AIAA 2014-1745 SpaceOps Conferences 5-9 May 2014, Pasadena, CA Proceedings of the 2014 SpaceOps Conference, SpaceOps 2014 Conference Pasadena, CA, USA, May 5-9, 2014, Paper DRAFT ONLY AIAA 2014-1745.

Collision Avoidance Operations in a Multi-Mission Environment

Manfred Bester,1 Bryce Roberts,2 Mark Lewis,3 Jeremy Thorsness,4 Gregory Picard,5 Sabine Frey,6 Daniel Cosgrove,7 Jeffrey Marchese,8 Aaron Burgart,9 and William Craig10 Space Sciences Laboratory, University of California, Berkeley, CA 94720-7450

With the increasing number of manmade object orbiting Earth, the probability for close encounters or on-orbit collisions is of great concern to spacecraft operators. The presence of debris clouds from various disintegration events amplifies these concerns, especially in low- Earth orbits. The University of California, Berkeley currently operates seven NASA spacecraft in various orbit regimes around the Earth and the Moon, and actively participates in collision avoidance operations. NASA Goddard Space Flight Center and the Jet Propulsion Laboratory provide conjunction analyses. In two cases, collision avoidance operations were executed to reduce the risks of on-orbit collisions. With one of the Earth orbiting THEMIS spacecraft, a small thrust maneuver was executed to increase the miss distance for a predicted close conjunction. For the NuSTAR observatory, an attitude maneuver was executed to minimize the cross section with respect to a particular conjunction geometry. Operations for these two events are presented as case studies. A number of experiences and lessons learned are included.

Nomenclature dLong = geographic longitude increment ΔV = change in velocity dZgeo = geostationary orbit crossing distance increment i = inclination Pc = probability of collision R = geostationary radius RE = Earth radius σ = standard deviation Zgeo = geostationary orbit crossing distance

I. Introduction PACECRAFT operators are concerned about close approaches between their spacecraft and other operational Sspacecraft or orbital debris, and need to be prepared to execute thrust or attitude maneuvers aimed towards reducing the risks of an on-orbit collision, if those capabilities exist. The space around Earth is rather crowded already, but even in less congested environments, there is a non-zero probability for collisions. Although only few

Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745 spacecraft currently operate in lunar orbits, a close encounter with another object could have catastrophic consequences.

1 Director of Operations, Space Sciences Laboratory, University of California, Berkeley, AIAA Senior Member. 2 Ground Systems Engineer, Space Sciences Laboratory, University of California, Berkeley, AIAA Member. 3 Mission Operations Manager, Space Sciences Laboratory, University of California, Berkeley. 4 Lead Flight Controller, Space Sciences Laboratory, University of California, Berkeley. 5 Lead Scheduler, Space Sciences Laboratory, University of California, Berkeley. 6 Mission Design Lead, Space Sciences Laboratory, University of California, Berkeley. 7 Navigation Lead, Space Sciences Laboratory, University of California, Berkeley. 8 Flight Dynamics Analyst, Space Sciences Laboratory, University of California, Berkeley. 9 Flight Dynamics Analyst, Space Sciences Laboratory, University of California, Berkeley. 10 Project Manager, Space Sciences Laboratory, University of California, Berkeley. 1 American Institute of Aeronautics and Astronautics

Copyright © 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. The University of California, Berkeley (UCB) currently operates seven NASA spacecraft – five in Earth and two in lunar orbits – from its multi-mission operations center at Space Sciences Laboratory (SSL).1 A summary of salient mission characteristics is provided in Table 1. The Time History of Events and Macroscale Interactions During Substorms (THEMIS) mission, a NASA Medium-class Explorer, is a five-spacecraft constellation launched in 2007 to study the physics of the aurora.2 Three of the original THEMIS spacecraft, also referred to as probes, currently operate in highly elliptical, low-inclination Earth orbits with perigees between 600 and 1,100 km, and apogees of 66,500-66,800 km. Primary concerns include periodic crossings of the geostationary belt as well as encounters with low-Earth objects near perigee. The remaining two THEMIS spacecraft departed Earth orbits in 2009 and arrived in lunar orbits in 2011, forming a new mission called the Acceleration, Reconnection, Turbulence, and Electrodynamics of the Moon’s Interaction with the Sun (ARTEMIS).3-5 The low-energy transfer with Earth and lunar gravity assists and with extended operations in lunar libration point orbits culminated in the direct insertion into highly elliptical lunar orbits. The THEMIS and ARTEMIS probes are spinning instrument platforms to measure fields and charged particles. The spin-plane wire booms that are part of the Electric Field Instrument (EFI) extend up to 50 m end-to-end. All five spacecraft carry propulsion systems and can execute ΔV maneuvers to avoid close approaches. The Nuclear Spectroscopic Telescope Array (NuSTAR), a NASA Small Explorer operating in low-Earth orbit (LEO), is a three-axis stabilized astrophysics observatory, consisting of a pair of hard X-ray telescopes with focusing optics, mounted at the end of a 10-m long mast.6 NuSTAR does not carry a propulsion system, hence the only practical way to reduce the risk of a collision is by performing an attitude maneuver to minimize the cross section for a particular conjunction geometry. Such an attitude maneuver is designed to orient the long axis of the observatory parallel or anti-parallel to the relative velocity vector between NuSTAR and the approaching object.

Table 1. Overview of Characteristics of NASA Missions Currently Operated at UCB.

THEMIS / ARTEMIS • Constellation of 5 spin-stabilized probes (P1-P5), spin rates 14-20 rpm • Electric & magnetic field and charged particle sensors, 5 instruments per probe • Mission science: Magnetospheric Physics and Heliophysics • Launch: February 17, 2007 • Two probes (ARTEMIS P1, P2) transferred to lunar orbits, beginning in 2009 • Lunar orbit insertion achieved in 2011 • Current Earth orbits (P3, P4, P5): 600-1,100 × 66,500-66,800 km, i = 8-13°

Image credits: NASA • Current lunar orbits (P1, P2): 50-1,200 × 15,900-17,500 km, i = 169° and 36° • Hydrazine propulsion systems with 4 thrusters for orbit and attitude control • Total mass per spacecraft: 127 kg (wet at launch), 78 kg (dry) • Spacecraft dimensions, tip-to-tip: 50 × 40 m in spin plane, 7 m along spin axis NuSTAR • Three-axis stabilized platform • Dual focusing hard X-ray telescopes, 10 m focal length • Mission science: Astrophysics • Launch: June 13, 2012 • Current Earth orbit: 613 × 630 km, i = 6° deg Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745 • Reaction wheels (4), no propulsion system, total mass 324 kg • Spacecraft dimensions: 11 m long axis, 4 m solar array, 1.2 m spacecraft body RHESSI • Sun-pointed, spinning platform, spin rate 12 rpm • Imaging spectrometer for X-ray and gamma ray wavelengths • Mission science: Heliophysics • Launch: February 5, 2002 • Current Earth orbit: 514 × 535 km, i = 38° deg • Magnetic torque bars (3), no propulsion system, total mass 300 kg • Spacecraft dimensions: 5.5 × 5.5 m solar arrays, 2.5 m long imager tube

2 American Institute of Aeronautics and Astronautics

The Ramaty High Energy Solar Spectroscopic Imager (RHESSI), another NASA Small Explorer, is a spinning, Sun-pointed observatory studying solar flares at X-ray and gamma ray wavelengths.7 It also operates in LEO and does not carry a propulsion system. The observatory’s cross section is not very strongly dependent on attitude. Due to torque limitations of the attitude control system the spacecraft can also not be reoriented efficiently on relatively short notice. RHESSI therefore does not have the ability to actively reduce the risk of a collision. This paper describes the planning activities and procedures that UCB developed to respond to conjunction warnings, as well as experiences and lessons learned with the process.

II. Operating in Crowded Space

A. General Concerns Operators of spacecraft in the crowded LEO environment must expect to be faced with close encounters without more than a week advance notice, sometimes even shorter. This situation was aggravated further as a result of several significant spacecraft break-up events. The most notable events creating large debris clouds were the Fengyun-1C disintegration on January 11, 2007, and the on-orbit collision of Iridium 33 and Cosmos 2251 on February 10, 2009.8,9 Due to differential precession, the debris clouds from these events have spread out around the Earth more or less evenly since the break-ups occurred. Pieces associated with these debris clouds alone account for more than 4,200 objects currently tracked by the United States Strategic Command (USSTRATCOM). Corresponding orbital elements are publicly available via the CelesTrak web site (www.celestrak.com). NuSTAR and RHESSI are operating within the crowded LEO regime. The three Earth orbiting THEMIS probes also pass through this regime near perigee of each orbit.

B. Special Concerns for THEMIS Of special concern for THEMIS are seasonal crossings of the geostationary belt. The three Earth orbiting THEMIS spacecraft, THEMIS P3, P4, and P5, also referred to as THEMIS D, E, and A, respectively, by their flight model numbers, operate in low-inclination orbits with perigees below and apogees above the geostationary altitude, as is illustrated in Figure 1. These crossings occur periodically when the distances of the ascending or descending nodes of the THEMIS orbits line up with the geostationary orbit.10 Such a case is shown in Figure 2 for THEMIS P3 (THEMIS D). Due to the low inclination and the rate of precession of the argument of perigee, these conditions typically last for 14-16 successive THEMIS orbits. Figure 3 indicates the seasonal evolution of these crossings for the current THEMIS orbit geometry. Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745

Figure 1. THEMIS P3-P5 (THEMIS D, E, A) spacecraft orbits. Current periods of these low-inclination orbits are about 22.2 h long. The grid represents the equator of the J2000.0 coordinate system. Circles are spaced by 2 RE. The geostationary orbit and its population can be recognized at a geocentric distance of 6.611 RE.

3 American Institute of Aeronautics and Astronautics

Figure 2. Example of the THEMIS P3 (THEMIS D) spacecraft crossing the geostationary belt. This alignment occurs when either the ascending or the descending node coincides with the distance R of the geostationary orbit.

Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745 Figure 3. THEMIS evolution of the inclination (left) and the geocentric nodal distances (right). Different colors represent the three spacecraft: P3 (D): dark blue, P4 (E): light blue, and P5 (A): purple. X-axis units are days counted since January 1, 2014. The vertical lines mark one year. The evolution of the nodal distances is dominated by the precession of the argument of perigee. Crossings of the geostationary orbit occur when either the ascending nodes (darker symbols) or the descending nodes (lighter symbols) coincide with the distance of the geostationary orbit at R = 42,164 km, as indicated by the horizontal line. A numerical example of a sequence of geostationary orbit crossings for THEMIS P3 (D) in July and August 2014 that correspond to the graphics shown in Figure 2 is given in Table A-1 in the Appendix. In this particular case, the spacecraft initially passes underneath the geostationary orbit at a distance Zgeo = -219 km, and approximately two weeks later crosses over the top at Zgeo = +206 km. Within this time frame the geostationary belt is crossed 14 times within ±200 km and with distance increments dZgeo of 25-31 km and longitude increments of about 25.9° per orbit, progressing in eastern direction, given the present THEMIS orbit geometry. 4 American Institute of Aeronautics and Astronautics

III. Collision Avoidance Operations NASA requires mandatory compliance with NPR 8715.6A, the NASA Procedural Requirements for Limiting Orbital Debris.11 NPR 8715.6A is applicable to spacecraft operating in Earth and lunar orbits, and states that conjunction assessment analyses shall be performed for all maneuverable Earth-orbiting spacecraft with a perigee altitude below 2,000 km or within 200 km of geosynchronous Earth orbit. All five THEMIS and ARTEMIS spacecraft, illustrated in their deployed configuration in Figure 4, are maneuverable. The Earth orbiting probes also meet both orbit conditions for which conjunction assessment is required. NuSTAR has maneuverability in attitude only. RHESSI has essentially no maneuverability, as far as mitigation of on-orbit collisions is concerned. Planning and execution of collision avoidance operations for these seven NASA spacecraft involves a complex interaction and coordination of a mission dependent subset of the following organizations: the Joint Space Operations Center (JSpOC) at Vandenberg Air Force Base (VAFB), USSTRATCOM, the Robotic Conjunction Assessment Risk Analysis (CARA) Team and the Space Science Mission Operations (SSMO) Project Office at NASA’s Goddard Space Flight Center (GSFC), the MArs (and Moon) Deepspace Collision Avoidance Process (MADCAP) Team and the NuSTAR Project Office at NASA’s Jet Propulsion Laboratory (JPL), the University of California, Los Angeles (UCLA), the California Institute of Technology (Caltech), and UCB. General and mission specific aspects of the processes involved are described in the following subsections.

A. Trajectory Determination UCB is responsible for maintaining accurate state vectors for all five THEMIS and ARTEMIS probes. The orbit determination (OD) process utilizes the Goddard Trajectory Determination System (GTDS).12 GTDS ingests two-way Doppler measurements from multiple ground stations and calculates accurate state vectors with a high-fidelity force model and a differential correction algorithm. For the two ARTEMIS probes, NASA’s Deep Space Network (DSN) also provides range measurements that are included in the OD process.13 OD arcs are typically seven days long, and the OD process runs at least once per week. Ephemerides are in turn generated by GTDS, and are used for all operational mission- planning functions at UCB.4 These ephemerides also include finite ΔV maneuver trajectory segments, and are made available to JSpOC and JPL to aid their conjunction screening processes. In addition, the UCB navigation team explicitly notifies both JSpOC and JPL when ΔV maneuvers are planned and executed. USSTRATCOM maintains a high-accuracy space object catalog (HAC) of orbital element sets, derived from observations with radar and optical sensors that are part of the Space Surveillance Network (SSN).14 Orbital element sets, including those for NuSTAR and RHESSI, are provided in two-line element (TLE) format. UCB downloads these TLE sets from the Space-Track.org web site for generation of all operational data products. The HAC also includes high-accuracy orbital elements for THEMIS A, D, and E. Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745 B. Conjunction Screening and Reporting Figure 4: THEMIS spacecraft with eight JSpOC screens ephemerides of operational spacecraft against the deployed science instrument booms (not HAC, but also uses Owner/Operator (O/O) provided ephemerides, to scale). The spin-plane wire booms of the when available.14,15 The screening process employs three different Electric Field Instrument (EFI) extend ±25 safety volumes as a filter to trigger activities, such as monitoring or m in X and ±20 m in Y direction. The tasking additional tracks to improve OD solutions. The coordinate instrument suite also includes a Fluxgate frame used to describe conjunction geometry is the radial (U), in- Magnetometer (FGM), mounted at the end track (V), and cross-track (W) system, centered on the primary of a 2-m long boom, a Search Coil (user) object. The safety volume depends on the type of orbit and Magnetometer (SCM) on a 1-m long boom, maneuverability of the user asset. For LEO objects, such as an Electrostatic Analyzer (ESA), and a pair NuSTAR and RHESSI, the tasking/alert volume is a box of size of Solid State Telescope (SST) sensors. UVW = (±0.5, ±5.0, ±5.0) km.14

5 American Institute of Aeronautics and Astronautics

The CARA Team works side-by-side with JSpOC personnel at VAFB to screen close approach predictions for all NASA robotic missions. Other CARA Team members at GSFC in turn perform additional analysis functions to quantify the risks of a potential collision. As part of this process, the CARA Team also generates automated conjunction assessment (CA) summary reports that include the Time of Closest Approach (TCA), the conjunction geometry (i.e. the miss distance vector in 14-16 the UVW frame), a calculated probability of collision (Pc), plus trending data for each conjunction. These reports are distributed to approved stakeholders, including UCB, several times per day. Conjunction warnings are typically provided up to 7-10 days ahead of the predicted TCA. The CARA Team also assists projects with specific analyses and recommendations for mitigating a selected number of higher-risk conjunctions. In addition to the conjunction geometry, the CA process also takes into account the accuracy of the OD solutions for both objects to determine Pc. A specific hard body radius (HBR) may be selected for each spacecraft, based on its physical size (also see Table 1 and Figure 4). For the Earth orbiting spacecraft operated by UCB, the CARA Team uses the following HBR values: THEMIS: 25 m in any attitude (worst case) NuSTAR: 10 m in any attitude (worst case), 5 m in optimized conjunction attitude (worst case solar array) RHESSI: 4 m in any attitude (worst case)

The HBR for the secondary object is added to obtain a combined HBR for the automated calculations of Pc. For small pieces of debris, a HBR value of 1 m is typically selected. For the two ARTEMIS spacecraft, UCB receives conjunction reports from the MADCAP Team at JPL on a daily basis.17 These reports include approach information between the two ARTEMIS probes, as well as between the ARTEMIS probes and other objects in lunar orbit. Since passive objects in lunar orbit are not tracked via radar, the accuracy for those conjunctions is lower. UCB provides inputs to MADCAP by uploading accurate mission ephemeris files for the two ARTEMIS spacecraft via the Deep Space Network (DSN) Service Preparation Subsystem (SPS) Portal on a weekly basis.

C. Operational Response Levels Operational responses to CARA (or MADCAP) conjunction warnings are categorized in three levels, based on the predicted value of Pc: -6 Level 1: 0 < Pc < 10 -6 -3 Level 2: 10 < Pc < 10 -3 Level 3: 10 < Pc < 1 -6 Any conjunction warning issued by the CARA Team with a Pc of less than 10 qualifies as a Level 1 event. Level 1 events are reported in weekly operations meetings as routine events, but no further action is taken unless a -6 refined conjunction warning indicates a Pc above 10 . However, opportunities for ground station or Space Network (SN) passes around TCA are identified, but no additional passes are scheduled yet. Trending of the miss distance vector components and of Pc is monitored. Level 2 events trigger notification of all Project stakeholders. Additional communications passes around TCA are scheduled, and communications passes prior to TCA are identified for special commanding, such as for uploading sequence tables to execute a collision avoidance maneuver, if needed. Trending of the miss distance vector Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745 components and of Pc is closely monitored. Level 3 events require planning and execution of a THEMIS ΔV maneuver or a NuSTAR attitude maneuver to reduce the probability of a collision. Additional tracking passes are scheduled for THEMIS prior to TCA to improve the OD accuracy to support the CARA Team with refining the predicted Pc. Command passes are scheduled to allow for uploading maneuver sequence tables. Maneuver plans are developed and discussed with all stakeholders, including the Principal Investigator (PI), NASA Mission Management, Project Management, UCB’s Multi-mission Operations Team (MMOT), and the CARA Team. PI and Project Management approval are required prior to executing any maneuvers. Final decisions for executing a THEMIS ΔV maneuver are made on a case-by-case basis in coordination with the CARA Team and the SSMO Project Office.

D. Level 3 Response Procedures As a first step in response to a Level 3 warning, the Mission Operations Manager (MOM) pages the Project Team and calls in flight controllers and navigation team members to support planning for a potential ΔV or attitude 6 American Institute of Aeronautics and Astronautics

maneuver to reduce the risk of a collision. Planning of collision avoidance operations is also coordinated with the CARA Team. The Project Manager (PM) notifies NASA Center and Headquarters management, as appropriate. At the discretion of the MOM, and depending on the timing between receiving a warning and TCA, the Mission Operations Center (MOC) is staffed and a corresponding maneuver is planned. The navigation team at UCB also developed procedures and scripts to execute THEMIS and ARTEMIS ΔV maneuvers and NuSTAR attitude maneuvers in near real-time, if needed, to mitigate the risk of collisions with only short advance notice. Suitable ground station or SN communications passes for primary and backup command opportunities are scheduled. The maneuver typically takes place no less than one orbit prior to TCA to allow for multiple communications opportunities, and is planned but not executed until the MOM receives final direction from the PI or his/her designee with concurrence from the PM or his/her designee. Once commanding is completed and spacecraft responses and maneuver execution are verified, the MMOT monitors spacecraft state of health (SOH), as practical, until after the conjunction event and reports the spacecraft status to the Project Team. The MOM is responsible for documenting all conjunction events in a master spreadsheet. For Level 3 events that result in an anomaly, the MOM leads a formal investigation and provides associated documentation. In such a case the anomaly is closed out via an Anomaly Technical Report that contains a detailed account of the sequence of the events related to the anomaly, a description of the associated symptoms, the probable underlying cause(s), the resulting action(s) taken, and the names and locations of the associated data files and logs. All of this information is preserved as part of the MOC reference library. Depending upon the level of severity, a mishap investigation may be initiated. Regardless of the response level, the MMOT uses any associated lessons learned during contingency operations to create new or update existing procedures as well as mission and flight rules in consultation with cognizant engineers and management. All significant anomalies and lessons learned are also entered into GSFC’s Spacecraft Orbital Anomaly Reporting System (SOARS) and the NASA Lessons Learned Information System (LLIS) databases. The following two sections describe case studies, one each for THEMIS and NuSTAR, in response to a Level 3 conjunction warning.

IV. Case Study 1 – THEMIS Collision Avoidance This section describes a case study of planning and executing a collision avoidance ΔV maneuver for one of the THEMIS probes to reduce the risk of a collision with another object.

A. Collision Risk Assessment On October 18, 2013 (DOY 291), UCB received a CA summary report that predicted a close encounter between THEMIS P3 (D) near its perigee with an object (hereafter referred to as Object B) from the Fengyun-1C debris cloud with a TCA on October 25 (DOY 298) at 07:49:11 UTC. The predicted miss distance was larger than 4,500 m -4 with a radial component below 500 m and a predicted Pc of nearly 10 . This triggered a Level 2 response with seven days ahead warning. The MOM notified stakeholders and the MMOT started to look into scheduling a communications pass around TCA. However, initial pass schedule analyses around TCA indicated that the spacecraft had no view of a supporting ground station, and that no Tracking and Data Relay Satellite (TDRS) passes were available in the SN schedule. Following a number of electronic mail exchanges between stakeholders, the UCB Navigation Lead then contacted the CARA Team to discuss providing an accurate O/O ephemeris to aid the CA process. This was the first time since launch that a close encounter was predicted for one of the THEMIS probes, and the first time that UCB Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745 had to coordinate planning and executing a ΔV maneuver with the CARA Team. The next CARA update for this conjunction was provided on October 20, when the miss distance increased to more than 5,500 m, but the Pc still held at Level 2. On October 22 (DOY 295), or three days prior to TCA, CARA generated a new update that now placed the miss distance within approximately four times the size of the THEMIS probe, and with a radial component less than its -3 size. The predicted Pc rose above 10 . This triggered a Level 3 response. The MMOT interacted with the SSMO Project Office and the CARA Team to assess the quality of the OD solutions that had significantly improved for both THEMIS P3 (D) and Object B. The last maneuvers on THEMIS P3 (D), a small orbit change maneuver with a ΔV of 8.8 cm/s and a small spin-rate control maneuver that imparted a ΔV of 1.3 cm/s, had both been completed on October 10, 2013 (DOY 283). UCB provided the latest ephemeris for THEMIS P3 (D) to CARA, and also scheduled additional tracks with two ground stations in the southern hemisphere. A TDRS pass around TCA was obtained by raising the criticality level for supporting this event.

7 American Institute of Aeronautics and Astronautics

A telephone conference was held on October 22 (DOY 295), and the following timeline was agreed upon: Later in the day, UCB would generate a new OD solution based on two-way Doppler tracks from four ground stations, including views from both the northern and southern hemisphere, and send an updated O/O ephemeris to CARA. UCB would also generate two suitable collision avoidance maneuver scenarios and send ephemerides for both, including finite thrust arcs, to CARA. CARA would then to forward these to JSpOC for screening early in the morning on October 23 (DOY 296). The maneuver would be waived off and alternate options would be considered if the JSpOC screening results came back with higher risk scenarios. Otherwise, the maneuver would be executed.

B. Thrust Maneuver Planning Process On October 23, 2013 (DOY 296), UCB refined the two maneuver scenarios with an execution near apogee, or 1.5 orbits prior to TCA. Since the orbit phasing between THEMIS P3, P4, and P5 was critical for science operations, the maneuver goals were chosen to create a ΔV parallel (Option A) or anti-parallel (Option B) to the line of apsides near apogee, using two tangential thrusters, T1 and T2, mounted in the spin plane of the spacecraft. This approach would not change the orbital period, but it would change the conjunction geometry at TCA. By contrast, firing the two axial thrusters, A1 and A2, in the present spacecraft attitude would change the orbit period and was therefore an undesirable option. It was determined that the Berkeley Ground Station (BGS), co-located with the MOC at UCB/SSL, had a view to support maneuver operations later in the same day at 23:50 UTC, and also with a back-up option on October 24 (DOY 297), one orbit later, or 0.5 orbits prior to TCA. Additional pre-maneuver and post- maneuver communications opportunities were available at the White Sands, New Mexico (WS1) and Santiago, Chile (AGO) ground stations. UCB provided the maneuver scenarios for Options A and B to the CARA Team, and CARA in turn analyzed both options and determined that the increase in radial miss distance would meet the recommended 6-σ criteria for the uncertainty of the radial component of the miss distance vector. CARA also submitted the post-maneuver ephemeris files for both options to JSpOC for additional screening to ensure that neither of these maneuver options would create new close conjunctions. The initial conditions and the maneuver goals for the selected collision avoidance maneuver Option A are summarized in Table 2. Note that the orbital period remained unchanged, so that the THEMIS P3-P4-P5 constellation alignment was unaffected.

Table 2. THEMIS P3 (D) Collision Avoidance Maneuver Parameters. Maneuver Parameter Initial Value Maneuver Goal Geocentric Perigee Radius 7343.76 km 7343.39 km Geocentric Apogee Radius 72917.71 km 72918.08 km Orbital Period 22.2241 h 22.2241 h RAAN 17.987° 17.990° Argument of Perigee 133.828° 133.824° Inclination 8.206° 8.206° ΔV − 7.4 cm/s Predicted Overall Miss Distance < 200 m > 5,000 m

Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745 C. Thrust Maneuver Execution Process In preparation for the maneuver execution, the MMOT scheduled multiple communications passes. The intent was to provide redundancy, in case a real-time contact dropped out, and also to provide options for recovery from last-minute spacecraft anomalies, such as an on-board processor reset. The following event timeline was executed on October 23/24/25 (DOY 296-298) as part of the collision avoidance maneuver operations. All times are given in UTC: 2013/296 18:00:00 - 18:30:00 Maneuver sequence table prepared for THEMIS P3 (D) 2013/296 19:00:00 - 20:00:00 Real-time simulation on the THEMIS flight simulator completed, Maneuver planning and simulation paperwork reviewed and signed by MMOT personnel 2013/296 23:00:00 - 23:20:00 Communications pass with the WS1 ground station, Sequence table uploaded,

8 American Institute of Aeronautics and Astronautics

Science instrument safing operations performed, Go/no-go poll prior to enabling the propulsion bus for autonomous maneuver execution 2013/296 23:35:00 - 00:54:15 BGS communications pass, Real-time spacecraft SOH monitoring, Final go/no-go poll prior to the thrust maneuver, Thrust maneuver execution (23:50:00 - 23:50:19), ΔV = 7.4 cm/s, Expended fuel 3.2 g, using tangential thrusters T1 and T2 with 6 pulses at a 40° thrust pulse width, nominal burn execution 2013/297 07:27:20 - 07:57:20 BGS pass for SOH monitoring and post-maneuver science instrument operations, First post-maneuver two-way Doppler track 2013/297 20:00:00 - 20:30:00 AGO pass for SOH monitoring and two-way Doppler track 2013/297 22:30:00 - 23:22:00 AGO pass for SOH monitoring and two-way Doppler track 2013/298 00:20:00 - 01:00:00 WS1 pass for engineering data recovery, including thrust history data needed for post-maneuver analysis and calibration 2013/298 06:52:00 - 07:12:00 BGS pass for engineering and science data recovery and two-way Doppler track 2013/298 07:34:00 - 08:04:00 TDRS 6 pass to monitor the conjunction, TCA at 07:34:11, no issues were noted. Post-maneuver OD solutions confirmed that the achieved miss distance was indeed larger than 5,000 m, as predicted. The overall interaction with the CARA Team worked very well. The collision avoidance maneuver planning and execution process was very successful and established valuable timeline examples and experiences to handle future events.

V. Case Study 2 – NuSTAR Collision Avoidance This section describes a case study of planning and executing a collision avoidance attitude maneuver for NuSTAR to reduce the risk of a collision with another object.

A. Collision Risk Assessment In case of NuSTAR, the CARA Team predicted a conjunction with another object (hereafter referred to as Object B) on July 1, 2013 (DOY 182) with a lead-time of 6.5 days. The predicted TCA was on July 7 (DOY 188) at 11:50:50 UTC. The predicted miss distance was above 1,000 m, but the radial component was below 100 m. The -4 calculated Pc was above 10 which triggered a Level 2 response. Over the course of the following days, the evolution of the conjunction geometry and the risk of a collision were monitored. The calculated Pc held rather steady, but on July 3 (DOY 184) reached the threshold for triggering a Level 3 event response. Multiple telephone conferences with stakeholder were scheduled, and the Project consulted with the CARA Team. On July 4 (DOY 185), the conjunction appeared to recede, dropping back to Level 2, but then on July 5 (DOY 186) it came back to Level 3. The Project therefore decided to perform an attitude maneuver to mitigate the risk of a collision. By July 6 (DOY 187), the Pc settled near the boundary between Level 2 and Level 3 responses with an overall miss distance below 300 m, placing Object B just Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745 within a 3-σ combined uncertainty of NuSTAR. For this particular conjunction, the CARA Team modeled the Pc by using a default HBR of 20 m. However, with the 5 m value that corresponds to the attitude with the lowest cross Figure 5. NuSTAR body coordinate system. section, the resulting P would drop by an order of magnitude. The +Z-axis is the boresight of the twin X-ray c telescopes. The +X-axis coincides with the long B. Attitude Maneuver Planning Process direction of the solar array, and +Y completes a The NuSTAR observatory has two main attitude control right-handed tripod. In stellar pointing mode, the modes, stellar pointing and inertial pointing, that are +Z-axis is aligned with the direction toward the employed during normal science operations. In stellar pointing observing target. The observatory then performs mode, the optical axis of the observatory is aligned with the a yaw rotation around +Z and rolls the solar astronomical X-ray target coordinates in +Z-direction of the array around +X so that it faces the Sun. body coordinate system, as shown in Figure 5. In this pointing 9 American Institute of Aeronautics and Astronautics

mode, the attitude control system then also performs an automated yaw rotation around +Z, and rolls the solar array around +X so that it faces the Sun. In inertial pointing mode the observatory points at a fixed, inertial position, much like in stellar pointing mode, but the automated yaw and solar array rotations are disabled. A combination of these two attitude control modes is used to place the observatory into an attitude that minimizes its cross section during a close conjunction. i.e. the long mast is oriented parallel or anti-parallel and the solar array edge-on to the relative velocity vector. In preparation for performing an attitude maneuver that places NuSTAR into the desired, low-cross-section attitude, UCB generates the relative velocity vector between NuSTAR and Object B (based on TLE sets downloaded from Space-Track.org) in Earth Centered Earth Fixed (ECEF) coordinates. This vector is in turn converted to Earth Centered Inertial (ECI J2000.0) coordinates, as required for conversion to a target attitude quaternion. Slew times are selected, based on knowledge of the observatory slew rates and the required slew arc lengths. The entire slew sequence is designed to arrive at the low-cross-section attitude one minute prior to TCA, and to depart one minute after TCA, so as to minimize the loss of science observations. It consists of four slew arc segments, as illustrated in Figure 6. Intermediate waypoints are needed to place the observatory into an attitude with a Sun Aspect Angle (SAA) of 90° in which the solar array faces the Sun squarely. These slew arcs have the following characteristics: 1) Stellar pointing slew from the current observing target to the SAA = 90° attitude (arc 1–2). 2) Inertial pointing slew (pure Eigen axis slew around the +X axis) from the SAA = 90° attitude to the low- cross-section attitude at TCA (arc 2–3). 3) Inertial pointing slew from the TCA attitude back to the SAA = 90° attitude (arc 3–2). 4) Stellar pointing slew from the SAA = 90° attitude back to the original observing target (arc 2–1). The third and forth slew arc segments are mirrors of the second and first segments, respectively. The main advantage of this scheme is that it utilizes existing flight software capabilities of the attitude control system, without the need to disable fault detection and correction (FDC) systems. Ground software and procedures to plan and implement these slew sequences were developed by the MMOT. Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745

Figure 6. Illustration of a NuSTAR attitude maneuver sequence. The solid black curve in this Aitoff projection indicates how the observatory is first slewed from the current science-observing attitude (1) to an intermediate attitude (2). The latter is always designed to fall onto a circle at 90° from the Sun that includes the north and south ecliptic poles (NEP, SEP). The exact location along this circle is calculated such that a second Eigen axis slew (roll) about the observatory’s X-axis then places the +Z or –Z-axis into the conjunction attitude (3), orienting the 10-m long mast parallel or anti-parallel to the relative velocity vector between the two objects. The red, green, and blue traces correspond to the boresight pointing of the three camera head units (CHUs) of the star tracker, indicating how Sun and Moon avoidance constraints are checked.

10 American Institute of Aeronautics and Astronautics

C. Attitude Maneuver Execution Process While the MOC has procedures in place to generate attitude quaternions and to build sequence tables for upload on short notice, it was decided that in this case the Science Operations Center (SOC) at Caltech should be kept in the loop in terms of processing the attitude maneuver sequence. The quaternions were given to Caltech, inserted into their planning system, and delivered back to UCB via a secure electronic interface. The advantage was that this way the SOC could cleanly maintain the observatory pointing timeline record that is essential for science data analysis. A sequence table for the collision avoidance slew sequence was built, tested on the flight simulator, and uploaded to the observatory on July 5 (DOY 186). The sequence of events (SOE) on July 7 (DOY 188), executed by the autonomous on-board sequence, is provided in Table 3.

Table 3. NuSTAR Collision Avoidance Attitude Maneuver Sequence of Events on July 7, 2013 (DOY 188). UTC Delta Time Event / Activity 11:00:07 TCA – 00:50:43 Begin of a stellar pointing slew from the current science target (Galactic Center Magnetar) to the intermediate SAA = 90° attitude, with the solar array facing the Sun. Slew arc length: 97.11°, slew duration: 1,826 s. 11:30:33 TCA – 00:20:17 Arrival at the intermediate attitude with the solar array facing the Sun at 90°, and edge on to the observatory’s Z-axis. In this case, the observatory +Z axis happened to point near the south ecliptic pole at slew completion. 11:31:37 TCA – 00:19:13 Begin of an inertial pointing slew (with the solar array frozen in place) to orient the Z-axis parallel to the relative velocity vector at TCA. Slew arc length: 52.26°, slew duration: 1,093 s. 11:40:00 TCA – 00:10:50 Begin of the scheduled TDRS 6 pass to monitor the observatory throughout the conjunction. 11:40:30 TCA – 00:10:20 Begin of the TDRS 6 real-time telemetry data flow to the MOC. 11:49:50 TCA – 00:01:00 Arrival at the collision avoidance attitude. In this attitude the surface normal of the solar array pointed 52° away from the Sun. 11:50:50 TCA – 00:00:00 Conjunction center time. 11:51:50 TCA + 00:01:00 Begin of an inertial pointing slew from the collision avoidance attitude back to the SAA = 90° attitude near the south ecliptic pole. 12:10:00 TCA + 00:19:10 End of the TDRS 6 pass. 12:10:02 TCA + 00:19:12 Arrival at the intermediate attitude near the south ecliptic pole, with the solar array facing the Sun at 90°, and still edge-on to the observatory’s Z- axis. 12:22:00 TCA + 00:31:10 Begin of a stellar pointing slew from near the south ecliptic pole to the original science target (Galactic Center Magnetar), including observatory yaw rotation, and solar array roll to keep the solar array facing the Sun. 13:00:00 TCA + 01:09:10 Arrival on the science target and egress from Earth occultation of the target.

Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745 The NuSTAR collision avoidance attitude maneuver was successfully executed. Again, the overall interaction with the CARA Team worked very well, and the planning and execution process was very successful and established valuable timeline examples and experiences for future events.

VI. Experiences and Lessons Learned This section describes a number of experiences and lessons learned that were gained during the complex collision avoidance operations scenarios.

A. Experiences During the first year of receiving regular CA summary reports from CARA, a total of 273 conjunctions across the five Earth orbiting spacecraft were tracked. The distribution between different spacecraft and as a function of response level is summarized in Table 4. A significant fraction, 104 events or 38.1%, were related to objects in the Fengyun-1C, Iridium 33, and Cosmos 2251 debris clouds. 11 American Institute of Aeronautics and Astronautics

Table 4. Summary of Conjunction Warnings Received between May 25, 2013 and May 16, 2014. Mission Level 1 Events Level 2 Events Level 3 Events Total Events THEMIS (P3, P4, P5) 116 1 1 118 NuSTAR 90 13 1 104 RHESSI 47 4 0 51 Total Events 253 18 2 273

Two conjunctions, one each for THEMIS P3 (D) and NuSTAR, were categorized as Level 3 events and resulted in planning and execution of evasive maneuvers to reduce the risk of a collision with another object. A number of predicted conjunctions were initially categorized as Level 2 events, but dissipated to Level 1 prior to TCA. No conjunction warnings were received to date for THEMIS crossings of the geostationary belt.

B. THEMIS Operations As discussed above, concerns for THEMIS conjunctions are not only close approaches with other objects near perigee, but are also related to crossings through the geostationary belt. UCB is in the process of implementing new software tools that will allow automated generation of geostationary belt crossing reports for all Earth orbiting THEMIS probes. These reports will be forwarded to the CARA Team prior to and during crossing seasons. The THEMIS thrust maneuver approach was designed to not change the orbital periods, so that the constellation geometry remained undisturbed. Procedures and software tools were originally developed to respond to conjunctions with geostationary or geosynchronous objects, but were found to be just as applicable to problems with encountering LEO objects. Scripts in the Interactive Data Language (IDL) were written to facilitate planning of THEMIS thrust maneuvers with conjunction warnings on short notice. Required input parameters are the maneuver execution time, the TCA, and the desired change in absolute position at TCA.18 Note that by firing thrusters in cross-track direction, the resulting position change at TCA will also have a significant cross-track component. Planned process improvements at UCB also include automated, daily delivery to the CARA Team of O/O ephemeris in Consultative Committee for Space Data Systems (CCSDS) Orbit Ephemeris Message (OEM) format.19 The main advantage will be that those ephemeris files also include finite burn arcs and predictive post-maneuver states, allowing the screening process to continue smoothly past ΔV maneuvers. While JSpOC is also directly notified of upcoming ΔV maneuvers, their OD process is mainly based on radar tracking and needs to be restarted with post-burn observations. It then takes at least several days to re-converge to an accurate OD solution suitable for screening of close encounters.

C. ARTEMIS Operations Spacecraft operators also have to worry about close encounters with natural debris. Such an event occurred on 5,20 October 14, 2010 on the ARTEMIS P1 (B) spacecraft during the lunar libration point (EM L2) orbit phase. As determined by post-event analysis, an EFI sensor sphere with a mass of 92 g at the end of the 25-m long -X wire boom was likely severed off by one or more micrometeorid impacts, associated with the Orionids meteor shower. Through conservation of linear momentum the event imparted an approximate ΔV of 5.7 cm/s on the spacecraft body, similar in magnitude to a typical stationkeeping maneuver in the sensitive EM L2 orbit. The anomaly was initially discovered as a residual range rate bias on subsequent two-way Doppler tracks. The resulting mass imbalance created asymmetrical thrust arms between the two tangential thrusters, T1 and T2, leading Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745 to a torque that caused significant spin rate changes during the Lunar Orbit Insertion (LOI) burn sequence. This unexpected problem needed to be taken into account in the planning and segmentation of the LOI burn arcs, as it affected the thrust pulse timing.4

D. NuSTAR Operations The NuSTAR attitude maneuver approach is elegant and simple, and was relatively easy to implement. The timeline presented in the second case study was followed flawlessly, and serves as a model for working around future conjunctions.

E. Resource Planning Receiving CARA CA summary reports and processing conjunction warnings for multiple missions up to four times per day became part of routine operations, but was not initially included in the mission operations budgets. Personnel resources are not insignificant to review CARA and MADCAP reports, to keep track of each predicted 12 American Institute of Aeronautics and Astronautics

conjunction in a multi-mission spreadsheet, to interact with scheduling and sequencing teams for planning additional communications coverage with Level 2 conjunctions, and to provide periodic status updates to Projects. In special cases meetings are scheduled to assess risks and discuss mitigation options. Additional resources need to be included in future budgets to better cover these activities.

F. Contingency Planning To further mitigate the risk of collisions, spacecraft operators should be prepared to contact operators of the other object if it turns out to be an operational satellite, so that collision avoidance operations may be coordinated. The latest version of the Space-Track.org web site provides contact information to support such interactions.

VII. Conclusion Operational procedures for collision avoidance maneuvers were successfully executed on two occasions to reduce the risk of an on-orbit collision between spacecraft operated by UCB and other orbiting objects. Complex processes to work effectively across multiple organizations were developed, and were applied in multiple operations scenarios. It was demonstrated that collision avoidance operations could be integrated into day-to-day activities without causing much disruption to science operations. Software tools to support collision avoidance operations were adequate, but will be refined further. A number of valuable experiences were gained and lessons were learned from handling a seemingly continuous stream of conjunction warnings towards protecting NASA assets and reducing the risks of creating additional orbital debris fragments.

Appendix

Table A-1. Example of THEMIS Geostationary Orbit Crossings. Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745

Acknowledgments The authors wish to thank the CARA Team at NASA/GSFC and a.i. solutions, Inc., in particular Lauri K. Newman, Ryan Frigm, Paul Frakes, Russel DeHart, Daniel Pachura, and Timothy Richardson for their dedicated support with collision analyses and risk assessment, and for commenting on the manuscript. We also wish to thank Richard Burns, Richard Harman, and Gregory Marr at the SSMO Project Office at NASA/GSFC for their support. The authors also appreciate collaboration with NuSTAR team members at Caltech (Karl Forster, Hiromasa Miyasaka, and colleagues), at JPL (Suzanne Dodd, Nazilla Rouse), and at Orbital ATK (David Oberg, Grace Baird, John Fulmer, and their colleagues). We also appreciate the support by our UCB operations team members John McDonald, Martha Eckert, Deron Pease, Renee Dumlao, and Warren Rexroad. Dr. Brandon Owens (formerly at

13 American Institute of Aeronautics and Astronautics

UCB, now at NASA Ames Research Center) was instrumental in developing the THEMIS collision avoidance maneuver procedure. THEMIS and ARTEMIS are operated by the University of California, Berkeley under NASA contract NAS5- 02099. NuSTAR work at the University of California, Berkeley is conducted under Caltech subcontract CIT-44a- 1085101 to NASA contract NNG08FD60C. The RHESSI mission is operated by the University of California, Berkeley under NASA contract NAS5-98033.

References 1Bester, M., Lewis, M., Roberts, B., Thorsness, J., McDonald, J., Pease, D., Frey, S., and Cosgrove, D., “Multi-mission Flight Operations at UC Berkeley – Experiences and Lessons Learned,” AIAA 2010 SpaceOps Conference Papers on Disk [CD-ROM], Huntsville, Alabama, April 25-30, 2010, Paper AIAA 2010-2198. 2Angelopoulos, V., The THEMIS Mission, Space Science Reviews, Vol. 141, 5-34, Springer, Dordrecht, 2008, DOI: 10.1007/s11214-008-9336-1. 3Angelopoulos, V., The ARTEMIS Mission, Space Science Reviews, Vol. 165, 3-25, Springer, Dordrecht, 2011, DOI: 10.1007/s11214-010-9687-2. 4Cosgrove, D., Frey, S., Marchese, J., Owens, B., and Bester, M., “ARTEMIS Operations from Earth-Moon Libration Orbits to Stable Lunar Orbits,” AIAA 2012 SpaceOps Conference Papers on Disk [CD-ROM], Stockholm, Sweden, June 11-15, 2012, Paper AIAA 2012-1296179. 5Bester, M., Cosgrove, D., Frey, S., Marchese, J., Burgart, A., Lewis, M., Roberts, B., Thorsness, J., McDonald, J., Pease, D., Picard, G., Eckert, M., and Dumlao, R., “ARTEMIS Operations – Experiences and Lessons Learned,” 2014 IEEE Aerospace Conference Papers on Disk [CD-ROM], Big Sky, Montana, March 1-8, 2014, Paper 2464. 6Kim, Y., Willis, J., Dodd, S., Harrison, F., Forster, K., Craig, W., Bester, M., and Oberg, D., “Nuclear Spectroscopic Telescope Array (NuSTAR) Mission,” 2013 IEEE Aerospace Conference Papers on Disk [CD-ROM], Big Sky, Montana, March 2-9, 2013, DOI: 10.1109/AERO.2013.6496933. 7Lin, R. P., Dennis, B. R., and Benz, A. O. (editors), The Reuven Ramaty High-Energy Solar Spectroscopic Imager (RHESSI) − Mission Description and Early Results, Solar Physics, Vol. 210, Kluwer Academic Publishers, Dordrecht, 2002. 8NASA Orbital Debris Program Office, Orbital Debris Quarterly News, Vol. 11 (2), 2-3, April 2007. 9NASA Orbital Debris Program Office, Orbital Debris Quarterly News, Vol. 13 (2), 1-2, April 2009. 10Frey, S., Bester, M., and Angelopoulos, V., “Crossings of Geostationary Orbit by THEMIS Scientific Satellites”, Proc. 23rd International Symposium on Space Flight Dynamics, Pasadena, California, October 29 - November 2, 2012. 11NASA Office of Safety and Mission Assurance, “NASA Procedural Requirements for Limiting Orbital Space Debris,” NPR 8715.6A, Washington, DC, May 14, 2009. 12Long, A. C., Cappellari, Jr., J. O., Velez, C. E., and Fuchs, A. J., “Goddard Trajectory Determination System (GTDS), Mathematical Theory,” Revision 1, FDD/552-89/001, NASA Goddard Space Flight Center, Greenbelt, Maryland, 1989, pp. 7-1 - 8-82. 13Roberts, B., Lewis, M., Thorsness, J., Picard, G., Lemieux, G., Marchese, J., Cosgrove, D., Greer, G., and Bester, M., “THEMIS Mission Networks Expansion – Adding the Deep Space Network for the ARTEMIS Lunar Mission Phase,” AIAA 2010 SpaceOps Conference Papers on Disk [CD-ROM], Huntsville, Alabama, April 25-30, 2010, Paper AIAA 2010-1934. 14Newman, L. K., The NASA Robotic Conjunction Assessment Process: Overview and Operational Experiences, Acta Astronautica, Vol. 66, 1253-1261, Elsevier, Amsterdam, 2010, DOI:10.1016/j.actaastro.2009.10.005. 15Frigm, R. C., Levi, J. A., and Mantziaras, D. C., “Assessment, Planning, and Execution Considerations for Conjunction Risk Assessment and Mitigation Operations,” AIAA 2010 SpaceOps Conference Papers on Disk [CD-ROM], Huntsville, Alabama, April 25-30, 2010, Paper AIAA 2010-1926. 16Chan, F. K., Spacecraft Collision Probability, The Aerospace Press, El Segundo, California, 2008. 17Berry, D. S., Guinn, J. R., Tarzi, Z. B., and Demcak, S. W., “Automated Spacecraft Conjunction Assessment at Mars and the Moon,” AIAA 2012 SpaceOps Conference Papers on Disk [CD-ROM], Stockholm, Sweden, June 11-15, 2012, Paper AIAA 2012-1262983. Downloaded by UNIV OF CALIFORNIA LOS ANGELES on July 23, 2014 | http://arc.aiaa.org DOI: 10.2514/6.2014-1745 18Interactive Data Language (IDL), Software Package, Ver. 7.1, ITT Visual Information Solutions, Boulder, Colorado, 2009. 19Consultative Committee for Space Data Systems (CCSDS), “Orbit Data Messages,” CCSDS 502.0-B-2, Blue Book, Reston, Virginia, 2009. 20Owens, B. D., Cosgrove, D. P., Marchese, J. E., Bonnell, J. W., Pankow, D. H., Frey, S., and Bester, M. G., “Mass Ejection Anomaly in Lissajous Orbit: Response and Implications for the ARTEMIS Mission,” Proc. 22nd AAS/AIAA Space Flight Mechanics Meeting, Charleston, South Carolina, January 29 - February 2, 2012, Paper AAS 12-181.

14 American Institute of Aeronautics and Astronautics