RU2280591C1 - Device for automatically shifting aircraft from stall and spinning conditions to standard mode of flight - Google Patents

Abstract

FIELD: aeronautical engineering.

SUBSTANCE: proposed device is provided with information sensors showing rates of yaw angle, pitch angle, roll angle and velocity head. Besides that device is provided with flight mode analyzer, logic unit, drives of rudder, ailerons and elevator and aircraft damping unit over entire ranges of angles of attack. Logic units include setters of threshold magnitudes of pitch angle, vertical overload and velocity head. Device is also provided with limiters of pitch angle and pitch angle error signals for rudder channel, roll signal limiter for aileron channel, rudder channel signal limiter, YES-NO comparison unit for preset and present magnitudes of velocity head, two YES-NO switches without self-reset, four OR-OR switches without self-reset and five summing-up amplifiers.

EFFECT: enhanced reliability and operational stability in all flight modes; enhanced controllability at all angles of attack.

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Description

The invention relates to the field of aviation technology and can be used as a system for automatically removing an aircraft from critical flight modes to a normal flight mode, to prevent entry into hazardous modes, to increase the reliability of withdrawal from these modes and flight safety.

A known system for withdrawing an airplane from corkscrew modes in which the steering surfaces are sequentially deflected. According to his own feeling in flight, the pilot determines the nature of the movement and performs the recommended deviations of the rudders to withdraw from critical flight modes (Sneshko Yu.I., Chetvergov V.N. Determination of the characteristics of an airplane corkscrew. Methods for determining the stability characteristics of an aircraft. - M.: Mashinostroenie, 1994 g.).

The disadvantage of this system is the probability of error in the analysis of the character of the movement by the pilot, an increase in the delay in the beginning of the withdrawal and, as a result, insufficient reliability of the withdrawal from the corkscrew mode.

A device is known for preventing entry into a corkscrew located on the edge of attack of an airplane or glider (Patent FR 2826632 A1, 06/29/2001, class IPC B 64 D 43.02).

The device includes a pressure sensor, a pressure difference sensor and a breaker. As soon as the pressure at the leading edge of the wing becomes less than the ambient pressure, the circuit breaker closes the electric circuit, which activates the horn, which signals the pilot about the risk of entering the tailspin.

However, when using such means signaling the approach to dangerous modes, there is also a high probability of a pilot's error when withdrawing from stall or corkscrew modes.

A known method and device for the withdrawal of maneuverable aircraft from a corkscrew according to the fifth method, which is mainly used in the presence of "second" stable balancing at large angles of attack (see Akhromeev VA, Goman MG, etc. "Automation of the withdrawal of maneuverable aircraft from the corkscrew mode, Technique of the Air Fleet, No. 3, 1991).

As an analyzer for hitting a plane in a corkscrew, an algorithm is proposed that uses the time during which the value of the angular yaw rate ω _y exceeds the value set in the case of rotational movement in the corkscrew, or the time during which the angle of attack exceeds the set value when the stabilizer is completely deflected in the case " falling leaf. "

To deduce from these modes, the dynamic or resonant buildup method is used, followed by the application of previously known methods to deduce from a corkscrew. Automation of the transition to the regular flight mode is not provided for in the device.

A device is known for preventing entry into a tailspin of a maneuverable aircraft, RF patent No. 601012, class. IPC B 64 C 13.18.

The device contains sensors of the angle of slip, angular velocity of the roll, the angle of attack, the analyzer of the corkscrew modes, the threshold value of the angle of attack, a comparison circuit, a subtracting element, a command generator with a logic unit, connected to an amplifying-converting device that converts signals and issues commands to reject control organs of the plane.

The subtracting element of the device generates a signal proportional to the increment of the angle of attack due to the angular roll velocity in accordance with the dependence, where the generated signal is proportional to the increment of the angle of attack only due to the presence of the slip angle and is obtained by appropriate scaling of the signal from the slip angle sensor. In the logic device, the output signal is sequentially compared with the master and with the signal of the slip angle sensor. A sequential deviation of the rudders occurs.

The device does not provide stable operation of the control system for all flight modes, it does not provide for automation of the transition to the normal flight mode, the time from a breakdown to a corkscrew to exit from it increases, and, consequently, a height difference, which reduces the reliability of the device and flight safety.

SUMMARY OF THE INVENTION

The objective of the invention is to provide a device that automatically displays an airplane from stall and corkscrew modes to regular flight mode, which provides increased reliability, stable operation of the control system in all flight modes, improved controllability at all angles of attack, including exceeding supercritical values.

In addition, the inventive device should have high efficiency output from corkscrew modes to normal flight mode of the aircraft by reducing altitude loss.

In accordance with the invention, the task is achieved in that a device for automatically removing an airplane from stall and corkscrew to normal flight mode, containing a block of information sensors for angular yaw, roll, pitch, located in the coordinate system associated with the plane, vertical overload, angle of attack, roll, pitch, speed head, flight mode analyzer, threshold value adjuster, logic unit, comparison device, summing devices connected to the steering wheel drives of aileron, elevator, and elevator, made connected with the damping unit of the aircraft, providing a stop rotation of the aircraft in the entire range of angles of attack. In this case, the first, second, third, fourth and fifth inputs of the damping unit are connected to the outputs of the sensors of the angular yaw rate, angular roll and pitch velocities, angles of attack and pressure head, respectively. The first, second and third outputs of the damping unit are connected in the logic unit with the first inputs of the first, second and third summing amplifiers of the channels of the rudder, ailerons, elevator, respectively. The outputs of the first, second, and third summing amplifiers are connected to the first inputs of the first, second, and third irrevocable switches of the "either-or" type, the second inputs of which are connected respectively to the outputs of the amplifiers of the rudder drives, ailerons, and elevator of the standard control system, and the outputs of these switches connected to the inputs of the steering gears of the rudder, ailerons and elevator.

In addition, the analyzer of the stall and corkscrew modes in the inventive device is made of a series-connected calculator of the angular yaw rate ψ, a high-pass filter, a block of the absolute value of the signal, a block for generating a command for controlling the logic block. In this case, the inputs of the calculator are connected to the outputs of the pitch angle, roll angle, yaw rate, angular roll speed sensors located in the coordinate system (ω _x , ω _y , ω _z ) associated with the aircraft. Moreover, the logic unit additionally includes pitch angle threshold adjusters, vertical overloads Nu _ass and speed head, signal limiters for pitch and roll angles, a yes-no comparison device of the set and current speed head value with the first input connected to the sensor pressure head, the second input - with a speed head setter, and the output connected through the first switch without self-return type "yes-no" with the first input of the fourth switch without self-return type "either-or". The second input of the fourth switch is connected to the output of the fourth summing amplifier.

To the first and second inputs of the fourth summing amplifier are connected a vertical overload sensor and a vertical overload master Nу, respectively. The third input of the fourth switch of the "either-or" type is connected via a pitch angle limiter to the output of the fifth summing amplifier, to the first and second inputs of which the pitch angle sensor and pitch angle adjuster are connected, respectively. The input of the second switch without self-return type "yes-no" is connected to the roll angle sensor, and the output is connected to the inputs of the third and fourth signal limiters on the ailerons and rudder channels, the outputs of which are connected respectively to the second inputs of the second and first summing amplifiers of the aileron and steering channels directions.

The output of the fourth irrevocable switch of the "either-or" type through the pitch misbalance signal limiter is connected to the second input of the third summing amplifier of the elevator channel, while the control inputs of the first, second switches without self-returning of the yes-no type, the first, second, and third irrevocable switches of the "either-or" type are connected to the output of the command generation block by the analyzer logic control block for executing the commands.

This embodiment of the inventive device allows for a minimum loss of altitude due to the continuous operation of the automatic control system in a closed loop to increase the reliability of the aircraft withdrawal from corkscrew to regular flight mode, to ensure stable operation of the control system in all flight modes, including at angles of attack greater than supercritical values, and improve flight safety.

The list of figures in the drawings.

The invention is illustrated by figures 1, 2, 3, 4, 5.

Figure 1 presents the block diagram of the device, where indicated:

1. Damping unit along three axes X / Y / Z;

2. Block logic;

3. Flight mode analyzer;

4. Block of information sensors;

5. DUS yaw;

6. CRS roll;

7. DUS pitch;

8. Vertical overload sensor;

9. The angle of attack sensor;

10. Pitch angle sensor:

11. Roll angle sensor;

12. Speed head sensor;

13. Pitch angle adjuster;

14. The adjuster vertical overload;

15. The first non-return switch of the type "or-or" drive rudder;

16. The second permanent switch type "or-or" drive ailerons;

17. The first summing power amplifier channel steering;

18. The second summing amplifier of the aileron channel;

19. The third summing power amplifier channel elevator;

20. The third signal limiter on the roll of the aileron channel;

21. The fourth signal limiter is directed along the steering channel;

22. Second pitch misalignment limiter;

23. The second switch without self-reset type "yes-no" roll;

24. The first signal pitch limiter;

25. The fourth switch of signals of the type "either-or";

26. The fifth summing amplifier of the pitch channel;

27. Power steering rudder standard system;

28. Amplifier drive ailerons standard system;

29. A yes-no comparison device of a given and current velocity head values;

30. The fourth summing amplifier current and preset values of vertical overload;

31. The calculator of the angular velocity of yaw;

32. High pass filter;

33. The block of the absolute value of the signal;

34. A block for generating commands for controlling a block of logic;

35. The third permanent switch type "or-or" elevator;

36. Power steering elevator standard system;

37. speed sensor;

38. Steering wheel rudder;

39. Steering drive ailerons;

40. Steering wheel elevator;

41. The first switch without self-reset type "yes-no."

Figure 2, 3 shows the results of mathematical modeling of the process of automatically bringing the aircraft into normal flight mode of the claimed device, connected to the output from a flat corkscrew by the analyzer command for 55 s.

Figure 4, 5 shows the process of automatically bringing the aircraft into normal flight mode when stalling under braking with γ _ass = 60 °, while figure 2, 3, 4, 5 indicate:

K - command from the analyzer;

F - signal from the output of the filter;

N - flight altitude;

V is the flight speed;

α is the angle of attack;

υ is the pitch angle;

γ is the angle of heel;

γ ω _x , ω _y , ω _z - the angular velocity of the roll, yaw and pitch.

Information confirming the possibility of carrying out the invention.

A device for automatically removing an airplane from stall modes and a corkscrew to a normal flight mode (see Fig. 1) contains a logic unit 2, an analyzer of stall modes and a corkscrew 3, a block of information sensors 4.

According to the invention, the device is made connected with the damping unit of the aircraft in three coordinate axes X, Y, Z 1. The first, second, third, fourth and fifth inputs of which are connected to the outputs of the sensors of the angular velocity of yaw 5, angular velocities of roll 6 and pitch 7 located along the coordinate system associated with the aircraft, the angles of attack 9 and the pressure head 12 of the block of information sensors 4, respectively. The first, second and third outputs of the damping unit are connected in the logic unit 2 with the first inputs of the first, second and third summing amplifiers of the channels of the rudder 17, ailerons 18, elevator 19, respectively.

The analyzer of the modes of stall and corkscrew 3 is made of a series-connected calculator of the angular velocity of yaw 31, a high-pass filter 32, a block of the absolute value of the signal 33, a block for generating a control command 34 of the logic block.

In this case, the yaw rate calculator 34 is connected to the outputs of the pitch angle sensors 10, roll angle 11, yaw rate 5, pitch angular velocity 7. Logic unit 2 contains pitch angle threshold adjusters 13, vertical overload Ny _back 14 and speed head 37, signal limiters at pitch angles — first 24 and second 22, yaw 20 and roll 21, summing amplifiers 17, 18, 19, 30, 26.

The yes-no comparison device of the set and current value of the pressure head 29 with the first input is connected to the pressure head sensor 12, the second input is connected to the speed head 37 and the output is connected via the first switch without self-return of the yes-no type 41 with the first input the fourth irrevocable switch of the type "either-or" 25, the second input of which is connected to the output of the fourth summing amplifier 30, the vertical overload sensor 8 and the transmitter are connected to the first and second inputs of the fourth summing amplifier 30 tikalnoy overload Ny 14 respectively.

The third input of the fourth switch of the type "either-or" 25 is connected through a signal limiter at pitch angle 24 to the output of the fifth summing amplifier 26, to the first and second inputs of which are connected the pitch angle sensor 10 and pitch angle adjuster 13, respectively.

The input of the second switch without self-return type "yes-no" 23 is connected to a roll angle sensor 11, its output is connected to the inputs of the third 20 and fourth 21 signal limiters on the channels of the ailerons and rudders, respectively.

The output of the first irrevocable switch of the type "either-or" 25 through the limiter of the error signal for the pitch difference 22 is connected to the second input of the third summing amplifier 19 of the elevator channel. The outputs of the first 17, second 18 and third 19 summing amplifiers are connected to the first inputs of the first 15, second 16 and third 35 non-return switches of the "either-or" type, the second inputs of which are connected respectively to the amplifiers of the rudder drives 27, ailerons 28 and elevator 36 standard control system, the outputs of these switches are connected to the inputs of the steering gears of the rudder 38, ailerons 39 and elevator 40.

In this case, the control inputs of the first 41, second 23 switches without self-resetting of the yes-no type, the first 15, second 16 and third 35 of the irrevocable switches of the type "either-or" are connected to the block for generating control commands of the logic unit 34.

The implementation of the computing units of the analyzer of the modes of stall and corkscrew 3, the damping unit 1 and the logic unit 2 in the digital version does not require large computing resources and does not present difficulties.

This embodiment of the inventive device will allow automatic withdrawal of the aircraft into the normal flight mode at the stage of determining the beginning of the stall mode or corkscrew, including at angles of attack exceeding supercritical values.

Device operation

The operation of the device for automatically removing the aircraft from the stall and corkscrew modes to the normal flight mode is carried out sequentially in three stages:

- determination of the beginning of the stall mode or corkscrew;

- stop rotation of the aircraft;

- the conclusion of the aircraft in flight mode with a trajectory angle close to the minimum possible (in absolute value).

The beginning of the stall or corkscrew mode is determined by analyzing the low-frequency component of the angular yaw rate, which is determined in the calculator (3) by the formula

где

Where

γ is the angle of heel,

ω _y , ω _z are the angular velocity of yaw and pitch in the coordinate system associated with the plane,

ψ is the angular velocity of yaw.

The angular yaw rate is composed of two components due to the movement around the center of mass of the aircraft and the movement of its center of mass. In normal flight modes, the frequency of these movements differs by more than an order of magnitude. In accordance with this, the amplitude of the angular velocity of yaw of the trajectory component of the motion (motion of the center of mass) is significantly smaller than the amplitude due to the motion around the center of mass.

When a stall or corkscrew mode occurs, the frequency and amplitude of oscillations caused by the movement of the center of mass increase significantly, which is a sign of the appearance of these modes. The introduction of a high-pass filter 32 into the signal of the angular velocity of yaw allows you to select a component related to the motion of the center of mass, and, having estimated its absolute value obtained in block 33, by comparing with a certain threshold value in block 34, generate a command to turn on the device, according to which the switches 15, 16 and 35 disconnect the control signals of the standard system 27, 28 and 36 from the inputs of the drives of the controls 38, 39 and 40 and connect the signals of the summing amplifiers 17, 18 and 19 of the device to them. The same command through the switch 41, the signal from the comparison device 29 is connected to the switch 25.

The selection of the high-pass filter parameters and the threshold value of the signal at which the device is turned on is carried out on the day of each type of aircraft by mathematical modeling and is confirmed in flight tests.

The algorithm of the damping unit 1 ensures that the rotation of the aircraft is stopped in the entire range of angles of attack by generating the necessary control signals for rudders, heights and ailerons.

After stopping rotation, the aircraft begins to switch to flight mode with a trajectory angle close to the trajectory angle when flying with maximum aerodynamic quality.

When rotation stops, the signal at the output of the absolute value block of the analyzer 34 becomes less than the specified threshold value and the analyzer issues a command to enable stabilization of the zero roll angle by connecting the roll angle signal 23 from the sensor 11 to the signal limiters 20 and 21 and the set vertical overload value by connecting the signal the mismatch between the actual and the set overload values by the switch 25 to the signal limiter 22, providing an increase in the trajectory angle of the aircraft. At the same time, there is an increase in flight speed and, accordingly, speed head. When the preset value of the pressure head is reached, the signals from the comparison device 29 by the switch 25 by connecting the error signal between the preset and actual pitch angles through the signal limiter 24, the longitudinal channel is switched to the stabilization mode of the specified pitch angle, the value of which is determined by the steady-state flight mode with an angle of attack slightly less angle of attack corresponding to the maximum value of aerodynamic quality.

The selection of the parameters of the device is made for each type of aircraft in accordance with its characteristics and configuration by mathematical modeling and is confirmed by flight tests.

The operability of the device is confirmed by the results of mathematical modeling of aircraft movement during stall and corkscrew.

In FIG. 2, 3, an example of a transition process of motion parameters during automatic output of an aircraft to a normal flight mode using the inventive device is presented, when the device is switched on to output from a corkscrew by the analyzer command for 55 s, and in FIG. 4, 5 - when stalling on braking with γ _ass = 60 °.

The claimed device has increased reliability, allows for minimal loss of altitude to ensure stable operation of the control system in all flight modes, including at angles of attack exceeding supercritical values.

The device uses existing commercially available sensors of angular velocities 5, 6, 7, angles 9, 10, 11, overload 8 and velocity head 12. Implementation of the computing units of the flight mode analyzer 3, damping unit 1 and logic unit 2 in the digital version does not require large computing resources and is not difficult.

Literature

1. The dynamics of the aircraft corkscrew. Kotik N.G. - M.: Mechanical Engineering, 1976, p.242.

2. Dynamics of longitudinal and lateral movement. Byushgens G.S., Studnev R.V. - M.: Mechanical Engineering, 1979, S. 326-343.

3. RF patent №601892, B 64 С 13/18, 01/30/94.

Claims (1)

A device for automatically removing an airplane from stall and corkscrew to normal flight mode, which contains a block of information sensors for angular yaw, roll and pitch, located in the coordinate system associated with the aircraft, as well as information sensors for vertical overload, attack angles, roll, pitch and high-speed pressure , flight mode analyzer, logic unit, rudder, ailerons and elevator drives, characterized in that it is equipped with an airplane damping unit providing rotation stop aircraft in the entire range of angles of attack, with the first, second, third, fourth and fifth inputs of the damping unit of the aircraft connected to the outputs of the sensors of the angular velocity of yaw, angular roll and pitch angles of attack and velocity head, respectively, and the first, second and third outputs connected in the logic block with the first inputs of the first, second and third summing amplifiers of the channels of the rudder, ailerons, elevator, respectively, the outputs of the first, second and third summing amplifiers are connected with the first inputs and the first, second and third irrevocable switches of the “either-or” type, the second inputs of which are connected respectively to the amplifiers of the rudder drives, ailerons and elevators of the standard control system, the outputs of these switches are connected to the inputs of the rudder drives, ailerons and elevators, the flight mode analyzer includes a yaw rate calculator, a high-pass filter, a signal absolute value block, a block for generating a command for controlling a logic block, an input calculating yaw rate (ψ) connected to outputs of the pitch angle sensors, roll angle, roll rate, yaw and pitch (ω _x, ω _y, ω _z), vertical acceleration, and in logic block includes setting elements threshold pitch angle, the vertical overloads of Nu _ass and speed head, signal limiters at pitch and roll angles, a yes-no comparison device of the set and current values of speed head, which is connected to the speed head sensor by the first input, and the speed head setter by the second input a, and the output through the first switch without self-return type "yes-no" - with the first input of the fourth switch without self-return type "either-or", the second input of which is connected to the output of the fourth summing amplifier, the sensor is connected to the first and second inputs of the fourth summing amplifier vertical overload and vertical overload adjuster Nу, respectively, the third input of the fourth switch of the type "either-or" is connected through the signal limiter in pitch angle with the output of the fifth summing amplifier, to the first and second at the inputs of which the pitch angle sensor and pitch angle adjuster are connected, respectively, the input of the second switch without self-return type "yes-no" is connected to the roll angle sensor, and the output is connected to the inputs of the third and fourth signal limiters on the ailerons and rudder channels, the outputs of which are connected respectively, to the second inputs of the second and first summing amplifiers of the channels of the ailerons and rudder, the output of the fourth irrevocable switch of the type "either-or" through the limiter signal misalignment on pitch connection nen with the second input of the third summing amplifier of the elevator channel, the control inputs of the first, second switches without self-return of the yes-no type, the first, second, and third irrevocable switches of the either-or type are connected to the output of the command block of the flight mode analyzer.

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