Relative Navigation Strategy About Unknown and Uncooperative Targets

We are proposing a mission concept for Active Debris Removal (ADR) of multiple uncooperative debris targets from Low Earth Orbit, using a minisat spacecraft ("mothership") carrying several Coulomb Drag based Plasma Brake¹ modules. The Plasma Brake technology enables very high downmass/upmass ratios for debris deorbiting. In optimal conditions, the ratio for such a mission could be up to 60, representing a radical increase in ADR performance of 1-2 orders of magnitude compared to the current state-of-the-art ADR missions, and drastically improving the cost efficiency compared to one-object-per-mission approaches. The  Plasma Brake is especially competitive in the ~700-1200 km altitude range. The applicable debris mass limit of this deorbiting method is dictated by the demisability of the targets and is roughly 500 kg / Plasma brake unit. It provides a scalable, cost-effective method for reaching a net-zero space debris operations in the long term.The mission scenario consists of transfers between targets and a sequence of mothership operations repeated at each target: rendezvouz, detection, approach, de-tumbling of the target, attachment of a Plasma Brake module to the target. The critical core technologies needed to realize the mission, and are currently at low TRL, include:Sensors, actuators and algorithms required for the proximity operations used to approach and de-tumble the target object, Capture/attachment technologies for uncooperative targets like: kinematic structures, Hoberman sphere mechanisms, electrostatic adhesive, Adaptation of Plasma Brake concept to work with uncooperative targets, Framework for choosing the targets of a a multi-object touring mission that maximizes the downmass/upmass ratio, including the navigation scheme and navigation algorithms. Most ADR technologies so far are only viable for deorbiting large pieces of debris. The proposed system allows the effective deorbiting of much smaller pieces of debris as well, maintaining its effectiveness indefinitely to clean up the debris that is required by zero-debris targets.The mission concept and development of the critical core technologies needed for the mission to TRL 4 (testing prototypes in laboratory environment) have been proposed in EIC Pathfinder call of October 2024 by the authors.¹Janhunen, P., Electrostatic plasma brake for deorbiting a satellite, J. Prop. Power, 26, 370-372, 2010, https://arc.aiaa.org/doi/10.2514/1.47537

Precise relative GPS navigation is essential technology for rendezvous and formation flying of spacecraft in Low Earth Orbit. Key design issues of precise relative GPS navigation software are studied and a novel formulation is proposed for the mission which requires high accuracy when the separation distance is up to several km. The navigation filter estimates float navigation solutions by extended Kalman filter with elaborate dynamics models, and resolves ambiguities of integer carrier phase biases to achieve high accuracy fix navigation solutions. A relative navigation software with the proposed formulation is implemented and evaluated in two different ways. One is a test using spacebourne GPS receivers and a GPS signal simulator to evaluate the performance and sensitivity of the software against variation of parameters. The other is a test using actual telemetry data from Gravity Recovery and Climate Experiment (GRACE) to demonstrate the software performance. The design, implementation, and results of the evaluation are presented and discussed on this paper.

Micrometeoroid/Orbital Debris (MMOD) Impact Detection and Location Using Fiber Optic Bragg Grating Sensing Technology

A well-known hazard associated with exposure to the space environment is the risk of failure due to an impact from a micrometeoroid and orbital debris (MMOD) particle. As NASA prepares to return astronauts to the moon with the Artemis program, the next generation of spacesuit known as the exploration extravehicular mobility unit (xEMU) is in development to support future extravehicular activities (EVAs). An MMOD impact to the spacesuit is of great concern to NASA as a large leak could prevent an astronaut from safely reaching an airlock in time resulting in a loss of life. In contrast to the legacy EMU, the xEMU must meet MMOD requirements for multiple environments including those in low earth orbit (LEO) as well as the meteoroid and secondary lunar regolith ejecta environments found on the lunar surface. The subject of this paper is an MMOD risk assessment of the internal xEMU configuration design developed by NASA Johnson Space Center (JSC) personnel. Assessment of the MMOD risk for an 8-hour, 2-person EVA in both LEO and on the lunar surface showed that the xEMU spacesuit meets the program technical requirement of 1 in 2500 failure odds.

Orbital debris has become an increasingly severe problem, and active debris removal (ADR) missions are gaining progressively more attention. The first step in an ADR mission would be approaching the debris, which is usually a non-cooperative target. In a non-cooperative rendezvous, obtaining continuous sensor measurements for navigation is quite difficult due to the loss of sensor lock or the effect of solar reflection on the optical camera. It is essential to ensure passive abort (PA) safety and to guarantee robustness against collisions in the case of off-nominal thruster burns caused by spacecraft anomalies such as navigation sensor or actuator failures. In the proximity phase, sensors and navigation algorithms need to be switched according to the distance between the chaser and the target, considering the field of view and range of the rendezvous sensors. At the sensor-switching point, the navigation values before and after the switchover are compared, and if the health of the sensor is confirmed, the system proceeds to the next approach sequence. This paper proposes a design method of robust relative trajectories in low Earth orbits that considers PA safety and the ease of troubleshooting and transition to the backup relative orbits in the spiral trajectory.

This paper is devoted to understanding relative navigation models that are used for space vehicles. The relative navigations models and approaches which are based on different systems (Inertial Navigation Systems (INS)&Global Navigation Satellite System (GNSS) , Laser&INS, Vision- Based, etc.) are compared. These models and approaches can be used individually or combined for solving relative navigation problems. Advantages and disadvantages of the models vary according to the usage area, platform type, and environment. Different methods and approaches exist in addition to different estimation and optimization algorithms for adaptation, control, and sensor fusion. Most of the models assume perfect attitude conditions. This study considers satellites' position estimates according to each other within formation on the Low Earth Orbit (LEO). Also, the aim of this article is to understand correlation between the relative navigation systems and the effectiveness of the algorithms which are used for estimating states during constellation or formation flight.

Trajectory optimization for multi-target Active Debris Removal missions

Ground-based laser momentum transfer concept for debris collision avoidance

Summary form only given. Low Earth Orbit (LEO) is becoming increasingly populated by orbital debris of all sizes and shapes that is hazardous to operational spacecraft. NASA has recently determined that a collision with debris between 0.5 mm and a few centimeters would be "mission ending". Recently, a novel technique to remediate small debris (less than 10 cm) from Low Earth Orbit (600-1100 km) has been proposed that uses injected high mass density dust particles of 25-50 micron radius to collide with orbital debris, transfer momentum and thus lower the apogee of the debris to an altitude that will naturally decay quickly. The dust is injected at orbital velocities in a polar orbit at the top of an altitude range targeted to be cleared. The dust then forms a ring around the Earth that can continue to exchange momentum with orbital debris in its path. One of the keys to the efficiency of this method is the physics of a hypervelocity impact, in which a very high-pressure shock wave is generated that can fracture, melt, vaporize and ionize the material. This high-pressure material that is not in a condensed state is ejected from the debris at large velocity, in effect acting like retro-rockets and greatly enhancing the exchange of momentum. The efficiency of the momentum transfer is measured by a parameter κ which is 1 for inelastic collisions and 2 for elastic collisions. Assuming a relative impact velocity of 10-15 km/s (twice the orbital speed) and assuming that the melting of aluminum (a standard orbital debris material) is sufficient to eject material from the target orbital debris we can estimate a value of κ to be about 18. The amount of mass necessary to perform the remediation is inversely proportional to κ and hence a solid understanding of the physics of this parameter is critical. We use CTH simulations and analytical theory to calculate the parameter κ, the velocity of the ejecta, and estimate the amount of ionization to occur. We will comment on the creation of electromagnetic radiation created by the impact generated plasma and the potential applications of the radiation.

The number of launches and objects in space has been growing fast in the last few years, particularly due to the growth of satellite mega-constellations. Defunct satellites and other space junk products collide and create a collisional cascade of smaller space debris. Space debris ablates and burns up in the atmosphere upon re-entry and thereby metals and rare materials are injected, some of which already exceed the natural input of exogenous material today.Quantifying the influx of these anthropogenic materials into the atmosphere is essential to address the possible environmental consequences, through constraining the physico-chemical atmospheric models. This quantification can be done using catalogs of spacecraft being launched, but not all manufacturers provide these data. Small micro-debris can be used as tracers of their larger counterparts through the collisional cascade, which would complement these existing catalogs, for the inventory of elemental compositions of human-made materials in Low EarthOrbit that will re-enter in the atmosphere.We propose in situ measurements of sub-micrometer and micrometer sized particles as tracers of the larger space debris, using in situ mass spectrometers with a velocity grid, that were originally designed for cosmic dust measurements.These instruments can measure the elemental composition (impact-speed dependent), mass distribution, surface charge, impact velocity vector, and time-resolved fluxes of dust and debris particles. Moreover, measuring the natural cosmic dust flux itself is necessary as a benchmark for the debris.In this talk we introduce in situ cosmic dust measurements in the past, the different measurement methods, and measurements of micrometer-sized space debris so far with “active” (time-resolved) and “passive” (sample return) methods. We elaborate on the particles we can expect to measure in orbit, and the science goals to be achieved through such measurements that are useful for both the assessment of the anthropogenic influx into the atmosphere and for space debris research in Low Earth Orbit. Elemental composition measurements of these micro-debris particles, combined with orbital velocity and location data, offer a new avenue for quantifying the chemical influx of anthropogenic material into Earth’s atmosphere, and for assessing more thoroughly the broader space debris populations.

A review of space robotics technologies for on-orbit servicing

To date, the International Space Station (ISS) has been one of the largest objects flown in lower earth orbit (LEO). The ISS utilizes high voltage solar arrays (160V) that are negatively grounded leading to pressurized elements that can float negatively with respect to the plasma. Because laboratory measurements indicate a dielectric breakdown potential difference of 80V, arcing could occur on the ISS structure. The paper identifies anodized aluminum surfaces, chromic acid and sulfuric acid anodized, on the ISS, which could arc due to plasma charging. An assessment of the effects of plasma-induced arcing on these ISS anodized surfaces, with and without concurrent micrometeoroid and orbital debris (MM/OD) impacts has been completed. The purpose of this study was to assess the loss of thermal control capability of the anodized surfaces on ISS if the Plasma Contactor Units (PCUs) were not continually operated as ISS is constructed and operated. The study considers six cases for the arcing Chromic Acid Anodized Aluminum (CAA) surfaces or Sulfuric Acid Anodized Aluminum (SAA) with no Micrometeoroid/Orbital Debris impacts (MM/OD), with MM/OD impacts that completely penetrated the anodize layer, and with MM/OD impacts that partially penetrated the anodize layer. The Marshall Space Flight Center (MSFC) performed analyses to provide the data giving the size and depth of penetration of MM/OD impacts. All six cases documented the thermal control properties of the anodized layers throughout the ISS life without PCU operations.

An analytical theory for avoidance collision between space debris and operating satellites in LEO

The Federal Communications Commission (FCC) defines regulations that ensure satellite communications systems seeking access to the U.S. market will properly implement post-mission disposal (PMD) plans to reduce the risk of creating orbital debris. In March 2020, the FCC initiated a processing round for additional applications and petitions for non-geostationary orbit (NGSO) fixed-satellite service (FSS) operations in the 10.7 - 12.7 GHz, 12.75 - 13.25 GHz, 13.8 - 14.5 GHz, 17.7 - 18.6 GHz, 18.8 - 20.2 GHz, and 27.5 - 30 GHz bands. Ten NGSO FSS communications systems operators (EOS Defense Systems, Kepler Communications, Kuiper, Mangata, New Spectrum Satellite, O3b, OneWeb, SpaceX, Telesat, and Viasat) submitted applications. In Part 25 of the Commission's rules, particularly 47 C.F.R. ? 25.114(d)(14), the FCC defines key requirements to limit debris, including guidelines for post-mission disposal and the limitation of accidental explosions. Prior to approving an application for access to the U.S. market, the FCC assesses the compliance of each applicant's system with the Commission's Part 25.114(d)(14) rules. This paper analyzes and validates the PMD and orbital debris mitigation strategies of the ten NGSO FSS communication systems that filed applications in the FCC's March 2020 Processing Round in order to guarantee that adequate plans are in place to prevent the creation of orbital debris, and thus prevent harm to other spacecraft and humans. The methodology includes obtaining relevant information regarding the orbits and operational strategies of the systems from their submitted Technical Narratives and Schedule S reports, and performing PMD analysis validation using NASA's Debris Assessment Software (DAS) to confirm pertinent information such as deorbit duration and required delta-v (Δv) for deorbiting maneuvers. Nine of the ten systems provided statements on their orbital debris mitigation strategies, including identifying the agency or government's orbital debris guidelines with which their system will comply (NASA, UK, Canada, etc.), Several systems under the jurisdiction of foreign administrations, including One Web and 03b, did not provide orbital debris mitigation information as they stated they will follow their administration's orbital debris guidelines rather than the FCC's Part 25 rules pertaining to technical orbital debris showings. Three of the systems (EOS Defense Systems, O3b, Mangata) either stated that they will use a graveyard orbit or expect based on orbital altitude that they will not re-enter, whereas seven systems (Kepler, Kuiper, OneWeb, New Spectrum Satellites, SpaceX, Telesat and Viasat) either stated that they will perform deorbit and re-entry maneuvers or it is expected the spacecraft will naturally deorbit based on the low earth orbit (LEO) configuration. The analysis and validation of the PMD for each system is discussed in detail in this work, with the ultimate goal of aiding the FCC in evaluating the legitimacy of each applicant's approach and maintaining a sustainable space environment for the future.

Space Debris Mitigation