Team Name: LOWCOINS University of Rome “La Sapienza”

LowCoINS SED – Final

BEXUS Student Experiment Documentation

SED Document ID: LowCoINS SED – Final.doc

Mission: BEXUS-6

Team Name: LOWCOINS University of Rome “La Sapienza”

Student team leader: Paolo Montefusco Team members: Maria Cristina Oliva

Emanuele Medaglia Experiment Title: Low.Co.I.N.S. – Low Cost Inertial Navigation System Version: Issue Date: Document Type: Valid from:

6 26.February 2009 Spec Issued by:

........................................................................ Experiment Scientist Approved by:

........................................................................ Payload Manager



Change Record

Version Date Changed chapters Remarks

0 2008-02-29 New Version Blank Book 1 2008-03-10 Marked Changes Student distribution 2 2008-04-14 1,2,3,4,8,9 PDR 3 2008-06-03 All CDR 3.1 2008-06-20 2.1.1, 2.1.2, 2.2, 2.3,

3.6, 7.1 CDR issue 2

4 2008-08-20 2, 3.1.1, 3.1.3, 3.1.7, 3.3, 3.4.4, 3.4.5, 4, 5.3, 5.5, 5.6

MTR

4.1 5

2008-08-25 2008-09-28

3.1.7, 3.3.2, 3.3.3, 9.1, 9.2, 9.3, 9.4, 9.53.1.3, 3.1.4, 3.1.6, 3.3.2, 3.3.4, 3.3.5, 3.3.6, 3.4, 9.5, 9.7.2, 9.8

MTR issue 2

Launch campaign

6 2009-02-26 6, 8, 9, 10.7.3 Final report



Abstract: The peculiarity of an Inertial Navigation System (INS) is that does not require any external reference in order to determine position, orientation, or velocity. INS is the only self-contained navigation system, therefore it is specially suitable for applications on rockets or balloons.

Low.Co.I.N.S is a low cost inertial navigation system based on a strapdown design that foresees the use of accelerometers and gyros rigidly connected to the vehicle. Inertial measurements are obtained by MEMS-based motion sensing devices. MEMS technology is improving day over day, and it will probably reach a sufficient level of accuracy in the near future to provide an effective low cost alternative in the production of inertial navigators where extremely high accuracy is not required. The limits of actual MEMS sensors are known and our experiment wants to test those limits trying to find out the maximum performances derivable from using them in extreme environment.

Keywords: Inertial navigation system, MEMS, Micro Electro Mechanical Systems, INS, Inertial Measurement Unit, IMU, Accelerometer, Gyro.



Table of Contents

TABLE OF CONTENTS ..................................................................................................... 4

INTRODUCTION ................................................................................................................... 7 1.1 Experiment objectives ......................................................................................... 7 1.2 Experiment overview ........................................................................................... 8 1.3 Team Organisation .............................................................................................. 9

2 INERTIAL NAVIGATION ........................................................................................... 10 2.1 Inertial navigation overview ............................................................................. 10 2.2 Inertial navigation process ................................................................................ 10 2.3 Mechanization ................................................................................................... 12

3 EXPERIMENT DESCRIPTION .................................................................................. 13 3.1 Hardware .......................................................................................................... 13

3.1.1 Electrical design ................................................................................. 13 3.1.2 Prototyping .......................................................................................... 16 3.1.3 Mechanical design .............................................................................. 18 3.1.4 Dimensions .......................................................................................... 23 3.1.5 Component list (detailed component list in annex) ............................. 24 3.1.6 Mass budget with single components .................................................. 24 3.1.7 Environment protection - Thermal ...................................................... 25

3.2 Power ................................................................................................................ 27 3.3 Data Handling ................................................................................................... 30

3.3.1 Data Acquisition ................................................................................. 30 3.3.2 Onboard software ............................................................................... 30 3.3.3 List of Telecommands .......................................................................... 32 3.3.4 Ground software ................................................................................. 35 3.3.5 Data Rate & Storage of Data .............................................................. 37 3.3.6 Communication, Data Link & Protocol .............................................. 38

3.4 Operation .......................................................................................................... 40 3.4.1 Experiment Connections ..................................................................... 40 3.4.2 Assembly and checkout of experiment................................................. 41 3.4.3 Optional Compass Calibration ........................................................... 41 3.4.4 Operations time line ............................................................................ 42 3.4.5 Maximum allowable launch delay time............................................... 43 3.4.6 Post Flight operations ......................................................................... 43



4 SYSTEM PERFORMANCE ANALYSIS ...................................................................... 44

5 PROJECT PLANNING (PHASE B AND C) ............................................................... 46 5.1 Resource estimation (costs) ............................................................................... 46 5.2 Support and Facilities ....................................................................................... 47 5.3 Time schedule .................................................................................................... 48 5.4 Task definition ................................................................................................... 48 5.5 Test plan ............................................................................................................ 49 5.6 Risk management ............................................................................................... 50

5.6.1 Risk analysis – Project development ................................................... 50 5.6.2 Risk Analysis – Safety.......................................................................... 50 5.6.3 Risk Analysis – Components ............................................................... 50

6 OUTREACH PROGRAM ............................................................................................ 53

7 MEETINGS AND REVIEWS .................................................................................... 54 7.1 Preliminary Design Review – PDR ................................................................... 54

8 EXPERIMENT REPORT ............................................................................................ 56 8.1 Experiment integration ...................................................................................... 56 8.2 Pre-flight tests ................................................................................................... 57 8.3 Launch day ........................................................................................................ 58

8.3.1 Flight phases ....................................................................................... 59 8.3.2 Thermal control: performance and considerations ............................ 66 8.3.3 Communications .................................................................................. 66 8.3.4 Data storage ........................................................................................ 67 8.3.5 Flight performance ............................................................................. 67

8.4 Post-processing ................................................................................................. 68 8.4.1 LowCoINS navigation performance .................................................... 68 8.4.2 LowCoINS as AHRS ............................................................................ 76 8.4.3 Future improvements & conclusions .................................................. 81

9 ABBREVIATIONS AND REFERENCES .................................................................... 83 9.1 Abbreviations .................................................................................................... 83 9.2 References ......................................................................................................... 85

10 ANNEX ........................................................................................................................ 86 10.1 Schematics ......................................................................................................... 86 10.2 Part list .............................................................................................................. 89 10.3 PCB ................................................................................................................... 90



10.4 Harness .............................................................................................................. 92 10.5 Housing design .................................................................................................. 93 10.6 Time Schedule ................................................................................................... 96 10.7 Outreach ............................................................................................................ 97

10.7.1 Workshop poster ................................................................................. 97 10.7.2 Published Article on “Memorie della Società Astronomica Italiana”98 10.7.3 Abstract for ESA Symposium on European Rocket and Balloon

Programmes and Related Research .................................................. 103 10.8 Test Report ...................................................................................................... 104

10.8.1 Overall hardware & software testing ............................................... 104 10.8.2 IMU testing – Sampling frequency .................................................... 105 10.8.3 IMU testing – Accelerometers calibration ........................................ 107 10.8.4 Magnetometers testing and calibration ............................................ 112 10.8.5 Pressure sensor calibration .............................................................. 115 10.8.6 Final Test: thermal-vacuum test with battery ................................... 116 10.8.7 Old Test Report ................................................................................. 118



INTRODUCTION

Inertial navigation techniques recently faced important developments with the introduction of very low cost sensors, collectively labelled as Micro Electro Mechanical Systems.

Our experiment is a test of our capability in the determination of the cinematic state of the Bexus payload gondola through the design and validation of an inertial measurement unit built with low cost commercial off-the-shelf components.

The sensors involved in the experiment have to provide measurements of accelerations, angular rates, magnetic field strength and direction, as well as atmospheric pressure and internal experiment temperature. The accuracy of these data makes possible the determination of the position and the attitude of the Bexus payload gondola with a precision that we want to evaluate and increase as much as possible.

MEMS sensors may not be able to provide inertial measurements with a sufficient level of accuracy needed in cutting-edge inertial navigators, however they might be accurate enough to allow for navigation where pinpoint accuracy is not required or where periodic re-alignment is made available through other measurements (eg. GPS and/or magnetometers).

1.1 Experiment objectives

Our main objective is to learn designing our first-ever “hands-on” project. Anyway, technical objectives we want to achieve are:

• To demonstrate the feasibility of a MEMS-based INS

• To estimate errors derived by using COTS components

• To compare results with more precise navigation system (GPS) provided by Bexus

• To investigate the hypothetical use as a possible backup in case of main navigation system unavailability

Moreover, through exhaustive post-processing, we want to perform the following analysis:

• Test of the solution validity with re-alignment (from “official” data performed at various times and mission phases)



• Behavior of the error equation for the INS obtained considering the “official data” as the true solution and linearizing around.

• Check of the sensors’ errors and calibration

1.2 Experiment overview

Our experiment consists mainly in the determination of the position and attitude of the payload gondola during the entire flight time. Accelerometers and gyros will sense acceleration and angular rates along three orthogonal axis, thus integrating those measurements in an opportune reference frame, the position and attitude of Bexus can be known once initial conditions have been specified.

Being a dead-reckoning process, where either attitude or position is evaluated using the state at the preceding time and adding the rate of change times the time interval, the calculated solution will sooner or later diverge from the true solution due to the accumulation of errors. Therefore it is necessary to periodically re-align the navigation platform using other references. These can be the Earth magnetic field vector that can provide the direction of the magnetic north as well as a reference around which is possible to determine the attitude. Barometric altimeter may be an effective reference and can be used to compare the calculated position along the vertical axis. Finally, during the post processing, also GPS data can be used for platform re-alignment as well as a benchmark.

Dead-reckoning also means that it is required to have an uninterrupted data flow from sensors, as if some measurements are lost, even for a small time interval, there is no more way to estimate correctly the cinematic state and the flight trajectory after the loss of data will be unrecoverable. That is why, even if we continuously downlink data to ground, we decided also to store all data gathered in a flash memory onboard. So, even a temporary loss of telemetry data stream will not affect the results of the experiment.

The reconstruction of the flight trajectory will be attended on the ground. Once the experiment has been recovered, the memory will be dumped and an exhaustive post-processing will occur with all the data gathered by the unit. During the post-processing the flight trajectory will be reconstructed in several methods and compared with GPS data, then exhaustive data analysis and comparison will be attended in order to find out the best solution.



1.3 Team Organisation

We all come from “Scuola di Ingegneria Aerospaziale” Univeristy of Rome “La Sapienza” where we are attending the last year at master degree level in Astronautics Engineering. In our academic experience we have had a “wide-range” overview of several aspects of engineering: starting with maths, physics, aerodynamics and astrodynamics, going through heat transfer and thermal control, attitude determination, guidance and navigation systems, space structures and materials, electronics and telecommunications. E.S.A. is now offering us the unique opportunity to work and face out with what we have studied so far.

Details about team-members tasks will be given later in chapter 5.



2 INERTIAL NAVIGATION

2.1 Inertial navigation overview

The operation of inertial navigation systems depends upon the Newton’s laws. They tell us that the external force acting on a body produce a proportional acceleration on it. Given the ability to measure that acceleration, it would be possible to calculate the change in velocity and position by performing successive mathematical integrations of the acceleration with respect to time. Acceleration can be determined using a device known as an accelerometer. An inertial navigation system usually contains three such devices, each of which is capable of detecting acceleration in a single direction. The accelerometers are commonly mounted with their sensitive axis mutually perpendicular. In order to navigate with respect to our inertial reference frame, it is necessary to keep track of the direction in which the acceleration are pointing. Rotational motion of the body with respect to the inertial reference frame may be sensed using gyroscopic sensors and used to determine the orientation of the accelerometers at all times. Given this information, it is possible to resolve the accelerations into the reference frame before the integration process takes place. Hence, inertial navigation in the process whereby the measurements provided by gyroscopes and accelerometers are used to determine their position of the vehicle in which they are installed. By combining the two sets of measurements, it is possible to define the translational motion of the vehicle within the inertial reference frame and so calculate its position within it.

Unlike many other types of navigation system, inertial system are entirely self-contained within the vehicle, in the sense that they are not dependent on the transmission of signals from the vehicle or reception from an external source. However, inertial navigation systems do rely upon the availability of accurate knowledge of vehicle position at the start of navigation. The inertial measurement are then used to obtain estimates of changes in position which take place thereafter.

2.2 Inertial navigation process

Therefore the inertial navigation system, INS, is made from a navigation computer and a set of gyroscopes and accelerometers that measure in Newton’s inertial axis. The group of inertial sensors is commonly called inertial measurement unit (IMU). Once aligned to a set of reference axes (such as the North-East-Down set), the sensors provide distance measurements and the navigation computer carries out the continuous dead reckoning calculations. The inertial sensors might be mounted in a set of gimbals so that



they stay level and head in a fixed direction no matter how the vehicle moves, so that they are mechanically isolated from the rotational motion of the vehicle. This construction is called a platform system. Alternatively, the instruments might be attached to the vehicle, in which case they measure its motion components in the vehicle axis set, and the system computes direction travelled in the reference axis by transforming the measurements from the vehicle axis to the reference axes. This is called stapdown system, jargon for instrument “strapped down” to the vehicle. These systems have removed most of the mechanical complexity of platform systems, but the major penalties incurred are the substantial increase in computing complexity and the need to use sensors capable of measuring much higher rates of turn. However, advances in computer technology combined with the development of suitable sensors have allowed such design to become realty. Our system is just a strapdown system.

To navigate inertially, we first measure the accelerations in the directions of the navigation axes, and if our instrument are not perfect, we might compensate their reading by removing bias or scale factor errors, perhaps known as a function of the system’s measured temperature.

Second, to find the vehicle’s vertical acceleration, we subtract gravity from the “Down” accelerometer output, perhaps using a gravity model to allow for the variation of gravity with latitude and longitude.

Third, we integrate the accelerations over a known time, once to get the velocity, twice to get the distance travelled. For periods of constant acceleration we can apply the equation of motion to find the distance travelled, s:

where a is the acceleration, v is the velocity after time t and v0 is the initial velocity.

Because integration is the process of summing the outputs at frequent, known interval, we must know the time interval accurately as it enters as a squared term in the distance computation.

Fourth, we measure the rotation rates directly with gyroscopes in a strapdown system. We then compensate for gyro bias and possibly scale factor errors, and we determinate a new heading.

Fifth, we compensate for earth rotation if we are in local level axes; otherwise the platform would be space stabilized and would seem to tilt in the vehicle axes set.

Finally, the combined distance and heading data give us an update dead reckoning position to display. Then we go back to the beginning and do it all over again, until the end of the journey.



2.3 Mechanization

The reduction of the navigation equations and their resolution may be done in different possible frame other than the inertial frame. This conversion in known as mechanisation. In this way it is possible to separate the vertical and horizontal channel behaviour making easier the analysis of the platform behaviour and test phase, and mainly isolating the intrinsic instability problem relegating it only to one channel. Just for this intrinsic instability of the vertical channel we decided to insert a pressure sensor in our experiment.

The variations in mechanizations are in the strapdown computational algorithms and not in the arrangement of the sensors or the mechanical layout of the system.



3 EXPERIMENT DESCRIPTION

3.1 Hardware

Our experiment consist mainly in a small and lightweight electronic board where all sensors and the relative control electronics component are mounted on. All measurements are gathered through a Microchip PIC microcontroller, stored into a Flash memory and then sent via RS-232 to E-link TM/TC module in order to be downlinked to ground. Moreover a rough estimate of the position will be carried onboard and sent to ground via telemetry. Once the experiment has been recovered, the memory will be dumped and an exhaustive post-processing will occur with all the data gathered by the unit.

3.1.1 Electrical design

The core of the experiment is the ANALOG DEVICES ADIS16355 Inertial Measurement Unit. This is an highly integrated sensor that has three accelerometers and three gyros, aligned over three orthogonal axis, with the relative signal amplification and conditioning embedded inside, and provides digital outputs over the SPI (Serial Peripheral Interface) 3-wire bus. Moreover, the device comes already factory calibrated and compensated over a wide temperature range (-40÷80°C).

The magnetic field sensors (single-axis FGM-1, two-axis FGM-2), built by “Speake & Co Llanfapley”, provides digital output, as the output signal is a square wave whose period is directly proportional to the field strength being measured. We chose to use digital output sensors because they can be easily interfaced with the microcontroller without need for any other components, moreover they offer high noise immunity level. The magnetometers are provided with a built in overwound coil intended for feedback systems, biasing or improve linearity of the sensors. Since there is a common ground, negative voltage are required in order to inject current in the coils in both directions.

The only analogue sensor is the Honeywell ASDX015A24R absolute pressure sensor that is connected to an analogue port of the microcontroller. However it provides a robust 4 V voltage span over the entire measurement range.

Every sensor is connected to a Microchip PIC 18F2620 microprocessor. It has 10 channel 10 bit ADC, hardware I2C, SPI, UART interfaces, as well as counter modules.



Data gathered from sensors are sent through RS232 interface to E-link TM/TC module. However the same RS232 interface can be used for service purpose during on ground tests.

All data are also stored onboard on four ATMEL AT45DB321D 32Mbit flash memories. Since they have 3-wire SPI interface compatible with ADIS16355 IMU, they can be easily connected together over the same SPI bus, providing a simple circuit layout design.

The power supply, provided by battery pack, is regulated to 5 V, 3.3 V and -5 V using a DC-DC converter.

In the following figure is represented an overall functional block diagram of the experiment.

Figure 1 – Functional block diagram

In annex, under drawings, is depicted a detailed electric schematic with every component needed. Decoupling and filtering capacitors are placed along the power supply line near every devices. ICL232 IC is required in order to have correct RS232 levels from TTL serial output coming from the microprocessor.

Magnetometers outputs are connected to the microprocessor CCP1 pin through a multiplexer (74HC153) and a 4-bit ripple counter (74HC93) used as frequency divider (by 16) required to increase magnetometer’s resolution.



IRL2203N is a low resistance logic level power mosfet used to switch on and off the heater. The thermal control is assured by 2 resistors controlled by the same logic output of the microcontroller through a buffer.

PCB drawings are in annex. Below few pictures of LowCoINS board.

Figure 2 - LowCoINS board



3.1.2 Prototyping

The development of the experiment started with a prototype that allowed us to develop and test the electronics. A modular approach has been adopted to integrate one by one every device needed in the experiment. Below is shown a picture of the early prototype:

Figure 3 - Prototype layout

The area dedicated for each component on the prototype is likely to be the same on the final PCB. In Circuit Serial Programming (ICSP) allows for easy program memory programming in software development phase. As stated earlier, a modular approach is adopted. SMD components are mounted on the prototype through dedicated PCB adapters.



Figure 4 - IMU Adapter

The pictures represents the IMU adapter module that allows for easy integration on the prototype board on standard 24 pin DIL socket. The same approach will be used for memories and for magnetometers that have non standard spacing between pins.

Figure 5 - Prototype Integration



Here few pictures of the growing of the prototype, below the prototype fully integrated.

3.1.3 Mechanical design

Our experiment will be contained in a aluminium box properly insulated with polystyrene panels. The housing has to withstand thermal and mechanical loads during the various mission phases. The box structure is basically a sandwich built with foam core and aluminium skins. The external aluminium case provides a first thermal and mechanical protection as well as an interface for experiment mounting on the payload gondola. The internal aluminium case provides support for internal components: battery pack and an electronic board.



Figure 6 – Experiment box description

Aluminium boxes are built with panels connected with “L-shaped” corners and rivets. Foam panels are placed in between and glued to the metal cases.



Figure 7 - Housing

The box is provided with brackets on the 4 sides. All of them are required to attach the experiment to calibration devices, mainly tilt and rate table. Connection to the payload gondola is assured by two beams placed under the experiment box. Two sets of mounting bar have been developed to match the gondola rails in both side of the gondola, the larger and the narrower (320 – 375 mm).



Figure 8 - Experiment with mounting bars

Detailed CAD drawings of the experiment box and attachment system are in Annex under “drawings” section. Please note that the dimensions reported in drawings slightly differ from actual total size of the box because they do not include corners and rivets thickness.

FEM structural analysis demonstrates that the experiment box can withstand to all loadings that it will encounter during the flight. The worst case mechanical load is encountered when the parachute opens, and the experiment will be loaded with 10 g along the Z-axis. The situation has been simulated using FEM software. The model is a box made of sandwich panels where the core has no bending stiffness; The constrains are no rotation and no translation for all the nodes in the bottom panel. The results are shown below:



Figure 9 - Total translation

The maximum total translation is 0.1 µm for the nodes that are in the centre of the top panel.

Stresses also are low, not exceeding 36 kPa, well below the yield stress of the aluminium.

Figure 10 - VonMises Stress



Loads due to thermal gradient has been evaluated, particularly stresses generated on constrains (where the box is connected with the payload gondola) by shrinks of the experiment box as the temperature decrease considerably. However, the values are so low that are not a reason of concern.

3.1.4 Dimensions

Considering the thickness of rivets and “L-shaped” corners, the overall experiment dimension is 21x21x17 cm. The final mass is 3300 g.

Figure 11 – Closed experiment box



3.1.5 Component list (detailed component list in annex)

QTY

LM2575T-5 1

LM3940IT-3.3 1

Analog Devices ADIS16355 1

Pressure sensor - ASDX015A24R 1

Magnetometer FGM 1 1

Magnetometer FGM 2 1

Microchip PIC 18F2620 1

ATMEL AT45DB321D 4

INTERSIL ICL232 1

Table 1 - Electronic component list

3.1.6 Mass budget with single components

Mass (grams)



Magnetometers - FGM 1 + FGM 2 5

PIC 18F2620 6

ATMEL AT45DB321D 1

LSH 20 battery pack (3 cell) 300

Housing 2500 (aluminium+ insulation) (21x21x17 cm - 0.2 cm thickness)

Electronic board with heater 200

TOTAL 3300

Table 2 - Mass budget



3.1.7 Environment protection - Thermal

In order to maintain electronic components in their operative optimal temperature ranges, to reduce drifts in bias and calibrations due to temperature changes, a thermal control system will be adopted to maintain the temperature inside the experiment box as constant as possible close to 0 °C. To maintain the internal temperature not below 0 °C, evaluation of the overall thermal dissipation from the experiment box to external environment that, in the worst case, can be as low as -90 °C is obtained through mathematical models and thermal test in thermal chamber.

Considering heat transfer that occurs by conduction through the walls of the experiment box, a value of the thermal power that flows out has been obtained considering a temperature difference between the inside and the outside of 90÷100 °C. Results shows that is necessary to provide a total thermal power of about 10 W in order to maintain the thermal gradient required. These calculations have been confirmed by experimental results, where a fixed amount of thermal power was supplied inside the experiment insulation box and the equilibrium temperature has been measured. Tests shows that a ∆T = 10 °C is obtained for every Watt provided, confirming the theoretical solution. However this occurs only in the worst condition and means that heater has to be controlled by the microprocessor, driving an on-off switch. The heater is placed above the component side of the electronic board.

Figure 12 - LowCoINS with the heater, ready to be mounted in the experiment box



Heat transfer by radiation has not been considered because we believe that will slightly lower the power need rather than increase it. Moreover, as the air flows out with increasing altitude, severe thermal gradient may occur due to the absence of free convection. Overheat problem has been considered for the early phase of the mission. Particularly, the heat generated by the experiment itself, can lead the internal temperature to rise while the balloon is still on the ground preparing to be launched. To avoid this event, a low power consumption mode is foreseen for the experiment during the launch preparation. However, tests demonstrate that even one to 2 W power do not heat up too much the experiment box.

The thermal control is obtained placing an heating plate above the PCB component side. Two independently driven resistors provide 10 W of power on an anodized aluminium plate. Anodized aluminium is used because of its high emissivity coefficient.



3.2 Power

Power budget for the experiment is shown in the following table.

Current Consumption

(mA) @ 5V



Magnetometers - FGM 1 + FGM 2 24

PIC 18F2620 2

ATMEL AT45DB321D 12

ICL 232 10

Contingency 50

TOTAL 150

Table 3 - Power budget

During test the experiment showed a power consumption of 70 mA at 11 V in normal operations and 1.15 A with heater on.

The power requirement for the electronics is less than 1 W, however this calculation does not take into account the power required for thermal control. As previously shown, 10 W are required to keep the temperature above 0 °C in the worst condition (external temperature = -90 °C).

Thus, the overall power required is about 10 W. We chose to use SAFT lithium-thionyl chloride (Li-SOCl2) LSH 20 cell. The nominal voltage and capacity for this cell are 3.6 V and 13 Ah respectively. Using three cell the nominal voltage reaches 10.8 V, so slightly less the one Ampere current is necessary to meet the power requirement.

The battery pack provides 10.8 V with a nominal capacity of 13 Ah, however cell voltage drop and actual capacity at working conditions have to be considered. First of all, effective cell voltage is lower than the nominal voltage due to its own internal resistance and varies with current and temperature. Assuming that cells operate at 0 °C at a discharge



current of 1 A, working point can be defined and cell voltage at working condition can be obtained.

Figure 13 - Cell voltage

The graph shows that cell voltage is approximately 3.25 V, that means 9.75 V for the entire battery pack, in turn almost 10 W are successfully generated by the battery.

Similarly, actual capacity can be obtained for the working condition.

Figure 14 - Capacity



As the graph shows, the capacity drops to 7 Ah. However, the capacity is widely sufficient to supply the experiment for 7 hours of operations with the heaters continuously switched on, situation that it will not likely occur.

Power supply to electronics is regulated with a 5V DC-DC switching regulator. A common National Semiconductor LM2575 fixed voltage buck regulator is used requiring only few components to build an efficient DC-DC converter. It operates at 52 Khz frequency and an additional LC filer is placed on the output for extra filtering. There are components requiring 3.3 V and is obtained using a low-dropout voltage regulator because of the small current (12 mA) required at this voltage. A negative (-5V) voltage has been implemented using a switching regulator and is required for magnetometers feedback coils.



3.3 Data Handling

3.3.1 Data Acquisition

The main task of the microprocessor is to acquire data from sensors. Sampling frequency is set to 20 Hz for most critical measurements. Data to be recorded are accelerations (3-axis), angular rates (3 axis), magnetic field (3-axis), atmospheric pressure and internal temperature. A counter generates an interrupt every 50 ms that makes the acquisition start, and all operations must be completed before the next interrupt occur.

ADIS16355 provides readings upon reception of two byte data containing appropriate “read” command, basically in the bytes has to be specified which measurement is required (X Acc, Y Acc, Z Acc, X Gyro, Y Gyro, Z Gyro). The unit replies sending two byte data containing the information required. The time needed to complete the acquisition of one measurement is the time needed to transfer 4 bytes over the SPI bus. Since the SPI bus clock speed is set to 1.25 MHz, 54.2 µs are necessary to complete one read operation. Six different measures have to be read from the IMU, so 307.2 µs are required to complete the process.

Measurements of the magnetic field component are carried out mainly measuring the period of the square wave output of the magnetometers. This is done by the CCP modules of the microprocessor that provides the value of an internal 16 bit counter every rising-edge (or every 16 rising-edges) of the incoming signal. The measurement of the period is obtained making the difference between the actual and the previous counter value. The process is regulated with dedicated interrupt routines, thus a value of the measurement is always ready for acquisition.

Pressure measurements are done mainly by the ADC module of the IMU, thus the time needed is the same for inertial sensors.

Temperature measurements are obtained by an LM335 analogue sensor and from ADIS16355 IMU that has got a temperature sensor inside. Thus, the acquisition time will not exceed 60 µs.

3.3.2 Onboard software

The onboard software is very simple and straightforward. There is a main loop that occur every 50 ms, regulated by a microcontroller timer, that holds all the principal tasks.



Figure 15 - Onboard software

The unit has two main state: “Flight Mode” and “Service Mode”. The Flight Mode is the default state. In this mode the unit acquires data from sensors, sends telemetry and writes to memory, if enabled. It is able to receive a limited number of telecommands, and the TC reception is assured by periodically polling the RX buffer. In Service Mode, the unit is able to receive long telecommands for calibrations purposes. In this state the TCs are handled with an Interrupt Service Routine (ISR) because the RX buffer is not big enough to receive the entire TC. This causes the telemetry packet to be corrupted when a TC is sent to the unit, and this is the main reason for the existence of two modes, “flight” and “service”. While in Service Mode, the unit can go to the “memory dump mode” state. In this case the data acquisition is halted until the unit remains in memory management mode.



Figure 16 - Flight Unit operational modes

There are three interrupt service routines:

• Real time clock ISR: the routine is entered to alert the main program that is time to acquire data from sensors. If 20 Hz sampling frequency is selected, the interrupt occurs every 50ms.

• Capture ISR : is an interrupt that occurs every rising edge of the incoming signal from magnetometers. It is used to make frequency measurements of magnetometers output.

• RX buffer ISR: is a routine that store the telecommand when received from e-link or service PC and makes it available to the main program for proper handling. This interrupt is enabled only in when the unit is in “Service Mode”

A C compiler (CCS) has been used to program the microcontroller.

3.3.3 List of Telecommands

LowCoINS experiment is able to receive either from e-link serial port or service serial port several telecommands. A detailed list of telecommands with the relative format is shown below.

When the unit is in “flight mode”, it is able to receive 2 bytes long TC without the need to fire an interrupt, as the microprocessor has a 2-bytes buffer on the receive pin. The first byte identifies the telecommand type, the following two bytes provide the value to be transferred where foreseen. Below, there is a table with a Short TC list.



Telecommand type 1st byte (Hex) 2nd byte Heater On 21 - Heater Off 22 - Heater Auto 23 - Set Temp. Low threshold 24 value Set Temp. High threshold 25 value Write to memory 50 stop writing to memory 55 AA erase memory 60 9F go to Service mode 80

For critical TCs such as “erase memory” and “stop writing” a second byte (one complement of the first) is transmitted in order to avoid unwanted transition in these states.

In Service mode the TC packet expected is 3 byte long. The first byte identifies the telecommand type, the following two bytes provide the value to be transferred where foreseen (eg. 1st byte=set bias for X gyro, 2nd & 3rd byte: bias correction value in two’s complement format).

Telecommand type 1st byte (Hex) 2nd byte 3rd byte Automatic Bias Null bias calibration 36 0 0 Precision Automatic Bias null cal. 37 0 0 X-Gyro Bias Calibration 30 value value Y-Gyro Bias Calibration 31 value value Z-Gyro Bias Calibration 32 value value X-Acc Bias Calibration 33 value value Y-Acc Bias Calibration 34 value value Z-Acc Bias Calibration 35 value value Dynamic Range/Digital Filter Control 38 value 0 Sample Rate 39 value 0 Enter memory dump mode 69 0 0 Read all memory 70 0 0 Read bank 1 70 1 0 Read bank 2 70 2 0 Read bank 3 70 3 0 Read bank 4 70 4 0 Read bank n, from page m, for k pages tbd



Thermal Control Telecommands

TCs regarding TCU are needed to switch the heater on and off manually and set the temperature threshold that trigger the heater switch when the TCU is in automatic mode.

Heater On/Off TCs switch the heater on/off regardless the actual temperature read on T0 and threshold temperatures. The Heater will remain in the on/off state until another on/off TC is received. When “Heater Auto” TC is received, the control of the heater is restored to the automatic mode and the heater will be switched on or off accordingly to the temperature thresholds.

TLT & THT (Temperature Low/High Threshold) TCs can set the temperature thresholds that trigger the heater switch. The TC is made by a first byte that specifies if the High or Low threshold has to be modified, the second byte specifies the value as two’s complement format over 8 bits, the 3rd byte doesn’t care.

Calibration Telecommands

Automatic Bias Null Calibration

A single-command, automatic bias calibration measures all three gyroscope output registers, then loads the three bias correction registers with values that return their outputs to zero (null).

Precision Automatic Bias Null Calibration

Another option for gyroscope calibration, which typically provides better accuracy, is with the single-command, precision autonull. This incorporates the optimal averaging time for generating bias correction factors for all three gyroscope sensors. This command requires approximately 30 seconds to complete. For optimal calibration accuracy, the device should be stable (no motion) for this entire period.

Manual Bias Calibration

Using this set of TCs, the user can manually set the bias correction for accelerometers or gyros. The TC format is as follows: 1st byte: set the inertial sensor to be calibrated, 2nd & 3rd bytes specifies the correction value in two’s complement format over 12 bit (accelerometers) or 13 bit (gyros).

Memory Dump Telecommands

Enter memory dump mode



Makes the unit to enter in memory dump mode; the telemetry transmission is suspended and a synchronization frame is transmitted every second.

Read All Memory The entire memory is dumped by the unit.

Read Bank N Only the bank N is dumped.

Read bank n, from page m, for k pages It allows for read a part of memory bank.

3.3.4 Ground software

Software for live data recording has been developed using National Instruments Labview 8.5. It allows to manage every aspect of LowCoINS experiment:

• Live data recording

• IMU management, sensor sampling rate, digital filtering, calibration and bias set up

• Memory management: memory erase, memory dump

• Thermal control unit management: allows for change temperature threshold that triggers the heater also during the flight, and allows for manual control of the heater.

• Estimated remaining battery charge

Three different versions have been developed:

1. Flight mode: to be used for live data recording during the flight. This version has the ability to send only the TCs needed during the flight, does not have the TCs for the calibration of the unit.

2. Service mode: to be used to calibrate the sensors and debug.

3. Memory Dump: to be used to read the data from the flash memory and save in the PC.

Below few screenshots of the different software:



Figure 17 – LowCoINS GSS Service Mode screenshot

Figure 18 - LowCoINS GSS Flight Mode



Figure 19 - LowCoINS GSS Memory Dump Mode

3.3.5 Data Rate & Storage of Data

In order to have a simpler data acquisition software every measurement is gathered with the same sampling frequency, however only nine measurement are downlinked and saved to memory with this frequency. The packet of data (frame) to be stored in the memory is organized as follow:

Figure 20 – Memory Data Frame

The last three bytes contain cyclically ten different data, the resulting update rate for these measurements is 2 Hz.



Figure 21 - Cyclic Data Bytes

Since the flash memories are organized in 8192 pages of 528 bytes size, each data frame is 24 bytes long, in this way 22 different frames exactly fit the page memory size of 528 bytes. Using four Atmel AT45DB321D 32Mbit memories the total amount of data stored reaches 10 hours of recordings.

The unit begins to record data in the memory upon reception of the “Start of acquisition” telecommand.

3.3.6 Communication, Data Link & Protocol

The experiment is connected to E-link TM/TC module through RS-232 interface. The primary function is to downlink data collected to the ground station. The telemetry frame layout is the same as the data frame to be recorded in memory showed earlier. The only difference is the presence of the byte 25 that contains a checksum for error detection. The checksum byte is simply the 8 bits sum of the 24 data bytes.

Figure 22 – Telemetry Data frame with Checksum byte



Serial communication is a standard 9600 bps, 1 bit stop, parity none, no hardware flow control setup. The protocol used to identify each data frame is the DLE-STX/ETX protocol. Each data frame starts with two control bytes DLE-STX (10-02 hex) and stops with DLE-ETX (10-03 hex). If in the data frame any byte is equal to “DLE” (10 hex), then it is sent twice. Adopting these rules, the start and the end of the frame can be recognised by the ground software without errors, as DLE-STX and DLE-ETX sequences can only identify the frame boundary since if they are encountered in the data frame an extra DLE is present.

Figure 23 - DLE-STX protocol example



3.4 Operation

3.4.1 Experiment Connections

The experiment presents 3 MIL connector on the top of the box SIGNAL, PWR, TEST.

Figure 24 - Experiment connectors panel

SIGNAL is used to connect the experiment to an RS-232 port of the E-link TM/TC interface. The connector used is MIL-C-26482, 8 pin socket, size 12, orientation normal, in order to fit the mating connector of the harness.

PWR is a 4 pin socket MIL connector (type: D38999/20WB4SA) used for experiment power up. It can be used either as a switch (with a special mating connector that has shorted pins) or as external power port.

TEST is a 4 pin socket MIL connector (type: D38999/20WB4SN) used to connect the experiment to a PC for service purpose and for a final pre-flight check. When this



connector is populated, the unit can receive telecommands only from the service line, while data from the unit will be transmitted on both (telemetry and service channel)

The connectors have three different keys in order to avoid mistakes.

3.4.2 Assembly and checkout of experiment

• Experiment integration on the payload gondola

• Misalignment check between gondola and experiment box

• PC connection to TEST port

• Connection of external power source to PWR

• Check for normal operations

• Perform a sensors calibration

• Check for correct readings

• Check for reception of telemetry packet by the E-link ground station

• Perform telemetry/telecommand functional test

• Send telecommand “write to memory”

• Send telecommand “stop writing to memory”

• Memory dump

• Memory erase

• Power cable disconnection (PWR)

• Service cable disconnection (TEST)

3.4.3 Optional Compass Calibration

In the days preceding the launch, we would like to perform a simple procedure to check the magnetometers calibration in a magnetic field different from which they were calibrated before. This procedure consist in place the gondola in the launch field in a levelled attitude and point it in the magnetic north direction. Then, it is rotated in 8



different position about the vertical axis 45° away each other. The operation should not take more than one and an half hour.

3.4.4 Operations time line

Experiment checkout T-1:10

• Connection of the computer to service port TEST.

• Experiment powers up on its own power source connecting the switch connector on PWR.

• Check for correct readings from sensors

• Gyros zero-point calibration (experiment requires to be steady for about 30 seconds)

• Erase memory command

• Check for correct readings from sensors

• Check for reception of telemetry packet by the E-link ground station

• Optional: send “write to memory” command

• Service cable disconnection

• Go for launch!

T-0:10

• Optional: Send telecommand “write to memory” (if not submitted before)

Since the memory capacity is 10 hours, the command to begin the writing to memory can be sent to the unit even hours before launch, without take the risk to send this TC via telemetry.

After landing if telemetry connection is still available, “stop writing to memory” telecommand will be sent. The unit now ends data acquisition and waits for recovery.

If telemetry channel is not available, unit ends data acquisition automatically when the memory is full.



3.4.5 Maximum allowable launch delay time

The experiment is switched on about one and an half hour before launch (in normal operation). In case the launch is postponed it is desirable to access the launch area and switch the experiment power supply to external power supply in order to save battery and keep the experiment warm. The operation can be accomplished using an external battery pack connected to the PWR port. Considering the experiment power consumption, this operation becomes necessary only if the launch is delayed by more than three hour. However, it depends upon the external ambient temperature, since if the weather is cold and the heater needs to be switched on while waiting for launch, the maximum allowed launch delay time can decrease considerably. As a guideline for the launch day we use 5 hours as the maximum on time period prior to launch. If this time period is exceeded, the substitution of the battery pack becomes necessary. We estimate that this operation takes about 20-30 minutes to complete. A complete memory erase can be operated at this time.

3.4.6 Post Flight operations

Disconnection of the connector PWR (power switch) should be operated by the recovery personnel if possible.

Memory dump will be held after recovery.



4 SYSTEM PERFORMANCE ANALYSIS

The accuracy to which a strapdown navigation system is able to operate is limited as a result of errors in the data which are passed to it prior to the commencement of navigation, as well as imperfections in various components which combine to make up the system. The sources of error may be categorised as follow:

• initial alignment errors; • inertial sensor errors; • computational errors.

In general, inertial navigation system performance is characterised by a growth in the navigation error from the position co-ordinate values which are initially assigned to it.

For simplicity, now we consider a two-dimensional case and then we assume navigation to take place in the local geographic reference frame. In this system, the x and z reference axes are coincident with the local horizontal and the local vertical respectively, and the navigation system provides estimates of velocity in each of these directions.

The error dynamics of the local geographic system is analysed for the condition where true body attitude is zero, that is, θ=0. In this case, the coupling between the channels is normally zero, allowing each channel to be analysed separately. In addition, it is assumed that the navigation system in mounted in a vehicle which is at rest to the Earth, or one which is travelling at a constant velocity with respect to the Earth. Under such conditions, the only force acting on the vehicle is the specific force needed to overcome the gravitational attraction of the Earth. In this situation, fxg=0 and fzg=g.

The errors equations, corrected to first order, for the vertical and horizontal channels, can now be written as shown in the following table.



Error source

Position error

x-axis

z-axis

Initial position errors δx0 δx0 -

δz0 - δz0

Initial velocity errors δvx0 δv0t -

δvz0 - Δvz0t

Initial attitude errors δθ0

Accelerometers biases δfxb

δfzb

Gyroscope biases δωyb

Table 4 - Propagation of errors in two-dimensional strapdown inertial navigation system

Substituting the values of biases provided by the IMU constructor in these equations, the position error would be as high as 20 km after one hour of flight! This result shows that an augmentation system is compulsory when using low cost inertial sensor for navigation purposes. Augmentation is achieved adding sensors that provides auxiliary data, in particular, a pressure sensor used as a barometric altimeter and magnetometers used as auxiliary source for attitude data. Even with these augmentation systems, the position error will grow at unacceptable level, so will be necessary a continuous integration with GPS data. The estimate of the overall position error of our system is one of the main goals of our experiment, so it cannot be estimate a priori.



5 PROJECT PLANNING (PHASE B AND C)

5.1 Resource estimation (costs)

A table with unit price for every components involved in the experiment as well as the supplier chosen in listed below:

Component Unit Price (€) QTY Total Price (€) Supplier ANALOG DEVICES ADIS16355 410,67 1 410,67 Digi-Key Microchip PICF2620 13,10 1 13,10 Distrelec ATMEL AT45DB161D 1,66 4 6,64 Digi-Key FGM-2 37,00 1 37,00 Speake Sensors FGM-1 21,00 1 21,00 Speake Sensors Honeywell ASDX015A24R 16,81 1 16,81 Digi-Key INTERSIL ICL232IPE 5,15 1 5,15 RS-Components MIL-C-26482 series I 24,48 1 24,48 Farnell-Italia LM2575T-5.0 1,87 1 1,87 RS-Components LM3940IT-3.3 1,70 1 1,70 Distrelec SAFT LSH20CNR 15,50 3 46,50 Blu Batterie TOTAL 584,92

Table 5 - Detailed component list

Housing, spare parts and minor components costs are not yet listed. All costs will be sponsored by university and Northrop Grumman Italia S.p.A..



5.2 Support and Facilities

A collaboration agreement between LOWCOINS Team and Northrop Grumman Italia S.p.A. has been arranged.

• The collaboration will provide us with:

o Facilities for experiment building & testing

o Thermal chamber

o Thermal-vacuum chamber

o Shaker

o Static and dynamic test

• Support for:

o Environmental and functional tests execution

o Analysis of test results

o Advices for mechanical and electronic design and manufacturing

Northrop Grumman Italia S.p.A. is our main sponsor and covered the majority of

the project expenses. We thank NGI for the support and experience that made available to us.



5.3 Time schedule

Overall time schedule of the project development is in annex.

5.4 Task definition

Paolo:

Electronic designer

Onboard software developer

Power Interface

Mechanical interface

Post-processing software programmer for attitude and position determination

Maria Cristina:

Ground station software developer for TM/TC handling

Real-time computation

Mechanical design

CAD designer

Webmaster


Emanuele:

Thermal interface

Thermal and Structural analysis


Team-members are currently working on the project at NGI facilities at full-time. Task definition with planning is in Annex under “Time Schedule” section.



5.5 Test plan

Extensive test plan has been scheduled with NGI, where instrumentations and devices for thermal, vacuum, vibration and dynamic tests is available. Final checkout test include: flight simulation in thermal-vacuum chamber, sensors calibration, misalignment verification with tilt and rate table.



5.6 Risk management

5.6.1 Risk analysis – Project development

Project ended.

5.6.2 Risk Analysis – Safety

Our experiment is safe for operators and third persons. It does not contain any RF devices and does not transmit at RF, no moving parts or hazardous materials. Risk of fire is null as the power available for the experiment is low, and battery cells are provided with an internal 5A fuse that effectively protect from short circuits and unexpected high current drain.

5.6.3 Risk Analysis – Components

A risk analysis at component level is shown below. The analysis is conducted considering all the components that takes place in each subsystem; The risk level for each component is evaluated giving risk “high” to the component that is “life critical” for the experiment. The following column explain what will happen if that component fails, then there is a note that explains in details consequences of a failure and the countermeasures adopted. The last column gives how much the component is reliable (based upon experience and our guess).

Subsystem Component Risk level

Consequences in case of failure

Note Confidence - reliability

Power Supply Battery high experiment failure SAFT battery

provides high reliability in extreme environment

very high

Battery connectors high experiment failure good connector with key (to prevent reverse polarity), retention mechanism (to prevent acidentally

high



disconnection) are needed. Experiment switch using MIL connector is necessary to have the sufficient level of reliability

Switching Regulators

high a failure on +5 V will cause experiment failure

All components involved in power supplies are overdimensioned (10 times or more)

very high

low a failure on -5 V will cause magnetometers reading to be less accurate (reduced linearity)

very high

low a failure on 3.3 V will cause the loss of memory

very high

Thermal Control

Microcontroller high experiment failure very high Resistors medium partial loss of TCU

cababilities there are 2 resistors wired in parallel, if one fails, the remaining will prevent the experiment from freezing. The resistors (maximum power) are overdimensioned

high

Mosfet medium partial loss of TCU cababilities

The resistors are driven with dedicated mosfet, if one fails the remaings keep operating. If a mosfet breaks remaing closed, the resistors itself will limit the current, however only a complete failure (2/2) will cause a possible experiment failure (overheat). The mosfets are widely overdimensioned (50 times or more)

high



Buffer medium complete/partial loss of TCU

prevents glitches to reach the microcontroller, will cause complete loss of TCU cababilities only in case of total IC failure (unlikely)

very high

Thermal Sensors medium partial loss of TCU The heater can be controlled manually from ground if faulty readings comes from the sensor, providing a sufficent level of redundancy (overheat prevention). The IMU internal temperature sensor might be used to manually control.

very high

Sensors IMU high experiment failure high Pressure Sensor medium reduced accuracy high Magntometers &

relative components (MUX - COUNTER)

medium reduced accuracy TBD

ICL 232 & TM/TC

connectors low loss of TM/TC The onboard

memory is used to prevent loss of telemetry to cause experiment failure. (redundancy)

very high

Memories low loss of memory The telemetry channel is used to prevent experiment failure in case the memories break. (redundancy)

very high



6 OUTREACH PROGRAM

The web page online to promote our experiment and to show the current status of the project is online at: http://digilander.libero.it/lowcoins/

A web page on university server is online:

http://w3.uniroma1.it/palmerini/altre_pag/Students_ENG.html

We participated and presented our project to a workshop called “Science and Technology through long duration balloons “ organized by INAF (Instituto Nazionale di Astrofisica spaziale e Fisica Cosmica – National Institute for Space Astrophysics and Cosmic Physics), ASI (Agenzia Spaziale Italiana – Italian Space Agency), La Sapienza – University of Rome, Tor Vergata – 2nd University of Rome. More information available at: http://projects.iasf-roma.inaf.it/Balloons/

The poster of the event is in Annex, under “Outreach” section.

The proceedings of the Workshop have been published on the “Memorie della Società Astronomica Italiana”. In Annex the published article.

Through our web page, we established contacts with:

Luis Eduardo Pacheco responsible of “StratoCat”, http://stratocat.com.ar/

It is an interesting webpage regarding ballooning throughout the world. We have been in contact with him before, during and after the launch campaign, giving him updates and pictures.

Hector Garcia de Marina from Spacefish Project and Mateusz Wolski from SCOPE team, both involved with the future Bexus campaign giving them advices, suggestions, information about the flight as well as flight data.

We exchanged our data with Turatemp/Turawind team.

A presentation of our experiment and the Bexus campaign has been given to students of our university within the course of “Aerospace guidance and navigation systems”.

An abstract has been submitted to participate to the 19th ESA Symposium on European Rocket and Balloon Programmes and Related Research that will be held in Bad Reichenhall, Germany, 7-11 June 2009. In annex the submitted abstract.



7 MEETINGS AND REVIEWS

7.1 Preliminary Design Review – PDR

Flight: BEXUS

Experiment: LowCoINS

Date: 24-04-08

1. Review Panel

Olle Persson (chair) Eurolaunch / SSC Bruno Sarti ESA

Andreas Stamminger Eurolaunch / DLR Koen de Beule ESA

Harald Hellmann Eurolaunch / DLR Helen Page (minutes) ESA

Markus Pinzer Eurolaunch / DLR Maria Roth DLR

2. General Comments

PDR presentation Good presentation

SED content Generally good. Delete the instructions, add and delete abbreviations. Major points covered but need to add more detail so that the reader understands the whole experiment.

3. Panel remarks

Team Organisation Small team but seem to work well together. Restructure description of tasks in SED.



Planning / Time schedule / Tests Breakdown time-planning into smaller sections. Develop test plan from existing description. Perform drop test – no need for vibration test.

Ensure that memory can be reset as it might be triggered during launch tests.

Electronic Schematic is good – include active thermal regulation. Thermal control is critical to find out whether experiment needs to be heated (impact on power requirement and mechanical interface) Would like to do thermal vacuum test.

Mechanics Think about fixing polystyrene to aluminium and attaching box to gondola.

Safety Need to carry out proper risk analysis in enough detail.

Outreach Develop some ideas and a better outreach plan by the CDR. Website idea good – make sure there are good images.

4. Final remarks

Preparation for CDR Perform risk analysis and thermal analysis

Develop planning for launch campaign and especially the flight sequence of events

Improve and adjust your time schedule (planning phase C and D)

Necessary actions Olle to provide more information about gondola, including turn rates, fixing points, etc..

Use ELINK, rather than EBASS



8 EXPERIMENT REPORT

8.1 Experiment integration

In the following images, it is shown the experiment integration on the Bexus 6 payload gondola:

The experiment is mounted through its 4 main brackets. Below the LowCoINS experiment with its axis orientation with respect to the gondola.



8.2 Pre-flight tests

A firmware upgrade of the LowCoINS experiment has been done upon our arrival in Kiruna to fix some minor bugs discovered after the shipment of the experiment to Esrange. A long run test (overnight) followed the firmware upgrade to re-test all the experiment functionality. No problems have been encountered within this test.

The magnetic field present at the launch site is stronger than the field that the experiment has experienced during the development in Rome, so a check of the 3-axis magnetometer embedded in the LowCoINS experiment has been conducted to verify their functionality. The test was accomplished on the launch site, far away from any major magnetic disturbances (hangars, vehicle) except for the unavoidable disturbances from the orefield below the under the launch site! No problems have been encountered within this test.



A communication test occurred after the experiment has been connected to the E-link TM/TC interface. The purpose of the test was to check that all telemetry data were received correctly from the ground station, and telecommands sent to the unit were correctly executed. No problems have been encountered within this test.

8.3 Launch day

The experiment has been switched on at 5:52:25 UTC, approximately 1 hour and 18 minutes before Bexus 6 launch occurred at 7:10 UTC. We experienced few minutes of panic 30 minutes before launch due to an unexpected loss of telemetry. However we were able to access the experiment while it was still on the launch area, connecting a computer to the test port of the experiment, and we saw it was working correctly. The interruption of the telemetry occurred because E-link was switched off due to an interference discovered by Esrange personnel. Short after the telemetry was restored and worked for the entire flight.



8.3.1 Flight phases

Launch

Below, the graphs show accelerometers reading when the launch occurred. The time instant the gondola took-off is recognizable looking at this data because:

X and Y axis accelerations have noticeable variation correlated with X and Y gyro output showing the gondola was oscillating when the balloon was released and Hercules was running behind it few seconds before the gondola lifted-off.

The variation of the Z axis accelerometer shows a big difference before and after the lift-off. It can be noticed the z-axis reading is within a band essentially defined by the measurement noise before the launch. When the gondola is released, the band is wider, and this is because of the balloon is hauling the gondola with a force that it is not time constant.

Obviously, it can be recognized watching at our pressure sensor data that starts recording a decreasing pressure value.



Ascent-float phase

An almost continuous oscillatory movement of the gondola around the vertical axis was recorded by the Z axis gyroscope. No major movements and/or accelerations were recorded during these phases



The cut down

The readings from the z axis accelerometer show the exact moment of the cut-down, it can be noticed that the acceleration goes suddenly to zero g while the gondola was in free falling, than the maximum acceleration recorded was -1.5 g when the parachute deployed. From the accelerometers along the Y and X axis and from the gyros readings can be noticed that the gondola is tumbling and oscillating a little when the parachute deployed.



The descent

After a small transient (probably when the gondola reached an almost constant speed descent rate) an oscillatory behavior of the vertical accelerometer can be noticed. Particularly, the acceleration along the vertical axis changes periodically between -0.5 and -1.5 g. The period is almost 1 Hz and can be due to the elastic energy accumulated by the flight train during the deceleration phase short after the parachute deployed. Thus, the flight train in this phase is acting as a spring that is periodically compressed and extended.



The landing

The maximum vertical acceleration recorded during the landing has been of almost 3 g. We can say that it has been a real soft landing, and later watching at the pictures showing the landing site we can understand why!



Figure 25 – Bexus 6 recovery



8.3.2 Thermal control: performance and considerations

Below it is reported a chart showing the experiment operating temperatures for the entire flight. The experiment temperature never fell below 12 °C, and there was no need for the heater to go on. It should be considered that the experiment has been designed to withstand thermal conditions much worse (-90°C and night flight) than what has been encountered during the flight, where the lowest temperature has been of -60°C and day flight. Certainly, the thermal input from the sun, even if shielded by a white “tend” and heat inputs from the nearby experiments and/or equipments contributed to limit the thermal dissipation of the experiment, leading to an entire flight without the active thermal control intervention.

8.3.3 Communications

It has been possible to follow almost the entire flight from the ground using the telemetry sent down via E-link interface. No major problems have been encountered and the telemetry was lost approximately at an altitude of 1800m prior to land . It is normal to lose the link with the balloon just before the landing because the line of sight from the ground antenna to the gondola is lost because of the curvature of the Earth! Moreover, before the expected loss of signal, two telecommands have been sent to and successfully received by the unit. The telecommands have been sent to reconfigure the heater threshold temperatures , they have been lowered to 3÷5 °C from the nominal 10÷12 °C. The main reason was to avoid the intervention of the heater during the time lapse between the landing and the recovery, in order to avoid a deep discharge of the batteries, and, at the same time, prevent the experiment from freezing. However the recovery occurred few hours after landing.



8.3.4 Data storage

Even if most of the data have been recorded in real time from the ground station, once the experiment returned to the base, the entire memory dump occurred, to retrieve the data from the final part of the flight, in order to have the entire flight data. Only two memories were full of data over the four memories available, because of the relatively short flight time. The system was dimensioned for 10 hours of recording.

8.3.5 Flight performance

The experiment worked as expected, no malfunctions have been encountered.

Flight performance

Power Power system successfully met flight power requirements

Thermal control Experiment temperatures remained within expected operative ranges

Telemetry Continuous data stream received by the ground station

Telecommand Two telecomannds sent to the unit and successfully executed before LOS

Sensor data All sensors’ data successfully gathered by the unit

Data storage 65 Mbits of data stored onboard



8.4 Post-processing

8.4.1 LowCoINS navigation performance

The process of integrating the measurements form the experiment to obtain position and velocity is achieved developing and executing a complex algorithm offline after the flight.

Below are reported the equations of the mechanization used to compute the position and velocity of the gondola. They own to the classical levelled mechanizations process, where all the measurements are projected into a reference frame NED (North-East-Down).

The results of the LowCoINS experiment used as a pure inertial navigation system are briefly reported hereafter:



As can be seen, just processing few minutes of data, the calculated position diverges immediately. This is due to the bias of the gyros resulting in wrong projection of the acceleration from the body frame to the navigation frame that lead to a wrong velocity determination and in turn a wrong position determination. Together with the bias of the gyros, also bias of the accelerometers and low sensitivity of the IMU contribute to the poor result. However this results were widely expected because inertial sensor for navigation need to be order of magnitudes better than those onboard LowCoINS.

The alternative is to integrate the position and velocity determination from the IMU with that from a GPS receiver, using a sensor fusion algorithm i.e a Kalman Filter.

Kalman filtering is a statistical technique that combines knowledge of the statistical nature of system errors with knowledge of system dynamics, as represented as a state space model, to arrive at an estimate of the state of a system. The state estimate uses a weighting function, called the Kalman gain, which is optimized to produce a minimum error variance. For this reason, the Kalman filter is called an optimal filter.

Figure 26 – Kalman filter process

Integration of GPS and INS was made with a configuration known as loosely integrated or loosely coupled. This configuration typically includes a GPS receiver, which measurements are position and velocity form rather than pseudorange and pseudorange rate, an IMU, navigation equations to convert the gyros and accelerometer measurement from the IMU to platform attitude, position and velocity.



Figure 27 – INS/GPS integration method

In this way it is possible to have a correction of the wrong position and velocity calculated by the IMU with a more precise GPS data, but in the meanwhile, having an estimate of the position and velocity of the Bexus gondola in the time lapse between two GPS fixes, because of the higher rate of the IMU.

The figure below shows the flight profile, with the comparison between LowCoINS filtered data and GPS data of Bexus.



A zoom of the first part of the flight allows to appreciate the real difference

between the two type of data.



The next figure shows the Bexus’ trajectory obtained by integrating INS/GPS data through the Kalman filter.



Moreover, the error on the altitude can be bounded not only with GPS data but also with the data from the pressure sensor. In this way it is possible to have a continuous measurement of the altitude even when the GPS is no longer available, i.e. during the descent below 1800 m.

The position error between two successive GPS updates is in the order of 25 m. This error increases as the time lapse is increased.



8.4.2 LowCoINS as AHRS

Earlier in this chapter, we saw the quite poor performances of the experiment used as a navigator. This is exclusively due to the poor quality of the sensors that are not good enough for inertial navigation purposes. However, they could be good enough to realize a low accuracy Attitude and Heading Reference System (AHRS). The attitude of the gondola can be very useful for a wide range of balloon borne experiments, and it is an information that is not available from the Bexus standard avionics. The working principle of the LowCoINS used as AHRS is to integrate the angular velocity from the gyros to determine the attitude, while bounding the unavoidable drift of the attitude, mainly due to gyros biases, using the magnetometers and/or accelerometers as an auxiliary system for the attitude determination. The fusion of the attitude data computed from the integration of the gyros output and from the manipulation of the information from accelerometers and magnetometers is achieved using a sensor fusion algorithm such as a Kalman filter.



Figure 28 – Kalman filter block diagram

The dynamic of the attitude is defined in the state matrix of the Kalman filter A, and the predicted state is updated with the observables or measures of the attitude from accelerometers and magnetometers. One method to determine the pitch and roll angle of the gondola is through the readings of the X and Y axis accelerometers

However this method is usable if and only if the dynamic acceleration on the

vehicle is almost zero, in turn only if the total acceleration measured by the 3-axis accelerometer is merely the gravitational acceleration g. This is the case of a balloon flight, that has a relatively slow dynamic, and has been widely confirmed by our data. Once the pitch and roll angles have been determined, the yaw or heading angle can be determined projecting the measurement of the magnetic field from the body or sensor frame in the levelled frame (horizontal plane) and making the arctangent of in-plane components.

In case the total acceleration sensed by the system it is not equal to the gravity, this

means that there is a dynamic acceleration acting on the vehicle and the pitch and roll



angles can no more be obtained from accelerometers but only from angular rate integration. In this case, the measurements (roll, pitch and yaw from accelerometers and magnetometers) are not available and the attitude is computed only integrating the turn rates from the gyros corrected with the latest bias available estimated by the Kalman filter.

Figure 29 – Bexus body frame

Hereafter are reported some flight phases. Below it is a graph showing the attitude of the gondola at lift-off. It can be noticed that the gondola is tilted about 15 deg about its pitch axis before returning in a levelled attitude at the moment of liftoff. This is due to the oscillation of the gondola as Hercules moves before the gondola is released. After the liftoff, the gondola begins to oscillate about its vertical axis. Note that the discontinuity in the yaw angle is due to the representation used for the yaw: the yaw in our assumption is between +/-180 deg, meaning that the solution jumps from 180 to -180 or vice-versa as the gondola points toward south. Zero degree for yaw means North, +90° means East, -90° means West.



During the ascent and float phase no major oscillation have been recorded. The

gondola remained levelled with oscillation around its vertical axis.



The attitude determination at the cut-down as well as at the landing, is uncertain and not reliable because of the large dynamic accelerations sensed by our sensors in these moments. It is noticeable a spike, particularly on the pitch and roll angles, corresponding to the time instants of the cut-down and the landing. Further tuning of the Kalman filter is required to have a better estimate of the attitude in these phases.

The following graph shows the attitude at the landing. Except for the spikes

exactly at the touchdown, the attitude estimation is meaningful, and can be noticed that the gondola remains a little bit tilted after the landing, probably because of the terrain under the gondola.



8.4.3 Future improvements & conclusions

A fine tuning of both the Kalman filters, the one used for the position and velocity determination and the one used for the attitude determination, is needed. Further improvements and tests on the covariance and filter gains are needed to improve the solutions. This is just a post-processing task.

Moreover lab tests are ongoing on the unit that flew onboard Bexus to have an evaluation of the accuracy of our computed solutions. This is particularly interesting for the use of LowCoINS as AHRS, as for position and velocity determination we already have a good benchmark provided by the GPS data from Bexus avionics. The tests are ongoing using a good navigation grade AHRS, used as reference, attached to the LowCoINS experiment. First results show that the accuracy of our system is in the region of 5 degree for the pitch and roll angles. The yaw is a bit worse, about 10 to 20 degrees. It should be considered that the tests conducted till now were inside a lab with many magnetic disturbances. Further tests will include a complete check outside the building in a magnetic disturbance free environment. We expect that with a carful post-processing the accuracy can be as good as 2-3 degrees for all axis.

The Bexus campaign has been an unique opportunity to realize something that was originally just an idea. This lead us to face off with real challenging problems to be



solved in very short time. The short time between the proposal and the flight, deeply influenced the design of the experiment itself, because we always had to keep things simple but working. However, working on a less-than-an-year project allowed us to be in all the project phases till the end, this is not always true for every education project in the space sector, that often takes years to be completed. This allowed us to use this experience and results for our final dissertation at university, and also allowed us to establish contact with the industry. The collaboration with an important company, also gave us a big input to learn and acquire experience. From the Bexus campaign we have learned that a careful planning is necessary to develop a project, going through all the typical deadlines of a project development (PDR,CDR, etc..).

Finally we would thank all the people that made LowCoINS possible, from those that gave us the opportunity to be in the Bexus programme, to those that have helped us during the development till the realization, and those people that practically made our ideas to fly in the stratosphere in a lovely arctic day of October.



9 ABBREVIATIONS AND REFERENCES

9.1 Abbreviations ADC Analog to Digital Converter ASI Agenzia Spaziale Italiana (Italian Space Agency) CAD Computer Aided Design CCP Capture Compare PWM CDR Critical Design Review CKSum Checksum COTS Commercial Off The Shelf EAR Experiment Acceptance Review EBASS Esrange Baloon Service System ESA European Space Agency FEM Finite Elements Method GPS Global Position System GSS Ground Station Software I2C Inter-Integrated Circuit ICSP In Circuit Serial Programming IMU Inertial Measurement Unit INAF Instituto Nazionale di Astrofisica spaziale e Fisica Cosmica

(National Institute for Space Astrophysics and Cosmic Physics) INS Inertial Navigation System ISR Interrupt Service Routine LOS Loss Of Signal LowCoINS Low Cost Inertial Navigation System MEMS Micro Electro Mechanical Systems NGI Northrop Grumman Italia PCB Printed Circuit Board PDR Preliminary Design Review PWR Power RF Radio Frequency SED Student Experiment Documentation SMD Surface Mounting Device SPI Serial Peripheral Interface TC Telecommand TCU Thermal Control Unit THT Temperature High Threshold



TLT Temperature Low Threshold TM Telemetry TTL Transistor Transistor Logic UART Universal Asynchronous Receiver-Transmitter



9.2 References

[1] EuroLaunch: BEXUS User Manual (2007)

[2] David H. Titterton and John L. Westone – “Strapdown Inertial Navigation Technology - Second Edition” – The Institution of Electrical Engineers

[3] Anthony Lawrence – “Modern Inertial Technology - Navigation, Guidance and Control - Second Edition” – Springer

[4] J. Millman and A. Grabel – “Integrated Electronics” – McGraw-Hill

[5] Elliott D. Kaplan and Christopher J. Hegarty – “Undestanding GPS - Principles and Applications - Second Edition” – Artech House

[6] Robert M. Rogers – “Applied mathematics in integrated navigation system” – AIAA Education Series

[7] Joao Luis Marins et al. – “An Extended Kalman Filter for Quaternion-Based Orientation Estimation Using MARG Sensors” – Proceedings of the 2001 IEEE/RSJ International Conference on Intelligent Robots and Systems

[8] Widyawardana Adiprawita et al. – “Development of AHRS (Attitude and Heading Reference System) for Autonomous UAV (Unmanned Aerial Vehicle)” – Proceedings of the International Conference on Electrical Engineering and Informatics



10 ANNEX

10.1 Schematics



10.2 Part list



10.3 PCB

Figure 30 – Components placement



Figure 31 – Top layer

Figure 32 – Bottom layer



10.4 Harness



10.5 Housing design

Figure 33 – Top view



Figure 34 – Side view (internal details)



Figure 35 – 3D view experiment box (mounting bar detail)



10.6 Time Schedule



10.7 Outreach

10.7.1 Workshop poster



10.7.2 Published Article on “Memorie della Società Astronomica Italiana”



10.7.3 Abstract for ESA Symposium on European Rocket and Balloon Programmes and Related Research



10.8 Test Report

10.8.1 Overall hardware & software testing

Short after the delivery of the flight unit PCB from Artel Srl, the company in charge for the PCB manufacture, we started a test campaign on the flight unit. After a simple check of the power supply lines, all the sensors have been integrated and fully tested. The onboard software together with the ground station software have been fully tested with several 10 hours long-runs. During the test all LowCoINS functionality have been tested. Each long-run has been followed by a complete memory dump, thus, the data dumped from memory have been compared with the recorded telemetry data to find mismatches. After the successful memory to telemetry comparison, the memory has been erased followed by a blank check. A problem came out during these tests on the flight unit number one: the memory bank nr. 3 and 4 didn’t contain coherent data, and data in the affected memories were written irregularly. The problem was a bad solder on the ground pin of those memory chip. The issue has been fixed and long-run test repeated several time to check the success of the repair. Below a picture of the flight unit number one under test, in foreground the external power supply used during long-run test showing the current consumption of the unit (70 mA).

Figure 36 - Test ongoing on flight unit Nr.1



10.8.2 IMU testing – Sampling frequency

ADIS16355 is the core of our experiment, so particular attention has been focused to this component to find out the good setup to use it in the best condition. This high integration sensor is fully programmable, and one of the main feature is that it is possible to setup the internal sampling frequency of the sensor. Acting on the sampling frequency, we found out that it were possible to increase the stability of the data outputted reducing the sampling frequency, narrowing the bandwidth, thus reducing the noise. The sensor has also a built in Barlett Window Digital Filter to increase output stability.

Figure 37 - Barlett Window FIR frequency response

It has been necessary to find out the best trade off between output stability and response time acting on the sampling frequency and filter setting. This has been accomplished making several acquisition campaign from the sensor for a fixed time window (10 minutes) with different IMU settings.



Figure 38 - Gyros output

Then the variance of each IMU output (accelerometers and gyros channels) has been calculated. Since we read the IMU register asynchronously with respect to its update rate, an high sampling rate is desirable to lower the error due to sampling delay, and at the same time a low sampling rate is desirable to lower the output noise. The best compromise has been found out to be 102 Hz with the filter set to 32 taps.



10.8.3 IMU testing – Accelerometers calibration

The performances of an accelerometer is usually investigated using a series of static and dynamic test procedures. A multi-position tests are undertaken using a precision dividing head. This type of equipment enables the sensitive input axis of an accelerometer to be rotated with respect to the gravity vector. Hence, the component of gravity acting along the input axis of the sensor may be varied very precisely.

Figure 39 - Tilt table

The purpose of the multi-position tests is to determine the following parameters of an accelerometer:

1. Scale-factor

2. Scale-factor linearity

3. Null bias error

4. Axis alignment error



The experiment is fasted firmly to a precision dividing head, and the outputs from accelerometers are recorded for four different attitudes of the sensitive axis corresponding to 0g, 1g, 0g, -1g acting along this axis. Rotating the tilt-table in order to point each sensitive axis alternately up and down (six point test) it is possible to estimate the factors mentioned earlier simply summing and differencing various combinations of accelerometer measurements.

Figure 40 - LowCoINS and the datalogging PC

Since the tilt table used for these procedures cannot tilt the vertical axis 180°, but it can only be tilted 90°, the calibration procedure turned into a five-position test. This imply that our calibration is excellent for attitudes of the IMU that are different from the upside-down situation, that it is reasonably true on a balloon flight.



Below a picture representing the orientation of the IMU axis in the different positions.

Figure 41 - Tilt sequence



The output from the accelerometers can be written:

where the diagonal elements of the m matrix represent the scale-factors of each accelerometer axis, the out-of-diagonal elements represent the misalignment of the axis, the b vector represent the bias. The m matrix and bias vector for our IMU came out to be:

BIAS (mg) X-axis -29,06775 Y-axis 2,02325 X-axis 42,3525

m MATRIX 1,00028957 0,005403187 -0,0079597 -0,0086356 1,000607436 0,00382778 0,0029567 -0,003000672 0,99820076

In the following table can be noticed the improvements of the data accuracy after the calibration for each axis in each of the five positions:



Raw uncorrected data

(g) Calibrated data

(g) Position 1 -0,021076 -1,41514E-18

-0,00174 0 1,04412 1

Position 2 -0,022091 0,001604329 -0,99733 -1,000032686 0,03918 -0,000147432

Position 3 -0,032839 0,001604329 1,00134 0,999967314 0,04522 -0,000147432

Position 4 0,96897 0,998395671 0,01068 3,26856E-05 0,03957 0,000147432

Position 5 -1,030311 -1,001604329 -0,006597 3,26856E-05 0,04544 0,000147432



10.8.4 Magnetometers testing and calibration

The magnetometers’ output is a square wave frequency modulated. What the microprocessor does is to measure the period of the output signal. Unfortunately, the response of the sensors is not linear, and an appropriate transfer function needed to be found. The response is linarized using the following formula:

The linearization technique we adopted is to compare the output periods at full scale negative and positive field with the zero field period.

The procedure we adopted is to align each sensitive axis alternatively towards the magnetic north and south. Thus, becoming Hmin=-24125 µTesla, Hmax=24125 µTesla which is the north component of the earth magnetic field in the test site. Then the sensitive axis is pointed towards east and Ho is assumed to be 0 µTesla. Even this affirmation is not completely true, since a not-zero component exists in the east-west direction, it is so small that can be neglected in our considerations.

In this way the c coefficients can be found and placed in the linearizing formula to obtain a linear response from the sensors.

The procedure is undertaken using a precision dividing head made of non ferrous materials (actually wood and plastic). The calibration has been attended outside the NGI plant where a north pointing arrow mark is placed on the ground.



Figure 42 - Non magnetic tilt table

The procedures lead to an overall heading error of 1.5°, but we believe it can still be lowered.

The calibration procedure is then repeated with the heater switched on in order to evaluate the deviation due to the high current flowing in the heater. The results show that the total deviation is no more than +2° in the first two quadrant and -2° in the second two.



Figure 43 - Calibration ongoing



10.8.5 Pressure sensor calibration

A test in vacuum chamber at constant temperature has been conducted to evaluate the performances of the Honeywell pressure sensor. The pressure in the chamber has been lowered down to 5 mbar and then raised quickly to ambient pressure after a long period of stabilization in near vacuum. The test results showed that the pressure sensor is within specification having a small average bias of 7 mbar.

Figure 44 - pressure sensor test



10.8.6 Final Test: thermal-vacuum test with battery

The last test we did is a thermal-vacuum test to simulate the condition we expect during the flight with the unit complete.

Figure 45 - LowCoINS in the thermal-vacuum chamber

The test condition foresees to lower the temperature down to -65 °C and then, after a stabilization period, lower the pressure to 5 mbar, holding these conditions for 5 hours (the expected maximum flight duration). Afterwards the pressure is restored to ambient pressure, then the temperature is raised to 25 °C. The total test duration is 8 hours. PCB, IMU and heater temperatures are measured and recorded by the experiment itself, the battery, internal air and external air temperatures are recorded with external thermocouples.



Figure 46 - Final Thermal-vacuum test

The test was successful, the unit succeeded to keep the temperature of the electronics above 10 °C for the entire period of the test. By the way, after the test, the unit remained switched on the whole night (at 25 °C) because we needed to dump the memory. When we returned the day after, LowCoINS was still working regularly. The total discharge of the battery during the test was a capacity of about 3.7 Ah against the 13 Ah total nominal battery capacity. Nevertheless, the expected capacity for the battery pack during the flight we assumed to be about 7 Ah, because of the de-rating due to the low operating temperature of 0 °C. This temperature was an estimate we did during the design phase of the experiment, and our guess was really good since the recorded battery temperature during the test oscillated between -2 °C and 5 °C. Even considering the battery capacity de-rated to 7 Ah (and it is a big underestimate, since the battery will not work for the entire flight time at 0 °C) the total experiment autonomy with its own battery pack will nearly reach 10 hours. The supply voltage was stable during the entire test and did not showed any marked change when the heater is switched on and off.



10.8.7 Old Test Report

10.8.7.1 Thermal Test with active thermal control

Below are reported results of a thermal test conducted in a common freezer at -20 °C. Four temperature sensors are placed on the prototype board two on the component side, one on the bottom and one on the heating panel.

Figure 47 – Temperature Sensors Position

The microprocessor switches the heater on whenever the temperature T0 (centre sensor component side) falls below 15 °C and turn the heater off above 20 °C. The heating element is composed by two high power resistor of 5.6 Ohm. The resistors are rated for a maximum power of 50 W, however in the experiment they are required to provide a power 10 times smaller, in fact the total heater power is set to 10 W.



Figure 48 - Prototype overview

The heater is turned on and off with a low resistance logic level power mosfet driven by a logic port of the microprocessor. The status is recorded in real time using a software that allows for watching temperatures, set the low and high threshold temperature and command the heater on and off either manually or automatically.



Figure 49 – Thermal Control Unit control software

The figure below shows temperature transients and the heater status versus time.



Figure 50 - Thermal test @ -20 °C

It can be noticed the heater duty cycle is about 30% at an actual temperature difference between the inside and the outside of the box of 35÷40 °C. Results are encouraging but upcoming thermal-vacuum tests will validate the results in absence of free-convection.

10.8.7.2 1st Thermal Test in thermal chamber

The first thermal test in thermal chamber down to -70 °C has been conducted in order to evaluate the overall insulation power of the experiment box. The test, conducted with a sea-level pressure, showed the effectiveness of the box. Four thermocouples were positioned around and on the box, particularly:

• Inner case

• Outer case

• Internal air

• External air



The temperature profile programmed foresaw a fast transient to -70 °C with the maximum allowed thermal gradient of 5 °C/min, then the low temperature is kept until equilibrium. The test showed the effectiveness of the box in terms of its insulation capability.

Figure 51 – Thermal test on empty experiment box

10.8.7.3 2nd Thermal Test in thermal chamber

The experimental setup is the same as the previous, but now the pressure is decreased down to 5 mbar short after the chamber temperature is stabilized to -70 °C. It can be noticed that as soon as the vacuum is created, the temperature inside the box decreases with a slower rate as expected.



Figure 52 – Second test in thermal chamber



Figure 53 – First and second test



10.8.7.4 3rd Thermal Test in thermal chamber

On June 12th 2008, a thermal-vacuum test has been conducted placing the electronic prototype inside the experiment box in order to validate the theoretical power consumption and requirements guess to maintain the electronics in a safe temperature range.

Figure 54 - Ongoing thermal test

The experimental setup is as follow:

• Two thermocouples placed on experiment box cases (inner wall and outer wall)

• Two thermocouples placed “in air” inside and outside the experiment box

On experiment prototype:



• Two LM335 temperature sensor placed on the prototype PCB on component side and soldering side

• One LM335 temperature sensor placed on the heating panel • One pressure sensor

Figure 55 - Experiment overview in the thermal chamber



Figure 56 - Experiment in the thermal chamber

The thermal-vacuum chamber has been programmed to pursuit the following situation:

• Decrease the temperature down to -65 °C • Decrease the pressure down to 5 mbar

Since it is not allowed to decrease the temperature and pressure simultaneously due to thermal chamber limitation, the test program foresaw this profile:

1. Temperature decrease to the set point with the maximum allowable speed 2. Temperature stabilization to the final temperature set point for 2 hours 3. Pressure decrease to the set point with the maximum allowable speed 4. 3 hours holding pressure and temperature to the set points 5. Pressure increase to ambient pressure with the maximum allowable speed 6. Temperature increase to ambient temperature with the maximum allowable

speed



The prototype communicates with a PC through RS-232 interface providing readings from temperature and pressure sensors and heater status. The thermal control software switches the heater on whenever the temperature sensors placed above the PCB (T0) falls below Tlow, than switches it off when the temperature rise above Thigh. Tlow and Thigh are user selectable sending appropriate telecommand to the unit. Moreover, it is possible to go to “manual heater control” and switch the heater on and off manually. The heater is placed above the PCB and it is able to dissipate about 10 W of power. The purpose of the test is to check if this power is sufficient to keep the temperature in the safe limit and to determine the heater duty cycle in order to evaluate the overall power consumption.

Results are shown below, where it is possible to see seven different phases:

• Phase A: Temperature transient • Phase B: Temperature stabilization • Phase C: Pressure transient • Phase D: Test phase with Tlow=10 °C ; Thigh=15°C • Phase E: Test phase with Tlow=10 °C ; Thigh=12°C • Phase F: Pressure transient • Phase G: Temperature transient



Figure 57 – Test results

Results analysis

Phase B: Temperature stabilization at -65 °C with ambient pressure

This is a phase that is not likely to occur because the lowest temperature is reached in presence of free convection with ambient pressure. This is the worst condition phase that should never occur in flight, anyway the results show that the unit succeeded to maintain the temperature above the lower limit. The resulting heater duty cycle is 75% with an heater switching frequency of 1.9 cycles/hour. Due to thermal inertia of the heating panel, the temperature (T0) slightly fall below 10 °C short after the heater is switched on,



however, it never falls below 8 °C. Internal air temperature is always lower than PCB (T0) temperature, but never below 5 °C.

Phase C-D: Vacuum condition at -65 °C. ; Tlow=10 °C ; Thigh=15°C

In phase C the thermal chamber begins to make vacuum inside. It should be noted that the internal air temperature suddenly falls because of the pressure adiabatic decrease. It should also be noted that outer case temperature and external air temperature start diverging, because the thermal conduction through air is inhibited by the near vacuum condition. The temperature difference reached is 10÷15 °C. The same happens inside the box, where internal air temperature is always lower than the PCB temperature (T0), the difference is once again 10÷15 °C.

The unit succeeded to keep the PCB temperature above 8 °C with a duty cycle of 45% and a switching frequency of 2.4 cycles/hour.

Phase E: Vacuum condition at -65 °C. ; Tlow=10 °C ; Thigh=12°C

The high temperature threshold has been lowered to 12 °C. The result is a finer temperature control with an heater duty cycle of about 45% and an heater switching frequency higher than the previous condition, to 4 cycles/hour. So, with a narrower temperature range is possible to have a thermal control more accurate without affecting the duty cycle.

Phase F: Pressure transient to atmospheric pressure ; Tlow=10 °C ; Thigh=12°C

It proved to be the most critical phase, because the thermal conductivity through air is restored and the PCB temperature (T0) remained low for a while, just as the heater is switched on, and the internal air temperature joins again the PCB temperature (T0). Once again it is a situation that will never occur in flight because the pressure and the temperature will increase simultaneously.

Conclusions

The unit kept successfully the temperature above the safe limit in all test phases, including the severest. The mean heater duty cycle has been approximately 50 %, thus for a five hours mission the heater should be on for half the time. This means that if the current consumption is 1 A with the heater on, the battery capacity required is 2.5 Ah. Considering



the battery de-rating with current and temperature, the LSH-20 batteries are able to provide 7 Ah of capacity while loaded continuously with 1 A at 0 °C. This means that the autonomy is more than twice the required. Results also show that the internal air temperature with vacuum falls to -5 °C and the inner wall falls below -10 °C (worst case). This has to be taken into account, particularly for battery operating temperature. A thin insulation panel should be placed between the batteries and the inner wall while mounting in order to insulate them from the “cold” wall. However, even though the batteries reach -20 °C, the de-rating is still acceptable since they are still able to provide 6 Ah capacity.

10.8.7.5 4th Thermal Test in thermal chamber

The experimental setup and procedure is the same as the 3rd test, but now is added a thermocouple near the temperature sensor T0 in order to verify the correct reading of the sensor. The temperature set point for heater control are set to 10÷12 °C and then lowered to 5÷7 °C in order to evaluate the power saving. All transient are similar to the previous case.



Figure 58 – Test results

Results show that in the first test phase (Tlow=10 °C ; Thigh=12 °C) the heater duty cycle is 44 % with an heater switching frequency of 3.6 cycles/hour. When the set points are lowered (Tlow=5 °C ; Thigh=7 °C) the heater duty cycle decreases to 39 % with an heater switching frequency of 3.3 cycles/hour. The lowest temperature reached in this case is 0 °C when the sea-level pressure is restored with the external temperature steady at -65 °C. The test showed that is possible to save some power lowering temperatures threshold, however the power saved is negligible.

Team Name: LOWCOINS University of Rome “La Sapienza”

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