Satellite Orbit and Ephemeris Determination using Inter Satellite Links

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Orbit ComputationInter Satellite Linksare solved using a fourth order Runge-Kutta algorithm. Note that the integration is performedin the earth fixed frame, thus, it is not necessary for the user to compute earth rotationparameters.The desired position at time t is obtained by integrating from the position at time t 0 which isgiven in the navigation message. The GLONASS navigation message valid fort − t0≤ 15 minutesEq. 4.4-2which means, the time of the reference state t 0 lies in the middle of the 30 minutes period ofvalidity.4.4.1.1 Extended GLONASS FormatIt is easy to augment the GLONASS message to enhance accuracy or adapt the message formore perturbed orbits, simply by allowing the acceleration to vary over time. An extendedGLONASS message using 12 Parameters could look likeXBroadcast( x , y ,z , xD, yD,zD,a ,a ,a ,a ,a ,a ,a ,a , a )= Eq. 4.4-3000000x0y0z0x1with the reference position and velocity being the same as in the GLONASS message, and theconstant residual acceleration being replaced byDD yDD zRe sRe sRe s= a= ax0y0z0+ a+ ax1y1z1( t − t0)( t − t0)( t − t )Dx= a + aEq. 4.4-40If even more adaptability to perturbations, or simply a longer period of validity is required,the navigation message could also be extended to 15 Parameters,XBroadcast( x , y ,z , xD, yD,zD,a ,a ,a ,a ,a ,a ,a ,a , a )= Eq. 4.4-5000000with the residual acceleration being modelled as a quadratic term.Re sRe sy0z0y1z1x0( t − t0) + ax2( t − t0)( t − t0) + ay2( t − t0)( t − t ) + a ( t − t ) 20z202y0z02> xRe s= ax0+ aEq. 4.4-6x1>> y>> z= a= a+ a+ aIn both cases, the same propagator is used as in the standard GLONASS message.x1y1y1z1z1x 2x 2y2y2z2z24.4.2 GPS Broadcast EphemerisIn contrary to the integrating-a-force-model based GLONASS broadcast ephemeris, the GPSstate propagator consists of a Keplerian orbit propagator accounting for secular and periodicPage 64R. Wolf
Inter Satellite LinksOrbit Computationperturbations. The following 15 ephemeris related parameters are part of the GPS navigationmessage.M 0∆neΩ 0i 0ωAOMEGADOT(=dΩ/dt)IDOT (=di/dt)CucCusCrcCrsCicCist oeMean anomaly at reference timeMean motion difference from computed valueEccentricitySquare root of semi-major axisLongitude of ascending node of orbital plane at weekly epochInclination angle at reference timeArgument of perigeeRate of right ascensionrate of inclination angleAmplitude of the cosine harmonic correction term to the argumentof latitudeAmplitude of the sine harmonic correction term to the argument oflatitudeAmplitude of the cosine harmonic correction term to the orbit radiusAmplitude of the sine harmonic correction term to the orbit radiusAmplitude of the cosine harmonic correction term to the angle ofinclinationAmplitude of the sine harmonic correction term to the angle ofinclinationreference time of ephemerisThe following computations are necessary, to derive the satellites position in an earth centredearth fixed reference frame.A =( A ) 2Semi-major axis Eq. 4.4-7n =0GM3AComputed mean motion Eq. 4.4-8R. Wolf Page 65
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Satellite Orbit and EphemerisDeterm
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Inter Satellite LinksAbstractGlobal
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Inter Satellite LinksTable of Conte
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Inter Satellite Links6.2.2 IGSO WAL
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Inter Satellite LinksList of Figure
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Inter Satellite LinksFigure 6-38 Tr
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Inter Satellite LinksList of Tables
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Inter Satellite LinksJ2Earths Oblat
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Inter Satellite LinksIntroduction1
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Inter Satellite LinksIntroductionGE
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Inter Satellite LinksISL Observatio
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Inter Satellite LinksState Estimati
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State EstimationInter Satellite Lin
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State EstimationInter Satellite Lin
Page 30 and 31: State EstimationInter Satellite Lin
Page 36 and 37: Orbit ComputationInter Satellite Li
Page 86 and 87: Software DescriptionInter Satellite
Page 108 and 109: Simulations and ResultsInter Satell
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Simulations and ResultsInter Satell
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Autonomous Onboard ProcessingInter
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ReferencesInter Satellite Linksinte
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Satellite Orbit and Ephemeris Determination using Inter Satellite Links

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