US-12625499-B2 - Method for estimating the rotor torques of an aircraft capable of hovering and control unit for an aircraft capable of hovering

US12625499B2US 12625499 B2US12625499 B2US 12625499B2US-12625499-B2

Abstract

A method for estimating rotor torques of an aircraft capable of hovering and comprising a plurality of rotors, which are rotatable under the action of respective rotor torques; and an engine, which is operatively connected to the rotors to provide them with an engine torque. Each rotor comprises a hub and a plurality of blades articulated on the respective hub in such a way that respective collective pitch angles are adjustable. The method comprises the steps of i) calculating a symmetric component on the basis of the engine torque; ii) receiving a signal associated with collective pitch angles; iii) calculating an asymmetric component on the basis of a pitch angle difference between the collective pitch angles; and iv) calculating each rotor torque as the algebraic sum of the symmetric component and the respective asymmetric component.

Inventors

Alberto Angelo Trezzini
Ahmad Mohamad HAIDAR

Assignees

LEONARDO S.P.A.

Dates

Publication Date: 20260512
Application Date: 20221020
Priority Date: 20211227

Claims (16)

1 . A method for estimating rotor torques of an aircraft capable of hovering; said aircraft comprising: a plurality of rotors, which are operatively connected to each other and rotatable relative to respective rotational axes under the action of respective rotor torques; and at least one first engine, which is operatively connected to said rotors and is adapted to provide said rotors with an engine torque; said rotors each comprising a hub and a plurality of blades articulated on said respective hub in such a way that respective collective pitch angles of said plurality of blades relative to the respective rotational axis are adjustable; said method comprising the steps of: i) calculating a symmetric component of said rotor torques on the basis of said engine torque; said symmetric component being equal to said rotor torques, when said collective pitch angles are equal to each other; characterized in that it comprises the further steps of: ii) receiving a signal associated with said collective pitch angles; iii) calculating an asymmetric component of each said rotor torques on the basis of at least a pitch angle difference between said collective pitch angles; and iv) calculating each said rotor torque as the algebraic sum of said symmetric component and said respective asymmetric component.
2 . The method according to claim 1 , wherein said rotors are associated with respective pitch angle differences; characterized in that said step iii) of calculating said asymmetric component of each said rotor torques comprises the steps of: v) determining a parameter associated with a variation of each said rotor torque with respect to a variation of said collective pitch angle of the same rotor; and vi) multiplying each said parameter by said respective pitch angle difference of the respective rotor.
3 . The method according to claim 2 , characterized in that said step v) comprises the step vii) of determining said parameters on the basis of at least two independent variables associated with the flight conditions of said aircraft.
4 . The method according to claim 3 , characterized in that said independent variables include: an angle corresponding to an orientation of said rotational axes of said rotors with respect to a reference system fixed to said aircraft; the airspeed of said aircraft; and/or said symmetric component calculated at said step i).
5 . The method according to claim 4 , characterized in that said step vii) comprises the further steps of: viii) determining each said parameter on the basis of said airspeed, if the angle of the respective rotor is lower than or equal to a threshold value; and ix) determining each said parameter on the basis of said symmetric component, if the angle of the respective rotor is greater than said threshold value.
6 . The method according to claim 1 , wherein said aircraft further comprises a second engine, which is operatively connected to said rotors and is adapted to provide said rotors with a further engine torque; characterized in that said step i) comprises the steps of: x) calculating a total torque produced by said first engine and second engine as the sum of said engine torque and said further engine torque; xi) subtracting from said total torque a first subtrahend term corresponding to the transmission losses due to the transmission of said engine torque and said further engine torque from said first engine and said second engine to said rotors; and/or subtracting from said total torque a second subtrahend term corresponding to loads imparted by said first engine and/or said second engine to accessories of said aircraft; and xii) dividing said sum obtained after said step xi) by the number of rotors of said aircraft.
7 . The method according to claim 2 , characterized in that the pitch angle difference of each said rotor is a difference between said respective collective pitch angle of the same rotor and a symmetric collective pitch angle; said symmetric collective pitch angle being the sum of collective pitch angles of all rotors divided by the number of said rotors.
8 . Control unit for an aircraft capable of hovering; said control unit being programmed to: calculate a symmetric component of rotor torques of rotors of said aircraft on the basis of an engine torque provided to said rotors by at least one first engine of said aircraft; said symmetric component being equal to said rotor torques, when collective pitch angles of respective rotors are, in use, equal to each other; said control unit being characterized in that it is further programmed to: receive a signal associated with said collective pitch angles; calculate an asymmetric component of each said rotor torque on the basis of at least a pitch angle difference of the respective rotor; and calculate each said rotor torque as the algebraic sum of said symmetric component and said respective asymmetric component.
9 . Control unit according to claim 8 , characterized in that it is configured to: determine a parameter associated with a variation of each said rotor torque with respect to a variation of said collective pitch angle of the same rotor; multiply each said parameter by said respective pitch angle difference.
10 . Control unit according to claim 9 , characterized in that it is configured to calculate each said parameter on the basis of at least two independent variables associated with the flight conditions of said aircraft.
11 . Control unit according to claim 10 , characterized in that said independent variables include: an angle corresponding to an orientation of said rotational axes of each said rotors with respect to a reference system of said aircraft; airspeed of said aircraft; and/or said symmetric component.
12 . Control unit according to claim 11 , characterized in that it is configured to: calculate each said parameter on the basis of said airspeed, if said respective angle is lower than a threshold value; and calculate each said parameter on the basis of said symmetric component, if said respective angle is equal to or greater than said threshold value.
13 . Control unit according to claim 12 , characterized in that it is configured to: calculate a total torque produced by said first engine and at least one second engine as the sum of said engine torque and a further engine torque of said second engine; subtract from said total torque a first subtrahend term corresponding to the transmission losses due to the transmission of said engine torques from said first and second engines to said rotors; and/or subtract from said total torque a second subtrahend term corresponding to loads imparted by said first and/or second engines to accessories of said aircraft; and divide said difference obtained after the subtraction of said first and/or said second subtrahend term from said total torque by the number of rotors of said aircraft.
14 . Control unit according to claim 9 , characterized in that it comprises a computational unit, a memory, and at least one interface unit electrically and operatively connectable to sensor means of said aircraft; said memory comprising, in turn, a database storing data correlating said parameter with a plurality of variables associated with the flight conditions of said aircraft.
15 . Aircraft capable of hovering, comprising: a plurality of rotors; said rotors being rotatable relative to respective rotational axes under the action of respective rotor torques; said rotors being operatively connected to each other; said rotors comprising each a hub and a plurality of blades articulated on said respective hub in such a way that respective collective pitch angles of said plurality of blades relative to the respective rotational axis are adjustable; at least one first engine, which is operatively connected to said rotors and is adapted to provide said rotors with an engine torque; first sensor means configured to measure said collective pitch angles; and second sensor means configured to measure said engine torque generated, in use, by said first engine; characterized in that it comprises a control unit according to claim 8 , which is operatively connected to said first and second sensor means.
16 . Aircraft according to claim 15 , characterized in that it comprises: third sensor means configured to measure the angles corresponding to an orientation of said rotational axes of said rotors with respect to a reference system of said aircraft and operatively connected to said control unit; and/or fourth sensor means configured to measure an airspeed of said aircraft and operatively connected to said control unit; and/or characterized in that it is a convertiplane or a helicopter; and/or characterized by comprising a second engine operatively connected to said rotors and adapted to provide said rotors with a further engine torque.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS This Patent Application is a U.S. National Phase Application under 35 U.S.C. § 371 of International Patent Application No. PCT/IB2022/060086, filed on Oct. 20, 2022, which claims priority from European Patent Application No. 21217859.4, filed on Dec. 27, 2021, all of which are incorporated by reference as if expressly set forth in their respective entireties herein. TECHNICAL FIELD The present invention relates to a method for estimating the rotor torques of an aircraft capable of hovering. The present invention also relates to a control unit for an aircraft capable of hovering. BACKGROUND Aircraft capable of hovering, such as convertiplanes or helicopters, are known comprising: at least two rotors, which are operatively connected to each other and controllable independently of each other;at least one engine, which is operatively connected to the rotors and is adapted to provide them with an engine torque; andtransmission units operatively connecting the engine to the rotors. In particular, the rotors are driven in rotation by respective rotor torques, which are in general different from the engine torque for several reasons. First, a portion of the engine torque is lost because of mechanical losses in the transmission units. In addition, a portion of the engine torque may be used to drive one or more accessories of the aircraft. The known aircraft further comprise physical sensors arranged at the two rotors and configured to directly measure the rotor torques. However, the rotor torques measured by such sensors (e.g., strain-based torque sensors) are not consistently reliable. In addition, the use of these physical sensors in aircraft results in increased installation and maintenance costs, as well as increased weight and complexity of the aircraft. This is especially relevant for the convertiplanes, the rotors of which are known to be tiltable relative to a reference system fixed with respect to stationary parts of the convertiplanes. Indeed, each time the rotors of the convertiplane are tilted, the physical sensors are tilted integrally with the rotors. This complicates the tilting movement of the rotors and the electrical connection of the sensors to the fixed parts of the convertiplane. Methods have therefore been developed to estimate the rotor torques acting on each of the two rotors without the physical sensors. In particular, the methods include the steps of: subtracting from the engine torque the mechanical losses due to the transmission of the engine torque from the engine to the rotors; andsubtracting from the engine torque the mechanical energy supplied by the engine to any accessories of the aircraft connected thereto. The result of these subtractions is the total available engine torque. The rotor torque is then estimated by dividing the total available engine torque by the number of rotors of the aircraft. In other words, the torque acting on each rotor is estimated by apportioning the total available engine torque produced by the engine equally between the two rotors. Such an estimate can be considered sufficiently accurate under the assumption that the rotor torques acting on the two rotors are equal or substantially equal to each other. However, the rotor torques of two independently controllable rotors may actually be significantly different from each other. For example, this may occur during particular manoeuvres of the aircraft. As a result, the rotor torque calculated by the known estimation methods fails to give an adequate indication of the value of the torque acting on each of the two rotors, when the assumption that the rotor torques acting on the two rotors are equal or substantially equal to each other is not valid. US-A1-2017336809 discloses a method for executing yaw control of an aircraft including two rotors. The method includes inducing helicopter yaw by creating a differential torque between the two rotors, wherein the creating of the differential torque comprises inducing a differential collective pitch to generate a differential thrust, and maintaining helicopter roll equilibrium during the inducing of the helicopter yaw by inducing a differential cyclic pitch to generate a differential lift offset. US-A1-2016122039 discloses a method for calculating torque through a rotor mast of a propulsion system of a tiltrotor aircraft including receiving the torque applied through a quill shaft of the rotorcraft. The quill shaft is located between a fixed gearbox and a spindle gearbox, and the spindle gearbox is rotatable about a conversion access. The torque through the rotor mast is determined by using the torque through the quill shaft and the efficiency loss value between the quill shaft and the rotor mast. In addition, a control method for turboshaft engine is disclosed in the paper “A novel control method for turboshaft engine with variable rotor speed based on the Ngdot estimator through LQG/LTR and rotor predicted torque feedforward”