Download BUSAT Air Force Office of Scientific Research

BUSAT
The Boston University Student-satellite for Applications and Training
A Proposal
to the
Air Force Office of Scientific Research (AFOSR)
University Nanosat Program
Submitted by
Boston University
Center for Space Physics
in response to the
Announcement of Opportunity
AFOSR BAA 2006-06, Modification 2
October 16, 2006
i
Abstract
The Trustees of Boston University (Boston, MA) are submitting a proposal named BUSAT in response
to the AFOSR University Nanosat Program (AFOSR BAA 2006-06, Modification 1). BUSAT is a
comprehensive, integrated nano-satellite project involving graduate students, undergraduate students,
and faculty in the engineering and science departments of Boston University. The purpose of the
proposed satellite is to perform measurements of the precipitating energetic electron fluxes from low
Earth orbit over the high latitude auroral zones and to simultaneously image the auroral emissions
caused by these electrons. The Boston University Center for Space Physics is a leading space weather
research institution and is involved in the study of auroral emissions with ground-based radars, optical
instruments, and magnetometers. Present efforts to characterize the electron energy spectrum from the
information obtained by these ground based instruments is a main thrust of this research but these efforts
are limited by lack of the measured electron precipitating fluxes directly. The proposed satellite will
permit this loop to be effectively closed. The satellite structure and bus rely on the cube-sat concept
developed by prior nanosat funded efforts at a number of universities and will rely specifically on the
experience gained in such a program at Taylor University (Uplands, IN). The communication
(command and telemetry) subsystem will be procured via a subcontract directly from TU. Many of the
TU designs will be utilized for BUSAT but will be developed and fabricated by student teams at BU.
The needed auroral imager and imaging electron spectrometer (IES) will also be fabricated by student
teams relying on the in-depth experience with such instruments by BU faculty.
Table of Contents
1
EXECUTIVE SUMMARY..............................................................................................................................................1
2
SCIENCE INVESTIGATION ........................................................................................................................................2
2.1
2.2
3
SUMMARY ...............................................................................................................................................................2
OBJECTIVE AND SIGNIFICANT ASPECTS ..............................................................................................................3
TECHNICAL PLAN........................................................................................................................................................3
3.1
3.2
3.2.1
3.2.2
3.2.3
3.2.4
3.2.5
3.2.6
3.3
3.3.1
3.3.2
3.3.3
3.3.4
3.4
3.4.1
3.4.2
3.4.3
3.5
3.6
3.7
MISSION DESIGN DESCRIPTION ............................................................................................................................3
SPACECRAFT DESCRIPTION ...................................................................................................................................4
Structure Subsystem.............................................................................................................................................5
Environmental Control Subsystem.......................................................................................................................6
Command and Data Handling.............................................................................................................................6
Power Subsystem .................................................................................................................................................8
Attitude and Control and Determination Subsystem............................................................................................9
Communication Subsystem ................................................................................................................................10
INSTRUMENTS DESIGN.........................................................................................................................................11
Novel Triaxial GMI Magnetometers ..................................................................................................................11
Low Resolution Auroral Spectrometer...............................................................................................................13
Imaging Electron Spectrometer (IES)................................................................................................................14
The Plasma Probe and VLF Receiver................................................................................................................16
EDUCATIONAL IMPACT ........................................................................................................................................16
Experience with Spacecraft Education ..............................................................................................................16
Student Design Teams........................................................................................................................................17
Summer Research Opportunities .......................................................................................................................18
SPACECRAFT INTEGRATION AND TEST PLAN ...................................................................................................18
PAYLOAD OPERATION AND CONTROL CENTER................................................................................................20
FLIGHT OPERATIONS ............................................................................................................................................20
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BUSAT:
Boston University Student-satellite for Applications and Training
Principal Investigator:
Prof. Theodore A. Fritz
PI Address:
Center for Space Physics
Boston University (BU)
725 Commonwealth Avenue
Boston, MA 02215
Phone:
(617) 353 - 7446
FAX:
(617) 353 - 6463
E-mail:
[email protected]
Boston University
Co-Investigators:
Professor Supriya Chakrabarti
Professor Michael Ruane
Professor Josh Semester
Professor William Oliver
Mr Jeffery Baumgardner
Graduate Student Investigator
and Project Manager:
Mr. David Voss
External Co-Principal Investigators:
Professor Hank D. Voss,
Taylor University
Physics Department
8651 E 700 S
Upland, IN 46989
Phone: (765)998-4843
email: [email protected]
ii
1
Executive Summary
BUSAT is a comprehensive, integrated satellite project proposed by Boston University
(Boston, MA) involving graduate students, undergraduate students, and faculty in response to the
AFOSR University Nanosat Program (AFOSR BAA 2006-06, Modification 1). The
central BUSAT mission objectives are to:
(1) significantly involve students in the design, fabrication, and operation of a lowearth-orbiting small satellite and its instrumentation (hereinafter referred to as the
Boston University Student-satellite for Applications and Training or BUSAT);
(2) develop a practical and useful ground operational facility (hereinafter referred to as
BU CENTRAL) where students will be trained in spacecraft operations, data
handling and analysis, and space weather forecasting;
(3) test the Auroral Specification and Forecast Model (ASFM), being developed within
the graduate program of the Department of Electrical and Computer Engineering
(ECE) at Boston University, which will reside and be run at BU CENTRAL using
inputs derived from data obtained by BUSAT;
(4) interface these efforts into an existing program within Boston University involving
the Center for Integrated Space Weather Modeling (CISM).
The motivation for the BUSAT instrumentation and mission profile is space weather forecasting.
This is a timely topic and its importance has recently been recognized internationally. Forecasting, when
it achieves sufficient reliability and accuracy, has the potential of guarding against on the order of
$100M/yr in assets susceptible to space weather effects.
Improved capabilities would lead to
considerable savings for geostationary and low-altitude satellites, and ground-based power stations and
communication systems. To that end, several US federal agencies are coordinating efforts to address this
issue which has become one of national policy; and has led the National Science Foundation to establish a
Center for Excellence headquartered at Boston University to develop space environment models and
forecast techniques in the form of the Center for Integrated Space Weather Modeling.
The BUSAT mission, a student nanosatellite, and BU CENTRAL, will acquire energetic electron
data at high latitudes and will simultaneously acquire optical images of auroral arcs from horizon to
horizon. This mission concept will be designed principally by BU students. These students will form
instrument and science teams which will have responsibility to produce operational instrument hardware.
The BUSAT program integrates many aspects important to space physics, from instrument design to basic
plasma physics to environmental phenomenology to applications and commercial payoff, that it can easily
become the centerpiece for both classroom instruction and graduate research topics.
The BUSAT management concept fosters a strong link between the aforementioned student teams
and their faculty mentors. The faculty mentors range from senior to junior status thus providing the
important cross-generation ties which are critical to the vitality of any technical discipline. For example,
the principal investigator (Professor Fritz) brings to the project almost 40 years of experience in
spacecraft and instrument design. He has been PI or co-I on numerous instruments flown on both rocket
and satellite programs, ranging from the University of Iowa Injun-3 satellite to the joint NASA/European
Space Agency four satellite Cluster mission. He is presently responsible for the development of a Loss
Cone Imager energetic particle experiment for the US Air Force Demonstration and Science Experiment
(DSX) satellite program taking place at Boston University using an instrument development team of
undergraduate and graduate students. He will be the overall BUSAT technical project leader.
Two entry-level professionals should also be noted. Mr. David Voss, a graduate student in the
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ECE department, has considerable experience while an undergraduate student at Taylor University in
Upland, IN in designing, building and flying satellite and balloon instrumentation, and integrating entire
spacecraft payloads both on budget and against oftentimes tight delivery schedules. The concept being
proposed for BUSAT relies on the modular design that was developed at Taylor University over previous
cycles of the AFOSR University Nanosat Program. Mr. Matt Zettergren, also a graduate student in the
ECE department, is developing the Auroral Specification and Forecast Model (ASFM) as the major thrust
of the research being done for his PhD thesis research. The ASFM uses ground-based optical and radar
data to characterize auroral forms and to predict the incident electron energy spectrum responsible for the
auroral emissions. BUSAT will provide the measurement of the incident electron energy spectrum
needed to assess the effectiveness of the ASFM model and verify its predictive extension to low-altitudes
and the ground.
BUSAT utilizes tested concepts of nanosat spacecraft design using modular stackable four-inch
cube designs developed at Taylor University in Upland, IN. The payload consists of four advanced, stateof-the-art and mature scientific instruments: (1) a high-resolution, high-dynamic range magnetometer, (2)
an optical and UV auroral imager, (3) a low-to-medium energy electron imaging spectrometer; (4) a
plasma probe and VLF receiver. The satellite will be in a near-polar, preferably sun-synchronous orbit
(approximately in the noon-midnight meridian) at low altitudes (300 - 2000 km). BUSAT will have a
permanent dipole magnet with active control to align its orientation along the local magnetic field
direction; spacecraft azimuth will be determined by an array of optical sun sensors.
BU CENTRAL will be located within the Boston University Center for Space Physics and will
serve as the spacecraft operational center, data handling/distribution center, and operational analysis
forecast center where the ASFM will be run. Specifically, within BU CENTRAL we will test the lowaltitude performance of the ASFM (through comparison with BUSAT data), design and test operational
algorithms to forecast ground-induced currents, and investigate the forecast prospects.
2 Science Investigation
2.1 Summary
The primary objective of the BUSAT mission is to make a significant contribution to the science of
space environment forecasting while concurrently training students in space hardware development and
space environment research and applications. The following important tasks fall under this heading:
• design and build a five-instrument satellite (BUSAT) to gather data for testing and
developing space-environment applications and related research;
• acquire data to evaluate the capability of the Auroral Specification and Forecast
Model (ASFM) to predict low-altitude space environment conditions;
• create a prototype space environment operational prediction center (BU
CENTRAL) in a university teaching-and-research setting.
BUSAT's secondary objective is to perform complementary and supporting research for space
environment applications. The following tasks fall under this heading:
• compare low-altitude data acquired by BUSAT instruments to identical
instrumentation flown the Polar spacecraft, and the Cluster spacecrafts; and the
future DSX satellite.
• investigate the usefulness for space environment predictions of satellite
observations of dayside auroras (develop forecasting implications);
• compare simultaneous, over-flight BUSAT magnetic field data with data from BU's highlatitude ground magnetometer chain, MACCS [http://www.bu.edu/dbin/csp/]
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2.2 Objective and Significant Aspects
Over the past thirty years, the space science community has learned much about the geospace
environment through space- and ground-based measurements, their ensuing data analysis, and theoretical
models. During this time, the field of space physics has naturally evolved from one that was initially
exploratory, to one that is now attempting to synthesize and quantify the inherently dynamic elements that
comprise the spatially extended but closely coupled magnetospheric system. One principle characteristic
of the magnetospheric system is its dynamism. The magnetospheric fields and particles respond in a
coupled fashion to time variations in the solar wind plasma and interplanetary magnetic fields. These
externally-induced changes in magnetospheric parameters can be profound: ions and electrons may be
substantially energized; boundary layers separating vastly different particle populations may move
rapidly; strong electrical currents may flow and dissipate energy in the auroral zone; etc. Much progress
over the past three decades of space physics research has allowed us to develop ever improving
quantitative and predictive models of this dynamic system.
2.3 Implemention Plan
The prime BUSAT objectives center around using students to build satellite hardware to enable
them to test the so far untested low-altitude (i.e., BUSAT altitude and ionosphere altitude) ASFM
predictions and to develop and test an algorithm that takes output from the ASFM, running with input
data from BUSAT, to predict the incident energy spectrum of the precipitating electrons responsible for
producing the auroral emissions being measured. In order to achieve the scientific objectives outlined
above the following complement of scientific instruments will be included as part of the BUSAT payload.
Table 1 - Summary of instruments on BUSAT and their measurement sensativity
Instrument
Magnetometer
Auroral Imager
IES
Langmuir Plasma Probe
VLF Receiver
Measurement
Magnetic Field
Auroral Emissions
Energetic Electrons
Thermal Electron and Ion Density
EM waves
Range
>10nT
390nm to 650nm
30 keV to 500 keV
0 to 6 eV
100Hz – 30 kHz
3 Technical Plan
3.1 Mission Design Description
The proposed satellite has seven major subsystems: Scientific Instruments, Communication,
Structural, Data System, Environmental, Attitude Control & Determination and Power System. These
subsystems will be integrated into a modular structure bus made out of Aluminum. A summary of the
subsystem can be seen in Table 2 and displays the power and mass throughout the major subsystems of
the spacecraft and their interfaces.
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Table 2 - Summary of BUSAT subsystems
System
Power
Number of Cubes
2
2
2
N/A
1
1
N/A
1
1
2
1
2
N/A
N/A
17
Battery box 1
Battery box 2
Control
Solar Arrays
DPU
DPU
ACS
Bar magnet control
Bar Magnet
Science
IES
Magnetometers
Auroral Imager
Plasma Probe/VLF
Com
FreeWave/Ham
Structure
Exoskeleton
Contingency N/A
Total
Power
N/A
N/A
1.5
N/A
2
4
N/A
2.8
3
2
0.5
3
N/A
2
20.8
Mass (kg)
1.7
1.7
1
1
0.75
1.5
5
1.5
1
1.5
0.6
2
7
2
28.3
3.2 Spacecraft Description
The university is a unique environment that requires simple interfaces between systems due to the
academic atmosphere. BUSAT’s mechanical design is based on the ideology of the CubeSat program and
the satellite bus designed at Taylor University. The CubeSat program, started at California Polytechnic
Institute and Stanford University, allows universities to design picosatellites of size 4”x4”x4” to launch as
secondary payloads. Using this concept Taylor University built and designed a satellite bus system where
the subsystems were cubes, or a multiple of the cube size. This greatly simplifies the design effort for
students by standardizing the mechanical and electrical interfaces. The subsystem cubes are held together
by an exoskeleton structure that provides structural support for the satellite and provides the interface to
the launch vehicle. BUSAT will take advantage of the LightBand deployment system that provides a
standardized mounting to a spacecraft or to another nanosatellite equipped with this feature.
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1
2
3
4
5
6
7
1 – IES
5 – Corner Brackets
2 – Modular Subsystem Cubes
6 – Bottom Plate
3 – Exoskeleton Walls
7 – Lightband adapter
4 – Solar Arrays
Figure 1 - Exploded view of BUSAT one depicting the modular structure
3.2.1 Structure Subsystem
Subsystem Cubes
The 4” x 4” x 4” aluminum subsystem module is the standard dimension on which the modular
design rests. It provides the structural element on which PC boards are mounted, batteries are secured,
and instruments are attached, as can be seen in Figure 2. Each subsystem module is capable of handling
the various environmental conditions placed upon it due to launch. Modules may be used individually for
PCB stacking (custom boards defined to be 3.5”x3.5”), or be combined to accommodate unique COTS
geometries. Each subsystem module provides an effective RF shield reducing electromagnetic noise
throughout the satellite. Low noise instrumentation may then be effectively isolated from noisy systems
such as communications and the data processing section. Each module is connected to the exoskeleton
during the assembly process. An effort will be made to examine the potential to interconnect subsystem
cubes together for three-dimensional stress translation through the subsystem cubes as well as the
exoskeleton. Unlike traditional satellites, where load is carried through large structural panels, this
concept allows for distributed loads to meet launch stresses. By stacking subsystems this concept is able
to utilize all of its internal space resulting in a compact, but efficient design. Each module has six walls
for attaching components, allowing for efficient use of the nanosatellite package. The dimensions of
BUSAT are approximately 15”x15”x 9.6”.
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Figure 2 - Cutaway of a subsystem cube showing stacked circuit
boards
Exoskeleton Structure and Solar Array
The primary structure (exoskeleton) of the
satellite surrounds the assembled subsystem cubes and is
responsible for shouldering much of the satellite load,
keeping the subsystem cubes firmly attached together,
and increasing the natural frequency of the satellite. The
0.5-inch thick-pocketed (weight-reduced) aluminum
plates provide primary structure through vertical and
lateral ribs. Attached to each wall of the exoskeleton is 0.0625-inch sheet of aluminum acting as a shear
plane to increasing the fundamental frequency at which the satellite will vibrate (>100 Hz). Secured to the
shear planes through thermally conducting standoffs are the solar arrays that are decoupled from the
structural load. Each solar array is constructed of a pocketed 0.25-inch aluminum sheet that runs the
complete height of the satellite. The array is designed for the GaAs cells to be arranged in multiple
strings. Each string will most likely consist of eight cells. Should something occur to a string each will be
isolated by a diode to avoid power drains by dysfunctional strings. The solar array interfaces with the
satellite through a single connector. The connector has two main positive and negative power rails within
to provide redundancy in case of a power spike. The idea behind the connector is that it allows the entire
solar panel to come off for testing and storage without interfering with other systems. Thus, the
modularity of the satellite is maintained not just within the cubes but also in the separate structures of the
satellite itself.
3.2.2 Environmental Control Subsystem
In low earth orbit, the satellite will be exposed to direct sunlight for most of a full orbit and the
additional reflection from the earth for approximately half the orbit. Under these conditions, the satellite
will not reach equilibrium either on the dayside or the night side. The temperature will vary from 270 K
just before dawn, to 290 K just before dusk. The equilibrium temperatures on the dayside and the night
side would be respectively 320 K and 215 K. While we expect the spacecraft surface to be outside the
270-290 K range, the thermal conductivity of the aluminum structure will prevent it from deviating
much. This exercise shows that a tolerable thermal environment can be maintained without a great deal
of effort. The only subsystems that have a thermal sensitivity are the Auroral Imager and the IES which
we would like to operate around -20 degrees C (253K). Thermal design of a small compact satellite is a
problem that has been encountered and solved numerous times in the past, so we do not expect to
encounter difficulties in this area.
3.2.3 Command and Data Handling
The command and data handling system of BUSAT will be designed to complement the high
degree of modularity offered by the mechanical system. The C&DH subsystem of the satellite can be
broken up into two logical levels: the microcontroller level, and the Single Board Computer (SBC) level.
A microcontroller is located on the common Interface Board of each subsystem and is connected to one or
more instruments via connections and hardware on the Interface Board. The connections provide up to six
analog-to-digital conversions, an RS-232 serial interface, and a one-wire serial interface. Through these
interfaces, the microcontroller on the Interface Board controls and collects data from the instruments,
storing the data on 512 KB of SRAM external to the microcontroller. The Interface Board and the
microcontroller on it are also used for the attitude control, power, and communications system. An I2C
bus allows for communication between each microcontroller, while a RS-485 serial bus provides a
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connection to the SBC level. The I2C bus is also a redundant bus for communication with the SBC level.
The central computing of the satellite is performed by a low powered (280 mA) SBC called a Vipers that
is manufactured by Arcom. Each Viper features an Intel PXA255 400MHz processor with 64 MB of
SDRAM and a CompactFlash drive with capacity of 1 GB. Each SBC runs a version of embedded Linux
developed by Arcom and comes with 5 USART serial ports (four RS- 232 and one RS-422/RS-485) and
16 digital I/O ports. The Viper is modified to allow access to the I2C bus.
Figure 3 - Logic diagram of the DPU system on the TEST satellite. A similar design will be implemented on BUSAT
On the functional level, the Viper is responsible for data collection from all instruments and
transferring it to the ground via the communication system. Communication with the microcontroller
level is performed via an RS-485 bus with an I2C bus for redundancy. The RS-485 bus is full duplex,
allowing for simultaneous transmission and reception, and has the capability of connecting up to 256
devices, which provides a large amount of expandability and modularity. A new microcontroller can be
added with relative ease by providing the Viper with its address.
Modular Software System
The software associated with the hardware of the command and data handling system is a vital
part of the satellite. The software is also modular, complementing the modularity of the mechanical
system. The software can be viewed from the same two logical levels: the microcontroller level, and the
SBC level. On the microcontroller level, many common software functions have been standardized to
form libraries that can be replicated throughout the subsystems. These functions have been grouped by
their applications, which include a real time clock, analog-to-digital conversion, external memory access,
one-wire communication, communication with the SBC level, and a variety of other applications. By
creating these libraries of functions, coding and debug time is greatly reduced while dependability of the
software increases. The SBC level has two main sections of software: one for communication with the
microcontroller level and one for communication with the ground station level. The main goal of
communicating with the microcontroller level is to retrieve the data collected from the instruments. Each
microcontroller has its own address and is polled periodically by the SBC. Then, a simple protocol is used
to transfer the data. After initial handshaking, the microcontroller sends a block of the data and the Viper
acknowledges reception, and this repeats until all the data is sent. The Viper then performs compression
(gzip and tar) on the data and stores it until contacted by the ground. The other section of software on the
7
SBC level communicates with the ground station. A program called serial transfer protocol (STP) controls
this communication. On either side of the communication, interface programs are responsible for
initiating STP and then controlling what is transferred through the connection.
3.2.4 Power Subsystem
BUSAT will employ a power monitoring and control system that is capable of handling the
various requirements placed upon it by the instruments, an onboard flight computer, and two
communications systems. The power subsystem will take advantage of the modularity provided by the
satellite and is divided into three sections. The power control system is housed in a single 4”x 4” x 4”
cube and consists of the four custom PCBs and one Interface Control Board.
Figure 4 - Block diagram of the power module
Two double cube modules house two separate battery packs of 11 Sanyo NiCad N4000DRL cells
that are each capable of providing a nominal capacity of 4Ah. BUSAT will design to the NiCad battery
but will look at current battery technology, such as Li Ion technology which has a significantly higher
energy density.
The first PCB in the Power Control module is the Interface Control Board. This board provides all
bus communications and I/Os necessary for controlling the power subsystem. The next circuit board on
the stack is the charging board. This board sits directly on top of the Interface Control Board, giving it
access to all A/D, I/O, and serial communications lines provided by the microcontroller. It’s external
connector provides an interface for both battery packs, and inhibit verification lines. This board supports
NiCaD charging via a switch mode topology controlled by the PIC to maximize power efficiency in the
system. The power subsystem is also responsible for monitoring the power health of the satellite. The
second power board is the voltage regulation board, which takes the unregulated power line from the first
board and passes it through 5 separate voltage regulators to provide the various voltages needed in the
satellite. The following voltages are provided to all subsystems: -5, 5, 8, 12, and unregulated. The
MIC4685 switching regulator in a step-down configuration is used for all five of the voltage regulators
due to its ability to support 3A of current, its relatively high 85% power conversion efficiency. Extra
storage capacitors are added onto each regulator to ensure that the voltage remains constant. Voltage and
current monitoring is performed on each line. An additional 5V regulator will be included for
contingency.
8
The third board on the stack is the power distribution board, which takes the 5 regulated voltages
and passes them through solid state relays to each of the subsystems. Every line out of the power
subsystem is controlled by a relay. Before leaving the switching board each subsystem power line is
passed through a resettable poly-fuse. Fusing the outputs ensures that the regulators will not be damaged
and also protects the wiring and PCBs. The power subsystem monitors the final subsystem 5V output
voltages to ensure that a subsystem is actually on.
The last board on the stack is the solar peak power tracker board (Figure 5). This board connects
to all four of the solar panels. There are two connections per panel for redundancy. Every
string has a blocking diode placed on the solar panels. The current provided by each panel is sensed by a
MAX471 and all of these are connected to a DS2450. The panels are all then tied together as one line that
is connected to a step down switching regulator. This regulator has its feedback off of the input voltage to
hold the solar array at a specific voltage, the max power point. Since the maximum power point for the
array will be dynamically changing due to the panels being on different axes and changing temperatures,
a digital pot is in the feedback divider for the PIC to continuously search for the peak power point. The
input voltage and the output voltage are both sensed via voltage dividers.
Figure 5 - Solar peak power tracker board
3.2.5 Attitude and Control and Determination Subsystem
The Attitude Control Magnet (ACM) subsystem consists of a permanent magnet, which can be
oriented in the spacecraft XY-plane by a stepped motor, a shaft encoder, two hall elements, a separate
electronics board, and hysteresis rods for damping.
The magnet will attempt to align itself with the local magnetic field, and excess mechanical
energy will be damped through hysteresis loss in ferromagnetic material (the hysteresis rods). Magnetic
alignment is a proven method of spacecraft attitude control, and the BUSAT Principal Investigator was
strongly involved with the Injun 3 satellite which very successfully demonstrated the simplicity and
efficiency of magnetic alignment and hysteresis damping as a means of attitude control (Fritz, 1965).
The magnet can be rotated in the XY-plane in small steps. This allows the spacecraft to be
positioned in different orientations with respect to the local magnetic field vector and without introducing
excess mechanical oscillations. The hall element measurements can be used to monitor the orientation of
the dipole axis in satellite coordinates as a backup to the shaft encoder measurements. Additionally, the
hall element measurements can be used to determine the decay of the core magnetization over time.
The electrical system consists of the motor control processor and digitization circuit on the
interface board for the hall element measurements. The SBC will send commands to the controller chip
(the HCTL-1100 is a motor control processor receiving feedback from the shaft encoder. It can be
controlled with simple commands to produce complex movements, if required. It communicates with the
microcontroller through an 8-bit interface). Since the ACM system will be non-operational for most of its
lifetime, it will be powered off while not in use, and only switched on by the power system interface
9
board for making a measurement or sending a command.
The Attitude Determination System (ADS) is based on the simultaneous measurements of zenith
angle of incidence of the sun on a large number of photodiodes at different orientations with respect to the
sun. We will use 24 diodes in 8 groups (sensor heads) of 3 mounted on complementary corners of the
satellite frame. Each diode will have a field of view of a 50 degree cone, and will be able to measure the
zenith angle of the sun to an accuracy of approximately 0.5 degrees. The data from the ADS consists of
light intensity measurements, which will be processed on the ground in order to obtain the attitude. The
ADS is a highly redundant system, and can sustain the failure of up to half of the photodiodes. The power
requirements of the system have been reduced by using a very simple power switching circuit, which
supplies power only when and where it is needed. In addition, the electronics board will be switched off
between measurement cycles to further reduce the power requirements.
Data acquisition from the ADS is simple. Every 1-5 seconds, the SBC will switch on the ADS
board. After power on and stabilization, the CPU will in sequence query the ADS for a measurement from
each diode. In each measurement step, the CPU will first send the ADS the number of the diode that is to
be measured. This will switch power to the sensor head containing that diode, and the diode will provide
a voltage for digitization. Once sufficient time has elapsed for the system to stabilize, the CPU will be
able to retrieve the digitized value of the diode voltage. The two other diodes in the sensor head are then
measured. This can happen very fast, since they are already powered.
The measured light intensity in the diodes will be digitized to 8 bits, giving a digitization accuracy
of 0.5 to 0.3 degrees, depending on the angle of incidence. These measurements will be transmitted to the
ground station for processing in the Attitude Determination Software. The complete cycle of powering on,
taking measurements from all diodes, and powering off takes approximately 0.1 seconds, so that the ADS
system will be idle and not consuming power for approximately 90%-98% of the time, reducing the
average power consumption by a factor of 10 to 50 over the peak power consumption.
The ADS will be constructed by a student team. The diodes will be procured pre-calibrated.
Sensor heads will be manufactured at the Physics Department Scientific Instrumentation Facility.
Electronics boards will be manufactured in the Electronic Design Facility at BU, and all electrical and
mechanical integration will be performed by students
3.2.6 Communication Subsystem
The communication system will be subcontracted to Taylor University and will include two
separate communication systems developed for the TEST satellite and flown on over 35 high altitude
balloon launches. The system takes advantage of a combination of a proven, reliable transceiver and an
experimental higher data-rate transceiver. The flight heritage transceiver flying is a TH-D7 dual band
amateur radio. The TH-D7 has a built in Terminal Node Controller allowing for serial connection from
the communications module microcontroller directly to the radio. The amateur radio will be primarily
used for transmitting a beacon with health and diagnostic information. This radio will be utilizing the
standard AX25 protocol allowing amateur radio operators around the world the capabilities of listening to
the satellite as well as posting the satellite’s health information online. The higher data-rate transceiver
being flown on BUSAT is a COTS 900 MHz spread spectrum transceiver produced and distributed by
FreeWave Technologies. Utilizing a proprietary hop table on an international frequency this COTS has
excellent signal to noise ratio (-106 dBm sensitivity) allowing for lower transmit power while maintaining
excellent data transfer. FreeWave also performs buffered 32 bit CRC transmission with notification when
connection is lost. The 900 MHz FreeWave is capable of ~115 Kbps of continuous throughput. Extensive
testing has been performed on the FreeWave transceiver through 35 high altitude BalloonSat launches at
Taylor University. During these flights in which the FreeWave was carried to heights of 100000 ft and
~115 miles away from the ground station, high-data rate communication was kept throughout the flight.
The FreeWave is directly connected to the satellite computer. The communication module is made of a
double cube and houses the FreeWaves and the TH-D7.
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3.3 Instruments Design
3.3.1 Novel Triaxial GMI Magnetometers
Magnetic fields will be sensed by a compact, low weight, and low power triaxial magnetometer.
These sensitive micro-magnetometers based on the giant magneto-impedance (GMI) effecti have not
previously been used in spacecraft. Sensors will be purchased from Madison Technology International,
Ltd. or constructed from commercial GMI wire.
3.3.1.1 Concept of operation
The GMI effect—discovered in 1993—arises in annealed magnetic amorphous wires from the
dependence of the skin depth upon magnetic permeability, which changes with the ambient field.ii Under
high frequency excitation (10MHz to 100 MHz) large impedance changes of the wire can be sensed with
good noise performance and robust temperature stability.
Figure 6 - Aichi Micro Intelligent Corp. GMI Field Sensor - 10mm x 31 mm
The GMI magnetometer will measure magnetic field
components by three identical sensor channels whose sensing
elements are aligned orthogonally. Several technologies exist for
fabricating this triaxial head. In one form, the sense elements can
be fabricated on a cube and short leads made to the electronics. In
another form the sense element for each channel is integrated with
its electronics on a small board. Three identical sensor boards are
arranged orthogonally. Either configuration can easily fit in the
basic cube of the nanosat.
3.3.1.2 Sensitivity
GMI wires can sense fields down to the nanoTesla range with bandwidths of 1 MHz iii. Sensor
heads, a few mm long, require only tens of mW. GMI sensors have resolution and sensitivity comparable
to fluxgate magnetometersiv, while requiring significantly less volume and power, and offering larger
bandwidths. A variety of sensor electronics have been developed for these applications.v
Noise performance for GMI sensors is already comparable to current fluxgates. Future noise
factor improvement may come from more careful annealing and fabrication control of the GMI wire.
Improved electronics will also reduce the circuit noise contributions. COTS applications, largely
targeting the automotive industry, have simply not needed excellent noise performance, so the technology
has not been pushed.
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Figure 7 - GMI Micro-Magnetometer Self-Noise Comparisons
3.3.1.3 Electronics
A GMI sensor is fundamentally a sensitive
impedance device. Supporting electronics provide a drive
signal, biasing, and detection of an analog output
proportional to the magnetic field along the sensor wire.
Processing electronics further filter the signal and convert it
to digital with standard ADCs.
Radiation hardness of a GMI magnetometer should
yield directly to good engineering practices. The wire itself
is unaffected by radiation. Sensor electronics are relatively simple, and have been implemented in COTS
CMOS surface mount devices for terrestrial and naval applications.
We will design, fabricate, calibrate and deliver the GMI triaxial magnetometers using standard
magnetic fiber with existing GMI magnetic sensor designs and interfaced with heading sensor electronics
with 16 bit A/D converters (per field component) to provide a digital readout of the X, Y, and Z
components of the magnetic field strength.
The magnetic sensors and electronics will be housed in a single, non-magnetic, spaceflight-ready cube
housing.
3.3.1.4 Mechanical Design
Two COTS technologies exist for fabricating the sensor heads, which comprise a length of GMI
wire surrounded by a bias coil. First, we can wind miniature coils (about 3 mm long and 2 mm in
diameter) and place short (3-4 mm) lengths of wire in the coil. The final four-lead sensor is potted for
ruggedness.
Alternately, Aichi Steel (Japan) produces a compact sensor that is intended for the
automotive navigation market or for magnetic sensing in cars. This sensor is packaged in a 2 mm x 3 mm
BGA surface mount chip that Aichi fabricates using semiconductor wafer processing. To create a bias
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coil, they embed a 2 mm length of amorphous wire within a bias coil realized as a solid state inductor
configuration (traces and vias form the bias coil).
The sensor and electronics will be packaged with dimensions of about 15 mm x 15 mm x 75 mm,
with a weight of about 100 g. There will be independent signals from each axis, independent control of
the bias coils, and power inputs to the electronics. One unit should be mounted on a boom, as far as
possible from the spacecraft with its fields. A second unit is planned for the spacecraft, to allow study of
the spacecraft fields, the fields at the loss cone detector, and as a redundant sensor if necessary. The low
weight, power, and cost make multiple sensors a simple option.
3.3.1.5 Sensor Technical Objectives
The magnetometers shall be able to provide measurement of the three dimensional field components with
the following specifications:
For 0-20 Hz, DC field (Space Weather application and attitude determination)
Range
100 to 10,000 nT
Resolution
1 nT
Self-noise
0.1 nT/sq. rt. Hz
@0.1 – 20 Hz
ADC resolution
16 bits
Field vector accuracy
<1 degree
3.3.2 Low Resolution Auroral Imager
A design for a compact spectrometer for observing auroral emissions in the wavelength range of
390nm to 650nm has been developed. This instrument will be built using off the shelf components that
have been modified to withstand the rigors of spaceborne operation. The instrument will fit inside a
rectangular prism 7.5cm x 20.0cm x 10.0cm or two of BUSAT modular cubes. The expected mass of this
instrument will be approximately 1.5kg.
The optical design presents of the spectrometer will employ a blazed transmission grating to
provide the required spectral dispersion. A 524 X 752 CCD binned 2x2 will be used for the detector. The
dispersion will be ~0.65 nm/pixel with an expected resolution of ~2.0nm. The slit will be ~ 0.1 mm wide
a by 25mm long. The objective lens will be a 16mm fl F/2.8 “fisheye” lens. This lens produces a 40mm
diameter circular fov of 180 deg. The proposed instrument will only use the central 25mm of this fov or
~115 deg. This fov will be re-imaged on to the 262 pixel spatial dimension of the CCD yielding a
projected pixel size of 0.36 deg along the orbital track by 0.44 deg. cross track (at 600km altitude this
corresponds to 3.8km by 4.6 km). A nominal integration time would be 1 second yielding a resolution
along the orbital track of ~10km . Longer exposure time will yield a better S/N but at the expense of
along the track resolution. An image is produced line by line along the satellite path that will be 115
degrees of selected auroral emission spectral lines.
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Figure 8 - Concept sketch of the Auroral Imager
3.3.3 Imaging Electron Spectrometer (IES)
BUSAT will employ passive attitude control through magnetic orientation the earth's dipole field
via a permanent magnet. The IES will be oriented on the spacecraft so that its field of view provides a
measurement of the complete energetic electron pitch angle distribution. Coupled with measurements of
the local magnetic field strength performed in situ by the magnetometer, electron pitch-angle distributions
can be determined from the look-direction resolved flux data. A major mission objective of BUSAT is the
study of the size and position of the auroral oval and polar cap. IES data can be used to determine the day
and night side position of the electron trapping boundary and thus the size of the polar cap. As noted in
section 1.x.x the determination of this asymmetry should also permit a determination of the geomagnetic
field topology.
The Instrument
The Imaging Electron Spectrometer (IES) is a particle telescope capable of sensing electrons in 9
different look directions, covering a full 180° field of view. Flux measurements as a function of pitchangle as well as energy can be achieved by the novel geometry of the IES sensor, shown in Figure 9. The
IES instrument, consisting of the sensor and electronics box is also shown in Figure 9.
Multi-element ion-implanted silicon strip detectors are used to sense energetic electrons. The IES
is sensitive to electrons of energy ranging from 30 keV to 500 keV. An aluminized mylar foil placed in
front of each strip detector eliminates protons of energies below 350 keV as well as a light response. Each
of the three strip detectors is positioned beneath a pinhole aperture. Each of the three detectors in turn
contains 3 elements capable of sensing incident electrons. This geometry is illustrated in Figure 13 for
one of the three detectors. An electron entering through the aperture will hit a detector element along its
line of flight. The angular view of each element is approximately 20° x 35° and the angular range of each
detector, limited by the aperture, is 60°. The 3 detectors thus provide 9 “look” directions which cover a
full 180° x 35° field of view.
Also located within the compact IES sensor “head” is an integrated circuit which collects the
charge accrued on each detector. Each channel of the 30-channel “RENA chip” is a charge-sensitive preamp and pulse shaper. Only 9 of the 30 available RENA channels are used, with a channel dedicated to
each one of the 9 detector look directions. The system is event-driven and can constantly monitor for
incident electrons.
14
Figure 9 - An exploded view of the IES is shown on the right, and a concept view of the crossection of a telescope in the
IES
Data Acquisition
A charge-to-pulse height proportionality is established by the charge-sensitive pre-amp input to
the RENA chip. The resulting pulse is then processed to extract the energy information. The 9-channels
are multiplexed through to an Analog-to-Digital converter (ADC) where the pulse-height or energy is
digitized to 8-bits. This energy value is look-direction coded by control logic, then appropriately binned
into one of 15 energy bins per look direction. 24-bit counters implemented in RAM are used for this
binning process. A micro-controlling element, implemented either in hardware logic or with the microprocessor performs the necessary binning operations. An integral count rate is determined through the use
of a level discriminator and 24-bit discrete counter in each of the 9 processing chains. During calibration
this threshold can be set so that it is above the system noise level but below the lowest energy of interest.
This “true” count rate can then be used to normalize data on the ground if high incident fluxes are
encountered. The expected data rate at the IES/CPU interface is therefore 16 x 9 x 24 bits or 3.5 kbits/sec.
Calibration of the pulse-height to energy proportionality is determined with an x-ray source.
Determination of energy loss in the entrance window (i.e.: Aluminum contact and aluminized mylar
combination) can then be made with a mono-energetic electron and ion beams. Determination of the IES
angular efficiency can be made with a rotating test platform. It is estimated that the highest pitch-angle
and energy-resolved flux that can be accurately measured by the IES is 50 x 106 electrons/cm2 sec ster,
limited by the speed of the ADC.
Heritage of the IES instrument
A version of the IES instrument proposed here became “flight proven” when it was flown on
board the PULSAUR II rocket, launched February 9, 1994 as part of a University of Bergen mission
investigating pulsating auroras. Telemetry data from the rocket indicated that the instrument performed
successfully in flight (Levine, 1994). Other versions of the IES are integrated into satellite payloads on
15
Polar, as part of the CEPPAD (Comprehensive Energetic Particle and Pitch Angle Distribution
Experiment) experiment (Blake et. al., 1994), and on Cluster, as part of the RAPID experiment (Wilken
et. al. 1993). All of these versions employ the same sensor head geometry, detection scheme and housing
structure. The only major differences are in the design of the support electronics.
The IES version proposed here utilizes the RENA chip as front-end support for the strip detectors.
A prototype of the IES sensor head incorporating the RNEA chip is being built at Boston University for
the DSX mission. Based on this, it is felt that the IES instrument can be wholly built and tested utilizing
graduate and undergraduate student involvement. It is a well understood design and is essentially flight
proven. A complete drawing package exists for every assembly, excluding the version-specific electronics
design and interface to the data system. Thus, no development costs for the majority of the instrument
design will be and flight spare detectors from the Cluster and Polar projects are available at BU
3.3.4 The Plasma Probe and VLF Receiver
The Plasma Probe and the VLF receiver are subcontracted through Taylor University and were
they were developed for the TEST satellite. The electronics for both instruments are integrated into a
single cube with the VLF receiver and the Plasma Probe sensor mounted external to the satellite. The
Plasma Probe is mounted on the bottom panel of the spacecraft. It measures the thermal plasma density
and temperature of the satellite environment. This low energy (0-6 eV) plasma probe circuit has had a
design heritage on the NASA DROPPS rocket and the TU Sat 1, a CubeSat designed and constructed at
Taylor University. The probe is programmed to various voltages and is connected to a log electrometer
and a low-noise instrumentation amplifier. The output of the amplifier subtracts the bias voltage and is
further amplified for a 0-5 V A/D conversion. Plasma probes are common on satellites because they
provide an excellent complement to the other instruments to better understand the environment the
satellite is traveling through. Knowing the environment of travel is a crucial step in interpreting the data
that is received from the other instruments. The Science Module Interface (SMI) board is responsible for
digitizing and storing the plasma probe data.
The charge build-up and lightning discharge from thunderstorms produce impulsive electric fields
and bursts of VLF radio energy. Ionosphere airglow observed above the thunderstorm region is associated
with strong electric fields that may be important to 1) the global electric circuit, 2) the making of F region
irregularities, and 3) the production of magnetospheric ducts that propagate VLF radio waves (see
wwwstar.stanford.edu/~vlf/). Lightning-induced Electron Precipitation (LEPs) from the VLF radio waves
may demonstrate a major loss of the earth’s radiation belt. The VLF receiver responds to this
electromagnet radiation in the VLF frequency range 500 HZ to 3kHz and 8 kHz to 33kHz. The VLF loop
antenna, mounted to the deployed thermal shield of the satellite is tuned to the Very Low Frequency
waves with the front end receiver circuit. After amplification they are passed through a switched capacitor
filter where each band and its peak amplification is defined. This filter applies a gain of 20 to the center
frequency of each band. The gain drops dramatically as the frequency changes from the center frequency.
When combined with the frequency nonspecific amplification of the circuit, the peak amplification for the
center of a band-pass is approximately 2000. Following the band pass filter stage each output charges a
capacitor in parallel with a bleed resistor. This allows for low-frequency digitization of the VLF wave
amplitude in each of the frequency bands.
3.4 Educational Impact
The nanosat project will have impacts across the curriculum at Boston University, affecting over 60
undergraduates from several undergraduate engineering programs, and graduate students.
3.4.1 Experience with Spacecraft Education
A nanosat education infrastructure is already in place, due to a prior multidisciplinary NASA
sounding rocket student launch project, SPECTRE, a Motorola-funded engineering curriculum project,
16
STARS (Spacecraft Teaching and Research for Students ), and a previous nanosat project in Astronomy,
Constellation Pathfinder.
SPECTRE - the Student-run Program for Exo-atmospheric Collecting Technologies and Rocket
Experiment, was a sounding rocket payload experiment in NASA's Student Launch Program. Electrical
and computer engineering seniors worked on the flight hardware as a continuing capstone design project
for five semesters, as part of an interdisciplinary student project team, while mechanical engineering
seniors worked on the structural frame, mounts, and enclosures. Students faced rich technical problems
and unique project management challenges arising from the multi-team, multi-semester nature of this
capstone senior design effort. The need to interface regularly with other students, multiple faculty, staff
engineers, and NASA review teams injected real-world pressures into the design course. SPECTRE was
launched on a Nike-Orion sounding rocket from Wallops Island on June 13, 2000. The payload was
recovered, and the data (recorded from payload telemetry) was analyzed by collaborating students from
Wellesley College. These final stages involved grad student leaders who had previously worked on
SPECTRE teams.
Figure 10 - Testing the assembled SPECTRE payload. On the right the
SPECTRE sounding rocket launch at Wallops Island
STARS was funded by a Motorola education grant in
2000, and led to the development of a freshman Introduction to
Engineering module on spacecraft design and control. In this
seven-week laboratory-based course, freshmen learn about
spacecraft instrumentation and actively participate by
integrating sensing, feedback and imaging systems to
implement a simple platform controller for remote sensing.
The freshman course has been taught every fall for six years,
and has reached about 250 students.
STARS funded several student undergraduate projects that led to small test apparatus related to
spacecraft instruments. For example, a 3-axis Helmholtz coil system was constructed with LabVIEW
control to apply a 3D variable field vector to a field sensing instrument. These projects gave
undergraduates hands-on design and fabrication experience, related to spaceflight applications.
Finally, STARS helped fund capstone senior design projects related to spacecraft testing and
subsystems that originated in Astronomy and Space Physics. Typical projects involved sun tracking
electronics, low noise sensor electronics, navigation systems, and dedicated testing systems.
Constellation Pathfinder - Professor of Astronomy Harlan Spence and his colleagues at the Center for
Space Physics had a 1998 Department of Defense funded project called Constellation Pathfinder. This
nanosatellite was to be deployed ultimately as one of a constellation of hundreds of nanosatellites orbiting
high above the Earth. This armada of nanosatellites, each about the size of a four-inch-thick Frisbee, was
designed to collect the data to produce complex, three-dimensional dynamic images of the earth's
magnetosphere, recording the effects of solar coronal mass ejections and tracking storms in space as they
happen.
3.4.2 Student Design Teams
Senior Design in the Electrical & Computer Engineering (ECE) Department consists of two 4
credit required courses that stress open-ended problem solving, team dynamics, written and oral
communication, and project planning and management. Annual enrollment is about 90 students. All
projects are done by teams of four students, usually organized to have a mix of electrical and computer
engineering majors. Projects are solicited from "customers" - local companies, government, public and
17
non-profit groups, faculty and individuals. Students have won numerous prizes including finalist in the
IEEE Computer Society International Design Competition (2006 and 2003), first place-winners in the
NSF Scientific Visualization Award (2003), and IEEE Northeast Design Contest 1st place winners.
In the first semester students are expected to develop a formal proposal, and design a solution. In
the second semester they fabricate a prototype, test their product, and document their efforts. This is an
ambitious schedule considering that most projects arise from the real problems of real customers. Some
problems being attacked again by a new team in the next semester. These reworked projects are different
from a few extended or "legacy" projects, e.g. the IEEE micromouse, or SPECTRE, that are deliberately
maintained over several semesters. In legacy projects new teams are expected to improve incrementally
on the prior design.
Senior design operates as a virtual company, with team leaders reporting vertically to the
professor. There is considerable and regular internal communication by members of the same team.
Regular design reviews, reporting and bench testing are used to keep teams on task.
3.4.3 Summer Research Opportunities
Design efforts continue in summer with students under project funding, funding from grants, or
NSF-funded Research Experiences for Undergraduates positions. The College of Engineering also
supplements summer work with stipends and housing awards. These programs would be able to
supplement the efforts of our nanosat team.
3.5 Spacecraft Integration and Test Plan
Students will develop the hardware for the BUSAT in a spiral design process using two types of
teams. Senior design teams will address specific subsystems using the normal, two semester format of the
capstone course, from September to May. System integration teams will operate each summer, MayAugust, to develop system level designs, to perform system integration, and to conduct functional testing
and critical design review. The Gantt chart Figure 11 identifies the major tasks milestones and schedule
for the planned BUSAT effort.
Senior design teams will follow the model used previously in our successful SPECTRE project.
Four-student teams, with members from the electrical, computer systems, mechanical, and aerospace
engineering programs, will design, build, and unit test subsystems for the Nanosat. Teams will have deep
resources of design tools, shop facilities, and faculty support. Senior Design Teams will have 24/7 access
to the senior design lab, and general access to any other instructional labs and shops in their departments.
They can also use the College of Engineering machine shops, the campus-wide Scientific Instrumentation
Facility, the Center for Space Physics, and the materials labs of the Photonics Center and Fraunhoffer
Center. The research labs of the Photonics Center are also generally available to the teams. Together
these labs and shops will support professional design, simulation, development, fabrication and testing of
our cube modules. Students also have full professional software support, through the campus networks,
ECE membership in the Microsoft Developers Network Academic Alliance, and extensive campus
license agreements with Mathworks (MATLAB, Simulink, all toolboxes), National Instruments
(LabVIEW, DAQ, ELVIS), Cadence (SPECCTRA, OrCAD Capture, Capture CIS, Autorouters, SPICE,
and Layout), and development tools for Freescale, TI, Analog Devices, Xilinx, and other major uC/uP,
RTOS, DSP, and FPGA/PLA hardware. In addition to the four assigned senior design teaching faculty,
the ECE and AME departments make faculty consultants available to the Senior Design Teams.
18
Figure 11 - MicProject schedule illustrating the design process for BUSAT
The student design spiral will start with System Integration Team I in summer 2007. Selected,
experienced undergrads and graduate students will begin the design process and establish system
modularity and interfaces in a set of requirements specifications. This set of requirements will be
assigned to several academic year Senior Design Teams. Those teams will design, fabricate, and test their
subsystems, under the guidance of the BUSAT Project Manager, the senior design course faculty, and
with the BUSAT faculty as “customers”. In summer 2008 the spiral repeats as student System
Integration Team II will review and improve the subsystems and conduct integration testing of the
subsystems. A revised set of requirements will be defined, for improvement of the initial subsystems and
to address any new subsystem requirements. In the second academic year, a set of Senior Design Teams
will work on those requirements for new subsystems with the faculty team as customers. Senior design
“legacy” projects, which have a prior design team record and prototypes, are common, and were used
extensively in SPECTRE. In summer 2009, the System Integration Team III will manage final integration
and conduct the Critical Design Review, leading to production of flight units and testing for flight
readiness. The use of BUSAT for the modeling and forecasting in ASFM will follow after launch.
Using Senior Design Teams allows Nanosat development on a small budget, and ensures that
students will have broad access to resources and faculty. Most usual senior design teams have a nominal
budget of $750, from department budgets and lab fees. The Nanosat teams will be able to purchase
equipment from the project budget. We will also seek corporate support and gifts in kind for hardware,
development resources, and access to industry facilities. The Boston area has several excellent test
facilities for thermal, vibration and EMI testing.
Approximately 60 undergraduates will participate in designing the Nanosat. The System
Integration Teams will be 6-8 students each summer. Distinct Senior Design Teams will be needed for
each instrument cube, and for command, control and communications. Not every unit will need a legacy
19
team in the second year. Likely teams are: (plasma probe, VLF, and communications are subcontracted
with Taylor University):
IES
Auroral Imager
Magnetometer
Exoskeleton
Power
DPU
Attitude
Operational Software
Quick Look Software
3.6 Payload Operation and Control Center
The payload operations and control center will be known as BUCENTRAL and will be located in a
dedicated control room in the Center for Space Physics. The communication link will incorporate a 2
meter antenna, located on the roof of the College of Engineering building. The antenna pointing will be
maintained by a simple PC, the Antenna Control Computer (ACC). The supporting receiving and
transmitting equipment will be found in the control room, where the Satellite Operations Computers
(SOC) will also be found.
The SOC will run the Satellite Operations Software (SOS). This software will be in control of the
data acquisition, satellite command generation and uplink, and command generation for the ACC.
The QLC will run the Quick Look Software (QLS). The QLS will allow the Satellite Operations
Team (SOT) to view the telemetered data immediately after data reception is completed. The software
will be a compilation of the programs developed and used during Integration & Testing (I&T) of the
individual instruments and subsystems. It will therefore from the beginning contain all the necessary tools
for rapid and optimal analysis and problem finding in each of the instruments and subsystems.
Developing the I&T software for reuse as QLS also represents a significant design simplification and cost
saving.
The raw telemetry data will be backed up as the data is received. A dedicated recorder will be
attached to the SOC. The commercially available software Retrospect Remote™ will be installed on the
SOC to perform backups automatically after each pass. Telemetry files will remain on the SOC disk until
removed by the operator, who will be responsible for transferring raw telemetry files to the Data Handling
and Distribution Computer (DHDC). The data acquisition process will run automatically as long as orbit
parameter files are available for pointing the antenna. The orbital parameters file can either be
downloaded by the operator, or it can be automatically generated on, and transferred from the DHDC.
Software is currently available for orbit calculations.
3.7 Flight Operations
After launch, the rocket mating system automatically separates BUSAT from the launch vehicle.
Orbit insertion and satellite separation will be accomplished without any intervention from BUSAT. Upon
insertion, the Attitude Control Magnet (ACM) system will begin the stabilization process. A countdown
timer in the satellite will be set for a minimum delay after separation. Upon expiration, the
communication and data subsystems will be activated. The data system will first run self-tests,
initialization routines, and system checkout routines. Checkout results will be stored and prepared for
ground transmission. After checkout completion, the data system will proceed with power turn-on
routines for engineering subsystems only. The ACM and ADS systems will be activated, and brought
operational. Hereafter, attitude and spacecraft housekeeping data will be recorded and transmitted
downlink.
Spacecraft operations will continue with the activation of the scientific instruments.
Instruments will be tested in different configuration functions and modes of operation. Each instrument
will receive both Real-Time uplink commands and commands from the stored command table (time
tagged commands and/or macros).
Throughout the flight operations phase, the ground station system will operate in a quick look
mode, where simple data displays will allow us to monitor the state-of-health of the payload. Telemetered
20
data will be used to test the DHDC and the DAC portions of the BUCENTRAL Flight operations should
last no longer than 30 days and flight operations could conclude in a shorter period of time if all systems
are judged to be operational and running routinely. Upon conclusion of the flight operations phase, we
will proceed with the Mission Operations and Data Analysis phase (MO&DA).
References:
Fritz, Theodore A. “The passive magnetic alignment of satellite INJUN 3” U. of Iowa Res. Rpt 65-21,
june 1965.
i
A.E. Mahdi, L. Panina, and D. Mapps, “Some new horizons in magnetic sensing: high Tc SQUIDs, GMR and GMI
materials,” Sensors and Actuators A, Vol. 105, (271-285), 2003.
ii
P. Ripka and L. Kraus, “Magnetoimpedance and magnetoinductance,” in Magnetic Sensors and Magnetometers, P. Ripka,
Ed., Norwood, MA: Artech House, (350-358), 2001.
iii
F. Humphrey, “Nearly 20 years of magnetic amorphous wire-an overview,” Journal of Magnetism and Magnetic Materials,
Vol. 249, 1-2, 2002.
iv
K. Mohri, T. Uchiyama, L.P. Shen, C.M. Cai, L.V. Panina, Y. Honkura, and M. Yamamoto, “Amorphous Wire and CMOS
IC-based Sensitive Micromagnetic Sensosrs Utilizing Magnetoimpedance (MI) and Stree-Impedance (SI) Effects,” IEEE
Transactions on Magnetics, Vol. 38, No. 5, (3063-3068), September 2002.
v
H. Hauser, L. Kraus, P. Ripka, “Giant Magnetoimpedance Sensors,” IEEE Instrumnetation & Measurement Magazine, 28-32,
June 2001.
Budget
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