Method for controlling an aircraft capable of hovering and relative aircraft

Patent metadata

Field	Value
Publication number	US-2023322400-A1
Application number	US-202118018961-A
Country	US
Kind code	A1
Filing date	Jul 30, 2021
Priority date	Aug 4, 2020
Publication date	Oct 12, 2023
Grant date	—

Abstract

A method for controlling an aircraft capable of hovering is described, comprising a first engine; a second engine; at least one rotor; and a transmission interposed between the first and second engine and the rotor; the transmission comprises a first and a second inlet connected respectively to a first outlet member of the first engine and to a second outlet member of the second engine; the method comprises step i) of placing the in a first configuration, in which the first and second engine make available a first and a second power value; or in a second configuration, in which the first engine (makes available a third power value greater than the first power value to the first inlet, and the second engine delivers a nil power value to the second inlet; the method also comprises, characterised in that it comprises the steps of ii) detecting a series of parameters associated with the operating conditions of the aircraft; and iii) enabling the transition of the aircraft from the first configuration to the second configuration, when the parameters assume respective first values.

First claim

1 . A method for controlling an aircraft ( 1 ) capable of hovering; said aircraft ( 1 ) comprising:
a first engine ( 10 a ); a second engine ( 10 b ); at least one rotor ( 3 ) operatively connected to said first engine and to said second engine ( 10 a , 10 b ); and a transmission ( 8 ) interposed between said first and second engine ( 10 a , 10 b ) and said rotor ( 3 ); said transmission ( 8 ) comprising a first and a second inlet ( 12 a , 12 b ) connected respectively to a first outlet member ( 11 a ) of said first engine ( 10 a ) and to a second outlet member ( 11 b ) of said second engine ( 10 b ); said method comprising step i) of placing said aircraft ( 1 ): in a first configuration, in which said first and second engine ( 10 a , 10 b ) make available respectively a first and a second power value (P1, P2) substantially equal to each other to the respective first and second inlet ( 12 a , 12 b ) of said transmission ( 8 ); the sum of said first and second power value (P1, P2) being equivalent to a power (P) necessary for the correct operation of at least said one rotor ( 3 ); or in a second configuration, in which said first engine ( 10 a ) makes available a third power value (P3) greater than said first power value (P1) to said first inlet ( 12 a ), and said second engine ( 10 b ) delivers a nil power value (P4) to said second inlet ( 12 b ); said third power value (P3) and said fourth power value (P4) being different from one another; said third power value (P3) being greater than said first power value (P1) and equal to said power (P); said method comprising the steps of: ii) detecting a series of parameters associated with the operating conditions of said aircraft ( 1 ); and iii) enabling the transition of said aircraft ( 1 ) from said first configuration to said second configuration, when said parameters assume respective first values; characterized in that said step iii) comprises step xiii) of checking the following conditions: the external temperature (OAT) is comprised between a first operating value and a second operating value; the pressure altitude is comprised between a third operating value and a fourth operating value: the density altitude is comprised between a fifth operating value and a sixth operating value: the height above ground level is comprised between 50 and 1000 feet; the height above sea level is comprised between 50 and 1000 feet; the angular speed of said rotor (3) is comprised between a seventh minimum operating value and an eighth maximum operating value, where the seventh value is equal to between 0.85 and 1 times the nominal angular speed and the eighth value is comprised between 1 and 1.15 times the aforesaid nominal angular speed; the indicated speed (IAS) is comprised between a ninth value and a tenth value, where the ninth value is comprised between 0.5 and 1.5 times the speed in level flight for which the minimum power at level and operating temperature is required and the tenth value is comprised between 1 and 2.5 times the aforesaid speed in level flight for which the minimum power is required; and the torque delivered by said first and second engine ( 10 a , 10 b ) is comprised between an eleventh operating value and a twelfth operating value, where the eleventh value is comprised between 0.5 and 1 times the torque necessary for a straight level flight at the aforesaid speed in level flight for which the minimum power is required and the twelfth value is comprised between 1 and 3 times the aforesaid necessary torque; said step iii) further comprising step xiv) of verifying that: the altitude above the minimum flight level is comprised between 50 and 1000 feet; and the variometric speed (Vz) is comprised between a thirteenth operating value and a fourteenth operating value, where the thirteenth operating value is comprised between 0.1 and 1 times the vertical speed of autorotation and the fourteenth value is comprised between 0.1 and 1 times the maximum variometric speed at the aforesaid speed for which minimum power is required; said method comprising the further steps of: performing a commanded nominal exit of said aircraft ( 1 ) from said second configuration to said first configuration; performing a commanded emergency exit of said aircraft from said second configuration to said first configuration; performing an automatic emergency transition of said aircraft ( 1 ) from said second configuration to said first configuration, when at least one of said parameters assumes the respective second value different from said first value; performing an automatic non-emergency transition of said aircraft ( 1 ) from said second configuration to said first configuration in case of minor anomalies of said aircraft ( 1 ) that do not require an automatic emergency exit; wherein said commanded non-emergency transition and said automatic non-emergency transition take place in longer times than said automatic emergency transition.

2 . The method according to claim 1 , characterized in that at least some of said parameters define a flight envelope of said aircraft ( 1 ); and/or at least further said parameters are indicative of a failure of said aircraft ( 1 ).

3 . The method according to claim 1 , characterized in that it comprises step iv) of preventing the transition of said aircraft ( 1 ) from said first configuration to said second configuration, when at least one of said parameters assumes a respective second value different from said first value; and/or
characterized in that it comprises step vi) of performing producing a commanded emergency transition of said aircraft ( 1 ) from said second configuration to said first configuration, when said parameters assume respective first values and following a first command.

4 . The method according to claim 1 , characterized in that it comprises the steps of:
vii) producing a trend of at least some of said parameters, when said aircraft ( 1 ) is in said second configuration; and viii) generating a warning signal if said trend leads said at least some parameters to assume second values different from said first values.

5 . The method according to any one of the preceding claims claim 1 , characterized in that it comprises the steps:
ix) detecting that said first engine ( 10 a ) starts failing; x) reducing the level and the vertical speed of said aircraft ( 1 ); and xi) increasing the power delivered by said second engine ( 10 b ).

6 - 7 . (Canceled)

8 . The method according to claim 1 , characterized in that it comprises step xv) of controlling said second engine ( 10 b ) so that the relative second outlet member ( 12 ) rotates at an angular speed greater than zero in said second configuration.

9 . The method according to claim 1 , characterized in that it comprises step xvi) producing a relative angular sliding between said second outlet member ( 11 b ) and said second inlet ( 12 b ) by means of a one-way clutch ( 13 ) when said aircraft ( 1 ) is in said second configuration and in case of failure of said second engine ( 10 b );
said one-way clutch ( 13 ) comprising a first element ( 16 ) rotatable integrally with said relative second inlet ( 12 b ) and a second element ( 15 ) rotatable integrally with said second outlet shaft ( 11 b ); said clutch ( 13 ) rotationally decoupling said second inlet ( 12 b ) and second outlet shaft ( 11 b ), when said first element ( 16 ) rotates at a lower speed than said second element ( 15 ), and rotationally coupling said second inlet ( 12 b ) and second outlet shaft ( 11 b ) when said first and second element ( 16 , 15 ) rotate at the same angular speed.

10 . The method according to claim 9 , characterized i

Assignees

• Leonardo Spa

Inventors

• Medici Luca
• Nannoni Fabio
• Braghiroli Marco
• Cogliati Andrea
• Frattini Tiziano
• Masellis Lorenzo
• Riviello Luca
• Scola Alessandro
• Tamborini Marco
• Pizzagalli Matteo

Classifications

• B64D35/08 Primary

  characterised by the transmission being driven by a plurality of power plants (for hybrid-electric power plants B64D35/022) · CPC title

• G05D1/854

  in response to motor or actuator failures · CPC title

• G05D1/0072

  to counteract a motor failure · CPC title

• B64D31/10 Primary

  for preventing asymmetric thrust · CPC title

• F02C7/36 Primary

  Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user ({F02C3/107 - F02C3/13 and} F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H) · CPC title