AAS 11-510 ARTEMIS MISSION OVERVIEW: FROM CONCEPT TO OPERATIONS David Folta* and Theodore Sweetser† ARTEMIS (Acceleration, Reconnection, Turbulence and Electrodynamics of the Moon’s Interaction with the Sun) repurposed two spacecraft to extend their useful science (Angelopoulos, 2010) by moving them via lunar gravity assists from elliptical Earth orbits to L and L Earth-Moon libration orbits and then to 1 2 lunar orbits by exploiting the Earth-Moon-Sun dynamical environment. This paper describes the complete design from conceptual plans using weak stability transfer options and lunar gravity assist to the implementation and operational support of the Earth-Moon libration and lunar orbits. The two spacecraft of the ARTEMIS mission will have just entered lunar orbit at this paper’s presentation. INTRODUCTION ARTEMIS, the Acceleration, Reconnection, Turbulence and Electrodynamics of the Moon’s Interaction with the Sun mission, repurposed two in-orbit NASA spacecraft to extend their useful science investigations.1,2,3 ARTEMIS uses simultaneous measurements of particles and electric and magnetic fields from two different trajectories to provide three-dimensional perspectives of how energetic particle acceleration occurs near the Moon's orbit, in the distant magnetosphere, and in the solar wind. The two spacecraft denoted P1 and P2, are from NASA’s Heliophysics Time History of Events and Macroscale Interactions during Substorms (THEMIS) constellation of five satellites shown in Figure 1 that were launched in 2007 and successfully completed their mission.4 The ARTEMIS mission transferred the two THEMIS spacecraft that were in the outer-most elliptical Earth orbits and, with lunar gravity assists, re-directed the spacecraft to both Earth-Moon L (EM L ) and L (EM 1 1 2 L ) libration point orbits via transfer 2 trajectories that exploited the multi-body dynamical environment.5,6 After the Earth- Moon libration point orbits were achieved and maintained for several months, both spacecraft were inserted into elliptical lunar orbits. The current baseline is a multi-year mission that began with orbit raise maneuvers in June 2009 that eventually targeted multiple lunar flybys in Figure1.TheARTEMISSpacecraft February and March of 2010 to place the spacecraft on their transfer trajectories. The P1 spacecraft entered Earth-Moon Lissajous orbit August 25th 2010 and P2 followed on October 22nd 2010. The trajectory design encompassed the entire near Earth environment: cis- lunar, Sun-Earth, Earth-Moon libration, and lunar orbits. Both spacecraft have been successfully * NASA/Goddard Space flight Center, Code 595, Greenbelt, MD, 20771, [email protected] † Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA 91109, [email protected] 1629 inserted into stable elliptical lunar orbits, one posigrade and the other retrograde, where they will remain for the foreseeable future. The ARTEMIS concept avoided extended shadow durations while in their Earth elliptical orbits that otherwise would have led to the demise of these THEMIS spacecraft due to re-entry requirements. The concept moved these spacecraft to a region near the Moon in order to improve the baseline for in-situ measurements of the Earth environment. Unfortunately, the limited remaining fuel did not support a direct transfer from the Earth elliptical orbits to lunar orbits. To overcome this problem a unique design was fashioned that incorporated apogee raising, lunar gravity assists, the favorable use of perturbations in Sun-Earth weak stability regions to raise perigee, and the utilization of the Earth-Moon libration orbits to reach the intended lunar orbits. This paper describes the complete design process from the conceptual plan using multi-body transfer options and the eventual use of lunar gravity assist to the implementation and operational support of the design. We discuss the impacts and limitations of the design with respect to the spacecraft constraints in terms of restricted Delta-velocity ((cid:39)v) directions and limitations of the propulsion system. We also present information on propellant usage and the sensitivity of controlling these unique orbits. We discuss the factors that contributed to the project's resounding success despite the high risks of the proposed implementation. The ARTEMIS mission is a collaborative effort between NASA’s Goddard Space Flight Center’s (GSFC) Navigation and Mission Design Branch (NMDB), the Jet Propulsion Laboratory (JPL) Mission Design and Navigation Section (MDNav), and the University of California at Berkeley (UCB) Space Sciences Laboratory (SSL). JPL provided the initial concept and the reference transfer trajectory from the elliptical orbit phase through libration orbit insertion and the details of the lunar orbit phase. GSFC’s NMDB provided the operational trajectory design to complete the mission and the maneuver and navigation support from pre-lunar gravity assists and Sun-Earth transfers to Earth-Moon libration maintenance and lunar orbit insertion. The UCB SSL Mission Operations Center (MOC) provides spacecraft operations support (command, telemetry, maneuver planning, and daily monitoring and maintenance) of all spacecraft. Tracking, telemetry, and command services are provided using the S-band frequency via various networks, including the Berkeley Ground Station (BGS), the Universal Space Network (USN), and the NASA Ground Network (GN) and Deep Space Network (DSN). The UCB SSL uses a GSFC software package for maneuver planning and navigation support via the THEMIS GSFC-UCB mission support partnership. Mission Concept and Design The ARTEMIS mission was approved in May 2008 by NASA's Heliophysics Senior Review panel as an extension to the THEMIS mission.3 The proposal encompassed a baseline design which showed that two of the THEMIS spacecraft could be placed into lunar orbit with the remaining fuel onboard. While the amount of fuel required to reach the Moon directly was available, it was insufficient to insert into a lunar orbit, so a low-energy trajectory design was used. This alternate design was first investigated at JPL using software tools that could model the Earth elliptical orbits, the multi-body environment for the transfer and Earth-Moon libration orbits, and the lunar orbits. The complete approach as shown in Figure 2 encompassed: (cid:120) Elliptical Earth Orbits o Raise apogee of P1 and P2 orbits 1630 o Utilize multiple lunar approaches to increase the spacecraft energy (cid:120) Trans-lunar trajectory o Use lunar fly-by(s) to send spacecraft towards the Earth-Sun Lagrange points o Free insertion into Earth-Moon Lissajous orbits (cid:120) Lissajous orbit phase o Place P1 and P2 into complementary Lissajous orbits and transfer to lunar orbit (cid:120) Lunar orbit phase o Insert P1 and P2 into long-term stable, complementary Lunar orbits (cid:120) Gravity Models and Assumptions (from start to lunar orbit insertion) (cid:120) Earth J2 and 8X8 models and point mass bodies: Sun and Moon (cid:120) Deterministic, impulsive maneuvers as needed to attain target goals (cid:120) 20x20 Lunar gravity and finite burns for lunar orbits. ARTEMIS takes advantage of 4-body dynamics to minimize (cid:39)v. The design raised the elliptical orbit apogee just enough for a lunar encounter. The lunar flyby provided the change in energy and direction to attain a transfer beyond lunar distance (but still loosely captured by the Earth-Moon system). For P1, one lunar flyby set up another flyby two weeks later and 180 degrees away across the lunar orbit. Once in the Sun-Earth environment, solar gravity perturbations raised perigee to lunar orbit distance for the flyby. After the timing is achieved with respect to the lunar position, the design approached the Moon along the Earth-Moon line, so that Earth gravity perturbations reduce lunar- relative energy. The Earth-Moon dynamics allowed ARTEMIS to enter, exit, and cross between Lissajous orbits around EM L and EM L with no or very small maneuvers. After lunar capture, 1 2 subsequent periselene maneuvers reduce aposelene to the desired altitude for the science. Figure 2. ARTEMIS Concept Design Figure 2. Original Mission Design 1631 Software Tools The software used to design this original concept was a combination of JPL’s LTool and Mystic.7 LTool’s trajectory design objects and differential corrector were used to find a trajectory candidate that assumed impulsive maneuvers in a multi-body environment, but ignored spacecraft thruster constraints. LTool applied successively tighter (cid:39)v constraints to reduce total (cid:39)v. In this design process we could check eclipse times. We also had to estimate the (cid:39)v cost of meeting thruster constraints while adding trajectory correction maneuvers (TCMs) to find a (cid:39)v-99 solution, which represents a statistical maximum (cid:39)v requirement to fly the mission. In the original concept, biased maneuvers in the –Z Sun-Earth rotating System direction were added at TCM locations to prove the feasibility of doing TCMs without violating thruster constraints. To better model the constraints and use higher fidelity models, analysis was switched to the Mystic software to minimize the total transfer (cid:39)v with thruster constraints and bias maneuvers as necessary. A parallel effort also used Mystic to replace impulsive orbit raise maneuvers (ORMs) with series of finite burns. Once the mission was given permission to proceed, GSFC became involved with the verification, mission design, and maneuver planning. At that time, a models and constants document was written to coordinate design efforts. GSFC and JPL software included full ephemeris models (DE421 file) along with third body perturbations including a solar radiation pressure acceleration based on the spacecraft mass and constant cross-sectional area (e.g. cannon ball model). A potential model for the Earth with degree and order eight was used. The operational plans are based on a variable step Runge-Kutta 8/9 or PrinceDormand 8/9 integrator. The libration point locations are also calculated instantaneously at the same integration interval. To compute maneuver requirements in terms of (cid:39)v, different strategies involve various numerical methods: traditional Differential Correction (DC) targeting with central or forward differencing, optimization using the Mystic Optimizer and the VF13AD optimizer from the Harwell library, and the Analytical Graphics Inc (AGI) /Satellite toolkit (STK) SQP optimizer. For the corrections scheme, equality constraints are incorporated, while for the optimization scheme, nonlinear equality and inequality constraints are employed. The software employed to met spacecraft constraints and orbit goals for operational trajectory design and maneuver planning included GSFC’s General Mission Analysis Tool (GMAT) (open source s/w), Analytical Graphic Inc’s (AGI) STK/Astrogator, and GSFC’s Goddard Maneuver program (GMAN).8 GMAN is high fidelity propulsion-modeling software that incorporates moments of inertia and detailed thruster models to simulate the kinematics and dynamics of the spinning ARTEMIS spacecraft. GMAN has been used successfully over 30 years to model spinning spacecraft and is being used for THEMIS support. Navigation Since the navigation solution is provided both by the UCB team and the GSFC Code 595 Flight Dynamics Facility, we were able to plan maneuvers with confidence.9,10 The observed navigation uncertainty was significantly smaller than the values used in the pre-flight assessment. The tracking of P1 and P2 was accomplished using the Deep Space Network (DSN), Universal Space Network (USN), and the antenna at UCB. More information on navigation can be found in Reference 9. The Goddard Trajectory Determination System (GTDS) was used for all operational navigation solutions. The GTDS least squares solution uncertainty, found from the comparison of overlap regions of the navigation solution, was estimated to be below 100 meters and 0.1 cm/s in all phases of the mission. Originally, as a conservative estimate for maneuver planning and error analysis, 1(cid:86) uncertainties of 1 km in position and 1 cm/s in velocity were used. We believe that 1632 the observed uncertainty values were optimistic and that they actually range in the 100s of meters and tens of cm/s. Additionally, it was difficult to ascertain the correct navigation error near maneuvers since both maneuver performance and attitude uncertainty were at the limit of observability. Throughout the transfer trajectory implementation process, navigation solutions were generated at a regular frequency of once every three days, while in the orbit raise and Lissjaous orbits daily solutions were generated. Post-maneuver navigation solutions were made available as soon as a converged solution was determined. The rapid response was to ensure that the maneuver had performed as predicted and that no unanticipated major changes to the design were necessary. These accuracies were obtained using nominal tracking arcs with alternating north and south stations. ARTEMIS SPACECRAFT OVERVIEW Each ARTEMIS spacecraft is spin-stabilized with a nominal spin rate of roughly 20 RPM. Spacecraft attitude and rate are determined using telemetry from a Sun sensor (SS), a three-axis magnetometer (TAM) used near Earth perigee, and two single-axis inertial rate units (IRUs). The propulsion system on each spacecraft is a simple monopropellant hydrazine blow-down system. The propellant is stored in two equally-sized tanks and either tank can supply propellant to any of the thrusters through a series of latch valves. Each observatory was launched with a dry mass of 77 kg and 49 kg of propellant, supplying a wet mass of 126 kg at beginning of life. Each spacecraft has four 4.4 Newton (N) thrusters – two axial thrusters and two tangential thrusters. The two tangential thrusters are mounted on one side of the spacecraft and the two axial thrusters are mounted on the lower deck, as seen in Figure 3. The thrusters fire singly or in pairs – in continuous or pulsed mode – to provide orbit, attitude, and spin rate control. Orbit maneuvers were implemented by firing the axial thrusters in continuous mode, the tangential thrusters in pulsed mode, or a combination of the two (beta mode). Since there are no thrusters on the upper deck, the combined thrust vector is constrained to the lower hemisphere of the spacecraft. ARTEMIS Spacecraft Maneuvers Constraints Each ARTEMIS probe’s spin axis is pointed within eight degrees of the south ecliptic pole. These spacecraft can implement a (cid:39)v (thrust direction) along the spin axis towards the south ecliptic pole direction or in the spin plane, but cannot produce a (cid:39)v in the northern hemisphere relative to the ecliptic. While the axial thrusters were used when necessary, these thrusters are not calibrated as well as the tangential (radial) thrusters. The pointing constraint limited the location of maneuvers so most maneuvers were Figure 3. ARTEMIS Spacecraft performed in a radial direction. For the lunar gravity assist and the multi-body dynamical environment, the trajectory was optimized using a nonlinear constraint that placed the (cid:39)v in the spin plane. The maneuver epoch was also varied to yield an optimal radial maneuver magnitude. In addition to the direction of maneuvers, another ‘error’ source also resulted in some interesting maneuver planning. This is the fact that, as a spinning spacecraft, a maneuver will be quantized into (cid:39)v pulses of ~1.5 cm/s each, with a start time that is dependent upon the Sun pulse in each spin. This meant that the achievable finite maneuver accuracy was dependent upon the (cid:39)v 1633 magnitude for each maneuver. Maneuvers were quantized by varying the maneuver epoch, but DSN coverage often led to this method not being easily enacted. Thus many early maneuvers were executed with the associated errors from spin pulse and timing. Later in the mission, the UCB operations team updated the onboard software to permit a variable spin-pulse to allow a moe accurate match to the required (cid:39)v, reducing the uncertainty in the (cid:39)v per maneuver to less than 1%. TEAM DISCIPLINES As mentioned in the background, ARTEMIS is a team effort and the process to plan and execute maneuvers demonstrates how that team process worked. Upon receipt of the daily orbit determination solution, a maneuver was computed for possible maneuver locations to meet the tracking and command load schedule. Even if an optimal (cid:39)v was found that minimized fuel use, the epoch of the maneuver needed to be contained within a scheduled tracking pass for the upload and verification in real time of maneuver execution. This meant that there were epochs and maneuver locations that did not meet truely optimal placement, but rather minimized (cid:39)v for the station contact. Maneuver plans were then generated as part of the optimization and transmitted to UCB SSL for further processing within the GMAN program to target these constrained optimal (cid:39)Vs. GMAN output was then sent to GSFC for verification of the maneuver plan and for an initial estimate of the next maneuver, since navigation and performance errors would result in the orbit eventually escaping. In a weekly setting, full team meetings were held for a complete presentation and discussion of trajectory design analysis. These analyses included; maneuver (cid:39)v estimates and contingency plans, Monte Carlo analysis, optimization results, navigation accuracies, spacecraft status, and tracking and contact schedules. The team not only discussed the analysis, but also cooperated in serving as reviewers of the plans as they related to the original concept design and support operations. MISSION PHASES With the conceptual design and mission support in place, the team of UCB-SSL, JPL, and GSFC began the endeavor to make this mission a reality. Beginning in July of 2009, the design was ready to be translated to high fidelity operational software that could model detailed spacecraft and thruster characteristics and all the multi-body regions that ARTEMIS would survey and transit. Orbit Raise Maneuvers Orbit Raise Maneuvers (ORMs) were required for each of the spacecraft.11 The P1 spacecraft was initially in a higher apogee orbit of 120,000 km than the P2 spacecraft at 80,000 km. With a low thrusting capability, multiple apoapsis raising maneuvers were executed to achieve the lunar flyby conditions necessary to place the spacecraft on their respective transfer trajectories. The number of maneuvers in the ORM sequence was a function of the current apoapsis, the propulsion system capability, station coverage, eclipse avoidance, and overall efficiency with long duration maneuvers near periapse. Several ORMs were executed as two maneuvers separated by the Earth shadow cone. All these orbital constraints along with the operational constraints of the ARTEMIS spacecraft resulted in a need to begin the orbit raise sequence in the appropriate time to achieve the final apoapsis distance and timing needed for arrival at the Moon for the planned flyby. There were several intermediate lunar approaches with distances on the order of 50,000 km to just over 11,000 km, during the ORM sequence. These encounters were designed so the resulting perturbations would improve the overall design. The P2 spacecraft had 1634 more fuel onboard since the THEMIS design moved the most efficient spacecraft, P1, to the higher apoapsis, thus P1 had less fuel remaining. Figure 4 shows the ARTEMIS P1 trajectory from the end of the nominal THEMIS mission through the first close lunar flyby. In the figure, the red line represents the ARTEMIS P1 trajectory and the gray circle indicates the Moon’s orbit. The plot is centered on the Earth and shown in the Sun-Earth synodic coordinate frame, which rotates such that the Sun is fixed along the negative X axis (to the left) and the Z axis is aligned with the angular momentum of the Earth’s heliocentric orbit. As time passes, the line of apsides of P1’s geocentric orbit rotates clockwise in the main figure. The insert in the bottom left shows P1’s motion out of the ecliptic plane, where the largest plane change was caused by a lunar approach in December 2009. The labels on the plot provide information about key events during this phase of the mission. The design of the P2 Earth orbits phase was similar, as shown in Figure 5, but lasted two months longer because it started from a smaller Earth orbit and a longer series of finite maneuvers needed to be included to raise the orbit. As we gradually came to realize, the reference trajectory design for the Earth orbit phase of both P1 and P2 would turn out to be significantly more complex than a simple series of maneuvers to replace the preliminary design’s single impulsive orbit raise maneuver. This complexity stemmed from: (1) probe operational constraints, (2) the tight (cid:39)v budget, (3) the precision phasing required to reach the designed low-energy transfers to the Moon, (4) lunar perturbations of intermediate orbits, and (5) the actual initial states for ARTEMIS P1, P2 in the summer of 2009. These actual initial states ended up significantly different from the initial states that were predicted in 2005-2007; this change was due to deterministic orbit-change maneuvers that occurred in 2008, mid-way through the THEMIS mission, to improve science yield for the second THEMIS tail. As expected, the actual orbit raise required perigee burns on multiple orbits due to the small thrust capability. The design of these burns was challenging because generally an optimal design of highly elliptical transfers is numerically difficult, and because lunar approaches created a complex three-body design space. Figure 4. P1 Earth Orbit Raising maneuver Sequence 1635 Figure 5. P1 Earth Orbit Raising maneuver Sequence Lunar Gravity Assist As part of the final transfer trajectory design, lunar gravity assists were required. These flybys targeted an encounter through the lunar B-Plane. The B-Plane parameters are shown in Table 2. Table 1. B-plane Target Parameters B-Magnitude B-Angle** Epoch P1 – First Flyby 21659.211 -96.44 January 31, 2010 08:10:13 P1 – Second Flyby 11931.05 -11.73 February, 15 2011 09:54:34 P2 – Single Flyby 16933.33 -6.23 March, 28 2010 07:58:11 ** B-plane reference Vector is lunar orbit normal Once the ORM sequence was completed, several revolutions of the Earth elliptical orbit were made available for correction maneuvers and for Flyby Targeting maneuvers (FTMs) before the lunar gravity assist. These maneuvers were optimized as a total maneuver sequence to minimize the overall targeting (cid:39)v. Transfer Trajectory Following the first close lunar gravity assist, the P1 spacecraft flew under the Moon’s orbit plane and performed a second gravity assist roughly 13 days later, as seen in the Sun-Earth rotating frame in Figure 6. A Deep Space Maneuver (DSM1) was performed 33 days later. DSM1 targeted through Earth periapse to the Earth-Moon libration insertion state. Following the Earth periapse, the P1 spacecraft once again transferred into the general vicinity of the Sun-Earth L 1 Lagrangian point. This region is also identified as a “weak stability boundary” region. At the final bend in the P1 trajectory, the spacecraft was at a maximum range of 1.50 million km from the Earth. At this point, the trajectory begins to fall back towards the Earth-Moon system on an 1636 unstable Lissajous manifold. A second deep space maneuver (DSM2) originally modeled was not required to target the Earth-Moon L Lagrangian point. A small Lissajous insertion orbit 2 maneuver was performed to insert P1 into the proper L Lissajous orbit. The P2 translunar 2 trajectory included a single lunar swingby, three deep space maneuvers, two Earth periapses, and the Lissajous orbit insertion maneuver.12,13 For both P1 and P2, we allocated 4% of the total propellant budget to perform any required trajectory correction maneuvers (TCMs) along the way. The trajectory design focused on achieving the Earth-Moon libration insertion conditions to permit the final stages of the ARTEMIS mission, which included a Lissajous orbit with a transfer to a high eccentric lunar orbit. The flyby targets were required to provide the correct energy needed to place the ARTEMIS spacecraft near the appropriate outgoing manifolds. Since the two spacecraft were originally designed for a different mission, a highly elliptical Earth orbit, and were already flying, fuel was (and is) extremely limited. Thus, within the unique operational constraints, accomplishment of the transfer goals with the minimum cost in terms of fuel was the highest priority. The total (cid:39)v for this phase of the mission was ~ 11 m/s for P1 and ~33 m/s for P2. Figure 6. P1 Transfer Trajectory 1637 Figure 7. P2 Transfer Trajectory Earth-Moon Libration Orbits It was already known that any change in energy from an unstable Earth-Moon libration point orbit would result in an orbit departure, either towards the Moon or in an escape direction towards the Earth or the Sun-Earth regions. The (cid:39)v required to effect these changes is very small, comparable to the small accelerations from solar radiation pressure, so natural perturbations will also result in these escape trajectories. To continue the orbit downstream and maintain the path in the vicinity of the libration point, this information can be exploited to selectively choose the goals that must be achieved to continue the orbit from one side of the libration orbit to the other.14,15,16 For the method applied directly to ARTEMIS the goals are directly related to the energy (velocity) at the x-axis crossing to simply wrap the orbit in the proper direction, always inward and towards the libration point. The targets used for the ARTEMIS optimal continuation method differed slightly between the EM L orbit and the EM L orbit.15 The continuation maintaining P1 while in orbit about EM L 2 1 2 used two different x-component-of-velocity targets, depending on which side of the orbit P1 was on. For example, targets on the far side of the orbit in the rotating frame (farther from the Earth) used an x-velocity target of -20 m/s at the x-axis crossing with a tolerance of 1 cm/s. Targets on the close side (nearer to the Earth) used x-axis crossing velocity targets of +10 m/s with a tolerance of 1 cm/s. Once in orbit about the EM L orbit the P1 targets were changed to make 1 ongoing operations similar to P2. These targets are +/- 10 cm/s at each crossing, a much smaller x-velocity target. The scheme here is to continuously target the next crossing downstream. Up to four crossings were used as the decrease in the stationkeeping (cid:39)v after meeting the third crossing target was usually below 0.1 cm/s and therefore unachievable by the spacecraft propulsion system. Depending on the location of the maneuver with respect to the lunar radius, the (cid:39)v also varied from maneuver to maneuver. 1638