RU2170410C1 - Flying vehicle navigation complex - Google Patents

Abstract

FIELD: navigational complexes of multi- mission fighters. SUBSTANCE: navigational complex includes inertial navigational system, satellite navigational system, navigational computer, navigations information recorder, internal navigational system error computer, storage unit, computer of corrections to velocity and azimuthal angle components of gyroplatform of inertial navigational system. Output of inertial navigational system is connected with first input of navigational computer and first input of recorder. Output of satellite navigational system is connected with second input of computer and second input or recorder. Output of recorder is connected with input inertial system error computer whose output is connected with of storage unit. Output of storage unit is connected with third input of navigational computer and input of velocity and azimuthal angle component correction computer whose output is connected with fourth input of computer. EFFECT: enhanced accuracy, trouble-free operation and informative capacity. 1 dwg

Description

 The invention relates to the field of navigation and can be used in navigation systems of aircraft, mainly multi-role fighters.

 Known navigation complex of the aircraft [1], containing a satellite navigation system (receiver SRNS), inertial navigation system (ANN), navigation computer (computer).

 A disadvantage of the known navigation complex is the lack of accuracy of the issued navigation data.

 The essence of the invention lies in the fact that the navigation system of the aircraft, comprising an inertial navigation system, satellite navigation system, navigation computer, further comprises a navigation information recorder, an error calculator for the inertial navigation system, non-volatile memory, a calculator for corrections to the components of speed and the azimuth angle of the gyro platform inertial navigation system.

 The output of the inertial navigation system is connected to the first input of the navigation calculator and the first input of the navigation information recorder, the output of the satellite navigation system is connected to the second input of the navigation calculator and the second input of the navigation information recorder, the output of the navigation information recorder is connected to the input of the inertial navigation system error calculator, the output is which is connected to the input of a non-volatile storage device, the output of which is connected third input and an input of the navigation calculator calculating corrections to the velocity components and azimuthal angle gyroplatform inertial navigation system, the output of which is connected to a fourth input of the calculator, whose output is the output 8 of navigation information.

The invention is illustrated in the drawing, on which are indicated:
1 - inertial navigation system;
2 - satellite navigation system;
3 - navigation computer;
4 - recorder of navigation information;
5 - error calculator inertial navigation system;
6 - non-volatile storage device;
7 - calculator corrections to the components of speed and azimuthal angle of the gyro platform inertial navigation system;
8 - output of navigation information.

 The invention does not depend on the type of inertial navigation system used. Improving the accuracy of determining navigation parameters is achieved by using systems with a built-in coordinate and heading calculator, and by using inertial course-and-track with free or corrected (semi-free) azimuth gyroplatform that do not have a coordinate calculator.

 Below, as an example, we consider the option of using an inertial heading vertical with an inertial gyro platform as an inertial navigation system 1.

 The navigation computer 3, the computer 5 errors of the inertial navigation system, the computer 7 amendments to the components of speed and the azimuthal angle of the gyro platform of the inertial navigation system are on-board electronic computers. The functions of the navigation calculator 3, calculator 5 errors of the inertial navigation system, calculator 7 corrections to the speed components and the azimuthal angle of the gyro platform of the inertial navigation system can also be performed by one on-board electronic computer.

 The navigation computer 3 is designed to carry out processing of the navigation information received from the inertial navigation system 1 and the satellite navigation system 2 during the flight of the aircraft. In this case, the data received from the inertial navigation system 1 is corrected according to the data of the satellite navigation system 2.

The result of processing the navigation information in the navigation computer 3 is the data on the geographical coordinates (latitude Φ and longitude λ) of the aircraft, its flight height (H), data on the components of the speeds of the aircraft (north V _N , east V _E , vertical V _H ), data on the course of the aircraft (Φ _East ).

 The navigational information recorder 4 is an information storage device into which data from the inertial and satellite navigation systems 1 and 2 are recorded throughout the flight. At the end of the flight, these data are used by the inaccuracy of the inertial navigation system calculator 5.

 Errors calculated by the inverter 5 inertial navigation system errors are recorded in non-volatile memory 6, in which this data is stored until the next flight.

 Calculator 7 corrections: to the velocity components and the azimuthal angle of the gyro platform of the inertial navigation system, based on the information about the errors of the inertial navigation system 1 stored in the non-volatile memory 6, it calculates the corrections that are received in the navigation calculator 3 and used to correct the data received from the inertial navigation system 1.

 The navigation system of the aircraft operates as follows.

Data on the components of the absolute speeds of the aircraft in the axes of the _{gyro platform} V _ξst , V _ηst , V _ξst , on the gyroscopic course ψ _g , on the roll angle γ, on the pitch angle υ from the output of the inertial navigation system 1 are received at the first inputs of the navigation calculator 3 and recorder 4 navigation information. Data on the geographical latitude Φ, on the geographical longitude λ, on the components of the relative speeds of the aircraft (aircraft speeds relative to the Earth) V _N (northern component), V _E (eastern component), V _H (vertical component) and data on the current time ( from the output of the satellite navigation system 2) go to the second inputs of the navigation computer 3 and the recorder 4 navigation information.

 The navigation computer 3, based on the data of the inertial navigation system 1, calculates the coordinates and flight altitude of the aircraft (integrates component speeds). Then, complex processing of the navigation data received from the inertial navigation system 1 and satellite navigation system 2 is performed. It is taken into account that the data of the satellite navigation system 2 are more accurate and their accuracy does not depend on the flight duration, but at the same time they contain high-frequency interference and there may be a situation of inoperability of the satellite navigation system 2 due to the presence of radio interference, antenna shading and other reasons. When the satellite navigation system 2 is inoperative, the operation of the navigation system is carried out only on the basis of data from the inertial navigation system 1.

 Throughout the flight, the navigation information recorder 4 records data received from the inertial navigation system 1 and from the satellite navigation system 2.

 The calculation of the coordinates and flight altitude of the aircraft (integration of component speeds) according to the inertial navigation system 1 is as follows.

где V_ξст,V_ηст - составляющие (по осям гироплатформы) абсолютной скорости, измеренные инерциальной навигационной системой 1.Before the flight of the aircraft, the navigation computer 3 calculates the initial value of the azimuthal angle of the gyro platform of the inertial navigation system (equal to the parking course of the aircraft) χ ₀ :

where V _ξst , V _ηst are the components (along the _{gyro platform} axes) of the absolute speed, measured by the inertial navigation system 1.

The components (along the _{gyro platform} axes) of the angular velocity of the Earth's rotation U _ξ0 , U _η0 , U _{ξ0 are calculated} :
U _ξ0 = UcosΦ ₀ cosχ ₀ ; (2)
U _η0 = UcosΦ ₀ sinχ ₀ ; (3)
U _0ζ = UsinΦ ₀ , (4)
where U is the angular velocity of the Earth’s rotation, Φ ₀ is the geographical latitude of the aircraft parking lot.

где e² = 0,0066934216.The initial values of the deviation of the horizontal axes of the gyro platform θ are calculated $\genfrac{}{}{0ex}{}{\text{M}}{\text{ξ0}}$ , θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{η0}}$ :

where e ² = 0.0066934216.

ω_ζф0= KV_ξст0-U_ζ0, (9)
где K - масштабный коэффициент, R - радиус Земли, на который настроена инерциальная навигационная система.The actual values of the compensated constant drift of the gyroscopes are calculated:

ω _ζφ0 = KV _ξst0 -U _ζ0 , (9)
where K is the scale factor, R is the radius of the Earth, which is tuned inertial navigation system.

ω_ζдоб0= -θ_η0U_ξ0+θ_ξ0U_η0, (12)
где R_ξ,R_η - радиусы кривизны земного эллипсоида по сечениям, проходящим через оси гироплатформы, R₀ - вспомогательный параметр.Additional drifts ω _ξdob0 , ω _ηdob0 , ω _ζdob0 , due to the methodological errors of the initial exhibition, are _calculated :

ω _{ζ add0} = _{-θ η0} U _ξ0 + θ _ξ0 U _η0 , (12)
where R _ξ , R _η are the radii of curvature of the earth's ellipsoid along the sections passing through the axis of the gyro platform, R ₀ is an auxiliary parameter.

где ΔV $\genfrac{}{}{0ex}{}{\text{M}}{\text{ξ}}$ ,ΔV $\genfrac{}{}{0ex}{}{\text{M}}{\text{η}}$ - погрешности (методические) инерциальной навигационной системы 1 в определении составляющих абсолютных скоростей, θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{ξ}}$ ,θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{η}}$ ,θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{ζ}}$ - погрешности (методические) инерциальной навигационной системы 1 в определении вертикали и курса.The corrections for the constant drift of the gyroscopes to the measured linear velocities ω _ξdr0 , ω _{ηdr0 are calculated} :
ω _ξдр0 = Rω _ηф0 ; (thirteen)
ω _ηдр0 = -Rω _ξф0 . (14)
The system of differential equations of gyro platform motion is solved under the influence of methodological errors:

where ΔV $\genfrac{}{}{0ex}{}{\text{M}}{\text{ξ}}$ , ΔV $\genfrac{}{}{0ex}{}{\text{M}}{\text{η}}$ - errors (methodological) of the inertial navigation system 1 in determining the components of the absolute speeds, θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{ξ}}$ , θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{η}}$ , θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{ζ}}$ - errors (methodological) inertial navigation system 1 in determining the vertical and heading.

The azimuthal gyroscope drift ω _{ζ0 is} calculated:
ω _ζ0 = ω _ζφ0 + ω _{ζ add0} . (16)
Moreover, the initial values of θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{ξ0}}$ , θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{η0}}$ parameters θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{ξ}}$ , θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{η}}$ are calculated by formulas (5) and (6), and the initial values ΔV $\genfrac{}{}{0ex}{}{\text{M}}{\text{ξ0}}$ , ΔV $\genfrac{}{}{0ex}{}{\text{M}}{\text{η0}}$ , θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{ζ0}}$ ΔV parameters $\genfrac{}{}{0ex}{}{\text{M}}{\text{ξ}}$ , ΔV $\genfrac{}{}{0ex}{}{\text{M}}{\text{η}}$ , θ $\genfrac{}{}{0ex}{}{\text{M}}{\text{ζ}}$ taken equal to zero.

Вычисляются абсолютные ускорения a_ξ,a_η,a_ζ :

где β = 0,05317,

= H - вертикальная скорость, a - большая полуось земного эллипсоида (a = 6378245 м).The absolute velocities for calculating V _ξ , V _η are calculated:
V _ξ = V _{ξst -ω} ξ to ₀ -ΔV $\genfrac{}{}{0ex}{}{\text{M}}{\text{ξ}}$ ; (17)
V _η = V _{ηst -ω} ηdob0 -ΔV $\genfrac{}{}{0ex}{}{\text{M}}{\text{η}}$ . (18)
The absolute angular velocities u _ξ , u _η are calculated:

The absolute accelerations a _ξ , a _η , a _ζ are calculated:

where β = 0.055317,

= H is the vertical velocity, a is the semimajor axis of the earth's ellipsoid (a = 6378245 m).

Вычисляются азимут гироплатформы χ и истинный курс φ_ИКВ :

φ_ИКВ= φ_г+χ. (28)
Вычисляются скорости в осях сопровождающего трехгранника V_N, V_E:
V_N= V_ξcosχ+V_ηsinχ; (29)
V_E= V_ξsinχ-V_ηcosχ-NUcosΦ_ИКВ. (30)
Вычисляются географическая широта Φ_ИКВ и географическая долгота λ_ИКВ :

После окончания полета вычислителем 5 погрешностей инерциальной навигационной системы производится комплексная обработка данных, собранных во время полета регистратором 4 навигационной информации. Комплексная обработка информации производится с использованием известных методов идентификации источников погрешностей инерциальных навигационных систем, например с использованием фильтра Калмана [2]. При этом расчеты производятся как в прямом, так и обратном времени. В процессе этой обработки производится расчет следующих погрешностей инерциальной навигационной системы 1:
ω_ξn, ω_ηn, ω_ζn - недокомпенсированные постоянные составляющие дрейфов гироскопов;
ω_ξξ, ω_ξη - дрейфы гироскопа с измерительной осью ξ , пропорциональные действующим ускорениям по осям ξ и η соответственно;
ω_ηξ, ω_ηη - дрейфы гироскопа с измерительной осью η , пропорциональные действующим ускорениям по осям ξ и η соответственно;
Δm_ξ, Δm_η - погрешности масштабных коэффициентов акселерометров;
ΔV_ξb, ΔV_ηb - скомпенсированные перед полетом постоянные дрейфы гироскопов.The radii of curvature of the meridian M, the first vertical N, and the parameter R ₀ are calculated:

The azimuth of the gyro platform χ and the true course φ of the _ICV are calculated:

φ _IKV = φ _g + χ. (28)
The velocities in the axes of the accompanying trihedron V _N , V _E are calculated:
V _N = V _ξ cosχ + V _η sinχ; (29)
V _E = V _ξ sinχ-V _η cosχ-NUcosΦ _IKV . (thirty)
The geographical latitude Φ _IKV and the geographic longitude λ _IKV are calculated:

After the flight, the calculator 5 errors of the inertial navigation system performs complex processing of data collected during the flight by the recorder 4 navigation information. Comprehensive information processing is carried out using well-known methods for identifying the sources of errors of inertial navigation systems, for example, using the Kalman filter [2]. In this case, the calculations are made both in direct and reverse time. In the process of this processing, the following errors of the inertial navigation system 1 are calculated:
ω _ξn , ω _ηn , ω _ζn are the _{uncompensated} constant components of the gyro drifts;
ω _ξξ , ω _ξη are the drifts of the gyroscope with the measuring axis ξ, proportional to the acting accelerations along the axes ξ and η, respectively;
ω _ηξ , ω _ηη are the drifts of the gyroscope with the measuring axis η, proportional to the effective accelerations along the axes ξ and η, respectively;
Δm _ξ , Δm _η - errors of scale factors of accelerometers;
ΔV _ξb , ΔV _ηb - compensated before the flight constant drifts of gyroscopes.

Additionally, using a fourth-order Kalman filter, the measurement vector of which uses the height and vertical speed obtained using information from the inertial navigation system 1 and satellite navigation system 2, estimates of the constant component of the error and the error of the scale factor of the vertical accelerometer (Δa _ζn , Δm _ζ )

The results of calculating the quantities ω _ξn , ω _ηn , ω _ζn , ω _ξξ , ω _ξη , ω _ηξ , ω _ηη , Δm _ξ , Δm _η , ΔV _ξb , ΔV _ηb , Δa _ζn , Δm _ζ from the calculator 5 errors of the inertial navigation system are non-volatile storage device 6.

Также на основании информации, записанной в энергонезависимом запоминающем устройстве 6, вычислитель 7 поправок к составляющим скорости и азимутальному углу гироплатформы инерциальной навигационной системы производит вычисление поправок ΔV_Nb, ΔV_Eb и Δχ_b исходя из следующих соотношений:

Результаты вычислений поправок ΔV_Nb, ΔV_Eb и Δχ_b из вычислителя 7 поправок к составляющим скорости и азимутальному углу гироплатформы инерциальной навигационной системы поступают на четвертый вход навигационного вычислителя 3. При этом также параметры ΔV_ξb и ΔV_ηb из энергонезависимого запоминающего устройства 6 на третий вход навигационного вычислителя 3. Навигационный вычислитель формирует скорости в осях сопровождающего трехгранника в виде:

При этом истинный курс формируется в виде:
φ_ИКВ= φ_ИКВ-Δχ_b. (36)
Затем вычисляются географические координаты летательного аппарата путем интегрирования угловых скоростей:

Вычисленные географические координаты, высота и составляющие скоростей летательного аппарата выдаются на выход 8 навигационной информации и передаются потребителям.Before the second and subsequent flights, the initial orientation angle χ _{0 of the} gyro platform of the inertial navigation system 1 in azimuth is formed in accordance with the ratio:

Also, based on the information recorded in the non-volatile memory 6, the calculator 7 corrections to the velocity components and the azimuthal angle of the gyro platform of the inertial navigation system calculates the corrections ΔV _Nb , ΔV _Eb and Δχ _b based on the following relationships:

The calculation results of the corrections ΔV _Nb , ΔV _Eb and Δχ _b from the calculator 7 corrections to the velocity components and the azimuthal angle of the gyro platform of the inertial navigation system are supplied to the fourth input of the navigation calculator 3. Moreover, the parameters ΔV _ξb and ΔV _ηb from the non-volatile memory 6 to the third input navigation computer 3. The navigation computer generates speeds in the axes of the accompanying trihedron in the form:

In this case, the true course is formed in the form:
φ _IKV = φ _IKV -Δχ _b . (36)
Then, the geographical coordinates of the aircraft are calculated by integrating the angular velocities:

The calculated geographical coordinates, altitude and speed components of the aircraft are issued to the output 8 of navigation information and transmitted to consumers.

 If the satellite navigation system 2 is inoperative (for example, due to satellite inoperability, the presence of radio interference, etc.), as evidenced by the absence of a service signal at the output of satellite navigation system 2, the calculation of geographical coordinates, altitude, speed components and aircraft heading based on data from an inertial navigation system only. But at the same time, they are adjusted based on the data received from the satellite navigation system 2 (until the satellite navigation system 2 is inoperative) and the corrections received from the calculator 7 corrections to the speed components and the azimuthal angle of the gyro platform of the inertial navigation system.

 A preliminary assessment of the results of the effectiveness of using the invention in the navigation system of a fighter aircraft showed that the implementation of the invention allows to reduce the errors of autonomous inertial reckoning of coordinates in flight by 3-4 times, the errors of gyrocompassing - by an order of magnitude.

 Thus, the present invention provides improved accuracy, fault tolerance and information content of the navigation system of the aircraft.

 The presented drawings and the description of the invention allow, using the existing element base, to manufacture it industrially and use it in the navigation systems of aircraft: multi-functional fighters, helicopters, etc., which characterizes the invention as industrially applicable.

Sources of information
1. Global satellite radio navigation system GLONASS. Ed. V.N. Kharisova, A.I. Perova, V.A. Boldin. M .: IPRZhR, 1998, p. 374, fig. 16.7.a.

 2. Flight tests of flight and navigation systems of aircraft and helicopters. E.G. Harin et al. M .: Mechanical Engineering, 1985, p. 46-49, 55-49.

Claims (1)

 An aircraft navigation system comprising an inertial navigation system, a satellite navigation system and a navigation computer, characterized in that it further comprises a navigation information recorder, an error calculator for an inertial navigation system, a non-volatile memory device, a calculator for corrections to the speed components and the azimuthal angle of the gyro platform inertial navigation system while the output of the inertial navigation system is connected to the first the input of the navigation computer and the first input of the navigation information recorder, the output of the satellite navigation system is connected to the second input of the navigation computer and the second input of the navigation information recorder, the output of the navigation information recorder is connected to the input of the inertial navigation system error calculator, the output of which is connected to the input of the non-volatile storage device, the output which is connected to the third input of the navigation computer and the input of the computer It is equal to the velocity components and the azimuthal angle of the gyro platform of the inertial navigation system, the output of which is connected to the fourth input of the computer, the output of which is the output of navigation information.