HYPER Industrial Feasibility Study Final Presentation Orbit Selection

Transcription

1 Industrial Feasibility Study Final Presentation Orbit Selection Steve Kemble Astrium Ltd. 6 March 2003

2 Mission Analysis Lense Thiring effect and orbit requirements Orbital environment Gravity Atmospheric effects Others Operational orbit design Options Launch Orbit selection Trade-offs Baseline orbit features 2 6 March 2003, Final Presentation, ESTEC

3 Lense Thiring effect and orbit requirements The latitude variation in the effect is extensive, with smaller amplitude near the Earth s equator and larger amplitude near the geographic poles deg latitude Relative amplitude w.r.t. 0 latitude in 1000km orbit Vertical axis corresponds to a location above the Earth s geographic North pole (or spin axis). Orbital radius does play a part in the amplitude, but is a much weaker term than the latitude effect deg latitude h=500 h=700 h=1000 h=1200 Therefore orbital design is weakly influenced by altitude requirements, but implies LEO 3 6 March 2003, Final Presentation, ESTEC

4 Lense Thiring effect and orbit requirements Z, polar A polar orbit measures the latitude dependent signature Yorbit X, equatorial Near polar orbits (ie sunsynchronous) offer an acceptable compromise Inclination Zorbit LT effect in orbit plane system Y, equatorial 2E-14 Orbit plane 1.5E-14 1E-14 LT amplitude 5E E-15-1E E-14-2E X Yorbit Zorbit -2.5E-14 True anomaly (deg) 4 6 March 2003, Final Presentation, ESTEC

5 Distrubance requirements Residual acceleration experienced by ASU to be less than 1.7e-8 m/s/s This is relative to the local free fall condition, ie from the resultant gravity field including all Earth harmonics and third body fields External, non-gravitational perturbations therefore contribute to this. Drag free control will compensate, but external perturbations should be minimised. Perturbations within the frequency range 0.03 to 0.3 Hz are of particular importance. 5 6 March 2003, Final Presentation, ESTEC

6 The Orbital environment A number of terms must be considered: Gravity (due to gradient effects within the payload and knowledge requirements) Earth harmonics Lunar/Solar Time dependent terms Atmospheric density Nominal density plus local free path effects Atmospheric winds Solar and Earth radiation pressures Electromagnetic effects 6 6 March 2003, Final Presentation, ESTEC

7 Gravity Earth s gravity can be represented as a sum of harmonic terms in latitude and longitude. Generally decreasing amplitude with order. Dominant term is J2 Figures show incremental effects between order 2, 4 and Longitude 324 (deg) Lat Longitude 324 (deg) Lat Longitude 324 (deg) Lat March 2003, Final Presentation, ESTEC

8 Gravity gradient Gravity gradient results in differential accelerations across a finite payload. Gradient varies with altitude and geocentric position. Central spherical term dominates Field dependence has similar charactersistics to LT effect Harmonic effects diminish in a similar manner to the gravity field effect 2.5 0deg offset from radial 2.60E E Gravity gradient (/s/s) 2.40E E E E E E deg offset from radial h=500 h=700 h=1000 H= Orbital altitude (km) 8 6 March 2003, Final Presentation, ESTEC

9 Knowledge of the gravity field Knowledge of the gravity gradient is required to better than 2.1e-11 /s/s at an altitude of 1000km. Therefore: Determine what order of gravity models are required to model to required accuracy Assess the uncertainties in the gravity models Maximum Fractional uncertainity Order 9 6 March 2003, Final Presentation, ESTEC

10 Knowledge of the gravity field (2) Field model order to meet knowledge requirement: Altitude (km) Order required Uncertainties at an altitude of 1000km: o Central field: 1e-14 o Order 2 components: 1e-13 o Remaining harmonic terms up to order 20: 1e March 2003, Final Presentation, ESTEC

11 Additional gravitational effects Sun and Moon Gradients are 8e-14 and 1.8e-13 respectively Ocean Tidal effects Gravity gradient effects at 1e-15 Earth s atmosphere This is included in the nominal GM term, but is in reality a dynamic term However mass is sufficiently small (less than one part per 1e6) that dynamic effects are not significant to gradient 11 6 March 2003, Final Presentation, ESTEC

12 Atmospheric effects Density varies significantly with altitude Also dependent on solar and magnetic activity, latitude, local solar time Eleven year solar cycle give order of magnitude variations Deceleration dependent of area/mass: following asumes E E Deceleration (m/s/s) 1.00E E E E E E-09 MSIS Max MSIS Mean 1.00E-10 Orbit altitude (km) 12 6 March 2003, Final Presentation, ESTEC

13 Atmospheric effects(2) Atmospheric density also shows higher frequency variations caused by motion of the satellite and the molecular mean free path Experimantal measurements suggest typically 10% variability in local dc value Frequency upper limit given by mean free path E E Force PSD (N/Hz) 1.00E E E E E-12 Maximum Mean Frequency (Hz) 1.00E E-16 Altitude (km) 13 6 March 2003, Final Presentation, ESTEC

14 Radiation pressure Nominal values of solar radiation in range 1316 to 1428 w/m^2 at 1AU Yields pressure of 4.5 micron/m^2 Earth radiation and albedo pressure also perturb satellite motion Infra-red radiation in range W/m^2 Albedo is dependent on local conditions and the satellite orbit. Dawn-dusk, Sun synchronous orbit effective maximum albedo is approximatly 0.17 Solar radiation pressure significantly greater than Earth effects 14 6 March 2003, Final Presentation, ESTEC

15 Radiation environment Result of magnetically trapped ions and electrons. Near polar, low Earh orbits encounter inner proton and electron belts Effect increases with orbit altitude (insignificant below 1000km) 4PI Total Dose in Si at centre of Al Spheres for various Sun- Synchronous altitudes for a 06:00 Ascending Node 1.00E E+07 Dose (Rads) 1.00E E E km 900km 1100km 1300km 1500km 1.00E E Z (mm) 15 6 March 2003, Final Presentation, ESTEC

16 Summary of orbital perturbations 1 = J2 2 = Higher harmonics 3 = Lunar gravity Acceleration nano m/s/s 4 = Solar gravity 5 = SRP 6 = ERP 7 = Albedo 8 = Max Drag 9 = Mean Drag k m km km 1000km 16 6 March 2003, Final Presentation, ESTEC

17 Orbit design: options Orbit Type Circular Elliptical Sun-synchronous Polar Observation implications LT effect periodic in magnitude and direction LT effect periodic but magnitude modulated by altitude dependent effect LT effect rotation axes have a component out of the orbit plane LT effect rotation axis contained in orbit plane Disturbances For sun-synchronous orbit, major disturbance (drag) has small variation over orbit. Significant drag variations over orbit period. Spacecraft implications Relatively constant environment over orbit period (true for sun-synchronous) Good ground station links for Molnya orbits. SSO inclination will experience apse rotation. Possibilities for near constant solar illumination geometries out of eclipse season. (select dawn-dusk orbit) Variable solar illumination geometry due to zero nodal regression March 2003, Final Presentation, ESTEC

18 Orbit design: options(2) Orbit Type Low altitude High altitude Observation implications LT effect amplitude high LT effect amplitude reduces slowly with altitude. Disturbances High atmospheric drag disturbances Low atmospheric drag disturbances Spacecraft implications Generally higher available spacecraft mass at orbit injection. Limited or no de-orbit effort required. Higher gravity gradients. Spacecraft mass at orbit injection generally reducing with orbit altitude. Improved ground station links. Longer de-orbit times. Reduced gravity gradient. Increased radiation dosage March 2003, Final Presentation, ESTEC

19 Payload mass, kg Launch Possible launchers are: Rockot Dnepr Cosmos i=63 i=73 i=82 i=86 Rockot has best mass performance Injection dispersions require approx 14 m/s correction i=97 i=94 i=100 i=98.1 Uncorrected orbits are feasible in terms of drift from nominal sun-synchronous (approx 10 deg node drift over 2 years) SSO Launch from Plesetsk allows transfer via 94 or 100 deg inclination orbits Circular orbit altitude, km 19 6 March 2003, Final Presentation, ESTEC

20 Baseline orbit selection issues Pre-selected: dawn-dusk, near circular, Sun-synchronous, LEO Orbit altitude is a critical parameter. Issues are: Injection mass from the launcher Atmospheric drag Gravity gradient knowledge Scientific observability Radiation End of mission de-orbit 20 6 March 2003, Final Presentation, ESTEC

21 Baseline orbit selection trade-offs Parameter 700km 1000km 1200km Comment Injection mass from the launcher 1000 kg 1000kg Estimated 800 to 850 kg No distinction between 700 and 1000km. Reduction at altitudes Atmospheric drag Max at 6.2e- 7 m/s/s Max at 4.5e- 8 m/s/s Max at 7.8e- 9 m/s/s greater than 1000km Order of magnitude reduction at higher altitude.(700 to 1000km) Drag at solar maximum is the dominant nongravitational perturbation at 700 and 1000km Gravity gradient knowledge Implies model at order 30 Implies model at order 20 Implies model at order 13 At 1200km Solar Radiation Pressure (SRP) at 2e-8 m/s/s exceeds atmospheric drag Higher altitude allows simplification of the modelling process 21 6 March 2003, Final Presentation, ESTEC

22 Baseline orbit selection trade-offs (2) Parameter 700km 1000km 1200km Comment Scientific observability LT effect magnitude reduced by 12%.c.f. 700km LT effect magnitude reduced by 19%.c.f. 700km Increasing altitude increases number of observations but reduces magnitude of the measured effect. Number of observations due to decrease in eclipse time increased by 9.9% Number of observations due to decrease in eclipse time increased by 17% Radiation De-orbit Dose with 5mm Al shield over 2 years: 8500 Rads Controlled De-orbit periods with 500 micron thrust typically 3 years Dose with 5mm Al shield over 2 years: Rads Controlled De-orbit periods with 500 micron thrust typically 10 years Dose with 5mm Al shield over 2 years: Rads Controlled De-orbit periods with 500 micron thrust typically 15 years Steeply increasing radiation dose at higher altitude De-orbit from higher altitude orbits required extended low thrust periods March 2003, Final Presentation, ESTEC

23 Baseline The reduced perturbation environment, acceptable launch mass and acceptable radiation environment give a preference for a 1009km altitude orbit over the original 700km. The magnitude of the LT effect is still acceptable at this higher altitude. Although a 1200km orbit allows even further reduction of perturbation effects, the reduction in available injection mass and also the more severe radiation environment count against this orbit. Therefore 1009km altitude is recommended as the best compromise. (1009km is used to obtain a 3 day repeat period) 23 6 March 2003, Final Presentation, ESTEC

24 Baseline features Orbit mean altitude (wrt spheroid radius 6378km): 1009 km Orbit mean radius 7387km Orbit period (node crossing): secs = mins Orbit eccentricity and perigee: Frozen orbit, perigee at antinode Orbit inclination: 99.5 deg Orbit ascending node: Dawn-dusk configuration Longitude motion per rev: deg Max eclipse period: 822 secs Eclipse season: 57 days 24 6 March 2003, Final Presentation, ESTEC

25 Baseline features Exact repeat orbit selected from opeartional considerations, (but not a strong driver) 25 6 March 2003, Final Presentation, ESTEC

26 Baseline modification implications The spacecraft experiences a more benign, non-gravitational perturbation environment Reduction in magnitude of the Lense-Thirring effect of approximately 12%. The number of days spent in eclipse reduces. Number of eclipse free days rises by 10% The radiation dose almost doubles, but over the two year period, the total dose is still low (15 krads) Launcher injection mass option with Rockot is almost unchanged The reduction in gravity gradient with altitude means that the problem of obtaining high accuracy knowledge of the field is alleviated March 2003, Final Presentation, ESTEC