AIAA 2007-6547
AIAA Guidance, Navigation and Control Conference and Exhibit 20 - 23 August 2007, Hilton Head, South Carolina
MEMS Rate Sensors in Space Becomes a Reality Dick Durrant* Systems Engineering & Assessment Ltd, Bristol – United Kingdom Stéphane Dussy† European Space Agency, Noordwijk – The Netherlands Brian Shackleton‡ SELEX Sensors and Airborne Systems Ltd, Edinburgh. – United Kingdom and Alan Malvern§ BAE SYSTEMS, Plymouth – United Kingdom
With ongoing European developments demonstrating the viability of MEMS based rate sensor units for space applications the application of this technology to planned missions is becoming a reality. It is therefore opportune to address the key benefits of the MEMS technology to space missions. This paper provides a user view of the MEMS Rate Sensor currently under development (SiREUS), with emphasis on the key features and performance (ie. what is such a unit good for and what does it not address). Based on this characteristics definition case studies are presented discussing the application of the SiREUS to three classes of real-world applications. These cover initial acquisition and fault monitoring/management, fault-monitoring/management of a planetary mission, and a geostationary station keeping application.
Nomenclature AAF AOCS ARW AV DPM DTG EPM FDI FOG GEO GMES GNC GTO
= = = = = = = = = = = = =
Anti Aliasing Filter Attitude and Orbit Control System Angular Random Walk Allan Variance Drift and Positioning Mode Dynamically Tuned Gyro Earth Pointing Mode Failure Detection and Isolation Fibre-Optic Gyro Geostationary Earth Orbit Global Monitoring for Environment and Security Guidance Navigation and Control Geostationary Transfer Orbit
IAAM ICD LEO LEOP MEMS NM ORM RBD SAM SBM SF STR
= = = = = = = = = = = =
Inertial Attitude Acquisition Mode Interface Control Document Low Earth Orbit Launch and Early Orbit Phase Micro Electro-Mechanical Systems Nominal Mode Orbit Raising Mode Rate Bias Drift Sun Acquisition Mode Stand-By Mode Scale Factor Star Tracker
*
SEA, Bristol Business Park, Coldharbour Lane, Bristol BS16 1SU, United Kingdom.
[email protected] ESA, Postbox 299, 2200 AG Noordwijk, The Netherlands.
[email protected] ‡ SELEX S&AS, Crewe Toll, Ferry Road, Edinburgh, EH5 2XS, United Kingdom.
[email protected] § BAE SYSTEMS, Clittaford Road, Plymouth, PL6 6DE, United Kingdom.
[email protected] †
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Copyright © 2007 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
I. Introduction The application of radiation hardened MEMS based rate sensors to space applications will become a reality over the next 2 years and will become increasingly prevalent thereafter. While such sensors are unlikely to achieve the performance requirements for high pointing accuracy science and remote sensing missions it is envisaged that early implementations will be immediately applicable to failure monitoring/failure management applications and, through continuing development, the application area will broaden into providing the necessary performance for the traditional range of telecoms missions and the expanding low cost smallsat market. Before the application of a MEMS Rate Sensor can be discussed it is first necessary to understand the key characteristics of such a unit and what application areas such a unit is good for in general terms. Within this paper the key characteristics are based on the ongoing European Silicon Rate Sensor development (SiREUS) as the leading example of the migration of this terrestrial technology into the space domain.
II. MEMS Rate Sensor Characteristics A. General For AOCS/GNC designers who have traditionally used mechanical gyros or emerging high accuracy solid-state gyros (eg. FOG), it is important to understand initially the general characteristics of a MEMS based sensor. The SiREUS is a medium class gyro (as discussed later in this paper) which may not challenge other gyros in respect of absolute performance, but provides a simpler and more cost effective solution in spacecraft design by comparison with many other classes of gyros. The characteristics which distinguish the sensor are as follows: Low Mass/Volume This significantly simplifies integration of the sensor compared to all other classes of gyro. Many such gyros have traditionally been two unit solutions with different (mechanical, thermal and electrical) interfacing characteristics for each unit. This complication can lead to long interconnects and EMC/EMI sensitivity which complicate the spacecraft design yet further. A more compact single unit sensor is therefore attractive in traditional satellite applications as well as within the increasing range of small/micro satellites. It also positions the sensor as a candidate sensor for lateral applications in the space robotics fields. Low Power A 3-axis MEMS based rate sensor requires approximately 4W from a 28V supply. This level of power means that in many satellite applications it can be left on continuously thereby simplifying the AOCS/GNC architecture. Previous consideration of gyros have required a trade-off between data availability and power consumption particularly during power up, not withstanding the added complication of drift settling or drift calibration before use. Continuous operation improves both overall effectiveness (eg. for fault monitoring and management) and also ensures that no reliability degradation is introduced by excessive power cycling. Low Cost of Ownership The inherent low cost of the MEMS based sensor provides an additional tool for AOCS/GNC designers in terms of permitting a cost effective approach to satellite fault monitoring and management. The need for exhaustive system modeling as a part of the non recurring engineering activity pre-launch is removed by the on board availability of vehicle angular rate parameters. There are also a number of inherent technical characteristics which the MEMS rate sensor can offer by way of advantages in the design of spacecraft AOCS.GNC: High Robustness/High Shock Tolerance Since there are no moving parts within the small and compact structure, there is an inherent robustness in the MEMS based sensor and there is high shock tolerance. This eases the choice of location within the spacecraft with no additional provision needed to attenuate the shock levels likely to be encountered at vehicle launch, stage separation or spacecraft/launcher separation. The MEMS technology heritage in missile and automotive applications demonstrates this inherent robustness. High immunity to out of plane vibration The inherent high tolerance to out of plane vibration of the vibrating ring technology within the MEMS rate sensor further eases choice of sensor location, as well as simplifying sensor assessment for particular missions by not requiring any specific calibration or correction. These core capabilities are already starting to attract significant attention for spacecraft designers and will be discussed later in the paper in a set of mini case studies for specific classes of applications.
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B. SiREUS User Perspective Prior to discussing the applications areas for a MEMS rate sensor in more detail it is useful to understand the user perspective the operation and perspective of such a unit. For the SiREUS unit the user perspective can be considered in four areas: • Mechanical Interface • Electrical Interface • Operational Interface System Interface • The first two are addressed by the unit ICD, and short form version of this can be provided to potential users in the form of an information sheet. This section provides an overview of the operational and system interfaces, where the user perspective is provided by a simplified functional and data model of the unit; this is shown in Figure II-1. Rate information is produced simultaneously about each of the three unit axes at the resonant frequency of the detector (nominally 14 KHz). This (noisy) rate information is filtered down to the required output bandwidth using a programmable rolling average filter. This filtering provides a significant noise improvement as well as performing an Anti-Aliasing Filter (AAF) function that ensures subsequent decimation to the final rate output bandwidth does not introduce noise from higher frequency products. Rolling average filter set to a bandwidth of A1, A2 or A3 Hz Filter Rate output frequency set to F1, F2, F3, or F4 Hz
User set output bandwidth to B1, B2, B3 or B4 Hz
Bit Shift Bit Shift
19 bit words at m Hz
Bit Shift
19 bit words at 14 KHz
19 bit words at m Hz
b bit words at m Hz
b bit words at m Hz
TM Output
Analogue Rate X output
Analogue Rate Y output
AROZ
16 bit words at 14 KHz
Telemetry Formatter
Downsample
19 bit words at 14 KHz
b bit words at m Hz
AROY
Z-axis detector Rate
16 bit words at 14 KHz
19 bit words at m Hz
TC Input
AROX
Y-axis detector Rate
19 bit words at 14 KHz
16 bit words at 14 KHz
Downsample
X-axis detector Rate
Downsample
User set rate resolution to 0.28, 2.2 or 17.6 arcsec/sec
Analogue Rate Z output
Figure II-1 SiREUS Functional and Data Model The user can set the SIREUS output bandwidth to one of 4 values, the required value of rolling average filter and output clock being selected automatically within the unit according to the values defined in Table II-1. User Set Output Bandwidth B1 1 Hz B2 2 Hz B3 5 Hz B4 10 Hz
Output Sample Rate (TM Message) 2 Messages/sec 4 Messages/sec 10 Messages/sec 20 Messages/sec
Table II-1 SiREUS Output Bandwidth Configuration Table The output sample rate is therefore synchronous across the three axes with a timing resolution determined by the detector resonant frequencies; worst case error is 2.55 µS. The output sample rate is set for the SiREUS unit and applied simultaneously to both the digital outputs (ie. the Telemetry Messages) and the Analogue Rate Outputs.
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The digital rate outputs have a resolution (1 LSB) of 0.275 deg/hr and a (19 bit) range of +/- 20 deg/s. The analogue rates also have a range of +/-20 deg/s with a resolution of 2.2 deg/hr or 141 deg/hr dependent on the applied rate to the particular axis. From an operational viewpoint the user interacts via defined Telecommand Messages to the unit and the reception of related Telemetry Messages from the unit. The format of these is well defined in the ICD and the baseline low level layers utilize either RS422 channels. The SiREUS unit has two user modes, OFF and Rate Mode, with the transitions as defined in Figure II-2 . At power-up the MRS will automatically perform the following initialisation operations by transferring all the detector parameters from the non-volatile (EEPROM) store into a protected SRAM within the FPGA/ASIC where they are made available to the detector control algorithms for the three detectors. The detector control algorithms are then initialized and the unit moves into Rate Mode.
OFF Power-Up Initialisation
Power-Down
Rate Mode
Fault Detected
Set BITE Status word
Configuration Command (HK Format) Configuration Command (Rate Bandwidth)
Update Output Format
Update Output Bandwidth
Calibration Commands Update Calibration Data Trim Command(ON) TMEN Trim Mode Trim Command(OFF)
Figure II-2 MRS Operational State Diagram In Rate Mode SiREUS outputs 3-axis rate values at the baseline output rates of 5 Hz for default messages containing rate and summary status for each axes. The output bandwidth can be changed by the user as discussed above and the status content of the HouseKeeping (HK) section of the message can be controlled by the user via Telecommand messages. Additional commands are provided to support the on-ground calibration and trimming of the MRS detector DC and AC control co-efficient values. These are shown in red in Figure II-2 and are not normally part of the user operations. However on-board access to all the detector parameters is possible, and can be used to support extended diagnostics and parameter updates.
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C. SiREUS User Model and Performance In order to support user system modeling a simple gyro model is under development of the measured angular rate values provided to the spacecraft AOCS/GNC computer referenced to the actual rate input measured along the sensitive axis; this model is summarized in Figure II-3. As discussed in §0 the basic rate information is generated at approx. 14 KHz and this is the starting point for the noise model; represented here by a 14 KHz white noise generator. Subsequent filtering permits the modeling of the white noise on the rate measurement and the angular white noise. While the SiREUS unit included temperature compensation algorithms to correct the movement in the rate bias with temperature there will be a residual temperature dependence that cannot be corrected. This is represented in the model as a temperature dependent function derived from a best fit to the unit test data. The unit scale factor is a simple multiplicative factor in the current model.
1
1
Rate
Scale Factor
1
1
s+1
z+0.5
White Noise Noise shaping @14Khz Rate
2 Temperature
Rolling Average
1 Decimation to Output Rate
Quantizer 19bit signed
Measured Rate
f(u) Bias Factor Add
Figure II-3. Simple Rate Sensor Model The first SiREUS complete unit is currently in manufacture and will be tested during September/October’07, however a good indication of the anticipated performance can be obtained from the prototype detector testing with breadboard implementations of the flight electronics designs. From this testing the Allan Variance plot providing an indication of the performance of the overall sensor in terms of Noise Equivalent rate (Allan Variance minimum) and Angular Random Walk (-0.5 slope of the AV plot). An example Allan Variance is shown in Figure II-4. Here it can be seen that the AV minimum equates to a Noise Equivalent Rate of less than 1 deg/hr (typically 0.8 deg/hr) with this occurring at integration times of about 100 seconds. This gives users a good indication that for typical applications where the AOCS computer will be working with an integrated rate value good system level performance can be achieved. Also from the typical plot shown the ARW can be derived to be about 0.045 deg/rt(hour). This level of performance places the SiREUS sensor in the range of applications covering initial acquisition, fault monitoring and fault management as discussed in previous paper3.
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Figure II-4. ARW Plot for Prototype Detectors/Electronics Telemetry Sample Rate: 10 HzRate Measurement: deg/hr Temperature Variation: approx 1 degC While the AV plot provides a good indication of performance the AOCS/GNC user generally will require a PSD plot or equivalent data in order to derive a sensor model, as discussed above, for the overall AOCS/GNC simulation. A typical SiREUS PSD plot is shown in Figure II-5 and indicates a flat noise response down to very low frequencies. The change in bias in the PSD at low frequencies is compensated within the SiREUS processing.
Figure II-5 PSD Plot for Prototype Detectors/Electronics Telemetry Sample Rate (Fs): 10 Hz PSD units: (deg/sec)2/Hz Single Sided PSD Temperature Variation: +/- 10 degC Finally rate error testing over the range of +20/-20 deg/sec provides a good level of confidence that the specified requirements of +/- 5000 ppm can be met and the goal of +/-1000 ppm should be achievable; a typical scale factor test plot is shown in Figure II-6.
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Rate Error Plot for Detector: #124 400
300
Rate Error (deg/hr)
200
-20
100
0 -15
-10
-5
0
5
10
15
20
-100
-200
-300 Rate error
+5000 ppm
-5000 ppm
-400 Rate Demand (deg/sec)
Figure II-6 Typical Scale Factor Plot for Prototype Detector/Electronics Additional test results are of interest to the SiREUS user, including: Switch-on to Switch-on variation in bias is less than 8 deg/hour. Rate Bias Drift of the order of 10 deg/hr for a variation in temperature of +/- 10 degC over a 24 hour period. These results, together with those from the current flight experiment ground testing, are being used to determine the parameters for the user gyro model described above. Overall the test results confirm the SiREUS are being a 10 deg/hr class of rate sensor that can be considered for a broad range of application areas as discussed in §III.
III. MEMS Rate Sensor Applications A. Applications Overview This section defines the approach to use the SiREUS unit within different classes of mission with representative near-term examples. This is presented in the form of a number of Case Study abstracts to illustrate the key benefits of the developed technology. These cover initial acquisition and fault monitoring/management of an Earth observation satellite, fault-monitoring/management of a planetary mission, and a geostationary station keeping application. For two of these the use of a MEMS gyro has either been baselined or is under consideration on near term missions. B. Case Study #1: Planetary Mission Failure Detection and Isolation (eg. Bepi-Colombo) Planetary science missions are often very demanding in terms of Failure, Detection and Isolation (FDI) because of the limited communications available, in particular the time lag hindering real time diagnostics. In such an environment a MEMS based sensor is particularly attractive offering very low power (continuous operation) with a performance level suitable for supporting spacecraft AOCS/GNC fault management. A very good example, where a MEMS Rate Sensor has been considered, is Bepi-Colombo. This is an ESA mission that will explore Mercury, the planet closest to the Sun. Europe's space scientists have identified the mission as one of the most challenging long-term planetary projects, largely because Mercury's orbit is so close to the Sun this makes the planet difficult for a spacecraft to reach and difficult to observe from a distance. Scientists want to study Mercury because of the valuable clues it will provide in understanding how planets form. Spacecraft launch is planned for 2013 with a 6 year cruise phase and 1 year in orbit around Mercury.
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Figure III-1
Bepi-Colombo mission
The use of SiREUS for a fault-monitoring and failure management role has been investigated. As such it would be operational during all critical periods of the mission and would determine whether the prime spacecraft control system is holding the spacecraft in the correct orientation. In this case study, the SiREUS unit would therefore be part of the spacecraft FDI sub-system, supporting rate damping and sun-acquisition in case an error in the prime operation is detected. C. Case Study #2: LEO acquisition and safe mode CryoSat is the first ESA satellite dedicated to the observation of the changes in the polar ice masses. Its three-axis stabilised local-normal pointing, with nose-down attitude and yaw-steering, relies on a challenging gyroless AOCS, based on star trackers and magnetometers for attitude determination. Precise orbit determination is obtained with a Doppler Orbit and Radio Positioning instrument (DORIS) that detects and measures the Doppler shift of signal broadcasted from a network of 50 radio beacons stations spread around the world. In addition, a set of coarse Earth and Sun sensors allows handling first attitude acquisition and safe mode. The demanding requirement on propellant consumption target is met thanks to a combined magnetic torque rods and cold-gas thrusters actuation.
Figure III-2
CryoSat-1 satellite during AIT procedures
After the launch of the CryoSat spacecraft was unfortunately aborted on 8 October 2005 due to a malfunction of its Rockot launcher, which resulted in the total loss of the spacecraft, it was decided to build and launch a CryoSat recovery mission, CryoSat-2, with limited redesign. One of the main modifications was the decision to embark a 8 American Institute of Aeronautics and Astronautics
MEMS gyro (SiREUS prototype) as a technology experiment in order to support the LEOP operations by providing the ground segment with angular rate measurements during the crucial phases of rate damping and coarse pointing mode. For the example of CryoSat, MEMS gyro acts as a monitoring device and allows remedying the well-known limitations of magnetometer-based rate measurements.
Figure III-3
CryoSat-2 mission
Previous gyro technologies, such as Dynamically Tuned Gyros (DTG), were considered as expensive and power consuming equipment. In addition such approaches can generate vibrations impacting the spacecraft and exhibit a poor reliability history, thus forcing a number of LEO satellites to adopt a gyroless concept. The MEMS Gyro technology allows does not suffer from these drawbacks, offering a reliable and affordable solution (with no moving parts) for low-cost LEO satellites. This is especially important during rate damping and first (sun or Earth) acquisition maneuvers, or during safe mode maneuvers. Such concepts are considered as promising concepts for other LEO satellites such as the Sentinel-3 mission. Sentinel-3 mission is currently developed under the European Union – ESA GMES (Global Monitoring for Environment and Security) programme, for a launch expected in 2011/2012. The mission is devoted to provision of ocean, ice and land surface monitoring services.
Figure III-4
Sentinel-3 mission for ocean, ice and land monitoring
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The Sentinel-3 satellite shown in Figure III-4 will operate in a sun-synchronous orbit at 814 km altitude. The current preliminary design identified the need for a package of three-axis coarse rate sensors, in the performance range of the MEMS gyro SiREUS. Such rate sensors are meant to be used during the initial rate dumping, until the sun acquisition and pointing, with the task to provide telemetry of the dynamics during deployment of the solar array. In addition they are meant to support slew maneuver for transition between modes and to ensure continuity of the mission during orbital maneuver in case of blinding of stellar sensors.
D. Case Study #3: GTO and GEO safe mode (eg. SpaceBus 4000) The SpaceBus 4000 AOCS is a three-axis body stabilization system for all mission phases, based on a simple and reliable four-wheel system providing high angular momentum storage capacity and continuous 3-axis control in the geostationary configuration. This AOCS has been developed by Thales Alenia Space since 1998, with a successful first in-flight use on Worldsat2 spacecraft, in February 2005 and a first ITAR-free version on APSTARVI spacecraft, also in orbit since April 2005.
Figure III-5
SpaceBus 4000C2 type Telecom Satellite
As detailed in a previous paper4, the SpaceBus 4000 AOCS mode logic is organized in functional modes. The Stand-By Mode (SBM) is the primary mode engaged after launcher separation. The Sun Acquisition Mode (SAM) is a mode which provides safe conditions to the platform in terms of power and thermal control. The Inertial Attitude Acquisition Mode (IAAM) is used in transfer orbit to perform the Star tracker (STR) acquisition, gyro drift calibration and reorientation towards the boost attitude to prepare for the Orbit Raising Mode (ORM). It is also the transition mode towards the Earth Pointing Mode (EPM) to reach and keep Earth pointing attitude. ORM allows for optimized orbit raising strategy through adequate apogee maneuvers. The Earth Pointing Mode (EPM), the Normal Mode (NM) and the Drift and Positioning Mode (DPM) are geostationary orbit operational modes that use gyro measurements only in back-up cases. A recent study3,4 demonstrated that the MEMS Gyro could replace the high performance gyro used on-board SpaceBus 4000, at the price of the following acceptable system modifications. (1) The target spin rate range in SAM [0.1°/s ; 0.75°s] shall be reduced to the upper part [0.5°/s ; 0.75°/s] in order to limit the nutation movement induced by the gyro drift. (2) The 0.3° pointing accuracy specification for the Attitude Keeping in IAAM during Van Allen belts crossing must be relaxed and adapted to each gyro model in case of control on gyros only. In fact, in case of loss of STR tracking, the attitude will be determined using gyros only and thus will drift with the gyro bias; but once the STR is operational again, the de-pointing is corrected within an acceptable duration, whatever the gyro model. (3) A new attitude determination function, with a hybridization of both STR and gyro data through a Kalman filter, shall be implemented, for a gyro utilization in Normal Mode. 10 American Institute of Aeronautics and Astronautics
(4) At AOCS level, only a few re-tunings of the digital modulators shall be performed, and noise rejection filters shall be implemented. Moreover, an integral term shall be added to the IAAM controller in order to cancel the static error, induced by the drift of the MEMS Gyro. It was also shown that the main difference between various gyro classes can be seen on the maximum durations of gyro use in order to respect the specified pointing accuracy, but the study4 demonstrated the applicability of the use of a medium class gyro (around 10 deg/hr) such as the emerging SiREUS unit. E. Further Development and Future Applications Generally the development of the SiREUS has migrated MEMS technology from the aerospace sector, while taking into account the detailed environmental requirements for space mission for the both the detector and the associated electronics (eg. Radiation hardening through component selection, radiation testing of the MEMS JFET amplifiers, shielding, temperature compensation etc). The results of this development have been extremely encouraging leading to the current implementation of a Flight Experiment a 10 deg/hr sensor, and unit qualification in place for completion by mid-2008. With the experience derived during the SiREUS development it is clear that further development at the detector level, with some associated electronics updates, could potentially enable a broader range of application areas to be addressed. In this respect it is envisaged that the development of technology for space MEMS gyroscope will continue to improve upon key features of the sensor. Further development can be envisaged in two areas. Evolutionary developments are envisaged based heavily on the experience to date with the development of the detectors for the SiREUS unit and can lead to incremental improvements in performance. For example charge trapping is one effect which causes a long term scale factor change. Potential means of eliminating this effect principally relate to MEMS processing improvements. Elimination of this effect will enable a scale factor accuracy of
