413-418In satellite geodesy, the motion of artificial satellites has a crucial importance and the satellite’s position is the core point in the geodetic network. The simplicity of the broadcast ephemeris model directly determines the rate and efficiency of real-time navigation and positioning. GNSS broadcast message is transmitted in ECEF system either (GLONASS) or are Keplerian-like orbital elements (GPS, BeiDou, Galileo). The main aim of this paper is to describe the magnitude of Keplerian elements and correction coefficients of the navigational satellite changes based on a GPS satellite. The author examines six Keplerian elements and six correction coefficients for three different intervals: 2-hour, 1-day and 5-day intervals for periods of 9 days, a year and 10 years respectively. The analysis shows the varying behaviour of the time-series of the analysed components on the basis of the on-board GPS message, both random-like and repetitive (seasonal) trajectory time series

Maciuk, Kamil

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Indian Journal of Geo Marine Sciences Vol. 49 (03), March 2020, pp. 413-418       Long and short-term changes of Keplerian elements and correction coefficients of GPS satellites Kamil Maciuk AGH University of Science and Technology, al. Mickiewicza 30 30-059 Kraków, Poland [E-mail: maciuk@agh.edu.pl]  In satellite geodesy, the motion of artificial satellites has a crucial importance and the satellite’s position is the core point in the geodetic network. The simplicity of the broadcast ephemeris model directly determines the rate and efficiency of real-time navigation and positioning. GNSS broadcast message is transmitted in ECEF system either (GLONASS) or are Keplerian-like orbital elements (GPS, BeiDou, Galileo). The main aim of this paper is to describe the magnitude of Keplerian elements and correction coefficients of the navigational satellite changes based on a GPS satellite. The author examines six Keplerian elements and six correction coefficients for three different intervals: 2-hour, 1-day and 5-day intervals for periods of 9 days, a year and 10 years respectively. The analysis shows the varying behaviour of the time-series of the analysed components on the basis of the on-board GPS message, both random-like and repetitive (seasonal) trajectory time series. [Keywords: Ephemeris, GPS, Keplerian orbital elements, Orbit, Orbital elements] Introduction The GPS broadcast ephemeris is the basis of satellite navigation and positioning1,2. GPS ephemeris are forecast, predicted or extrapolated satellite orbit data, which are transmitted to the receiver from the satellite via a navigation message3. The GPS navigational message is available in real-time, therefore it constitutes the essence of standalone positioning as well of RTK measurements, which are today in use in various areas, e.g. geodesy4–6, crustal deformations analysis7, spatial analysis8,9 or thermography10,11. GPS broadcasts a message as Keplerian ephemeris parameters12. Keplerian motion is called simplified satellite orbiting and the problem is called the two-body problem. An artificial Earth satellite (AES) moves in a central force field and its mass is negligibly small compared to the mass of the Earth (𝑀). According to Newton’s second law of motion13, satellite motion is described as14:  𝑟ሷ ൌ െ ఓ௥మ௥௥ … (1)  where 𝑟ሷ  is the acceleration of the motion (second order differentiation of vector 𝒓 with respect to time) and 𝜇 (ൌ 𝐺𝑀, 𝐺 is the universal gravitational constant) is the Earth’s so-called gravitational potential. Equation (1) describes the motion of the satellite in the gravity field of the central body 𝑀, on the assumption of its symmetry (sphere). Acceleration magnitude depends on the quotient of the statement 1 𝑟ଶ⁄  — it is inversely proportional to the distance between the central body (Earth) and the satellite. The transformation of equation (1) to the general form of a two-body problem without taking into consideration the influence of the satellite on the central body 𝑀 (the so-called limited case of the two-body problem) and with assumption of the position consistency of the 𝑀 body against the origin of the coordinate system leads to the equation:  𝒓ሷ ൅ ఓ௥మ𝒓௥ ൌ 0 … (2)  Equation (2) is a second order differential equation. The geometric centre of the central body overlap with the origin of the coordinate system vector’s 𝑟 coordinates will be defined as 𝑥, 𝑦, 𝑧 and the vector of differential equation (2) may be substituted with three scalar second order differential equations:  𝑥ሷ ൅ 𝜇𝑟ଷ 𝑥 ൌ 0  𝑧ሷ ൅ 𝜇𝑟ଷ 𝑧 ൌ 0  𝑦ሷ ൅ ఓ௥య 𝑦 ൌ 0 … (3)  where 𝑟 ൌ ඥ𝑥ଶ ൅ 𝑦ଶ ൅ 𝑧ଶ. The analytical solution of equation (2) leads to the Keplerian motion defined by six parameters15 presented in Table 1. INDIAN J. GEO-MAR. SCI., VOL. 49, NO. 03, MARCH 2020   414The first two parameters (𝑎 and 𝑒) define the size and shape of elliptical orbits. The argument of perigee (𝜔) defines the location of the body against the equator plane (Fig. 1). The orientation of the circulating body against the equator plane defines the inclination angle (𝑖) and the right ascension of the ascending node (Ω). The above parameters define others, such as altitude or orbital period (half a sidereal day)17. Keplerian elements can be characterised by annual, semi-annual or draconitic courses18. Keplerian parameters can be divided into two groups20: dynamical constants (𝑎, 𝑒, 𝜏) and spatial orientation of the angles of eclipse (𝜔, 𝑖, Ω). Perturbation forces acting on the orbital motion of the AES cause a difference between the actual position of the satellite and the position calculated on the basis of Keplerian elements21, therefore this position cannot be described by equation (2). Therefore, equations of motion can be described as14,22:  𝑟ሷ ൌ െ𝜇 ௥௥య ൅ 𝑎ሺ𝑡; 𝑟, 𝑟ሶ ,𝑑ଵ,𝑑ଶ, … ,𝑑௠ሻ … (4)  where:  𝑟ሷ —  geocentric acceleration vector of the satellite 𝑎 —  perturbing acceleration 𝑟 —  geocentric position vector of the satellite 𝑟ሶ —  geocentric velocity vector of the satellite 𝑑௝ — dynamical parameters, 𝑗 ൌ 1 …𝑚 The first element in equation (4) is the force exerted by the central body, whereas vector  represents the sum of all perturbing accelerations affecting the AES’s motion. Satellite perturbing forces may be described by a series of instantaneous Keplerian orbits deviates from true trajectory. The set of orbital elements assigned to the satellite vector of state on each epoch 𝑡 may be described by the equation:  ሼ𝑟ሺ𝑡ሻ, 𝑟ሶሺ𝑡ሻሽ ↔ ሼ𝑎ሺ𝑡ሻ, 𝑒ሺ𝑡ሻ, 𝑡ሺ𝑖ሻ,Ωሺ𝑡ሻ,𝜔ሺ𝑡ሻ, 𝜏ሺ𝑡ሻሽ   … (5)  These elements are osculatory to the real orbit on a particular epoch 𝑡22. Accelerations caused by the gravity of other celestial bodies and forces not related to gravitational forces have an influence on the movement of the satellite. The role of these forces is negligibly small compared to the accuracy of the determined position. Table 2 presents the greatest perturbing forces acting on the AES: The second zonal harmonic arising from the Earth’s flatness is the biggest perturbing force. The influence of the Moon and the Sun is one order of magnitude smaller. The rest of the perturbing forces in the navigational satellite motion computation due to their insignificant size may be ignored.  Materials and methods Two different approaches are followed in computing earth-centred, earth-fixed (ECEF) coordinates from GNSS broadcast orbits. Based on Keplerian elements transmitted in the navigational message, ECEF coordinates are computed as an analytical function with a 2-hour update rate of broadcast data14. This approach is adopted by GPS, Galileo, BeiDou, IRNSS and QZSS24,25. The second approach is based on the numerical integration of nine elements of the vector of state (three positions,   Fig. 1 — The Keplerian elements of an elliptical orbit19.  Table 1 — Keplerian orbital parameters16. Parameter Notation a semi-major axis of orbital ellipse e numerical eccentricity of ellipse τ epoch of perigee passage ω argument of perigee i inclination of orbital plane Ω right ascension of ascending node  Table 2 — Summary of approximate perturbing forces for GPS satellite23. Source Maximum perturbing acceleration [m/s2] Maximum excursion growth in one hour [m/s2] Earth-mass attraction 5.65ꞏ10-1 — Second zonal harmonic 5.3ꞏ10-5 300 Lunar gravity 5.5ꞏ10-6 40 Solar gravity 3ꞏ10-6 20 Fourth zonal harmonic 10-7 0.6 Solar radiation pressure 10-7 0.6 Gravity anomalies 10-8 0.06 All other forces 10-8 0.06  MACIUK: KEPLERIAN ELEMENTS OF GPS SATELLITES   415velocities and accelerations). The vector of state is transmitted every 30 minutes and is in use in GLONASS, GAGAN and EGNOS25. In this work, the author analysed changes of data transmitted in GPS broadcast messages (Table 3). The parameters presented in Table 3 allow the coordinates of the satellite in the ECEF coordinate system to be computed27, and the details of the computation have already been widely described in the literature, e.g. Beutler et al.15, Ogaza28, Li29, and Hugentobler & Montenbruck30.  Results and discussion In this paper, six elements of Keplerian orbit and six correction element changes were analysed. These values were collected from GPS navigational messages between 1064 and 1616 GPS week (years 2000-2010, 552 weeks, almost 10.6 years) of GPS satellite PRN 10. An analysis was conducted for three different periods:   9 days with a 2-hour interval (1 - 9 Sep 2010)  1 year with a 1-day int. (Jan - Dec 2010)  10 years with a 5-day int. (May 2000 - Dec 2010)  Figure 2 presents the behaviour of Keplerian elements in 2-hour intervals with 9-day coverage. Short-periodic perturbations dominate the semi-major axis 𝑎, eccentricity 𝑒 and argument of perigee 𝜔 with a polynomial trend for the entire analysed period. In the case of the mean anomaly of reference time 𝑀଴, there is clearly seen a repeatable, sawtooth pattern with a close to 12-hour long period (half a sidereal day – GPS satellite evolution period). An analysis of the year-long time-series with a 1-day interval shows similar, periodical fluctuations for the same parameters Table 3 — Ephemeris data definitions26. Notation Data description M଴ Mean anomaly of reference time ∆n Mean motion difference from computed value e Eccentricity √a Square root of the semi-major axis Ω଴ Longitude of ascending node of orbit plane at weekly epoch i଴ Inclination angle at reference time ω Argument of perigee Ωሶ  Rate of right ascension idot Rate of inclination angle C୳ୡ Amplitude of the cosine harmonic correction term to the argument of latitude C୳ୱ Amplitude of the sine harmonic correction term to the argument of latitude C୰ୡ Amplitude of the cosine harmonic correction term to the orbit radius  C୰ୱ Amplitude of the sine harmonic correction term to the orbit radius C୧ୡ Amplitude of the cosine harmonic correction term to the angle of inclination C୧ୱ Amplitude of the sine harmonic correction term to the angle of inclination t୓୉ Reference time ephemeris IODE Issue of data (ephemeris)    Fig. 2 — Time-series of Keplerian elements in 2 hours intervals. INDIAN J. GEO-MAR. SCI., VOL. 49, NO. 03, MARCH 2020   416presented in Figure 3. In the case of the semi-major axis and longitude of the ascending node 𝛺଴, the patterns have a linear trend with single pitches. In the case of three parameters (𝑖, 𝜔, 𝑒), there is a long-term polynomial trend and the mean anomaly of reference time 𝑀଴ pattern has a random changes ranging from -3 to 3. Figure 4 presents the behaviour of Keplerian elements in 5-day intervals covering 10 years of observations. The time-series of three elements (𝑖, 𝜔, 𝑒) present polynomial behaviour with imposed sinusoidal fluctuations with annual (𝜔, 𝑒) and semi-annual (𝑖) period oscillations. In the case of the semi-major axis 𝑎 and longitude of the ascending node 𝛺଴, the time-series demonstrates a sawtooth graph plot with an annual period. In the case of the mean anomaly of reference time 𝑀଴, the graph oscillates within a range of -3 to 3 with a random-like, close to annual period. The behaviour of correction coefficients is presented in   Fig. 3 — Time-series of Keplerian elements in daily intervals    Fig. 4 — Time-series of Keplerian elements in 5 days intervals MACIUK: KEPLERIAN ELEMENTS OF GPS SATELLITES   417Figures 5-7. Each coefficient has periodical sinusoidal fluctuations with half a sidereal day. An analysis of annual graphs of four coefficient 𝐶௥௖, 𝐶௥௦, 𝐶௨௦, 𝐶௨௖ time series presents periodical fluctuations with a nearly 13.5-day period. The coefficients 𝐶௜௖, 𝐶௜௦ time series present a pseudorandom noise course both for the year period (1-day interval) and for 10 years (5-day interval). The coefficients 𝐶௥௖, 𝐶௥௦, 𝐶௨௦, 𝐶௨௖  time series in the 10-year period have a sinusoidal course with a 175-day period, which is very close to half of a draconitic year.  Conclusion The paper discusses the orbits of GPS satellites and the author looks at the Keplerian element changes based on the navigational message. There is an analysis of six Keplerian elements and six correction coefficient time series in three different intervals: 2 hours, 1 day and 5 days. The analysis shows the varying behaviour of the analysed time series both in short- and long-time periods, the majority of which is characterised by seasonal, repetitive changes. Short-term sinusoidal fluctuations are imposed on long-term close to annual, semi-annual and even draconitic periods, which can be extrapolated and whose values may be anticipated.  Acknowledgment Statutory research number 16.16.150.545.  References 1 Wang, A., Chen, J. & Wang, J., Precision Analysis of CNAV Broadcast Ephemeris and Its Impact on the User Positioning. In: China Satellite Navigation Conference (CSNC) 2017 Proceedings: Volume I edited by Sun, J., Liu, J., Yang, Y., Fan, S. & Yu, W., (Springer Singapore) 2017, 431–440.   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Long and short-term changes of Keplerian elements and correction coefficients of GPS satellites

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Long and short-term changes of Keplerian elements and correction coefficients of GPS satellites

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