Robust and precise baseline determination of distributed spacecraft in LEO

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.

The European Space Agency's satellite ARTEMIS will support data links between spacecraft in low earth orbit, ARTEMIS in geostationary orbit and ground stations in Europe. The links between spacecraft in low earth orbit and ARTEMIS will operate at S-band (2 GHz), Ka-band (25 GHz) or optical wavelength (800 nm). This paper addresses the S-band and Ka-band links, which both operate via a steerable 2.85 m diameter reflector antenna on ARTEMIS. The link design has been performed to minimize the signal degradations, and has led to rather demanding requirements on the flatness of the amplitude and group delay distortion versus frequency, on the phase noise and on the linearity of the payload. The measured values for these parameters are presented. In S-band, the antenna pointing is performed in open loop according to a preloaded table, while in Ka-band a closed loop pointing system is implemented and open loop pointing is back-up. The design, test and operation of the antenna pointing system are presented. The European Space Agency (ESA) participates in the Satellite Network Interoperability Panel (SNTP) together with NASA and NASD A. SNTP has established a set of interoperability parameters making it possible for low earth spacecraft from the three agencies to operate through the relay satellites of NASA, NASDA and ESA. ARTEMIS will be the first ESA satellite to offer data relay services, and its performances are compared to the interoperability parameters established by SNIP. Inter Orbit Links Spacecraft in low earth orbit (LEO) are restricted in their ability to communicate with the ground, because of the limited time a spacecraft in LEO is visible from a single ground station. ARTEMIS, located in the geostationary orbit, will significantly increase the communication time between the spacecraft in LEO and a single ground station. This is illustrated in Fig. 1. When a spacecraft is flying over the cross hatched areas in Fig. 1 at 300km altitude, no contact is possible via ARTEMIS. This area is slightly increased for a spacecraft at 1000km altitude as marked by the contour. Spacecraft can maintain contact to ground via ARTEMIS, when they are over SouthAmerica, eastern part of North-America, the Atlantic, Europe, Africa, most of Asia, the Indian Ocean and the polar regions The data relay function is provided by communication links between the ground and ARTEMIS named feeder links and links between ARTEMIS and the LEO spacecraft named inter orbit links (IOL). The feeder link operates at 20 GHz from ARTEMIS to ground (i.e. return feeder link) and at 30 GHz from ground to ARTEMIS (i.e. forward feeder link). The IOL can operate in three different bands namely S-Band (i.e 2.025 to 2.110 GHz for the forward IOL from ARTEMIS to LEO spacecraft and Copyright © 1997 by J. Sandberg. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission.

Predictions of induced radioactivity for spacecraft in low Earth orbit

In the high-precision application of Global Navigation Satellite System (GNSS), integer ambiguity resolution is the key step to realize precise positioning and attitude determination. As the necessary part of quality control, integer aperture (IA) ambiguity resolution provides the theoretical and practical foundation for ambiguity validation. It is mainly realized by acceptance testing. Due to the constraint of correlation between ambiguities, it is impossible to realize the controlling of failure rate according to analytical formula. Hence, the fixed failure rate approach is implemented by Monte Carlo sampling. However, due to the characteristics of Monte Carlo sampling and look-up table, we have to face the problem of a large amount of time consumption if sufficient GNSS scenarios are included in the creation of look-up table. This restricts the fixed failure rate approach to be a post process approach if a look-up table is not available. Furthermore, if not enough GNSS scenarios are considered, the table may only be valid for a specific scenario or application. Besides this, the method of creating look-up table or look-up function still needs to be designed for each specific acceptance test. To overcome these problems in determination of critical values, this contribution will propose an instantaneous and CONtrollable (iCON) IA ambiguity resolution approach for the first time. The iCON approach has the following advantages: (a) critical value of acceptance test is independently determined based on the required failure rate and GNSS model without resorting to external information such as look-up table; (b) it can be realized instantaneously for most of IA estimators which have analytical probability formulas. The stronger GNSS model, the less time consumption; (c) it provides a new viewpoint to improve the research about IA estimation. To verify these conclusions, multi-frequency and multi-GNSS simulation experiments are implemented. Those results show that IA estimators based on iCON approach can realize controllable ambiguity resolution. Besides this, compared with ratio test IA based on look-up table, difference test IA and IA least square based on the iCON approach most of times have higher success rates and better controllability to failure rates.

: This paper presents the result of the SOCRATES (Shuttle Orbiter Contamination Representation Accounting for Transiently Emitted Species) code which has been developed to assess the effects of contamination of measurements aboard spacecraft in low earth orbit. SOCRATES is a Monte Carlo code which includes in its present version scattering, collisions leading to kinetic-to- vibrational energy transfer, and reactive collisions. The application of this code to actual measurements aboard spacecraft in low earth orbit makes it possible to evaluate data obtained on these platforms with a view towards extracting the data of interest from contaminated signals.

Effects of the South Atlantic Anomaly (SAA) on spacecraft in low Earth orbit (LEO) are well known and documented. The SAA exposes spacecraft in LEO to high dose of ionizing radiation as well as higher than normal rates of Single Event Upsets (SEU) and Single Event Latch-ups (SEL). CubeSats, spacecraft built around 10 x 10 x 10 cm cubes, are even more susceptible to SEUs and SELs due to the use of commercial off-the-shelf components for electronics and payload instrumentation. Examination of the SAA using both data from the Defense Meteorological Satellite Program (DMSP) and a new set of models for the flux of particles is presented. The models, AE9, AP9, and SPM for energetic electrons, energetic protons and space plasma, were developed for use in space system design. These models introduce databased statistical constraints on the uncertainties from measurements and climatological variability. Discussion of the models’ capabilities and limitations with regard to LEO CubeSat design is presented.

This paper presents a general approach to solve multi-client space logistics path optimization problems in the case of low-thrust propulsion. The methodology accounts for key time-dependent factors, including secular J 2 and drag perturbations, eclipse propulsion constraints, and servicer fuel mass depletion. A low-thrust transfer strategy, consisting of three phases (thrust-coast-thrust), is employed, utilizing a drift orbit to correct the difference in Right Ascension of the Ascending Node. The drift orbit parameters are optimized by using a nonlinear programming algorithm to minimize the fuel cost associated with the transfer while satisfying a maximum time of flight constraint. An optimization is conducted for all possible transfers between the satellites of a dataset on a two-dimensional discrete grid of initial servicer mass and departure time. This procedure creates two four-variable arrays representing the cost of each possible visitation in terms of fuel consumption and time of flight. The arrays are then interpolated to efficiently solve a general path optimization problem by using a genetic algorithm. Two test cases are analyzed for datasets consisting of 12 and 20 satellites in low Earth orbit, respectively. The first addresses an optimal open tour problem where the servicer must visit all satellites once while minimizing total fuel consumption. The second addresses an on-orbit refueling problem where the servicer must refuel as many satellites as possible. The mission scenario is generalized to consider a priority index associated with each client; servicing time and fuel mass constraints are also taken into account.

Distributed spacecraft systems (DSS) involving SmallSats in low Earth orbit are gaining significant interest both for Earth observation and on-orbit servicing purposes. However, the miniaturized low-cost components of SmallSats are susceptible to faults, making DSS prone to failures that are detrimental to its overall system performance. In this work, we address the problem of providing resiliency to potential failures for a fleet of spacecraft that are performing on-orbit inspection. The proposed methodology guarantees graceful degradation of the inspection performance even against the worst-case failures, through selection and assignment of formation orbits that are resilient to it. We define quantitative metric to measure collaborative inspection performance, taking into consideration both the information gain and control cost, and formulate worst-case failure of ℓ-spacecraft that maximally undermine it. The main algorithm searches through the space of formation orbits, sampling through a set of orbits with evaluation of its inspection performance in the presence of worst-case failures. The algorithms are designed to be computationally efficient and have linear scaling to the number of spacecraft and orbits, making it applicable to real-time planning for spacecraft swarms. The effectiveness of the proposed approach is validated through simulation experiments on a design reference mission involving five CubeSats inspecting a target spacecraft in low Earth orbit.

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

Atomic oxygen and molecular nitrogen are the major constituents in the residual atmosphere at low Earth orbital altitudes, and they collide with spacecraft surfaces at relative velocities of ∼7.8 km s−1. The energy associated with these hyperthermal collisions is in excess of many bond dissociation energies and may help promote materials degradation by allowing barriers to reaction or to collision‐induced dissociation (CID) to be overcome. Spacecraft in low Earth orbit (LEO) are also exposed to high fluxes of vacuum ultraviolet (VUV) radiation, which may degrade materials through various photochemical mechanisms. Fluorinated ethylene‐propylene copolymer (FEP Teflon®) is commonly used on spacecraft. Many researchers have studied the individual and/or combined effects of atomic oxygen, VUV light, and CID on FEP Teflon®, yet the detailed degradation mechanisms of FEP Teflon® in LEO are still a subject of debate. Although not as ubiquitous as a spacecraft material, polymethylmethacrylate (PMMA) has been studied as a model and control polymer because of its well known propensity to “unzip” upon exposure to VUV radiation. A combination of beam‐surface scattering, quartz crystal microbalance (QCM), and surface‐recession experiments were conducted to study the effects of various combinations of O atoms (in the ground O(3P) state), Ar atoms, and VUV light on FEP Teflon® and PMMA. A laser‐breakdown source was used to create hyperthermal beams containing O and O2 or argon. A deuterium lamp provided a source of VUV light. O atoms with 4 eV of translational energy or less did not react with a pristine FEP Teflon® surface. Volatile O‐containing reaction products were observed when the O‐atom energy was higher than 4.5 eV, and the signal increased with O‐atom energy. Significant FEP Teflon erosion (∼20% of Kapton® H) was observed when it was exposed to the hyperthermal O/O2 beam with an average O‐atom energy of 5.4 eV. FEP Teflon® and PMMA that were exposed to VUV light alone yielded volatile products and mass loss. Similarly, CID by Ar also yielded volatile products and mass loss, when the Ar energy was higher than 8 eV. However, the erosion caused by VUV light and/or CID is not significant compared to that caused by O/O2. There were no observed synergistic effects of VUV light and O/O2 exposure.

The aim of this article is to provide a review of the architecture and business for the first commercial propellant depots to be deployed in space within a decade. The capability to refuel spacecraft in low Earth orbit (LEO) underpins a paradigm shift that considerably decreases cost and increases the mass of spacecraft hardware possible per launch because of the reduction of onboard propellant requirements. This same refueling capability also enables repeated long-duration high-thrust missions for commerce, exploration, and security to be carried out at superior price-performance, resulting from extensive reuse of in-space vehicles and systems. Shackleton Energy Company is establishing initial propellant depots in LEO using propellants launched from Earth to commence sales and deliveries within 5 years from program start, followed by deliveries of water-derived propellants from the lunar poles within an additional 5 years. By sourcing the propellant from the Moon's lower energy gravity well, significant reductions in operating costs are possible, with additional infrastructure costs amortized over multiple sales cycles. The most readily accessible and operationally robust source of cryogenic liquid oxygen and liquid hydrogen is from the craters situated at the poles of the Moon, in the original form of water ice.

The European Space Agency’s (ESA) Space Debris Office provides a service to support operational collision avoidance activities. This support currently covers ESA’s missions Cryosat-2, Sentinel-1A and -2A, the constellation of Swarm-A/B/C in low-Earth orbit (LEO), as well as missions of third-party customers. In this work, we describe the current collision avoidance process for ESA and third-party missions in LEO. We give an overview on the upgrades developed and implemented since the advent of conjunction summary messages (CSM)/conjunction data messages (CDM), addressing conjunction event detection, collision risk assessment, orbit determination, orbit and covariance propagation, process control, and data handling. We pay special attention to the effect of warning thresholds on the risk reduction and manoeuvre rates, as they are established through risk mitigation and analysis tools, such as ESA’s Debris Risk Assessment and Mitigation Analysis (DRAMA) software suite. To handle the large number of CDMs and the associated risk analyses, a database-centric approach has been developed. All CDMs and risk analysis results are stored in a database. In this way, a temporary local “mini-catalogue” of objects close to our target spacecraft is obtained, which can be used, e.g., for manoeuvre screening and to update the risk analysis whenever a new ephemeris becomes available from the flight dynamics team. The database is also used as the backbone for a Web-based tool, which consists of the visualization component and a collaboration tool that facilitates the status monitoring and task allocation within the support team as well as communication with the control team. The visualization component further supports the information sharing by displaying target and chaser motion over time along with the involved uncertainties. The Web-based solution optimally meets the needs for a concise and easy-to-use way to obtain a situation picture in a very short time, and the support for third-party missions not operated from the European Space Operations Centre (ESOC). Finally, we provide statistics on the identified conjunction events, taking into account the known significant changes in the LEO orbital environment and share ESA’s experience along with recent examples.

Carrier-phase integer ambiguity resolution (IAR) is the key to highly precise, fast positioning and attitude determination with Global Navigation Satellite System (GNSS). It can be seen as the process of estimating the unknown cycle ambiguities of the carrier-phase observations as integers. Once the ambiguities are fixed, carrier phase data will act as the very precise range data. Integer aperture (IA) ambiguity resolution is the combination of acceptance testing and integer ambiguity resolution, which can realize better quality control of IAR. Difference test (DT) is one of the most popular acceptance tests. This contribution will give a detailed analysis about the following properties of IA ambiguity resolution based on DT: All these conclusions are verified based on the single and combination GNSS simulation experiments. The experiment results indicate the correctness of these conclusions. These properties demonstrate the special properties of DT-IA estimator, and also provide the research frame to investigate other IA estimators. They are helpful to realize better use of IA estimators in quality control and precise positioning in future.

Several existing and emerging applications of Space Situational Awareness (SSA) relate directly to spacecraft Rendezvous, Proximity Operations, and Docking (RPOD) and Formation / Cluster Flight (FCF). Observation correlation of nearby objects, control authority estimation, sensor-track re-acquisition, formation re-conguration feasibility, ‘stuck’ thrusters, and worst-case passive safety analysis are some areas where analytical reachability methods have potential utility. Existing reachability theory is applied to RPOD and FCF regimes. An optimal control policy is developed to maximize the reachability set and optimal control law discontinuities (switching) are examined. Necessary conditions for maximum position reachability are developed, allowing for a reduction in reachable set computation dimensionality. The nonlinear relative equations of Keplerian motion are introduced and used for all reachable position set determinations. The linearized ClohessyWiltshire equations of motion are normalized to accentuate relative control authority for spacecraft propulsion systems at both Low Earth Orbit (LEO) and Geostationary Earth Orbit (GEO). Several examples with traditional and low thrust propulsion systems in LEO and GEO are explored to illustrate the eects of relative control authority on the timevarying reachability set surface using the nonlinear equations of motion. Both monopropellant spacecraft at LEO and Hall thruster spacecraft at GEO are shown to be strongly actuated while Hall thruster spacecraft at LEO are found to be weakly actuated. Weaknesses with the current implementation are discussed and future numerical improvements and analytical eorts are discussed.

The PROBA-2 microsatellite of the European Space Agency (ESA) is equipped with a novel single frequency global positioning system (GPS) receiver, which combines low resource requirements with a high navigation performance. Aside from supporting the spacecraft and mission operations, the Phoenix GPS receiver on PROBA-2 is used to study realtime and offline navigation using single-frequency GPS measurements. A 1 m accuracy is targeted for both applications which has so far been the domain of dual-frequency receivers. Following a mission overview, the achievable performance for ground-based orbit determination and real-time onboard navigation is discussed for different processing concepts based on flight data collected during the first half year of mission operations. It is shown that a 1 m or better accuracy can be achieved in either case, despite sub-optimal GPS tracking conditions induced by the mission specific attitude profile. Besides self-consistency checks and comparisons with independent GPS based orbit determination solutions, the accuracy is evaluated using satellite laser ranging (SLR) measurements as an external reference. The results show that single-frequency GPS tracking is sufficient to meet the navigation requirements of even advanced remote sensing missions in low Earth orbit, if their accuracy potential is properly exploited.