Minimize Aalto-1

Aalto-1: The Finnish Student Nanosatellite

Aalto University of Aalto, Finland has launched a student nanosatellite project called Aalto-1. The satellite is being built mainly by students in project assignments and thesis work. The satellite project is coordinated by the Department of Radio Science and Engineering and supported by Space Technology teaching. 1) 2) 3) 4) 5)

Project cooperation is also provided by various Departments of Aalto University: Department of Automation and Systems Technology, Department of Communications and Networking, Department of Applied Mechanics; and in addition the Department of Physics of the University of Helsinki (HY), Department of Physics and Astronomy of University of Turku (UTU), VTT (Technical Research Center of Finland, Helsinki), FMI (Finnish Meteorological Institute), Aboa Space Research Oy, Oxford Instruments Analytical Oy and other Finnish companies.

The core consortium is comprised of: 6) 7)

- Aalto University (satellite bus, ground segment, project PI)

- VTT Technical Research Centre of Finland (Spectrometer payload)

- University of Helsinki (RadMon payload)

- University of Turku (RadMon payload)

- Finnish Meteorological Institute (Electrostatic Plasma Brake payload).

A feasibility study and the preliminary design of the satellite have been made by Aalto University students during the spring semester 2010. The nanosatellite is based on the 3U CubeSat standards. The overall objective is to demonstrate the feasibility of a MEMS Fabry-Perot spectrometer for space applications.

The Aalto-1 mission goals are: 8)

• To design, build and operate first Finnish Earth Observation (EO) nanosatellite

• Technology demonstration of a very small imaging spectrometer for spaceborne EO

• Technology demonstration of a very small radiation detector for future satellites

• Development and demonstration of a deorbiting device for nanosatellites based on e-sail concept and measurement of its performance

• Promotion of engineering education in Finland with the aid of a satellite project.


Figure 1: Artist's rendition of the Aalto-1 nanosatellite in orbit (image credit: Aalto University)

Legend to Figure 1: The bottom side of the nanosatellite contains the spectrometer lens, digital camera lens, S-band communication antenna and access port of the satellite. The satellite is covered by solar panels.



The nanosatellite is based on a 3U CubeSat form factor with a size of 34 cm x 10 cm x 10 cm and a mass of ~ 4 kg. Depending on the final orbit, it will have an average power production of 4.8 W. The design life is 2 years.

The mechanical structure of the satellite consists of two subsystem PCB fastener stacks, which are joined by specially designed stack plates, and the main outer frame. The outer frame itself consists of two separate aluminum parts, both 1.5 mm thick, fastened together into one tubelike structure.

ADCS (Attitude Determination and Control Subsystem): The project selected the iADCS -100 system, designed and built by BST ( Berlin Space Technologies) GmbH, Berlin-Adlershof, Germany. The iADCS-100 is a semi-autonomous ADCS , providing an attitude knowledge and pointing accuracy of < 1º during the science observation phase and less than 10º during plasma brake operation, with several operation modes for the entire mission, with only rough positional information required. The IADCS-100 uses magnetorquers, MEMS gyros, miniature reaction wheels, sun sensors and a star tracker for accurate attitude knowledge. The subsystem fits into < 1/2U space. The star tracker is the primary sensing device during the imaging mode, with the reaction wheels enabling precise pointing, while the magnetorquers provide the “raw” and less precise pointing. The IT03 GPS receiver of Fastrax is employed for positional knowledge.

The OBC (On-Board Computer) design is based on an ARM 9 processor (AT91RM9200) utilizing an external RAM. This computer is equipped with EEPROM, DataFlash, NAND Flash, and microSD for data storage. The OBC is designed to run a customized GNU/Linux operating system and uses several strategies to increase the system reliability. The OBC features a dual boot memory, an external I2C bus controller and LVDS transceivers, it utilizes two 32 MB SDRAM modules. 9)

EPS (Electrical Power Subsystem): The satellite is covered with solar cells on almost all available surfaces, forming three long-side solar panels, and one smaller on the nadir side (during the remote-sensing phase of the mission). These panels produce roughly an average power of 4.5 W for an ideal 500 km orbit in the entirety of the mission, with a battery as part of theEPS to store the collected energy for the eclipse phase in the orbit.


Figure 2: Schematic view of the Aalto-1 nanosatellite (image credit: Aalto-1 consortium)

RF communications: The TT&C data are transmitted in VHF/UHF. The COM (Communications system) consists of VHF/UHF transceiver providing the command link between the satellite and ground station. It provides also the radio beacon signal. An ADS (Antenna Deployment System) deploys the VHF/UHF antennas when the satellite is released from the launch vehicle.

The payload data are downlinked in S-band in the frequency range of 2.4-2.45 GHz. The S-band system of Aalto-1 is illustrated in Figure 4. The satellite segment consists of interface towards the OBC, S-band transmitter circuit and S-band patch antenna. Between the satellite and the ground station is the radio wave propagation medium that is an important factor in the design of radio links. The ground segment consists of ground station equipment, such as the receiving antenna, the antenna rotating and control system, low noise amplifier located after the antenna, the receiver and the whole ground station control software. 10)

The selected architecture for the Aalto-1 S-band transmitter consists of the Texas Instruments TI CC2500 transceiver. The device provides a data rate of 1 Mbit/s, the RF output power is up to + 33 dBm, the power consumption is 4.2 W (max) and the data throughput/station pass is ~41.6 MB.


Figure 3: Overview of the Aalto-1 S-band system elements (image credit: Aalto University)


Figure 4: Schematic view of the Aalto-1 communication system (image credit: Aalto University)


Launch: A launch of the Aalto-1 nanosatellite is planned for late 2014 (launch arrangements are in progress).

Orbit: Sun-synchronous orbit, altitude of 500-900 km, a“midday - midnight” orbit provides ideal lighting conditions for the payload spectral imager.



Sensor complement: (AaSI, RadMon, EPB)

Aalto-1 is equipped with several scientific instruments. The miniture imaging spectrometer is the main payload of the satellite. It is intended primarily as a technology demonstration and, if operating successfully, will be used for actual scientific remote-sensing purposes.

AaSI (Aalto-1 Spectral Imager):

The AaSI instrument is under development at VTT Technical Research Center of Finland. The spectrometer is based on a tunable FPI (Fabry-Pérot Interferometer), a piezo-actuated MEMS device. The staring AaSI instrument is able to record 2D spatial images at one to three selected wavelength bands simultaneously. The interferometer consists of just two highly reflecting surfaces separated by a tunable air gap. To measure more than one channel at once, the multiple orders of the spectrometer’s transmission function can be matched to the sensitivities of the normal CCD/CMOS image sensor channels, such as red, green and blue pixels of a Bayer pattern RGB sensor or the different CCD’s of a 3 CCD detector system. 11) 12) 13)

AaSI is controlled in a closed capacitive feedback loop by three different piezo actuators. With these actuators the air gap can be adjusted from ~0.5-3 µm, with a spectral range from 500 -900 nm (using an RGB detector). Filter apertures of 7 or even 19 mm can be reached with the piezo-actuated FPI (Fabry-Pérot Interferometer). This design has already flown on UAVs (Unmanned Aerial Vehicles); it worked well, reaching a spectral resolution of 7-10 nm.

The MEMS version is based on a different concept, where the interferometer is a completely monolithic structure which has no discrete actuators: the second mirror is bent by an electrostatic force; thereby the air gap is adjusted. This design has not flown yet, but a design with a similar basic technology has been used in Vaisala’s CARBOCAP® sensor since 1997. Currently apertures of 0.5 - 2 mm with a wavelength range of 435 - 570 nm can be reached with the MEMS technology, but numbers are improving constantly and the spectral resolution is already equal to that of the piezo-actuated version.

The MEMS technology is based on the use of dielectric Bragg mirrors, which provide excellent optical throughput but have a limited spectral operational range, typically ±1 % about the center wavelength. So far, VTT has used metallic mirrors in the PFPI (Piezo-actuated Fabry-Pérot Interferometer), which provide a wide spectral range of 400-1000 nm but an optical transmission of only 25 to 35 %, thus reducing the optical throughput. However, Bragg mirrors can also be used in piezo-actuated devices which makes the optical throughput of the two technologies equal.




Spectral range

500 - 900 nm

VIS range

Spectral resolution

10-30 nm

@FWHM, spectral step 1 nm

Ground pixel size

~240 m x 240 m

@ 700 km orbital altitude

Spectral channels

6 to 20

60+ channels possible

FOV (Field of View)

10º x 10º

120 km swath width @ 700 km

SNR @ 3 ms&20 nm @FWHM

> 50

SNR requirement is defined for June and latitude of Helsinki (60) and for albedo 30%.

Power usage

< 4 W

Peak power

Instrument mass, size

600 g, 9.7 cm x 9.7 cm x 4.8 cm



CMOS sensor with 2048 x 2048 pixels

CMOSIS CMV4000, binned to 512 x 512 resolution (spectral image size)

VIS Camera

Image size of RGB VIS camera

1910 x 1270 pixels

Imagery is used for georeference of AaSI data


30º x 19º


Table 1: Operational parameters

PFPI (Piezo-actuated FPI) module design: The PFPI module is a key component in AaSI, composed of the FPI itself and the surrounding support structure, which also houses the necessary electronics for the piezo-actuation. The FPI mirrors are made with silver coating (thickness 50 nm), on top of which a 50 nm layer of silicon dioxide (SiO2) is deposited as a protective layer. A thin layer of titanium (4 nm) is used as an adhesive to fix all layers together. These mirrors are then fixed (Figure 5) to the piezo actuators with UV-curable adhesive (Dymax OP-61).

The PFPI used in AaSI has already passed space qualification testing and the instrument has passed the critical design review. The construction of the complete qualification model is currently under way and the flight model is expected to be completed during the autumn of 2013. 14)


Figure 5: Left: The PFPI-006-V035 right after the piezo-actuators have been glued to the mirrors. Right: The FPI mounted to the support structure with five silicone pads (shown in red), image credit: Aalto University


Figure 6: Working principle of the Fabry-Perot interferometer (image credit: VTT)

The source data rate of AaSI is highly dependent on the number of wavelength bands and the image size chosen. As such, the amount of data per picture is widely variable, and can be anything from 2 MB to > 500 MB (max).

The design of the AaSI is such, that the imager is controlled by the OBC (master) via a redundant I2C bus, both for commands and imager housekeeping data, while the imagery is being transmitted via a 3-wire SPI (Serial Peripheral Interface) to the OBC.

In addition, the AaSI instrumentation includes a separate VIS camera with a FOV of 30º x 19º (parallel accommodation of both instruments and nadir pointing). The objective of the VIS camera is to confirm the location of the AaSI imagery, and to determine whether it is sensible to downlink the high rate data, due to e.g. cloud cover in the target area.

The main functional blocks of AaSI are shown in Figure 7. There are three main parts: The FPI controller, the image sensor memory controller and the main microcontroller which acts as the main interface (I/F) between the OBC and AaSI. AaSI is also directly connected to the EPS (Electrical Power System) of the satellite by two voltage lines +5 V and +12 V.


Figure 7: Block diagram of the AaSI electronics (image credit: VTT)

Both image sensors have a global shutter function, so every pixel is exposed simultaneously. The image sensor data is transferred over four LVDS lines to the memory controller FPGA, which moves the data to the buffer memory. In AaSI a 32 MB SDRAM is used as the buffer memory, and it can hold the data of 16 megapixels. The same FPGA will also be used to send the image data to the OBC over and SPI-over-LVDS (Low Voltage Differential Signaling) link. The main microcontroller manages all the communication with the OBC and the synchronization of FPI and image sensor functions. It will also manage the necessary housekeeping data. The controller is connected to the satellite-wide I2C bus, which is used for command and telemetry transfer. The maximum power consumption is 3 W from the +5 V supply voltage line and 4 W from the +12 V line. The total power consumption depends on the operational mode of AaSI.

In-orbit spectral calibration is planned to be done by using known bright spectral features (e.g. the Sahara desert) and measuring the spectrum around strong absorption peaks (e.g. O2 absorption at 750-760 nm). Also on-board calibration using the 500 and 900 nm filter edges will be possible.


Figure 8: Schematic view of the imaging instrumentation (image credit: VTT)

Legend to Figure 8: The larger instrument on top of the box is the AaSI, while the smaller instrument at the bottom is the VIS camera.

The main advantages of the AaSI concept are the small size and the spectral programmability, which provides flexibility and reduced data rate when the application is well defined. A successful space qualification and orbit demonstration will enable development of more advanced instruments based on piezo and MEMS Fabry-Pérot interferometer technologies.


RadMon (Radiation Monitor):

RadMon is a radiation monitor, jointly built by the University of Helsinki (UH) and the University of Turku (UT). The instrument is a miniature spaceborne particle telescope, designed to detect particle (electron and proton) fluxes at the LEO altitude of Aalto-1. RadMon consists of two adjacent detector layers, a thin Si detector to measure the particle energy loss, and a thick CsI (TI) scintillator with a PD (Photodiode) readout to measure the residual energy of the particles stopping in the telescope (Figure 9). The particle energy ranges covered by RadMon are 10-200 MeV [0.7-10 MeV] for protons [electrons]. The measurements consist of particle counting rates (i.e., fluxes) in several energy passbands inside the available energy range (Ref. 4).


Figure 9: Schematic view of the RadMon detector assembly (image credit: UH, UT)

Legend to Figure 9: a) A cut of the RadMon detector unit showing the Si detector (blue chip) and the CsI (TI) scintillator (red cube). Particles enter the system through the 300 µm thick Al window (gray) and a signal from both detectors is required for counting the particle. The housing (brown) is made of brass. b) sketch of the RadMon unit including the stack of three PCBs (analog, digital and power supply, from top down) and the detector unit.

The RadMon design is divided into four subsystems:

• Detector unit

• Analog electronics board for performing the amplification of the detector signals and converting them into digital form

• Digital FPGA-based board for processing of the signals and conversion into the final data product, i.e., electron and proton counting rates in several energy passbands

• PSU (Power Supply Unit).

The instrument utilizes the available UART (Universal Asynchronous Receiver/Transmitter) bus for data transfer, commands and telemetry, as well as I2C as a backup. The down-loadable data (both scientific and housekeeping) is estimated to be around 1 MB per 24 hours of operation.


EPB (Electrostatic Plasma Brake):

The electric solar wind sail is a space propulsion method, invented in Finland at FMI (Finnish Meteorological Institute ). 15) The electrostatic plasma brake is a variant of the concept which consists of a single gravity-stabilized tether, intended to deorbit a satellite, to avoid space debris after the mission. The Electric Sail Experiment onboard Aalto-1 is intended to:

• demonstrate the deployment of a conducting thin multiline tether

• measure the electrostatic force exerted on the tether by the ram flow of the ionospheric plasma in different positive and negative tether voltages

• reduce the satellite altitude and so to demonstrate the usefulness of the plasma brake as a satellite deorbiting device.

To measure the expected µN scale electrostatic force, the voltage is turned on always in the same phase of the tether's rotation (e.g. always when the tether is moving towards the ram flow). After several spins, the effect accumulates enough to cause a detectable change in the tether's and satellite's spin rate, from which the force can be calculated. Over a longer timescale, the effect of the force can be deduced from a lowering of the satellite orbit.

EPB is a system, consisting of a 100 m long and 25-50 µm wide tether. As this tether moves with respect to the ionosphere of the LEO spacecraft, it experiences a Coulomb drag force. The tether, itself is composed of several strands, which are arranged in a specific order to minimize the risk of total failure due to strand severing (Ref. 4). 16) 17)



Figure 10: Various elements of the EPB instrumentation (image credit: UH,UT)

Legend to Figure 10: a) redundant strands ensure a level of tolerance to tearing; b) the tether reel storage shown on a part of the PCB with 53 mm diameter in size; c) an example of the cold cathode electron gun of size of 42 mm x 17.2 mm.

The tether is stored in a reel within the spacecraft (Figure 10 b, and will be reeled out with the help of a motor at a speed of 1-3 mm/s. The tip of the tether includes a small mass (0.5-1 gram) for safe initiation of the deployment.

The tether will be tested in the positively and negatively charged modes in order to estimate how the Coulomb drag will act in both modes, with the negative mode requiring only the satellite's surface to be conducting. The positive mode will require a special miniaturized cold cathode electron gun to complete the circuit (Figure 10 c).

The control of EPB's tether deployment as well as interaction with the rest of the spacecraft systems is handled by a separate control electronics device and a central FPGA. It interfaces via a redundant I2C to the OBC, but actualizes only a few functions, and sends a relatively small amount of data (< 10 kB per 24 hours) to the OBC.



Aalto-1 critical mission phases:

Before starting the science phase in which the payloads will perform their targeted science observations, the satellite will go through the commissioning phase (Figure 11, Ref. 8).

The commissioning phase will consist of booting up the satellite, making contact with ground and performing the required technical demonstration for the spectral imager and RADMON. The science phase is divided into two specific parts based on attitude and the downlink communication requirements.


Figure 11: Overview of the planned mission phases of Aalto-1 (image credit: Aalto University)

The technology demonstration phase has no time limit. Once the technology demonstration goals have been met and in general the commissioning phase has gone according to plan, the science phase can begin. This phase is divided into two particular phases, as seen in Figure 11, the first of which will consist of science observations performed with the AaSI and the RADMON instruments, and the latter for operating the plasma brake. This division of the science phase is due to the drastically different attitude, data rate and power requirements of the different payloads. The science observations phase is estimated to last 6–12 months, depending on the satellite and the sensor complement performance.

While performing the first part of the science phase, the satellite will assume either a nadir pointing or target tracking attitude and operate the AaSI and the RADMON instruments. Both instruments will rely during this time exclusively on the S-band downlink for transferring their collected data to the ground segment; they are limited by how much data can possibly be downlinked during the contact periods, in particular the imagery of the AaSI instrument. A preliminary estimate gives 29 – 49 MB/day for the S-band for an orbital altitude of 500 – 900 km.

An average image size expected to be taken by the spectral imager is estimated to be around 7.8 MB (512 x512 pixels with 15 channels of 16 bit each), which is for a minimum of 6 months of science observations phase. Thus, the AaSI instrument will take images several times per day, with a typical dynamical power usage profile as shown in Figure 12 of a few minutes duration.


Figure 12: Imaging sequence (not to scale) of the AaSI and the associated total satellite power profile at an orbital altitude of 700 km (image credit: Aalto University)

Legend to Figure 12: Only a few minutes are needed for the entire observation sequence. The GPS cold-start is by far the longest item in this sequence, the rest requiring only a few seconds per operation block. As the AaSI technology develops further, the power and time requirements will most likely be lower. The power values shown represent the total satellite power during the various operational blocks within the imaging sequence.

RADMON: The RADMON instrument is expected to operate during ~80% of the satellite's orbital period; it is switched off only during station passes (communications) and during an imaging sequence. The RADMON observations have a constant perpendicular angle with respect to the nadir and the satellite's velocity vector directions, the instrument will gather a steady average data rate of ~ 2 MB/day. The main targets of interest are the planet's high latitudes and the SAA (South Atlantic Anomaly) region (Figure 13).


Figure 13: Intended main targets for RADMON; a) 10 or more MeV proton fluxes at a 700 km orbit height. b) 700 keV or more fluxes at the same height, produced by the ESA SPENVIS (Space Environment Information System), image credit: Aalto University

The primary imaging targets are in Finland, both due to accessibility as well as good reference availability. After the AaSI instrument technology has been demonstrated, VTT expects at least one datacube time series to be imaged during the June-August growing season for the science observation phase. In addition, a “dark” image from the night side of the orbit and another one from near the noon-equator will be taken to assess the minimum and maximum lighting conditions of the imager.

EPB (Electrostatic Plasma Brake) demonstration: The plasma brake will be operated in the second part of the science phase, namely the plasma brake demonstration phase. During this phase, the spectral imager will not be used due to drastically different satellite attitude parameters compared to the science observations phase, while the RADMON will be used sporadically and depending on the system's instantaneous power demands.


Average power

Time taken

1) Satellite assumes a correct sun pointing attitude

3.2 W

few minutes

2) Spin-up to 200º/s

2.6 W

158 h

3) Tether reel-out to initial 10 m at 1mm s-1 rate (EstCube-1 CubeSat length)

5 W

~4 h

4) Initial experiment with 10 m of tether deployed in both positive and negative modes (with the radiation monitor on)

5 W

1 month

5) Reeling out the tether to 100m full deployment length

5 W

~ 34 h

6) Actual plasma brake experiment can begin, with the tether this time only in negative mode (with the radiation monitor on)

5 W

~ 1 year

7) Continuing until the satellite has deorbited, once the primary experiment has been completed (optional)

5 W

~ 1 year

Table 2: The electrostatic plasma brake demonstration phase steps

The EPB experiment itself has the goal to demonstrate the technology behind the tether and how well it interacts with the LEO ionosphere around the planet. When moving with respect to the plasma of the ionosphere, it will create a Coulomb drag force due to the charged particles interacting with the equivalently charged tether, with the resulting force vector along the ionosphere velocity vector (Figure 14 a).

As the drag is exerted on the tether and so the rest of the satellite, it will discernibly change its orbit parameters by slowing it down. With regular position and velocity determination, such as GPS, the drag effect created by the ionosphere can, thus, be measured over the period of this last phase, and so the experiment works nominally.


Figure 14: a) An example of the satellite moving in the magnetic field of the Earth. b) Illustrating how the tether is active only in the vicinity of the poles (within an angle of 20º), marked as angle Φ in a), image credit: Aalto University

Both positive and negative tether charging modes will be tested during this phase, as described in Table 2, and the negative mode requires only a voltage source and the tether itself, while the positive mode will need additionally 1–3 electron guns to achieve the necessary charge levels in the tether; the tether in this mode gathers electrons from the plasma and shoots them out using the electron gun.


2) Miniature Imaging Spectrometer for Aalto-1 Nanosatellite, 1st IAA Conference on University Satellite Missions and Cubesat Workshop, Rome, Italy, January 24-29, 2011, URL:

3) Jaan Praks, Antti Kestilä, Martti Hallikainen, Heikki Saari, Jarkko Antila, Pekka Janhunen, Rami Vainio, “Aalto-1 - an experimental nanosatellite for hyperspectral remote sensing,” Proceedings of IGARSS (International Geoscience and Remote Sensing Symposium), Vancouver, Canada, July 24-29, 2011

4) Antti Kestilä, Tuomas Tikka, Pyry Peitso, Jesperi Rantanen, Antti Näsilä, Kalle Nordling, Jaan Praks , Heikki Saari, Rami Vainio, Pekka Janhunen, Martti Hallikainen, “Science Operations of the Remote Sensing Nanosatellite AALTO-1,” Proceedings of the 4S (Small Satellites Systems and Services) Symposium, Portoroz, Slovenia, June 4-8, 2012

5) “Aalto-1: the first Finnish nanosatellite,” Feb, 5, 2012, URL:

6) “Aalto-1 Finnish Hyperspectral Remote Sensing Nanosatellite: Current Progress,” Fourth European CubeSat Symposium, ERM (Ecole Royale Militaire), Brussels, Belgium, Jan.30-Feb. 1, 2012


8) A. Kestilä, T. Tikka, P. Peitso, J. Rantanen, A. Näsilä, K. Nordling, H. Saari, R. Vainio, P. Janhunen, J. Praks, M. Hallikainen, “Aalto-1 nanosatellite – technical description and mission objectives,” Geoscientific Instrumentation Methods and Data Systems, Vol. 2, 2013, pp. 121-130, URL:

9) E. Razzaghi, A. Yanes, J. Praks, M. Hallikainen, “Design of a reliable On-Board Computer for Aalto-1 nanosatellite mission,” Proceedings of the 2nd IAA Conference on University Satellite Missions and CubeSat Workshop, IAA Book Series , Vol. 2, No 2, Editors: Filippo Graziani, Chantal Cappelletti, Rome, Italy, Feb. 3-9, 2013, paper : IAA-CU-13-08-03

10) Jaakko Jussila, “S-band transmitter for Aalto-1 nanosatellite,” Thesis submitted for examination for the degree of Master of Science in Technology, Aalto University, School of Electrical Engineering, Espoo, Finland, May 1, 2013, URL:

11) Anttti Näsilä, “Validation of Aalto-1 Spectral Imager Technology to Space Environment,” Thesis submitted for examination for the degree of Master of Science in Technology, Aalto University, School of Electrical Engineering, Espoo May 1, 2013, URL:

12) Antti Näsilä, Heikki Saari, Jarkko Antila, Rami Mannila, Antti Kestilä, Jaan Praks, Heikki Salo, Martti Hallikainen, “Miniature SpectraAalto-1 Spectral Imagerl Imager for the Aalto-1 Nanosatellite,” Fourth European CubeSat Symposium, ERM (Ecole Royale Militaire), Brussels, Belgium, Jan.30-Feb. 1, 2012

13) Antti Näsilä, Heikki Saari, Jarkko Antila, Antti Kestilä, Jaan Praks, Martti Hallikainen, “Miniature Imaging Spectrometer for Aalto-1 Nanosatellite,” First IAA Conference on University Satellite Missions and Cubesat Workshop, Rome, Italy, Jan. 28,2011, URL:

14) K. Viherkanto, H. Saari, A. Näsilä, “AaSI – Aalto-1 Spectral Imager development status,” 5th European CubeSat Symposium, Royal Military Academy, VKI (Von Karman Institute), Brussels, Belgium, June 3-5, 2013

15) P. Janhunen, A. Sandroos, “Simulation study of solar wind push on a charged wire: basis of solar wind electric sail propulsion,” Annales Geophysicae, EGU, Vol 25, pp. 755-767, 2007, URL:

16) P. Janhunen, P. K. Toivanen, J. Polkko, S. Merikallio, P. Salminen, E. Haeggström, H. Seppänen, R. Kurppa, J. Ukkonen, S. Kiprich, G. Thornell, H. Kratz, L. Richter, O. Krömer, R. Rosta, M. Noorma, J. Envall, S. Lätt, G. Mengali, A. A. Quarta, H. Koivisto, O. Tarvainen, T. Kalvas, J. Kauppinen, A. Nuottajärvi, A. Obraztsov, “Electric solar wind sail: Toward test missions,” Invited Article, Review of Scientific Instruments, Vol. 81, 2010, URL:

17) Osama Khurshid, Pekka Janhunen, Matthias Buhl, Aarto Visala, Jaan Praks, Martti Hallikainen, “Attitude Dynamics Analysis of Aalto-1 Satellite during deorbiting Experiment with plasma brake,” Proceedings of the 63rd IAC (International Astronautical Congress), Naples, Italy, Oct. 1-5, 2012, URL:

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.

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