Leonardo

A method for assisting the implementation of a real or simulated manoeuvre for a helicopter (1) comprising an autopilot (25) is described, comprising the steps of commanding an autopilot (25) to control the first, second and third flight control (15, 16, 17) so as to execute a descent and landing manoeuvre in spot hover conditions (LDG); comparing the first signal (S7) with a first target value (Vz_set) of descent speed (Vz) and operating the first flight control (15) so as to maintain the speed (Vz) equal to the first target value (Vz_set) during the landing manoeuvre in spot hover conditions (LDG); comparing the second signal (S14) with a second target value (γ_set) of the yaw angle (γ) and operating the third flight control (17) so as to maintain the yaw angle (γ) equal to the second target value (γ_set); and comparing the third signal (S15, S16) with a target landing position (Px_set, Py_set) and operating the second flight control (16) so as to set the helicopter (1) to the target landing position (Px_set, Py_set) during the descent and landing manoeuvre (LDG).

Tool (14) wear status monitoring software (1) being loadable in, and executable by, electronic processing resources. The tool (14) wear status monitoring software (1) being designed to cause, when executed, said electronic processing resources to become configured to receive, from one or more sensors (11) arranged within a working environment (10), one or more acoustic signals indicative of acoustic quantities of a sound caused by the operation of a tool (14) within said working environment (10). In addition, the tool (14) wear status monitoring software (1) is designed to cause, when executed, said electronic processing resources to become configured to compute (block 4) a graphical representation (6), in particular a spectrogram, of the acoustic quantities of the sound caused by the operation of said tool (14) based on the received acoustic signals indicative of said acoustic quantities. The tool (14) wear status monitoring software (1) is further designed to cause, when executed, the electronic processing resources to become configured to determine a wear status (9) of the tool (14) by running a predictive neural model (7) on the computed graphical representation (6); the predictive neural model (7) being trained (block 8) to predict a wear status (9) to be associated to a graphical representation (6) of acoustic quantities of a sound. Furthermore, the tool (14) wear status monitoring software (1) is designed to cause, when executed, the electronic processing resources to become configured to provide an output based on the determined wear status (9) of the tool (14).

A method is described for estimating the pitch angle (γ) of the blades (25) of a rotor (10, 10') of an aircraft (1) capable of hovering comprising a motor member (16) with a stator (14) and an output shaft (18) rotatable about the stator (14); a hub (19) rotatable about the axis (B) and operatively connected with the output shaft (18) so as to be rotatable with an angular velocity (ω) about the axis (B); at least two blades (25) rotatable about the axis (B) and about corresponding second axes (D) of extension of the blades (25), so as to adjust the pitch angles (γ) of the blades (25) between a first value (α) and a second value (β) different from each other; the method comprises the steps of: i) computing a third value associated with the torque (τ_EM) generated by the motor (16); ii) computing a fourth value associated with the torque (τ_tr) required, in use, by the rotor (10, 10'); iii) computing a fifth value associated with the thrust coefficient (C_P) and a sixth value associated with the thrust coefficient (C_T) of the rotor (10, 10'), also based on the fourth value; iv) comparing the fifth and sixth values with seventh and eighth threshold values (Th_P, Th_T); and v) estimating a ninth value of the pitch angle (γ), based on the result of step iv). (Figure 2)

A system (20) for detecting anomalies of a transmission assembly (8) of an aircraft (1). The aircraft (1) comprises a lubrication system (10), which comprises a tank (11) containing a lubricating fluid, a nozzle (12) configured to supply the lubricating fluid to the transmission assembly (8) and a fluid line (13), which fluidly connects the tank (11) to the nozzle (12). The detection system (20) comprises a first device (22), which detects a first quantity (n) associated with particles transported by the lubricating fluid along the fluid line (13), a second device (23), which detects a second quantity (ai) associated with vibrations of the transmission assembly (8) and an electronic control unit (24). The electronic control unit (24) verifies whether one between the first quantity (n) and the second quantity (ai) satisfies a first control criterion, whether the other between the first quantity (n) and the second quantity (ai) satisfies a second control criterion when the first control criterion is not satisfied and determines the presence of a possible anomaly of the transmission assembly (8) and/or provides for corrective actions when neither control criteria are satisfied.

A rolling bearing (20) for a helicopter (1) is described, comprising a first ring (21) extending around an axis (A) and a second ring (22) extending around the axis (A) and radially external to the first ring (20). The rolling bearing (20) further comprises a plurality of rolling bodies (23, 24) cooperating with said first and second rings (21, 22), so as to roll on respective raceways (25, 26, 27, 28) of said first and second rings (21, 22) and to allow the relative rotation between the first and the second ring (21, 22) around the axis (A); and a cage (29) configured to maintain the rolling bodies (23, 24) angularly distanced from each other around the axis (A). The cage (29) comprises a cavity (30) configured to be selectively engaged by a sensor (50); and facing the first ring (21) and/or the second ring (22).

An aircraft (1, 1', 1") is described with a fuselage (2) extending along a first longitudinal axis (Y) of the aircraft (1, 1', 1"); a pair of half-wings (3, 8) projecting sideways, cantilevered, from the fuselage (2); a pair of first rotors (20a, 20b; 21a, 21b; 22a, 22b) that revolve about a second axis (B, C; D, E; F, G) and can be tilted independently of each other about a third axis (H; L; J); a pair of first motors (30a, 30b; 31a, 31b; 32a, 32b) that can be operated to rotate first rotors (20a, 20b; 21a, 21b; 22a, 22b) independently of each other; the aircraft (1, 1', 1") can be moved between: a first forward flight configuration and a second take-off/landing configuration in which the second axes (B, C; D, E; F, G) are arranged transverse to the first axis (Y); first wing areas (60a, 60b; 61a; 62a, 62b; 70a, 70b; 72a, 72b) operationally connected to respective first rotors (20a, 20b; 21a; 21b; 22a, 22b); the first wing areas (60a, 60b; 61a, 61b; 62a, 62b; 70a, 70b; 72a, 72b) being tiltable with respect to the third axis (H; L; J) as a result of the inclination of the respective first rotors (20a, 20b; 21a; 21b; 22a, 22b) about the third axis (H; L; J) independently of each other.

A half-wing (3, 3', 3", 3‴; 8; 9) for a convertible aircraft (1) capable of hovering is described, comprising: a first portion (50) that can be attached to a fuselage (2) of the aircraft (1); a second portion (51) that can be tilted with respect to the first portion (50) around a first axis (H; I; J); a first rotor (20a, 20b; 21a, 21b; 22a, 22b) integral with the second portion (51) and revolving about a second axis (B, C; D, E; F, G) transverse to the first axis (H; L; J); the second portion (51) being tiltable with respect to the first portion (50) by a first angle (α) between a first and a second position associated, respectively, with a first forward flight configuration and a second take-off/landing configuration of the aircraft (1); the half-wing (3, 3', 3", 3‴; 8, 9) comprises a wing area (60) defining a free wing tip (52) of the half-wing (3) and operationally connected to the second portion (51); each first rotor (20a, 20b; 21a, 21b; 22a, 22b) is interposed between the first wing area (60) and the first portion (50) along the first axis (H; L; J).

There is disclosed a method of controlling a helicopter (1) comprising: at least one thermal engine (11); at least one rotor (3), which is operatively connectable to thermal engine (11) to be drive, in use, at a set rotational speed (ω<SET>); an emergency system (40) comprising: at least one electric motor (41), which is operatively connectable to rotor (3) and an electrical power source (42), which is electrically connected to electric motor (41); the method comprises the steps of i) operatively connecting thermal engine (11) to rotor (3) and keeping electric motor (41) disconnected by rotor (3) or idle, in a normal operating condition of thermal engine (11); ii) detecting if thermal engine (11) is in at least partial failure condition; iii) accelerating electric motor (41) up to when said current rotational speed (ω) equals set value (ω<SET>), in case of at least partial failure of said thermal engine (11); step iii) comprises the steps of iv) sensing a plurality of parameters (ω, ω<ERR>, ϕ, ϕ<ERR>, IIT, τ<τ/s>, τ<MGB>, Power<τ/s>, Power<MGB>) of thermal engine (11) and rotor (3); and v) identifying an at least partial failure of thermal engine (11), on the basis of parameters (ω, ω<ERR>, ϕ, ϕ<ERR>, IIT, τ, P. (Figure 1)

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