RU2644632C1 - Small-sized navigational complex - Google Patents

Abstract

FIELD: navigation.

SUBSTANCE: invention relates to the field of aircraft planes navigation using compound method of navigation, functionally combining inertial method of navigation, satellite way of navigation and air-speed way of navigation, and also relates to the navigators for the control and management of aircraft planes. Proposed small-sized navigation system contains a satellite navigation system (SNS) receiver, an integrated sensor box, air signal system (ASS), three-component magnetometric sensor (MS), limit switch of strut compression, cluster for determining the reliability of SNS signals, a computing section, range unit of deviation ratio of MS, range unit of heading of aircraft, cluster of reserve navigation, adaptation unit to bumpiness.

EFFECT: technical result achieved at implementation of the proposed device, is to reduce the time of initial exhibition, increase reliability and increase the accuracy of determining the true track.

1 cl, 1 tbl, 1 dwg

Description

The invention relates to the field of navigation of aircraft (LA) using an integrated navigation method that functionally combines an inertial navigation method, a satellite navigation method and an air-speed navigation method, as well as navigation devices for monitoring and controlling aircraft.

Known small-sized strapdown inertial navigation system (SINS) of medium accuracy, adjustable from the air signal system, described in patent No. 2502049, publ. 12/20/2013, bull. No. 35, adopted as a prototype.

The system contains, in particular, a block of sensitive elements (CE) of medium accuracy, consisting of three accelerometers and three angular velocity sensors along three orthogonal axes, two computing platforms, an airborne signal system, a heading error detection unit, and a wind speed determining unit associated with the system satellite navigation.

Medium-precision SINSs are implemented by two computational navigation platforms, each of which has its own control law (damping of inertial errors), which depends on the parameters of the carrier motion, namely, on the components of the horizontal accelerations of the carrier. In this case, the first platform provides the calculation of the pitch and roll angles of the carrier, while the second - the course angle and the projection of the projections of the carrier speeds and its geographical coordinates, taking into account predefined and saved estimates of the wind speed and its direction. Each platform has its own control law. One of them is a traditional unperturbed computing platform, but with its own acceleration damping. The second performs error damping by the difference in the speed readings measured by the INS inertial navigation system and the air signal system (SHS). In this case, preliminary, in the presence of SNA signals, the error of not showing the ANN in azimuth, the speed and direction of the wind are determined.

In the known invention, insufficient attention has been paid to questions of the reliability of information and the reliability of a small navigation complex (MNC). The operating experience of the SNA shows that in practice inertial-satellite complexes use inaccurate information from the SNA receivers and, as a result, the complex malfunctions. It should be noted that determining the course of an aircraft by the signals of the SNA and SHS does not give a sufficiently high accuracy, which leads to the appearance of an additional error when calculating the coordinates of the aircraft. Failure of the navigation system or power sources leads to the loss of all navigation measurements.

The aim of the invention is to expand the functionality, increase the reliability and improve the accuracy of determining navigation parameters in the event of loss of signals from the receiver of the SNA.

This goal is achieved through the use of MNCs containing an integrated sensor unit (IDB), an airborne signal system (AHS), a computing unit and a satellite navigation system (PSNS) receiver, in which a three-component magnetometric sensor (TMD), a landing gear compression limit switch, a unit for determining the reliability of SNA signals, a unit for determining deviation coefficients (BOKD), a unit for determining the course (BOK) of an aircraft, a backup navigation unit (BRN), a turbulence adaptation unit (BAP), a second output d PSNS, the output of the IDB and the first output of the SHS are connected to the first second and third inputs of the computing unit, the second output of the PSNS is connected to the input of the unit for determining the reliability of the SNA signals, the series-connected TMD and BOKD are connected to the fourth input of the computing unit, the first is connected to the fifth input of the computing unit BRN output, sequentially connected BOK connected to the sixth input of the computing unit, the landing gear compression end switch, the input of which is connected to the first PSNS output, connected to the seventh input the output of the unit for determining the reliability of the SNA signals, series-connected BAT and BRN are connected to the second output of the SHS, the second input of the BRN is connected to the second output of the BOKD, the output of the OLS in the main and alternative modes is the output of the computing unit, in standby mode - the second output of the BRN.

The invention is illustrated in the drawing, which presents a structural diagram of the inventive device.

The device contains a receiver 1 of a satellite navigation system (PSNS), an integrated unit of 2 sensors, a system of 3 air signals, a three-component magnetometric sensor 4, a limit switch 5 compression chassis rack, unit 6 determine the reliability of the signals of the SNA, computing unit 7, block 8 for determining the deviation coefficients MD , block 9 determining the course (SIDE) of the aircraft, block 10 backup navigation (BRN), block 11 adaptation to turbulence (BAP).

The output signals of the device are the angles of the spatial orientation of the aircraft, components of the speed of the SNA and geodetic latitude, longitude and altitude. The following is a description of the operation of MNCs.

In block 9 determining the course at the stage of taxiing and take-off, the path angle of the aircraft is determined by the SNA signals. The measured direction angle until the aircraft is separated from the runway is the true aircraft course at which the initial aircraft course is set:

where W _N and W _E are the ground speeds of the aircraft in the direction of the northern meridian and eastern parallel, obtained from the SNA.

The moment of separation of the aircraft from the runway is fixed by the limit switch of compression of the landing gear 5, which gives the command to fix the value of the course of the aircraft.

. Используя полученные измерения, определяют коэффициенты магнитной девиации магнитометрического датчика. Определение коэффициентов девиации, представляющих собой шесть параметров Пуассона, осуществляют способом, приведенным в работе [1]. При этом используют упрощенную модель, которая позволяет учесть влияние второй по значимости четвертной девиации на ошибку определения магнитного курса.After climbing, a complete circle is made over the airfield or in a predetermined area where there are no magnetic anomalies with a constant roll angle, during which, in block 8 for determining the MD deviation coefficients, the values of the measured magnetic course are stored using TMD 4 from

. Using the obtained measurements, the coefficients of the magnetic deviation of the magnetometric sensor are determined. Determination of deviation coefficients, which are six Poisson parameters, is carried out by the method described in [1]. In this case, a simplified model is used, which allows one to take into account the influence of the second most important quarter deviation on the error in determining the magnetic course.

Simplified Poisson Equations

Where

The coefficients p, q, r take into account the magnetic field of the magnetically solid iron, which forms a constant magnetic field of the aircraft and retains its magnetism due to the large coercive force. Coefficients a, b, d, e take into account the magnetic field of the magnetically soft iron aircraft, forming an alternating magnetic field, depending on the course, not retaining its magnetism due to the small coercive force and causing an alternating inductive magnetic field of the aircraft. The problem of determining and compensating for the deviation of magnetometric sensors is solved by the iteration method. A complete algorithm for determining and writing off deviation is given in [1].

In the main mode, in the computing unit 6 of the OLS, the signals from the receiver 1 of the SNA are used to correct the angles of spatial orientation, the values of the earth speeds and the location coordinates of the aircraft determined by the signals of the IDB 2. At the same time, minimizing the difference in speeds and location coordinates calculated using the SNA and the signals The ISD, using the open Kalman filtering scheme, estimate the errors in determining the coordinates, velocities, spatial orientation angles and the errors of inertial sensors of the ISD.

Estimated parameters included in the vector X:

where ΔB, ΔL, Δh are the errors in determining the geodetic latitude, longitude and altitude of the aircraft,

ΔV _N , ΔV _E , ΔV _U - errors in determining the projections of the ground speeds of the aircraft,

Ψ ₁ , Ψ ₂ , Ψ ₂ - errors in determining the angles of a computing (platform) coordinate system relative to the navigation SK,

Δ a _x , Δ a _y , Δ a _z , δ a _x , δ a _y , δ a _z - constant and random error components of linear acceleration sensors,

Δω _x , Δω _y , Δω _z , δω _x , δω _y , δω _z are constant and random components of the errors of the angular velocity sensors.

The expressions describing the Kalman Filter are given below:

In the process of operation of the SNA receiver in block 5, the verification of the reliability of the signals continuously monitors the reliability of the signals from the receivers of the SNA. Monitoring the output parameters of the SNA is carried out at two levels.

At the first level ("rough control") determine the latitude, longitude and altitude with a given accuracy. In this case, the thresholds for the coordinates are determined based on the area limited by the maximum possible range and altitude. Speed thresholds are controlled modulo speed, which should be within the operational range.

At the second level, control is carried out on a moving observation interval, where speed measurements are monitored using a shifting buffer of type “running line” BV, in which the last N measurements of the speed module are stored. When a new measurement arrives, the content is shifted and the variation of the velocity module is calculated, and if the variation of the velocity module exceeds the specified threshold δ | V |> Por (V), a malfunction indicator Pr = 1 is formed. Coordinate measurement control is carried out by determining the increment of the path, using the shifting buffer BD type "running line", which stores the last N measurements of the increment of the path. When a new dimension arrives, the content is shifted and the variation of the path increment is calculated. If the variation in the increment of the path increment exceeds a predetermined threshold δD> Por (D), a malfunction sign Pr = 1 is formed. The control of the issuance of the same parameter values from the SNA is carried out to “n” matches, upon reaching which a symptom of malfunction Pr = 1 is formed. A complete algorithm for determining the reliability of signals from the SNA receiver is given in [2].

At the same time, in calculating unit 7, projections of wind speed, airspeed, current coordinates based on airspeed signals and barometric altitude according to SHS 3 signals are determined. Algorithms for estimating airspeed errors and determining winds are given in [5].

An alternative mode is activated in the absence of reliable signals from the SNA receiver. In this case, in the FC residual equations, the values of the velocities and coordinates from the SNA are replaced by the currently corrected airspeed and altitude values. To calculate the earth's speed, the last calculated values of the projections of the wind speed at the time of switching off the SNA are used.

Calculation of wind speed:

Standby mode works in hot standby. During normal operation, the verification of the values of the spatial orientation angles is carried out according to the BRN signals, which has its own emergency power source. The course is determined by the signals from the magnetic sensor, taking into account the coefficients of magnetic deviation. The height is determined by the barometric height sensor BRN.

In the work of the BRN, two orientation models are used. Model 1 is designed to account for changes in the orientation of the apparatus at the sampling steps of gyroscope measurements. It is dynamic, deterministic and is presented in the parameters of Rodrigue Hamilton:

Here q _i is the quaternion of the orientation of the apparatus; q _ωi is the quaternion of angular velocities; i is the number of the discrete time instant; Δt is the measurement discretization step; q ₀ is the quaternion of orientation for the initial moment of time.

Quaternion q determines the orientation of the associated coordinate system of the apparatus relative to the inertial coordinate system, the role of which is played by the normal earth coordinate system. From the point of view of the theory of strapdown inertial systems, ratios are a one-step orientation algorithm.

Model 2 is designed to correct model 1. The state vector of model 2 takes into account three components - roll, pitch and speed of the device relative to the ground. Its state vector is subject to evaluation by the current values of the signals from the accelerometers. Model 2 is stochastic and static, since its state is taken into account for the current moment of discrete time t _i . Relations for model 2 have the form

.

.

- его априорное нормальное распределение; V - земная скорость; w_i - вектор возмущений; x_iq - вспомогательный вектор,Here x _i is the state vector;

- its a priori normal distribution; V is the earth's speed; w _i is the perturbation vector; x _iq is an auxiliary vector,

calculated according to the quaternion of the orientation of model 1 using the matrix A of rotation of the associated coordinate system relative to the inertial

ϑ = arcsin ( a (1,2)), γ = -acrtg ( a (3,2) / a (2,2)).

- априорное нормальное распределение вектора состояния модели 2 для момента времени t_i;

- апостериорное распределение, подлежащее оцениванию;

- априорное распределение для следующего момента времени t_i+1. Тогда алгоритм БРН, решаемый на одном интервале дискретизации измерений Δt, представляется в виде последовательности следующих шагов.We denote

- a priori normal distribution of the state vector of model 2 for time t _i ;

- a posterior distribution to be evaluated;

- a priori distribution for the next time t _{i + 1} . Then the BRN algorithm, solved on one measurement discretization interval Δt, is represented as a sequence of the following steps.

вектора состояния модели 2 с учетом

и текущих измерений гироскопов и акселерометров.Step 1. Determination of posterior density statistics

state vector of model 2 taking into account

and current measurements of gyroscopes and accelerometers.

.Step 2. Calculation of Estimation of Quaternion Orientation

.

для следующего момента дискретного времени.Step 3. Calculation of the a priori quaternion orientation

for the next moment of discrete time.

Step 4. Calculation of statistics of a priori density

Let us dwell on the features of performing the calculations in step 1 and step 4.

имеет вид известных соотношений байесовского оценивания вектора состояния по вектору его дискретных измерений.1. Definition of statistics of posterior density

has the form of well-known relations of Bayesian estimation of the state vector by the vector of its discrete measurements.

- оценка вектора наблюдений, которая имеет видHere

- assessment of the observation vector, which has the form

включает в себя расчет априорного математического ожидания и ковариационной матрицы. Априорное математическое ожидание компоненты скорости определяется с учетом допущения о постоянстве скорости и принимается равным ее оценке:

. Априорные математические ожидания тангажа и крена

,

вычисляются по кватерниону

Априорная ковариационная матрица приближенно принимается равной апостериорной ковариационной матрице:

. Применение более сложных соотношений для ее расчета представляется неоправданным в силу приближенности модели 2.4. Calculation of statistics of a priori density

includes calculation of a priori mathematical expectation and covariance matrix. The a priori expectation of the velocity component is determined taking into account the assumption that the velocity is constant and is taken equal to its estimate:

. A priori mathematical expectations of pitch and roll

,

calculated by quaternion

The a priori covariance matrix is approximately taken equal to the posterior covariance matrix:

. The use of more complex relations for its calculation seems unjustified due to the proximity of model 2.

These calculations are performed at each step Δt as new measurements of gyroscopes and accelerometers arrive. In standby mode, the aircraft heading, roll, pitch and barometric altitude are determined. Complete algorithms of adaptive inertial course-of-flight, the work of the BRN are given in [3, 4].

In block 11, the adaptation to turbulence is optimized. Full expressions for overloads depending on aircraft flight parameters:

The degree of influence of the terms in the right-hand sides of (1) depends on the flight mode. In a straight horizontal flight with constant speed, n _x = sin (ϑ), n _y = cos (ϑ) cos (γ), n _z = -cos (ϑ) sin (γ). These ratios are used in the simplest correction options when the accelerations created by the aircraft are much less than gravitational ones.

The terms (V _z ω _y -ω _z V _y ) / g, (V _x ω _z -ω _x V _z ) / g, (V _y ω _x -ω _y V _x ) / g are due to the appearance of Coriolis forces and are significant for U-turns.

,

,

имеют значимость при появлении линейных ускорений по связанным осям ЛА.Terms

,

,

are significant when linear accelerations appear along the associated aircraft axes.

At relatively small angles of attack and slip, the speed is directed mainly along the construction axis.

,

,

Then (1) is presented in a simplified form

n _y = cos (ϑ) cos (γ) + Vω _z / g,

n _z = -cos (ϑ) sin (γ) -ω _y V / g.

Any deviation of the overload modulus from unity and any inaccuracy (1) is taken into account by a decrease in confidence in the observations of overloads by increasing variances in the covariance matrix R _i , in accordance with a given law. In this case, the correction intensity decreases, and, consequently, its errors are reduced.

In acceleration and braking modes, the derivative of speed makes a significant contribution to the change in apparent acceleration.

Given that the change in the speed of the aircraft relates to long-period motion, the derivative of the speed is distinguished from the speed estimate using a smoothing low-pass filter

.

.

Here T is the low-pass filter time constant.

в выражении для

влияет на оценивание тангажа.From the consideration of Jacobian H _i it follows that the estimation of the speed of the aircraft occurs when performing turns, when one or both angular velocities ω _z , ω _y are present. Moreover, the terms ω _x V / g, -ω _y V / g in the ratios for n _y , n _z provide an estimate of the roll. It follows from the first equation in (1) that the term

in the expression for

affects pitch estimation.

для повышения точности оценивания тангажа.In the flight areas during take-off and landing, when the speed changes most intensively, and the aircraft does not perform turns, the speed is not estimated, but

to increase the accuracy of pitch estimation.

To this end, use the measurement of speed by the system of air signals (SHS). SHS generates a true airspeed measurement V _TA .

выполняют аналогично (2) с помощью фильтра нижних частот в дифференцирующем устройстве:Modern SHS have a fairly high accuracy. Derivation

perform similarly (2) using a low-pass filter in a differentiating device:

.Note that in this case, constant mismatches between the earthly and true airspeed do not introduce errors into the definition of the derivative. Dynamic measurement errors V _TA at low altitudes of takeoff and landing are negligible. SHS noise errors are smoothed out by the filter. Therefore, with sufficient accuracy, it is legitimate to put

.

Additional optimization of the filter coefficients is performed taking into account the level of sensor errors. By analyzing the errors of inertial sensors using the spectral power density and Alan dispersion, quantization noise, random walk (drift), instability of zero bias (flicker noise), random walk (speed drift), multiplicative systematic error, and sinusoidal noise are distinguished.

Considering the fact that the main contribution to the orientation errors of the strapdown BRN is made by the displacements of the zeros of the gyroscopes, the adjustment of the Kalman filter coefficients is performed on the set of training sequences formed for a set of combinations of displacement signs.

The filter coefficients are optimized as follows.

For each flight, nine training sequences are formed.

Variants of the signs of the displacements of the zeros of the gyroscopes are presented in table 1, where c ₀ is the absolute value of the displacement. The value of c ₀ is set taking into account the accuracy class of disposable gyroscopes.

,

- коэффициенты нелинейной функции

, определяющей диагональные элементы ковариационной матрицы ошибок наблюдения R.In total, there are six coefficients to be tuned in the BRN algorithm: q ₁ , q ₂ , q ₃ - diagonal elements of the covariance matrix of perturbations Q, and a = d ₁ ,

,

are the coefficients of the nonlinear function

defining the diagonal elements of the covariance matrix of observation errors R.

The quality criterion J ₁ is the weighted mean square error of roll and pitch orientation, averaged over time and over the set of all nine training sequences.

- среднеквадратическая ошибка оценивания тангажа:

- среднеквадратическая ошибка оценивания крена: α_ϑ=0.5 и α_γ=0.5 - весовые коэффициенты; J={a, b, k, q₁, q₂, q₃} - множество из шести искомых коэффициентов алгоритма.Here

- standard error of pitch estimation:

- the standard error of the roll estimation: α _оцени = 0.5 and α _γ = 0.5 - weight coefficients; J = { a , b, k, q ₁ , q ₂ , q ₃ } is the set of six desired coefficients of the algorithm.

на скользящем интервале небольшой длины, порядка 1-2 секунд. При превышении некоторого заданного порога

принимают решение о наличии турбулентности.The identification of turbulence, in the sense of revealing its presence, is performed by the standard deviation

on a sliding interval of small length, about 1-2 seconds. If a certain threshold is exceeded

decide on the presence of turbulence.

The optimization of the coefficients is carried out in three stages.

.1. Numerical minimization criterion of quality minJ ₁ and determining the coefficients {a _1, b _1, k _1, q _11, q _21, q _31} for flight in a calm atmosphere,

.

.2. Numerical minimization of the quality criterion minJ ₁ and determination of the coefficients { a ₂ , b ₂ , k ₂ , q ₁₂ , q ₂₂ , q ₃₂ } for flights in turbulence conditions,

.

3. Definition of the procedure for calculating the dispersions d, q ₁ , q ₂ , q ₃ , satisfying with sufficient accuracy flights in a calm atmosphere, and in turbulence.

,

.Most simply, this procedure is implemented by linear interpolation of the filter coefficients according to the results of steps 1, 2 and current values

,

.

The output signals of the roll, course and pitch received from the BRN operating in the hot reserve are compared with similar signals from the OLS operating in the alternative mode, which allows additional control over the accuracy and reliability of the obtained spatial orientation angles.

Small-sized navigation system can be used both on highly maneuverable and on low-maneuverable aircraft. Reducing the time of the initial exhibition is achieved by determining the course of the aircraft during taxiing and take-off. The increase in reliability occurs due to the use of SHS in case of loss of signals from the SNA and the use of a backup navigation system, as well as checking the reliability of signals from the receiver of the SNA and checking orientation angles by comparing the signals of the inertial system with the signals of the corners of the backup system. Improving the accuracy of determining the true course is ensured by writing off the deviation in flight and constant adaptive correction of the roll and pitch angles according to signals from accelerometers, as well as by optimizing the filtration coefficients depending on turbulence, and by tuning on a variety of training sequences. Note that the cancellation of deviation can be carried out periodically, in the absence of replacement of onboard and overhead equipment of the aircraft.

The proposed technical device can be implemented by using basic elements for computer systems of on-board digital computers and basic elements used in existing NK aircraft.

Literature

1. Patent No. 2550774. A method for determining and compensating deviation of magnetometric sensors and a device for its implementation, IPC G01C 21/08, bull. No. 13, 2015. Authors Zaets V.F., Kulabukhov V.S., Kachanov B.O., Tuktarev N.A., Grishin D.V.

2. Patent No. 2585051. A method for monitoring data from satellite navigation systems and a device for its implementation, IPC G01S 19/08. Bull. No. 15, 2015. Authors Zaets V.F., Kulabukhov V.S., Kachanov B.O., Tuktarev N.A.

3. Patent No. 2564380. Correction method for strapdown inertial navigation system, IPC G01C 21/06. Bull. No27, 2015. Authors Zaets V.F., Kulabukhov V.S., Kachanov B.O., Tuktarev N.A., Grishin D.V.

4. Patent No. 2564379. Strap-on inertial kursertical, IPC G01C 21/16, bull. No27, 2015. Authors Zaets V.F., Kulabukhov V.S., Kachanov B.O., Tuktarev N.A.

5. Patent No. 2579550. A method for determining the error of measuring air speed and a device for its implementation, IPC G01P 21/00. Bull. No. 10, 2016. Authors Zaets V.F., Korsun O.N., Kulabukhov V.S., Tuktarev N.A., Lysyuk O.P.

Claims (1)

A small-sized navigation complex (MNC) containing an integrated sensor unit (IDB), an airborne signal system (AHS), a computing unit and a satellite navigation system (PSNS) receiver, characterized in that a three-component magnetometric sensor (TMD), an end switch landing gear compression, SNA signal reliability determination unit, deviation coefficient determination unit (BOKD), aircraft heading determination unit (BOK), backup navigation unit (BRN), turbulence adaptation unit (BAP), sec the first PSNS output, the IDB output and the first SHS output are connected to the first second and third inputs of the computing unit, the second PSNS output is connected to the input of the SNA signal reliability determination unit, the TMD and BOKD are connected in series to the fourth input of the computing unit, and the fifth input of the computing unit is connected the first BRN output, to the sixth input of the computing unit are connected serially connected BOK, the limit switch of the compression of the landing gear, the input of which is connected to the first output of the PSNS, to the seventh input under the output of the unit for determining the reliability of SNA signals is switched on, serially connected BAT and BRN are connected to the second output of the SHS, the second BRN input is connected to the second output of the BOKD, the output of the OLS in the main and alternative modes is the output of the computing unit, in standby mode the second output of the BRN.

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