- Email: [email protected]
Design of a micro-satellite for precise formation flying demonstration
Acta Astronautica 59 (2006) 862 – 872 www.elsevier.com/locate/actaastro
Design of a micro-satellite for precise formation flying demonstration Richard Sanchez∗ , Patrice Renard EADS-Astrium, 31 av. des Cosmonautes 31402 Toulouse Cedex 4, France Available online 30 September 2005
Abstract This paper presents a joint CNES-EADS Astrium contribution to the ESA’s SMART-2 project. SMART-2, 2nd mission of the Satellite Missions for Advanced Research and Technology program, slated for launch in 2006, will test key technologies needed to develop two ambitious ESA missions: • LISA (Laser Interferometry Space Antenna), an ESA cornerstone mission dedicated to the detection and observation of gravitational waves; to be launched in 2011, • DARWIN, another ESA cornerstone mission dedicated to the search of Earth-like planets; to be launched in 2015. In Phase A study of this demonstrator, one of the options contemplated by ESA was considering two formation-flying satellites. In that sense, and in order to both reduce and share cost, CNES proposed with the technical support of EADS-Astrium, to build one of them from its Myriade micro-satellite product line, mainly used for LEO scientific applications. The study carried out has permitted to validate the concept of using a low-cost micro-satellite in a scientific interplanetary mission requiring not more than 10 m inter-satellite position accuracy! © 2005 Elsevier Ltd. All rights reserved.
1. Scope of the Sat definition for SMART-2
three major functions:
The study was focused on the Darwin technology demonstration part of the SMART-2 mission, with objective to validate the capability of two satellites to fly with accurate location from each other. To achieve this objective, the proposed micro-satellite had to achieve
• Radio-frequency metrology; • Optical metrology; • Guidance, navigation and control (GNC).
∗ Corresponding author.
E-mail addresses: [email protected] (R. Sanchez), [email protected] (P. Renard). 0094-5765/$ - see front matter © 2005 Elsevier Ltd. All rights reserved. doi:10.1016/j.actaastro.2005.07.039
On one-hand the metrology was providing the GNC with distance measurement within the required accuracy. On the other hand, critical sensor/actuators technologies were to be implemented in the GNC to control such accuracy.
R. Sanchez, P. Renard / Acta Astronautica 59 (2006) 862 – 872
Thus, the major challenges for the Myriade platform were mainly related to • the accommodation of the radio-frequency units and the optical bench of the high precision optical metrology (HPOM), • the management of the coupling effect between the attitude control system and the metrology unit measurements (with end-to-end performances requirement of 10 m inter-satellite position accuracy), • the use of electrical propulsion system (FEEP), • the configuration during orbit transfer (concept of cargo and Sat passenger).
863
Towards optical target
Towards optical target
∆ OPD
2. Darwin mission requirements summary This chapter describes the major Darwin phases. 2.1. Formation deployment This phase covers the operations from separation of the eight Darwin spacecraft from the launcher(s) until a coarse constellation is created with the accuracy of 10 m at inter-satellite distance > 250 m. At the end of this phase, the flyers and the master spacecraft have reached their nominal positions relative to the hub with a 1 cm3 uncertainty; a configuration allowing subsequent laser metrology enabling and fringe acquisition. The modes described hereafter use of RF sensing only for constellation navigation. 2.2. Fringe acquisition The first step after deployment consists in stabilizing the baseline and improving the lateral positioning by using laser metrology (ranging and coarse lateral sensor); in order to enable inter-satellite beam exchanges and to reach the inter-satellite distance (ISD) stability necessary for fringe acquisition. The absolute pointing of the flyer spacecraft is also improved by shifting from the star sensors to the telescope signal. At GNC level, the optical path delay (OPD) is due to two independent errors (RSS combination): • An uncertainty on the inter-satellite distances <1 cm; • A flyer transverse offset relative to the interferometer plane defined by the observation direction
Fig. 1. Correct and tilted constellation configuration (2D example) creating external optical path delay (OPD).
<1 cm. Indeed this offset induces an OPD error (so-called “external OPD”) due to the tilt of the baselines relative to the observation direction (telescopes line-of-sight) as shown in Fig. 1. The uncertainty on OPD obtained from RF sensing and coarse transverse optical metrology is then 1.4 cm. 2.3. Nulling interferometer mission During normal operation, the beams transmitted via the telescope flyers must be co-phased for recombination in the central hub. The OPD between the beams from a pair of telescope flyers must be <20 nm, split in 5 nm allocated to control error and 15 nm to instrument systematic errors. The interfering beams must also have balanced 1 ( 1000 th) intensities, which requires to have beam superimposition on the fringe tracker with the same 1 accuracy ( 1000 th) of the diffraction spot of 1.22 /D, where D is the telescope diameter). This corresponds to 0.08 mas on the detector for 1.5 m telescopes. 2.4. Imaging interferometer mission It is expected that OPD and pointing requirements will be significantly relaxed for imaging (by at least a
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factor of 10). Indeed, conventional OPD requirements for interferometric imaging are in the range of /10 to /100, and beam superposition error only affects 1 1 fringe contrast. Consequently, an accuracy of 10 – 100 of the diffraction spot is sufficient. As a starting point, 1 100 th is assumed. The major difference compared to Nulling is that the interferometer baselines shall be continuously varied (length and orientation in the interferometer plane) during observations to achieve complete scanning of the Fourier plane within a reasonable duration.
3. SMART-2 and Darwin compared requirements The study considered that all technologies required for Darwin would not be tested on SMART-2. Some of them would be verified by ground-based measurements. The flight tests would concentrate on formation deployment, reconfiguration and precision formation flying for the two scientific modes, Nulling and Imaging. Although the analysed SMART-2 option was composed of only two satellites, all the formation flying key technologies required for DARWIN, could, however, be validated. The foreseen sequence of modes was: Orbit transfer: • Transfer phase from Earth up to L1 Lagrange point; Formation deployment: • Stabilization of attitude and on-station control • Initial formation acquisition and collision avoidance quite similar to Darwin; • Coarse baseline control mode using star tracker and RF subsystem measurements: range: 25–250 m; range accuracy: 1 cm; attitude accuracy/inertial frame: 0.5◦ . Nulling interferometer mission: • Acquisition of the 1-cm position accuracy in 3D. With only two satellites this would require line-ofsight measurements with a refined accuracy from a transverse laser metrology system (e.g. a divergent laser associated with a CCD detector).
Fringe acquisition: • Length freezing mode based on laser metrology, down to the accuracy needed for fringe sensor acquisition. • Fine control mode using the fringe sensor. Imaging interferometer mission: • Reorientation and resizing between Nulling mission observations and during Imaging mission observations. Table 1 compares GNC requirements values between Darwin and SMART-2.
4. SMART-2 technology requirements 4.1. RF metrology The RF metrology system is a major technology step forward required for Darwin, the performances of which cannot be assessed on ground because of the difficulty to reproduce the conditions of operation in space (relative dynamics, RF environment). RF metrology shall therefore be validated to the maximum extent by the SMART-2 demonstration. The RF metrology will be used in all Darwin-related modes and for the initial deployment. The “basic hardware set” of the RF subsystem proposed in the study was a pair of “transceivers” (emitter/receiver) and a set of emit/receive antennas flown on the two satellites with the following two functions: • 2-Way pseudo-ranging functions: each transceiver emits a GPS-like signal carrier modulated with a binary code (the navigation message). The opposite transceiver receives this signal, demodulates the binary code and measures its delay relative to an internally-generated code. When multiplied by the speed velocity, this delay provides a pseudorange measurement. Their respective pseudo-range measurements are then exchanged between the transceiver (this is autonomously achieved through the navigation signal) to resolve the clock bias between them. The ISD information is therefore
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Table 1 SMART-2 and Darwin compared GNC performance requirements Mode
Parameter
Darwin requirement
SMART-2 requirement
Deployment mode (coarse optical sensors) Baseline control mode (RF metrology sensors)
Minimum inter-sat. distance Sun direction
20 m 40◦ half-cone from solar array normal 25–250 m, selectable <1 cm & 1 mm/s 3D <1 arcsec & 0.1 mm/s
20 m 45◦ half-cone from solar array normal 25–250 m, selectable 1 cm, az & el err. <0.25◦ (half cone) Use off-the-shelf sun sensors formation roll <0.5◦
400 s 0.1 Hz
400 s 1–10 Hz
4 m at = 1 m 50 m at = 4 m 0.4 m/s at = 1 m 5 m/s at = 4 m ±80 m
5–50 m
Fringe acquisition (optical metrology perfos)
Relative distance Relative positioning Absolute pointing Scan duration Fringe sensor sampling rate Scan step Scan rate
ISD absolute mismatch ISD stability ISD drift OPD stability Global tilt Nulling interferometry (optical metrology perfos)
Imaging interferometry (optical metrology perfos)
Absolute attitude to the flyers Flyer relative angular velocity Relative hub-flyer attitude Relative hub-flyer positions Relative hub-flyer velocity Baseline change rate Duration of imaging periods Resizing
available on both satellites at any time. This function requires dedicated emission/reception RF antennas and relevant transceivers. In order to cover the whole space, two antennas placed in opposite
50 m/s
0–2 cm
2.8 m 0.28 m/s
>5.50 m >50 m
5 nm 0.002 mas @ 250 m (corresponding to 5 nm OPD) 8 mas
±1 m 1◦
0.1 mas/s
Not applicable
60 mas 1 cm 3D 1 mm/s 3D
15 arcsec 0–2 cm ISD az & el: 0.25◦ half cone ISD : 1 mm/s
<3.75 mm/s
<3.75 mm/s
TBD s
TBD s
ISD : 25–250 m and return ISD <1 m in 1 s Duration 16 h
Same as Darwin
10 arcsec
locations need to be implemented on each satellite (Fig. 2); • Transverse position (azimuth/elevation) determination function: this function is in charge of
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LTPter LTP Mas LiLisa sater
Precision Flying Axis (Xsat)
Da rw in Darwin DTµµSat PSaµSat t Flye r FlFlyer ye r
Z Transmit Receive View from X Fig. 2. RF antenna implementation on SMART-2.
determining the direction of the other satellite (defined in S/C frame through elevation and azimuth angles). The principle of so-called “RF goniometry” is to measure the direction of the incoming RF signal using the differential phase of the carrier signal received by three antennas (Fig. 2). The accuracy is directly proportional to the distance between antennas, and the main error source is multipath (reflection of the RF signal on nearby obstacles), so the antennas need to be implemented close to the edges of a satellite wall with minimum obstacles between them. When combined with the measurements of an absolute attitude sensor (e.g. a star tracker) on both satellites, this function allows re-building of the 3D relative motion of one spacecraft relative to the other, in the reference inertial frame. Let us note that having a star tracker on each satellite and RF goniometry on only one end is sufficient to determine the relative position. RF goniometry on both ends would also provide the relative attitude between the S/C, of no interest for SMART-2. Therefore, the “minimum” RF metrology configuration proposed for SMART-2 would implement the RF goniometry function only on the LISA satellite, to maximize possible antenna separation and reduce subsystem mass/power demand on the Darwin satellite. 4.2. Optical metrology SMART-2 had to demonstrate the feasibility of formation flying with accuracy in the nm range. Such
extreme performances call for precise and complex systems. For this purpose, different optical equipments with optimised dynamical range and accuracy had to be accommodated. Two different metrology levels have to be considered in a Darwin type experiment. The internal metrology that freezes the constellation configuration, and the external metrology which ensures the correct pointing of the interferometer onto the guide star. SMART-2 experiment was representative of the first type of metrology. The line-of-sight acquisition and pointing could be done with high accuracy by means of classical equipments such as CCD camera and low power laser source. These equipments were covering the 25–250 m distance range between the two satellites. The inter-satellite distance measurement accuracy was much more demanding and was driving directly the choice of laser metrology technology. The three main parameters sizing the laser were: the length of coherence of the laser source, the 1 m accuracy and the capability to generate different wavelengths for absolute distance measurements. For SMART-2, a low coherence source associated with two fixed delay lines with lengths of 25 and 250 m to cover the ISD range were selected (Fig. 3). The Sat metrology hardware was then able to provide the following accuracies: • Coarse lateral position of the two satellites, approximately 1 mm; • Fine lateral and longitudinal distance between the two satellites, 10 m; • Fringe position in the interferometer defined by an internal fiber reference and the counter satellite, 1 m. It was also possible to develop a simple fringe-tracking sensor, but this approach was found to be less representative. 4.3. Micro-Newton propulsion system The micro-thrusters used for drag-free and attitude control during the LISA demonstration could also be used for accurate pointing and relative position of the two satellites of the Darwin demonstration. As a consequence, the N-thrusters on the Sat had to
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Fringe Fr Friinnggee Tracker Tr ke r Tr ac acker White iteLight Li ght source sour ce
ReLe fere Reference ngnc th e Length Re fere nc e Le ngt h Delay Line
Lo ngi ng itud Longitudinal tudiinnaall Sensor Se ns Se ns or or (10 µm) (1 0 µm µm)) (10
La serSource So ur ce Laser ser S our ce
UHF Link
Fine Lateral Sensor (10 µm)
Co ar se Lateral La ra ll Coarse Co ar se La te tera Sensor Se Se ns nsor or
Collimated Laser
DARWIN µsat
LISA mini-satellite
Fig. 3. Optical metrology implementation on SMART-2.
102
Smart-2 specification Darwin specification Solar Pressure qthrusters=1 qthrusters=0.1 (LISA) qthrusters=0.01
100
ISD error (µm)
provide a 6-DOF control capability with tolerance to the failure of at least one thruster (or one cluster if arranged in clusters). The required authority was in the range of 100 N. Another challenge was the capability of the Sat to support long consumption from FEEP and the use of this electrical propulsion as unique propulsion module for all modes, in particular on station setting after the orbit transfer. The thrust noise was the driving factor as opposed to thrust vector misalignment, bias and scale factor. In order to derive thruster noise requirements, a preliminary allocation of the ISD error contributors was considered assuming equal quadratic weighting:
10-2
10-4 10-6 10-8 10-3
10-2
10-1
100
101
Controller cut-off frequency (Hz)
ISD = ISDsensors + ISDsolar_pressure + ISDthrusters √ 1/ 2
1/2
√ 1/ 2
1/2
Fig. 4. Thruster noise impact on the inter-satellite distance (ISD).
(1) The thruster noise transmission in the GNC loop was evaluated as a function of the control bandwidth.
Fig. 4 compares OPD requirements for SMART-2 and for Darwin (corrected by the SMART-2 and Darwin masses ratio, and assuming the same error allocation for Darwin), and the OPD caused by solar pressure
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Fig. 6. SMART-2 Sat in orbit configuration.
• the dominant disturbance is the thruster noise con√ sidering requirements of 0.1 N/ Hz, as for the LISA demonstration. The resulting control bandwidth of about 0.01 Hz (about 0.5 Hz for Darwin) is significantly higher but still compatible with the thruster technology. 5. SMART-2 Sat configuration The mechanical architecture coped with the Myriade structure to enhance the re-use and associated cost gain. It is worth noticing that the design was defined with:
Fig. 5. SMART-2 Sat in stowed configuration.
√ and various thrust noise levels (specified in N/ Hz) as functions of the controller cut-off frequency. The major conclusions were: • the rejection of the solar pressure can be achieved with very low control bandwidth: <0.001 Hz (about 0.05 Hz for Darwin);
• two batteries; • two UHF transponders and relevant antennas (optional); • two RF metrology units and relevant antennas (with possibility to have them internally redunded to provide additional mass margin). On the other hand, it was assumed that the same antenna was used for signal transmission and reception, and that only one antenna was required for the Darwin satellite (baseline for pseudo-ranging function only); finally, two identical antennas were embedded to perform omni coverage; • one emergency transponder and one antenna (optional).
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6. SMART-2 Sat functional description 6.1. Overview The main challenge was to deal with a Sat concept together with highly challenging performances. The result led to implementing at Sat level all functions relevant to scientific mission performances, and to relying on another satellite (minisat) for all mission services (Earth link and transfer phase handling for instance). This approach also offers new perspectives for other missions by using a system made of several low-cost Sats in orbit, controlled in position, associated with a minisat (mother ship) with secured architecture offering reliable links with Earth and extended lifetime for mission follow-on. The SMART-2 Sat electrical architecture described in Fig. 8 was built on the Myriade avionic design. The Myriade redundancy philosophy is nominally to embed only redundant functions for the TM/TC communications and some key devices inside the main on board computer (OBC). This philosophy was kept and remained in line with ESA recommendations that did not require redundant units. However, some redundancies were proposed for some critical functions as pseudo-ranging function and inter-communications for TM/TC link (Table 2). 6.2. Power system Table 3 shows the power budget used for the sizing of the battery and the solar array. Using a recurrent Solar array from Myriade provided a margin of 17%. 6.3. Data management
Fig. 7. SMART-2 Sat internal layout.
The computed mass budget shows 144 kg associated to a system margin of 14%. Views of the SMART-2 micro-satellite are presented in Figs. 5–7.
The on board data-handling sub-system was achieved around the central OBC. All the payload units were connected through the standard I/F (RS422 and UART RS422) in order to exchange the telecommand and the telemetry data. Other equipments such as FEEP, S band and UHF units were connected to the RS422 I/F as well. As all data exchanges were identified at lower rate (1 Hz), the centralized architecture around the OBC was fully compatible with the Myriade data-handling capacity. Depending on the detailed real-time constraints of each sub-system (house keeping and
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Fig. 8. SMART-2 Sat electrical architecture.
payload), the detailed allocation of each unit vs each OBC I/F still had to confirm the real-time performances. This architecture also could be simplified as the S band transponder was an option, and as the UHF transponders could be removed if the RF payload unit would be able to support the TM/TC link. 6.4. AOCS The selected star tracker was a conventional medium field-of-view sensor providing autonomously three-axis attitude determination relative to the inertial reference with a better accuracy about the two axes perpendicular to the star tracker line-of-sight. A factor of five geometrical amplification for the attitude estimation error around the line-of-sight was assumed.
Star tracker of type SODERN SED 16, or ASC from the Denmark Technical University, were compliant for the best axes with the ESA specification (10 arcsec @ 3). The preliminary simulations of the SMART-2 ACS/GNC system designed for the Darwin demonstration allowed to achieve the following points: • the complex transition between coarse RF metrology to fringe acquisition is feasible with the available technology; • fine formation-flying and OPD performance requirements are met with margins for the two investigated fringe acquisition and OPD control concepts (with and without delay line); • the proposed architecture with maximum decoupling between mini and micro satellites is satisfactory;
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Table 2 SMART-2 Sat redunded and non-redunded functions Redunded function
Justification
Units
Inter S/C comm’s
In order to prevent collision and distance limit exceed
Redundant UHF RX/TX
RF metrology
Same justification as per inter S/C coms
Redundant RF metrology unit
FEEP
Same justification as per inter S/C coms
Four FEEP embedded and system robust to the lost of one
Non-redunded function
Justification
Emergency comm’s with ground segment
As this function should have been used only following a major failure on the use of the inter S/C link, the need of a redundant transponder was considered to cover double failure which was largely out of scope
Power distribution
Myriade design
Star tracker
Myriade design
Star sensor
Myriade design
OBC
Myriade design
Table 3 SMART-2 Sat power budget Equipment
Optical Metrology RF Metrology DC/DC converter Total Payload Power AOCS FEEP Command & control
Thermal Harness Total
Qty
LEOP
Safe mode
Deploy.
1
Observ.
Reorient.
Included margin (%)
35.1
35.1
15
1
0.00
18.0
18.0
18.0
18.0
20
1
0.0
5.4
5.4
15.9
15.9
na
PCDU Star tracker
1 1
0.00 9.90 0.00 0.00
On-Board processor UHF emitter/receiver S-band transmitter S-band receiver Heaters
1 1 1 1
23.40 9.90 0.00 48.00 5.25
23.40 9.90 8.40 48.00 5.25 3.60
69.00 9.90 8.40 48.00 5.25 3.60
69.00 9.90 8.40 48.00 5.25 3.60
15.00 0.80
6.00 3.36 6.00 0.80
3.00 0.80
3.0 1.7
3.00 0.90
25.70
102.71
102.35
148.85
148.05
10 5 20 5 20 20 20
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• the performance drivers have been identified: coupling of laser metrology sensors (longitudinal and lateral) with the pointing error, N thruster noise and response time, fine longitudinal metrology resolution. • the lateral formation-flying performance (as measured by the fine lateral sensor) is hidden by the impact of Sat pointing stability error. 6.5. Propulsion The SMART-2 requirements matched the N FEEP thrust capacity for all satellite phases, authorizing thus a fully recurring FEEP system. The FEEP thrusters were sized by reconfiguration manoeuvres and orbit corrections rather than by external disturbances (and therefore not by satellite configuration). It was demonstrated that a thrust of 100 N, including significant margins (to account for modelling uncertainties and different manoeuvre profiles) was sufficient in particular with power sizing. 150 N was, however, specified in the ESA phase A study, to take into account the needs for the LISA demonstration, and was kept as baseline. The evaluated Nitrogen mass was safely sized at 5 kg for a 2-year mission. Due to that low value, cold gas thrusters were also seen as good candidates. 7. Conclusion The study carried out by astrium on request of CNES has permitted to rapidly design the mechanical, thermal, electrical architectures, and most of all the avionics, of a micro-satellite meeting the
requirements of a Darwin key technologies demonstrator, in a SMART-2 option made of two formationflying satellites. It was clearly demonstrated that the use of sats for interplanetary missions is of interest, at the cost of a so-called “service satellite” in the range of a minisat, to support in-orbit transfer and TM/TC link with Earth stations. While the SMART-2 second satellite is already scheduled for LISA technology demonstration, this micro-satellite based on the CNES Myriade product line presents a solution optimising the cost for the Darwin features to be demonstrated. The main critical items: RF metrology, optical metrology and FEEP have been studied thoroughly and their implementations in the Sat validated. Finally, this study has proven the feasibility of using recurrent platforms for modular scientific applications with very stringent requirements, such as an inter-satellite distance of 10 m for instance. Such result is encouraging the use of low-cost solutions even in the case of high-valued missions. Further reading [1] L. Vaillon, J. Lebas, E. Sein, B. Calvel, Formation flying system design for the Darwin demonstration of the Smart-2 mission, International Symposium on Formation Flying, Mission & Technologies, October 2002.
Design of a micro-satellite for precise formation flying demonstration Richard Sanchez∗ , Patrice Renard EADS-Astrium, 31 av. des Cosmonautes 31402 Toulouse Cedex 4, France Available online 30 September 2005
Abstract This paper presents a joint CNES-EADS Astrium contribution to the ESA’s SMART-2 project. SMART-2, 2nd mission of the Satellite Missions for Advanced Research and Technology program, slated for launch in 2006, will test key technologies needed to develop two ambitious ESA missions: • LISA (Laser Interferometry Space Antenna), an ESA cornerstone mission dedicated to the detection and observation of gravitational waves; to be launched in 2011, • DARWIN, another ESA cornerstone mission dedicated to the search of Earth-like planets; to be launched in 2015. In Phase A study of this demonstrator, one of the options contemplated by ESA was considering two formation-flying satellites. In that sense, and in order to both reduce and share cost, CNES proposed with the technical support of EADS-Astrium, to build one of them from its Myriade micro-satellite product line, mainly used for LEO scientific applications. The study carried out has permitted to validate the concept of using a low-cost micro-satellite in a scientific interplanetary mission requiring not more than 10 m inter-satellite position accuracy! © 2005 Elsevier Ltd. All rights reserved.
1. Scope of the Sat definition for SMART-2
three major functions:
The study was focused on the Darwin technology demonstration part of the SMART-2 mission, with objective to validate the capability of two satellites to fly with accurate location from each other. To achieve this objective, the proposed micro-satellite had to achieve
• Radio-frequency metrology; • Optical metrology; • Guidance, navigation and control (GNC).
∗ Corresponding author.
E-mail addresses: [email protected] (R. Sanchez), [email protected] (P. Renard). 0094-5765/$ - see front matter © 2005 Elsevier Ltd. All rights reserved. doi:10.1016/j.actaastro.2005.07.039
On one-hand the metrology was providing the GNC with distance measurement within the required accuracy. On the other hand, critical sensor/actuators technologies were to be implemented in the GNC to control such accuracy.
R. Sanchez, P. Renard / Acta Astronautica 59 (2006) 862 – 872
Thus, the major challenges for the Myriade platform were mainly related to • the accommodation of the radio-frequency units and the optical bench of the high precision optical metrology (HPOM), • the management of the coupling effect between the attitude control system and the metrology unit measurements (with end-to-end performances requirement of 10 m inter-satellite position accuracy), • the use of electrical propulsion system (FEEP), • the configuration during orbit transfer (concept of cargo and Sat passenger).
863
Towards optical target
Towards optical target
∆ OPD
2. Darwin mission requirements summary This chapter describes the major Darwin phases. 2.1. Formation deployment This phase covers the operations from separation of the eight Darwin spacecraft from the launcher(s) until a coarse constellation is created with the accuracy of 10 m at inter-satellite distance > 250 m. At the end of this phase, the flyers and the master spacecraft have reached their nominal positions relative to the hub with a 1 cm3 uncertainty; a configuration allowing subsequent laser metrology enabling and fringe acquisition. The modes described hereafter use of RF sensing only for constellation navigation. 2.2. Fringe acquisition The first step after deployment consists in stabilizing the baseline and improving the lateral positioning by using laser metrology (ranging and coarse lateral sensor); in order to enable inter-satellite beam exchanges and to reach the inter-satellite distance (ISD) stability necessary for fringe acquisition. The absolute pointing of the flyer spacecraft is also improved by shifting from the star sensors to the telescope signal. At GNC level, the optical path delay (OPD) is due to two independent errors (RSS combination): • An uncertainty on the inter-satellite distances <1 cm; • A flyer transverse offset relative to the interferometer plane defined by the observation direction
Fig. 1. Correct and tilted constellation configuration (2D example) creating external optical path delay (OPD).
<1 cm. Indeed this offset induces an OPD error (so-called “external OPD”) due to the tilt of the baselines relative to the observation direction (telescopes line-of-sight) as shown in Fig. 1. The uncertainty on OPD obtained from RF sensing and coarse transverse optical metrology is then 1.4 cm. 2.3. Nulling interferometer mission During normal operation, the beams transmitted via the telescope flyers must be co-phased for recombination in the central hub. The OPD between the beams from a pair of telescope flyers must be <20 nm, split in 5 nm allocated to control error and 15 nm to instrument systematic errors. The interfering beams must also have balanced 1 ( 1000 th) intensities, which requires to have beam superimposition on the fringe tracker with the same 1 accuracy ( 1000 th) of the diffraction spot of 1.22 /D, where D is the telescope diameter). This corresponds to 0.08 mas on the detector for 1.5 m telescopes. 2.4. Imaging interferometer mission It is expected that OPD and pointing requirements will be significantly relaxed for imaging (by at least a
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factor of 10). Indeed, conventional OPD requirements for interferometric imaging are in the range of /10 to /100, and beam superposition error only affects 1 1 fringe contrast. Consequently, an accuracy of 10 – 100 of the diffraction spot is sufficient. As a starting point, 1 100 th is assumed. The major difference compared to Nulling is that the interferometer baselines shall be continuously varied (length and orientation in the interferometer plane) during observations to achieve complete scanning of the Fourier plane within a reasonable duration.
3. SMART-2 and Darwin compared requirements The study considered that all technologies required for Darwin would not be tested on SMART-2. Some of them would be verified by ground-based measurements. The flight tests would concentrate on formation deployment, reconfiguration and precision formation flying for the two scientific modes, Nulling and Imaging. Although the analysed SMART-2 option was composed of only two satellites, all the formation flying key technologies required for DARWIN, could, however, be validated. The foreseen sequence of modes was: Orbit transfer: • Transfer phase from Earth up to L1 Lagrange point; Formation deployment: • Stabilization of attitude and on-station control • Initial formation acquisition and collision avoidance quite similar to Darwin; • Coarse baseline control mode using star tracker and RF subsystem measurements: range: 25–250 m; range accuracy: 1 cm; attitude accuracy/inertial frame: 0.5◦ . Nulling interferometer mission: • Acquisition of the 1-cm position accuracy in 3D. With only two satellites this would require line-ofsight measurements with a refined accuracy from a transverse laser metrology system (e.g. a divergent laser associated with a CCD detector).
Fringe acquisition: • Length freezing mode based on laser metrology, down to the accuracy needed for fringe sensor acquisition. • Fine control mode using the fringe sensor. Imaging interferometer mission: • Reorientation and resizing between Nulling mission observations and during Imaging mission observations. Table 1 compares GNC requirements values between Darwin and SMART-2.
4. SMART-2 technology requirements 4.1. RF metrology The RF metrology system is a major technology step forward required for Darwin, the performances of which cannot be assessed on ground because of the difficulty to reproduce the conditions of operation in space (relative dynamics, RF environment). RF metrology shall therefore be validated to the maximum extent by the SMART-2 demonstration. The RF metrology will be used in all Darwin-related modes and for the initial deployment. The “basic hardware set” of the RF subsystem proposed in the study was a pair of “transceivers” (emitter/receiver) and a set of emit/receive antennas flown on the two satellites with the following two functions: • 2-Way pseudo-ranging functions: each transceiver emits a GPS-like signal carrier modulated with a binary code (the navigation message). The opposite transceiver receives this signal, demodulates the binary code and measures its delay relative to an internally-generated code. When multiplied by the speed velocity, this delay provides a pseudorange measurement. Their respective pseudo-range measurements are then exchanged between the transceiver (this is autonomously achieved through the navigation signal) to resolve the clock bias between them. The ISD information is therefore
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Table 1 SMART-2 and Darwin compared GNC performance requirements Mode
Parameter
Darwin requirement
SMART-2 requirement
Deployment mode (coarse optical sensors) Baseline control mode (RF metrology sensors)
Minimum inter-sat. distance Sun direction
20 m 40◦ half-cone from solar array normal 25–250 m, selectable <1 cm & 1 mm/s 3D <1 arcsec & 0.1 mm/s
20 m 45◦ half-cone from solar array normal 25–250 m, selectable 1 cm, az & el err. <0.25◦ (half cone) Use off-the-shelf sun sensors formation roll <0.5◦
400 s 0.1 Hz
400 s 1–10 Hz
4 m at = 1 m 50 m at = 4 m 0.4 m/s at = 1 m 5 m/s at = 4 m ±80 m
5–50 m
Fringe acquisition (optical metrology perfos)
Relative distance Relative positioning Absolute pointing Scan duration Fringe sensor sampling rate Scan step Scan rate
ISD absolute mismatch ISD stability ISD drift OPD stability Global tilt Nulling interferometry (optical metrology perfos)
Imaging interferometry (optical metrology perfos)
Absolute attitude to the flyers Flyer relative angular velocity Relative hub-flyer attitude Relative hub-flyer positions Relative hub-flyer velocity Baseline change rate Duration of imaging periods Resizing
available on both satellites at any time. This function requires dedicated emission/reception RF antennas and relevant transceivers. In order to cover the whole space, two antennas placed in opposite
50 m/s
0–2 cm
2.8 m 0.28 m/s
>5.50 m >50 m
5 nm 0.002 mas @ 250 m (corresponding to 5 nm OPD) 8 mas
±1 m 1◦
0.1 mas/s
Not applicable
60 mas 1 cm 3D 1 mm/s 3D
15 arcsec 0–2 cm ISD az & el: 0.25◦ half cone ISD : 1 mm/s
<3.75 mm/s
<3.75 mm/s
TBD s
TBD s
ISD : 25–250 m and return ISD <1 m in 1 s Duration 16 h
Same as Darwin
10 arcsec
locations need to be implemented on each satellite (Fig. 2); • Transverse position (azimuth/elevation) determination function: this function is in charge of
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LTPter LTP Mas LiLisa sater
Precision Flying Axis (Xsat)
Da rw in Darwin DTµµSat PSaµSat t Flye r FlFlyer ye r
Z Transmit Receive View from X Fig. 2. RF antenna implementation on SMART-2.
determining the direction of the other satellite (defined in S/C frame through elevation and azimuth angles). The principle of so-called “RF goniometry” is to measure the direction of the incoming RF signal using the differential phase of the carrier signal received by three antennas (Fig. 2). The accuracy is directly proportional to the distance between antennas, and the main error source is multipath (reflection of the RF signal on nearby obstacles), so the antennas need to be implemented close to the edges of a satellite wall with minimum obstacles between them. When combined with the measurements of an absolute attitude sensor (e.g. a star tracker) on both satellites, this function allows re-building of the 3D relative motion of one spacecraft relative to the other, in the reference inertial frame. Let us note that having a star tracker on each satellite and RF goniometry on only one end is sufficient to determine the relative position. RF goniometry on both ends would also provide the relative attitude between the S/C, of no interest for SMART-2. Therefore, the “minimum” RF metrology configuration proposed for SMART-2 would implement the RF goniometry function only on the LISA satellite, to maximize possible antenna separation and reduce subsystem mass/power demand on the Darwin satellite. 4.2. Optical metrology SMART-2 had to demonstrate the feasibility of formation flying with accuracy in the nm range. Such
extreme performances call for precise and complex systems. For this purpose, different optical equipments with optimised dynamical range and accuracy had to be accommodated. Two different metrology levels have to be considered in a Darwin type experiment. The internal metrology that freezes the constellation configuration, and the external metrology which ensures the correct pointing of the interferometer onto the guide star. SMART-2 experiment was representative of the first type of metrology. The line-of-sight acquisition and pointing could be done with high accuracy by means of classical equipments such as CCD camera and low power laser source. These equipments were covering the 25–250 m distance range between the two satellites. The inter-satellite distance measurement accuracy was much more demanding and was driving directly the choice of laser metrology technology. The three main parameters sizing the laser were: the length of coherence of the laser source, the 1 m accuracy and the capability to generate different wavelengths for absolute distance measurements. For SMART-2, a low coherence source associated with two fixed delay lines with lengths of 25 and 250 m to cover the ISD range were selected (Fig. 3). The Sat metrology hardware was then able to provide the following accuracies: • Coarse lateral position of the two satellites, approximately 1 mm; • Fine lateral and longitudinal distance between the two satellites, 10 m; • Fringe position in the interferometer defined by an internal fiber reference and the counter satellite, 1 m. It was also possible to develop a simple fringe-tracking sensor, but this approach was found to be less representative. 4.3. Micro-Newton propulsion system The micro-thrusters used for drag-free and attitude control during the LISA demonstration could also be used for accurate pointing and relative position of the two satellites of the Darwin demonstration. As a consequence, the N-thrusters on the Sat had to
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Fringe Fr Friinnggee Tracker Tr ke r Tr ac acker White iteLight Li ght source sour ce
ReLe fere Reference ngnc th e Length Re fere nc e Le ngt h Delay Line
Lo ngi ng itud Longitudinal tudiinnaall Sensor Se ns Se ns or or (10 µm) (1 0 µm µm)) (10
La serSource So ur ce Laser ser S our ce
UHF Link
Fine Lateral Sensor (10 µm)
Co ar se Lateral La ra ll Coarse Co ar se La te tera Sensor Se Se ns nsor or
Collimated Laser
DARWIN µsat
LISA mini-satellite
Fig. 3. Optical metrology implementation on SMART-2.
102
Smart-2 specification Darwin specification Solar Pressure qthrusters=1 qthrusters=0.1 (LISA) qthrusters=0.01
100
ISD error (µm)
provide a 6-DOF control capability with tolerance to the failure of at least one thruster (or one cluster if arranged in clusters). The required authority was in the range of 100 N. Another challenge was the capability of the Sat to support long consumption from FEEP and the use of this electrical propulsion as unique propulsion module for all modes, in particular on station setting after the orbit transfer. The thrust noise was the driving factor as opposed to thrust vector misalignment, bias and scale factor. In order to derive thruster noise requirements, a preliminary allocation of the ISD error contributors was considered assuming equal quadratic weighting:
10-2
10-4 10-6 10-8 10-3
10-2
10-1
100
101
Controller cut-off frequency (Hz)
ISD = ISDsensors + ISDsolar_pressure + ISDthrusters √ 1/ 2
1/2
√ 1/ 2
1/2
Fig. 4. Thruster noise impact on the inter-satellite distance (ISD).
(1) The thruster noise transmission in the GNC loop was evaluated as a function of the control bandwidth.
Fig. 4 compares OPD requirements for SMART-2 and for Darwin (corrected by the SMART-2 and Darwin masses ratio, and assuming the same error allocation for Darwin), and the OPD caused by solar pressure
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Fig. 6. SMART-2 Sat in orbit configuration.
• the dominant disturbance is the thruster noise con√ sidering requirements of 0.1 N/ Hz, as for the LISA demonstration. The resulting control bandwidth of about 0.01 Hz (about 0.5 Hz for Darwin) is significantly higher but still compatible with the thruster technology. 5. SMART-2 Sat configuration The mechanical architecture coped with the Myriade structure to enhance the re-use and associated cost gain. It is worth noticing that the design was defined with:
Fig. 5. SMART-2 Sat in stowed configuration.
√ and various thrust noise levels (specified in N/ Hz) as functions of the controller cut-off frequency. The major conclusions were: • the rejection of the solar pressure can be achieved with very low control bandwidth: <0.001 Hz (about 0.05 Hz for Darwin);
• two batteries; • two UHF transponders and relevant antennas (optional); • two RF metrology units and relevant antennas (with possibility to have them internally redunded to provide additional mass margin). On the other hand, it was assumed that the same antenna was used for signal transmission and reception, and that only one antenna was required for the Darwin satellite (baseline for pseudo-ranging function only); finally, two identical antennas were embedded to perform omni coverage; • one emergency transponder and one antenna (optional).
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6. SMART-2 Sat functional description 6.1. Overview The main challenge was to deal with a Sat concept together with highly challenging performances. The result led to implementing at Sat level all functions relevant to scientific mission performances, and to relying on another satellite (minisat) for all mission services (Earth link and transfer phase handling for instance). This approach also offers new perspectives for other missions by using a system made of several low-cost Sats in orbit, controlled in position, associated with a minisat (mother ship) with secured architecture offering reliable links with Earth and extended lifetime for mission follow-on. The SMART-2 Sat electrical architecture described in Fig. 8 was built on the Myriade avionic design. The Myriade redundancy philosophy is nominally to embed only redundant functions for the TM/TC communications and some key devices inside the main on board computer (OBC). This philosophy was kept and remained in line with ESA recommendations that did not require redundant units. However, some redundancies were proposed for some critical functions as pseudo-ranging function and inter-communications for TM/TC link (Table 2). 6.2. Power system Table 3 shows the power budget used for the sizing of the battery and the solar array. Using a recurrent Solar array from Myriade provided a margin of 17%. 6.3. Data management
Fig. 7. SMART-2 Sat internal layout.
The computed mass budget shows 144 kg associated to a system margin of 14%. Views of the SMART-2 micro-satellite are presented in Figs. 5–7.
The on board data-handling sub-system was achieved around the central OBC. All the payload units were connected through the standard I/F (RS422 and UART RS422) in order to exchange the telecommand and the telemetry data. Other equipments such as FEEP, S band and UHF units were connected to the RS422 I/F as well. As all data exchanges were identified at lower rate (1 Hz), the centralized architecture around the OBC was fully compatible with the Myriade data-handling capacity. Depending on the detailed real-time constraints of each sub-system (house keeping and
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Fig. 8. SMART-2 Sat electrical architecture.
payload), the detailed allocation of each unit vs each OBC I/F still had to confirm the real-time performances. This architecture also could be simplified as the S band transponder was an option, and as the UHF transponders could be removed if the RF payload unit would be able to support the TM/TC link. 6.4. AOCS The selected star tracker was a conventional medium field-of-view sensor providing autonomously three-axis attitude determination relative to the inertial reference with a better accuracy about the two axes perpendicular to the star tracker line-of-sight. A factor of five geometrical amplification for the attitude estimation error around the line-of-sight was assumed.
Star tracker of type SODERN SED 16, or ASC from the Denmark Technical University, were compliant for the best axes with the ESA specification (10 arcsec @ 3). The preliminary simulations of the SMART-2 ACS/GNC system designed for the Darwin demonstration allowed to achieve the following points: • the complex transition between coarse RF metrology to fringe acquisition is feasible with the available technology; • fine formation-flying and OPD performance requirements are met with margins for the two investigated fringe acquisition and OPD control concepts (with and without delay line); • the proposed architecture with maximum decoupling between mini and micro satellites is satisfactory;
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Table 2 SMART-2 Sat redunded and non-redunded functions Redunded function
Justification
Units
Inter S/C comm’s
In order to prevent collision and distance limit exceed
Redundant UHF RX/TX
RF metrology
Same justification as per inter S/C coms
Redundant RF metrology unit
FEEP
Same justification as per inter S/C coms
Four FEEP embedded and system robust to the lost of one
Non-redunded function
Justification
Emergency comm’s with ground segment
As this function should have been used only following a major failure on the use of the inter S/C link, the need of a redundant transponder was considered to cover double failure which was largely out of scope
Power distribution
Myriade design
Star tracker
Myriade design
Star sensor
Myriade design
OBC
Myriade design
Table 3 SMART-2 Sat power budget Equipment
Optical Metrology RF Metrology DC/DC converter Total Payload Power AOCS FEEP Command & control
Thermal Harness Total
Qty
LEOP
Safe mode
Deploy.
1
Observ.
Reorient.
Included margin (%)
35.1
35.1
15
1
0.00
18.0
18.0
18.0
18.0
20
1
0.0
5.4
5.4
15.9
15.9
na
PCDU Star tracker
1 1
0.00 9.90 0.00 0.00
On-Board processor UHF emitter/receiver S-band transmitter S-band receiver Heaters
1 1 1 1
23.40 9.90 0.00 48.00 5.25
23.40 9.90 8.40 48.00 5.25 3.60
69.00 9.90 8.40 48.00 5.25 3.60
69.00 9.90 8.40 48.00 5.25 3.60
15.00 0.80
6.00 3.36 6.00 0.80
3.00 0.80
3.0 1.7
3.00 0.90
25.70
102.71
102.35
148.85
148.05
10 5 20 5 20 20 20
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• the performance drivers have been identified: coupling of laser metrology sensors (longitudinal and lateral) with the pointing error, N thruster noise and response time, fine longitudinal metrology resolution. • the lateral formation-flying performance (as measured by the fine lateral sensor) is hidden by the impact of Sat pointing stability error. 6.5. Propulsion The SMART-2 requirements matched the N FEEP thrust capacity for all satellite phases, authorizing thus a fully recurring FEEP system. The FEEP thrusters were sized by reconfiguration manoeuvres and orbit corrections rather than by external disturbances (and therefore not by satellite configuration). It was demonstrated that a thrust of 100 N, including significant margins (to account for modelling uncertainties and different manoeuvre profiles) was sufficient in particular with power sizing. 150 N was, however, specified in the ESA phase A study, to take into account the needs for the LISA demonstration, and was kept as baseline. The evaluated Nitrogen mass was safely sized at 5 kg for a 2-year mission. Due to that low value, cold gas thrusters were also seen as good candidates. 7. Conclusion The study carried out by astrium on request of CNES has permitted to rapidly design the mechanical, thermal, electrical architectures, and most of all the avionics, of a micro-satellite meeting the
requirements of a Darwin key technologies demonstrator, in a SMART-2 option made of two formationflying satellites. It was clearly demonstrated that the use of sats for interplanetary missions is of interest, at the cost of a so-called “service satellite” in the range of a minisat, to support in-orbit transfer and TM/TC link with Earth stations. While the SMART-2 second satellite is already scheduled for LISA technology demonstration, this micro-satellite based on the CNES Myriade product line presents a solution optimising the cost for the Darwin features to be demonstrated. The main critical items: RF metrology, optical metrology and FEEP have been studied thoroughly and their implementations in the Sat validated. Finally, this study has proven the feasibility of using recurrent platforms for modular scientific applications with very stringent requirements, such as an inter-satellite distance of 10 m for instance. Such result is encouraging the use of low-cost solutions even in the case of high-valued missions. Further reading [1] L. Vaillon, J. Lebas, E. Sein, B. Calvel, Formation flying system design for the Darwin demonstration of the Smart-2 mission, International Symposium on Formation Flying, Mission & Technologies, October 2002.