates a forward thrust force on the aircraft 1, as represented by arrow 36 in Figures 2 and 3. In particular, the propeller 16 contributes to generating the forward thrust of the aircraft 1 during the forward flight phase. In the embodiment shown in Figures 2 and 3, the propeller 16 is arranged on the nose 5 of the aircraft 1.

[0029] The aircraft 1 further includes onboard equipment, or equipment 14, which requires cooling during the predominantly vertical flight phase. The equipment 14 may include one or more engines, such as turbogenerators, one or more batteries, or avionics equipment. Typically, the equipment 14 is housed in the fuselage 3 of the aircraft 1, between the nose 5 and the tail 7.

[0030] The aircraft 1 also has, near the propeller 16, at least one air intake 22. The air intake 22 is connected to the equipment 14, for example, via an air duct (not shown). Typically, this air duct exits the fuselage 3 through an air outlet (not shown). For example, the air intake 22 may be of the type known as a NACA air intake or of the type known as a SCOOP air intake.

[0031] Aircraft 1 may further include a heat exchanger (not shown), connected to equipment 14. This heat exchanger is arranged to facilitate the dissipation of heat from equipment 14 by the airflow from the air inlet 22.

[0032] The device 10 includes a control system 12, arranged to receive as input a trajectory command 20 and to output an actuation command for the set of actuators 18.

[0033] Typically, the trajectory command 20 is entered by an aircraft pilot 1, from an input human-machine interface (not shown). For example, this interface may include one or more control sticks. During the predominantly vertical flight phase, the trajectory control 20 may include a pitch rate, a roll rate, a yaw rate, a horizontal speed, and / or a vertical speed. The pitch rate, roll rate, and yaw rate are each determined relative to the aircraft frame of reference. The horizontal speed and vertical speed are each determined relative to the ground frame of reference.

[0034] During the essentially vertical flight phase, the control system 12 receives as input the trajectory command 20 and outputs an actuation command for the actuator assembly 18. This actuation command may include a rotation control of one or more of the rotors in this assembly. For example, for one of these rotors, the rotation control may provide as output a speed value or direction of rotation, a tilt value of the axis of rotation, or a blade pitch angle value. In addition, the actuation control may include a deflection control of one or more of the control surfaces of the actuator assembly 18. For example, for one of these surfaces, the deflection control may provide a deflection angle value.

[0035] The control system 12 is further arranged to receive as input a cooling requirement for the equipment 14 and to output a rotation command for the propeller 16 during the predominantly vertical flight phase. For example, this rotation command can provide as output a rotational speed value or a blade pitch angle value for the propeller 16.

[0036] The rotation of the propeller 16 generates an airflow in the air inlet 22, as designated in Figures 2 and 3 by reference numeral 32. This airflow is directed to the equipment 14 in order to cool it. The more the equipment 14 needs to be cooled, the more the control system 12 increases the rotational speed of the propeller 16. In this way, even during a predominantly vertical flight phase, the equipment 14 receives an airflow with a flow rate sufficient to meet its cooling requirements. Furthermore, this flow rate can be controlled in real time.

[0037] The rotation of the propeller 16 generates effects on the forces and moments acting on the aircraft 1. The rotation of the propeller 16 primarily increases the thrust force acting on the aircraft 1. This rotation also increases the roll moment of the aircraft 1 (anti-torque moment). Furthermore, this rotation causes the formation of an air vortex, which exerts a lateral force on the wings and the vertical portion of the tail assembly 9, thereby generating an additional yaw moment.

[0038] In addition, in the event of wind around the aircraft 1, the rotation of the propeller 16 induces additional effects, which depend on the angle of incidence and the angle of sideslip of the aircraft 1. These additional effects may also depend on the relative position of the center of rotation of the propeller 16 and the center of gravity of the aircraft 1, in particular along the vertical axis of the aircraft 1.

[0039] The additional effects induced by the rotation of the propeller 16 in the case of external wind are detailed below.

[0040] The center of rotation of propeller 16 is called center C. The axis X is the axis passing through center C and parallel to the longitudinal axis of aircraft 1. The axis X corresponds to the axis of rotation of propeller 16. The axis Y is the axis passing through center C and parallel to the lateral axis of aircraft 1. The axis Z is the axis passing through center C and parallel to the vertical axis of aircraft 1.

[0041] When the angle of incidence and the angle of sideslip are negative, the rotation of the propeller 16 generates, on the center of rotation of the propeller 16, a thrust force along the X axis, as well as an anti-torque moment around this X axis.

[0042] When the angle of incidence is non-zero and the angle of sideslip is zero, the rotation of the propeller 16 generates, on the center of rotation of the propeller 16, a thrust force along the X axis, normal and lateral forces along the Y axis, an anti-torque moment about the X axis, as well as a moment called "IP" about the Z axis. A moment called "IP" is a cyclic moment, the period of which corresponds to one revolution of the propeller.

[0043] When the angle of incidence is zero and the angle of sideslip is non-zero, the rotation of the propeller 16 generates, on the center of rotation of the propeller 16, a thrust force along the X axis, a normal force along the Y axis, a force along the Z axis, an anti-torque moment around the X axis, as well as an IP moment around the Y axis.

[0044] When the angle of incidence and the angle of sideslip are not harmed, the rotation of the propeller 16 generates, on the center of rotation of the propeller 16, a thrust force along the X axis, a normal force along the Y axis, a force along the Z axis, an anti-torque moment along the X axis as well as IP moments around the Y and Z axes.

[0045] These forces and moments acting on the center C of rotation of the propeller 16 result in effects on the forces and moments acting on the center of gravity of the aircraft 1.

[0046] When the propeller 16 is rotating, the effects described above concerning the forces and moments applied to the aircraft 1 tend to cause it to deviate from the trajectory targeted by the trajectory control 20. The control system 12 according to the invention is arranged to determine the actuation command of the actuator assembly 18 in such a way as to compensate for these effects and to ensure that the aircraft 1 maintains a trajectory conforming to the trajectory control 20.

[0047] In this perspective, the control system 12 is arranged to determine the actuation command for the actuator assembly 18 as a function of the rotation of the propeller 16. The control system 12 may be based on measurements from sensors and / or directly use values ​​from the rotation command of the propeller 16 to determine the effects of the propeller 16's rotation to be compensated for. The control system 12 includes a servo function, which may be of the PID type, for example. The actuation command may include modifying the rotation command of one or more of the rotors, or modifying the deflection command of one or more of the control surfaces.

[0048] In this way, the control system 12 ensures that the equipment 14 is cooled efficiently while guaranteeing that the trajectory predicted in the trajectory control 20 is followed. In particular, the equipment 14 is cooled efficiently without the need to integrate an additional fan on the aircraft 1. Besides the Thanks to the reduction in size and weight due to the absence of the fan itself, the invention can, depending on the configuration, allow for a reduction in the dimensions of the air inlet, air outlet, and heat exchanger used to cool the onboard equipment. This, in turn, can also reduce drag during high-speed flight.

[0049] Figures 2 and 3 illustrate an example of an implementation of the control system according to the invention. In this example, the control system 12 receives a trajectory command 20 corresponding to a stationary motion. This command includes zero values ​​for roll, pitch, and yaw rates (in the aircraft frame of reference), as well as zero values ​​for vertical and horizontal speeds (in the ground frame of reference).

[0050] In [Fig.2], the control system 12 does not yet take into account a rotation of the propeller 16 to determine the actuation command of the actuator assembly 18. The control system 12 determines this actuation command so that the thrust force of the rotors, represented by arrow 38, is substantially equal to the weight of the aircraft 1, represented by arrow 34. Here, the pitch angle of the aircraft 1 is zero.

[0051] Furthermore, the control system 12 rotates the propeller 16 to cool the equipment 14 by means of the airflow 32. Here, this rotation generates the thrust force represented by the arrow 36. This thrust force is directed forward of the aircraft 1, along the longitudinal axis of the aircraft 1. Under the effect of the rotation of the propeller 16, the aircraft 1 tends to move forward. The aircraft 1 therefore has a non-zero actual horizontal speed, contrary to what is predicted by the trajectory control 20.

[0052] In [Fig. 3], the control system 12 determines the actuation command for the actuator assembly 18 as a function of the rotation of the propeller 16, so as to compensate for the effects generated by this rotation. Here, the control system 12 detects a discrepancy between a measurement of the actual horizontal speed of the aircraft 1, for example provided by a sensor, and the zero horizontal speed value provided by the trajectory control 20. The control system 12 then determines an actuation command comprising pitching the aircraft 1 up, that is, increasing its pitch angle, so that the resultant of the rearward thrust force of the rotors compensates for the resultant of the forward thrust force of the propeller 16.This actuation control includes, in particular, increasing the thrust force of the rotors arranged on the front wing 11, for example by increasing their rotational speed, relative to the thrust force of the rotors arranged on the rear wing 13.

[0053] The pitch angle of aircraft 1 is increased by a positive value A2. The nose 5 of aircraft 1 is oriented upwards. The direction of the thrust force 38 of the rotors is inclined at an angle A2 to the vertical direction, towards the rear.

[0054] The thrust force 38 of the rotors can then be broken down into a vertical component, represented by arrow 39, and a horizontal component, represented by arrow 40. Similarly, the thrust force 36 of the propeller 16 can be broken down into a vertical component and a horizontal component.

[0055] The vertical component 39 of the thrust force 38 of the rotors has a direction opposite to that of the weight of the aircraft 1. The actuation control is determined so that the resultant of this vertical component 39, the weight 34 and the vertical component of the thrust force 36 of the propeller 16 is substantially zero.

[0056] The horizontal component 40 of the thrust force 38 of the rotors has a direction opposite to that of the horizontal component of the thrust force 36 of the propeller 16. The actuation control is determined so that the resultant of these two components is zero.

[0057] In addition, the actuation control may include compensating for the roll, pitch, and yaw moments induced by the rotation of the propeller 16. The actuation control may include compensating for the roll moment by generating a differential in thrust force between the rotors located on the right and the rotors located on the left of the aircraft 1, for example, by varying their respective rotational speeds. The actuation control may further include compensating for the pitch moment by generating a differential in thrust force between the rotors located on the front wing 11 and the rotors located on the rear wing 13. The actuation control may also include compensating for the yaw moment by rotating some of the rotors clockwise and others counterclockwise.

[0058] The trajectory of aircraft 1 is therefore not affected by the increment of rotation of propeller 16 required for cooling purposes.

[0059] This actuation control results in an unusual positioning of the aircraft 1. Indeed, ordinarily, in the essentially vertical flight phase, it is preferred that the nose of the VTOL aircraft be slightly inclined downwards, for example by about 3° relative to a horizontal plane, in order to improve visibility for the pilot.

[0060] The magnitude of the vertical component 39 is equal to the magnitude of the thrust force 38, multiplied by the cosine of angle A2. The magnitude of the horizontal component 40 is equal to the magnitude of the thrust force 38, multiplied by the sine of angle A2. Thus, an inclination of only a few degrees by an angle A2 is sufficient to generate, for the rotors, a rearward horizontal thrust component 40 sufficient to counteract the forward thrust of the propeller 16. This inclination does not require a significant increase in the magnitude of the thrust force 38 to obtain an upward vertical component 39 sufficient to oppose the weight of the aircraft 1.

[0061] For example, consider an aircraft weighing approximately 2000 kg equipped with a propeller with a diameter of approximately 1.75 m. The control system rotates this propeller at a speed of 800 rpm during the predominantly vertical flight phase, requiring a power output of approximately 4 kW. This rotation generates an airflow with a speed of approximately 8 m / s, sufficient to cool the onboard equipment. The propeller rotation also generates a forward thrust force of approximately 350 N, which can be counteracted with a bank angle A2 of only approximately 1°.

[0062] The invention is not limited to the embodiments described above, but encompasses all variants conceivable by those skilled in the art. In particular, the aircraft may include an additional air intake, connected to the engine driving the propeller, so that, when this propeller is rotating, the resulting airflow also cools this engine. Furthermore, the control system may include an additional feedback loop, connected to the temperature of the equipment to be cooled, so as to determine the propeller's rotational direction.

Claims

Demands

1. Vertical takeoff and landing aircraft, the aircraft comprising: - an air inlet (22), connected to equipment (14) to be cooled; - a horizontal drive propeller (16), arranged to generate an airflow in the air inlet (22); - a control system (12), arranged to receive as input a trajectory command and to output an actuation command for a set of actuators (18), at least some of which are capable of generating an essentially vertical motion; characterized in that, in the essentially vertical flight phase, the control system (12) is further arranged to receive as input a cooling requirement and to output a rotation command for the propeller (16), and the control system (12) is arranged to determine the actuation command as a function of the rotation command for the propeller (16).

2. Aircraft according to claim 1, wherein the control system (12) is arranged to determine the actuation control according to a PID type servo function.

3. Aircraft according to any one of claims 1 and 2, wherein the actuator assembly (18) includes a vertical drive rotor and the actuation control includes a rotation control of this vertical drive rotor.

4. Aircraft according to any one of the preceding claims, wherein the actuator assembly (18) includes a control surface and the actuation control includes a deflection control of this control surface.

5. Aircraft according to the preceding claim, wherein the control surface comprises an aileron surface, a rudder surface or a elevator surface.

6. Aircraft according to any one of the preceding claims, wherein the control system (12) is further arranged to determine the actuation command as a function of an angle of incidence or a sideslip angle of the aircraft.

7. Aircraft according to any one of the preceding claims, wherein the actuation control comprises changing the inclination of a longitudinal axis of the aircraft relative to a horizontal plane.

8. Aircraft according to any one of the preceding claims, wherein the aircraft has, in essentially vertical flight phase, a forward speed strictly less than 8 m / s.

9. Method of piloting a vertical takeoff and landing aircraft in a predominantly vertical flight phase, the aircraft comprising a control system (12) arranged to: - receive as input a trajectory command, and receive as input a cooling requirement for equipment (14) to be cooled; - output a rotation command for a horizontal drive propeller (16), so as to generate an airflow in an air inlet (22) connected to the equipment (14) to be cooled; - determine an actuation command for a set of actuators (18) as a function of the trajectory command and the rotation command for the propeller (16); - output the actuation command to the set of actuators (18).

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