RU2248304C2 - Method and device for limitation of angle of attack and overloading of aircraft - Google Patents

Abstract

FIELD: aircraft control units.

SUBSTANCE: proposed device includes correcting unit for limitation of control signal at lead; magnitude of lead is calculated in correcting unit on principle of transfer function. Proposed method consists in introduction of limitation at lead on control signal of elevator control system till moment of attaining maximum angle of attack; magnitude of lead is calculated on principle of transfer function.

EFFECT: enhanced safety of aircraft maneuvering.

3 cl, 10 dwg, 2 tbl

Description

The invention relates to methods and devices for controlling an airplane. The task to which the claimed invention is directed is to increase the safety of maneuvering the aircraft by introducing into the aircraft pitch control system (elevator, rotary jet nozzles of aircraft engines, aircraft engine thrust vector or other steering elements controlling the aircraft pitch or its angle of attack) additional corrective device. (An abbreviated form for describing the control will be used next - elevator instead of elevator, rotary jet nozzles of aircraft engines, thrust vector of aircraft engines, or other steering elements controlling the aircraft pitch or angle of attack.) By adjusting the elevator control signal, it allows the system control quickly and accurately bring the aircraft to the maximum permissible angle of attack (α max) if necessary.

The method of limiting the angle of attack and overload of the aircraft is characterized in that the control signal of the aircraft elevator control system, coming from the control that controls the material object - the elevator of the aircraft, introduces a restriction with anticipation until the aircraft reaches its maximum angle of attack while the lead value is calculated according to the principle of the transfer function — the first-order inertial link, the input signal of which is a value proportional to the angles the speed of the elevator or control element of the aircraft elevator control system, and the outgoing one is the lead, and after removing the lead on the control signal, the latter gets the opportunity to change to a value corresponding to the maximum allowable angle of elevator, which corresponds to the maximum allowable angle of attack at a steady flight mode, in which the angle of attack does not change and is completely balanced, corresponding to the maximum value set by the elevator.

The device for limiting the angle of attack of an aircraft differs in that a correction device is included in the control system of the elevator control of the aircraft, introducing a restriction on the control signal with lead, and the value of lead is calculated in the correction device according to the principle of the transfer function — an inertial link of the first order, the input signal of which - a value proportional to the angular velocity of the elevator or control of the elevator control system of the aircraft, and - the value of pre-emption; and devices that allow the corrective device to change the control signal, for example, summing devices, thrust devices or actuators, which, together with the corrective device, affect the aircraft elevator control system, limiting the control signal, which, in turn, restricts the elevator elevator movement .

In addition, the inventive device is characterized in that it receives a parameter proportional to the angular velocity of the elevator of the aircraft or the angular velocity of the elevator of the aircraft from sensors of these quantities associated with the control or elevator or with any element of the elevator control system aircraft, which is a conductor of the control signal.

List of drawings:

Figure 1: Structural-functional diagram of the calculator of the transfer function of the lead value Δ δ (the lead value Δ δ is the angle of deviation of the elevator Δ δ proportional to Δ α _{2 of the} upcoming inertial deviation of the angle of attack α above a given value α _specified (see Fig. .6)).

1. The summing block (C).

2. Product block (multiplier) N1 (U1).

3. Product block (multiplier) N2 (U2).

4. Block integrating link (I3).

5. Product block (multiplier) N3 (U3).

6. Product block (multiplier) N4 (U4).

7. The calculator of the transfer function (VP).

Figure 2: Structural-functional diagram of the calculator of the value of the lead Δ δ (VVU).

8. Software device (PRU).

9. The calculator gain (VKU).

10. Calculator feedback coefficient (VOS).

11. The calculator of the time interval τ (BB).

12. A differentiating device (DU) can be omitted if the parameter δ ’(t) equal to the angular velocity of the elevator is fed to the input of the VP.

12a) An optional conversion device (NPC) that allows only negative Δ δ values to pass through.

13. The calculator of the lead value (VVU).

Figure 3: Structural-functional diagram of the aircraft elevator control system using a corrective device:

14. The calculator of the maximum deviation of the elevator (VPO).

15. The adder corrective device (SKU).

16. Transforming device (PU), passing only negative values Δ δ _to - correction value.

17. The corrective device (KUS).

18. The governing body (OS) in the system of longitudinal aircraft control (pitch) (steering wheel, aircraft control stick).

19. The position sensor of the control (DOU).

20. The totalizer of the control system (CMS).

21. Actuator (power steering, booster, autonomous steering machine, etc.) (IU).

22. Aircraft elevator (RV).

Figure 4: Diagram of the calculator of the maximum deviation of the elevator δ max.

23. A functional block that reproduces the nonlinear function δ max1 (B1).

24. The calculator of the value of δ max2 (B2).

25. A comparing device (CS) that selects the minimum (maximum modulo) value from the parameters δ max1 and δ max2.

Figure 5: Block diagram of the aircraft elevator control system using a corrective device:

26. The angle of attack sensor or the system of determining the angle of attack (DUA).

27. An airborne signal system or a central line of speed and altitude, or an on-board computer or other system that processes data on altitude and flight speed (SHS).

28. The fuel gauge system and / or signaling devices for the presence of cargo on the plane and / or the setting device for the amount of cargo on the plane (ST).

29. Air pressure receiver (LDPE).

6: Graph of the process of deriving the angle of attack of the aircraft α to the maximum allowable value.

Figure 7: Diagram of the aircraft elevator control system using a corrective device made in one building with a CMS:

30. Corrective device (KUS), made in one case with the WM.

Fig. 8: Diagram of a elevator control system using a corrective device made in one housing with a CMS.

Fig.9: The algorithm of the program for the on-board computer. Program Stages.

31. Obtaining on-board computer (BC) from aircraft systems the parameters of the number M of flight, air pressure at flight altitude P _II , the mass of fuel on the plane m _T , the mass of cargo m _G , flight speed V, angle of attack α, the angular position of the control unit (helm etc.) specifying the position of the elevator δ, the first derivative of the parameter δ, derivatives of the parameter α, signals from the limit switches, the fuel gauge system and other flight parameters.

32. The calculation of the parameters of the gain of the transfer function of the limiter K, its time constant T, the lead time Δ δ, the duration of the time interval τ, the maximum allowable steering angle deviation δ max1 and δ max2 and program synchronization.

33. Comparison of δ max1 and δ max2.

34. The choice of the value of δ max from δ max1 and δ max2.

35. Calculation of δ + Δ δ -δ max.

36. Determining the sign of the quantity δ + Δ δ -δ max and determining the operating mode of the device according to the conditions in table 1. Depending on the mode, the gain of the transfer function K and its time constant T are determined in accordance with table 2.

37. Issue (change) in SLA 20 of the signal Δ δ _to (negative or equal to zero) and the transition to the beginning of the program.

Figure 10: Diagram of a mechanical aircraft control system (elevator).

38. Mechanical actuator (MIU).

39. Limit switch (KB).

40. Thrust device (UU).

41. The stock of the mechanical actuator (ШУ).

42. An emphasis of the persistent device (U).

The composition and interconnections of the elements are as follows.

The inventive device includes a typical set of elements of an aircraft control system from the helm to the elevator and a control signal adjusting device 17 and an adder 20 for remote control wiring connected to the system (Fig. 3). Or a correcting device for the control signal 17 and a mechanical actuating device 38 acting on the thrust device 40 connected to the wiring of the mechanical control system (Fig. 10).

The corrective device 17 (Fig. 3) includes a calculator of the lead value 13, a calculator of the maximum deviation of the elevator 14, an adder of the correcting device 15, the converting device 16. All elements are mutually connected to each other according to Fig. 3. The calculator of the lead amount 13 consists of the following parts: VP 7, switchgear 8, BB 11, VKU 9, BOC 10 and optional elements of the remote control and the control unit (figure 2). If the VP receives the parameter δ-angle of the RV installation from the corresponding position sensors of the elevator or from the position sensor of the steering column (located on any of the control system elements that conduct the control signal δ), then it must be differentiated in ДУ 12 to obtain the derivative δ '(t) , which should be fed to the input of the air conditioner 7. If the air conditioner receives the parameter δ '(t) from the corresponding sensors located on any of the elements of the control system that conducts the control signal δ, then it should be fed to the input of the air conditioner 7 without any changes vany in control. To increase the reliability of the system, it is possible to apply an element of an NPU 12a, which allows only negative values of the parameter Δ δ, the values of lead, to be passed. All elements of the IVU are interconnected in accordance with figure 3.

VP 7 (figure 1) consists of elements C (position 1), U1 (position 2), U2 (position 3), I3 (position 4), U3 (position 5), U4 (position 6). All elements of the VP are mutually connected in accordance with figure 1.

VPO 14 (figure 4) consists of elements B1 (position 23), B2 (position 24), SU (position 25). All elements of the malware are interconnected in accordance with figure 4.

KUS 17 (figure 10), working in conjunction with mechanical wiring, is connected to MIU 38, which has a rod 41 with KB 49 located on it. A stop device 40 is included in the elevator control system, which can interact with MIU according to this description.

KUS 17 (figure 5) in any design can be connected with the DUA 26 to obtain the parameter angle of attack α, with the SHS 27 to obtain flight parameters, with CT 28 to obtain the parameters of the fuel mass and cargo on board the aircraft. SHS 27 is associated with LDPE 29.

The remote control system may have the following elements: OS 18, DOU 19, SUS 20, IU 21, RV 22, interconnected according to Fig.5.

The mechanical control system may have the following elements: OA 18, UU 40, IU 21, RV 22, interconnected according to Fig.10.

Given the urgent requirements of FIPS staff regarding the description of the relationships of elements, I quote the following data. The following is a list of control wiring (system) elements (EPUs) with the pitch (or elevator) of the aircraft in the case of a mechanical aircraft control system (SULA). (An airplane is an aircraft (LA)). The OS can be presented in the form of a helm and / or control column and / or aircraft control stick and / or automatic control unit (AU) (if any) (autopilot).

In the case of using control with a mechanical wiring (system) (Fig. 10), the OS is mechanically connected in series using known connecting elements (SE), such as: bolts, nuts, cotter pins, studs, brackets, eyes, gears, sliders, chain transmissions, helical gears, gears, shafts, rods, etc. (hereinafter (SE)) with EPUs connected in series in any sequence: traction, rocking, shafts, lunettes, brackets, chains, cables, gears, levers, etc. EPU transmit using mechanical displacement (MP) (linear, rotational, translational, etc.) the control signal δ (US) on UU 40. UU can be made in the form of any element (set of elements) from the list related to the above explained the concept of an electronic control unit with a stop mounted on it, or with a lever attached to it with a stop, or a rocking chair with a stop, etc. (The connection of the stop with this element can be by means of SE or welding, soldering, gluing, etc.). The control unit is sequentially mechanically connected by means of SC with the ECU front and rear (at both ends) (i.e., from two sides). EPU, following UU, transfer MP to UI 21 (if it is installed): booster, hydraulic booster, electric drive, electric lift, autonomous steering machine, hydraulic drive, pneumatic drive or their combinations, etc. Further, the DUT is connected mechanically with the help of SE with the following EPCs. In the absence of PS in the SULA, the EPUs are associated with the subsequent EPUs using SE. EPUs are connected in series with each other and the last of them is connected by means of SE with elevator (RV) 22. Thus, all EPUs are sequentially interconnected by means of SE and form a circuit of series-connected elements: ОУ, ЭПУ, УУ, ИУ and / or EPU, RV.

All EPUs are mounted on the aircraft and connected to it by means of KE fasteners such as SE, lunettes, brackets, levers, bearings, bolts, screws, rivets, cotter pins, nuts, etc. They provide one degree of freedom for the MP of these elements, ensuring the passage of the CSS in the BAR. I will add that the pilot (person) acts on the OS, associated with it with the pilot’s hand by clamping it between the fingers. The pilot is connected with the OS only during the flight and the movement of the aircraft on the ground. If the aircraft is unmanned, then the functions of the pilot are performed by an automatic device (AU), mechanically connected to the control unit, replacing the pilot and the control unit in the above scheme and forming a new circuit of series-connected elements: automatic device, control unit, control unit, control unit and / or control unit, electronic control unit.

If the SULA is electric or electronic, then the AU can be connected to the wires (electrically conductive, isolated from each other) of this SULA using the plug connectors of SR.

For an electric or electronic SDULA (ESULA), the op-amp is connected (mechanically with the help of SE and EPU) in series with the DOU, which converts the MP OS into an electric or electronic-digital DC δ in the form of voltage or current frequency, or in the form of electronic-bit digital information.

As DOW, you can use: potentiometric, inductive, induction sensors, microsin, capacitive, pulse sensors that convert the MP to an electrical signal, electromechanical digital devices, analog-to-digital converters and other converters of the MP to electric (E) or electronic-digital (EC) signal, or their analogs and / or combinations thereof. For example, like an F-22. For example, a potentiometric or induction or capacitive sensor connected to an analog-to-digital converter, as a result of which the op amp MP is converted to a digital code, the bit (or digital) value of which is proportional to the op amp MP.

The DOU is connected with the help of SR, to which the output US is supplied from the DOU, with electrically conductive EPUs: electrically conductive insulated metal wires (electric (E) current conductors).

(All ECUs that transmit DCs without conversion, regardless of the type of CSs (mechanical, E, or EC), are referred to in this description as ECUs.) The ECUs next to the DOW are connected by means of SR to a corrective device (ACC) and can be connected to ACC (according to figure 3, 5, 7, 8). EPU on these figures are shown in the form of arrows.

KUS can be made in the form of an on-board computer (BC) (Fig. 7) or its part, or an E or EC block (EB) built on semiconductor, electric lamp elements and / or microcircuits. KUS is connected by means of SR to the next KUS EPU. According to these EPUs, the DC is transmitted to the DUT, with which they are connected with the help of the SR, in which the DC is converted into the MP of the rod of the DUT. In E or EC SULA (or EPU) US is not MP, but E or EC EC (EUS) (electrical or electronic digital control signal). Therefore, the DUT converts the DC not from the type of MP, but from the type of ESA into the MP of the rod or shaft of the DUT. Each DUT has a rod or shaft that performs MP proportional to the value of the DC. For example, the power boosts of the F-22 aircraft. Mechanical devices are connected to the DUT with the help of solar cells connected to the RS with the help of solar cells. All of the above elements are fixed on the plane (have a connection) with the help of FE, SE, clamps, screws, thread-poppies, etc. All EPUs (mechanical, E, or EC) are connected in series with each other from the OS to the RV and mounted on an airplane, forming a chain of elements. However, duplication of the electronic control unit by parallel connection with the same electronic control unit is allowed. The number of parallel elements is not limited. For example, from DOW to KUS and from KUS to DUT it is possible to lay not one electric wire, but 2, 3, 4, 10, 128, 365, etc. (parallel). In E or EC SULA KUS does not have to be connected directly with EPUs coming from DOW. If the CAS 20 is not part of the CSC of Fig. 3, then the ECUs following the DOW are connected with the help of the control system with the CMS and with the help of the CMS with the CSC (or rather, with the CMS of the CCP 15).

Thus, the CMS should have 3 connections through 3 SRs or without SRs: with EPUs following from DOW, with EPUs following to DUT and VVU, with EPUs connecting PU and SLS. In addition, the EPU following from the DOW should have a connection (with the help of the SR) with the ACS, which is part of the control system.

As a control system, you can use magnetic amplifiers, a summing block based on operational amplifiers, electronic amplifiers, potentiometric circuits, polarized relays, bridge circuits, a summing block based on microcircuits or analog elements, arithmetic devices for digital BC, summing blocks made separately from BC, arithmetic devices made separately from the digital BC. For a mechanical SULA with mechanical control units, the control unit has an emphasis which is rigidly connected to the control unit or on a hinge. If the emphasis is on the hinge, then it is made in the form of a rocker on the bracket, due to which, each new position of the EPU will correspond to only one new position of the emphasis. Emphasis UU interacts with MIU, which is not associated with UU. It has a retractable rod with an emphasis, or a rotatable shaft with an emphasis, which, if necessary, can abut against the emphasis UU. MIU is fixed on the plane with the help of FE. It can be an electric motor using a worm, slide, chain or other connection with or without a gearbox, moving a shaft or rod, gear, chain or other device with a stop located on it (her). (For example, the MP-100 mechanism, widely used in aviation.)

All these elements form MIU and are mechanically connected with it (or with its body). MIU can be an electrohydraulic actuator moving any of the above elements (rod, shaft, etc.).

MIU is connected by means of SR with electric wires connected by means of SR to the correction device KUS 17 (Fig. 10). On the rod (shaft, etc.) of the MIU, a KB 39 limit switch connected to the rod of the MIU is installed. When the rod (shaft, etc.) abuts the MIU against the support of the MIU, the HF electric circuit closes. This E circuit is connected with KB and with the KUS E conductors using SR or soldering or other known conductive methods. As a result, any SULA, regardless of type, is a conductor of the DC that goes from the OS to the RC (through elements belonging to the SULA and transmitting the US to each other). The CSS is transferred from one EPU to another EPU and can be distorted (corrected) only in the CMS or in the CU depending on the type of the SULA. In addition, they can be connected mechanically or in any other way to the SLA, or using CE or SE to any of the devices listed above, any devices used in aviation.

For example: spring loading devices simulating aerodynamic loading can be attached to the op-amp or electronic control gear in order to create the pilot's sensation of physical effort at the helm. For example, the mechanisms of the trimmer effect imitating the change in the load on the helm can be attached to the SULA. For example, mechanisms for changing the gear ratios of the SULA can be attached to the electronic control gear, in order to change the gear ratios of the SULA depending on the height and speed of flight. For example, automatic stability enhancing machines that adjust the DC depending on the change in the angle of attack, or dampers that correct the US depending on the angular velocity of rotation of the aircraft, can be attached to the electronic control gear. For example, autopilot devices can be attached to the ECU, correcting or defining the DC depending on the autopilot operating mode (for example, autopilot steering machines, for example, autopilot actuators, for example, summing devices), connected as described above for the control system to and from the ECU to autopilot using insulated conductive wires electrically connected to the adder and autopilot, i.e. transmitting electrical signals from the autopilot to the autopilot adder associated with the electronic control unit. In addition, all EPUs of a mechanical SULA can be made of metal or other durable materials based on any chemical elements listed in the periodic table and / or their combinations in any proportions as desired by the designer. The shape, color, smell, taste and relative position of these elements do not matter.

When transmitting analog electronic signals from devices, sensors, etc. to EC devices, BC, etc. and from EC devices, BC to analog devices, sensors, etc. analog-to-digital (or digital-to-analog) converters (ADCs) are used. ADCs are also used when using electronic computers (computers). A set of EPUs from OS to RV, which also includes all of the above elements (OS, EPU, DOU, SUS, IU, RV, UU, etc.), in this description is called control wiring.

For its operation, the control system can receive all or some of the following parameters: M (Mach number) of the flight, R _p - atmospheric pressure overboard, δ '(t) - derivative of the state equation (angular velocity of the elevator or helm or a signal proportional to it ) or δ (t) —US (the angular position of the elevator or steering wheel or a signal proportional to it), α (t) —the angle of attack, V — flight speed, ρ — air density, m _g — mass of cargo, m _t — mass fuel.

Due to the fact that the device can be designed for a typical flight mode and for a typical load, it can do without most of these parameters. For example, without R _c , M, m _t , m _g , V, ρ, α. The parameter δ '(t) can be obtained by differentiating the parameter δ (t). The KUS parameter δ (t) can be received from the sensors connected with it installed on the plane using the FE or in another way and mechanically connected using the FE or in another way (for example: glued, welded, inserted into special grooves, etc.) with OU, RV or EPU. The parameter CSC can be obtained from the BC, the flight control and navigation system PNK, from the DOE, from the electric wires (EP) going from the DOW to the CMS, CMS or to the CSC. To obtain this parameter, the control switchgear is connected to any of the above sensors (which can be of any type, as will be indicated below or as indicated for DOW sensors, for example: potentiometric, induction, capacitive, pulsed, etc.), elements, devices, PNK , A computer using electric wires through a ballast or without a ballast. If the control panel is a digital computer or BC, and the parameter coming from the aforementioned any element is analog, then the electric wires are connected to this element and the ADC through or without SR. The ADC is connected by electric wires to the control system via or without SR. Thus, the aforementioned element is connected by wires to the ADC, and the ADC is connected by wires to the ACC. For example: a DOU or a potentiometric sensor associated with an op-amp, or an induction sensor connected to a mechanical electronic control unit, is connected by electric wires to the ADC through the ballast, and the ADC through the ballast to the KUS. As sensors of the parameter δ, you can use any of the devices listed in the above list for sensors of the DOW devices. For example: potentiometric, inductive, capacitive, induction sensor, pulse, etc. All or any of the electrical wires, devices and sensors mentioned in this description can be connected to each other without the use of SR using soldering (hot or cold), etc. The control unit receives the parameter δ (or δ (t)) from the PNK, on-board computer, or other things only if these devices have this parameter. To obtain δ (t) (or δ), these devices themselves must be connected by electric wires to the sensors δ (t) located in any of the places already listed. These sensors can be of any of the types listed.

In addition to the parameter δ (t), the control panel can receive the parameter δ ’(t) from the sensors installed on the aircraft and connected by the CE, welded, glued, inserted into the grooves, etc. with OS or EPU or YIU or RV. The parameter sensor δ ’(t) (or δ’) can be connected to the op-amp or electronic control unit or the IU or the RV using rods, rockers, gear, chain, worm, slider, spline, etc. As the sensor δ ’(t), a tachogenerator or other known device can be used to convert the MP to an electrical or electronic-digital signal proportional to the derivative of the MP. In this case, the sensor δ ’(t) is connected by electric wires through the ballast to the control panel. If the KUS is digital, then the δ ’(t) sensor is connected by electric wires to the ADC through the SR, and the ADC by electric wires from the KUS through the SR.

Example: RV shaft or mechanical control unit or steering wheel is connected through a gear, worm, slide gear, using traction, rocking chair, shaft, chain, their combinations, etc. with the shaft of a tachogenerator or other sensor δ ’(t). This sensor is connected by electric wires to the ADC through the SR. ADC through SR is connected with the control system (digital). If the KUS is analog, then the electric wires from the sensor go directly to the KUS.

Any compounds in this description may not use SR. Instead, E signals can be transmitted in any other conductive manner. For example, through soldered wires or along the E conductive tracks on the board or from the legs of the microcircuit soldered into a special board to the E conductive channels forming on these same boards numerous parallel tracks called buses (for example, PCI or ISA or AGP bus) on the boards special computing devices, called computers abroad (in developed countries).

In all the relationships indicated in the description, the use of intermediate signal amplifiers is not excluded. This rule applies to the entire description. In the case of their use, amplifiers should be connected to electric wires. In electric SLMS there is another possibility of obtaining the parameter δ (t) or δ ’(t) by the control system, since the control system can be connected with electric control units. EPUs are a conductor of CSS. CSS is δ (t) (or proportional to δ (t)).

If the control system is connected to the control unit through the control system, then it can receive δ (t) directly from the control panel or the remote control, since the control panel or remote control are conductors of the parameter δ (t). For this, the control switchgear must be connected through the SR with the help of electric wires to the electronic control unit or DOW and use it as a sensor.

If the control system is connected directly to the control system, and not through the control system, then it will receive the parameter δ (t) from the DOW (Fig. 7). If the control panel receives δ (t) from a sensor or device or electronic control unit, then it can differentiate this parameter within itself to obtain δ ’(t). In this case, obtaining the parameter δ ’(t) from other sensors or devices is not necessary.

If necessary, the control and control system can receive the parameters M, Rn, V, ρ from the airborne signal system (for example, SHS) or from the speed and height central line of the type of central air conditioner, etc. In this case, the control system must be connected to these devices (systems) using electrical wires using (or without) SR and / or ADC, if necessary. Aneroid and / or gauge devices, with which these devices are connected using tubes with choke connectors (SRA), are used as sensitive elements of a central heating circuit or internal combustion system. (for inquiries, see "Handbook of an Aviation Engineer" V. G. Alexandrov, editor Lazarevich, "Transport" Publishing House, 1973, p. 246).

It is possible that these parameters are obtained for the control system and directly from the LDPE air pressure receivers, aneroid devices, on-board PNK, BC, on-board computers, etc., with which the control panel must be connected in this case with electric wires and / or tubes using SR and / or choke ball joints respectively. If the KUS is a digital computer or an analog device, then these signals must be converted to a digital code or an analog signal using the same devices listed above for the DOW and / or ADC, which can be connected to aneroid or pressure gauges and / or their combinations and / or sensors of the above parameters.

For example, an air pressure receiver (LDPE) is connected by a tube to an aneroid membrane connected mechanically to a potentiometric (or inductive, or electromechanical digital device or SHS) device, which is connected by wires through an SR to an analog-to-digital converter of the ADC, which is connected by electric wires to a control system through an SR . Other combinations are possible.

If the control system receives signals not directly from the SHS, DAC, PNC or computer, but directly, then these signals can be converted to the control system according to the formulas known from the last century into the values of pH, M, etc.

If the ECCUS receives these parameters in an electronic digital form, then it is not necessary to convert them through the ADC. In this case, the digital control systems are connected by electric wires through the SR to the SHS or CVS, PNK, computers or so on.

The parameter CSC can be obtained from the angle of attack sensor (CAS), with which the CSC must be connected in this case using electric wires through the ballast (or without). If the parameter α is generated in the DUA in analog form, and the control system is a digital computer, then in this case the electric wires from the DUA must be connected to the ADC (via or without ballast), and the ADC is connected to the ADC by wires (through the ballast or without).

It is possible to obtain α from PNK, on-board digital computer (BCM) and other devices with which in this case the control panel must be connected by electric wires through or without SR. The parameter m _{t of the} control system is received from the fuel gauge system (for example, SECS), with which the control system is connected by electric wires through the SR. If the control system is digital, then the wires from the fuel gauge system are connected to the analog-to-digital converter, which is connected to the control panel (see "Aircraft Engineer's Handbook") also by electric wires through the SR. The parameter m _g can be entered manually by members of the flight crew before flying when installing cargoes, special cargoes on special fasteners or without them on the plane.

For manual entry, the aircraft is equipped with special master devices, into which information about the mass of the load is entered either by a switch (can be wired), or using special buttons or from the keyboard of computer devices.

To automatically enter the parameter m _g on the aircraft, limit switches are installed in the cargo attachment points, closing their circuits when special loads are installed (for example, limit switches in the attachment points on the Mriya aircraft for transporting a special container). The control panel is connected by electric wires (through or without ballasts) to these devices through which information about the load is sent to the control panel. If necessary, these devices can be connected with the ADC, and the ADC with KUS electric wires (with or without SR). To generate a signal proportional to the mass of the cargo, these devices can have inside torque sensors, potentiometers, capacitive sensors, inductive, any sensors listed in this description that are configured to generate signals proportional to the mass of the cargo when the aircraft is on the ground and devices for storing and storing this information. For example, a biscuit switch installed in the cockpit can close circuits with specially configured potentiometers or transformers installed in them, from which a voltage proportional to the mass of the load, which is written on the outside of the switch, will be removed. For example, 1 ton, 2 tons, 10 tons, etc. (which will correspond to the output voltage of the switch 1 volt, 2 volts, 10 volts, etc.). To generate a signal proportional to the mass of the cargo, these devices can simply close the electrical circuits associated with the control panel, the appearance of the signal in which the control panel should be regarded as the appearance of the corresponding cargo on board.

KUS can be performed in the form of an analog, digital computer, calculating and solving devices based on bridge circuits, rotating transformers, operational computers and other elements for KUS (ESC). The concept of ESC includes all of these elements. KUS can be divided into the following parts:

1. Calculator δ _max - the maximum deviation of the elevator (steering wheel, control wiring, OS, etc.) VPO 14.

2. The adder KUS (SKU 15).

3. The calculator Δ δ-values of lead (VVU) 13.

4. The converting device (PU) 16.

5. The composition of the KUS can include SU C 20, but you can make it separately from the KUS.

VPO and VVU are connected with SKU. SKU is also connected with the electronic control unit (with control wiring) (for receiving the signal δ) and with the control unit. PU can be associated with VVU and WM.

VPO 14 can be divided into:

- the functional unit δ _max1 (B1), reproducing the nonlinear function δ depending on the number M in accordance with formula 25. In this formula, all parameters depend on the number M of the flight. A signal proportional to M (hereinafter in the description the word proportional will be omitted), this unit takes from the control panel, of which it is a part, and is connected to it by wires that transmit the parameter M (in the form of an E or EC signal);

- calculator δ _max2 (B2) 24, connected to the _{control panel} (of which it is a part) with wires and through which it receives signals M, P _n , m _g and m _t . A computer assembled on the basis of ESCs (for example, special amplifiers, bridge circuits, function blocks, etc.) performs calculations using the formulas (26, 27, 28);

- a comparator (CS) 25, connected to the calculator δ _max1 and the calculator δ _max2 wires and comparing the parameters δ _max1 , δ _max2 received from them. The value closest to zero is output from the CS comparing device to the SKU 25 connected to it. If B2 is absent in the KUS (calculator δ _max2 ), then the calculator δ _max1 (B1) is directly connected to the SKU, and the CS comparing device is not used.

SKU can be assembled on the basis of ESC (for example, operational amplifiers, bridge circuits and connected by wires with VPO (with SU comparing device) or with B1 (calculator δ _max1 ), with calculator Δ δ (WWU 13), with KUS, part of which it is and from which it receives the signal δ, with the input of the PU 16 (converting device), which passes signals of only one sign through itself.

The conversion device PU 16 can be assembled on the basis of diode elements and / or operational amplifiers and other ESCs and is connected with the CMS.

The calculator Δ δ (VVU) is connected to the control system, control panel, to the control panel (of which it is a part) by wires and receives from them all or some of the following signals: δ '(t) (or δ, which is differentiated to obtain δ' (t) inside ДУ 12), М, Р _H , m _T , m _Г , α. It is also connected to the output of the conversion device PU wire, through which it receives an output signal from it.

Especially for FIPS staff, I recall that both the Δ δ calculator (VVU) and the δ _max calculator (VPO) are connected by wires to the devices, sensors, instruments, elements specified in this description, from which they receive the parameters δ, δ ', M , Р _H , m _Г , m _T , α (this is in case FIPS employees do not want to recognize the fact that the calculator Δ δ (VVU) and the calculator δ _max (VPO) are part of the control system, and the control system receives all of these parameters from the sensors specified in this description, control wiring elements, devices, etc. And, therefore, calculate these and receive all these parameters since they relate to elements listed in the description, sensors, devices (wire) or other known producers of the art.

The calculator Δ δ (VVU) can be connected with the ДУА, with the air signal system using the wires in the aforementioned way, by which it receives the parameters α, Р _H , М, V, ρ, as well as with the aircraft fuel system, from which it receives the parameter m _T .

The calculator δ _{max is} also connected to the same systems as the calculator Δ δ. The calculator Δ δ can be connected by wires to the elevator position sensor, which is mechanically (that is, by any method listed in this description) connected to the elevator or to any ECU or op-amp or to the DOU itself. From them he receives a parameter proportional to δ. It can be connected with a rudder angular velocity sensor located in the same place and connected in the same way as described above for the RV position sensor, but outputting not the parameter δ, but the parameter δ '(t).

The calculator of the amount of lead of the VVU can conditionally be divided into the following parts:

- calculator gain VKU 9,

- calculator feedback coefficient BOC 10,

- calculator of the transfer function of the VP 7,

- software device PRU 8,

- if necessary, to obtain a signal δ ’(t) from δ (t), a differentiating remote control device can be used.

In addition, the combination of a differentiating and inertial link of the first order results in a boost link. This fact makes it possible to use the signal δ instead of the input signal δ ’(t), and instead of the calculator Δ δ, not a first-order inertial link, but a boosting link. Entrepreneurial Russian aircraft manufacturers, this fact will probably avoid licensing fees (unfortunately).

- time calculator τ (BB) (optional) 11,

- an optional conversion device 12a,

- ДУ 12 is connected with the control panel, of which it is a part, with electric wires and receives parameter δ from it. Converting it to δ ’(t), the remote control transmits it by wire to the VP, with which the remote control is connected by wires. If the KUS receives δ ’(t), then the VP is directly connected to the KUS, of which it is a part, by wires through which the parameter δ’ (t) is received from it, and the remote control is not applied. In this description, the word wire means not only a metal wire with external insulation, but also any conductive element (for example, a conductive path etched on a circuit board, sprayed, glued, otherwise applied, etc.).

- VKU is assembled on the basis of ESC and connected with the control system, of which it is a part, with wires through which it can receive the parameters M, R _H , m _G. VKU makes calculations in accordance with formulas (23.2-22). VKU may not perform calculations according to formulas (23.2-22) if the device is designed for a typical flight mode. In this case, the gain of the VCU (K _us ) can be calculated at the design stage and set constant. In addition, the VCU can be made calculating the gain depending on only some parameters, for example, only on P _H and / or M. The remaining parameters can be set constant as for a typical flight mode. The VKU gives a signal proportional to the gain coefficient K _us in the VP, with which it is connected by wires or other electrically conductive method specified in the description. (For example, a voltage proportional to K _us ).

The switchgear is assembled on the basis of ESC and connected to the control circuit, of which it is a part, by electrically conductive wires, through which, if necessary, it can receive the parameter α (E or EC signal proportional to the angle of attack of the aircraft α) and differentiate it repeatedly to obtain the parameters α, α '(t), α' '(t) (E or EC signals proportional to α, α' (t), α '' (t)) using its internal differentiating devices with which it is connected by electrically conductive wires. Obtaining the parameter α is optional. The switchgear can receive its signal from the output of the switchgear 16, with which it is connected by wires (the same conversion device that transmits a signal of only one sign). This signal is used by the software device to generate the signal for switching on the operating mode of the VP No. 1. (According to table 1). For example, the signal from the converting device generates a signal to turn off the first limiting mode in the VI). To process this signal, you can use ESCs, relays, polarized relays, relay amplifiers with or without self-locking, etc.). In general, the connection of the VVU elements is as follows: remote control, control gear, VV, VKU, VOS are connected to the KUS, of which they are a part, and receive from it all or some of the parameters they need, received by the KUS from the associated devices, sensors, elements, systems etc. specified in this description. Communication is carried out by any known method of transmitting E or EC signals. For example, on E wires. DU is associated with S. C is associated with U1. U1 is connected with U2, U2 is connected with I3. I3 is associated with U4 and with SKU or with NPU. U4 is associated with U3. U3 is connected in a feedback type with C. In addition, BB is associated with a switchgear, a switchgear is connected with U1. VKU is connected with U2. VOS is connected with U4. PRU also has a connection with U3. For a mechanical SULA, the switchgear receives a signal from KB 39, with which it is connected by electric wires (KB is located on the stem SHU 41, which is part of the MIU). Upon receipt of this signal, the switchgear transmits to the VI, with which it is connected by an electric wire, the signal for activating mode N1 according to the description tables 1 and 2. The switchgear withstands this mode (and / or signal) for a time τ. Formula 8 for calculating τ is presented below.

VV is assembled on the basis of ESC and is connected with KUS, of which it is a part, e. wires through which it can receive the parameters M, P _H , m _G and all the parameters listed in this description. BB makes calculations in accordance with formulas (8.4-22).

The explosive may not perform calculations according to formulas (8.4-22) in the event that the device is designed for a typical flight mode. In this case, the parameter τ can be calculated at the design stage and set constant. In addition, an explosive can be made to calculate the parameter τ depending only on certain parameters, for example, only on P _H and / or M. The remaining parameters can be set constant as for a typical flight mode. The explosive generates a time delay value for issuing a control signal from the switchgear to the air conditioner. The explosive is connected to the switchgear with wires or other electrically conductive method specified in the description. For example, it is possible to output a voltage proportional to τ from the explosive in the switchgear via the electric wires by which the explosive is connected to the switchgear.

The duration τ is determined by the explosive (calculator τ), with which the switchgear is connected by a wire, τ can be calculated by the on-board computer using formulas (8.4-22). Communication between the BC and any element of the VVU is possible, in addition, the VVU and the KUS can themselves be implemented as an on-board computer or a program for it. Perhaps the use of BC instead of any (any) element (s) KUS. For example, BC can be used instead of explosives and control gear, and both the value of τ and the time delay itself can be formed in the BC, which will give signals to the air conditioner (U1 and U3) to switch the operating mode of the control switchgear (air conditioner) from the first (No. 1) in the second (No. 2) in the form of amplification factors K1 and K2, respectively (in U1 K1, and in U3 K2) after time τ after the start of operation of mode No. 1. The gain K1 and K2 can be transmitted in the form of E or EC signals, voltages, and any other known FIPS employees by the electrically conductive method to the AM from a BC or a switchgear via ADC (in this description, digital-to-analog converters, like analog-to-digital converters, are designated as ADCs) or without ADCs depending on the implementation method on the wires with which they are connected (specifically connected).

τ can be calculated in advance for a typical application and stored in a switchgear, and the use of a τ calculator is not required. After the time τ has elapsed after the start of operation of mode number 1, according to the description tables, the switchgear issues a signal to the associated wires of the VP to switch the mode from 1 to mode 2 (and / or turns on mode No. 2). To form a time delay τ in the switchgear, one can use ESCs, polarized relays, relay amplifiers, timer devices, and other well-known elements and devices (up to electric motors with cams and a contactor or limit switch, potentiometric devices, timer devices, etc.).

To generate a signal for switching mode 1 to mode 2, instead of τ, you can use signal α and its derivatives (according to the table). The duration of the N2 mode does not play a big role, and it can be assigned equal to 1 second. For its development, you can still use all of the above elements. The duration of mode 2 of the switchgear can be determined not by time, but all by the same parameter α. To enable mode 2, the simultaneous existence of the conditions α ’≤ 0 and α’ ’<0. The disappearance of any of these conditions serves as a signal to turn off mode 2.

For example, when a signal is received from the control panel, the control panel issues a signal to turn on mode 1. After a time τ, it gives a signal to the VI to turn on mode 2. After another second, turn off mode No. 2 (and switch to working mode) according to tables 1 and 2 and 6.

For example, when a signal is received from a converting device, the control gear generates a signal to turn on mode No. 1 when receiving a signal α, the first and second derivatives of which are less and / or equal to 0 (α '≤ 0 and α' '<0), mode No. 2 is turned on . (In this case, the corresponding signal from the switchgear goes through the wires to the VP connected with the switchgear wires). When α ’’ changes its sign, the PRU will turn off mode No. 2 (since the condition α ’≤ 0 and α’ ’<0 ceases to exist). Mode No. 2 of the control gear can include only from mode No. 1.

When mode 2 is turned off, the VP switches to operating mode. Mode No. 3, described in the table, is not provided for consideration by the FIPS employee and is designed only for the desire of specific manufacturers. Mode number 3 can not be used. The software device may not be used. The switchgear can be made in the form of a simple circuit with an electric motor, in which, upon receipt of a signal from a KB or a conversion device, the relay controlling the power supply of this electric motor (ED) becomes self-locking and supplies power to the ED. It rotates a shaft connected through a gearbox to a cam mechanism from a set of cams. The cam mechanism has a shaft with cams located on it, as well as limit switches that these cams act on. Cams during rotation of the shaft rotated by the motor through the gearbox (these are all connected elements) act on the corresponding limit switches, which, closing and / or opening the circuits of the corresponding circuits connected with the VI, provide the corresponding signals for switching on or off the VI modes, and It is also possible to remove the self-locking relay.

The time calculator τ (BB) is connected to the control panel (of which it is a part) with wires and can receive signals M, P _H , m _T , m _{G from it} . He performs calculations according to the formula (8). It is assembled using an ESC and connected to the switchgear by wires, through which it transfers the parameter τ to the switchgear. EXPLOSIVES can not be used, and the parameter τ can be calculated for a typical flight mode at the design stage and introduced into the control gear.

In this description, the switchgear can be considered as a case to which it is connected (not electrically, that is, the wires should not create a “short circuit” between each other on the switchgear case), all of the above wires through the ballast with which it is connected, and from these ballasts there are wires to all consumers (elements) with which they (some of these wires) are connected inside the control panel.

The VVU also includes a VOS feedback gain calculator. The VOS is assembled on the basis of the ESC and is connected to the control panel, of which it is a part, with wires, through which it can receive parameters m _T , M, P _n , m _G , VOS calculates according to the formulas: (24.3-22), VOS may not perform calculations using formulas (24.3-22) if the calculation is for a typical flight mode. In this case, the gain of the OSI can be calculated at the design stage and set constant and embedded in the OSI or VP (U2). In addition, BOC can be made by calculating a coefficient depending on some parameters, for example, only P _H and / or M. The remaining parameters can be set constant as for a typical flight mode. BOC produces a signal proportional to the feedback coefficient for the transfer function of the calculator Δ δ in the VP with which it is connected, by wires. (For example, a voltage proportional to K _oc ).

The remote control (which may be part of the VVU) is a differentiating unit that differentiates the input signal δ and transmits the signal δ (t) resulting from the differentiation to the VI. The need for the use and communication of remote controls were discussed above. The remote control is assembled on the basis of ESC.

VP is assembled on the basis of ESC and connected to the control system, of which it is a part, by wires through which it can receive the following parameters: δ (t), signals of the software device K1 and K2, parameters of VKU and VOS, VP can be connected with remote control, control gear , KUS, VKU, VOS, SKU, NPU with sensors and devices described in this description. For example, with a sensor δ (t). VP performs the role of a device that processes the input signal δ '(t) (or δ) according to the law of the transfer function — a first-order inertial link with a variable gain of a link and a time constant T. It is known from the basics of automation that these parameters can be changed by changing gain factors K _yc and feedback gain K _oc . VP can be performed on the basis of operational amplifiers, electronic computing units, analog circuits, ESC, etc. The VP can consist of the following interconnected elements: a summing block (C) connected to the remote control or the control panel and receiving from it the parameter δ '(t). This block is connected with the block of product No. 1 (U1) (multiplier), which is connected with the multiplier No. 2 (block of product), (U2), which is connected with the integrating unit block (IZ) with the integrating block, which is connected with the output of the VP ( it is the way out of the VP). The exit from the airspace is associated with SKU (if you do not apply NPU). In addition, the exit from the integrating link is connected by feedback type with the multiplier No. 4 (U4) (product block), which is connected with the multiplier No. 3 (U3) (product block), which is connected by the type of feedback with the summing VP unit (C )

Multiplier No. 1 (U1) is associated with a switchgear.

Multiplier No. 3 (U3) is associated with a switchgear.

Multiplier No. 2 (U2) is connected with VKU.

Multiplier No. 4 (U4) is associated with VOS.

All the connections listed in this description, the VP can be made in the form of E wires, in the form of a separate wiring diagram and other methods known to FIPS staff. All blocks and / or elements listed in this description for KUS, VP, PRU, etc., can be performed in the form of one or more related blocks. For example, in the form of a single block, you can perform KUS, VP, multiplier No. 1 with multiplier No. 2, multiplier No. 3 with multiplier No. 4. In C (summing block VP), the feedback signal is subtracted from multipliers No. 3 and 4 from the signal δ '(t). The resultant signal goes to the multipliers No. 1 and 2, after which the resulting signal is integrated in the integrating link (integrating unit) (FROM). In the multipliers, the input signals are multiplied by the signals coming from the associated PRU, VKU and VOS. From VKU a signal is proportional to the gain K _us (formula 23). From BOC receives a signal proportional to K _a (Formula 24). Signals from the switchgear are proportional to the numbers shown in table 2, depending on the operating mode of the software device. Instead of multipliers, it is possible to use relays breaking and / or switching electric circuits inside the VP together with signal amplifiers.

KUS and / or all the elements indicated in this description (except mechanical elements) can be performed using a digital computing device based on integrated circuits and / or otherwise, and assembled in a single unit (or several interconnected units). This block (blocks) in order to obtain all of the parameters listed above in this description from the above sensors, devices and systems must be connected to all of them using the above elements (for example, E wires with SR). In addition, if the signals from sensors, systems, devices do not come in digital, but in analog form, then they must be converted to digital form using analog-to-digital converters, which must be connected with sensors, systems, devices and with a digital computing unit (in blocks) using wires. In addition, the digital unit must be connected either to the DUT, or to the MIU, or to the CMS. All the elements indicated in the description can be performed as program elements for the on-board computer. For this, the on-board computer must receive from the sensors, systems, devices listed above, all and / or some of the parameters listed in this description. For example: M, P _H , m _T , m _G , α, δ, δ '(t), etc.

To convert to digital form, sensors, systems, devices must be connected by wires to the ADC, and those, in turn, with the on-board computer. In the memory of the on-board computer there should be a program that performs calculations according to formulas (1-30) and uses data in the computer's memory according to the aerodynamic coefficients used in these formulas. (These data can be entered at the aircraft production stage). As a result of the calculations, the computer must either give a signal proportional to the corrected position of the elevator δ _k on the DUT according to formula 30, or in the MIU (in the case of using a mechanical BUL) δ _{max mia} according to the formula: δ _{max mia} = δ _max -Δ δ (where Δ δ is the lead value calculated in the VVU, and δ _max is the maximum permissible position of the RV calculated in VPO 14), or in the CMS (in case the CMS is part of the BMS and is not part of the BC or CCP) Δ δ _k according to formula 29. Moreover, Δ δ _k cannot be positive. If Δ δ _k > 0, then the BC or KUS resets the value of the parameter Δ δ _k and gives the value Δ δ _k = 0.

A digital computing unit (if used instead of a BC) must be connected according to the same principles with the same sensors, systems, devices, actuators, summing devices, MIU, according to the same formulas and perform calculations using the same formulas and algorithms as and for BK or KUS. The program algorithm for the CD is shown in Fig.9 and is described in more detail in this description. The control system can be performed in the same case as the control system, then the general connection scheme of the control system and the control system will look like this: the control system is connected to the control system, which is connected to the control units, which are connected in series with each other and to the control system and control system, which is again connected to the control system, which sequentially connected with each other and with the DUT, which is connected with the help of solar cells with RV. In this case, the control system is connected with the sensors, devices and systems listed in this description, and can receive from them the parameters M, P _H , m _T , m _G , α, δ '(t).

A description of the behavior of the aircraft in the process of reaching the maximum angle of attack is as follows.

The device operates in such a way that the process of bringing the aircraft to the maximum allowable angle of attack will look like in Fig.6. At time t1, the pilot begins to move the helm to increase the angle of attack (α _specified ), while the two-stage limiter calculates the upcoming throw of the angle of attack, and if the magnitude of this throw is such that the angle of attack α can exceed the limit value (α _max ), then the limiter introduces the first step of limiting the position of the elevator (at time t2 of FIG. 6), and the aircraft continues to increase the angle of attack α by inertia. At time t3, the angle of attack α comes to the upper point of the oscillatory process. Thanks to the limiter, it coincides with the maximum allowable value ((α _max ). The angle of attack stops growing and the first stage of restriction is turned off. The elevator gains freedom of movement to the second stage of limitation. In this case, the elevator at the second stage of limitation will specify exactly that angle of attack ( α _set ), which the plane has already reached (with a slight error) _.This angle of attack is equal to the maximum allowable value (α _max ).

At time t4, the growth rate of the angle of attack α '(t) is zero, and the angle of attack α is approximately equal to the specified value (α _given ). Consequently, there is no inertia and unbalanced moments of forces that could bring the plane out of this state. There is a steady balance in the angle of attack. Thus, the aircraft reaches the maximum allowable angle of attack accurately, quickly and without subsequent oscillations (Fig.6).

The scope is as follows.

This invention can be practically applied as an additional device to the control system on airplanes and unmanned aerial vehicles. On aircraft with mechanical control wiring, electrohydromechanical or electromechanical elevator limiters can be used (FIG. 10). On airplanes with remote control wiring, electronic elevator position limiters can be used (FIGS. 5, 7). The claimed invention can be applied on airplanes having, in addition to the elevator, also a controlled stabilizer. In this case, the input parameters of the device must be converted according to the conversion formulas. You can apply this invention in the form of a signal to the pilot about approaching the maximum angle of attack.

The operation and principle of operation of the elements of the device are as follows.

The main element of the device is a calculator of the lead value Δ δ VVU. The main mode of operation of the VVU is the “P” mode (working). In this mode, the VVU calculates the lead Δ δ in accordance with the transfer function W (P) and with formulas (1-22).

- передаточная функция вычислителя Δ δ (1)

- transfer function of the calculator Δ δ (1)

- постоянная времени для функции W(P) (2)Where

is the time constant for the function W (P) (2)

K = -R · A / q is the gain for the function W (P) (K> 0) (3)

RA _mz $\genfrac{}{}{0ex}{}{\text{ω}}{}$ ^z -A _y $\genfrac{}{}{0ex}{}{\text{α}}{}$ - aerodynamic coefficient (4)

q = A _mz $\genfrac{}{}{0ex}{}{\text{α}}{}$ -A _y $\genfrac{}{}{0ex}{}{\text{α}}{}$ A _mz $\genfrac{}{}{0ex}{}{\text{ω}}{}$ ^z - aerodynamic coefficient (5)

(6, 7, 8)

(6, 7, 8)

where arctan (-L / N) ∈] π / 2 ... π [; (9)

A = [cos (L · τ) + (N / Lq / (R · L)) · sin (L · τ)] · e ^{-N · τ} ; (10)

A _mz $\genfrac{}{}{0ex}{}{\text{δ}}{}$ = m _z $\genfrac{}{}{0ex}{}{\text{δ}}{}$ Ρ · V ² · S · b _a / (2 · J _z ) - aerodynamic coefficient (11)

A _mz $\genfrac{}{}{0ex}{}{\text{ω}}{}$ ^z -m _z $\genfrac{}{}{0ex}{}{\text{ω}}{}$ ^z · ρ · V ² · S · b _a / (2 · J _z ) - aerodynamic

coefficient (12)

A _mz $\genfrac{}{}{0ex}{}{\text{α}}{}$ -m _z $\genfrac{}{}{0ex}{}{\text{α}}{}$ Ρ · V ² · S · b _a / (2 · J _z ) - aerodynamic

coefficient (13)

A _y $\genfrac{}{}{0ex}{}{\text{α}}{}$ = -C _y $\genfrac{}{}{0ex}{}{\text{α}}{}$ Ρ · V ² · S / (2 · m · V) - aerodynamic coefficient (14)

To calculate the aerodynamic coefficients And _mz $\genfrac{}{}{0ex}{}{\text{δ}}{}$ , A _mz $\genfrac{}{}{0ex}{}{\text{ω}}{}$ ^z , A _mz $\genfrac{}{}{0ex}{}{\text{α}}{}$ , A _y $\genfrac{}{}{0ex}{}{\text{α}}{}$ You can use the following formulas:

A _mz $\genfrac{}{}{0ex}{}{\text{δ}}{}$ = m _z $\genfrac{}{}{0ex}{}{\text{δ}}{}$ 0.7 · P _N · M ² · S · b _a / J _z ; (fifteen)

A _mz $\genfrac{}{}{0ex}{}{\text{ω}}{}$ ^z = -m _z $\genfrac{}{}{0ex}{}{\text{ω}}{}$ ^z · 0.7 · P _H · M ² · S · b _a / J _z ; (16)

A _mz $\genfrac{}{}{0ex}{}{\text{α}}{}$ = -m _z $\genfrac{}{}{0ex}{}{\text{α}}{}$ 0.7 · P _H · M ² · S · b _a / J _z ; (17)

A _y $\genfrac{}{}{0ex}{}{\text{α}}{}$ = -C _y $\genfrac{}{}{0ex}{}{\text{α}}{}$ 0.7 · P _H · M ² · S / (m · V); (18)

A _y $\genfrac{}{}{0ex}{}{\text{α}}{}$ = -C _y $\genfrac{}{}{0ex}{}{\text{α}}{}$ 0.7 · P _H · M · S / (m · a); (19)

a = 286 + P _H · 5.4 · 10 ^-1 ; (22)

The following parameters are indicated in these formulas:

- W (P) - designation of the transfer function of the calculator.

- The parameter Δ δ is the amount of lead. It corresponds to the inertial deviation of the angle of attack Δ α _casting (above the rudder angle of attack angle α _ass ) when the elevator stops at the first stage of restriction. (In FIG. 6, the _throw Δ α is equal to the distance between the given angle of attack α _ass and the actual angle of attack α at time t3). The parameter Δ δ is the output parameter of the transfer function W (P).

- The parameter δ ’(t) is the angular velocity of the elevator. This is the input parameter of the transfer function W (P).

- The parameter <τ> for formulas 10 and 8 is the duration of the time interval from time t2 to time t3 (theoretical).

- m is the mass of the aircraft, fuel and load,

- p _n - atmospheric pressure at altitude,

- A _mz $\genfrac{}{}{0ex}{}{\text{ω}}{}$ ^z , A _mz $\genfrac{}{}{0ex}{}{\text{α}}{}$ , A _y $\genfrac{}{}{0ex}{}{\text{α}}{}$ , A _mz $\genfrac{}{}{0ex}{}{\text{δ}}{}$ - aerodynamic coefficients,

- С _уα is the derivative of the lift coefficient with respect to the angle of attack (known aerodynamic coefficient),

(известный аэродинамический коэффициент),- m _z $\genfrac{}{}{0ex}{}{\text{ω}}{}$ ^z is the derivative of the longitudinal moment coefficient of the aircraft with respect to the angular velocity of the aircraft

(known aerodynamic coefficient)

- m _z $\genfrac{}{}{0ex}{}{\text{α}}{}$ - the derivative of the longitudinal moment coefficient of the aircraft with respect to the angle of attack of the aircraft α (known aerodynamic coefficient),

- m _z $\genfrac{}{}{0ex}{}{\text{δ}}{}$ - the derivative of the coefficient of the longitudinal moment of the plane with respect to the elevator angle of the aircraft height δ (known aerodynamic coefficient),

- ρ is the air density at the height of the aircraft,

- b _a - the length of the average aerodynamic chord of the aircraft,

- J _z is the longitudinal moment of inertia of the aircraft. It can be approximated as a function of the mass of fuel and cargo: J _z = J _z0 + K _g · m _g · K _t · m _t , where K _g and K _t are proportionality coefficients. You can use more complex formulas.

- M is the Mach number of the flight,

- S is the wing area of the aircraft,

- a is the speed of sound at altitude,

- V - flight speed.

The transmitter of the transfer function (position 7) (VP) is constructed so that its elements form a device that converts its input signal δ ’(t) to Δ δ according to the mathematical law corresponding to the transfer function - the inertial link of the first order. A signal proportional to δ ‘(t) is supplied to the input of the calculator Δ δ, which can be obtained either from the remote control 12 or from special sensors that can be connected to any element of the elevator control system (wiring), including a radio relay or an op-amp.

One of the many ways to implement the inertial link of the first order is the circuit depicted in figure 1. At the same time, in the summing block (C) (Fig. 1, position 1), the signal from the multiplier No. 3 (US) (Fig. 1, position 5) is subtracted from the signal δ '(t). The signal converted in the VP adder (position 1) enters the multiplier No. 1 (position 2), where it is multiplied by the signal K1 from the switchgear (position 8). The signal converted in the multiplier No. 1 enters the multiplier No. 2 (position 3), where it is multiplied by the signal K _must from VKU (position 9). The signal converted in the multiplier No. 2 enters the integrating unit block (IZ) (integrating unit) (position 4), where it is integrated over time. The signal transformed in the integrating unit enters the output of the VP and into the multiplier No. 4 (position 6), where it is multiplied by the signal K _os from the BOC (position 10). The signal converted in the multiplier No. 4 enters the multiplier No. 3 (position 5), where it is multiplied by the signal K2 from the switchgear. The signal converted in the multiplier No. 3 enters the summing block of the VP, where it is subtracted from the input signal δ '(t).

The switchgear generates K1 signals equal to (0; 1/3; 1) for multiplier No. 1 and signals K2 equal to (1; 100) for multiplier No. 3 in accordance with tables 1 and 2, depending on the operating mode of the device. The operating mode of the switchgear is determined in accordance with tables 1 and 2.

To calculate the time τ, you can use EXPLOSIVES (position 11), performing calculations according to formulas 6-9. Depending on the result of these calculations, the explosive gives control signals to the switchgear for switching modes.

VKU (position 9) performs the calculations according to formula 23 and generates a signal To _us (gain) in the multiplier No. 2.

K _us = K / T (23)

BOC (position 10) performs the calculations according to the formula 24 and gives to the multiplier No. 4 a signal To _OS (feedback coefficient).

K _OS = 1 / K (24)

From the VP (calculator Δ δ), the signal enters the control system (Fig. 3, position 15), where this signal is added to the signal δ proportional to the angular position of the PB or the position of the op-amp, and the signal δ _max received from the calculator is subtracted from the result δ _max (position 14). The calculator δ _max produces a signal δ _max as follows: a function block that reproduces the function δ _max1 -f (M) receives a signal proportional to the Mach number of flight M from device 27 (system). Depending on the magnitude of this signal, the functional unit 23 generates a value of δ _max1 and outputs the value to the comparator 25. For each type of aircraft there is a different relationship between the flight number M and the maximum angle of attack α _max , aerodynamic coefficients m _z0 , m _zα , m _zδ . For each type of aircraft the value α _mо is _known - the angle of attack of the zero pitch moment.

The value of δ _max1 can be determined using formula (25).

This formula should be used to determine the function δ _max1 = f (M). It is this function that function block 23 should play. The calculator of the value δ _max2 (position 24) receives signals proportional to M, air pressure P _H , fuel mass m _T , cargo mass m _G from the devices with which it is connected. He makes calculations according to the formulas (26, 27, 28),

m = m _C + m _T + m _GH ; (28)

_F (относительное расстояние аэродинамического фокуса от передней кромки средней аэродинамической хорды крыла вдоль оси Х самолета) зависит от числа М полета.- where is the parameter

_F (relative distance of the aerodynamic focus from the leading edge of the middle aerodynamic chord of the wing along the X axis of the aircraft) depends on the number M of the flight.

- The parameter n _{at max} - the maximum permissible vertical overload of the aircraft - is determined individually for each type of aircraft and depends only on the design features of the aircraft or on the physical abilities of the human pilot.

- Parameter g - gravitational acceleration (constant).

- S is the wing area of the aircraft.

- относительное расстояние от центра тяжести самолета до передней кромки средней аэродинамической хорды крыла вдоль оси Х самолета.-

- the relative distance from the center of gravity of the aircraft to the leading edge of the middle aerodynamic chord of the wing along the X axis of the aircraft.

- m _with - the mass of an empty plane.

- m is the mass of the aircraft.

The calculator of the value δ _max2 24 outputs the value obtained as a result of the calculations, proportional to the value δ _max2 , to the comparator 25.

In the comparison device 25, the parameters δ _max1 and δ _{max2 are compared} . The smallest of them in absolute value (the largest in absolute value, since δ _max1 and δ _{max2 are} negative), the value is issued in SKU 15 and is called δ _max .

In SKU 15, the signals Δ δ from the calculator of the lead value 13 with the value of δ - the preset elevator position, which can be obtained from DOW 19, the wiring elements of the elevator control system (pitch), from sensors and other elements mentioned in this description, are added. In SKU there is also a subtraction of the signal δ _max received from the calculator 14.

Thus, in SKU calculations are performed according to the formula (29):

Δ δ + δ -δ _max = Δ δ _k , (29)

where Δ δ _k is the result of these calculations (correction value).

The value Δ δ _k formulated in SKU 15 is transferred from it to the conversion device associated with it (position 16).

The converting device 16 transmits to the output from the device 16 without changing the signals corresponding to the negative value of the correction value, and resets (does not pass) the signals corresponding to the positive value of the correction value. The signal from the converting device 16 enters the adder 20 (CMS) and the calculator of the correction value 13 (specifically, the software device 8) (PRU). Perhaps the use of signal amplifiers for software device 8, in which the correction signal entering the software device 8, would be amplified. For a software device, the value of the correction value is not important. Only the fact of the presence or absence of a negative signal of the correction value is important. Upon receipt of a signal corresponding to a negative value of the correction value, the software device generates signals for the air conditioner (transfer function computer) (item 7), corresponding to the inclusion of the first stage of restriction in accordance with the table (mode No. 1 is turned on) K1 = 0, K2 = 1. This occurs at time t ₂ (Fig.6). Mode No. 1 allows the aircraft by inertia to reach the maximum angle of attack α _max (at time t ₃ ). In this case, the angle of attack of the aircraft α will be higher than the angle of attack α _given , which is set by the elevator (Fig.6, the time interval from t ₂ to t ₃ ).

The signal from the converting device 16 (CU) to the adder 20 (CMS) corrects the control signal δ that goes through all elements of the elevator control system from the OS 18 to the elevator 22. In the CMS 20, the correction value Δ δ _k is subtracted from the signal control δ according to the formula (30):

δ _to = δ -Δ δ _to . (thirty)

From the control system 20, the corrected control signal δ _k enters the actuator 21. It sets the PB in such an angular position that corresponds to the corrected control signal δ _k = δ -Δ δ _k . In addition, it is possible to use the signal δ _to differentiate it in the remote control 12 to obtain the parameter δ '(t).

In the case of using a mechanical aircraft control system using mechanical rods and rockers, the operation of devices 13 and 14 is similar to the above devices. The difference is as follows.

The summing device (figure 10, position 15) (SKU) subtracts from the signal δ _max coming from the device 14, the signal Δ δ coming from the device 13 (VVU). The resulting signal, proportional to the value δ _{max max} = δ _max- δ, is transmitted to the MIU 38. The MIU moves its rod, which limits the stroke of the stop device 40, mechanically connected with the elevator control system. The placement of the MIU and the stroke of its rod should be adjusted so that when the rod abuts against the stop of the thrust device, the elevator travel would be limited at such angles δ, the value of which would be equal to δ _max -Δ δ. The thrust device is mechanically connected to the actuator 21 (IU), which is associated with the elevator. At the end of the MIU rod there is a limit switch 39, which, when the rod stops at the stop 42 of the thrust device 40, gives a signal to the switchgear 8 for switching on the mode No. 1 (K1 = 0, K2 = 1). Mode No. 1 allows the aircraft by inertia to reach the maximum angle of attack (α _max (at time t ₃ ). Moreover, the angle of attack of the aircraft α will be higher than the angle of attack α _specified , which is set by the elevator (Fig.6, time interval from t ₂ to t ₃ ).

The operating mode No. 2 (for a mechanical or electrical remote control system) can be turned on according to the conditions indicated in table 1. This is either the expiration of the time interval τ (formula 8) from the moment mode No. 1 is turned on, or the condition for mode No. 1 is turned off while three conditions are fulfilled simultaneously α '' (t) <0; α '(t) ≤ 0; α ≤ α _max . This occurs at time t ₃ (Fig.6). In mode No. 2, the switchgear 8 generates K1 = 1; K2 = 100 according to table 2 and gives these coefficients to the multipliers No. 1 and No. 3, respectively. Mode No. 2 (transition to the second stage of restriction) allows the elevator to reach the angle δ _max . Due to this, the aircraft will be fixed at the maximum angle of attack at that moment in time t ₄ , when the growth of the angle of attack will stop. The equality of the given and actual angle of attack at α '(t) = 0 means the absence of further fluctuations in the angle of attack.

After one or two seconds, mode No. 2 is turned off and there is a transition to mode “P” (operating mode K1 = 1; K2 = 1). In this case, the aircraft can remain at the maximum angle of attack. This occurs at time t ₅ (Fig.6). In the general case, during the “P” operating mode, the device monitors the moment when it is necessary to turn on the mode No. 1 (the first level of restriction). If the pilot, already at the maximum angle of attack, tries to increase the angle of attack even more, the first stage of restriction will again turn on (mode No. 1). In this case, the elevator angle δ in both the “P” mode, and in the mode No. 1, and in the mode No. 2 will be equal to δ _max (Fig.6, the time interval from t ₅ onwards).

The conditions for switching modes are summarized in table 1. When the aircraft power is turned on, the device operates in “P” mode, entering modes No. 1 and No. 2 (the first and second stages of restriction) only if necessary. To increase the reliability of the device, mode No. 3 is introduced. The mode switching circuit is shown in FIG. 11. Switching modes is carried out by changing the coefficients K1 and K2 in the switchgear 8.

The algorithm of the program for the on-board computer is as follows.

If an on-board computer is used as a correction device (KUS), the on-board computer receives all or some of the above parameters (step 31). Further, the control system (on-board computer) performs the calculations using the description formulas (step 32), using the data and data for the modes specified in tables 1 and 2 (when the system starts up, the “P” mode is set). At this stage of the algorithm, the KUS program also computes the lead value Δ δ. Since the transfer function (PF) is used for this — a first-order inertial link, one can use the method of successive approximations (iterations) to calculate it using the time increment Δ t equal to the real (synchronized) one. Synchronization can be carried out using the computer’s ability to determine real time by the computer’s system timer. Synchronization can be carried out at another stage of the program. For example, at step 37.

The output parameter Δ δ of the transfer function W (P) can be obtained by the method of successive approximations (iterations). Below is a section of a program in the C ++ programming language that calculates the output parameter Δ δ of the transfer function W (P) depending on the input parameter δ ’(t) - a parameter proportional to the angular velocity of the elevator. In this program, the variable i sets the number of approximations equal to 5. In the interest of speed and accuracy, this variable can be set different (for example, 3 or 6, etc.).

...

for (i = 1, i≤ 5; i ++)

{

xx1 = KB / T-x1 / T;

x1 = (xx1 + xxo) / 2 Δ t + x0;

}

x0 = x1;

xx0 = xx1;

...

In this program, the parameters x0 are the output signal of the transfer function Δ δ at the current time. Parameter x1 is the same after a period of time Δ t. The parameter Δ t is the time parameter through which the output parameter of the transfer function is calculated. Parameter xx0 is the rate of change of the output parameter Δ δ at the current time. Parameter xx1 is the same after a period of time Δ t. The time interval Δ t should not be large for the accuracy of the calculations. Normal can be considered Δ t equal to 0.1 seconds. The parameter K is the gain of the inertial link (transfer function). The parameter T is its time constant. Parameter B is the input signal of the transfer function in the current time period equal to δ ’(t). The parameters xx0 and x0 when turning on the power of the on-board computer are equal to zero.

Steps 33 and 34. A comparison of the values of δ _max1 and δ _max2 in the _{control panel} is performed with the aim of choosing the ones closest to zero (the highest, since δ _max1 and δ _max2 are negative values). In this case, you can use the logical operator "more" (or less). For example, in C ++:

...

if (δ _max2 > δ _max1 )

δ _max = δ _max2 ;

else

δ _max- δ _max1 ;

...

Next, the program adds the parameters δ, Δ δ and subtracts the parameter δ _max from the sum (step 35), where Δ δ _k is the correction value.

Δ δ _k = δ + Δ δ -δ _max

Step 36. Next, the program determines the presence or absence of conditions for switching to another operating mode (according to table 1) and, depending on this, changes the gain K of the transfer function and the time constant T of the transfer function (according to table 2).

Step 37. Issue (change) in the control system of the signal Δ δ _to (or negative or equal to zero). Then the transition to the beginning of the program.

For an on-board computer working in conjunction with a mechanical aircraft control system, the program algorithm will differ in that at step 35 the value Δ δ _maxMIU = δ _max -Δ δ is calculated and the result will be output to the MIU. In addition, at step 36, the computer (program) checks for the presence of a signal from the limit switch 39 (KB) located at the end of the stem 41 of the MIU 38 (Fig. 10). The appearance of this signal is a condition for switching to mode No. 1. At all stages of the program, the MIU extends its stem to a distance proportional to the value of δ _{max max mi} = δ _max -Δ δ, which it receives from the computer at step 37. The stroke of the rod is adjusted so that the thrust device abutting against the MIU stem limits the elements elevator control system and does not allow the elevator to reach an angle exceeding (in absolute value) an angle equal to Δ δ _max -Δ δ.

The operation of the device is as follows.

The device operates in such a way that if the negative sum of the values of δ and Δ δ is less than or equal to the negative value of δ ' _max , then the first stage of restriction will turn on ("mode 1"). In this case, the value of δ will be limited by the value of δ _max -Δ δ by subtracting the value of Δ δ _k from the value of δ in the control system 20 for the electronic remote control system, or by abutting the stop of the thrust device into the MIU 38 rod for a mechanical control system. The angle of attack of the aircraft by inertia increases as in Fig.6. And at time t ₃ it becomes approximately equal to the maximum allowable angle of attack α _max . In this case, the increase in the angle of attack of the aircraft α stops at the upper point of the oscillatory process.

The theoretical duration of the time interval from t ₂ to t ₃ is τ and is calculated by the formula (8). The time t ₃ can also be determined by the value of α and its derivatives (first and second) according to table 1. In order for the aircraft to have a maximum angle of attack for the required time, the device switches to the second stage of restriction (mode No. 2) in the interval time from t ₃ to t ₅ . One second is enough to go to the second level of restriction. The transition is achieved by zeroing the value of Δ δ in the corrective device (CUS) in the time interval from t ₃ to t ₅ . In this case, the value of δ will be limited only by the value of δ _max by subtracting the value of Δ δ _k from the value of δ in the control system (position 20), or by abutting the stop of the thrust device into the MIU stock. If the pilot needs a limiting angle of attack, then he will set the limiting position of the op-amp. RV will receive freedom of movement to a value of δ _max . In this case, the rod of the thrust device will abut the rod of the MIU at a value of δ equal to δ _max , or in SLA 20, the value Δ δ _k equal to (δ -δ _max ) will be subtracted from the value of δ (if the negative value of δ _max is greater than the negative value δ). At time t ₄ PB will be in the position δ _max . Due to this, the angle of attack of the aircraft will retain its limiting value due to the aerodynamic dependence of α on the angular position of the RS δ and due to the lack of inertia of growth α (α '(t) = 0).

VP (position 7) is an inertial link of the first order.

PRU (position 8, figure 2) controls the coefficients K1 and K2, thereby changing the gain of the calculator K and its time constant T.

They provide 4 limiter operation modes:

- working / P / - calculation of the lead value,

- first / 1 / - inclusion of the first stage of restriction,

- second / 2 / - transition to the second stage of restriction,

- the third / 3 / (optional) - increases the reliability of the device, slowing down the process of reducing the output signal of the calculator.

The conditions for switching on and off the various operating modes of the limiter are summarized in table 1:

Mode number Inclusion moment Duration Shutdown moment 1 - δ -Δ δ> -δ max
- For mechanical control wiring: operation of the limit switch 39 on the rod MIU 38 (Fig. 10).
- For wiring control: the appearance of an electrical signal Δ δ _to <0 from PU 16
(figure 3). Until the mode 1 is turned off. (Theoretical duration is <τ>)
Instead of condition 1a, we can use condition 1b (Switching to mode 2) 1a) Or the simultaneous fulfillment of three conditions:
- α ’’ (t) <0
- α ’(t) <0
- α ≤ α_max
(Switch to mode 2).
1b) Or according to the duration of mode No. 1, equal to τ (Transition to mode 2).
2) Or: removal of efforts by the pilot from the helm. (On the helm column, you can place the limit switches, fixing the application of effort by the pilot).
(Switch to P mode). 2 1a) Turning off mode 1 by simultaneously fulfilling three conditions:
- α ’’ (t) <0
- α ’(t) <0
- α ≤ α_max
1b) Or, according to the duration of regime No. 1, equal to τ 1 second The transition to the operating mode <P> after one second. Mode number Inclusion moment Duration Shutdown moment 3 The appearance of an increasing (decreasing in absolute value) negative signal Δ δ at the output of the VP calculator in the <P> mode. (When -Δ δ ’(t) <0 in mode P.) Mode 1 takes precedence over mode 3. Until –Δ δ ’(t) <0 or until mode 1 is turned on. Or: Enabling mode 1. (Mode 1 takes precedence over modes P and 3.)
2) Or: Turning on the P mode with an increase of -Δ δ. (Go to <P> with -Δ δ '(t)> 0). R Turning on the power supply of the limiter or the absence of any of modes 1, 2, 3. Not limited. Or: Enabling mode 1. (Mode 1 takes precedence over Mode 3.)
2) Or: Turn on mode 3.

The values of the gain of the transmitter of the transfer function K, its time constant T, the coefficients K1 and K2, depending on the operating mode of the limiter, are summarized in table 2:

Mode TO T K1 K2 R TO T 1 1 1 K ∞ 0 1 2 0.01K 0.01T 1 100 3 K 3T 1/3 1

The method of limiting the maximum angle of attack and vertical overload of the aircraft

The method of limiting the maximum angle of attack and vertical overload of the aircraft is the impact of the KUS 17 on the aircraft control system. KUS 17 has 2 steps to limit the position of the PB. The value of the angle of attack aerodynamically depends on the angle of installation of the PB δ.

The first stage of restriction is introduced when the position of the PB δ, equal to the value of δ _max minus the value of Δ δ (lead value). This preemptive restriction is valid until t ₃ (Fig. 6), while the inertial growth of the angle of attack continues. (Theoretically, the duration of this restriction is τ). At time t _3, this growth stops. The second stage of restriction is introduced at time t ₃ . In this case, the control system imposes a restriction on the angular position of the PB δ equal to δ _max . The duration of this period (mode No. 2) can be equal to one second. After turning off the second stage of the limitation KUS 17 goes into "working" mode "P". Moreover, if the pilot continues to move the OS 18 to increase the angle of attack, in the “operating” mode “P”, the control panel determines that δ is less than or equal to the value of δ _max minus the value of Δ δ (all of these values are negative), and therefore again includes first step of limitation. However, the value of Δ δ will be equal to zero and, therefore, the value of δ will be equal to δ _max . Further analysis of the operation of this device will show that under such conditions under any of the operating modes of the control system, δ will not go beyond δ _max , and the angle of attack of the aircraft α will not exceed α _max .

Thus, the control system acts on the aircraft control system so that a restriction of δ _max with a lead Δ δ is introduced on the course of the aircraft. The lead value Δ δ is determined by the control system in accordance with the transfer function - the inertial link of the first order. The input value of this function is the angular velocity of the elevator or a value proportional to it (for example, the angular velocity of the control or the velocity of any element of the elevator control system).

The gain K and the time constant T of the transfer function W (P) depend on the parameters specified in this description, which can be measured by the aircraft's on-board sensors or can be chosen constant for a typical flight mode.

The lead Δ δ is temporary. The time for introducing and removing the restriction on the control signal δ is determined in the control system according to the principles set forth in this description and shown in table 1. The lead-in occurs at the moment when the value δ reaches the value δ _max minus Δ δ. Anticipation lasts until the angle of attack of the aircraft reaches the upper point of the oscillatory process close to the value of the maximum angle of attack (from the condition of the maximum allowable angle of attack for aerodynamic reasons, or from the condition of maximum overload). After removing (turning off) the lead on the control signal δ, the control signal is able to change its value to the value δ _max (the range is expanded). In this case, a change in the control signal to a value of δ _max does not lead to an increase in α (angle of attack of the aircraft) above the value of α _max (maximum allowable α). Indeed, in this case, α will already be greater than or equal to the angle of attack defined by the elevator α _given . So if α changes, then only downward, because its growth rate (inertia) is already extinguished, and the aerodynamic forces (from the elevator) will tend to bring into correspondence α, real with α, given the elevator. And α, given by the elevator, can only be less than or equal to α _max .

This method can also be used as a signal to the pilot about approaching the limiting angles of attack. At the same time, at time t _{2, the} device will give a value Δ α ₁ equal to the throw angle of attack α above a predetermined angle of attack α _given for time t ₃ . The formula for the transfer function and its gain will take the form:

K = -R · A · A _mz $\genfrac{}{}{0ex}{}{\text{δ}}{}$ / q ² , (32)

where Δ α ₂ is a value equal to the throw of the angle of attack α, above a given angle of attack α _given at time t ₃ , calculated at time t ₂ . 6, this value is equal to the distance from the first stage of restriction to the second stage of restriction or the distance from line α to line α _given at time t ₃ .

In addition, the following transfer function may be most useful:

K = -R · (A-1) · A _mz $\genfrac{}{}{0ex}{}{\text{δ}}{}$ / q ² ; (34)

where Δ α ₁ is a value equal to α _given minus α at the current time.

At any moment of time (for example, t ₂ ), the quantity Δ α ₁ + Δ α ₂ will show how much the angle of attack α grows, if at that moment the elevator is stopped.

Claims (3)

1. The method of limiting the angle of attack and overload of the aircraft, characterized in that the control signal of the aircraft elevator control system, coming from the control that controls the material object - the elevator of the aircraft, introduces a restriction with pre-emptive until the aircraft reaches the maximum angle of attack, while the lead value is calculated according to the principle of the transfer function — the inertial link of the first order, whose input signal is a value proportional to the global speed of the elevator or control element of the aircraft elevator control system, and the outgoing one is the lead, and after removing the lead on the control signal, the latter gets the opportunity to change to a value corresponding to the maximum allowable angle of elevator installation, which corresponds to the maximum allowable angle of attack on steady-state flight mode, in which the angle of attack does not change and is completely balanced, corresponding to the maximum value set by the elevator. 2. A device for limiting the angle of attack of an aircraft, characterized in that a correction device is included in the control system of the elevator control of the aircraft, introducing a restriction on the control signal with lead, and the lead value is calculated in the correction device according to the principle of the transfer function — the inertial link of the first order, whose input signal is a value proportional to the angular velocity of the elevator or control element of the aircraft elevator control system, and the output box - the amount of lead, and devices that allow the correction device to change the control signal, for example, summing devices, thrust devices or actuators, which, together with the correction device, affect the aircraft elevator control system, limiting the control signal, which, in turn, restricts movement elevator control. 3. The device according to claim 2, characterized in that it receives a parameter proportional to the angular velocity of the elevator of the aircraft or the angular velocity of the elevator of the aircraft from sensors of these quantities associated with the control or elevator or with any element of the system control the elevator of the aircraft, which is a conductor of the control signal.

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