KR100298260B1 - Method for controlling angle of attack of airplane - Google Patents

Abstract

PURPOSE: A method for controlling the angle of attack of an airplane is provided to offer a relatively big change to the displacement angle of an elevator when a present angle of attack is over a critical angle of attack. CONSTITUTION: A stall angle of attack is determined. A critical angle of attack less than the stall angle of attack and a reference angle of attack less than the critical angle of attach are obtained. The displacement angle of an elevator is controlled depending on a first controlling value in proportion to a pitch angular velocity and a vertical overload of an airplane(512) and in inverse proportion to a present displacement and a present angle of attack. If the present angle of attack is over the critical angle of attack, the first control value is adjusted to be subtracted by a first value. If the present angle of attack is larger than the reference angle of attack and less than the critical angle of attack, the first control value is adjusted to be subtracted by a second value in proportion to the present angle of attack.

Description

How to control the angle of attack of the aircraft

The present invention relates to a method of controlling an attack angle of an aircraft.

Aircraft in flight must maintain proper wing lift at all times. If you lose momentary lift, the aircraft will stall and risk falling. Since the lift force changes according to the angle of attack, the angle of attack must be maintained within a certain range to prevent stall.

1 is a view illustrating a control angle of attack of a conventional aircraft. Referring to FIG. 1, the displacement angle δ _e of the elevator is controlled in a state in which the current angle of attack of the aircraft 108 is not fed back. In the addition and subtraction unit 112, the product (Kq · q) of the pitch angular velocity q of the aircraft 108 and the feedback gain Kq 110 is equal to the vertical overload Δnz and the gain Knz 111. It is added up with the product (Knz · Δnz). The addition and subtraction unit 104 multiplies the current displacement (δ _cc ) 101 of the steering column and the conversion gain (Kccn) 103 from this summed value (Kq · q + Knz · Δnz) (Kccn · δ _cc). Subtract). That is, the output C ^* of the addition subtraction unit 104 is obtained according to Equation 1 below.

[Equation 1]

C ^* = Kq q q Knz Δnz-Kccn δ _cc

The output C ^* of the additive subtraction unit 104 is multiplied by the feedback proportional gain K _FB 105 and the proportional integral gain 1 1 / S 106. That is, the proportional integral output is C ^* _KFB (1 + 1 / S). Where S represents the Laplace operator. In the addition / subtraction unit 107, the product of the pitch angular velocity q and the damper gain K _damp 109 (K _damp q) and the proportional integral output C ^* K _FB (1 + 1 / S) are obtained. after summing, it subtracts the product (K _{cc FF} · δ) of the current displacement of the steering column (δ _cc) (101) and the feed-forward (feed-forward) gain (K _FF) (102). That is, the displacement angle δ _e of the elevating rudder output from the add / subtract unit 107 is obtained according to the following expression (2).

[Equation 2]

^{_{δ e = C * · K FB}} · (1 + 1 / S) + K damp · q - K FF · δ cc

In the angle of attack control algorithm as described above, since the current angle of attack α of the aircraft 108 is not fed back, the accuracy of the control may be reduced.

FIG. 2 shows a load feel system added to the control method of FIG. 1. FIG. 3 shows the relationship of the force Fcc applied according to the displacement δ _cc of the steering column in the system of FIG. 2. 2 and 3, the spring 23 provided between the lower part of the steering column 21 and the support 24 is the opposite direction of travel of the steering column 21 with a force Fcc proportional to the current angle of attack. It is designed to pull the steering column 21. When the pilot pushes the handle of the steering column 21, the steering column 21 rotates about the rotation center 22. At this time, the force applied to the lower portion of the steering column 21 to move to the maximum movement distance (δ _ccmax ) from the origin, is applied to the displacement (δ _cc ) of the steering column at a predetermined slope in the range of F _cc1 to F _cc2 . Proportional. Accordingly, the pilot can feel whether the current angle of attack reaches the boundary angle of attack smaller than the stall angle by the weighted force.

In the rod fill system shown in Figs. 2 and 3, the force exerted on the steering column 21 is simply proportional to the current angle of attack, so that the pilot may not be able to accurately sense when the force is at the boundary angle of attack. Accordingly, inexperienced pilots may not take any action until the aircraft reaches stall angle.

4 is a view illustrating an angle of attack of another conventional aircraft. Referring to FIG. 4, the displacement angle δ _e of the elevator is controlled while the current angle of attack α of the aircraft 412 is fed back.

In the addition / subtraction unit 416, the product (Kqq) of the pitch angular velocity q of the aircraft 412 and the feedback gain Kq 414 is equal to the vertical overload Δnz and the gain Knz 415. It is added up with the product (Knz · Δnz). Acceleration peaks 404, the product of the sum current displacement of the steering column in (Kq · q + Knz · △ nz) (δ cc) (401) and the conversion gain (Kccn) (403) (Kccn · δ _cc Subtract). That is, the output C ^* of the addition / subtraction unit 404 is obtained according to the following equation (3).

[Equation 3]

C _* = Kq q q Knz Δnz-Kccn δ _cc

In the angle of attack determining unit 419, the boundary angle of attack α _ALT smaller than the stall angle of attack α α _STL and the boundary angle of attack α _ALT according to the velocity M and the flap angle δ _FLAP of the aircraft 412. Determine a reference angle of attack (α ₀ ) that is less than. The switching conversion unit 417 obtains a conversion value K _TR for the current angle of attack α. The conversion value K _TR is 1 when α ≤ α _{0, and} when α ₀ <α <α _ALT 0 is obtained when α _ALT ≤ α. In the multiplication unit 405, the output C ^* of the addition / subtraction unit 404 is multiplied by the converted value K _TR . That is, the output of the multiplication unit 405 is K _TR · C ^* as a result of the output C ^* of the addition / subtraction unit 404 multiplied by the conversion value K _TR .

The acceleration peaks 418, the difference between the current angle of attack (α) and the reference angle of attack (α ₀₎ - a (α α ₀₎ are calculated. In the addition / subtraction unit 411, the product of the current displacement δ _cc of the steering column and the conversion gain Kccα1 410 is subtracted from the output α-α ₀ of the addition subtraction unit 418. That is, the output of the addition-subtraction section 411 is α-α ₀ -Kccα1 · δ _cc . In the addition / subtraction unit 406, the output α-α ₀ -K _ccα1 · δ _cc of the addition subtraction unit 411 is subtracted from the output K _TR · C ^* of the multiplication unit 405. That is, the output of the addition / subtraction unit 406 is K _TR · C ^* -α + α ₀ + K _ccα1 · δ _cc .

The output of the adder / subtracter 406 is multiplied by the feedback proportional gain (K _FB ) 407 and the proportional integral gain (1 + 1 / S) 408. That is, proportional integral output Then, Equation 4 below is established.

[Equation 4]

^{_{= [K TR · C * -}} α + α 0 + K ccα1 · δ cc] · K FB · (1 + 1 / S)

The adder / subtracter 409 outputs the product of the pitch angular velocity q and the damper gain K _damp 413 (K _damp q) and the proportional integral output. After summing up, the product (K _FF · δ _cc ) of the current displacement (δ _cc ) 401 of the steering column and the feed forward gain (K _FF ) 402 is subtracted. That is, the displacement angle δ _e of the elevating rudder output from the add / subtract unit 409 is obtained according to Equation 5 below.

[Equation 5]

δ _e = + K _damp · q-K _FF · δ _cc

According to the angle of attack control method shown in FIG. 4, the displacement angle δ _e of the elevating rudder is controlled only at a constant ratio according to the current angle of attack α. Accordingly, even when the angle of attack α is greater than or equal to the boundary angle α _ALT , the aircraft 412 still has room to stall because it does not relatively change the displacement angle δ _e of the elevator.

SUMMARY OF THE INVENTION An object of the present invention is to provide a method of controlling the angle of attack of an aircraft, which can make a relatively large change in the displacement angle of the elevator when the current angle of attack becomes equal to or more than the boundary angle of attack.

1 is a control block diagram showing the angle of attack control method of a conventional aircraft.

FIG. 2 is a schematic diagram of a load feel system added to the control method of FIG. 1. FIG.

FIG. 3 is a graph showing the relationship between the forces exerted on the displacement of the steering column in the system of FIG.

Figure 4 is a control block diagram showing the angle of attack control method of another conventional aircraft.

5 is a control block diagram showing the angle of attack control method of the aircraft according to the present invention.

6 is a schematic diagram of a load fill system in addition to the control method of FIG.

7 is a characteristic diagram of the force exerted upon displacement of a steering column in the system of FIG.

8 is a characteristic view of the force applied to the steering column according to the angle of attack in the system of FIG.

9 is a graph of displacement of the steering column versus time applied to the experiment of the control method of FIG.

FIG. 10 is a graph showing a displacement angle of a hoist with respect to time applied to the graph of FIG. 9; FIG.

FIG. 11 is a graph showing an angle of attack of a lift over time applied to the graph of FIG.

12 is a graph showing vertical overload with respect to time applied to the graph of FIG.

<Explanation of symbols for the main parts of the drawings>

101, 401, 501 ... the current displacement of the steering column,

102, 402, 502 ... feed forward gain,

103, 403, 410, 503, 510, 522 .... conversion gain,

405, 505, 525 ...

504, 506, 509, 518, 511, 516, 518, 521, 523, 526 ....

C ^* ... first control value,

105, 407, 507 ... feedback proportional gain,

106, 408, 508 ...... proportional integral gain, δ _e .....

108, 412, 512 .... aircraft, q .., pitch angular velocity,

109, 413, 513 ... damper gain,

110, 414, 514 ...... pitch angular velocity feedback gain,

△ nz ..... Vertical Overload, 111, 415, 515 ....... Vertical Overload Gain,

α .... current angle of attack, 417, 527 ....

α _ALT ...... Boundary angle of attack, α ₀ ....... Reference angle of attack,

K _TR ........ Switching conversion value, 419, 519 .....

M ..... aircraft speed, δ _FLAP ...... flap angle,

21, 61 .... steering column, 22, 62 ... rotation center,

23, 63, 65 ...... Spring, 24, 64 ..... Support,

δ _cc ....... the current displacement of the steering column, Fcc ...... the force on the steering column,

δ _ccmax ..... the maximum displacement of the steering column, _ΔFcc .... the combined force,

Fcc ₂ ....... force at maximum displacement, α _STL ... stall angle,

520, δ _ccsoft ... boundary displacement of the steering column,

524 ... Switching.

The angle of attack control method of the aircraft of the present invention for achieving the above object comprises the steps of (a) setting the stall angle of attack (α _STL ) that can cause stall of the aircraft. (b) A boundary receiving angle α _ALT smaller than the stall attack angle α _STL and a reference angle α ₀ smaller than the boundary receiving angle α _ALT are obtained. (c) in accordance with the first control value C ^* proportional to the pitch angular velocity q and the vertical overload Δnz of the aircraft and inversely proportional to the current displacement δ _cc of the steering column and the current angle of attack α. , Control the displacement angle δ _e of the elevating rudder. (d) If the current angle of attack α is equal to or greater than the boundary angle of attack α _ALT , the first control value C ^* is adjusted to be subtracted by the first value δ _αSTL . And (e) if the current angle of attack α is greater than the reference angle of attack α ₀ and less than the boundary angle of attack α _ALT , then proportional to the current angle of attack α in a range less than the first value α α _STL . The first control value C ^* is adjusted to be subtracted by a second value.

According to the angle of attack control method of the aircraft of the present invention, according to the execution of the step (d), when the current angle of attack α becomes equal to or more than the boundary angle of attack α _ALT , a relatively large change in the displacement angle δ _e of the elevator hit Can be given.

Preferably, (f) pulling the steering column in a direction opposite to the direction of travel of the steering column with a force Fcc proportional to the current angle of attack α; And (g) if the current angle of attack (α) reaches the boundary angle of attack (α _ALT ), performing step (f) by adding a separate force (ΔFcc). This allows the pilot to accurately sense how much force is applied to the steering column to reach the alert angle α _ALT .

Hereinafter, preferred embodiments of the present invention will be described in detail.

Figure 5 shows the angle of attack control method of the aircraft according to the present invention. Referring to FIG. 5, the displacement angle δ _e of the elevating rudder output from the adder / subtracter 509 is proportional to the first control value C ^* output from the adder / subtracter 506. The first control value C ^* is proportional to the pitch angular velocity q and the vertical overload Δnz of the aircraft 512 and is dependent on the current displacement δ _cc of the steering column and the current angle of attack α. Inversely If the current angle of attack α is equal to or greater than the boundary angle of attack α _ALT , the first control value C ^* is adjusted to be subtracted by the first value δ _αSTL . If the current angle of attack α is larger than the reference angle of angle α ₀ and smaller than the boundary angle of attack α _ALT , the current angle of attack α is subtracted by a second value proportional to the current angle of attack α in a range smaller than the first angle of angle α _ATL . Adjust the first control value C ^* so that it is adjusted. In this way, if the current angle of attack α is equal to or more than the boundary angle of attack α _ALT , the first control value C ^* is adjusted to be subtracted by the first value δ _αSTL larger than the second value, so that the displacement angle of the elevator δ _e ) can give a relatively large change.

In the addition / subtraction unit 516, the product (Kq · q) of the pitch angular velocity q of the aircraft 512 and the feedback gain Kq 514 is a vertical overload Δnz and its gain Knz 515. It is added to the product (Knz · Δnz). Acceleration peaks 504 is the product of the sum current displacement of the steering column in (Kq · q + Knz · △ nz) (δ cc) (501) and the conversion gain (Kccn) (503) (Kccn · δ _cc Subtract). That is, the output C ^* of the adder / subtracter 504 is obtained according to Equation 6 below.

[Equation 6]

C ^* = Kq q q Knz Δnz-Kccn δ _cc

In the angle of attack determining unit 519, the angle of attack angle α _ALT smaller than the stall angle of attack α α _STL and the boundary angle α _ALT depending on the speed M and the flap angle δ _FLAP of the aircraft 512. Determine a small reference angle of attack α ₀ . The switching conversion unit 417 obtains a conversion value K _TR for the current angle of attack α. The converted value K _TR is 1 when α ≤ α _{0, and} when α ₀ <α <α _ALT It becomes 0 if (alpha) _{ALT <(} alpha). In the multiplication unit 505, the output C ^* of the addition / subtraction unit 504 is multiplied by the converted value K _TR . That is, the output of the multiplication unit 505 is K _TR · C ^* as a result of the output C ^* of the addition / subtraction unit 504 multiplied by the converted value K _TR .

The acceleration peaks 526, the difference between the current angle of attack (α) and the reference angle of attack (α ₀₎ - a (α α ₀₎ are calculated. In the addition / subtraction unit 521, the displacement difference δ _cc -δ _ccsoft is output by subtracting the boundary displacement δ _ccsoft 520 of the steering column from the current displacement δ _cc 501 of the steering column. Therefore, the output δ _αSTL of the addition / subtraction unit 523 is obtained by the following equation.

[Equation 7]

δ _αSTL = α-α _ALT -K _ccα2 (δ _cc -δ _ccsoft )

In Equation 7, if the conversion gain K _ccα2 522 of the displacement difference δ _cc -δ _ccsoft is the maximum displacement of the steering column corresponding to the stall angle α _STL , δ _ccmax , Obtained by the formula The switching conversion unit 527 obtains a conversion value (1-K _TR ) with respect to the current angle of attack α. The conversion value (1-K _TR ) is ₀ when α ≤ α _{0, and} when α ₀ <α <α _ALT It becomes 1 if (alpha) _ALT <(alpha). The output δ _α ^* of the switching unit 524 becomes the output δ _αSTL of the addition / subtraction unit 523 when α _ALT ≦ α. The multiplication unit 525 multiplies the output δ _α ^* of the switching unit 524 by the conversion value 1-K _TR from the switching conversion unit 527 and inputs the multiplication unit 506 to the subtraction unit 506. Therefore, the first value of the output of the multiplication unit 525 applied when α _ALT ≤ α is the output δ _αSTL of the addition / subtraction unit 523.

The acceleration peaks 518, a difference between the current angle of attack (α) and the reference angle of attack (α ₀₎ - a (α α ₀₎ are calculated. In the addition / subtraction unit 511, the product of the current displacement (δ _cc ) 501 of the steering column and the conversion gain _Kccα1 510 is subtracted from the output α-α ₀ of the addition / subtraction unit 518. That is, the output δ _αALT of the addition / subtraction unit 511 is obtained by the following equation (8).

[Equation 8]

δ _αALT = α-α ₀ -K _ccα1 , δ _cc

In Equation 8, the conversion gain K _ccα1 of the current displacement δ _cc 501 is Obtained by the formula The switching conversion unit 527 obtains a conversion value (1-K _TR ) with respect to the current angle of attack α. The conversion value (1-K _TR ) is ₀ when α ≤ α _{0, and} when α ₀ <α <α _ALT It becomes 1 if (alpha) _ALT <(alpha). The output δ _α ^* of the switching unit 524 becomes the output δ _αALT of the add / subtract unit 511 when α ₀ <α <α _ALT . The multiplication unit 525 multiplies the output δ _α ^* of the switching unit 524 by the conversion value 1-K _TR from the switching conversion unit 527 and inputs the multiplication unit 506 to the subtraction unit 506. Therefore, the second value which is the output of the multiplier 525 applied when α ₀ <α <α _ALT is δ _αALT · Obtained by the formula On the other hand, there is no output of the multiplier 525 that is applied when α ≦ α ₀ .

In the adder / subtracter 506, the output of the multiplier is subtracted from the output K _TR · C ^* of the multiplier 505. Therefore, the output C ^** of the addition / subtraction unit 506 applied when α ≦ α ₀ is obtained by the following equation (9).

[Equation 9]

C ^** = K _TRC ^*

In addition, the output C ^** of the addition / subtraction unit 506 applied when α ₀ <α <α _ALT is obtained by the following equation (10).

[Equation 10]

In addition, the output C ^** of the addition / subtraction unit 506 applied when α _ALT ≦ α is obtained by the following equation (11).

[Equation 11]

C ^** = K _TR · C ^* _{-δ αSTL}

The output C ^** of the adder / subtracter 506 is multiplied by the feedback proportional gain (K _FB ) 507 and the proportional integral gain (1 + 1 / S) 508. That is, proportional integral output applied when α ≤ α ₀ Is obtained by the following equation (12).

[Equation 12]

In addition, the proportional integral output applied when α ₀ <α <α _ALT Is obtained by the following equation (13).

[Equation 13]

And, proportional integral output applied when α _ALT ≤ α Is obtained by the following equation (14).

[Equation 14]

The addition / subtraction unit 509 outputs the product (K _damp · q) of the pitch angular velocity q and the damper gain K _damp 513 and the proportional integral output. After summation, the product (K _FF · δ _cc ) of the current displacement (δ _cc ) 501 of the steering column and the feedforward gain (K _FF ) 502 is subtracted. That is, the displacement angle δ _e of the elevating rudder applied when α ≦ α ₀ is obtained by Equation 15 below.

[Equation 15]

δ _e = K _TR · C ^* · K _FB · (1 + 1 / S) + K _damp · q-K _FF · δ _cc

In addition, the displacement angle δ _e of the elevating rudder applied when α ₀ <α <α _ALT is obtained by the following equation (16).

[Equation 16]

-K _FF · δ _cc

Then, the displacement angle δ _e of the elevating rudder applied when α _ALT ≤ α is obtained by the following equation (17).

[Equation 17]

δ _e = [K _TR · C ^* _{-δ αSTL} ] · K _FB · (1 + 1 / S) + K _damp · q-K _FF · δ _cc

As such, when the current angle of attack α is equal to or greater than the boundary angle of attack α _ALT , the second value Since the first control value C ^* is adjusted to be subtracted by a larger first value δ _αSTL , it is possible to make a relatively large change in the displacement angle δ _e of the _hoist . Thus, there is no room for the aircraft 512 to stall.

6 shows a load fill system added to the control method of FIG. 5. FIG. 7 shows the characteristics of the force Fcc applied according to the displacement δ _cc of the steering column 61 in the system of FIG. have. FIG. 8 shows the characteristics of the force Fcc applied to the steering column 61 according to the angle of attack α in the system of FIG. 6.

6, 7 and 8, the first spring 63 installed between the lower portion of the steering column 61 and the support 64 is a lower portion of the lower portion of the steering column 61 with a force Fcc proportional to the current angle of attack. It is designed to pull the steering column 61 in the direction opposite to the advancing direction. Further, the second spring 65 has a range from the boundary position δ _ccsoft where the displacement δ _cc of the steering column 61 corresponds to the boundary angle of attack α _ALT to the maximum position δ _cc = δ _ccmax . Is designed to generate a force equal to ΔFcc. Here, the maximum position δ _ccmax is a position corresponding to the stall attack angle α _STL . The force Fcc pulling the steering column 61 is represented according to Equation 18 below.

Equation 18

The first term in Equation 18, the first spring 63, the displacement (δ _cc ) of the steering column 61 in the range from the origin (δ _cc = 0) to the maximum position (δ _cc = δ _ccmax ) Is the product of the modulus of elasticity K ₁ and the current displacement δ _cc of the steering column 61. In addition, the second term, the displacement of the steering column 61 (δ _cc ) of the second spring 63, in the range from the boundary position (δ _cc = δ _ccsoft ) to the maximum position (δ _cc = δ _ccmax ) It is the product of the modulus of elasticity K ₂ and the current displacement δ _cc of the steering column 61.

When the pilot pushes the handle of the steering column 61, the steering column 61 rotates about the rotation center 62. At this time, the force of Fcc ₁ is initially applied at the origin (δ _cc = 0). While the steering column 61 moves from the origin δ _cc = 0 to the boundary position δ _ccsoft corresponding to the boundary angle of attack α _ALT , a force of K ₁ δ _cc is applied. In order for the steering column 61 to move from the boundary position δ _ccsoft to the maximum position δ _ccmax corresponding to the stall angle α _STL , the force Fcc ₃ must be greater than the force of Fcc ₂ . That is, the force of K ₁ · δ _cc + K ₂ · δ _{cc to which} the force of ΔFcc is added is taken up to Fcc ₄ . That is, the pilot can accurately sense whether the steering column 61 has reached the boundary angle of attack α _ALT by the force _ΔFcc added when the steering column 61 is at the boundary position δ _ccsoft . Accordingly, even if the pilot is inexperienced, the aircraft can be prevented from reaching the stall angle of attack (α _STL ).

Experimental results for the control algorithm of FIG. 5 are shown in FIGS. 9, 10, 11 and 12. Referring to the figures, the displacement (δ _cc ) of the steering column is maintained after being changed from the origin to 0.14 m (see FIG. 9). Accordingly, the displacement angle δ _e of the elevating rudder is maintained after changing from 0 ° to -10 ° (see FIG. 10). In addition, the vertical overload Nz is reduced at a constant rate after changing from 1.0 g to 1.7 g (see Fig. 12). Referring to FIG. 11, the stall angle α _STL 111 is 16 °, the boundary angle α _ALT 113 is 12 °, and the reference angle α ₀ 114 is 9 °. At this time, the actual angle of attack α 112 of the aircraft is changed to about 15 ° smaller than the stall angle α _STL 111 at 5 ° and then gradually approaches the boundary angle α _ALT 113.

As described above, according to the method of controlling the angle of attack of the aircraft according to the present invention, when the current angle of attack becomes greater than the boundary angle of attack, a relatively large change is made in the displacement angle of the elevator, thereby reducing the possibility of stalling the aircraft.

In addition, since the change in force applied to the adjustment column is increased at the angle of attack, even an inexperienced pilot can accurately sense that the angle of attack has been reached.

The present invention is not limited to the above embodiments, and modifications and improvements are possible at the level of those skilled in the art.

Claims (11)

(a) setting a stall angle of attack (α _STL ) that can cause stall of the aircraft; (b) obtaining a boundary angle of attack (α _ALT ) smaller than the stall angle of attack (α _STL ) and a reference angle of attack (α ₀ ) smaller than the boundary angle of attack (α _ALT ); (c) Elevate according to the first control value (C ^* ) proportional to the pitch angular velocity (q) and vertical overload (Δnz) of the aircraft and inversely proportional to the current displacement (δ _cc ) of the steering column and the current angle of attack (α). Controlling the displacement angle δ _{e of} ; (d) _adjusting the first control value C ^* to be subtracted by a first value δ _αSTL if the current angle of attack α is greater than or equal to the boundary angle α _ALT ; And (e) If the current angle of attack α is larger than the reference angle of attack α ₀ and smaller than the boundary angle of attack α _ALT , the current angle of attack is proportional to the current angle of attack α in a range smaller than the first value of α α _STL . And adjusting the first control value C ^* to be subtracted by two values. The method of claim 1, further comprising: (f) pulling the steering column in a direction opposite to the direction of travel of the steering column with a force Fcc proportional to the current angle of attack α; and (g) if the current angle of attack (α) reaches the boundary angle of attack (α _ALT ), further comprising the step of performing the step (f) by adding a separate force (ΔFcc). . The method of claim 1, wherein in the steps (d) and (e), the angle of attack angle difference between the first value δ _αSTL and the second value is obtained by subtracting the boundary angle of attack α _ALT from the current angle of attack α. angle of attack according to the aircraft, characterized in that proportional to (α-α _ALT ). 4. An _aircraft according to claim 3, wherein in steps (d) and (e), the first value δ _αSTL and the second value are inversely proportional to the current displacement δ _cc of the steering column. Angle of attack control method. The method of claim 4, wherein in the step (d), the first value (δ _αSTL) is, characterized in that in proportion to the boundary displacement of the steering column (δ _ccsoft) corresponding to the boundary angles of attack (α _ALT) How to control the angle of attack of the aircraft. The method of claim 5 wherein in step (d), the first value (δ _αSTL) the current displacement (δ _cc) obtained by subtracting the displacement of the boundary displacement (δ _ccsoft) of the steering column in the difference between the steering column ( angle of attack angle control method of the aircraft, characterized in that inversely proportional to (δ _cc -δ _ccsoft ). The method according to claim 6, wherein the conversion gain (K _ccα2 ) of the displacement difference (δ _cc -δ _ccsoft ) is the maximum displacement of the steering column corresponding to the stall angle α _STL (δ _ccmax ), The angle of attack control method of the aircraft, characterized in that obtained by the equation. 8. The angle of attack of the aircraft according to claim 7, wherein in the step (d), the first value δ _αSTL is obtained by the formula α-α _ALT -K _{ccα 2} · (δ _cc -δ _ccsoft ). Control method. 6. The method according to claim 5, wherein in step (e), the conversion gain K _ccα1 of the current displacement δ _cc of the steering column is The angle of attack control method of the aircraft, characterized in that obtained by the equation. 10. The angle of attack of an aircraft according to claim 9, wherein in the step (e), the second value is proportional to a value (δ _αALT ) obtained by the formula α − α _ALT -K _{ccα 1} · δ _cc . Control method. The method of claim 10, wherein in the step (e), the second value is δ _αALT · Angle of attack control method of the aircraft, characterized in that determined by the equation.