RU2504815C2 - Method of aircraft control and device to this end - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: method of control over aircraft with two and more engines consists in differential feed of fuel. Fuel feed is performed by main fuel pumps and extra fuel system driven by actuating spring of one of the main engines or by motor and electric or air operated stabilisation control system. Proposed device consists of differential fuel feed system, direction and/or pitch control stabilisation system and set of gyroscope transmitters.

EFFECT: bank control in hovering.

13 cl, 9 dwg

Description

The invention relates to aircraft with two or more engines, mainly turbojet engines and with a thrust vector.

It is known to control the direction of a twin-engine aircraft using a chain of engines mounted to the right and left of the longitudinal axis of the aircraft, see Kotik MG "The dynamics of the take-off and landing of aircraft", Moscow: Engineering, 1984, pp. 118-119. It is unknown, but in principle it is possible to control the aircraft in pitch if two or more engines are located above and below the aerodynamic focus (although when designing an aircraft, aircraft designers try not to position the engines in this way). However, controlling the draw with the help of the main fuel pumps is difficult due to their large error. In addition, this will lead to accelerated wear of pump performance control mechanisms.

Also known are aircraft controlled by a deflected thrust vector.

The purpose of this invention is not only controlling a twin-engine aircraft in one of the parameters (direction or pitch), but with three or more engine airplanes in two of these parameters using a different rod of engines, but also, more importantly, stabilization in these parameters, i.e. failure from keels, vertical rudders, stabilizers, horizontal rudders, mechanisms and rods for controlling them. This will sharply reduce the aerodynamic drag of the aircraft, increase the weight return, and reduce the cost of the aircraft.

In addition, an equally important goal of the invention is the implementation of tracking (proportional) control of the spatial position of the aircraft in the tail-down position for takeoff and landing and in this position of promising aircraft. Without tracking control of the direction and pitch of the aircraft, take-off and landing in the tail-down position become so difficult that they are practically impossible. With this system, such take-off and landing on designated devices for this Pat. Russia No. 2335437 (two pillars with a chain between them). This will make it possible to abandon the landing gear, that is, it will make it possible to lighten, harden and reduce the cost of the aircraft even more. This will also reduce the wing area (especially the scope), since take-off and landing will be made from hovering mode, and, therefore, a large wing area, flaps, air brake and brake parachute become unnecessary.

In addition, the aim of the present invention is to control roll in hover mode when the ailerons are not working. The present invention provides for conventional, that is, non-tracking, but integral roll control (as in a normal flight). However, tracking roll control is also possible, for which the roll control system can be constructed in the same way as, for example, the pitch control system described below. But doing this is not recommended, since in this case the roll control range will be limited to some extent by moving the control knob.

This method of controlling an aircraft with two or more engines, which consists in differential fuel supply to the engines, differs in that this supply is carried out, along with the main fuel pumps of the engines, from the additional fuel system (s) driven by the drive spring of one of the main motors or from an electric motor and controlled from a gyroscopic stabilization-control system of electric or pneumatic type.

A device for implementing this method consists of a system (systems) of differential fuel supply, stabilization control system (s) in the direction and / or pitch and a block of gyroscopic sensors.

Option 1. The simplest differential fuel supply system for controlling in one parameter (direction or pitch) consists of a pump and a controlled three-way valve (flow distributor). It is better to drive the pump of such a system from an electric motor, since the drive from the drive spring of one of the main engines is fraught with the fact that the regulation efficiency when changing engine speed will change from insufficient to “idle” to excess to “maximum”.

For stabilization control in two parameters (direction and pitch), two such systems are needed. Moreover, they can have a common fuel pump. That is, such a system consists of one pump and two controlled three-way valves or three or more controlled valves.

Such a system has one feature - it is desirable for it to have a separate nozzle or nozzle ring in the combustion chamber, since when connecting to nozzles of the main fuel system, more precisely, with systems of two or more engines, the pressure difference in them and even the pressure difference in the combustion chambers will interfere with the stable distribution of fuel flows. That is, this method assumes rather high requirements for the synchronous operation of two, and even more so, three engines.

The following three options are free from this feature.

Option 2. In it, the differential fuel supply system for controlling in one parameter (direction or pitch) consists of two pumps with adjustable capacity. Moreover, both pumps can be controlled by one actuator, see figure 5.

Option 3. In it, the differential fuel supply system for controlling in one parameter (direction or pitch) consists of two pumps with electric motors with power regulators, for example, thyristor, see Fig.6.

Option 4. In it, the differential fuel supply system for controlling in one parameter (direction or pitch) consists of two or more main fuel pumps shunted by controlled jets, and after the pump there is a pressure sensor connected to the control input “less” of the stabilization system amplifier control, see Fig.8.

This option requires a small, about 1%, margin in performance of the main fuel pumps. And it also requires a linear characteristic of the controlled jets, that is, with the same deviation of the regulator “more or less”, the fuel supply should change by the same amount, otherwise the total total thrust of all engines will slightly change during the regulation process. This characteristic is achieved by the profile of the nozzle needle.

With any of the options for differential fuel supply, the control stabilization system for one parameter consists of a control knob 3 connected via a switch K 4 to the input of the proportional-integral controller 5 or to the input of the proportional controller 6, after which the outputs of both controllers through the same switch K 4 are connected to the input of the adder 7, with the other input of which is also connected a gyro sensor 8, and the output of the adder is connected to the input of the amplifier 9, the output of which, in turn, is connected to a different differential fuel supply system mechanism 10 and with a symmetric diode thyristor unit 11, from the output of which a signal for various purposes is supplied to the actuators of the nozzles of the controlled vector 13, and for these same inputs of these mechanisms a signal for different purposes (there and back) from the end switches of the sensor 13 the position of the actuator of the differential fuel supply system, and depending on the type of actuator, it can have a negative return through its position sensor communication with an input amplifier 9, see FIG. 2.

Moreover, in the electrical version of automation, the integrating part of the controller is a gyroscope of the same type as the gyroscopic sensor, but its rotor is stationary, and is controlled by a servo drive from the signal of the control unit.

The block of a symmetric diode thyristor may be a symmetric diode thyristor, but to increase the power, such a block can be a circuit of two parallel counter circuits from a series-connected diode and thyristor, to the control electrodes of which two chains from a series-connected dynistor and resistor are connected respectively.

In the pneumatic version of automation, the unit of the symmetric diode thyristor is two bypass valves in two lines for different purposes.

The signs “+” and “-” when controlling direct current can really correspond to electric plus and minus, but when controlling alternating current, especially three-phase, or with pneumatic automation, they mean signals of different movements for actuators (back and forth).

This system works as follows: the signal of the control unit ZU 3 in normal flight is supplied through the switch K 4 in position 2 to the proportional-integral controller 5 and then from the controller to the adder 7, which also receives the signal from the gyroscopic sensor G 8. The resulting signal is amplified by the amplifier At 9 and further it is fed to the actuator IM 10 of the differential fuel supply system (in Fig. 2 it is a three-way valve 2) and to the SDT 11 symmetric diode thyristor unit. Depending on the type of IM 10, there may be feedback from the sensor Ika of its position 13 to the input of the amplifier.

With such a signal transmission, the system, firstly, with the help of the G 8 sensor reacts to a change in the position of the aircraft and stabilizes the flight according to a given parameter (direction or pitch). Moreover, if the efficiency of the draw rod is insufficient, then the position sensor 13 of the actuator with one of its limit switches (hereinafter referred to as the “limit switch”, not shown separately) through the resistors R slowly moves the nozzle thrust vector in the desired direction using the actuators of the IC 12 nozzles (the latter may be 2, 3 or more).

And secondly, the system works out in an integrated mode (that is, like on a regular airplane) the pilot’s commands issued with the help of the memory charger 3, moreover, in two stages: small corrections are worked out with the help of a different rod, that is, with the help of a three-way crane 2, and sharp commands cause a large signal from the amplifier U 9, which exceeds the response threshold of the unit of the symmetric diode thyristor 11 and passes directly to the actuators of the nozzles of the IC 12, causing a sharp response.

However, the system has one more purpose: for a vertical landing by the tail-down type, the pilot, on approaching the landing device, makes a “slide”, or rather, a “cobra” and switches the K 4 switch to the “1” position in the vertical position along the horizon. In this position, the pitch control stabilization system switches to servo control from proportional controller P 6. What does “servo control” mean? This means that when the airplane control stick is deviated by a certain amount, the airplane does not start to increase or decrease the pitch incrementally, as it would be on a conventional airplane, but takes in a close to vertical position some inclination in the pitch direction, which automatically stabilizes, It does not depend on gusts of wind and other external influences, but depends only on the tilt of the control handle. For example, they turned the knob forward 10 degrees - the plane leaned forward 1 degree, rejected 20 degrees - the plane leaned 2 degrees, etc. It is even better to make this dependence non-linear, for example, the pen is rejected by 10 degrees - the plane deviated by 0.5 degrees, rejected by 20 degrees - the plane deviated by 2 degrees, rejected by 30 degrees - the plane deviated by 4 degrees, etc. This will allow the pilot to make very fine control near the vertical position. Such non-linearity can be achieved either by selecting the appropriate output characteristics of the master, or the characteristics of the amplifier, or regulator 6, or by installing a non-linear electronic converter anywhere between them (for example, the current-voltage characteristics of all diodes and transistors near zero are non-linear, that is, the basis of such a converter can be, for example, a high-voltage selenium rectifier or several series-connected diodes or transistors, to which an initial bias voltage is applied tions from the resistor bridge).

Of particular note is the direction control system - with the same K 4 switch near the vertical position, the pilot can switch it from direction sensors that do not work adequately near the vertical to the roll sensor (see below). In this case, one horizontal gyroscope provides good control of the plane tilting to the right and left from the vertical, that is, as if controlling the pedals in the direction, and also following - that is, when the pedals deviate, the plane tilts slightly to the left or right of the vertical position and maintains this position regardless external influences.

Deviating in this way by a fraction of a degree from the vertical, the aircraft can gently shift horizontally left-right and back and forth, steering to the landing gear.

To control a stabilization control system or two such systems (in direction and pitch), theoretically, two gyroscopes in orthogonal planes are enough. But the system will turn out simpler and more reliable if three gyroscopes are used, see Fig. 7.

This unit of gyroscopic sensors consists of three mutually perpendicular gyroscopes: horizontal 14, longitudinal 15 and transverse 16 (respectively, the three main planes of the aircraft), the output of the pitch sensor 17 of the horizontal gyroscope is input to the pitch amplifier 18, and the output from this amplifier is fed to the adder of the control system pitch (see paragraph 8), the output from the roll sensor 19 of the same gyroscope is fed to one of the contacts of the switch K 4, direction sensors 20 and 21 of the longitudinal and transverse gyroscopes are connected to the inputs their amplifiers direction 22 and 23, and the outputs of these amplifiers are connected to the input of the adder direction 24, the signal from which is fed to the switch K 4, and the vertical sensors 25 and 26 of the longitudinal and transverse gyroscopes are connected to the input of the adder vertical 27, the signal from which is fed to the control inputs are “larger” than the mentioned direction amplifiers, and from the common contact “o” of the switch K 4 the signal enters the adder of the direction control system (see point 8), and, in addition, the output from the roll sensor 19 of the horizontal gyro is fed to the control input “greater” than the pitch amplifier 18 and / or the output of the vertical adder 27 is fed to the input of the pitch amplifier 18 (shown by a dotted line), see FIG. 7.

This unit works along the pitch line as follows: a pitch deviation signal from the pitch sensor 17 of the horizontal gyroscope 14 is fed to its amplifier 18. In the absence of a roll, the gain is minimal (for example, equal to one). When the roll appears, the pitch signal will decrease, approximately like the cosine of the roll angle. To compensate for this attenuation, a signal from the roll sensor 19 is supplied to the input “more” of the amplifier 18, and the gain of the amplifier increases, it is desirable that the input-output characteristic of the amplifier be selected so that the pitch signal does not depend on the roll, i.e., is inversely proportional the cosine of the roll. Theoretically, such a system is not functional with a roll of exactly 90 degrees. However, this does not have practical significance, since such a regime is always actively controlled by the pilot and does not need stabilization. And secondly, at the slightest deviation from the absolute 90 degrees, the system starts working again.

But for the sake of fidelity, if tests require this, you can do something else: you can apply a total signal from the vertical adder 27 to the input of the pitch amplifier, since the constructive vertical becomes pitch with a roll of 90 degrees. You can apply both of these roll compensation options at the same time.

The signal about the deviation of the direction also depends on the roll and also depends on the pitch, that is, it depends on any deviation from the vertical. Therefore, the signals of the two direction amplifiers 22 and 23 can be amplified by deviating from the vertical using a control signal from the adder of the vertical 27.

Since at some spatial positions one of the two gyroscopes 15 and 16 becomes insensitive either to the direction or to the vertical (when the plane of the gyroscope coincides with one of these structural planes), then two vertical sensors are used, whose signals are summed and therefore averaged, and two sensors directions whose amplified signals also add up and therefore are averaged. To compensate for the attenuation of the direction signal, the amplification factor of the amplifiers is also used according to the signal of deviation from the vertical. Linear characteristics of the sensors are desirable. And finally, the amplified signal from the adder direction 24 is fed into the system of figure 2 for stabilization control.

And in the vertical hovering mode, the roll signal from the sensor 19 becomes the direction signal. This is done by the contact group of the switch K 4 mentioned earlier.

Another variant of the gyroscopic sensors unit is possible, in which the signals of the direction sensors are first summed and then amplified by one common direction amplifier. True, in order not to violate the unification of amplifiers, the signal from the adder should be attenuated by a divider or resistor by half. This option corresponds to paragraph 13 of the "claims", it is not graphically illustrated.

To control the roll in hovering mode, a control method has been applied for an aircraft with two or more engines with a deviating thrust vector, which consists in the fact that for the roll control, the thrust vectors of two or more engines are deflected so that the resulting force acting on the center of mass is zero.

For this, the aircraft has an appropriate system, which consists of a roll adjuster, the signal from which goes to the adders, which also receive a common signal from the thrust vector control system, and the signals from these adders are amplified by amplifiers, from the inputs of which are fed to the nozzle actuators.

The system will be much simpler, more reliable and more accurate if only an even number of engines participate in it. In this case, two.

Upon reaching a sufficiently high speed, the system becomes unnecessary and can be turned off by the signal of the speed sensor, which, using a relay, turns off the power or the signal of the sensor.

Figure 1 shows a possible version of the aircraft, for which the base can be MIG-29 or SU-27, rear view. The third engine under the fuselage is clearly visible.

Figure 2, 5, 6, 8 shows four options for a differential fuel supply system, where: 1 - pump, 01 - main fuel pump, 2 - three-way valve, 10 - actuator, 29 - thyristor engine control units, 30-controlled jets with their actuators IM 1 and IM 2, 31 - pressure sensor R. The signs "+" and "-" in Fig. 6 show that the same signal accelerates the rotation of one engine and slows down the rotation of the other.

Figure 2, in addition, shows the stabilization control system (discussed earlier).

Figure 3, 4 shows a possible variant of a three-way valve - flow distributor, where: 32 is a cylindrical cavity in which the plate 33 is fixed on the axis. Fuel is supplied through the expanding pipe 34, which is divided by the plate and the slot 35 into two side pipes 36, 37 .

Figure 7 shows a block of gyroscopes (discussed earlier).

9 shows a roll control system. It consists of a roll adjuster ZKR 3 connected to two adders 38, 39, which also receives a control signal from the pitch control system. The signals from the adders are fed to the amplifiers of the left and right engines 40, 41 and then to the actuators of the nozzles 12. The signal from the pitch control system is fed to the lower engine without an additional signal through the amplifier 42.

Differential fuel supply systems work like this:

- figure 2, 3, 4 - the actuator IM 10 rotates the plate 33 and the fuel in a predetermined ratio is distributed among the pipes 36, 37;

- figure 5 - the actuator IM 10 with the same thrust shifts the control levers of the controlled fuel pumps 1, and so that, closing one pump, it opens the other;

- in Fig. 6, the signal from the stabilization control system is supplied to the thyristor controllers 29 of the electric motors 28, so that one engine is accelerated and the other is decelerated. Accordingly, the performance of the pumps 1;

- in Fig. 8, the signal from the stabilization control system enters the controlled jets 30, so that one nozzle opens and the other closes. Accordingly, the fuel supply to the main fuel lines of these engines increases or decreases.

The system in Fig. 9 works as follows: the roll adjuster ZKR 3 gives two signals - “left” and “right”. They are fed to the adders 38, 39 of the left and right engines, where pitch signals are mixed, and then fed to the nozzles. Depending on the sign of the signal (“up” or “down”), the nozzles deviate in different ways and create a heeling moment.

The invention is intended for the modernization of MIG-29 and SU-27 aircraft and will make them fifth-generation aircraft. Expected characteristics: supersonic cruising speed, atmospheric acceleration speed - 4 M, ceiling - 25,000 meters, rate of climb - supersonic, permissible overload - 20 g (with a new anti-reloading suit and reinforced spar).

Claims (13)

1. The method of controlling an aircraft with two or more engines, which consists in differential fuel supply to the engines and characterized in that this supply is carried out, along with the main fuel pumps of the engines, also from the additional fuel system (s) driven by the drive spring of one of main engines or from an electric motor and controlled from a gyroscopic stabilization-control system of electric or pneumatic type. 2. The device for implementing the method according to claim 1, consisting of a system (systems) of differential fuel supply, stabilization control system (s) in the direction and / or pitch and a block of gyroscopic sensors. 3. The device according to claim 2, in which the differential fuel supply system for controlling in one parameter (direction or pitch) consists of a pump and a controlled three-way valve (flow distributor). 4. The device according to claim 2, in which the differential fuel supply system for regulation in both parameters (direction or pitch) consists of one pump and two controllable three-way valves or three or more controllable valves. 5. The device according to claim 2, in which the differential fuel supply system for controlling in one parameter (direction or pitch) consists of two pumps with adjustable capacity. 6. The device according to claim 2, in which the differential fuel supply system for controlling one parameter (direction or pitch) consists of two pumps with electric motors with power controllers. 7. The device according to claim 2, in which the differential fuel supply system for controlling in one parameter (direction or pitch) consists of two or more main fuel pumps shunted by controlled jets, and after the pump there is a pressure sensor connected to the control input "less »Amplifier stabilization control system. 8. The device according to claim 2, in which the control stabilization system according to one parameter consists of a control unit (3) connected via a switch (K 4) to the input of the proportional-integral controller (5) or to the input of the proportional controller (6), after whereby the outputs of both controllers through the same switch (K 4) are connected to the input of the adder (7), the gyro sensor (8) is also connected to the other input, and the output of the adder is connected to the input of the amplifier (9), the output of which, in turn, connected to the actuator differential fuel supply system (10) and with a block of symmetric diode thyristor (11), from the output of which a signal for various purposes goes to the actuators of the nozzles of the controlled vector (13), and a signal for different purposes (back and forth) arrives at the same inputs of these mechanisms limit switches of the sensor (13) for the position of the actuator of the differential fuel supply system, and depending on the type of actuator, through the sensor of its position, it has negative feedback from the input Ithel (9). 9. The device according to claim 8, in which, in the electrical version of the automation, the integrating part of the controller is a gyroscope of the same type as the gyroscopic sensor, but its rotor is stationary, and is controlled by a servo drive from the signal of the control unit. 10. The device according to claim 8, in which the symmetrical diode thyristor unit is a symmetric diode thyristor or a circuit of two parallel counter circuits of diode and thyristor connected in series, to the control electrodes of which are connected two chains of dynistor and resistor respectively . 11. The device according to claim 8, in which, in the pneumatic version of the automation, the block of the symmetric diode thyristor is two bypass valves in two lines for different purposes. 12. The device according to claim 8, in which the block of gyroscopic sensors consists of three mutually perpendicular gyroscopes: horizontal (14), longitudinal (15) and transverse (16), the output of the pitch sensor (17) of the horizontal gyroscope is fed to the input of the pitch amplifier (18 ), and the output from this amplifier is fed into the adder of the pitch control system, the output from the roll sensor (19) of the same gyroscope is fed to one of the switch contacts (K 4), direction sensors (20, 21) of the longitudinal and transverse gyroscopes are connected to their inputs directional amplifiers (22, 23 ), and the outputs of these amplifiers are connected to the input of the direction adder (24), the signal from which goes to the switch (K 4), and the vertical sensors (25, 26) of the longitudinal and transverse gyroscopes are connected to the input of the vertical adder (27), the signal which is fed to the control inputs “more” of the mentioned direction amplifiers, and from the common contact “o” of the switch (K 4), the signal goes to the adder of the direction control system, and, in addition, the output from the roll sensor (19) of the horizontal gyroscope is fed to the control input "Bol Chez "pitch amplifier (18) and / or output vertical adder (27) is input to a pitch amplifier (18) (shown in phantom). 13. The device according to claim 8, in which the block of gyroscopic sensors consists of three mutually perpendicular gyroscopes: horizontal (14), longitudinal (15) and transverse (16), the output of the pitch sensor (17) of the horizontal gyroscope is input to the pitch amplifier 18, and the output from this amplifier is fed to the adder of the pitch control system, the output from the roll sensor (19) of the same gyroscope is fed to one of the switch contacts (K 4), the direction sensors (20, 21) of the longitudinal and transverse gyroscopes are connected to the inputs of the direction adder ( 24), and output e of the adder is connected to the input of the directional amplifier (22), the signal from which goes to the switch (K 4), and the vertical sensors (25, 26) of the longitudinal and transverse gyroscopes are connected to the input of the vertical adder (27), the signal from which is fed to the control the inputs are “larger” of the mentioned direction amplifiers, and from the common contact “o” of the switch (K 4), the signal goes to the adder of the direction control system, and, in addition, the output from the roll sensor (19) of the horizontal gyroscope is fed to the control input “more” I pitch (18) and / or the vertical output adder (27) is input to a pitch amplifier (18).

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