System and a method for controlling pitching stabilizer means of an aircraft
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
B64C-013/00
B64C-013/50
F16H-003/72
B64C-013/42
B64C-027/82
출원번호
US-0267158
(2014-05-01)
등록번호
US-9643717
(2017-05-09)
우선권정보
FR-13 01034 (2013-05-03)
발명자
/ 주소
Grohmann, Boris
Buro, Thomas
출원인 / 주소
Airbus Helicopters
대리인 / 주소
Brooks Kushman P.C.
인용정보
피인용 횟수 :
1인용 특허 :
7
초록▼
A control system (20) for controlling pitching stabilizer means of an aircraft, said system (20) being provided with at least one outlet shaft (21) and with a first actuator (31) and a second actuator (36). The first and second actuators (31 and 36) are different, and the first actuator (31) is a sl
A control system (20) for controlling pitching stabilizer means of an aircraft, said system (20) being provided with at least one outlet shaft (21) and with a first actuator (31) and a second actuator (36). The first and second actuators (31 and 36) are different, and the first actuator (31) is a slow actuator having a first driving portion that moves at a first speed, the second actuator (36) being a fast actuator having a second driving portion that moves at a second speed faster than the first speed, said control system (20) comprising a control device (50) connected to the first actuator (31) and to the second actuator (36) in order to cause said outlet shaft (21) to be driven either by the first driving portion (32) and/or by the second driving portion (37).
대표청구항▼
1. A control system for controlling pitching stabilizer means of an aircraft, the system being provided with at least one outlet shaft for turning the stabilizer means, the system having a first actuator and a second actuator, the first and second actuators being different, wherein the first actuato
1. A control system for controlling pitching stabilizer means of an aircraft, the system being provided with at least one outlet shaft for turning the stabilizer means, the system having a first actuator and a second actuator, the first and second actuators being different, wherein the first actuator has a first driving portion that moves at a first speed, the second actuator has a second driving portion that moves at a second speed greater than the first speed, the control system comprising a control device connected to the first actuator and to the second actuator in order to cause the outlet shaft to be driven either by the first driving portion or by the second driving portion, or else by both the first driving portion and the second driving portion, the control system including a memory having a first relationship for electrically controlling the first actuator while limiting the authority of the first actuator for turning the outlet shaft over a first angular range, and a secondary relationship for electrically controlling the second actuator while giving the second actuator full authority for turning the outlet shaft over a second angular range, the first angular range being less than the second angular range, wherein the control system includes a differential device for connecting the first actuator and the second actuator to the outlet shaft, wherein the differential device comprises: a ring gear mechanically engaged with the first driving portion,a sun gear mechanically engaged with the second driving portion, anda planet carrier carrying a plurality of planet gears, each interposed between the sun gear and the ring gear, the planet carrier meshing with the outlet shaft. 2. A system according to claim 1, wherein the memory includes a third relationship for controlling the first actuator while giving full authority to the first actuator to turn the outlet shaft over the second angular range in the event of the second actuator failing. 3. A system according to claim 2, wherein the system includes an alert device for signaling failure of the second actuator to a pilot in order to enable the pilot to adapt piloting to the first speed of the first actuator. 4. A system according to claim 1, wherein the memory includes a fourth relationship for use when the first actuator has failed in order to control the second actuator while giving the second actuator full authority over turning the outlet shaft over the second angular range in order to reach a predetermined position of the stabilizer means depending on the stage of flight, and in order to adjust the position of the stabilizer means dynamically about the predetermined position. 5. A system according to claim 1, wherein the first actuator develops a first power output and the second actuator develops a second power output, the second power output being greater than the first power output. 6. A stabilizer assembly having pitching stabilizer means comprising at least one stabilizer surface, wherein the assembly includes a control system according to claim 1 having an outlet shaft connected to each stabilizer surface in order to turn each stabilizer surface about an axis of rotation. 7. An aircraft having a rotary wing, wherein the aircraft includes a stabilizer assembly according to claim 6. 8. A method of stabilizing an aircraft according to claim 7, the method comprising the steps of quickly positioning the stabilizer means in a predetermined position depending on the stage of flight by using the second actuator, and operating the stabilizer means slowly for stabilizing the aircraft by dynamically adjusting the position of the stabilizer means about the predetermined position by using the first actuator. 9. A method according to claim 8, wherein the first speed of the first actuator is designed to cause the stabilizer means to move at an angular speed lying in the range 0.1°/s to 2°/s, and the second speed of the second actuator is designed to cause the stabilizer means to move at an angular speed lying in the range 7°/s to 14°/s. 10. A method according to claim 8, wherein in the event of a failure of the second actuator, the stabilizer means is positioned in the predetermined position by using the first actuator, and an alert is generated. 11. A method according to claim 8, wherein the stabilizer means are positioned: in a first predetermined position when the aircraft is flying in a zone that is subjected to a ground effect;in a second predetermined position when the aircraft is not present in the zone and when the aircraft possesses a forward speed below a threshold; andin a third predetermined position when the aircraft is not present in the zone and when the aircraft possesses a forward speed greater than the threshold. 12. An aircraft comprising: a stabilizer assembly having a pitch stabilizing surface connected to an outlet shaft, wherein rotation of the outlet shaft about an axis of rotation turns the pitch stabilization surface;a first actuator having a first driving portion to move at a first speed, the first driving portion drivably connected to the outlet shaft;a second actuator having a second driving portion to move at a second speed, the second speed being greater than the first speed, the second driving portion drivably connected to the outlet shaft;a control device connected to the first and second actuators for controlling the first and second actuators to drive the outlet shaft, the control device (i) electrically controlling the first actuator to drive the outlet shaft over a first angular range by limiting the authority of the first actuator, and (ii) electrically controlling the second actuator to drive the outlet shaft over a second angular range while giving the second actuator full authority for turning the outlet shaft, wherein the first angular range is less than the second angular range; anda differential device connecting the first and second actuators to the outlet shaft, the differential device having a ring gear, a sun gear, and a planet carrier;wherein the ring gear is meshed with the first driving portion of the first actuator, and the ring gear is coaxial with the outlet shaft;wherein the sun gear is meshed with the second driving portion of the second actuator, and the sun gear is coaxial with the outlet shaft; andwherein the planet carrier is constrained to rotate with the outlet shaft, the planet carrier having a plurality of planet gears, each planet gear positioned between and meshed with the sun gear and the ring gear. 13. The aircraft of claim 12 wherein each planet gear rotates about an axis that is parallel with the axis of rotation of the outlet shaft. 14. A method of stabilizing an aircraft comprising: electrically controlling a first actuator to drive an outlet shaft over a first angular range while limiting the authority of the first actuator to operate a stabilizer assembly slowly to stabilize the aircraft by dynamically adjusting a position of a pitch stabilizer surface about a predetermined position, the first actuator being controlled to move a first driving portion of the first actuator at a first speed to drive the outlet shaft, the outlet shaft being connected to the stabilizer assembly having the pitch stabilizer surface such that rotation of the outlet shaft about an axis of rotation turns the pitch stabilizer surface;electrically controlling a second actuator to drive the outlet shaft over a second angular range while giving full authority to the second actuator to quickly position the pitch stabilizer surface in the predetermined position based on a stage of flight of the aircraft, the second actuator being controlled to move a second driving portion of the second actuator at a second speed to drive the outlet shaft;in response to a failure event of the second actuator, electrically controlling the first actuator to drive the outlet shaft over the second angular range while giving full authority to the first actuator to operate the stabilizer assembly to position the pitch stabilizer surface in the predetermined position; andin response to a failure event of the first actuator, electrically controlling the second actuator to drive the outlet shaft over the second angular range while giving full authority to the second actuator to operate the stabilizer assembly to dynamically adjust the position of the itch stabilizer surface about the redetermined position;wherein the first speed of the first driving portion is less than the second speed of the second driving portion; andwherein the first angular range is less than the second angular range. 15. The method of claim 14 wherein, in response to the aircraft being in a first stage of flight defined by the aircraft being between the ground and a height defined by a diameter of a main lift rotor of the aircraft, the second actuator is electrically controlled to move the pitch stabilizer surface to a first predetermined position, and upon the pitch stabilizer surface reaching the first predetermined position, electrically controlling the first actuator to move the pitch stabilizer surface to maintain a constant longitudinal attitude of the aircraft; wherein, in response to the aircraft being in a second stage of flight defined by the aircraft being above the height and having a forward speed less than a threshold, the second actuator is electrically controlled to move the pitch stabilizer surface to a second predetermined position, and upon the pitch stabilizer surface reaching the second predetermined position, electrically controlling the first actuator to move the pitch stabilizer surface to maintain the constant longitudinal attitude of the aircraft; andwherein, in response to the aircraft being in a third stage of flight defined by the aircraft being above the height and having a forward speed greater than the threshold, the second actuator is electrically controlled to move the pitch stabilizer surface to a third predetermined position, and upon the pitch stabilizer surface reaching the third predetermined position, electrically controlling the first actuator to move the pitch stabilizer surface to maintain the constant longitudinal attitude of the aircraft. 16. The method of claim 15 wherein, in response to the aircraft being in an initial stage on the ground, the second actuator is electrically controlled to move the pitch stabilizer surface to an initial position corresponding to a zero angle relative to horizontal; wherein the first predetermined position of the pitch stabilizer surface lies between the initial position and the second predetermined position;wherein the third predetermined position of the pitch stabilizer surface corresponds to the initial position; andwherein the threshold is set such that the aircraft is hovering in the second stage of flight and is flying forward in the third stage of flight.
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이 특허에 인용된 특허 (7)
Kubica, Fran.cedilla.ois, Aircraft with electrical fly-by-wire controls, equipped with an automatic pilot.
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