CubeSail: A low cost CubeSat based solar sail demonstration mission

CubeSail is a nano-solar sail mission based on the 3U CubeSat standard, which is currently being designed and built at the Surrey Space Centre, University of Surrey. CubeSail will have a total mass of around 3 kg and will deploy a 5 × 5 m sail in low Earth orbit. The primary aim of the mission is to demonstrate the concept of solar sailing and end-of-life de-orbiting using the sail membrane as a drag-sail. The spacecraft will have a compact 3-axis stabilised attitude control system, which uses three magnetic torquers aligned with the spacecraft principle axis as well as a novel two-dimensional translation stage separating the spacecraft bus from the sail. CubeSail’s deployment mechanism consists of four novel booms and four-quadrant sail membranes. The proposed booms are made from tape-spring blades and will deploy the sail membrane from a 2U CubeSat standard structure. This paper presents a systems level overview of the CubeSat mission, focusing on the mission orbit and de-orbiting, in addition to the deployment, attitude control and the satellite bus.     

Modified gradiometer technique applied to Double Star (TC-1)

On TC-1 (Tan Ce 1), the equatorial spacecraft of the Double Star mission, a strong spin-synchronized magnetic interference from the solar panels was observed. In-flight correction techniques for spinning spacecraft that are based on minimizing spin tones in the spin-aligned component and in the magnitude of the ambient magnetic field are therefore not possible in this case. However, due to the fortunate situation that the spacecraft carries two flux-gate magnetometers on the same boom (at 0.5 m distance from each other), the spacecraft field effects could be removed from the spin-averaged data to achieve 0.2 nT relative accuracy, by using a gradiometer technique. Methodology and results are presented. The obtained accuracy allows the use of the data in multi-spacecraft studies together with the Cluster satellites.     

A multiple-CubeSat constellation for integrated earth observation and marine/air traffic monitoring

CubeSats has evolved from purely educational tools, to useful platforms for technology demonstration and many practical applications. This paper reviews a CubeSat constellation mission involving 3 CubeSats launched into orbit on Sep. 25th 2015, aiming to demonstrate the integrated application of low-cost CubeSat technologies with distributed payloads using a group of satellites, as well as to demonstrate several new technologies. The mission scenario, the satellite system design, the innovative technologies and instruments or devices used on the CubeSats and the in-orbit experimental results and the payload data analysis, as well as some experiences and lessons learned, are presented and summaried.     

Deorbiter CubeSat mission design

This paper presents the mission design for a CubeSat-based active debris removal approach intended for transferring sizable debris objects from low-Earth orbit to a deorbit altitude of 100 km. The mission consists of a mothership spacecraft that carries and deploys several debris-removing nanosatellites, called Deorbiter CubeSats. Each Deorbiter is designed based on the utilization of an eight-unit CubeSat form factor and commercially-available components with significant flight heritage. The mothership spacecraft delivers Deorbiter CubeSats to the vicinity of a predetermined target debris, through performing a long-range rendezvous maneuver. Through a formation flying maneuver, the mothership then performs in-situ measurements of debris shape and orbital state. Upon release from the mothership, each Deorbiter CubeSat proceeds to performing a rendezvous and attachment maneuver with a debris object. Once attached to the debris, the CubeSat performs a detumbling maneuver, by which the residual angular momentum of the CubeSat-debris system is dumped using Deorbiter’s onboard reaction wheels. After stabilizing the attitude motion of the combined Deorbiter-debris system, the CubeSat proceeds to performing a deorbiting maneuver, i.e., reducing system’s altitude so much so that the bodies disintegrate and burn up due to atmospheric drag, typically at around 100 km above the Earth surface. The attitude and orbital maneuvers that are planned for the mission are described, both for the mothership and Deorbiter CubeSat. The performance of each spacecraft during their operations is investigated, using the actual performance specifications of the onboard components. The viability of the proposed debris removal approach is discussed in light of the results.     

Mars Phobos and Deimos Survey (M-PADS) – A martian Moons orbiter and Phobos lander

We describe a Mars ‘Micro Mission’ for detailed study of the martian satellites Phobos and Deimos. The mission involves two ∼330 kg spacecraft equipped with solar electric propulsion to reach Mars orbit. The two spacecraft are stacked for launch: an orbiter for remote investigation of the moons and in situ studies of their environment in Mars orbit, and another carrying a lander for in situ measurements on the surface of Phobos (or alternatively Deimos). Phobos and Deimos remain only partially studied, and Deimos less well than Phobos. Mars has almost always been the primary mission objective, while the more dedicated Phobos project (1988–89) failed to realise its full potential. Many questions remain concerning the moons’ origins, evolution, physical nature and composition. Current missions, such as Mars Express, are extending our knowledge of Phobos in some areas but largely neglect Deimos. The objectives of M-PADS focus on: origins and evolution, interactions with Mars, volatiles and interiors, surface features, and differences. The consequent measurement requirements imply both landed and remote sensing payloads. M-PADS is expected to accommodate a 60 kg orbital payload and a 16 kg lander payload. M-PADS resulted from a BNSC-funded study carried out in 2003 to define candidate Mars Micro Mission concepts for ESA’s Aurora programme.     

A new modeling method of solar energetic proton events for ISO specification

Solar energetic protons degrade performance and reliability of spacecraft systems due to single-event effects, total dose effects and displacement damage in electronics components including solar cells. On designing a solar cell panel, a total fluence of solar energetic protons (SEPs) which cause solar cell damage is needed to estimate power loss of the solar cells over mission life. Nowadays a solar panel area of spacecraft is increasing as spacecraft mission life becomes longer (15–18 years). Thus an accurate SEP model is strongly required for the cost-minimum design from the aerospace industry. The SEP model, JPL-91 proposed by Feynman et al., is currently used widely for solar cell designing. However, it is known that the JPL-91 model predicts higher fluences of protons than values actually experienced in space, especially after 7 years on orbit. In addition, the model is based on several assumptions, and also needs Monte-Carlo simulations for calculating fluences. In this study, we propose a new method for modeling SEPs especially focused on solar cell degradation. The newly-proposed method is empirical, which constructs a model based directly upon proton flux measurement data taken by instruments onboard spacecraft. This method has neither assumptions nor dependence on SEP event selection, both of which are needed in JPL-91. The model fluences derived from this method show lower fluences in longer missions compared to JPL-91. This method has been proposed to ISO (International Organization for Standardization) and has been discussed for a new standard SEP model.     

Small satellite survey mission to the second Earth moon

This paper presents an innovative space mission devoted to the survey of the small Earth companion asteroid by means of nano platforms. Also known as the second Earth moon, Cruithne, is the target identified for the mission. Both the trajectory to reach the target and a preliminary spacecraft budget are here detailed. The idea is to exploit high efficient ion thrusters to reduce the propellant mass fraction in such a high total impulse mission (of the order of 1e6 Ns). This approach allows for a 100 kg class spacecraft with a very small Earth escape energy (5 km²/s²) to reach the destination in about 320 days. The 31% propellant mass fraction allows for a payload mass fraction of the order of 8% and this is sufficient to embark on such a small spacecraft a couple of nano-satellites deployed once at the target to carry out a complete survey of the asteroid. Two 2U Cubesats are here considered as representative payload, but also other scientific payloads or different platforms might be considered according with the specific mission needs. The small spacecraft used to transfer these to the target guarantees the manoeuvre capabilities during the interplanetary journey, the protection against radiations along the path and the telecommunication relay functions for the data transmission with Earth stations. The approach outlined in the paper offers reliable solutions to the main issues associated with a deep space nano-satellite mission thus allowing the exploitation of distant targets by means of these tiny spacecraft. The study presents an innovative general strategy for the NEO observation and Cruithne is chosen as test bench. This target, however, mainly for its relevant inclination, requires a relatively large propellant mass fraction that can be reduced if low inclination asteroids are of interest. This might increase the payload mass fraction (e.g. additional Cubesats and/or additional scientific payloads on the main bus) for the same 100 kg class mission.     

Magnetic torquer induced disturbing signals within GRACE accelerometer data

The GRACE (Gravity Recovery And Climate Experiment) gravity field satellite mission was launched in 2002. Although many investigations have been carried out, not all disturbances and perturbations upon satellite instruments and sensors are resolved yet. In this work the issue of acceleration disturbances onboard of GRACE due to magnetic torquers is investigated and discussed. Each of the GRACE satellites is equipped with a three-axes capacitive accelerometer to measure non-gravitational forces acting on the spacecraft. We used 10 Hz Level 1a raw accelerometer data in order to determine the impact of electric current changes on the accelerometer. After reducing signals which are induced by highly dominating processes in the low frequency range, such as thermospheric drag and solar radiation pressure, which can easily be done by applying a high-pass filter, disturbing signals from onboard instruments such as thruster firing events or heater switch events need to be removed from the previously filtered data. Afterwards the spikes which are induced by the torquers can be very well observed. Spikes vary in amplitude with respect to an increasing or decreasing current used for magnetic torquers, and can be as large as 20 nm/s². Furthermore, we were able to set up a model for the spikes of each scenario with which we were able to compute model spike time series. With these time series the spikes can successfully be removed from the 10 Hz raw accelerometer data. Spectral analysis of the time series reveal that an influence onto gravity field determination due to these effects is very unlikely, but can theoretically not be excluded.     

A deorbiter CubeSat for active orbital debris removal

This paper introduces a mission concept for active removal of orbital debris based on the utilization of the CubeSat form factor. The CubeSat is deployed from a carrier spacecraft, known as a mothership, and is equipped with orbital and attitude control actuators to attach to the target debris, stabilize its attitude, and subsequently move the debris to a lower orbit where atmospheric drag is high enough for the bodies to burn up. The mass and orbit altitude of debris objects that are within the realms of the CubeSat’s propulsion capabilities are identified. The attitude control schemes for the detumbling and deorbiting phases of the mission are specified. The objective of the deorbiting maneuver is to decrease the semi-major axis of the debris orbit, at the fastest rate, from its initial value to a final value of about 6471?km (i.e., 100?km above Earth considering a circular orbit) via a continuous low-thrust orbital transfer. Two case studies are investigated to verify the performance of the deorbiter CubeSat during the detumbling and deorbiting phases of the mission. The baseline target debris used in the study are the decommissioned KOMPSAT-1 satellite and the Pegasus rocket body. The results show that the deorbiting times for the target debris are reduced significantly, from several decades to one or two years.     

Solar proton events recorded in the stratosphere during cosmic ray balloon observations in 1957–2008

Long-term balloon observations have been performed by the Lebedev Physical Institute since 1957 up to the present time. The observations are taken several times a week at the polar and mid latitudes and allow us to study dynamics of galactic and solar cosmic ray as well as secondary particle fluxes in the atmosphere and in the near-Earth space. Solar energetic particles (120) – mostly protons – (SEP) events with >100 MeV proton intensity above 1 cm⁻² s⁻¹ s⁻¹ were recorded during 1958–2006. Before the advent of the SEP monitoring on spacecraft these results constituted the only homogeneous series of >100 MeV SEP events. The SEP intensities and energy spectra inferred from the Lebedev Physical Institute observations are consistent with the results taken in the adjacent energy intervals by the spacecraft and neutron monitors. Joint consideration of the SEP events series recorded by balloons and by neutron monitors during solar cycles 20–23 makes it possible to restore the probable number of events in solar cycle 19, which was not properly covered by observations. Some correlation was found between duration of SEP event production in a solar cycle and sunspot cycle characteristics.     

Systems design of a hybrid sail pole-sitter

This paper presents the preliminary systems design of a pole-sitter. This is a spacecraft that hovers over an Earth pole, creating a platform for full hemispheric observation of the polar regions, as well as direct-link telecommunications. To provide the necessary thrust, a hybrid propulsion system combines a solar sail with a more mature solar electric propulsion (SEP) thruster. Previous work by the authors showed that the combination of the two allows lower propellant mass fractions, at the cost of increased system complexity. This paper compares the pure SEP spacecraft with the hybrid spacecraft in terms of the launch mass necessary to deliver a certain payload for a given mission duration. A mass budget is proposed, and the conditions investigated under which the hybrid sail saves on the initial spacecraft initial mass. It is found that the hybrid spacecraft with near- to mid-term sail technology has a lower initial mass than the SEP case if the mission duration is 7 years or more, with greater benefits for longer duration missions. The hybrid spacecraft with far-term sail technology outperforms the pure SEP case even for short missions.     

Synergistic approach of asteroid exploitation and planetary protection

The asteroid and cometary impact hazard has long been recognised as an important issue requiring risk assessment and contingency planning. At the same time asteroids have also been acknowledged as possible sources of raw materials for future large-scale space engineering ventures. This paper explores possible synergies between these two apparently opposed views; planetary protection and space resource exploitation. In particular, the paper assumes a 5 tonne low-thrust spacecraft as a baseline for asteroid deflection and capture (or resource transport) missions. The system is assumed to land on the asteroid and provide a continuous thrust able to modify the orbit of the asteroid according to the mission objective. The paper analyses the capability of such a near-term system to provide both planetary protection and asteroid resources to Earth. Results show that a 5 tonne spacecraft could provide a high level of protection for modest impact hazards: airburst and local damage events (caused by 15–170 m diameter objects). At the same time, the same spacecraft could also be used to transport to bound Earth orbits significant quantities of material through judicious use of orbital dynamics and passively safe aero-capture manoeuvres or low energy ballistic capture. As will be shown, a 5 tonne low-thrust spacecraft could potentially transport between 12 and 350 times its own mass of asteroid resources by means of ballistic capture or aero-capture trajectories that pose very low dynamical pressures on the object.     

A small mission for in situ exploration of a primitive binary near-Earth asteroid

We present a concept for a challenging in situ science mission to a primitive, binary near-Earth asteroid. A sub-400-kg spacecraft would use solar electric propulsion to rendezvous with the C-class binary asteroid (175706) 1996 FG3. A campaign of remote observations of both worlds would be followed by landing on the ∼1 km diameter primary to perform in situ measurements. The total available payload mass would be around 34 kg, allowing a wide range of measurement objectives to be addressed. This mission arose during 2004 from the activities of the ad-hoc Small Bodies Group of the DLR-led Planetary Lander Initiative. Although the particular mission scenario proposed here was not studied further per se, the experience was carried over to subsequent European asteroid mission studies, including first LEONARD and now the Marco Polo near-Earth asteroid sample return proposal for ESA’s Cosmic Vision programme. This paper may thus be of interest as much for insight into the life cycle of mission proposals as for the concept itself.     

Metrology on-board PROBA-3: The shadow position sensors subsystem

PROBA-3 is an ESA mission aimed at the demonstration of formation flying performance of two satellites that will form a giant coronagraph in space. The first spacecraft will host a telescope imaging the solar corona in visible light, while the second, the external occulter, will produce an artificial eclipse. This instrument is named ASPIICS (Association of Spacecraft for Polarimetric and Imaging Investigation of the Corona of the Sun). To accomplish the payload's scientific tasks, PROBA-3 will ensure sub-millimeter reciprocal positioning of its two satellites using closed-loop on-board metrology. Several metrology systems will be used and the Shadow Position Sensor (SPS) subsystem senses the penumbra around the instrument aperture and returns the 3-D displacement of the coronagraph satellite, with respect to its nominal position, by running a dedicated algorithm. In this paper, we describe how the SPS works and the choices made to accomplish the mission objectives.     

Interplanetary spacecraft navigation using pulsars

We demonstrate how observations of pulsars can be used to help navigate a spacecraft travelling in the solar system. We make use of archival observations of millisecond pulsars from the Parkes radio telescope in order to demonstrate the effectiveness of the method and highlight issues, such as pulsar spin irregularities, which need to be accounted for. We show that observations of four millisecond pulsars every seven days using a realistic X-ray telescope on the spacecraft throughout a journey from Earth to Mars can lead to position determinations better than ∼20 km and velocity measurements with a precision of ∼0.1 ms⁻¹.     

Prototype design and testing of a Venus long duration,high altitude balloon

This paper describes the design, fabrication and testing of a full scale prototype balloon intended for long duration flight in the upper atmosphere of Venus. The balloon is 5.5 m in diameter and is designed to carry a 45 kg payload at an altitude of 55 km. The balloon material is a 180 g/m² multi-component laminate comprised of the following layers bonded together from outside to inside: aluminized Teflon film, aluminized Mylar film, Vectran fabric and a polyurethane coating. This construction provides the required balloon functional characteristics of low gas permeability, sulfuric acid resistance and high strength for superpressure operation. The design burst superpressure is 39,200 Pa which is predicted to be 3.3 times the worst case value expected during flight at the highest solar irradiance in the mission profile. The prototype is constructed from 16 gores with bi-taped seams employing a sulfuric acid resistant adhesive on the outside. Material coupon tests were performed to evaluate the optical and mechanical characteristics of the laminate. These were followed by full prototype tests for inflation, leakage and sulfuric acid tolerance. The results confirmed the suitability of this balloon design for use at Venus in a long duration mission. The various data are presented and the implications for mission design and operation are discussed.     

Assessing attitude error of FORMOSAT-3/COSMIC satellites and its impact on orbit determination

An attitude determination and control system (ADCS) is critical to satellite attitude maneuvers and to the coordinate transformation from the inertial frame to the spacecraft frame. This paper shows specific sensors in the ADCS of the satellite mission FORMOSAT-3/COSMIC (F3/C) and the impact of the ADCS quality on orbit accuracy. The selection of main POD antenna depends on the beta angles of the different F3/C satellites (for FM2 and FM4) during the inflight phase. In particular, under the eclipse, alternative attitude sensors are activated to replace the Sun sensors, and such a sensor change leads to anomalous GPS phase residuals and a degraded orbit accuracy. Since the nominal attitude serves as a reference for ADCS, the 3-dimensional attitude-induced errors in reduced dynamic orbits over selected days in 2010 show 9.35, 10.78, 4.97, 5.48, 7.18, and 6.89 cm for FM1–FM6. Besides, the 3-dimensional velocity errors induced by the attitude effect are 0.10, 0.10, 0.07, 0.08, 0.09, and 0.10 for FM1–FM6. We analyze the quality of the observed attitude transformation matrix of F3/C and its impact on kinematic orbit determination. With 249 days of GPS in 2008, the analysis leads to the following averaged 3-dimensional attitude-induced orbit errors: 2.72, 2.62, 2.37, 1.90, 1.70, and 1.99 cm for satellites FM1–FM6. Critical suggestions of geodetic payloads for the follow-on mission of F3/C are presented based on the current result.     

Impact effects from small size meteoroids and space debris

When the impact risk from meteoroids and orbital debris is assessed the main concern is usually structural damage. With their high impact velocities of typically 10–20 km/s millimeter or centimeter sized objects can puncture pressure vessels and other walls or lead to destruction of complete subsystems or even whole spacecraft. Fortunately chances of collisions with such larger objects are small (at least at present). However, particles in the size range 1–100 μm are far more abundant than larger objects and every orbiting spacecraft will encounter them with certainty. Every solar cell (8 cm² area) of the Hubble Space Telescope encountered on average 12 impacts during its 8.25 years of space exposure. Most were from micron sized particles.