RU2553776C1 - Method of determining angular position of controlled device hanged to aircraft carrier - Google Patents

Abstract

FIELD: instrument engineering.

SUBSTANCE: invention relates to the field of instrument engineering and can be used in placement of strap-down inertial navigation systems (SINS) of the controlled device (CD). The method comprises preliminary processing of signals corresponding to the angular velocities measured by the three-component angular velocity sensors (AVS) mounted on the DC and the aircraft carrier (AC), on AC manoeuvre of "snake" type, based on the method of Kalman filtration. Upon entering to the computer CD of the appropriate command AC performs a manoeuvre of "snake" type, while executing of matching of vectors of measurements AVC CD and AVC AC starts, on which the weighted mean square error of misalignment of the measurements AVC CD regarding measurements of AVC AC minimises. The alignment process is recurrent. At each step the current measurements of sensor signals are used, and the next approximation of the rotation matrix is calculated, which determines the misalignment of axes of the block of AVC CD relative to the axes of the block of AVC AC. In case of the absence of measurement errors the process terminates with complete alignment of the axes of the blocks in which the measurements of AVC CD are accurately translated into the measurements of AVC AC. According to the final rotation matrix the resulting assessments of misalignment angles between the axes of AVC CD and AVC AC are calculated, which determine the orientation of the product relative to the carrier.

EFFECT: invention enables to alert CD in a short time, in particular not exceeding 5 seconds, with the required accuracy.

Description

The invention relates to the field of instrumentation and can be used in the exhibition of strapdown inertial navigation systems of a controlled apparatus (BINS UA).

The known method (patent RU 2348010 C1, IPC G01C 21/16, publ. 02/27/2009) determine the initial exhibition of the instrument coordinate system of the strapdown inertial block (BIB) of a controlled object installed on the launcher (PU), relative to the base (starting) coordinate system , materialized stabilized course vertical (KB) platform, also installed on the PU. The essence of this method lies in the fact that carry out the deployment of the launcher with BIB and expose it to the first given elevation and azimuth angles; measure the BIB accelerometers in a position of the launcher in a relatively inactive position relative to the Earth on a time interval from t0 to tn; determine in the computing device (WU) n increments of each of the projections of the apparent velocity vector (VKS) on the axis of the instrument coordinate system (UCS) for the known given time intervals from t0 to tj (j = 1, ..., n) and get the projection of the VKS on the axis CPM; evaluate each received projection of the VKS on the UCS axis and calculate the assessment of each projection of the apparent acceleration vector (ICU) on the UCS axis; according to the estimates of the projections of the VKU, the value of each projection of the VKU on the UCS axis is predicted at some given instant of time T1; carry out measurements with angle sensors of the vertical axis over the same time interval from t0 to tn and determine n values of each of the Euler angles; evaluate each of the Euler angles, according to the obtained estimates of the Euler angles, predict the value of the Euler angles at the same given time instant T1, from the obtained values of the Euler angles determine the angular position of the associated coordinate system (SSC) relative to the BSK; according to the readings of the accelerometers, the vertical direction determines the projection of the VKU on the axis of the BSK; the angular position of the CCK relative to the BSK and the projections of the VKU on the axis of the BSK determine the projection of the VKU on the axis of the SSK; then carry out a U-turn of the launcher with BIB and set it to the second predetermined elevation and azimuth angles; using the signals from the BIB accelerometers, as well as the signals of the accelerometers and angle sensors, repeating the same operations as at the first elevation and azimuth angles, the projection of the second VKU on the UCS and SSC axes at another specified time point T2 and the SSC angular position are determined in the WU relative to BSK; from the obtained values of the projections of two vector quantities - two VKUs on the axis of the UCS and the SSC, the angular position of the SSC relative to the SSC is determined in the WU and, taking into account the known angular position of the SSC with respect to the BSC, the above determination of the angular position of the SSC with respect to the BSC is performed; on the projections of the VCS on the UCS axis predicted for a given other moment of time T2, the angular position of the UCS relative to the horizon plane of the base (starting) coordinate system is specified in the WU.

The disadvantage of this method is the impossibility of using it for the SINS display of a controlled device in flight from the base ANN of the carrier.

There is also a method described in the article Savelyev V.M., Antonov D.A. Exhibition of strapdown inertial navigation system of an unmanned aerial vehicle on a moving base. Electronic journal "Transactions of Moscow Aviation Institute", issue number 45 (adopted by us as a prototype).

According to the known method, the initial exhibition of SINS of a controlled apparatus (UA), in particular an unmanned aerial vehicle (UAV), is carried out by joint processing by Kalman filtering of course angles, roll, pitch, as well as speeds and coordinates of the carrier aircraft (SN) and mounted on the external UAV suspension.

The disadvantage of this method is the long time, of the order of several minutes, required to estimate the mismatch angles between the axes of the connected coordinate systems of the SN and the UAV mounted on the external sling of the UAV, which is not acceptable for the SINS exhibition of a combat controlled vehicle.

The aim of the invention is the ability to determine the angle of the mismatch of the axes of the associated coordinate systems UA and SN in flight, which are used for the initial exhibition, in a short time (no more than 5 seconds) and with the required accuracy.

This goal is achieved due to the fact that according to the claimed method for determining the angular position of a controlled vehicle (UA) suspended from a carrier aircraft (SN) using the output signals of the strapdown inertial system (SINS) UA and SINS SN for joint processing by Kalman filtering, As the output signals of the SINS, signals corresponding to angular velocities measured by three-component angular velocity sensors (DLS) installed on UA and SN are used, and angular velocity measurements are made ny using CH maneuver type "snake", and based on them, by matching the angular velocity vector and the V CH, a determination error angle angular velocity vectors UA and CH.

To determine the orientation of the product relative to the carrier, i.e. axes, it is necessary to determine the rotation matrix, determined by the three angles of mismatch of the coordinate axes of the SINS UA and SN. Their determination is performed by matching the angular velocity vectors UA and SN.

To do this, in the BINS UA must receive signals measuring the angular velocity of the SN with a frequency of at least 20 Hz. At the same time, within 5 seconds the SN must perform a maneuver of the “snake” type, or its part, with an amplitude of angular velocities at the heading and roll of the order of 7-10 [deg / s].

The determination of the mismatch angles begins with the command about the beginning of the maneuver entering the calculator UA, and ends after 5 seconds. The correctness of the solution is controlled by the magnitude of the residual error of the mismatch of the angular velocity vectors.

The determination of the mismatch angles is performed by discrete estimation of the state vector consisting of three desired angles, for which an a priori normal distribution is specified. The main provisions of the method are as follows.

The determined state vector is taken as

Here ψ is the yaw axis mismatch angle, ϑ is the yaw axis mismatch angle, γ is the roll mismatch angle of the axes, i is the number of the discrete measurement moment of time, measured from the moment the maneuver begins.

Denote: ω _x1 , ω _y1 , ω _z1 - angular velocity of the carrier aircraft; ω _x2 , ω _y2 , ω _z2 are the angular velocities of the SINS UA. The vector of observations of the angular velocities of the SINS UA, denoted by Z, is connected with the angular velocity vector of the SN using the rotation matrix denoted by E:

Here V _i is the vector of measurement errors with the covariance matrix R.

The components of vector (1) are assumed to be random variables, and their change processes are Markov ones.

Discrete models of their measurement are described by stochastic difference equations of the first order.

Here T _ψ , T _ϑ , T _γ - correlation time constants; w _ψi , w _ϑi , w _γi are random processes of discrete white noise with a given covariance matrix Q

.

.

The matrix of the discrete model of the object has the form

The Jacobi matrix of the observation vector Z has the form

.

.

Here f _x , f _y , f _z are the functions indicated in (2).

. Процесс определения представляется рекуррентным алгоритмом, решаемым в реальном времени вычислителем БИНС УА.The task is to estimate vector (1) from observations (2) under the condition of given covariance matrices R, Q and an a priori distribution $${\text{õ}}_{\text{0}}\in \text{N {}{\overline{\text{x}}}_{\text{0}}\text{,}{\overline{\text{Ð}}}_{\text{0}}\text{}}$$

. The determination process is represented by a recurrent algorithm, solved in real time by the BINS UA calculator.

The process of determining the axis mismatch angles is presented step by step:

Step 0

и матриц R, Q.Setting a priori distribution: $${\text{õ}}_{\text{0}}\in \text{N {}{\overline{\text{x}}}_{\text{0}}\text{,}{\overline{\text{Ð}}}_{\text{0}}\text{}}$$

and matrices R, Q.

Step 1.

 Beginning of the filtration cycle at the command of the beginning of the maneuver.

Enter the next measurement reading DUS BINS UA and DUS BINS SN.

Step 2

The forecast of the mathematical expectation of vector (1) by solving equations (3) for w _ψi = 0, w _ϑi = 0, w _γi = 0.

Step 3

Formation of matrices F and H by (4) and (5).

Step 4

Prediction of the covariance matrix.

$$\overline{\text{P}}=F\widehat{P}{F}^T+\text{Q}$$

Step 5

Determination of filter gain.

$$K=\overline{P}{H}^T{(HP{H}^T+R)}^{-\mathrm{one}}$$

Step 6

Determination of the posterior covariance matrix.

. $$\widehat{P}=(I-KH)\overline{P}{(I-KH)}^T+KR{K}^T$$

.

Step 7

по (2) при V_i=0.Definition of an estimation of a vector of measurements $$\overline{Z}$$

by (2) for V _i = 0.

Step 8

и оценки вектора (1).Definition of residuals. $$Z-\overline{Z}$$

and estimates of the vector (1).

$$\widehat{x}=\overline{x}+K(Z-\overline{Z})$$

Step 9

Determination of the mismatch vector of angular velocities SN and SINS UA.

Step 10

Recursion of posterior distribution.

, $$\overline{\text{P}}=\widehat{P}$$

. $$\overline{\text{x}}=\widehat{x}$$

, $$\overline{\text{P}}=\widehat{P}$$

.

Step 11

Go to step 1, provided that the filtering time has not expired.

Step 12

на последней секунде маневра. В данном примере пороговое значение, по которому принимается решение о правильной оценке углов рассогласования осей достаточно принять равным 0.1 [град/с].Verification of the correctness of the solution by checking the smallness of the average module of the mismatch of angular velocities $$\left|\overline{\epsilon }\right|$$

at the last second of the maneuver. In this example, the threshold value by which a decision is made on the correct assessment of the misalignment angles of the axes is sufficient to be taken equal to 0.1 [deg / s].

Thus, the use of the invention allows to alert a controllable device in a short time, in particular not exceeding 5 seconds, with the required accuracy.

Claims (1)

A method for determining the angular position of a controlled apparatus (UA) suspended from a carrier aircraft (SN) using the output signals of the strapdown inertial system (SINS) of UA and SINS SN for joint processing by the Kalman filtering method, characterized in that the signals are output as SINS corresponding to the angular velocities measured by three-component angular velocity sensors (DLS) installed on UA and SN, and the angular velocities are measured using the snake type maneuver, and based on them, by matching the angular velocity vectors of UA and SN, determine the angles of mismatch of the angular velocity vectors of UA and SN.