Minimize MicroSCOPE


MicroSCOPE (Micro-Satellite a traînée Compensée pour l'Observation du Principe d'Equivalence) is an approved CNES/ESA gravity-research minisatellite mission, put forward in 2000 by ONERA (Office National d'Etudes et de Recherches Aérospatiales), Châtillon, France, and by OCA (Observatoire de la Côte d'Azur) in Grasse, France. The mission is a collaboration between CNES, ZARM laboratory (University of Bremen), PTB (Physikalisch-Technische Bundesanstalt) of Braunschweig, Germany, OCA and ONERA. 1) 2) 3)

The objective is to conduct a fundamental physics experiment, namely to test the general theory of relativity. This concept is also known as the EP (Equivalence Principle) test. The EP postulates a perfect proportionality between the inertial mass and the gravitational mass of a body. A direct consequence of this Equivalence Principle is the `universality of free fall' such that all objects fall with exactly the same acceleration in the same gravity field. EP turned out to be a major principle of Albert Einstein's general theory of relativity, formulated in 1907 and in 1911.

The microscope mission concept to test EP is similar to the original STEP (Satellite Test of the Equivalence Principle) mission design of 1989 (a NASA/ESA proposal that wasn't realized). The major simplification of MicroSCOPE is the use of a room-temperature instrument (instead of a cryogenic payload operating at 2 K for STEP). - The best accuracy value of the EP achieved by experiments on Earth is 10-13, the MicroSCOPE mission wants to improve this value by two orders of magnitude to 10-15.. Leading-edge technologies are implemented, these include: ultra-sensitive accelerometers, ionic thrusters, and simultaneous spin control and drag-free systems. 4) 5) 6) 7) 8) 9) 10) 11)

The mission performance relies on very high technologies not yet tested all together: full 3-axes and continuous drag-free satellite control, the hybridization of star sensor and accelerometer measurements for attitude control, micro-newton (µN) cold gas thrusters, milli-Kelvin (mK) passive thermal control of the payload, inflight calibration, and a femto-g accurate electrostatic differential accelerometer for the payload.

In the overall organization, CNES is the lead agency in charge of the microsatellite (system, test and integration) and the S/C operations. ESA is responsible for the procurement of the FEEP thrusters. Mission and data analysis functions are performed in close collaboration with ONERA and OCA (Observatoire de la Côte d'Azur)/GEMINI. The ZARM (Center for Applied Space Technology and Microgravity) of Bremen, Germany is contributing to the project (assisting with accelerometer qualification testing in their drop tower and drag-free system analysis) through DLR funding. 12)

Note: In the text as well as in the various references, the spelling of the terms “MicroSCOPE” and “Microscope” are being used freely with the same meaning.


The main satellite/payload design challenges and requirements:

The MicroSCOPE satellite plays the role of a space laboratory devoted to a very complex experiment of physics, and for this it must point the instrument on 3 axis, protect it against non-gravitational forces, and ensure an ultra-stable thermal environment, in particular around the frequency fep which is the frequency of excitation of the equivalence principle.

On the other hand, the restricted funding of the mission that belongs to the CNES Myriade program does not allow to develop specific satellite equipment or to make complex choices of architecture.

• First of all, the reduction of the mean level of applied forces on the instrument introduces a new function: the acceleration control (referred to as drag-free control). Since there is no better instrument than an inertial sensor to measure the external forces, this function uses SAGE (Space Accelerometer for Gravitation Experimentation) as sensor.

- On the other side of the process, a very fine actuator is needed to counteract the perturbations (atmospheric drag, radiation pressure, electromagnetic forces) the mean value of which is not higher than some tens of µN. MicroSCOPE will use a proportional micro propulsion system which is being based on the FEEP (Field Effect Electrical Propulsion) principle.

- The FEEP thrusters are grouped in 4 Electrical Propulsion System Assemblies (EPSA, Figure #). The combination of the six degrees of freedom (attitude and linear acceleration) in one function leads create a new subsystem, called AACS (Attitude and Acceleration Control Subsystem). The attitude and drag-free control requirements call for motion reduction noises as low as 10-12 ms-2 Hz-1/2 (i.e., accelerometer resolution) -- to obtain a test resolution of better than 10-15.

• A second major design driver is the need to have the satellite mass center close to each of the 4 proof masses of the instrument: therefore a set of two sensor units (SUs) are chosen along with the corresponding FEEUs (Front End Electric Units) in a payload structure which is inside the satellite structure.

- The payload position inside the satellite puts severe requirements on the mechanical, and moreover on the thermal design, induced by the thermal sensitivity of the instrument: the required stability around fep is 1mK at SU interface, 10 mK at the FEEU interface.

• A third design driving point is the need to guarantee an ultra-stable dynamic environment to the instrument. The requirement on the attitude rate stability in rotating mode around fep is equivalent to a requirement in attitude stability of the instrument better than 0.166 µrad at a very low frequency. As this order of magnitude is far below the thermoelastic deformation between the instrument plate and the measurement axes of every kind of attitude sensors that can be laid out, it is clearly a challenging requirement for the AACS.

• Finally, the mission requires to reduce or calibrate any kind of dynamic perturbation on the proof-masses, combined with the position of the payload inside the satellite structure and thus close to the platform equipment, create another category of drivers in the architecture: the reduction of internal micro-perturbations. There are three dominating categories of perturbations: a) mechanical (direct forces due to matter displacement), b) gravitational (change of the induced gravity gradient when the satellite shape changes), and c) magnetic (resulting from the electrical activity in the satellite circuits and equipment).


Figure 1: SAGE instrument orientation during flight (image credit: ONERA)


Figure 2: Schematic of SAGE measurement orientation in orbit (image credit: ONERA)

In the MicroSCOPE experiment, the Earth is the gravitational source about which free fall motion of two masses, composed of different materials, is observed and controlled taking care that both masses are submitted exactly to the same gravitational field. The controlled electrostatic field, added to break the experimentation symmetry by forcing the masses to remain on the same orbit is accurately measured: a defect of symmetry gives rise to evidence of an EP violation.

Due to the noise level of the best current ultra-precise accelerometers (10-12 ms-2 Hz-1/2), the aimed accuracy of the test can only be achieved by filtering the signal over a long period of time Ti. Typically there will be two modes:

• Inertial mode: Ti = 120 orbits; fep ≈ 1.7 x 10-4 Hz

• Rotating mode: Ti = 20 orbits; fep ≈ 1.0 x 10-3 Hz


Figure 3: Artist's view of the MicroSCOPE spacecraft (image credit: CNES)



The Myriade series microsatellite bus of CNES is being used (of DEMETER and PARASOL heritage). The microsatellite bus consists of a box-like honeycomb-and-plate structure. Lateral panels are assembled by four L-spar support structures, and can be opened independently during integration. The Z panels are dedicated to CGPS (Cold Gas Propulsion System), and the Y panels accommodate the rest of the equipments except Star Tracker Optical Heads which are located on the –X wall as close as possible of the PAS (Payload Assembly Subsystem) in order to allow a good natural alignment stability of Star Tracker measurement axes with respect to the instrument spin axis. 13)

The design of Myriade (actually a minisatellite in this implementation) is a compromise between high performance, efficiency, robustness and cost. The architecture of the satellite is based on a platform with generic functional chains (AOCS, Energy, Communication, Computer, Structure, Thermal Control), and on a decoupled payload located on the upper part of the platform structure.

The solar array (AsGa cells) with an area of 1.6 m2 consists of two deployed panels, providing a power of 140 W (shared between the electric propulsion system, the instrument and the S/C bus). Particular attention is paid to the thermo-elastic behavior of the S/C as well as to possible magnetic disturbances to the test masses.

The satellite is either spin-stabilized with the spin axis directed normal to the orbit plane - or inertially (3-axis) stabilized. The main functions of the AACS (Attitude and Acceleration Control Subsystem) are to:

- To control the attitude of the spacecraft

- To control the linear accelerations of the satellite (drag compensation)

- To produce the linear and angular excitations of the satellite which are necessary for instrument calibration.

Attitude sensing is provided by sun sensors, a 3-axis magnetometer, and a star tracker (in the anti-sun direction). Attitude actuation and drag compensation is provided by DFACS (Drag-Free Attitude Control System) as described below in the sensor complement. In addition, actuation may be provided by magnetorquers and reaction wheels (when needed like in the orbit injection phase).

The AACS uses the scientific instrument (SAGE) as a sensor for angular and linear accelerations. In mission mode, the only other attitude sensor is a star tracker, µASC (micro Advanced Stellar Compass) of DTU (Technical University of Denmark) with two optical devices, i.e. with two CHU (Camera Head Units).

MicroSCOPE is bigger and heavier than the standard Myriade product line. The S/C mass budget is about 290 kg (hence, MicroSCOPE is a minisatellite), the mission design life is one year. The thermal control of the minisatellite is performed entirely by passive methods.

The external layout (Figure 4) is mainly constrained by antennas accommodation and AACS equipment I/F requirement in order to minimize the modifications with respect to AOCS standard non-mission modes. For centering and symmetry reasons, the solar generator is separated in two identical wings of one panel each, mounted on Y panels and directed toward +X. During launch, the wings are folded; the release is provided by 3 pyrolock mechanisms and the deployment by 2 Carpenter blades. These components are already used on the standard Myriade product line. Solar array driving mechanisms are not needed due to the good energy budget of the dawn/dusk orbit.


Figure 4: The MicroSCOPE spacecraft in deployed configuration (image credit: CNES, Ref. 14)

Payload interface concept: A specific PAS (Payload Assembly Subsystem) has been defined (Figure 5) to achieve the temperature stability requirements of SU and FEEU. PAS is the structure supporting the SU and the FEEU, it includes also the thermal control hardware. The requirements of PAS shall:

- Allow the centering of proof masses with respect to the satellite spin axis

- Offer a very high mechanical stability during the entire mission

- Provide a thermal stability to SU better than 2 mK at fep (peak to peak value)

- Protect the SU from magnetic perturbations

- Guarantee the thermal control of FEEU (10:45ºC operating range and taking into account their 6 W power dissipation each).

The PAS is mounted on the anti-sun panel of the satellite (-X) to take advantage of the high natural external flux stability of this side along the 6 hour local time orbit, and to avoid thermal cycling from Earth's radiative flux. The PAS is made of a two-stage structure. The FEEUs are centered on the first stage where thermal stability requirements is a sinus of half amplitude 10 mK at fep.

The accelerometer sensor units (SUs) are centered on the second stage by four brackets where thermal stability requirement is ten times better (less than 1 mK at fep)), so that the first stage thermal stability can be used for pre-regulation. Radiative insulation of the PAS is achieved using MLI (Multi Layer Insulation). The total mass of the PAS approaches 50 kg, and the volume is 54 cm in diameter by 50 cm in height.


Figure 5: Schematic of the PAS configuration (image credit: CNES)


Figure 6: The thermo-mechanical PAS concept (image credit: CNES)


Figure 7: Layout of the PAS design (image credit: CNES)

Equipment layout: The internal equipment layout is based on three main considerations: mass symmetry, thermal symmetry, and thermal stability. The PAS is located in the center of the spacecraft and equipment masses are balanced on the four lateral panels to have the PAS spin axis close to the satellite center of mass. Dissipations of equipment are balanced on opposite lateral panels (±Ys) so that external fluxes absorbed by the radiators at fep are in phase opposition. Hence, the resulting conductive thermal flux at the PAS interface is minimized. Equipment with lower thermal dissipation loads are fixed on the other panels with poor thermal rejection capacity. The deorbiting system is located on the +Xs panel and deployed at the end of the mission.


Figure 8: Internal configuration of the MicroSCOPE spacecraft (image credit: CNES, Ref. 14)

Minisatellite mass, power

290 kg, ~ 140 W

Overall volume deployed
Overall volume at launch (folded configuration)

138 cm (X) x 2470 cm (Y) x 155 cm (Z)
138 cm x 104 cm x 158 cm

Centering of instrument spin axis w.r.t. center of mass

< 5 mm

Main inertia axis along the X axis

40 kgm2

Thermal stability @ fep on FEEU interface
Thermal stability @ fep on SU interface

< 10 mK
< 1 mK

Thermoelastic stability @ fep star tracker / instrument

< 12 µrad in inertial mode

Table 1: Overview of key satellite parameters

RF communications: The functional chain is the S-band communication link of the Myriade product line with a data rate of 625 kbit/s. The on-board solid-state recorder has a capacity of 1 Gbit (the total source data rate, including housekeeping data, is about 1 kbit/s). The spacecraft operations are conducted at CNES. The science mission center is at ONERA.


Figure 9: Instrument axis definition in the spacecraft (image credit: CNES)

Status of project: 14)

• The MicroSCOPE project approval was given in 2004.

• The MicroSCOPE PDR (Preliminary Design Review) was held in March 2011.

The long time spent between these dates is justified by a major change in the propulsion system (initial propulsion system using FEEP (Field Effect Electric Thrusters) has been replaced by cold gas thrusters requiring a considerable rebuilt of the satellite.

During this time, the development of the payload kept on through several EMs (Engineering Models). The qualification model of SU has been manufactured (Figure 20); the qualification campaign started in December 2010 and it is scheduled to end before the end of 2011.

• In the spring of 2011, the project advanced to Phase B.


Launch: A launch of MicroSCOPE is planned for the timeframe 2015 from Kourou.

Orbit: Sun-synchronous quasi-circular dawn-dusk orbit, altitude = 720 km, eccentricity = 5 x 10-3, inclination = 98º, LTAN (Local Time on Ascending Node) = 6:00 hours or 18:00 hours. The low eclipse dawn/dusk orbit is chosen to maintain the payload in a very stable thermal environment.
Note: The nominal altitude (720 km) is a compromise between several factors, among which mainly the need of the highest possible gravity signal (favored by low orbits) and the minimization of the atmospheric drag and other perturbations (better if the orbit is higher).


Sensor/payload complement: (SAGE, DFACS, CGPS)

The objective of the EP test is to expose two test masses to the same gravitational force to see if they are both affected in the same way. This, of course, requires concentric masses, with their centers of mass precisely aligned. However, as they cannot be idealized point masses, they will in fact be affected by the local gravity gradient.

Experiment principle: The experiment consists in controlling the free fall motions of two quasi-cylindrical masses made of different materials on the same orbit inside a dedicated satellite. They are subjected to the same gravitational field. To achieve this in the variable field of Earth's gravity, two restrictions are placed on the accelerometer design: the masses must be concentric to share a common center of gravity, and the shape of the masses must be chosen so that the gravity gradient effects are analogous on the two masses. 15) 16) 17) 18) 19) 20) 21) 22) 23) 24)

The satellite constitutes for the experimental masses a shield from the atmospheric drag and from the Sun and Earth radiation pressures. It carries a specific electrical propulsion system to compensate these surface forces and to follow the masses.


SAGE (Space Accelerometer for Gravity Experiment):

SAGE is the differential accelerometer designed and developed by ONERA/DMPH (Département Mesures Physiques), Châtillon, France. SAGE consists of a set of two cylindrical concentric electrostatic differential accelerometers designed and developed at ONERA. The SAGE design is of STAR and SuperSTAR (Super Space Three-axis Accelerometer for Research mission) heritage, flown on the CHAMP and the GRACE missions, respectively. The combination of the two differential accelerometers (SAGE-EP and SAGE-REF) constituting the scientific instrument is also referred to as T-SAGE for Twin-SAGE due to the fact that it consists of two differential accelerometers.

An electrostatic accelerometer consists, fundamentally, of a proof mass (PM) suspended in a highly stable electrode cage. The principle of operation is to measure the electrostatic forces required to maintain the position of the proof mass with respect to the electrodes. The proof or test mass of each accelerometer is maintained along the three orthogonal axes at the center of the fused silica instrument cage by electrostatic forces. Electrodes, engraved in the cage wall, are used for the capacitive sensing of the mass position and attitude. Both test masses are controlled with respect to the cage frame: the sum of the forces is maintained to a null value by the satellite drag compensation system; the thrusters of the electric propulsion system act to move the instrument silica frame following the masses. The difference of the electrostatic forces is then observed in the orbital path for the search of the EP-violating signal. The electrode configuration allows six-degree-of-freedom measurements and control of each test mass.

Rotational axis of S/C

About any axis ( all axes x, y, z, are identical)


Max value at DC

Stability at fep

Angular velocity

10-6 rad/s, [or 3.9 and 4.9 x10-3 rad/s (spin)]

10-6 rad/s Hz-1/2

Angular acceleration

2 x 10-6 rad s-2

2 x 10-8 rad s-2 Hz-1/2

Linear acceleration

3 x 10-8 ms-2

3 x 10-10 ms-2 Hz-1/2

Table 2: SAGE velocity and acceleration specifications

In the case of SAGE, two inertial sensors are positioned to give their proof-masses the same center of gravity, to form one differential accelerometer. To obtain such a configuration the masses are cylindrical, with carefully machined dimensions, to approximate spherical moments of inertia, to reduce the gravity gradient disturbing effects.


Figure 10: Illustration of the SAGE instrument (image credit: ONERA, Ref. 24)

When considering a perfect instrument, the measurement of a difference in the acceleration of two concentric test mass in a perfect Earth geodetic gives the evidence of the Equivalence Principle violation. Indeed when orbiting around the Earth the two concentric test masses see the same gravity and thus by equivalence the same acceleration independently on their mass or composition.

However nothing is perfect and one has to consider the machining and integration accuracy of the parts that will lead to a relative miss-centering of the two masses. If 20 µm mis-centering is considered, the gravitational effect on the measured acceleration is of more than 8 x 10-15 ms-2 for only one contributor. Obviously the test mass centering is calibrated in orbit with 0.1 µm accuracy helping to reject by a factor 200 this contributor. Nevertheless, on ground a centering of 20 µm along all axes has to be performed.

As shown in Figure 11, the centering of test masses is ensured by electrostatic positioning. The same electrodes serve for the position detection through capacitive sensing at 100 kHz and for the positioning through applied voltages on electrodes. The applied voltages are calculated by a digital controller (Digital Signal Processor TMS21020) taking into account the six degrees of freedom of each test mass. Thus, the test masses are servo-controlled to be motion less with respect to the electrode setting seen as the reference frame.

The 4 cylinders of the electrode set are placed on a gold-coated silica reference ‘Hat’ as in Figure 12. Then the relative centering of the two test-masses of each sensor unit depends on the quality of the assembly.


Figure 11: Payload reference sensor core (EP sensor core is identical but with a outer test mass made of titanium alloy), image credit: ONERA


Figure 12: Qualification Model during integration (image credit: ONERA)

The servo-control of the test mass works on the principle of balancing the capacitances for each degree of freedom. In the Figure , the balancing of the capacitances C1 and C2 used for the X control will induce a displacement of the test-mass along X proportional to the cone angle. This particular effect is also considered in addition to the machining accuracy of the electrode set and the reference ‘Hat’.


Figure 13: Effect of the TM conicity on the TM centering (in yellow) when measuring X displacement with electrodes C1 and C2 (in blue), image credit: ONERA

The two differential accelerometers on the satellite are identical except for the material of their masses. The instrument providing the science baseline (proof mass) has both masses in platinum-rhodium, while in the EP test instrument has the external mass in titanium and the internal in platinum-rhodium. The titanium mass has a nominal length of 79.9 mm, outer radius of 35 mm, and mass of 0.364 kg, while the smaller platinum mass has nominal dimensions of 43.51 mm in length, 20 mm in outer radius, and a mass of 0.473 kg.

Two differential accelerometers in SAGE:

• Two masses of identical material (PtRh) for test accuracy verification

• Two masses of different material (PtRh/TA6V) for the EP test.

The two inertial sensors are concentric coaxial cylinders, with a common center of mass between the two proof masses. The cylinder axis corresponds to the sensitive measurement axis (X). Each mass is maintained centered in six degrees of freedom by means of electrostatic forces from a surrounding cage of electrodes in gold-coated silica. The only physical contact between the mass and its surrounding cage is a 5 µm diameter gold wire, which is necessary to maintain its electrical charge stable, and also applies a high frequency voltage used for the capacitive position sensing.

The electrodes, working in pairs, are used for both displacement detection, by differential capacitive measurement, and position control, by electrostatic force actuation. Position sensing and actuation are designed to be linear along the instrument sensitive axis, the cylinder axis, in such a way to permit fine frequency signal analysis of the data.


Figure 14: Internal configuration of the SAGE sensor unit (image credit: ONERA)


Figure 15: SAGE electrode configuration for proof mass position measurement (image credit: ONERA)


Figure 16: Prototype of one electrostatic accelerometer core (image credit: ONERA)

The SAGE instrumentation consists of three components: (SU, FEEU, and ICU)

SU (Sensor Unit) core = differential accelerometer, consisting of:

- Two SUMI (SU on a Mechanical Interface)

- Each SU = 2 test masses centered to 20 µm

- 1 mass corresponds to 1 inertial sensor (defines measurement frame)

Each SU comprises the concentric pair of test masses which motion is to be compared along the orbit travel. The masses are surrounded by a silica core, graved and gold coated to provide the necessary electrodes for the electrostatic control. The entire core is enclosed in a tight housing performing 10-5 Pa vacuum to minimize the residual parasitic forces on the masses like the radiometer effect or the outgassing.


Figure 17: Illustration of the T-SAGE instrument composition (image credit: ONERA)

FEEU (Front End Electronics Unit) which contains the low noise analog electronics required for proof mass levitation, including the ADCs (Analog Digital Converters), DACs (Digital Analog Converters), and position sensors. There is 1 FEEU for each SU.

Each FEEU includes the capacitive sensing of the pair of masses, the reference voltage sources and the analog electronics to generate the electrical voltages applied on the electrodes. Each FEEU is associated to one SU. Associated to the SU in a thermal “cocoon”, the thermal environment provided by the satellite is a few mK for the FEEU and a fraction of 1 mK at orbital frequency for the SU. This helps to provide very steady voltages and operation conditions for the measurements.

ICU (Interface Control Unit) which contains the remaining electronics for SU operation, specifically the proof mass position control loop, as well as the systems for general experiment control and the satellite interface. The components are:

- 1 ICU for each FEEU (2 stacked together)

- 1 DSP (Digital Signal Processor) + 2 FPGA (Field Programmable Gate Array) for test mass control laws and data conditioning for the OBC (Onboard Computer)

- 2 Power Control Units (1 redundant) to convert the sat 28 V to stable secondary voltages (±45 V, ±15 V, +5 V, 3.3 V)

Both ICUs are integrated into a common device called ICUME (Interface Control Unit Mechanical Ensemble) as shown in Figure 17, providing the interface with the spacecraft.


Figure 18: Control loop scheme of the EP test axis (image credit: ONERA)


Figure 19: Sketch of the two SU sensors (image credit: ONERA)


Figure 20: Photo of the SU qualification model (imafe credit: CNES, Ref. 14)


Figure 21: Cross-sectional view of the SAGE sensor unit (SU), image credit: ONERA

Both differential accelerometer cores are integrated in tight vacuum housings with thermal insulation and magnetic shielding provided. The instrument temperature variation is < 0.1º C per orbit. The entire instrument is positioned at the center of mass of the spacecraft to reduce any torque demand onto the electric propulsion system. The sensitive axes of the accelerometer are oriented in the orbital plane of the S/C x-axis, with the center of the test masses on the rotating axis of the S/C (normal to the orbital plane). The total instrument mass is 40 kg, the power is 20 W for each of the differential accelerometers.

Operations of the accelerometers: After all initial tests and calibrations, the EP experiment is realized with the first differential accelerometer in inertial and spinning attitudes and with two angular phases along the orbit. The test is concluded with a calibration. The same procedure is repeated with the second differential accelerometer. 25)


Figure 22: Test of the EP in space with the MicroSCOPE concept (image credit: ONERA)


DFACS (Drag-Free Attitude Control System):

The mission requirements have several consequences on the satellite design. The spacecraft must protect the payload from all nongravitational forces perturbing the EP measurement; hence, an active control of accelerations and attitude of satellite is necessary to reach the required background level of acceleration.

The MicroSCOPE documentation provides several names for this control system. There is the designation DFACS as well as AACS (Acceleration and Attitude Control System).

DFACS is the attitude actuator and measurement system of the MicroSCOPE mission providing accurate measurements of the spacecraft accelerations in 6-DoF (Six Degrees of Freedom). DFACS employs a CGPS (Cold Gas Propulsion System).

Note: The initial design of DFACS (aka AACS) featured a FEEP (Field Emission Electric Propulsion)-type system -to be developed by Alta S. p. A. of Pisa, Italy. However, in 2009, the FEEP system was replaced by CGPS, developed by CNES with a major ESA contribution; its design is mainly based on the MPS (Micro Propulsion System) of ESA's GAIA mission, developed by TAS-I (Thales Alenia Space-Italia), Ref. 14). 26)

DFACS modes of operation:

In the MicroSCOPE mission, the AACS (Attitude and Acceleration Control System) replaces the Myriade generic AOCS (Attitude and Orbit Control System) because the the real-time acceleration control of satellite is performed instead of the orbit control. However the two systems are very similar.

The AACS architecture is based on five modes (Figure 23):

MLT is the launch mode, all the equipments are off.

MAS (MicroSCOPE Acquisition & Safe-hold) mode. MAS employs the following attitude sensors and actuators (magnetometer, sun sensors, magnetorquers, reaction wheel) to obtain a stable sun-pointing attitude autonomously. MAS corresponds to the Myriade product line standard acquisition / safe mode.

MGT2 (MicroSCOPE Geomagnetic Transition mode). MGT2 is a transition mode from acquisition to normal mode providing a coarse pointing. The attitude guidance is conical and rotating with a pro-grade spin rate of 2 vo (the satellite makes one revolution per orbit in an orbital frame). The control is based on the measurement of the magnetic field and its comparison with the on-board model.

Note: Normally, the Myriade series spacecraft (DEMETER, PARASOL, etc. ) spin at the orbital frequency providing a geocentric pointing, whereas MicroSCOPE spins at twice the orbital frequency (that is what the “2” of MGT2 stands for), with a roll bias.

MSP (MicroSCOPE Stellar and Propulsion) mode, specific to MicroSCOPE only. Its control loop uses the star tracker as sensor, the CGPS thrusters as actuators, and three software functions (attitude and angular rate filter, control, command distribution). It has three functions:

- Ensure a transition from MGT2 to mission pointing in MCA (inertial or spinning)

- Allow technological operations (such as testing the instrument, or perform a maneuver to avoid the moon blooming in the star tracker)

- It constitutes an intermediate safe mode from MCA in case of an anomaly due to the instrument.

MCA (Acceleration Control Mode). This drag-free mission mode is specific to MicroSCOPE allowing the EP measurement. Its main objective is to provide linear acceleration control, with a very high accuracy (10-12 ms2 residual acceleration at fep), in addition to the Myriade series traditional 3-axes attitude control. Figure 24 illustrates the control loops in MCA mode:

- Linear measures provided by SAGE are directly sent to the controller while the angular measures are introduced in the control loop after hybridization with star tracker measurements

- In MCAi (inertial) the attitude command is inertial, while in MCAs (spin) there is a commanded rate of 5 x 10-3 rad/s around the X-axis.

- MCAc is the calibration mode in inertial coordinates. Instrument calibration (scale factors and common-to-differential modes coupling, as well as FEEP calibration) is being conducted by sequentially commanding a low-frequency (0.005 Hz) sine excitation one each DoF.

The main issue is the quality of attitude measurements at fep. While the target is 0.166 µrad of attitude stability at fep on SAGE, the thermoelastic distortion of the platform between SAGE and the µASC (star tracker) is typically 5 µrad.


Figure 23: Overview of AACS/DFACS modes and transitions (image credit: ONERA)


Figure 24: Block diagram of the MCA control loop (image credit: CNES)

The DFACS system performance was already assessed and verified in a simulator, conducted by Astrium SAS, France.

Note: the terms AACS, and DFACS are being used in various papers to define all the same thing.

AACS (Acceleration and Attitude Control System), aka DFACS. AACS is used to suppress the non-gravity forces applied to the test masses. Due to the extreme precision required by the mission, this drag-free control system shall provide the satellite with very high performances in particular in the frequency of the EP check (fep). The AACS uses the inertial sensor unit's acceleration measurements (which is also the scientific instrument) and CGPS micro-propulsion.


Figure 25: Artist's view of the DFACS jet cones (image credit: CNES)


CGPS (Cold Gas Propulsion System):

The CGPS is composed by two identical and independent subsystems called CGPSS (Cold Gas Propulsion Sub-System) which are accommodated on –Z and +Z panels (Figure 8). Each CGPSS (Figure 26) is composed of 4 modules:

• GDM (Gas Distribution Module) stores and maintains the gas at its operational range (pressure and temperatures).

• PRM (Pressure Regulation Module), provides the gas distribution to the thrusters, it contains all the equipment units necessary to ensure the pressure regulation of the CGPS.

• TRM (Thrust Regulation Module) contains 4 nominal and 4 redundant MT (Micro Thrusters).

• ECM (Electronics Control Module) contains the electronics items necessary to provide the power supply to all the CGPSS modules, control the TRM thrust, and ensure the avionic I/F with the OBC.

Each GDM is composed by 3 Arde D5048 carbon overlapped pressure vessels filled with 8.25 kg of gaseous Nitrogen stored at the maximum pressure of 345 bar.

PRM is composed of existing off-the-shelf equipment, only the pressure regulator needs to be delta qualified to withstand the maximum pressure of 345 bar.

CGPS is developed by CNES, ESA is providing the TRM and ECM devices; the CGPS design is mainly based on the GAIA MPS microthrusters developed by TAS-I.


Figure 26: Overview of the CGPSS configuration (image credit: CNES)

The candidate MTs (Micro Thrusters) are those developed by TAS-I in the frame of the GAIA project (launch planned for 2013); their qualification has been achieved in 2011. The MTs operate in a close-loop configuration using a miniaturized MFS (Mass Flow Sensor) as thrust measurement and piezoelectrics actuator to modify the nozzle section and modulate the gas flow.

The candidate ECM is a new MicroSCOPE development based on the existing equivalent electronic module of GAIA; the MT driver boards and DC/DC boards are the same as those of GAIA; the CPU card and the I/O card have been designed specifically for MicroSCOPE.

To improve the ECM time response, a specific new algorithm including an anti-hysteric controller, has been under study by TAS-I. This package is planned to be implemented in ECM of MicroSCOPE.

Two other technologies, developed in Europe, are under consideration for MT and its control electronic: 1) a Bradford concept controlled through a pressure sensor, and 2) the MEMS technology developed by Nanospace. Their performances (Isp, noise, thrust range, resolution, etc..) should be compliant to the MicroSCOPE main requirements; an ITT (Invitation to Tender) has been issued by ESA to select one of them.

The CGPS shall be operated on a continuous basis, because the noise generated by the pulsed thrusters is not compliant with the noise specification and would saturate the T-SAGE range.

Thrust range

1-300 µN


2 µN

Thrust axial noise

< 3.22 µN rms in the (0.001:10) Hz bandwidth.

Thrust linearity

< 5%

Response time

250 ms at 1σ) in total thrust performance range

Table 3: Main requirements of the CGPS system


Figure 27: Illustration of the cold gas configuration on the spacecraft (image credit: CNES)


Auxiliary payloads on MicroSCOPE: (GNSS subsystem, IDEAS)

GNSS subsystem:

The GNSS subsystem, based on two GPS antennas and a new equipment, is a low-cost software Galileo/GPS receiver under development at TES. It is a technological passenger which will be used for on ground orbit determination in addition to Doppler ranging.

IDEAS (Innovative DEorbiting Aerobrake System):

CNES is abiding by the “Space Debris Code of Conduct” rule, adopted in 2004, which recommends that after the end of operational life, the orbital life span of a LEO spacecraft should be < 25 years. 27)

MicroSCOPE adopts a passive deorbiting subsystem based on two sails deployed at the end of the mission by two Gossamer arms. In this way the surface/mass ratio is increased and the altitude of the orbit is naturally reduced by the atmospheric drag. The inflating of the arms is ensured by gaseous Nitrogen stored at high pressure in a dedicated titanium vessel. A passive solution has been preferred to an active (solid propulsion) implementation because of its low development cost and its adaptability to the existing design of MicroSCOPE.

A drag strategy, based on deployable wings with an inflatable mast, was selected. Two 4.5 kg wings made of an aluminized Kapton membrane (100 gram/m2 density) will be deployed by a central inflatable mast - providing a total mean drag surface of 6.3 m2. The deorbit assembly has a total mass of 12 kg.

To avoid the generation of long orbital decay, the membrane is reinforced by an embedded mesh. A patented inflated mast concept of Astrium Space Transportation will be used. From the folded state, deployment is controlled by TADECS (Tetragonal Accordion Deployment Control System). The inflation system is based on cold gas technology. It should be flexible enough to be usable for applications of various sizes.


Figure 28: Deorbiting configuration of the spacecraft with wings and inflatable mast (image credit: CNES)


Figure 29: Schematic view of the deployment control principle (image credit: Astrium ST)

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26) The information was provided by Valerio Cipolla of CNES, France.

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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.