Small Satellite Missions for Earth Observation

This chapter gives a programmatic and technical overview of the MYRIADE micro satellites line of product MYRIADE developed by CNES. The characteristics and mission topics of satellites under CNES responsibility are presented, for in flight, and under development systems. The main drivers of the roadmap for next years are addressed as a Conclusion.

This chapter presents Examples of the benefits of flying small satellites with other satellites, large and small. Scientists worldwide are beginning to take advantage of the opportunities for coincidental science offered by the international Earth observing constellation known as the A-Train. The A-Train comprises two large satellites (EOS-Aqua and EOS-Aura) flying with 4 small satellites (CloudSat, CALIPSO, PARASOL, and Glory). By flying in a constellation, each mission benefits from coincidental observations from instruments on the other satellites. The A-Train can be envisioned as a single, virtual science platform with mulTiple instruments. This chapter describes the challenges of operating an international constellation of independent satellites from the US and Europe to maximize the coincidental science opportunities while at the same time minimizing the operational interactions required between the teams. Constellation operations must be safe, but the team members must be able to conduct their respective science missions independently. The A-Train mission teams have also demonstrated that flying as members of a constellation does not take away the flexibility to accommodate new requirements.

This chapter is a description of the INTAμSat-1 first Earth observation mission objectives and its development status. With a mass around 100 kg and a 60 × 60 × 90 cm size compatible as an auxiliary payload for VEGA, Soyuz-ST and Dnepr, the project is running now the Phase-C and will be ready for launch by the middle of 2011. This new INTA small satellite programme initiated in October 2005, belongs to the well known μSAT class (up to 150 kg) as a further step after the Nanosat programme success: first Nanosat-1 with a 19 kg mass and still working OK in orbit that was launched by Ariane-5 in December 2004; and Nanosat-1B with 23.5 kg set for launch in a DNEPR by July 2009 from Baikonur. In fact both programmes will be running in parallel at INTA with next Nanosat-2 tentatively set for launch by 2013. It will be an evolved 20–50 kg s/c with improved service module resources and a separated payload modular design. After the preliminary progress on the different platform subsystems design, it was decided by the middle of Phase-B in October 2007 that the first μSAT demonstration flight will be devoted to an experimental Earth observation payload with several instruments, together with some technological experiments. Both the Service Module (SVM) and Payload Module (PLM) are fully developed at INTA, with contributions from several Universities and few small Spanish companies. INTAμSat-1 PLM will carry the following instruments: (1) CINCLUS, a 30 m GSD pushbroom TDI camera with a wide swath, that will try to collect information on the water quality over more that 150 reservoirs spread out over the Spanish territory, with a repetition cycle of 9 days; (2) MS-WAC, a 10 m GSD wide swath MS camera with 4–5 channels that will provide full coverage of Spain also in 9 days; PAU, a combined reflectometer–radiometer that using GPS reflected signals will try to measure the sea surface height variations and salinity. All this PL data will be transmitted down to Maspalomas ground receiving station in X-band at 40–80 Mbps, while the Control Centre (CC) will use the same Nanosat CC facilities for S-band TT&C available at INTA headquarters in Torrejón de Ardoz, Madrid.

VENμS (Vegetation and Environment on a New Micro Satellite) is an Earth observation demonstration mission developed in cooperation by FRANCE and ISRAEL in the framework of the European Global Monitoring for Environment and Security Program (GMES).

This program gathers both a scientific mission (earth images at high resolution and high repetitivity) and a technological mission (low orbit control by IHET – Israeli Hall Effect Thruster). It is due to fly at mid of 2011, for an overall duration of 3.5 years.

This chapter presents all the aspects of the VENμS mission: missions’ description, missions’ implementation, missions’ concurrence, ground image processing, ground control segment, satellite and payloads descriptions. It will also detail the cooperation aspects.

This chapter aims at proposing a small mission devoted to observe the mid-low latitude regions. The selected orbit, called multi-sun-synchronous (MSS), represents an innovation with respect to the classical sun-synchronous orbit. The advantage offered by this kind of orbit, more suitable for observing tropical regions, will be evaluated and described. The satellite will be equipped with three optical sensors: a medium-high spatial resolution VIS-NIR multi-spectral sensor, allowing the surface monitoring and land-use and land-cover studies; a medium spatial-resolution 3-band thermal (MIR-TIR) sensor allowing the surface temperature (LST, SST) estimate and hot-spot (fires, volcanic eruption, etc.) detection; a panchromatic VIS-NIR camera for night-time observation able to reveal artificial and natural lights. Further, such an orbit allows the observation of the same region of the Earth at different local-times. In this way, the diurnal cycle of surface temperatures can be reconstructed with a 2.4-hour local-time interval. The amount of data acquired by an equatorial station has been estimated. Orbital perturbations effects have been taken into account in order to verify the feasibility of the mission; the Results have been reported in another chapter.

PRISMA, that stands for “PRecursore IperSpettrale della Missione Operativa” is an Earth Observation mission funded by Italian Space Agency (ASI) and is the follow-on of previous programmes, both national (HypSeo, phase A/B) and international (JHM phase A, in cooperation with Canada). The Mission will supply worldwide imaging access, with a specific area of interest covering the whole European and Mediterranean region, by providing with unique fusion of hyperspectral and panchromatic imagines. The program is conceived as a precursor of a future operational mission and, consequently, it is mainly driven by the needs of the Italian user community that, for the exploitation of the PRISMA data, will develop a set of dedicated applications. In fact, according to its “public good” nature, PRISMA will focus on the needs of the Italian institutional and research entities. Hyperspectral data, combined with panchromatic imagines, will help to efficiently address the selected applications, such as those related to quality and protection of the environment, sustainable development, climate change, strong economy. The quality of the images and products will be superior to previous and competing systems for what concerns geometry (also thanks to fusion of panchromatic and hyperspectral data), spectral and radiometry.

Concept of new low-orbit system “Radiomet” is offered. This system is intended to global radio occultation monitoring and uses signals of radio navigation satellites passing through lower atmosphere and ionosphere. Authors of the concept – Russian organizations FSUE “RISDE” and IRE RAS.

“Radiomet”, in comparison with existing system “Formosat-3” (COSMIC), should provide more than twice higher quantity of atmospheric parameters measurements per day, higher speed of final data deliver and some other significant improvements.

Plasma-wave experiment on the micro-satellite “Chibis-M” is aimed at the solution of fundamental problem – a study of the interrelation of the plasma- wave processes connected with the manifestation in the ionosphere of solar – magnetosphere – ionosphere – atmosphere connections and the parameters of space weather. Specific fundamental problem is the search for universal laws governing transformation and dissipation of plasma-wave energy in the magnetosphere-ionosphere system. The solution of this problem will be achieved employing the coordinated Procedure: (1) Study in situ of the fluctuations of electrical and magnetic field, the parameters of thermal and epithermal plasma in the ionosphere near layer F during different helio- and geomagnetic conditions. (2) Study of the geomagnetic and geophysical parameters on the ground-based observatories with the time scales from 10

–1

to 10

–3

s. (3) Study of the interrelation of electromagnetic phenomena (spectra of ULF/VLF- waves) in different regions of near-earth space by means of via the comparative analysis of the wave measurements of those carry out simultaneously on different spacecrafts and ground geophysical stations. Micro-satellite “Chibis-M” now designed in IKI. Total mass of “Chibis-M” with support systems construction and scientific instruments is 40 kg.

SEPSAT (Spherical EUV- and Plasma Spectrometer-Satellite) is a nanosatellite that will observe several parameters of space weather, especially solar Extreme Ultraviolet Radiation (EUV), the higher atmosphere and effects of solar fluctuation on parameters of the ionosphere. This helps to improve our knowledge of integrity and accuracy of GNSS signals, the origin and variation of EUV radiation and finally the composition of the solar and upper atmosphere. To detect and observe these parameters a concept for a nanosatellite was designed. SEPSAT is a boom stabilized, 35 × 35 × 35 cm³ satellite, weighing just 15 kg. Most of the subsystems are based on developments of Technical University of Berlin (TU Berlin), where the focus lies on the design of small satellites and their miniaturized components. The SEPSAT payload is a low-cost spherical EUV- and plasma spectrometer.

A constellation of small satellites is well suited to making measurements of the near-Earth space environment given the spatial and temporal variability that preclude solely independent, localized measurements. In this chapter we discuss some of the factors that must be considered when designing a system for space weather monitoring and the scientific investigation of the near-Earth environment. Some important parameters include: (1) the number flux and energy of energetic particles in the auroral region and/or the radiation belts, (2) the number density temperature and composition of neutrals and ions in the upper atmosphere, (3) the wind speed in the upper atmosphere, (4) the ion drift velocity and/or DC component of the electric field, and (5) small scale fluctuations in the ionosphere that lead to radio scintillation. An integrated solution will be addressed that spans the range of potential science objectives, enabling sensor technologies, triple-cubesats (<5 kg) and nanosatellite (<50 kg) system design, secondary launch options, and the associated data processing and distribution plan. In addition, details related to a JHU/APL effort with a planned 2010 mission schedule to serve the requirements of user communities for near-Earth space environment and space weather data will also be discussed.

The technique called “Low-Low Satellite-to-Satellite Tracking” makes use of two-satellite loose formations for detecting the Earth’s gravity field and its space/time variations: the effect of the geopotential shows up as a variation of the inter-satellite distance, which is measured by a suitable metrology. Accelerometers are utilized on each satellite for measuring and separating the effect of the non-gravitational forces. Thales Alenia Space Italia (TAS-I) has studied for the European Space Agency a gravimetric mission of this kind, in which the inter-satellite distance variation is measured by a laser interferometer. The reference mission scenario that has been defined and studied consists of two satellites flying along the same circular orbit at 10 km relative distance and 325 km altitude. The formation control for this mission shall be designed to work in synergy with the drag-free control (necessary for providing quiet operational environment to the accelerometers), to not interfere with the scientific measurement and to minimize the use of the thrusters. Another control system is in charge of maintaining the fine pointing of the interferometer laser beam from one satellite to the other. This chapter summarizes the main results of the studies performed by TAS-I and its team on these subjects 1 2 3.

The aim of the ESA study on concepts for demonstration of advanced techniques and technologies on an EO small mission is to assess ideas for EO missions compatible with implementation on a small satellite such as PROBA, and which may benefit from Formation Flying. “Small Mission” here implies small satellite platforms in the 150–200 kg spacecraft mass range. The output of the study is a definition of various small satellite missions and their required developments, including costs. The study has been led by Astrium Ltd, with support from Astrium SAS, Astrium GmbH, ENVEO, GMV and Verhaert Space. Following initial selection in the first part of the study, 3 candidates have been analysed in more detail as discussed below.

Trajectory design for formation-based bistatic SAR missions is presented for both close (cross-track interferometry) and large (radargrammetry) formations. Trajectory design process is simplified by using analytical relative motion models that include Earth oblateness perturbations and optimize mission performance. In addition, proposed design reduces trajectory establishment and maintenance efforts. examples are briefly summarized where best relative trajectory is derived starting from application models.

The chapter presents the latest results in the design of FAST-D, the Dutch micro-satellite for the Dutch–Chinese FAST (Formation for Atmospheric Science and Technology demonstration) formation flying mission. Over the course of the 2.5 year mission, the two satellites, FAST-D and FAST-T, will demonstrate various new technologies and perform observations of atmospheric aerosols and seasonal variations of height profiles in the cryosphere using spectropolarimeter and altimeter payloads on both spacecraft. A conceptual design for the Dutch spacecraft, FAST-D, is presented. Special focus is laid on the design of the attitude determination and control subsystem and on the space-based computing experiments to be performed on this spacecraft. Furthermore, new results in the development of the science payloads on FAST-D, the aerosol characterisation instrument SPEX (Spectropolarimeter for Planetary Exploration) and the altimeter SILAT (Stereo Imaging Laser Altimeter), are described. For SPEX, several design changes have been made to make the instrument more compatible with the FAST mission. For SILAT, an instrument re-design for Earth missions is presented, which results in considerable mass savings as compared to the earlier design.

Japan Canada Joint Collaboration Satellites (JC2Sat) is a joint project between the Canadian Space Agency (CSA) and the Japan Aerospace Exploration Agency (JAXA). The main objective of the project is to design, build, launch and operate two 18 kg nearly identical nanosatellites that will be launched as a stack and separated in space to demonstrate the feasibility of Autonomous Formation Flight (AFF) based on aerodynamic differential drag only. The specific configuration of the JC2Sat nanosatellites serves as an ideal technology demonstration platform for the Miniature far Infra-Red Radiometer (Mirad) instrument developed jointly by CSA and Institut National d’Optique (INO). Each nanosatellite carries a Mirad instrument for the purpose of Earth’s limb sounding. Based on un-cooled far infrared micro-bolometers, these low mass and low power payloads will co-register the limb profiles in the emission bands of greenhouse gases CO

2

and H

2

O, respectively centered at wavelengths of 15 and 25 μm. The development of JC2Sat is carried out by a united small team consisting of engineers and researchers from both CSA and JAXA. JC2Sat is planned to be ready for launch in 2011. Mission duration is defined to be one year. At the time of writing, the project is in phase C.

Microsatellites are already playing a major role in operational optical-imaging remote sensing applications,

viz.

the Disaster Monitoring Constellation (DMC) and RapidEye. However, such satellites are currently limited to providing useful images over cloud free regions during daylight. Whilst the use of constellations does increase the probability of achieving useful images within short timescales over most regions, there are still areas, such as the tropics, that have significant cloud cover, or extreme latitudes, which can have low light levels, where the all weather, day/night imaging capability of Synthetic Aperture Radar (SAR) would be beneficial.

Current SAR satellites typically have complex, large volume, large mass and high power payloads. Usually they operate singly with relatively long repeat cycles. Recent missions have shown that SAR payloads can be flown on smaller satellites, but there are still significant size/cost barriers to being able to set up affordable SAR satellite constellations. To this end, a feasibility study has been undertaken to examine incorporating a low cost L-band SAR payload into a 100 kg class microsatellite – primarily for disaster monitoring and forestry/flood monitoring. The radar makes use of continuous wave (CW) technology and bi-static pairs, to provide modest resolution (30–100 m) images over 30 km swaths.

The chapter treats the mission design for the Dutch-Chinese FAST (Formation for Atmospheric Science and Technology demonstration) mission. The space segment of the 2.5 year mission consists out of two formation flying micro-satellites. During the mission, new technologies will be demonstrated and, using spectropolarimeter and altimeter payloads on both spacecraft, observations will be performed characterizing atmospheric aerosols and seasonal variations of height profiles in the cryosphere. The mission is divided into four phases, each with a different orbital geometry. The rationale for and the orbital geometry during these phases as well as the transitions between the phases are treated in detail. A major complication to the mission design is the amount of data that can be sent to the ground. Since only two moderately capable ground stations, one in Delft and one in Beijing, are baselined, as much data processing as possible has to be performed onboard to allow high duty factors for the science payloads. When this is not possible, alternative payload operation modes have to be sought with which maximum scientific data return can be obtained through as little payload operation as possible. Both options are dealt with in the chapter.

SPRITE-SAT is a micro satellite in the size of 50 cm cube and weighing 45-kg, designed and developed by Tohoku University. Its mission objective is to conduct scientific observation of atmospheric luminous emissions called “sprites” and terrestrial Gamma-ray flushes. Both are recently discovered phenomena and their mechanisms are still under the veil. SPRITE-SAT was developed to achieve significant observations to determine clear models of these mysterious phenomena. On January 23rd, 2009, SPRITE-SAT was successfully launched by JAXA’s H-IIA rocket as a piggyback payload of Greenhouse Gas Observation Satellite (GOSAT). The spacecraft is now in a sun-synchronous polar orbit with 670 km altitude form the Earth’s surface.

In 2006, Ecole Polytechnique Fédérale de Lausanne (EPFL) in cooperation with other Swiss academic and industrial partners started the SwissCube project based on the CubeSat program. The primary objective of the SwissCube project is to provide a dynamic and realistic learning environment for undergraduates, graduates and to foster the development of small satellite technologies. Besides the educational objectives, the SwissCube mission objective is to observe the airglow phenomenon (intensity) over selected latitudes and longitudes for a period of 4 months (possibly up to 1 year). The payload consists of a telescope which takes images of the airglow emissions. The telescope has a length of 50 mm. At one end, a CMOS detector captures images with a resolution of 188 × 120 pixels and a pixel size of 24 μm via a focusing optics. A bandpass filter centered at 767 nm, with a bandwidth of 20 nm, selects the desired wavelength of the airglow. At the other end, a baffle protects the optical system and the detector from straylight.

This chapter describes the mission of the satellite and provides information on the Earth observation payload.

International cooperation in space exploration and education can be a benefit but also a challenge to the accomplishment of a project. The German Russian Education Satellite (GREsat) is a cooperation between the German Center of Applied Space Technology and Microgravity (ZARM), the Russian Institute of Space Device Engineering and the Keldysh Institute of Applied Mathematics (KIAM). The main objective of the program is the education of students from both universities and the technology verification of the satellite subsystems. An annual exchange of students could be established with the support of the German Academic Exchange Service (DAAD).

The chapter will roughly outline the mission of GRESat and takes a deeper look onto the new technologies and concepts, that shall be demonstrated with this satellite. The cooperation of the partners also lead to a cooperation for the provision of components for the satellite bus. One of the main aspects will be a heterogeneous redundancy of the onboard data processing by using two different computers. Here the communication between the computers and the mode management are difficult topics to adress, when the cooperation of these units shall work without problems. The communication system for the command and data interface to the ground station is also planned as a redundant system, as Globalstar and ORBCOMM services shall be accessible from the satellite. That enables the test of both communication systems and provides a failure tolerant data interface. The chapter will outline what challenges the layout of the communication protocoll between both onboard computers has faced, and how the application of two independant satellite communication systems could be designed to benefit from the redundant hardware and services.

Another part of the on-orbit verification will be the usage of an AMR magnetometer and new micro torquers. These devices are acting as shared resources, so both onboard computers are allowed to access them for attitude control maneuvers. While the sensor delivers its data constantly to both computers, the magnetic torquers have to be managed for an exclusive access of only one computer at the same time. A defined mode management and proper inter-computer communication shall ensure the proper function of these subsystems.

Due to the mission outline as a technology verification and student education project, another mission objectives is also the test of novel attitude control algorithms by using magnetic sensors and actuators. A fallback algorithm (B-dot) is used to stabilize the spacecraft in a non conform situation. This shall guarantee a good connectivity for the command and data interface with the ground station via Globalstar and ORBCOMM. Any further mode changes and failure handling can then be prepared offline on ground and uploaded to the satellite afterwards.

Currently, space missions must take into account a relatively new threat, which is represented by space debris, i.e. the orbital residuals of past artificial satellites and rockets. This problem has arisen in the last 25 years and requires specific strategies for mitigation, with the main intent of avoiding collisions between orbital debris and spacecrafts. Space debris monitoring and orbit determination is an essential premise to this task.

In the last decade, the Group of Astrodynamics of the University “La Sapienza” (GAUSS) of Rome has been involved in the optical surveillance of space debris. This activity was carried out using amateur astronomers’ facilities and led to participating to the IADC (Inter-Agency Space Debris Coordination Committee) observation joint campaign. In 2006 GAUSS completed the first Italian observatory, fully dedicated to space debris monitoring. In this context, GAUSS students have analyzed the feasibility of two different formation flight missions, tailored to detecting space debris and, more specifically, focused on the detection of objects along MEO and LEO orbits. This paper deals with the possibility to carry out an optical surveillance of space objects by using a ground-based telescope and a microsatellites equipped with an optical payload. The feasibility of placing an optical payload (capable of detecting space debris) onboard the microsatellite UNISAT5 is investigated.

MOVE (Munich Orbital Verification Experiment) is a program of the Institute of Astronautics (LRT) at the Technische Universität München (TUM), which aims on building pico-satellites with university students mainly for educational purposes. First-MOVE shall create a robust platform as a starting point for sophisticated satellite missions of the institute in the future. In the paper, the state of development is described, but emphasis is on the requirements for high reliability of the First-MOVE satellite and how robustness drives the actual design of the satellite.

This chapter deals with the design of the EduSAT microsatellite: a small educational satellite developed by the Group of Astrodynamics of University of Roma “Sapienza” (GAUSS), on the basis of its previous experience. The UNISAT program (UNIversity SATellite) started at School of Aerospace Engineering of Roma in the early nineties. The EduSAT Project is funded and coordinated by Italian Space Agency with the aim to promote space education among high school students and to support the qualification and scientific careers of young people (university students, PhD students and young researchers). Another target of this program is to develop a small space mission for low cost scientific experiments and technological tests in orbit. The launch of EduSAT microsatellite is scheduled in 2010: a cluster launch in Low Earth Orbit, performed by Russian-Ukrainian DNEPR Launch Vehicle. This chapter synthesizes project motivations, program organization and describes system architecture and satellite main subsystems design.

On October 22nd 2008, the VERTICAL (VERification and Test of the Initiation of CubeSats After Launch) experiment was flown on the REXUS 4 sounding rocket mission at Esrange in Kiruna, Sweden. The experiment’s objective was to verify critical hardware to be used on the MOVE

CubeSat

in a space environment. The items to be verified were multiple micro switches from different manufacturers and a

solar panel deployment mechanism

developed at TUM. The deployment mechanism is triggered by a melt wire. During launch, the switches are depressed by a plate which is retracted once the rocket is near its apogee. This simulates the satellite’s ejection from the launch vehicle. The verification sequence was executed as planned during the 10-min flight and the experiment was safely recovered. The acquired data suggests that the deployment mechanism can be used as is and COTS of verified quality micro switches will survive LEOP conditions and are suitable for further testing, addressing their long-term reliability.

The laboratory facility named LuVeX is developed for motion control algorithms verification of the formation flying purposed for the Earth remote sensing. The mock-up description, the parameter determination of its individual elements, description of the realized algorithms for its thruster engine control are given in the chapter. The mathematical simulation results of a mock-up motion controlled with the algorithms which provide the required mock-up formation motion in given trajectories are presented. Several configurations of the formation and formation’s elements coupling methods are shown. Required motion precision and its dependence on formation parameters are examined.

This chapter describes the development of satellite control algorithms implemented into FPGA (Field Programmable Gate Array) hardware logic, which is gathering a great interest in applying for central computing systems of small satellites with high computational demands. In order to maximize the parallel processing capability of the FPGA, the satellite control algorithms are classified into subsystems and are implemented as a parallel running multi-agent system. The developed asynchronous parallel reactive system is capable of real-time processing and reasoning, which enables the implementation of the multi-agent system in FPGAs. This system introduces a reactive subsumption system, which has been ever proven its potential for autonomous systems, and is exemplary applied for the small satellite

Flying Laptop.

This concept offers further possibilities to be extended as an on-board autonomous system based on a Belief-Desire-Intention architecture. The performance of the implemented control algorithm is evaluated and its potential for autonomous space systems is investigated.

NanoSiGN will be the first nanosatellite with permanent 3-axes stabilization dedicated to the scientific interpretation of GNSS (Global Navigation Satellite System) dual-frequency signals in the low Earth orbit. It will carry and operate a complex GNSS receiver and antenna system for Precise Orbit Determination (POD) and GNSS ionospheric remote sensing based on measurements with a dual-frequency GNSS receiver for space applications whose design is based on Commercial-of-the-shelf (COTS) technologies and which is therefore free of International Traffic in Arms Regulations (ITAR) limitations. The NanoSiGN design builds on the know-how of the longtime TUBSAT series and on volume- and power-saving technologies from the recent faulttolerant BeeSat design by the TU Berlin. It is a robust design set up to be adaptable to other payloads. Despite its comparatively high volume of 350 × 350 × 350 mm³ NanoSiGN will be a nanosatellite of not more than 20 kg. It takes advantage of new and innovative technologies sized for nanosatellites such as micro-reaction wheels and a star tracker for precise attitude determination and control.

In this chapter recent progress in the development of the Vegetation Instrument (VI) for the PROBA-V mission is presented. PROBA-V is an earth observation mission, which will ensure continuity of the actual SPOT/ VEGETATION mission until Sentinel-3 will become operational.

The consortium will consist of VITO as prime investigator, VES as mission prime, OIP as prime contractor for the payload and Xenics and AMOS as most important subcontractors for the payload.

The payload is a multispectral spectrometer with 4 spectral bands. To guarantee daily coverage of the earth, a very large swath of 2,250 km is envisaged. The ground resolution of the instrument will be at least 1,000 m.

The payload will consist of 3 SIs (Spectral Imagers), each containing a very compact TMA (Three Mirrors Anastigmat) telescope, and a large VNIR and SWIR FPA (Focal Plane Array). The three SIs will be mounted on a common optical bench, and aligned to each other to cover the full swath.

Feasibility studies have revealed two major criticalities in the payload, being the TMA manufacturing and alignment, and the SWIR FPA manufacturing. To reduce potential risks during the very short development time of the payload, two pilot projects were already started. This chapter presents the first achievements of both development projects.

There is an increasing demand for a low cost, day-night, all weather spaceborne imaging capability using synthetic aperture radar (SAR) on small satellites. Traditional pulsed SAR payloads have been too expensive and too power demanding to be employed on low-cost microsatellite platforms. Recent developments based on Continuous Wave (CW) techniques have proved successful in minimizing the cost, mass and power of SARs for small airborne platforms (UAVs). However, when considering the use of CW techniques for space based SARs, other considerations come into place: A major one being need of high isolation between transmit and receive antennas, in short, a bi-static configuration. Conventional receiver designs based on analogue demodulation techniques remain inflexible and are complicated to change to suit changing operational requirements in a dynamic bi-static satellite formation. This chapter focuses on the research being carried out in the Surrey Space Centre (SSC) on the design and development of a software defined linear frequency modulated (LFM) CW SAR receiver that can be used on bi-/multi-static microsatellites for remote sensing.

Recent advancement on detectors, optics fabrication, metrology and interference filters are the basis of a new and compact hyperspectral instrument to map vegetation and soil in applications requiring fast revisit time and medium spatial resolution.

The instrument stems from the design of the Proba-V, a small satellite developed to ensure continuation of the Spot-Vegetation products. The instrument presented in this paper has been optimized to accommodate a detector with a slightly different size that provides the necessary format for a hyperspectral instrument.

With a mass of only 5 kg, a power consumption of 5 W, an overall size of 300 × 220 × 160 mm (W × L × H), and with a staggering 2,400 spectral channels covering the VNIR spectral range, the instrument provides unprecedented ratio between mass and resolution, opening the possibility to perform hyperspectral imaging to both small satellites for Earth Observation and small interplanetary probes.

The paper presents the design, performance of this instrument, and the development status of the critical technologies, namely optics manufacturing, development of the linear variable spectral filter, and of the detector.

The earth observation micro-satellite “TAIKI” is 50 kg satellite which has low-cost and small bus-subsystem. TAIKI is characterized by a low-cost spaceborne small hyperspectral sensor “HSC-III”. HSC-III is targeted at the performances of 30 m ground sampling distance, visible and near infrared wavelength range, 10 nm spectral resolutions, 61 spectral bands and 10 kg weight. HSC-III consists of the telescope, the imaging spectrometer, the electrical system, the on-orbital calibration equipment. The telescope has a pupil diameter of 0.2 m, and has two mirror configuration of Ritchey-Chretien type. The spectrometer has the transmitting grating with the slit and relay lens unit, and array sensor using back-illumination type CMOS image sensor. As a SNR model of HSC-III, we did some calculations and concluded that SNR is approximately 340. Last year, we succeeded to develop the breadboard model of HSC-III optics instrument, and we obtained result of more requirement specification. Also, we have developed the on-orbit spectral calibration equipment. It achieved 0.02 nm of spectral calibration accuracy.

In the following we review the optimization for microsatellite deployment of a highly integrated payload suite comprising a high resolution camera, an additional camera for stereoscopic imaging, and a single photon counting laser altimeter. This payload suite, the “Stereo Imaging Laser Altimeter” SILAT has been designated for deployment aboard the FAST microsatellite formation mission for Earth observation. This instrument suite has been designed for a Jupiter mission, but has been redesigned and optimized for an Earth observation mission. This chapter reviews the simulated Earth observation performance, the design modifications made for the mission and the optimization of the design for microsatellite use. Mass and power budgets are used to demonstrate the changes and the performance analysis is represented trough the simulation results. It is expected that the optimization will reduce the mass of the instrument by approximately 20% without compromising the performance of the instrument. In addition, results from breadboarding experiments of individual instrument components will be presented to show the progress from design optimization towards the FAST flight model.

Propose a plan of spaceborne inverse synthetic aperture radar (SISAR) imaging to space object based on present techniques and equipments, and provide practical approach from on-board-computer (OBC) to payload to carry out mission requirements. Space objects are classified into three categories through the analyses of motion characters, three-axis-stabilized satellite, spin-stabilized satellite and space debris. Then simulation models are established respectively to achieve imaging investigations. The result shows that objects belonged to three-axis-stabilized and spin-stabilized satellite can be reconstructed clearly if proper parameters are set, and space debris can be detected and estimated correctly. The primary results verify the feasibility of SISAR system.

An X-band Antenna Pointing Mechanism has been developed in order to improve the power efficieny of data downlinking on small, low cost Science and Earth Observation satellite missions. This permits the accompanying payload chain to support high speed data storages, processing, and downlinking in X-band. The system is scalable, and can be implemented in redundant or single string configuration. The unique aspect of the downlink chain is the implementation of an antenna pointing mechanism. Such systems have not been used before on small satellites before, but it permit the system designer to either maximise the downlink data throughput, or minimise the power resources on the spacecraft. As such, the small antenna pointing mechanism is an enabling technology in commercialisation of small EO spacecraft missions.

RCST-HIT (Research Centre of Satellite Technology, Harbin Institute of Technology) is now in the process of designing a new Plug-and-Play (PnP) satellite, carrying X-band, HH polarization spaceborn Synthetic Aperture Radar (SAR), offering the 1 meter resolution image in ScanSAR mode, less than 300 km sun- synchronism orbit. The satellite contains a light weight high accuracy Paraboloid deployable reflector antenna. The SAR component electronic beam steering capability is achieved by using a feed array in the focal plane. This project will make the small SAR satellite advance the modular satellite with open standards and interfaces, self describing components, and an auto-configuring system. And the integration design based Plug-and play is simplified and testing tasks can be automated.

Already the first satellite of the Technical University Berlin, TUBSAT-A (launched in 1991) contained a University built CCD star sensor with a unique technique to reduce the sensor data by three orders of magnitude. In the following 15 years of operation this technique proved to be very successful. On the other hand, TUBSAT-A was one of the first satellites to discover and investigate the impact of proton radiation that is responsible for the hotspot phenomenon.

Based on this extensive experience a simple hot spot filter was conceived and successfully tested on MAROC-TUBSAT that was launched in 2001 and is still operational.

A star sensor of the third generation is scheduled to fly on the Indonesian missions LAPAN-A2 and AMSAT-ORARI in the beginning of 2010. It excels, based on the on board experience of the previous designs by its compact size (one unit of 700 g), high signal to noise ratio even in the presence of hotspots and robustness in star extraction and identification.

The chapter will present the results of the in-orbit experiments with TUBSAT-A and MAROC-TUBSAT and discuss the design principles and improvements that lead to the design of the latest model.

Accurate attitude control is as important for a small satellite as it is for a big one. TNO has been engaged in micro sunsensor developments for a couple of years now and has reached a situation where some critical decisions will have to be taken. Several options are still open but as research resources are limited some possible solutions will have to be discarded or further investigations postponed shortly. This chapter describes the current status as well as possible developments for the future which would lead to sensors that are small and reliable enough to provide a grown up attitude sensing solution for even the smallest satellites.

The chapter presents the design of the attitude and orbit control system (AOCS) of the small satellite TET-1 (Technology verification career) as an example of a cost effective but still robust and reliable AOCS for small satellites. The AOCS of TET-1 is fully three-axis stabilized. Particular attention is paid to the implementation of robust and fault tolerant design of the AOCS. The redundancy management concept and robust control algorithms are presented. Furthermore the fault detection, isolation and recovery (FDIR) mechanism, which are implemented in various functional levels, are shown. Finally the chapter presents the test strategy for verifying the proposed fault tolerant design.

This chapter presents the interfaces that exist between the TNO, TU-Delft and UTwente developed micro propulsion system experiment (T3μPS) and the Delfi-n3Xt satellite currently under development at TU-Delft. Delfi-n3Xt is successor to Delfi-C3, the first Dutch student nano-satellite. Delfi-n3Xt, like its predecessor is a 3-unit cubesat, but is more advanced in that it is equipped with an active three-axis attitude determination and control system, eclipse power and a high speed data link.

The T

3

μPS is especially developed to provide for a means of propulsion fit for use in nano- and microsatellites. Such a means is considered essential for e.g. future missions of clusters of satellites which require relative orbit control. It employs a cold gas thruster delivering a thrust in the order of a few up to several tens of millinewton. As propellant, the system uses cool nitrogen gas, stored in solid form in one or more cool gas generators.

The T

3

μPS hardware is assembled on a single printed circuit board easing integration of the system in Delfi-n3Xt.

When operating, the T

3

μPS may disturb the proper operation of Delfi-n3Xt.This is because of thrust misalignment. This disturbance needs to be compensated for to keep the solar panels of the satellite sun-pointed.

The chapter presents the Delfi-n3Xt satellite as well as the T

3

μPS experiment and describes the interfaces between the two. Furthermore, it investigates the implications on other satellite subsystems with particular emphasis on the ADCS. Finally, some Conclusions and recommendations are given.

Propulsion systems are required by several space applications, e.g. Formation Flying, when position and attitude control are foreseen and shall be very fine. The system design can present many challenges regarding system mass and performances: a great number of equipments and thrusters may be required and low and variable thrust values could be needed to manage disturbances on orbit. Throttleable thrusters often lead to an increase of complexity and system size, against small-spacecraft constraints. In this chapter a novel micro-propulsion system is proposed, based on MEMS micro-thrusters developed in house by Carlo Gavazzi Space S.p.A. The system design architecture refers to a baseline configuration where each thrust valve commands a torque around an assigned direction. Optimization for two Italian missions (micro and nanosatellite) is analysed. Both applications require high-accuracy attitude control, involving low thrust values (order of 10 mN) and low impulse bits (order of 10

–5

Ns). With respect to the first application, a comparison with a reaction-wheel based system in terms of mass and power has been made, showing a power gain of about 5 W against a propellant mass of about 1.5 g for each attitude manoeuvre (200 s each) in LEO orbit.

This chapter outlines GNSS receiver requirements driven by navigation needs identified from applications in future ESA Low Earth Observation (LEO) satellites. In the domain of operational EO, the GMES mission Sentinel-3 (S3) poses the most demanding positioning requirements in real-time in order to aid the altimeter tracking and also in on-ground post-processing for POD. The Sentinel-1 (S1) navigation requirements are also demanding in post-processing in view of differential interferometry applications with its SAR payload. Sentinel-2 has more relaxed navigation needs as compared to S1 and S3. A series of these Sentinels is planned to be manufactured with a common receiver and a common processing procurement approach. Sentinel-3, and possibly other Sentinels too, will require Galileo-compatible receivers in-orbit from 2017. The Sentinel-5 precursor (S5p) mission for air-quality and the candidate mission to the 7th Earth Explorer (i.e. Biomass, CoreH2O, Premier) require GNSS solutions, but the navigation requirements are less demanding than for S3 and S1. Within meteorological missions, the GNSS Radio Occultation instrument on Post-EPS will need to meet very demanding along-track velocity requirements and requires a high number of channels compatible with multiple GNSS. Regarding Earth science missions, the high-resolution gravity monitoring with Post-GOCE implies demanding post-processing relative position and velocity noise measurements. As a general trend, besides a multi-GNSS capability (GPS and Galileo, at least), the capability of on-board POD, already partly considered in S3, is likely to become more of a common feature for all missions. The robustness and reliability of the navigation solution will also become increasingly important. This chapter also describes the key technology “building blocks”, like the AGGA-4 digital GNSS processor to fulfill the requirements in these future missions.

This chapter outlines the benefits of Galileo for future satellite missions. In this context, it will address the Galileo services, signals and frequencies. Further, aspects of interoperability with other GNSS and their potential impact on space applications will be discussed.

The differential processing of carrier phase measurements from Global Navigation Satellite Systems (GNSS) is a well know technique for relative positioning in terrestrial and aeronautical applications. Over the past decade intensive research has been conducted to demonstrate its suitability for high-precision relative navigation of spacecraft in low Earth orbit (LEO). This chapter describes the fundamental concepts of differential GNSS and its application to spacecraft formation flying. A review of past achievements is given and the practical aspects of differential GPS are discussed for real-time and offline navigation in the upcoming PRISMA and TerraSAR-X missions.

The problem of relative navigation of formation flying satellites for remote-sensing applications is addressed in this chapter. Specifically, a filtering scheme is proposed for the determination of relative position and velocity in real time, relying on the cascade combination of an Extended Kalman Filter and a kinematic filter. Since high measurement accuracy is desired both filters process dual-frequency double-difference carrier-phase GPS observables. In addition, the dynamic filter state is augmented with the Vertical Total Electron Content. Filtering scheme performance in terms of accuracy and robustness are evaluated by Monte Carlo simulations on orbital scenarios relevant to considerable Earth remote sensing applications.results shows that centimetre-level average accuracy can be achieved in real-time and it can be maintained in presence of poor GPS satellite geometry.