IPC분류정보
국가/구분 |
United States(US) Patent
등록
|
국제특허분류(IPC7판) |
|
출원번호 |
UP-0573524
(2005-07-27)
|
등록번호 |
US-7702429
(2010-05-20)
|
우선권정보 |
FR-04 08862(2004-08-13) |
국제출원번호 |
PCT/FR2005/001950
(2005-07-27)
|
§371/§102 date |
20070209
(20070209)
|
국제공개번호 |
WO06/024745
(2006-03-09)
|
발명자
/ 주소 |
- Lavergne, Fabien
- Villaume, Fabrice
|
출원인 / 주소 |
|
대리인 / 주소 |
|
인용정보 |
피인용 횟수 :
2 인용 특허 :
7 |
초록
▼
The invention relates to an electric flight control system for aircraft elevators. According to the invention, the flight control system can be controlled in terms of load factor or rate of pitch. The inventive system comprises built-in protections in relation to load factor, incidence and pitch att
The invention relates to an electric flight control system for aircraft elevators. According to the invention, the flight control system can be controlled in terms of load factor or rate of pitch. The inventive system comprises built-in protections in relation to load factor, incidence and pitch attitude.
대표청구항
▼
The invention claimed is: 1. An electric flight control system for the control in terms of load factor of the elevators of an aircraft, said elevators being controlled by a control section that compels said elevators to take a deflection position dependent on an electrical signal δmc represen
The invention claimed is: 1. An electric flight control system for the control in terms of load factor of the elevators of an aircraft, said elevators being controlled by a control section that compels said elevators to take a deflection position dependent on an electrical signal δmc representative of a controlled value of the angle of deflection δm of said elevators, said system comprising: a first calculation section for calculating, on the basis of an electrical signal nzc representative of a controlled value of said load factor, a first electrical signal {dot over (γ)}c representative of a controlled value of the derivative, with respect to time, of the aerodynamic slope γ of said aircraft; a first constituent device, which: is configured to receive at its input said first electrical signal {dot over (γ)}c; comprises a first protection section configured to maintain said first electrical signal {dot over (γ)}c between a first minimum value and a first maximum value; on the basis of said first electrical signal {dot over (γ)}c, determines at least a second electrical signal αc, representative of a corresponding controlled value of the incidence α, and a third electrical signal θc, representative of a corresponding controlled value of the longitudinal attitude θ; comprises a second protection section configured to maintain said second electrical signal αc between a second minimum value and a second maximum value; and delivers at least said third electrical signal θc to its output; and a second constituent device, which is configured to receive at its input at least said third electrical signal θc or a fourth electrical signal θd similar to said third electrical signal θc; comprises a third protection section configured to maintain said third or fourth electrical signal between a third minimum value and a third maximum value; and is configured to deliver at its output a fifth electrical signal which constitutes said signal δmc, representative of the corresponding controlled value of the angle of deflection δm of said elevators. 2. The system as claimed in claim 1, further comprising a first switching section configured to take: either a first position for which the output of said first constituent device is connected to the input of said second constituent device, so that said third electrical signal θc is then transmitted to said second constituent device; or a second position for which the input of said second device receives said fourth electrical signal θd, similar to said third electrical signal θc and produced on the basis of a sixth electrical signal qd, representative of a desired value for a rate of pitch q. 3. The system as claimed in claim 1, wherein said first constituent device determines, in addition to said second electrical signal αc and said third electrical signal θc, a seventh electrical signal qc, representative of a corresponding controlled value of the rate of pitch q, and a first switching section configured to transmit said seventh electrical signal qc to said second constituent device. 4. The system as claimed in claim 3, wherein said first constituent device delivers, for the seventh electrical signal qc, an approximate value equal to that of said first electrical signal {dot over (γ)}c. 5. The system as claimed in claim 2, further comprising: an automatic pilot configured to deliver a first controlled load factor signal nzc; a manual piloting member configured to deliver, by switching, either a second controlled load factor signal nzc or said sixth electrical signal qd, representative of a desired value for the rate of pitch q; and a second switching section for: transmitting to said first constituent device either the first controlled load factor signal delivered by said automatic pilot, or the second controlled load factor signal delivered by said manual piloting member; or else transmitting said sixth electrical signal qd to a first integration section configured to form the fourth electrical signal θd, representative of a desired value for an attitude θ, said first switching section being configured to transmit to said second constituent device said fourth and sixth electrical signals θd and qd, instead of said third and seventh electrical signals θc and qc produced by said first constituent device. 6. The system as claimed in claim 1, wherein, to determine said second electrical signal αc on the basis of the first signal {dot over (γ)}c, said first constituent device comprises a second calculation section for calculating the expression αc=({dot over (γ)}c−F65 )/Gy in which Fγ and Gγ are functions of the state of the aircraft with F γ = g · cos ( γ ) V + 1 2 ρ m · V · S · Cz a = 0 and G γ = 1 2 · ρ m · V · S · ∂ Cz ∂ α | α = 0 + T m · V where g is the acceleration due to gravity, γ the aerodynamic slope, V the speed of the aircraft, ρ the density of the air, m the mass of the aircraft, S the reference area of the aircraft, Czα=0 the coefficient of lift of the aircraft for a zero incidence, ∂ Cz ∂ α ❘ α = 0 the gradient of the aerodynamic coefficient of lift as a function of the incidence and T the thrust of the aircraft. 7. The system as claimed in claim 3, wherein said first constituent device comprises a second integrator section configured to integrate said seventh electrical signal qc and a first summator for forming the sum of the integral delivered by said second integrator section and of said second electrical signal αc, so as to form said third electrical signal θc. 8. The system as claimed in claim 5, wherein said second constituent device, either on the basis of said third electrical signal θc and of the seventh electrical signal qc, originating from said first constituent device, or on the basis of said fourth signal θd and of said sixth signal qd originating from said manual piloting member, well as current values qr and θr of the rate of pitch q and of the longitudinal attitude θ, determines an eighth electrical signal {dot over (q)}c, representative of a corresponding controlled value of the pitch acceleration {dot over (q)}, then, on the basis of this eight electrical signal {dot over (q)}c said second constituent device determines said fifth electrical signal δmc. 9. The system as claimed in claim 8, wherein said second constituent device calculates said eight electrical signal {dot over (q)}c, through the relation {dot over (q)}c=K1.θv−K2.θr+K3.qv−K4.qr where θv said third or fourth electrical signal, θr the current value of the longitudinal attitude θ, qv said sixth or seventh electrical signal, qr the current value of the rate of pitch q, K1, K2, K3 and K4 being constant coefficients. 10. The system as claimed in claim 8, wherein, to determine said fifth electrical signal δmc on the basis of said eighth electrical signal {dot over (q)}c, said second constituent device comprises a third calculation section that calculates the expression δmc=({dot over (q)}c−Fq)/Gq in which Fq and Gq are functions of the state of the aircraft with F q = 1 I y · 1 2 · ρ · v 2 · S · l · Cm δ m = 0 + 1 I γ · T · b · cos ( τ ) and G q = 1 I y · 1 2 · ρ · v 2 · S · l · ∂ Cm ∂ δ m | δ m = 0 where Iy is the pitch inertia, ρ the density of the air, V the speed of the aircraft, S the reference area of the aircraft, l the reference length of the aircraft, Cmδm=0 the coefficient of pitch, T the thrust, b the lever arm of the engines, τ the angle of longitudinal trim of the engines and ∂ Cm ∂ δ m ❘ δ m = 0 the effectiveness of the elevators.
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