Executive Summary. SENER document number: SEN-APOD-EXECSUM Issue: 1 Revision:

Transcription

1 Autonomous Precise Orbit and Attitude Determination System (APOD) Executive Summary Written ESA T.O. S. de Miguel G. Ortega Name Signature Date SENER document number: SEN-APOD-EXECSUM Issue: 1 Revision: B Date: 30 June 2003 File: APOD_Executive_Summary_issue1_revB.doc Number of Pages: 39 EUROPEAN SPACE AGENCY CONTRACT REPORT The work described in this report was done under ESA contract. Responsibility for the contents resides in the author or organisation that prepared it. The copyright in this document is vested in SENER Ingenieria y Sistemas. This document may only be reproduced in whole or in part, or stored in a retrieval system, or transmitted in any form, or by any means electronic, mechanical, photocopying or otherwise, either with the prior permission of SENER Ingeniería y Sistemas or in accordance with the terms of ESTEC/Contract no 15667/01/NL/LvH

2 Page: 2 of 38 LIST OF ISSUES Issue Rev Date Section Change 1-20 June 2003 All Initial issue A 25 June All Minor changes B 30 June All Minor changes

3 Page: 3 of 38 ABSTRACT This document is the executive summary for the Autonomous Precise Orbit and Attitude Determination Unit (APOD) development project. This work has been done by SENER for the ESA General Support Technology Program in the frame of a GSTP-3 contract. The activity concerns the design, development, integration, manufacturing, test, and evaluation, of a model unit in charge of providing fully autonomous capability to establish orbit and attitude determination for a spacecraft. The APOD development was structured in two phases. Phase 1 dealt with the specification, analysis and design of the FM model. Phase 1 concluded with the definition of a DM model to be manufactured in phase 2 for demonstration purposes. Phase 2 covered the DM chassis development, the procurement of the hardware and the manufacturing of the DM unit. The navigation software was developed and validated in phase 2. This document presents an overview of the proposed FM design and the details of the DM model manufactured. The software development cycle and some final validation results are also shown. In software the main effort was dedicated to the attitude determination algorithms, to design and code the VxWorks real-time program, and to implement the MIL-STD-1553 interface for TM/TC. In hardware the main effort was dedicated to the DM model manufacturing. This included the adaptation of available hardware, and the production of a power supply VME board using COTS. The DM box is a rugged unit that is mechanized in Aluminum and specifically designed to integrate the APOD electronics. An excerpt of the obtained performance results is included herein to show the good functioning of the on-board software when running on a high fidelity non-real time Matlab/Simulink simulator. A practical and reliable multipath model has been developed. Finally, the on-board software was tested with the APOD hardware in the loop. The real-time test bed uses dspace for simulation of the environment (e.g. GPS phase measurements). The main conclusion is that the developed extended Kalman filter for attitude determination combining GPS and a gyro inertial measurement unit works very well. Furthermore, the functional validation of the APOD real-time software on the target hardware is achieved. The APOD phase A development has finished on a very good starting point to initiate the future development of a qualification model.

4 Page: 4 of 38 TABLE OF CONTENTS TABLE OF FIGURES INTRODUCTION CONTEXT DELIVERABLE ITEMS APOD SYSTEM OVERVIEW APOD GENERAL DESCRIPTION OPERATIONAL ENVIRONMENT HARDWARE DESIGN HARDWARE ITEMS IDENTIFICATION IMU sub-unit GPS sub-unit Processor sub-unit I/O sub-unit Power supply Antennae subsystem SYSTEM ARCHITECTURE Processing and communications Internal connexions and cabling OPERATIONAL SCENARIOS APOD modes ALGORITHM DESIGN ORBIT DETERMINATION ATTITUDE DETERMINATION Attitude determination algorithms Kalman filter design TEST PROGRAM DELIVERABLE MODELS Development Model Engineering Qualification Model Proto Flight Model Flight Models FLIGHT MODEL PHYSICAL CONFIGURATION FM SYSTEM PERFORMANCES FM EXTERNAL INTERFACES Power DEVELOPMENT MODEL DM DESCRIPTION DM BOX DESIGN ADM hardware items identification Power supply board Ruggedisation of the I/O board Thermal and stress analysis Materials and processes... 24

5 Page: 5 of 38 7 SOFTWARE DEVELOPMENT SIMULATION RESULTS SENSOR AND ERROR MODELS GPS sensor model The carrier phase multipath error model The carrier phase white noise error model The total carrier phase error model IMU model RESULTS FOR THE NADIR POINTING SCENARIO Case 1 hybrid mode and 2.47mm RMS carrier phase difference errors Case 2 hybrid mode and 4.94mm RMS carrier phase difference errors Case 3 GPS mode and 4.94mm RMS carrier phase difference errors Case 4 Inertial mode RESULTS FOR SCENARIOS WITH ATTITUDE MANOEUVRES Inclusion of a torque feedforward term REAL-TIME VALIDATION TESTS TEST BED SETUP FUNCTIONAL VALIDATION - SAMPLE RESULTS IMU in-the-loop tests IMU off-the-loop tests Automatic code generation validation approach SUMMARY... 37

6 Page: 6 of 38 TABLE OF FIGURES FIGURE 3-1 APOD FM SUB-UNITS FOR THE LEO CONFIGURATION FIGURE 3-2 ANTENNAE AND BASELINE NAMING CONVENTIONS FIGURE 3-3 APOD MODES AND STATES FIGURE 5-1 3D VIEW OF THE FM UNIT MECHANICAL DESIGN FIGURE 5-2 MECHANICAL DRAWINGS OF THE FM DESIGN FIGURE 6-1 APOD DM BOX DESIGN FIGURE 6-2 APOD DM CUSTOM POWER SUPPLY BOARD (SENER) FIGURE 6-3 THERMAL RESULTS (TEMPERATURE) COMING FROM THERMAL ANALYSIS FIGURE 6-4 STRESS RESULTS (DISPLACEMENTS) COMING FROM A RANDOM VIBRATION ANALYSIS FIGURE 7-1 SOFTWARE DEVELOPMENT APPROACH FIGURE 8-1 MULTIPATH ERROR PROFILE FOR ONE ANTENNA FIGURE 8-2 PHASE DIFFERENCE ERROR BETWEEN TWO ANTENNAE FIGURE 8-3: EXAMPLE OF A PHASE DIFFERENCE MEASUREMENT ERROR PROFILE OVER ONE ORBIT FIGURE 8-4 APOD SIMULATION RESULTS NADIR POINTING HYBRID MODE (I) FIGURE 8-5 APOD SIMULATION RESULTS NADIR POINTING HYBRID MODE (II) FIGURE 8-6 APOD SIMULATION RESULTS NADIR POINTING GPS MODE FIGURE 8-7 APOD SIMULATION RESULTS NADIR POINTING INERTIAL MODE FIGURE 8-8 APOD SIMULATION RESULTS PITCH MANOEUVRE (I) - HYBRID MODE FIGURE 8-9 APOD SIMULATION RESULTS YAW MANOEUVRE HYBRID MODE FIGURE 8-10 APOD SIMULATION RESULTS YAW MANOEUVRE HYBRID MODE + FEEDFORWARD TERM FIGURE 9-1 APOD HIL TEST BED SETUP FIGURE 9-2 APOD OUTPUT FOR A REAL-TIME LABORATORY TEST USING THE IMU HARDWARE AND THE EMULATED GPS DATA FOR A LEO SPACECRAFT WITH SOME DYNAMIC MANOEUVRES FIGURE 9-3 TESTS IN GPS MODE DSPACE RESULTS FIGURE 9-4 TESTS IN GPS MODE SIMULINK RESULTS TABLE OF TABLES TABLE 5-1 PERFORMANCES... 21

7 Page: 7 of 38 ACRONYMS ACS ADOP APOD ADM AFM APFM A/S BIT BSP C/A CEL CoG COTS CPU ECEF ECI EEPROM EM FLPP FOG FM HDOP GDOP GEO GEU GPS GSFC GST HW IERS ITRF LEO LNA LOS LVLH MoI N/A OBC OS PDOP PFM PMC PPS RFQ RLG RMS S/A SBC Attitude Control System Attitude Dilution of Precision Autonomous Precise Orbit and Attitude Determination Unit APOD Development Model APOD Flight Model APOD Proto Flight Model Anti Spoofing Built-In-Test Board Support Package Coarse Acquisition Celestial (frame) Center of Gravity Commercial Off The Shelf Central Processing Unit Earth Centered Earth Fixed Earth Centered Inertial (=CEL) Electrically Erasable Programmable Memory Engineering Model Future Launchers Preparatory Program Fibre Optic Gyroscope Flight Model Horizontal Dilution of Precision Geometric Dilution of Precision Geostationary Earth Orbit Gyroscope Electronic Unit Global Positioning Service Goddard Space Flight Center Greenwich Sidereal Time Hardware International Earth Rotation Service International Terrestrial Reference Frame Low Earth Orbit Low Noise Amplifier Line Of Sight Local Vertical Local Horizontal Moment of Inertia Not Applicable On-Board Computer Operating System Position Dilution of Precision Proto Flight Model PCI Mezzanine Card Pulse Per Second Request for Quotation Ring Laser Gyro Root Mean Squared Selective Availability Single Board Computer

8 Page: 8 of 38 S/C SIL SNR SoW SPS SV SW TBC TBD TT UART UT UTC VDOP WGS WRT u,v,w X,Y,Z Spacecraft Software In The Loop Signal to Noise Ratio Statement of Work Standard Positioning Service Space Vehicle (GPS) Software To Be Confirmed To Be Determined Terrestrial Time Universal Asynchronous Receiver/Transmitter Universal Time Universal Time Coordinated Vertical Dilution of Precision World Geodetic System With respect to CEL frame unitary vectors Coordinate frame axes

9 Page: 9 of 38 1 Introduction 1.1 Context The present work is framed in the scope of a study for the European Space Agency (ESA) to develop the APOD, an onboard unit for precise autonomous orbit and attitude determination. The unit will include a multi-antenna GPS receiver and an inertial measurement unit. The APOD development has been structured in two phases. Phase 1 dealt with the specification, analysis and design of the unit models. Phase 1 concluded with the definition of the DM model to be manufactured. The second phase covered the DM chassis development, the procurement of the hardware and the manufacturing of the unit. The navigation software was developed and tested in phase Deliverable items The project has provided a great deal of valuable documentation that provides solid ground for further development. In addition to standard reporting documents such as Executive Summary, Progress Reports, Minutes of Meeting, Progress Meeting presentations, and Engineering Coordination Memos, this program has yielded the deliverable items below: Documents: - System Requirements Specification for the AFM (SRS) - System Design Document for the AFM (SSDD) - Test Plans (TP) - OBSW Design Document (SDD) - APOD Simulator User s Manual - Test Report (TR) - APOD Future Phases (power point presentation) - Mechanical and electrical drawings (in autocad and pdf form) Software: - OBSW for the APOD DM (developed algorithms in C-code and VxWorks source code) - Simulator SW in Simulink Hardware: APOD Development Model (mechanized box with integrated electronics) 2 APOD system overview 2.1 APOD general description The APOD Flight Model consists of a model unit in charge of providing fully autonomous capability to perform precise orbit and attitude determination for Low Earth Orbit (LEO) spacecraft (S/C). The modularity and flexibility of the AFM allows the tailoring of the unit to other applications such as launchers and re-entry. The AFM may be adapted to meet various platform interfaces while maintaining the same essential architecture. The APOD system is composed of the APOD unit plus the antennae subsystem.

10 Page: 10 of 38 The APOD unit is composed of the following sub-units: APOD unit subsystem Inertial measurement sub-unit (IMU), Attitude capable GPS embedded receiver sub-unit, Processor sub-unit, I/O sub-unit, Power sub-unit The APOD sub-units are enclosed in a rigid chassis to form the APOD unit. 2.2 Operational environment The environmental conditions for APOD are defined in the system requirements specification (SRS). 3 Hardware design 3.1 Hardware items identification The APOD unit is composed of the following hardware items: - IMU sub-unit: composed of the gyro assembly subsystem (and the accelerometer assembly subsystem for the APOD configurations that require it). - GPS sub-unit: composed of the GPS receiver. - SBC board: to perform the control and real-time management tasks. It will also perform the communications with the GPS. - I/O boards: one mezzanine board is inserted in each of the two PMC slots located on the SBC board: one PMC ensures the communications with the platform (by 1553 interface); the second mezzanine provides the communications with the IMU and the output serial data link. - Power sub-unit: composed by a DC/DC converter to provide the power required by the APOD sub-units.

11 APOD Executive Summary Page: 11 of 38 Gyro assembly - rates Backplane Power Supply - DC/DC conv. - EMI filter +28VDC VME Internal Bus GPS - position - velocity - raw data - GPS time Processor Unit - task scheduling - mode management - communications - computations - data storage (EEPROM) - BIT (power-up & periodic) - time synchronisation PCI I/O - read: - commands & sensors - transmit: - navigation data - attitude data - time data - status data - miscellaneous data Clock Cmd Data 1553 Data (RS-422) RS-232 Ethernet Antennae Development Figure 3-1 APOD FM sub-units for the LEO configuration IMU sub-unit The IMU sub-unit consists of two independent subsystems packaged into separate enclosures: - a gyro assembly, - an accelerometer assembly Each one includes the sensor block element and the associated electronics. Note that an accelerometer market survey has been performed. The APOD configuration addressed in this technical document, however, is tailored to LEO and the IMU sub-unit consists of the gyro assembly alone Gyro assembly The pre-selected gyro assembly is the Astrium MP FOG IMU (see appendix). This inertial gyro sensor assembly provides mountings for a mutually orthogonal triad of gyros. The gyro assembly provides the following functions: - To provide angular body rate measurements for attitude determination (propagation) - To guarantee a free-inertial attitude solution during GPS outage - To assist the GPS receiver with dynamic outputs if deemed necessary (inertial rates)

12 Page: 12 of 38 A full gyro characteristics evaluation matrix was produced as part of the market survey. Although MEMs IMUs are being developed the maturity of the space qualifiable product is still not ufficient Accelerometer assembly The present design does not include accelerometers, although these may be incorporated for specific applications (e.g. launchers, reentry). A full accelerometer characteristics evaluation matrix was also generated in the program GPS sub-unit The selected GPS sub-unit is a L1 C/A code GPS receiver with 12 channels, performs continuous tracking, and has 3 correlators per channel. The GPS receiver is designed to acquire, track and navigate using GPS satellite SPS signals. This receiver is under development by Astrium and a DM breadboard has been manufactured. Both for kinematic point solution and for dynamic filtered solution, valid (monitored by RAIM) measurements from all tracked GPS satellites are used. The MosaicGNSS Rx flyer specifies a single receiver FM board, which has roughly the size of today's digital breadboard board. Using higher integrated components, a single receiver FM board comprises one RF front-end, one DC/DC converter and the digital board. The receiver consists of two parts, the Sensor Module, which actually receives and processes the GPS signals, and the Navigation Module, which performs navigation planning and solution. The FM model provides a 1Hz PulsePerSecond (PPS) output synchronized with GPS time. At the moment, the (PPS) output is not implemented in the DM breadboard. The receiver provides raw data (carrier phase differences for each pair of antennae) for attitude determination. It also provides miscellaneous information (ephemeris,fault detection, etc). The selected GPS receiver has a time to first fix (TTFF) of less than 4 minutes for a warm start and less than 8 minutes for a cold start. Other technical specification details are found in the product documentation. The Alcatel Topstar 3000 GPS receiver is another good alternative for the GPS receiver. This receiver has almost identical performances as the quoted performances for the Astrium GPS receiver but has the advantage that it has already flown in space. More than 12 FMs have been delivered since 2000 for commercial space applications. It should be noted that the presented APOD FM design is such that it allows to incorporate either the Astrium GPS receiver or the Alcatel receiver keeping identical mechanical design and electrical architecture. Note that the Topstar-SBC interface would be also through a RS-422 interface as currently envisioned Processor sub-unit Processing resources The APOD FM CPU board selected is the S210 from Aitech. This board includes the high performance Motorola's MPC750 PowerPC processor which delivers unmatched computing power while maximum power dissipation is 5 watts. This single board computer (SBC) includes two PMC connectors for two standard mezzanine boards. One of the PMC connectors would be used for the MIL-STD-1553 bus interface.

13 Page: 13 of 38 I/O sub-unit The FM APOD I/O sub-unit is composed of two mezzanine boards inserted onto the SBC. These interface boards ensure the communications between APOD-platform and IMU-SBC. The specifications for the I/O sub-unit are consistent with the interface requirements specified in reference document [RD-1]. The interface of the I/O sub-unit with the SBC is through a PCI bus. The design approach is to use a MIL-1553 PMC board with one redundant 1553 channel to communicate with the platform, and another PMC board for data link and IMU communication via serial link interface. If required, the 1553 PMC may be modified to include an additional channel to communicate the IMU by a MIL-1553 bus interface instead of via serial interface Power supply An internal power supply is included to adapt the 28VDC input voltage to the voltage required by the SBC and the IMU. The power supply has self-monitoring capabilities such that a failure of this module is detected. The power supply accepts an ON/OFF command to power ON/OFF the APOD by command. The power supply may be designed to switch off the GPS or IMU as needed for power savings, as done for the DM Antennae subsystem The use of four (4) active patch antennae are recommended; the antennae should be mounted on a ground plane to reduce multipath. Helical type antennae may be used for better multipath reduction. The antennae may be disposed in a coplanar configuration as shown in Figure 3-2. L2 a Slave 1 Slave 2 L1 Baseline 1 X b a Master Y a Y b nominal - orbit normal Baseline 2 Slave 3 X a L3 Baseline 3 View from positive yaw body axis (Z b ) Figure 3-2 Antennae and baseline naming conventions

14 Page: 14 of System architecture Processing and communications The APOD unit employs a 3-slot VME standard backplane in which a space qualifiable SBC board is inserted. The 2 spare slots are left for future expansion and improvements; this allows a more flexible design, while the weight and consumption of the spare slots are not significant. As mentioned, the SBC includes a serial 422 lane for interface with the GPS receiver. Two PMC cards are stacked onto the SBC to provide the I/O functions. One of the PMC provides the redundant MIL-1553 bus interface with the platform. The other PMC ensures the interface with the IMU sub-unit and provides the serial data interface with the platform. The system diagram in Figure 3-1 illustrates the system top-level architecture. Note that for convenience only one I/O block diagram is depicted in the figure. A power supply is inserted for DC/DC conversion Internal connexions and cabling The FM design concept is based on a VME bus backplane architecture supplemented by two PMC interface boards. This setup yields a high degree of flexibility and modularity; in effect, different I/O internal and/or external interfaces may be supported by just changing or customising the two embedded PMC mezzanine boards. The PMC expansion slots incorporated in the S210 board also come with user defined pins. Each of the 2 standard PMC slots have three independent connectors, two for PCI signals and one for user defined signals (up to 64 user define signals). With this powerful combination (S210 + PMC boards) all the elements may be connected directly to the VME Backplane. As shown in the S210 board functional block diagram below, the PMC slots are connected to PO and P2 connectors of the VME bus. This allows to implement either RS-422 connexions and MIL-STD connexions using the VME backplane. With the present design, the RS-422 lines are routed from the PMC boards to a VME connector, and get to the IMU or GPS connector through the VME backplane. The RS-442 link for communications with the IMU is implemented in the PMC board located in slot #1. Using the PMC I/O pins, the signals reach the P2 VME connector and from there to the IMU connector located in the backplane. The MIL-STD-1553-bus output interface is implemented in the PMC board located in slot #2. The 1553 lines go through the PMC connector to the P0 VME connector and from there to the platform connector located in the backplane. With this architecture the system may adapt easily to support different IMU or GPS interface protocols by inserting different PMC boards, hence with no impact in the architecture. If a PMC board is changed, the VME board software would be updated with the drivers of the new boards. The hardware and software architecture would remain identical thus reducing the time and risk for new developments.

15 Page: 15 of Operational scenarios APOD modes The APOD unit may be in the following modes: - Off - Initialisation - Alignment - Hybrid - GPS - Inertial - Propagation For a description of each of the modes please consult the technical documentation. The Figure 3-3 illustrates the APOD modes and transitions between modes in graphical form. Note that the states for the Hybrid and GPS modes are also shown. The states that APOD first enters in a mode or submode are marked with dots. APOD modes and states OFF TC TC INITIALISATION TC TC Automatic or TC TC TC TC TC ALIGNMENT Automatic or TC HYBRID Automatic or TC TC or Automatic (GPS recovery) TC or Automatic (GPS failure) TC Initialisation Auto Auto Alignment (att) Data Acquisition Auto INTENSIVE Auto TC Initialisation Auto Sleep Alignment (att) Data Auto Acquisition Auto ECONOMIC TC Initialisation Auto Auto Data Acquisition INERTIAL TC TC or TC Automatic (IMU + GPS recovery) TC or TC or Automatic Automatic (IMU failure) (IMU + GPS failure) TC or Automatic (IMU recovery) PROPAGATION TC or Automatic (GPS recovery) TC or Automatic (IMU failure) Auto TC Initialisation Auto Alignment (att) Data Acquisition Auto INTENSIVE TC or Automatic (GPS failure) TC or Automatic (IMU recovery) GPS Auto Auto TC Initialisation Auto Sleep Alignment (att) Data Auto Acquisition Auto ECONOMIC TC TC TC Figure 3-3 APOD modes and states

16 Page: 16 of 38 4 Algorithm design For the studied application the attitude and orbit determination are independent problems and as such the APOD treats each estimation problem independently. 4.1 Orbit determination Several methods have been investigated. The proposed methods are subject to modifications or refinement depending on simulation results. The orbit determination process relies primarily on the GPS measurements, but also on an on-board orbit model. The APOD may use either the point PVT solution (in order to have independent measurements) or the dynamic filtered solution also available from the selected GPS receiver. No orbit determination algorithms were implemented as they are provided by the GPS receivers in the market. The study of precise position determination will be further explored once a candidate mission is defined. This feature is left for future research. 4.2 Attitude determination The spacecraft attitude can be determined with GPS when three or more antennae are available where the difference of the carrier phase measurements between them are observed. However, GPS on itself is not considered reliable enough for this purpose thus GPS is used in conjunction with other sensors and enhanced through Kalman filtering. The first problem to be solved is the integer ambiguity resolution. Once the partial carrier phase differences are known, the attitude problem may be addressed. An extended Kalman filter has been developed for the GPS/gyro hybridisation. This filter employs a new ambiguity resolution technique and a well established tight integration of the GPS carrier phase measurements and the IMU angular rate measurements Attitude determination algorithms The general attitude determination approach based on linearised least squares is the pre-selected method. This method provides an optimal estimation for non-correlated measurements although for convergence of the correct solution, it requires a good initial attitude. Although APOD may accept initialisation from external platform sensors, the requirement of fully autonomous determination still stands. Therefore, the initial attitude obtained in the acquisition mode (right after the ambiguities are resolved) must be good enough for the carrier phase based attitude least squares algorithm to converge. For the least squares algorithms an initial error within 10 degrees for each axis is deemed sufficient Kalman filter design For the estimation of attitude, an Extended Kalman filter discrete formulation is implemented which exploits the benefits of combining the GPS and IMU sensors. The dynamic model of the Kalman filter is a kinematic model based on the gyroscope measurements. The type of measurements combined are as follows: GPS measurements, which provide instantaneous simultaneous (quasi simultaneous for a multiplexed system) measurements of the L1 carrier phases IMU measurements which provide the S/C angular body rates wrt an inertial reference system expressed in the body frame (after convenient adaptation of the output)

17 Page: 17 of 38 The filter processes the raw measurements from the GPS receiver every time data is available (typically 1Hz to 10 Hz for carrier measurements). Consequently, in nominal conditions with good visibility a drift-free attitude update is obtained at least every second. The estimated state error residual is computed based on the measurement update, and the state and covariance matrix are propagated to the next epoch using the IMU measurements. The body rate data is used at a rate of 10Hz to propagate the kinematic rotational equations. This sample rate is to be confirmed based on simulations and on the IMU effective bandwidth and supported rates. For reentry or aerocapture applications higher update rates may be required. Therefore, the APOD software developed for the DM has been programmed to function at a 50Hz update rate. The acquisition of the first attitude solution requires an initial guess of the attitude to linearise the observation equations about it. The APOD unit is designed such that the estimate can be obtained either from an external source (star tracker, magnetometer + sun sensor), or by an autonomous initialisation process. The output of the filter are the estimated attitude quaternion and gyro biases. The attitude rate quaternion is computed directly from the estimated attitude quaternion. To minimise computational burden and for simplicity the filter state vector is first proposed to have only seven states: the attitude quaternion with respect to an inertial reference frame (only 3 first elements) and the gyro biases (3 elements). According to Lefferts et Al (1982), the formulation of the covariance propagation and measurement update equations using the six-dimensional state vector precludes singularities. Time propagation of the quaternion state estimate and the covariance matrix is performed using the kinematic model and the angular body rate measurements. 4.3 Test program A suggested test plan has been produced for future reference. The plan considers the manufacturing of five formal APOD units: - one APOD Development Model (ADM), - one APOD Engineering Qualification Model (AQM), - one APOD Protoflight Model (APFM) - two APOD Flight Models (AFM). The qualification of the design and applied processes would be performed in three steps: a. ADM which would be submitted to functional and performance tests, in order to validate the algorithms and the electrical and software architecture, including HIL real-time tests in operational conditions. b. AQM which will be submitted to vibration qualification test, EMC (full test), physical check, and functional testing, in order to validate the FM mechanical design as well as the electrical and software approach. c. PFM which will be submitted to functional test, EMC (partial test) and to a qualification thermal vacuum test to qualify the ability of the design to withstand the thermal environment. It will be also submitted to vibration acceptance test. This qualification test program ensures verification of all the critical requirements. FMs will also be tested but under acceptance levels and durations before final delivery.

18 Page: 18 of 38 For each model, some or all of the following test groups will be performed, as detailed in the Test Plan: Functional Performance Physical Properties EMC Vibration Thermal Vacuum 4.4 Deliverable models Development Model The ADM will exhibit the following characteristics: It will be manufactured using low-cost COTS space qualifiable representative hardware in order to permit a deep evaluation of the overall concept focusing on software and electrical issues. It will enable to provide a functional assessment and some preliminary performance data. It will be a representative platform to test new scenarios and ideas in real-time conditions Engineering Qualification Model The EQM will exhibit the following characteristics: It will be populated using, as far as possible, same components as in PFM/FM in form, fit and function. However, some may be of standard commercial quality. The EQM will be assembled in a box fully representative of the PFM/FM to perform qualification with respect to Physical Measurements, Vibration and EMC testing Proto Flight Model The PFM model will exhibit the following characteristics: It will be manufactured following full flight standards and procedures. It will be populated using flight standard parts. It will be fully representative of the flight model, being in fact considered as the first flight model. The mechanical design of the box will be the same of the flight unit. The PFM will be subject to a full qualification test program except with respect to EMC and vibration testing that will be performed on the ABM. In the case of vibration testing this model will be submitted to acceptance vibration tests (acceptance levels, acceptance durations) Flight Models The FM models will exhibit the following characteristics: They will be manufactured following full flight standards and procedures. The FMs will be subject to an acceptance test program (acceptance levels, acceptance duration) In case of modifications introduced after equipment qualification with the ABM and PFM, a qualification extension may be required to be performed on the FM. 5 Flight Model The guideline followed for the design of the APOD FM unit (AFM) has been to house in a single package all the components of the equipment, i.e.:

19 Page: 19 of 38 IMU gyro sub-unit GPS sub-unit DC/DC converter Power PC SBC The APOD unit is manufactured with Al6082T6 aluminium (or equivalent) CNC machined; external surfaces will be black anodised and all internal surfaces will be then treated with Alodine All different pieces will be bolted together by means of stainless steel bolted joints. Average thickness of the walls will be 1.5 mm. Twelve mounting ears give provision for M4 screws to assembly the unit onto the satellite panel. 5.1 Physical configuration The following figure shows the unit conceptual design. Top cover has been removed for illustration purposes Figure 5-1 3D view of the FM unit mechanical design The unit is divided in two cavities, fitted to house the different sub-elements of AFM. Lower cavity is designed to house the GPS receiver Lower and upper cavity are separated via an internal mid plate. On the upper cavity it is housed the main mass contributor for the unit; the IMU unit, weighting 3 kg. The AFM unit on its current design provides room for three 6U VME slots; although only one is used at the moment. The implementation of the current IMU unit selected does not allows the use of 2 of the 3 slots. The design is, nevertheless, ready for future trends where the use of a smaller IMU is envisaged A 3 slots VME backplane is bolted to the unit base plate; and that is where the Power PC is connected. On the upper cavity the DC/DC converter is housed. For configurations requiring an accelerometer unit this would be housed in the upper cavity.

20 Page: 20 of 38 This design approach, on the other hand, allows the implementation of different GPS units (Alcatel, Surrey) or even IMU s without varying the unit envelope and only with minimal modification on the inside of the unit. Software modifications, however, would be relevant if these alternative sensors are employed.. The following figure shows the main dimensions of the AFM unit. Top cover has been removed for illustration purposes. Figure 5-2 Mechanical drawings of the FM design 5.2 FM system performances The APOD system performances are expected to exceed the requirements established in the SRS. The expected minimum performances (root mean square) of the unit are included below. Output Data APOD Mode HYBRID GPS INERTIAL Position 10m 10m 100m (after 24h) Velocity 0.01m/s 0.01m/s 0.2m/s (after 24h) Acceleration (non g) 0.01 m/s 2 N/A 0.01m/s 2 Attitude 0.1º 0.3º 1º (after 1h) Attitude rate 0.01º/s 0.2º/s 0.01º/s Time of fix 1000ns 1000ns N/A UTC time (PPS) 1000ns 1000ns N/A

21 Page: 21 of 38 Table 5-1 Performances 1 These values assume each of the three antennae baseline is <1 metre. Note that the attitude determination accuracy can be improved with longer antennae baselines. If this is not feasible, the APOD may interface with a star tracker instead of with the three slave antennae. 5.3 FM external interfaces The APOD has three interfaces with the platform and are implemented through the I/O subunit: one 1553 interface for TM/TC, a RS-422 interface for the Data delivery link and a TTL interface for PPS. The data link interface is a differential synchronous serial interface that is in accordance with the EAI Standard RS422. It has two lines for the data transmission and two lines for the data clock. The APOD has one RS-232 interface and an Ethernet interface for debugging purposes. The GPS receiver provides the power to the active antennae selected for the application. 5V power is typical Power The APOD is powered from a 28VDC spacecraft power bus and it is able to switch ON/OFF by command. The DC/DC converters provide the voltages needed for the sub-units included in the APOD housing: processor and I/O sub-units, GPS and IMU. The Power interface includes a differential common EMI filter. The typical MIL-STD-461 perturbations are supported by the unit. 6 Development Model 6.1 DM description The APOD DM system is composed of two main elements: the AGGA-2 Astrium GNSS DM breadboard receiver and the APOD DM box including the IMU and all the required electronics. 6.2 DM box design Figure 6-1 APOD DM box design 1 The performances are to be verified by further simulation and HIL tests

22 Page: 22 of 38 Figure 6-1 shows the DM unit conceptual design. The system has been designed using a 3D CAD software to ensure the successful integration of all the units housed inside. From the 3D model the mechanical drawings for manufacturing were generated. Right cover has been shifted up for illustration purposes. The unit is divided in two cavities. The left cavity houses the VME rack with 4 slots and the three VME boards (processor board, I/O board, and power supply board). The right cavity houses the IMU unit. The cable harness is not shown. A 4 slot 6U VME backplane is bolted to the unit base plate; and that is where the Power PC is connected. The box structure is fixed by means of 6 main bolts to a baseplate ADM hardware items identification - IMU sub-unit: composed of the gyro assembly subsystem and an accelerometer assembly. - GPS sub-unit: composed of the AGGA-2 GNSS GPS breadboard receiver. - SBC rugged VME board: to perform the control and real-time management tasks. - I/O commercial VME board: to provide the communications with the IMU, the GPS, and the output serial data link. Unit was adapted to fit a conduction cooled and rugged design commercial PMC board: one mezzanine board is inserted in one PMC slot located on the SBC board to ensure the communications with the platform (by 1553 interface). - Power sub-unit: composed by two DC/DC converters to provide the power required by the APOD sub-units. The power subunit was built by SENER and integrated as a VME board Power supply board Since an adequate commercial off the shelf power supply board was not available at the time of the DM design, it was decided to design and specify a custom PCB board compatible with VME specifications. The PCB was manufactured such that two low cost DC/DC converters could be integrated on it. Figure 6-2 shows the geometric model of the manufactured power supply board. The SENER power supply board includes circuitry to enable additional serial communication channels. In addition the board provides the system with power saving features that could be most interesting for low power space applications. Indeed, in the current DM configuration, the APOD system powers off the IMU when in GPS mode for significant power savings. The system switches on the IMU when the APOD is changed back to hybrid or inertial modes. The design of a similar power supply VME board for the EQM and FM models is hence a possibility that may be considered.

23 Page: 23 of 38 Figure 6-2 APOD DM custom power supply board (SENER) Ruggedisation of the I/O board The air cooled VME board for I/O communications had to be adapted to fit in a conduction cooled rugged box. This effort required the insertion of a Gap Pad for thermal continuity from the board components to the box structure and the integration of wedge locks for vibration endurance Thermal and stress analysis Even if the DM is not expected to undergo qualification tests, the team used the latest FEM and CAD simulation tools for analysis. This is an initial exercise that would be redone with higher attention for the design of an eventual flight model. Figure 6-3 Thermal results (temperature) coming from thermal analysis

24 Page: 24 of 38 Figure 6-4 Stress results (displacements) coming from a random vibration analysis Materials and processes All materials and processes using on the manufacturing and integration of DM unit are selected among those widely used in the industry. The unit housing is made of aluminium AA6082T6 (or equivalent) CNC machined; being the surface finishing a coating with conversion coating per MIL-C-5541 class 1A. 7 Software development A significant part of the APOD software was developed using the latest autocode tools (RTW/Embedded Coder 4.0). The software model in Simulink was converted to ANSI C and cross compiled to a PowerPC space qualifiable board for real-time validation. The target application runs on VxWorks. The development approach was the one depicted in Figure 7-1. Note that only test platforms 1, 2 and 3 were used in the project. Platform 3 was used for development and test of the final APOD software for the target HW and O.S.. The main findings were 1) autocoded software model integration effort into VxWorks was moderate or low, 2) manual coding was still quite prominent, 3) even for this application with a low autocode/manual code ratio, the productivity gain was remarkable (possibly around 30%). The use of a UML/SDL tool would have dramatically increased the productivity, in addition to enhance maintainability. The effort involved at the OS task level was remarkable (semaphore definition, synchronization, demodulation of sensor data). The manual coding of the software model implemented in Simulink was 17% of the total lines of code (excluding comments). Note also that the whole software model autocode accounted for 40% of the total APOD software. For this development a state machine was implemented in Stateflow and code was automatically generated using StateflowCoder. No FDIR was designed nor implemented.

25 Page: 25 of 38 ESA INPUT APPLICATION DEFINITION SENER Ingenieria y Sistemas, 2002 Test platform 1 Sensor Data Actuator Data SIMULATION OF REALITY Spacecraft dynamics Space environment Actuator and sensor models ENVIRONMENT CODE ON-BOARD SW simulated measurements control commands Non-Real-Time Tests FDIR SW Mode Logic SW TM/TC SW Data Fusion and Filtering APOD SW Attitude Determination Orbit Determination AOCS SW Control Rapid Prototyping Auxiliary Code RTW SFC FLIGHT CODE Prototype Automatic Code Generation Real-Time Platform Flight RT SW Environment RT SW DSPACE HW I/O HW I/O Test platform 2 Func. & Perf. Real-Time tests Auxiliary Code Test platform 3 Environment RT SW DSPACE HW I/O TARGET (no SIL*) Automatic Code Generation FLIGHT CODE Target HW Target O.S. *SIL= Sensor In the Loop Test platform 4 VxWorks/Tornado (Functional ) HOST Tilt Table GPS Sim TARGET (Perf. with SIL*) TARGET Func. & Perf. Real-Time tests on target Physical & Environmental tests on target Figure 7-1 Software development approach

26 Page: 26 of 38 8 Simulation results For the algorithm and attitude determination software validation the system was tested in a non-realtime high fidelity simulator in Matlab/Simulink which contains a model of the GPS constellation. In this model not only the GPS constellation dynamics are modeled but also the main GPS sensor functions such as satellite dynamic selection based on ADOP criteria, and raw data measurements. The tests have been conducted for the earth pointing satellite application, for an orbit with a semimajor axis of 7200km and an inclination of 45º. The reentry and aerocapture applications were unfortunately left for future work as they would require adaptation of the developed navigation algorithms and additional simulation verification effort for new reality scenarios including relevant linear accelerations. Extensive simulation has proven the stability and accuracy of the developed Kalman filter algorithms. Only a number of tests are summarized herein. The results next presented are organized as follows: - Sensor and error models definition used during simulations - Performance of attitude determination for a nadir pointing scenario. - Performance of attitude determination for a several manoeuvres scenarios - Evaluation of a feedforward term using the actuator torque as an input to the APOD 8.1 Sensor and error models GPS sensor model The modelled GPS receiver tracks up to six GPS satellites simultaneously when functioning in attitude mode where the ADOP is used as the selection criterion. At each time step the true phase differences are calculated where the measurement errors are added which consists of a deterministic multipath model and white receiver noise The carrier phase multipath error model A 3D height profile, which represents the phase measurement error as a function of the elevation and azimuth of the GPS line of sight vector w.r.t. antenna plane, is created for each antenna. They are created such that the errors are greater when the elevation of the line of sight vectors are low (though not necessarily monotonically increasing with decreasing elevation). A factor is introduced which is multiplied by the profiles in order to adjust the error magnitude for the convenience of the user. An example of the multipath profile is shown in Figure 8-1. In Figure 8-2 the carrier phase error for a given baseline is shown. This error results from the difference of two multipath profiles that correspond to different antennae. During simulations it was observed that the multipath error range greatly influenced the performances of the attitude determination filter, as expected Multipath Error [m] Elevation [deg] Azimuth [deg] Figure 8-1 Multipath error profile for one antenna

27 Page: 27 of Phase Difference Error [m] Elevation [deg] Azimuth [deg] 45 0 Figure 8-2 Phase difference error between two antennae The carrier phase white noise error model For simulations the white noise in carrier phase is modeled as a gaussian zero mean random variable with standard deviation equal to 1.67mm, that is 5mm for 3 sigma. The delivered simulator includes a GUI such that the user may easily increase or decrease the measurement white noise by using a multiplying factor. Several tests have been conducted at SENER with higher white noise errors obtaining little decrease in performance. As opposed to multipath, the white noise is quite well filtered by the Kalman mechanization The total carrier phase error model An example of the total measurement errors (multipath + noise) over one orbit for one receiver channel is shown in Figure 8-3. This figure just serves to give an impression of the resulting measurement error profiles which are used for the tests. 20 Error profile (multipath + white noise) Measurement errors [mm] Figure 8-3: Example of a phase difference measurement error profile over one orbit IMU model The IMU model used for the final validation simulations corresponds to the Astrium MP (Middle Performances) Eurofog IMU. This IMU has a bias of approximately 0.1º/h with a repeatability of 0.2º/h, an angular random walk <0.01 / h. (1σ), a scale factor linearity of 100ppm (3σ). The gyro axes misalignement is also modelled. The sample rate was assumed to be 10Hz, supported by the unit.

28 Page: 28 of Results for the nadir pointing scenario Case 1 hybrid mode and 2.47mm RMS carrier phase difference errors Figure 8-4 APOD simulation results Nadir pointing Hybrid mode (I) The following table summarizes the obtained attitude determination errors in degrees: Euler Angle RMS Error MEAN MAX Yaw Pitch Roll

29 Page: 29 of Case 2 hybrid mode and 4.94mm RMS carrier phase difference errors Figure 8-5 APOD simulation results Nadir pointing Hybrid mode (II) The following table summarizes the obtained attitude determination errors in degrees: Euler angle RMS Error MEAN MAX Yaw Pitch Roll

30 Page: 30 of Case 3 GPS mode and 4.94mm RMS carrier phase difference errors Figure 8-6 APOD simulation results Nadir pointing GPS mode The following table summarizes the obtained attitude determination errors in degrees: Euler angle RMS Error MEAN MAX Yaw Pitch Roll

31 Page: 31 of Case 4 Inertial mode Figure 8-7 APOD simulation results Nadir pointing Inertial mode The following table summarizes the obtained attitude determination errors in degrees: Euler angle RMS Error MEAN MAX Yaw Pitch Roll Results for scenarios with attitude manoeuvres The APOD attitude determination software was also tested for simulated attitude manoeuvres. The observed performance of APOD was adequate. The obtained results for the hybrid, GPS and Inertial modes are illustrated in the following figures. The next figure shows the APOD performance for an attitude manoeuvre from 0º to 70º in pitch conducted in only 15 seconds. The maximum error was about 0.45º in pitch which is caused by the fact that the Kalman filter does not take into account the control torques during the manoeuvre. There is also a raise of the roll and yaw angular errors during the manoeuvre. The error in yaw continued to raise after the manoeuvre until 0.2º. This is caused by a sudden increase in multipath errors as the spacecraft changes its attitude and consequently also the GPS lines of sight such that a worse case of multipath arises. Other simulated manoeuvres have shown that the opposite also occurs where the change in spacecraft attitude results in lower multipath errors.

32 Page: 32 of 38 Figure 8-8 APOD simulation results Pitch manoeuvre (I) - Hybrid mode Inclusion of a torque feedforward term SENER has evaluated a simple method to improve the attitude determination performance during manoeuvres. As expected, one may improve the attitude propagation accuracy by introducing a feedforward term providing the control torque profile to the APOD during manoeuvres. This element acts as a lead term by which the APOD software estimates the delta angular velocity expected due to the control torque. For this exercise it was assumed that the spacecraft inertia and the executed torques are well known. The improvement in performance is evident as illustrated in the yaw manoeuvres depicted in Figure 8-9 and Figure 8-10 corresponding to a case with the feedforward term deactivated and activated respectively. These promising results indicate the convenience of considering this technique in future programs. To better illustrate the power of this method the Figure 8-9 and Figure 8-10 below show the compared performances for a simulated 70º yaw manoeuvre without multipath errors. In the future the simulations should be conducted with a more detailed error model for the satellite inertia and the actuators.

33 Page: 33 of 38 Figure 8-9 APOD simulation results Yaw manoeuvre Hybrid mode Figure 8-10 APOD simulation results Yaw manoeuvre Hybrid mode + Feedforward term

34 Page: 34 of 38 9 Real-time validation tests After the integration and individual testing of all the DM box sub-units a final validation test phase was performed. As agreed with ESA the real-time validation effort was limited to the verification of the main functional and interface requirements for the APOD DM system. This included the functional validation of the attitude determination on-board software, the verification of the 1553 TM/TC link, and the evaluation of the correct functioning of the RS422 output data link. 9.1 Test bed setup To validate the APOD OBSW on the target DM configuration the APOD system was tested with the VME boards and the IMU hardware in the loop. The GPS receiver carrier phase measurements for a LEO orbit spacecraft were emulated from dspace using the Simulink reality simulator. The Figure 9-1 is a schematic of the test setup. Reality Model dspace GNSS Sensor Real-Time Emulation Data (RS-422l) APOD DM (IMU=ON) 1553 TM/TC Data Link (RS-422l) Platform Simulator and Recorder PC Figure 9-1 APOD HIL test bed setup 9.2 Functional validation - sample results The real-time test integration has ended with the system providing adequate results. Essentially three sort of tests have been performed in the real-time test bed. a) Onboard software tests with IMU in the loop, b) Onboard software tests with IMU off the loop c) 1553 bus link tests (mode selection, and parameter modification) The latter were conducted with the APOD in GPS mode and a sample of the results are provided in IMU in-the-loop tests The sequence for the IMU in-the-loop tests was as follows: a) 2 minutes in hybrid mode b) 1minute in GPS mode c) 1 minute in Inertial mode d) 1 minute in GPS mode e) 5 minutes in hybrid mode In step e) the APOD DM box was moved to simulate spacecraft manoeuvres. Three types of manoeuvres were emulated: i) a roll manoeuvre from horizontal level to +35º back to level and to 35º ii) a pitch manoeuvre from horizontal level to +40º

35 Page: 35 of 38 iii) a yaw manoeuvre from horizontal level to +20º back to original position and +20º The results depicted in Figure TBD indicate that the APOD is correctly interfacing with the IMU. Note that during the forced manoeuvres the Euler angle estimations are updated correctly. As the IMU is mounted with a 45º inclination inside the APOD model Note that the roll manoeuvre applied to the APOD DM box is equivalent to a yaw manoeuvre in IMU axes due to the 45º tilt of the IMU mounting plate. Note also that the observed change in roll angle for manoeuvre i) is about 25º, as expected due to the 45º tilt of the IMU. Indeed the 35º angle of the applied motion multiplied by cos(45º) is about 25º. Figure 9-2 APOD output for a real-time laboratory test using the IMU hardware and the emulated GPS data for a LEO spacecraft with some dynamic manoeuvres IMU off-the-loop tests The system was also tested on the real-time test bed without the IMU in the loop to verify the software functional correctness in GPS mode. The results were satisfactory. The Figure 9.5 below shows the Euler angle errors obtained when running the software on the target DM system and using the emulated GPS measurements of a LEO spacecraft through dspace.

36 Page: 36 of 38 Figure 9-3 Tests in GPS mode - dspace results Automatic code generation validation approach The Figure 9.3 depicts the Euler angle errors obtained from the DM output data link running on the target APOD system. The Figure 9.4 is a scope of the results in Simulink when using the identical GPS sensor measurements. For the real-time tests the GPS measurements were emulated in the dspace environment and transmitted through a serial connection to the APOD DM box. It is interesting to note that the results in the Simulink environment and in the Tornado/VxWorks/dSPACE environment are identical. Hence the process of automatic coding of the software model was deemed as correct. Of course, the validation of the software generation process will need to be more thorough for a final intended application. Due to the development nature of the project the effort devoted to the validation was coherent with the resources available. Figure 9-4 Tests in GPS mode - Simulink results