Attitude Control System for CubeSats – Summary from
Literature
Introduction to Attitude Control in CubeSats
CubeSats are compact satellites used for academic and industrial space missions. After
deployment, they tend to spin uncontrollably due to separation forces. This uncontrolled rotation
(tumbling) must be corrected through a process called detumbling. For this, an Attitude
Determination and Control System (ADCS) is essential.
Traditional ADCS Setup
A common setup includes magnetorquers for initial detumbling and reaction wheels for fine
orientation control. Magnetorquers are magnetic actuators that generate torque using Earth's
magnetic field. Reaction wheels adjust orientation by spinning and exchanging angular
momentum with the spacecraft.
However, reaction wheels take up space, and about half of ADCS failures are due to
mechanical issues in moving parts like these wheels. Moreover, magnetorquers are often built
as long cylindrical rods, which are bulky for CubeSats.
Hybrid ADCS Architecture
To overcome these limitations, a hybrid ADCS was proposed. It combines:
●​ Asymmetric planar magnetorquers, which are flat coils printed directly onto the PCB
using copper traces.​
●​ COTS (Commercial Off-The-Shelf) reaction wheels, which are compact and
cost-effective.​
This combination helps in strong detumbling and precise orientation while saving space. The
planar coils have non-uniform winding widths to enhance torque performance and can function
independently if the reaction wheels fail, adding fault tolerance to the system.
Coil Configuration and Control
The planar coils are arranged along the satellite's X, Y, and Z axes to provide full 3D control.
Each coil is controlled by adjusting the current through it. The shape and width of these copper
traces affect performance — wider traces allow more current and reduce resistance but
increase heating. Two coil types were studied:
●​ C2 coils (linear geometry): Higher current, faster detumbling, more torque, but more
power and heat.​
●​ C5 coils (quadratic geometry): Lower current and torque, slower detumbling, but better
energy efficiency.​
Planar coils enable a compact design but consume around 3× more power than traditional
torque rods.​
​
Power and Performance Trade-offs
Power usage increases with more coils. In series configurations, the current stays the same but
voltage increases; in parallel, voltage is constant but current rises. A good coil setup gives a
high magnetic dipole moment while keeping power low.
Performance improves with:
●​ More coils​
●​ Better geometry​
●​ Lower resistance​
The goal is to maximize torque and minimize power consumption. These trade-offs make
planar coils a compelling choice despite their lower efficiency.
Control Algorithm: Projection-Based B-dot Control
The ADCS uses a projection-based B-dot algorithm, ideal for detumbling. It observes how
fast Earth’s magnetic field changes around the CubeSat and generates opposing torques using
magnetorquers. This algorithm projects the angular velocity onto the magnetic field direction to
optimize control.
It’s simple, computationally light, and suited for CubeSats with limited onboard computing. But
there are challenges: if the magnetic dipole or angular velocity aligns with Earth’s field, the
system may fail to generate useful torque or current. These conditions are rare and temporary
because Earth’s field keeps changing as the satellite moves.
Combined Operation with Reaction Wheels
During detumbling, reaction wheels may take over once the satellite is stable. When these
wheels reach their speed limits (saturation), magnetorquers are reactivated to dump
momentum. However, the papers did not describe how desaturation was specifically
implemented.
In one study, both magnetorquers and reaction wheels were run together to measure power
usage and detumbling speed. This hybrid mode helps balance performance and energy use.
Thermal Modeling
Thermal analysis was done only for heat generated by the coils, excluding other components
like processors or batteries. Wider traces (lower track width ratio) reduce resistance and
increase current, leading to higher temperatures. Proper thermal management is necessary to
ensure safe operation.​
​
System Performance and Achievements
The proposed system achieved fast and stable detumbling for large CubeSats using embedded
planar coils and off-the-shelf reaction wheels. It saves internal space, is modular, and
outperforms several commercial and research systems. It also supports reconfiguration,
adapting power levels and heat output to the mission’s needs.
Future directions suggest combining this ADCS with small thrusters for even more advanced
maneuvers in micro-spacecraft.
Identified Gaps and Limitations
1.​ B-dot Limitations: Works well early in detumbling but less effective in steady-state.
Could benefit from adaptive or scheduled control enhancements.​
2.​ No Momentum Dumping Strategy: Lacks detailed handling of reaction wheel
desaturation.​
3.​ Isolated Thermal Model: Ignores internal heat sources, requiring a full thermal system
model.​
4.​ High Power Draw: Planar coils consume significantly more power. Strategies like PWM
or current shaping could help reduce average draw.​
5.​ No Sensor Fusion or Magnetometer Drift Handling: Real-time feedback from
magnetometers isn't validated. Sensor fusion using gyros and sun sensors could
improve accuracy.​
6.​ No Fault Detection or Recovery: Currently, the system lacks fault-tolerant logic for
automatically switching control modes upon failure.​
​
Reliability Physics of Momentum Wheels
Momentum wheels are crucial for precision control and depend heavily on the bearings and
their lubrication. A typical wheel includes:
●​ Bearings (with ball, retainer, and race)​
●​ Brushless DC motor​
●​ Non-magnetic stainless steel wheel​
●​ Lightweight outer shell​
Critical design note: The three bearing elements (ball, retainer, race) must not contact
directly — such contact causes friction, wear, and failure due to fatigue. To prevent this:
●​ A lubrication film is used, based on electrohydrodynamic principles.​
Overall reliability of the reaction wheel is tied closely to the performance and longevity of its
bearings and lubrication system.
FORESAIL-1 CubeSat – Magnetorquer-Based Attitude
Control and Spin Stabilization
FORESAIL-1 is a 3U CubeSat developed by the Finnish Centre of Excellence with both
scientific and technological objectives. Scientifically, it aims to study energetic particle
precipitation into Earth’s atmosphere and solar energetic neutral atoms. Technologically, it
serves as a testbed for a plasma brake, which is a device used to slow down the satellite for
deorbiting—contributing to space sustainability by preventing space debris. This mission is part
of a broader multi-mission effort to support safer, cleaner, and longer-lasting space missions.
Mission Requirements
The satellite is required to operate in two spin-stabilized modes. The Precise Spin Mode
demands a spin rate of 24°/s with a tight tolerance of ±0.24°/s, primarily for maintaining
accurate pointing of the particle telescope at the Sun. The High Spin Mode, on the other hand,
involves spinning at 130°/s to enable the deployment of the plasma brake. Maintaining attitude
stability during these fast spin modes is critical, especially for the low-spin mode that demands
high precision.
System Design and Engineering Approach
The team began with requirement analysis, identifying the exact performance and precision
needs of both spin modes. A trade-off analysis was then conducted between different actuator
types. Magnetorquers were chosen over alternatives like reaction wheels due to their simplicity,
reliability, low power consumption, and compactness—ideal for small satellites like CubeSats.
For the magnetorquer design, air-core coils (coils without magnetic cores) were selected.
These were controlled using H-bridge circuits, which allow the direction of current to be
reversed, and filters to smooth signal fluctuations. The team studied whether to connect the
coils in series or parallel, optimizing for efficiency and control authority. By modeling the coils
as equivalent electrical circuits, they were able to determine the optimal driving frequency,
i.e., how fast the current should switch on and off to produce effective torque.
To enhance the magnetic dipole generation efficiency, extrapolation techniques based on
model predictions were applied. This helped in optimizing how effectively the magnetorquers
could generate torque under different operating conditions.
Custom Manufacturing and Integration
One of the unique features of FORESAIL-1 is its custom in-house manufacturing process.
The team used a dedicated 3D printer setup to build the magnetorquers. This not only reduced
costs but also ensured that the components were precisely tailored to the satellite’s size and
performance constraints. The magnetorquers were tightly integrated into the embedded system
hardware, maintaining a compact and efficient design.
Testing and Validation
The performance of the attitude control system was validated through a three-tier testing
strategy:
1.​ Hardware Performance Tests – Real-world testing of the built magnetorquers to ensure
they met expected torque outputs.​
2.​ Circuit Simulations – Virtual models of the electrical circuits were tested to predict
behavior under various input conditions.​
3.​ Attitude Simulations – End-to-end simulations of the satellite's attitude dynamics in
orbit were run to verify that the system could achieve the desired spin rates and maintain
pointing accuracy.​
Why Spinning Matters in CubeSats
Spinning plays a critical role in CubeSat attitude control and system design. First, it provides
passive stability via angular momentum. Like a gyroscope, a spinning satellite resists
changes in its orientation, maintaining a stable pointing direction unless acted upon by external
forces. This is especially useful for maintaining the telescope’s alignment with the Sun.
Second, spinning supports precise pointing. Instruments like particle telescopes require
alignment with specific celestial targets (e.g., the Sun), and spin stabilization allows consistent
pointing without complex control systems.
Third, spinning is necessary for the deployment of tethers or booms. In FORESAIL-1, the
plasma brake is deployed using centrifugal force generated by spinning the satellite at 130°/s.
Finally, spinning is simple and energy-efficient. Compared to three-axis stabilization
systems—which require multiple sensors and actuators—spin-stabilization is easier to
implement and consumes less power. This is especially beneficial during early mission phases
like detumbling after launch.
However, spin stabilization is not suitable for missions requiring very high pointing accuracy,
such as Earth observation or detailed imaging, and it may limit communication if the antenna
must face a specific direction.
it's clear that high spin rates (like 130°/s) are technically feasible. However, the real challenge
lies in maintaining a precise and stable low-spin rate, such as 24°/s with a ±0.24°/s
tolerance—especially over time and in the presence of disturbances like magnetic field
fluctuations or aerodynamic drag.
Practical Challenges in Spin Control
The biggest technical hurdle is not just achieving the required spin speeds but doing so reliably
and keeping them stable within tight tolerances. The spin-up process—increasing the
satellite’s rotation rate—can take significant time depending on the actuator’s strength. Longer
spin-up durations also increase the risk of power strain, control instability, and delays in critical
mission operations like instrument activation or tether deployment.​
Attitude Modes and Spin Control Strategy for
FORESAIL-1 CubeSat
The FORESAIL-1 CubeSat operates in multiple attitude modes, each tailored to a specific
mission phase or payload requirement. All these modes rely on accurate spin rate control and
precise orientation of the satellite’s spin axis.
1. Detumbling Mode
Purpose:​
This is the first mode activated immediately after deployment and is also used in emergency
recovery scenarios. The satellite may be tumbling unpredictably due to mechanical ejection
forces, rocket vibrations, or mounting tolerances. The primary goal here is to reduce the
unwanted rotation and bring the satellite to a controlled and known state.
Design Goal:​
The detumbling system must reduce the spacecraft’s spin rate to below 1°/s, providing a
conservative and safe baseline for initiating precise operations like payload experiments.
Importance:​
A tumbling satellite cannot perform payload operations or deploy its systems reliably.
Detumbling ensures stability so that other attitude modes can function correctly. This is
especially important for experiments like the plasma brake deployment, which requires
controlled rotational states to work effectively.
Design Challenge:​
The detumbling system must be robust enough to handle worst-case scenarios, such as initial
spin rates that match or exceed those used for the plasma brake experiment (~130°/s). It should
bring these down reliably under the 1°/s threshold.
2. Particle Telescope Operation Mode
Purpose:​
Used during the scientific observation phase involving the particle telescope payload.
Requirements:
●​ Spin-Up Maneuver: The satellite must spin up to a stable rate to maintain attitude
stability.​
●​ Precise Pointing: The spin axis must be carefully aligned, often toward the Sun or
another celestial reference, for accurate data collection.​
This mode benefits from the stable spin to ensure consistent orientation during long observation
periods.
3. Plasma Brake Operation Mode
Purpose:​
This mode is activated to test the plasma brake, a deorbiting technology.
Requirements:
●​ Similar to the telescope mode, it requires a spin-up maneuver.​
●​ Precise pointing of the spin axis is critical to ensure proper deployment and operation
of the tether-based brake.​
Given that this experiment demands the highest spin rates among all modes, the spacecraft
must first be detumbled before transitioning into this mode.
Pointing Maneuvers Between Modes
Switching between operational modes (e.g., from detumbling to telescope or plasma brake
mode) involves re-aligning the spacecraft’s spin axis. This ensures that each payload
functions optimally, with its required field of view or orientation maintained.
These transitions must be carefully planned and executed by the control system to avoid
destabilizing the spacecraft.
Attitude Control System (ACS) Design Overview
To enable the above modes and transitions, the spacecraft is equipped with a dedicated Attitude
Control System. This system is responsible for generating torque, delivering it accurately, and
maintaining control through well-defined strategies.
Key Subsystems:
1.​ Actuators: Devices that produce the required torque. FORESAIL-1 uses a combination
of magnetorquers (air-core coils) and reaction wheels.​
2.​ Driving Hardware: Power electronics such as H-bridge circuits are used to control the
flow and direction of current in the coils, enabling torque generation.​
3.​ Control Strategy: This includes the algorithms and logic that determine what torque
outputs are needed to achieve or maintain a desired orientation or spin rate.​
Magnetorquer Configuration and Modeling
FORESAIL-1 uses air-core magnetorquers, which are coils without any magnetic core. This
design avoids magnetic interference and is lightweight, which is ideal for CubeSats.
The coils are modeled as equivalent electrical circuits to simulate and optimize performance.
The team also explored whether connecting coils in series or parallel would yield better
performance and control authority.
Driving hardware involves:
●​ H-Bridge circuits to manage bidirectional current flow.​
●​ Filters and driving frequencies optimized through simulations to maximize torque
efficiency.​
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CubeSat Mission to Study Solar Particles
Mission Objective
The goal of this CubeSat mission is to develop and validate a miniaturized, directional
particle detection system capable of sensing suprathermal and solar energetic particles
(SEPs). These high-energy particles pose significant radiation threats to spacecraft and
astronauts. Understanding and predicting SEP-associated radiation environments requires the
detection of peak flux intensities, maximum particle energies, and total fluence (total particles
over time).
Ground-Based Directionality Test
Purpose
To simulate how a CubeSat’s angle-sensitive detector (e.g., PIN diode or SSD) responds to
particle flux from various directions—replicating in-orbit performance using a ground test.
Procedure Overview
1.​ Detector Setup:​
Use a solid-state detector (SSD) or PIN photodiode connected to a basic data
acquisition system (e.g., Arduino, Raspberry Pi). Collimation (e.g., pinhole, shielding) is
added to restrict the field of view (FOV), making the detector sensitive to particle
direction.​
2.​ Radiation Source:​
Use a weak beta source (like Strontium-90, with radiation approval) or a low-energy
X-ray tube. As an alternative, a laser or LED can simulate angular light input when
testing detection angle response.​
3.​ Rotation Platform:​
The detector is mounted on a rotating platform (servo motor or 3D-printed gimbal). It is
rotated from -90° to +90° in defined steps (e.g., every 10°).​
4.​ Data Logging:​
Measure and record detector counts or analog signals at each angle. A peak count
should occur when aligned with the source, with decreasing counts off-axis—confirming
angle sensitivity.​
This test validates that the CubeSat detector can measure directional changes in particle
flux—essential for real-time SEP direction estimation in orbit.
Scientific and Technical Goals
The mission aims to:
●​ Track onset, peak, and decay phases of SEP events.​
●​ Determine energy spectra of incident particles.​
●​ Understand acceleration mechanisms such as shock waves from flares or CMEs
(coronal mass ejections).​
●​ Predict and study interplanetary shocks and solar wind dynamics.​
Payloads will be used to observe changes in particle energy, angle of arrival, and density over
time. For low Earth orbit (LEO), higher altitudes enable plasma monitoring using instruments like
particle spectrometers and Langmuir probes. These tools help assess atmospheric drag, a key
parameter influenced by ionospheric oxygen density.
Particle Sources and Hazards in LEO
In LEO, particle flux can:
●​ Damage satellite electronics.​
●​ Cause surface charging.​
●​ Disrupt power and communication systems.​
●​ Affect atmospheric density and drag.​
These particles originate from solar storms and plasma sheets in the magnetosphere,
depositing energy in the 90–300 km range—intersecting CubeSat altitudes and posing
operational risks.
Observation Techniques
Optical Techniques
●​ Aurora/Airglow Imaging: Used for atmospheric observation, with info about shape,
color, and spatial distribution.​
●​ Spectral Imagery: Demands larger instruments.​
●​ Remote Sensing: Images auroras remotely for visual space weather indicators.​
●​ Spectrometer + Particle Detector Combo: Suited for particles under 30 eV.​
●​ Best Imaging Sensor: Onyx Teledyne E2V imager.​
Calibration and Orbit Selection
●​ Photometric Calibration: ROLO and POLO methods, leveraging the Moon’s
photometric stability.​
●​ Instrument Channels: RGB or black-and-white setups possible.​
●​ Orbit: Sun-synchronous orbit (SSO) at ~600 km altitude with 70°–80° inclination offers
optimal lighting and coverage.​
Detector and Instrument Design
The Suprathermal Ion Sensor is proposed for miniaturized, energy-resolved particle detection
on CubeSats.​
​
What Are Suprathermal (ST) Ions?
●​ ST ions have energies between solar wind plasma and typical energetic particles.​
●​ Range: ~10 eV to hundreds of keV.​
●​ Important as seed particles for SEPs.​
●​ Hard to measure: too high for solar wind instruments, too low for traditional detectors.​
Detection Solution
●​ Use Electrostatic Analyzers (ESA) with broad energy coverage.​
●​ Acts like a spectrograph for ions.​
●​ Provides high time, mass, and charge resolution.​
●​ Retains good sensitivity due to longer sampling per E/q value.​
Reference Mission: CuSP (CubeSat to Study Solar
Particles)​
​
Mission Parameters
●​ Type: 6U CubeSat (30×20×10 cm)​
●​ Mass: 14 kg​
●​ Stabilization: Three-axis, Sun-pointed​
●​ Orbit: Heliocentric (Earth-escape via NASA’s EM-1)​
Onboard Instruments
1.​ SIS (Suprathermal Ion Sensor)​
○​ Electrostatic Analyzer + Microchannel Plate​
○​ Measures angular/energy distributions of ST ions​
○​ Focused on Parker spiral particle flow​
2.​ MErIT (Miniaturized Electron and Ion Telescope)​
○​ Solid-state energy tracking (ΔE & E)​
○​ Detects 2–150 MeV/nucleon particles​
○​ Measures elemental compositions (H to Fe)​
3.​ VHM (Vector Helium Magnetometer)​
○​ Tracks magnetic fields​
○​ Essential for correlating ion data with solar wind conditions​
Instrument Case Study: STATIC
Introduction
STATIC stands for SupraThermal and Thermal Ion Composition. It studies cold and accelerated
ions in the upper atmosphere.
What It Measures
●​ Ion species: H⁺, He⁺, He²⁺, O⁺, O₂⁺, CO₂⁺​
●​ Energy range: 0.1 eV to 30 keV​
●​ Coverage: 360° × 90° field of view​
●​ Applications: Tracks atmospheric escape on Mars, ion acceleration processes​
Working Mechanism
1.​ Electrostatic Analyzer​
○​ Sorts ions by energy​
2.​ Time-of-Flight (TOF) Analyzer​
○​ Measures ion mass using −15 kV acceleration​
○​ Detects timing via secondary electrons and microchannel plates​
○​ Resolution: 22.5° angular bins​
Features
●​ Data includes ion mass, energy, arrival angle, and detection time​
●​ Uses electrostatic and mechanical attenuators to adapt to changing plasma densities​
●​ Can handle density changes up to 1000× (10³ dynamic range)​
​
CubeSat Payload and Mission
Planning: Concepts for Space Weather
and Thermospheric Monitoring
1. Spacecraft Orientation and Coordinate System
The CubeSat’s orientation is defined by a solar-centric coordinate system where the +x axis
points toward the Sun. The +y axis lies in the ecliptic plane, perpendicular to +x, while the +z
axis points toward the ecliptic north pole. This configuration allows sensors mounted on the
spacecraft to observe not just the nominal Parker spiral direction (where solar particles
stream out), but also the anti-Parker spiral and ecliptic north, enabling a wide field of view for
monitoring energetic particles.
2. Mission Operations and Management
The CubeSat is designed for a minimum operational lifetime of three months, which is
considered sufficient to achieve the core science objectives. It leverages Commercial
Off-The-Shelf (COTS) components to reduce development costs and time. The spacecraft
benefits from heritage designs while integrating tailored parts to maintain performance and
reliability. Development is overseen by experienced institutions including SwRI, NASA GSFC,
and JPL, following a typical CubeSat Work Breakdown Structure (WBS) optimized for
low-cost and moderate-risk missions.
3. Scientific and Technological Significance
The mission is dual-purpose. Scientifically, it aims to investigate how suprathermal (ST) ions
evolve into solar energetic particle (SEP) populations, which is crucial for protecting
astronauts and space infrastructure. Technologically, the CubeSat serves as a flight-proven
platform for future ST ion sensor development, offering a scalable solution for heliophysics
missions.
4. Feasibility of Detecting Suprathermal Ions in LEO
Detecting ST ions in Low Earth Orbit (LEO) with a 3U CubeSat is technically possible but
comes with several constraints. ST ion populations do exist in the near-Earth environment,
especially during solar events and at high latitudes, such as auroral zones. Compact
instruments like miniaturized electrostatic analyzers (ESAs) and solid-state detectors can
fit within the payload volume of a 3U CubeSat.
Supporting Missions:
●​ SAMPEX (1992–2004) detected energetic particles in a polar orbit.​
●​ CubeSat-scale detectors have flown on missions like CeREs and AeroCube.​
However, there are major challenges:
●​ Earth’s magnetic field in LEO deflects many incoming solar ions.​
●​ High background noise from trapped radiation, secondary particles, and electron fluxes
can contaminate readings.​
●​ CubeSats have limited surface area and geometry factor, leading to low sensitivity.​
●​ Constraints on power, telemetry, and field-of-view further limit performance.​
In summary, detection is feasible under specific conditions, like during solar particle events
or in high-inclination orbits (>70°), but demands precise instrument design and is best suited
for targeted event-based observations rather than continuous global monitoring.
5. Instrument Overview
Key Payload Types:
●​ Mass Spectrometer: Measures the mass-to-charge ratio of particles via ionization and
fragmentation, yielding a mass spectrum for chemical composition analysis.​
●​ Electrostatic Analyzer (ESA): Offers a 360° × 90° field of view, useful for mapping
charged particle distributions.​
●​ Spectrometer: Separates and quantifies a particular physical property (e.g., light
wavelength) across a spectrum.​
These instruments together enable real-time monitoring of the upper atmosphere, especially
the thermosphere, which is highly reactive to space weather events like solar flares.
Monitoring the Thermosphere
The thermosphere, a region of Earth’s upper atmosphere (150–400 km), is sensitive to solar
radiation, geomagnetic activity, and charged particle precipitation. Monitoring this layer
helps predict satellite drag, communication issues, and GNSS signal interference.
Traditional Monitoring Parameters:
1.​ Atomic Oxygen Green Line (557.7 nm) – Indicates oxygen population and energy
input.​
2.​ Total Electron Content (TEC) – Measured via radio signals; reflects ionospheric
changes but gives limited thermospheric data.​
However, these two are not enough for a complete picture.
New Approach: Red Line Polarization as a Third Parameter
Researchers propose using polarization of the atomic oxygen red line (630.0 nm) as a third
monitoring metric. Polarization refers to the orientation of light waves, and in this case, may
result from anisotropic collisions in the thermosphere.
This effect, driven by asymmetrical particle collisions, provides unique insight into the energy
dynamics and particle interactions in the upper atmosphere. Historical attempts to detect this
polarization go back to the International Geophysical Year (1957–1958) but yielded
inconsistent results. Modern instruments now offer a chance to revisit and validate these
findings.
Importance of Real-Time Thermospheric Monitoring
Effects of Solar Activity:
●​ Increased UV and particle input heats the thermosphere, causing it to expand (dilate).​
●​ This leads to higher density at satellite altitudes, which:​
○​ Increases drag,​
○​ Accelerates orbital decay,​
○​ Requires more fuel for corrections.​
Hence, real-time thermospheric data is essential for LEO satellite resilience and long-term
mission planning.
Current Gaps in Upper Atmosphere Modeling
Limitations:
●​ Solar X-rays, EUV, and particle precipitation are poorly monitored in real time.​
●​ SuperDARN radar network maps ionospheric electric fields but only when irregularities
exist.​
●​ Physics-based atmospheric models often fail during magnetic storms or at high
latitudes, leading to orbit prediction errors exceeding 20 km in just 24 hours.​
Thermospheric Emissions and Diagnostic Potential
Thermospheric oxygen emissions result from deactivation of excited oxygen in the F-region.
The two key emissions are:
●​ Green line (557.7 nm) from the 1S → 1D transition.​
Red line triplet (630.0, 636.4, 639.2 nm) from 1D → 3P transitions.​
Excitation Mechanisms:
●​ Energetic electron impact,​
●​ Photodissociation of O₂,​
●​ Nitrogen reactions,​
●​ Dissociative recombination,​
●​ Cascading from upper energy states.​
These processes are mostly isotropic by day but can be anisotropic at night, especially from
magnetically guided particles — making nighttime red line polarization a strong candidate for
diagnostics.​
​
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Motivation: Why Study the Upper Atmosphere?
The upper atmosphere, including the thermosphere, ionosphere, and exosphere, plays a crucial
role in satellite drag, space weather, and communication systems. Understanding how this
region responds to solar activity (like geomagnetic storms) is vital for:
●​ Accurate orbit predictions​
●​ Reliable GPS and communication signals​
●​ Better atmospheric and space weather models​
​
However, current measurements are limited. We lack sufficient real-time data, especially
for the thermosphere, which restricts our understanding of upper atmospheric dynamics​
​
Existing Measurement Techniques
Currently, scientists rely on:
●​ Total Electron Content (TEC) – Measures ionospheric charge density.​
●​ 557.7 nm green light from atomic oxygen – A known auroral emission that increases
during geomagnetic storms.​
While useful, these methods alone don’t provide a full picture of the upper atmosphere,
especially the thermosphere.
Proposed New Technique: Polarized Red Light
A novel method under exploration involves detecting polarization in red light (630 nm) emitted
by atomic oxygen at night. This light is influenced by:
●​ Energetic particles hitting the atmosphere​
●​ Their direction and energy​
●​ Auroral activity​
Historically, scientists debated whether this red light is truly polarized—results were
inconclusive. With modern instruments, researchers aim to re-investigate this, potentially
unlocking a powerful new way to observe thermospheric dynamics.
The Challenge of Atomic Oxygen (AO) in Space
When spacecraft orbit in Low Earth Orbit (LEO, ~180–650 km), they face:
Space Hazards
Effects
Radiation
Degrades electronics and materials
Extreme Temperatures
Causes thermal stress
Atomic Oxygen (AO)
Causes corrosion and material
erosion
Space Debris
Collision risk
Why is AO Dangerous?​
AO is formed when solar UV radiation breaks apart O₂ molecules. This
reactive form of oxygen damages spacecraft by:
●​ Corroding surfaces​
●​ Thinning protective layers​
●​ Causing mass loss​
Early missions (like the Shuttle) observed AO damage on materials such as Kapton film.
Real-Time AO Monitoring: 3Cat-1 CubeSat
Traditional AO studies have limitations:
●​ Focus on external materials​
●​ Post-mission data analysis​
●​ Ground testing that doesn’t match real space conditions​
3Cat-1, a CubeSat from the Technical University of Catalonia, addresses these gaps by
enabling real-time AO monitoring in space.
Mission Details:
●​ Orbit: Polar orbit (550–650 km)​
●​ Launch: Planned in the second half of 2016​
●​ Material Tested: TIPS-Pentacene – used in Organic Field-Effect Transistors (OFETs)​
How AO Damage is Measured
3Cat-1 carries a MEMS-based sensor with a thin pentacene layer. As AO reacts with the
material:
●​ Its mass reduces​
●​ This changes the vibration frequency of the MEMS cantilever​
A Pulsed Digital Oscillator (PDO) keeps the MEMS vibrating and measures frequency
changes digitally. This setup:
●​ Is energy-efficient​
●​ Detects mass loss as small as 0.5 nanograms​
●​ Sends live data back to Earth​
​
CubeSat Payload for Atmospheric Composition:
CubeSatTOF
To improve chemical analysis of the upper atmosphere, researchers have developed
CubeSatTOF—a compact time-of-flight (TOF) mass spectrometer for CubeSats.
Why Study Chemical Composition?
Understanding the exosphere’s gas content helps with:
●​ Monitoring space weather​
●​ Tracking atmospheric escape​
●​ Studying effects of solar flares and burning space debris​
​
How CubeSatTOF Works
CubeSatTOF uses a direct open source inlet system, avoiding problems in earlier designs:
System Type
Problem
Closed source Particles hit walls → break apart
Open source
Needed high voltage for ion bending
Direct open source allows neutral particles to go directly into the analyzer, avoiding collisions
and preserving their integrity.
Working Principle:
1.​ Neutral particles are ionized with an electron beam.​
2.​ Ions are sent through a TOF path.​
3.​ Their travel time reveals their mass-to-charge ratio (m/z).​
4.​ A reflector (ion mirror) corrects energy spread.​
5.​ Ions hit a detector (MCP), producing a mass spectrum.​
​
Why CubeSatTOF is Powerful
●​ Detects particles from m/z 1 to 200​
●​ High mass resolution (can differentiate isotopes)​
●​ Real-time operation (measurements every 100 ms)​
●​ Operates at satellite speeds (~7.6 km/s)​
●​ No mass bias​
This makes it ideal for observing minor species, radicals, and real atmospheric conditions that
influence satellite drag and space weather.
CHESS Mission: A Step Further
The CHESS (Constellation of High-performance Exosphere Science Satellites) mission aims to
launch a network of CubeSats with advanced spectrometers like CubeSatTOF. Objectives
include:
●​ Real-time monitoring of the upper atmosphere​
●​ Mapping ion-neutral transitions​
●​ Understanding particle behavior at the exobase​
●​ Space weather forecasting using real data​
​
​
​
Overview of the Lucky 7 CubeSat Experiment
Lucky 7 is a compact and cost-effective 1U CubeSat (10 × 10 × 10 cm), designed to
demonstrate that even a small satellite can effectively study cosmic radiation in Low Earth Orbit
(LEO). It was launched on June 27, 2019, into a polar orbit approximately 520 km above Earth.
The main goal of the mission was to test whether such a simple satellite could accurately
measure cosmic radiation, making space science more accessible through miniaturization and
affordability.
How Radiation is Measured Onboard
Lucky 7 uses a scintillator detector to measure radiation. When high-energy particles like
cosmic rays hit the scintillator material, it emits a flash of light. This light is then detected by a
PIN diode, which converts it into an electrical signal for data processing.
To provide accurate spatial and temporal context, the satellite includes a GPS receiver. This
receiver logs the exact time and location each time a radiation event is recorded. As a result,
every measurement includes:
●​ The strength of the radiation (count rate),​
●​ The time and position of the event.​
​
Key Achievements of Lucky 7
The Lucky 7 CubeSat successfully collected radiation data that matched well with existing
models, confirming that the onboard detectors worked as expected. This validation proved that
small, low-cost CubeSats can generate scientifically reliable data. The mission showed that a
basic detector, combined with precise time and location tracking, can be enough to perform
meaningful space radiation research.
Lucky 7 Mission Objectives
The mission was designed to:
●​ Test a simple and affordable radiation detection system,​
●​ Accurately measure radiation and log time and position,​
●​ Improve and validate numerical radiation models by comparing real-time
measurements with existing predictions.​
​
Sources of Space Radiation
In space, radiation comes from three major sources:
1. Galactic Cosmic Rays (GCR)
These are high-energy particles from outside our solar system, often produced by distant
supernovae. GCRs are mostly protons (about 85% of baryons), along with some heavier ions
and a small fraction of electrons (~2%). They are extremely energetic and can penetrate
spacecraft shielding and biological tissue.
2. Solar Cosmic Rays (SCR)
Produced by the Sun, especially during solar flares and coronal mass ejections (CMEs).
Although typically less energetic than GCRs, SCRs are dangerous because they can arrive in
sudden bursts, particularly during solar storms. These events can last from a few hours to
several days, with particle energies sometimes exceeding 1 GeV.
3. Van Allen Radiation Belts
These are two donut-shaped zones of trapped charged particles around Earth, formed by the
interaction of GCRs and SCRs with Earth's magnetic field.
●​ The inner belt primarily consists of protons formed by neutron decay.​
●​ The outer belt contains electrons and solar-originated particles.​
An important anomaly is the South Atlantic Anomaly (SAA), where the inner radiation
belt dips closest to Earth (~200 km altitude) due to a deformation in Earth’s magnetic
field.​
Radiation Detectors Used in CubeSats
CubeSats have strict limitations on size, weight, and power, so they rely on compact,
energy-efficient radiation detectors, particularly solid-state dosimeters. These include:
●​ Scintillators: Emit light when struck by radiation. A second sensor detects this light to
infer radiation presence.​
●​ PIN Diodes: Semiconductor devices that directly detect charged particles. They are
low-power and easy to integrate.​
●​ RadFETs: Radiation-sensitive Field Effect Transistors. These detect cumulative radiation
dose by changing their electrical properties over time.​
Managing Temperature and Budget Challenges
Temperature fluctuations in space affect detector sensitivity. However, CubeSats usually cannot
afford active temperature control. Instead, they monitor the temperature and later correct the
radiation readings based on how the temperature affects the detector.
Most CubeSats, including Lucky 7, use Commercial Off-The-Shelf (COTS) components to
keep costs low. These parts are mass-produced and affordable but not necessarily
space-grade.
Technical Breakdown of Lucky 7 CubeSat
Component
Function
Onboard
Computer
Two CPUs for redundancy; manages data, communication, and
control
Radio Modems
Enables communication with ground stations
Experimental
Board
Hosts piNAV GPS, piDOSE detector, a spectrometer, and a low-res
camera
Power System
Uses 10 gallium arsenide solar cells and a LiFe battery pack
Antennas &
Sensors
One side holds the GPS antenna, camera, and UHF
communication antennas
Structure
Aluminum frame with PCB external panels; offers basic radiation
shielding
Radiation Detection Challenges in Lucky 7
The primary radiation detector was piDOSE, a silicon-based dosimeter that successfully
recorded valuable radiation data. It used a PIN diode and a CsI:TI scintillator to detect
high-energy particles.
The satellite also had a spectrometer with two modes:
1.​ High-time resolution mode: Counted particles every 500 ms — this worked well.​
2.​ Energy spectrum mode: Intended to detect the energy range of particles. However, due
to an incorrect setup and a software bug, this mode failed, as the detector became
saturated and the configuration couldn’t be changed post-launch.
Table: Component Performance
Feature
Status
piDOSE radiation detector
Worked and provided usable data
Spectrometer (energymod
Failed due to configuration issues
GPS receiver
Provided accurate time/location
Power system (solar)
Enabled continuous operation
Shielded structure
Protected electronics from radiation
How piDOSE-DCD Detector Works
The piDOSE-DCD combines two components:
●​ A PIN diode, which detects photons (light),​
●​ A CsI:TI scintillator crystal (4×8×8 mm), which emits photons when hit by radiation.​
When radiation strikes the scintillator, it glows. The PIN diode then picks up this glow, signaling
a radiation event. This method extends the range of particle energies that the detector can
measure.​
​
​
Electromagnetic Interference (EMI) Mitigation
EMI can interfere with radiation measurements, especially in compact CubeSats where
antennas, transmitters, and detectors are tightly packed. Lucky 7 used three key strategies:
1.​ Placing the piDOSE in a 2 mm aluminum shielded box,​
2.​ Installing EMI filters in the electronics,​
3.​ Turning off the transmitter while taking radiation measurements to prevent
interference.​
​
Internal Layout and Mounting of piDOSE
The piDOSE detector was mounted just beneath the top panel of the satellite. The satellite’s
aluminum frame (2 mm thick) provided structural integrity and radiation shielding. The top panel
was made from 1.6 mm FR4 PCB, housing both the GPS and UHF antennas.
Internally, all electronics, including the OBC and detectors, were assembled on FR4 PCBs, a
common circuit board material. Aluminum partitions separated various electronics to offer
structural support and shielding.
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Calibration of piDOSE: Detector Testing for Accuracy
To ensure the piDOSE radiation detector's accuracy, initial calibration was performed using
standard gamma-ray sources—Cobalt-60 and Cesium-137. These isotopes emit radiation with
well-known characteristics, making them ideal for verifying the detector's response. However,
since space radiation includes high-energy electrons and protons, more advanced calibration is
planned. Future tests will use electron and proton beams to better simulate the actual space
environment and validate the detector’s real-world performance.
piDOSE CubeSat Radiation Measurements (2019–2020)
Between summer 2019 and spring 2020, the piDOSE-equipped CubeSat gathered 10,000
minutes of radiation data. The satellite had no attitude control, so it rotated freely in orbit. This
rotation affected sensor orientation, as confirmed by variations in solar panel outputs and GPS
signal strength (see Fig. 6 & Fig. 7 in the original source).
A GPS simulator test also confirmed that the satellite could achieve meter-level positioning
accuracy in Low Earth Orbit (LEO), which is notable for such a small platform.
Visualizing the Data
An interactive map was developed (Fig. 8) to display:
●​ Radiation counts per minute (CPM),​
●​ Satellite position and timestamps,​
●​ Sensor temperature.​
​
These values were also mapped onto Google Earth (Figs. 9–11). A major radiation peak
occurred over the South Atlantic Anomaly (SAA), where readings reached up to 115,000
CPM—highlighting this region as a known hotspot for trapped radiation.
Comparing Measurements to Radiation Models
1. Galactic Cosmic Radiation (GCR) Modeling
GCR flux was simulated using the Matthiä et al. (2013) model, accounting for:
●​ Geomagnetic shielding via vertical cut-off rigidity values (0 GV at poles to ~15 GV at
the equator),​
●​ Earth shielding, using a 0.69 reduction factor to account for the Earth’s occlusion,​
●​ Energy range from hydrogen to nickel nuclei, up to 1 TeV/n.​
Simulation tools included:
●​ PLANETOCOSMICS (rigidity map generation),​
●​ IGRF-13 and Tsyganenko magnetic models for magnetic field strength calculations.​
​
Issues identified included:
●​ Use of vertical rigidity only, instead of full 3D directional models.​
●​ Exclusion of secondary particles produced within the CubeSat.​
●​ No modeling of albedo particles reflected from Earth’s atmosphere.​
2. Radiation Belts (SAA and Polar Regions)
Measured radiation in the SAA confirmed very high CPM due to trapped protons, consistent with
models. Polar regions also showed elevated readings due to weaker geomagnetic shielding and
the presence of trapped electrons. These results align well with standard trapped radiation
models, such as AE-8 or AP-8.
piDOSE CubeSat System Overview
Payload and Onboard Systems
●​ Main Payload: piDOSE detector using a PIN diode and scintillator combination.​
●​ Electromagnetic Interference (EMI): Became an issue, so the detector was enclosed in
an aluminium shield and EMI filters were added.​
●​ Calibration Sources Used: Cobalt-60 and Cesium-137.​
●​ Future Testing: Electron and proton beam exposure.​
Supporting Systems
●​ OBC (Onboard Computer): Two computers with 2 Mbytes of Ferroelectric RAM.​
●​ Radio Communication: UHF radio modem integrated into the OBC.​
●​ Experimental Boards:​
○​ piNAV GPS receiver: Tracks satellite position.​
○​ piDOSE detector: Measures radiation.​
●​ Spectrometer:​
○​ High-time-resolution mode for particle count.​
○​ Energy spectrum mode to identify particle energy (some issues occurred).​
●​ Low-Resolution Camera: Takes photos during flight.​
Power and Structure
●​ Power: 10 Gallium Arsenide solar cells, LiFe battery, and regulator.​
●​ Panel Mounting: Solar panels on 5 of 6 CubeSat faces.​
●​ Structure: Aluminium frame with PCB panels and a 2 mm aluminium casing.​
CubeSat Atmospheric Science Mission Design
EUV Radiation and Atmospheric Impact
Extreme Ultraviolet (EUV) radiation originates from the Sun's corona and chromosphere,
covering the 1–120 nm range. It includes spectral lines from elements such as H, He, O, Na,
Mg, Si, and Fe.
Upon reaching Earth:
●​ EUV is absorbed above 80 km altitude, primarily heating the thermosphere.​
●​ It also ionizes atmospheric gases, creating free electrons, forming the ionosphere.​
Significance of EUV Monitoring
EUV radiation varies:
●​ In minutes (solar flares),​
●​ Over days (solar rotation),​
●​ Over years (solar cycles).​
These fluctuations cause the upper atmosphere to expand or contract, which can:
●​ Disrupt satellite communications and GPS,​
●​ Change drag in LEO, accelerating orbital decay.​
Forces in LEO
The primary forces acting on a CubeSat in LEO are:
●​ Gravitational force,​
●​ Atmospheric drag (most significant for decay),​
●​ Solar radiation pressure.​
OwlSat: CubeSat for Solar EUV Monitoring
Mission Objective
OwlSat is designed to measure EUV radiation from the Sun and compare it with the satellite’s
position and velocity in orbit. It follows a 1U structure based on EnduroSat’s aluminium 6061
CubeSat frame.
Scientific Instruments
●​ EUV Sensors: 4 sensors from Gigahertz-Optik (1–200 nm range).​
○​ Always face the Sun,​
○​ Mounted externally,​
○​ Controlled using an attitude control system.​
●​ Accelerometer: Measures satellite velocity.​
●​ GPS Receiver: Determines position in orbit.​
Data Flow and Transmission
●​ EUV sensor data → science board → main computer.​
●​ GPS and accelerometer data → secondary board → main computer.​
●​ Data is stored onboard and sent back to Earth.​
Communication Limits
●​ Transmission rate: 9600 baud.​
●​ Max data: ~1 MB/day.​
●​ Operational region: ±10° latitude (approx. 11 minutes/orbit) to stay within bandwidth
limits.​
CIRCE Mission: Nighttime Ionosphere Study
Overview
CIRCE (Coordinated Ionospheric Reconstruction CubeSat Experiment) is a joint US/UK
mission using two 6U CubeSats flying in a lead-follow formation in LEO. Its goal is to study
the Equatorial Ionization Anomaly (EIA) — an area with enhanced nighttime electron density
near Earth’s magnetic equator.
Instruments Onboard
●​ From the US (Naval Research Lab):​
○​ Tri-TIP Photometers: Detect 135.6 nm UV emissions from atomic oxygen,
mapping electron distribution.​
●​ From the UK (Dstl):​
○​ Ion/Neutral Mass Spectrometer,​
○​ Triple-frequency GPS receiver (for total electron content),​
○​ Radiation environment monitor.​
Scientific Method
●​ Combines:​
○​ UV photometry from four angles,​
○​ In-situ plasma measurements,​
○​ GPS-based ionospheric data.​
These are processed using tomographic reconstruction — similar to a CT scan — to create
2D maps of ionospheric electron density.​
​
CIRCE Mission Overview
The Co-ordinated Ionospheric Reconstruction CubeSat Experiment (CIRCE) is a
collaborative science mission that uses two CubeSats in a lead-trail formation to study the
ionosphere — a charged region of Earth’s upper atmosphere that plays a crucial role in satellite
communication, GPS reliability, and space weather forecasting.
Mission Objectives
Objective
Purpose
Study nighttime ionosphere
Understand ionospheric behavior after
sunset, especially near the equator.
Monitor Equatorial Ionization Anomaly (EIA)
Investigate its effects on radio signals and
satellite communications.
Create 2D maps of electron density
Reconstruct detailed nighttime ionospheric
maps using UV and GPS data.
Improve prediction models
Enhance space weather forecasts and
improve satellite operation reliability.
Importance of Studying the Ionosphere
●​ The ionosphere is affected by solar activity like solar flares, which disrupt satellite
communication and GPS systems.​
●​ The Equatorial Ionization Anomaly (EIA) causes strong electron density changes that
impact signal propagation over the equator.​
●​ CIRCE focuses on mapping this region using UV light and GPS signals, which provides
data previously difficult to gather from space.​
Tri-TIP Sensors: Remote Sensing of UV Light
The Tri-TIP (Triple Tiny Ionospheric Photometer) is a compact UV sensor used to measure
faint emissions from oxygen atoms in the ionosphere. These emissions — particularly at 135.6
nm — help estimate electron density in the F-region (~150–500 km altitude).
How Tri-TIP Works
Component
Function
Hinged Mirror Cover
Protects optics pre-launch, flips open in space to become part of
the mirror system.
Mirrors
Direct UV light into the system.
SrF₂ Filter
Filters out unwanted light wavelengths.
Beam Splitter
Splits light into multiple paths for measurement.
Photomultiplier Tubes
(PMTs)
Extremely sensitive detectors powered at 1100 V, used to detect
faint UV signals.
Sunlight Protection
Prevents sensor damage with a sun-sensitive shutter.
●​ ​
Tri-TIP has a field of view of 3° × 7.25° (nadir and 45° angles) and 0.2° × 7.25° (limb
view) for high-resolution edge measurements.​
●​ These sensors operate best at night, using ionospheric "nightglow" to derive electron
density.​
Tri-TIP Sensor Orientation & Satellite Configuration
Each CubeSat carries two Tri-TIP sensors. Their placement enables multiple-angle viewing for
3D electron density reconstruction:
Sensor Viewing Angles
Satellite
Sensor View
Orientation
Lead CubeSat 45° behind (wake side) Diagonal toward rear
Lead CubeSat 17° behind (limb)
Along Earth’s horizon
Trail CubeSat
45° forward (ram side)
Matches lead's view
Trail CubeSat
Straight down (nadir)
Directly below
This multi-angle setup allows scientists to triangulate ionospheric conditions like a CT scan
in space.
Day/Night Operations and Differential Drag
●​ Nighttime: Tri-TIP sensors collect ionospheric data when UV emissions are most
visible.​
●​ Daytime: Trail satellite performs a 180° yaw maneuver every other orbit to operate
other instruments (like INMS and TOPCAT) and gather particle data.​
●​ Differential drag is used to maintain 250–500 km spacing between the satellites by
adjusting their orientation and aerodynamic drag.​
TIP: Why Far Ultraviolet (FUV) Light?
●​ FUV light (like 135.6 nm) is blocked by Earth’s lower atmosphere and must be studied
from space.​
●​ At night, oxygen ions in the ionosphere recombine with electrons, releasing FUV light.​
●​ These emissions help estimate electron density, which is essential for modeling space
weather effects.​
UK’s Contribution: The IRIS Suite
The UK provides the IRIS (In-situ and Remote Ionospheric Sensing) suite for CIRCE. It
includes three payloads aimed at collecting particle, radiation, and GPS signal data:
IRIS Payloads
Payload
Institution
Function
INMS (Ion and Neutral Mass
Spectrometer)
UCL Mullard Space
Science Lab
Measures ion and neutral particle
composition.
RadMon (Radiation Monitor)
SSTL & University of
Surrey
Measures radiation levels to assess
satellite exposure.
TOPCAT (GPS Receiver)
University of Bath
Uses GPS signal delays to map total
electron content (TEC).
IRIS data will:
●​ Improve understanding of atmospheric drag and thermospheric chemistry.​
●​ Identify radiation hotspots and inform satellite orbit and shielding design.​
●​ Validate the MIDAS tomography algorithm for mapping the topside ionosphere using
GPS TEC data.​
Each CIRCE satellite includes an IRIS unit: the lead satellite houses IRIS at the front (~2U ×
1U), and the trail satellite carries it at the rear.
Combined Impact of Tri-TIP & IRIS
The integration of Tri-TIP (UV photometry) with IRIS (in-situ sensing and GPS data) enables
comprehensive ionospheric studies:
●​ Tri-TIP provides remote UV-based electron density mapping.​
●​ IRIS adds real-time contextual data (particles, radiation, TEC).​
●​ Together, they deliver multi-angle, high-resolution, and multi-modal observations of
the ionosphere, helping refine models and improve space-weather resilience.​
IRIS Suite on CIRCE​
Overview: What is IRIS?
The IRIS suite (In-situ and Remote Ionospheric Sensing) is a set of compact scientific
instruments developed in the UK to study the ionosphere, a charged region of Earth’s upper
atmosphere. This region affects satellite drag and radio communications. IRIS is part of the
CIRCE (Coordinated Ionospheric Reconstruction CubeSat Experiment) mission, which
involves two CubeSats flying in formation—one leading and the other trailing—in low Earth orbit
(LEO). Each carries an identical IRIS unit.
IRIS Instrument Components
1. INMS – Ion and Neutral Mass Spectrometer
Developed by University College London (MSSL), the INMS is designed to measure both
neutral gases and charged particles in the upper atmosphere. It helps in identifying the
atmospheric composition and tracking changes due to space weather or orbital position.
Instrument Working:
The INMS consists of a collimator, ioniser, and a spectrometer. The ioniser includes an
electron source, energy selector, and beam steerer that directs 50 eV electrons into a charge
exchange region.
It operates in two modes:
●​ Voltage applied: Rejects incoming charged particles; neutral particles are ionized and
analyzed.​
●​ No voltage: Allows charged particles into the spectrometer; neutral particles pass
through unmeasured.​
​
This switching enables INMS to distinguish between particle types.
Deployment in CIRCE:
Each CIRCE satellite has one INMS:
●​ The lead satellite’s INMS is fixed in the ram-facing direction (the direction of motion).​
●​ The trailing satellite’s INMS operates intermittently (when Tri-TIP is inactive), rotating
twice per orbit to measure in various orientations.​
Scientific Utility:
INMS helps:
●​ Identify gas types in the thermosphere.​
●​ Compare neutral vs. charged particle populations.​
●​ Understand altitude- and time-dependent atmospheric variability.​
●​ Enhance drag models and space weather predictions.​
2. RadMon 3.0 – Radiation Monitor
Created by Surrey Satellite Technology Ltd and University of Surrey, RadMon 3.0 monitors
radiation exposure in orbit. It is crucial for understanding the radiation environment around
500 km altitude in LEO, which is important for protecting satellite electronics.
Key Measurements:
●​ Total Ionizing Dose (TID)​
●​ Proton and heavy ion flux​
●​ Radiation dose rate​
Internal Design:
Sensor
Purpose
Placement
3 × RadFETs
Long-term TID tracking
1 internal, 2
external
UV Photodiode
Measures instantaneous radiation dose
rate
Internal
Large-area PIN
Diode
Measures flux of protons and heavy ions
Internal
Why Two RadMon Units?
Each satellite in CIRCE has one RadMon. This allows for:
●​ Cross-calibration​
●​ Tracking dynamic changes in the radiation belts​
●​ Continuous comparison and increased measurement fidelity​
Scientific Relevance:
●​ Monitors solar storm effects and particle spikes​
●​ Improves space radiation models and forecasts​
●​ Assesses risk to COTS components in future missions​
3. TOPCAT – GPS Receiver
Designed by the University of Bath, the TOPCAT receiver tracks GPS signal delays caused
by free electrons in the ionosphere. These delays are used to compute the Total Electron
Content (TEC) along the signal path.
How it Works:
Using phase delays from GPS signals, TOPCAT feeds data into a tomographic model called
MIDAS. This model reconstructs a 3D electron density map of the topside ionosphere and
plasmasphere.
Purpose:
●​ Enhances ionospheric modeling accuracy​
●​ Helps in tracking ionospheric variability​
●​ Supports mission planning by predicting signal distortion zones​
Why IRIS Matters
The IRIS suite answers three key scientific questions:
1.​ Upper Atmospheric Behavior:​
IRIS investigates the variability in atmospheric drag, thermospheric chemistry, and
space weather impacts on the ionosphere.​
2.​ Radiation Hazard Identification:​
With RadMon, it identifies high-radiation zones that may threaten satellites, guiding the
selection of shielding strategies and safer orbits.​
3.​ 3D Ionospheric Mapping:​
Through GPS TEC data and MIDAS inversion, IRIS creates high-resolution
ionospheric maps, improving prediction of space weather effects.​
​
Integration and Contribution to CIRCE
Although each IRIS instrument is valuable individually, together they enrich CIRCE’s overall
mission, especially in supporting its primary payload, the Tri-TIP UV photometer. IRIS adds
environmental context (like gas composition and radiation levels) to the Tri-TIP
measurements, enhancing the accuracy of electron density reconstructions.
Two IRIS units are integrated into CIRCE:
●​ Lead satellite: IRIS occupies the front 2U × 1U section.​
●​ Trailing satellite: IRIS is placed in the rear 2U × 1U segment.​
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TOPCAT II – GPS-Based Ionospheric Mapping Instrument
TOPCAT II is a key instrument on the CIRCE CubeSats designed to measure the Total Electron
Content (TEC) in the ionosphere. It does this by analyzing delays in GPS signals caused by
ionospheric plasma. These delays provide valuable data for constructing high-resolution 4D
maps (3D space + time) of electron density in the ionosphere.
How It Works
GPS satellites constantly transmit dual-frequency signals (L1 and L2). When these signals
travel through the ionosphere to reach CIRCE, they are delayed depending on the number of
free electrons along the path. TOPCAT II measures this signal delay and uses it to calculate the
TEC — the total number of electrons between the GPS satellite and the CIRCE CubeSat.
By collecting multiple such measurements from different GPS satellites at various angles and
times, TOPCAT can build a detailed time-evolving map of electron distribution in the ionosphere.
What’s Inside TOPCAT II
TOPCAT II includes the following main components:
Component
NovAtel OEM719 GPS
Receiver
Description
Modified for space applications; capable of tracking
fast-moving GPS signals at ~7 km/s.
Antcom G5 Dual-Frequency Aerospace-grade antenna that captures L1, L2, and L5 GPS
Antenna
signals.
Controller Board
Manages the instrument’s operations and provides automation
flexibility.
Why TOPCAT II Matters
●​ Improves ionospheric imaging: Enhances vertical resolution in TEC maps.​
●​ Supports space weather and atmospheric science: Helps study how the ionosphere
reacts to solar activity.​
●​ Cost-effective: Built from COTS (commercial-off-the-shelf) components, making it a
budget-friendly research tool.​
●​ Complements CIRCE mission: Works in tandem with instruments like Tri-TIP and
RadMon to provide a fuller picture of the near-Earth environment.​
Ionospheric F Region (150–400 km Altitude) – Energy Sources and
Dynamics
The ionospheric F region is a critical part of the upper atmosphere, rich in charged particles and
influenced by both solar radiation and Earth’s magnetic field. Here’s how it behaves:
Daytime Energy Source: EUV Radiation
During the day, this region is primarily energized by extreme ultraviolet (EUV) radiation from
the Sun, particularly wavelengths below 90 nanometers. This energy ionizes atmospheric
gases, turning neutral atoms into ions and free electrons.
Nighttime Energy Source: Charged Particles
At night, in the absence of EUV radiation, energy comes from charged particles — mostly
electrons and protons from the solar wind and Earth’s magnetosphere. These particles enter the
atmosphere and continue to cause ionization, heating, and chemical reactions, similar to what
sunlight does during the day.
Collisions and Chemical Reactions
This region is dense enough for frequent collisions:
●​ Elastic collisions: Particles bounce off one another without changing their internal
states.​
●​ Inelastic collisions: Involve energy exchange or chemical transformations.​
A variety of chemical reactions occur between neutral molecules and ions, playing a
role in the evolution of the ionospheric environment.​
Role of Earth’s Magnetic Field
The Earth’s magnetic field controls how particles move in this region. It:
●​ Directs the motion of charged particles.​
●​ Influences their distribution and transport across the globe.​