Mars Odyssey Navigation

Transcription

1 Mars Odyssey Navigation Moriba Jah Jet Propulsion Laboratory California Institute of Technology April 2, 2002 Page - 1

2 Spacecraft Mission Investigate the Martian environment on a global scale, over a period of 917 Earth days. Serve as a relay for information to Earth, following the science phase. April 2, 2002 Page - 2

3 Spacecraft Mission constraints To achieve the mission, the spacecraft must: Be injected into an orbit with a period of less than 22 hours, while having a 300 km periapse altitude (+/- 25 km) and an inclination of 93.5º (+/- 0.2º), including MOI burn execution errors. This is equivalent to hitting a golf ball from NY to Paris and making it in the hole in only 4 swings. (achieved: 18:36 period km and 93.51º) Employ aerobraking over a 3-month period (walk-in, main phase, end-game/walk-out) in order to maximize payload mass and minimize propellant expense. By the end of aerobraking, stabilize in a 400 km circular, frozen, sun-synchronous orbit with a 2PM LMST AEQUAX. April 2, 2002 Page - 3

4 Trajectory Selection: Pork-Chop plot Courtesy of Rodney Anderson April 2, 2002 Page - 4

5 Interplanetary Trajectory Earth at Arrival 1.02 AU Vernal Equinox Launch: 07-APR AU 45.9 TCM-5 E-7 h Mars Arrival 24-Oct-2001 TCM-4 E-12 d 10 day time ticks Mars at Launch TCM-1 L + 46 d TCM-2 L + 86 d April 2, 2002 Page - 5 TCM-3 E - 37 d

6 Collecting Navigation Data April 2, 2002 Page - 6

7 Collecting Navigation Data Radiometric Data Types Doppler Measurements are comparisons of transmitted frequency (from ground station or spacecraft) with received frequency on ground; typical frequencies are at S-band (2 GHz) and X-band (7-8 GHz) Highly reliable; used in all interplanetary missions to date Range Measurements are typically two-way light time for radio signal to propagate between ground stations and spacecraft with a turnaround time; typical frequencies are also at S- and X-band Used in nearly all interplanetary missions since late 1960s April 2, 2002 Page - 7

8 Range and Doppler Tracking April 2, 2002 Page - 8

9 Radial vs Angular Measurements For most interplanetary missions, S/C position uncertainty is much smaller in Earthspacecraft ( radial ) direction than in any angular ( plane-of-sky ) direction Radial components of position and velocity are directly measured by range and Doppler observations In absence of other data, angular components are much more difficult to determine -- they require either changes in geometry between observer and spacecraft or additional simultaneous observer, neither of which is logistically simple to accomplish Angular errors are more than 1000 x radial errors even under the most favorable conditions (see below) when depending on range and Doppler measurements Spacecraft Position Uncertainty Ellipsoid z σ Declination σ Right Ascension Range r Declination y σ Radial 1999 Capability Position Velocity Radial Error 2 m 0.1 mm/s Angular Error (at 1 AU) 3 km* 0.1 m/s *Equivalent to angle subtended by quarter atop Washington Monument as viewed from Chicago Right Ascension x April 2, 2002 Page - 9

10 Navigation Data Types Delta Differential One-Way Range (ΔDOR) ΔDOR is a measurement technique that utilizes two ground stations to simultaneously view the spacecraft and then a known radio source (quasar or another S/C) to provide an angular position determination Two stations viewing the same signal allows for geometric plane-of-sky angular position measurement (Differential) By viewing two sources, common errors cancel and the angular separation can be calculated (Delta) April 2, 2002 Page - 10

11 Very Long Baseline Interferometry - ΔDOR April 2, 2002 Page - 11

12 ΔDOR Campaign Project requirement to use ΔDOR as independent data type VLBI implementation effort led by Jean Patterson and Jim Border 9 successful MGS demonstrations (Jan 2001) 5 more scheduled on MGS (Aug-Sep 2001) North-South baseline only geometric opportunity for majority of cruise Provides critical plane-of-sky information East-West measurements possible beginning in October Campaign began as soon as geometrically possible Two measurements per week started 04-June-01 All opportunities successful (except for E-W baseline low elevation ) Total of 45 measurements scheduled (40 N-S, 5 E-W) Traditional S/C-Quasar-S/C Measurements Measurement Accuracy 0.12 nsec (0.27 km - 1σ) April 2, 2002 Page - 12

13 DSN Viewperiods DSN Rise/Set Time of Day - Day 01, 52 Park Orbit Inc 00:00 Goldstone Rise Canberra Rise Madrid Rise Goldstone Set Canberra Set Madrid Set 22:00 20:00 Madrid Viewperiod 18:00 16:00 14:00 Canberra Viewperiod 12:00 10:00 08:00 06:00 Goldstone - Canberra Overlap 04:00 MOI No Coverage 02:00 Goldstone Viewperiod 00:00 07-Apr May Jun Jun Jul Aug Sep Oct-01 April 2, 2002 Page - 13 Date

14 Navigation Processes Trajectory/Mission Design Orbit Determination Maneuver Design & Analysis April 2, 2002 Page - 14

15 Trajectory Targeting Process Targets are designed pre-launch, updated as necessary Cruise Targets (encounter at Mars) usually defined with Closest Approach metrics Orbiter: Radius (Altitude) of Periapse, Inclination, Time Lander: Entry Radius, Entry Latitude, Entry Flight Path Angle, Time Can be expressed in other coordinates (B-plane) Aerobraking trajectory defined by a corridor Corridor defined by spacecraft and trajectory constraints Dynamic Pressure (structural), Heat Rate (Thermal), Density (Trajectory) Target Altitude and Time at Periapsis Mapping Orbit Targets are usually orbital elements Semi-Major Axis, Eccentricity, Longitude on Asc/Desc Node Node often described via True or Local Mean Solar Time Orbit can be described via orbit Beta angle April 2, 2002 Page - 15

16 Our Targeting Plane: B-plane April 2, 2002 Page - 16

17 Orbit Determination Orbit Determination is the process of adjusting trajectory models/apriori information to best match the observed tracking data, and quantify the error associated with the trajectory estimate The collected tracking data are the actual or Observed measurements Trajectory models produce predicted or Computed measurements Data Residuals = Observed Computed OD method is to minimize residuals by adjusting the trajectory models Minimized in a weighted least-squares sense (square-root information filter) OD filter accounts for measurement and apriori state parameter accuracies OD products: OPTG & SPK P-file April 2, 2002 Page - 17

18 What will our spacecraft experience? Satellite motion is determined by a number of forces that act on the spacecraft: Gravitational Forces Central body force Third-body force (other planets, moons) Central body gravity field asymmetries General relativistic effects Non-gravitational Forces Thruster Firings Trajectory correction maneuvers (TCMs) Attitude control thrusting Angular Momentum Desaturations (AMDs) Solar Radiation Pressure Aerodynamic Drag Gas Leaks April 2, 2002 Page - 18

19 Spacecraft Configuration (cruise) +Z Earth +X Sun Solar Array Normal April 2, 2002 Page - 19

20 Thruster Configuration RCS-3 TCM-3 TCM-2 RCS-2 AACS Cruise Coordinate Frame (Same as Mechanical Frame) X c RCS-4 TCM-4 Y c TCM-1 RCS-1 April 2, 2002 Page - 20

21 Models That May Be Estimated Trajectory Force Models Initial S/C position and velocity (State at Epoch) 6 components of cartesian state Any S/C thrusting events 3 components (ΔVx,ΔVy,ΔVz or ΔV, RA, DEC) for each discrete event Many events over course of cruise trajectory: TCMs, AMDs Solar Radiation Pressure Dependent on attitude profile and component orientation (solar panel) Specular Diffuse Planet and Satellite Ephemerides and Gravity Fields Gravity Field of Mars: MGS75C Atmospheric Density Due to drag pass during aerobraking April 2, 2002 Page - 21

22 Models That May Be Estimated Measurement or Signal Path Models Earth Platform Parameters Tracking Station Locations Earth Rotation and Pole Nutation (Timing and Polar Motion) Tracking Data Calibration Parameters Signal delays induced by Ionosphere and Troposphere Measurement Biases Range Biases due to hardware delays One-way doppler bias due to oscillator frequency drift April 2, 2002 Page - 22

23 Trajectory Prediction All planning is based on predictive capabilities, not real-time spacecraft location Trajectory Prediction involves accurately modeling and estimating all past events, as well as predicting all future events During the cruise to Mars, Nav must model all future events such as: Solar Pressure - Attitude profile and component orientations Thrusting - Angular Momentum Desaturations, or thruster slews Unmodeled forces must eventually be compensated with maneuvers Solar pressure mismodeling can contribute ~ 10,000 km trajectory error AMD mismodeling can contribute ~ 7,000 km trajectory error These effects are inexpensive at TCMs-1,2, but can be costly at TCMs-3,4 April 2, 2002 Page - 23

24 Low-Torque Attitude Low-torque configuration starting at MOI - 50 days Reduces desat frequency from ~1/day to ~1/week Desat ΔV per event drops from ~ 8 mm/s to ~2 mm/s Deterministic trajectory change per event decreases significantly Minimizes predict bias error At the time of TCM-4 Design (MOI-16 days) the deterministic altitude change remaining due to predicted AMDs : Original Torque Profile: -80 km (Altitude Drop) Low-Torque Profile: 5 km (Altitude Raise) April 2, 2002 Page - 24

25 Maneuver Design Trajectory Correction Maneuvers (TCMs) Clean up Injection Errors from Upper Stage Remove Injection Bias Correct Targeting Errors Maneuver execution errors Orbit Determination errors Satisfy Planetary Quarantine (PQ) Requirements Achieve Injection Conditions April 2, 2002 Page - 25

26 Maneuver Analysis Statistical propellant usage calculated via Monte-Carlo analysis based on the nominal trajectory, and expected trajectory dispersions, due to Launch vehicle injection dispersions Orbit Determination errors Maneuver execution errors Usually quoted as ΔV 99 (99% of cases require no more than) PQ analysis is the calculation of aimpoint biases required to ensure that the probability of impacting a planetary body is sufficiently small Probability of Impact calculated on each trajectory leg Includes probability of not being able to perform another maneuver Based on expected trajectory dispersions Generally presented in terms of B-plane aimpoints and dispersion ellipses April 2, 2002 Page - 26

27 Planetary Protection Requirements COSPAR 1964: a sterilization level such that the probability of a single viable organism aboard any spacecraft intended for planetary landing or atmospheric penetration would be less than 1 x 10-4 a probability limit for accidental planetary impact by unsterilized fly-by or orbiting spacecraft of 3 x 10-5 or less At that time, it was thought that Mars had a life-harboring environment Liquid water on the surface Water ice caps Atmospheric pressure ~ 85 mbar This led COSPAR to assign a probability of 1.0 that a terrestrial organism would grow on the planet NASA s requirements for the Viking missions: 10-3 or less of contaminating Mars. Combination of the following probabilities: Survival of organisms in space vacuum, temperature, and UV flux Arrival of organisms at Mars Survival or organisms through atmospheric entry Release of organisms from the lander Growth and proliferation of terrestrial organisms on Mars April 2, 2002 Page - 27

28 Planetary Protection Requirements NASA s Revisions 1988: Category I: Spacecraft targets such as the Moon or Sun Category II, III, IV: Flybys, orbiters, landers, and probes sent to planets or targets with increasing exobiological interest Category V: Sample return missions Specific Missions: Viking 1 and 2 Landers: Substantial heating to produce P ~ 10-6 or less of contamination Mars Observer: Category III orbiter Launch aimpoint bias P ~ 10-5 or less Spacecraft maneuvers P ~ 10-4 or less Orbit maintained until Dec. 31, 2008; P > 0.95 of impact until Dec. 31, 2038 Mars Global Surveyor: Category III orbiter Mars Pathfinder: Category IV lander Mars 96: Category IV lander April 2, 2002 Page - 28

29 Mars Odyssey Navigation Navigation Major Events Injection TCM-1 TCM-2 TCM-3 TCM-4 TCM-5 (Contingency) MOI Period Reduction Maneuver April 2, 2002 Page - 29

30 Interplanetary Trajectory Earth at Arrival 1.02 AU Vernal Equinox Launch: 07-APR AU 45.9 TCM-5 E-7 h Mars Arrival 24-Oct-2001 TCM-4 E-12 d 10 day time ticks Mars at Launch TCM-1 L + 46 d TCM-2 L + 86 d April 2, 2002 Page - 30 TCM-3 E - 37 d

31 The B-plane April 2, 2002 Page - 31

32 Mars Odyssey Navigation April 2, 2002 Page - 32

33 Mars Odyssey Navigation TCM-1 Execution Date: 23-May-01 April 2, 2002 Page - 33

34 Mars Odyssey Navigation TCM-2 Execution Date: 02-July-01 April 2, 2002 Page - 34

35 Mars Odyssey Navigation TCM-3 Execution Date: 17-Sept-01 April 2, 2002 Page - 35

36 Mars Odyssey Navigation TCM-4 Execution Date: 12-Oct-01 Target! Alt:!300 km! Inc:!93.47!! Current Estimate (OD034)! Alt:!324.1±11 km! Inc:!94.10 ±0.2!!!! Current Miss (Est-Target)! Alt:!+24 km! Inc:!+0.6!! TCM-4 to Correct Miss ΔV:!0.08 m/s! April 2, 2002 Page - 36

37 Data Type Contributions to the Solution OD Knowledge at the time of TCM-4 Design (3σ) April 2, 2002 Page - 37

38 MOI Configuration Thrust Vector Velocity April 2, 2002 Page - 38

39 Mars Odyssey Navigation MOI and PRM MOI Burn to Oxidizer depletion to minimize Capture Orbit Period Main Engine Thrust: N Oxidizer mass available: kg ==> 1183 sec burn Design Start time: 24-OCT :26:19 UTC - ERT Magnitude: 1426 m/s Pitch rate: deg/sec (44.1 deg in 1183 sec) Expected Capture Orbit 300 km post-moi periapsis altitude 19.9 hour period PRM Period Reduction Maneuver Scheduled for 3rd Periapsis after MOI (P4) Perform PRM (if necessary) to ensure completion of Aerobraking If post-moi orbit period < 22 hrs => No PRM If post-moi orbit period > 22 hrs => PRM to reduce period to 20 hrs April 2, 2002 Page - 39

40 Mars Orbit Insertion Burn Start +00:00:00 Enter Earth Occult +00:09:37 Enter Solar Eclipse +00:09:59 Exit Solar Eclipse +00:11:46 Periapsis km +00:12:51 Burn End +00:19:44 Exit Earth Occult +00:29:24 April 2, 2002 Page - 40

41 MOI - View from Earth Burn Start +00:00:00 Goal: Altitude: 300 km ± 25 km Inclination: 93.5 ± 0.2 Enter Earth Occult +00:09:37 Achieved: Altitude: km Inclination: Exit Earth Occult +00:29:24 April 2, 2002 Page - 41

42 Aerobraking Nav Prediction Accuracy Requirement Must predict Periapsis Time to within 225 sec Must predict Periapsis Altitude to within 1.5 km Capability Altitude requirement easily met with MGS gravity field (Nav Plan) Timing requirement uncertainty dominated by assumption on future drag pass atmospheric uncertainty Atmospheric Variability Total Orbit-to-Orbit Atmospheric variability: 80% (MGS: 90%) Periapsis timing prediction To first order, the expected change in orbit period per drag pass will indicate how well future periapses can be predicted This simplifying assumption is supported by OD covariance analysis April 2, 2002 Page - 42

43 Nav Predict Capability Example Total expected Period change for a given drag pass is 1000 seconds Atmosphere could change density by 80% Resulting Period change could be off by 80% = 800 sec If orbit Period is different by 800 seconds, then the time of the next periapsis will be different by 800 seconds This fails to meet the 225 sec requirement Large Period Orbits Period change per rev is large Therefore can never predict more than 1 periapsis ahead within the 225 sec requirement with any confidence Small Period Orbits Period change per rev is small (for example 30 seconds) Therefore can predict several periapses in the future to within the 225 second requirement Example: 80% uncertainty (24 sec) will allow a 9 rev predict April 2, 2002 Page - 43

44 Aerobraking Navigation Process Long Orbits Drag Pass A1 A2 Collect Tracking Data P1 P2 P3 Analysis And Uplink Tp < 225 sec Tp > 225 sec Drag Pass (No Comm) A1 P1 Collect Tracking Data Nav Analysis Sequence Update & Uplink April 2, 2002 Page - 44

45 Navigation Process Short Orbits P1 A1 P2 A2 P3. Pn An Pn+1 Tp < 225 sec Tp > 225 sec Collect Tracking Data Nav Analysis Sequence Update & Uplink April 2, 2002 Page - 45

46 What contributed to MOI success? A Baseline set of Navigation solution strategies were identified Varied data arcs, data types, data weights, parameter estimates, a-prioris These solutions were regularly performed and trended Built a time history of trajectory solutions Trended evolution of parameter estimates and encounter conditions Lessons learned from MCO and MPL Regularly demonstrate consistency to Project and NAG Weekly Status Reports Daily Status after TCM-4 (MOI-12 days) Daily Show Shadow navigators Independent solutions run by Sec312 personnel (Bhaskaran, Portock) April 2, 2002 Page - 46

47 Conclusions Questions, comments, etc. April 2, 2002 Page - 47