Minimize PICASSO

PICASSO (Pico-Satellite for Atmospheric and Space Science Observations)

Overview    Spacecraft    Launch    Sensor Complement   Ground Segment   References

PICASSO is a project of ESA ( European Space Agency) led by the Belgian Institute for Space Aeronomy, in collaboration with VTT Technical Research Center of Finland Ltd, Clyde Space Ltd. (UK) and the CSL (Centre Spatial de Liège), Belgium. The goal is to develop and operate a scientific 3U CubeSat. With a payload composed of two scientific instruments, it will contribute to the determination of the ozone distribution in the stratosphere, the temperature profile up to the mesosphere and the electronic plasma characterization in the ionosphere. 1) 2)

Anticipating the fact that sensor miniaturization and tiny satellites will play an ever-increasing role in remote sensing and in situ measurements of the atmosphere of the Earth and of other celestial bodies, PICASSO has three strategic objectives:

1) To perform true science within a scope compatible with the technological constraints and current limitations of CubeSat technology

2) To anticipate the future of remote sensing and in situ measurements, for Earth and other planets, through miniaturization

3) To demonstrate that tiny satellites can achieve a very high ratio of "science data versus cost".

The PICASSO flight aims at a polar orbit at 500 km, thus enabling a mission with a nominal lifetime of about two years. In addition, with an inclination of 98º, the coverage of the atmosphere will be almost complete.

PICASSO is a nanosatellite demonstration mission of BIRA initiated in 2010. It is currently administered by ESA within the frame of the General Support Technology Program (GSTP) and of the Technology Research Program (TRP). In addition to BIRA, which acts as project prime and scientific PI (Principal Investigator) of the two instruments, the PICASSO consortium includes three European partners: 3) 4)

• Clyde-Space Ltd. (Glasgow, UK), is in charge of the platform development, of the payload integration, of the ground-based station and of the satellite control and monitoring

• VTT (Technical Research Center of Finland), Espoo, Helsinki, Finland, in charge of developing the VISION instrument

• The CSL (Centre Spatial of Liège), Belgium, responsible for the technical coordination and for the PA/QA (Product Assurance/Quality Assurance) aspects of the project.

PICASSO is an ambitious 3U CubeSat implementing a science mission aimed at studying ozone in the stratosphere, the air temperature profile up to the mesosphere, and the electron density and temperature in the ionosphere. The PICASSO next-generation mission allows mission-specific customization; a dedicated payload computer, to process and compress payload data for transmission; customized deployable solar panels, providing significant power generation in addition to enabling a minimized power system volume.

The mission will provide valuable scientific data from complex instruments, and will serve as an indicator of the capabilities and value of a modern CubeSat mission, in addition to providing valuable Lessons Learned for future platform designs. As such, PICASSO will serve as an ESA in-orbit-demonstrator of CubeSat technology and a trailblazer for small science missions.

 

Spacecraft:

The nanosatellite is built using a standard 3U CubeSat structure of size 30 cm x 10 cm x 10 cm and a mass of ~3.8 kg. In this configuration, 1U volume is reserved for the payload accommodation. The orbital lifetime of the PICASSO mission is expected to be 29 months.

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Figure 1: Artist's rendition of the deployed PICASSO nanosatellite (image credit: BIRA)

• Triple unit CubeSat (one unit left for the payload)

• Four deployable 2-unit long solar panels

• Attitude control:

- Inertial flight, one face towards the Sun

- Pointing accuracy: ±0.5° (knowledge: ±0.2°)

- Magneto-torquers and three dynamical wheels

• Mass estimated at ~4 kg

• Power Budget: 9.7 W generated, 6.5 W consumed

• RF communications: (Uplink: VHF; downlink: UHF + S-band)

- Data transmissions of up to 204 MB/day

Table 1: The platform parameters

Platform: The complexities of the mission necessitated greater latitude for customization of the structure, and flexibility during integration. After a review of the structures commercially available at the time it was determined there was nothing that would be suitable for this mission. Clyde Space therefore designed a custom CubeSat structure (Figure 2), which has now been adapted for the wider commercial market in order to support other missions facing similar challenges. The structure allows removal of individual faces while the inner stack remains in place. The solar panels can also be removed without the need to remove end plates. A solid section in one panel provides enhanced radiation protection for VISION. The internal ribs allow placement of the PC104 stack at any desired location within the structure. Similarly, customizable side plates permit the positioning of an antenna deployment module at any required location within the CubeSat. This structure is also very light, particularly given its strength, weighing 332 g including rails, end plates, ribs, rods, L-pieces and standoffs.

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Figure 2: Illustration of the PICASSO 3U CubeSat bus (image credit: Clyde Space)

The platform features four deployable solar panels, UHF/VHF and S-band communications, two on board computers, a high performance ADCS (Attitude Determination and Control Subsystem), and two scientific instruments for ESA-grade science data.

The key pointing requirement for this mission is to ensure that the VISION instrument stays aligned to the Sun during the entire duration of the Sun's occultations by the Earth's atmosphere. This requires the pointing accuracy of VISION relative to the sun to be ≤1º . To meet this requirement, the ADCS must be able to have sufficient pointing knowledge and control performance to provide this resolution of control. In addition, the alignment between the ADCS and VISION must be measured and corrected for.

The ADCS subsystem architecture is illustrated in Figure 3. It includes:

• ADCS Motherboard (ADCS MB)

• Reaction wheel set (RWS)

• Fine Sun sensor (FSS)

• Coarse Sun sensors (CSS)

• Star tracker (STT)

• Magnetorquers (MTQ)

• GNSS system (GPS).

The ADCS motherboard benefits from the heritage of the ADCS board that has successfully flown on UKube-1. An FPGA-based processing architecture has been specifically selected to ensure a system that is more robust to radiation events. The central Actel FPGA interfaces to the sensors and actuators, while a secondary processor acts as a watchdog, can place the spacecraft into a safe mode and can also be used to provide emergency detumbling of the spacecraft should the need arise.

The ADCS will utilize the standard Clyde Space three-axis CubeSat reaction wheel system. Each reaction wheel is capable of providing a torque of up to 2mNm; however, for this mission the torque is limited to 0.23 mNm in order to enable a finer pointing control. The wheels will provide a total angular momentum of 3.53 mNms in an angular velocity range of ± 7500 RPM.

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Figure 3: Schematic of the ADCS architecture (image credit: Clyde Space)

The FSS is mounted on the Sun pointing face of the satellite and is used as the primary sensor during the sunlit period of the orbit, while the star tracker is used in support of orientation determination during eclipse. The magnetometers are used to sense the magnitude and direction of the magnetic field on the body of the CubeSat.

Voice coils embedded into the solar panels represent the MTQ. These devices generate a magnetic dipole that interacts with the Earth's magnetic field, generating a mechanical torque. These actuators are used to detumble the spacecraft, to provide coarse pointing acquisition, and to manage the RWS angular momentum.

The ADCS is validated during ground testing using Clyde Space's HIL (Hardware-In-the-Loop) simulator. The complete HIL setup is a high fidelity, six degrees of freedom, spacecraft dynamical model interfaced directly with the ADCS hardware on which the attitude control algorithms run. The setup allows validation of the autonomous attitude control software and hardware for all phases of the mission.

Critically, real data from the sensors and actuators are used to simulate the entire mission. This allows many of the un-modelled dynamics that, because of the presence of unknown parameters, do not have clear mathematical formulations. These can include interference on the magnetometers from magneto-torquer output and magnetometer reading; reaction wheel velocity and gyros output; sensor noise and pointing accuracy; and others. Consequently, the use of this system level HIL test vastly reduces the impact of ‘non-ideal' operation of system hardware on the performance of the ADCS control algorithms, and therefore de-risks the potential for attitude control problems on-orbit.

EPS (Electric Power Subsystem): Power conditioning is carried out using an off-the-shelf third generation Clyde Space EPS (Figure 4), with 3 maximum power point tracking battery charge regulators providing protected primary power lines: 3.3 V at 4.5 A, 5 V at 4.5 A, 12 V at 1.5 A, and unregulated battery V at 4.5 A. The system also incorporates the flight activation system of separation microswitches, and the flight interface with an external 5 V USB charge and remove before flight pin. It also incorporates 10 power distribution switches to control the loads. The EPS board will be powered throughout the orbit, with a nominal consumption of 200 mW.

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Figure 4: The Clyde Space EPS board (image credit: Clyde Space)

The EPS also makes use of a 30 Wh battery – another Clyde Space standard product with spaceflight heritage on numerous missions. Based on Lithium Polymer technology, the battery cells are arranged 2S3P, with charging EoC voltage of 8.2 V.

During sunlit periods, the solar panels shall generate power to run the platform and charge the battery. A modified version of the standard CS product, four 2U deployable solar panels are mounted to the structure on their short edge. Their deployed configuration is shown in Figure 1. These panels host the SLP (Sweeping Langmuir Probe) devices whose fixtures run alongside the solar cells. Body-mounted solar panels are also used to generate power from Earth's albedo and to maximize power generation during tumbling.

In most cases, deployable solar panels would necessitate the use of a larger EPS variant, with a daughterboard featuring additional BCRs (Battery Charge Regulators) to accommodate the additional solar panels. However, given the complexity of the platform design, volume within the spacecraft is at a premium. To address this, the solar panels will feature a custom string configuration and harnessing solution in order to eliminate the need for the additional daughterboard within the CubeSat stack.

CDHS (Command and Data Handling Subsystem): There are two computers onboard PICASSO: the OBC (On-Board Computer) and PLC (Payload Computer). The OBC is the primary intelligence board and manages spacecraft operations. The PLC is dedicated to controlling the payloads and managing their data all the way to downlink.

The VISION instrument produces vast quantities of data, approximately 8 GB per observation. Transmitting the raw data in its entirety to the ground is infeasible with the anticipated link budget, so a high-performance secondary computer is incorporated on board, dedicated to processing the payload data prior to transmission.

The PLC uses a Xiphos Q7S processor card adapted for a CubeSat form factor, accommodating a Xilinx Zynq-7020 all-programmable system-on-chip with a dual-core ARM Cortex-A9 core clocked at 766 MHz. The card provides 512+256 MB of low power RAM and up to 32 GB of non-volatile storage on SD (Secure Digital)cards. Typical power consumption for the PLC is 1 W.

RF communications: The satellite's communication system is composed of two radios: The VHF/UHF Transceiver (VUTRX), which is mainly dedicated to TMTC (Telemetry and Telecommand), and the STX (S-band Transmitter) for high data-rate payload data downlink. The radios are connected to their respective antennas (deployable dipoles for the VUTRX, patch for the STX) and will communicate with dedicated antennas on the ground as depicted in Figure 5.

VUTRX uses an off-the-shelf CPUT (Clyde-Space UTRX half duplex UHF Transceiver), providing a VHF uplink and UHF downlink at 9.6 kbit/s both ways using modified CCSDS packets and multi-access protocols, providing down/uplink of 2.15 MB/day. When not transmitting, the transceiver enters a Morse Code Beacon mode and broadcasts identification and basic health data for tracking. The VU Transceiver interfaces to two VU Whip Antennas deployed from the ADM (Antenna Deployment Module).

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Figure 5: Space-ground communication architecture (image credit: Clyde Space)

The high volume of data generated by the VISION instrument necessitates use of greater downlink capabilities than provided by the VUTRX alone. A CPUT STX (S-band Transmitter), as flown on Clyde Space's previous satellite UKube-1, was chosen to facilitate the high data-rate transfers of payload data as well as enhanced telemetry data. This provides downlink at variable rates up to 2 Mbit/s QPSK (Quadra-Phase Shift Keying), utilizing the Intelsat encoding standard and modified CCSDS packets. The transmitter shall interface to a 7 dBi S-band patch antenna (Figure 6) located on the Earth-facing side of the satellite.

The S-band system will enable a downlink of 102 MB/day in the spacecraft's nominal sun-pointing mode. The satellite can also be reorientated into a ground station-pointing mode, to enable a downlink of up to 204 MB/day when required.

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Figure 6: Photo of the S-band patch antenna (image credit: Clyde Space)

OBSW (On-Board Software): The platform OBC runs the Bright Ascension Generation 1 Onboard Software. This component-based software has an underlying framework including OS and hardware abstract libraries, as well as support for FreeRTOS and POSIX/Linux. The software components support Clyde Space platform subsystems including integrated EPS, battery, solar panels, ADCS, VUTRX and the standard Clyde Space payload protocol. Activities including telemetry sampling, pooling, monitoring, logging, etc., as well as automated activities that are event-, time- or orbit-triggered are also supported.

The PLC software is composed of the following modules:

• Main software process

• SLP control and processing

• VISION control and processing

• GPS interface

• OBC interface

• STX control and processing.

The ground segment for PICASSO is divided into two functional areas: the Mission Operations Control and the Scientific Control Center. The Mission control center will be supplied with the Bright Ascension Generation 1 Ground Software (GNDSW) developed in unison with the OBSW for harmonious integration. The GNDSW reaps of heritage from UKube-1.

 

Launch: Launch arrangements are being made after the CDR (Critical Design Review) at the end of 2015. A launch of PICASSO as a secondary payload is expected for 2016.

Orbit: Sun-synchronous orbit, altitude of ~500 km, inclination = 98º and a LTAN around 10:30 hours.

 


 

Sensor complement (VISION, SLP)

VISION (Visible Spectral Imager for Occultation and Nightglow)

VISION, developed by VTT, targets primarily the observation of the Earth's atmospheric limb during orbital Sun occultation. By assessing the radiation absorption in the Chappuis band for different tangent altitudes, the vertical profile of the ozone is retrieved. A secondary objective is to measure the deformation of the solar disk so that stratospheric and mesospheric temperature profiles are retrieved by inversion of the refractive ray-tracing problem. Finally, occasional full-spectral observations of polar auroras are also foreseen. 5)

VISION is a hyperspectral imager in the visible and the near-infrared domain, between 430 and 800 nm. It is capable of taking 2D snapshots at freely selectable wavelengths within this range; the spectral selection is performed by a tunable FPI (Fabry-Perot Interferometer), the basic concept of which is shown in Figure 7 (Ref. 3).

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Figure 7: VISION functional diagram (image credit: VTT, PICASSO collaboration)

The key parameters of the VISION instrument are:

- Spectral filters: up to 3 modes

- Detector: Commercial CMOS 2048 x 2048 RGB

- Field of View: 2.5°

- Spectral range: 400-800 nm, FWHM: < 10 nm

- Heritage: AaSI (Aalto-1 Spectral Imager) on board of Aalto-1 of Aalto University, Finland.

The science goals are:

• Polar and mid-latitude stratospheric ozone vertical profile retrieval (via spectral observation of Sun occultations in the Chappuis band)

• Upper atmosphere temperature profiling based on the Sun refractive flattening: "Atmospheric Refractivity from Inversion of Dilution." The objective is to assess the refracted Sun shape in order to retrieve stratospheric and mesospheric temperature profiles by inverting the ray-tracing problem. Full spectral scan observations of nightglows and polar auroras are also foreseen if the dynamic range of the instrument allows it.

As shown in Figure 8, the apparent shape of the solar disk shrinks along the vertical dimension (relative to the Earth image). This deformation comes from the fact that rays emanating from the bottom of the Sun image propagate into denser atmospheric layers than those emanating from top. By solving the inverse ray-tracing problem of the photons propagation in the atmosphere, the mesospheric and stratospheric temperature profiles can be retrieved. Spectacular results have recently been obtained by the SOFIE instrument aboard the AIM (Aeronomy of Ice in the Mesosphere) spacecraft by detecting the edges of the solar disk and the related refraction angle. The VISION project expects to improve the method by making use of the full solar disk.

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Figure 8: Illustration of Sun image flattening (image credit: PICASSO collaboration)

The miniaturized VISION instrument is the primary instrument of the PICASSO mission. The optimized VISION design is expected to result in stratospheric ozone measurements to 5% accuracy with a vertical resolution of 2 km after post-processing. Thus, VISION will be one of the first CubeSat optical payloads capable of providing real atmospheric science data.

The solar brightness is attenuated by scattering, absorption and diffusion processes along the optical path in the atmosphere. By studying the attenuation spectrally, profiles of physical properties and chemical components in the atmosphere can be retrieved. Results from the numerical simulation of the attenuation of the photon number at different tangent heights are illustrated in Figure 5 – here, the upper curve corresponds to a 50 km altitude, decreasing in 5 km increments down to a 5 km altitude, which is represented by the lower curve. The photon depletion around 600 nm corresponds to photon absorption by the ozone in the so-called Chappuis band. Resolving it gives access to the ozone profile.

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Figure 9: Spectral photon depletion (5 km to 50 km altitude in 5 km increments – lower to upper curve), image credit: PICASSO collaboration)

 

SLP (Sweeping Langmuir Probe)

The SLP instrument is designed and developed at BIRA to observe the electron density and temperature with the following scientific objectives: 6)

- Investigate the ionosphere-plasmasphere coupling

- Study of the aurora structures

- Survey of the polar cap arcs. Monitor density irregularities in the polar cap ionosphere and relate those to signatures of the polar cap arcs, (e.g. those found in Cluster data).

- Study of ionospheric dynamics.

Instrument: SLP includes four cylindrical probes (needle-like Langmuir probes) whose electrical potential is swept in such a way that both electron temperature and electron density can be derived. In addition, since at least two probes will be out of the spacecraft's wake, differential measurements will be performed in order to increase the accuracy. The probes will be mounted on 5 cm long booms, at the extremity of 20 cm long solar panels deployed perpendicularly to the body of the spacecraft.

To avoid spacecraft charging, which leads to erroneous measurement data, the probes will be swept with a duty cycle of less than 5%. In addition, to avoid probe surface contamination and ageing, a high voltage will be applied to the probes when necessary. The raw data will be processed on board, using a dedicated FPGA. Only the compressed parameters of interest will be sent back to the ground station via an S-band link.

Given the high inclination of the orbit, the SLP instrument will allow a global monitoring of the ionosphere. Therefore, PICASSO will enable the study of space weather phenomena such as ionosphere-plasmasphere coupling, the subauroral ionosphere and corresponding magnetospheric features, auroral structure, polar cap arcs, ionospheric dynamics via coordinated observations with EISCAT's heating radar, and turbulence in the partially ionized ionosphere.

Parameter

Minimum

Maximum

Electron density

109/m3

1012/m3

Electron temperature

1300 K

2700 K

Measured current

50 x 10-12 A

8 x 10-6 A

Spatial sampling

N/A

160 m

Sampling frequency

50 Hz

N/A

Instrument mass

N/A

150 g

Power consumption

N/A

1.5 W

PCB (Printed Circuit Board) size

N/A

95 mm x 95 mm

Probe length (boom inclusive)

N/A

100 mm

Probe diameter

N/A

2 mm

No of probes

4

Deployment system

Solar panels

Table 2: Requirements of the SLP instrument parameters

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Figure 10: Block diagram of the electronics part of SLP (image credit: BIRA)

The SLP instrument provides in-situ observations. The principle of SLP's measurement is based on conventional Langmuir probe theory. By sweeping the potential of the probes with respect to the plasma potential, the instrument measures the current in the three regions: ion saturation, retardation and electron saturation regions. The typical characteristics of such a probe are illustrated in Figure 6 (Ref. 3).

The ion density is derived from the ion saturation region, where the potential of the probes is sufficiently negative to repel electrons and attract only true ions. The electron temperature and S/C potential are retrieved from the retardation region, where the potential of the probes is close to that of the plasma so that both ions and electrons are attracted. The electron density is derived from the electron saturation region, where the potential of the probes is sufficiently positive to repel ions and attract only electrons.

In nominal mode, SLP sweeps the potential of the probes from -7 V to +7 V with respect to the plasma potential in order to retrieve the electron density and temperature, together with the spacecraft potential and the ion density (when it is large enough). In another mode, the instrument measures only in the electron saturation region at a higher rate, measuring electron density with better spatial resolution. This operating principle is more advanced than the idea of using 4 probes at four different fixed bias potentials in the electron saturation region, since it allows a more detailed analysis and permits the determination of electron temperature as well. In addition, with this operating principle, there is no need for an electron gun.

The probes are thin cylindrical titanium rods, mounted on the deployable solar panels, which act as deployable booms. This configuration ensures that at least one probe is out of the S/C's wake at any time, in addition to providing redundancy.

 

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Figure 11: Typical Langmuir probe characteristics in linear (left) and logarithmic (right) current axis (image credit: BIRA)

 


 

Ground segment:

Clyde Space is constructing a Ground Station (GS) at their new headquarters in the center of Glasgow, Scotland, which will be used for PICASSO mission operations. The ground station will be composed of a pair of UHF and VHF Yagi-Uda type antennas capable of both uplink and downlink as well as a 3 m mesh dish for S-band downlink operations. The ground station will be capable of tracking satellites using orbit prediction from TLEs, and will be set up for fully autonomous up/downlink operations to maximize the use of passes.

 


1) D. Fussen, J. De Keyser, M. De Mazière, D. Pieroux, H. Lamy, S. Ranvier, E. Dekemper, A. Merlaud , E. Neefs, O. Karatekin, Z. Ping Z, V. Dehant, M. Van Ruymbeke, J. P. Noël, "The Picasso Mission," Proceedings of the 4S (Small Satellites Systems and Services) Symposium, Portoroz, Slovenia, June 4-8, 2012

2) P. Cordoen, J. De Keyser, Ph. Demoulin, D. Fussen, D. Pieroux, S. Ranvier, "PICASSO - Pico-Satellite for Atmospheric and Space Science Observations," 6th European CubeSat Symposium, Estavayer-le-Lac, Switzerland, October 14-16, 2014

3) Bena Mero, Kevin A. Quillien, Malcolm McRobb, Simone Chesi, Ross Marshall, Alasdair Gow, Craig Clark, M. Anciaux , P. Cardoen, J. De Keyser, Ph. Demoulin, D. Fussen, D. Pieroux, S. Ranvier, "PICASSO: A State of the Art CubeSat," Proceedings of the 29th Annual AIAA/USU Conference on Small Satellites, Logan, Utah, USA, August 8-13, 2015, paper: SSC15-III-2, URL: http://digitalcommons.usu.edu/cgi/viewcontent.cgi?article=3179&context=smallsat

4) "Picasso CubeSat," ESA, January 28, 2015, URL: http://www.esa.int/spaceinimages/Images/2015/01/Picasso_CubeSat2

5) "VTT develops a miniaturized spectral camera for ESA's Picasso nanosatellite mission," VTT, 2015, URL: http://www.vttresearch.com/services/smart-industry/process-and-analytical-measurement/vtt-develops-a-miniaturized-spectral-camera-for-esa%E2%80%99s-picasso-nanosatellite-mission

6) S. Ranvier, P. Cardoen, J. De Keyser, D. Pieroux, "A Novel Langmuir Probe Instrument for CubeSats," 5th European CubeSat Symposium, Royal Military Academy, VKI (Von Karman Institute), Brussels, Belgium, June 3-5, 2013
 


The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: "Observation of the Earth and Its Environment: Survey of Missions and Sensors" (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates (herb.kramer@gmx.net).

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