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RockSat 2010 Final Report
RocketSat VI
The objective of RSVI is to measure large aerosol particle density and
charge from 75 to 95 km to characterize the numerical density distribution of
these particles and the charge of the particles in the region of the atmosphere
above Wallops Flight Facility.
University of Colorado at Boulder Students
Chris Koehler
Kendra Kilbride
Brian Sanders
Zoltan Sternovsky
Colorado Space Grant Consortium
University of Colorado at Boulder
5-7-2010
University of Colorado at Boulder
RockSat 2010
5-7-2010
1.0
Mission Statement
The objective of RSVI is to measure large aerosol particle density and
charge from 75 to 95 km to characterize the numerical density distribution of
these particles and the charge of the particles in the region of the atmosphere
above Wallops Flight Facility.
The two main goals of the RocketSat VI project is to measure the amount
of meteoric smoke particles in the atmosphere above Virginia, and to determine if
the particles have a positive or negative charge. This data will be used to compare
to other similar studies to see if the findings are consistent. Additionally, we can
use the information to support or refute scientific models of the global shift of the
particles during certain seasons. The objectives for the team are to design and
implement a supporting payload for the detectors, measure the attitude of the
rocket during flight, and retrieve all data when the payload is returned. This
mission will allow the team to implement a new flash memory for the RocketSat
missions, and will also teach the team about attitude representation.
2.0
Mission Description and Requirements
2.1
Mission Background
Meteors enter the atmosphere and burn up every day, leaving behind several tons
of small smoke particles in the upper mesosphere and lower thermosphere [1]. These
particles add significantly to the layers of metals already found in the atmosphere, which
consists of mostly sodium and iron, but also contain potassium and magnesium. It is
unclear what type of metal is left behind from meteoroids, but scientists suspect that they
are typically positively charged sodium ions. These ions form larger aerosol particles
(diameter > 1 nanometer) when ice collects around the sodium ion. This also causes the
particle to descend slightly, so that the particles are concentrated through a range of the
atmosphere, approximately 75 km to 95 km.
The meteoritic smoke particles have numerous effects on the upper atmosphere.
As previously mentioned, these particles increase the overall metal budget, or amount of
metal, in the mesosphere. The concentrations of metals can affect radar waves in the
mesosphere [4]. The metal budget is also important to atmospheric scientists because it
affect tidal and gravity activities that dictate global circulation [2]. The increasing
amount of large particles in the mesosphere also increases the number of solid surfaces
on which reactions can occur. Scientists are still investigating the major effects of these
particles on the atmosphere.
These sodium ions in the mesosphere are believed to be the positively charged
core for ice particles that form in the altitude range from 75 to 95 km in the
atmosphere. These ice particles are the result of sub freezing temperatures and an
increase in water vapor. As the sodium ions descend in the atmosphere the ice
particles continue to gain new layers of ice and the particle grows substantially by
the time it reaches 75 km above the surface. The ice particles can reach up to 50
nm in size. The formation of these particles is linked to the increase in global
warming effects on the planet. The increased amount of methane causes an
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RockSat 2010
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increase in the water vapor in the atmosphere, and the increased carbon dioxide
causes a reduction of upper atmospheric temperatures. This combination results
in a much higher formation rate of these ice particles. However, the ice particles
do not form over all locations on Earth. The particles have been measured to be
more present in the more polar latitudes of the Earth. Beyond the formation and
location, little is known about these particles. In situ measurements of the
particles have not been made on more than a handful of occasions.
2.2
Mission Requirements
The requirements for the project were to detect the number large aerosol
particles and measure the charge of the particles. Initial design requirements for
the system were imposed by the RockSat Program and Wallops Flight Facility.
The project level and system level requirements are shown in the following tables.
The Subsystem requirements tables are shown in the Appendices.
Project Level Requirements
Requirement
Description
The payload shall measure meteoritic
smoke particle density
P1
P2
The payload shall measure meteoritic
smoke particle charge
Mission Statement
The payload shall be designed to conform
to the RockSat 2010 User's Guide as set
forth by Wallops Flight Facility
Mission Statement
P3
System Level Requirements
Requirement
Description
The payload shall be able to detect
meteoritic smoke particles
S1
The payload shall be able to measure
charge of meteoritic particles
S2
S3
S4
S5
Parent Req.
Mission Statement
The payload shall be able to determine the
angle of attack of the detector
Parent Req
P1
P2
P1
The payload shall convert detector output
to a digital signal
S6
The payload shall store the digital signal
The payload shall perform in situ
measurements from 75-95 km
P1, P2
S6.1
The system shall record measurements for
the entire duration of the flight
S6
S7
The payload shall not alter the particles
during measurement
P1, P2
S8
The system shall fly on a sounding rocket
with an apogee of at least 95 km
P3
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The payload shall survive the flight
S8.1
S8.2
S8
The payload shall be able to perform with
a temperature range of 60-100 degrees F
P1, P2
The system shall characterize the flight
environment
S9
P1, P2
S10
The system shall make the measured data
available after collection
P1, P2
S11
The system shall be electrically isolated
from the rocket
P3
S12
The system shall be contained in a canister
with a diameter of 9.3 inches and a height
of 9.5 inches.
P3
S13
The payload shall not weigh more than 20
lbs
P3
S14
The system shall have a center of gravity
that lies within a 1x1x1 inch envelope of
the geometric centroid of the integrated
RockSat payload canister
P3
The structural requirements from the system requirements can be seen
through P3 and S11-S14. The most important parts of these requirements is that
the design must fit inside an aluminum canister provided by RockSat, and that the
payload cannot weigh more than 20 lbs. Because the project consists of mostly
electronics, and the science instruments are not contained within the canister, it
was decided that the RockeSat project would only use half of a canister. The
design, coordinated with canister partners (Virginia Tech), would there for be
required to use half of the weight. Additionally, requirement S14 had to be
coordinated so that the RocketSat assembly had the cg in the middle of the stack,
so that when combined with Virginia Tech, the cg would fall into the envelope
specified by the requirement.
CDH also had several important system requirements S4, S5, S9-11. WFF
required that the payload be electronically isolated from the rocket. A new
requirement was that, instead of the electronics having no voltage or current on
them before launch, WFF needed to know the status of the payload at all times.
Additionally, a command line was made available, which imposed the
requirements that the payload could be triggered by the RBF pin and g-switch,
and also be turned on and off by the Wallops command line. Payload
requirements for the CDH system are that it will store all signals on board and
allow them to be retrieved by connecting to the electronics after the payload has
been retrieved. The CDH needs sensors to monitor the flight environment, and
sensors that will measure the attitude of the rocket.
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The Science system requirements are critical to the mission.
Requirements S1 and S2 summarize the entire purpose of this system.
Additionally, to collect the data for the mission, the measurements must be made
in-situ, which is what leads to the requirement for a sounding rocket flight. The
science instruments will have to be able to detect the particles but not interfere
with the charge. Additionally, the instruments must be able to only detect the
large particles, which means they will have to have a way to deflect the smaller
particles, including ions and electrons.
2.3
Concept of Operations
The payload will be launched on a Terrier-improved Orion Rocket out of
Wallops Flight Facility. The rocket will reach an altitude of approximately 120
km, and then land in the Atlantic Ocean, where it will be retrieved. The payload
will be turned on approximately 3 minutes before launch by the command line.
The payload will begin sampling all sensors on board. After the rocket launches,
all electronics will be latched on, and will continue taking data. No particle
detection is expected during burnout, due to the flow that will be going over the
detectors. After burnout, the most important area of the ascent will be from 75-95
km. The rocket will land in the ocean, and electronics will continue to run until
the payload is retrieved, or until the battery power runs out.
3.0
Payload Design
3.1
Overall System Description
The RocketSat VI system will be utilizing two
graphite patch detectors and the electronics boards designed by
the RocketSat VI Electrical team. The detectors will be
mounted flush with the skin of the rocket for direct exposure
to the atmosphere. The detectors will be connected to the
electronics using RF cables. All electronics components will
be secured in an aluminum canister provided by Wallops
Flight Facility (WFF). The system will be recording data for
the full battery life (no sequence is used to power down). This
is because power down is not required for this mission. Data
is stored on board will be read from the electronics memory
after the payload is retrieved.
A diagram of the payload integrated into the rocket is
shown in the figure below. The detectors are positioned in
ports two canisters below the RocketSat canister. There are
two cables being used to connect to the detector. Each cable
connects to the back of the detector through an SMA
connector. The cables will then run up the longerons and into
the canister. Because each CVA is enclosed in an aluminum
box, a BNC connector will be used to interface with the
electronics.
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RSVI Can
Cable
connecting
can and
detector
Detector
Figure 1. Layout of
canister and
detectors in rocket
5-7-2010
3.2
Detector Design
A modified electrometer called the Colorado Dust Detector will be used to detect
these large aerosol particles. This electrometer is also called as a graphite patch detector,
due to the large section of graphite located in the middle of the white plastic, as seen in
Fig. 2.
Figure 2. Graphite patch detector with aluminum casing
Figure 3. Diagram of Detector
This detector works very similarly to a regular electrometer. When a particle
comes into contact with the graphite, the charge is deposited onto the graphite in the form
of a current. The current signal is then transferred through the back of the graphite with a
wire and is transmitted through a cable to electronics that convert and store the signal.
For this type of detector, a higher current corresponds to a greater number of particles
impacting the detector. This detector also requires direct exposure to the atmosphere,
which means for rocket flights it must be mounted to the skin of the rocket [1].
The detector is designed specifically for larger aerosol particles. To avoid
detection of small ions and particles with less mass, there are four magnets in the white
plastic around the graphite. These magnets help repel charged particles with a radius less
than 1 nm because small particles have less momentum than larger particles due to the
smaller mass. Also, smaller particles such as stray ions are not as insulated as meteoritic
smoke particles, which have a shell of ice formed around the nucleus of charge [3]. The
distance that the charged particles are deflected by is defined by the Larmor radius [1], or
gyroradius, equation:
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RockSat 2010
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𝑟𝑔 =
𝑚∗𝑣𝑝𝑒𝑟𝑝
|𝑞|∗𝐵
(1)
where rg is the radius of the particle, m is the mass of the particle, vperp is the velocity
perpendicular to magnetic field, q is the charge of the particle, and B is the magnetic
field.
This equation defines the radius of the particles that will be deflected, and is
based on the strength of the magnetic field produced by the magnets. The magnetic field
is not 100% effective, though, due to the fact that as particles collide in the air, small ions
can be deflected with enough speed to penetrate the magnetic field and hit the patch.
Also, the formation of the boundary layer around the detector can affect how many
particles impact the detector. This issue is accounted for in the data analysis, discussed in
the Analysis section.
The flight configuration being used for RocketSat VI has been used for a
previous experiment on the MAGIC Sounding Rocket Mission. The detectors were
donated to the team by Professor Zoltan Sternovsky of CU Boulder. The detectors were
effective for this mission, and the results of the MAGIC mission are discussed in
Expected Results.
3.1.1
Boundary Layer Effects on Particle Detection
Due to the supersonic speed of the rocket, there was concern about the data
collected being affected by the boundary layer formed on the detectors. This layer would
reduce the efficiency of the detectors, therefore providing inaccurate results. A Direct
Simulation Monte Carlo (DSMC) code was used to simulate the airflow around the
detectors with inputs specifying the altitude and flight conditions. The simulation
showed that particles with a diameter greater than 5 nm were not affected by the
boundary layers formed on the detectors [1]. To account for the particles with a diameter
greater than 1 nm (particles of interest) but less than 5 nm, a relatively small voltage bias
of 2V is added into the data collection from the detectors. This voltage bias is important
to account for aerosol particles affected by the boundary layer, as well as small ions that
may not be deflected by the detectors due to particle collisions.
3.1.2
Attitude Determination
The attitude of the rocket plays an important role in determining the numerical
density of the particles at each altitude. While the boosters on the rocket are firing, there
is no particle detection. This is because all of the airflow is parallel to the detectors. The
rocket is being forced in one direction from the engines, which doesn’t allow any flow
perpendicular to the detectors, allowing for particle detection. After burnout, the rocket
is on a parabolic trajectory. At this point, it is no longer being forced in the direction of
the nose cone. The rocket has velocity in the x and y directions, and it is the velocity in
the x direction that is critical for the detectors. The longitudinal velocity, no longer
forced in any direction, allows air to flow perpendicular to the detector. As the rocket
spins, each detector will be exposed to this airflow. However, since the airflow on the
detector is constantly changing, the attitude of the rocket needs to be measured at all
times during data collection to determine how much airflow is hitting the detector, or the
effective area of the detector [3].
The system devised for measuring attitude is using two MEMS analog
gyroscopes to measure pitch, yaw, and roll. These components output the angular rate of
the rocket, meaning that they are relative sensors as opposed to absolute. Accelerometers,
measuring all three axes, will also be used to measure the attitude of the rocket. The
gyroscope and accelerometer output will be integrated to determine the absolute attitude
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and velocity of the payload. An attitude algorithm will be used in analysis that will use
quaternion models to determine the attitude of the rocket.
3.1.3 Error Analysis
Some of the detectors being used during flight will output shifted data when used
over a long period of time. In particular, the MEMS gyroscopes are known for having
very high drift rates, especially on board sounding rockets. Some time will be spent to
ensure that the data obtained during flight is not skewed by error from the detectors.
The amount of error outputted from the gyroscopes will be measured by a spin
test. This test will spin the payload, with the gyroscopes integrated, along the z-axis for a
length similar to the length of flight. During the whole test the gyroscopes will be
experiencing the same angular rate. After the testing, the data stored from the gyroscopes
will be analyzed to see how much they drifted from the known angular rate of the spin
table. This error analysis will later be implemented into the attitude determination code.
The error will be utilized as a filter, to eliminate some of the error from the flight data.
3.2 Electrical System
The electronics system must be capable of handling multiple data sets during
flight for data analysis and control the payload. The system must be able to determine
the attitude and velocity of the rocket during at least the ascent of the flight. The
electronics will use gyroscopes to measure the pitch, yaw, and roll, and accelerometers to
determine x, y, and z axis accelerations. The objective is to achieve six degrees of
freedom attitude determination. It will also retrieve and store signals from the detectors
to a nonvolatile memory for future analysis. As mentioned, a similar flight configuration
has been used before, and Current to Voltage Amplifiers (CVAs, also known as
Transimpedance amplifiers) will be used to convert the current output from the detector
to a voltage signal. All signals must be stored into a flash memory so that they can be
retrieved upon recovery of the payload. Additionally, the detectors must be sampled at a
rate of 1 kHz to achieve the precision required for attitude determination. Additionally,
pressure and temperature will be measured to monitor the flight environment. The
system is divided into a total of five electronics boards: the AVR board, the SCIENCE
board, the Mesospheric Particle Observation (MEPO) board, and two CVAs. The AVR,
MEPO, and SCIENCE board each have a separate battery, for system redundancy. If a
battery connection fails for one board, it can still draw power from the other boards. A
single 9V battery would be sufficient to power all three boards for the entire 15 minute
flight. The AVR board will be recording pressure, temperature, and accelerometer data.
The AVR is a general board that has been used on previous RocketSat flights and
is used in the RockOn Workshop. It has flight heritage, but modifications were required
for this year’s mission. The accelerometers were replaced with versions that could detect
a wider range of accelerations. This is due to the fact that the previous RocketSat flight
experienced a shock that exceeded the range of the accelerometers and potentially
knocked out the power on the board. It was difficult to determine the magnitude of the
shock and therefore prove it was the cause of the power interrupt. The z-accelerometer is
mounted separately from the AVR board and is connected using discrete wires. The
board also had to be modified to be compatible with a command line that would power
the payload on before launch. The command line is designed to meet the Wallops
requirement. Wallops must be able to turn the payload on and off using a single
command line for payloads wanting early activation in order to provide a safe
environment for possible launch delays and removing the rocket from the launch pad.
Typically the payload is triggered on by both the removal of the Remove Before Flight
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(RBF) pin and the g-switch (closed upon launch), however the gyroscopes being used for
attitude data need to collect initial conditions of the rocket. The g-switch system has a
delay to where the payload does not begin collecting data for more than 2 seconds after
launch, by which time the rocket has already left the rail and could have changed attitude.
The command line required an additional design and section on the AVR board. An
initial revision of the board is shown in Fig. 4.
Figure 4: Initial revision of the AVR board. The microcontroller is in the upper left corner
of the board, and DB9 connectors for power and board to board connections are the two
connectors on the lower edge of the board.
The board uses 3.3V and 5V regulators to control power from a 9V battery,
which is then used to power all of the components. An Atmega 32 microcontroller is
used as the host to control the components on each board and most importantly sampling
and handling data. The microcontroller is implemented using C basic as the core architect
and Python to present a user friendly GUI to retrieve data, analyze latch status, and
change timer settings. The software is developed in AVR studio provided by Atmel.
The CVAs were provided by our faculty mentor, Zoltan Sternovsky. The CVA
will take the current signal from the detector and amplify it, because it is so small, in the
region of nano-amps. The CVA will then convert the amplified signal to a voltage signal,
which is the output of the board. Each board has an output centered on 2.5V, and ranges
from 0V to 5V. As previously discussed, one CVA has a voltage bias of 2V for analysis
purposes. The schematics of each type of CVA were provided by Scott Knappmiller, and
each circuit schematic (bias and no bias) will have to be analyzed to back out the original
current signal for data analysis. The CVAs and their corresponding schematics are
shown in Fig. 5. These boards are extremely sensitive and external signals can
completely overshadow the actual signal. This means that the CVAs must be enclosed in
some type of Faraday cage. Aluminum boxes have been used on previous flights, so
smaller versions will be used to prevent any external signals from interfering with the
detector signal inputted into the CVAs.
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Figure 5: The CVAs being used for flight and their corresponding schematics, provided by
Z. Sternovsky and S. Knappmiller
The MEPO board will be receiving the analog voltage signals from the CVAs,
converts the analog signal to digital and then stores it onto a 64Mbit flash memory. This
board is also controlled with a microprocessor. The board uses an external analog to
digital converter to match the resolution of the current to voltage amplifier, and increase
the sample rate to 1kHz to meet the sampling requirements for the detectors. The
external analog to digital converter communicates to the host microcontroller through a
serial to peripheral interface (SPI). The PCB of the MEPO is shown in Fig. 6.
Figure 6. MEPO board with hardware latch installed and tested
The SCIENCE board will be flying the gyroscopes. Each gyroscope has been
mounted on a smaller breakout board, and each gyroscope board is mounted on two 1x5
headers to interface with the SCIENCE board. The PCB for the SCIENCE board is
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shown in Fig. 7. This is the newest revision of the board and has not been populated at
this time.
Figure 7: MEPO board before population
Similar to the AVR and MEPO, the SCIENCE board also uses the same
microcontroller and flash memory. The gyroscopes will be sampled at a rate of 300 Hz,
which is enough to account for the minimum sample rate to avoid aliasing. . The
LPY5150 and LPR5150 gyroscopes are both dual axis and MEMs technology. The
maximum spin rate for the gyroscopes before saturation is 6000 degrees per second. The
maximum expected spin rate experience during the rocket flight is 6Hz from previous
flight data. 6Hz correspond to a spin rate of 2160 degrees per second which is well within
the range and capability of the LPY5150 and LPR5150. This board also has a hardware
latch to activate early, due to the fact that the initial conditions of the rocket are critical
for analysis purposes.
3.3 Structural System
The structure design for RocketSat VI is dictated by the constraints imposed by
the flight provider, Wallops Flight Facility. The structure will be enclosed in an
aluminum canister, secured to an internal structure within the rocket. Multiple
universities utilize the canister system, and this year RocketSat VI will be sharing the
canister with another university. The result is that the payload must weigh 6.55 pounds
and have a height no greater than 6 inches, half the height of the canister. The design
consists of two Makrolon plates, which is the same material that has been used on
previous flights. The structure houses five electronics boards and four batteries. The
Makrolon plates have a diameter of 9”, which leaves space between the edge of the plates
and the inner diameter of the canister, which is 9.5”. The stack will be connected to the
lid of the canister and extend downwards to a height of 5”, where the stack will be
connected to the other university’s stack. A model of the layout of the stack is shown in
Fig. 8.
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Figure 8. A SolidWorks drawing of our half of the canister payload and the stack outside of
the canister
Each Makrolon plate is separated by 2” stainless steel standoffs to allow for
enough clearance for the electronics boards. The plates are complete circles without
additional holes for wireways because the plates have extra clearance from the edge of
the plates to the canister.
The top plate on the stack will hold four lithium 9V batteries, the SCIENCE
board, and the z accelerometer. The batteries will be secured individually by brass
brackets, and secured with hot glue before flight. Three of the batteries will be connected
to boards; the fourth is to balance out the others to keep the center of gravity in the
middle of the stack. The SCIENCE board is mounted so that it is centered on the plate.
This is so that the gyroscopes can provide data from the center of the rocket, otherwise
conversions would be needed to translate the location, reducing the accuracy of the
results. The plate layout is shown in Fig. 9. This board is mounted above the zaccelerometer, which is connected to the AVR board on the plate below.
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Figure 9. Top plate in stack with batteries, z-accelerometer, and SCIENCE board
The lower plate contains the AVR, MEPO, and CVA boards. The MEPO board
is mounted in the center of the plate, elevated slightly off the surface with nylon spacers.
The AVR board is mounted directly above the MEPO board, again centered on the plate.
The CVAs are mounted on opposite sides of the plate, and are connected to the MEPO
board with wire connections secured by staking with hot glue, shown in Fig. 10.
Figure 10. Second plate in stack with the AVR board mounted above the MEPO board, and
two CVAs on either side of the MEPO
The detectors will be mounted flush with the skin of the rocket. This required a
special interface so that the detectors would be compatible with the Wallops rocket ports,
and also one that would ensure the rocket was still water tight, as the payload section will
land in the ocean. The mounting flange for the detector was designed by WFF, and the
RocketSat Structures team was in charge of machining it to the specified tolerance from
Wallops, 1/1000”. Both flanges have been completed, and a fit check was performed to
ensure the detector fit in the flange. The result is shown in Fig. 11.
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Figure 11. Fit check for the detector and the flange machined in house
Each detector will be connected to a separate CVA using an RF coaxial cable.
The cable will be screwed into the back of the detector with an SMA connection. The
cable will run up from the port with the detector to the payload canister, and is soldered
to the CVA. The cable will be secured to the stack on two standoffs to ensure that if the
cable is pulled it will not lose connection with the board.
4. Expected Results
As previously discussed, a similar flight configuration has been used on a
sounding rocket. This mission collected data in a polar region in Sweden. Fig. 12 shows
the results of the MAGIC sounding rocket experiment combined with several other
experiments collecting the same data. The dotted lines indicate a negative current.
RocketSat VI hopes to
collect data with a trend similar to
that measured by the MAGIC
experiment from the flight in June.
We are also expecting that the
numerical density of the particles
will be much lower than those
measured at higher latitudes, based
on the models developed for the
particle distribution globally.
These models predict the particles
are concentrated near the poles, but
this has not been fully verified.
The payload is scheduled to launch
June 24, 2010. The teams are
currently working on subsystems
testing to prepare for systems
testing and integration.
Figure 12. Results of multiple sounding rocket
flights measuring the numerical density of
meteoritic smoke particles [1]
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4.0
Student Involvement
PROJECT MANAGER
Emily Logan – Project Manager, Aerospace Engineering sophomore, working on the
team Fall 2009 – Summer 2010
The project manager is in charge of the entire RocketSat team. They help with
finding a mission for the project, and help the team through all design and building
processes. It is important for the project manager to have a baseline understanding of all
subsystems in the project, because they are also acting as a Systems Engineer for the
project. The manager also looks for funding and manages the budget for the project.
They set the schedule, and are the primary contact for industry and Student Coordinators.
MISSION SPECIALISTS
Kirstyn Johnson - Mission Specialist lead, Aerospace Engineering freshman, Fall 2009 –
Summer 2010
Andrew Fruge - Mission Specialist, Aerospace Engineering freshman, Spring 2010
Alex Harvey - Mission Specialist, Aerospace Engineering sophomore, Spring 2010
Marcell Smalley - Mission Specialist, Aerospace Engineering freshman, Fall 2009 –
Summer 2010
Ryder Whitmore - Mission Specialist, Aerospace Engineering freshman, Spring 2010
Kyle Wolma - Mission Specialist, Aerospace Engineering sophomore, Fall 2009 – Spring
2010
Marcus Flores – Mission Specialist, Aerospace Engineering freshman, Fall 2009
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The mission specialists focus on the science portion of the payload. They are in
charge of learning and understanding the concepts and theories behind our mission and
using this information to design the data collection system. The mission specialists
consider all aspects of the flight that could affect how the data is collected, as well as the
instruments themselves and how they will affect the data. A significant amount of testing
has to be performed on the science system to characterize how the system will affect the
data, and the mission specialists design and perform these tests. Finally, the mission
specialists analyze the flight data and account for all inaccuracies in the measurements
ELECTRICAL SYSTEMS
Frankie Ning- Electrical Systems lead, Electrical Engineering junior, Fall 2009 –
Summer 2010
Jared Yenzer- Electrical Systems lead, Electrical Engineering freshman, Fall 2009 –
Summer 2010
Jordan Dickard- Electrical Systems, Aerospace Engineering freshman, Spring 2010
Michael Murry- Electrical Systems, Aerospace Engineering sophomore, Fall 2009
David Reid- Electrical Systems, Aerospace Engineering freshman, Spring 2010
The electrical team designs the circuits for the payload that will store all
measured signals during flight. They have to account for power for each circuit board,
and ensure that sample rates and data storage are sufficient to provide useful data to the
team. The team is working with three separate boards that will collect data on
temperature, pressure, attitude, and induced current on the science instruments. The main
AVR is an old design, but the science and data acquisition board are being designed from
scratch. The team will build, test and debug these boards throughout the semester. They
are also working on selecting the right cable to transmit the signal from the detectors to
the circuit boards.
STRUCTURES
Chris Young- Structures lead, Aerospace Engineering freshman, Fall 2009 – Spring 2010
James Bader- Structures, Aerospace Engineering freshman, Spring 2010
Nathaniel Frank- Structures, Aerospace Engineering freshman, Spring 2010
Chris LaPanse- Structures, Aerospace Engineering junior, Fall 2009 – Spring 2010
Michael Rice- Structures, Mechanical Engineering junior, Fall 2009
The structures team is in charge of designing the structure to meet the
requirements of the flight provider and to support the electrical and science subsystems.
They design the structure in SolidWorks, and must select all materials to ensure the
structure does not fail. This year, the structures team had to work out a way to fit all of
the electronics required for the mission in half of the flight canister, as the team is sharing
the canister with another university. Also, the team must work with the flight provider to
be able to mount the science instruments so that they are exposed to the atmosphere. The
structures team will machine the flight layout, and test the structure for shock and
vibrations.
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RockSat 2010
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5.0
Testing Results
5.1
Electronics Testing
5.1.1 AVR Board
The AVR board has undergone two revisions, and the second revision has
passed POST testing, and the team has successfully stored data in the flash
memory and read it back. This proves that the design for the AVR board is
working, and the final revision for the PCBs will incorporate all of the
modifications instated with the second revision. This includes the addition of
several resistors on the board. The hardware latch was also modified with a
resistor. These design changes were implemented for the final revision of the
PCBs.
5.1.2 MEPO and SCIENCE Testing
The MEPO and SCIENCE boards have also been tested with the new flash
memory, and also using external ADCs. New code had to be implemented to use
University of Colorado at Boulder
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both the AVR converter and the external converter. The second revision testing
of these boards showed that there was bus contention between the Flash and the
ADC. A transmission gate was built on a breadboard and integrated into the PCB.
This gate made it possible to switch between reading from the ADC and the
Flash. This gate will be implemented on the PCB with an IC gate. Once the
boards were tested with the transmission gate, the ADC sampled and data was
read from the flash successfully. A realterm printout of these results is shown in
the following figure.
5.1.3 CVA Testing
The Current to Voltage Amplifiers (CVAs) were tested to determine
functionality and ensure that the payload had backups. Two of the provided
boards had blown opamps. The other four were characterized in the lab to
determine the bias on the input, and how the output varies with input. The results
from three circuit tests are shown in the following figure. It can be seen that all of
the boards vary with input at 1V output per 1nanoamp of input. Additionally,
circuits in the 80s have a 2.5 V bias on the input, and the 60s circuits have a bias
on both the input and the output.
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RockSat 2010
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Circuit 87. 2.5 V bias on the input
Circuit 84. 2.5 V bias on the input
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Circuit 62. 2.5 V bias on the output and input
The CVAs require ±15V rails to power the opamp and buffer. This power
will be provided by the MEPO using a DC/DC converter. The CVAs were tested
for functionality connected to the MEPO rails, and the board was powered and
functioned as expected, similar to the characterization found using a power
supply.
5.2
Detector Functionality Testing
The detectors were initially tested for functionality by attempting to
produce electrons in a vacuum to determine if the signal was discernable through
any ambient noise. The test setup was putting the detector in a bell jar, pumping
it down to 1 millibar, and running a high amount of current through a tungsten
wire, which would emit electrons. The detector would be tested with and without
magnets to check that the magnets are effective. The test, although inconclusive
overall due to blown CVAs and that the bell jar was not pumped down enough,
provided several insights to the system setup. The use of shielded cables is
critical for prevent signal interference. Without all connections between the CVA
and detector shielded, the output would be a 60 Hz wall signal that was sensitive
to air movement and vibrations. The test also revealed that the CVAs were not
working.
This test was completed to demonstrate whether the graphite patch
detectors were able to pick up a charge from a source, via the graphite, and
convert this charge to a current. The ability of the detectors to pick up a charge
from what it interfaces with, and then convert this to a current, is crucial to the
mission. During flight, if the detectors are not able to pick up a charge from
atmospheric particles, then the mission will be unsuccessful. Therefore, it is vital
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to know whether the detectors are working properly. This test was also useful in
determining whether the detectors and the CVAs (current to voltage amplifiers)
worked in unison to take the current produced from the graphite patch detectors
and convert it to a voltage.
The detector was initially tested by connecting one of the detectors to an
oscilloscope with a co-axial cable. The detector was allowed to pick up a signal
from just the ambient noise of the room. As seen from Figure 1 below, the Vpp
(voltage peak-to-peak) was 1.00 V. Also, the frequency of this signal was at about
60 Hz, which is believed to correspond to the frequency of the power coming
from the wall outlets.
Figure 13. Ambient Noise Test
There were a few other tests that were done to compare their oscilloscope
output to the “ambient noise” test. A soldering iron was used to test how the
electronics from other appliances affected the charge that the detector picked up.
The Vpp from the soldering iron test (Figure 2) was higher than from just the
ambient noise test, with a value of 14.80 V.
Figure 14. “Soldering Iron” Test
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The detector was also placed near a set of wall outlets to see what signal it
would pick up at this location. The Vpp from this location (Fig. 14) was 14.40 V.
Figure 15. Wall Outlet Signal
The two signals from the soldering iron and the wall outlets were very
similar in terms of their square waveform and their Vpp. This can be justified since
the soldering iron used the wall outlets as its power source.
The changes in signal that can be seen from each of these three tests
supports the conclusion that the graphite patch detectors are picking up a
charge/signal from around the room. As soon as the detector was unhooked from
the CVAs, the Vpp that was read was 20 mV. This change in Vpp between the
connection of the CVAs and the detector, and then subsequently the loss of the
connection, leads to the conclusion that the detector was the one picking up the
signal. Also, it was apparent that the coaxial cable being used for this connection
was not contributing to picking up a signal. Any disturbance in the cable did not
affect the waves seen on the oscilloscope. Also, if the cable was picking up noise
in addition to the graphite patch detector, then it would have still transmitted a
substantial signal to the oscilloscope when the detector was disconnected from it.
Overall, this test can be used as evidence that the graphite patch detectors are
ready for launch, and are expected to output a signal during the entirety of flight.
5.3
Structures Testing
Structures will be performing a vibe test with the integrated payload. The
purpose of this test is to check the ballast design, and ensure that the plates will
not crack. Additionally, the test will check power connections between boards
and from the batteries to the boards, to ensure that everything is secure and will
not be interfered with during flight.
A shock test will also be used to simulate any shocks the payload may
experience during flight, particularly when the second engine ignites. The entire
payload will be dropped from a small height so that the stack is decelerated. The
accelerometers will record the magnitude of this deceleration, and the test will be
repeated until a deceleration of 40g is measured. This will test the power
connections, and potentially the software latch written for this year’s flight.
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6.0
Mission Results (4-??? pages)
This is your second major section. This section should detail your results with
pictures, tables, graphs etc. This section should leave the reader with a sense of
how well your payload performed and why it performed the way it did (or didn’t).
This is the most important section of your report and should communicate how
well your payload performed. What was learned? Did the payload function as
designed. If the payload did not function as designed, failure analysis should be
completed for lessons learned. Will the data you retrieved be used for a larger
purpose? (i.e publication)
7.0
Conclusions (0.5 to 1 page)
This section should make conclusions from Section 6.0. This section may be
difficult for those that did not get any data. If that is the case, use this section to
discuss your payload failure and how your team determined the cause of failure (a
summary of your full failure analysis in Section 6.0.)
8.0
Potential Follow-on Work (0.5 to 1 page)
This section should discuss briefly how this mission could continue and if it is
worth continuing. A brief list of improvements or ways to learn and discover
further information should be listed.
9.0
Benefits to the Scientific Community (0.5 to 1 page)
This section is pretty obvious. Discuss how and why this mission could be
applied to current research in its field. Don’t be afraid to go out on a limb with
your statements.
10.0
Lessons Learned (1 to 2 page(s))
Use this section to discuss what you learned from this experience and what you
would not let happen if this payload were flown again. This discussion should
stay focused on your mission and not so much on the program logistics. This
section is extremely important because future participants will probably look at
your mistakes and try to learn from them. Lessons learned are an essential part of
the final documentation of a project to ensure the advancement of those involved
as well. JPL puts a large value on Lessons Learned in its programs.
11.0
Appendices
11.1
Subsystem Requirements
Detector Subsystem Requirements
Requirement
Description
The payload shall have a detector that measures
particles every meter from 75-95 km
DS1
Parent Req
S1
The detector shall be able to sample at a rate of 1 kHz
DS1.1
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DS1
5-7-2010
The detector shall be activated at launch
DS1.2
DS1
The detector shall only register particles with a
diameter larger than 1 nm
DS2
S1
The detector shall produce a different output for
positive and negative charge
DS3
S2
The payload shall determine the attitude of the
detector with a precision to at least 1 degree
DS4
DS4.1
DS5
DS5.1
DS6
S3
The attitude determination system shall be activated at
launch
DS4
The attitude shall be detected every 1 meter from 75
to 95 km
S3
The payload shall be able to sample the attitude at a
rate of 1 kHz
DS5
The system shall amplify current reading and convert
to voltage
CDH6
The system shall have current/voltage amplifier
outputs +/- 10 V
DS6.1
DS7
DS6
The attitude determination system shall be internal to
the payload
S3
The attitude determination system shall not interfere
with internal electronics
DS8
S3
The power subsystem shall provide 250 mW of power
to the attitude determination system
DS9
Structures Subsystem Requirements
Requirement
Description
The system shall use building materials that can
withstand a temperature of at least 100degC
S3, DS4
ST1
S8.1
Parent Req
The sytem shall use building materials that can
withstand Gloading of at least 25 G
ST2
ST3
S8.1
The system shall support CDH boards with
nonconductive material
CDH6
The system shall secure the CDH boards to the canister
ST4
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S8.1
5-7-2010
The system shall secure connections between boards
in a way that ensures gforces will not disrupt power
ST5
CDH Subsystem Requirements
Requirement
Description
Accelerometers (xyz axis) shall record +/- 50 G
significant events
CDH1
S8.1, S1-S5
Parent Req
S9
Shall be sensitive from 0 t0 5V with 2.5V offset
CDH1.1
CDH1.2
CDH2
Shall have a sensitivity of 27mV/G
Accelerometers (xyz axis) shall record +/- 1.7G
vibrations
CDH1
CDH1
S9
Shall be sensitive from 0 t0 5V with 2.5V offset
CDH2.1
CDH2.2
CDH3
Shall have a sensitivity of 1V/G
The payload shall record temperature from -40 degC
to 125 degC
CDH2
CDH2
S9
Shall use analog output voltage .1-1.75 V range
CDH3.1
CDH3
The payload shall record pressure from 0-15 psi
CDH4
CDH4.1
S9
Shall output from 0 t0 5V with a sensitivity of .267V/psi
precision
CDH4
The system shall be able to sample at 1 kHz
CDH5
CDH5.1
DS1.1, DS6.1
The ADC shall have a throughput rate much greater
than 1kHz
CDH5
The SPI bus must be able to process 16kbit/s
CHD5.2
CDH5
The system shall be able to store all digital values
received from detector, attitude determination
system, and all sensors in flash memory
CDH6
CDH6.1
S10
The system shall use SPI bus from ADC microcontroller
to flash memory
CDH6
The system shall read bipolar -10V to 10V analog
signals from the current to voltage amplifier
CDH6.2
CDH6.3
CDH6, DS6
The system shall be able to convert the analog signals
to digital signals
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CDH6
5-7-2010
The system shall record all freedoms of attitude of the
payload
CDH7
DS4, DS7
The system shall measure the zenith and polar angles.
CDH7.1
CDH7
CDH7.2
The attitude determination system shall sample faster
than 6Hz
CDH7
CDH8
The smoke particle detector shall output a nano-amp
of current
DS1
CDH8.1
The nano-amp shall be converted and amplified to an
output voltage of +/- 10v
CDH8
The system shall activate prior to launch through the
removal of the RBF pin and command line from WFF,
or through the triggering of the G-switch
CDH9
11.2
S6
Final Electronics Schematics
Please try to keep with this format so it is easier for the reviewers. Please keep same font
and section names. These reports are due no later than July 16, 2010 at 5:00 PM. Reports
will be reviewed by Chris or Shawn. You may be asked for revision changes before they
are posted online in final form. Please contact Chris Koehler or Shawn Carroll with any
questions
or
concerns
at
[email protected]
303-378-4765
or
[email protected] 720-234-4902. We understand that documentation isn’t the
most exciting thing in the world, but completing this report will not only help bring
closure to your project, but will preserve your effort. This completed report is a great
resume builder, and will be of great benefit to future program participants.
University of Colorado at Boulder
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5-7-2010