RU86326U1 - ON-BOARD DIGITAL ANALOGUE ADAPTIVE AIRCRAFT CONTROL SYSTEM - Google Patents

Abstract

An on-board digital-to-analog adaptive control system for the aircraft, comprising a pitch angle adjuster and a first comparison unit, a second comparison unit, a roll angle adjuster and a third comparison unit in series, as well as a free gyroscope unit, a damping gyroscope unit, an actuator and a first, second and the third summing amplifiers, the outputs of which are connected to the first, second and third inputs of the actuator, respectively, and the second inputs of the first, second of the first and third comparison units are connected respectively to the first, second and third outputs of the free gyroscope unit, characterized in that it comprises series-connected flight speed sensors, a first multiplication unit, the second input of which is also connected to the output of the flight speed sensor, second multiplication unit, adder , a unit for generating the reciprocal of the signal and a scaling unit, connected in series with the flight altitude sensor and the first division unit, the output of which is connected to the second input of the adder, followed by the newly connected sensor of the current mass value and the second division unit, the output of which is connected to the third input of the adder, a reference signal adjuster, the first, second and third outputs of which are connected respectively to the second input of the second multiplication unit, the second input of the first division unit and the second input of the second division unit , heading angle adjuster, connected by the output to the first input of the second comparison unit, a digital multiplication unit, the first, second and third inputs of which are connected respectively to the outputs of the first, second and

Description

The utility model relates to airborne systems for automatic control of aircraft (control of angular parameters and speeds), which are characterized by substantially unsteady flight conditions.

Closest to the proposed solution is the "Angular motion control system of an unmanned aerial vehicle" [1]. It contains a series-connected pitch angle adjuster and a first comparison unit; second comparison unit; serially connected roll angle adjuster and a third comparison unit, as well as a unit of free gyroscopes, the first, second and third outputs of which are connected, respectively, with the second inputs of the first, second and third comparison units. The “System” also includes a block of damping gyroscopes, the first, second and third summing amplifiers and an actuator. In this case, the first, second and third outputs of the block of damping gyroscopes are connected to the inputs of the first, second and third summing amplifiers, respectively, and the outputs of these summing amplifiers are connected to the first, second and third inputs of the actuator, respectively.

The essential features of the prototype, which coincides with the essential features of the proposed technical solution, is that the on-board digital-to-analog adaptive control system for the aircraft contains a series-connected pitch angle adjuster and a first comparison unit, a second comparison unit, a series-connected roll angle adjuster and a third comparison unit, and also a block of free gyroscopes, a block of damping gyroscopes, an actuator and the first, second and third summing amplifiers, outputs the second are connected to the first, second and third inputs of the actuator, respectively, and the second inputs of the first, second and third blocks of comparison are connected, respectively, with the first, second and third outputs of the block of free gyroscopes.

The disadvantages of the known system are limited functionality in the conditions of non-stationary parameters of the aircraft’s motion, primarily the speed and altitude and inertial mass characteristics of the aircraft. This disadvantage is explained by the impossibility of the optimal choice of constant values of gear ratios of the control system with a sufficiently deep change in the characteristics of the control object. The constancy of gear ratios not only worsens the quality of transients in the channels of stabilization and control, but in extreme cases (a very deep change in characteristics) can lead to loss of stability of the aircraft.

The technical problem solved in the proposed control system is the expansion of the system’s functionality with the provision of an acceptable control quality in essentially unsteady flight conditions.

The specified technical result is achieved by the fact that in the known system for controlling the angular movement of an unmanned aerial vehicle, comprising serially connected pitch angle adjuster and a first comparison unit; second comparison unit; serially connected roll angle adjuster and a third comparison unit, as well as a block of free gyroscopes, a block of damping gyroscopes, an actuator and first, second and third summing amplifiers, the outputs of which are connected to the first, second and third inputs of the actuator, respectively, and the second inputs of the first, the second and third comparison blocks are connected, respectively, with the first, second and third outputs of the block of free gyroscopes; in addition, a speed sensor connected in series a flight, a first multiplication unit, the second input of which is also connected to the output of the flight speed sensor, a second multiplication unit, an adder, an inverse signal generating unit and a scaling unit, a flight altitude sensor and a first division unit, the output of which is connected to the second input of the adder, a mass current sensor and a second division unit, the output of which is connected to the third input of the adder, a reference signal setter, the first, second and third outputs of which are connected in series respectively, to the second input of the second block of multiplication, the second input of the first block of division and the second input of the second block of division, the angle gauge connected to the output of the first input of the second block of comparison, the digital block of multiplication, the first, second and third inputs of which are connected, respectively , to the outputs of the first, second and third comparison units, the fourth, fifth and sixth inputs are connected, respectively, to the first, second and third outputs of the scaling unit, and the first, second and third outputs are connected, respectively With the first, second, and third inputs of the digital-to-analog converter, the first, second, and third outputs of which are connected to the first inputs of the first, second, and third summing amplifiers, and the digital-to-analog multiplication unit, whose first, second, and third inputs connected, respectively, to the first, second and third outputs of the block of damping gyroscopes, the fourth, fifth and sixth inputs are connected, respectively, to the fourth, fifth and sixth outputs of the scaling unit, and the first, second and third outputs are connected s with the second inputs, respectively, of the first, second and third summing amplifiers.

The distinguishing features of the proposed technical solution are the following: a series-connected flight speed sensor, a first multiplication unit, the second input of which is also connected to the output of the flight speed sensor, a second multiplication unit, an adder, a signal return value generating unit and a scaling unit, a flight altitude sensor and the first division unit, the output of which is connected to the second input of the adder, connected in series to the sensor of the current mass value and the second division unit, the course of which is connected to the third input of the adder, a reference signal adjuster, the first, second and third outputs of which are connected, respectively, to the second input of the second multiplication unit, the second input of the first division unit and the second input of the second division unit, the heading angle adjuster connected to the first the input of the second comparison unit, a digital multiplication unit, the first, second and third inputs of which are connected, respectively, to the outputs of the first, second and third comparison units, the fourth, fifth and sixth inputs are connected, with Responsibly, to the first, second and third outputs of the scaling unit, and the first, second and third outputs are connected, respectively, to the first, second and third inputs of the digital-analog converter, the first, second and third outputs of which are connected to the first inputs, respectively, of the first , second and third summing amplifiers, and a digital-to-analog multiplication unit, the first, second and third inputs of which are connected, respectively, to the first, second and third outputs of the block of damping gyroscopes, the fourth, fifth and sixth input s are connected, respectively, with the fourth, fifth and sixth outputs of the scaling unit, and the first, second and third outputs are connected with the second inputs, respectively, of the first, second and third summing amplifiers.

The proposed construction of an on-board digital-to-analog adaptive control system for an aircraft allows a wide enough range of parameter changes that determine its dynamic properties (flight speed and altitude, mass-inertial characteristics) to provide the required quality of control processes. Ultimately, this makes it possible to carry out a stable flight of the aircraft and the required accuracy of a given trajectory of its movement.

The use of the proposed on-board digital-to-analog adaptive control system is especially effective on high-altitude aircraft with high flight speeds and high fuel mass consumption.

The proposed technical solution is illustrated by the diagrams of figures 1, 2.

Figure 1 presents the structural diagram of the system.

Figure 2 presents the structural diagram of the actuator.

The on-board digital-to-analog adaptive control system (Fig. 1) contains series-connected flight speed sensor 1 (DSP), the first multiplication unit 11 (1st control unit), the second input of which is also connected to the output of the flight speed sensor 1, the second multiplication unit 14 (2nd Control unit), adder 15 (C), the unit for generating the reciprocal of the signal 16 (BFOVS) and the scaling unit 17 (BM), the flight altitude sensor 2 (DVP) and the first division unit 12 (1st DB), the output of which is connected to the second the input of the adder 15, series-connected sensor tech a higher mass value 3 (DTM) and a second division unit 13 (2nd DB), the output of which is connected to the third input of the adder 15, the reference signal generator 10 (AIA), the first, second and third outputs of which are connected, respectively, to the second input of the second block multiplication 14, the second input of the first division unit 12 and the second input of the second division unit 13, serially connected pitch angle adjuster 19 (ZUT) and the first comparison unit 22 (1st BS), serially connected course angle adjuster 20 (ZUK) and the second comparison block 23 (2nd BS), connected in series e roll angle adjuster 21 (ZUKr) and a third comparison unit 24 (3rd BS), a digital multiplication unit 25 (CBU), the first, second and third inputs of which are connected, respectively, to the outputs of the first 22, second 23 and third 24 comparison units, the fourth, fifth and sixth inputs are connected, respectively, to the first, second and third outputs of the scaling unit 17, and the first, second and third outputs are connected, respectively, to the first, second and third inputs of the digital-to-analog converter 26 (DAC), the first, the second and third outputs of which are connected to the first the moves, respectively, of the first 7 (1st control system), second 8 (2nd control system) and the third 9 (3rd control system) of the summing amplifiers, and the digital-to-analog multiplication unit 18 (TsABU), the first, second and third inputs of which are connected, respectively, to the first, second and third outputs of the block of damping gyroscopes 4 (BDG), the fourth, fifth and sixth inputs are connected, respectively, to the fourth, fifth and sixth outputs of the scaling unit 17, and the first, second and third outputs are connected to the second inputs, respectively, of the first 7, second 8 and third 9 summing amplifiers, outputs to of which are connected to the first, second and third inputs of the actuator 5 (DUT), and the second inputs of the first 22, second 23 and third 24 comparison units are connected, respectively, with the first, second and third outputs of the free gyroscopes unit 6 (BSG). The aircraft (LA) in figure 1, and also in figure 2 is shown conditionally.

In accordance with the diagram - figure 2, the actuator 5 of the on-board digital-to-analog adaptive control system comprises serially connected the first adder 27 and the first steering gear 28, serially connected the first inverter 29, the second adder 30 and the second steering gear 31 and the second inverter connected in series 32, the third adder 33 and the third steering gear 34, wherein the first input of the actuator is connected to the second inputs of the second 30 and third 33 adders, the second input of the actuator is connected the first input of the first adder 27, the input of the first inverter 29 and the third input of the third adder 33, and the third input of the actuator is connected to the second input of the first 27 and third input of the second 30 adders and the input of the second inverter 32. Blocks 27, 29, 30, 32 and 33 conditionally combined in block 35 - block kinematic wiring (BKR).

Thus, the on-board digital-to-analog adaptive control system of the aircraft contains two circuits:

1 - control loop and stabilization of the aircraft, which includes links 4 ... 9, 19 ... 24;

2 - adaptation circuit, which includes links 1 ... 3, 10 ... 18 and 25, 26.

The control and stabilization circuit operates as follows.

The input signals of the actuator 5 σ _ϑ , σ _ψ , σ _γ , are formed in the summing amplifiers 7, 8, 9 according to the laws:

where i _a, i _n, i _e, ρ _a, ρ _n, ρ _e - gear ratio control and stabilization of the positional signals and the damping circuit defined system adaptation circuit; ω _x , ω _y , ω _z , are the angular speeds of rotation of the aircraft, which are the output signals of the block of damping gyroscopes 4.

Positional signals (ϑ-ϑ _у ), (ψ-ψ _у ), (γ-γ _у ) are generated in comparison blocks 22, 23, 24 by the signals of pitch angle, heading and roll indicators 19, 20, 21 and signals proportional to the angles orientation of the aircraft - to the output signals of the block of free gyroscopes 6.

Channel-by-channel control signals σ _ϑ , σ _ψ , σ _γ in the kinematic wiring block of the actuator 5 are distributed among the steering gears of the aircraft according to the relevant laws; for example, for an aircraft with three steering surfaces, located according to the scheme shown in figure 1, 2 in the aircraft block, these laws have the form [2, p. 7]:

where σ ₁ , σ ₂ , σ ₃ , are the input signals of the steering drives 28, 31, 34, which are part of the actuator 5; k _ψ , k _γ are the kinematic wiring coefficients implemented in the adders 27, 30, 33 of the kinematic wiring block (figure 2). The determination of the kinematic wiring coefficients can be performed, for example, in accordance with the formulas:

Where , - aerodynamic efficiency coefficients of the first and second rudders (with the rudder arrangement in FIGS. 1, 2) relative to the roll axes and the course of the aircraft. Formulas (3) are obtained under the assumption that the efficiency coefficients for the second and third rudders are equal.

The adaptation circuit works as follows.

In a first approximation, the equations of angular motion of the aircraft in the pitch channel, for example, according to [3, p.616] are described as:

where a, b are the dynamic coefficients of the aircraft for damping and rudder efficiency, respectively,

δ _in - the angle of deviation of the elevators by the actuator (steering gear) of the aircraft.

Assuming for simplicity of exposition of further reasoning, the actuator is ideal, i.e. putting δ _in = σ _ϑ , from equations (1), (4) you can get a description of the regulation processes in a closed loop "control system - aircraft" in the form:

or

The characteristic equation of a closed-loop control system in accordance with (6) has the form:

Equation (7) allows you to compare the parameters of the control law (1) i _in and ρ _in with the dynamic coefficients of the aircraft a and b. To ensure the constancy of the stability characteristics and the quality of transient processes, it is necessary to ensure that the stability conditions (invariance) of the coefficients of the characteristic equation are satisfied during the flight, i.e.

In equations (5) and, accordingly, (8), the coefficient a characterizing the inherent damping of the aircraft is small enough, varies slightly, and is not characteristic, dominant in comparison with the parameter bρ _в , which determines the damping of the closed system as a whole. Therefore, the conditions for maintaining the required indicators of stability and quality of transients, based on (8), can be defined as

where b is the coefficient of efficiency of the controls (steering surfaces) of the aircraft, determined in accordance with [3, p. 618] the formula

- the derivative of the aerodynamic efficiency coefficient for the deviation of the rudders,

S, b _a - characteristic geometric dimensions (area and average aerodynamic chord) of the aircraft,

I _z - moment of inertia of the aircraft relative to the transverse axis,

q - velocity head:

ρ is the air density at the current flight altitude, ρ = ρ (H),

ν is the flight speed.

If we assume that the aerodynamic coefficient is mainly a function of the number M (the ratio of flight speed to sound speed) and can be assumed constant for a sufficiently wide range of altitudes and flight speeds, it can be considered stable in the vicinity of the balancing values of the current orientation angles of the aircraft.

The moment of inertia I _z for unmanned aerial vehicles also varies slightly. Thus, the main reason for the non-stationary dynamic coefficient b (10) is the velocity head q. Based on equations (9) and on the basis of equations (10), (11), taking into account the above, the adaptation function λ must be formed in the form

λ = Aq,

Where

Moreover, for the regimes with velocity head values belonging to non-calculated values q <q _min and q> q _max , the quantity λ is limited to the corresponding values, i.e.

From equations (9), adaptation algorithms are obtained:

As indicated above, conclusions (4) ... (13) are made for the first approximation in the descriptions of the equations of motion of an aircraft. More accurate relationships are given by the following adaptation laws for all three control and stabilization channels:

where H is the flight altitude, m is the current value of the mass of the aircraft.

Formula dependencies (14) are implemented in the scheme of the proposed device. The values of ν, H, m are measured by sensors 1 ... 3, the value of ν ^{2 is} calculated in the first block of multiplication 11, the terms of the denominator of expressions (14) are calculated in the second block of multiplication 14 and in the first 12 and second 13 blocks of division, to the second inputs of which from the master the reference signals 10 are supplied by the constants B ₁ , B ₂ , B ₃ . The denominator is computed in an adder 15, an inverse value of its - in the block forming the reciprocal of the signal 16, a scaling unit 17 calculates the value of i _in, i _n, i _e, ρ _a, ρ _n, ρ _e by multiplying the reciprocal of the denominator, the constant scale values A _{1c, n, e} , A _{2c, n, e} .

In the digital multiplication unit 25, the digital values of the positional control signals obtained in the first 22, second 23, and third 24 comparison units are multiplied by gear ratios i _в , i _н , i _э . In the digital-to-analog Converter 26, the obtained values are converted into analog signals and fed to the inputs of the first 7, second 8 and third 9 summing amplifiers, the second inputs of which are fed analog signals, which are the results of multiplication in the digital-analog unit of multiplication of 18 digital values ρ _in , ρ _n , ρ _e to the analog signals of angular velocities ω _x , ω _y , ω _z from the output of the block of damping gyroscopes 4.

Most accurately, rather complicated adaptation laws based on relations (14) can be realized if there is a digital computer system on board the aircraft. All blocks of the control device are well-known and can be implemented on the elements of automation and computer technology, for example, according to [5, p. 103; 6, p. 55].

Thus, the proposed on-board digital-to-analog adaptive control system for the aircraft allows expanding the control capabilities of the aircraft and providing the required stability margins of the stabilization loop and the required quality of transients by adapting the gear ratios of the control channels to flight conditions.

Information sources:

1. The control system of the angular movement of an unmanned aerial vehicle. RF patent No. 2234117, 2003, cl. G05D 1/08.

2. A device for generating control signals for an aircraft control system. RF patent No. 2238582, 2003, cl. G05D 1/08.

3. Aerodynamics, stability and controllability of supersonic aircraft. Ed. G.S. Byushgens. M.: Science, Fizmatlit, 1998.

4. V.A. Bodner. Theory of automatic flight control. M .: Nauka, 1964, p. 178.

5. A.U. Yalyshev, O. I. Razorenov. Multifunctional analog control devices for automation. M. Engineering, 1981.

6. V. B. Smolov. Functional information converters. L .: Energy Publishing House, Leningrad Branch, 1981.

Claims (1)

An on-board digital-to-analog adaptive control system for the aircraft, comprising a pitch angle adjuster and a first comparison unit, a second comparison unit, a roll angle adjuster and a third comparison unit in series, as well as a free gyroscope unit, a damping gyroscope unit, an actuator and a first, second and the third summing amplifiers, the outputs of which are connected to the first, second and third inputs of the actuator, respectively, and the second inputs of the first, second of the first and third comparison units are connected respectively to the first, second and third outputs of the free gyroscope unit, characterized in that it comprises series-connected flight speed sensors, a first multiplication unit, the second input of which is also connected to the output of the flight speed sensor, second multiplication unit, adder , a unit for generating the reciprocal of the signal and a scaling unit, connected in series with the flight altitude sensor and the first division unit, the output of which is connected to the second input of the adder, followed by the newly connected sensor of the current mass value and the second division unit, the output of which is connected to the third input of the adder, a reference signal adjuster, the first, second and third outputs of which are connected respectively to the second input of the second multiplication unit, the second input of the first division unit and the second input of the second division unit , heading angle adjuster, connected by the output to the first input of the second comparison unit, a digital multiplication unit, the first, second and third inputs of which are connected respectively to the outputs of the first, second and of the comparison blocks, the fourth, fifth and sixth inputs are connected, respectively, to the first, second and third outputs of the scaling unit, and the first, second and third outputs are connected respectively to the first, second and third inputs of the digital-to-analog converter, the first, second and third outputs of which connected to the first inputs of the first, second and third summing amplifiers, respectively, and a digital-to-analog multiplication unit, the first, second and third inputs of which are connected respectively to the first, second and third outputs ladies damping unit gyroscopes, fourth, fifth and sixth inputs connected respectively with the fourth, fifth and sixth outputs of the scaling unit and the first, second and third outputs are connected to second inputs of respective first, second and third summing amplifiers.