Launch Opportunity Research Articles

Scale Up and Coupling of the MOXIE Solid Oxide Electrolyzer for Both Terrestrial and Space Applications

The Mars Oxygen ISRU Experiment (MOXIE) is a first of its kind demonstration of in-situ resource utilization technology to produce propellant and breathable oxygen from the Mars ambient carbon dioxide. The use of Lunar and Martian resources represents a significant opportunity to reduce the cost of launch from Earth, enabling propellant production for space refueling, and allowing for life support of manned missions to the Lunar and Martian surfaces.Since developing the Solid OXide Electrolysis (SOXE) stacks for the Mars 2020 MOXIE program in 2017, the OxEon team has made significant advancements in scale and capabilities of the SOXE stack technology used in that system. Newer variants have a five-fold larger cell area and a 6.5-fold increase in cells per stack, for a stack scaled 33-times the 0.5% scale of the device in MOXIE. Six of these OxEon mission-scale SOXE stacks will produce 30 tons of propellant oxygen to fuel a Mars Ascent Vehicle (MAV) in the 19-month window between landing an unfueled MAV pre-supply mission and the next launch opportunity for the first crewed Mars Mission, meeting target requirements for a return mission. OxEon has built and demonstrated systems with mission-scale stacks for both Lunar and Martian applications. A system for the production of propellant H2 and O2 from Lunar ice was successfully tested in a cryo-vac chamber at the Colorado School of Mines in 2022. Another mission-scale demonstration system is scheduled to demonstrate production of O2 and methane from Martian H2O and atmospheric CO2 at Jet Propulsion Laboratory in 2022.In addition to SOXE/SOFC systems, the OxEon technology portfolio also consists of a low energy plasma reformer capable of producing sygnas from a variety of hydrocarbons and a modular-scale Fischer-Tropsch reactor for the production of liquid fuels from syngas. These technologies allow OxEon to provide a wide range of terrestrial energy solutions. A DOE Bioenergy Technologies Office (BETO) funded system is currently being fabricated to demonstrate the conversion of both CO2 and CH4 in anaerobic digester gas to liquid transportation fuels. This will be accomplished using an SOEC system to convert steam and CO2 to synthesis gas. OxEon’s plasma reformer will be used to convert methane to synthesis gas, using the plasma to catalyze the reaction of methane with steam and oxygen, supplied in part by the byproduct oxygen from the electrolysis system. The syngas from the plasma reformer and electrolysis systems are then combined to produce liquid fuels in the Fischer-Tropsch (FT) reactor. Figure 1

Effect of Double Impulse Trajectory Transfer from Earth to Mars Mission

The interplanetary mission from Earth to Mars takes at minimum energy launch opportunity with a single impulse transfer. Due to some unavoidable circumstances, if the launch takes at a later date, other than the minimum energy launch opportunity then the energy requirement increases. Also if the terminal states, approach the condition of opposition, the energy requirements rapidly increase and become infeasible too, resulting in restrictions on the launch window. This is due to the relative inclinations of the two planets. To overcome this situation, the possibility of applying an intermediatory impulse between the two terminals, named as double or two impulse transfers is considered. This paper gives the necessary details for the double impulse transfer and draw attention to its advantages.

Expanding Interplanetary Transfer Opportunities from Geostationary Transfer Orbits via Earth Synchronous Orbits

Geostationary transfer orbits (GTOs) are geocentric orbits widely considered for kick-stage operations of satellites for deep space missions. A piggyback spacecraft departing from GTOs requires a low transfer cost. The type of GTO and launch timings of piggyback missions, on the other hand, are typically determined by the primary mission that carried the piggyback spacecraft. This research introduces a new transfer strategy that coordinates piggyback spacecraft’s departure timing from Earth to deep space via an Earth synchronous orbit (ESO) after departure from a GTO. It also enables the spacecraft to change its velocity direction by introducing Earth gravity assists. Connecting several GTOs and interplanetary trajectories via ESOs reduces the required for the transfers. It leads to using transfer opportunities previously considered unsuitable for missions or miniaturizing the kick motor. As a result, ESOs allow for a broader launch opportunity for deep space piggyback probes. In addition, the feasibility of ESOs is shown numerically by comparing direct and ESO-assisted transfers from GTOs to Mars with the required .

Mission Architecture to Characterize Habitability of Venus Cloud Layers via an Aerial Platform

Venus is known for its extreme surface temperature and its sulfuric acid clouds. But the cloud layers on Venus have similar temperature and pressure conditions to those on the surface of Earth and are conjectured to be a possible habitat for microscopic life forms. We propose a mission concept to explore the clouds of Venus for up to 30 days to evaluate habitability and search for signs of life. The baseline mission targets a 2026 launch opportunity. A super-pressure variable float altitude balloon aerobot cycles between the altitudes of 48 and 60 km, i.e., primarily traversing the lower, middle, and part of the upper cloud layers. The instrument suite is carried by a gondola design derived from the Pioneer Venus Large Probe pressure vessel. The aerobot transmits data via an orbiter relay combined with a direct-to-Earth link. The orbiter is captured into a 6-h retrograde orbit with a low, roughly 170-degree, inclination. The total mass of the orbiter and entry probe is estimated to be 640 kg. An alternate concept for a constant float altitude balloon is also discussed as a lower complexity option compared to the variable float altitude version. The proposed mission would complement other planned missions and could help elucidate the limits of habitability and the role of unknown chemistry or possibly life itself in the Venus atmosphere.

Modelling the interaction of the Astro Bio Cube Sat with the Van Allen’s Belt radiative field using Monte Carlo transport codes

PurposeThe AstroBio Cube Satellite (ABCS) will deploy within the inner Van Allen belt on the Vega C Maiden Flight launch opportunity of the European Space Agency. At this altitude, ABCS will experience radiation doses orders of magnitude greater than in low earth orbit, where CubeSats usually operate. The paper aims to estimate the irradiation effect on the ABCS payload in the orbital condition, their possible mitigation designing shielding solutions and performs a preliminary representativity simulation study on the ABCS irradiation with fission neutron at the TAPIRO (TAratura Pila Rapida Potenza 0) nuclear research reactor facility at ENEA.MethodsWe quantify the contributions of geomagnetically trapped particles (electron and proton), Galactic Cosmic Rays (GCR ions), Solar energetic particle within the ABCS orbit using the ESA’s SPace ENVironment information system. FLUKA (Fluktuierende Kaskade—Fluctuating Cascade) code models the ABCS interaction with the orbital source.ResultsWe found a shielding solution of the weight of 300 g constituted by subsequent layers of tungsten, resins, and aluminium that decreases on average the 20% overall dose rate relative to the shielding offered by the only satellite’s structure. Finally, simulations of neutron irradiation of the whole ABCS structure within the TAPIRO’s thermal column cavity show that a relatively short irradiation time is requested to reach the same level of 1 MeV neutron Silicon equivalent damage of the orbital source.ConclusionsThe finding deserves the planning of a future experimental approach to confirm the TAPIRO’s performance and establish an irradiation protocol for testing aerospatial electronic components.

The Ariel Instrument Control Unit

Ariel, the Atmospheric Remote-sensing Infrared Exoplanet Large-survey mission (Tinetti 2019; Puig et al. 2018; Pascale et al. 2018), has been selected in March 2018 by ESA for the fourth medium-class mission (M4) launch opportunity of the Cosmic Vision Program, with an expected lift off in late 2028. It is the first mission dedicated to measuring the chemical composition and thermal structures of the atmospheres of hundreds of transiting exoplanets, enabling planetary science far beyond the boundaries of our own Solar System. Its Payload (P/L) (Eccleston and Tinetti 2018; Eccleston et al. 2017; Middleton et al. 2019), has been designed to perform transit spectroscopy from space during primary and secondary planetary eclipses in order to achieve a large unbiased survey concerning the nature of exoplanets atmospheres and their interiors, to determine the key factors affecting the formation and evolution of planetary systems (Tinetti et al. 2017, 2018). Ariel will observe hundreds of warm and hot transiting gas giants, Neptunes and super-Earths around a wide range of host star types, targeting planets hotter than $\sim $ 600 K to take advantage of their well-mixed atmospheres. It will exploit primary and secondary transit spectroscopy in the 1.10 to 7.80 μm spectral range and broad-band photometry in the optical (0.50 - 0.80 μm) and Near IR (0.80 - 1.10 μm) . One of the two instruments of the Ariel Payload is the Fine Guidance System (FGS), including three photometric channels (two used for guiding as well as science) between 0.5-1.1 μm plus a low resolution NIR spectrometer for 1.1-1.95 μm range. Along with FGS an IR Spectrometer (AIRS) (Amiaux et al. 2017) is foreseen, providing low-resolution spectroscopy in two IR channels: Channel 0 (CH0) for the 1.95 − 3.90 μm band and Channel 1 (CH1) for the 3.90 − 7.80 μm range. Finally, an Active Cooler System (ACS) including a Ne Joule-Thomson cooler is adopted to provide active cooling capability to the AIRS detectors working at cryogenic temperatures. AIRS is located at the intermediate focal plane of the telescope and common optical system and it hosts two HgCdTe-based hybrid IR detectors and two cold front-end electronics (CFEE) for detectors control and readout. Each CFEE is driven by a Detector Control Unit (DCU) part of AIRS but hosted within and managed by the Instrument Control Unit (ICU) of the Payload (Focardi et al. 2018). ICU is a warm unit residing into the S/C Service Module (SVM) and it is based on a cold redundant configuration involving the Power Supply Unit (PSU) and the Commanding and Data Processing Unit (CDPU) boards; both DCUs are instead cross-strapped and can be managed by the nominal or the redundant (PSU+CDPU) chain. ICU is in charge of AIRS management, collecting scientific and housekeeping (HK) telemetries from the spectrometer and HK from the telescope (temperatures readings), the P/L Optical Bench (OB) and other Subsystems (SS), thanks to a warm slave unit (TCU, Telescope Control Unit) interfaced to the ICU. Science and HK telemetries are then forwarded to the S/C, for temporary storage, before sending them to Ground. Here we describe the status of the ICU design at the end of B1 Phase, prior to the Mission Adoption Review (MAR) by ESA, with some still open architectural choices to be addressed and finalised once selected the ICU industrial Prime contractor.

Zero-pressure balloons trajectory prediction: Duster flight simulations

DUSTER (Dust from the Upper Stratosphere Tracking Experiment and Retrieval) is a balloon-born instrument designed to collect dust particles with sizes less than 30 µm in the upper stratosphere (30 – 40 km), to be analysed in laboratory. In support to the DUSTER flight campaigns, it is crucial to predict the balloon trajectories according to atmospheric circulation and to estimate the landing. For a feasibility study of an Antarctic flight campaign we predict the DUSTER trajectory with a statistical approach, considering safety issues and the pivotal instrument recovery, to retrieve the collected samples. To this aim, we used the NOAA (National Oceanic and Atmospheric Administration) meteorological data and the HYSPLIT software (HYbrid Single Particle Lagrangian Integrated Trajectory). The Antarctic Polar Vortex, every year in the period December – January, generates a counter clockwise circulation (averaged wind speeds around 36 m/s from 29 to 35 km of altitudes). We evaluated the best launch opportunity windows for DUSTER to take advantage of the Polar Vortex starting from historical data. The NOAA meteorological data (Global reanalysis of atmospheric data – GFS Analysis and Forecasts products of the National Centers for Environmental Predictions - NCEP) from 1994 to 2015 allowed us to select the period after the 10th of January as the launch window maximizing the DUSTER time of flight. We used HYSPLIT program to project the balloon simulated trajectories on ground, providing critical information for the recovery operations. Successively, with the aim of becoming autonomous within the DUSTER team, i.e. taking care of the complete procedure to realize a DUSTER flight campaign, we developed a dedicated software, the DUSTER Flight Simulation Predictor (DFSP) that considers the ascent, floating and descent phases and the meteorological conditions. DFSP considers the balloon dynamics, the heat flow transfer between the balloon and the atmosphere, and the lifting variable gas mass to simulate the complete flight. We are now able to obtain balloon trajectory simulations with a flight time up to 240 h, starting and ending in any Earth geographical position. In the following paper, a comparison between real DUSTER flight data in 2009 and simulated DFSP data will be shown.

Multidisciplinary design and optimization of an innovative nano air launch vehicle with a twin-fuselage UAV as carrier aircraft

The population of nano satellites has increased dramatically in recent years, but there are few launch vehicles providing dedicated launch opportunities for them. This paper presents the conceptual design of an innovative nano air launch vehicle, dedicated to providing exclusive launch opportunity to nano satellites weighing 10 kg or less. This nano launcher is a three-stage launch vehicle carried by a twin-fuselage unmanned aerial vehicle and released at an altitude of 10 km from a velocity of 0.5Ma. The air launch vehicle is innovative in two senses: firstly, it adopts a novel air launch method. Before separation, the launch vehicle is adjusted to a large angle of attack by the carrier aircraft. The launch vehicle only needs to keep attitude stable after separation, and sequentially when first stage ignites, the orientation of the thrust is inclined upward so that the thrust can provide enough vertical force to pull up the flight path angle rapidly, without aid from large wings; secondly, a terminal lumped control unit is applied to undertake the flight control task on its own. The terminal lumped control unit is a combination of four thrusters that share the liquid propellant with the last stage's main engine. With the terminal lumped control unit, the air launch vehicle needs no movable attitude control instruments, including thrust vectoring gimbals and movable wings. This innovative flight control strategy makes the launch vehicle very simple in structure and light in weight. Multidisciplinary modeling of the nano launcher is performed and genetic algorithm is applied for optimization. Among all the disciplines, trajectory optimization is key in checking the feasibility of the design, for which the Radau pseudospectral method is successfully implemented to obtain the optimal ascent trajectory. The optimal nano launcher has a gross take-off weight of 1365 kg. Finally, a sensitivity analysis is performed to check the robustness of the result using Monte Carlo simulation method.

Integration and testing of the Mini-EUSO telescope

The Mini-EUSO telescope, approved by ASI (Italian Space Agency) and Roscosmos (Russian Space Agency) as long-term program of space experiments to be launched on the Progress spacecraft during the Expedition 60/61 in 2019, will be placed on a nadir-facing, UV transparent window on the ISS Russian Zvezda module looking down on Earth. As one of the JEM-EUSO collaboration missions, its first objective is to observe UV light produced by UHECR (Ultra High Energy Cosmic Ray) events from upper atmosphere (~ 400 km altitude), moreover, from this point of view, Mini-EUSO will perform, for the first time, high-resolution mapping of night-earth UV (300 - 400 nm) emission, focusing on terrestrial background sources, TLEs (Transient Luminous Events), bioluminescence, meteoroids burning in atmosphere, search for SQM (Strange Quark Matter) signals, as well as detecting and mapping space debris. The integration of the instrument, currently underway at the University of Rome Tor Vergata, is at an advanced stage in order to be compliant with a launch opportunity in June 2019. The on ground test results will be presented.

Concurrent AIV as a method to hard tailor test- and model philosophies in times of need

MASCOT, a small 11 kg prototype Asteroid Lander on-board JAXA’s Hayabusa2 space probe, was launched on December 3rd, 2014, and arrived at its destined target asteroid (162173) Ryugu on June 27, 2018. MASCOT was separated from its mother spacecraft and successfully landed on October 3rd, 2018, accomplishing the first ever landing of a European spacecraft on the surface of an asteroid. To catch this particular launch opportunity its development timeline needed to be heavily compressed. In particular, the kick-off for hardware production was released in February 2012, only 2 years before the initially planned delivery of the flight unit. Due to this compact schedule, current and well established verification processes could not be followed in order to finalize the project in the given time. But by applying a unique mix of conventional and tailored model philosophies it was possible to dynamical adapt the test program to accomplish for the shortest planning and a suitable weighing of costs and risks. A strategy of Concurrent Assembly, Integration and Verification (C-AIV) helped to identify and mitigate design and manufacturing issues and shortened the test timeline further from a general 4–5 year C/D-phase down to 2,5 year C/D-phase. This short article outlines the general idea of the applied method which could be used by AIV and System Engineers in a general tailoring process for projects of similar nature which could be run in an alternative and much faster way, if the circumstances would call for it.•Concurrent AIV, a new agile methodology to hard tailor test and model philosophies for space projects is presented,•The methodology is based on parallelization of test activities, creation of independent unique test threads and synergizing their dependencies at key points,•On the baseline of the asteroid lander MASCOT, this methodology has been successfully applied to shorten the overall test and implementation schedule to only 2.5 years.

Preliminary orbit design for a Mars sample return mission

Mars sample return missions are the focus of international deep space exploration planning in recent years. In this paper, the orbit design of the different flight phases of a Mars sample return mission is studied with consideration of practical engineering constraints, using 2028 launch opportunity as an example; based on comparative analysis of different schemes, feasible orbit designs are recommended for future Mars sample return missions. Preliminary analysis and design of the mission orbit are carried out and the analysis results of the main flight phases are provided in the paper, and the discussion focuses on launch opportunity, Earth to Mars transfer, Mars orbit insertion, rendezvous and docking in Mars orbit, and Mars to Earth transfer. It is shown in the paper that the launch opportunities for Mars sample return missions emerge every two years, and the launch windows in 2028 opportunity concentrate in the November; considering the high risk of the research and development of heavy lift rockets, the scheme that the descent and ascent module, the orbiter and reentry module are launched separately using two available rockets is recommended; the Mars to Earth return opportunities emerge about every two years, and the return windows open about two years after the Earth departure.

Hierarchical approach for fast searching optimal launch opportunity in a wide range

Launch opportunity search is crucial for preliminary design of interplanetary trajectories. However, it is difficult to obtain the optimum solution efficiently when the range of search is wide. In this paper, a new fast search algorithm based on a modified hierarchical approach is proposed. At the top level, a simplified Pork-Chop diagram is plotted by sampling interpolation to delimit the candidate regions of the optimal solution. Each candidate region contains only one optimal solution, which fundamentally avoids the local optimal problem of genetic algorithms. At the bottom level, a hybrid optimization approach combining the genetic algorithm and the conjugate directions method is used to solve the accurate optimal solution of each candidate region and obtain the global optimal launch opportunity in the whole wide range of search. Simulation results show that the proposed hierarchical approach can quickly and robustly find the global optimal launch opportunity in a wide search space.

A novel orbiter mission concept for venus with the EnVision proposal

In space exploration, planetary orbiter missions are essential to gain insight into planets as a whole, and to help uncover unanswered scientific questions. In particular, the planets closest to the Earth have been a privileged target of the world's leading space agencies. EnVision is a mission proposal designed for Venus and competing for ESA's next launch opportunity with the objective of studying Earth's closest neighbor. The main goal is to study geological and atmospheric processes, namely surface processes, interior dynamics and atmosphere, to determine the reasons behind Venus and Earth's radically different evolution despite the planets' similarities. To achieve these goals, the operational orbit selection is a fundamental element of the mission design process. The design of an orbit around Venus faces specific challenges, such as the impossibility of choosing Sun-synchronous orbits. In this paper, an innovative genetic algorithm optimization was applied to select the optimal orbit based on the parameters with more influence in the mission planning, in particular the mission duration and the coverage of sites of interest on the Venusian surface. The solution obtained is a near-polar circular orbit with an altitude of 259 km that enables the coverage of all priority targets almost two times faster than with the parameters considered before this study.

IRMA: a Graphical Tool for Interplanetary Mission Design

Designing an interplanetary mission is a complex task and requires the choice of the launch opportunity and of the exact launch and arrival dates. Depending on these choices, the trajectory must be defined and, in case of continuous thrust, also the thrust profile needs to be optimized.. Traditionally, these choices are made using some plots which allow to find a good compromise between the travel duration and the cost of the mission, which is often expressed in terms of initial mass in Earth orbit (IMLEO). IRMA (InterPlanetary Mission Analysis) code, based on the MATLAB®environment, is here described. It allows to deal with both impulsive propulsion (using the patched conics approach) and low continuous thrust (Solar or Nuclear electric or propellantless, like solar sails). A specific solver, based on indirect optimization techniques, has been developed specifically for this program, but IRMA can be used also as an interface for standard solvers, based on direct methods, like the FALCON.m code.

The ARIEL Instrument Control Unit design

The Atmospheric Remote-sensing Infrared Exoplanet Large-survey mission (ARIEL) is one of the three present candidates for the ESA M4 (the fourth medium mission) launch opportunity. The proposed Payload will perform a large unbiased spectroscopic survey from space concerning the nature of exoplanets atmospheres and their interiors to determine the key factors affecting the formation and evolution of planetary systems. ARIEL will observe a large number (>500) of warm and hot transiting gas giants, Neptunes and super-Earths around a wide range of host star types, targeting planets hotter than 600 K to take advantage of their well-mixed atmospheres. It will exploit primary and secondary transits spectroscopy in the 1.2-8 um spectral range and broad-band photometry in the optical and Near IR (NIR). The main instrument of the ARIEL Payload is the IR Spectrometer (AIRS) providing low-resolution spectroscopy in two IR channels: Channel 0 (CH0) for the 1.95-3.90 um band and Channel 1 (CH1) for the 3.90-7.80 um range. It is located at the intermediate focal plane of the telescope and common optical system and it hosts two IR sensors and two cold front-end electronics (CFEE) for detectors readout, a well defined process calibrated for the selected target brightness and driven by the Payload's Instrument Control Unit (ICU).

Overview of China’s 2020 Mars mission design and navigation

Scheduled for an Earth-to-Mars launch opportunity in 2020, the China’s Mars probe will arrive on Mars in 2021 with the primary objective of injecting an orbiter and placing a lander and a rover on the surface of the Red Planet. For China’s 2020 Mars exploration mission to achieve success, many key technologies must be realized. In this paper, China’s 2020 Mars mission and the spacecraft architecture are first introduced. Then, the preliminary launch opportunity, Earth–Mars transfer, Mars capture, and mission orbits are described. Finally, the main navigation schemes are summarized.

The effect of the displacement damage on the Charge Collection Efficiency in Silicon Drift Detectors for the LOFT satellite

The technology of Silicon Drift Detectors (SDDs) has been selected for the two instruments aboard the Large Observatory For X-ray Timing (LOFT) space mission. LOFT underwent a three year long assessment phase as candidate for the M3 launch opportunity within the ``Cosmic Vision 2015 – 2025'' long-term science plan of the European Space Agency. During the LOFT assessment phase, we studied the displacement damage produced in the SDDs by the protons trapped in the Earth's magnetosphere. In a previous paper we discussed the effects of the Non Ionising Energy Losses from protons on the SDD leakage current. In this paper we report the measurement of the variation of Charge Collection Efficiency produced by displacement damage caused by protons and the comparison with the expected damage in orbit.

The PLATO 2.0 mission

PLATO 2.0 has recently been selected for ESA's M3 launch opportunity (2022/24). Providing accurate key planet parameters (radius, mass, density and age) in statistical numbers, it addresses fundamental questions such as: How do planetary systems form and evolve? Are there other systems with planets like ours, including potentially habitable planets? The PLATO 2.0 instrument consists of 34 small aperture telescopes (32 with 25 sec readout cadence and 2 with 2.5 sec candence) providing a wide field-of-view (2232 deg2) and a large photometric magnitude range (4-16 mag). It focusses on bright (4-11 mag) stars in wide fields to detect and characterize planets down to Earth-size by photometric transits, whose masses can then be determined by ground-based radial-velocity follow-up measurements. Asteroseismology will be performed for these bright stars to obtain highly accurate stellar parameters, including masses and ages. The combination of bright targets and asteroseismology results in high accuracy for the bulk planet parameters: 2%, 4-10% and 10% for planet radii, masses and ages, respectively. The planned baseline observing strategy includes two long pointings (2-3 years) to detect and bulk characterize planets reaching into the habitable zone (HZ) of solar-like stars and an additional step-and-stare phase to cover in total about 50% of the sky. PLATO 2.0 will observe up to 1,000,000 stars and detect and characterize hundreds of small planets, and thousands of planets in the Neptune to gas giant regime out to the HZ. It will therefore provide the first large-scale catalogue of bulk characterized planets with accurate radii, masses, mean densities and ages. This catalogue will include terrestrial planets at intermediate orbital distances, where surface temperatures are moderate. Coverage of this parameter range with statistical numbers of bulk characterized planets is unique to PLATO 2.0.

Quantum tests of the Einstein Equivalence Principle with the STE–QUEST space mission

We present in detail the scientific objectives in fundamental physics of the Space–Time Explorer and QUantum Equivalence Space Test (STE–QUEST) space mission. STE–QUEST was pre-selected by the European Space Agency together with four other missions for the cosmic vision M3 launch opportunity planned around 2024. It carries out tests of different aspects of the Einstein Equivalence Principle using atomic clocks, matter wave interferometry and long distance time/frequency links, providing fascinating science at the interface between quantum mechanics and gravitation that cannot be achieved, at that level of precision, in ground experiments. We especially emphasize the specific strong interest of performing Equivalence Principle tests in the quantum regime, i.e. using quantum atomic wave interferometry. Although STE–QUEST was finally not selected in early 2014 because of budgetary and technological reasons, its science case was very highly rated. Our aim is to expose that science to a large audience in order to allow future projects and proposals to take advantage of the STE–QUEST experience.

Mars Science Laboratory Heatshield Development, Implementation, and Lessons Learned

The Mars Science Laboratory is a scientific mission that placed a 900 kg rover with 10 science instruments on the surface of Mars. A key component for landing on the surface of Mars is the aeroshell, which decelerates the entry vehicle to just 1% of its initial kinetic energy at atmospheric entry. The forward portion of the aeroshell is the heatshield, which must be designed to withstand the pressure, heating, and shear stresses it experiences during the atmospheric deceleration process. The original thermal protection system chosen for the heatshield, SLA-561V, was the material used on all previous Mars aeroshells but required new qualification testing. Unfortunately, the material failed testing in the more extreme environment three months after the aeroshell and project critical design reviews. A rapid decision was required to switch to another material under test by the Orion program called phenolic impregnated carbon ablator. With 18 months left to design, fabricate, test, and deliver this new heatshield for the 2009 launch opportunity, the heatshield team had to rapidly respond to the challenge and work in a manner that ensured success. Lessons learned can be had from both the failure of the original SLA-561V design and the success of the phenolic impregnated carbon ablator implementation.