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ROUTINE OPERATIONAL AND HIGH-PRECISION ORBIT DETERMINATION OF ENVISAT

R. Zandbergen’, M.Otten’ , P.L.Righetti2 , D. Kuijpeti, and J.M. Dow’,

‘ESAlEuropean Space Operations Centre, Darmstadt, Germany ‘GMV SA, based at ESOC

3Logica UK, based at ESOC

ABSTRACT

ESA’s Earth observation satellite Envisat was successfully launched on 1 March 2002 by an Ariane-5 launcher, and ESOC immediately took over the task of determining and predicting the orbit using S-band tracking data, and optimising the manoeuvre sequence to bring the spacecraft into an orbit accurately phased with ERS-2. On-board, Envisat carries, among others, a radar altimeter, a DORIS instrument and a laser retro-reflector array (SLR). Data from these instruments are being used at ESOC for high-precision orbit determination, for verification of the routine orbit determination and for cross-comparison with orbits computed on-board by the DORIS navigator and with those delivered with the Envisat products. This paper presents the first consolidated results obtained for Envisat routine and high-precision orbit determination. All orbit determination and control activities were performed with the software package NAPEOS, which was developed in-house. 0 2003 COSPAR. Published by Elsevier Science Ltd. All rights reserved.

INTRODUCTION The Navigation Package for Earth Orbiting Satellites (NAPEOS) was developed at ESOC during the last few

years, with the aim of supporting precise orbit determination primarily of Earth observations satellites. It can be applied both for orbit determination, prediction and control as used during satellite operations, and for high- precision orbit determination and geodetic parameter estimation, aiming at centimetre accuracy.

Designed for many different missions and tracking observation types, it was first used for ESA’s Envisat mission, through Launch and Early Orbit Phase (LEOP), routine operations and high-precision orbit determination (POD). For operations use, the software had been validated during the launch preparation campaign, while in POD the software could be validated by cross-comparison both with the existing ESOC software used for ERS-2 and GPS, and with results obtained by other groups.

ENVISAT LAUNCH AND EARLY ORBIT PHASE The Envisat satellite was launched at 1:07 UTC on the 1” of March, from the Kourou European spaceport by

an Ariane 5 launcher. In order to support the initial satellite acquisition and orbit determination activities, a joint network of 9 ESA, NASA and CNES stations provided S-band range, Doppler and angles measurements during the first 3 days. This setup was similar to the one already used for the launch of the two ERS satellites.

The early orbit determination of the satellite has been performed in three steps. After the very first pass, an along-track injection error, expressed as a time offset value (TOV), was estimated through a fast comparison of the tracking data against the nominal injection orbit. After one orbit, a rough orbit determination, relying mainly on the angular measurements and on the previously estimated TOV was carried out. After three revolutions a first reliable orbit determination and prediction was performed, based on range and Doppler measurements and on the previously determined rough orbit. Starting from this solution it was possible to start the definition of the manoeuvre strategy for bringing the satellite into its operational orbit.

Adv. Space Res. Vol. 31, No. 8, pp. 1953-1958.2003 Q 2003 COSPAR. Published by Elsevier Science Ltd. All rights reserved Printed in Great Britain 0273-l 177/03 $30.00 + 0.00

1954 R. Zandbergen er al.

For the Envisat mission an operational orbit very similar to the one used for ERS-2 has been selected, i.e. near- circular with frozen eccentricity and a repeating ground track - 501 orbits every 35 days - and sun- synchronous. Moreover, the Envisat satellite follows exactly the same ground-track as ERS-2, but leading it by half an hour, due to the different local times at descending nodes of the two satellite orbits (10:00 for Envisat and -lo:30 for ERS-2).

For this particular launch date a very long drift, of around 119” along the orbit, was required for achieving the desired separation in argument of latitude between the two satellites. The duration of the drift phase was set to 35 days, in order to limit the fuel consumption - and thus the impact on the satellite operational life.

The operational orbit was achieved on 5 April 2002 and since then the satellite has been accurately maintained within 1 km from its nominal ground-track. The maintenance strategy is based on the compensation of altitude decay and inclination drift, the former through in-plane manoeuvres - around one per month, assuring the ground-track maintenance at the equator - and the latter through out-of plane manoeuvres at the ascending node - around three per year, assuring the ground-track maintenance at the poles. A correct allocation of the ivailable dead-band at the equator and at the pole assures the maintenance of the ground track at all latitudes. Moreover, the frozen eccentricity conditions are maintained by splitting the altitude maintenance maneuver in two bums of different size separated by half orbit and executed at an adequate argument of latitude.

OPERATIONAL S-BAND ORBIT DETERMINATION Starting from the third day after launch, sufficient tracking data was available for performing the orbit

determination on a routine basis. Three days of range and, Doppler measurements from the Kiruna S-band antenna, located in North Sweden, are used. For the satellite dynamics, a 36/36 gravity field, the gravity of sun and moon and a constant area model for atmospheric drag and solar pressure are taken into account. Besides the Envisat state vector, one drag coefficient per day and a range and range-rate bias per station are estimated. For the on-board transponder a temperature-dependent delay is modeled.

With this set-up very satisfactory performances both in terms of execution time - around 10 minutes - and accuracy in determination - of the order of few meters - and propagation - well below 100 meters after one day - are achieved. Table 1 shows some very preliminary statistics, valid in the absence of orbit manoeuvres.

Table 1. Preliminary along-track orbit prediction error statistics (m).

Days 1 3 6

Error RMS (m) 25 150 700

Figure 1 shows some statistics of the difference between the ESOC high precision orbit (see next section) and the operational orbit. As the error in the high precision orbit is of the order of centimetres, the difference can be interpreted as the accuracy of the operational orbit.

Operational orbit accuracy (rms)

5 4.5

4 3.5

e 3 3 2.5 E 2

1.5 1

0.5 0 9-Apr 4-May 29-May 23-Jun 18-Jul 12-Aug 6-Sep I-Ott

date

Fig. 1. Envisat operational orbit accuracy

Routine Operational ENVISAT Orbit Determination 1955

ENVISAT HIGH-PRECISION ORBIT DETERMINATION ERS-2 and Envisat POD is now performed by the recently established Navigation Support Office of ESOC.

The purpose of this work is two-fold, First, it is part of the research in the fields related to high-precision modeling of orbit perturbation forces and tracking observations of all types. This research ultimately aims at maintenance and improvement of the operational systems, to support future missions with new or more stringent requirements. Secondly, it allows an immediate verification of the accuracy of the orbits computed for spacecraft operations, and of the orbits used for the geodetic application of the mission data. Both aspects have received attention in previous publications of similar work done for ERS-2 (e.g. Zandbergen et al., 199?). The benefit from performing POD has heen repeatedly demonstrated, for example in the improved accuracy obtained in routine operational orbit determination and prediction since the launch of ERS-1 in 199 1.

High precision models and method The Envisat high precision orbits are determined using five-day data arcs with a one-day overlap at the

beginning and end of the determination arc. The resulting middle three days are written to the final high precision orbit file. The models and estimated parameters used in this process are summarised in Table 2. The total number of estimated satellite parameters for each five-day arc is 47.

Table 2. Models and estimated parameters used in ESOC POD of Envisat.

Type Model Geopotential GRIM5Cl degree and order 70 Third bodies Sun, Moon (52) and all planets (JPL DE200) Aerodynamic ANGARA (MSIS 90, HWM-93) Solar Radiation ANGARA Albedo/IR. Radiation ANGARA Station positions (SLR, DORIS) ITRF 2000 position and velocity SLR and DORIS corrections IERS-96 and for DORIS also from data Estimated Parameters Initial position and velocity Drag Four parameters per day (6-hours) Solar Radiation Once per arc l-CPR acceleration Once per day, along and cross track DORIS range rate bias Per pass per station

An internal analysis using NAPEOS has shown that the GRIM%Cl gravity model (Gruber et al., 2000) results in the lowest SLR and DORIS residuals for Envisat at the time of writing. GRIM%Cl is a combined solution using both satellite and ground-based measurements, and claiming a surface accuracy of 5.5 mgal. Although GRIMS-Cl is complete to degree and order 120, no improvement in the orbit and the residuals is visible using terms above degree 70. Currently available gravity models based on CHAMP data have not yet shown an improvement in the SLR and DORIS residuals for Envisat.

An important new feature within the Envisat POD at ESOC is the modeling of the non-gravitational accelerations. For this a new software system ANGARA (Analysis of Non-Gravitational Accelerations due to Radiation and Aerodynamics) (Fritsche, et al, 1998) has been applied to Envisat (Doombos, et al., 2002). ANGARA uses a 2-phase approach, where in the first phase a detailed geometric model (see Figure 1) is analysed using either an integral method or Monte Carlo analysis. This analysis results in a file containing the normalised aerodynamic and radiation coefficients as a function of attitude. This file is then used in the orbit determination together with information about the attitude and position of the satellite relative to the Earth and the Sun.

ITRF 2000 station position and velocity are used for the SLR and DORIS stations. Several stations have undergone changes in their position since the release of ITRF 2000 and station positions are therefore regularly updated when required.

All corrections applied to the SLR and DORIS measurements e.g. troposphere, centre of mass and station displacement in NAPEOS follow the IERS-96 convection (McCarthy, 1996). For the DORIS measurements the correction for centre of mass, troposphere and ionosphere are already included in the data and are used instead.

1956 R. Zandbergen et al.

Fig. 2. Envisat ANGARA model used for non-gravitational forces.

For each five-day arc the following parameters are estimated: initial state vector, a drag scale factor for each six hours (linear), solar radiation scale factor (per arc), along- and cross-track l-cpr accelerations (per day) and for each DORIS station a range-rate bias per pass.

ORBIT DETERMINATION ACCURACY Availability of a high-accuracy independent measurement type can be used to validate the accuracy of the

determined satellite orbit. For Envisat, the altimeter data would be the best choice for this, but at the time of writing this data is not yet available in a suitable format and quantity. Instead, other quality indicators can be used. The consistency of the orbits is tested by comparing overlapping arcs, and by looking at the SLR and DORIS residuals for each arc. Also, different solutions are compared with each other, for example using only SLR data or using both SLR and DORIS data. Furthermore, the SLR/DORIS orbits are compared against high precision orbits of other institutes.

Orbit comparisons Two different solutions were used: from CNES and from the Delft Institute for Earth-Oriented Space

Research (DEOS). Both centres are involved, together with ESOC, in Envisat calibration and validation. CNES is responsible for the DORIS instrument and generates high precision orbits using DORIS and SLR measurements. The DEOS high precision orbits are also based on ‘DORIS and SLR data. The period selected for the orbit comparison is from 25 July 2002 to 24 August 2002, almost one Envisat repeat cycle. This is the first cycle for Envisat in which continuous DORIS data is available. Table 3 lists the radial RMS difference statistics between the three high precision solutions.

Table 3. Radial RMS difference (cm) between ESOC, Fig. 3. Difference between ESOC and CNES orbit (m). CNES and DEOS orbits. The asterisk indicates an +Along, *Cross, o Radial. Envisat manoeuvre. CNES orbit is not available for 9 hours around the manoeuvre.

Period ESOC vs ESOC vs CNES vs CNES DEOS DEOS 0.1

25.7-28.7 1.98 1.86 2.14 28.7-31.7 2.13 2.32 2.13 31.7 -03.8 2.64 2.66 3.94 0 03.8 -06.8 3.05 3.39 2.03 06.8-09.8 1.87* 7.24* 1.93* 09.8 - 12.8 2.05 2.47 2.16 -0.1 12.8 - 15.8 2.23 2.27 2.13 15.8 - 18.8 3.43 2.46 2.96 18.8-21.8 2.27 2.32 2.80 -01 21.8 -24.8 2.54 2.68 2.54

140' I time from 2002~8/09 , W:CZ

Routine Operational ENVISAT Orbit Determination I957

Clearly visible from the table is the very good radial agreement between all three solutions. For almost all the comparisons the radial difference is of the order of 2 - 3 cm. Interesting to see is the radial difference of 7.24 cm between ESOC and DEOS during the 06.8 - 09.8 arc which contains an Envisat manoeuvre. This large RMS difference is only due to the period directly around the manoeuvre, about 2 orbits before and after. After this period the RMS difference is again of the order of 2 cm. This radial difference is not visible in the comparison with CNES because the CNES high precision orbit is not available for 9 hours around the manoeuvre.

This good agreement also exists for the transverse and cross track differences between the three solutions. Figure 3 shows for a three-day arc the radial, transverse and cross-track difference between the ESOC and CNES high precision solution. For the complete period (from 25.7 - 24.8) typical RMS values for transverse are 7 - 12 cm and for cross track 5 - 8 cm between all three solutions.

Internal comparison of the SLR/DORIS solution with SLR only solution for this period shows radial RMS differences of the order of 3 cm. This very good agreement is partly due to the very high number of SLR passes available for this period combined with relative calm atmospheric conditions. For other periods where DORIS data has been made available, usually slightly higher RMS differences (4 cm) are found between the two solutions.

Tracking data residual analysis Figure 4 shows some characteristics of the SLR data between 25 July 2002 and 24 August 2002. The total

RMS for all stations for this period is 3.41 cm with individual stations values between 2.26 - 6.19 cm. The total percentage of used SLR measurements for this period is 93.6 % with individual station contributions ranging from 17% (1873) to 100% (7810).

Individual Laser Station Contribution

Fig. 4. Top: the individual SLR station contribution and the used measurements for the comparison arc (in percentages). Bottom: SLR station RMS residuals for the comparison arc.

The DORIS residuals for this period are shown in Table 4. The average RMS value for all DORIS stations for this period is 0.52 mm/s with individual station values between 0.40 - 0.65 mm/s. Clearly visible from Table 4 is the very small variation over time in the DORIS residuals. Furthermore, there does not seem to be any geographical correlation between the DORIS residuals.

1958 R. Zandbergen et al.

Table 4. DORIS residual RMS (mm/s) and radial overlap (cm) for ESOC POD orbits. The asterisk indicates an Envisat manoeuvre.

Period: 25.7 - 28.7 - 31.7 - 03.8 - 06.8 - 09.8 - 12.8 - 15.8 - 18.8 - 21.8 28.7 31.7 03.8 06.8 09.8 12.8 15.8 18.8 21.8 24.8

DORIS: 0.508 0.518 0.529 0.522 0.514* 0.514 0.514 0.524 0.525 0.509 Overlap: ----- 2.08 2.13 2.74 3.17* 0.94 0.83 1.32 1.08 1.05

Radial overlap Table 4 also lists the radial overlap for the five-day arcs. The overlap period is 2 days: the last two days of the

previous arc against the first two days of the next arc. Some care should be used when analysing these radial overlap values as the least square process used for the orbit determination tends to result in larger errors at the beginning and end of each arc. Still, because the dynamic modelling and the used measurements for the overlap period are the same, this value gives an indication of the consistency of the orbit determination and thus an indication of the minimal radial error in the final solution.

CONCLUSIONS AND OUTLOOK ESOC has the capability to provide operational flight dynamics support and compute precise orbit solutions

with a common software package: NAPEOS. The successful support provided by NAPEOS both during the critical LEOP and the ensuing months of routine operations highlight the operational maturity of NAPEOS. Moreover, the comparison of the operational orbit with the precise orbit solution shows an accuracy much better than required.

Starting a few months after launch, precise orbit solutions have been provided on a routine basis using NAPEOS with high-precision SLR and DORIS tracking data. These solutions are characterised by very small residuals and consistent, small arc overlap differences, which are good indications of the high quality of the solutions. Comparisons of ESOC’s precise orbit solutions with solutions computed by CNES and DEOS have shown radial differences of the order of 2-3 centimeters, but of course these solutions share a number of common models. The radial orbit determination accuracy for each of these solutions may therefore be considered to be below 5 cm.

ln the next years the NAPEOS software will be used for the LEOP and routine operations support of all upcoming earth observation mission that will be supported by ESOC, starting with CryoSat (2004), MetOp (2005, LEOP only) and GOCE (2006). POD activities will also be carried out for both CryoSat and GOCE.

NAPEOS will also be used for high-precision orbit determination of GNSS constellations using geodetic- quality ground receiver data, as well as of LEO satellites carrying on-board GPS receivers (e.g. GOCE). The capability of processing all relevant types of GNSS data is currently being implemented and tested.

REFERENCES

Doombos E., R. Scharroo, H. Klinkrad, R. Zandbergen and B. Fritsche, Improved modelling of surface forces in the orbit determination of ERS and Envisat, Canadian Journal of Remote Sensing, Vol. 28, No.4, pp. 535- 543,2002.

Fritsche, B., M. Ivanov, A. Kashkovsky, G. Koppenwallner, A. Kudryavtsev, U. Voskoboinikov, and G. Zhukova, Radiation pressure forces on complex spacecraft, HTG, Germany and ITAM Russia, ESOC contract 11908/96/D/IM, 1998.

Gruber Th., A. Bode, Ch. Reigber, P. Schwintzer, R. Biancale, R. Lemoine., GRIMS-Cl: Combination solution of the global gravity field to degree and order 120, Geophys. Res. Lett. Vol. 27, pp. 4005- 4008 ,200O.

McCarthy D.D. (ed.), lERS Conventions (1996), ZERS Technical No@ 21, U.S. Naval Observatory, July 1996. Zandbergen R., J.M. Dow, M. Romay Merino, R. Piriz and F. Martinez Fadrique, ERS-1 and ERS-2 tandem

mission: Orbit Determination, Prediction and Maintenance, Adv. Space Res. Vol. 19, No. 11, pp. 1649-1653, 1997.

E-mail address of R. Zandbergen Rene.Zandberaen@esa.int Manuscript received 21 November 2002; revised 30 January 2003, accepted 14 March 2003