CCDs have been used for many years for Hyperspectral imaging missions and have been extremely successful. These include the Medium Resolution Imaging Spectrometer (MERIS) [1] on Envisat, the Compact High Resolution Imaging Spectrometer (CHRIS) on Proba and the Ozone Monitoring Instrument operating in the UV spectral region. ESA are also planning a number of further missions that are likely to use CCD technology (Sentinel 3, 4 and 5). However CMOS sensors have a number of advantages which means that they will probably be used for hyperspectral applications in the longer term.

There are two main advantages with CMOS sensors: First a hyperspectral image consists of spectral lines with a large difference in intensity; in a frame transfer CCD the faint spectral lines have to be transferred through the part of the imager illuminated by intense lines. This can lead to cross-talk and whilst this problem can be reduced by the use of split frame transfer and faster line rates CMOS sensors do not require a frame transfer and hence inherently will not suffer from this problem. Second, with a CMOS sensor the intense spectral lines can be read multiple times within a frame to give a significant increase in dynamic range.

We will describe the design, and initial test of a CMOS sensor for use in hyperspectral applications. This device has been designed to give as high a dynamic range as possible with minimum cross-talk. The sensor has been manufactured on high resistivity epitaxial silicon wafers and is be back-thinned and left relatively thick in order to obtain the maximum quantum efficiency across the entire spectral range

A novel detector, incorporating e2v’s L3 CCD (L3Vision™) [1] technology for use in LIDAR (Light Detection And Ranging) applications has been designed, manufactured and characterised. The most critical performance aspect was the requirement to collect charge from a 120μm square detection area for a 667ns temporal sampling window, with low crosstalk between successive samples, followed by signal readout with sub-electron effective noise. Additional requirements included low dark signal, high quantum efficiency at the 355nm laser wavelength and the ability to handle bright laser echoes, without corruption of the much fainter useful signals.

The detector architecture used high speed charge binning to combine signal from each sampling window into a single charge packet. This was then passed through a multiplication register (Electron Multiplying Charge Coupled Device) operating with a typical gain of 100X to a conventional charge detection circuit. The detector achieved a typical quantum efficiency of 80% and a total noise in darkness of < 0.5 electrons rms. Development of the detector was supported by ESA (European Space Agency).

Solar Dynamics Observatory (SDO) is imaging the Sun in many wavelengths near simultaneously and with a resolution ten times higher than the average high-definition television.

In this paper we describe our innovative systems approach to the design of the CCD cameras for two of SDO’s remote sensing instruments, the Atmospheric Imaging Assembly (AIA) and the Helioseismic and Magnetic Imager (HMI). Both instruments share use of a custom-designed 16 million pixel science-grade CCD and common camera readout electronics. A prime requirement was for the CCD to operate with significantly lower drive voltages than before, motivated by our wish to simplify the design of the camera readout electronics. Here, the challenge lies in the design of circuitry to drive the CCD's highly capacitive electrodes and to digitize its analogue video output signal with low noise and to high precision. The challenge is greatly exacerbated when forced to work with only fully space-qualified, radiation-tolerant components. We describe our systems approach to the design of the AIA and HMI CCD and camera electronics, and the engineering solutions that enabled us to comply with both mission and instrument science requirements.

MTF (Modulation Transfer Frequency) of a detector is a key parameter for imagers. When image is not moving on the detector, MTF can be measured by some methods (knife edge, slanted slit,…). But with LEO satellites, image is moving on the surface of the detector, and MTF has to be measured in the same way: that is what we call "dynamic MTF". CNES (French Space Agency) has built a specific bench in order to measure dynamic MTF of detectors (CCD and CMOS), especially with component working in TDI (Time delay and integration) mode. The method is based on a moving edge, synchronized with the movement of charges inside the TDI detector. The moving part is a rotating cube, allowing a very stable movement of the image on the surface of the detector The main difficulties were:

- stability of the rotating speed

- synchronization between cube speed and charge transfer inside the detectors

- synchronization between cube position and data acquisition.

Different methods have been tested for the displacement of the knife edge:

- geometrical displacement

- electrical shift of the charge transfer clocks.

Static MTF has been performed before dynamic measurements, in order to fix a reference measurement, Then dynamic MTF bench has been set up. The results, for a TDI CCD show a very good precision. So this bench is validated, and the dynamic MTF value of the TDI CCD is confirmed.

Three mirror anastigmat (TMA) telescope designs [1] had been implemented in different projects ranging from the narrow Field-Of-View large instruments as Quickbird (2° FOV) [2] to smaller telescopes as JSS 12° FOV developed for RapidEye mission [3].

This telescope configuration had been also selected for the PROBA-V payload, the successor of Vegetation, a multispectral imager flown on Spot-4 and subsequently on Spot-5 French satellites for Earth Observation and defence. PROBA-V, small PROBA-type satellite, will continue acquisition of vegetation data after the lifetime of Spot-5 expires in 2012.

The PROBA-V TMA optical design achieves a 34° FOV across track and makes use of highly aspherical mirrors. Such a telescope had become feasible due to the recently developed Single Point Diamond Turning fabrication technology. The telescope mirrors and structure are fabricated in aluminium and form an athermal optical system.

This paper presents the development of the compact wide FOV TMA, its implementation in PROBA-V multispectral imager and reviews optics fabrication technology that made this development possible. Furthermore, this TMA is being used in combination with a linear variable filter in a breadboard of a compact hyperspectral imager. Moreover, current technology allows miniaturization of TMA, so it is possible to use a TMA-based hyperspectral imager on a cubesat platform.

PILOT is a balloon borne experiment, which will measure the polarized emission of dust grains, in the interstellar medium, in the sub millimeter range (with two photometric channels centered at 240 and 550 μm).

The primary and secondary mirror must be positioned with accuracies better than 0.6 mm and 0.06°. These tolerances include environmental conditions (mainly gravity and thermo-elastic effects), uncertainties on alignments, and uncertainties on the dilatation coefficient. In order to respect these tolerances, we need precise characterization of each optical component. The characterization of the primary mirror and the integrated instrument is performed using a dedicated submillimeter test bench.

A brief description of the scientific objectives and instrumental concept is given in the first part. We present, in the second and in the third part, the status of these ground tests, first results and planned tests.

Multivariable parametric cost models for space telescopes provide several benefits to designers and space system project managers. They identify major architectural cost drivers and allow high-level design trades. They enable cost-benefit analysis for technology development investment. And, they provide a basis for estimating total project cost. A survey of historical models found that there is no definitive space telescope cost model. In fact, published models vary greatly [1]. Thus, there is a need for parametric space telescopes cost models. An effort is underway to develop single variable [2] and multi-variable [3] parametric space telescope cost models based on the latest available data and applying rigorous analytical techniques.

Specific cost estimating relationships (CERs) have been developed which show that aperture diameter is the primary cost driver for large space telescopes; technology development as a function of time reduces cost at the rate of 50% per 17 years; it costs less per square meter of collecting aperture to build a large telescope than a small telescope; and increasing mass reduces cost.

The Broadband Radiometer (BBR) is an instrument being developed for the ESA EarthCARE satellite. The BBR instrument objective is to provide measurements of the reflected short-wave (0.25-4.0 μm) and emitted long-wave (4.0-50 μm) top of the atmosphere (TOA) radiance over three along-track views (forward, nadir and backward). The instrument has three fixed telescopes, one for each view, each containing a broadband detector. The BBR instrument is led by SEA in the UK with RAL responsible for the BBR optics unit (OU) while EADS Astrium is the EarthCARE prime contractor. A detailed description of the instrument is provided in [1].

The BBR detectors consist in three dedicated assemblies under the responsibility of INO. The detectors development started in 2008 and led to the design and implementation of a new gold black deposition facility at INO [2], in parallel with the preliminary and detailed design phases of the detector assemblies. As of today, two breadboard models and one engineering model have been delivered to RAL. In the BBR OU each detector mechanically interfaces with the telescope and electrically with the front-end electronics (FEE). The detectors’ development is now at the Critical Design Review (CDR) level.

This paper first provides a description of the detector design along with its principles of operation. It further presents and discusses measurement and analysis results for the performance characterization of the engineering model in the context of the applicable requirements. Detector-level qualification planning is finally discussed.

The achievement of deep ultraviolet (UV) focal plane arrays (FPA) is required for both solar physics [1] and micro electronics industry. The success of solar mission (SOHO, STEREO [2], SDO [3]…), has shown the accuracy of imaging at wavelengths from 10 nm to 140 nm to reveal effects occurring in the sun corona. Deep UV steppers at 13 nm are another demanding imaging technology for the microelectronic industry in terms of uniformity and stability. A third application concerns beam shaping of Synchrotron lines [4]. Consequently, such wavelengths are of prime importance whereas the vacuum UV wavelengths are very difficult to detect due to the dramatic interaction of light with materials.

The fast development of nitrides has given the opportunity to investigate AlGaN as a material for UV detection. Camera based on AlGaN present an intrinsic spectral selectivity and an extremely low dark current at room temperature. We have previously presented several FPA dedicated to deep UV based on 320 x 256 pixels of Schottky photodiodes with a pitch of 30 μm [4, 5]. AlGaN is grown on a silicon substrate instead of sapphire substrate only transparent down to 200 nm.

After a flip-chip hybridization, silicon substrate and AlGaN basal layer was removed by dry etching. Then, the spectral responsivity of the FPA presented a quantum efficiency (QE) from 5% to 20% from 50 nm to 290 nm when removing the highly doped contact layer via a selective wet etching. This FPA suffered from a low uniformity incompatible with imaging, and a long time response due to variations of conductivity in the honeycomb. We also observed a low rejection of visible. It is probably due to the same honeycomb conductivity enhancement for wavelength shorter than 360 nm, i.e., the band gap of GaN.

We will show hereafter an improved uniformity due to the use of a precisely ICP (Inductively Coupled Plasma) controlled process. The final membrane thickness is limited to the desertion layer. Neither access resistance limitation nor long response time are observed. QE varies from 5% at 50 nm to 15% at 6 nm (85% more when taking into account the filling factor). Consequently, we can propose prototypes concerning not only “solar blind” camera optimized for narrow band in the near UV range (between 280 nm and 260 nm), but also devices with spectral range extended in the deep UV (290 nm to 10 nm). Both detectors are available for an optical budget evaluation.

LYRA is a solar radiometer, part of the PROBA-2 micro-satellite payload (Fig. 1). The PROBA-2 [1] mission has been launched on 02 November 2009 with a Rockot launcher to a Sun-synchronous orbit at an altitude of 725 km. Its nominal operation duration is two years with possible extension of 2 years. PROBA-2 is a small satellite developed under an ESA General Support Technology Program (GSTP) contract to perform an in-flight demonstration of new space technologies and support a scientific mission for a set of selected instruments [2]. PROBA-2 host 17 technological demonstrators and 4 scientific instruments. The mission is tracked by the ESA Redu Mission Operation Center.

One of the four scientific instruments is LYRA that monitors the solar irradiance at a high cadence (> 20 Hz) in four soft X-Ray to VUV large passbands: the “Lyman-Alpha” channel, the “Herzberg” continuum range, the “Aluminium” and “Zirconium” filter channels. The radiometric calibration is traceable to synchrotron source standards [3]. LYRA benefits from wide bandgap detectors based on diamond. It is the first space assessment of these revolutionary UV detectors for astrophysics. Diamond sensors make the instruments radiation-hard and solar-blind (insensitive to the strong solar visible light) and, therefore, visible light blocking filters become superfluous. To correlate the data of this new detector technology, silicon detectors with well known characteristics are also embarked. Due to the strict allocated mass and power budget (5 kg, 5W), and poor priority to the payload needs on such platform, an optimization and a robustness of the instrument was necessary. The first switch-on occured on 16 November 2009. Since then the instrument performances have been monitored and analyzed during the commissioning period. This paper presents the first-light and preliminary performance analysis.

Proba-V payload is a successor of the Vegetation instrument, a multispectral imager flown on Spot-4 and subsequently on Spot-5, French satellites for Earth Observation and defence. The instrument, with its wide field of view, is capable of covering a swath of 2200 km, which, in combination with a polar low Earth orbit, guarantees a daily revisit.

The lifetime of Spot-5 expires in early 2013, and to ensure the continuity of vegetation data, BELSPO, the Belgian Federal Science Policy Office, supported the development of an instrument that could be flown on a Proba type satellite, a small satellite developed by the Belgian QinetiQ Space (previously known as Verhaert Space).

The challenge of this development is to produce an instrument responding to the same user requirements as Vegetation, but with an overall mass of about 30 kg, while the Vegetation instrument mass is 130 kg. This development had become feasible thanks to a number of new technologies that have been developed since the nineties, when Vegetation was first conceived, namely Single Point Diamond Turning fabrication of aspherical mirrors and efficient VNIR and SWIR detectors.

The Proba-V payload is based on three identical reflective telescopes using highly aspherical mirrors in a TMA (Three Mirrors Anastigmat) configuration. Each telescope covers a field of view of 34° to reach the required swath.

One of the challenges in the development of the PROBA-V instrument is the efficient reduction of stray light. Due to the mass and volume constraints it was not possible to implement a design with an intermediate focus to reduce the stray light. The analysis and minimization of the in-field stray light is an important element of the design because of the large FOV and the surface roughness currently achievable with the Single Point Diamond Turning.

This document presents the preliminary baffle layout designed for the Three Mirrors Anastigmatic (TMA) telescope developed for the Proba-V mission. This baffling is used to avoid 1^st order stray light i.e. direct stray light or through reflections on the mirrors. The stray light from the SWIR folding mirror is also studied. After these preliminary analyses the mechanical structure of the TMA is designed then verified in term of vignetting and stray light.

Regarding Earth observation missions, it has become unnecessary to point out the importance of making available wide field of view optical instruments for the purpose of spectral imaging.

Taking advantage of the pushbroom instrument concept with its linear field across the on-ground track, it is in particular relevant to consider front-end optical configurations that involve an all-reflective system presenting inherent and dedicated advantages such as achromaticity, unobscuration and compactness, while ensuring the required image quality over the whole field.

The attractiveness of the concept must be balanced with respect to the state-of-the-art mirror manufacturing technologies as the need for fast, broadband and wide field systems increases the constraints put on the feasibility of each individual component.

As part of an ESTEC contract, AMOS designed, manufactured and tested a breadboard of a four-mirror wide field telescope for typical Earth observation superspectral missions.

The initial purpose of the development was to assess the feasibility of a telecentric spaceborne three-mirror system covering an unobscured rectangular field of view of 26 degrees across track (ACT) by 6 degrees along track (ALT) with a f-number of 3.5 and a focal length of 500 mm and presenting an overall image quality better than 100 nm RMS wavefront error within the whole field.

This paper presents the mission profile as well as the optical configuration of the space-borne AsteroidFinder telescope. Its main objective is to retrieve asteroids with orbits interior to the earth’s orbit. The instrument requires high sensitivity to detect asteroids with a limiting magnitude of equal or larger than 18.5mag (V-Band) and astrometric accuracy of 1arcsec (1σ). This requires a telescope aperture greater than 400cm², high image stability, detector with high quantum efficiency (peak > 90%) and very low noise, which is only limited by zodiacal background. The telescope will observe the sky between 30° and 60° in solar elongation. The telescope optics is based on a Cook type TMA. An effective 2°×2° field of view (FOV) is achieved by a fast F/3.4 telescope with near diffraction-limited performance. The absence of centre obscuration or spiders in combination with an accessible intermediate field plane and exit pupil allow for efficient stray light mitigation. Design drivers for the telescope are the required point spread function (PSF) values, an extremely efficient stray light suppression (due to the magnitude requirement mentioned above), the detector performance, and the overall optical and mechanical stability for all orientations of the satellite. To accommodate the passive thermal stabilization scheme and the necessary structural stability, the materials selection for the telescope main structure and the mirrors are of vital importance. A focal plane with four EMCCD detectors is envisaged. The EMCCD technology features shorter integration times, which is in favor regarding the pointing performance of the satellite. The launch of the mission is foreseen for the year 2013 with a subsequent mission lifetime of at least 1 year.

High resolution Optical imaging instruments based on aperture synthesis have been developed over the last decades with the aim of reaching angular resolution in the nano radian range. These different instruments [1] take avantage of the property of the Zernike van Cittert theorem to recover the intensity distribution of the object from its spatial coherence analysis. With this method, the instrument can never select the light coming only from one of the pixels composing the full object because the measurements are being carried out on the Fourier spectral domain.

The Lunar Orbiter Laser Altimeter (LOLA) instrument [1-3] on NASA’s Lunar Reconnaissance Orbiter (LRO) mission, launched on June 18th, 2009, from Kennedy Space Center, Florida, will provide a precise global lunar topographic map using laser altimetry. LOLA will assist in the selection of landing sites on the Moon for future robotic and human exploration missions and will attempt to detect the presence of water ice on or near the surface, which is one of the objectives of NASA’s Exploration Program.

Our present knowledge of the topography of the Moon is inadequate for determining safe landing areas for NASA’s future lunar exploration missions. Only those locations, surveyed by the Apollo missions, are known with enough detail. Knowledge of the position and characteristics of the topographic features on the scale of a lunar lander are crucial for selecting safe landing sites. Our present knowledge of the rest of the lunar surface is at approximately 1 km kilometer level and in many areas, such as the lunar far side, is on the order of many kilometers. LOLA aims to rectify that and provide a precise map of the lunar surface on both the far and near side of the moon.

LOLA uses short (6 ns) pulses from a single laser through a Diffractive Optical Element (DOE) to produce a five-beam pattern that illuminates the lunar surface. For each beam, LOLA measures the time of flight (range), pulse spreading (surface roughness), and transmit/return energy (surface reflectance). LOLA will produce a high-resolution global topographic model and global geodetic framework that enables precise targeting, safe landing, and surface mobility to carry out exploratory activities. In addition, it will characterize the polar illumination environment, and image permanently shadowed regions of the lunar surface to identify possible locations of surface ice crystals in shadowed polar craters.

Trace gases and their isotopic ratios in planetary atmospheres offer important but subtle clues as to the origins of a planet's atmosphere, hydrology, geology, and potential for biology. An orbiting laser remote sensing instrument is capable of measuring trace gases on a global scale with unprecedented accuracy, and higher spatial resolution that can be obtained by passive instruments.

For Earth we have developed laser technique for the remote measurement of the tropospheric CO2, O2, and CH4 concentrations from space. Our goal is to develop a space instrument and mission approach for active CO2 measurements. Our technique uses several on and off-line wavelengths tuned to the CO2 and O2 absorption lines. This exploits the atmospheric pressure broadening of the gas lines to weigh the measurement sensitivity to the atmospheric column below 5 km and maximizes sensitivity to CO2 changes in the boundary layer where variations caused by surface sources and sinks are largest. Simultaneous measurements of O2 column use a selected region in the Oxygen A-band. Laser altimetry and atmospheric backscatter can also be measured simultaneously, which permits determining the surface height and measurements made to thick cloud tops and through aerosol layers.

We use the same technique but with a different transmitter at 1.65 um to measure methane concentrations. Methane is also a very important trace gas on earth, and a stronger greenhouse gas than CO2 on a per molecule basis. Accurate, global observations are needed in order to better understand climate change and reduce the uncertainty in the carbon budget. Although carbon dioxide is currently the primary greenhouse gas of interest, methane can have a much larger impact on climate change. Methane levels have remained relatively constant over the last decade but recent observations in the Arctic have indicated that levels may be on the rise due to permafrost thawing. NASA’s Decadal Survey underscored the importance of Methane as a greenhouse gas and called for a mission to measure CO2, CO and CH4. Methane has absorptions in the mid-infrared (3.3 um) and the near infrared (1.65 um). The 3.3 um spectral region is ideal for planetary (Mars) Methane monitoring, but unfortunately is not suitable for earth monitoring since the Methane absorption lines are severely interfered with by water. The near infra-red overtones of Methane at 1.65 um are relatively free of interference from other atmospheric species and are suitable for Earth observations. The methane instrument uses Optical Parametric Generation (OPG) along with sensitive detectors to achieve the necessary sensitivity. Our instrument generates and detects tunable laser signals in the 3.3 or 1.65 um spectral regions with different detectors in order to measure methane on Earth or Mars. For Mars, the main interest in methane is its importance as a biogenic marker.

During the previous 15 years, Rendezvous and Docking Sensors (RVS) were developed, manufactured and qualified. In the mean time they were successfully applied in some space missions: For automatic docking of the European ATV “Jules Verne” on the International Space Station in 2008; for automatic berthing of the first Japanese HTV in 2009, and even the precursor model ARP-RVS for measurements during Shuttle Atlantis flights STS-84 and STS-86 to the MIR station. Up to now, about twenty RVS Flight Models for application on ATV, HTV and the American Cygnus Spacecraft were manufactured and delivered to the respective customers.

RVS is designed for tracking of customer specific, cooperative targets (i.e. retro reflectors that are arranged in specific geometries). Once RVS has acquired the target, the sensor measures the distance to the target by timeof- flight determination of a pulsed laser beam. Any echo return provokes an interrupt signal and thus the readout of the according encoder positions of the two scan mirrors that represent Azimuth and Elevation measurement direction to the target. [2], [3].

The capability of the RVS for 3D mapping of the scene makes the fully space qualified RVS to be real 3D Lidar sensors; thus they are a sound technical base for the compact 3D Lidar breadboard that was developed in the course of the Imaging Lidar Technology (ILT) project.

The James Webb Space Telescope (JWST) is a successor of the Hubble Space Telescope and the Spitzer Space Telescope to be launched in 2015 for 5 years long (10 years goal) science mission.

The European Space Agency (ESA) is providing the Near Infrared Spectrograph (NIRSpec) developed by EADS Astrium GmbH to fly on the James Webb Space Telescope (JWST). NIRSpec covers the 0.6-5.0 µm domain. It will be primarily operated as a multi-object spectrograph, using a MEMS micro-shutter array (MSA) provided by NASA to select multiple objects from the field of view at an intermediate image plane formed by the NIRSpec fore-optics. The MSA apertures form multiple entrance slits of the spectrometer section.

EUCLID, a medium-class mission candidate of ESA's Cosmic Vision 2015–2025 Program, currently in Definition Phase (Phase A/B1), shall map the geometry of the Dark Universe by investigating dark matter distributions, the distance-redshift relationship, and the evolution of cosmic structures. EUCLID consists of a 1.2 m telescope and two scientific instruments for ellipticity and redshift measurements in the visible and nearinfrared wavelength regime. We present a design concept of the EUCLID mission which is fully compliant with the mission requirements. Preliminary concepts of the spacecraft and of the payload including the scientific instruments are discussed.

TNO has developed, built and tested the Wave Front Sensor (WFS) for ESA’s Gaia mission. The WFS will help Gaia create an extraordinarily precise three-dimensional map of more than one billion stars in our Galaxy. Part of ESA's Cosmic Vision programme, Gaia’s build is led by EADS Astrium and is scheduled for launch in 2012. The Wave Front Sensor will be used to monitor the wave front errors of the two main telescopes mounted on the GAIA satellite. These mirrors include a 5-degree of freedom (DOF) mechanism that can be used to minimize the wave front errors during operation. The GAIA-WFS will operate over a broad wavelength (450 to 900 nm) and under cryogenic conditions (130 to 200 K operation temperature). The WFS uses an all reflective, a-thermal design and is of the type of Shack-Hartmann. The boundary condition for the design is that the focal plane of the WFS is the same plane as the focal plane of the GAIA telescopes. The spot pattern generated after a micro lens array ( MLA) by a star is compared to the pattern of one of the three calibration sources that is included in the WFS, allowing in flight calibration. We show the robust and lightweight opto mechanical design that is optimised for launch and cryogenic operation. Details are given on its alignment and commissioning. The WFS is able to measure relative wave front distortions in the order of lambda/1000, and can determine the optimum position of the focal plane with an accuracy of 50 μm

Future NASA exploratory missions to the Moon and Mars will require safe soft-landings at the designated sites with a high degree of precision. These sites may include areas of high scientific value with relatively rough terrain with little or no solar illumination and possibly areas near pre-deployed assets. The ability of lidar technology to provide three-dimensional elevation maps of the terrain, high precision distance to the ground, and approach velocity can enable safe landing of large robotic and manned vehicles with a high degree of precision. Currently, NASA-LaRC is developing novel lidar sensors aimed at meeting NASA’s objectives for future planetary landing missions under the Autonomous Landing and Hazard Avoidance (ALHAT) project [1]. These lidar sensors are 3-Dimensional Imaging Flash Lidar, Doppler Lidar, and Laser Altimeter. The Flash Lidar is capable of generating elevation maps of the terrain identifying hazardous features such as rocks, craters, and steep slopes. The elevation maps collected during the approach phase between 1000 m to 500 m above the ground can be used to determine the most suitable safe landing site. The Doppler Lidar provides highly accurate ground velocity and distance data allowing for precision navigation to the selected landing site. Prior to the approach phase at altitudes of over 15 km, the Laser Altimeter can provide sufficient data for updating the vehicle position and attitude data from the Inertial Measurement Unit. At these higher altitudes, either the Laser Altimeter or the Flash Lidar can be used for generating a contour map of the terrain below for identifying known surface features such as craters for further reducing the vehicle relative position error.

A near-flight prototype of a pulsed UV laser has been developed for the Mars Organic Molecule Analyzer (MOMA) of the ExoMars mission. The laser head is based on a Nd:YAG oscillator with subsequent frequency quadrupling and emits nanosecond pulses with an energy of > 300 μJ at a wavelength of 266 nm. The design is compact and lightweight. Tests in relevant environment regarding temperature, vibration, and radiation have been performed.

Laser applications become more and more interesting in contemporary missions such as earth observations or optical communication in space.

One of these applications is light detection and ranging (LIDAR), which comprises huge scientific potential in future missions. The Nd:YAG solid-state laser of such a LIDAR system is optically pumped using 808nm emitting pump sources based on semiconductor laser-diodes in quasi-continuous wave (qcw) operation. Therefore reliable and efficient laser diodes with increased output powers are an important requirement for a spaceborne LIDAR-system. In the past, many tests were performed regarding the performance and life-time of such laser-diodes. There were also studies for spaceborne applications, but a test with long operation times at high powers and statistical relevance is pending.

Other applications, such as science packages (e.g. Raman-spectroscopy) on planetary rovers require also reliable high-power light sources. Typically fiber-coupled laser diode modules are used for such applications. Besides high reliability and life-time, designs compatible to the harsh environmental conditions must be taken in account. Mechanical loads, such as shock or strong vibration are expected due to take-off or landing procedures. Many temperature cycles with high change rates and differences must be taken in account due to sun-shadow effects in planetary orbits. Cosmic radiation has strong impact on optical components and must also be taken in account. Last, a hermetic sealing must be considered, since vacuum can have disadvantageous effects on optoelectronics components.

The Absolute Scalar Magnetometer (ASM) developed by CEA-LETI/CNES is an optically pumped ⁴He magnetic field sensor based on the Zeeman effect and an electronic magnetic resonance whose effects are amplified by a laser pumping process [1-2]. Consequently, the role of the laser is to pump the ⁴He atoms at the D₀ transition as well as to allow the magnetic resonance signal detection.

The ASM will be the scalar magnetic reference instrument of the three ESA Swarm satellites to be launched in 2012 in order to carry out the best ever survey of the Earth magnetic field and its temporal evolution. The sensitivity and accuracy of this magnetometer based on ⁴He optical pumping depend directly on the characteristics of its light source, which is the key sub-system of the sensor.

We describe in this paper the selected fiber laser architecture and its wavelength stabilization scheme. Its main performance in terms of spectral emission, optical power at 1083 nm and intensity noise characteristics in the frequency bands used for the operation of the magnetometer, are then presented. Environmental testing results (thermal vacuum cycling, vibrations, shocks and ageing) are also reported at the end of this paper.

An extended qualification program has proven the quality of a previously evaluated semiconductor laser diode, which is intended to be used in a subsystem for the GAIA mission. We report on results of several reliability tests performed in subgroups. The requirements of the procurement specification with respect to reliability and desired manufacturing processes were confirmed. This is an example for successful collaboration between component supplier, system integrator and payload responsible party.

The next Japanese solar mission, SOLAR-C, which has been envisaged after successful science operation of Hinode (SOLAR-B) mission, is perusing two plans: plan-A and plan-B, and under extensive study from science objectives as well as engineering point of view. The plan-A aims at performing out-of-ecliptic observations for investigating, with helioseismic approach, internal structure and dynamo mechanisms of the Sun. It also explores polar regions where fast solar wind is believed to originate. A baseline orbit for plan-A is a circular orbit of 1 AU distance from the Sun with its inclination at around or greater than 40 degrees. The plan-B aims to study small-scale plasma processes and structures in the solar atmosphere which attract researchers’ growing interest, followed by many Hinode discoveries [1], for understanding fully dynamism and magnetic nature of the atmosphere. With plan-B, high-angular-resolution investigation of the entire solar atmosphere (from the photosphere to the corona, including their interface layers, i.e., chromosphere and transition region) is to be performed with enhanced spectroscopic and spectro-polarimetric capability as compared with Hinode, together with enhanced sensitivity towards ultra-violet wavelengths. The orbit of plan-B is either a solar synchronous polar orbit of altitude around 600 km or a geosynchronous orbit to ensure continuous solar observations. After the decision of any one of the two plans, the SOLAR-C will be proposed for launch in mid-2010s.

In this paper, we will present a basic design of one of major planned instrumental payload for the plan-B: the Solar Ultra-violet Visible and near IR observing Telescope (hereafter referred to as SUVIT). The basic concept in designing the SUVIT is to utilize as much as possible a heritage of successful telescope of the Solar Optical Telescope (SOT) aboard Hinode [2]. Major differences of SUVIT from SOT are the three times larger aperture of 1.5 m, which enables to collect one order of magnitude more photons than SOT, relatively shorter telescope length of 2.8 m to accommodate a launcher’s nosecone size for possible dual-satellite-launch configuration, and much wider observing wavelength from UV (down to 250 nm) through near IR (up to 1100 nm). The large aperture is essentially important to attain scientific goals of the plan-B, especially for accurate diagnostics of the dynamic solar chromosphere as revealed by Hinode, although this make it difficult to design the telescope because of ten times more solar heat load introduced into the telescope. The SUVIT consists of two optically separable components; the telescope assembly (TA) and an accompanying focal plane package equipped with filtergraphs and spectrographs. Opto-mechanical and -thermal performance of the TA is crucial to attain high-quality solar observations and here we present a status of feasible study in its optical and thermal designing for diffraction-limited performance at visible wavelength in a reasonably wide field of view.

The PROBA2 mission has been launched on 2nd November2009 with a Rockot launcher to a Sunsynchronous orbit at an altitude of 725 km. Its nominal operation duration is two years with possible extension of 2 years. PROBA2 is a small satellite developed under an ESA General Support Technology Program (GSTP) contract to perform an in-flight demonstration of new space technologies and support a scientific mission for a set of selected instruments. The mission is tracked by the ESA Redu Mission Operation Center.

Classical externally-occulted coronagraphs are presently limited in their performances by the distance between the external occulter and the front objective. The diffraction fringe from the occulter and the vignetted pupil which degrades the spatial resolution prevent useful observations of the white light corona inside typically 2-2.5 solar radii (Rsun). Formation flying offers and elegant solution to these limitations and allows conceiving giant, externally-occulted coronagraphs using a two-component space system with the external occulter on one spacecraft and the optical instrument on the other spacecraft at a distance of hundred meters [1, 2]. Such an instrument ASPIICS (Association de Satellites Pour l’Imagerie et l’Interférométrie de la Couronne Solaire) has been selected by the European Space Agency (ESA) to fly on its PROBA-3 mission of formation flying demonstration which is presently in phase B (Fig. 1). The classical design of an externally-occulted coronagraph is adapted to the formation flying configuration allowing the detection of the very inner corona as close as ∼0.04 solar radii from the solar limb. By tuning the position of the occulter spacecraft, it may even be possible to reach the chromosphere and the upper part of the spicules [3]. ASPIICS will perform (i) high spatial resolution imaging of the continuum K+F corona in photometric and polarimetric modes, (ii) high spatial resolution imaging of the E-corona in two coronal emission lines (CEL): Fe XIV and He I D3, and (iii) two-dimensional spectrophotometry of the Fe XIV emission line. ASPIICS will address the question of the coronal heating and the role of waves by characterizing propagating fluctuations (waves and turbulence) in the solar wind acceleration region and by looking for oscillations in the intensity and Doppler shift of spectral lines. The combined imaging and spectral diagnostics capabilities available with ASPIICS will allow mapping the velocity field of the corona both in the sky plane (directly on the images) and along the line-of-sight by measuring the Doppler shifts of emission lines in an effort to determine how the different components of the solar wind, slow and fast are accelerated. With a possible launch in 2014, ASPIICS will observe the corona during the maximum of solar activity, insuring the detection of many Coronal Mass Ejections (CMEs). By rapidly alternating high-resolution imaging and spectroscopy, CMEs will be thoroughly characterized.

New opportunities are emerging for the implementation of hardware sub-systems based on OptoElectronic Devices (OED) for space application. Since the end of this decade the main players for space systems namely designers and users including Industries, Agencies, Manufacturers and Laboratories are strongly demanding of adequate strategies to qualify and validate new optoelectronics products and sub-systems [1]. The long term space application mission will require to address either inter-satellite link (free space communication, positioning systems, tracking) or intra-satellite connectivity/flexibility/reconfigurability or high volume of data transfer between equipment installed into payload.

Optical technology developed for ground and submarine telecommunications is becoming of strong interest for next generation satellites. In addition to inter-satellite laser communications and LIDAR’s, new applications are being considered such as on-board distribution and processing of microwave signals, fiber sensors or gyroscopes as well. Whereas common optical / optoelectronic components are known to be weakly sensitive to radiations, the essential optical amplifiers are strongly degraded in such an environment because of the RIA (Radio-Induced-Absorption) experienced by the Erbium-Doped Fiber (EDF) itself [1-3]. This degradation is mainly caused by the presence of co-doping ions, such as Aluminium or Germanium, inserted in the fibre to assist the inclusion of the Erbium ions in the silica matrix or to provide to the optical fibre its guiding properties.

Although it has been recently postponed due to high cost and risks, nulling interferometry in space remains one of the very few direct detection methods able to characterize extrasolar planets and particularly telluric ones. Within this framework, several projects such as DARWIN [1], [2], TPF-I [3], [4], FKSI [5] or PEGASE [6], [7], have been proposed in the past years. Most of them are based on a free flying concept. It allows firstly to avoid atmosphere turbulence, and secondly to distribute instrumental function over many satellites flying in close formation. In this way, a very high angular resolution can be achieved with an acceptable launch mass. But the price to pay is to very precisely position and stabilize relatively the spacecrafts, in order to achieve a deep and stable extinction of the star. Understanding and mastering all these requirements are great challenges and key issues towards the feasibility of these missions. Thus, we decided to experimentally study this question and focus on some possible simplifications of the concept.

Since 2006, PERSEE (PEGASE Experiment for Research and Stabilization of Extreme Extinction) laboratory test bench is under development by a consortium composed of Centre National d’Etudes Spatiales (CNES), Institut d’Astrophysique Spatiale (IAS), Observatoire de Paris-Meudon (LESIA), Observatoire de la Côte d’Azur (OCA), Office National d’Etudes et de Recherches Aérospatiales (ONERA), and Thalès Alénia Space (TAS) [8]. It is mainly funded by CNES R&D. PERSEE couples an infrared wide band nulling interferometer with local OPD and tip/tilt control loops and a free flying Guidance Navigation and Control (GNC) simulator able to introduce realistic disturbances. Although it was designed in the framework of the PEGASE free flying space mission, PERSEE can adapt very easily to other contexts like FKSI (in space, with a 10 m long beam structure) or ALADDIN [9] (on ground, in Antarctica) because the optical designs of all those missions are very similar. After a short description of the experimental setup, we will present first the results obtained in an intermediate configuration with monochromatic light. Then we will present some preliminary results with polychromatic light. Last, we discuss some very first more general lessons we can already learn from this experiment.

Wavefront filtering is mandatory in the realisation of nulling interferometers with high star light suppression capability required to detect extrasolar planets, such as the one foreseen for the ESA Darwin mission.

This paper presents the design, manufacturing, and test results of single mode fibres to be used as wavefront filters in mid-infrared range. Fibres made from chalcogenide glass and silver halide crystals were produced. The first class can serve as wavefront filters up to a wavelength of 11 microns, while silver halide fibres can be used over the full Darwin wavelength range from 6.5 to 18 micron. The chalcogenide glass fibres were drawn by double crucible method whereas polycrystalline fibres from silver halides were fabricated by multiple extrusion from a crystalline preform.

Multi-layer AR-coatings for fibre ends were developed and environmentally tested for both types of fibres. Special fibre facet polishing procedures were established, in particular for the soft silver halide fibre ends. Cable design and assembly process were also developed, including termination by SMA-connectors with ceramic ferrules and fibre protection by loose PEEK-tubings to prevent excessive bending and chemical attacks for fibres.

The wavefront filtering capability of the fibres was demonstrated on a high quality Mach-Zehnder interferometer. Two different groups of laser sources were used to measure the wavefront filtering of the fibres by using a CO-laser for testing in the lower sub-band and a CO₂-laser to check the upper sub-band.

Measurements of the fibres far field intensity distribution and transmission were performed for numerous cable samples. Single mode behaviour was observed in more than 25 silver halide fibre cables before AR-coating of their ends, while after that 17 cables were compliant with all technical requirements. Residual cladding modes existing in short single mode fibres were effectively removed by applying of a proper absorbing jacket to the fibre's lateral surface and by adding an oversized output aperture in front of fibre ends.

Several fibres were exposed to gamma radiation of total dose of 25, 50, and even 500 krad. No deterioration was found on AR-coated fibre ends and on fibre material. Five fibres were irradiated by proton radiation of 10MeV energy and 10¹⁰ p/cm² equivalent fluence. Several fibres were cooled down to 10 K by plunging them in a dipstick into liquid Helium. Silver halide fibres survived that test when cables were properly assembled. The brittle chalcogenide glass fibres were much more sensitive to thermal gradients and the related cables did not survive the thermal shock.

Critical issues have been revealed in multiple drawings of chalcogenide glass fibres where core and cladding composition were not stable at some fabrication stages - resulting in a poor single mode guiding. Much better results have been achieved with polycrystalline fibres from silver halides made with a small core and low NA and enabling single mode guiding in the mid infrared.

The search for Earth-like exoplanets, orbiting in the habitable zone of stars other than our Sun and showing biological activity, is one of the most exciting and challenging quests of the present time. Nulling interferometry from space, in the thermal infrared, appears as a promising candidate technique for the task of directly observing extra-solar planets. It has been studied for about 10 years by ESA and NASA in the framework of the Darwin and TPF-I missions respectively [1].

Nevertheless, nulling interferometry in the thermal infrared remains a technological challenge at several levels. Among them, the development of the "modal filter" function is mandatory for the filtering of the wavefronts in adequacy with the objective of rejecting the central star flux to an efficiency of about 10⁵. Modal filtering [2] takes benefit of the capability of single-mode waveguides to transmit a single amplitude function, to eliminate virtually any perturbation of the interfering wavefronts, thus making very high rejection ratios possible.

The modal filter may either be based on single-mode Integrated Optics (IO) and/or Fiber Optics. In this paper, we focus on IO, and more specifically on the progress of the on-going “Integrated Optics” activity of the European Space Agency.

Launched on November 2^nd, 2009, SMOS (Soil Moisture, Ocean Salinity) is the second Earth Explorer Opportunity mission developed as part of ESA’s Living Planet Programme. It demonstrates a completely new type of instrument – a large, deployable synthetic-aperture microwave radiometer [1].

RUAG Space, Switzerland, as a subcontractor of EADS Astrium, Spain, has provided the instrument’s fibreoptic harness, which interconnects the central data processor with all 69 microwave receivers, as well as 12 auxiliary units on board. For reasons explained in Section 3, SMOS is the first European mission extensively using both fibre-optic clock distribution and data transmission in space.

In Section 2, we present an overview of the scientific goals of SMOS, and describe the payload’s basic function. There from we derive the rationale and the design of the fibre-optic harness (Section 3). In Section 4 all development, manufacturing, and test activities are summarised, which culminated in the successful delivery of all flight units to EADS Astrium by October 2006. We present the major test results obtained with the flight harness (Section 5), and conclude with a short summary of the higher-level activities, which lead to successful launch and commissioning of the SMOS satellite (Section 6).

For the next generation of HighThroughPut (HTP) Telecommunications Satellites, space end users’ needs will result in higher link speeds and an increase in the number of channels; up to 512 channels running at 10Gbits/s. By keeping electrical interconnections based on copper, the constraints in term of power dissipation, number of electrical wires and signal integrity will become too demanding. The replacement of the electrical links by optical links is the most adapted solution as it provides high speed links with low power consumption and no EMC/EMI.

But replacing all electrical links by optical links of an On Board Payload (OBP) is challenging. It is not simply a matter of replacing electrical components with optical but rather the whole concept and architecture have to be rethought to achieve a high reliability and high performance optical solution. In this context, this paper will present the concept of an Innovative OBP Optical Architecture.

The optical architecture was defined to meet the critical requirements of the application: signal speed, number of channels, space reliability, power dissipation, optical signals crossing and components availability. The resulting architecture is challenging and the need for new developments is highlighted. But this innovative optically interconnected architecture will substantially outperform standard electrical ones.

The flexibility required for future telecom payloads will require of more digital processing capabilities, moving from conventional analogue repeaters to more advanced and efficient analog subsystems or DSPbased solutions. Aggregate data throughputs will have to be handled onboard, creating the need for effective, ADC/DSP and DSP/DAC high speed links. Broadband payloads will have to receive, route and retransmit hundreds of channels and need to be designed so as to meet such requirements of larger bandwidth, system transparency and flexibility.[1][2]

One important device in these new architectures is analog to digital converter (ADC) and its equivalent digital to analog converter (DAC). These will be the in/out interface for the use of digital processing in order to provide flexible beam to beam connectivity and variable bandwidth allocation.

For telecom payloads having a large number of feeds and thus a large number of converters the mass and consumption of the mixer stage has become significant. Moreover, the inclusion of ADCs in the payload presents new trade-offs in design (jitter, quantization noise, ambiguity). This paper deals with an alternative solution of these two main problems with the exploitation of photonic techniques.

NASA and other national agencies ask the National Research Council (NRC) once every decade to look out ten or more years into the future and prioritize research areas, observations, and notional missions to make those observations. The latest such scientific community consultation referred to as the Decadal Survey (DS), was completed in 2007 [1]. DS thematic panels developed 35 missions from more than 100 missions proposed, from which the DS Executive Committee synthesized 17 missions, with suggested order presented in three time-phased blocks. The first block with aim for near term launch (2010-2013) included four missions. The Climate Absolute Radiance and Refractivity Observatory (CLARREO) mission is one of them.

The CLARREO mission was classified as a Small Mission to be contained in a 300 M US$ budgetary envelope. CLARREO will provide a benchmark climate record that is global, accurate in perpetuity, tested against independent strategies that reveal systematic errors, and pinned to international standards. The long term objective thus suggests that NOAA or NASA will fly the CLARREO instrument suite on an operational basis following the first scientific experiment

The CLARREO missions will conduct the following observations:

1. Absolute spectrally-resolved measurements of terrestrial thermal emission with an absolute accuracy of 0.1 K in brightness temperature (3σ or 99% confidence limits.) The measurements should cover most of the thermal spectrum.

2. Absolute spectrally-resolved measurements of the solar radiation reflected from Earth. The measurements should cover the part of the solar spectrum most important to climate, including the near-ultraviolet, visible, and near-infrared.

3. Independent measurements of atmospheric temperature, pressure, and humidity using Global Positioning System (GPS) occultation measurements of atmospheric refraction.

4. Serve as a high accuracy calibration standard for use by the broadband CERES instruments on-orbit.

Following the DS conclusion, and considering the early development stage of the mission, NASA funded three Instrument Incubator Programs (IIP) to push instrument concepts to a higher level of maturity. A joint proposal between University of Wisconsin (UW) and Harvard University was selected to address the first above objective and part of the fourth one in the corresponding spectral region. In order to achieve this goal, four complementary technologies are to be developed [2]:

(1) On-orbit Absolute Radiance Standard (OARS), a high emissivity blackbody source that uses multiple miniature phase-change cells to provide a revolutionary on-orbit standard with absolute temperature accuracy proven over a wide range of temperatures.

(2) On-orbit Cavity Emissivity Modules (OCEMs), providing a source (quantum cascade laser, QCL, or “Heated Halo”) to measure any change in the cavity emissivity of the OARS.

(3) On-orbit Spectral Response Module (OSRM), a source for spectral response measurements using a nearly monochromatic QCL source configured to uniformly fill the sensor field-of-view.

(4) Dual Absolute Radiance Interferometers (DARI), providing spectral coverage from 3.3 to 50 μm that can be inter-compared to dissect any unexpected systematic errors in overlapping spectral regions.

ABB's GFI (Generic Flight Interferometer) has been selected as the favoured architecture for the DARI, mainly due to the maturity of the design and its space heritage. A GFI with commercial grade components was optimised for the selected spectral range. The architecture of the GFI will ensure a high response stability between calibrations.

A novel hyperspectral imaging technique based on Fourier Transform analysis applied to a low finesse scanning Fabry-Perot (F-P) interferometer has been demonstrated in the visible and NIR regions. The technique allows the realization of a lightweight and compact instrument yet having high throughput with respect to classical instruments based on dispersive means. Experiments carried out in the visible region have demonstrated hyperspectral imaging capability with a spectral resolution of 2 nm @ 532 nm and an image resolution limited by the CCD (696x512 pixel). In the NIR (0,9-1,7 μm) region a spectral resolution of 8 nm @1064nm and an image resolution limited by the CCD (320x256 pixel) has been obtained. The potentiality in spectroscopic applications like remote gas sensing has been demonstrated as well as accurate thermal imaging capabilities. Unlike classical hyperspectral instruments, based on dispersive means or on tuneable band-pass filters, the efficiency of our F-P based system is very high (about 30% of the photons collected by the objective reach the CCD) allowing much faster and/or better quality hyperspectral images. In the present experiments the speed was by far limited by the acquisition speed of the CCD sensors. Furthermore the device is very compact and is placed between the objective and the CCD of a standard imaging system: in this configuration the field of view of the instrument is only limited by the same objective that in the present system is interchangeable. Because of its roughness, compactness, lightweight and luminous efficiency, the device is a good candidate for airborne or space borne hyperspectral applications.

PHEBUS (Probing of Hermean Exosphere by Ultraviolet Spectroscopy) is a double spectrometer for the Extreme Ultraviolet range (55-155 nm) and the Far Ultraviolet range (145-315 nm) dedicated to the characterization of Mercury’s exosphere composition and dynamics, and surface-exosphere connections. PHEBUS is part of the ESA BepiColombo cornerstone mission payload devoted to the study of Mercury. The BepiColombo mission consists of two spacecrafts: the Mercury Magnetospheric Orbiter (MMO) and the Mercury Planetary Orbiter (MPO) on which PHEBUS will be mounted. PHEBUS is a French-led instrument implemented in a cooperative scheme involving Japan (detectors), Russia (scanner) and Italy (ground calibration). Before launch, PHEBUS team want to perform a full absolute calibration on ground, in addition to calibrations which will be made in-flight, in order to know the instrument’s response as precisely as possible. Instrument overview and calibration philosophy are introduced along with the first lights results observed by a first prototype.

The coupling between imaging and spectrometry has proved to be one of the most promising way to study remotely planetary objects [1][2]. The next step is to use this concept for in situ analyses. MicrOmega IR has been developed within this scope. It is an ultra miniaturized near-infrared hyperspectral microscope dedicated to in situ analyses, selected to be part of the ESA/ExoMars rover and RKA/Phobos Grunt lander payload. The goal of this instrument is to characterize the composition of samples at almost their grain size scale, in a nondestructive way. Coupled to the mapping information, it provides unique clues to trace back the history of the parent body (planet, satellite or small body) [3][4].

Fiber Sensor Demonstrator (FSD) developed by MPB Communications (MPBC) is the first demonstration of a full fiber-optic sensor network in the space environment on a satellite. FSD has been launched on ESA’s Proba-2 satellite in November 2009. FSD contains twelve temperature sensors to measure the temperature at different locations in the satellite, and one High-Temperature sensor to measure the transient high temperature in the thruster, as well as one pressure sensor to measure the xenon tank pressure. First set of on-orbit test data were obtained in January 2010. The FSD unit successfully established the communication with Proba-2. The temperature of FSD unit was also acquired through a AD590 sensor inside the unit. The measurements of all the optical fiber sensor lines will be evaluated after the testing results obtained. The FSD contains twelve specially-packaged FBG temperature sensors to measure the temperature at different locations in the propulsion system and the spacecraft over the range of −60°C to +120°C. A high-temperature sensor is provided to measure the transient temperature response of the thruster to beyond 350°C. There is also an innovative P/T sensor that provides both temperature and pressure measurements of the Xe propellant tank. The preliminary data of on-orbit functional testing and temperature measurements are provided mainly in Section 6.

Monitoring needs of spacecraft are rapidly increasing due to new and more challenging missions, along with demands to reduce launching costs by minimizing the manufacture, assembly, integration and test time and employing new low weight materials balanced by the need for maximizing system lifetime while maintaining good reliability. Conventional electronic sensors are characterized by their low multiplexing capability and their EMI/RF susceptibility and it is in this scenario that Fiber Optic Sensors (FOS) in general, and more specifically Fiber Bragg Grating (FBG) technology offers important benefits, improving in various ways the already deployed sensing subsystems (e.g. reducing the weight associated with sensor cabling, increasing the number of sensing points) and enabling new monitoring applications that were not possible by using conventional sensing technologies.

This work presents the activities performed and the lessons learnt in the frame of ESA’s ARTES-5 project “Fiber Optic Sensing Subsystem for Spacecraft Health Monitoring in Telecommunication Satellites”. This project finished in July 2009, with the implementation and testing of two different demonstrators employing FBG sensor technology: FBG sensors for temperature monitoring in high voltage environments, and in particular in several parts of electric propulsion subsystems [1], and FBG sensors for thermal monitoring of array-antennas during RF testing [2].

In addition, the contacts performed with different actors within the space community allowed the identification of a special area of interest for the substitution of regular thermocouple instrumentation by FBG technology for thermal vacuum ground testing of satellites.

We present our work about the development of two complementary interrogation schemes based on fiber optic sensing for the use of structural and thermal monitoring of Ariane launchers. The advantages of fiber optic sensing in particular light-weight, immunity to electromagnetic interferences and the possibility of sensor distribution along optical fibers are driving factors for utilization of this technology in space crafts [1]. The edge-filter (EF) and scanning-laser (SL) interrogators for determination of the mean wavelength of fiber Bragg grating (FBG) sensors have been implemented as two separate demonstrators. Within this paper we describe the functional principles of both interrogators. Furthermore we present test results where the developed systems have been used for readout of FBG sensors which are implemented in an Ariane structural demonstrator during thermal, thermal-vacuum and vibration tests. Functionality of both systems is demonstrated and their potential for further development towards space qualified systems is shown.

Since the performance characteristics of the two systems are different from each other, they are dedicated for different sensing applications on a launcher. The EF sensor interrogator provides a sample rate of 20 kHz at a number of 4 connected sensors and supports parallel readout and aliasing free operation. Therefore it is best suited for high priority measurement. Structural monitoring which requires the acquisition of real time sensor information in order to support control of the launcher is one operation area for a future EF system. The SL interrogator provides an overall measurement rate of 1 kHz at a number of 24 connected sensors distributed on three sensor channels. It can be adapted to any sensors that have design wavelengths lying within the output spectrum of the laser diode. Furthermore the number of overall sensors to be read out with this system can be adapted easily. Thermal mapping of satellite panels is one possible future application for the SL interrogator.

ChemCam is an instrument suite on the Mars Science Laboratory (MSL) mission that will launch to Mars in 2011. MSL is a rover-type lander that is capable of exploring large territories over the mission lifetime and includes a number of instruments for analysing rocks and soil. ChemCam includes a laser induced breakdown spectroscopy (LIBS) [1] instrument that samples the surface chemistry of target objects within about 10 m of the rover without having to physically move to the target to obtain emission spectra in the 240 nm to 800 nm range. The ChemCam laser and sensing telescope are mounted on the rover Remote Sensing Mast (RSM) and have 360 degrees of azimuthal range, and 180 degrees of vertical range, allowing sampling of any object within range and line-of-sight of the mast top. This capability can be used to select targets for further analysis by other MSL instruments.

The LIBS portion of ChemCam is split between the top of the RSM and inside the rover body. The laser and the telescope are located atop the mast and rotate to select and observe targets. The three spectrometers (UV, VIS, and NIR) are located inside the rover body, along with a demultiplexer (demux) that splits the signal into the three bands. The signal from the telescope is transmitted to the demux by the fiber optic cable that is the subject of this paper. The fiber optic cable (FOC) is a single 5.7-m long, broadband, mult-mode fiber that connects the telescope and demux and is exposed to the full martian environment in some places and subjected to significant temperature gradients as it runs from interior areas to exterior areas.

Laser spectral analysis systems are increasingly being considered for in situ analysis of the atomic and molecular composition of selected rock and soil samples on other planets [1][2][3]. Both Laser Induced Breakdown Spectroscopy (LIBS) and Raman spectroscopy are used to identify the constituents of soil and rock samples in situ. LIBS instruments use a high peak-power laser to ablate a minute area of the surface of a sample. The resulting plasma is observed with an optical head, which collects the emitted light for analysis by one or more spectrometers. By identifying the ion emission lines observed in the plasma, the constituent elements and their abundance can be deduced. In Raman spectroscopy, laser photons incident on the sample surface are scattered and experience a Raman shift, exchanging small amounts of energy with the molecules scattering the light. By observing the spectrum of the scattered light, it is possible to determine the molecular composition of the sample.

For both types of instruments, there are advantages to physically separating the light collecting optics from the spectroscopy optics. The light collection system will often have articulating or rotating elements to facilitate the interrogation of multiple samples with minimum expenditure of energy and motion. As such, the optical head is often placed on a boom or an appendage allowing it to be pointed in different directions or easily positioned in different locations. By contrast, the spectrometry portion of the instrument is often well-served by placing it in a more static location. The detectors often operate more consistently in a thermally-controlled environment. Placing them deep within the spacecraft structure also provides some shielding from ionizing radiation, extending the instrument’s useful life. Finally, the spectrometry portion of the instrument often contains significant mass, such that keeping it off of the moving portion of the platform, allowing that portion to be significantly smaller, less massive and less robust.

Large core multi-mode optical fibers are often used to accommodate the optical connection of the two separated portions of such instrumentation. In some cases, significant throughput efficiency improvement can be realized by judiciously orienting the strands of multi-fiber cable, close-bunching them to accommodate a tight focus of the optical system on the optical side of the connection, and splaying them out linearly along a spectrometer slit on the other end.

For such instrumentation to work effectively in identifying elements and molecules, and especially to produce accurate quantitative results, the spectral throughput of the optical fiber connection must be consistent over varying temperatures, over the range of motion of the optical head (and it’s implied optical cable stresses), and over angle-aperture invariant of the total system. While the first two of these conditions have been demonstrated[4], spectral observations of the latter present a cause for concern, and may have an impact on future design of fiber-connected LIBS and Raman spectroscopy instruments. In short, we have observed that the shape of the spectral efficiency curve of a large multi-mode core optical fiber changes as a function of input angle.

The size and the weight of state of the art spectrometers is a serious issue regarding space applications. SWIFTS (Stationary Wave Integrated Fourier Transform Spectrometer) is a new FTS family without any moving part. This very promising technology is an original way to fully sample the Fourier interferogram obtained in a waveguide by either a reflection (SWIFTS Lippmann) or counter-propagative (SWIFTS Gabor) interference phenomenon. The sampling is simultaneously performed the optical path thanks to "nano-detectors" located in the evanescent field of the waveguide. For instance a 1.7cm long waveguide properly associated to the detector achieves directly a resolution of 0.13cm-1 on a few centimetre long instruments. Here, firstly we present the development status of this new kind of spectrometers and the first results obtained with on going development of spectrometer covering simultaneously the visible domain from 400 to 1000 nm like an Echelle spectrometer. Valuable technologies allows one to extend the concept to various wavelength domains. Secondly, we present the results obtained in the frame of an activity funded by the European Space Agency where several potential applications in space missions have been identified and studied.

We present SPEX, the Spectropolarimeter for Planetary Exploration, which is a compact, robust and low-mass spectropolarimeter designed to operate from an orbiting or in situ platform. Its purpose is to simultaneously measure the radiance and the state (degree and angle) of linear polarization of sunlight that has been scattered in a planetary atmosphere and/or reflected by a planetary surface with high accuracy. The degree of linear polarization is extremely sensitive to the microphysical properties of atmospheric or surface particles (such as size, shape, and composition), and to the vertical distribution of atmospheric particles, such as cloud top altitudes. Measurements as those performed by SPEX are therefore crucial and often the only tool for disentangling the many parameters that describe planetary atmospheres and surfaces. SPEX uses a novel, passive method for its radiance and polarization observations that is based on a carefully selected combination of polarization optics. This method, called spectral modulation, is the modulation of the radiance spectrum in both amplitude and phase by the degree and angle of linear polarization, respectively. The polarization optics consists of an achromatic quarter-wave retarder, an athermal multiple-order retarder, and a polarizing beam splitter. We will show first results obtained with the recently developed prototype of the SPEX instrument, and present a performance analysis based on a dedicated vector radiative transport model together with a recently developed SPEX instrument simulator.

A new concept for an imaging spectrograph operating at grazing incidence and stigmatic in a large field-of-view is presented. It can acquire both spectral and spatial information at the instrumental focal plane. The design is based on a crossed telescope and on a spherical variable-line-spaced grating. A laboratory prototype working in the spectral range 4–20 nm is presented.

The objective of this report is to present the study performed for a specially demanding solution of a spectrometer [1] design based on a transmissive holographic grating especially designed to actuate as the dispersion element. The main driver of the design has been to obtain a device with a clear reduction in mass, power and mechanical envelope with respect to the previous configuration based on a prism and/or Echelle grating. This simplification is produced mainly at expense of the waveband range. This study has been carried out by Laboratorio de Instrumentación Espacial (LINES) optical designers from Instituto Nacional de Técnica Aeroespacial (INTA).

The increasing demand on high speed communication networks has stimulated the development of optical free space data transmission during the last years. TESAT has developed a laser communication terminal (LCT) that fulfills the need of a power efficient system whose capability has been successfully demonstrated at bidirectional space-to-space links and bidirectional space-to-ground links (SGLs) at a data rate of 5.625 GBit/s with a homodyne detection scheme and a BPSK modulation format. In comparison to a direct detection system, the homodyne detection scheme works as a bandpass filter. The transmission is immune to false light and even data transmission with the sun in the receiver field of view (FOV) is possible.

Compared to common RF transmission which is implemented on spacecrafts for data transmission, optical transmission provides not only higher transmission rates (factor 10) but also shows excellent security features since the laser beams directivity making it immune to interception.

In-situ interplanetary science missions constantly push the spacecraft communications systems to support successively higher downlink rates. However, the highly restrictive mass and power constraints placed on interplanetary spacecraft significantly limit the desired bandwidth increases in going forward with current radio frequency (RF) technology. To overcome these limitations, we have evaluated the ability of free-space optical communications systems to make substantial gains in downlink bandwidth, while holding to the mass and power limits allocated to current state-of-the-art Ka-band communications systems.

A primary component of such an optical communications system is the optical assembly, comprised of the optical support structure, optical elements, baffles and outer enclosure. We wish to estimate the total mass that such an optical assembly might require, and assess what form it might take. Finally, to ground this generalized study, we should produce a conceptual design, and use that to verify its ability to achieve the required downlink gain, estimate it’s specific optical and opto-mechanical requirements, and evaluate the feasibility of producing the assembly.

Pyrotechnics is in widespread use in both launchers and satellites. The devices are relatively simple, light and compact and provide near-instantaneous response with very little input energy.

Lasers with high peak power pulses are commonly used in spaceflight missions for a wide range of applications, from LIDAR systems to optical communications. Due to the high optical power needed, the laser has to be located on the exterior of the satellite or coupled through a series of free space optics. This presents challenges for thermal management, radiation resistance, and mechanical design. Future applications will require multiple lasers located close together, which further complicates the design. Coupling the laser energy into a fiber optic cable allows the laser to be relocated to a more favorable position on the spacecraft. Typical fiber optic termination procedures are not sufficient for injection of these high-power laser pulses without catastrophic damage to the fiber endface. In the current study, we will review the causes of fiber damage during high-power injection and discuss our new manufacturing procedures that overcome these issues to permit fiber use with high reliability in these applications. We will also discuss the proper methods for launching the laser pulses into the fiber to avoid damage and how this is being implemented for current spaceflight missions.

We present an all-fiber continuous optical frequency comb-generator (OFCG) able to generate over 6 nm (750 GHz) at 1560 nm using a combination of electro-optic and acousto-optic modulations. As opposed to numerous experimental setups that use the longitudinal modes of an optical cavity to generate continuous optical frequency combs, our setup doesn’t need any active stabilization of the cavity length since we use the intrinsically high stability of radiofrequency sources to generate the multiple lines of the comb laser. Moreover, compared to the work of ref [1], the hybrid optical modulation we use allows to suppress the problem of instability due interferences between the generated lines. We notice that these lines benefit from the spectral quality of the seed laser because the spectral width of the synthesized hyperfrequency and radiofrequency signals are generally narrower than laser sources.

The Greenhouse gases Observing SATellite (GOSAT) was developed to contribute to monitoring of carbon dioxide and methane from space [1]. The mission objectives are global greenhouse gas measurements from space with precision of 1 % for CO₂ and 2 % for CH₄ in seasonal mean. The GOSAT carries Thermal And Near infrared Sensor for carbon Observation (TANSO) for precise measurement of greenhouse gases. Main instrument is Fourier Transfer Spectrometer (TANSO-FTS) to observe atmospheric absorption spectra of CO₂ and CH₄ with high spectral resolution of 0.2 cm^-1, broad wavelength coverage of 0.76 − 14.3 microns, wide swath of 790 km and frequent revisit of 3 days. Cloud and Aerosol Imager (TANSO-CAI) is simultaneously on board for cloud detection and correction of optical thin cirrus and aerosol interference within the FTS instantaneous field of view. The GOSAT satellite was launched by H2A-15 rocket on January 23, 2009. The Level 1B products of calibrated spectra were released from September 2009 in public. The Level 2 products of CO₂ and CH₄ column densities were released from February 2010 [2]. The normal observation data is acquired over one year regularly from April 2009. The mission lifetime is 5 years.

Resulting from a call for ideas for the 7th Earth Explorer mission, the Fluorescence Explorer (FLEX) mission [1] was selected for assessment as one of the six Earth Explorer missions to be studied within Phase 0. After the review of the study outcome by the Earth Science Advisory Committee (ESAC), FLEX was not selected for Phase A study. Although not selected, ESAC has expressed a clear recommendation to make an in-orbit demonstration of the measurement of vegetation fluorescence from space. Since FLEX involved four different instruments that would be required by an autonomous satellite to measure the chlorophyll fluorescence signal and to make the necessary atmospheric corrections, a new strategy has been proposed, which should make use of existing missions. The aim of this feasibility study is to identify a suitable instrument concept that is compatible with the expected resource limitations.

A number of Earth-observation missions, particularly those aimed at monitoring atmosphere composition, require hyperspectral instruments measuring Earth reflectance and emission at very fine spectral resolution. Examples include instruments for ESA’s Sentinel 5 mission, operating in the short-wave IR (SWIR) spectral region to quantify greenhouse gases, and also in the near-IR region (Oxygen-A band) to provide data on clouds, aerosols and atmosphere pressure. Conventional dispersive spectrometers, typically using diffraction gratings, are needed to provide data at moderate spatial resolution over large swath widths. Fine spectral resolution, together with the system aperture needed for radiometric resolution, tends to demand large beam diameters at gratings; this can lead to excessively large spectrometers, driven by the apertures of gratings and the associated collimators and camera lenses. This presents problems for accommodation of space-based spectrometers, particularly on small platforms.

Apertures, and complete spectrometer sizes, are reduced if the gratings provide high angular spectral dispersion (radians/nm) – this is a property of “immersed” gratings ^1,2. An immersed grating is a reflecting grating formed on a prism of refracting material, in which the incident and diffracted beams are both in the refracting medium. Instruments will typically use immersed gratings in silicon for short-wave IR bands, and silica for near-IR bands. This paper describes example designs for spectrometers using immersed gratings, and early results from a current development of immersed gratings are outlined.

Gaia is an ambitious ESA mission to chart a three-dimensional map of our Galaxy, the Milky Way, in the process revealing the composition, formation and evolution of the Galaxy. Gaia will provide unprecedented positional and radial velocity measurements with the accuracies needed to produce a stereoscopic and cinematic census of about one billion stars in our Galaxy. The payload consists of 2 Three Mirror Anastigmat (TMA) telescopes (aperture size ∼1.5 m x 0.5 m), 3 instruments (astrometer, photometer and spectrometer) and 106 butted CCDs assembled to a 0.9 Giga-Pixel focal plane.

The Radial Velocity Spectrometer (RVS) of Gaia measures the red shift of the stars in the spectral band between 847 nm and 874 nm. The spectrometer is a fully refractive optics consisting of 2 Fery prisms, 2 prisms, a pass band filter and a blazed transmission grating (instrument mass about 30 kg). It is located in the vicinity of the focal plane and illuminates 12 of the 106 Charge Coupled Devices (CCDs).

Gaia is in the implementation phase, the launch of the 2120 kg mass satellite is planned in Dec. 2012.

The optical part of the Raman Laser Spectrometer (RLS) instrument for ExoMars consists of an excitation laser, an optical harness, an optical head and a spectrometre. The optical harness delivers the green radiation generated by the laser to the optical head which, in turn, focuses the laser radiation on the sample of interest and collects the Raman emission from the sample. The optical head then separates excitation light and Raman emission by a filter setup and sends the isolated Raman signal to a reception fiber, which delivers it to the spectrometer of the instrument. This paper concentrates on the innovative technologies applied for the excitation path of the instrument, the laser, the optical harness with its new compact fiber optic connectors and the Raman optical head; and describes their design, the design driving requirements and the status these units have reached by now. The spectrometer of the system with its transmission grating design will be presented separately.

With the raise of resolution, optical payloads are becoming increasingly sensitive to satellite jitter. An approach where the entire spacecraft is pointed with great accuracy requires sophisticated and expensive bus design. In an effort to lower the overall cost of space missions that require highly stable line-of-sight pointing, a method of separating the bus and the payload with low frequency isolators is proposed. This isolation system can block the transmission of disturbance and allow relatively large bus motion. However, if the isolator is linear then there is a trade-off between isolation and static deflection as the launch and the on-orbit stage have difference requirements on the isolation frequency. Otherwise, an extra locking system should be appended to protect the payload before getting into orbit, as the STABLE isolation system^[1] and the MIM isolation system^[2] did.

To overcome this limitation, an alternative approach is to design a nonlinear isolator with high-static stiffness during launch and low dynamic stiffness on orbit. Several specially designed nonlinear isolators have achieved low dynamic stiffness with large static load capacity. Virgin^[3] considered a structure made from a highly deformed elastic element to achieve a softening spring. Platus^[4] exploited the buckling of beams under axial load in a specific configuration to achieve a negative stiffness in combination with a positive stiffness, and hence low-dynamic stiffness. Others have achieved the same by connecting linear springs with positive stiffness in parallel with elements of negative stiffness^[5]~[7].

In the present study, a bifunctional isolator has been developed for optical payloads. The isolator have good performance both during launch and on orbit because of its specially designed nonlinear stiffness and damping. The isolator works in a linear part with low stiffness and small damping ratio under the micro-vibration and microgravity on orbit. The transmissibility requirement and the displacement restriction during launch are satisfied by tuning the nonlinear stiffness and damping parameters. A group of sample isolators are designed tested both statically and dynamically.

The ground resolution of remote sensing satellite has been raised from hundreds of meters to less than one meter in recent few decades. As a result, the precision optical payload becomes more and more sensitive to structure vibrations of satellite buses. Although these vibrations generally have extremely low magnitude, they can result in significant image quality degradation to an optical payload. The suggestion of using vibration isolators to isolate payload from the satellite bus has been put forward in 1980s’^[1]. Recently, WorldView-2 achieved its perfect image quality via using a set of low frequency isolators^[2]. Recently, some of the optical payload manufacturers begin to provide vibration isolators as standard parts together with their main products . During the prototype testing of an earth resource satellite, the image of the optical payload was found to jitter for 5~10 pixels due to disturbances transmitted from the satellite bus structure. Test results indicated that the acceleration level of the vibration was of mG magnitude. To solve the problem, a highly sensitive vibration isolation system was developed to reduce the transmission of disturbances. Integrated isolation performance tests showed that the image jitter can be decreased to below 0.3 pixels.

The “Association de Satellites Pour l'Imagerie et l'Interferometrie de la Couronne Solaire”, ASPIICS, selected by ESA for the PROBA-3 mission, heralds the next generation of coronagraph for solar research, exploiting formation flying to gain access to the inner corona under eclipse-like conditions for long periods of time. A detailed description of the ASPIICS instrument and of its scientific objectives can be found in [1].

ASPIICS is distributed on the two PROBA 3 spacecrafts (S/C) separated by 150 m. The coronagraph optical assembly is hosted by the “coronagraph S/C” protected from direct solar disk light by the occulting disk on the “occulter S/C”.

The most critical issue in the design of a solar coronagraph is the reduction of the stray light due to the diffraction and scattering of the solar disk light by the occulter, the aperture and the optics. In the present article, we deal with two of these issues:

- The analysis of the stray light inside the telescope.

- The optimization of the external occulter edge, in order to eliminate the Poisson spot behind the occulter and to lower the stray light level going through the entrance pupil of the telescope.

This work was performed in the framework of the ESA STARTIGER program which took place at the Laboratoire d’Astrophysique de Marseille (LAM) during a 6-month period from September 2009 to March 2010.

In general, it is a very complicated task to combine the above two stray light issues together in the simulation and design phase as it requires to consider the propagation inside the telescope of the light diffracted by the external occulter. Actually, the present literature only reports diffraction calculations performed for simple occulting systems (i.e., two disks and serrated disk). A more pragmatic approach, also driven by the tight schedule of the STARTIGER program, is to separate the two contributions, and perform two different stray light analyses. This paper is dedicated to the description of both analyses: in particular, the first part is dedicated to the evaluation of the stray light inside the telescope, assuming a simple disk as occulter, and a preliminary baffle design is presented; the second part describes the investigation on the geometry of the external occulter, with a detailed description of the laboratory setup that has been designed and implemented to compare together several types of occulting systems.

To achieve superb stability in cryogenic optical systems, NOVA-ASTRON generally designs optical instruments on the basis of a 'no adjustments' philosophy. This means that in principle no corrections are possible after assembly. The alignment precision and consequently the performance of the instrument is guaranteed from the design, the tolerance analysis and the detailed knowledge of the material behavior and manufacturing process. This resulted in a higher degree of integrated optomechanical-cryogenic design with fewer parts, but with a higher part complexity. The 'no adjustments' strategy is successful because in the end the risk on instrument performance and project delays is much reduced. Astronomical instrument specifications have become more challenging over the years. Recent designs of the European Southern Observatory Very Large Telescope Interferometer (ESO VLTI) 4 Telescope combiner MATISSE include hundreds of optical components in a cryogenic environment. Despite the large number of optical components the alignment accuracy and stability requirements are in the order of nanometers. The 'no adjustments' philosophy would be too costly in this case, because all components would need to meet extremely tight manufacturing specifications. These specifications can be relaxed dramatically if cryogenic mechanisms are used for alignment. Several mechanisms have been developed: a tip-tilt mirror mechanism, an optical path distance mechanism, a slider mechanism, a bistable cryogenic shutter and a mirror mounting clip. Key aspects of these mechanisms are that the optical element and mechanism are combined in a compact single component, driven by e.g. self braking piezo actuators in order to hold position without power. The design, realization and test results of several mechanisms are presented in this paper.

The EarthCARE satellite mission objective is the observation of clouds and aerosols from low Earth orbit. The key spatial context providing instrument within the payload suite of 4 instruments is the Multi-Spectral Imager (MSI), previously described in [1]. The MSI is intended to provide information on the horizontal variability of the atmospheric conditions and to identify e.g. cloud type, textures, and temperature. It will form Earth images at 500m ground sample distance (GSD) over a swath width of 150km; it will image Earth in 7 spectral bands: one visible, one near-IR, two short-wave IR and three thermal IR. The instrument will be comprised of two key parts:

• a visible-NIR-SWIR (VNS) optical unit radiometrically calibrated using a sun illuminated quasivolume diffuser and shutter system

• a thermal IR (TIR) optical unit radiometrically calibrated using cold space and an internal black-body. This paper, being the first of a sequence of two, will provide an overview of the MSI and enter into more detail the critical performance parameters and detailed design the MSI TIR optical design.

The TIR concept is to provide pushbroom imaging of its 3 bands through spectral separation from a common aperture. The result is an efficient, well controlled optical design without the need for multiple focal plane arrays. The designed focal plane houses an area array detector and will meet a challenging set of requirements, including radiometric resolution, accuracy, distortion and MTF.

The EarthCARE satellite mission objective is the observation of clouds and aerosols from low Earth orbit. The payload will include active remote sensing instruments being the W-band Cloud Profiling Radar (CPR) and the ATLID LIDAR. These are supported by the passive instruments Broadband Radiometer (BBR) and the Multispectral Imager (MSI) providing the radiometric and spatial context of the ground scene being probed. The MSI will form Earth images over a swath width of 150 km; it will image the Earth atmosphere in 7 spectral bands. The MSI instrument consists of two parts: the Visible, Near infrared and Short wave infrared (VNS) unit, and the Thermal InfraRed (TIR) unit. Subject of this paper is the VNS unit.

In the VNS optical unit, the ground scene is imaged in four spectral bands onto four linear detectors via separate optical channels. Driving requirements for the VNS instrument performance are the spectral sensitivity including out-of-band rejection, the MTF, co-registration and the inter-channel radiometric accuracy. The radiometric accuracy performance of the VNS is supported by in-orbit calibration, in which direct solar radiation is fed into the instrument via a set of quasi volume diffusers.

The compact optical concept with challenging stability requirements together with the strict thermal constraints have led to a sophisticated opto-mechanical design.

This paper, being the second of a sequence of two on the Multispectral Imager describes the VNS instrument concept chosen to fulfil the performance requirements within the resource and accommodation constraints.

Geo-Oculus is a GEO-based Earth observation mission studied by Astrium for ESA in 2008-2009 to complement the Sentinel missions, the space component of the GMES (Global Monitoring for Environment & Security). Indeed Earth imaging from geostationary orbit offers new functionalities not covered by existing LEO observation missions, like real-time monitoring and fast revisit capability of any location within the huge area in visibility of the satellite. This high revisit capability is exploited by the Meteosat meteorogical satellites, but with a spatial resolution (500 m nadir for the third generation) far from most of GMES needs (10 to 100 m). To reach such ground resolution from GEO orbit with adequate image quality, large aperture instruments (> 1 m) and high pointing stability (<< 1 μrad) are required, which are the major challenges of such missions.

To address the requirements from the GMES user community, the Geo-Oculus mission is a combination of routine observations (daily systematic coverage of European coastal waters) with "on-demand" observation for event monitoring (e.g. disasters, fires and oil slicks). The instrument is a large aperture imaging telescope (1.5 m diameter) offering a nadir spatial sampling of 10.5 m (21 m worst case over Europe, below 52.5°N) in a PAN visible channel used for disaster monitoring. The 22 multi-spectral channels have resolutions over Europe ranging from 40 m in UV/VNIR (0.3 to 1 μm) to 750 m in TIR (10-12 μm).

The advent of giant, formation-flight, externally-occulted solar coronagraphs such as ASPIICS (Association de Satellites Pour l’Imagerie et l’Interférométrie de la Couronne Solaire [1,2,3,4]) selected by the European Space Agency (ESA) for its third PROBA (Project for On-Board Autonomy) mission of formation flying demonstration (presently in phase B) and Hi-RISE proposed in the framework of ESA Cosmic Vision program, presents formidable challenges for the study and calibration of instrumental stray light. With distances between the external occulter (EO) and the optical pupil (OP) exceeding hundred meters and occulter sizes larger than a meter, it becomes impossible to perform tests at the real scale. The requirement to limit the over-occultation to less than 1.05 Rsun, orders of magnitude to what has been achieved so far in past coronagraphs, further adds to the challenge. We are approaching the problem experimentally using reduced scale simulators and present below a progress report of our work.

The increasing interest on space telescopes for scientific applications leads to implement the manufacturing technology of the most critical element, i.e. the primary mirror: being more suitable a large aperture, it must be lightweight and deployable. The presented topic was originally addressed to a spaceborne DIAL (Differential Absorption LIDAR) mission operating at 935.5 nm for the measurement of water vapour profile in atmosphere, whose results were presented at ICSO 2006 and 2008.

Aim of this paper is to present the latest developments on the main issues related to the fabrication of a breadboard, covering two project critical areas identified during the preliminary studies: the design and performances of the long-stroke actuators used to implement the mirror active control and the mirror survivability to launch via Electrostatic Locking (EL) between mirror and backplane. The described work is developed under the ESA/ESTEC contract No. 22321/09/NL/RA.

The lightweight mirror is structured as a central sector surrounded by petals, all of them actively controlled to reach the specified shape after initial deployment and then maintained within specs for the entire mission duration. The presented study concerns: a) testing the Carbon Fiber Reinforced Plastic (CFRP) backplane manufacturing and EL techniques, with production of suitable specimens; b) actuator design optimisation; c) design of the deployment mechanism including a high precision latch; d) the fabrication of thin mirrors mock-ups to validate the fabrication procedure for the large shells.

The current activity aims to the construction of an optical breadboard capable of demonstrating the achievement of all these coupled critical aspects: optical quality of the thin shell mirror surface, actuators performances and back-plane - EL subsystem functionality.

Large aperture telescope technology (LATT) is a design study for a differential lidar (DIAL) system; the main investigation being into suitable methods, technologies and materials for a 4-metre diameter active mirror that can be stowed to fit into a typical launch vehicle (e.g. ROKOT launcher with ∼ 2.1-metre diameter cargo) and can self-deploy – in terms of both leaving the space vehicle and that the mirrors unfold and self-align to the correct optical form within the tolerances specified. The primary mirror requirements are: main wavelength of 935.5 nm, RMS corrected wavefront error of λ/6, optical surface roughness better than 5 nm, areal density of less than 16 kg/m² and 1-2 mirror shape corrections per orbit. The primary mirror consists of 7 segments - a central hexagonal mirror and 6 square mirror petals which unfold to form the 4-meter diameter aperture. The focus of the UK LATT consortium for this European Space Agency (ESA) funded project is on using lightweighted aluminium or carbon-fibre-composite materials for the mirror substrate in preference to more traditional materials such as glass and ceramics; these materials have a high strength and stiffness to weight ratio, significantly reducing risk of damage due to launch forces and subsequent deployment in orbit.

We present an overview of the design, which includes suitable actuators for wavefront correction, petal deployment mechanisms and lightweight mirror technologies. Preliminary testing results from manufactured lightweight mirror samples will also be summarised.

The need for both high quality images and light structures is a constant concern in the conception of space telescopes. In this paper, we present an active optics system as a way to fulfill those two objectives. Indeed, active optics consists in controlling mirrors’ deformations in order to improve the images quality [1]. The two main applications of active optics techniques are the in-situ compensation of phase errors in a wave front by using a corrector deformable mirror [2] and the manufacturing of aspherical mirrors by stress polishing or by in-situ stressing [3]. We will focus here on the wave-front correction. Indeed, the next generation of space telescopes will have lightweight primary mirrors; in consequence, they will be sensitive to the environment variations, inducing optical aberrations in the instrument.

An active optics system is principally composed of a deformable mirror, a wave front sensor, a set of actuators deforming the mirror and control/command electronics. It is used to correct the wave-front errors due to the optical design, the manufacturing imperfections, the large lightweight primary mirrors’ deflection in field gravity, the fixation devices, and the mirrors and structures’ thermal distortions due to the local turbulence [4]. Active optics is based on the elasticity theory [5]; forces and/or load are used to deform a mirror. Like in adaptive optics, actuators can simply be placed under the optical surface [1,2], but other configurations have also been studied: a system’s simplification, inducing a minimization of the number of actuators can be achieved by working on the mirror design [5]. For instance, in the so called Vase form Multimode Deformable Mirror [6], forces are applied on an external ring clamped on the pupil. With this method, there is no local effect due to the application of forces on the mirror’s back face. Furthermore, the number of actuators needed to warp the mirror does not depend on the pupil size; it is a fully scalable configuration.

The insertion of a Vase form Multimode Deformable Mirror on the design of an optical instrument will allow correcting the most common low spatial frequency aberrations. This concept could be applied in a space telescope. A Finite Element Analysis of the developed model has been conducted in order to characterize the system’s behavior and to validate the concept.

The Italian Space Agency selected the imaging interferometer ALISEO (Aerospace Leap-frog Imaging Stationary interferometer for Earth Observation) as the main payload for a technological optical mission based on the small satellite MIOsat. The simple design of such an instrument, based on Sagnac configuration, makes it a promising for Earth observation missions.

The ALISEO instrument acquires an image of 10 Km by 10 Km with a spatial resolution better than 10 m and a spectral resolution of 200 cm-1 (7 nm @ 0.6 μm) in the 0.4 – 1 μm spectral range.

ALISEO does not employ any moving part to generate the phase delays between the two interfering beams. The sensor acquires target images modulated by a pattern of autocorrelation functions of the energy coming from each scene pixel, and the resulting fringe pattern remains fixed with respect to the instrument’s field-of-view. The complete interferogram of each target location is retrieved by introducing a relative source-observer motion, which allows any image pixels to be observed under different viewing-angles corresponding to different Optical Path Differences (OPDs).

In this paper various optical configurations are analyzed in order to meet the mission requirements. Optical configurations are discussed taking into account: detector size, spatial resolution, and entrance pupil aperture. The proposed configurations should avoid vignetting, reduce geometric and chromatic aberrations, and comply with the size and weight constrains requested by space mission. Optical configurations, based on both refractive and reflective focusing elements, are presented and discussed. Finally, some properties pertaining to the selected Sagnac configuration are discussed in conjunction with spectral estimations and data processing.

Large field of view surveys with a high density of objects such as high-z galaxies or stars benefit of multi-object spectroscopy (MOS) technique. This technique is the best approach to eliminate the problem of spectral confusion, to optimize the quality and the SNR of the spectra, to reach fainter limiting fluxes and to maximize the scientific return. Next generation MOS for space like the Near Infrared Multi-Object Spectrograph (NIRSpec) for the James Webb Space Telescope (JWST) require a programmable multi-slit mask. The European EUCLID mission has also considered a MOS instrument in its early study phase. Conventional masks or complex fiber-optics-based mechanisms are not attractive for space. The programmable multi-slit mask requires remote control of the multi-slit configuration in real time. A promising possible solution is the use of MOEMS devices such as micromirror arrays (MMA) [1,2,3] or micro-shutter arrays (MSA) [4]. MMAs are designed for generating reflecting slits, while MSAs generate transmissive slits. MSA has been selected to be the multi-slit device for NIRSpec and is under development at NASA's Goddard Space Flight Center. In Europe, an effort is currently under way to develop single-crystalline silicon micromirror arrays for future generation infrared multi-object spectroscopy [5]. By placing the programmable slit mask in the focal plane of the telescope, the light from selected objects is directed toward the spectrograph, while the light from other objects and from the sky background is blocked.

Visitech is an engineering company experienced in developing DMD solution for industrial customers. The Laboratoire d’Astrophysique de Marseille (LAM) has, over several years, developed different tools for modeling and characterization of MOEMS-based slit masks, especially during the design studies on JWSTNIRSpec [6,7]. ESA has engaged with Visitech and LAM in a technical assessment of using a Digital Micromirror Devices (DMD) from Texas Instruments for space applications (for example in ESA EUCLID mission). The DMD features 2048 x 1080 mirrors on a 13.68μm mirror pitch (left-hand side of Fig. 1). Typical operational parameters of this device are room temperature, atmospheric pressure and mirrors switching thousands of times in a second, while for MOS applications in space, the device should work in vacuum, at low temperature, and each MOS exposure would last for typically 1500s with micromirrors held in a static state (either ON or OFF) during that duration. A specific thermal / vacuum test chamber has been developed for test conditions down to -40°C at 10^-5 mbar vacuum. Imaging capability for resolving each micro-mirror has also been developed for determining any single mirror failure. Dedicated electronics and software allows us to hold any pattern on the DMD for duration of up to 1500s.

We present the summary of this ESA study, the electronic test vehicle as well as the cold temperature test set-up we have developed. Then, results of tests in vacuum at low temperature, including low temperature stress test, low temperature nominal test, thermal cycling, and life test are presented. Results after radiation (TID and proton), and vibration and shock are also shown.

Optical instruments for space missions work in hostile environment, it’s thus necessary to accurately study the effects of ambient parameters variations on the equipment.

In particular optical instruments are very sensitive to ambient conditions, especially temperature. This variable can cause dilatations and misalignments of the optical elements, and can also lead to rise of dangerous stresses in the optics. Their displacements and the deformations degrade the quality of the sampled images.

In this work a method for studying the effects of the temperature variations on the performance of imaging instrument is presented. The optics and their mountings are modeled and processed by a thermo-mechanical Finite Element Model (FEM) analysis, then the output data, which describe the deformations of the optical element surfaces, are elaborated using an ad hoc MATLAB routine: a non-linear least square optimization algorithm is adopted to determine the surface equations (plane, spherical, n^th polynomial) which best fit the data. The obtained mathematical surface representations are then directly imported into ZEMAX for sequential raytracing analysis. The results are the variations of the Spot Diagrams, of the MTF curves and of the Diffraction Ensquared Energy due to simulated thermal loads.

This method has been successfully applied to the Stereo Camera for the BepiColombo mission reproducing expected operative conditions.

The results help to design and compare different optical housing systems for a feasible solution and show that it is preferable to use kinematic constraints on prisms and lenses to minimize the variation of the optical performance of the Stereo Camera.

This paper presents the results obtained in the frame of an ESA-funded project called “Screening and Preevaluation of Shortwave Infrared Laser Diode for Space Application” with the objective of verifying the maturity of state of the art SWIR DFB lasers at 2.1μm to be used for space applications (mainly based on the occultation measurement principle and spectroscopy). The paper focus on the functional and environmental evaluation test plan. It includes high precision characterization, mechanical test (vibration and SRS shocks), thermal cycling, gamma and proton radiation tests, life test and some details of the Destructive Physical Analysis performed. The electro-optical characterization includes measurements of the tuning capabilities of the laser both by current and by temperature, the wavelength stability and the optical power versus laser current.

In the framework of the ADM-Aeolus satellite mission, successful test campaigns have been performed in ESTEC’s laser laboratory, and the efficiency of several mitigation techniques against Laser-Induced Contamination (LIC) have been demonstrated for the ALADIN laser. These techniques include the standard contamination control methods of materials identification with particular tendency to cause LIC, reduction of the outgassing of organic materials by vacuum bake-out and shielding of optical surfaces from the contamination sources. Also novel mitigation methods such as in-situ cleaning via partial pressures, or the usage of molecular absorbers were demonstrated. In this context, a number of highly sensitive optical measurement techniques have been developed and tested to detect and monitor LIC deposits at nanometre level.

Applying freeform optics in high-end optical systems can improve system performance while decreasing the system mass, size and number of required components. Their widespread application is however held back by the lack of a suitable metrology method. TNO, TU/e and NMi VSL have therefore developed the NANOMEFOS measurement machine [1], capable of universal non-contact and fast measurement of aspherical and freeform optics ranging from convex to concave and from flat to freeform, up to ⊘500 mm, with an uncertainty below 30 nm (2σ).

A freeform optical surface is typically defined as any surface that does not have an axis of rotational symmetry. These surfaces provide additional degrees of freedom that can lead to improved performance compared to systems that make use solely of conventional optics.

Optical instrumentation on-board satellites suffer degradation due to the hostile conditions of space environment. Space conditions produce instrumentation performances changes causing a decrease or a cancellation of their features. Particularly, space environment conditions have a significant influence on the optical properties of glasses which are part of space optical systems.

Space environment characteristics which effects on the optical system have to be taken into account are: outgassing, volatile components, gas or water vapor which form part of the spacecraft materials, vacuum, microgravity, micrometeorites, space debris, thermal, mechanical and radiation environment and effects of the high atmosphere [1].

This work is focused on analyzing temperature variations and ultraviolet (UV) and gamma radiation effects on the optical properties of several glasses used on space applications.

Thermal environment is composed of radiation from the Sun, the albedo and the Earth radiation and the radiation from the spacecraft to deep space. Flux and influence of temperature on satellite materials depend on factors as the period of year or the position of them on the space system. Taking into account that the transfer mechanisms of heat are limited by the conduction and the radiation, high gradients of temperature are obtained in system elements which can cause changes of their optical properties, birefringence… Also, these thermal cycles can introduce mechanical loads into material structure due to the expansion and the contraction of the material leading to mechanical performances degradation [2].

However, it is the radiation environment the main cause of damage on optical properties of materials used on space instrumentation. This environment consists of a wide range of energetic particles between keV and MeV which are trapped by the geomagnetic field or are flux of particles that cross the Earth environment from the external of the Solar System [3].

The damage produced by the radiation environment on the optical materials can be classified in two types: ionizing or non-ionizing. This damage may produce continual or accumulative (dose) alterations on the optical material performances, or may produce alterations which not remain along the time (transitory effects). The effects of the radiation on optical materials can be summarized on changes of optical transmission and refractive index, variation of density and superficial degradation [4-6].

Two non-invasive and non-destructive techniques such as Optical Spectrum Analyzer and Spectroscopic Ellipsometry [7] have been used to characterize optically the three kinds of studied glasses, CaF₂, Fused Silica and Clearceram.

The study of the temperature and radiation effects on the glasses optical properties showed that the gamma radiation is the principal responsible of glasses optical degradation. The optical properties of the Clearceram glass have been affected by the gamma irradiation due to the absorption bands induced by the radiation in the visible spectral range (color centers). Therefore, an analysis about the behavior of these color centers with the gamma radiation total dose and with the time after the irradiation has been carried out in the same way that it is performed in [8].

It is essential to achieve fine pointing stability in a space mission aiming for high resolutional observations. In a future Japanese solar mission SOLAR-C, which is a successor of the HINODE (SOLAR-B) mission, we set targets of angular resolution better than 0.1 arcsec in the visible light and better than 0.2 - 0.5 arcsec in EUV and X-rays. These resolutions are twice to five times better than those of corresponding instruments onboard HINODE. To identify critical items to achieve the requirements of the pointing stability in SOLAR-C, we assessed in-flight performance of the pointing stability of HINODE that achieved the highest pointing stability in Japanese space missions. We realized that one of the critical items that have to be improved in SOLAR-C is performance of the attitude stability near the upper limit of the frequency range of the attitude control system. The stability of 0.1 arcsec (3σ) is required in the EUV and X-ray telescopes of SOLAR-C while the HINODE performance is slightly worse than the requirement. The visible light telescope of HINODE is equipped with an image stabilization system inside the telescope, which achieved the stability of 0.03 arcsec (3σ) by suppressing the attitude jitter in the frequency range lower than 10 Hz. For further improvement, it is expected to suppress disturbances induced by resonance between the telescope structures and disturbances of momentum wheels and mechanical gyros in the frequency range higher than 100 Hz.

In the frame of the Dutch Prequalification for ESA Programs(PEP), as part of the efforts to design an integrated optical attitude control subsytem (IOPACS), a consortium of TNO and several SME's in the Netherlands have been working on a novel type of startracker called MABS (Multiple Aperture Baffled Startracker). The system comprises a single cast metal housing with four reflective optical telescopes which use only structural internal baffling. Inherent to the design are a very high stability and excellent co-alignment between the apertures, a significant decrease in system size and low recurring production cost. The concept is a radical change from more common multiple startracker setups. The presentation will concentrate on the validity of the concept, the predicted performance and benefits for space applications, the produced breadboard and measured performances as well as the costing aspects.

What started as an academic development in frame of the Dutch MicroNed program five years ago should culminate in the introduction of the smallest digital sunsensor available in (and probably outside) of Europe at the ICSO. At the ICSO, TNO plans to show for the first time a working mini-DSS, based on the APS+ chip. The sunsensor has been optimized for low power, low recurring costs and high repeatability in production. In order to achieve this, several innovations have been included and verified in a diverse range of supporting programs. During the presentation, trades performed that lead to the current setup, as well as the properties of the sensor system and interfaces will be discussed.

Autonomous navigation of Satellite and constellation has series of benefits, such as to reduce operation cost and ground station workload, to avoid the event of crises of war and natural disaster, to increase spacecraft autonomy, and so on. Autonomous navigation satellite is independent of ground station support. Many systems are developed for autonomous navigation of satellite in the past 20 years. Along them American MANS (Microcosm Autonomous Navigation System) ^[1] of Microcosm Inc. and ERADS ^[2]，[3] (Earth Reference Attitude Determination System) of Honeywell Inc. are well known. The systems anticipate a series of good features of autonomous navigation and aim low cost, integrated structure, low power consumption and compact layout. The ERADS is an integrated small 3-axis attitude sensor system with low cost and small volume. It has the Earth center measurement accuracy higher than the common IR sensor because the detected ultraviolet radiation zone of the atmosphere has a brightness gradient larger than that of the IR zone. But the ERADS is still a complex system because it has to eliminate many problems such as making of the sapphire sphere lens, birefringence effect of sapphire, high precision image transfer optical fiber flattener, ultraviolet intensifier noise, and so on. The marginal sphere FOV of the sphere lens of the ERADS is used to star imaging that may be bring some disadvantages., i.e. , the image energy and attitude measurements accuracy may be reduced due to the tilt image acceptance end of the fiber flattener in the FOV. Besides Japan, Germany and Russia developed visible earth sensor for GEO ^[4]、[5]. Do we have a way to develop a cheaper/easier and more accurate autonomous navigation system that can be used to all LEO spacecraft, especially, to LEO small and micro satellites? To return this problem we provide a new type of the system—CANS (Compact Autonomous Navigation System) ^[6].

GAIA is one of the cornerstone ESA missions which aims at compiling a catalogue of about one billion stars of our galaxy. Reaching the highly demanding scientific requirements lead ASTRIUM engineers to design a mechanically and thermally ultra-stable instrument. This is the reason why, thanks to its physical properties, the SiC turned out to be indispensable.

The GAIA payload includes the following hardware which is mainly made of SiC i) the 3 meters quasi octagonal torus structure, ii) two identical 1.5 meters TMA type telescopes, iii) the central sub-assembly which holds several folding mirrors and the “Radial Velocity Spectrometer”, iv) the focal plane and v) the “Basic Angle Monitoring”.

Due to the required large size (1 - 3 meters class), accuracy and shape complexity, developing and manufacturing these SiC parts was a real challenge for BOOSTEC. It is reviewed in this paper.

As shown and discussed on a Sagem poster presented at the ICSO 2010 conference [1], scientific or commercial earth observation space instruments are more and more taking advantage of the remarkable properties of Silicon Carbide in term of hardness, stiffness and thermal stability combined with a reasonable density which are indeed of primary importance for all space applications.

Sagem-REOSC High Performance Optics Unit works on the polishing, coating and integration technologies of SiC mirrors since more than ten year through various successful space programs for various customers: INSAT 3D scan mirror, ROCSAT II and SPIRALE main telescopes, GAIA large primary mirrors and Auto-collimation flats, …).

This paper aims to provide to the international space community an exhaustive vision of the work performed by Sagem-REOSC on the polishing, coating and integration of the three Three Mirror Anastigmats of the NIRSpec spectrographic instrument which is the main ESA contribution to the JWST.

We carried out various tests of 800-mm-diameter aperture, lightweight optics that consisted wholly of carbon fiber-reinforced SiC composite, called HB-Cesic. A cryogenic optical test was performed on the primary mirror to examine any CTE irregularity as a surface change, and only small deformations were observed. The primary mirror was assembled with a convex secondary mirror into an optical system and tested under vacuum at the 6-m-diameter radiometer space chamber at Tsukuba Space Center of JAXA, where we have prepared interferometric metrological facilities to establish techniques to test large optical systems in a horizontal light-axis configuration. The wavefront difference between under vacuum and under atmosphere was confirmed to be less than 0.1 λ at λ=633 nm, if we realigned the optical axis of the interferometer and flat mirror under vacuum. We also demonstrated a stitching interferometry using the Φ800-mm optics by rotating a mask wheel of subapertures in front of the optical reference flat. The wavefront stitched from eight individual measurements of Φ275-mm subapertures differs from the full-aperture measurement without the mask by about 0.1 λ nm RMS, which showed the technique could able to be applied to test large telescopes especially for infrared wavelength region.

Future scientific space missions require ever more demanding large optics that work at cryogenic temperatures. In the frame of a Darwin assessment study conducted under ESA contract by TAS, the need of future very lightweight cryogenic mirrors with superior optical quality has been identified. Such mirrors need to be of size up to 3.5 m in diameter, with a mass of less than 250 kg (i.e. 25 kg/m²) and possess excellent optical quality at cryogenic temperature down to 40 K.

The World Space Observatory Ultraviolet (WSO/UV) is a multi-national project lead by the Russian Federal Space Agency (Roscosmos) with the objective of high performance observations in the ultraviolet range. The 1.7 m WSO/UV telescope feeds UV spectrometers and UV imagers. The UV spectrometers comprise two high resolution Echelle spectrographs for the 100 − 170 nm and 170 − 300 nm wavelength range and a long slit spectrograph for the 100 − 300 nm band. All three spectrometers represent individual instruments that are assembled and aligned separately. In order to save mass while maintaining high stiffness, the instruments are combined to a monoblock. Cesic has been selected to reduce CTE related distortions of the instruments.

In contrast to aluminium, the stable structure of Cesic is significantly less sensitive to thermal gradients. No further mechanism for focus correction with high functional, technical and operational complexity and dedicated System costs are necessary. Using Cesic also relaxes the thermal control requirements of ±5°C, which represents a considerable cost driver for the S/C design.

The WUVS instrument is currently studied in the context of a phase B2 study by Kayser-Threde GmbH including a Structural Thermal Model (STM) for verification of thermal and mechanical loads, stability due to thermal distortions and Cesic manufacturing feasibility.

High resolution stellar interferometers are very powerful efficient instruments to get a better knowledge of our Universe through the spatial coherence analysis of the light. For this purpose, the optical fields collected by each telescope T_i are mixed together. From the interferometric pattern, two expected information called the contrast C_ij and the phase information φ_ij are extracted. These information lead to the V_ij, called the complex visibility, with V_ij=C_ijexp(jφ_ij). For each telescope doublet T_iT_j, it is possible to get a complex visibility V_ij. The Zernike Van Cittert theorem gives a relationship between the intensity distribution of the object observed and the complex visibility. The combination of the acquired complex visibilities and a reconstruction algorithm allows imaging reconstruction.

To avoid lots of technical difficulties related to infrared optics (components transmission, thermal noises, thermal cooling…), our team proposes to explore the possibility of using nonlinear optical techniques. This is a promising alternative detection technique for detecting infrared optical signals. This way, we experimentally demonstrate that frequency conversion does not result in additional bias on the interferometric data supplied by a stellar interferometer.

In this presentation, we report on wavelength conversion of the light collected by each telescope from the infrared domain to the visible. The interferometric pattern is observed in the visible domain with our, so called, upconversion interferometer. Thereby, one can benefit from mature optical components mainly used in optical telecommunications (waveguide, coupler, multiplexer…) and efficient low-noise detection schemes up to the single-photon counting level.

Since several years, perspectives and assets offered by photonic technologies compared with their traditional RF counterparts (mass and volume reduction, transparency to RF frequency, RF isolation), make them particularly attractive for space applications [1] and, in particular, telecommunication satellites [2]. However, the development of photonic payload concepts have concurrently risen and made the problem of the ability of optoelectronic components to withstand space environment more and more pressing. Indeed, photonic components used in such photonic payloads architectures come from terrestrial networks applications in order to benefit from research and development in this field.

This paper presents some results obtained in the frame of an ESA-funded project, carried out by Thales Alenia Space France, as prime contractor, and Alter Technology Group Spain (ATG) and Universidad Politecnica de Madrid (UPM), as subcontractors, one objective of which was to assess commercial high frequency optical intensity modulators for space use through a functional and environmental test campaign. Their potential applications in microwave photonic sub-systems of telecom satellite payloads are identified and related requirements are presented. Optical modulator technologies are reviewed and compared through, but not limited to, a specific figure of merit, taking into account two key features of these components : optical insertion loss and RF half-wave voltage. Some conclusions on these different technologies are given, on the basis of the test results, and their suitability for the targeted applications and environment is highlighted.

In this paper we present the results obtained within the context of the ESA-funded project Programmable Optoelectronic Adaptive Element (AO/1-5476/07/NL/EM). The objective of this project is the development of adaptive (reconfigurable) optical elements for use in space applications and the execution of preliminary qualification tests in the relevant environment.

The different designs and materials that have been considered and manufactured for a 2D beam steerer based on passive matrix liquid crystal programmable blaze grating will described and discussed.

CNES (French spatial agency) will provide the AltiKa high resolution altimeter, Doris instrument and the LRA (Laser Retroreflector Array) for SARAL (Satellite with Argos and AltiKa) in cooperation with ISRO (Indian space agency).

The LRA is a passive equipment reflecting the laser beams coming from the Earth ground stations. Computing the send-return time travel of the laser beams allows the determination of the satellite altitude within an accuracy of a few millimeters. The reflective function is done by a set of 9 corner cube reflectors, with a conical arrangement providing a 150 degrees wide field of view over the full 360 degrees azimuth angle.

According to CNES optomechanical specifications, the LRA has been developed by SESO (French optical firm). SESO has succeeded in providing the corner cube reflectors with a very stringent dihedral angle error of 1.6 arcsec and an accuracy within ±0.5 arcsec. During this development, SESO has performed mechanical, thermal and thermo-optical analyses. The optical gradient of each corner cube, as well as angular deviations and PSF (Point Spread Function) in each laser range finding direction, have been computed. Mechanical and thermal tests have been successfully performed. A thermo-optical test has successfully confirmed the optical effect of the predicted in-flight thermal gradients.
Each reflector is characterized in order to find its best location in the LRA housing and give the maximum optimization to the space telemetering mission.

Lightweight X-ray Wolter optics with a high angular resolution will enable the next generation of X-ray telescopes in space. The International X-ray Observatory (IXO) requires a mirror assembly of 3 m² effective area (at 1.5 keV) and an angular resolution of 5 arcsec. These specifications can only be achieved with a novel technology like Silicon Pore Optics, which is developed by ESA together with a consortium of European industry.

Silicon Pore Optics are made of commercial Si wafers using process technology adapted from the semiconductor industry. We present the manufacturing process ranging from single mirror plates towards complete focusing mirror modules mounted in flight configuration. The performance of the mirror modules is tested using X-ray pencil beams or full X-ray illumination. In 2009, an angular resolution of 9 arcsec was achieved, demonstrating the improvement of the technology compared to 17 arcsec in 2007. Further development activities of Silicon Pore Optics concentrate on ruggedizing the mounting system and performing environmental tests, integrating baffles into the mirror modules and assessing the mass production.

The name “eROSITA” stands for extended Roentgen Survey with an Imaging Telescope Array. The general design of the eROSITA X-ray telescope is derived from that of ABRIXAS. A bundle of 7 mirror modules with short focal lengths make up a compact telescope which is ideal for survey observations. Similar designs had been proposed for the missions DUO and ROSITA but were not realized due to programmatic shortfall. Compared to those, however, the effective area in the soft X-ray band has now much increased by adding 27 additional outer mirror shells to the original 27 ones of each mirror module. The requirement on the on-axis resolution has also been confined, namely to 15 arc seconds HEW. For these reasons the prefix “extended” was added to the original name “ROSITA”. The scientific motivation for this extension is founded in the ambitious goal to detect about 100,000 clusters of galaxies which trace the large scale structure of the Universe in space and time.

The X-ray telescope of eROSITA will consist of 7 identical and co-aligned mirror modules, each with 54 nested Wolter-1 mirror shells. The mirror shells are glued onto a spider wheel which is screwed to the mirror interface structure making a rigid mechanical unit. The assembly of 7 modules forms a compact hexagonal configuration with 1300 mm diameter (see Fig. 1) and will be attached to the telescope structure which connects to the 7 separate CCD cameras in the focal planes. The co-alignment of the mirror module enables eROSITA to perform also pointed observations.

The replication process described in chapter III allows the manufacturing in one single piece and at the same time of both the parabola and hyperbola parts of the Wolter 1 mirror.

The Wide Field X-ray Telescope (WFXT) is a medium class mission proposed to address key questions about cosmic origins and physics of the cosmos through an unprecedented survey of the sky in the soft X-ray band (0.2-6 keV) [1], [2]. In order to get the desired angular resolution of 10 arcsec (5 arcsec goal) on the entire 1 degrees Field Of View (FOV), the design of the optical system is based on nested grazing-incidence polynomial profiles mirrors, and assumes a focal plane curvature and plate scale corrections among the shells. This design guarantees an increased angular resolution also at large off-axis positions with respect to the usually adopted Wolter I configuration. In order to meet the requirements in terms of mass and effective area (less than 1200 kg, 6000 cm² @ 1 keV), the nested shells are thin and made of quartz glass. The telescope assembly is composed by three identical modules of 78 nested shells each, with diameter up to 1.1 m, length in the range of 200-440 mm and thickness of less than 2.2 mm. At this regard, a deterministic direct polishing method is under investigation to manufacture the WFXT thin grazing-incidence mirrors made of quartz. The direct polishing method has already been used for past missions (as Einstein, Rosat, Chandra) but based on much thicker shells (10 mm ore more). The technological challenge for WFXT is to apply the same approach but for 510 times thinner shells. The proposed approach is based on two main steps: first, quartz glass tubes available on the market are ground to conical profiles; second the pre-shaped shells are polished to the required polynomial profiles using a CNC polishing machine. In this paper, preliminary results on the direct grinding and polishing of prototypes shells made by quartz glass with low thickness, representative of the WFXT optical design, are presented.

The New Hard X-ray Mission (NHXM) project will be operated by 2017 and is currently undergoing a Phase B study, under the coordination of the Italian Space Agency (ASI). The project is being proposed by an international team in the context of the ESA Call CV M3 as a Small Mission program, with a large Italian participation. It is based on 4 hard X-ray optics modules, each formed by 60 evenly spaced multilayer coated Wolter I mirror shells. An extensible bench is used to reach the 10 m focal length. The Wolter I monolithic substrates with multilayer coating are produced in NiCo by electroforming replication. Three of the mirror modules will host in the focal plane a hybrid a detector system (a soft X-ray Si DEPFET array plus a high energy CdTe detector). The detector of the fourth telescope will be a photoelectric polarimeter with imaging capabilities, operating from 2 up to 35 keV. The total on axis effective area of the three telescopes at 1 keV and 30 kev is of 1500 cm² and 350 cm² respectively, with an angular resolution of 20 arcsec HEW at 30 keV. In this paper we report on the design and development of the multilayer coated X-ray mirrors based on NiCo shells.

We describe the realization of a 5 km free space coherent optical link through the turbulent atmosphere between a telescope and a ground target. We present the phase noise of the link, limited mainly by atmospheric turbulence and mechanical vibrations of the telescope and the target. We discuss the implications of our results for applications, with particular emphasis on optical Doppler ranging to satellites and long distance frequency transfer.

We present two different laser ranging systems under development, both based on the use of a high frequency modulated beam. The first range meter makes no use of interferometry: only the phase of the return beam is detected, in a way that rejects cyclic errors due to optical and electronic crosstalk. An Allan deviation slightly better than 10nm has been obtained with this simple system. The other range meter should provide better resolution, at the expense of a somewhat more sophisticated procedure, as it involves both time-of-flight and interferometry measurements.

Design, integration, test setup, test results, and lessons-learnt of a high precision laser metrology demonstrator for dual absolute and relative laser distance metrology are presented. The different working principles are described and their main subsystems and performance drivers are presented. All subsystems have strong commonalities with flight models as of LTP on LISA Pathfinder and laser communication missions, and different pathways to flight models for varying applications and missions are presented. The setup has initially been realized within the ESA project "High Precision Optical Metrology (HPOM)", originally initiated for DARWIN formation flying optical metrology, though now serves as demonstrator for a variety of future applications. These are sketched and brought into context (PROBA-3, IXO onboard metrology, laser gravimetry earth observation missions, fundamental science missions like LISA and Pioneer anomaly).

Formation flying, with ESA’s mission PROBA-3, is providing the chance of creating a giant solar coronagraph in Space. The scientific payload, the solar coronagraph ASPIICS, has been selected in January 2009 [1]. The advantages of formation flying are: 1) larger dimensions for the coronagraph, which leads to better spatial resolution and lower straylight level and 2) possibility of continuous observations of the inner corona. The PROBA-3/ASPIICS mission is composed of two spacecrafts (S/Cs) at 150 meters distance, the Occulter-S/C (O-S/C) which holds the external occulter, and the Coronagraph-S/C (C-S/C) which holds the main instrument, i.e. the telescope. In addition of the scientific capabilities of the instrument, it will continuously monitor the exact position and pointing of both S/Cs in 3D space, via two additional metrology units: the Shadow Position Sensor (SPS) and the Occulter Position Sensor (OPS). In this paper we are presenting the metrology of this formation flying mission combining the outputs of the above mentioned sensors, SPS and OPS. This study has been conducted in the framework of an ESA “STARTIGER” initiative, a novel approach aimed at demonstrating the feasibility of a new and promising technology concept (in our case formation flying applied to solar coronagraphy, cf. [2, 3]) on a short time scale (six months study).

Lunar Lander and Rover of China will be launched in 2013. It will finish the mission targets of lunar soft landing and patrol exploration. Lunar Rover has forward facing stereo camera pair (Hazcams) for hazard avoidance. Hazcams calibration is essential for stereo vision. The Hazcam optics are f-theta fish-eye lenses with a 120°×120° horizontal/vertical field of view (FOV) and a 170° diagonal FOV. They introduce significant distortion in images and the acquired images are quite warped, which makes conventional camera calibration algorithms no longer work well.

A photogrammetric calibration method of geometric model for the type of optical fish-eye constructions is investigated in this paper. In the method, Hazcams model is represented by collinearity equations with interior orientation and exterior orientation parameters [1] [2]. For high-precision applications, the accurate calibration model is formulated with the radial symmetric distortion and the decentering distortion as well as parameters to model affinity and shear based on the fisheye deformation model [3] [4]. The proposed method has been applied to the stereo camera calibration system for Lunar Rover.

The X-ray optics is a key element of various X-ray telescopes, X-ray microscopes, as well as other X-ray imaging instruments. The grazing incidence X-ray lenses represent the important class of X-ray optics. Most of grazing incidence (reflective) X-ray imaging systems used in astronomy but also in other (laboratory) applications are based on the Wolter 1 (or modified) arrangement. But there are also other designs and configurations proposed, used and considered for future applications both in space and in laboratory. The Kirkpatrick-Baez (K-B) lenses as well as various types of Lobster-Eye optics and MCP/Micropore optics serve as an example. Analogously to Wolter lenses, the X-rays are mostly reflected twice in these systems to create focal images. Various future projects in X-ray astronomy and astrophysics will require large segments with multiple thin shells or foils. The large Kirkpatrick-Baez modules, as well as the large Lobster-Eye X-ray telescope modules in Schmidt arrangement may serve as examples. All these space projects will require high quality and light segmented shells (bent or flat foils) with high X-ray reflectivity and excellent mechanical stability. The Multi Foil Optics (MFO) approach represent a promising alternative for both LE and K-B X-ray optical modules. Several types of reflecting substrates may be considered for these applications, with emphasis on thin float glass sheets and, more recently, high quality silicon wafers. This confirms the importance of non- Wolter X-ray optics designs for the future.

Future large space X-ray telescopes (such as IXO) require precise and light-weight X-ray optics based on numerous thin reflecting shells. Novel approaches and advanced technologies are to be exploited and developed. In this contribution, we refer on results of tested X-ray mirror shells produced by glass thermal forming (GTF) and by shaping Si wafers. Both glass foils and Si wafers are commercially available, have excellent surface microroughness of a few 0.1 nm, and low weight (the volume density is 2.5 g cm^-3 for glass and 2.3 g cm^-3 for Si). Technologies are needed to be exploited; how to shape these substrates to achieve the required precise Xray optics geometries without degradations of the fine surface microroughness. Although glass and recently silicon wafers are considered to represent most promising materials for future advanced large aperture space Xray telescopes, there also exist other alternative materials worth further study such as amorphous metals and glassy carbon [1]. In order to achieve sub-arsec angular resolutions, principles of active optics have to be adopted.

The Smart X-ray Optics (SXO) Basic Technology project started in April 2006 and will end in October 2010. The aim is to develop new technologies in the field of X-ray focusing, in particular the application of active and adaptive optics. While very major advances have been made in active/adaptive astronomical optics for visible light, little was previously achieved for X-ray optics where the technological challenges differ because of the much shorter wavelengths involved.

The field of X-ray astronomy has been characterized by the development and launch of ever larger observatories with the culmination in the European Space Agency’s XMM-Newton and NASA's Chandra missions which are currently operational. XMM-Newton uses a multi-nested structure to provide modest angular resolution (∼10 arcsec) but large effective area, while Chandra sacrifices effective area to achieve the optical stability necessary to provide sub-arc second resolution. Currently the European Space Agency (ESA) is engaged in studies of the next generation of X-ray space observatories, with the aim of producing telescopes with increased sensitivity and resolution. To achieve these aims several telescopes have been proposed, for example ESA and NASA’s combined International X-ray Observatory (IXO), aimed at spectroscopy, and NASA’s Generation-X. In the field of X-ray astronomy sub 0.2 arcsecond resolution with high efficiency would be very exciting. Such resolution is unlikely to be achieved by anything other than an active system. The benefits of a such a high resolution would be important for a range of astrophysics subjects, for example the potential angular resolution offered by active X-ray optics could provide unprecedented structural imaging detail of the Solar Wind bowshock interaction of comets, planets and similar objects and auroral phenomena throughout the Solar system using an observing platform in low Earth orbit.

A major aim of the SXO project was to investigate the production of thin actively controlled grazing incident optics for the next generation of X-ray space telescopes. Currently telescope systems are limited in the resolution and sensitivity by the optical quality of the thin shell optics used. As part of its research programme an actively controlled prototype X-ray thin shell telescope optic of dimensions 30x10cm has been developed to bench test the technology. The design is based on thin nickel shells bonded to shaped piezo-electric unimorph actuators made from lead zirconate titanate (PZT).

A Compact Real-Time Optronic SAR Processor has been successfully developed and tested up to a Technology Readiness Level of 4 (TRL4), the breadboard validation in a laboratory environment. SAR, or Synthetic Aperture Radar, is an active system allowing day and night imaging independent of the cloud coverage of the planet. The SAR raw data is a set of complex data for range and azimuth, which cannot be compressed. Specifically, for planetary missions and unmanned aerial vehicle (UAV) systems with limited communication data rates this is a clear disadvantage.

SAR images are typically processed electronically applying dedicated Fourier transformations. This, however, can also be performed optically in real-time. Originally the first SAR images were optically processed. The optical Fourier processor architecture provides inherent parallel computing capabilities allowing real-time SAR data processing and thus the ability for compression and strongly reduced communication bandwidth requirements for the satellite.

SAR signal return data are in general complex data. Both amplitude and phase must be combined optically in the SAR processor for each range and azimuth pixel. Amplitude and phase are generated by dedicated spatial light modulators and superimposed by an optical relay set-up. The spatial light modulators display the full complex raw data information over a two-dimensional format, one for the azimuth and one for the range. Since the entire signal history is displayed at once, the processor operates in parallel yielding real-time performances, i.e. without resulting bottleneck. Processing of both azimuth and range information is performed in a single pass.

This paper focuses on the onboard capabilities of the compact optical SAR processor prototype that allows in-orbit processing of SAR images. Examples of processed ENVISAT ASAR images are presented. Various SAR processor parameters such as processing capabilities, image quality (point target analysis), weight and size are reviewed.

For observation of gravitational waves at frequencies between 30 μHz and 1 Hz, the LISA mission will be implemented in a triangular constellation of three identical spacecraft, which are mutually linked by laser interferometry in an active transponder scheme over a 5 million kilometer arm length. On the end point of each laser link, remote and local beam metrology with respect to inertial proof masses inside the spacecraft is realized by the LISA Optical Bench. It implements further- more various ancillary functions such as point-ahead correction, acquisition sensing, transmit beam conditioning, and laser redundancy switching.

A comprehensive design of the Optical Bench has been developed, which includes all of the above mentioned functions and at the same time ensures manufacturability on the basis of hydroxide catalysis bonding, an ultrastable integration technology already perfected in the context of LISA's technology demonstrator mission LISA Pathfinder. Essential elements of this design have been validated by dedicated pre-investigations. These include the demonstration of polarizing heterodyne interferometry at the required Picometer and Nanoradian performance levels, the investigation of potential non-reciprocal noise sources in the so-called backlink fiber, as well as the development of a laser redundancy switch breadboard.

LISA, the first space project for detecting gravitational waves, relies on two main technical challenges: the free falling masses and an outstanding precision on phase shift measurements (a few pm on 5 Mkm in the LISA band). The technology of the free falling masses, i.e. their isolation to forces other than gravity and the capability for the spacecraft to precisely follow the test masses, will soon be tested with the technological LISA Pathfinder mission. The performance of the phase measurement will be achieved by at least two stabilization stages: a pre-stabilisation of the laser frequency at a level of 10^-13 (relative frequency stability) will be further improved by using numerical algorithms, such as Time Delay Interferometry, which have been theoretically and numerically demonstrated to reach the required performance level (10^-21).

Nevertheless, these algorithms, though already tested with numerical model of LISA, require experimental validation, including ‘realistic’ hardware elements. Such an experiment would allow to evaluate the expected noise level and the possible interactions between subsystems. To this end, the APC is currently developing an optical benchtop experiment, called LISA On Table (LOT), which is representative of the three LISA spacecraft.

A first module of the LOT experiment has been mounted and is being characterized. After completion this facility may be used by the LISA community to test hardware (photodiodes, phasemeters) or software (reconstruction algorithms) components.

Detection and observation of gravitational waves requires extreme stability in the frequency range 0.03 mHz to 1 Hz. The Laser Interferometer Space Antenna (LISA) mission will attain this by creating a giant interferometer in space, based on free floating proof masses in three spacecrafts. Due to orbit evolution and time delay in the interferometer arms, the direction of transmitted light changes. To solve this problem, a picometer stable Point-Ahead Angle Mechanism (PAAM) was designed, realized and successfully tested. The PAAM concept is based on a rotatable mirror. The critical requirements are the contribution to the optical path length (less than 1.4 pm / rt Hz) and the angular jitter (less than 8 nrad / rt Hz). Extreme dimensional stability is achieved by manufacturing a monolithical Haberland hinge mechanism out of Ti₆Al₄V, through high precision wire erosion. Extreme thermal stability is realized by placing the thermal center on the surface of the mirror. Because of piezo actuator noise and leakage, the PAAM has to be controlled in closed-loop. To meet the requirements in the low frequencies, an active target capacitance-to-digital converter is used. Interferometric measurements with a triangular resonant cavity in vacuum proved that the PAAM meets the requirements.

Atom interferometry has hugely benefitted from advances made in cold atom physics over the past twenty years, and ultra-precise quantum sensors are now available for a wide range of applications [1]. In particular, cold atom interferometers have shown excellent performances in the field of acceleration and rotation measurements [2,3], and are foreseen as promising candidates for navigation, geophysics, geo-prospecting and tests of fundamental physics such as the Universality of Free Fall (UFF). In order to carry out a test of the UFF with atoms as test masses, one needs to compare precisely the accelerations of two atoms with different masses as they fall in the Earth’s gravitational field. The sensitivity of atom interferometers scales like the square of the time during which the atoms are in free fall, and on ground this interrogation time is limited by the size of the experimental setup to a fraction of a second. Sending an atom interferometer in space would allow for several seconds of excellent free-fall conditions, and tests of the UFF could be carried out with precisions as low as 10^-15 [4].

However, cold atoms experiments rely on complex laser systems, which are needed to cool down and manipulate the atoms, and these systems are usually very sensitive to temperature fluctuations and vibrations. In addition, when operating an inertial sensor, vibrations are a major issue, as they deteriorate the performances of the instrument. This is why cold atom interferometers are usually used in ground based facilities, which provide stable enough environments. In order to carry out airborne or space-borne measurements, one has to design an instrument which is both compact and stable, and such that vibrations induced by the platform will not deteriorate the sensitivity of the sensor.

We report on the operation of an atom interferometer on board a plane carrying out parabolic flights (Airbus A300 Zero-G, operated by Novespace). We have constructed a compact and stable laser setup, which is well suited for onboard applications. Our goal is to implement a dual-species Rb-K atom interferometer in order to carry out a test of the UFF in the plane. In this perspective, we are designing a dual-wavelength laser source, which will enable us to cool down and coherently manipulate the quantum states of both atoms. We have successfully tested a preliminary version of the source and obtained a double species magneto-optical trap (MOT).

The Space Optical Clocks project aims at operating lattice clocks on the ISS for tests of fundamental physics and for providing high-accuracy comparisons of future terrestrial optical clocks. A pre-phase-A study (2007- 10), funded partially by ESA and DLR, included the implementation of several optical lattice clock systems using Strontium and Ytterbium as atomic species and their characterization. Subcomponents of clock demonstrators with the added specification of transportability and using techniques suitable for later space use, such as all-solid-state lasers, low power consumption, and compact dimensions, have been developed and have been validated. This included demonstration of laser-cooling and magneto-optical trapping of Sr atoms in a compact breadboard apparatus and demonstration of a transportable clock laser with 1 Hz linewidth. With two laboratory Sr lattice clock systems a number of fundamental results were obtained, such as observing atomic resonances with linewidths as low as 3 Hz, non-destructive detection of atom excitation, determination of decoherence effects and reaching a frequency instability of 1×10^-16.

Active Optics methods, based on elasticity theory, allow the aspherisation of optical surfaces by stress polishing but also active aspherisation in situ. Researches in this field will impact the final performance and the final cost of any telescope or instrument. The stress polishing method is well suited for the superpolishing of aspheric components for astronomy. Its principle relies on spherical polishing with a full-sized tool of a warped substrate, which becomes aspherical once unwarped. The main advantage of this technique is the very high optical quality obtained either on form or on high spatial frequency errors. Furthermore, the roughness can be decreased down to a few angstroms, thanks the classical polishing with a large pitch tool, providing a substantial gain on the final scientific performance, for instance on the contrast on coronagraphic images, but also on the polishing time and cost. Stress polishing is based on elasticity theory, and requires an optimised deformation system able to provide the right aspherical form on the optical surface during polishing. The optical quality of the deformation is validated using extensive Finite Element Analysis, allowing an estimation of residuals and an optimisation of the warping harness. We describe here the work realised on stress polishing of toric mirrors for VLT-SPHERE and then our actual work on off axis aspherics (OAA) for the ASPIICS-Proba3 mission for solar coronagraphy. The ASPIICS optical design made by Vives et al is a three mirrors anastigmat including a concave off axis hyperboloid and a convex off axis parabola (OAP). We are developing a prototype in order to demonstrate the feasibility of this type of surface, using a multi-mode warping harness (Lemaitre et al). Furthermore, we present our work on variable OAP, meaning the possibility to adjust the shape of a simple OAP in situ with a minimal number of actuators, typically one actuator per optical mode (Focus, Coma and Astigmatism). Applications for future space telescopes and instrumentation are discussed.

The fairing of the launcher selected for the Space Infrared telescope for Cosmology and Astrophysics (SPICA) mission is not compatible with a primary mirror of 3.5m in diameter. Thus three alternative optical designs of the SPICA Telescope Assembly (STA) with a primary mirror of reduced size were defined and their theoretical optical performances assessed. The impact of the size reduction on the STA optical performances was then quantified. Based on the results of the study, we defined a STA optical design optimum in terms of optical performances and of accommodation of instruments in the STA focal surface.

The purpose of this paper is to present the optical system developed for the Wide Field imaging Camera - WFI that will be integrated to the CBERS 3 and 4 satellites (China Brazil Earth resources Satellite). This camera will be used for remote sensing of the Earth and it is aimed to work at an altitude of 778 km. The optical system is designed for four spectral bands covering the range of wavelengths from blue to near infrared and its field of view is ±28.63°, which covers 866 km, with a ground resolution of 64 m at nadir. WFI has been developed through a consortium formed by Opto Electrônica S. A. and Equatorial Sistemas. In particular, we will present the optical analysis based on the Modulation Transfer Function (MTF) obtained during the Engineering Model phase (EM) and the optical tests performed to evaluate the requirements. Measurements of the optical system MTF have been performed using an interferometer at the wavelength of 632.8nm and global MTF tests (including the CCD and signal processing electronic) have been performed by using a collimator with a slit target. The obtained results showed that the performance of the optical system meets the requirements of project.

High Energy Astrophysics (HEA) encompasses a broad range of astrophysical science, with sources that include stars and stellar clusters, compact objects (black holes, neutron stars, and white dwarfs), supernova remnants, the interstellar medium, galaxies and clusters of galaxies, Active Galactic Nuclei (AGN), and gamma ray bursters, as well as a variety of fundamental physical processes. The physics involved includes extremes of gravity, density and magnetic field and is often inaccessible via any other waveband. HEA investigates and answers crucial questions in all fields of contemporary astrophysics.

Unlike the focusing of radio and optical light, X-rays are brought to focus through shallow, grazing incident angles. The analogy of skimming a stone across a pond is appropriate in describing how X-rays are focused. The higher the energy of the X-ray photon the shallower the incident angle must be, thereby introducing the requirement of longer focal lengths for focusing high-energy X-rays (E > 10 keV). This technical challenge has hindered scientific advancement in the high-energy regime, while at lower X-ray energies the community has prospered immensely with spectacular data from focusing observatories like XMM-Newton, Chandra, and Suzaku. Now, with ASTRO-H, the community will reap similar rewards from the tremendous improvement in spatial and spectral resolution at high energies. ASTRO-H is a JAXA mission. More information can be found on the ASTRO-H web site [1].

Because of the grazing-angle optics, high-energy X-ray instruments have a long focal length. The Hard X-ray Imager (HXI) of ASTRO-H has its detector housed in a boom that will extend by about 6 m in orbit so that a focal length of 12 m can be achieved for that instrument. This long structure will inevitably oscillate and flex, especially when passing across the orbital day/night boundary. In order to retain the essential imaging resolution, it is important that the boom has a metrology system that measures this flexion in order to allow post-acquisition compensation in generating the science images. In the current paper, we describe a possible Alignment Monitoring System (AMS) to measure in real time the relative position of the boom. The AMS will be an important element to guaranty that the ASTRO-H observatory will meet its performance requirements.

The Canadian Space Agency has the intention of providing the AMS to the ASTRO-H mission. The current paper reports a study that was conducted to support that intention.

The name “eROSITA” stands for extended Roentgen Survey with an Imaging Telescope Array. The general design of the eROSITA X-ray telescope is derived from that of ABRIXAS. A bundle of 7 mirror modules with short focal lengths make up a compact telescope which is ideal for survey observations. Similar designs had been proposed for the missions DUO and ROSITA but were not realized due to programmatic shortfall. Compared to those, however, the effective area in the soft X-ray band has now much increased by adding 27 additional outer mirror shells to the original 27 ones of each mirror module. The requirement on the on-axis resolution has also been confined, namely to 15 arc seconds HEW. For these reasons the prefix “extended” was added to the original name “ROSITA”. The scientific motivation for this extension is founded in the ambitious goal to detect about 100,000 clusters of galaxies which trace the large scale structure of the Universe in space and time.

The X-ray telescope of eROSITA will consist of 7 identical and co-aligned mirror modules, each with 54 nested Wolter-1 mirror shells. The mirror shells are glued onto a spider wheel which is screwed to the mirror interface structure making a rigid mechanical unit. The assembly of 7 modules forms a compact hexagonal configuration with 1300 mm diameter (see Fig. 1) and will be attached to the telescope structure which connects to the 7 separate CCD cameras in the focal planes. The co-alignment of the mirror module enables eROSITA to perform also pointed observations.

The replication process described in chapter III allows the manufacturing in one single piece and at the same time of both the parabola and hyperbola parts of the Wolter 1 mirror.

The Linearly Variable Filters (LVF) are complex optical devices that, integrated in a CCD, can realize a "single chip spectrometer". In the framework of an ESA Study, a team of industries and institutes led by SELEX-Galileo explored the design principles and manufacturing techniques, realizing and characterizing LVF samples based both on All-Dielectric (AD) and Metal-Dielectric (MD) Coating Structures in the VNIR and SWIR spectral ranges. In particular the achieved performances on spectral gradient, transmission bandwidth and Spectral Attenuation (SA) are presented and critically discussed. Potential improvements will be highlighted. In addition the results of a feasibility study of a SWIR Linear Variable Filter are presented with the comparison of design prediction and measured performances. Finally criticalities related to the filter-CCD packaging are discussed.

The main achievements reached during these activities have been:
- to evaluate by design, manufacturing and test of LVF samples the achievable performances compared with target requirements;
- to evaluate the reliability of the projects by analyzing their repeatability;
- to define suitable measurement methodologies

Future NASA exploratory missions to the Moon and Mars will require safe soft-landings at the designated sites with a high degree of precision. These sites may include areas of high scientific value with relatively rough terrain with little or no solar illumination and possibly areas near pre-deployed assets. The ability of lidar technology to provide three-dimensional elevation maps of the terrain, high precision distance to the ground, and approach velocity can enable safe landing of large robotic and manned vehicles with a high degree of precision. Currently, NASA-LaRC is developing novel lidar sensors aimed at meeting NASA’s objectives for future planetary landing missions under the Autonomous Landing and Hazard Avoidance (ALHAT) project. These lidar sensors are 3-Dimensional Imaging Flash Lidar, Doppler Lidar, and Laser Altimeter. The Flash Lidar is capable of generating elevation maps of the terrain identifying hazardous features such as rocks, craters, and steep slopes. The elevation maps collected during the approach phase between 1000 m to 500 m above the ground can be used to determine the most suitable safe landing site. The Doppler Lidar provides highly accurate ground velocity and distance data allowing for precision navigation to the selected landing site. Prior to the approach phase at altitudes of over 15 km, the Laser Altimeter can provide sufficient data for updating the vehicle position and attitude data from the Inertial Measurement Unit. At these higher altitudes, either the Laser Altimeter or the Flash Lidar can be used for generating a contour map of the terrain below for identifying known surface features such as craters for further reducing the vehicle relative position error.

Conference slides for: The APS+ and intelligent active pixel sensor centered on low power

Ozone loss in the stratosphere, especially its drastic reduction in the Antarctic, has been the main driver for the middle atmosphere research during the last two decades [1]. A good understanding of the relevant processes has been obtained and the dynamical-chemical models are mature enough so that credible predictions of the future of the ozone layer can be made. However, i.e. the effect of the climate change on the pole-to-pole circulation is not clear. The proposed Spectral Imaging of Middle Atmosphere for Climate Change (SIMACC) concept aims to significantly contribute to answering the many open questions in understanding the climate change. The main objective of the proposed SIMACC mission is to investigate the changes the middle atmosphere will undergo during the ongoing rapid climate change [2]. VTT has developed novel Fabry-Perot –based hyperspectral imagers that are to be used in SIMACC [3]-[5]. In these spectral imagers the Fabry-Perot interferometer (FPI) is used at multiple orders simultaneously with each order matched to a specific pixel filter. In the visible (VIS) and very near infrared (VNIR) wavelengths (400 – 1100 nm) the standard Bayer patterned RGB color sensor was used providing the possibility for measuring three spectral bands with one exposure.

The International X-ray Observatory IXO, a candidate follow-on mission for XMM/Newton and CHANDRA, has been studied in Europe in the frame of two parallel industrial assessment studies commissioned by ESA as part of the Cosmic Vision 2015-2025 programme. One of these studies was successfully carried out by EADS Astrium. The present paper is focussed on the results obtained by Astrium GmbH in this study, i.e. on the optical, mechanical and thermal design of the IXO telescope mirror and on the predicted telescope performance.

The IXO telescope mirror consists of nearly 1700 Silicon Pore Optics (SPO) mirror modules [1] accommodated on eight identical petal-shaped support structures. These petals are themselves supported by an optical bench. Design drivers proved to be the allowable mass, the required optical throughput as well as the axial and lateral stiffness required by the launch vehicle. To ensure light-weighting, low thermal gradients as well as robustness during AIV, the petals and the optical bench are deliberately made of a high-modulus carbon fibre material with high thermal conductivity. The telescope optical performance achievable with this design has been analysed based on rigorous finite element, thermal and optical modelling. It has been found to be fully compliant with the required optical throughput and image quality, thus enabling the scientific benefit of the IXO mission.

The mirrors of the International X-ray Observatory (IXO) consist of a large number of high quality segments delivering a spatial resolution better than 5 arcsec. A study concerning the slumping of thin glass foils for the IXO mirrors is under development in Europe, funded by ESA and led by the Brera Observatory. We are investigating two approaches, the “Direct” and “Indirect” slumping technologies, being respectively based on the use of convex and concave moulds. In the first case during the thermal cycle the optical surface of the glass is in direct contact with the mould surface, while in the second case it is the rear side of the foil which touches the master. Both approaches present pros and cons and aim of this study is also to make an assessment of both processes. The thin plates are made of D263glass produced by Schott and are 0.4 mm thick, with a reflecting area of 200 mm x 200 mm; the mould are made of Fused Silica. The adopted integration process foresees the use of reinforcing ribs for bonding together the plates and forming in that way a rigid and stiff stack of segmented mirror shells; the stack is supported by a thick backplane. During the bonding process the plates are constrained to stay in close contact with the surface of the master (i.e. the same mould used for the hot slumping process) by the application of a vacuum pump suction. In this way the spring-back deformations and low frequency errors still present on the foil profile after slumping can be corrected. In this paper we present some preliminary results achieved during the first part of the project.

Global atmospheric chemistry monitoring from space is a key to follow up air quality and provide the repeatable global coverage needed to explain the different mechanisms explaining global warming effects. Space observations with high revisit frequency allow following air mass and pollution clouds movements, both over continents and oceans. Whereas sounders like IASI (Infrared Atmospheric Sounding Interferometer) on board MetOp satellite are mostly designed for meteorological applications, CNES started studying alternative concepts to dedicate new instruments to some selected molecular species. A patent released in 1998 by CNES [1] is at the basis of a static configuration of Michelson interferometers, sampling each interferogram at once. New perspectives to reduce existing instruments in terms of mass, volume and complexity and to create constellations of small satellites or micro-satellites were opened. CNES either proposed this concept for CO2 concentration monitoring (Carbosat and Minicarb missions) in the near infrared or CO and O3 atmospheric profiles retrieval (SIFTI instrument on TRAQ mission [2]) in the thermal infrared. Though TRAQ was finally not selected by ESA, a validation of this concept has been undertaken in parallel with SIFTI phase A studies with the development of a breadboard called MOPI (which stands for ‘Maquette Optique de Performances Infrarouges’ in French, approximately translated into ‘Optical Breadboard for Infrared Performances’). In this paper, we will first describe the instrumental concept. The second part will detail the different conception choices made to design MOPI and, finally, the third part will present the next steps awaited for this breadboard.

The Multi Element Telescope for Imaging and Spectroscopy (METIS) is a coronagraph onboard of Solar Orbiter. It will perform simultaneous observations at HeII Lyman-α line, HI Lyman-α line and in visible. To achieve such capability, instrument mirrors need to be coated by multilayer (ML) structures with high efficiency at all three spectral ranges. Coatings with higher performances with respect standard Mo/Si are desirable. An instrument prototype of METIS has just flown onboard of a NASA sounding rocket: in this case, optics were coated with Mg/SiC MLs. Better performances have been obtained in terms of reflectivity, but long term stability of this coating is an open problem. Moreover the harsh conditions of the environment met during the Solar Orbiter mission given by plasma particles and high temperature could affect the lifetime of the optical components on the long term. We present the design and reflectivity tests of multilayer structures in which performances improvement is obtained by the use of novel capping layers. All multilayers are tuned at 30.4nm line but the design also maximize the performances at 121.6nm and 500 – 650 nm visible range. Analysis of Solar Orbiter environment have been carried on in order to point out the main damaging sources for the nanostructures. Computer simulations with a devoted software have been performed to preliminary evaluation of the possible instabilities in multilayers. Experimental tests for investigating the effects of the thermal heating and particles bombardments in the reflectivity performances have been planned.

The Large Optical Test and Integration Site (LOTIS) Facility at Lockheed Martin Space Systems Company in Sunnyvale, CA has been design specifically to accommodate assembly, integration, and testing of optical payloads from as small as 1 meter class apertures to as large as 6 meters. The facility has successfully reached initial operational capability and a basic overview of the LOTIS Facility has been previously reported including wavefront performance [1], its potential for optical payload testing has been previously reported [2]. The facility has been engineered from the foundation up to provide an ideal single location for optical payloads including details such as crane access, clean room space and levels, vibration levels, and atmospheric turbulence levels within the facility especially the interior of the vacuum chamber containing the 6.5 meter collimator. This paper will present and overview of the facility in general and then present results of the vibration isolation bench performance and resulting line of sight jitter of the collimator, as well as atmospheric turbulence measurements within the chamber.

This paper describes aircraft flight campaign test results for the Imaging spectral signature instrument (ISSI) breadboard developed in the ESA contract 19754/06/NL/PA. The ISSI project was inspired by the Informationefficient hyperspectral imaging sensor (ISIS) [1]. The development and design of the ISSI breadboard is described in [2]. ISSI is a line imaging programmable correlation spectrometer. ISSI is implemented with a front objective, two spectrographs, a liquid crystal display (LCD) spatial light modulator (SLM) and a line sensor. A line imaged by the front objective is dispersed by the first spectrograph on the LCD. Any transmission pattern can be programmed on the LCD. The modulated image is then re-gathered by the second spectrograph to a line on the CCD line detector. Effectively the system performs a dot product between the transmission vector on the LCD and the spectral signature vector of the imaged pixel, where the spectral bins are the components of the vector. Different hyperspectral correlation algorithms, which contain this dot product, can be implemented.

Next-generation infrared astronomical instrumentation for ground-based and space telescopes could be based on MOEMS programmable slit masks for multi-object spectroscopy (MOS). This astronomical technique is used extensively to investigate the formation and evolution of galaxies. MOS with multi-slits is the best approach to eliminate the problem of spectral confusion, to optimize the quality and the SNR of the spectra, to reach fainter limiting fluxes and to maximize the scientific return both in cosmology and in legacy science. Major telescopes around the world are equipped with MOS in order to simultaneously record several hundred spectra in a single observation run. Next generation MOS for space like the Near Infrared Multi-Object Spectrograph (NIRSpec) for the James Webb Space Telescope (JWST) require a programmable multi-slit mask. The EUCLID mission from the European Space Agency (ESA) will study the dark universe by characterizing a very high number of galaxies in shape and in spectrum; this mission has also considered a MOS instrument in its early study phase. Conventional masks or complex fiber-optics-based mechanisms are not attractive for space. The programmable multi-slit mask requires remote control of the multi-slit configuration in real time.

This presentation reports the laboratory tests and calibrations of a novel electro-optically tunable Lyot filter for imaging spectro - polarimetry. The filter’s high-resolution bandpass (λ/▵λ = 3.5 E+3) is centred on filter Fe XIV, 530 3 nm (green line) solar corona emission line Performances of this Liquid Crystal Tunable filter and Polarimeter Results F-P Etalon 530.3 line. Tunable-(LCTP) are also compared with the performances of a classical very high-resolution (λ/▵λ = 2.5 E+4) Fabry-Perot (FP) etalon. The results of the tests show even if the F-P etalon has a higher instrumental resolution, the LCTP has the same effective spectral resolution thanks to its electro-optical fine-tuning capability in wavelength. In addition, the LCTP has the advantage over the F-P etalon of having no mechanically moving parts, polarimetric capability and to do imaging. This project is part of the ESA Startiger activities and part of the preliminary studies for the ASPIICS coronagraph for the PROBA-3 formation flying mission.

In this paper, we report on an on-going feasibility study of a thermal spectro-imager for future space missions to small bodies in the Solar System. Thanks to the recent development of uncooled micro-bolometer technology, we decided to use such a detector as a baseline for our study. For this feasibility study, we have defined four objectives using this detector to perform : (i) uncalibrated images, (ii) calibrated radiometric images, (iii) spectroscopic measurements, and (iv) a definition of the instrumental concept including the optical and mechanical design. Here we present our progress on the first two objectives of this work.

The compact infrared camera (CIRC) is a technology demonstration instrument that Japanese Aerospace Exploration Agency (JAXA) is developing by using commercial-off-the-shelf (COTS) technology in recent years. CIRC is an infrared camera equipped with an uncooled infrared array detector (microbolometer). CIRC will be utilized for various technology demonstrations. This paper introduces the current development status of CIRC and the initial results of characteristics of the CIRC ground test model (Engineering Model; EM) that has been studied with laboratory experiments.