CASSIOPE (Cascade SmallSat and Ionospheric Polar Explorer)

CASSIOPE is a new generation of small-satellite and multifunctional platform technology demonstration program of CSA (Canadian Space Agency) with the goal to serve both, namely scientific and commercial support applications in a variety of future Canadian space missions. The program is based on a PPP (Public Private Partnership) consortium - and developed under a Contribution Agreement between CSA, TPC (Technology Partnerships Canada), the University of Calgary, and MDA (MacDonald, Dettwiler and Associates). 1) 2) 3) 4) 5) 6) 7) 8)

The objectives call for a three-part mission:

1) Courier in the sky service: This involves an experimental payload, named Cascade, to demonstrate the world's first commercial space-based digital courier service based on store & forward file transfer techniques for message transfer. Cascade is designed to deliver data files of a size between 50 - 500 Gbyte at transmission speeds of 1.2 Gbyte/s over Ka-band when a ground station is within view of the spacecraft.

2) The science mission, called ePOP (Enhanced Polar Outflow Probe), is comprised of eight instruments. The objective is to measure the interaction of the Earth's upper atmosphere with the solar wind (space weather monitoring with greatly improved prediction capability). The ePOP team is comprised of scientists and engineers from Canadian universities and three research organizations: the University of Calgary (lead organization), York University, the Universities of Alberta, Athabasca, Saskatchewan, Western Ontario, and New Brunswick. The Communications Research Center (CRC), located in Ottawa, as well as ISAS (Institute of Space and Astronautical Science) of, Tokyo, Japan and the US NRL (Naval Research Laboratory) are also partners in the project.

3) Platform demonstration: The mission serves as a testbed for a Canadian spacecraft bus and subsystem functionality. The generic platform is considered for total mass budgets of 350-500 kg. The main goal of the program is to develop a minisatellite bus that may be used for future missions, including Earth observation, communications, space research and exploration.

CSA selected MDA Ltd. of Richmond, BC as the prime contractor for the space and ground infrastructure of the CASSIOPE Mission, including system engineering, design, assembly, integration, testing, launch and operation of the spacecraft. Within this context, Bristol Aerospace of Winnipeg, Manitoba, was selected to provide a preliminary design of a cost-effective multipurpose platform.

Under the Contribution Agreement, the University of Calgary will have ownership of the e-POP instrument payload while MDA will operate it, the CASCADE payload, and the CASSIOPE bus.


Figure 1: Artist's rendition of the CASSIOPE spacecraft (image credit: MAC, CSA)


The CASSIOPE system is composed of a minisatellite space segment and the associated ground segment. The so-called SmallSat bus structure is a CSA-sponsored initiative, a generic design approach for multimission functionality under development at Bristol Aerospace, a division of MAC (Magellan Aerospace Corporation). The spacecraft bus is referred to as MAC-200 (a generic bus of SciSat-1 flight heritage - which used the MAC-100 bus). The bus utilizes a fully redundant cross-strapped architecture and is capable of providing mission lifetimes of up to 7 years. CASSIOPE is the first mission to use the MAC-200 bus. Note: In the meantime, the MAC-200 bus is also known in Canada as the MMSSB (Multi-Mission Small Satellite Bus). 9) 10) 11) 12) 13)

The spacecraft structure is made up of aluminum honeycomb panels with aluminum face sheets, held together with machined aluminum brackets and stringers. Aluminum inserts are bonded into the panels to fasten down bus and payload units. The structure is held together with stainless steel fasteners. The spacecraft is of hexagonal shape with a diameter of 1.8 m and a height of 1.4 m. The spacecraft has a total mass of ~490 kg and a design life of 2 years.

The spacecraft is 3-axis stabilized. The ADCS (Attitude Determination and Control Subsystem) is designed as a zero-momentum system. Attitude is sensed with the µASC (Micro Advanced Stellar Compass) of DTU (Danish Technical University), CSS (Coarse Sun Sensors), and magnetometers; actuation is provided by 4 reaction wheels and torque rods. The MAC-200 bus provides an autonomous, active safe-hold mode to keep the spacecraft power-positive in the vent of an onboard anomaly detection. A passive safe hold can also be provided that can function without the ADCS. Use of nadir-ram or latitude-longitude target slewing.


Figure 2: The µASC of DTU (image credit: Bristol Aerospace)


Figure 3: Torque rod of SSTL (image credit: Bristol Aerospace)


Figure 4: Illustration of the CASSIOPE spacecraft in stowed configurarion (image credit: Bristol Aerospace, CSA MDA)

Thermal subsystem: Thermal control of the bus is primarily passive, supplemented with software-controlled operational heaters operating off the non-essential power bus. Thermostatically controlled survival heaters operate off the essential power bus. The equipment heat is rejected from strategically placed radiator surfaces along the bus hex sidewalls. MLI (Multilayer Insulation) radiatively isolates the spacecraft from solar heating and deep space cooling, and component-to-component heat transfer, as required.


Figure 5: Functional block diagram of the general MAC-200 configuration (image credit: Bristol Aerospace)

The C&DH (Command & Data Handling) subsystem consists of:

- A standard cPCI (Compact Peripheral Component Interface) bus that allows for future expendability of all cards on the bus and selection of external card manufacturers that support the cPCI bus

- Standard enclosure/backplane

- The C&DH subsystem is comprised of a controller card (CC), a data handling card, A PSC (Power Supply Card), an IOC (Input/Output Card), and MIC (Mission Interface Card). The MIC is specific to each mission and can be adapted, removed, or several can be included according to mission specific requirements.

- The controller card has a PowerPC processor with several MB of RAM and EEPROM, which provides ample processing power and memory margins for current and future software.

- Development of all software (flight, ADCS, and payload handling) on a single platform (PowerPC with VxWorks).

- The CASSIOPE mission does not have a propulsion system.

The EPS (Electric Power Subsystem) uses triple-junction surface-mounted solar cells (InGaP/GaAs/Ge) of Emcore with a power conversion efficiency of 27% (panels on 5 sides of the S/C). The battery is based on multiple strings of Lithium-ion cells placed 8 in series to generate a fully charged bus voltage of 33.6 V. The EPS provides an unregulated bus voltage at 28V ±6 V DC (payload power of up to 130 W average).

RF communications: Communications are provided via S-band (TT&C + ePOP) and Ka-band (350 Mbit/s). The protocol is CCSDS compatible. The spacecraft supports half-duplex transmission and reception of user data (large data packages) to and from the CX ground terminal over the Ka-band link. TT&C data rates are 125 kbit/s uplink and 1.6 - 4 Mbit/s in downlink.

Spacecraft stabilization

3-axis stabilized, zero momentum

Momentum management

Reaction wheels and magnetic torque rods

Attitude control

±0.04º (3σ) in pointing, including biases
±0.15º (3σ) in ground point tracking, including biases

Attitude knowledge

±0.015 º (3σ)

Orbit knowledge

Position: ±10 m (3σ) with GPS, velocity: +0.25 m/s (3σ) with GPS

C&DH architecture

Power PC, up to 1600 MIPS, VxWorks and C++, cPCI peripherals

Data downlink

S-band, bit rate programmable to 4 Mbit/s, CCSDS with optional Reed-Solomon and convolutional encoding

Command uplink

S-band, 4 kbit/s

Cascade data downlink

Ka-band, > 300 Mbit/s

E-POP dtata storage

Up to 1.5 GB/orbit

Spacecraft mass, size, power

~500 kg, 1.8 m Ø x 1.4 m, up to 600 W (average)

Table 1: Performance parameters of the MAC-200 bus


Figure 6: Photo of the CASSIOPE minisatellite prior to launch (image credit: University of Calgary)


Launch: The CASSIOPE minisatellite was launched on Sept. 29, 2013 on a Falcon-9 launch vehicle of SpaceX. The launch site was VAFB, CA.The Falcon-9 used for this launch the new Merlin 1D engines which have not flown so far. This new configuration of the Falcon is designated as Falcon 9 v1.1. 14) 15)

In 2011, CASSIOPE was in storage in Montreal awaiting its launch date, which will ultimately depend on the success of the upcoming Falcon launches.

The secondary payloads on this flight are:

• CUSat-1 and CUSat-2, microsatellites (each of ~41 kg) of Cornell University, Ithaca, N.Y.

• DANDE (Drag and Atmospheric Neutral Density Explorer), a microsatellite (<50 kg) of the University of Colorado at Boulder.

• POPACS (Polar Orbiting Passive Atmospheric Calibration Sphere). Three passive POPACS satellites were separated from a 3-Unit CubeSat dispenser installed on the second stage. The dispenser system released three independent spherical satellites immediately after deployment into an identical orbit.

POPACS is a collaborative a 3U CubeSat mission of several US universities and entities including: MSU (Morehead State University), Gil Moore (POPACS Project Director), the University of Arkansas, PSC (Planetary Systems Corporation), Silver Spring, MD, MSU (Montana State University), Drexel University (Philadelphia), et al.

Orbit: Elliptical polar orbit, 324 km x 1500 km, inclination =80º, period = 103 minutes (14 orbits/day).



Sensor complement: (ePOP, CX)

ePOP (Enhanced Polar Outflow Probe):

ePOP includes an innovative scientific probe carrying a suite of eight scientific instruments. The ePOP team is led by the University of Calgary (Institute of Space Research) and sponsored by CSA and NSERC (Natural Sciences and Engineering Research Council) of Canada. Other Canadian members of the team are: York University, University of Alberta, Athabasca University, University of Saskatchewan, University of Western Ontario, and the University of New Brunswick. In addition, three research organizations are partners of the project: the Communications Research Centre, Ottawa, JAXA/ISAS (Institute of Space and Astronautical Science), Tokyo, Japan, and the NRL (Naval Research Laboratory), Washington, D.C., USA.

The ePOP mission objective is to investigate atmospheric and plasma flows and related wave-particle interaction and radio wave propagation in the topside ionosphere. Specific objectives are to quantify the micro-scale characteristics of plasma outflow and related micro- and meso-scale plasma processes in the polar ionosphere, explore the occurrence morphology of neural escape in the upper atmosphere, and to study the effects of auroral currents on plasma outflow. The ePOP payload represents Canada's first mission contribution to ILWS (International Living With a Star) initiative. 16) 17) 18) 19) 20)

The ePOP package comprises three interconnected elements:

- a) In-situ observation on a polar orbit. The focus of ePOP is on quantitative in-situ measurements of charged particle distributions, waves, and fields, at the highest possible spatial-temporal resolution, and imaging and tomographic measurements of the large-scale auroral morphology and ionospheric topology.

- b) Coordinated ground-based observations

- c) Related modeling and data assimilation studies.


Figure 7: Overview of e-POP instruments on CASSIOPE (image credit: University of Calgary)



e-POP instruments: (IRM, SEI, FAI, NMS, MGF, RRI, GAP, CER)

IRM (Imaging and Rapid-scanning Ion Mass Spectrometer):

IRM was designed and developed at the University of Calgary (Peter Amerl and Andrew W. Yau.). The objective is to observe the mass composition and 3-D velocity distributions (energy and angular) of ions in the energy range of 0.1-100 eV and 1-40 amu/q. IRM employs several circuit boards that support a semi-autonomous ion detection and imaging. It consists of a semi-toroidal electrostatic deflection system, a pair of fast-switching time-of-flight electrodes, a hemispherical electrostatic analyzer, a microchannel plate (MCP) detector, and a 64 pixel anode. Its output data rate is driven by the incident ion flux and varies from 0-300 kwords/s.

IRM is an improved version of the Nozomi TPA (Thermal Plasma Analyzer) instrument and its predecessors. The IRM sensor is cylindrical and is deployed on an 8 cm boom. Its entrance aperture has a 360º FOV ( Field of View) perpendicular to the deployment axis and close to 120º FOV along the deployment axis.


Figure 8: Photo of the IRM sensor head and electronics unit (image credit: University of Calgary)


SEI (Suprathermal Electron Imager):

SEI was developed, built and tested at the University of Calgary (PI: David Knudsen). SEI measures the 2-D energy and angular distributions of thermal electrons and soft electrons in the range 1-200 eV/q with particular emphasis on photoelectrons in the 1 to 50 eV range, which are believed to play an important role in the polar wind outflow. The instrument is a derivative of a design by the University of Calgary flown on a sounding rocket campaign. SEI is implemented as a hemispherical electron analyzer with MCP electron amplification producing images on a phosphor screen. The latter images are optically coupled to a frame-transfer CCD located in the electronics module.

The SEI sensor is mounted on a boom of 80 cm in length to place it outside the spacecraft sheath. Its skin bias will be adjustable in its thermal mode to counteract the spacecraft potential and to control the input particle flux. SEI has a FOV of ±15º in elevation and 360º in azimuth. Its entrance aperture plane will be within 15º of the local magnetic field at high latitude.

SEI employs a 256 x 256 pixel CCD detector, which will be optically coupled to the electronics unit and de-coupled from the sensor bias voltages. It uses an MCP to amplify the incident electron azimuth/energy distribution and image it onto a phosphor screen that will be fiber-optically coupled to the CCD sensor. It is capable of operating at an imaging rate of 100 Hz.


Figure 9: Photo of the SEI sensor head, boom, and electronics unit (image credit: University of Calgary)


FAI (Fast Auroral Imager):

FAI was developed by a team of experts in partnership from the University of Calgary, Routes AstroEngineering, Burley Scientific, Keo Consultants Inc, and JENOPTIK Optical Systems, Inc. (Leroy Cogger). FAI images the auroral emission in the near infrared (NIR) and visible (VIS) spectral range. The major scientific objectives of the mission are associated with the Earth's high latitude atmosphere and ionosphere. The first thrust is to investigate the outflow of plasma from the polar regions and the related processes of micro-scale ion acceleration and wave-particle interaction, and auroral excitation. The second thrust is the study of 3D ionospheric irregularities using both active and passive radio techniques. The third emphasis is on the escape of neutral particles from high latitudes caused by temperature enhancements as well as by non-thermal processes. 21) 22)

While the spatial resolution capability for most auroral imagers over the past two decades has been of the order of tens of km, the need in the ePOP mission was for an order of magnitude improvement. Likewise, a similar improvement in repetition rate was required. The challenge was met by carefully selecting the spectral elements to image and by taking advantage of the best available technology within the constraint of a very limited budget.

The instrument is a coaligned dual-head CCD imaging camera with detectors for NIR and narrow-band VIS (630 nm), intended to operate only over the polar regions of the orbit. The sensitivity of the cameras was maximized through the choice of a CCD detector and optics module. A thinned, backside-illuminated, high quantum efficiency and low noise CCD was selected, the E2V CCD67 in a 256 x 256 pixel array. The quantum efficiency is 0.8 at 630 nm. With two-stage thermoelectric cooling (TEC) of the AIMO (Asymmetric Inverted Mode Operation) device, the dark current can be kept insignificant for normal instrument operating temperatures. An f/4 telecentric lens system is being used in the optics unit with a 5:1 fiber-optic taper to provide an effective f-number of 0.8. Typically, FAI takes imagery at a rate of about 60 0.1 second exposures per minute with the NIR camera, and 1/2 second exposure per minute with the VIS camera.


SV camera (630 nm)

SI camera (NIR)

Pixel array

128 x 128

256 x 256

Projected pixel size

4.7 km

2.6 km (from apogee)

Projected pixel size

0.4 km

0.4 km (from perigee)

Exposure time

0.5 s

0.1 s

Exposure interval

30 s

1 s

Sensitivity (SNR > 3)

200 rayleigh

100 rayleigh (557.7 nm equivalent)

Table 2: Selected default parameters (operation modes) for each camera to meet the science requirements

The default operating modes are given in Table 2 for each camera. Calculations of projected pixel size are shown for a satellite orbit apogee of 1500 km and perigee of 300 km. The assumed altitude of the aurora is 220 km for the 630 nm and 110 km for the near infrared emissions. A single pixel of size 26 µm x 26 µm limits the optical resolution of the camera and subtends an angle of about 0.1º.


Figure 10: Cutaway view of the FAI camera unit (image credit: University of Calgary)


Figure 11: Photo of the FAI camera unit (image credit: University of Calgary)

FAI will aid in the investigation of the relationship between auroral emissions and ion bulk upflow, heating, acceleration, field-aligned currents, and plasma waves. In more specific terms, the FAI was designed to support the investigation of the following:

- The small-scale mechanisms for auroral excitation

- The phenomenon of black aurora and its relation to streaming electrons and ions

- The optical signatures of large-scale convection and their relation to radar measurements

- The connections between aurora and ion energization and bulk upflow.

The FAI imager system will produce 16 bit digital images of the near infrared band at one image per second (again taking advantage of the non-rotating platform on CASSIOPE), and the 630 nm wavelength at two images per minute, giving adequate temporal resolution to investigate the above scientific objectives.


NMS (Neutral Mass and Velocity Spectrometer):

NMS was developed by JAXA/ISAS, Japan (PI: Hajime Hayakawa). NMS measures the density and velocity of neutral atmospheric species using an open-source electron impact ionization chamber and a microchannel plate (MCP) imaging detection subassembly. It employs a planar entrance aperture with a plasma filter to repel low energy charged particles.

NMS is only being operated in the perigee and the near-perigee phase of the orbit (300-400 km) where the atmospheric density is highest. In orbit, neutral particles are rammed into the entrance slit. Internally, NMS employs a high-voltage electron gun to ionize the neutral particles, and an electric-field deflection and focusing system to image the ions onto the MCP detector, which amplifies the ion image by placing then onto a CCD detector of 256 x 256 elements. The CCD data is binned and compressed within the instrument to reduce the data rate.


Figure 12: Photo of the NMS mounted on the e-POP deck of CASSIOPE (image credit: University of Calgary)


MGF (Magnetic Field Instrument):

MGF was developed by Magnametrics of Ottawa ain partnership with Bennest Enterprises and Narod Geophysics (Lead: Don Wallis). The primary objective of MGF is the characterization of electric currents flowing to and from the high latitude (auroral) ionosphere. The goal of the MGF is the detection of 0.1 µAm-2 currents or stronger at scale sizes of 100 m or larger. This calls for an instrument with resolution of ~0.1 nT and a sample rate of 160 per second. These currents perturb the terrestrial magnetic field causing it to tilt and thereby changing the mapping of in-situ measurements to the auroral phenomena below. This tilt is variable and up to 2º or 3º, and causes the magnetic footprint of the spacecraft to be up to 5 km away from the undisturbed location for every 100 km of altitude above the E-region.

The success of MGF is contingent upon determining the spatial distributions of the field aligned currents in relation to the measurements of other on-board instrumentation. The only means of making this determination is through the perturbations of the geomagnetic field associated with these currents.

For this reason a specially designed fluxgate vector magnetometer is employed. MGF employs dual sensors on a 80 cm carbon fiber boom, at two separations to estimate and correct for the spacecraft influences. The instrument measures the vector magnetic field (in 3 digital output channels) with a precision of 21 bit, a resolution of 0.0625 nT and a dynamic range of ±65,536 nT. The sampling rate is 160 Hz. The MGF relies on a unique ranging design that was previously adopted successfully in the design of the CANOPUS magnetometers. 23) 24)


Figure 13: View of the dual ring dual wound MGF sensor (image credit: Magnametrics)


Figure 14: Photo of the MGF boom and sensors (image credit: University of Calgary)


RRI (Radio Receiver Instrument):

RRI developed at CRC (Communications Research Centre),Ottawa, Canada (Lead: Gordon James). The instrument measures radio waves in the ULF and HF frequency range (up to 18 MHz). RRI is a four-channel ULF/HF digital radio receiver is operating in burst mode when it is in the beams of cooperating HF ground radars such as CADI (Canadian Advanced Digital Ionosondes) and SuperDARN (Super Dual Auroral Radar Network), an international mostly ground-based distributed radar network for studying the Earth's upper atmosphere, ionosphere, and connection into space. The burst mode lasts during the acquisition phase of the orbital overflight (several minutes) of the ground radars. 25) 26) 27) 28) 29)

The RRI is being fed by four 3 m monopoles with preamplifiers. From below 100 Hz to about 3 MHz, the RRI measures the electric fields of spontaneous waves. Between about 10 kHz and 18 MHz, the receiver measures the electric fields of waves created by ground transmitters, such as ionosondes, HF radars and ionospheric heaters. The scientific objectives of the RRI program are to improve understanding in the following areas:

- The morphology and dynamics of density structure in the ionosphere

- The generation of spontaneous radio emissions created by auroral processes

- The nonlinear plasma physics of the HF-modified ionosphere

- The physics and metrology of radiowave scattering, diffraction and refraction.

The RRI measures and record various parameters associated with RF wave electric fields E incident upon the spacecraft. The magnitudes of E in the case of discrete modulated CW signals emitted by artificial sources on the ground below require to be accompanied by information about the wave polarization or DOA (Direction of Arrival) of the incident waves. The detection of the Doppler shift of artificial signal carrier frequencies constitutes an independent tool for checking the direction of propagation of waves. The absolute time stamp (to an accuracy of ±8 µs) permits the determination of the absolute signal delay time of waves from the ground.

The RRI instrument design features a dynamic range of 120 dB above an input threshold of 0.3 µV. This total range is achieved by dividing it into three overlapping preamplifier gain settings, each of 72 dB. Bandwidths up to a maximum of 30 kHz are available in each of the four receiver channels. Both the amplitude and relative phase of an incoming electric field can be measured. After suitable amplification, the signals are digitized and then processed by a digital down converter and decimating low-pass filter. The resultant digital data are time-tagged to an accuracy < ±8 µs. The system has four identical signal processing channels, of which two are shown in Figure 15. Each channel is tied to a single monopole antenna.


Figure 15: Functional block diagram of two of the four RRI channels (image credit: CRC)


Figure 16: Photo of the RRI Digital Radio Receiver Module with antennas stowed, mounted on CASSIOPE (image credit: University of Calgary)


GAP (GPS Attitude and Positioning Experiment):

GAP is of the UNB (University of New Brunswick). The objective is to measure S/C velocity and attitude as well as TEC (Total Electron Content) of the ionosphere. The relative phase delay of signals in both the L1 and L2 bands from a GPS satellite occulted by the limb ionosphere provide large-scale (1000's of km) information on how the total electron content responds to magnetospheric perturbations. GAP consists of two components: 30) 31) 32)

- GAP-A: Three single-frequency (L1) GPS receivers and four patch antennas. This package is being used to provide an accurate absolute time reference, spacecraft 3-D attitude, and post-processed spacecraft position and velocity.

- GAP-O: The objective is to provide ionospheric tomography observations. GAP-O consists of a dual-frequency (L1/L2) GPS receiver with a phase data sampling rate of > 10 Hz, sufficient for tomographic analysis.


Figure 17: Photo of the GAP GPS antennas (image credit: University of Calgary)

The four GAP-A patch antennas are mounted on opposite sides of the S/C in the zenith direction to minimize multipath reflections from the S/C structure and to maximize the baseline between the antennas - to improve the accuracy of the 3-D attitude.

For GAP-O, the helical antenna is being placed on the X side of the S/C. Phase measurements are acquired for several minutes (5-7) per occultation from a GPS satellite in occultation using the dual-frequency receiver.

GAP distributes also its time signals to certain ePOP instruments for precise time stamping of science data via the DHU (Data Handling Unit) of ePOP.

GAP employs five differential Global Positioning System receivers and associated antenna complement to provide the e-POP payload with high-resolution spatial positioning information, flight-path velocity determination, and real-time, high-stability timing. In addition, by measuring the arrival times of the various GPS signal wave fronts at each antenna against a very stable time base, the relative range between antennas can be determined, yielding real-time spacecraft attitude determination. One of the GAP antennas is mounted on the anti-ram side of the spacecraft and dedicated to ionospheric radio occultation measurements in which the relative phase delay of the measured L1 and L2 signals (at frequencies of 1.57542 GHz and 1.2276 GHz, respectively) from different satellites of the GPS constellation will be used to determine the electron density profile of the ionosphere using tomographic techniques. The other four antennas are mounted on the anti-nadir face of the spacecraft, with bore-sights normal to the spacecraft surface, and are used primarily for the timing, position, and velocity measurements.


Figure 18: Block diagram of GAP (image credit: UNB)


CER (Coherent Electromagnetic Radiation):

CER was developed at NRL (Naval Research Laboratory) in Washington, D.C. (Paul Bernhardt). The objective is to determine ionospheric total electron density (TEC) by using a three-frequency beacon; cooperative ionospheric observations with fixed ground receivers. The instrument is of CERTO (Coherent Electromagnetic Radio Tomography Instrument) heritage flown on ARGOS of DoD (launch Feb. 23, 1999) and on PICOSat (launch Sept. 30, 2001) of DoD.

CER consists of a tri-frequency beacon and a three-frequency, circular polarized antenna on a 68.6 cm boom. CER provides a global ionospheric map to aid the prediction of radio-wave scattering. CER transmits three beacon signals near 150, 400, and 1066 MHz (VHF/UHF bands), for reception by a network of ground receiving stations located around the globe. Data are collected from these receiving stations and analyzed, producing 2-D ionospheric density maps.


Figure 19: Photo of the CER antenna (image credit: NRL)



CX (Cascade Experimental), pilot phase of a new information delivery service

CX is an advanced satellite communication demonstration payload with twofold mission objectives:

- a) it must develop and demonstrate the key Cascade enabling technologies

- b) it must demonstrate the feasibility of the unique very large file end-to-end transfer method, envisioned in support of a future Cascade digital courier service.

Key enabling Cascade technologies to be demonstrated include: Ka-band RF transmit and receive chains, 350 Mbit/s digital space-qualified modulators and demodulators, and high-capacity, low-power and low-mass onboard bulk data storage (1+ Tbit of onboard memory). CASCADE uses the unique GigaPackage ExpressTM store-and-forward technology developed by MDA (pickup and delivery of huge data packages: 50-500 GByte/day).

The CX users are responsible for proof of the Cascade technologies. The CX users request data delivery service via the CX ground terminal. In turn, they provide user data over a high-speed line to the CX ground terminal for uplink service to the spacecraft. 33)

The CX Cascade system provides:

- Daily accessibility to any point on the globe excluding the polar regions

- High bandwidth data links

- Large capacity storage onboard the spacecraft

- Service on a regular and reliable basis

Over the years, the CX design has evolved into a new system called PCPMU (Payload Controller, Processor and Memory Unit).


Figure 20: Conceptual view of the Cascade delivery service (image credit: MDA) 34)


PCPMU (Payload Controller, Processor and Memory Unit):

The innovative and compact PCPMU design challenges the conventional approach to PE (Payload Electronics) by incorporating a modular architecture utilizing highly integrated, multi-purpose components (based on reconfigurable, state-of-the-art FPGA technology) and standardized interfaces. This approach introduces a scalable, multi-mission capability providing for a rapid, low-cost PE assembly, test and integration that results in a reduction in complexity, mass and volume, and a corresponding increase in reliability. 35)

The PCPMU architecture is based on the use of DMFC (Digital Multi-Function Card), a software reconfigurable digital signal processing card that can be programmed to perform a variety of mission specific functions.

The PCPMU comprises two main sections: a Controller/Processor section that is mission specific, and a DSU (Data Storage Unit) section that is common for all configurations.

• The Controller/Processor section, based on the DMFC, can be configured to provide a variety of core functions: overall payload control and interface to bus and payload instrument, DC power conditioning, echo or replica data processing for SAR (Synthetic Aperture Radar), or image data compression for optical missions, and on-board data processing (modulation and demodulation) for store-and-forward communication missions.

• The DSU (Data Storage Unit) is a highly scalable unit, also based on DMFC that provides control and data management function, and integrates at least two memory cards, each with a capacity of 1 Tbit at EOL (End-of-Life). Additional memory boards can be added to achieve capacity of up to 8 Tbit at EOL. This storage unit offers significantly reduced power consumption compared to the traditional solid state mass memory solutions as the unit uses non-volatile memory and can be switched off completely when not in use.

The Communication PCPMU architecture is developed to provide on-board data processing and storage for the CASCADE mission – a fleet of low-cost LEO small satellites for high bandwidth, store-and-forward communications. CASCADE enables the transfer of large digital data files (10s to 100s of GBytes) from/to anywhere in the world, including polar areas, within hours. It uses Ka-band links at aggregate data rate of up to 1.4 Gbit/s (QPSK modulation), or up to 2.1 Gbit/s using higher order modulations.

The Communications PCPMU consists of a PU (Processing Unit) and a DSU (Data Storage Unit). The PU comprises five identical, independent MDMs (Modulator/Demodulator) assemblies, and a pair of (prime and redundant) FGMs (Frequency Generator Modules) and EPCs (Electric Power Control) cards.

The communication DSU functionality is similar to the optical DSU. The CDI (Controller and Data Interface) card, besides data routing, formatting, and file management, also performs payload operational control, interfaces with the spacecraft bus over the TT&C interface, and acts as a master on the internal CAN (Controller Area Network) bus. The communication PCPMU architecture is presented in Figure 21.



Figure 21: Communication PCPMU architecture (image credit: MDA)

The CASSIOPE mission will carry store-and-forward communication payload enabling 0.7 Tbit/s data transfer rate and will have data storage capacity of 1 Tbit. Replacing the conventional PE (Payload Electronics) with the PCPMU configuration will bring mass and peak power consumption down by 64% and 58%, respectively.

The PCPMU has a mass of 21 kg and a peak power consumption of 135 W (64 W average).


Figure 22: Schematic view of CASSIOPE payload electronics (image credit: MDA)


Figure 23: Photo of the CASCADE demodulator subassembly (image credit: MDA)

The PCPMU architecture and the design methodology focuses on creating reconfigurable Payload Electronics (PE) based on card/module reuse and utilization of reconfigurable cards that minimizes non-recurring engineering at the payload and overall mission level.

The PCPMU design utilizes a combination of space qualified and screened OTS (Off-The-Shelf) parts. A significant contributor to the performance and low price points of the PCPMU is the application of a flight proven, NASA approved process that enables cost effective application of OTS parts for space.



Ground segment:

The ground segment consists of MCC (Mission Control center), the CX ground terminal, the CX SCC (Service Control Center) and GSE (Ground Support Equipment). In addition, there is the SOC (Science Operations Center) in support of the ePOP payload. SOC will be operated at the University of Calgary. The SOC schedules and coordinates the ePOP instrument data collection. SOC receives the ePOP data via S-band downlink over Mission Control and over Ka-band via the CX ground terminal.


Figure 24: Overall architecture of the CASSIOPE mission

1) Andrew Yau, Gordon James, Gregory Enno, Robert Hum, Peter Duggan, Mark Senez, Ziad Ali, Gilles Brassard, Berthier Desjardins, Luc Dubé, Richard Giroux, Dave Beattie, Ian Walkty, “Multidimensional Challenges and Benefits of the CASSIOPE Mission,” Proceedings of the IAA Symposium on Small Satellite Systems and Services (4S), Rhodes, Greece, May 26-30, 2008, ESA SP-660, August 2008

2) A. W. Yau, E. P. King, R. Hum, H. G: James, W. Ressl, S. Oldham, A. Dubeau, Z. Ali, W. Liu, R. Shelly, “The Canadian Enhanced Polar Outflow Probe on the Canadian CASSIOPE SmallSat,” Proceedings of IAC 2004, Vancouver, Canada, Oct. 4-8, 2004, IAC-04-IAA.

3) G. B. Giffin, W. Ressl, A. W. Yau, E. P. King, “CASSIOPE: A Canadian Smallsat-Based Space Science and Advanced Satcom Demonstration Mission,” Proceedings of AIAA/USU Conference on Small Satellites, Logan, UT, Aug. 9-12,2004, SSC04-VI-5

4) G. Giffin, K. Magnussen, M. Wlodyka, J. Bravman, “CASCADE: A SmallSat System Providing Global, High Quality Movement of Very Large Dat Files,” Proceedings of IAC 2004, Vancouver, Canada, Oct. 4-8, 2004, IAC-04-M.4.06


6) J. Bates, “Canada Funds Small Satellite Platform Demonstration,” Space News, Feb. 16, 2004, p. 8

7) A. W. Yau, H. G. James, “The Canadian CASSIOPE Enhanced Polar Outflow Probe (e-POP) for High-Resolution Observations of SpaceWeather Processes,” 58th IAC (International Astronautical Congress), International Space Expo, Hyderabad, India, Sept. 24-28, 2007, IAC-07- B.2.06

8) “CASSIOPE/e-POP Fact Sheet,” University of Calgary, URL:

9) S. Page, R. Harris, I. Walkty, D. Beattie, H. Dahl, “The Development of the MAC-200 Small Satellite Bus for the Canadian Space Agency,” Proceedings of the 13th Canadian Astronautics Conference, ASTRO 2006, Montreal, QC, Canada, organized by CASI (Canadian Astronautics and Space Institute), April 25-27, 2006

10) A. Denis, S. Page, I. Walkty, “The Evolution of a SciSat-1 Spacecraft to Provide a Generic Small Satellite Bus for the Canadian Space Agency,” Proceedings of IAC 2004, Vancouver, Canada, Oct. 4-8, 2004, IAC-04-IAA.4.11.2

11) W. Peruzzini, G. Brassard, “Canadian Multi-Mission Small and Micro Satellite Buses,” Proceedings of IAC 2004, Vancouver, Canada, Oct. 4-8, 2004, IAC-04-IAA.


13) Y. Soucy, A. Woronko, P. Tremblay, M. O’Grady, “Force limited vibration testing applied to the CASSIOPE spacecraft,” Proceedings of ASTRO 2010, 15th CASI (Canadian Aeronautics and Space Institute) Conference, Toronto, Canada, May 4-6, 2010

14) Patrick Blau, “SpaceX successfully Launches Upgraded Falcon 9 on 1st Demonstration Flight,” Spaceflight 101, Sept. 29, 2013, URL:

15) Marc Boucher, “Canada's CASSIOPE Satellite Nearing Liftoff,” SpaceRef, June 26, 2012, URL:

16) “e-POP - Enhanced Polar Outflow Probe,” URL:


18) H. G. James, P. A. Bernhardt, R. B. Langley, C. L. Siefring, A. W. Yau, “Radio-science experiments with the Enhanced Polar Outflow Probe satellite payload using its RRI, GAP and CERTO instruments,” URSI GA2005 New Delhi, India, Oct. 23-29, 2005, URL:

19) M. Nejad Ensan, D. G. Zimcik, “Response of the Enhanced Polar Outflow Probe (e-POP) Instrument Under Shock Loading,” Proceedings of the 11th International LS-DYNA Users Conference, Detroit, MI, USA, June 6-8,, 2010, URL:

20) A.W. Yau, C. Alonso, G.A. Enno, M. Grigorian, R.H. Hum, H.G. James, R. Giroux, P. Langlois, B. Poller, M. Senez, “The Canadian CASSIOPE Small Satellite Mission: The Enhanced Polar Outflow Probe and CASCADE Technology Demonstration Payloads,” Proceedings of the 4S (Small Satellites Systems and Services) Symposium, Portoroz, Slovenia, June 4-8, 2012

21) L. Cogger, S. Murphree, T. Trondsen, A. Yau, D. Asquin, B. Gordon, P. Marchand, D. Ng, G. Burley, M. Lessard, “The Fast Auroral Imager Experiment to Investigate the Dynamics of Nighttime Optical Aurora as part of the CASSIOPE/ePOP Mission,” Proceedings of the 13th Canadian Astronautics Conference, ASTRO 2006, Montreal, QC, Canada, organized by CASI (Canadian Astronautics and Space Institute), April 25-27, 2006


23) D. D. Wallis, B. B. Narod, J. R. Bennest, “The ePOP / CASSIOPE Magnetic Field Instrument,” Proceedings of the 13th Canadian Astronautics Conference, ASTRO 2006, Montreal, QC, Canada, organized by CASI (Canadian Astronautics and Space Institute), April 25-27, 2006


25) H. G. James, W. H. H. J. Lunscher, “The Radio Receiver Instrument on CASSIOPE/ePOP,” Proceedings of the 13th Canadian Astronautics Conference, ASTRO 2006, Montreal, QC, Canada, organized by CASI (Canadian Astronautics and Space Institute), April 25-27, 2006

26) H. G. James, C. L. Siefring. E. P. King, ”The Radio Receiver Instrument on the e-POP small satellite”, Proceedings of the 12th Conference on Astronautics, Astro2002, CASI, Ottawa, Canada, Nov. 12-14, 2002.

27) G. C. Hussey, H. G. James, G. J. Sofko, “Radar Observations of the Ionosphere from Space: A Preliminary Instrument Proposal,” Proceedings of the 13th Canadian Astronautics Conference, ASTRO 2006, Montreal, QC, Canada, organized by CASI (Canadian Astronautics and Space Institute), April 25-27, 2006



30) Don Kim, “The e-POP GAP project,” URL:


32) Don Kim, Richard B. Langley, “GPS RTK-Based Attitude Determination for the e-POP Platform onboard the Canadian CASSIOPE Spacecraft in Low Earth Orbit,” TimeNav'07 (ENC GNSS 2007), 29 May - 1 June 2007, Geneva, Switzerland, URL:

33) B. Gordon, L. Piché, J. Bourdeau, S. Bazley, K. Magnussen, G. Giffin, L. Laba, J. Tulip, “A High Speed, High Capacity, Data Storage Unit based on Digital FLASH Technology for Next Generation Space Missions including CASSIOPE/CASCADE,” Proceedings of the 13th Canadian Astronautics Conference, ASTRO 2006, Montreal, QC, Canada, organized by CASI (Canadian Astronautics and Space Institute), April 25-27, 2006


35) Sasa T. Trajkovic, George Tyc, Kenneth V. James, Daniel J. Schulten, Peter Allan, Ed Ahad, Richard Allen, Simon Ladouceur, Martin Cote, “PCPMU: A Modular, Multi-Use Payload Electronics Architecture for Affordable, Responsive Missions,” 7th Responsive Space Conference, April 27–30, 2009, Los Angeles, CA, USA, AIAA-RS7-2009-3006, paper URL:, presentation 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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