RU2377507C1 - Facility for gross weight control of aircraft fuel - Google Patents

Abstract

FIELD: measuring technology.

SUBSTANCE: invention refers to aeronautical engineering and, particularly to systems of complex control of weight and centring of passenger and transport airplanes of classic scheme wherein fuel is placed in wing, while load is placed in fuselage. The facility of gross weight control of airplane fuel consists of sensors of angle of attack, Max number, altitude, interceptor's deviation angles, wing flap deviation angles, slat deviation angle, rate of list, rate of pitch and rate of yawing connected to the input of a multiplier, of a differentiator, and of a switchboard, the input of which is connected to the output of a divider. The switchboard is connected to a memory block, the input of which is connected with the output of circuit "OR", the inputs of the latter are connected to comparing blocks. Additionally the switch board is connected to a functional converter which together with the comparing block evaluates plane fuel weight and centring. Also the comparing block consists of connections with sensors of aerodynamic indices, angular motions of a plane, and rudders deviation, with the multiplier, differentiators, a subtraction circuit, an adder and a block of list moment determination.

EFFECT: upgraded validity and accuracy of fuel weight control, increased reliability and safety of flight, and improved operation characteristics of airplane.

3 cl, 6 dwg

Description

The invention relates to the field of aviation technology and can be used for integrated control of the fuel systems of passenger and transport aircraft of the classical scheme, in which the fuel is located in the wing and the load is in the fuselage. It can also be used to determine the current value of the fuel weight in flight, the transverse alignment of the aircraft, ensure flight safety and operational control of the basic aerodynamic characteristics of the aircraft: range, altitude, speed.

A known system for diagnosing faults in numerical control devices using a programmable controller that contains a remote host computer that performs the diagnostic procedure by reading the input and output signals of the programmable controller and finds out the malfunction by processing the specified input and output signals [Pat. 90/02366 PCT (WO). MKI G05B 23/02. Fault Diagnostic System / RJ. Inventions of the countries of the world. No. 9. 1990]. The system has broad capabilities for general failure detection. However, for its operation, it is necessary to have input and output signals of the analyzed system, which for the aircraft fuel systems takes place only for ground, pre-flight measurement of the amount of fuel from the tanker. For flight control of fuel weight and alignment, it is necessary to use a fuel gauge sensor in the aircraft tank [1, p.225-229], which has limited reliability, as do engines (fuel consumers), pipelines, fuel tanks, electrical wiring, and fuel weight indicators. Errors of fuel weight control are known for such ground control ("Gimli Glider"). The method for determining the weight of fuel with a fuel gauge is important for the quality of control [2, p. 99, table 7.1].

A device for flight and preflight monitoring of a capacitive fuel gauge containing a relay, buttons for its remote activation on the body of the fuel weight indicator and reference capacitors [Fuel control and change system SUIT4-5, SUIT4-5A, SUIT4-5B, SUIT4-5V. Technical Operation Manual 6T1.620.020. RE. p. 47, fig. 22; Konstantinov V.D. Aviation equipment and its operation. Ed. VVIA them. N.E. Zhukovsky, 1966. S. 171, Fig. 10.7]. The fuel gauge sensor is included in the bridge circuit of the tracking system connected to the input of the pointer. The capacity of one of the capacitors of this device corresponds to the capacity of the fuel gauge sensor when the tank is empty, and the capacity of the other corresponds to the capacity of the fuel gauge sensor when filling the tank by 2/3 of its maximum volume. When buttons are pressed instead of the capacitance of the fuel gauge sensor, the relay includes reference capacitors in the bridge measuring circuit of the fuel gauge. The deviation of the pointer arrow should consistently display the zero value of the fuel weight and 2/3 of the maximum fuel weight. Control is reduced to determining the serviceability of only a bridge circuit, a tracking system and a pointer. Fuel weight, sensors, signal lines, fuel distribution in tanks (centering) are not checked. The device does not control the alignment of the aircraft. For typical fuel gauges of the SUIT8 type (reliability Р1 (2) = 0.998824, Т1 = 1700 h) when monitoring the weight of fuel on an IL-86 airplane (34 sensors, maximum reliability of monitoring Рк (2) = 0.99932) reliability [3] detection of a fuel system failure in 2 hours of flight will be Rd1 (2) = 0.633894, which corresponds to the time of reliable control Td1 = 4.39 hours.

A device for flight control of the weight of fuel of an aircraft containing a capacitive fuel meter and a summing fuel flow meter [Aviation devices. Ed. S.S.Dorofeeva. M .: Military. publ., 1992. S. 236-241]. Devices duplicate each other. The capacitive fuel gauge sensor is a capacitor, the plates of which are located in the fuel tank and are coaxial profiled pipes. The totalizing fuel flow meter includes an impeller sensor connected to a magnetic induction tachometer located in the fuel pipe between the fuel tank and the airplane engine. The integrator at the output of the tachometer and the subtraction circuit with the setter of the initial amount of fuel allow you to determine the remaining fuel in the tank. When the tank is filled with fuel, the capacitance of the capacitor changes and at the output of the bridge circuitry, a capacitor is connected to one of the arms, a signal proportional to the fuel level in the aircraft tank appears. At the same time, the fuel consumption from the tank is measured by the flow meter and its integrated value is subtracted from the known initial amount of fuel on the master. The device has a comprehensive nature, as it also serves to control the engine of the aircraft for fuel consumption from the tank. The totalizing flow meter allows you to determine the remaining amount of fuel in the tank. In flight, the pilot constantly monitors the difference between the readings of the duplicating fuel meter and the summing flow meter. A working fuel system of the aircraft should be minimal [4, p.237-242]. If there is a mismatch in the testimony, the pilot may find himself in a difficult situation of uncertainty in the state of serviceability of the system, including the aircraft engine, fuel, fuel tanks, pipelines, electrical wiring, fuel gauge and totalizing flow meter [5]. Given the very slow nature of fuel consumption in flight and the relatively low reliability of the complex multicomponent fuel system of an airplane, P2 (2) = 0.998824 × 0.99757 = 0.996397 (T2 = 554 hours), where P3 (2) = 0.99757 - the reliability of the totalizing flow meter, there are pilot errors due to the imperfection of such flight control. This is largely due to the indirect nature of monitoring the weight of the fuel, both by the fuel meter (by the level of fuel in the tank) and by the totalizing flow meter (by the difference between the declared volume and the volume integrated by the flow meter). The reliability of controlling the weight of the fuel by the fuel meter here will be Рд2 (2) = 0.674411, and by the summing flow meter - Рд3 (2) = 0.326382, which for the fuel meter corresponds to the time of reliable failure detection Тд3 = 1.79 hours, which is less than the flight time . Determining the alignment of the aircraft requires additional pilot attention.

Known for the most widely used test equipment of fuel and flow systems in the pre-flight state of the aircraft [Aviation devices. Ed. S.S.Dorofeeva. M .: Military. publ., 1992. p.248]. The device consists of a capacitance meter for dry fuel gauge sensors, a self-balancing AC bridge, the two arms of which are the rheostat of the test equipment, and the other two are the rheostat in the device under test. Measurement of electric capacitance is carried out using a variable capacity magazine, in parallel with which the tested sensor capacitance is connected. On the front panel of the equipment is a cassette for attaching a fuel gauge pointer, warning lights and system mode switches. The capacitance of the fuel gauge sensors is measured and the health of the tracking systems and indicators is assessed. To check the flowmeter without an impeller sensor (it is difficult to dismantle it from the fuel system pipe), an inductive-pulse device is used, similar in principle to the operation of the flow sensor device and design. The magneto-induction tachometer is driven by an electric motor, the adjustable rotation speed of which simulates various hourly fuel consumption on an airplane. Such control is carried out only on the ground in a partially dismantled system, has a high degree of reliability of detection of sensor failure, but is very time-consuming [6, p. 28; 7, p. 257].

The closest of the known technical solutions is a device for flight control of the load weight of the aircraft [US Pat. 2260179 RF. MKI G01G 19/07, G06G 7/70. Device for flight control of aircraft load weight / V.Yu. Chernov // B.I. 2005, No. 25], which contains an angle of attack sensor, a Mach number sensor, a speed sensor, a height sensor, an angle sensor for deflecting the spoilers, an angle sensor for the deflection of the flaps, an angle sensor for the deviation of the slats, an angle sensor for the deviation of the flaps, a sensor for the angle of deviation of the elevator, an angle sensor stabilizer deviations, centering sensor, chassis release sensor, pitch speed sensor, roll speed sensor, yaw rate sensor, multiplier, adder, differentiator, divider, memory unit, first, second and third comparison blocks, OR circuit, comm tripod, functional converter and pitch moment determination unit, first, second, third, fourth, fifth, sixth, seventh, eighth, ninth, tenth, eleventh, twelfth, thirteenth inputs of which are connected respectively to the outputs of the angle of attack sensor, Mach number sensor, sensor speed sensor, height sensor, sensor for deflection angles of spoilers, sensor for deflection angles for flaps, sensor for angle deflection for slats, sensor for angle deflection for flaps, sensor for angle of deviation for elevator, sensor for angle of deviation mash, centering sensor, landing gear sensor, pitch speed sensor, the first input of the adder through a differentiator is connected to the pitch speed sensor, the second input to the output of the multiplier, the inputs of which are connected to the roll speed sensor and the yaw rate sensor, the first input of the switch, as well as the input the first comparison unit, connected to the output of the adder, the second input of the switch - with the output of the divider, the first input of which, like the input of the second comparison unit is connected to the output of the pitch moment determination unit, the third input of the switch a - with the output of the memory block, the fourth input of the switch, like the first input of the memory block - with the output of the OR circuit, the inputs of which are connected respectively to the outputs of the first and second comparison blocks, the first output of the switch with the second input of the divider, the second output of the switch with the second the input of the memory block, the third output of the switch with the inputs of the functional Converter, designed to calculate the inverse function, and the third block of comparison. In the pitch moment determination unit, the current value of the pitch moment acting on the aircraft is calculated. At the same time, its pitch angular acceleration is determined. The load weight of the aircraft is determined by the known functional relationship with the moment of inertia along the transverse axis of the associated coordinate system. Weight control is implemented in the comparison device.

A disadvantage of the known prototype device is the inability to determine the weight of the fuel in flight. The reason that impedes the obtaining of the technical result indicated below when using the known prototype device is the use of dynamic and kinematic relations describing the movement of the aircraft along the transverse axis of the associated coordinate system. In contrast to the weight of the load, the effect of changes in fuel weight on the inertial characteristics of the aircraft is minimal.

The main task, the solution of which is claimed by the claimed object device is the flight control of the weight of the fuel of the aircraft.

The technical result is to increase the reliability and accuracy of control of the fuel system, the reliability of the assessment in flight of the actual weight of the fuel, changing in flight. The use of a different physical principle and standard inertial equipment for fully autonomous operation of the device increases the safety of the fuel system without installing electric fuel meter sensors in the tanks, eliminates electrical wiring, controls alignment and reduces the number of tight joints on the fuel tanks.

The specified technical result is achieved by the fact that in the device for flight control of the weight of the fuel of the aircraft containing the sensor of the angle of attack, the sensor of the Mach number, the sensor of speed, the height sensor, the sensor of the deflection angle of the spoilers, the sensor of the deflection angle of the flaps, the sensor of the deflection angles of the slats, roll speed sensor a pitch speed sensor and a cutting speed sensor, the output of which is connected to the input of the multiplier, a differentiator, a switch, the first input of which is connected to the output of the divider, the first input connected to the first output by the switch, the second input of the switch is connected to the output of the memory block, the first input of which, like the third input of the switch, is connected to the output of the OR circuit, the inputs of which are connected to the outputs of the first and second comparison blocks, the second output of the switch is connected to the second input of the memory block, the third the output of the switch is connected to the input of the functional converter, the third comparison unit, a slip angle sensor, a rudder angle sensor, an aileron angle sensor, a temperature sensor, a second multiplier, are introduced swarm and third differentiators, subtraction circuit, four-input adder, roll moment determination unit, first, second, third, fourth, fifth, sixth, seventh, eighth, ninth, tenth, eleventh, twelfth and thirteenth inputs of which are connected respectively to the sensor outputs angle of attack, Mach number sensor, speed sensor, height sensor, sensor for deflection angles of the spoilers, sensor for deflection angles for flaps, sensor for angles for deviation of slats, sensor for sliding angle, sensor for angles of deviation of the rudder, sensor and the deviation angles of the ailerons, roll speed sensor, yaw rate sensor, temperature sensor, wherein the output of the yaw rate sensor is connected to the input of the second multiplier and the input of the second differentiator, the output of the roll speed sensor is connected to the first differentiator and the second multiplier, the output of the roll moment detection unit is connected to the first summing input of the adder and the input of the second comparison unit, the second subtracting input of the adder, like the first subtracting input of the subtraction circuit, is connected to the output of the first multiplier, second the input of which is connected to the output of the pitch speed sensor, the third subtracting input of the adder is connected to the output of the second multiplier, the fourth summing input of the adder is connected to the output of the second differentiator, the output of the adder is connected to the second input of the divider, the output of the first differentiator is connected to the summing input of the subtraction circuit, the output of which connected to the inputs of the first comparison unit and the fourth input of the switch, the input of the third differentiator is connected to the third output of the switch, and the output to the input of the third block with equalities.

The set of essential features of the invention provides a technical result achieved during the implementation of the invention is a device for flight control of the weight of the fuel of the aircraft. Moreover, the invention consists in determining the weight of the fuel of the aircraft through its moment of inertia along the longitudinal axis of the associated coordinate system. For this, the dynamics of the angular motion of the aircraft along the longitudinal axis of the associated coordinate system is considered [8, p.20, 39]

where I _X , I _Y , I _Z - the moments of inertia of the aircraft, respectively, along the longitudinal, normal, transverse axes;

I _XY - centrifugal moment of inertia;

ω _X , ω _Y , ω _Z - the angular velocity of the aircraft, respectively, along the longitudinal, normal, transverse axes;

,

- угловое ускорение самолета по продольной и нормальной оси;

,

- angular acceleration of the aircraft along the longitudinal and normal axis;

M _X - moment of external forces relative to the longitudinal axis - roll moment.

The roll moment M _{X is} calculated in the roll moment determination unit based on the signals of the angle of attack sensor, Mach number sensor, speed sensor, height sensor, interceptor deflection angle sensor, flap deflection angle sensor, slat deflection angle sensor, glide angle sensor, rudder angle sensor , aileron deflection angle sensor, roll speed sensor, pitch speed sensor, yaw rate sensor and temperature sensor. Given the ratio of moments of inertia of the aircraft of the classical scheme:

expressions (1) can be written in the form where only I _X depends on the fuel weight m _T ,

получаемThen, given the roll moment M ' _X and the roll acceleration

we get

The determination of the moment of inertia by (4) is possible only under the conditions:

, M_Xmin - пороговые значения углового ускорения

и момента внешних сил М'_X, для которых возможно определение момента инерции (4). Они определяются по (4) из условия точности решения задачи контроля. Если оба условия (5) выполняются, то момент инерции определяется по (4). Если же хотя бы одно из условий (5) не выполняется, то момент инерции определяется по значению вычисленному в предшествующий момент времени, когда условия (5) выполнялись. Информация о предшествующем значении момента инерции хранится в блоке памяти, а условия (5) определяются первым и вторым блоками сравнения, которые управляют работой блока памяти и коммутатором. Последний выдает либо текущее значение момента инерции с делителя, реализующего выражение (4), либо значение запомненное в блоке памяти. Момент инерции I_X по продольной оси связанной системы координат, из-за малого радиуса инерции нагрузки m_K в фюзеляже самолета, приближенно включает только две составляющие так, чтоWhere

, M _Xmin - threshold values of angular acceleration

and the moment of external forces M ' _X , for which it is possible to determine the moment of inertia (4). They are determined by (4) from the condition for the accuracy of solving the control problem. If both conditions (5) are satisfied, then the moment of inertia is determined by (4). If at least one of conditions (5) is not satisfied, then the moment of inertia is determined by the value calculated at the previous time moment when conditions (5) were satisfied. Information about the previous value of the moment of inertia is stored in the memory unit, and conditions (5) are determined by the first and second comparison units that control the operation of the memory unit and the switch. The latter gives either the current value of the moment of inertia from the divider that implements expression (4), or the value stored in the memory block. The moment of inertia I _X along the longitudinal axis of the associated coordinate system, due to the small radius of inertia of the load m _K in the fuselage of the aircraft, approximately includes only two components so that

where I _X (m _T ) is the moment of inertia of the fuel in the wings of the aircraft;

I _X (m _K ) - moment of inertia of the load in the fuselage of the aircraft.

The signal at the switch output, proportional to I _X , can be a control signal for estimating the fuel weight m _T. The third comparison unit, together with the third differentiator, implements the condition dI _X / dt≤0. It corresponds to a monotonic decrease in the function I _X (m _T ) at fuel consumption in flight and its symmetrical generation from the tanks of the right and left wing [9, p. 240, the first row of the table of moments of inertia], which corresponds to the normal transverse centering of the aircraft. The moment of inertia I _X (m _T ) for the approximate linear model of the wing with a uniform distribution of the weight of the fuel here has the least value only if the weight is distributed symmetrically over its span.

At the same time, knowing the dependence I _X (m _T ) for a given type of aircraft, the value of the fuel weight m _T = I ^-1 _X in the functional converter is calculated as the inverse function of the dependence of the moment of inertia on the weight of the fuel on the plane. The output signal of the functional converter allows the pilot to determine the current value of the fuel weight. So if the moment of inertia is approximated by a polynomial of the form

then in the functional converter the ratio

where a ₁ ÷ a ₃ are the spline coefficients of the approximation of the known dependence I _X (m _T ).

The specified technical result in particular cases is achieved by the fact that in the device for flight control of the fuel weight of the aircraft according to claim 1, the roll moment determination unit comprises a control unit and a roll moment calculator, the first, second, third, fourth, fifth inputs of which are connected respectively to the first , second, third, fourth, fifth inputs of the roll moment determination unit, the sixth and seventh inputs of which are connected respectively to the first and second inputs of the control unit, the output of which forms a control bus connected to the sixth input of the roll moment calculator, the seventh, eighth, ninth, tenth, eleventh, twelfth inputs of which are connected to the eighth, ninth, tenth, eleventh, twelfth, thirteenth inputs of the roll moment determination unit, the output of which is the output of the roll moment calculator. The essence of a special case of the implementation of the roll moment determination unit is to separate the functions of computing control and the continuous calculation of the moment, respectively, in the control unit and the roll moment calculator. The control unit receives signals at its first and second inputs proportional to the deviation angles of the slats and wing flaps of the aircraft. By their magnitude, it generates control signals at the first, second, third or fourth outputs of the control bus u. The control unit is necessary for changing the coefficients in the unit for calculating the roll moment at different ratios of the angles of the slats and flaps of the wing mechanization in take-off, climb, cruising and landing airplane flight modes.

The specified technical result in particular cases is achieved by the fact that in the device for flight control of the fuel weight of the aircraft according to claim 2, the roll moment calculator contains eighteen variable coefficient blocks, seventeen multipliers, two dividers and an adder so that the first input of the roll moment calculator is connected to the first the inputs of the first, second, third, fourth, fifth, sixth, seventh, eighth blocks of variable coefficients, the second input of the roll moment calculator is connected to the first inputs of the ninth, tenth, one of the twelfth, thirteenth, fourteenth blocks of variable coefficients and second inputs of the first, fourth blocks of variable coefficients, the third input of the roll moment calculator is connected to two inputs of the third multiplier, the first inputs of the second and third divider, the second inputs of which are connected respectively to the outputs of the fourth and fifth multipliers , the outputs of the second and third divider are connected to the inputs of the second adder, the fourth input of the roll moment calculator is connected to the second inputs of the ninth, tenth, od of the eleventh, twelfth, thirteenth, fourteenth blocks of variable coefficients and the first input of the fifteenth block of variable coefficients, the fifth input of the roll moment calculator is connected to the first input of the sixteenth block of variable coefficients, the second input of which, like the second inputs of the second, third, fifth, sixth, seventh, of the eighth, seventeenth and eighteenth blocks of variable coefficients, connected to the sixth input of the roll moment calculator, and the output to the input of the sixth multiplier, the second input of which is connected to the output of the fifth block of variable coefficients, and the output with the input of the seventh multiplier, the second input of which is connected to the output of the twelfth block of variable coefficients, the output of the seventh multiplier is connected to the input of the second adder, the seventh input of the roll moment calculator is connected to the first input of the eighth multiplier, the second input of which is connected with the output of the eleventh block of variable coefficients, and the output with the first input of the ninth multiplier, the second input of which is connected to the output of the fourth block of variable coefficients, d of the ninth multiplier is connected to the first input of the tenth multiplier, the second input of which is connected to the output of the third block of variable coefficients, the output of the tenth multiplier is connected to the input of the second adder, the eighth input of the roll moment calculator is connected to the first input of the seventeenth block of variable coefficients, the output of which is connected to the input of the eleventh a multiplier, the second input of which is connected to the output of the fourteenth block of variable coefficients, the output of the eleventh multiplier is connected to the first input of the twelfth a multiplier, the second input of which is connected to the output of the eighth block of variable coefficients, and the output is connected to the input of the second adder, the ninth input of the roll moment calculator is connected to the first input of the eighteenth block of variable coefficients, the output of which is connected to the thirteenth multiplier, the second input of which is connected to the output of the ninth block variable coefficients, the output of the thirteenth multiplier is connected to the first input of the fourteenth multiplier, the second input of which is connected to the output of the seventh block of variable coefficients, and the output is with the input of the second adder, the tenth input of the roll moment calculator is connected to the first input of the fifteenth multiplier, the second input of which is connected to the output of the sixteenth multiplier, and the output is with the first input of the fifth multiplier, the second input of which is connected to the output of the second block of variable coefficients, the first the input of the sixteenth multiplier is connected to the output of the first block of variable coefficients, the second input to the output of the tenth block of variable coefficients, the eleventh input of the roll moment calculator is connected to the input seventeenth multiplier, the second input of which is connected to the output of the thirteenth block of variable coefficients, and the output is connected to the first input of the fourth multiplier, the second input of the fourth multiplier is connected to the output of the sixth block of variable coefficients, the output of the second adder is connected to the input of the eighteenth multiplier, the second input of which is connected to the output the third multiplier, and the output, as well as the output of the fifteenth block of variable coefficients, the second input connected to the twelfth input of the roll moment calculator, is connected to the input s nineteenth multiplier, the multiplier output is output nineteenth roll moment calculator.

The roll moment calculator provides the determination of the roll moment M _X acting on the aircraft according to the signals of the angle of attack sensor, Mach number sensor, speed sensor, height sensor, deflector angle sensor, flap angle sensor, flap angle sensor, slope angle sensor, angle sensor deviation of the rudder, aileron deviation angle sensor, roll speed sensor, yaw rate sensor, pitch speed sensor, temperature sensor. The calculation is carried out with variable coefficients of the influence of flight parameters for the main deviation angles of the slats and flaps of the wing mechanization. The commands for changing the coefficients are received by the roll moment calculator via the control bus from the control unit output, in which the standard configurations of the deviation angles of the wing slats and flaps of the aircraft are evaluated.

The roll moment is calculated by the following relations [10-11]:

where m _X is the moment coefficient; ρ is the air density; V is the flight speed; S is the wing area; l is the wingspan [8, p.21]. The moment coefficient m _X includes [8, p. 45; 10, p. 129; 11, p. 357, etc.]:

- составляющие коэффициента момента соответственно от углов отклонения элеронов, от углов отклонения руля направления, углов отклонения интерцепторов, угла скольжения, скорости крена, скорости рыскания самолета. Сумма составляющих коэффициентов получается на втором сумматоре вычислителя момента крена, входы которого включают все составляющие коэффициентов выражения (10), а выход, через восемнадцатый и девятнадцатый умножители, квадратор на третьем умножителе, и пятнадцатый блок переменных коэффициентов, формирует момент крена по выражению (9). Произведение 0,5 Sl в выражении (9) получается с помощью коэффициентов передачи названных умножителей и пятнадцатого блока переменных коэффициентов, который учитывает зависимость плотности воздуха ρ от температуры Т воздуха.Where

- components of the coefficient of the moment, respectively, from the angles of deviation of the ailerons, from the angles of deviation of the rudder, the angles of deviation of the spoilers, the angle of slip, roll speed, yaw rate of the aircraft. The sum of the component coefficients is obtained on the second adder of the roll moment calculator, the inputs of which include all the components of the coefficients of expression (10), and the output, through the eighteenth and nineteenth multipliers, the quadrator on the third multiplier, and the fifteenth block of variable coefficients, forms the roll moment according to the expression (9) . The product of 0.5 Sl in expression (9) is obtained using the transmission coefficients of the said multipliers and the fifteenth block of variable coefficients, which takes into account the dependence of the air density ρ on the air temperature T.

Coefficient m _{X. EL} , as a function of angles of attack, deviations of ailerons and Mach numbers, is formed in the seventh, ninth, eighteenth blocks of variable coefficients, thirteenth and fourteenth multipliers

where a _0.7 , a _1.7 , and _2.7 ; a _0.18 , a _1.18 , a _2.18 - variable coefficients that vary according to the control signals at the second inputs of the seventh and eighteenth blocks of variable coefficients according to the control signal u from the sixth control input of the calculator; a _0.9 , a _1.9 , a _2.9 - variable coefficients that change in the ninth block of variable coefficients for a signal proportional to the flight altitude.

The coefficient m _X.PH , as a function of angles of attack, deviations of the rudder and the Mach number, is formed in the eighth, fourteenth, seventeenth blocks of variable coefficients, the eleventh and twelfth multipliers

where a _0.8 , a _1.8 , a _2.8 ; a _0.17 , a _1.17 , a _2.17 - variable coefficients that vary according to the control signals at the second inputs of the eighth and seventeenth blocks of variable coefficients according to the control signal u from the sixth control input of the calculator; a _0.14 , a _1.14 , and _2.14 - variable coefficients that change in the fourteenth block of variable coefficients for a signal proportional to the flight altitude.

Coefficient m _X.IN , as a function of angles of attack, deviations of interceptors and Mach number is formed in the fifth, twelfth, sixteenth blocks of variable coefficients and the sixth and seventh multipliers

where a is _0.5 , and _1.5 , and _2.5 ; a _0.16 , a _1.16 , a _2.16 - variable coefficients that vary according to the control signals at the second inputs of the fifth and sixteenth blocks of variable coefficients according to the control signal u from the sixth control input of the calculator; and _0.12 , a _1.12 , a _2.12 - variable coefficients that change in the twelfth block of variable coefficients for a signal proportional to the flight altitude.

, как функция углов атаки, скольжения и числа Маха формируется в третьем, четвертом, одиннадцатом блоках переменных коэффициентов и восьмом, девятом, десятом умножителяхCoefficient

, as a function of the angle of attack, slip and Mach number is formed in the third, fourth, eleventh blocks of variable coefficients and the eighth, ninth, tenth multipliers

where a _0.3 , a _1.3 , a _2.3 - variable coefficients that vary by control signals at the second inputs of the third block of variable coefficients by a control signal u from the sixth control input of the calculator; and _0.4 , and _1.4 , and _2.4 - variable coefficients that change in the fourth block of variable coefficients for a signal proportional to the angle of attack; a _0.11 , a _1.11 , a _2.11 - variable coefficients that change in the eleventh block of variable coefficients for a signal proportional to the flight altitude.

, как функция углов атаки, скорости крена и числа Маха формируется в первом, втором, десятом блоках переменных коэффициентов, пятом, пятнадцатом, шестнадцатом умножителях и третьем делителеCoefficient

, as a function of angles of attack, roll speed and Mach number is formed in the first, second, tenth blocks of variable coefficients, the fifth, fifteenth, sixteenth multipliers and the third divider

where a _0.2 , a _1.2 , and _2.2 are variable coefficients that vary according to the control signals at the second input of the second block of variable coefficients according to the control signal u from the sixth control input of the calculator; a _0.1 , a _1.1 , and _2.1 - variable coefficients that change in the first block of variable coefficients by a signal proportional to the angle of attack; a _0.10 , a _1.10 , a _2.10 - variable coefficients that change in the tenth block of variable coefficients for a signal proportional to the flight altitude.

, как функция углов атаки, скорости рыскания и числа Маха формируется в шестом, тринадцатом блоках переменных коэффициентов, четвертом, семнадцатом умножителях и втором делителеCoefficient

, as a function of angles of attack, yaw rate and Mach number is formed in the sixth, thirteenth blocks of variable coefficients, the fourth, seventeenth multipliers and the second divider

where a _0.6 , a _1.6 , and _2.6 are variable coefficients that vary according to the control signals at the second input of the sixth block of variable coefficients according to the control signal u from the sixth control input of the calculator; and _0.13 , and _1.13 , and _2.13 are variable coefficients that change in the thirteenth block of variable coefficients for a signal proportional to the flight altitude. The numerical values of all the variable coefficients a _{i, j} i = 0 ÷ 2, j = 1 ÷ 18 are obtained by the aerodynamic characteristics of a particular aircraft [see, for example 11, p. 350], on which the inventive device is installed.

The analysis of the prior art by the applicant has established that there are no analogues that are characterized by sets of features that are identical to all the features of the claimed device for flight control of aircraft fuel weight, therefore, the claimed invention meets the “novelty” condition.

Search results for known technical solutions in this and related fields of technology in order to identify features that match the distinctive features of the claimed invention from the prototype have shown that they do not follow explicitly from the prior art.

From the prior art determined by the applicant, the influence of the transformations provided for by the essential features of the claimed invention on the achievement of the indicated technical result is not revealed and the invention is not based on:

- supplementing a known device - an analogue with any known part, attached to it according to well-known rules, to achieve a technical result, in respect of which the effect of this addition is established;

- replacing any part of the device - analogue with another known part to achieve a technical result, in respect of which the effect of such an addition is established;

- the exclusion of any part of the device - analogue with the simultaneous exception due to its presence of the function, and the achievement of the usual result for such an exception;

- increasing the number of elements of the same type to enhance the technical result due to the presence in the device of just such elements;

- the implementation of the known device is an analogue or part thereof from a known material to achieve a technical result due to the known properties of the material;

- the creation of a device consisting of known parts, the choice of which and the relationship between them are based on known rules and the technical result achieved is due only to the known properties of the parts of this device and the connections between them;

- a change in the quantitative sign (s) of the device and the provision of such signs in the relationship or a change in the type of relationship, if the fact of the influence of each of them on the technical result is known and new values of these signs or their relationship could be obtained on the basis of known dependencies, therefore, the claimed invention meets "Inventive step".

The invention is illustrated by drawings, where figure 1 shows a structural diagram of a device for flight control of the weight of the fuel of the aircraft. In figure 1, the following notation:

1 - angle of attack sensor;

2 - Mach number sensor;

3 - speed sensor;

4 - height sensor;

5 - sensor deflection angles of interceptors;

6 - flap deflection angle sensor;

7 - sensor deviation angles slats;

8 - roll speed sensor;

9 - pitch speed sensor;

10 - yaw rate sensor;

11-1, 11-2 - the first, second multipliers;

12-1, 12-2, 12-3 - the first, second, third differentiators;

13 - switch;

14-1 - the first divider;

15 - memory block;

16 is an OR diagram;

17-1, 17-2, 17-3 - the first, second, third block of comparison;

18 - functional converter;

19 - sensor angle of slip;

20 - angle deviation angle sensor;

21 - aileron deflection angle sensor;

22 - temperature sensor;

23-1 - the first adder;

24 is a subtraction scheme;

25 - block determining the moment of roll.

Figure 2 shows the structural diagram of the block 25 determining the moment of heel according to claim 2 of the formula, where the designations are:

25 - block determining the moment of roll;

26 - control unit;

27 - roll moment calculator.

Figure 3 shows the structural diagram of the calculator 27 determine the moment of heel according to claim 2 of the formula, where the following notation:

27 - roll moment calculator;

11-3, 11-4, 11-5, 11-6, 11-7, 11-8, 11-9, 11-10, 11-11, 11-12, 11-13, 11-14, 11- 15, 11-16, 11-17, 11-18, 11-19 - the third, fourth, fifth, sixth, seventh, eighth, ninth, tenth, eleventh, twelfth, thirteenth, fourteenth, fifteenth, sixteenth, seventeenth, eighteenth, nineteenth multipliers;

14-2, 14-3 - second, third dividers;

23-2 - the second adder;

28-1, 28-2, 28-3, 28-4, 28-5, 28-6, 28-7, 28-8, 28-9, 28-10, 28-11, 28-12, 28- 13, 28-14, 28-15, 28-16, 28-17, 28-18 - first, second, third, fourth, fifth, sixth, seventh, eighth, ninth, tenth, eleventh, twelfth, thirteenth, fourteenth, fifteenth, sixteenth, seventeenth, eighteenth blocks of variable coefficients.

Figure 4 shows the structural diagram of the switch 13 according to claim 1 of the formula, where the following notation:

13 - switch;

29-1, 29-2, 29-3 - the first, second, third pair of normally closed contacts.

Figure 5 shows the structural diagram of the control unit 26 according to claim 1 of the formula, where the following notation:

17-4, 17-5, 17-6, 17-7, 17-8, 17-9, 17-10 - the fourth, fifth, sixth, seventh, eighth, ninth, tenth comparison blocks;

30-1, 30-2, 30-3, 30-4 - the first, second, third, fourth schemes I.

Figure 6 shows the structural diagram of the block 28-1 ÷ 18 variable coefficient according to claim 3 of the formula, where the following notation:

23-3 - the third adder;

11-20 - the twentieth multiplier;

31 - source of a constant signal;

32-1, 32-2, 32-3 - the first, second, third amplifiers with a variable gain.

A device for flight control of the fuel weight of an airplane contains an angle of attack sensor 1, a Mach 2 sensor, a speed sensor 3, an altitude sensor 4, a 5 deflector angle sensor 5, a flap deflection angle sensor 6, a slat deflection angle sensor 7, a roll speed sensor 8, a sensor 9 pitch speed, yaw rate sensor 10, the output of which is connected to the inputs of the first 11-1 and second 11-2 multipliers, the second inputs of which are connected respectively to the outputs of the pitch speed sensor 9 and the roll speed sensor 8, the first differentiator 12-1, input for which is connected to the output of the roll speed sensor 8, a second differentiator 12-2, the input of which is connected to the output of the yaw rate sensor 10, a switch 13, the first input of which is connected to the output of the first divider 14-1, the first input connected to the first output of the switch 13, the second input of the switch 13 is connected to the output of the memory unit 15, the first input of which, like the third input of the switch 13, is connected to the output of the OR circuit 16, the inputs of which are connected to the outputs of the first 17-1 and second 17-2 of the comparison unit, the second output of the switch 13 connected to the second the input of the memory unit 15, the third output of the switch 13 is connected to the input of the functional transducer 18 and the input of the third differentiator 12-3, the output of which is connected to the third comparison unit 17-3, the roll moment determination unit 25, the first, second, third, fourth, fifth, the sixth, seventh, eighth, ninth, tenth, eleventh, twelfth and thirteenth inputs of which are connected respectively to the outputs of the sensor 1 of the angle of attack, sensor 2 of the M number, sensor 3 of speed, sensor 4 of height, sensor 5 of the deflection angle of the spoilers, sensor 6 of the deflection angle flaps, sensor 7 angles of deviation of slats, sensor 19 of the angle of slip, sensor 20 of angles of deviation of the rudder, sensor 21 of angles of deviation of ailerons, sensor 8 of heel speed, sensor 10 of yaw rate 10, temperature sensor 22, the output of block 25 for determining the moment of heel is connected to the first by the summing input of the adder 23-1 and the input of the comparison unit 17-2, the second subtracting input of the adder 23-1, like the first subtracting input of the subtraction circuit 24, is connected to the output of the first multiplier 12-1, the third subtracting input of the adder 23-1 is connected to the output tue of the multiplier 11-2, the fourth summing input of the adder 23-1 is connected to the output of the second differentiator 12-2, the output of the adder 23-1 is connected to the second input of the divider 14-1, the output of the first differentiator 12-1 is connected to the summing input of the subtraction circuit 24, the output of which is connected to the inputs of the first block 17-1 comparison and the fourth input of the switch 13.

The roll moment determination unit 25 comprises a control unit 26 and a roll moment calculator 27, the first, second, third, fourth, fifth inputs of which are connected to the first, second, third, fourth, fifth inputs of the roll moment determination unit 25, the sixth and seventh inputs of which connected respectively to the first and second inputs of the control unit 26, the output of which forms a bus and controls connected to the sixth input of the roll moment calculator 27, the seventh, eighth, ninth, tenth, eleventh, twelfth inputs of which are connected respectively with the eighth, ninth, tenth, eleventh, twelfth, thirteenth inputs of the roll moment determination unit 25, the output of which is the output of the roll moment calculator 27.

The calculator 27 of the roll moment contains eighteen blocks of variable coefficients 28-1 ÷ 18, seventeen multipliers 11-3 ÷ 19, two dividers 14-2 ÷ 3 and the adder 23-2. The first input of the roll moment calculator 27 is connected to the first inputs of the blocks 28-1, 28-2, 28-3, 28-4, 28-5, 28-6, 28-7, 28-8 variable coefficients, the second input of the moment calculator 27 the roll is connected to the first inputs of blocks 28-9, 28-10, 28-11, 28-12, 28-13, 28-14 variable coefficients and the second inputs of blocks 28-1, 28-4 variable coefficients, the third input of the calculator 27 of the moment the bank is connected to two inputs of the third multiplier 11-3, the first inputs of the second 14-2 and third 14-3 divider, the second inputs of which are connected respectively with the outputs of the fourth 11-4 and fifth 11-5 multipliers, the outputs of the second 14-2 and third 14-3 dividers are connected to the inputs of the second adder 23-2, the fourth input of the roll moment calculator 27 is connected to the second inputs of blocks 28-9, 28-10, 28-11, 28-12, 28-13 , 28-14 variable coefficients and the first input of the block 28-15 variable coefficients, the fifth input of the calculator 27 of the roll moment is connected to the first input of the block 28-16 variable coefficients, the second input of which, like the second inputs of the blocks 28-2, 28-3, 28-5, 28-6, 28-7, 28-8, 28-17, 28-18 variable coefficients, connected to the sixth input of the calculator 27 roll moment, and the output to the input of the sixth smart a life-ghost 11-6, the second input of which is connected to the output of the block 28-5 variable coefficients, and the output is connected to the input of the seventh multiplier 11-7, the second input of which is connected to the output of the block 28-12 variable coefficients, the output of the seventh multiplier 11-7 is connected to the input of the second adder 23-2, the seventh input of the roll moment calculator 27 is connected to the first input of the eighth multiplier 11-8, the second input of which is connected to the output of the variable coefficient block 28-11, and the output to the first input of the ninth multiplier 11-9, the second input which is connected to the output of the block 28-4 trans of the coefficients, the output of the ninth multiplier 11-9 is connected to the first input of the tenth multiplier 11-10, the second input of which is connected to the output of the block 28-3 variable coefficients, the output of the tenth multiplier 11-10 is connected to the input of the second adder 23-2, the eighth input of the calculator 27, the roll moment is connected to the first input of the variable coefficient block 28-17, the output of which is connected to the input of the eleventh multiplier 11-11, the second input of which is connected to the output of the variable coefficient block 28-14, the output of the eleventh multiplier 11-11 is connected to the first input m of the twelfth multiplier 11-12, the second input of which is connected to the output of the block of 28-8 variable coefficients, and the output is connected to the input of the second adder 23-2, the ninth input of the roll moment calculator 27 is connected to the first input of the block of 28-18 variable coefficients, the output of which connected to the thirteenth multiplier 11-13, the second input of which is connected to the output of the block 28-9 variable coefficients, the output of the thirteenth multiplier 11-13 is connected to the first input of the fourteenth multiplier 11-14, the second input of which is connected to the output of the block 28-7 variable coefficients, and the output is with the input of the second adder 23-2, the tenth input of the roll moment calculator 27 is connected to the first input of the fifteenth multiplier 11-15, the second input of which is connected with the output of the sixteenth multiplier 11-16, and the output is with the first input of the fifth multiplier 11-5 the second input of which is connected to the output of block 28-2 variable coefficients, the first input of the sixteenth multiplier 11-16 is connected to the output of block 28-1 of variable coefficients, the second input is to the output of block 28-10 of variable coefficients, the eleventh input of calculator 27 of the roll moment is connected with input m of the seventeenth multiplier 11-17, the second input of which is connected to the output of the block of 28-13 variable coefficients, and the output is connected to the first input of the fourth multiplier 11-4, the second input of the fourth multiplier 11-4 is connected to the output of the block of 28-6 variable coefficients, output the second adder 23-2 is connected to the input of the eighteenth multiplier 11-18, the second input of which is connected to the output of the third multiplier 11-3, and the output, as well as the output of the block 28-15 variable coefficients, the second input connected to the twelfth input of the calculator 27 roll moment, connected to inputs evyatnadtsatogo multiplier 11-19, the multiplier 11-19 nineteenth output is the output of the calculator 27, roll moment.

The switch 13 contains three pairs of normally closed contacts 29-1, 29-2, 29-3 so that the first input of the switch 13 is connected to the second and third output through the normally closed contacts 29-2, 29-3, and the fourth input of the switch 13 through normally closed contact 29-1 is connected to the first output of the switch 13, the third input is the control input of the synchronous switching of normally closed contacts 29-1, 29-2, 29-3, at which only the second input of the switch 13 is connected to its third output.

The control unit 26 contains seven comparison blocks 17-4 ÷ 10 and four circuits 30-1 ÷ 4 I. The first input of the control unit 26 is connected to the inputs of the comparison blocks 17-4, 17-5, 17-6, the second input of the control unit 26 is connected with the inputs of blocks 17-7, 17-8, 17-9, 17-10 comparison. The first input of the 30-1 AND circuit, like the first input of the 30-2 I circuit, is connected to the output of the comparison unit 17-4, the second input is connected to the output of the comparison unit 17-8, and the output is connected to the first output of the control bus. The second input of the circuit 30-2 And is connected to the output of the comparison unit 17-7, and the output is connected to the second output of the control bus. The first input of circuit 30-3 AND is connected to the output of comparison unit 17-5, the second input is connected to the output of comparison unit 17-9, and the output is connected to the third output of the control bus. The first input of circuit 30-4 AND is connected to the output of comparison unit 17-6, the second input is connected to the output of comparison unit 17-10, and the output is connected to the fourth output of the control bus.

Block 28-1 ÷ 18 variable coefficients contains a constant signal source 31, a multiplier 11-20, an adder 23-3 and amplifiers 32-1 ÷ 3 with variable gain. The first, second, third, fourth control inputs of the amplifiers 32-1 ÷ 3 are connected to the same inputs of the control bus and at the second input of the blocks 28-1 ÷ 18 variable coefficients. The fifth input of the variable gain amplifier 32-1 is connected to the output of the constant signal source 31, and the output is connected to the first input of the adder 23-3. The fifth input of the variable gain amplifier 32-2, like the inputs of the multiplier 11-20, is connected to the first input of the variable gain block 28-1 ÷ 18, and the output is connected to the second input of the adder 23-3. The fifth input of the amplifier 32-3 with a variable gain is connected to the output of the multiplier 11-20, and the output is connected to the third input of the adder 23-3, the output of which is the output of blocks 28-1 ÷ 18 variable coefficients.

The practical implementation of the device for flight control of the fuel weight of the aircraft is possible on the elements of the analog and digital circuitry base [12]. However, preference here should be given to additional software of a serial computing flight control system such as VSUP-85 of a complex of standard digital flight-navigation equipment KSCSPNO (for example, Tu-204/214, Il-96 aircraft) [13, p. 63; 14, p.141], which already contains all the necessary measuring means of the proposed device. So the sensor of the angle of attack 1, the sensor of 2 Mach numbers, the sensor of 3 speeds, the sensor of 4 heights, the sensor of 19 sliding angles, the sensor 22 of temperature are implemented in the SVS-85 air signal system, and the sensor 8 of the roll speed, sensor 9 of pitch speed, sensor 10 of speed yaw - in the strap-down laser inertial system I-42-1C on angular velocity sensors KM-11. The exchange of information between these systems, the sensor of 5 angles of deviation of the spoilers, the sensor of 6 angles of deviation of the flaps, the sensor of 7 angles of deviation of the slats, the sensor of 20 angles of deviation of the rudder, the sensor of 21 angles of deviation of the ailerons and the TsVM-80-40001 in KSKSPNO is carried out in a serial bipolar code asynchronously in in accordance with GOST 18977-79 and is presented on the control panel and display PUI-85M. In this case, the multipliers 11-1, 11-2, differentiators 12-1, 12-2, the switch 13, the divider 14-1, the memory block 15, the circuit 16 OR, the comparison blocks 17-1, 17-2, 17-3, a functional converter 18, an adder 23-1 and a roll moment determination unit 25 with a control unit 26 and a roll moment calculator 27 with their connections are implemented programmatically in the TsVM-80, which ensures high accuracy and reliability of the device. For a more modern integrated electronic display and signaling system KSEIS-148 and VSS-100 airplane navigation system (an-148 airplane) with VTs-3 computers, an information complex of high-speed parameters IKVSP, a strap-down system of course and vertical LCR-93, implementation of the device gives even higher results [15].

A device for flight control of the weight of the fuel of the aircraft works as follows. Signals proportional to angle of attack α, Mach number M, speed V, height H, angle δ _And deflection of the interceptors, angle δ _{Z of} deflection of the flaps, angle δ _{P of} deflection of the slats, angle β of slip, angle δ _{H of the} deviation of the rudder, angle δ of _{E of the} deviation ailerons, angular velocity ω _X roll, angular velocity ω _Y yaw and temperature T, respectively, are supplied to the first, second, third, fourth, fifth, sixth, seventh, eighth, ninth, tenth, eleventh, twelfth, thirteenth inputs of the moment determination unit 25 roll, where is the calculation the current value of the moment M _X roll acting on the aircraft along its longitudinal axis. A signal proportional to the magnitude of this moment, from the output of the roll moment determination unit 25, is supplied to the first summing input of the first adder 23-1 and the input of the comparison unit 17-2. At the same time, a signal proportional to the pitch angular velocity ω _Z is fed to one input of the first multiplier 11-1, the other input of which receives a signal proportional to the yaw rate ω _{Y from the} yaw rate sensor 10, and a signal proportional to ω _Z ω is generated at the output _Y It arrives at the second subtracting input of the first adder 23-1 and the first subtracting input of the subtraction circuit 24. The signal proportional to the angular velocity ω _{X of the} roll from the roll speed sensor 8 is supplied to the first differentiator 12-1 and one input of the second multiplier 11-2, the other input of which receives a signal proportional to the angular velocity ω _Y from the yaw rate sensor 10. The output of the multiplier 11-2 produces a signal proportional to ω _X ω _Y , which is fed to the third subtracting input of the adder 23-1, the fourth summing input of which receives a signal differentiated in the differentiator 12-2, proportional to the angular velocity ω _Y. So at the output of the adder 23-1, a signal proportional

where (I _Z -I _Y ) ₀ is the difference between the moments of inertia I _Z and I _{Y of the} aircraft independent of the weight m _{T of} fuel; I _XY0 is the centrifugal moment of inertia I _{XY of the} aircraft independent of the weight m _{T of} fuel. The relation is obtained from expression (1) when taking into account the dependence of its moments of inertia on the weight m _{T of} fuel, which is known for the type of aircraft. According to expression (1), the part depending on the weight m _{T of} fuel is also determined in the form

, что проверяется в блоке 17-1 сравнения, то сигнал, пройдя через четвертый вход и первый выход коммутатора 13, поступает на первый вход делителя 14-1. Если одновременно выполняется и условие M_X≥M_Xmin, которое проверяется в блоке 17-2 сравнения, то делитель 14-1 выполняет деление сигнала с выхода сумматора 23-1, пропорционального выражению (17), на сигнал, имеющийся на первом входе делителя 14-1 так, что получается сигнал, пропорциональный моменту инерцииA signal proportional to the angular accelerations in parentheses is obtained at the output of the subtraction circuit 24. In the event that this signal is less

that is checked in block 17-1 comparison, the signal, passing through the fourth input and the first output of the switch 13, is fed to the first input of the divider 14-1. If the condition M _X ≥M _Xmin , which is checked in comparison block 17-2, is simultaneously satisfied, then divider 14-1 divides the signal from the output of adder 23-1, which is proportional to expression (17), by the signal available at the first input of divider 14 -1 so that a signal is obtained proportional to the moment of inertia

. При симметричной выработке топлива из правых и левых крыльевых баков самолета отрицательная скорость изменения момента инерции İ_x должна быть ограничена весьма малой величиной. Несимметричная выработка топлива из-за отказа топливного насоса, нарушения герметичности баков, неисправности топливной системы одного из двигателей, приводящие к неравномерному потреблению топлива и нарушению центровки, приводят к повышению момента инерции I_X, изменению величины и знака его производной, что и фиксируется блоком 17-3 сравнения. Контроль веса топлива на выходе функционального преобразователя 18 производится как в автоматическом, так и полуатоматическом режиме, когда ведется визуальное сопоставление пилотом показаний заявляемого устройства с топливомером или суммирующим расходомером, фиксируемое бортовым регистратором.From the output of the divider 14-1, it enters the first input of the switch 13 and through its normally closed contacts to the second and third outputs. From the second output of the switch 13, the signal is supplied to the second signal input of the memory unit 15, where it is repeated without storing, since there is no signal at the first control input of the memory unit 15. From the third output of the switch 13, the signal is fed to a functional converter 18 that implements the inverse function (8) of the dependence of the moment of inertia I _X (m _T ) on the weight of the fuel. At the same time, the signal from the third output of the switch 13 through the differentiator 12-3 is fed to the input of the comparison unit 17-3. At the same time, the condition of a change in the moment of inertia limited in speed of normal fuel production is checked

. In case of symmetric fuel production from the right and left wing tanks of the aircraft, the negative rate of change of the moment of inertia İ _x should be limited to a very small value. Asymmetric fuel production due to fuel pump failure, tank leakage, fuel system malfunction of one of the engines, leading to uneven fuel consumption and misalignment, increase the moment of inertia I _X , change the magnitude and sign of its derivative, which is fixed by block 17 -3 comparisons. The weight control of the fuel at the output of the functional converter 18 is performed both in automatic and semi-automatic mode, when the pilot compares the readings of the inventive device with a fuel meter or accumulating flow meter, recorded by the on-board recorder.

If one or both of the conditions (5) is not satisfied, in the comparison blocks 17-1 and 17-2, a signal appears at the output of one or both comparison blocks 17-1 and / or 17-2, which, having passed the OR circuit 16, is fed to the first input of the block 15 memory and the third control input of the switch 13. At the same time, all contact pairs of the switch 13 are switched so that the signals of the divider 14-1 do not go to its third output, and hence the inputs of the functional Converter 18 and the comparison unit 17-3. The signal at the first control input of the memory unit 15 transfers it to the storage mode of the value of the signal received at its second signal input at the previous time. A normally open contact of the switch 13 connects its second input to the third output and the stored signal value proportional to the moment of inertia I _X (m _T ) calculated at the previous moment of time enters the functional converter 18 and through the differentiator 12-3 to the comparison unit 17-3 . The flight weight control device operates in memory mode. This mode saves the crew continuous information about the fuel weight m _T until the moment until conditions (5) are met again.

The duration of this mode depends on the accuracy of the measurement of parameters, the dynamics of the aircraft and its control by the pilot.

Block 25 determine the moment of heel works as follows. The signals of the sensor 1 of the angle of attack, sensor 2 of the Mach number, sensor 3 of speed, sensor 4 of height, sensor 5 of the angle of deviation of the spoilers are received respectively at the first, second, third, fourth, fifth inputs of the block 25 for determining the roll moment and the same inputs of the calculator 27 for the roll moment. At the same time, from the sensor 19 of the sliding angle sensor 20 of the angles of deviation of the rudder, sensor 21 of the angles of deviation of the ailerons, sensor 22 of the temperature signals are respectively transmitted to the eighth, ninth, tenth, thirteenth inputs of the roll moment determination unit 25, to the eleventh and twelfth inputs of which the outputs respectively signals of the roll speed sensor 8 and the yaw rate sensor 10. The signals from the eighth, ninth, tenth, eleventh, twelfth and thirteenth inputs of the roll moment determination unit 25 are supplied to the seventh, eighth, ninth, tenth, eleventh and twelfth inputs of the roll moment calculator 27, which are controlled by the signal u at its sixth input. The sixth input of the calculator 27 of the roll moment receives the control signal u via the control bus from the control unit 26. The signal u is generated in the control unit 26 for changing variable coefficients in the calculator 27 of the roll moment. To this end, the first and second inputs of the control unit 26 receive signals from the sixth and seventh inputs of the roll moment determination unit 25, respectively. They are obtained at the outputs of respectively the flap deflection angle sensor 6 and the flap deflection angle sensor 7. This takes into account the effect of changes in the aerodynamics of the aircraft during wing mechanization during takeoff, climb, cruising and landing.

, поступает на вход сумматора 23-2, аналогично на вход с выхода делителя 14-3 поступает сигнал, пропорциональный

The calculator 27 roll moment works as follows. A signal proportional to the angle of attack α from the first input of the calculator 27 of the roll moment is fed to the first inputs of blocks 28-1, 28-2, 28-3, 28-4, 28-5, 28-6, 28-7, 28-8 variable coefficients. In this case, the dependence of the components of the moment coefficient (10) on the angle of attack of the aircraft is taken into account. The signal proportional to the number M, from the second input of the calculator 27 of the roll moment is fed to the first inputs of variable coefficients 28-9, 28-10, 28-11, 28-12, 28-13, 28-14, the second inputs of blocks 28-1 , 28-4 variable coefficients. A signal proportional to V from the third input of the roll moment calculator 27 is supplied to both inputs of the 11-3 multiplier, the first inputs of the second 14-2 and the third 14-3 divider, the second inputs of which respectively receive the signals from the outputs of the multipliers 11-4, 11- 5. The output of the divider 14-2 signal proportional

, is fed to the input of the adder 23-2, similarly, a signal proportional to the input from the output of the divider 14-3

, поступает на вход сумматора 23-2. Сигнал, пропорциональный углу δ_Н отклонения руля направления, с восьмого входа вычислителя 27 момента крена поступает на первый вход блока 28-17 переменных коэффициентов. Выходной сигнал блока 28-17 переменных коэффициентов поступает на вход умножителя 11-11, на другой вход которого поступает выходной сигнал блока 28-14 переменных коэффициентов. Выходной сигнал умножителя 11-11 поступает на вход умножителя 11-12, на другой вход которого поступает выходной сигнал блока 28-8 переменных коэффициентов. Выходной сигнал умножителя 11-12, пропорциональный m_X.РН, поступает на вход сумматора 23-2. Сигнал, пропорциональный углу δ_Э отклонения элеронов, с девятого входа вычислителя 27 момента крена поступает на первый вход блока 28-18 переменных коэффициентов. Выходной сигнал блока 28-18 переменных коэффициентов поступает на вход умножителя 11-13, на другой вход которого поступает сигнал с выхода блока 28-9 переменных коэффициентов. Выходной сигнал умножителя 11-13 поступает на вход умножителя 11-14, на другой вход которого поступает выходной сигнал блока 28-7 переменных коэффициентов. Выходной сигнал умножителя 11-14, пропорциональный т_Х.ЭЛ, поступает на вход сумматора 23-2. Сигнал, пропорциональный скорости ω_X крена, с десятого входа вычислителя 27 момента крена поступает на вход умножителя 11-15, на другой вход которого поступает выходной сигнал умножителя 11-16. При этом на один вход умножителя 11-16 поступает выходной сигнал блока 28-1 переменных коэффициентов, а на другой - выходной сигнал блока 28-10 переменных коэффициентов. Выходной сигнал умножителя 11-15 поступает на вход умножителя 11-5, на другой вход которого поступает выходной сигнал блока 28-2 переменных коэффициентов. Далее выходной сигнал умножителя 11-5 через второй вход делителя 14-2 участвует в получении на входе сумматора 23-2 сигнала, пропорционального

Сигнал, пропорциональный скорости ω_Y рыскания, с одиннадцатого входа вычислителя 27 момента крена поступает на вход умножителя 11-17, на другой вход которого поступает выходной сигнал с блока 28-13 переменных коэффициентов. Выходной сигнал умножителя 11-17 поступает на один вход умножителя 11-4, на другой вход которого поступает выходной сигнал блока 28-6 переменных коэффициентов. Далее выходной сигнал умножителя 11-4 через второй вход делителя 14-2 обеспечивает получение на входе сумматора 23-2 сигнала, пропорционального

Выходной сигнал сумматора 23-2, пропорциональный коэффициенту момента m_X (10), поступает на вход умножителя 11-18, где он умножается на выходной сигнал умножителя 11-3 так, что выходной сигнал умножителя 11-18 пропорционален

Этот сигнал поступает на вход умножителя 11-9, на другой вход которого поступает выходной сигнал блока 28-15 переменных коэффициентов. На второй вход блока 28-15 поступает сигнал, пропорциональный температуре Т наружного воздуха на высоте Н полета самолета так, что выходной сигнал блока 28-15 переменных коэффициентов пропорционален плотности ρ воздуха на высоте полета. Выходной сигнал умножителя 11-19 пропорционален моменту крена M_X (9).The signal proportional to the height H, from the fourth input of the calculator 27 of the roll moment is fed to the second inputs of the blocks 28-9, 28-10, 28-11, 28-12, 28-13, 28-14 variable coefficients and the first input of the block 28-15 variable coefficients. The signal proportional to the angle δ _{and the} deviation of the interceptors from the fifth input of the calculator 27 of the roll moment is fed to the first input of the block 28-16 variable coefficients, the second input of which, like the second inputs of the blocks 28-2, 28-3, 28-5, 28-6, 28-7, 28-8, 28-17, 28-18 variable coefficients, a control signal u is received from the sixth input of the calculator 27 of the roll moment. The output signal of the block 28-16 variable coefficients is fed to the input of the multiplier 11-6, the second input of which receives the output signal of the block 28-5 variable coefficients. The output signal of the multiplier 11-6 is fed to the input of the multiplier 1-7, the other input of which receives the output signal of the block 28-12 variable coefficients. From the output of the multiplier 11-7, a signal proportional to m _X.IN is fed to the input of the adder 23-2. A signal proportional to the angle β of sliding from the seventh input of the calculator 27 of the roll moment is fed to the input of the multiplier 11-8, the other input of which receives the output signal of the block 28-11 variable coefficients. The output signal of the multiplier 11-8 is fed to the first input of the multiplier 11-9, the second input of which receives the output signal of the block 28-4 variable coefficients. From the output of the multiplier 11-9, the signal goes to the input of the multiplier 11-10, the other input of which receives the output signal of the block 28-3 variable coefficients. The output signal of the multiplier 11-10, proportional

arrives at the input of the adder 23-2. A signal proportional to the rudder angle δ _{H of} the rudder from the eighth input of the calculator 27 of the roll moment is fed to the first input of a block of 28-17 variable coefficients. The output signal of the block 28-17 variable coefficients is fed to the input of the multiplier 11-11, the other input of which receives the output signal of the block 28-14 variable coefficients. The output signal of the multiplier 11-11 is fed to the input of the multiplier 11-12, the other input of which receives the output signal of the block 28-8 variable coefficients. The output signal of the multiplier 11-12, proportional to m _X.RN , is fed to the input of the adder 23-2. A signal proportional to the aileron deviation angle δ _E from the ninth input of the calculator 27 of the roll moment is fed to the first input of a block of 28-18 variable coefficients. The output signal of the block 28-18 variable coefficients is fed to the input of the multiplier 11-13, the other input of which receives a signal from the output of the block 28-9 variable coefficients. The output signal of the multiplier 11-13 is fed to the input of the multiplier 11-14, the other input of which receives the output signal of the block 28-7 variable coefficients. The output signal of the multiplier 11-14, proportional to t _{X. EL} , is fed to the input of the adder 23-2. A signal proportional to the roll speed ω _X from the tenth input of the calculator 27 of the roll moment is fed to the input of the multiplier 11-15, the other input of which receives the output signal of the multiplier 11-16. In this case, the output signal of the block 28-1 of variable coefficients is supplied to one input of the multiplier 11-16, and the output signal of the block of 28-10 variable coefficients to the other. The output signal of the multiplier 11-15 is fed to the input of the multiplier 11-5, the other input of which receives the output signal of the block 28-2 variable coefficients. Next, the output signal of the multiplier 11-5 through the second input of the divider 14-2 is involved in receiving at the input of the adder 23-2 a signal proportional

A signal proportional to the yaw rate ω _Y , from the eleventh input of the calculator 27 of the roll moment, is fed to the input of the multiplier 11-17, the other input of which receives the output signal from the block 28-13 variable coefficients. The output signal of the multiplier 11-17 is fed to one input of the multiplier 11-4, the other input of which receives the output signal of the block 28-6 variable coefficients. Further, the output signal of the multiplier 11-4 through the second input of the divider 14-2 provides a signal proportional to the input of the adder 23-2

The output signal of the adder 23-2, proportional to the moment coefficient m _X (10), is input to the multiplier 11-18, where it is multiplied by the output of the multiplier 11-3 so that the output of the multiplier 11-18 is proportional

This signal is fed to the input of the multiplier 11-9, the other input of which receives the output signal of the block 28-15 variable coefficients. At the second input of block 28-15, a signal is proportional to the temperature T of the outdoor air at the flight height H of the aircraft so that the output signal of block 28-15 of variable coefficients is proportional to the density ρ of air at the height of flight. The output signal of the multiplier 11-19 is proportional to the roll moment M _X (9).

The switch 13 operates as follows. When condition (5) is fulfilled, normally closed contacts 29-1 provide a signal from the fourth input of the switch 13 to its first output. This ensures that the output signal of the subtraction circuit 24 is connected to the divider 14-1. At the same time, the signal from the first input of the switch 13 through normally closed contacts 29-2, 29-3 enters its second and third outputs. The output signal of the divider 14-1 is supplied to the functional Converter 18 and through the differentiator 12-3 to the comparison unit 17-3. The output signal from the memory unit 15 at the second input of the switch 13 is disconnected from the third output. If the condition (5) is not fulfilled, the signal from the OR circuit 16 comes to the third control input of the switch 13. There is a synchronous switching of normally closed contacts 29-1, 29-2, 29-3. This ensures that the output of the divider 14-1 is disconnected from the third output of the switch 13 and the output signal of the memory unit 15 is connected to it. The signal to the third output of the switch 13 comes from the second input through the normally open contacts 29-3.

The control unit 26 operates as follows. A signal proportional to the slope deviation angle δ _{P is} supplied to the first input of the control unit 26 and then to the inputs of the comparison blocks 17-4, 17-5, 17-6. Blocks 17-4, 17-5, 17-6 comparisons are configured respectively for three possible main ranges of values of the angles δ _P deviations of the wing slats of the wing mechanization: medium, zero and maximum. A signal proportional to the flap deflection angle δ _{Z is} fed to the second input of the control unit 26 and then to the inputs of the comparison blocks 17-7, 17-8, 17-9, 17-10. Blocks 17-9, 17-7, 17-8, 17-10 of comparison are configured respectively for four possible main ranges of values of angles δ ₃ deviations of the flaps of the wing mechanization: zero, minimum, average and maximum. The thresholds for the operation of blocks 17-4, 17-5, 17-6, 17-7, 17-8, 17-9, 17-10 of the comparison cover all ranges of possible values, respectively, of the angles δ _П , δ _{З of the} deviation of the wing slats and wing flaps. The main combinations of the angles δ _П , δ _{З of the} deviations of the slats and flaps are distinguished using schemes 30-1, 30-2, 30-3, 30-4 I. These combinations affect the values of the variable aerodynamic coefficients implemented in the calculator 27 of the roll moment. If the signal proportional to the slope deviation angle δ _P is in the average value range and triggers the comparison unit 17-4, and the signal proportional to the flap deviation angle δ _З is also in the average value range and triggers the block 17- 8 of the comparison, the output signals of the comparison blocks 17-4, 17-8, arriving at the inputs of the circuit 30-1 And, will lead to the appearance of a signal at its output. This signal is fed to the first output of the control bus u - the output of the control unit 26. The signal at the first output of the control bus u corresponds to the take-off configuration of the wing of the aircraft and the corresponding values of the variable aerodynamic coefficients in the calculator 27 roll moment. If the signal proportional to the slope deviation angle δ _P is in the average range of values and causes the comparison unit 17-4 to trigger, and the signal proportional to the flap deviation angle δ _{Z of} the flaps is in the minimum value range and triggers the comparison block 17-7, then the output signals of the comparison blocks 17-4, 17-7, having arrived at the inputs of the 30-2 AND circuit, will lead to the appearance of a signal at its output. This signal is fed to the second bus and control output — the output of the control unit 26. The signal at the second output of the control bus u corresponds to the configurations of the climb and the approach wing of the aircraft and the corresponding values of the variable aerodynamic coefficients in the calculator 27 roll moment. If the signal proportional to the slope deviation angle δ _P is in the minimum value range and causes the comparison unit 17-5 to be triggered, and the signal proportional to the flap deviation angle δ _З is in the zero value range and triggers the comparison unit 17-9, then the output signals of the comparison blocks 17-5, 17-9, having arrived at the inputs of the 30-3 AND circuit, will lead to the appearance of a signal at its output. This signal is fed to the third bus and control output — the output of the control unit 26. The signal at the third output of the control bus u corresponds to the cruise configuration of the aircraft wing and the corresponding values of the variable aerodynamic coefficients in the calculator 27 of the roll moment. If the signal proportional to the slope deflection angle δ _P is in the maximum value range and causes the comparison unit 17-6 to be triggered, and the signal proportional to the flap deflection angle δ _З is also in the maximum value range and causes the comparison block 17-10 to fire , then the output signals of the comparison blocks 17-6, 17-10, coming to the inputs of the circuit 30-4 And, will lead to the appearance of a signal at its output. This signal is supplied to the fourth output of the control bus u - the output of the control unit 26. The signal at the fourth output of the control bus u corresponds to the landing configuration of the aircraft wing and the corresponding values of the variable aerodynamic coefficients in the calculator 27 of the roll moment.

Block 28-1, ... 28-18 variable coefficient works as follows. A signal proportional to the parameter α, H, M flight or the deviation angle of the aerodynamic surface is fed to the first signal input of the variable coefficient block 28-1, ..., 28-18 and then to the fifth input of the amplifier 32-2 with a variable gain and the multiplier 11- twenty. The first, second, third, fourth control inputs of the amplifiers 32-1, 32-2, 32-3 with a variable gain on the control bus u receive signals from the second control input of the block 28-1, ..., 28-18 of a variable coefficient. They are formed in the control unit 26 and depend on the configuration of the wing. At the fifth signal input of the amplifier 32-1 with a variable gain, a constant signal is supplied from the output of the constant signal source 31. Upon receipt of the control signal at one of the control inputs of the amplifier 32-1 with a variable gain, its output signal changes so as to correspond to the value a _{0, j of the} coefficient of the approximating polynomial (11) - (16). From the output of the amplifier 32-1 with a variable gain, the signal is fed to the first input of the adder 23-3. Synchronously, the output signal of an amplifier 32-2 with a variable gain will form a value a _{1, j} x of the coefficient of the approximating polynomial (11) - (16), which is fed to the second input of the adder 23-, depending on the parameter x and the control signal at one of the control inputs 3. Synchronously, according to the input signal at the fifth signal input of the 32-3 amplifier with a variable gain received from the output of the multiplier 11-20, a square-dependent parameter x ² and a control signal will be formed on one of the control inputs _{, the} coefficient a _{2, j} x ² coefficient approximating polynomial (11) - (16), which is fed to the third input of the adder 23-3. The adder 23-3 combines the three components of the approximating polynomial and at its output a signal is obtained proportional to the component of the variable moment coefficient.

As follows from the above, the achievement of the technical result of flight control of the fuel weight of the aircraft is ensured by introducing into the device adopted as a prototype and containing an angle of attack sensor, a Mach number sensor, a speed sensor, a height sensor, an interceptor deflection angle sensor, a flap deflection angle sensor, a deflection angle sensor slats, roll speed, yaw, pitch sensors, multiplier, switch, divider, memory block, OR circuit, comparison blocks, functional converter, new elements, such as sensors the glide angle, the rudder deflection angle sensor, the aileron deflection angle sensor, the temperature sensor, the roll moment determination unit with the roll moment calculator, the second multiplier and differentiator, the second adder for four inputs with the connections indicated in the claims. A comparison of the parameters characterizing the claimed invention and the prototype allows us to conclude that with a large number of matching elements and connections, it is not able to control the weight of the fuel. Analogs are able to control the weight of the fuel, but their reliable operation is ensured only by checking bridge circuits, tracking systems and a pointer. In the most trusting scheme of a complex fuel meter with a flow meter, fuel weight control is achieved by a significant complication of the device, the use of a large number of sensors in fire hazardous compartments of the fuel system. The most reliable and accurate ground check with the help of test equipment, but it is lengthy and possible when dismantling sensors with empty aircraft tanks. It significantly complicates the operation of the aircraft. The accuracy of the claimed device to the greatest extent depends on achievable at the present stage the accuracy of the assessment of aerodynamic parameters, which is 3-5% [16, 17]. It is capable of reliably P4 (2) = 0.999600 for average sensors flight time of 2 hours for the sensors of a typical complex of standard digital flight and navigation equipment (KSCSPNO), to control the weight of the fuel. At the same time, the reliability of failure detection of a SUIT type fuel meter (P1 (2) = 0.998824) here will be Рд4 (2) = 0.746416. This corresponds to the average time of reliable control TD4 = 6.84 hours, which is longer than the flight time. A similar indicator of a complex fuel meter with a totalizing flowmeter Рд2 (2) = 0.674411 and the average time - ТД2 (2) = 5.08 hours. It is important that in the inventive device it is possible to reliably control precisely the weight (mass) of the fuel, and not its level in the tanks, as in classic fuel meters, with all known methodological errors and difficulties in identifying a fuel system failure [2, 4]. Weight control through the estimation of the moment of inertia here also allows one to evaluate the lateral alignment of the aircraft, the control of which, for example, due to fuel pump failures, is relevant for flight safety taking into account a number of accidents [5, 18-20]. The physical properties (density, temperature) of the fuel on the operability of the device here does not seem to be as important as in classic fuel meters that measure, for example, liquid gas [2]. It has the greatest completeness of coverage of the most important elements of the fuel system with the smallest weight, since sensors, data buses of the on-board network of KSCSPNO are used. When implemented on board, there is a parallel to the classical, completely independent from it chain of control of the weight of the fuel and the fuel system on a different, with respect to the verified, physical principle of its operation. The inventive device can reduce the weight of the fuel system, reduce the number of redundant sensors in the tanks and fuel lines, the accompanying electrical signal circuits on the wing. This increases fire safety, since aerometric, inertial sensors used in the device are remotely located from tanks and fuel lines, which, under conditions of increased vibrations, overloads, and elevated temperatures, are more likely to leak the more seams and mounting holes, including and for classic fuel gauges or for impellers of flowmeters. For long-range aircraft (such as TU-95, TU-142, the number of tanks is more than 100), the importance of reducing the weight of sensors and signal circuits, increasing the reliability of weight control with high fire safety is especially important. With the modern characteristics of a digital computer, the amount of memory, speed and complexity of software in high-level languages are not significant for the implementation of the device, and together with the prototype it allows creating an integrated system for optimizing flight modes of increased reliability and efficiency on the basis of the equipment already on board.

Thus, the above information proves that when implementing the claimed invention, the following conditions are met:

- a tool embodying the device of the invention in its implementation, is intended for use in aircraft and, in particular, for integrated control of the weight of the fuel control systems of passenger and transport aircraft of the classical scheme. It can be used to determine and control in flight the current weight of the aircraft fuel and its lateral alignment;

- for the claimed invention in the form described in the independent claim, the possibility of its implementation using the described or other means known prior to the filing date of the application has been confirmed;

- a tool embodying the claimed invention in its implementation, is able to provide the specified technical result.

Therefore, the claimed invention meets the condition of patentability "industrial applicability".

Information sources

1. Leshchiner LB, Ulyanov I.E. Design of aircraft fuel systems. M .: Mechanical Engineering, 1975, 344 p.

2. Vorobyov V.G., Glukhov V.V., Groholsky A.L. et al. Aviation devices and measuring systems. Training for high schools of GA. Ed. V.G. Vorobyov. M .: Transport, 1981, 391 p.

3. Golinkevich T.A. Reliability assessment of electronic equipment. M .: Sov. Radio, 1969, 176 pp.

4. Aviation devices. Ed. S.S.Dorofeeva. M .: Military. Publ., 1992, 496 p.

5. Orlov B.A. Notes of the test pilot. M .: Aviko Press, 1994, 171 p.

6. Smirnov N.N., Itskoviich A.A. Maintenance and repair of aviation equipment as of. M .: Transport, 1987, 272 p.

7. Technical operation of aviation equipment: Textbook. for universities / V.G. Vorobyov, V.D. Konstantinov, V.G. Denisov and others; under the editorship of V.G. Vorobyova - M .: Transport, 1990, 296 p.

8. Flight Dynamics of Transport Aircraft / Ed. A.Ya. Zhukova. M .: Transport, 1996, 326 p.

9. Butenin N.V., Lunts Ya.L., Merkin D.R. The course of theoretical mechanics. T.2. M .: Nauka, 1971, 464 p.

10. Flight tests of aircraft / K.K. Vasilchenko, V. A. Leonov, I. M. Pashkovsky, B. K. Poplavsky. M.: Mechanical Engineering, 1996, 720 p.

11. Katkov M.S. Continuous adaptive management systems with identifiers. M .: MNI, World of books, 1992, 386 p.

12. Analog and Digital Integrated Circuits: Reference Guide / Ed. S.V.Yakubovsky. M .: Radio and communications, 1984, 432 p.

13. Avionics of Russia. Encyclopedic Reference / Under the general. ed. S.D.Bodrunova. SPb. National Association of Aircraft Manufacturers, 1999, p. 780.

14. Altukhov V.Yu., Stadnik V.V. Gyroscopic instruments, automatic on-board aircraft control systems and their technical operation. M .: Engineering, 1991, 160 p.

15. Galunenko A.V., Godunov V.A., Groshev V.V., Panich V.P., Strelkov V.T. Aircraft navigation computing system // World of Avionics, 2005, No. 2, p. 48-50.

16. Zaitseva N.N. Passenger aircraft ERBAS INDUSTRY A310. M .: TsAGI, 1990, 115 p.

17. Zaitseva N.N. Passenger aircraft ERBAS INDUSTRY A320. M .: TsAGI, 1993, 104 pp.

18. www.nnews.ru/2001/7/5/russia/869.php3

19. http://vfnik.chat.ru/1990k.htm

20. http://vfnik.chat.ru/1995k.htm

Claims (3)

1. A device for flight control of the weight of the fuel of an airplane, comprising an angle of attack sensor, a Mach number sensor, a speed sensor, an altitude sensor, a deflector angle sensor, a flap deflection angle sensor, a slat deflection angle sensor, a roll speed sensor, a pitch speed sensor and a speed sensor yaw, the output of which is connected to the input of the multiplier, a differentiator, a switch, the first input of which is connected to the output of the divider, the first input connected to the first output of the switch, the second input of the switch is connected to the output of the memory unit, the first input of which, like the third input of the switch, is connected to the output of the OR circuit, the inputs of which are connected to the outputs of the first and second comparison unit, the second output of the switch is connected to the second input of the memory unit, the third output of the switch is connected to the input of the functional converter, a third comparison unit, characterized in that a slip angle sensor, a rudder angle sensor, aileron deviation angle sensor, a second multiplier, a second and third differentiators, a subtraction scheme are introduced into it , an adder for four inputs, a roll moment determination unit, the first, second, third, fourth, fifth, sixth, seventh, eighth, ninth, tenth, eleventh, twelfth and thirteenth inputs of which are connected respectively to the outputs of the angle of attack sensor, Mach number sensor, speed sensor, height sensor, deflector angle sensor, flap deflection angle sensor, slat deflection angle sensor, glide angle sensor, rudder deflection angle sensor, aileron deflection angle sensor, speed sensor roll, yaw rate sensor, temperature sensor, wherein the output of the yaw rate sensor is connected to the input of the second multiplier and the input of the second differentiator, the output of the roll speed sensor is connected to the first differentiator and the second multiplier, the output of the roll moment detection unit is connected to the first totalizing input of the adder and the input of the second comparison unit, the second subtracting input of the adder, like the first subtracting input of the subtraction circuit, is connected to the output of the first multiplier, the second input of which is connected to the output of the sensor with pitch axis, the third subtracting input of the adder is connected to the output of the second multiplier, the fourth summing input of the adder is connected to the output of the second differentiator, the output of the adder is connected to the second input of the divider, the output of the first differentiator is connected to the summing input of the subtraction circuit, the output of which is connected to the inputs of the first comparison unit and the fourth input of the switch, the input of the third differentiator is connected to the third output of the switch, and the output is connected to the input of the third comparison unit. 2. A device for flight control of the fuel weight of an airplane according to claim 1, characterized in that the roll moment determination unit comprises a control unit and a roll moment calculator, the first, second, third, fourth, fifth inputs of which are connected respectively to the first, second, third, the fourth, fifth inputs of the roll moment determination unit, the sixth and seventh inputs of which are connected respectively to the first and second inputs of the control unit, the output of which forms a control bus connected to the sixth input of the roll moment calculator, seventh, the eighth, ninth, tenth, eleventh, twelfth inputs of which are connected respectively with the eighth, ninth, tenth, eleventh, twelfth, thirteenth inputs of the roll moment determination unit, the output of which is the output of the roll moment calculator. 3. The device for flight control of the weight of the fuel of the aircraft according to claim 2, characterized in that the roll moment calculator contains eighteen blocks of variable coefficients, seventeen multipliers, two dividers and an adder so that the first input of the roll moment calculator is connected to the first inputs of the first, second, third, fourth, fifth, sixth, seventh, eighth blocks of variable coefficients, the second input of the roll moment calculator is connected to the first inputs of the ninth, tenth, eleventh, twelfth, thirteenth, fourteenth blocks of variable coefficients and second inputs of the first, fourth blocks of variable coefficients, the third input of the roll moment calculator is connected to two inputs of the third multiplier, the first inputs of the second and third divider, the second inputs of which are connected respectively to the outputs of the fourth and fifth multipliers, the outputs of the second and third divider are connected with the inputs of the second adder, the fourth input of the roll moment calculator is connected to the second inputs of the ninth, tenth, eleventh, twelfth, thirteenth, fourteen of the fifth block of variable coefficients and the first input of the fifteenth block of variable coefficients, the fifth input of the roll moment calculator is connected to the first input of the sixteenth block of variable coefficients, the second input of which, like the second inputs of the second, third, fifth, sixth, seventh, eighth blocks of variable coefficients, is connected with the sixth input of the roll moment calculator, and the output with the input of the sixth multiplier, the second input of which is connected to the output of the fifth block of variable coefficients, and the output with the input of the seventh multiplies spruce, the second input of which is connected to the output of the twelfth block of variable coefficients, the output of the seventh multiplier is connected to the input of the second adder, the seventh input of the roll moment calculator is connected to the first input of the eighth multiplier, the second input of which is connected to the output of the eleventh block of variable coefficients, and the output to the first the input of the ninth multiplier, the second input of which is connected to the output of the fourth block of variable coefficients, the output of the ninth multiplier is connected to the first input of the tenth multiplier, the second input to is connected to the output of the third block of variable coefficients, the output of the tenth multiplier is connected to the input of the second adder, the eighth input of the roll moment calculator is connected to the first input of the seventeenth block of variable coefficients, the second input of which is connected to the sixth input of the roll moment calculator, and the output to the input of the eleventh multiplier the second input of which is connected to the output of the fourteenth block of variable coefficients, the output of the eleventh multiplier is connected to the first input of the twelfth multiplier, the second input which is connected to the output of the eighth block of variable coefficients, and the output to the input of the second adder, the ninth input of the roll moment calculator is connected to the first input of the eighteenth block of variable coefficients, the second input of which is connected to the sixth input of the roll moment calculator, and the output to the thirteenth multiplier, the second the input of which is connected to the output of the ninth block of variable coefficients, the output of the thirteenth multiplier is connected to the first input of the fourteenth multiplier, the second input of which is connected to the output of the seventh block of variable coefficients, and the output is with the input of the second adder, the tenth input of the roll moment calculator is connected to the first input of the fifteenth multiplier, the second input of which is connected to the output of the sixteenth multiplier, and the output is with the first input of the fifth multiplier, the second input of which is connected to the output of the second block variable coefficients, the first input of the sixteenth multiplier is connected to the output of the first block of variable coefficients, the second input is the output of the tenth block of variable coefficients, the eleventh input of the calculator I of the roll moment is connected to the input of the seventeenth multiplier, the second input of which is connected to the output of the thirteenth block of variable coefficients, and the output is connected to the first input of the fourth multiplier, the second input of the fourth multiplier is connected to the output of the sixth block of variable coefficients, the output of the second adder is connected to the input of the eighteenth multiplier, the second input of which is connected to the output of the third multiplier, and the output, as well as the output of the fifteenth block of variable coefficients, is calculated by the second input connected to the twelfth input ator torque roll, connected to the input of the nineteenth multiplier, the multiplier output is output nineteenth roll moment calculator.

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