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What were the precedessor missions ofr ATHENA at L2?
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ESA now has much engineering and operational experience regarding Lagrange Point 2 missions, with the successful flight of Herschel, Planck and Gaia, and the forthcoming missions Euclid and Plato.
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Herschel | Planck | Gaia | Euclid | Plato
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CDFReport_Athena.pdf
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CDF Report
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When did the ATHENA study take place?
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This study has been requested by SRE-FM and financed by the General Studies Program (GSP). It was carried out in the CDF by a team of specialists from ESTEC and ESOC in 9 sessions starting with a kick-off on the 16th September 2014 and ending with an Internal Final Presentation on the 23rd October 2014.
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From 16th September 2014 to 23rd October 2014
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CDFReport_Athena.pdf
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CDF Report
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Which launcher will ATHENA use?
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A summary of the mission architecture that resulted from the CDF evaluation is given in the IDEF0 shown in the following figure. More details can be found in the Concept of Operations document RD[7]. The Mission involves an Ariane 5 (now 6) launch to a large-amplitude Halo orbit around L2
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Ariane 5
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CDFReport_Athena.pdf
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CDF Report
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What is the wet mass at launch of ATHENA?
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The launch wet mass of ~6.1t is significantly higher than ATHENA_L1 (~4t), which can be understood considering the significantly higher MAM mass, and also the increased mass of the PL, with associated knock-on effects for the SC.
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~6.1t
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CDFReport_Athena.pdf
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CDF Report
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Why are space-based observations in the X-Ray band needed?
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Because most of the baryonic component of the Universe is locked up in hot gas at temperatures of around a million degrees, and because of the extreme energetics of the processes close to the event horizon of black holes, understanding the Hot and Energetic Universe requires space-based observations in the X-ray band.
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Because most of the universe is locked up in hot gas | because of the extreme energetics of the processes
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CDFReport_Athena.pdf
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CDF Report
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What is ATHENA observing?
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ATHENA will predominantly perform pointed observations of celestial targets. There will be around 300 such observations per year, with durations ranging from 1 ks to 1 Ms, with typical duration of 100 ks per pointing. This routine observing plan will be interrupted by ToO (e.g. GRBs and other transients) observations at a typical rate of twice a month.
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Celestial targets | Pointed observations of celestial targets
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CDFReport_Athena.pdf
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CDF Report
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How can the ATHENA orbit be reached?
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Such an orbit can be reached via a so called ‘free’ transfer trajectory, not requiring any deterministic orbit insertion manoeuvre after Earth departure. The SC travels on the so called stable manifold toward its operational orbit about SEL2. A typical transfer trajectory on the stable manifold of the target orbit is depicted in Figure 4-2. The full stable manifold of the target orbit is shown. Some parts of the manifold intersect with the near-Earth environment (the Earth is at the origin), where the launcher can place the SC on the stable manifold of the target orbit.
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via a so called 'free' transfer trajectory | via a so called 'free' transfer trajectory, not requiring any deterministic orbit insertion manoeuvre after Earth departure
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CDFReport_Athena.pdf
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CDF Report
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What are the advantages of the orbit?
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The chief advantages of orbits about SEL2 are (I) a constant thermal environment, since they can be designed to be eclipse free, and (II) a limited communication distance. Another advantage for astronomy missions is that the Sun, Earth and Moon are all located in one hemisphere as seen from the SC.
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a constant thermal environment | limited communication distance | Sun, Earth and Moon are all located in one hemisphere
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CDFReport_Athena.pdf
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CDF Report
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Which elements are contained in the configuration of the ATHENA instrument?
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The configuration of the instrument includes the cryostat, the electronics for the detector, the electronics to operate the cryo-coolers and control electronics. The detector is located inside a Dewar, which is cooled by a set of cryo-coolers with its dedicated drive electronics. The Dewar, together with its cryo-coolers is called the X-IFU cryo. The Front End Electronics (FEE) is galvanically well connected to the detector. The amplified signals from the FEE are passed to the Digital Electronics / Event Processing (DE/EP) unit where the data is digitized, the feedback signal is generated and the (demodulated) triggered event data, including triggers from the anti-coincidence detector are selected and subsequently processed to extract the relevant event parameters (e.g. energy, time stamp, and event grade).
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cryostat |electronics for the detector | electronics to operate the cryo-coolers | control electronics
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CDFReport_Athena.pdf
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CDF Report
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What is the effective area?
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The effective area is the area that must be used when calculating the physical properties of sources in the sky (e.g. flux, surface brightness). Reflectivity and vignetting, among other effects, cause the geometric area of a telescope to be reduced to a smaller "effective area".
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area that must be used when calculating physical properties of sources in the sky
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CDFReport_Athena.pdf
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CDF Report
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How is the image quality defined for ATHENA?
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For ATHENA the image quality was specified in terms angular resolution, and characterised with the Half Energy Width (HEW) of the Point Spread Function (PSF) at different energy point sources. Table 7-1 aggregates the angular resolution requirements for the different energy ranges.
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in terms of angular resolution | characterised with the Half Enerhy Width (HEW)
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CDFReport_Athena.pdf
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CDF Report
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Why can ATHENA tests not be done with the complete spacecraft?
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Furthermore the FPM is shown as a separate unit, and then the middle part of the FMS. This middle part of the FMS does not have any electronic units instrumentation, it is only the structure between the upper and lower part of the SC. It does include some internal baffles. The separation in this fashion is needed for AIV and transportation purposes. The SC is very large, and test centres for various tests (e.g. mechanical, thermal, functional) cannot all be performed with the complete SC assembly.
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the SC is very large
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CDFReport_Athena.pdf
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CDF Report
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Why were ceramic materials not used?
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Besides those design parameters, also the manufacturability plays a major role in choosing the MS material. In terms of potential MS materials only titanium and CFRP were considered. Ceramic materials were a priory ruled out for reasons of too high sensitivity to shock loads, i.e. too low fracture toughness. Since the MMA is positioned nearby the SC interface the shock loads due to clamp band release will be high and might be detrimental for a ceramic MS.
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for reasons of too high sensitivity to shock loads | for reasons of too high sensitivity to shock loads i.e. too low fracture toughness
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CDFReport_Athena.pdf
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CDF Report
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What is the material of the ATHENA mirror structure?
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Although titanium is inferior to CFRP for what concerns specific stiffness, CTE and ΔCTE, titanium was still chosen as the baseline material for the mirror structure. The main reason was the easier, cheaper and more precise manufacturing procedure. A titanium MS could be manufactured by milling of a monolithic forging or a set of identical forged segments. Large titanium forgings have been manufactured in the US for the F22 raptor, see Figure 9-2a. Another viable option would be to manufacture it using Additive Layer Manufacturing (ALM), also referred to as 3D printing, see Figure 9-2b.
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titanium | titanium was still chosen as the baseline material for the mirror structure
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CDFReport_Athena.pdf
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CDF Report
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How is the ATHENA mirror structure manufactured?
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The architecture of the MMA is depicted in Figure 9-4. The mirror structure (MS) is made of titanium and is either milled out of a monolithic titanium forging or is manufactured by 3D-printing (Additive Layer Manufacturing, ALM). The MS is populated with 15 rows of MMs and is attached to the SVM drum by means of HDRMs. During launch all loads will pass through the HDRM bipods while in orbit the bipods are disconnected by 6 Non-Explosive Actuators (NEA). From that moment onwards the MS will be kept in position by a hexapod mechanism which consists of 6 actuators mounted parallel to the bipod struts.
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milled out of monolithic titanium forging | manufactured by 3D-printing (Additive Layer Manufacturing, ALM)
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CDFReport_Athena.pdf
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CDF Report
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How much power is needed for thermal control?
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With the proposed 12 heater zones, it is possible to maintain the MMs within the required temperature range of 20±1 ˚C with a Active Thermal Control consumption of ~2500W. The requirement to keep the temperature differences in radial direction within the requirement of ΔT < 3K can be achieved by tuning the dissipated heater power for each zone. This gradient can also be decreased by further optimising the heater zoning, as suggested by the dashed lines in the zoning scheme described above
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Active Thermal Control consumption of ~2500W
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CDFReport_Athena.pdf
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CDF Report
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Why is the battery not sized for the launch phase?
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It is not necessary to size the battery for the launch phase plus TCM#1 in series, because the intervening time period (min. 24 hours) will be long enough to fully recharge the battery.
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the intervening time period (min 24 hours) will be long enough to fully recharge the battery
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CDFReport_Athena.pdf
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CDF Report
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What is the size of the solar array on ATHENA?
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This results in a total of 30m2 array area, 154 kg total mass. This is assumed to be implemented as 2 fixed deployable wings, each 15m2, 77 kg.
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30m2 |2 fixed deployable wings, each 15m2
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CDFReport_Athena.pdf
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CDF Report
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How is the instrument line of sight defined?
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It is important also to define the instrument line of sight, which is the line passing through the centre of the detector and the mirror node (the centre of the mirror assembly).
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line passing through the centre of the detector | line passing through the centre of the detector and the mirror node (centre of the mirror assembly)
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CDFReport_Athena.pdf
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CDF Report
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When will ATHENA be launched?
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The main challenge for this project is to find technical solutions fitting to the cost frame work. Next to that all technologies with low TRL need to be advance such that they reach TRL 6 (RD[47]) before the start of the implementation phase. This could become difficult, although the launch is only required in 2028, because some parts of the X-Ray Field Unit are reported to be only at TRL 2 presently.
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in 2028 | launch is only required in 2028
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CDFReport_Athena.pdf
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CDF Report
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How long will the NG-CryoIRTel mission last?
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The operational mission is 3 years in duration + 2 years potential extension, observing core and observatory targets within +/- 10deg viewing zone.
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3 years | 3 years in duration + 2 years potential extension
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CDFReport_NGCryoT.pdf
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CDF Report
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Who supported ESA in the NG-CryoIRTel study?
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The study was requested by ESA Science Directorate SRE-FM and funded by the General Studies Programme GSP. The study was carried out in 8 sessions, starting with a kick off on the 13th November 2014 and ending with an Internal Final Presentation on the 18th December 2014 by an interdisciplinary team of specialists from ESTEC and ESOC and supported by SPICA/SAFARI experts from JAXA and SRON.
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SPICA/SAFARI experts | SPICA/SAFARI experts from JAXA and SRON
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CDFReport_NGCryoT.pdf
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CDF Report
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Who funded the NG-CryoIRTel study?
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Requested by SRE-FM and funded by the General Studies Program (GSP) the NG-CryoIRTel study was performed in the Concurrent Design Facility (CDF) in eight sessions with participation of JAXA and SRON, starting with a kick-off on 13 November 2014 and finishing with an internal final presentation on 18 December 2014. The sessions were supplemented with various smaller working sessions and meetings.
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the General Studies Program (GSP)
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CDFReport_NGCryoT.pdf
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CDF Report
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What wavelengths can be observed by NG-CryoIRTel?
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NG-CryoIRTel will cover the full 20 to 210 μm wavelength range, including the missing 28 μm to 55 μm octave which is out of the Herschel and JWST domains with unprecedented sensitivity and spatial resolution. Hence, NG-CryoIRTel will be the only observatory of its era to bridge the wavelength gap between JWST and ALMA, and carry out unique science programs.
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20-210μm | the full 20-210μm wavelength range
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CDFReport_NGCryoT.pdf
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CDF Report
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Where will NG-CryoIRTel be launched from?
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The launch is envisioned on a Japanese H-X launch vehicle from the Tanegashima spaceport in Japan. The H-X launcher can lift more than 3,700 kg (payload excluding payload adapter) into the transfer orbit towards SEL2. The initial ascent is into a circular parking orbit with an inclination of 30 Deg RD[4]. The drift duration in the circular parking orbit until the final upper stage burn for a transfer towards SEL2 determines the final argument of perigee of the departure orbit. The drift duration can be selected to optimise the yearly launch window duration. The powered ascent phase is followed by an upper stage re-orientation phase in case a specific separation attitude is required, e.g. Sun pointing of the solar panels prior to separation.
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Tanegashima | Tanegashima spaceport in Japan
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CDFReport_NGCryoT.pdf
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CDF Report
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When will the NG-CryoIRTel be transferred to its operational orbit?
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The commissioning of the S/C and its transfer to the operational orbit should be completed within the first three months. And before starting the nominal operations phase, the S/C needs to finish its cooling process and go through a two months Instrument Performance Verification and Science Demonstration phase.
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within the first three months
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CDFReport_NGCryoT.pdf
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CDF Report
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What is the size of the telescope?
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The NG-CryoIRTel study was done in two stages. The main part of the study was dedicated to the baseline design of a 2 m size telescope at 6K and the corresponding assessment of all its subsystems.
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2m | 2m size telescope
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CDFReport_NGCryoT.pdf
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CDF Report
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What temperature does the detector need to be cooled to?
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The science case, as detailed in the sections hereunder, calls for photometry and medium to high resolution spectroscopy between 20 and 210 μm. Detectors operating at these wavelengths require active cooling to 6K and 2K (Si:As and Si:Sb technology) and to 50mK (TES Transition Edge Sensor technology).
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6K | 6K and 2K | active cooling to 6K and 2K
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CDFReport_NGCryoT.pdf
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CDF Report
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What is the purpose of the thrust cone structure?
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The primary structure of the SVM is composed of a thrust cone which also constitutes the load path between the PLM and the launcher, and shear panels. The thrust cone structure will interface the H-X launcher interface at 2360mm diameter. The secondary structure made by the upper and lower platform and side panels accommodates all spacecraft equipment and the cryo-cooler units. The lower platform serves as sunshield and solar panels. Solar cells are mounted on the panel in- and out-side the thrust cone. The octagonal shaped spacecraft body has 940mm height dimension. The upper diameter of the thrust cone of 3m diameter is mainly driven by the optical design dimension. The thrust cone will then have an inverted shaped cone. The outer diameter of the upper and lower platform dimensions is limited to 4.5m by the H-X fairing (which allows a max. diameter of 4.6m).
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constitutes the load path | constitutes the load path between the PLM and the launcher| will interace the H-X launcher interace |will interface the H-X launcher interface at 2360mm
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CDFReport_NGCryoT.pdf
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CDF Report
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Where is the sunshield panel mounted?
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At a distance of 200 mm above the spacecraft to launch vehicle interface the sunshield panel is mounted horizontally to the thrust cone. In fact this panel consists of two parts, one positioned outside of the thrust cone and one positioned inside the thrust cone. The sunshield is also used as a solar panel and therefore its entire area is equipped with solar cells. On the upper side of the thrust cone the SVM top panel is mounted on the outside of the thrust cone. On the inside the thrust cone is open.
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at a distance of 200mm above the spacecraft to launch vehicle interface
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CDFReport_NGCryoT.pdf
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CDF Report
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Why are the bipods made out of CFRP?
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The bipods which hold the PLM assembly are made of GFRP to reduce the conductive heat loads induced by the warm SVM. The length of the large bipod struts that connect to the TOB is 1.43 m. The short bipod struts have a length of 0.53 m and connect to the bottom side of the metering structure. Thin CFRP bipods are placed parallel to the GFRP main bipod struts to hold the PLM in position in orbit when the main bipod struts will be conductively decoupled at the PLM/SVM interface. The concept is similar to the bipod concept used on the GAIA spacecraft (see small picture inset in Figure 9-2). The metering structure itself is a monocoque CFRP beam-type structure which together with the TOB forms the backbone of the PLM telescope structure.
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to reduce the conductive heat loads | to reduce the conductive heat loads induced by the warm SVM
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CDFReport_NGCryoT.pdf
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CDF Report
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What does the NG-CryoIRTel refocusing mechanism do?
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The refocusing mechanism provides the motion along multiple degrees of freedom (DoF), in general 5, in order to accurately locate the secondary mirror of the telescope and keep it stable. This refocusing mechanism can therefore correct possible misalignments coming from the assembly residual errors, deformations due to environmental effects from ground to orbit etc. Its main specifications define the resolution, range, accuracy with respect to a reference position, lifecycle (number of full strokes) and operative temperature range. Other important characteristics are the capability to produce a limited heat dissipation, survive launch loads, compactness and light weight.
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Provide motion | Provide motion along mutliple degrees of freedom (DoF) in general 5
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CDFReport_NGCryoT.pdf
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CDF Report
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What are the advantages of the Hexapod solution?
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For the current study, the Hexapod solution is selected. The advantages of greater modularity and stiffness have been regarded as particularly attractive.
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Greater modularity | Greater stiffness | Greater modularity and stiffness
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CDFReport_NGCryoT.pdf
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CDF Report
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How does the baselinde design provide accuracy?
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The baseline design does not use displacement sensors. The accuracy is provided by the stability of the steps of the motor, and the position accuracy of the reference points defined by the limit switches.
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stability of the steps | stabilitiy of the steps of the motor | stability of the steps of the motor and the position accuracy of the reference points
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CDFReport_NGCryoT.pdf
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CDF Report
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How can plume impingement of the thrusters be avoided?
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In order to avoid the plume impingement of the thruster into the instrument, all thrusters are pointing opposite to the payload as shown in Figure 13-1 and this therefore results in an unbalanced solution (similarly to Herschel).
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pointing opposite to the payload
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CDFReport_NGCryoT.pdf
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CDF Report
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What is the category of the NG-CryoIRTel mission?
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The NGCryoIRTel mission can be classified as a category A (spacecraft-Earth surface distance < 2∙106 km) mission according to ECSS standards. Furthermore the radio service used for telemetry, tracking and command (TT&C) is classified as Space Research (SR) service.
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category A | category A (spacecraft-Earth surface distance < 2∙106 km) | category A mission according to ECSS standards
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CDFReport_NGCryoT.pdf
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CDF Report
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Where are the ground-stations needed for NG-CryoIRTel mission?
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In order to be able to transmit the large volume of scientific telemetry produced every day one daily communication pass between an Earth ground-station and the telescope of 8 hours is assumed [MR-OGS-140]. Due to high amplitude halo orbit, this can only be guaranteed using at least two ground stations: one on the northern and one on the southern hemisphere. These ground stations will need to be switched on a seasonal schedule.
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on on the northern | one on the southern hemisphere| one on the northern and one on the southern hemisphere
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CDFReport_NGCryoT.pdf
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CDF Report
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What size do the groundstations need to be?
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Due to the long range and associated high path losses on the RF-signals (~225dB in S-Band, ~236dB in X-Band and ~248dB in K-Band), the assumed ground stations must have a diameter of at least 35m in order to guarantee the necessary transmission and reception gains.
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at least 35m
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CDFReport_NGCryoT.pdf
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CDF Report
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When shall the instruments be delivered?
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The flight units of the instruments shall be delivered at least 30 months before the start of the launch campaign
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30 months before the start of the launch campaign
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CDFReport_NGCryoT.pdf
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CDF Report
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Where is the satellite AIV taking place?
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Baseline for the study is that the satellite AIV is performed in Europe. The satellite structural qualification is performed with a Structural Qualification Model which is then refurbished to a Cryogenic Qualification Model. The cryogenic tests could be performed in CSL as for Herschel and Planck.
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in Europe | in CSL
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CDFReport_NGCryoT.pdf
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CDF Report
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What program is the MarsFAST rover part of?
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In this framework, the ESA Mars Robotic Exploration Preparation (MREP) program is now exploring the possibility for Europe to contribute to a mission with NASA reusing the Sky-crane concept, following the successful landing of NASA MSL mission. The European contribution will mainly consist of a surface element, the MarsFAST rover, able to demonstrate European fast and autonomous mobility technologies.
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Mars Robotic Exploration Preparation (MREP)
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CDFReport_MarsFast.pdf
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CDF Report
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Who requested the MarsFAST study?
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The MarsFAST rover study has been requested by the ESA Science Directorate (SRE-FMP) and financed as part of the General Studies Program. MarsFAST is a "fast" mobile and autonomous 150 kg class rover which represents a building block for the potential Mars Sample Return (MSR) mission, as well as providing in-situ science in support to future Mars robotic exploration.
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ESA | ESA Science Directorate (SRE-FMP)
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CDFReport_MarsFast.pdf
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CDF Report
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When can dust storms occur during the MarsFAST mission?
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Note: Local dust storms could occur at any time during the MarsFAST surface lifetime. This reference timeline assumes a contingency allocation of 14 sols (i.e. MarsFAST hibernation mode) at any time during Nominal Operations. Should a local dust storm happen at any other surface mission phase (e.g. Deployment, Egress, Commissioning and Early Operations), the mission timeline can be adjusted accordingly.
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any time | any time during the MarsFAST surface lifetime
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CDFReport_MarsFast.pdf
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CDF Report
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What is the goal of the MarsFAST mission?
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The technical demonstration goal of this mission, to cover a long distance drive within a few weeks only, supports the scientific goal to investigate multiple and different places at similar boundary (= weather) conditions. However, the possible requirement to enter “difficult” terrain like the edges of dune field or potentially soft old river beds must be carefully analysed and traded against the inherent risks.
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to cover a long distance drive |to cover a long distance drive within a few weeks | to investigate multiple and different places |to investigate multiple and different places at similar boundary (=weather) conditions
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CDFReport_MarsFast.pdf
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CDF Report
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Where is the panoramic camera mounted?
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The camera package combines two different systems. The panoramic stereoscopic camera contains a pair of identical wide-angle cameras (WAC) with a field of view (FoV) of 48 degree. The high resolution camera (HRC) with a FoV of 8.8 deg is designed for imaging specific features on the landscape. Both cameras will be mounted on a deployable mast and point in the same direction.
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on a deployable mast
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CDFReport_MarsFast.pdf
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CDF Report
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Why can the sample material not be exposed to daylight?
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After collection of the sample material from a depth greater than 2 cm it must not be exposed to daylight. The UV component of the light would empty the charges relatively fast and destroy the contained information. The sample will be exposed to different laser sources at different wavelengths within the instrument. The release of charges creates photons that are analysed and their amount reflects the total dose accumulated after last exposure.
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UV component of the light would empty the charges | UV component of the light would empty the charges relatively fast and destroy the contained information
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CDFReport_MarsFast.pdf
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CDF Report
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What is the purpose of the tunable laser spectrometer?
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The tunable laser spectrometer will detect trace concentration of water and volatiles. The laser light sources are tunable laser diodes; each laser diode is tuned to a spectral region which includes absorption lines of interest for the relevant gases. Several diodes can be used to allow optimum sensitivity to different gases, e.g. water and methane.
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detect trace concentration of water | detect trace concentration of water and volatiles
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CDFReport_MarsFast.pdf
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CDF Report
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When will the cruise phase take place?
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The cruise phase would begin when the spacecraft separates from the launch vehicle and end prior to entry, descent, and landing (EDL). The ESA rover and platform would be enclosed inside an aeroshell during cruise.
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when the spacecraft separates from the launch vehicle | when the spacecraft separates from the launch vehicle and prior to entry, descent and landing (EDL)
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CDFReport_MarsFast.pdf
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CDF Report
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Why does the platform have cut-outs?
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The platform itself (see figures) would consist of a roughly hexagonal pallet with a honeycomb top deck and a six-outrigger support structure. The landing mechanism would consist of six deployable hinged panels consisting of honeycomb crushable material with six edge-protection airbags that would deploy behind the crushable panels during final descent. These airbags would be retracted after landing to allow the egress hardware to be deployed. The platform has cut-out areas to accommodate the descent stage hardware, and these cut-outs have been increased in size near the rover to accommodate better the rover egress hardware. Depending on potential NASA science payloads, additional adjustments in the platform (including possible cut-outs) may need to be made.
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to accompdat the descent stage hardware
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CDFReport_MarsFast.pdf
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CDF Report
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What provides power to the rover during cruise?
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During cruise, the cruise stage solar arrays would be used to supply power to the ESA rover. At 24 hours prior to EDL, the ESA rover would be required to survive on its own power (with no additional power from the cruise/EDL/platform system) until the rover could deploy its solar arrays no earlier than one hour post landing. It is assumed that there would be enough time post landing and prior to sunset for both the platform and the rover to power their respective batteries for surviving the first night.
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cruise stage | cruise stage solar arrays
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CDFReport_MarsFast.pdf
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CDF Report
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Which heritage can be used in the design of the Mars lander thermal control system?
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The lander thermal control system would leverage MER and MSL heritage with previous studies done for the MAX-C/ExoMars concept. Survival heating power requirements during winter (depending on latitude and payload requirements) may drive the design to use RHUs, and future studies would address this issue.
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MER and MSL heritage | MER and MSL heriateg with previous studies done for the MAX-C/ExoMars concept
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CDFReport_MarsFast.pdf
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CDF Report
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When can rover operations start?
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Rover operations can only start when enough power is available for the mode and the environmental temperature allows the required equipment to be within the operating temperature limits, specially for those not pre-heated.
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when enough power is available
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CDFReport_MarsFast.pdf
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CDF Report
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What speed can by achieved by the MarsFAST rover?
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During this demonstration the rover speed of 1.8 cm/s is representative of the required speed of the fast roving capability. Demonstrating this capability over 10 sols after a previous mission duration of 100 sols, provides a representative validation under the given environmnental conditions, such as solar cell efficiency degradation, wear of the wheels and mechanisms and battery performance.
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1.8cm/s
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CDFReport_MarsFast.pdf
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CDF Report
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Why is the rover top part larger than the bottom part?
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The locomotion system is connected to the rover body in 3 locations shown in Figure 8-6. This interface location drives the rover body bottom part dimension of 790mm by 500mm (see chapter 8.4). 5 layers of solar panels are placed on top of the rover body. The area of the rover top part is larger than the bottom part to support the solar panels stack namely 910mm by 500mm (see chapter 8.4). Front part of the rover accommodates robotic arm, OSL payload, camera mast assembly and LGA. Meteo Pack assembly, hazard camera and LGA are located on the back side of the rover. Camera mast supports three different cameras: stereo-, navigation- and high resolution camera. The mast supports also the HGA of 35cm diameter.
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to support the solar panels | to support the solar panels stack namely 910mm by 500mm
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CDFReport_MarsFast.pdf
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CDF Report
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What releases the HDRM mechanism?
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The NEA HDRM is an electrically initiated, one-shot release mechanism that has the ability to carry a very high tensile preload until commanded to release. The preload is applied through a release rod held in place by two separable spool halves which are in turn held together by tight winding of restraining wire. The restraint wire is held in place by redundant electrical fuse wires; actuation of either circuit allows release, assuring maximum reliability. When sufficient electrical current is applied, the restraint wire unwinds allowing the spool halves to separate releasing the release rod and the associated preload.
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actuation of either circuit
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CDFReport_MarsFast.pdf
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CDF Report
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Where is the rover attached?
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The rover body is attached in the corners at discrete hold-down points by 4 bipod supports. These are arranged such that the normals to the planes through each bipod pass through a common central point to provide an iso-static (or kinematic) mounting system. This minimises constraint loads at the interface arising from differential strains between the Rover and Lander at separation. Each bipod provides constraints in the vertical and tangential translational degrees of freedom. 3 bipods would provide a true iso-static mounting system (6 DOFs constrained), however the 4 point arrangement has been baselined because the 2 front corner bipods provide a stiff support to the side bogie pivots and because of the difficulty in accommodating the 3rd bipod interface at the rear close to the rear bogie telescopic deployment mechanism.
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in the corners |in the corners at discrete hold-down points
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CDFReport_MarsFast.pdf
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CDF Report
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How are the panels deployed?
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The actual deployment of the panels is done by spring actuated hinges. The torque to actuate the hinges shall overcome the mass of the panels and the resistive torques in hinges. The resistive torques of the hinges will possibly be increased by a sealing system which should prevent any dust or sand accumulating in the journal bearings. Therefore a resistive torque of 2 Nm has been assumed per hinge line. Normally a torque order of magnitude lower would be expected.
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by spring actuated hinges
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CDFReport_MarsFast.pdf
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CDF Report
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What size of obstacle can the rover travel over?
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In theory, the rover is able to travel over rocks or obstacles which are the size of the wheel diameter (6 wheeled platform). Following this rule it‟s easy to conclude that the bigger the wheel, the fewer obstacles the rover needs to avoid which will end up in shortening the path to get to final destination.
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rocks or obstacles which are the size of the wheel diameter
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CDFReport_MarsFast.pdf
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CDF Report
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What distance can the MarsFAST rover travel?
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From the mission requirements it is derived that for the worst case scenario the rover will need to traverse a total ground track distance per sol of 255 m (for a traversability factor of 1.4), which considering 4h of traverse per sol requires the rover to navigate on Mars at an average speed of 1.8 cm/s. This and the rest of the subsystem requirements that are derived from here are detailed in the table below:
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255m | per sol 255m
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CDFReport_MarsFast.pdf
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CDF Report
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Where will MarsFAST land?
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The baseline mission scenario was to land between latitude -5° to 20° and to operate for 180 SOLs without RHU. There were no constraints on the longitude, therefore areas on Mars with low surface thermal inertia and high Albedo will have large temperature excursions which can be beyond the capabilities of existing and conceptual solar mission designs with RHUs. Therefore, it is required to limit the landing sites to areas between latitude -5° to 20° with a thermal inertia above 150, which is currently the baseline on Exomars. Otherwise, the energy consumption to maintain the units above their minimum operating temperature could be too high for the power sub-system.
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between latitude -5° to 20°
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CDFReport_MarsFast.pdf
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CDF Report
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Dataset Owner(s):
expert.ai Research Lab
License/Terms of Use
This dataset is licensed under the Creative Commons Attribution 4.0 International License (CC BY 4.0) available at https://creativecommons.org/licenses/by/4.0/legalcode.
How to cite
To cite this research please use the following:
@inproceedings{10.1145/3477495.3531697,
author = {Garcia-Silva, Andres and Berrio, Cristian and Gomez-Perez, Jose Manuel and Mart\'{\i}nez-Heras, Jose Antonio and Donati, Alessandro and Roma, Ilaria},
title = {SpaceQA: Answering Questions about the Design of Space Missions and Space Craft Concepts},
year = {2022},
isbn = {9781450387323},
publisher = {Association for Computing Machinery},
address = {New York, NY, USA},
url = {https://doi.org/10.1145/3477495.3531697},
doi = {10.1145/3477495.3531697},
abstract = {We present SpaceQA, to the best of our knowledge the first open-domain QA system in Space mission design. SpaceQA is part of an initiative by the European Space Agency (ESA) to facilitate the access, sharing and reuse of information about Space mission design within the agency and with the public. We adopt a state-of-the-art architecture consisting of a dense retriever and a neural reader and opt for an approach based on transfer learning rather than fine-tuning due to the lack of domain-specific annotated data. Our evaluation on a test set produced by ESA is largely consistent with the results originally reported by the evaluated retrievers and confirms the need of fine tuning for reading comprehension. As of writing this paper, ESA is piloting SpaceQA internally.},
booktitle = {Proceedings of the 45th International ACM SIGIR Conference on Research and Development in Information Retrieval},
pages = {3306–3311},
numpages = {6},
keywords = {space mission design, reading comprehension, open-domain question answering, neural networks, language models, dense retrievers},
location = {Madrid, Spain},
series = {SIGIR '22}
}
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