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Register a new userAMBRE: A Compact Instrument to Measure Thermal Ions, Electrons and Electrostatic Charging Onboard Spacecraft
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B. Lavraud
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The Active Monitor Box of Electrostatic Risks (AMBER) is a double-head thermal electron and ion electrostatic analyzer (energy range 0-30 keV) that was launched onboard the Jason-3 spacecraft in 2016. The next generation AMBER instrument, for which a first prototype was developed and then calibrated at the end of 2017, constitutes a significant evolution that is based on a single head to measure both species alternatively. The instrument developments focused on several new subsystems (front-end electronics, high-voltage electronics, mechanical design) that permit to reduce instrument resources down to ~ 1 kg and 1.5 W. AMBER is designed as a generic radiation monitor with a twofold purpose: (1) measure magnetospheric thermal ion and electron populations in the range 0-35 keV, with significant scientific potential (e.g., plasmasphere, ring current, plasma sheet), and (2) monitor spacecraft electrostatic charging and the plasma populations responsible for it, for electromagnetic cleanliness and operational purposes.
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The OSIRIS-REx Thermal Emission Spectrometer (OTES) will provide remote measurements of mineralogy and thermophysical properties of Bennu to map its surface, help select the OSIRIS-REx sampling site, and investigate the Yarkovsky effect. OTES is a Fourier transform spectrometer covering the spectral range 5.71 - 100 {mu}m (1750 - 100 cm-1) with a spectral sample interval of 8.66 cm-1 and a 6.5-mrad field of view. The OTES telescope is a 15.2-cm diameter Cassegrain telescope that feeds a flat-plate Michelson moving mirror mounted on a linear voice-coil motor assembly. A single uncooled deuterated L-alanine doped triglycine sulfate (DLATGS) pyroelectric detector is used to sample the interferogram every two seconds. Redundant ~0.855 {mu}m laser diodes are used in a metrology interferometer to provide precise moving mirror control and IR sampling at 772 Hz. The beamsplitter is a 38-mm diameter, 1-mm thick chemical vapor deposited diamond with an antireflection microstructure to minimize surface reflection. An internal calibration cone blackbody target provides radiometric calibration. The radiometric precision in a single spectrum is <= 2.2 x 10-8 W cm-2 sr-1/cm-1 between 300 and 1350 cm-1. The absolute integrated radiance error is <1% for scene temperatures ranging from 150 to 380 K. The overall OTES envelope size is 37.5 x 28.9 x 52.2 cm, and the mass is 6.27 kg. The power consumption is 10.8 W average. The OTES was developed by Arizona State University with Moog Broad Reach developing the electronics. OTES was integrated, tested, and radiometrically calibrated on the Arizona State University campus in Tempe, AZ.
The Large Yield Radiometer (LYRA) is an XUV-EUV-MUV (soft X-ray to mid-ultraviolet) solar radiometer onboard the European Space Agency PROBA2 mission that was launched in November 2009. LYRA acquires solar irradiance measurements at a high cadence (nominally 20 Hz) in four broad spectral channels, from soft X-ray to MUV, that have been chosen for their relevance to solar physics, space weather and aeronomy. In this article, we briefly review the design of the instrument, give an overview of the data products distributed through the instrument website, and describe the way that data are calibrated. We also briefly present a summary of the main fields of research currently under investigation by the LYRA consortium.
There is a long history of radio telescopes being used to augment the radio antennas regularly used to conduct telemetry, tracking, and command of deep space spacecraft. Radio telescopes are particularly valuable during short-duration mission critical events, such as planetary landings, or when a mission lifetime itself is short, such as a probe into a giant planets atmosphere. By virtue of its high sensitivity and frequency coverage, the next-generation Very Large Array would be a powerful addition to regular spacecraft ground systems. Further, the science focus of many of these deep-space missions provides a ground truth in the solar system that complements other aspects of the ngVLAs science case, such as the formation of planets in proto-planetary disks.
One source of noise for the Laser Interferometer Space Antenna (LISA) will be time-varying changes of the space environment in the form of solar wind particles and photon pressure from fluctuating solar irradiance. The approximate magnitude of these effects can be estimated from the average properties of the solar wind and the solar irradiance. We use data taken by the ACE (Advanced Compton Explorer) satellite and the VIRGO (Variability of solar IRradiance and Gravity Oscillations) instrument on the SOHO satellite over an entire solar cycle to calculate the forces due to solar wind and photon pressure irradiance on the LISA spacecraft. We produce a realistic model of the effects of these environmental noise sources and their variation over the expected course of the LISA mission.
The paper is concerned with examining the effects that design-for-demise solutions can have not only on the demisability of components, but also on their survivability that is their capability to withstand impacts from space debris. First two models are introduced. A demisability model to predict the behaviour of spacecraft components during the atmospheric re-entry and a survivability model to assess the vulnerability of spacecraft structures against space debris impacts. Two indices that evaluate the level of demisability and survivability are also proposed. The two models are then used to study the sensitivity of the demisability and of the survivability indices as a function of typical design-for-demise options. The demisability and the survivability can in fact be influenced by the same design parameters in a competing fashion that is while the demisability is improved, the survivability is worsened and vice versa. The analysis shows how the design-for-demise solutions influence the demisability and the survivability independently. In addition, the effect that a solution has simultaneously on the two criteria is assessed. Results shows which, among the design-for-demise parameters mostly influence the demisability and the survivability. For such design parameters maps are presented, describing their influence on the demisability and survivability indices. These maps represent a useful tool to quickly assess the level of demisability and survivability that can be expected from a component, when specific design parameters are changed.
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