The Deformable Mirror Demonstration Mission
(DeMi) On-Orbit Analysis
by
Jennifer N. Gubner
B.A. in Physics, Wellesley College (2018)
Submitted to the Department of Aeronautics and Astronautics
in partial fulfillment of the requirements for the degree of
Master of Science in Aeronautics and Astronautics
at the
MASSACHUSETTS INSTITUTE OF TECHNOLOGY
June 2021
© 2021 Massachusetts Institute of Technology. All rights reserved.
Author . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Department of Aeronautics and Astronautics
May 18, 2021
Certified by. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Kerri L. Cahoy
Associate Professor, Aeronautics and Astronautics
Thesis Supervisor
Accepted by . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Zoltan Spakovsky
Professor, Aeronautics and Astronautics
Chair, Graduate Program Committee
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The Deformable Mirror Demonstration Mission (DeMi)
On-Orbit Analysis
by
Jennifer N. Gubner
Submitted to the Department of Aeronautics and Astronautics
on May 18, 2021, in partial fulfillment of the
requirements for the degree of
Master of Science in Aeronautics and Astronautics
Abstract
The Deformable Mirror Demonstration Mission (DeMi) is a 6U CubeSat mission to
demonstrate the use of a 140 actuator microelectromechanical system (MEMS) de-
formable mirror (DM) and a closed-loop adaptive optics (AO) system in space. DeMi
launched to the International Space Station (ISS) on the NG-13 Cygnus resupply
mission on February 15, 2020 and was deployed from the ISS into a 51° inclination,
423 km average altitude low-Earth orbit on July 13, 2020. The expected mission life-
time of DeMi was 6 months, however DeMi continues to be operational 9 months post
deployment. During its lifetime, DeMi has completed several internal observations
with the DM and both imagers using the internal laser source. The team is now work-
ing toward external observations of stars and demonstrations of closed-loop wavefront
control. The biggest driver of mission success is spacecraft and component health.
Looking at spacecraft data over time helps to characterize spacecraft performance
and inform adjustments to the lifetime estimate. Additionally, telemetry analysis can
alert the operations team of anomalies and provide useful information for resolving
those anomalies. This thesis analyzes the spacecraft telemetry received between July
13, 2020 and April 4, 2021 and discusses trends and anomalies in the data. This work
provides an overview of spacecraft and payload on-orbit health to date and provides
recommendations on paths forward for anomaly resolution.
Thesis Supervisor: Kerri L. Cahoy
Title: Associate Professor, Aeronautics and Astronautics
3
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Acknowledgments
I would like to extend a very special thank you to my advisor, Professor Kerri Cahoy,
who has given me incredible guidance and support throughout my graduate education
and my undergraduate involvement with STAR Lab. Thank you for helping me grown
as a student and a person and allowing me to take on responsibilities within the DeMi
project.
I would also like to thank the DeMi team for their support, advice, and encourage-
ment. I would specifically like to thank Rachel Morgan, Joey Murphy, Bobby Holden,
Christian Haughwout, Paula do Vale Pereira, Greg Allan, and Ewan Douglas. With-
out the help and support of each of you, this work would be impossible. Thank you
for making my experience on the team so enjoyable and educating. Thank you to
John Merk and Danilo Roascio for your management of the project.
Thank you to my friends and family, especially Sydney, for putting up with my
weird hours for operating the spacecraft and for listening to me talk about the project
to almost no end.
I would also like to thank Blue Canyon Technologies for their support and guidance
on the spacecraft bus systems. I would specifically like to thank Steve Stem and Bryan
Button for answering my many questions and helping me understand the spacecraft.
The DeMi project is sponsored by DARPA and has been managed by Aurora
Flight Sciences, a Boeing Company.
5
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6
Contents
1 Introduction 21
1.1 Project Motivation . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21
1.1.1 Exoplanet Direct Imaging . . . . . . . . . . . . . . . . . . . . 21
1.1.2 Adaptive Optics Systems . . . . . . . . . . . . . . . . . . . . . 22
1.1.3 Mission Objectives and Requirements . . . . . . . . . . . . . . 24
1.2 Mission Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
1.2.1 Introduction to CubeSats . . . . . . . . . . . . . . . . . . . . 24
1.2.2 DeMi Configuration . . . . . . . . . . . . . . . . . . . . . . . . 25
1.2.3 Concept of Operations (ConOps) . . . . . . . . . . . . . . . . 27
1.2.4 Orbit and Lifetime Estimates . . . . . . . . . . . . . . . . . . 28
1.3 Thesis Contributions . . . . . . . . . . . . . . . . . . . . . . . . . . . 29
2 Command and Data Handling Configuration 31
2.1 Wallops Pipeline Configuration . . . . . . . . . . . . . . . . . . . . . 33
2.2 UHF Low-Rate Configuration . . . . . . . . . . . . . . . . . . . . . . 34
2.3 Ground Station Statistics . . . . . . . . . . . . . . . . . . . . . . . . 37
3 Payload Operations Analysis 39
3.1 Deformable Mirror and Driver Status . . . . . . . . . . . . . . . . . . 39
3.2 Flight Computer Status . . . . . . . . . . . . . . . . . . . . . . . . . 42
3.3 Imagers and Laser Status . . . . . . . . . . . . . . . . . . . . . . . . . 44
3.4 Payload Operations Progress . . . . . . . . . . . . . . . . . . . . . . . 47
3.5 Overall Payload Health Discussion . . . . . . . . . . . . . . . . . . . 48
7
4 Bus Operations Analysis 53
4.1 Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53
4.2 Temperatures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56
4.3 Electric Power System . . . . . . . . . . . . . . . . . . . . . . . . . . 62
4.4 Radio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64
4.5 Attitude Determination and Control Systems . . . . . . . . . . . . . . 65
4.5.1 Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67
4.5.2 Pointing Attempts and Results . . . . . . . . . . . . . . . . . 82
4.6 Bus Resets . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 85
4.7 Event Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 89
4.8 Overall Bus Health Discussion . . . . . . . . . . . . . . . . . . . . . . 91
5 Results, Discussion, and Future Work 93
5.1 Results and Discussion . . . . . . . . . . . . . . . . . . . . . . . . . . 93
5.2 Future Work . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 94
5.2.1 Operations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 95
5.2.2 Telemetry Analysis . . . . . . . . . . . . . . . . . . . . . . . . 96
A List of Acronyms 97
B Beta Angle and Solar Illumination Script 99
C Bus Reset Reconfiguration Script 103
8
List of Figures
1-1 A generic representation of an AO system. A perturbed wavefront
enters the optical system and is reflected through the system by a
DM. A portion of the light (amount determined by the beamsplitter)
is transmitted to an image sensor and the rest of the light is reflected
to the WFS. The WFS records the incoming wavefront and a closed-
loop algorithm supplies a corrective shape to the DM. The corrected
wavefront is recorded by the high resolution camera [15]. . . . . . . . 23
1-2 Image of the fully integrated spacecraft. The bus is the main box that
houses all of the components and the bus electronics are in the top
right corner, taking up about 1.5U of the spacecraft. The payload is
housed in the remaining L-shaped volume. The spacecraft baffle, in
black, is mounted to the top left of the internal volume. Next to the
baffle is the payload electronics stack in green. The electronics stack
includes the two Raspberry Pi compute modules and the DM driver
boards. The payload optical bench is mounted in the lower part of the
bus volume and the two camera boards are mounted to the right of the
payload optical bench. This image is taken from [18]. . . . . . . . . . 26
1-3 Optical path of the DeMi payload, taken from Morgan, et al. [19]. . . 27
9
1-4 Graphic of the concept of operations for the DeMi mission. DeMi
aboard the NG-13 ISS resupply mission inside a Cygnus spacecraft on
February 15, 2020. DeMi was stored aboard the ISS until July 13, 2020
when it was deployed into a 51° inclination orbit at a mean altitude
of 423 km above the Earth. After deployment, solar array actuation,
and detumble, the spacecraft was commissioned for regular operations.
These operations include pre-operation, or calibration, internal oper-
ation using an internal laser as the target of interest, and external
operation using an astronomical target. . . . . . . . . . . . . . . . . . 28
1-5 Block diagram of the different modes of operation available on the
DeMi mission. After launch and commissioning, the satellite pro-
gressed into standby mode waiting for the operation commands. From
there, the operations team can command the satellite to take calibra-
tion images, followed by either internal or external mode. Both modes
have several experiment options that are executed depending on what
the team is trying to observe. Finally, the data is compressed and
downlinked via one of the radios on the bus. Diagram taken from [18]. 29
1-6 Plot created by graduate student Joey Murphy showing the orbital
decay of the DeMi mission through April 22, 2021 along with the orbital
decay of a previous CubeSat mission, ASTERIA, with a similar form-
factor. After 283 days in orbit, DeMi has decayed by about 20 km
and it is estimated that approximately 700 days remain before DeMi
de-orbits. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
10
2-1 The MIT ground station configuration during a pass with DeMi. On
the screen are several COSMOS tools, along with pass statistics tools
and antenna control tools. The top right window, “Command and
Telemetry Server,” shows the connection statistics, the “Command
Sender” window (hidden behind the waterfall plot and the "Script
Runner") is where the user can configure and send a command to the
satellite, the window at the center, the “Script Runner,” allows the user
to run scripts that automate and control the timing of commands and
telemetry, and the window hidden directly behind the "Script Runner"
window, “Packet Viewer,” lets the user see the various states and val-
ues in a specific telemetry packet. There are other available COSMOS
tools not shown here [16]. . . . . . . . . . . . . . . . . . . . . . . . . 32
2-2 Diagram of the connection to and from the DeMi satellite via the Cadet
radio. The NASA Wallops Flight Facility has a ground station antenna
that is configured to interface with DeMi’s Cadet radio and route the
telemetry and commands between the DeMi satellite and the MIT
ground station. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33
2-3 Image of the MIT ground station antenna on the roof of building 37.
The MIT ground station uses a Lithium radio to communicate with the
DeMi satellite with an uplink frequency of 449.775MHz and a downlink
frequency of 401MHz. . . . . . . . . . . . . . . . . . . . . . . . . . . 34
2-4 Diagram of the connection to and from the DeMi satellite via the
Lithium radio. MIT has a receive and transmit antenna located on
the roof of building 37. The team formats commands and receives
telemetry from the ground station computer on campus and the data
is transmitted through the antenna. . . . . . . . . . . . . . . . . . . . 35
11
2-5 Distribution of the number of passes by ground station. The WFF
ground station is the NASA Wallops Flight Facility ground station
and the campus ground station is the MIT ground station located on
the roof of building 37. This data is from the start of the mission in
July, 2020 to April 5, 2021. . . . . . . . . . . . . . . . . . . . . . . . . 37
2-6 Distribution of data volume, in 𝐾𝑖𝑙𝑜𝑏𝑦𝑡𝑒𝑠, received per pass by ground
station through April 5, 2021. . . . . . . . . . . . . . . . . . . . . . . 38
3-1 Figure of the DeMi electronics stack. The top two boards are the
Raspberry Pi flight computers and the bottom three boards are the
DM driver boards. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43
3-2 The top subplot shows Payload 1 Raspberry Pi temperature plotted as
the black points with the throttling limit shown as the gold horizon-
tal line and commanded Payload 1 operations shown as blue vertical
dotted lines. The operations line at the beginning of August demon-
strated successful commanding of the DM. The bottom subplot shows
the beta angle (red) in degrees and the percentage of an orbit that
DeMi is illuminated by the sun (light blue). . . . . . . . . . . . . . . 45
3-3 The top subplot shows Payload 2 Raspberry Pi temperature plotted as
the black points with the throttling limit shown as the gold horizontal
line and commanded Payload 2 operations shown as green vertical
dotted lines. The bottom subplot shows the beta angle (red) in degrees
and the percentage of an orbit that DeMi is illuminated by the sun
(light blue). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46
3-4 Shack-Hartmann wavefront sensor temperatures with commanded Pay-
load 1 operations shown as the blue dotted vertical lines. The imager
sensor temperature is shown as the orange points, and the imager body
temperature is shown as the blue points. . . . . . . . . . . . . . . . . 47
12
3-5 Image plane camera temperatures with commanded Payload 2 oper-
ations shown as the green dotted vertical lines. The imager sensor
temperature is shown as the orange points, and the imager body tem-
perature is shown as the blue points. . . . . . . . . . . . . . . . . . . 48
4-1 All of the spacecraft data partitions SD card usage shown as a percent-
age over time. Black is the total usage, red is the payload usage, blue
is the FSW usage, orange is the SOH usage, green is the line usage,
and purple is the table usage. . . . . . . . . . . . . . . . . . . . . . . 54
4-2 SD percent usage of the 5 spacecraft data partitions along with total
usage shown in separate subplots with different y-axis scales to show
details. The colors represent the same data values as in Figure 4-1. . 55
4-3 Photo showing the placement of the thermistor on the payload optical
bench. (a) The red box shows the zoomed in location for the temper-
ature sensor placement. (b) The temperature sensor is the small black
wire and probe, shown inside the red circle, that is placed between the
field mirror and the small off-axis parabolic mirror. This was the safest
location for placing the temperature probe close to the DM, which is
in the top left of the zoomed-in red box. . . . . . . . . . . . . . . . . 57
4-4 Payload heater placement on the underside of the optical bench. The
two kapton heaters are shown inside the yellow rectangles. The heater
in the left part of the image is placed underneath the DM and the other
heater is placed close to underneath the Shack-Hartmann wavefront
sensor (SHWFS). The component placements on the optical bench are
shown in Figure 1-3. These heaters are automatically triggered on or off
when the payload temperature sensor reaches a threshold temperature
specific to the mode of operation. . . . . . . . . . . . . . . . . . . . . 57
13
4-5 Payload optical bench temperature over the mission duration plotted
with heater status and operating mode. Heater setpoints for each oper-
ating mode are also shown. The blue points represent periods of heaters
in operating mode and the red points represent periods of heaters in
survival mode. The X’s represent heaters off and the dots represent
heaters on. The operating mode heater setpoints are shown in blue
with the on setpoint at 15 °C and the off setpoint at 20 °C. The sur-
vival mode heater setpoints are shown in red with the on setpoint at
10 °C and the off setpoint at 15 °C. . . . . . . . . . . . . . . . . . . . 58
4-6 Payload optical bench temperature shown as the black points with
commanded operations by both flight computers shown as the vertical
lines.The bottom subplot shows the beta angle (red) in degrees and
the percentage of an orbit that DeMi is illuminated by the sun (light
blue). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59
4-7 Payload/bus interface heater and temperature sensor location. The
yellow rectangles indicate the location of the two kapton heaters and
the red circle indicates the location of the temperature sensor. . . . . 60
4-8 Payload/bus interface temperature over the mission duration plotted
with heater status and operating mode. Heater setpoints for each op-
erating mode are also shown. The blue points represent periods of
operating mode and the red points represent periods of survival mode.
The X’s represent heaters off and the dots represent heaters on. The
operating mode heater setpoints are shown in blue with the on set-
point at 15 °C and the off setpoint at 20 °C. The survival mode heater
setpoints are shown in red with the on setpoint at 10 °C and the off
setpoint at 15 °C. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61
4-9 Payload/bus interface temperature shown as the black points in the top
subplot. The bottom subplot shows the beta angle (red) in degrees and
the percentage of an orbit that DeMi is illuminated by the sun (light
blue). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
14
4-10 Battery voltage and current over mission duration. Battery voltage is
plotted in blue in the top plot with the left axis indicators. Battery
current is shown in orange with the right axis indicators. The bottom
subplot shows the beta angle (red) in degrees and the percentage of an
orbit that DeMi is illuminated by the sun (light blue). . . . . . . . . . 63
4-11 Battery temperatures in the top subplot. The bottom subplot shows
the beta angle (red) in degrees and the percentage of an orbit that
DeMi is illuminated by the sun (light blue). . . . . . . . . . . . . . . 64
4-12 Bus voltage through April 4, 2021 shown with battery voltage for ref-
erence. The periods of higher voltage (in the 19V range) are periods
when the solar arrays are facing the sun and the periods of lower volt-
age (in the 12V range) are when the spacecraft is operating off of
battery power. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65
4-13 Cadet radio temperature history. Temperatures remain between 10 °C
and 40 °C for all of the mission data except one anomalous point on
January 19, 2021 where it is recorded as 307 °C. . . . . . . . . . . . . 66
4-14 Cadet radio temperature shown in the top subplot. The bottom sub-
plot shows the beta angle (red) in degrees and the percentage of an
orbit that DeMi is illuminated by the sun (light blue). . . . . . . . . . 66
4-15 Spacecraft body frame. +x is from the solar panel plane into the
satellite. +z is the face of the satellite with the payload and star
tracker aperture. Image courtesy of Blue Canyon Technologies and
adapted from [8]. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67
4-16 Reaction wheel speed given in RPM over the mission, shown with com-
manded pointing periods as the dotted red vertical lines. . . . . . . . 69
4-17 Reaction wheel drag with commanded pointing periods shown as the
dotted red vertical lines. . . . . . . . . . . . . . . . . . . . . . . . . . 70
4-18 Reaction wheel temperatures shown in the top subplot. The bottom
subplot shows the beta angle (red) in degrees and the percentage of an
orbit that DeMi is illuminated by the sun (light blue). . . . . . . . . . 71
15
4-19 Star tracker detector temperature in the top subplot. The bottom
subplot shows the beta angle (red) in degrees and the percentage of an
orbit that DeMi is illuminated by the sun (light blue). . . . . . . . . . 72
4-20 Star tracker baffle temperature in °C shown as the black dots. The bot-
tom subplot shows the beta angle (red) in degrees and the percentage
of an orbit that DeMi is illuminated by the sun (light blue). . . . . . 72
4-21 The median mean star tracker background brightness level plotted with
detector temperature as the color. The horizontal dotted gold line is
the saturation point for the star tracker detectors as 1023 counts. . . 73
4-22 Star tracker estimated quaternions where Q4 is the scalar term. The
red dashed vertical lines show commanded pointing attempts. . . . . 74
4-23 Star tracker estimated right ascension, declination, and roll. The red
dashed vertical lines show commanded pointing attempts. . . . . . . . 75
4-24 Star tracker estimated rates for X, Y, and Z in the spacecraft body
frame. The red dashed vertical lines show commanded pointing attempts. 75
4-25 Raw sun sensor data for all sun sensor diodes. . . . . . . . . . . . . . 76
4-26 Raw sun sensor diode data from each of the 4 diodes in each of the
three sun sensor packages. Sun sensor diodes 1-4 correspond to package
1, diodes 5-8 correspond to package 2, and diodes 9-12 correspond to
package 3. The bottom-most subplot shows the beta angle (red) in
degrees and the percentage of an orbit that DeMi is illuminated by the
sun (light blue). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77
4-27 Number of diodes that “see” the sun per sun sensor package plotted
with commanded pointing attempts.The red dashed vertical lines show
commanded pointing attempts. The bottom-most subplot shows the
beta angle (red) in degrees and the percentage of an orbit that DeMi
is illuminated by the sun (light blue). . . . . . . . . . . . . . . . . . . 78
4-28 Sun vector status. This is an indication of how many diodes see the sun. 79
4-29 Sun body vector estimated from sun sensor data. . . . . . . . . . . . 80
16
4-30 IMU rate estimates in radians per second for X, Y, and Z in the space-
craft body frame. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80
4-31 GPS validity. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81
4-32 Estimated magnetic field vector in the spacecraft body frame based on
magnetometer data. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82
4-33 Total system momentum in Nms. The system momentum is dumped
by the torque rods when it exceeds threshold value. . . . . . . . . . . 83
4-34 Sun in payload FOV keepout event check. The red dashed vertical
lines show commanded pointing attempts. The triggering points are
when the status changes from 1 to 0. A state of 1 means the event
check in enabled, a state of 0 means the event check is disabled. . . . 84
4-35 Spacecraft ADCS mode with pointing command periods requested
by the operations team. FINE_REF_POINT is when the space-
craft is actively pointing at a target, such as the ground station, and
SUN_POINT is the default state where the solar panels are pointed
at the sun. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 85
4-36 Commanded attitude quaternions. The red dashed vertical lines show
commanded pointing attempts. Q4 is the scalar term and is always
positive. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86
4-37 Commanded x, y, and z acceleration. The red dashed vertical lines
show commanded pointing attempts. . . . . . . . . . . . . . . . . . . 87
4-38 Commanded x, y, and z rates. The red dashed vertical lines show
commanded pointing attempts. . . . . . . . . . . . . . . . . . . . . . 88
4-39 Bus resets, shown as the pink vertical lines, and high rate run count
of the flight software, shown as the black dotted lines. The high rate
run count is run by the flight software and is reset to 0 every time the
software (or bus) reboots. . . . . . . . . . . . . . . . . . . . . . . . . 90
17
4-40 Cadet radio lockup event check with bus resets shown as the pink
vertical lines. All of the points on the 0.0 line that come right after the
pink bus reset lines are not actual event check triggers, they are just
incidents of the event check reverting back to default state upon reset.
The points in mid August, mid September, and early January do not
align with observed bus resets and are possibly Cadet lockup events. . 91
18
List of Tables
3.1 Table of Payload 1 operations captured between July 13, 2020 and
April 4, 2021. Operations include poke tests to varying percentages
of actuation, calibration dark and bias frames, spotfield images of the
laser on the detector of the SHWFS, and DM commands like wfs and
zernikes that apply shapes to the DM. . . . . . . . . . . . . . . . . . 50
3.2 Table of Payload 2 operations captured between July 13, 2020 and
April 4, 2021. Operations include individual actuator pokes to vary-
ing percentages of actuation, point spread function (PSF) images of
the laser on the detector of the image plane camera, and DM zernike
commands that apply shapes to the DM. . . . . . . . . . . . . . . . . 51
4.1 Table of the dates of bus resets between July 13, 2020 and April 4,
2021. May of the resets were due to South Atlantic Anomaly (SAA)
radiation strikes, a few were commanded resets, and the others have
undetermined causes. . . . . . . . . . . . . . . . . . . . . . . . . . . . 89
4.2 DeMi event checks to ensure fault tolerance during operations. . . . . 90
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Chapter 1
Introduction
This thesis is a continuation of the project originally developed by Dr. Anne Marinan
in her 2016 Ph.D. thesis [13] and is an extension of the author’s undergraduate thesis
[9]. It builds upon the prior work presented with the exciting addition of on-orbit
performance analysis1.
1.1 Project Motivation
1.1.1 Exoplanet Direct Imaging
Since it was first proposed in 1953 by H.W. Babcock, adaptive optics (AO) has been
increasingly used on ground-based telescopes to correct for the effects of atmospheric
turbulence on observations [4]. With the miniaturization of AO components, these ca-
pabilities can be applied to space-based experiments such as exoplanet direct imaging.
While imaging astrophysical objects in space is usually not affected by atmospheric
turbulence, other AO correctable factors can still affect the resolution and contrast
of the image. These factors include dynamic noise generated by the reaction wheels
of the spacecraft, thermal variations within the spacecraft due to changes in solar
illumination throughout its orbit, and optical misalignment or imperfections inherent
to the system or induced by launch or deployment [2].
1Sections 1.1 and 1.2 of this chapter contain similar content to the author’s undergraduate thesis,
[9], as this thesis is a continuation of the same project.
21
AO correctional capabilities in space are particularly useful in several main ap-
plications [12]. One use of AO is to improve coupling of signals to fiber for ground-
to-satellite and satellite-to-satellite laser communication, enabling higher data rate
communication signals. AO can be applied to high-contrast imaging of dim celestial
objects. The DeMi mission can be categorized in the celestial object imaging cat-
egory, as it is intended to demonstrate capabilities for direct imaging of Earth-like
exoplanets. Specifically, AO can be the critical difference in reaching the necessary
contrast ratio to image Earth-like exoplanets that are typically 1010 times fainter than
their host star [20][22].
The correctional capabilities of AO systems can also enable the use of larger and
less expensive optics on space telescopes because optical imperfections can be actively
corrected. The use of AO can improve imaging performance and stretch the limits of
what scientists can observe.
1.1.2 Adaptive Optics Systems
DeMi’s AO system has three main components: a deformable mirror, a wavefront
sensor, and a camera. The component at the center of the mission and the center
of the AO system is the deformable mirror (DM). The job of the DM is to supply
physical corrections to the path length of the incoming wavefront. There are differ-
ent types of DMs with different sizes, actuation methods, and configurations. Some
examples of these are discussed in [11] for ground-based systems. The DM on DeMi
is a microelectromechanical systems (MEMS) DM. MEMS DMs are convenient for
use in space because of their compact size, low weight and power, low actuator mass,
and high actuator density [2]. The wavefront sensor (WFS) captures and records
the incoming wavefront, as shown in Figure 1-1. The goal of the WFS is to pro-
vide an estimate of the shape and a measurement of the deviations in the wavefront.
There are different types of wavefront sensors, including curvature sensors, pyramid
sensors, and Shack-Hartmann wavefront sensors [23]. DeMi has a Shack-Hartmann
WFS (SHWFS), which uses an array of microlenslets and a CMOS imager to capture
its measurements. Once the wavefront is captured by the WFS, a closed-loop con-
22
trols system reconstructs the incoming wavefront, measures the wavefront errors, and
supplies a correction shape to the DM. When the DM receives the correctional shape
from the WFS algorithm, it actuates the various sections of the mirror to reshape it-
self into the corrected shape. This correctional cycle repeats itself to continually make
adjustments to the wavefront as the light enters the system. Finally, the image plane
sensor captures a point spread function (PSF) of the reference source that incorpo-
rates the changes made by the DM and the wavefront sensing algorithm. Figure 1-1
shows a schematic of a typical AO system. See Section 1.1.3 for more information on
the specific DeMi configuration.
Figure 1-1: A generic representation of an AO system. A perturbed wavefront enters
the optical system and is reflected through the system by a DM. A portion of the light
(amount determined by the beamsplitter) is transmitted to an image sensor and the
rest of the light is reflected to the WFS. The WFS records the incoming wavefront
and a closed-loop algorithm supplies a corrective shape to the DM. The corrected
wavefront is recorded by the high resolution camera [15].
23
1.1.3 Mission Objectives and Requirements
Prior to the DeMi mission, high actuator count MEMS DMs, which are a key com-
ponent to a space-based AO system, were not qualified or characterized for use in
space. The objective of DeMi, a 6U (30 cm × 20 cm × 10 cm) CubeSat mission, is
to demonstrate the use of AO in space, as well as to characterize, and increase the
technology readiness level (TRL) of a 140-actuator MEMS DM. References [2], [9],
and [17] have more details on the specific science requirements of the mission.
The components on the DeMi payload and the mission’s science requirements
informed the spacecraft environmental, power, and pointing requirements. In order
to operate the payload components, the payload optical bench must remain between
0 °C and 30 °C. In order to minimize thermal-induced misalignments of the optical
system, the payload should maintain a temperature of 20± 4°C during operations [1]
. Throughout the entirety of the mission, the payload must remain within the survival
temperature range of −20 °C to 70 °C. In addition to maintaining the temperatures
of the spacecraft, the bus must provide power to the payload on a 5V line and two
separate 3.3V lines. The spacecraft battery voltage should not be below 11.5V. The
pointing stability of the spacecraft should be less than 10′′ in all three axes [14].
1.2 Mission Overview
1.2.1 Introduction to CubeSats
CubeSats were first introduced in 1999 and are a type of small satellite, whose stan-
dard was developed by the California Polytechnic State University (Cal Poly), de-
signed to improve access to space by reducing the cost and development time required
for space missions [6]. The specification for CubeSats provides the size of each satellite
based on the number of units, or U’s of volume. Each U is a 10 cm× 10 cm× 10 cm
cube. The DeMi CubeSat is a 6U and is configured in a 30 cm × 20 cm × 10 cm
shape with about 4.5U of space dedicated to the scientific payload and the other 1.5U
dedicated to the spacecraft bus electronics and attitude determination and control
24
systems (ADCS). With the growing popularity of CubeSats, the methods of deploy-
ment of these satellites have also become increasingly standardized. Some of the DeMi
mission requirements are therefore derived from the CubeSat design and deployment
standards.
1.2.2 DeMi Configuration
The DeMi spacecraft can be broken out into two main parts: the spacecraft bus and
the payload. The spacecraft bus was provided by Blue Canyon Technologies (BCT)
and is a 6U XACT bus. The payload, designed and built by the team at MIT is
contained in approximately 4.5U of the 6U bus.
The bus provides all of the critical support for the mission. In order to maintain
the appropriate temperatures, the bus is equipped with several temperature sensors
to record thermal variations throughout the spacecraft and two sets of two kapton
heaters. One set is mounted on the interface between the bus and the payload, and
the other set is mounted on the underside of the payload optical bench. The specific
configuration of heaters is discussed in more detail in Section 4.2. The main spacecraft
battery can be charged up to 12.3V by the two solar panels. To provide communica-
tion between the satellite and the spacecraft operators, the spacecraft uses one of two
UHF radios. A Cadet radio is used for high data rate communication with the NASA
Wallops Flight Facility ground station, and a Lithium radio is used for communication
with the MIT ground station. See Chapter 2 for more details on the communication
configuration. The bus is also responsible for the attitude determination and control
(ADCS) of the spacecraft. To provide the necessary pointing accuracy, the bus uses 3
reaction wheels, a star tracker, three packages of 4 sun sensor diodes, an inertial mea-
surement unit (IMU), a GPS system, torque rods, and a magnetometer. Section 4.5
discusses the on-orbit performance of these ADCS components.
The DeMi payload includes an optical bench with all of the mounted optical
components, a baffle to block out stray light at the external aperture, a payload
electronics stack, and two camera boards. Figure 1-2, taken from [18], shows the
fully integrated payload and bus system with the baffle in the top left, the payload
25
electronics stack to the right of the baffle, the payload optical bench on the bottom
and the camera boards mounted to the right of the optical bench.
Figure 1-2: Image of the fully integrated spacecraft. The bus is the main box that
houses all of the components and the bus electronics are in the top right corner, taking
up about 1.5U of the spacecraft. The payload is housed in the remaining L-shaped
volume. The spacecraft baffle, in black, is mounted to the top left of the internal
volume. Next to the baffle is the payload electronics stack in green. The electronics
stack includes the two Raspberry Pi compute modules and the DM driver boards.
The payload optical bench is mounted in the lower part of the bus volume and the
two camera boards are mounted to the right of the payload optical bench. This image
is taken from [18].
The optical bench consists of a series of off-axis parabolic mirrors (OAPs), field
mirrors, two complementary metal-oxide semiconductor (CMOS) cameras, a Shack-
Hartmann Wavefront Sensor (SHWFS) and a Boston Micromachines Corporation
(BMC) 140-actuator MEMS DM. The optical layout is shown in Figure 1-3. The
payload directs the light source, either internal or external, through a beamsplitter
and uses part of the beam for the closed-loop wavefront correction and the other part
to image the point spread function (PSF). All of the components and operations are
controlled by two Raspberry Pi Compute Module 3 flight computers. References [1],
[9], and [19] discuss the optical design in more detail.
26
Figure 1-3: Optical path of the DeMi payload, taken from Morgan, et al. [19].
1.2.3 Concept of Operations (ConOps)
DeMi is configured to be able to perform both internal and external observations dur-
ing its lifetime. The internal observations are used to characterize the DM and to test
the closed-loop wavefront control system with a 635 nm internal laser. The external
observations are intended to demonstrate the use of the AO system on astronomical
targets. Before each science operation, the operations team checks the battery volt-
age levels and the payload temperatures, in addition to the spacecraft pointing, if
doing an external observation. After the checks, the operators send a command to
the satellite to perform an operation either immediately or for execution at a later
time. At the specified time, the spacecraft powers on the required components for
the specific mode of operation. If commanded to do so, the payload will take baseline
measurements and image frames to measure dark current and sensor bias measure-
ments prior to capturing an image of the target. Figure 1-4 shows a diagram of the
mission concept of operations and Figure 1-5 shows a block diagram of the various
27
modes of operations.
Figure 1-4: Graphic of the concept of operations for the DeMi mission. DeMi aboard
the NG-13 ISS resupply mission inside a Cygnus spacecraft on February 15, 2020.
DeMi was stored aboard the ISS until July 13, 2020 when it was deployed into a 51°
inclination orbit at a mean altitude of 423 km above the Earth. After deployment,
solar array actuation, and detumble, the spacecraft was commissioned for regular
operations. These operations include pre-operation, or calibration, internal operation
using an internal laser as the target of interest, and external operation using an
astronomical target.
1.2.4 Orbit and Lifetime Estimates
Prior to launch, the DeMi mission had an estimated lifetime of about 6 months.
The lifetime estimate comes from a combination of component lifetime estimates,
radiation tolerances, thermal cycling lifetimes, and orbital decay historical data for
CubeSats. DeMi was deployed into a circular orbit from the ISS with an initial mean
altitude of 423 km and an inclination of 51°. As of April 22, 2021, DeMi is at a
mean altitude of 403 km, having dropped 20 km in 9 months. Figure 1-6, created by
graduate student Joey Murphy, shows a comparison between the ASTERIA CubeSat
mission and DeMi. Both missions have similar form-factors and thus, it is expected
that DeMi will follow an orbital decay rate similar to the ASTERIA mission. Based
on the historical data from the orbital decay of ASTERIA, it is expected that DeMi
has about 700 days left before it de-orbits. However, it is possible that some of the
critical science components will not last as long.
28
Figure 1-5: Block diagram of the different modes of operation available on the DeMi
mission. After launch and commissioning, the satellite progressed into standby mode
waiting for the operation commands. From there, the operations team can command
the satellite to take calibration images, followed by either internal or external mode.
Both modes have several experiment options that are executed depending on what
the team is trying to observe. Finally, the data is compressed and downlinked via
one of the radios on the bus. Diagram taken from [18].
1.3 Thesis Contributions
This thesis analyzes the DeMi spacecraft on-orbit data between deployment on July
13, 2020 and April 4, 2021 for both the payload and the spacecraft bus. Chapter 2
provides an overview of the command and data handling system that enables the
retrieval of the data from the spacecraft. Chapter 3 looks at the payload performance
including sensor status and completed operations to-date. Chapter 4 discusses the
performance of the spacecraft bus including thermal management, power distribution,
and attitude determination and control. Finally, Chapter 5 summarizes the findings
from the spacecraft data and outlines the future work for the DeMi mission.
29
Figure 1-6: Plot created by graduate student Joey Murphy showing the orbital decay
of the DeMi mission through April 22, 2021 along with the orbital decay of a previous
CubeSat mission, ASTERIA, with a similar form-factor. After 283 days in orbit, DeMi
has decayed by about 20 km and it is estimated that approximately 700 days remain
before DeMi de-orbits.
30
Chapter 2
Command and Data Handling
Configuration
Command and data handling infrastructure is one of the most important subsystems
for any space mission. Without it, there would be no way to command operations
or return data to the ground. Command and data handling (C&DH) is the system
that manages the sending and receiving of commands and telemetry to and from the
spacecraft. The DeMi mission uses two different onboard radios that interface with
two different ground stations on the East Coast of the United States. Additionally,
the DeMi mission uses Ball Aerospace’s open-source COSMOS software [16], which
provides a helpful interface to check connections with the satellite and radio, format
commands to send to the satellite, format and process telemetry, execute automated
scripting to send commands and check telemetry, and record entire communication
sessions. Figure 2-1 shows some of the COSMOS tools that the DeMi team uses
during passes with the satellite. Because the DeMi team contracted Blue Canyon
Technologies (BCT) to provide the 6U bus, the team received a BCT-modified ver-
sion of COSMOS upon delivery of the bus that already included the command and
telemetry configurations for the bus-specific items. In addition, the COSMOS version
received from BCT was configured to interface with data from both of DeMi’s radios:
a Lithium radio and a Cadet radio. Each radio type has a corresponding ground
station setup described in more detail throughout this chapter.
31
Figure 2-1: The MIT ground station configuration during a pass with DeMi. On
the screen are several COSMOS tools, along with pass statistics tools and antenna
control tools. The top right window, “Command and Telemetry Server,” shows the
connection statistics, the “Command Sender” window (hidden behind the waterfall
plot and the "Script Runner") is where the user can configure and send a command
to the satellite, the window at the center, the “Script Runner,” allows the user to run
scripts that automate and control the timing of commands and telemetry, and the
window hidden directly behind the "Script Runner" window, “Packet Viewer,” lets
the user see the various states and values in a specific telemetry packet. There are
other available COSMOS tools not shown here [16].
32
2.1 Wallops Pipeline Configuration
The Cadet radio on the DeMi CubeSat was configured to have a direct link with
the NASA Wallops Flight Facility (WFF) ground station located in Wallops Island,
Virginia. This is a UHF 3Mbps connection with a downlink frequency of 468MHz
and an uplink frequency of 449.775MHz. The WFF ground station is connected
over the network to the ground station computer located at MIT. It is from the MIT
ground station computer that the team formats commands and monitors the incoming
telemetry, however the WFF ground station actually sends the commands over radio
frequency (RF) up to the spacecraft and receives the telemetry from the spacecraft
over RF before routing it to the MIT ground station computer that is equipped with
the COSMOS tool to process the telemetry. A diagram of this connection is shown
in Figure 2-2.
Figure 2-2: Diagram of the connection to and from the DeMi satellite via the Cadet
radio. The NASA Wallops Flight Facility has a ground station antenna that is con-
figured to interface with DeMi’s Cadet radio and route the telemetry and commands
between the DeMi satellite and the MIT ground station.
The WFF ground station was intended to be the high data rate ground station
used for file downlinks from the satellite. Unfortunately, the WFF ground station
went off-line due to mechanical failures in September of 2020 and did not return
to operations until March 22, 2021. Between July and September of 2020, when
both DeMi and the WFF ground station were operational, the DeMi team was able
to use the WFF ground station for over 40 overpasses to command the satellite and
downlink telemetry. Since March 22, 2021, the DeMi team has successfully completed
4 WFF passes, one of which produced a mission record data volume of 3,115,984 bytes
33
received. Because the high data volume passes are supported by WFF, which has been
offline for a majority of the DeMi mission thus far, the DeMi team has been actively
working on ways to supplement the lower rate downlinks. These include modifying
the MIT ground station software, working on logic to request smaller portions of files,
and actively looking for other ground stations that are able and willing to support
DeMi at the Cadet uplink and downlink frequencies. More information about these
modifications can be found in Subsection 2.2, UHF Low-Rate Configuration.
2.2 UHF Low-Rate Configuration
The other ground station, configured in large part for a DeMi communication link by
MIT graduate student Joey Murphy, is located on the roof of building 37 at MIT.
This ground station is configured for UHF uplink and downlink and has been used for
communications with other satellites, including MiRaTa [5]. An image of the antenna
is shown in Figure 2-3. This communication link uses a Lithium radio with an uplink
frequency of 449.775MHz and a downlink frequency of 401MHz to communicate with
MIT. The Lithium has a data rate of 9600 bps and uses GMSK modulation scheme
[3]. Figure 2-4 shows a diagram of the UHF low-rate configuration.
Figure 2-3: Image of the MIT ground station antenna on the roof of building 37. The
MIT ground station uses a Lithium radio to communicate with the DeMi satellite
with an uplink frequency of 449.775MHz and a downlink frequency of 401MHz.
34
Figure 2-4: Diagram of the connection to and from the DeMi satellite via the Lithium
radio. MIT has a receive and transmit antenna located on the roof of building 37. The
team formats commands and receives telemetry from the ground station computer on
campus and the data is transmitted through the antenna.
Initially, the MIT ground station was intended for telemetry monitoring and oper-
ation scheduling on-board the DeMi spacecraft. Scheduling operations and receiving
state of health telemetry should nominally require relatively low data rates. After the
temporary decommissioning of the WFF ground station, the MIT ground station be-
came the only working link with DeMi and was then required for file downlinks as well
as the previously mentioned telemetry monitoring and operation scheduling. As dis-
cussed in Section 2.1, the file downlink process involves downlinking a large quantity
of data. With only the low data rate configuration available for the DeMi mission,
there were many configuration adjustments that were needed to accommodate the
downlink of data.
One of the first major changes to the configuration was the way the ground station
software handles the data packets. When the DeMi team switched to using the MIT
ground station to receive data from the satellite, for some unknown reason the data
packets coming from the Lithium radio were not reaching the COSMOS processing
software, even though the Cadet radio data packets had no issue and the rest of
the Lithium telemetry packets were being sent to COSMOS. The DeMi team is still
looking into why this is happening for the Lithium connection and specifically the
data packets only. In the meantime, Joey Murphy and the DeMi team put a lot
of effort into processing the incoming data packets right as it is received through
35
the COM port on the ground station computer. Joey created a tool, called the COM
Sniffer by the DeMi team, which intercepts the telemetry stream so that it can process
the data packets before routing the data through to COSMOS. This tool is able to
successfully intercept the chunks of data sent by the satellite, parse the chunks, and
save them in a similar structure to that used by COSMOS.
In addition to the changes for handling data downlinks on the ground station
computer, the team had to make changes to the communication architecture in or-
der to command a file downlink from the Lithium. Not only do both radios have
significant differences in available data rates, but they also have different maximum
command and telemetry packet lengths. This difference in command packet length
led to problems when requesting a file for downlink because that specific command
exceeded the maximum command length on the Lithium. As a result, MIT gradu-
ate student Bobby Holden and the DeMi team created a work-around in which the
team uploads several shell scripts to the payload that move the file and execute the
downlink when it is run. While this fix works and is the best option we have at this
moment for Lithium-specific data downlinks, it is a multi-step process that usually
requires several passes to upload and execute and is very inefficient as compared with
the single downlink request command.
The temporary loss of the WFF ground station also brought about other ineffi-
ciencies in the DeMi team’s ability to receive data. One of the major inefficiencies
arose out of the inability to easily downlink directory listings from the satellite, as
these are treated the same as a file downlink and would require the multi-step shell
script uplink process described above. Because the DeMi team was not able to easily
view the existing file list for downlinks, and the command to downlink a file requires
knowing the exact name and location of the file, the DeMi team implemented a re-
naming strategy to re-name the top 9 desired files with generic names that are easier
to request without errors.
36
2.3 Ground Station Statistics
As of April 5, 2021, the DeMi mission has completed 534 successful passes, where
success means that data has been received from the satellite. Due to the unavailability
of the WFF ground station for a large portion of the mission so far, the majority of
DeMi’s passes have been taken by the MIT ground station. Figure 2-5 shows a
breakdown of the number of successful passes taken by each ground station. As of
April 5, 2021, the campus ground station at MIT is nearing 500 DeMi passes and the
WFF ground station is nearing 50.
Figure 2-5: Distribution of the number of passes by ground station. The WFF ground
station is the NASA Wallops Flight Facility ground station and the campus ground
station is the MIT ground station located on the roof of building 37. This data is
from the start of the mission in July, 2020 to April 5, 2021.
Throughout the mission, the campus ground station has collected on average
about 31 kB of data from the satellite each pass with a maximum of 125.748 kB from
a single pass. The WFF ground station has collected 402 kB of data on average each
pass, and holds the record for the most data volume collected on a single DeMi pass
at 3115.984 kB. A distribution of data volume received per pass by ground station is
shown in Figure 2-6.
Chapter 3 discusses how some of this data is used to understand payload health
behavior over the mission lifetime.
37
Figure 2-6: Distribution of data volume, in 𝐾𝑖𝑙𝑜𝑏𝑦𝑡𝑒𝑠, received per pass by ground
station through April 5, 2021.
38
Chapter 3
Payload Operations Analysis
The DeMi payload consists of the payload optical bench, a baffle for external observa-
tions, an electronics stack, and two camera boards. The optical bench configuration,
which was described in Chapter 1, holds all of the optical components including both
imagers and the deformable mirror (DM). The electronics stack consists of two main
flight computers, referred to as Payload 1 and Payload 2, and several DM driver
boards. The electronics stack is discussed in more detail in Section 3.2. The two
camera boards are connected to the electronics stack and the two imagers by cables,
but they are not mounted with the electronics stack or on the optical bench.
The reader should keep in mind that the data presented here is only the data
received from the satellite, and is therefore limited in quantity. Due to the ground
station configuration and pass priorities, the operators have not downnlinked all of
the payload playback telemetry data. Therefore, the data presented here may not
completely represent the payload behavior on-orbit. However, the data provides in-
sight into the status of the payload health. The data is plotted with times shown in
UTC.
3.1 Deformable Mirror and Driver Status
The main science component of interest for the DeMi mission is the DM. The DM is
the basis for the wavefront control demonstration and for the technology demonstra-
39
tion. Without the DM, the mission’s science objectives are unattainable. Because
DM operation is so critical to mission success, the driver boards that control the DM
actuation are also critical. The driver boards were designed in-house by MIT gradu-
ate student Christian Haughwout, whose master’s thesis covers the details of design
and implementation [10].
Once the satellite was deployed and commissioned, the team began checking out
all of the on-board systems. One of the first tests the operations team attempted was
a poke test. The poke test is a diagnostic test used during ground and environmental
testing, and the team has a solid understanding of the baseline poke test performance.
A poke test is an operation where each DM piston is actuated to a specified percentage
of the maximum displacement, and the actual displacement of each piston is measured
by the Shack-Hartmann wavefront sensor (SHWFS) and the payload flight electronics.
By running the poke test, the team can confirm that the internal laser, the SHWFS,
the DM, and the payload flight electronics are working as expected. Additionally, this
test produces a smaller data file because the displacement measurements are returned
rather than an actual image, making this test a good choice for on-orbit testing of
component functionality and for measuring DM performance.
On August 4, 2020, shortly after the beginning of the mission, the team suc-
cessfully captured and downlinked a poke test that confirmed DM actuation, laser
operation, and SHWFS operation. In data analysis done by MIT graduate student
Rachel Morgan, deflection of the mirror in space was compared with ground testing,
and the results from in space agreed with ground data with a median |𝑔𝑟𝑜𝑢𝑛𝑑−𝑠𝑝𝑎𝑐𝑒| of
𝑔𝑟𝑜𝑢𝑛𝑑
5.35%. On September 11, 2020, the team captured an image with the image plane
camera that confirmed the laser was still working and that the image plane camera
was functional.
Shortly after these images were taken and downlinked, the Wallops Flight Facility
went offline and remained offline for a significant portion of the mission to-date. The
DeMi operations team continued attempting to collect data, but downlinking the new
data was a very slow process and resulted in a lot of inconclusive results due to lack
of data. On March 25, 2021, the team was able to confirm that the laser and SHWFS
40
were still operational, however data from December, 2020 and March, 2021 that the
team received only recently, showed no signs of DM actuation.
Currently, the team’s top priority is figuring out why the DM actuation is not
working as expected and how it can be resolved. The team is actively working on
ruling out issues with:
• The laser
• The imagers
• The DM
• The DM driver boards
• The power delivery to the DM
• Command delivery to the DM
• The Raspberry Pi flight computers
• The Raspberry Pi interface with the DM driver boards
So far, imager and laser operation has been confirmed through the capturing of spot-
field images on the SHWFS and PSFs on the image plane camera. It is possible that
the issue could be with the DM. The DM could be sensitive to the space environment
which, over time, may have caused damage to the component. If this is the only issue,
the team may not be able to conclusively confirm it, but it will become increasingly
likely as the other possible issues are ruled out. Currently, the team is working on
confirming DM driver board operation as well as power and command delivery to
the DM. The original intention was to include a complete capture of this data in
the payload housekeeping telemetry. However, due to time constraints, not all DM
telemetry measurements were implemented before delivery, but several implementa-
tions are planned for future uplink. Several scripts to record power delivery to the
DM are in development and will soon be uplinked to the satellite. Once the team is
41
able to return the contents of the logs created by these scripts, the power delivery
and DM commanding can be confirmed or addressed.
Another system that could be contributing to DM actuation failure is the Rasp-
berry Pi flight computers. Through analysis of on-chip temperature measurements,
discussed in more detail in Section 3.2, it appears that the temperatures of the Rasp-
berry Pis have been consistently above a widely recognized limit where the CPU
begins to throttle its processing rates. It is possible that the flight computers are not
behaving as expected, and the DM is not receiving properly formatted commands.
To rule this out as an issue, the team will be performing a series of thermal tests
on the ground and will be looking more closely at the flight computer temperatures
during operations from the environmental testing campaign. Future testing will also
be done on the Raspberry Pi interface with the DM and driver boards if the power
analysis and flight computer temperature analysis come back inconclusive.
3.2 Flight Computer Status
The two flight computers for the DeMi payload are Raspberry Pi Compute Module 3’s.
They were selected for their small size, configurability, and flight heritage. Figure 3-1
shows the DeMi electronics stack. The top two boards are the Raspberry Pis and
the bottom boards are DM driver boards. The top Raspberry Pi board is referred to
as Payload 2 and the bottom Raspberry Pi is referred to as Payload 1. Each of the
two flight computers are equipped with two SD cards, A and B. The primary flight
computer configuration we have been using is Payload 1 on SD card A. Payload 1
is in charge of capturing images on the Shack-Hartmann wavefront sensor (SHWFS)
and Payload 2 is responsible for capturing images on the image plane sensor. Both
Payload 1 and Payload 2 can control the DM and perform closed-loop wavefront
control, however the two flight computers cannot operate at the same time.
Throughout most of the mission, the DeMi team has been using the Payload 1
flight computer to try to capture images on the SHWFS. Section 3.4 discusses these
operations in more detail. Two telemetry items of particular interest are the payload
42
Figure 3-1: Figure of the DeMi electronics stack. The top two boards are the Rasp-
berry Pi flight computers and the bottom three boards are the DM driver boards.
43
flight computer Raspberry Pi temperatures. Each Raspberry Pi is equipped with
an on-chip silicon temperature sensor that measures the temperature of each flight
computer. These temperatures are included in the payload housekeeping telemetry
packets, which are returned to the operations team when the flight computers are
powered on. During payload operations, the Raspberry Pi temperatures should not be
above the CPU throttling temperature of 80 °C, where the on-board clock frequency
is throttled down to minimize thermal loading. Figure 3-2 shows the Payload 1
flight computer temperature over time with the throttling limit shown as the gold
horizontal line and the commanded Payload 1 operation times shown as the vertical
blue dashed lines. Figure 3-3 shows the same, but for Payload 2. During most of the
payload operations, except for the successful DM actuation on August 4, 2020, the
Raspberry Pi temperatures appear to be at or over the throttling limit. On August
4, 2020, the temperatures appear to be cooler and closer to 70 °C. Throttling may
be contributing to some of the DM issues discussed in Section 3.1. To rule out flight
computer throttling as a contributing factor to the DM performance issue, the on-orbit
Raspberry Pi temperatures will be compared with the Raspberry Pi temperatures
taken during ground performance and environmental testing. Additionally, the team
will work to better monitor Raspberry Pi temperatures prior to commanding science
operations to mitigate taking images when the temperatures are close to the known
throttling limit.
3.3 Imagers and Laser Status
While the DM has not shown recent evidence of actuation, the team has confirmed
that the SHWFS, the image plane sensor, and the laser are still functional.
Although complete DM power telemetry is not yet available in the payload house-
keeping telemetry, temperature data from both imager sensors and housings is avail-
able. Figure 3-4 and 3-5 show temperatures from the SHWFS and image plane sensor
respectively. The orange points represent sensor temperatures and the blue points
represent body temperatures. In most cases, sensor and body temperatures agree
44
Figure 3-2: The top subplot shows Payload 1 Raspberry Pi temperature plotted
as the black points with the throttling limit shown as the gold horizontal line and
commanded Payload 1 operations shown as blue vertical dotted lines. The operations
line at the beginning of August demonstrated successful commanding of the DM. The
bottom subplot shows the beta angle (red) in degrees and the percentage of an orbit
that DeMi is illuminated by the sun (light blue).
45
Figure 3-3: The top subplot shows Payload 2 Raspberry Pi temperature plotted
as the black points with the throttling limit shown as the gold horizontal line and
commanded Payload 2 operations shown as green vertical dotted lines. The bottom
subplot shows the beta angle (red) in degrees and the percentage of an orbit that
DeMi is illuminated by the sun (light blue).
46
to within 5 °C. The values on the -1 °C line are when the payload was on, but the
wavefront controller and cameras had not been turned on yet. The values on the 0 °C
line are from when the cameras were turned on but had not been used yet to take
an image. The temperatures from all of the commanded operations appear to be
between 20 °C and 35 °C for the SHWFS and between 15 °C and 30 °C for the image
plane. The operating range for the cameras we are using on DeMi is between 0 °C and
50 °C [21], so the cameras have been within the operating limits during all operations.
Figure 3-4: Shack-Hartmann wavefront sensor temperatures with commanded Pay-
load 1 operations shown as the blue dotted vertical lines. The imager sensor temper-
ature is shown as the orange points, and the imager body temperature is shown as
the blue points.
3.4 Payload Operations Progress
Overall, the operations team has commanded and successfully generated data files
for almost 45 science operations as of April, 2021. These commanded operations
include poke tests, images on the SHWFS, images on the image plane sensor, and
calibration images which include bias and dark frames. Tables 3.1 and 3.2 show all
of the successfully captured images and data files. Not all of the images and data
collected have yielded successful results, but the images and data were successfully
captured by the payload.
47
Figure 3-5: Image plane camera temperatures with commanded Payload 2 operations
shown as the green dotted vertical lines. The imager sensor temperature is shown as
the orange points, and the imager body temperature is shown as the blue points.
The successful DM actuation data from August 4th was a commanded poke test
to 30% actuation. Additionally, the team has downlinked images from the SHWFS
which showed that the sensor and laser are working, but that the DM actuation is
not.
3.5 Overall Payload Health Discussion
Overall, the payload remains semi-operational. The payload computers are both re-
sponding to commands from the operations team, the cameras are functional, the
laser is working, and telemetry is successfully being transmitted to the ground sta-
tions. Unfortunately, the team has been unable to confirm DM actuation since the
WFF ground station returned to operations.
As discussed in [9] and [18], the main objectives for the DeMi mission are to:
1. Demonstrate on-orbit performance of a 140 actuator Boston Micromachines
Corporation DM,
2. Raise the technology readiness level (TRL) of the microelectromechanical sys-
48
tems (MEMS) DM to a 7,
3. Correct static and dynamic wavefront errors to less than 100 nm RMS, and
4. Measure low order aberrations to 𝜆 precision and 𝜆 accuracy.
50 10
DeMi has met objectives 1 and 2, but has yet to complete 3 and 4. Having measured
DM actuator displacement on orbit successfully, the team hopes to continue with
demonstrating wavefront correction. The DM actuation is a critical component of
the wavefront correction demonstration, and that is why all of the team’s efforts
are focused on getting the DM operational again. When, or if, the DM becomes
operational again, the team will demonstrate wavefront correction with the internal
laser before moving to demonstration on external targets. A future plan for the DeMi
mission is also to try to capture an image of the moon or other astrophysical target
on the image plane camera, although this operation would not be related to any of
the mission objectives that take top priority.
For the payload to be successful, the spacecraft bus needs to be healthy and oper-
ating as expected. Chapter 4 analyzes spacecraft health and performance throughout
the mission and discusses how certain payload operations may contribute to or be
affected by spacecraft operations.
49
Payload 1 Operations
Date of Operation (UTC) Type of Operation
8/4/20 19:55 poke test (30%)
8/21/20 07:29 dark, bias
8/22/20 06:39 spotfield, dark, bias
8/22/20 13:19 poke test (15%)
8/25/20 06:08 poke test (30%)
8/25/20 12:30 poke test (30%)
9/2/20 02:34 dark, bias
9/2/20 09:01 spotfield
9/3/20 01:52 poke test (15%)
10/19/20 15:24 poke test (30%)
10/19/20 15:37 spotfield
10/20/20 14:36 poke test (15%)
10/21/20 09:06 poke test (15%)
10/21/20 13:42 poke test (30%)
10/21/20 14:01 bias, dark
10/22/20 12:58 poke test (15%)
11/30/20 14:53 poke test (30%)
12/1/20 14:04 poke test (30%)
12/2/20 13:26 poke test (30%)
12/2/20 19:23 poke test (30%)
12/2/20 21:12 poke test (30%)
12/3/20 14:00 poke test (30%)
12/11/20 14:22 poke test (30%)
12/12/20 15:01 poke test (30%)
3/13/21 04:02 wfc
3/15/21 04:54 zernike
3/15/21 22:02 poke test (90%)
3/16/21 02:49 spotfield
3/22/21 00:39 spotfield
3/22/21 20:35 spotfield
3/24/21 00:29 spotfield
3/25/21 01:02 spotfield
Table 3.1: Table of Payload 1 operations captured between July 13, 2020 and April 4,
2021. Operations include poke tests to varying percentages of actuation, calibration
dark and bias frames, spotfield images of the laser on the detector of the SHWFS,
and DM commands like wfs and zernikes that apply shapes to the DM.
50
Payload 2 Operations
Date of Operation (UTC) Type of Operation
9/11/20 23:08 psf
9/20/20 19:04 zernike
9/20/20 19:06 zernike
10/1/20 21:22 zernike
10/2/20 14:00 zernike
10/2/20 14:01 poke actuator (30%)
10/2/20 14:02 poke actuator (30%)
12/12/20 15:04 psf
4/2/21 05:04 poke actuator (30%)
4/2/21 06:40 poke actuator (30%)
4/3/21 12:19 zernike
4/3/21 12:20 zernike
4/3/21 12:22 zernike
Table 3.2: Table of Payload 2 operations captured between July 13, 2020 and April
4, 2021. Operations include individual actuator pokes to varying percentages of ac-
tuation, point spread function (PSF) images of the laser on the detector of the image
plane camera, and DM zernike commands that apply shapes to the DM.
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Chapter 4
Bus Operations Analysis
As discussed in Chapter 1, the DeMi payload is integrated into a Blue Canyon Tech-
nologies (BCT) 6U XACT bus. The BCT bus provides the DeMi mission with all
of the necessary communication, power, thermal, attitude determination and control,
and command and data handling subsystem interfaces. Because the BCT bus acts as
the life support system for the satellite, it is critical that all of the bus components
function as required throughout the mission lifetime. This chapter looks at all of the
critical spacecraft bus components and analyzes the bus performance throughout the
mission to-date.
The reader should keep in mind that the data presented here is only the data
received from the satellite, and is therefore limited in quantity. Due to the ground
station configuration and pass priorities, the operators have not downnlinked all of
the spacecraft health data. Therefore, the data presented here may not completely
represent the spacecraft behavior on-orbit. However, the data provides insight into
the status of the payload health. The data is plotted with times shown in UTC.
4.1 Storage
One of the main purposes of the bus flight electronics is to receive data from the
DeMi payload, format the data into telemetry packets, and downlink the data as re-
quested by the operations team on the appropriate radio. Additionally, the bus flight
53
electronics store important health related data in order to downlink the spacecraft
health status to the operations team upon request. The spacecraft data is split up
into 5 partitions: payload, flight software (FSW), state of health (SOH), tables, and
line. Payload is all of the data received from the payload. This includes images and
payload housekeeping data that the payload sends to the bus to transmit back to
Earth. The FSW partition holds information about the current flight software and
its configuration. SOH contains information on the state of health of the spacecraft
and its operation. Line data is used for the star tracker buffer, and tables are stored
information about the data configuration on the satellite. Figure 4-1 shows the usage
of all of the partitions on the SD cards together on one plot. In addition to each
individual partition, the spacecraft reports on the total usage.
Figure 4-1: All of the spacecraft data partitions SD card usage shown as a percentage
over time. Black is the total usage, red is the payload usage, blue is the FSW usage,
orange is the SOH usage, green is the line usage, and purple is the table usage.
To maintain storage capabilities on the SD cards, when the payload data reaches
90% usage, the flight software will delete the oldest payload data until the payload
usage returns to 80%. From the second subplot in Figure 4-2, you can see that at the
end of January, the payload usage reached the limit and the flight software performed
the expected clean-up of the data. Currently, the DeMi team is considering manually
54
clearing out some of the payload data in order to preserve the more important files
and images. Below the payload partition, the FSW and SOH usages look sporadic
and seem to remain in between 80% and 90% usage. There is likely no discernable
trend with this data because it is constantly being stored on the SD cards and only
a small portion of the data has been received by the ground station. If we were
to receive all of the spacecraft data, we would likely see a quick increase and clear
out of the SOH and FSW partitions in a similar pattern as the payload partition.
FSW and SOH usage data also indicates that those partitions are being cleared out
appropriately once they reach 90%. The total usage reflects all of the clearing of
payload, FSW, and SOH data and appears to remain under 70% usage. Overall, the
SD card management appears to be performing as expected.
Figure 4-2: SD percent usage of the 5 spacecraft data partitions along with total
usage shown in separate subplots with different y-axis scales to show details. The
colors represent the same data values as in Figure 4-1.
55
4.2 Temperatures
As stated in Chapter 1, thermal management of the spacecraft is critical for the
safe operation of the payload. The acceptable temperature ranges are defined by
the payload and bus operational and survival limits. The acceptable temperature
range for payload operation is between 0 °C and 30 °C, with an ideal temperature
of 20± 4°C. However, the payload can survive, with all components off, between -
10°C and 65 °C. These ranges are limited by the most thermally sensitive component
on the payload structure. The DM has the lowest maximum operating temperature
of 30 °C and the cameras have the highest minimum operating temperature of 0 °C.
Additionally, the bus operational temperature range is from -20°C to 60 °C.
The bus is responsible for measuring the spacecraft temperatures at various points
on the structure and turning on or off the spacecraft heaters when appropriate. One
of the most important temperature sensors on the spacecraft is located on the pay-
load optical bench. Figure 4-3 shows the placement of this temperature sensor on
the payload. This sensor ensures that the optical bench, and most importantly the
deformable mirror (DM), remains within the required temperature limits at all times.
Additionally, when the temperature reading reaches threshold values, the heaters are
triggered on or off. Figure 4-4 shows the heater placements on the underside of the
payload optical bench. During spacecraft survival mode, where none of the payload
components are on, the heater-on setpoint is 10 °C and the heater-off setpoint is 15 °C.
During payload operational mode, where the payload components are on and poten-
tially capturing images, the heater-on set point is at 15 °C and the heater-off setpoint
is at 20 °C. The ideal temperature for payload operations is 20± 4°C.
Figure 4-5 shows the payload optical bench temperature and heater status through-
out the mission. The blue points represent times when the payload heaters were in
operational mode and the red points represent survival mode. The X’s represent
heater off points and the dots represent heater on. The operating mode heater set-
points are shown in blue at 15 °C on and 20 °C off. The survival mode heater setpoints
are shown in red at 10 °C on and 15 °C off. From the available data, the payload tem-
56
Figure 4-3: Photo showing the placement of the thermistor on the payload optical
bench. (a) The red box shows the zoomed in location for the temperature sensor
placement. (b) The temperature sensor is the small black wire and probe, shown
inside the red circle, that is placed between the field mirror and the small off-axis
parabolic mirror. This was the safest location for placing the temperature probe
close to the DM, which is in the top left of the zoomed-in red box.
Figure 4-4: Payload heater placement on the underside of the optical bench. The two
kapton heaters are shown inside the yellow rectangles. The heater in the left part of
the image is placed underneath the DM and the other heater is placed close to under-
neath the Shack-Hartmann wavefront sensor (SHWFS). The component placements
on the optical bench are shown in Figure 1-3. These heaters are automatically trig-
gered on or off when the payload temperature sensor reaches a threshold temperature
specific to the mode of operation.
57
perature does not drop below 10 °C at any point. Additionally, the heaters are off for
a majority of the available data. Furthermore, the heaters are never on in survival
mode above the survival heaters off setpoint and the same goes with the operating
mode heaters. Finally, the payload optical bench temperature remains between 10 °C
and 25 °C, which is within the limits of 0 °C to 30 °C required to operate the optical
components. At most times, the optical bench temperature remains in the 20± 4°C
range during operating mode.
Figure 4-5: Payload optical bench temperature over the mission duration plotted with
heater status and operating mode. Heater setpoints for each operating mode are also
shown. The blue points represent periods of heaters in operating mode and the red
points represent periods of heaters in survival mode. The X’s represent heaters off
and the dots represent heaters on. The operating mode heater setpoints are shown in
blue with the on setpoint at 15 °C and the off setpoint at 20 °C. The survival mode
heater setpoints are shown in red with the on setpoint at 10 °C and the off setpoint
at 15 °C.
Figure 4-6 shows the payload optical bench temperature in the top plot and solar
illumination and beta angle in the bottom plot. Solar illumination and beta angle is
calculated by a script that was created by MIT graduate student Joey Murphy and is
included in Appendix B. The top plot also shows vertical dotted lines that represent
the times of commanded science operations for both payload flight computers. It
appears that during periods of operations, the payload optical bench temperature
58
runs slightly higher than normal payload non-operation. On the bottom plot, solar
illumination is given in terms of percent of orbit illuminated by the sun and beta
angle is given in degrees. It appears that payload operations seem to have more of
an impact on payload optical bench temperature than solar illumination in orbit.
Figure 4-6: Payload optical bench temperature shown as the black points with com-
manded operations by both flight computers shown as the vertical lines.The bottom
subplot shows the beta angle (red) in degrees and the percentage of an orbit that
DeMi is illuminated by the sun (light blue).
In addition to the payload optical bench thermal management system, the bus
provides thermal management for the payload/bus interface. The term interface is
used here as the point of contact between the optical bench and the spacecraft bus.
On the underside of the payload optical bench, three titanium standoffs attach the
optical bench to the bus, as seen in Figure 4-4. These standoffs allow the payload to
be significantly thermally isolated from the bus and allow some flexibility for thermal
deformations [7]. The payload/bus interface is also equipped with a temperature
sensor and two kapton heaters. These thermal components can be seen in Figure 4-
7. The payload/bus interface heaters have the same operational and survival mode
on/off setpoints as the payload heaters.
Figure 4-8 shows the payload/bus interface temperature and heater status through-
59
Figure 4-7: Payload/bus interface heater and temperature sensor location. The yellow
rectangles indicate the location of the two kapton heaters and the red circle indicates
the location of the temperature sensor.
out the mission. The blue points represent times when the payload heaters were in
operational mode and the red points represent survival mode. The X’s represent
heater off points and the dots represent heater on. The operating mode heater set-
points are shown in blue at 15 °C on and 20 °C off. The survival mode heater setpoints
are shown in red at 10 °C on and 15 °C off. From the available data, the interface
temperature does not drop below 0 °C at any point and seems to remain at or be-
low 20 °C. Additionally, the heaters are on in operating mode for a majority of the
available data. Similarly to the optical bench thermal management performance,
the interface heaters are never on above the corresponding heater operating mode
setpoints.
Figure 4-9 shows the payload/bus interface temperature plotted with solar illu-
mination and beta angle respectively. In this case, thermal variations due to eclipse
status of the spacecraft may have more of an effect on the temperature. However,
with limited data it is hard to say for sure. Since the heaters are on for a majority
60
Figure 4-8: Payload/bus interface temperature over the mission duration plotted
with heater status and operating mode. Heater setpoints for each operating mode
are also shown. The blue points represent periods of operating mode and the red
points represent periods of survival mode. The X’s represent heaters off and the dots
represent heaters on. The operating mode heater setpoints are shown in blue with the
on setpoint at 15 °C and the off setpoint at 20 °C. The survival mode heater setpoints
are shown in red with the on setpoint at 10 °C and the off setpoint at 15 °C.
61
of this dataset, they can be thought of as a constant heat source and so most of
the temperature variations likely come from orbital configuration changes, or thermal
discharge from spacecraft electronics.
Figure 4-9: Payload/bus interface temperature shown as the black points in the top
subplot. The bottom subplot shows the beta angle (red) in degrees and the percentage
of an orbit that DeMi is illuminated by the sun (light blue).
4.3 Electric Power System
The electric power system (EPS) on the DeMi spacecraft is responsible for providing
power to all of the spacecraft and payload components. The EPS consists of a battery,
two solar panels, and a power routing board with voltage lines at 3.3V, 5V, 8V, and
12V.
The battery holds a maximum charge of 12.9V based on design documents, but
an observed maximum of 12.4V. Figure 4-10 shows both the battery voltage and
the battery current draw, along with spacecraft orbit percent sunlit and beta angle.
Voltage is plotted in blue on the left axis and current is plotted in orange on the
right. The battery current is positive when the battery is being discharged, and is
62
negative when the battery is being charged by the solar arrays. The received on-orbit
spacecraft data indicates that battery voltage has never gotten over 12.4V and has
never gone under 11.5V, staying within the battery power requirements. Additionally,
the periods of negative current seem to align well with the periods of full power in
the batteries and the periods of positive current align well with periods of decreased
battery state of charge.
Figure 4-10: Battery voltage and current over mission duration. Battery voltage is
plotted in blue in the top plot with the left axis indicators. Battery current is shown
in orange with the right axis indicators. The bottom subplot shows the beta angle
(red) in degrees and the percentage of an orbit that DeMi is illuminated by the sun
(light blue).
In addition to charge level constraints, the batteries operate under thermal con-
straints. In order to safely charge, the batteries must remain between 0 °C and 45 °C.
To safely discharge, the batteries must remain between -20°C and 60 °C. Figure 4-11
shows battery temperatures over the current mission lifetime. The battery tempera-
ture has stayed well within the tighter temperature range of 0 °C to 45 °C throughout
the received mission data.
63
Figure 4-11: Battery temperatures in the top subplot. The bottom subplot shows the
beta angle (red) in degrees and the percentage of an orbit that DeMi is illuminated
by the sun (light blue).
Finally, the unregulated spacecraft voltage is shown in Figure 4-12 with battery
voltage for reference. When the solar arrays are facing the sun and the spacecraft
is charging, the unregulated spacecraft voltage reads in the 19V range. When the
spacecraft is operating on battery, the unregulated spacecraft voltage should match
the battery voltage.
4.4 Radio
As discussed in Chapter 2, there are two radios on the DeMi CubeSat: the Cadet
and the Lithium. The Cadet is the high data rate UHF radio that interfaces with
the NASA Wallops Flight Facility (WFF) ground station. The Lithium is a UHF
radio that interfaces with the MIT ground station. Unfortunately, temperatures
from the Lithium radio are not available in the spacecraft data. There is, however,
a history of Cadet temperatures on orbit that are shown in Figure 4-13. Cadet
radio temperatures look normal with the exception of one data point from January
19, 2021, when the temperature data shows 307 °C. It is possible that during the
64
Figure 4-12: Bus voltage through April 4, 2021 shown with battery voltage for ref-
erence. The periods of higher voltage (in the 19V range) are periods when the solar
arrays are facing the sun and the periods of lower voltage (in the 12V range) are
when the spacecraft is operating off of battery power.
conversion of the raw data to the temperature values, an error occurred. However, the
team has been unable to locate exactly what caused the anomalous value. Because
temperature measurements directly before and directly after this anomalous point
show temperatures in the 30’s, this outlier is likely not a major cause for concern.
Zooming in on the temperature range between 10 °C and 40 °C, the fluctuations
of Cadet temperature are more visible. Figure 4-14 shows these fluctuations along
with solar illumination and beta angle.
4.5 Attitude Determination and Control Systems
The attitude determination and control system (ADCS) components are primarily
used for orienting the satellite appropriately. This is to help processes like spacecraft
battery charging and pointing at targets for external observations. The DeMi team
has attempted several pointing operations, but has yet to take full advantage of the
ADCS system on DeMi by taking observations of external targets.
Many of the commands to point the satellite, as well as many of the telemetry
65
Figure 4-13: Cadet radio temperature history. Temperatures remain between 10 °C
and 40 °C for all of the mission data except one anomalous point on January 19, 2021
where it is recorded as 307 °C.
Figure 4-14: Cadet radio temperature shown in the top subplot. The bottom subplot
shows the beta angle (red) in degrees and the percentage of an orbit that DeMi is
illuminated by the sun (light blue).
66
items that capture how the spacecraft is pointing and moving, are given in terms
of the spacecraft body frame. Figure 4-15 shows how the spacecraft body frame
coordinate system is oriented relative to the satellite.
Figure 4-15: Spacecraft body frame. +x is from the solar panel plane into the satellite.
+z is the face of the satellite with the payload and star tracker aperture. Image
courtesy of Blue Canyon Technologies and adapted from [8].
4.5.1 Components
The BCT bus is equipped with a suite of ADCS components to detumble the space-
craft upon deployment and point the spacecraft in the desired configuration. These
components include 3 reaction wheels, a star tracker, 3 packages of 4 sun sensor
diodes, an inertial measurement unit (IMU), a GPS, a three-axis magnetometer, and
3 torque rods.
Reaction Wheels
One of the main components used to orient the spacecraft is the set of reaction
wheels. The DeMi spacecraft has three reaction wheels and their speeds in rotations
per minute over the mission are shown in Figure 4-16. In addition to showing reaction
wheel speeds, the figure shows the periods where the spacecraft was commanded to
67
point by the operators. Many of the big jumps in speed correspond to pointing
commands given by the operations team.
Figure 4-17 shows the reaction wheel drag along with the commanded pointing
periods. Measuring the drag is a good way to understand reaction wheel performance
degradation throughout the mission. In DeMi’s case, the wheel drag has remained
fairly low and relatively stable throughout the mission. The highest wheel drag seen
to date is around 65 rad s−2. We would start to be concerned if drag estimates were
in the 100’s range for a long period of time. Furthermore, the drag only remained at
elevated levels for a short period of time.
Looking at the reaction wheel temperatures, shown in Figure 4-18, we can see a
harsh and persistent temperature jump between February 10th and March 2nd, 2021.
During this period, the temperature data jumped up to around 100 °C and 80 °C.
Despite close analysis of the February 9th data, the author has been unable to find a
reason for this jump in temperature. Interestingly, this behavior is also present for all
of the other ADCS component temperatures, shown in Figures 4-19 and 4-20. All of
the temperature readings that showed this anomalous behavior go through the same
A/D converter, so it is likely that the temperature jump was not representative of the
true temperatures during this period. Furthermore, the anomaly appears to resolve
itself on March 2nd, which corresponds to an observed bus reset. This reset likely
cycled whatever was causing the false readings. Section 4.6 contains more information
on the observed bus resets throughout the mission.
Star Tracker
The DeMi spacecraft uses a star tracker to measure its orientation and provide critical
information to the ADCS components. Figures 4-19 and 4-20 show the star tracker
detector temperature and star tracker baffle temperature respectively. Each is also
plotted with solar illumination and beta angle in subplots. Just like with the reaction
wheel temperatures, the star tracker temperatures also display the anomalous increase
in temperature from February 10th and March 2nd, 2021.
Figure 4-21 shows the star tracker mean background brightness level, or back-
68
Figure 4-16: Reaction wheel speed given in RPM over the mission, shown with com-
manded pointing periods as the dotted red vertical lines.
69
Figure 4-17: Reaction wheel drag with commanded pointing periods shown as the
dotted red vertical lines.
70
Figure 4-18: Reaction wheel temperatures shown in the top subplot. The bottom
subplot shows the beta angle (red) in degrees and the percentage of an orbit that
DeMi is illuminated by the sun (light blue).
71
Figure 4-19: Star tracker detector temperature in the top subplot. The bottom
subplot shows the beta angle (red) in degrees and the percentage of an orbit that
DeMi is illuminated by the sun (light blue).
Figure 4-20: Star tracker baffle temperature in °C shown as the black dots. The
bottom subplot shows the beta angle (red) in degrees and the percentage of an orbit
that DeMi is illuminated by the sun (light blue).
72
ground noise level, plotted with detector temperature. The saturation point for the
star tracker is at 1023 counts and is shown as the dashed horizontal gold line. For a
majority of the DeMi mission, the median star tracker background level has remained
below 200 counts, and there is no significant trend upwards in the median levels.
Figure 4-21: The median mean star tracker background brightness level plotted with
detector temperature as the color. The horizontal dotted gold line is the saturation
point for the star tracker detectors as 1023 counts.
Finally, Figures 4-22, 4-23, and 4-24 show star tracker estimates of spacecraft ori-
entations and motions throughout the mission. These plots show estimated quater-
nions, RA and declination, and rate estimates respectively.
Sun Sensors
The DeMi satellite is equipped with 3 packages of 4 sun sensor diodes. The sun
sensors help to provide coarse pointing information by measuring the amount of light
hitting each of the diodes. Sun sensor package 1 is on the -X face of the spacecraft,
package 2 is on the -Z face, and package 3 is on the +Z face. See Figure 4-15 for
clarification on spacecraft orientation. Figure 4-25 shows all of the sun sensor diode
illuminations. This is raw data from each diode shown in counts.
Figure 4-26 shows each sun sensor diode individually along with beta angle and
73
Figure 4-22: Star tracker estimated quaternions where Q4 is the scalar term. The red
dashed vertical lines show commanded pointing attempts.
74
Figure 4-23: Star tracker estimated right ascension, declination, and roll. The red
dashed vertical lines show commanded pointing attempts.
Figure 4-24: Star tracker estimated rates for X, Y, and Z in the spacecraft body
frame. The red dashed vertical lines show commanded pointing attempts.
75
Figure 4-25: Raw sun sensor data for all sun sensor diodes.
solar illumination in the bottom subplot. Diodes 1-4 belong to package 1, 5-8 belong
to package 2, and 9-12 belong to package 3. The diodes in each individual package
appear to follow similar patterns to the rest of the diodes in the package.
With the raw sun sensor diode data, the spacecraft determines how many diodes
are actually seeing the sun based on the counts reaching a threshold value. Figure 4-
27 shows the breakdown of the number of diodes per package that are being used at
a given time on the spacecraft. The maximum is 4 and the minimum is 0. Package 1
frequently observes the sun from all 4 diodes, which makes sense because this is the
sun sensor package that faces the same direction as the solar panels. Package 3 is on
the same face as the external aperture for the payload and the star tracker, so it should
ideally never observe the sun. From the received spacecraft data, sun sensor package
3 only saw the sun two times, which were after a commanded pointing attempt. For
this specific pointing attempt, we can confirm that the spacecraft oriented itself such
that the sun was within the payload FOV keepout.
The sun sensors are also used to estimate the sun body vector. Figure 4-28 shows
how accurate the vector is over time. A value of “GOOD” indicates that 3-4 diodes
can see the sun, “COARSE” means 1-2 diodes see the sun, and “BAD” means that 0
diodes see the sun.
76
Figure 4-26: Raw sun sensor diode data from each of the 4 diodes in each of the
three sun sensor packages. Sun sensor diodes 1-4 correspond to package 1, diodes 5-8
correspond to package 2, and diodes 9-12 correspond to package 3. The bottom-most
subplot shows the beta angle (red) in degrees and the percentage of an orbit that
DeMi is illuminated by the sun (light blue).
77
Figure 4-27: Number of diodes that “see” the sun per sun sensor package plotted
with commanded pointing attempts.The red dashed vertical lines show commanded
pointing attempts. The bottom-most subplot shows the beta angle (red) in degrees
and the percentage of an orbit that DeMi is illuminated by the sun (light blue).
78
Figure 4-28: Sun vector status. This is an indication of how many diodes see the sun.
Finally, using data from the sun sensor diodes, the spacecraft can estimate the
sun body vector. This estimation is shown in Figure 4-29.
Inertial Measurement Unit (IMU)
DeMi’s inertial measurement unit (IMU) is another ADCS component used to esti-
mate rates of motion in X, Y, and Z. Figure 4-30 shows these rate estimates along
with commanded pointing attempts. These rates are similar to those estimated by
the star tracker.
GPS
The GPS is used to confirm spacecraft position and to sync the onboard clocks.
Figure 4-31 shows the times when the GPS information was valid, meaning GPS lock
had been received within the last 90 seconds. In the early period of the mission, before
the bus reconfiguration script in Appendix A was created, the operations team was
not consistently enabling GPS, so the GPS information was not valid for a majority
of the early operations. After November, however, the operations team made sure to
re-enable the GPS after a reset of the spacecraft.
79
Figure 4-29: Sun body vector estimated from sun sensor data.
Figure 4-30: IMU rate estimates in radians per second for X, Y, and Z in the spacecraft
body frame.
80
Figure 4-31: GPS validity.
Magnetometers
The ADCS system also consists of a magnetometer, which estimates the magnetic
field in the spacecraft body frame. Figure 4-32 shows the estimated magnetic field
vector based on magnetometer data. The data shown is relatively inconclusive and
uninformative because only data from pass times have been collected and the data
has not been compared to any magnetic field model. Future work with MIT graduate
student Nicholas Belsten will look at how the measured magnetic field compares to
the world magnetic model (WMM) and how the various spacecraft components may
be impacting the measured magnetic field.
Torque Rods and System Momentum
Throughout the duration of the mission, the 3 torque rods have been enabled and op-
erating in automatic mode. The torque rods are used systematically to dump system
momentum when needed as specified by the program requirements. The torque rods
have not been actively commanded and should not be nominally commanded by the
operators throughout the mission.
Figure 4-33 shows the total system momentum. The total spacecraft momentum
is a very important measurement because over a threshold value, the ADCS can no
81
Figure 4-32: Estimated magnetic field vector in the spacecraft body frame based on
magnetometer data.
longer perform attitude control of the spacecraft. The threshold for DeMi is set by
the total momentum that a single reaction wheel can generate.
4.5.2 Pointing Attempts and Results
Payload Field of View Keepout
An issue that the DeMi team has been dealing with since initial contact is that the
signal sometimes fades in and out during a pass. We think that this is due to the
spacecraft roll, which is intended to keep the satellite thermally balanced and equally
heated from the sun. While the rotation is necessary for the satellite, sometimes it
causes the antennas to roll away from the ground station during passes. This causes
the signal to weaken or even disappear for a bit of time. Because our passes are only
typically on the order of only 10 minutes, the time lost due to roll can be significant.
To try and combat these effects, the team has tried to send pointing commands to
the satellite to actively point the antennas at the ground station for the duration of
the pass and then go back to pointing the solar arrays at the sun.
82
Figure 4-33: Total system momentum in Nms. The system momentum is dumped by
the torque rods when it exceeds threshold value.
For many of these attempted ground station points, the satellite has gone into safe
mode shortly after the command is executed. It took the team a while to understand
why the pointing commands were not resulting in perfect passes, but ultimately it
was discovered that the satellite was orienting such that the sun entered the payload
field of view (FOV) keep out zone, which is 30° from boresight. When the sun enters
the FOV keepout, the FOV keepout event check is triggered, which puts the satellite
into safe mode and points the solar arrays back at the sun. Section 4.7 goes into more
detail on event checks. Figure 4-34 shows the total number of times the FOV keepout
event check was triggered along with the commanded pointing. Event check state of
1 indicates that the event check is enabled. This means that the spacecraft is actively
checking that a condition is or is not met. In this case, the event check is checking
that the sun remains outside of the 30° keepout angle from the payload boresight. An
event check state of 0 indicates that the event check is disbled. When the FOV event
check is triggered, meaning the sun is within the specified keepout zone, the event
check state goes from 1 to 0. As is evident in the figure, every time the event check
was triggered, or the state switched from 1 to 0, the operations team had commanded
spacecraft pointing. This further confirms that the pointing commands sent by the
DeMi team were causing the spacecraft to orient itself in a non-ideal way. For this
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reason, the process for spacecraft pointing will need to be altered so as to not trigger
this event check.
Figure 4-34: Sun in payload FOV keepout event check. The red dashed vertical
lines show commanded pointing attempts. The triggering points are when the status
changes from 1 to 0. A state of 1 means the event check in enabled, a state of 0 means
the event check is disabled.
Currently, the team is working on resolving the commanded pointing issue by
exploring a few ideas. One idea that we are looking into is being more explicit about
the rotation about each axis on the spacecraft. MIT graduate student Patrick McKeen
is actively working on this problem. Another idea, that we may use instead of or in
conjunction with the explicit rotation, is to disable the payload FOV keep out event
check while doing these commands. A keepout angle of 30° for the payload is very
conservative and we do not plan on taking any images during the pointing. These two
methods, and any other ideas that may come up, will be tested on the engineering
model with simulated spacecraft behavior before trying it on the actual spacecraft.
Overall Spacecraft Pointing
Some of the plots throughout this chapter have included red dotted vertical lines
to indicate commanded pointing. This is when the DeMi operators have sent up
commands for DeMi to point itself. So far, the DeMi team has mainly been trying to
point the satellite such that the antennas on board face in the optimal direction for
passes. Figure 4-35 shows the spacecraft ADCS mode over time with an overlay of
the pointing commands. Many of the pointing command periods overlap or appear to
trigger the "FINE_REF_POINT" mode, which is where the spacecraft is performing
84
fine pointing to a target. The "SUN_POINT" mode is where the solar panels are
facing the sun. "SUN_POINT" is the default ADCS mode for the spacecraft and is
entered after a safe mode is triggered.
Figure 4-35: Spacecraft ADCS mode with pointing command periods requested by
the operations team. FINE_REF_POINT is when the spacecraft is actively pointing
at a target, such as the ground station, and SUN_POINT is the default state where
the solar panels are pointed at the sun.
Figures 4-36, 4-37, and 4-38 show the commanded attitude quaternions, the com-
manded rates in degrees per second, and the commanded acceleration respectively
along with the periods of commanded pointing attempts. Once again, the pointing
attempts correspond to many of the changes in the quaternions, rates, and accelera-
tions.
4.6 Bus Resets
Between deployment in July and April 4th, 2021, the DeMi spacecraft has experienced
17 bus resets. A few of these resets have been commanded, many others have been
due to South Atlantic Anomaly (SAA) radiation strikes, and others have been caused
by undetermined anomalies. Table 4.1 shows the dates and causes of all of the resets
through April 4, 2021. Figure 4-39 shows the bus resets throughout the mission. The
black dotted lines are the high rate run counts which represent how long the flight
85
Figure 4-36: Commanded attitude quaternions. The red dashed vertical lines show
commanded pointing attempts. Q4 is the scalar term and is always positive.
86
Figure 4-37: Commanded x, y, and z acceleration. The red dashed vertical lines show
commanded pointing attempts.
87
Figure 4-38: Commanded x, y, and z rates. The red dashed vertical lines show
commanded pointing attempts.
88
software has been running. Every time the bus is reset, shown by the pink vertical
lines, the run count starts back at 0.
Date of Reset Reason for Rest
7/31/20 commanded
8/7/20 unknown cause
8/9/20 unknown cause
8/28/20 commanded
9/4/20 unknown cause
9/23/20 unknown cause
9/24/20 unknown cause
10/3/20 unknown cause
10/25/20 SAA
10/31/20 SAA
12/15/20 SAA
12/25/20 SAA
1/17/21 SAA
1/21/21 SAA
3/2/21 SAA
3/9/21 SAA
3/31/21 unknown cause
Table 4.1: Table of the dates of bus resets between July 13, 2020 and April 4, 2021.
May of the resets were due to South Atlantic Anomaly (SAA) radiation strikes, a few
were commanded resets, and the others have undetermined causes.
When the bus resets occur, the bus and payload heaters get set to survival, the
spacecraft points the solar arrays at the sun, the Lithium radio beacons, the event
checks get reset to their default state, and power to the payload gets cut. To au-
tomate and standardize the reconfiguration process of the satellite, meaning to turn
on the heaters, re-enable the appropriate event checks, and enable certain spacecraft
operations, the author wrote a script that takes all of the reconfiguration actions.
This script can be found in Appendix C.
4.7 Event Checks
Some of the DeMi spacecraft fault protection is defined by a list of event checks.
Event checks are conditions that, when enabled, are constantly being checked by the
89
Figure 4-39: Bus resets, shown as the pink vertical lines, and high rate run count of
the flight software, shown as the black dotted lines. The high rate run count is run
by the flight software and is reset to 0 every time the software (or bus) reboots.
spacecraft. If the condition is not met, the spacecraft will take an action to resolve
the issue, reset the spacecraft, or put the spacecraft into safe mode. DeMi has 26
defined event checks that cover temperature limits, voltage limits, battery state of
charge, radio lockup, loss of communication, sun in the payload FOV, and booting
the payload to the appropriate flight computer and SD card. The DeMi-specific event
checks are shown in Table 4.2. Some of the event checks have been added to the same
row as similar event checks because they are checking the same telemetry item, but
respond differently. Those different responses are designated by “OR” in the Response
When Triggered column.
Event Check Response When Triggered
Cadet radio lock-up Cadet radio power cycle
Sun in payload FOV Transitions to safe mode
Payload 1 or Payload 2 communication stops Transitions payload to safe mode OR to shutdown
Check if payload 1 booted to SD card A/B Cycle payload 1 power
Check if payload 2 booted to SD card A/B Cycle payload 2 power
Payload temperatures out of operational bounds Payloads to safe mode OR payloads power down
Payload temperatures out of survival bounds Power down both payloads
Table 4.2: DeMi event checks to ensure fault tolerance during operations.
The only two event checks that have been triggered unexpectedly are the payload
FOV event check, discussed in the pointing attempts section, and the Cadet radio
lock-up event check. The cadet lock up event check is shown in Figure 4-40 along
with bus reset periods. When the bus is reset for any reason, the Cadet lock up event
check is disabled by default. Therefore, all of the points on the disabled (0) line that
90
correspond with the bus reset lines are not actual radio lockups. The period at the
end of July to the beginning of August was a period of time when the DeMi team
did not know that the event check needed to be re-enabled. After the beginning of
August, once the team began regularly using the bus reconfiguration script, only 3
Cadet radio lockup events were present in the available data. One of them occurred
in mid-August, one in mid-September, and one at the beginning of January.
Figure 4-40: Cadet radio lockup event check with bus resets shown as the pink vertical
lines. All of the points on the 0.0 line that come right after the pink bus reset lines are
not actual event check triggers, they are just incidents of the event check reverting
back to default state upon reset. The points in mid August, mid September, and
early January do not align with observed bus resets and are possibly Cadet lockup
events.
4.8 Overall Bus Health Discussion
Overall, the spacecraft is performing very well. All of the temperatures have remained
within bounds, other than the anomalous offset experienced by the ADCS compo-
nents. Additionally, the spacecraft is responding to pointing commands from the
ground, however user error is causing safe mode triggers. These commanded pointing
errors will be addressed by the DeMi team over the coming months. Additionally, all
of the event checks are working as expected and performing the necessary functions
to keep the spacecraft healthy and operational.
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Chapter 5
Results, Discussion, and Future Work
5.1 Results and Discussion
As discussed in Chapter 2, the NASA Wallops Flight Facility (WFF), which serves
as DeMi’s high data rate ground station, was offline for a majority of the mission
duration to date. Luckily, as of the writing of this thesis, the WFF ground station is
back online and the team was able to resume passes. With the project budget, we were
allocated 100 passes from the WFF and we have used close to 50% of those allocated
passes. During the time that WFF was not operational, the DeMi operations team
made several modifications to the MIT campus ground station to supplement passes
and work on receiving the larger files that would have been downlinked through WFF.
These modifications are still in work and will continue to be useful because they allow
the operations team to use the campus ground station for more than just monitoring
spacecraft health.
In Chapter 3, payload on-orbit operations are discussed. MIT graduate student
Rachel Morgan has confirmed that the payload has captured and returned several
successful images that show the deformable mirror (DM), laser, and cameras are
working. However, the successful images were taken early in the mission. After
the WFF ground station went offline, the team was unable to downlink additional
images and properly test the payload components. Since WFF came back online,
the team has tried to capture more images to demonstrate successful operation of
93
all of the payload components, but those efforts have been unsuccessful. Now, the
team is working on resolving issues with the deformable mirror by uplinking scripts
to the payload that will record DM voltages to a file that the team can downlink
through the WFF ground station. Additionally, the team is exploring options for
attempting operations on the backup SD card or on the Payload 2 flight computer.
Both cameras and the laser still appear to be working, however the DM commands
have been unsuccessful since WFF returned to operations. The Raspberry Pi flight
computer temperatures have been largely elevated and will be monitored more closely
in the future.
Overall, the DeMi spacecraft bus, discussed in Chapter 4, remains healthy and
operational. There have been no major anomalies, just a few minor issues with the
radios, the payload FOV keepout during commanded pointing, and some anomalous
temperature measurement data from the ADCS components. Payload temperatures
have remained within operational bounds, and the batteries are charging and dis-
charging as expected. The attitude determination and controls system (ADCS) is
functioning appropriately and the team will likely be attempting more refined point-
ing attempts in the coming months. The spacecraft has experienced 17 bus resets
from deployment on July 13th, 2020 through April 4, 2021. Many of these bus resets
are likely due to radiation strikes from the South Atlantic Anomaly (SAA) region
over South America.
5.2 Future Work
As of April 4, 2021, the DeMi mission has been operational for several months past
the originally planned mission lifetime. The DeMi operations team has a lot of work
ahead of them to try to meet all of the mission objectives and to keep all of the
spacecraft and payload components operating nominally. The team will continue to
monitor the telemetry items presented throughout this thesis and will continue to
resolve problems as they arise.
94
5.2.1 Operations
Presently, the DeMi team’s main focus is resolving the issues with the DM actuation.
In order to successfully meet all of the mission’s science objectives, the team needs
to understand why the DM actuation attempts have been unsuccessful recently. The
operations team is currently pursuing several paths for resolving the issue. The team
is working on uploading scripts to the payload that will record power to the DM as
it is turned on and actuated and will confirm that properly formatted commands are
being sent to and received by the DM. These scripts will save those measurements to
a file that the team can downlink to try to rule out any issues with the power supply
or command formatting. Additionally the team may try to command DM actuation
with the backup SD card, SD B, to see if that makes a difference. Finally, more
research and analysis is being done on the Raspberry Pi behaviors at high tempera-
tures, as discussed in reference to Figure 3-2 and 3-3. The team is looking through
environmental testing data and performing more ground testing with a Raspberry Pi
and the engineering model of the payload.
After the DM issue is resolved or understood completely, the next priority is to get
the commanded pointing working without triggering the payload field of view (FOV)
keep out event check. Testing is currently being done on the engineering model with
a simulated spacecraft to try to recreate some of the event check triggering we were
seeing in orbit and then correct it. It is possible that the pointing commands should
be more explicit and specify rotation of all the axes to actively avoid getting the sun
in the keepout zone. This is work that is currently being done by MIT graduate
student Patrick McKeen. Another option that the team is considering is disabling
the event check for the FOV keepout while we attempt pointing the radios toward
the ground station. This should be done only if prior analysis is done to ensure that
the payload FOV is not stuck facing the sun for the duration of the pass.
Finally, some of the modifications that the team made during the WFF downtime
need to be uploaded to the Payload 2 computer and the backup SD cards. The team
is working on uploading these “fixes” periodically and in parallel with some of the
95
troubleshooting discussed above. Some of the fixes are necessary for carrying out all
of the testing that needs to be done.
The operations team is testing these changes in as flight-like of a configuration
as possible on the engineering model. However, due to configuration differences, the
team has so far been unable to connect or simulate a Payload 2 flight computer.
5.2.2 Telemetry Analysis
As for the telemetry analysis, future work is likely to include more analysis of the
temperature anomaly seen in Figure 4-13 of the Cadet radio temperature data set.
This analysis will help confirm that the anomaly was not representative of true tem-
peratures and give insight into where the issue originated. Additionally, the team may
try to better understand the cause of the anomalous ADCS component temperature
offset from February, 2021 so that the behavior can be prevented in the future. More
analysis of the magnetometer data and the satellite’s impact on measured magnetic
field is planned with the help of MIT graduate student Nicholas Belsten. There is
also more work to be done to understand the state of health of the bus electronics,
including timing, GPS locks, and stored command processing. Finally, more analy-
sis will be done on space weather throughout the mission and how that might have
had an impact on some of the telemetry presented here, including bus resets with
unknown causes.
96
Appendix A
List of Acronyms
ADCS: Attitude Determination and Control System
AO: Adaptive optics
BMC: Boston Micromachines Corporation
CalPoly: California Polytechnic State University
CAD: Computer-aided design
C&DH: Command and Data Handling
CMOS: Complementary metal-oxide semiconductor
ConOps: Concept of operations
DeMi: Deformable Mirror Demonstration Mission
DM: Deformable mirror
EPS: Electric Power System
FOV: Field of View
FSW: Flight Software
IMU: Inertial Measurement Unit
ISS: International Space Station
MEMS: Microelectromechanical systems
MIT: Massachusetts Institute of Technology
OAP: Off-axis parabolic mirror
PSF: Point spread function
RF: Radio Frequency
97
SAA: South Atlantic Anomaly
SHWFS: Shack-Hartmann wavefront sensor
SOH: State of Health
STAR Lab: Space Telecommunications, Astronomy and Radiation Laboratory
TRL: Technology Readiness Level
TVAC: Thermal vacuum
WFF: Wallops Flight Facility
WFS: Wavefront sensor
WMM: World Magnetic Model
98
Appendix B
Beta Angle and Solar Illumination
Script
1 # Script created by Joey Murphy and modified by Jenny Gubner for use in
this thesis
2 import matplotlib.pyplot as plt
3 import matplotlib.dates as mdates
4 from skyfield.api import Topos, load, Time, EarthSatellite
5 from skyfield.nutationlib import iau2000b_radians
6 from skyfield.elementslib import osculating_elements_of
7 from skyfield.constants import ERAD
8 from numpy import count_nonzero, cos, sin, arcsin, degrees, average,
sqrt, square, clip
9 from scipy.signal import savgol_filter
10 import datetime
11 import math
12
13 # Calculate the Solar beta angle for a satellite at time t (in degrees),
and the corresponding
14 # proportion of sunlit time (expressed as a percentage).
15 def solar_beta(sat, sun, earth, t):
16 sun_astro = earth.at(t).observe(sun)
17 sun_ra, sun_dec, sun_distance = sun_astro.radec()
18 elements = osculating_elements_of(sat.at(t))
99
19 inclination = elements.inclination.radians
20 raan = elements.longitude_of_ascending_node.radians
21 pre_beta = cos(sun_dec.radians) * sin(inclination) * sin(raan −
sun_ra.radians) + \
22 sin(sun_dec.radians) * cos(inclination)
23 beta = arcsin(clip(pre_beta, −1, 1))
24 sun_prop = sun_proportion(beta, elements)
25 return degrees(beta), sun_prop
26
27 # https://stackoverflow.com/questions/48235232/valueerror−math−domain−
error−from−math−acos−and−nan−from−numpy−arccos
28 # For a given solar beta angle, return the corresponding theoretical
proportion
29 # of time in sun per orbit
30 def sun_proportion(solar_beta, osc_elements):
31 sm_axis = osc_elements.semi_major_axis.m # in meters, consistent
with ERAD
32 pre_sun = sqrt(1 − square(ERAD / sm_axis)) / cos(solar_beta)
33 pre_sun = clip(pre_sun, −1, 1)
34 return (0.5 + 1 / math.pi * arcsin(pre_sun)) * 100
35
36 ts = load.timescale(builtin=True)
37 TLE = open(path + "/demi_latest.txt").readlines()
38 satellite = EarthSatellite(TLE[1],TLE[2],TLE[0],ts)
39 eph = load(path + "/de421.bsp")
40 sun = eph[’sun’]
41 earth = eph[’earth’]
42 sunlit_by_days = []
43 solar_beta_by_days = []
44 solar_prop_by_days = []
45
46 #Set up Beta Angle Days
47 start_time = ts.utc(2020,7,13)
48 days = []
49 for i in range(1, 366):
50 next_day = ts.tt_jd(start_time.tt + i)
100
51 dayrange = ts.utc(*list(next_day.utc[:−1]),range(0,86400,10))
52 dayrange._nutation_angles = iau2000b_radians(dayrange)
53 illum_array = satellite.at(dayrange).is_sunlit(eph)
54 sunlit_pct = count_nonzero(illum_array) / len(illum_array) * 100.0
55 sunlit_by_days.append(sunlit_pct)
56 beta, sunprop = solar_beta(satellite, sun, earth, dayrange)
57 #if i % 10 == 0:
58 #print(i, beta[0], sunprop[0])
59 solar_beta_by_days.append(average(beta))
60 solar_prop_by_days.append(average(sunprop))
61 days.append(i)
62
63 # https://stackoverflow.com/questions/46633544/smoothing−out−curve−in−
python
64 # https://en.wikipedia.org/wiki/SavitzkyGolay_filter
65 # https://stackoverflow.com/questions/20618804/how−to−smooth−a−curve−in−
the−right−way
66 yhat = savgol_filter(sunlit_by_days, 7, 2) # window size, polynomial
order
67 start = start_time.utc_datetime()
68 end = start + datetime.timedelta(days=365)
69 year_drange = mdates.drange(start, end, datetime.timedelta(hours=24))
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Appendix C
Bus Reset Reconfiguration Script
The script presented here has been modified to anonymize telemetry
items.
1 ######## Script to Re−Configure After Bus Reset #########
2 ######## Written by Jenny Gubner, Sept 15, 2020 #########
3 ######## Logic updated by Jenny on Sept 22, 2020 #########
4
5 ######## Re−enable Event Checks A, B, C, D, E #########
6 # Taken from enable_event_checks.rb script
7 ecs_to_enable = [A, B, C, D, E]
8
9 for i in ecs_to_enable
10 while(tlm("UUT EVENT_CHECK CHECK_ENABLED_PACK_BIT#{i}") == 0)
11 start_cmd_count = tlm("UUT COMMAND_TLM RECEIVED_COUNT").to_i
12 cmd("UUT ENABLE with NUM #{i}")
13 while(tlm("UUT TLM COUNT").to_i < start_cmd_count +2 and tlm("UUT
EVENT_CHECK CHECK_ENABLED_PACK_BIT#{i}") == 0)
14 wait(1)
15 end
16 end
17 p "Done enabling EC #{i}"
18 end
19
20 ####### Turn on Cadet #######
103
21 # Taken from quick_health_check.rb and fsw_pl1_downlink_slow.rb
22 while not (tlm("UUT POWER CADET_8V") == "ON" and tlm("UUT POWER CADET_5V
") == "ON" and tlm("UUT POWER CADET_3V3") == "ON")
23 start_cmd_count = tlm("UUT TLM COUNT").to_i
24 cmd("UUT MACRO with ID A")
25 while(tlm("UUT TLM COUNT").to_i < start_cmd_count +2 and (tlm("UUT
POWER CADET_8V") == "ON" and tlm("UUT POWER CADET_5V") == "ON" and
tlm("UUT POWER CADET_3V3") == "ON") == false)
26 wait(30)
27 end
28 p "Done turning on Cadet"
29 end
30
31 ####### Set Heater Config to Operational #########
32 # Taken from set_heater_setpoints.rb script and quick_health_check.rb
33 while not (tlm("UUT POWER HEATER_CONFIG1") == "OPERATING" and tlm("UUT
POWER HEATER_CONFIG2") == "OPERATING")
34 start_cmd_count = tlm("UUT TLM COUNT").to_i
35 cmd("UUT HEATER_CONFIG with CONTROLLER_NUM A, CONFIG OPERATING")
36 while(tlm("UUT TLM COUNT").to_i < start_cmd_count +2 and (tlm("UUT
POWER HEATER_CONFIG1") == "OPERATING" and tlm("UUT POWER
HEATER_CONFIG2") == "OPERATING") == false)
37 wait(1)
38 end
39 p "Done setting heaters to operational"
40 end
41
42 # Turn on payload if within temperature bounds
43 # Not yet implemented
44
45
46 ####### Re−Enable Stored Commands ########
47 #Taken from quick_health_check.rb and
48 stored_cmd_status = tlm("UUT TLM STORED_ENABLED")
49 while not tlm("UUT TLM STORED_ENABLED") == ’EN’
50 start_cmd_count = tlm("UUT TLM COUNT").to_i
104
51 cmd("UUT STORED_CMD_ENABLE with ON_OFF ON")
52 while(tlm("UUT TLM COUNT").to_i < start_cmd_count +2 and tlm("UUT TLM
STORED_ENABLED") != true)
53 wait(1)
54 end
55 p "Done enabling stored commands"
56 end
57
58 # Re−upload macro B
59 # Not yet implemented
60
61 # Re−Upload macro C
62 # Not yet implemented
63
64 # Disable Beacon if you feel it is safe
65 # Not yet implemented
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