Control system for cooling on-board equipment of a vertical takeoff and landing aircraft

Abstract

A vertical takeoff and landing aircraft. The aircraft comprises an air inlet (22), connected to equipment (14) to be cooled. The aircraft further comprises a horizontal drive propeller (16), arranged to generate an air flow in the air inlet (22). The aircraft further comprises a control system, arranged to receive as input a trajectory command and to emit as output an actuation command for a set of actuators. At least some of the set of actuators are capable of generating an essentially vertical movement. In the essentially vertical flight phase, the control system is further arranged to receive as input a cooling requirement and to emit as output a rotation command for the propeller (16). In the essentially vertical flight phase, the control system is further arranged to determine the actuation command as a function of the rotation command for the propeller (16). Abstract figure: [Fig.3]

Description

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

[0001] The invention relates to the field of vertical take-off and landing aircraft, also referred to 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 on-board equipment of a VTOL aircraft, when the latter is in the vertical flight phase.

[0002] A VTOL aircraft generally comprises a propulsion system having a plurality of rotors, each rotor being configured to rotate about an axis of rotation relative to a respective stator. The rotors, when rotated, are capable of jointly producing an essentially vertical movement of the aircraft, in particular for the takeoff and landing phases thereof. A VTOL aircraft can take off from and land on a reduced ground infrastructure. This makes its use particularly suitable in highly constrained environments, such as cities for example, and makes it possible to reduce the footprint required for its operation.

[0003] An aircraft, particularly a VTOL aircraft, includes numerous on-board devices that require cooling during flight. For example, these devices 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 air flow coming from these air inlets. For example, to cool the engine turning a propeller, a solution of the state of the art consists of having, downstream of the propeller, an air inlet connected to the engine, so that the air flow generated by the rotation of the propeller cools the engine.

[0005] During takeoff and landing, a non-VTOL aircraft benefits from a forward speed capable of generating, in its air intakes, an air flow whose flow rate is sufficient to cool the on-board equipment. Typically this forward speed is greater than or equal to 8 m / s.

[0006] On the other hand, a VTOL aircraft in the takeoff and landing phase 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 has an insufficient flow rate to effectively cool the on-board equipment.

[0007] A solution of the state of the art consists of placing one or more fans near the equipment to be cooled. These fans ensure, by their rotation, an air flow at the necessary flow rate, even when the VTOL aircraft is in a phase of essentially vertical flight. This solution is not, however, satisfactory, because the presence of the fans induces additional mass and bulk for the aircraft. Furthermore, the cooling circuit comprising these fans may be oversized for high-speed flight phases, during which the flow rate of the air flow around the aircraft is high. This oversizing may result in additional drag. Furthermore, the fans represent a potential source of additional failure for the aircraft.

[0008] The invention improves the situation. To this end, it proposes a vertical takeoff and landing aircraft. The aircraft comprises an air intake, connected to equipment to be cooled. The aircraft further comprises a horizontal drive propeller, arranged to generate an air flow in the air intake. The aircraft further comprises a control system, arranged to receive as input a trajectory command and to emit as output an actuation command for a set of actuators. At least some of the set of actuators are capable of generating an essentially vertical movement. In the essentially vertical flight phase, the control system is further arranged to receive as input a cooling requirement and to emit as output a rotation command for the propeller. In the essentially vertical flight phase, the control system is further arranged to determine the actuation command as a function of the rotation command for the propeller.

[0009] The invention makes it possible to effectively cool on-board equipment, even when the aircraft is in a phase of essentially vertical flight. It is thus possible to do without fans, or at least to reduce their size and power. The reliability of the cooling of on-board equipment is improved.

[0010] Furthermore, 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 substitutable, are set out below: - the control system is arranged to determine the actuation command according to a PID type servo function; - the actuator assembly comprises a vertical drive rotor and the actuation control comprises a rotation control of this vertical drive rotor; - the actuator assembly comprises a control surface and the actuation control comprises a deflection control of this control surface; - the control surface comprises an aileron surface, a rudder surface or an elevator surface; - the control system is further arranged to determine the actuation command as a function of an angle of incidence or a sideslip angle of the aircraft; - the actuation control comprises modifying the inclination of a longitudinal axis of the aircraft relative to a horizontal plane; - the aircraft has, in the essentially 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 in an essentially vertical flight phase. The aircraft comprises a control system, arranged to receive as input a trajectory command, and to receive as input a cooling requirement of equipment to be cooled. The control system is further arranged to output a rotation command of a horizontal drive propeller, so as to generate an air flow in an air inlet connected to the equipment to be cooled. The control system is further arranged to determine an actuation command of a set of actuators as a function of the trajectory command and the rotation command of the propeller. The control system is further arranged to output the actuation command to the set of actuators.

[0013] Other characteristics and advantages of the invention are set out in detail in the following description, given with reference to the appended drawings, in which: - [Fig.l] is a schematic diagram of the operation of a control system of the VTOL aircraft according to the invention; - [Fig.2] represents an aircraft according to one embodiment of the invention, in side view; - [Fig.3] is analogous to [Fig.2].

[0014] The drawings and the description below contain, for the most part, elements of a certain character. 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] [Fig. 1] represents a schematic view of the operation of a piloting device 10 for a vertical take-off and landing aircraft, or aircraft 1, according to the invention. Figures 2 and 3 represent an embodiment of the 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 further characterized by a third axis, perpendicular to its longitudinal axis and to its lateral axis 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 art, 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 reference frame attached to the aircraft 1, which is called the "aircraft reference frame" in the remainder of the description.

[0018] The aircraft 1 is generally symmetrical, along a plane comprising its longitudinal axis and its vertical axis. This plane is called the "plane of symmetry" of the aircraft 1 in the remainder of the description. The plane comprising the longitudinal axis and the lateral axis of the aircraft 1 is called the "transverse plane". The angle formed between the transverse plane of the aircraft 1 and a horizontal plane is called the "attitude angle".

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

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

[0021] The fuselage 3 of the aircraft 1 has a front part, shaped as a nose 5, and a rear part, shaped as a tail 7, mutually opposite. 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 a tail 9, arranged on the tail 7 of the fuselage 3. The tail 9 has a so-called "vertical plane" part, which generally extends along the plane of symmetry of the aircraft 1, and a so-called "horizontal plane" part, which generally extends along the transverse plane of the aircraft 1.

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

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

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

[0025] The aircraft 1 comprises 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 may thus comprise one or more vertical drive rotors, which are called rotors in the remainder of the description. The rotors, when rotated, are capable of jointly producing an essentially vertical movement of the aircraft, in particular for the takeoff and landing phases thereof. Each rotor is configured to rotate around 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 may be modifiable relative to the direction of the vertical axis of the aircraft 1. Each rotor further has a plurality of blades, which may be of 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 comprise 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 fairing of the aircraft 1 for rotation. The control surfaces participate in 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 comprise aileron surfaces provided on the wings of the aircraft 1, which participate in controlling the stability of the aircraft 1 in roll.The control surfaces may further comprise rudder surfaces provided on the vertical plane portion of the empennage 9, which participate in controlling the stability of the aircraft 1 in yaw. The control surfaces may further comprise elevator surfaces provided on the horizontal plane portion of the empennage 9, which participate in controlling the stability of the aircraft 1 in pitch.

[0028] The 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 be variable pitch. When the propeller 16 is rotating, it generates a forward thrust force for the aircraft 1, as represented by the arrow 36 in FIGS. 2 and 3. In particular, the propeller 16 participates in generating the forward thrust of the aircraft 1 during the forward flight phase. In the embodiment shown in FIGS. 2 and 3, the propeller 16 is arranged on the nose 5 of the aircraft 1.

[0029] The aircraft 1 further comprises on-board equipment, or equipment 14, which needs to be cooled during the essentially vertical flight phase. The equipment 14 may comprise one or more engines, such as turbogenerators, one or more batteries, or even 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 inlet 22. The air inlet 22 is connected to the equipment 14, for example via an air duct (not shown). Typically, this air duct opens from the fuselage 3 via an air outlet (not shown). For example, the air inlet 22 may be of the type known as a NACA air intake or of the type known as a SCOOP air intake.

[0031] The aircraft 1 may further comprise a heat exchanger (not shown), connected to the equipment 14. This heat exchanger is arranged so as to facilitate the dissipation of heat from the equipment 14 by the flow of air coming from the air inlet 22.

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

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

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

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

[0036] The rotation of the propeller 16 generates an air flow in the air inlet 22, as designated in FIGS. 2 and 3 by the reference 32. This air flow is conducted to the equipment 14 so as to cool the latter. The more the equipment 14 needs to be cooled, the more the control system 12 increases the rotation speed of the propeller 16. In this way, even in the essentially vertical flight phase, the equipment 14 receives an air flow whose flow rate is sufficient to cover its cooling needs. Furthermore, the control of this flow rate can be done in real time.

[0037] The rotation of the propeller 16 generates effects on the forces and moments that apply to the aircraft 1. The rotation of the propeller 16 has the main effect of increasing the thrust force applied to the aircraft 1. This rotation also has the effect of increasing the rolling moment of the aircraft 1 (anti-torque moment). This rotation also has the effect of causing the formation of an air vortex, which exerts a lateral force on the wings and the vertical plane portion of the empennage 9 so as to generate an additional yaw moment.

[0038] Furthermore, 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 sideslip angle 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 event of external wind are detailed below.

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

[0041] When the angle of incidence and the sideslip angle are equal, 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 sideslip angle 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 around the X axis, as well as a so-called "IP" moment around the Z axis. A so-called "IP" moment is a cyclical moment, the period of which corresponds to one turn of the propeller.

[0043] When the angle of incidence is zero and the sideslip angle 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 sideslip angle are not affected, 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 applied to the center C of rotation of the propeller 16 result in effects on the forces and moments applied to 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 set of actuators 18 so as to compensate for these effects and so that the aircraft 1 maintains a trajectory in accordance with the trajectory control 20.

[0047] In this perspective, the control system 12 is arranged to determine the actuation command of the set of actuators 18 as a function of the rotation of the propeller 16. The control system 12 can be based on measurements from sensors and / or directly use values ​​of the rotation command of the propeller 16 to determine the effects of the rotation of the propeller 16 to be compensated. The control system 12 comprises a servo-control function, which can be of the PID type, for example. The actuation command can comprise 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 provided in the trajectory control 20 is respected. In particular, the equipment 14 is cooled efficiently without the need to integrate an additional fan on the aircraft 1. In addition to the saving in mass and space due to the absence of the fan itself, the invention can allow, depending on the configuration, to reduce the dimensions of the air inlet, the air outlet and the heat exchanger provided for cooling the on-board equipment. Thus, drag in high-speed flight can also be reduced.

[0049] Figures 2 and 3 illustrate an example of 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 movement. This command includes zero values ​​of roll, pitch and yaw rates (in the aircraft frame of reference), as well as zero values ​​of vertical speed and horizontal speed (in the terrestrial 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 set of actuators 18. The control system 12 determines this actuation command in such a way that the thrust force of the rotors, represented by the arrow 38, is substantially equal to the weight of the aircraft 1, represented by the arrow 34. Here, the attitude angle of the aircraft 1 is zero.

[0051] Furthermore, the control system 12 rotates the propeller 16 so as to cool the equipment 14 by the air flow 32. Here, this rotation generates the thrust force represented by the arrow 36. This thrust force is directed towards the front 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 real horizontal speed value, contrary to what is provided by the trajectory control 20.

[0052] In [Fig. 3], the control system 12 determines the actuation command of 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 notes a difference 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 up the aircraft 1, that is to say increasing its attitude angle, so that the resultant of the thrust force of the rotors towards the rear compensates for the resultant of the thrust force of the propeller 16 towards the front.This actuation command comprises, in particular, increasing the thrust force of the rotors arranged on the front wing 11, for example by increasing their rotation speed, relative to the thrust force of the rotors arranged on the rear wing 13.

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

[0054] The thrust force 38 of the rotors is then broken down into a vertical component, represented by the arrow 39, and a horizontal component, represented by the 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 command is determined so that the resultant of this vertical component 39, of the weight 34 and of 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 command is determined so that the resultant of these two components is zero.

[0057] Furthermore, the actuation control may comprise compensating for the roll, pitch and yaw moments induced by the rotation of the propeller 16. The actuation control may comprise compensating for the roll moment by generating a thrust force differential between the rotors arranged on the right and the rotors arranged on the left of the aircraft 1, for example by means of a variation in their respective rotational speeds. The actuation control may further comprise compensating for the pitch moment by generating a thrust force differential between the rotors arranged on the front wing 11 and the rotors arranged on the rear wing 13. The actuation control may further comprise compensating for the yaw moment by rotating a portion of the rotors clockwise and the other portion counterclockwise.

[0058] The trajectory of the aircraft 1 is thus not affected by the rotation increment of the propeller 16 necessary for cooling requirements.

[0059] This actuation command 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 approximately 3° relative to a horizontal plane, in order to improve visibility for the pilot.

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

[0061] For example, consider an aircraft of 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 essentially vertical flight phase, which requires a power of the order of 4 kW. This rotation generates an air flow with a speed of approximately 8 m / s, suitable for cooling the on-board equipment. The rotation of the propeller also generates a forward thrust force of approximately 350 N, which can be compensated with an inclination of an 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 comprise an additional air intake, connected to the engine rotating the propeller, so that, when this propeller is rotating, the resulting air flow also cools this engine. Furthermore, the control system may comprise an additional servo loop, connected to the temperature of the equipment to be cooled, so as to determine the rotation control of the propeller.

Claims

Claims

1. Vertical take-off 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 air flow in the air inlet (22); - a control system (12), arranged to receive as input a trajectory command and to emit as output an actuation command of a set of actuators (18), at least some of which are capable of generating an essentially vertical movement; 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 emit as output a rotation command of the propeller (16), and the control system (12) is arranged to determine the actuation command as a function of the rotation command of the propeller (16).

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

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

4. Aircraft according to one of the preceding claims, in which the set of actuators (18) comprises a control surface and the actuation control comprises a deflection control of this control surface.

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

6. Aircraft according to 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. An aircraft according to any preceding claim, wherein the actuation control comprises modifying the inclination of a longitudinal axis of the aircraft relative to a horizontal plane.

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

9. Method for piloting a vertical takeoff and landing aircraft in an essentially 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 of equipment (14) to be cooled; - emit as output a rotation command of a horizontal drive propeller (16), so as to generate an air flow in an air inlet (22) connected to the equipment (14) to be cooled; - determine an actuation command of a set of actuators (18) as a function of the trajectory command and the rotation command of the propeller (16); - emit as output the actuation command to the set of actuators (18).

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