Localization of Gas Leakage in Spacecraft Structures Using an Improved Beamforming Algorithm

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

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

: 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.

<p>The aurora can have strong electric fields and currents associated with it, which deposit a significant amount of energy in the neutral upper atmosphere through heating. Such heating must be included in global atmospheric models used to study thermospheric dynamics, coupling between the atmospheric layers, climate, and drag on spacecraft and space debris in low Earth orbit. However, the heating rate is poorly quantified, and often spatial structure is not well represented. Heating is typically estimated by measuring the ionospheric electric field using radar, which is then combined with measurements or estimates of the neutral wind velocity and Pedersen conductivity to calculate a Joule heating rate. However, such measurements of the electric field necessarily neglect small scale spatial and temporal variability through their relatively coarse resolution and averaging. The Joule heating rate is proportional to the square of the electric field, and therefore the spatial and temporal averaging can lead to a significant underestimate of the Joule heating rate. As a step towards improving estimates of neutral heating, we have developed a technique to invert spectrographic measurements of aurora to observe the thermospheric neutral temperature altitude profile at high temporal resolution. Application of the technique to an auroral event shows substantial Joule heating adjacent to an arc where the E-field must be strong, as well as heating embedded within an auroral curl, which we associate with an intense field-aligned current.</p>

The aurora is a significant source of heat in the high latitude upper atmosphere: tens of gigawatts of energy from the solar wind is deposited as heat in the thermosphere and ionopshere (Østgaard et al 2002). Strong electric fields and currents associated with the aurora cause heating through friction between ions and neutrals (Joule heating) and resistive heating by magnetic field-aligned currents (Lanchester et al 2002). This energy input must be included in whole-climate models and models used to predict additional drag on spacecraft and space debris in Low Earth Orbit during geomagnetic storms. We present results from a new technique to measure neutral temperatures in fine-scale aurora at high spatial and temporal resolution (tens of milliseconds) using simultaneous images of emissions in two different parts of the auroral molecular nitrogen spectrum from the University of Southampton’s Auroral Structure and Kinetics (ASK) multi-spectral imager. The technique measures the neutral temperature at the altitude of the auroral emissions so the observations require careful interpretation to separate local neutral temperature changes from spatial variation in auroral altitude across each image. Height profiles of the neutral temperature are therefore obtained using a third image in an atomic oxygen emission to determine the energy of the auroral electron precipitation and hence estimate the altitude of the aurora and temperature measurement. The resulting profiles show rapid neutral temperature changes on the order of several hundred Kelvin across E-region altitudes (between 100 km and 160 km). The coolest temperatures are found within the brightest regions of the aurora, whereas higher temperatures are typically associated with the edges of arcs where the electric field is expected to be strongest. The neutral temperature profiles are compared to ion temperature profiles from the European Incoherant Scatter (EISCAT) Svalbard radar (co-located with ASK) to better understand Joule heating and ion-neutral coupling in the extreme electrodynamic environment surrounding the aurora.

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.

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.

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 possibility of implementing a modeling of the flight of a small spacecraft in low earth orbit based on a client-server architecture was researched. Issues such as transmitting and displaying the system state, aggregating states at different points in time, and executing user code were discussed. The conducted studies demonstrate the benefits of using a client-server implementation in solving the problem of simulating the flight of a small spacecraft in near-earth orbit in a world of increasing interest in satellite developments.

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.

Long duration spacecraft in low earth orbit, such as the International Space Station Alpha (ISSA), are highly susceptible to hypervelocity impacts by pieces of debris from past earth-orbiting missions. With this increased likelihood of debris impact over time comes a responsibility on the part of a spacecraft design engineer to quantify, and subsequently reduce, the hazardous effects on the spacecraft and its crew should a penetration occur. Among the various hazards that accompany the penetration of a pressurized manned spacecraft module are critical crack propagation in the module wall (i.e. the so-called 'unzipping' of the module) and depressurization-related phenomena, such as crew hypoxia and uncontrolled thrust due to air rushing out of the module wall hole. These phenomena are directly related to the hole size and crack lengths that result in a spacecraft wall following a penetration. As a result, accurate models for hole size and crack size resulting in thin-walled spacecraft structures following hypervelocity impact penetration are required for accurate spacecraft quantitative risk assessments. This paper presents the results of a study whose objectives were to develop semi- empirical models of hole size and tip-to-tip crack length for some of the multi-wall shielding systems being developed for ISSA. The empirical models were developed using light gas gun test data at impact velocities around 6.5 km/sec, and inhibited shaped charge test data at an impact velocity of 11.3 km/sec. The significance of the work performed is that these models can be incorporated directly into a survivability analysis to determine whether or not module unzipping would occur under a specific set of impact conditions. In addition, the prediction of hole size can be used as part of a survivability analysis to determine the time available for module evacuation prior to the onset of incapacitation due to air loss.

The tracking and data relay satellite (TDRS) in geosynchronous orbit (GEO) can provide ranging and communication services for user spacecraft in low Earth orbit (LEO) through the inter‐satellite link (ISL). The time delay of TDRS ISL transponder in orbit differs from its pre‐launch state and changes due to the complex space environment, adding significant errors to the user spacecraft's orbit determination based on four‐way ranging data. The authors have developed a calibration system and proposed an efficient on‐orbit calibration method for the ISL transponder's time delay in TDRS. Experimental results demonstrate the effectiveness of the authors’ method, and the RMS of calibration residuals is less than 0.5 m.

This paper considers a spacecraft pursuit-evasion problem taking placeinlow earthorbit. The problem is formulated as a zero-sum differential game inwhich there are two players, a pursuing spacecraft that attempts tominimize a payoff, and an evading spacecraft that attempts to maximizethe same payoff. We introduce two associated optimal control problemsand show that a saddle point for the differential game exists if andonly if the two optimal control problems have the same optimalvalue. Then, on the basis of this result, we propose twocomputational methods fordetermining a saddle point solution: asemi-direct control parameterizationmethod (SDCP method), which is based on a piecewise-constant controlapproximation scheme, and a hybrid method,which combines the new SDCP method with the multiple shootingmethod. Simulation results show that the proposed SDCP and hybridmethods are superiorto the semi-direct collocation nonlinear programming method(SDCNLP method), which is widely used to solve pursuit-evasionproblems in the aerospace field.