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A composite optical bench made up of Carbon Fiber Reinforced Polymer (CFRP) skin and aluminum honeycomb has been developed for the Tunable Magnetograph instrument (TuMag) for the SUNRISE III mission within the NASA Long Duration Balloon Program. This optical bench has been designed to meet lightweight and low sensitivity to thermal gradient requirements, resulting in a low Coefficient of Thermal Expansion (CTE). In addition to the flight model, a breadboard model identical to the flight one has been manufactured, including embedded fiber Bragg temperature and strain sensors. The aim of this is to explore if the use of distributed fiber Bragg gratings (FBGs) can provide valuable information for strain and temperature mapping of an optical instrument on board a space mission during its operation as well as its on-ground testing. Furthermore, surface-mounted strain FBG sensors and thermocouples have been installed in the optical bench for intercomparison purposes. This paper presents the results obtained from a thermal vacuum test consisting of three thermal cycles with stabilization steps at 100 °C, 60 °C, 20 °C and −20 °C. Experimental results provide information about how FBG embedded temperature sensors can provide a proper and quick response to the temperature changes of the optical bench and that embedded FBG strain sensors are able to measure micro-deformation induced in a close-to-zero CTE optical bench. [ABSTRACT FROM AUTHOR]

Liquid crystal variable retarders (LCVRs) will be used for the first time in a space instrument, the Solar Orbiter mission of the European Space Agency, as polarization states analyzers (PSAs). These devices will determine the Stokes parameters of the light coming from the Sun by temporal polarization modulation, using the so-called modulation matrix O. This is a matrix constituted by the first rows of properly selected PSA Mueller matrices. Calibrating a space instrument, in particular, finding O, is a critical point because in a spacecraft there is no possibility of physical access. Due to the huge difficulty in calibrating the complete instruments in all possible scenarios, a more complete calibration of the individual components has been done in ground in order to make extrapolations to obtain O in-flight. Nevertheless, apart from the individual calibrations, the experimental errors and nonideal effects that inhibit the system to reach the designed and theoretical values must be known. In this work, description and study of these effects have been done, focusing on the nonideal effects of the LCVRs and the azimuthal misalignments between the optical components of the PSA during the mechanical assembly. The Mueller matrix of a representative LCVR has been measured and mathematically decomposed by logarithm decomposition, looking for values of circular birefringence and fast axis angle variations as a function of voltage. These effects, in the absence of other nonidealities, affect the polarimetric performance, reducing the polarimetric efficiencies in some cases until 11%. Nevertheless, in this case, they are negligible if compared to the other nonideality studied, which are the azimuthal misalignments between the PSA optical components. The study presented in this work is key to extrapolate the PSA O matrix if the expected instrumental set-point temperatures are not reached in flight and can be used for the design and implementation of other polarimetric instruments. [ABSTRACT FROM AUTHOR]

Liquid crystal variable retarders (LCVRs) have been extensively used as light polarization modulators for ground-based polarimetric applications. Shortly, LCVRs will be used as polarization state analyzers in two instruments onboard the Solar Orbiter mission of the European Space Agency. Both ground- and space-based polarimeters require LCVR response time values that fulfill the required image acquisition rate of the polarimetric measurements. Therefore, it is necessary to have a reliable method to measure the LCVR optical retardance response times. Response times are usually estimated via optical methods using crossed or parallel polarizers. Nevertheless, these methods measure light intensity transitions to infer the response time instead of directly measuring the changes in the optical retardance. In this work, an experimental setup that uses a Soleil-Babinet variable compensator is proposed. On one hand, this allows one to study the effect of the nonlinear dependence of the light intensity on the optical retardance in the response time determination, which is neglected in most works. On the other hand, the use of the variable compensator allows one to measure the LCVR response times in the highest sensitivity areas of the system that minimizes the uncertainty of the measurement. The six transitions for the Polarimetric and Helioseismic Imager instrument modulation scheme of a representative LCVR have been measured. Based on the results, the optimized conditions to measure response times are found, which can be achieved by using the variable compensator and an IR wavelength (λ = 987.7 nm) as proposed in the experimental setup. [ABSTRACT FROM AUTHOR]

The authors present the results of the ellipsometric characterization of thin layers of bismuth and aluminum oxide deposited over the waist of a tapered optical fiber by pulsed laser deposition. The characteristics of the deposits are studied by spectroscopic ellipsometry. From the effective thicknesses determined by the ellipsometric characterization, it is shown by simulations that surface plasmon resonances (SPRs) can occur in the fiber device, and it is demonstrated experimentally. These results show the feasibility of employing bismuth as a plasmonic material in SPR fiber sensors based on doubly-deposited uniform-waist tapered optical fibers, which show excellent performance and versatility. [ABSTRACT FROM AUTHOR]

The Full Disk Telescope is part of the Polarimetric Helioseismic Instrument on board the future Solar Orbiter ESA/NASA mission. It will provide full-disk measurements of the photospheric magnetic field vector and line-of-sight velocity, as well as the continuum intensity in the visible wavelength range. Along this mission, it is expected that thermal drifts will induce image focus displacements. Consequently, providing an autofocus system is mandatory to prevent image degradation. The refocusing system is based on an autonomous image contrast analysis and it allows for a lens displacement in order to locate the best focus position. The figure of merit chosen for the image quality evaluation is presented. The influences of attitude instability and mechanical uncertainties are considered in a refocusing process simulation. In addition, an engineering model of the mechanism is tested at flight operating conditions. To check its performance, an optical interrogation system is set up. Determination of accuracy and repeatability of the mechanism positioning is experimentally evaluated and discussed according to the ISO standard. The results show that the proposed refocusing system is sufficiently robust against the expected image shifts and mechanical instabilities. [ABSTRACT FROM AUTHOR]

The Multi Element Telescope for Imaging and Spectroscopy (METIS) is one of the remote sensing instruments to be onboard the future NASA/ESA Solar Orbiter mission. The science nominal mission orbit will take the spacecraft from 0.28 to 0.95 astronomical units from the Sun, setting challenging and variable thermal conditions to its payload. METIS is an inverted-occultation coronagraph that will image the solar corona in the visible and UV wavelength range. In the visible light path a Polarization Modulation Package (PMP) performs a polarimetric analysis of the incoming solar light. This PMP is based on liquid crystal variable retarders (LCVR) and works under a temporal modulation scheme. The LCVRs behavior has a dependence on temperature and, as a consequence, it is critical to guarantee the PMP performance in the mission thermal environment. Key system specifications are the optical quality and the optical retardance homogeneity. Moreover, the thermally induced elastic deformations of the mechanical mounts and the LCVRs shall not produce any performance degradation. A suitable thermal control is hence required to maintain the system within its allowed limits at any time. The PMP shall also be able to reach specific set-points with the power budget allocated. Consequently, and in order to verify the PMP thermal design, we have experimentally reproduced the expected thermal flight environment. Specifically, a thermal-vacuum cycle test campaign is run at the different mission operational conditions. The purpose is both to check the stability of the thermal conditions and to study the optical quality evolution/degradation. Within this test transmitted wavefront measurements and functional verification tests have been carried out. To do that we adapted an optical interrogation scheme, based on a phase shifting interferometric technique, that allows for inspection of the PMP optical aperture. Finally, measurements obtained at non-operational temperature conditions are also shown. These results demonstrate that the device meets the specifications required to perform its operational role in the space mission environment. [ABSTRACT FROM AUTHOR]