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NASA NTRS2016

The Laser Ranging Experiment of the Lunar Reconnaissance Orbiter: Five Years of Operations and Data Analysis

We describe the results of the Laser Ranging (LR) experiment carried out from June 2009 to September 2014 in order to make one-way time-of-flight measurements of laser pulses between Earth-based laser ranging stations and the Lunar Reconnaissance Orbiter (LRO) orbiting the Moon. Over 4,000 hours of successful LR data are obtained from 10 international ground stations. The 20-30 centimeter precision of the full-rate LR data is further improved to 5-10 centimeter after conversion into normal points. The main purpose of LR is to utilize the high accuracy normal point data to improve the quality of the LRO orbits, which are nomi- nally determined by the radiometric S-band tracking data. When independently used in the LRO precision orbit determination process with the high-resolution GRAIL (Gravity Recovery and Interior Laboratory) gravity model, LR data provide good orbit solutions, with an average difference of approximately 50 meters in total position, and approximately 20 centimeters in radial direction, compared to the definitive LRO trajectory. When used in combination with the S-band tracking data, LR data help to improve the orbit accuracy in the radial direction to approximately 15 centimeters. In order to obtain highly accurate LR range measurements for precise orbit determination results, it is critical to closely model the behavior of the clocks both at the ground stations and on the spacecraft. LR provides a unique data set to calibrate the spacecraft clock. The LRO spacecraft clock is characterized by the LR data to a timing knowledge of 0.015 milliseconds over the entire 5 years of LR operation. We here present both the engineering setup of the LR experiments and the detailed analysis results of the LR data.

Mao, Dandan↗

NASA NTRS2017

An Alternative Approach to Human Servicing of Crewed Earth Orbiting Spacecraft

As crewed spacecraft have grown larger and more complex, they have come to rely on spacewalks, or Extravehicular Activities (EVA), for assembly and to assure mission success. Typically, these spacecraft maintain all of the hardware and trained personnel needed to perform an EVA on-board at all times. Maintaining this capability requires up-mass, volume for storage of EVA hardware, crew time for ground and on-orbit training, and on-orbit maintenance of EVA hardware. This paper proposes an alternative methodology, utilizing either launch-on-need hardware and crew or regularly scheduled missions to provide EVA capability for space stations in low Earth orbit after assembly complete. Much the same way that one would call a repairman to fix something at their home these EVAs are dedicated to maintenance and upgrades of the orbiting station. For crew safety contingencies it is assumed the station would be designed such the crew could either solve those issues from inside the spacecraft or use the docked Earth to Orbit vehicles as a return lifeboat, in the same manner as the International Space Station (ISS) which does not rely on EVA for crew safety related contingencies. This approach would reduce ground training requirements for long duration crews, save Intravehicular Activity (IVA) crew time in the form of EVA hardware maintenance and on-orbit training, and lead to more efficient EVAs because they would be performed by specialists with detailed knowledge and training stemming from their direct involvement in the development of the EVA. The on-orbit crew would then be available to focus on the immediate response to any failures such as IVA systems reconfiguration or jumper installation as well as the day-to-day operations of the spacecraft and payloads. This paper will look at how current unplanned EVAs are conducted on ISS, including the time required for preparation, and offer an alternative for future spacecraft. As this methodology relies on the on-time and on-need launch of spacecraft, any space station that utilized this approach would need a robust transportation system, possibly including more than one launch vehicle capable of carrying crew. In addition, the fault tolerance of the future space station would be an important consideration in how much time was available for EVA preparation after the failure. Ideally the fault tolerance of the station would allow for the maintenance tasks to be grouped such that they could be handled by regularly scheduled maintenance visits and not contingency launches. Each future program would have to weigh the risk of on-time launch against the increase in available crew time for the main objective of the spacecraft. This is only one of several ideas that could be used to reduce or eliminate a station's reliance on rapid turnaround EVAs using on-board crew. Others could include having shirt-sleeve access to critical systems or utilizing low pressure temporarily pressurized equipment bays.

Mularski, John R.↗

NASA NTRS2019

NASA's Orbital Debris Optical Program: ES-MCAT Updated and Upgraded

The 1.3m ES-MCAT telescope (or MCAT for short) now has a proven capability for observing objects from Low- Earth Orbit (LEO) out to Geosynchronous (GEO) orbit, and the ability to run all systems autonomously. A GEO survey, the initial focus for MCAT, will commence in late 2019 to map out the current state of the GEO population as input for the ORbital Debris Engineering Model (ORDEM 4.x). This survey will statistically sample the GEO belt (0 to ~15 deg orbital inclinations) to detect both known and unknown targets. If a break-up occurs, additional surveys of the break-up field can be followed for discovery and investigations of daughter debris fragments from the parent satellite. Discovery can be accomplished by tracking orbits near to and including the parent object’s orbit. Targeted observations of debris can be taken with a suite of broadband filters for characterizing individual objects by ratetracking their known or calculated orbital elements (Two-Line Element sets, TLEs). Several modifications and upgrades have been made to the instrumentation and systems originally installed in 2015 and are reported here. In 2018, MCAT’s primary mirror was recoated with a high-end protected, enhanced silver by the ZeCoat Corporation. The CCD chip was replaced in the Spectral Instruments camera with a broad-band antireflective coated chip. The automated weather systems have been modified from the original system, removing some weather sensors and installing replacements that are better suited to Ascension’s weather and environment. A new 2.5-m ObservaDome replaced the Astrohaven dome on the nearby tower platform that will house an auxiliary 0.4-meter telescope. Finally, in 2019, the Observatory Control System was upgraded to 2.0 which includes additional flexibility for automating data collection and reduction. With these updates completed, MCAT is now well on track to reach Full Operational Capability (FOC) in 2019 for its survey, rate-track, and TLE tracking capabilities.

Lederer, S. M.↗

NASA NTRS2019

Orbital Debris Radar Measurements from the Haystack Ultra-Wideband Satellite Imaging Radar (HUSIR): 2014-2017

For many years, the NASA Orbital Debris Program Office (ODPO) has partnered with the U.S. Department of Defense and the Massachusetts Institute of Technology Lincoln Laboratory (MIT/LL) to collect data on the orbital debris environment using the Haystack radar. These measurements are used to characterize the small debris environment in low Earth orbit (LEO), down to a noise-limited size of approximately 5 mm—depending on altitude. The Haystack radar operated by MIT Lincoln Lab underwent upgrades starting in May 2010, with operations resuming in 2014 as the Haystack Ultra-wideband Satellite Imaging Radar (HUSIR). Hence, the data collected beginning in 2014 represents the first dataset available from this upgraded sensor. HUSIR is the primary source of data used by the ODPO to statistically sample orbital debris in the 5-mm to 10-cm size regime in LEO and is a key source of data to build and validate the NASA Orbital Debris Engineering Model. In this paper, we will present recent results from measurements performed during the US Government fiscal years 2014 – 2017. Using the NASA Size Estimation Model, a method based on laboratory radar measurements of debris, we will compare the size distributions of selected orbital debris populations over this 4-year period and flux measurements of orbital debris greater than 1 cm.

Murray, James↗

NASA NTRSDate not supplied

Near Rectilinear Halo Orbit Determination with Simulated DSN Observations

This paper presents the results of a high-fidelity simulation of spacecraft orbit determination in a near rectilinear halo orbit (NRHO). Others in the literature have examined this problem with linear covariance analysis, but the highly-nonlinear dynamics of this orbit challenge the assumptions underlying such analyses. The present work builds on similar analysis performed by other authors to contribute a fuller understanding of the operational requirements for NRHO navigation. The present work serves as a check to the assumptions of previous studies and an independent verification of those results. The results from the literature are extended by quantifying the space of orbital states from which a spacecraft with given control authority can safely return to the nominal path. Spacecraft state uncertainty estimates are evaluated as a function of time. Simulated range and range-rate measurements with the Deep Space Network (DSN) ground stations are used to model orbit determination accuracy. Orbit maintenance maneuvers are performed using both short-horizon and long-horizon stationkeeping targeting. Monte Carlo analysis of orbit determination and stationkeeping is performed. This paper quantifies the achievable state uncertainty with deep space network (DSN)-only range and range-rate observations. This paper also addresses requirements on the frequency of DSN observation periods and correlates ground contact frequency with navigation accuracy. The results of several related studies are presented and discussed: the effect of missing ground station passes, the effect of missing stationkeeping maneuvers, the sensitivity of the spacecraft state estimate to realistic error sources, and stationkeeping propellant budget.

Nathan L Parrish↗

NASA NTRSDate not supplied

Optical Characterization of DebriSat Fragments in Support of Orbital Debris Environmental Models

The NASA Orbital Debris Program Office (ODPO) develops, maintains, and updates orbital debris environmental models, such as the NASA Orbital Debris Engineering Model (ORDEM), to support satellite designers and operators by estimating the risk from orbital debris impacts on their vehicles in orbit. Updates to ORDEM utilize the most recent validated datasets from radar, optical, and in situ sources to provide estimates of the debris flux as a function of size, material density, impact speed, and direction along a mission orbit. On-going efforts within the NASA ODPO to update the next version of ORDEM include a new parameter that highly affects the damage risk – shape. Shape can be binned by material density and size to better understand the damage assessments on spacecraft. The in situ and laboratory research activities at the NASA ODPO are focused on cataloging and characterizing fragments from a laboratory hypervelocity-impact test using a high-fidelity, mock-up satellite, DebriSat, in controlled and instrumented laboratory conditions. DebriSat is representative of present-day, low Earth orbit satellites, having been constructed with modern spacecraft materials and techniques. The DebriSat fragment ensemble provides a variety of shapes, bulk densities, and dimensions. Fragments down to 2 mm in size are being characterized by their physical and derived properties. A subset of fragments is being analyzed further in NASA’s Optical Measurement Center (OMC) using broadband, bidirectional reflectance measurements to provide insight into the optical-based NASA Size Estimation Model. Additionally, pre-impact spectral measurements on a subset of DebriSat materials were acquired for baseline material characterization. This paper provides an overview of DebriSat, the status of the project, and ongoing fragment characterization efforts within the OMC.

Heather M Cowardin↗

NASA NTRS2021

Optical Characterization of DebriSat Fragments in Support of Orbital Debris Environmental Models

Heather M. Cowardin↗

NASA NTRS1995

Orbits Close to Asteroid 4769 Castalia

We use a radar-derived physical model of the kilometer sized, uniformly rotating asteroid 4769 Castalia (1989 PB) to investigate close orbit dynamics. We establish that a Jacobi integral exists for particles orbiting this asteroid, and the attendant zero-velocity surfaces are generated and examined. We find families of periodic orbits and determine their stability. We find that all synchronous orbits and direct orbits within 3 mean radii of Castalia are unstable and are subject to impact or escape from Castalia. We find that retrograde orbits are mostly stable and allow particles to orbit closely to the asteroid surface.

Asteroid↗

NASA NTRS1993

The Role of Anomalous Satellite-Fixed Accelerations in TOPEX/Poseidon Orbit Maintenance

Shortly after the launch of TOPEX/Poseidon on 10 August 1992 orbit determination indicated orbital decay levels approximately 60 times larger than could be explained by atmospheric drag. Outgassing, a complex process of molecular releases from satellite non-metallic parts, was the most likely source of these early decay rates. The high decay levels steadily declined during the first six weeks while a planned sequence of six orbit adjust maneuvers placed the satellite in the operational orbit to precisely overfly a predetermined repeat ground track. At the same time, on-going orbit trend analysis revealed the presence of residual along-track forces comparable to atmospheric drag which clearly exhibited a body-fixed origin. These anomalous forces cause either orbital decay or boost, depending on the satellite attitude and solar array articulation mode. As such, these along-track forces can either add to, or oppose, the resident orbital decay due to drag...

Bhat, R.S.↗

NASA NTRS2017

The Design of the Reference Orbit for NISAR, the NASA-ISRO Synthetic Aperture Radar Mission

The NISAR mission plans to use a 12-day-repeating sun-synchronous orbit for repeat-pass interferometry at multiple time scales using SAR data. For the interferometry to work the radar measurements must be made from within a critical baseline, which happens if all of the orbits are maintained to be within a fixed tube around a reference orbit. This paper describes the choice of dynamical models used in defining such a reference orbit, the perturbative effects of dynamics not considered in the repeat orbit, and the process of designing the orbit to repeat. We also describe our method for sharing the repeat orbit among multiple mission participants who use different models and software.

Hatch, Sara J.↗

NASA NTRS2017

The Design of the Reference Orbit for NISAR, the NASA-ISRO Synthetic Aperture Radar Mission

Hatch, Sara J.↗

NASA NTRS2022

Orbital Trade Study for the PREFIRE Mission

The Polar Radiant Energy in the Far-Infrared Experiment (PREFIRE) is a low-cost cubesat-based Earth Science mission selected for rapid implementation through NASAs Earth Ventures-Instrument program. Two cubesats carrying thermal infrared spectrometers (TIRS) in polar orbits will test long-held assumptions regarding radiative balance at long wavelengths where the cold-dry Arctic and Antarctic surfaces expel significant amounts of the earth's radiant energy. Thermal radiometry with better than one micron sampling at wavelengths beyond 15 microns will provide critical observational bounds for inputs into glacier melt and climate change models through improvements in surface emissivities, longwave greenhouse effect parameterizations, and cloud presence/types. Additional information on potential far-infrared surface/atmosphere feedbacks during rapid melt processes may be captured with sub-diurnal revisits of targeted areas. A baseline mission with two cubesats in different orbits provides a means for optimization of sub-diurnal sampling for science purposes. In this presentation we provide details of orbit studies that inform expected sampling metrics. Candidate orbits are modeled in SGP and then TIRS sampling is projected into surface footprints with appropriate geo-location. Matching algorithms then compile statistics for time-delayed revisits between and within the two orbits. Binning of results by latitude allows for abstraction of precession issues as well as for correlations with surface types as defined by radiometric classification schemes. Multiple acceptable pairs of science-driven orbits were found that optimize sampling in areas known to experience rapid melt events. The final choice of PREFIRE orbital parameters includes the systems trade-space, particularly the power budget, as well as the availability of launch opportunities.

"L'Ecuyer, Tristan"↗

NASA NTRSDate not supplied

On-Orbit Performance of the NOAA-21 Advanced Technology Microwave Sounder (ATMS)

The Advanced Technology Microwave Sounder (ATMS) instrument provides sounding measurements of Earth’s atmosphere to collect temperature and water vapor data for NOAA’s Joint Polar Satellite System (JPSS) program. ATMS is a total-power passive microwave radiometer with 22 channels spanning a frequency range from 22 to 183 GHz. A general description of the ATMS instrument is detailed in [1]. Three ATMS units are currently on-orbit. The Suomi National Polar-orbiting Partnership (SNPP) unit was launched in 2011 and the NOAA-20 (previously JPSS-1) unit was launch in 2017. Both SNPP and NOAA-20 ATMS units are currently operational and the data is used for numerical weather prediction (NWP) models. The NOAA-21 (formerly JPSS-2) ATMS unit was launched on November 10, 2022 from Vandenberg Space Force Base in California [2]. The ATMS instrument began generating radiance data on November 21, 2022. At the time of this abstract NOAA-21 ATMS is currently completing on-orbit commissioning and checkout activities; it is expected that the commissioning activities will be complete by the IGARSS 2023 conference. This paper will focus on the on-orbit performance of the NOAA-21 ATMS instrument based on measurements and activities from the commissioning period. Comparisons will be made to specifications and to expected performance based on pre-launch test activities. Pre-launch JPSS-2 (NOAA-21) ATMS performance is detailed in [4]. Additionally, the NOAA-21 performance will be compared to the on-orbit performance of the SNPP and NOAA-21 ATMS units. The post-launch performance of SNPP ATMS is detailed in [1][5][6][7]. The post-launch performance of NOAA-20 ATMS is detailed in [5][7][8]. On-orbit performance is evaluated through data collects from nominal mission operations as well as specific post-launch tests and spacecraft maneuvers. Nominal mission operations can be used to evaluate parameters such as instrument thermal stability, Noise Equivalent Delta Temperature (NEDT)/radiometric sensitivity, geolocation, inter-channel noise correlation, striping, and radiometric bias and stability. Shortly after instrument activation, the instrument thermal stability as measured from onboard temperature sensors is evaluated to demonstrate stable and steady-state behavior. Methods for computing NEDT are described in [9][10][11]. Passive geolocation measurements are performed using the coastline inflection point (CIP) method, described in [8][12]. The striping index, used as a metric to quantify striping, is the ratio of along-track variance to cross-track variance of the observed brightness temperature [4][8][14]. Inter-channel noise correlation has been previously reported for SNPP and NOAA-20 [1][8]. Dedicated post-launch tests and spacecraft maneuvers are utilized during the commissioning phase for further evaluation of on-orbit instrument performance. These activities provide information on the optimal space view selection, noise power spectral density (PSD), gain stability, scan bias, interference from onboard transmitters, reflector emissivity, active geolocation, and radiometric bias and stability detection methods. The optimal space view selection is used to determine which of ATMS space view sectors is preferred for minimizing contamination of the space view [1]. Point and stare testing is used to generate noise PSD and gain stability information, as described in [1][8]. Spacecraft roll and pitch maneuvers allow the ATMS to view different zones, including deep space, and the data can be used to provide information on scan biases and antenna sidelobe contamination [1][8]. The pitch maneuver can also be used to assess interference from onboard Ka-band transmitters and reflector emissivity [8]. Active geolocation is evaluated as a method for geolocation, described in [15].

Edward Kim↗

NASA NTRSDate not supplied

Flagship-Class Uranus Orbiter and Probe Using Aerocapture

Exploration of the Ice Giants, especially Uranus, via orbiter and atmospheric probes, is required to answer pressing science questions that have been raised in the latest Decadal Survey. As the Ice Giants are the farthest planets from Earth, traditional fully-propulsive orbit insertion missions have transit times to the planetary bodies nearing 13-15 years and require a large amount of propellant (wet mass percentages of around 60-70%) for the orbit insertion maneuver, leaving less mass for the scientific payload and a planetary probe. Aerocapture uses aerodynamic forces generated by flight within a planetary atmosphere to decelerate and achieve orbit insertion. Aerocapture has been considered for several past missions but it has not been demonstrated. However, recent developments in thermal protection systems (TPS), guidance and control (G&C), and interplanetary navigation capabilities show the potential for using rigid, heritage entry vehicle configurations already flown at other planetary bodies for Ice Giants aerocapture. Aerocapture can robustly deliver spacecraft to Ice Giant orbits, while substantially increasing on-orbit payload mass (more than 40%) that can be used for a robust atmospheric entry probe. Additionally, the aerocapture maneuver would reduce the interplanetary transit time by 2-5 years (15-30%) relative to fully-propulsive orbit insertion. Recent work has shown that a flagship-class mission can be conducted in a shorter time than fully-propulsive missions if using aerocapture.

S Dutta↗

NASA NTRSDate not supplied

HelioSwarm: Transfer Trajectory and P/2 Lunar Resonant Orbit Design for a Multi-Satellite Observatory

The HelioSwarm mission will fly a multi-satellite observatory in a P/2 lunar resonant orbit to study plasma turbulence in the solar wind. Lunar resonant orbits enable long-term access to high Earth altitudes. The relative dynamics of a multi-satellite observatory introduce novel considerations for lunar resonant orbit design. Additional maneuvers in the post lunar encounter transfer enable heightened control of the lunar resonant orbit initial state compared to previous missions. Adjusting the mission orbit’s initial conditions enables refinement of the orbit design based on requirements imposed by the relative dynamics of the observatory. refinement of the orbit design based on requirements imposed by the relative dynamics of the observatory.

cislunar↗

NASA NTRSDate not supplied

Deep Space navigation for the BioSentinel spacecraft science orbit

BioSentinel is an astrobiology small spacecraft mission. The payload consists of two parts, the first has optical and microfluidics sensors, and the second is a Linear Energy Transfer spectrometer that has the objective to measure deep space radiation from events such as coronal mass ejections. The goal of the mission is to observe potential DNA damage due to the radiation in heliocentric space on the living organism Saccharomyces cerevisiae, which is a budding yeast. Two types of this living organism are included in the payload. The first is a natural type that is more radiation tolerant, while the second is a mutant strain that has a deficiency in a gene that allows DNA repair once damage occurs. The impact caused by the radiation on the DNA is compared to an identical sample aboard the International Space Station, as well as another identical sample at a laboratory on the ground. The BioSentinel mission consists of a 6U CubeSat currently ,as of January 2024, active in heliocentric orbit. The spacecraft was launched aboard the first SLS flight as part of the Artemis-I campaign in November 2022. After successful deployment from the launch vehicle, it performed a lunar flyby with an altitude of 406 km. The delta-V imparted by the flyby provided the necessary energy to achieve a heliocentric orbit, in an Earth-trailing pattern. The navigation analysis consisted of a Kalman-filter that utilized data from the Deep Space Network and the ESA Estrack network. All those antennas were needed since the Artemis-1 campaign included the deployment of several other cubesats, therefore the scheduling process required more antenna assets than usual due to simultaneous demands from various missions. The processed tracking data was later also refined with a smoother in order to obtain a more accurate solution. The type of tracking data included TCP, Sequential Range, Doppler and Range formats. The solar radiation pressure coefficient, as well as the delta-V from the deployment and the flyby were modeled to obtain suitable solutions that could decrease the position and velocity uncertainties at several steps along the mission concept of operations. The final product each time resulted in updated ephemeris files that were used by the mission and the antenna networks as the mission progressed. Once in the final science orbit, the utilized antennas are only from the DSN network and the data format is bounded to just TCP. Regular orbit determination is performed, every two weeks. The spacecraft is in a nominal well-known orbit, performing regular operations. This paper includes an analysis of the final science orbit, the techniques and procedures utilized to perform orbit determination and a description of the overall navigation campaign produced during the mission and, more specifically, during the final science operations in Deep Space.

BioSentinel↗

NASA NTRSDate not supplied

Assessing Utility of Different Orbits for a Lunar PNT Constellation

"As the Artemis program establishes a long-term presence on the lunar surface, lunar surface users will need an accurate source of Position, Navigation, and Timing (PNT). Under the LunaNet framework, NASA has proposed the Lunar Navigation System (LNS), a constellation of satellites in lunar orbit which will broadcast LunaNet’s Augmented Forward Signal (AFS) [1] to the lunar surface. LNS will have the objective of providing user position accuracies of 10 m and timing accuracies of 10 microseconds on the lunar surface and up to 200 km above the surface. Several classes of orbits are possibilities for the LNS constellation, including circular orbits, eccentric orbits, and various families of 3-body orbits. Each of these orbit classes have both advantages and disadvantages in their utility as PNT constellations. In this study we examine a wide variety of potential LNS constellations and evaluate their performance against several metrics, including: a. Dilution of Precision (DOP) performance over the lunar service volume and the South Pole region, during both the constellation buildup and after the constellation is completed. b. Number of satellites required to achieve required DOP performance. c. Relative dynamics of LNS satellites with respect to surface users. d. Signal in space error (SISE) performance, assuming orbit determination that includes the Global Positioning System (GPS), crosslinks, and monitor sites on the surface of the moon. e. Constellation insertion, station keeping, and disposal costs. f. Specific Concept of Operations (CONOPS) and engineering challenges associated with establishing and operating the constellations. (E.g., frequent station keeping, variable signal strength, divergent CONOPS for different planes of the constellation.) The system’s performance is considered both independently and in combination with other proposed systems, such as NASA’s Lunar Communication Relay and Navigation System (LCRNS) and ESA’s Moonlight constellation. The study concludes that a 3-plane Walker constellation at 6 lunar radii (slightly higher than GPS’s 4.2 earth radii from a geometry perspective, but significantly closer in an absolute sense) is a good candidate for the LNS constellation, providing the user with good SISE and DOP performance, while not requiring excessive station keeping or complex CONOPS. However, other viable concepts exist and could be chosen if early South Pole coverage is a driving factor in the system design."

navigation↗

NASA NTRS2023

Slow Solar Wind Connection Science during Solar Orbiter's First Close Perihelion Passage

The Slow Solar Wind Connection Solar Orbiter Observing Plan (Slow Wind SOOP) was developed to utilize the extensive suite of remote-sensing and in situ instruments on board the ESA/NASA Solar Orbiter mission to answer significant outstanding questions regarding the origin and formation of the slow solar wind. The Slow Wind SOOP was designed to link remote-sensing and in situ measurements of slow wind originating at open–closed magnetic field boundaries. The SOOP ran just prior to Solar Orbiter's first close perihelion passage during two remote-sensing windows (RSW1 and RSW2) between 2022 March 3–6 and 2022 March 17–22, while Solar Orbiter was at respective heliocentric distances of 0.55–0.51 and 0.38–0.34 au from the Sun. Coordinated observation campaigns were also conducted by Hinode and IRIS. The magnetic connectivity tool was used, along with low-latency in situ data and full-disk remote-sensing observations, to guide the target pointing of Solar Orbiter. Solar Orbiter targeted an active region complex during RSW1, the boundary of a coronal hole, and the periphery of a decayed active region during RSW2. Postobservation analysis using the magnetic connectivity tool, along with in situ measurements from MAG and SWA/PAS, showed that slow solar wind originating from two out of three of the target regions arrived at the spacecraft with velocities between ∼210 and 600 km s −1 . The Slow Wind SOOP, despite presenting many challenges, was very successful, providing a blueprint for planning future observation campaigns that rely on the magnetic connectivity of Solar Orbiter.

Slow solar wind↗