University of Michigan Post-Launch Assessment Review
Project Wolverine:
Butterfly Valve Drag Variation for Mile High Apogee
2011-2012
University Student Launch Initiative
Department of Aerospace Engineering
University of Michigan
3012 Francois-Xavier Bagnoud Building
1320 Beal Avenue
Ann Arbor, MI 48109-2140
Table of Contents
1. Summary ..................................................................................................................................... 3
1.1 Team Name and Location ..................................................................................................... 3
1.3.1 Vehicle Size ................................................................................................................... 3
1.3.2 Motor Choice ................................................................................................................. 3
1.4 Science Experiment .............................................................................................................. 3
1.5 Altitude Reached ................................................................................................................... 3
2. Vehicle ........................................................................................................................................ 4
2.1 Vehicle Summary.................................................................................................................. 4
2.2 Data Analysis and Results of Vehicle ................................................................................... 4
3. Experiment .................................................................................................................................. 8
3.1 Experiment Summary ........................................................................................................... 8
3.2 Scientific Value ..................................................................................................................... 8
3.3 Data Analysis and Results .................................................................................................... 8
3.4 Visual Data Observed ......................................................................................................... 10
4. Lessons Learned and Experience .............................................................................................. 11
5. Educational Engagement .......................................................................................................... 12
6. Budget ....................................................................................................................................... 13
6.1 Budget Overview ................................................................................................................ 13
6.2 Project Budget Details ........................................................................................................ 14
1. Summary
1.1 Team Name and Location
The University of Michigan Rocket Engineering Association (MREA) from Ann Arbor,
Michigan proposes Project Wolverine.
1.3.1 Vehicle Size
Outer diameter: 5.52”
Inner diameter: 5.36”
Total length: 114.00”
Mass with Motor: 40 lbs
1.3.2 Motor Choice
The motor type for the vehicle is a Cesaroni L-1395 motor (75mm diameter) with 4,895
N-s of total impulse. This determination was made based on Rocksim simulations, hand
calculations, and multiple test launches.
1.4 Science Experiment
The projected experiment for MREA’s Project Wolverine is two butterfly valves
(separated 180 degrees) that vary pressure drag. During the ascent, the butterfly valves
will be actuated based on real time data from GPS and flight computers to vary the
pressure drag on the rocket. The goal of this experiment is to put a high impulse motor
in the rocket that ensures the rocket will reach at least one mile at apogee, yet have the
rocket only reach one mile at apogee due to the increased drag from the butterfly
valves.
1.5 Altitude Reached
5549 ft
2. Vehicle
2.1 Vehicle Summary
The vehicle proposed to answer the requirements of NASA’s USLI is pictured in Figure 1 below.
The composite makeup of the rocket includes wound vulcanized fiber for all body tubes,
fiberglass for the fins, and standard 6061 aluminum structurally supporting the drag induction
mechanism where appropriate. The 36 pound (dry) vehicle is 114 inches long with a 5.5 inch
main body tube. Side-cans with a 2.5 inch diameter run just over half of the body length and
house two butterfly valves. Electronics contained within the avionics bay send pulse width
modulated control signals to a single servo, actuating the counter-rotating butterfly valves
located at the maintenance separation. The angle of these valve alter the drag profile of the
rocket and hence can be used in order to achieve a given target altitude. One of the most
important artifacts of this system that must be considered is the boundary layer progression in the
side cans given differing velocities and butterfly flap angles. This boundary layer progression
can induce phenomena such as choked flow, causing false pressure sensor readings as well as
negating the controller as we will see later in test result analysis. Another important
consideration is that this design is single axisymmetric, necessitating robustly assessed stability
given lateral wind vectors over at least half the vehicle body. The vehicle has been built to fly at
least five full launches and recoveries over a wide variety of environmental conditions to build
up confidence in the design and ability to achieve goals.
Set of four
orthogonal fins
Side-can opening
Avionics bay
Maintenance separation
Apogee separation
Figure 1: Launch Vehicle
2.2 Data Analysis and Results of Vehicle
Successes
The choice to use Blue Tube 2.0 as the material for the airframe as well as the cans was
determined to be an extremely good decision. In multiple launches in various humidity and
temperatures ranging from below freezing to near 80 degrees, the vehicle was solid, yet able to
absorb a fair share of wear. Even in the extreme case of falling under only a drogue parachute all
the way from apogee to the ground (at a rate of nearly 70 ft/sec), the vehicle sustained minimal
structural damage. The tubes did not exhibit any great tendency to warp or crack, and they were
very easy to work with (epoxying together, cutting, sanding).
The vehicle performed consistently and as designed during boost phase in all launches,
exhibiting a high level of stability and the ability to handle different launch angles. This may be
attributed to the addition of the Mass Ballast System that added about 3.5 lbs of weight to an
otherwise completed rocket in the form of 5 concrete rings located inside the nose cone. Adding
this mass brought the pre-launch stability margin of the rocket up to about 2.0, a safe margin that
provided a little cushion for any instabilities generated by the unorthodox “side-can” design of
the vehicle. Without these concrete rings it is doubtful that the rocket would have displayed such
robustness and stability in all flights.
At apogee, the vehicle successfully separated every time, deploying a drogue chute and
beginning its descent. The size of the chamber used to house the drogue parachute and associated
protective packing material (“dog barf”) was repeatedly proved to be adequate – the parachute
deployed each launch and sustained no major scorching or other damage as a result of the
ejection charge explosion.
During descent, the rocket’s recovery system worked flawlessly for 3 out of 4 launches. During
the second test launch, the main parachute failed to deploy, causing the rocket to fall from
apogee to ground under only the drogue chute. Luckily, due to the drogue and some soft mud,
the rocket sustained no major damage. Instead, the most important consequence of this failure
was the creation of a new set of rules for folding and packing the parachute, as well as an
increase in the length of shock cord in the rocket. Both of these fixes ensured the successful
deployment of the main chute at both the third and fourth launches. Quantitatively, actual
descent rates under both drogue and main chutes gathered from the final launch were very close
to predicted:
Predicted (ft/sec)
Actual (ft/sec)
36” Drogue
74
70
Cert-3 XL Main Chute
18
21
Failures
There were two repeated failures that resulted from the design of the avionics bay (AvBay),
concerning specifically the system that equalized the pressures inside and outside the avionics
bay so that the altimeters could accurately measure pressure. The first failure that occurred more
than once was a downward spike in the altitude curve at apogee, as visible in Figure 2. This
anomaly was attributed to the improper sealing of the avionics bay from the aft ejection charge
that separated the vehicle at apogee. Despite the addition of gaskets and another fiberglass plate
inside the avionics bay after the second test launch, this problem persisted through our final
launch in Alabama. Although it was never determined that this failure was actually detrimental
to any of the electronics in the avionics bay or the vehicle in general, the inability to totally seal
the AvBay from the ejection charges produced an incorrect altitude curve that repeatedly marred
the flight data.
Figure 2: Huntsville, AL launch showing erroneous spike in pressure-based altitude curve (blue) at apogee
The second failure related to the AvBay was a repeated disparity in altitude measurements
between the accelerometer in the Raven flight computer and both altimeters (one in the Raven,
the other being the Stratologger). At motor burnout, only some 3-4 seconds into flight, there was
at least a 200-meter difference between the two altitude measurements, with the pressure-based
data from the altimeter indicating the lower altitude of the two – probably the more incorrect
altitude. On a larger scale, the altitude curve derived from the altimeters looks very strange up to
and at motor burnout, as there is a sudden increase in slope at burnout (
Figure 3 &
Figure 4, blue line). This sudden increase in slope would normally indicate a
positive acceleration, however the vehicle should not be positively accelerating at all after the
engine has ceased to burn. On the other hand, the altitude data from the accelerometer looks
much nicer and agrees with calculations for an approximate altitude at burnout (
Figure 3 &
Figure 4, black line).
Figure 3: 3rd test launch data
Figure 4: Huntsville, AL launch data
Due to the above considerations, it was determined after the third launch that these
measurements from the altimeters were faulty and contained significant error during the initial
portion of the ascent. Besides this failure having dramatic and extremely detrimental effects on
the payload portion of the rocket (discussed in 3.2) it also prompted a look and redesign of the
pressure equalization system between the AvBay and the outside of the vehicle. For the first
three launches, the vehicle had four, orthogonal quarter-inch holes drilled in both the airframe
and the AvBay internal body tube to equalize pressure for the altimeters, which were open to the
environment inside the AvBay. It was suggested that perhaps there was airflow either over or
through these holes during boost that was causing errant measurements from the altimeter.
Before the launch in Alabama, wind tunnel testing was performed on the actual AvBay to try and
relocate these holes in order to ensure accurate measurements. In reaction to the wind tunnel
work the holes were made smaller and moved as far away from any obstacles as possible.
Despite these adjustments, the problem persisted and actually worsened during the flight in
Alabama (
Figure 4). There has not been any definitive conclusion about
why this failure occurred, however, there is no doubt that the vehicle did not perform exactly as
intended as a large consequence of the inability to accurately determine altitude based on static
pressure during and immediately after boost – the most dynamic part of flight and most critical
for adjusting drag.
3. Experiment
3.1 Experiment Summary
Our payload will use the principles of PID control theory to govern the aforementioned
mechanism designed to induce pressure drag as a means of regulating vehicle altitude. The
control system will be actuated at a pre-designated trigger velocity to ensure that our flight speed
has passed the highly unstable transonic regime, where shock formation on our control surfaces
could lead to instabilities. The most significant objective of our controller is induction of
pressure drag in the mean energy solution path such that both apogee-amplifying and apogeedepreciating perturbations are recoverable through the majority of flight. The controller should
be robust enough to recover altitude goals over various launch environment conditions expected
during operation in testing and in competition. Drag should be calculated dynamically during
flight, and the controller should respond to physical system changes in no more than 50
milliseconds.
Of utmost importance in payload success is that drag is actuated exclusively opposite to the
vehicle velocity vector, and that no moments are created around any other axis than the
longitudinal body axis of our vehicle. The control system must demonstrate sufficient
disturbance rejection over the range reasonable environmental perturbations encountered during
flight at any stage, and recover within 2% of the goal altitude.
3.2 Scientific Value
Due to the nature of our flight goals, and the high level of a prior knowledge we have regarding
the nature of our flight, we have implemented a PID algorithm with a slight modification. As
opposed to direct gain scheduling, we modify our target altitude dislocating it from our nominal
target altitude to ensure that the mean energy path solution is attained as opposed to the
controller quickly damping to the goal altitude. The extent and means by which this dislocation
occurs is described in detail below, and is based off of several vehicle state variables. This
design was selected due to the fact that the control system has no means to add energy back into
the system, making energy management an extremely important aspect of the controller design.
If the nominal flight path is attained within a few moments of the controller activating, there
would be no means for the controller to reject any disturbance which would reduce the apogee
altitude. The risk inherent with this classical PID design in failing the criterion of being able to
recover within 2% of the goal altitude was determined to be too great to allow an unmodified
PID algorithm to control the ascent of our rocket.
3.3 Data Analysis and Results
The primary purpose of our USLI launch was
to reach a mile altitude, and perform a full
scale vehicle test. As with the previous two
test launches, the vehicle weighed in at 16.3
kg and was propelled via a Cesaroni L1395
motor which provided 4900 N-s of impulse
over 3.6 seconds.
This launch took place on April 22, 2012 in
Huntsville, Alabama. In the early afternoon
our vehicle began its 18 second ascent. Our
pressure port size had been reduced in an
attempt to alleviate some of the issues
experienced during the last two test launches
with pressure equalization lag.
However, as the flight data in Figure 6
indicated, our pressure port sizing issue had
not been alleviated, but rather aggravated.
Our rocket apogee was inconclusive, and lay
between 1580 and 1683 meters. What likely
contributed most significantly to this pressure
deviation was the burnout altitude projected
at apogee. According to our barometric
sensors, which are very heavily
weighted in altitude calculations, we were
only at 91 meters of altitude at motor
burnout. All simulations and accelerometer
based altitude readings indicated that this
Figure 5: Launch Vehicle Shortly After Liftoff on April 22,
2012
Max velocity
at motor
burnout
Latently
regressive
BATES
burn profile
Accelerometer
(black) vs.
Altimeter (blue)
altitude
Large pressure
equalization
lag
Pressure from
ejection charge
indicates altimeter
seal issues
Figure 6: Raven Flight Computer Data from Flight on April 22, 2012
altitude was closer to 350 meters. Due to our system having the largest ability to influence drag
early on in the flight, this effect was extremely detrimental. Our final conclusion is that there is a
fundamental flaw in the setup of pressure entrance into our avionics bay. Larger pressure holes in
various locations may alleviate this issue although that does come at the cost of a larger chance
for dynamic pressure fluctuations to interfere with guidance. In the future it is strongly suggested
that a guidance system based on inertial rather than external measurement is used.
3.4 Visual Data Observed
The major finding during this test and launch process has been that a DART, or any target
modifying control system is extremely sensitive to incorrect flight data. The inherent risk of such
a system is present in twofold, one being that when the system is modifying the target the most
heavily, data must very accurate. The second is that a system such as ours relies on a controller
that becomes less effective as we approach our target. A perhaps more thorough revision of the
controller would include an undershoot command, that rectifies itself via injection of energy into
the flight system to increase apogee.
Speaking more directly to the control system itself, gain scheduling effects have been mimicked
with general success via a modified target. Had the controller had the correct data during the
entire flight, there is a good chance we would have seen altitudes that very nearly achieved our
goals, as we can see by the effect of the controller in our final test flight with just a single point
of what can be considered consistent data. However by comparing the flight results with what
would have been achieved via a simple Zeiger-Nicholas tuned PID controller, we can see the
DART controller did provide some substantial gains. Further work regarding a DART controller,
or a controller that employs continuous gain scheduling, would involve looking at which state
variables can most accurately be used to modify a given arbitrary target, and how those
modifications can be based on vehicle characteristics in a general way. Overall this study has
shown that there is promise that continuous gain scheduling, if built into the dynamics and via a
priori knowledge of a vehicle, can attain better results that a simple PID controller.
4. Lessons Learned and Experience
Before the competition we attempted three separate test flights, but we were unable to collect
data from our control system on any of these test flights. We feel like we could have benefitted
from data collection to see if the control system was working correctly. This showed us how
important test flights are in the testing of systems.
We attempted to control our altitude by using two butterfly valves that vary pressure drag. The
valves would be actuated based on real time data from multiple onboard electronics shortly after
motor burn out. However at competition it was determined that the butterfly valves did not
actuate as expected. After doing some analysis we have determined that it was due to inaccurate
number readings. At apogee the accelerometer we were using to read off of said 5185ft when in
actuality our Raven altimeter read 5523ft and our competition altimeter read 5549ft. We believe
the reason for the valves not performing as expected is due to some fundamental flaw in the
pressure-hole system resulting in inaccurate readings from both altimeters. This has been an
issue that we have been trying to resolve since our third full scale test launch. Our currently
assumption is that having the avionics bay enclosed in another hollow cylinder has allowed
pressure to leak into the avionics bay.
Even though the system did not work as desired the experience helped prepare us as future
engineers. By having these issues we were able to go through the design-fabrication-test cycle
multiple times, testing our problem solving skills. Each time we learned more about how the
onboard systems work in relation to each other and the environment.
5. Educational Engagement
The Michigan Rocket Engineering Association’s first run of a new outreach project was a
deemed a success in a few ways. Our members taught students at Pinckney Community High
School the basics of rocketry in a three week seminar in which students built rockets in teams as
a part of a TARC-like competition. These students in turn showed a true interest in the subject
wanting to start a TARC team and have MREA members mentor the team. The school labeled
the seminar a success and invited MREA back to teach again next year. The goals of the USLI
educational outreach project were to inspire an interest in rocketry and engineering in general as
well as to establish an ongoing project that could continue in years to come. MREA has
completed these goals on a small scale and hopes to continue this project on a larger scale. Plans
for a manuscript that describes in detail how to run this seminar are in the works and MREA
hopes to distribute the information to other schools so that they may run their own seminar
without having our members there in person. MREA hopes to continue this project in the fall
with a new set of students, perfecting this seminar through trial and error.
6. Budget
6.1 Budget Overview
MREA has spent just over $6,200 on the entire USLI project. Actual flight hardware has cost
MREA just over $3,200, which puts MREA well under the competition rule of $5,000 of flight
hardware on the launch pad. Additionally, MREA spent almost $2,400 transporting 12 team
members to Huntsville for the competition, and MREA spent about $600 on the outreach project
with Pinckney High School. Table 1 and Figure 7 detail all of the expenditures MREA has
incurred during the USLI project.
Subsystem
Structural Components
Controls Structural Components
Avionics Components
Propulsion Components
Travel Expenses
Outreach Project
Shipping Costs
TOTAL
Total Cost (USD)
722.58
289.43
600.05
1216.91
2443.43
604.37
387.13
6263.90
Table 1: Project Cost Breakdown by Subsystem and Category
Shipping
Costs
6%
Outreach
Project
10%
Controls
Structural
Components
5%
Structural
Components
11%
Avionics
Components
10%
Travel
Expenses
39%
Propulsion
Components
19%
Figure 7: Visual Representation of Project Costs
In order to fund the project, MREA received numerous grants and contributions from
organizations around the University of Michigan. These organizations included the Aerospace
Engineering Department at the University, student councils, and Raytheon Missile Systems who
provided a $1,000 grant in support of MREA’s USLI efforts. Finally, MREA also had its twelve
travelling team members contribute to defray the travel costs. Overall, MREA had ample
funding to support the USLI.
Income Source
Starting Balance
UM Aerospace Engineering Dept. Grant
Raytheon Company Grant
UM Central Student Government Grant
University of Michigan Engineering Council Grant
Team Travel Contributions
Category
Starting Balance
Grants
Grants
Grants
Grants
Refund
TOTAL
Amount
1956.67
2500.00
1000.00
700.00
1000.00
500.00
7851.67
Table 2: Project Income Sources
6.2 Project Budget Details
Included in the following tables are details on the individual components that were purchased for
the Structural, Controls Structural, Avionics, and Propulsion subsystems. Additionally, outreach
project expenses and travel expenses are listed.
Structural Components
Upper Body Tube
Outer Fuselage/Side Cans
Can Coupler
Upper/Lower Body Tubes
Misc. Lab Supplies
Blue Tube Coupler 2.56"
Blue Tube Coupler 5.5"
Main Parachute
Plywood
Home Depot Supplies
RadioShack Supplies
Supplier
Apogee Rocketry
Apogee Rocketry
Apogee Rocketry
Apogee Rocketry
Ace Hardware
Apogee Rocketry
Apogee Rocketry
LOC Precision
Home Depot
Home Depot
Circuit City
Unit Cost (USD)
56.95
26.95
9.25
56.95
93.05
9.25
18.95
189.00
$34.97
25.51
31.17
Quantity
1
5
4
2
N/A
2
1
1
N/A
N/A
N/A
TOTAL
Table 3: Structural Subsystem Component Costs
Total Cost (USD)
56.95
134.75
37.00
113.90
93.05
18.50
18.95
189.00
34.97
25.51
31.17
722.58
Controls Structural Components
Flat Washer
Lock Washer
Screw
Rods
Hex Locknuts
Hex Nuts
Drag Flap Thrust Plate (Ring)
Avionics Bay
HS-5645MG Digital Torque
Voltage Regulator Resistor
Diode
Gears
Eyebolts
Supplier
McMaster-Carr
McMaster-Carr
McMaster-Carr
McMaster-Carr
McMaster-Carr
McMaster-Carr
McMaster-Carr
Apogee Rocketry
ServoCity
Mouser Electronics
Mouser Electronics
McMaster Carr
McMaster-Carr
Unit Cost (USD)
3.37
4.03
9.80
6.21
5.01
4.49
51.30
38.95
46.99
0.19
0.34
36.48
2.76
Quantity
1
1
1
1
1
1
1
1
1
12
6
3
2
TOTAL
Total Cost (USD)
3.37
4.03
9.80
6.21
5.01
4.49
51.30
38.95
46.99
2.28
2.04
109.44
5.52
289.43
Table 4: Controls Structural Subsystem Component Cost
Avionics Components
Drag Flap Thrust Plate (Ring)
732 Ohm Resistor
240 Ohm Resistor
300 Ohm Resistor
Capacitor
Capacitor
Arduino Female Terminal
Connectors
Accelerometer Voltage Regulator
Arduino Voltage Regulator
Connectors
Accelerometer
LED Indicator
LED Indicator
LED Mount
Fixed Terminal block
Public Missile Fin-D-07
Key Switches
E-Charge Canisters
E-Charge Canisters
Perfect Flight Altimeter
Voltage Regulator Resistor
Diode
Micro-SD Card Breakout Board
Steel Bore Miter Gear
3"x6"x3" Aluminum Block
Servo Hub
Coupler Hub
Screws
Key Switches
Drag Computer
Supplier
McMaster-Carr
Mouser Elec.
Mouser Elec.
Mouser Elec.
Mouser Elec.
Mouser Elec.
Mouser Elec.
Mouser Elec.
Mouser Elec.
Mouser Elec.
Sparkfun Elec.
Mouser Elec.
Mouser Elec.
Mouser Elec.
Mouser Elec.
Public Missiles
McMaster-Carr
Apogee Rocketry
Apogee Rocketry
Perfect Flight
Mouser Electronics
Mouser Electronics
Adrafruit Industries
McMaster-Carr
McMaster-Carr
Servo City
Servo City
Servo City
McMaster-Carr
Mouser Elec.
Unit Cost (USD)
51.30
0.06
0.13
0.13
0.31
0.98
Quantity
1
12
12
12
12
12
Total Cost (USD)
51.30
0.72
1.56
1.56
3.72
11.76
2.06
0.60
1.70
1.05
28.95
1.39
0.59
0.26
0.41
19.15
12.95
10.00
9.99
67.96
0.19
0.34
15.00
18.58
44.12
9.99
7.99
0.35
6.65
29.95
4
6
6
6
1
2
2
4
4
4
2
5
5
1
12
6
1
3
1
1
1
4
4
1
TOTAL
8.24
3.60
10.20
6.30
28.95
2.78
1.18
1.04
1.64
76.60
25.90
50.00
49.95
67.96
2.28
2.04
15.00
55.74
44.12
9.99
7.99
1.40
26.58
29.95
600.05
Table 5: Avionics Subsystem Component Costs
Propulsion Components
Motor Mount Tube
Motor Mount Centering Rings
Motor Retention
Motor Thrust Plate
5 Grain Motor Retention
Pro 75 Case Spacer
Flight Test Motor #1
Flight Test Motor #2
Flight Test Motor #3
USLI Launch Motor
Supplier
Public Missiles
LOC Precision
Apogee Rocketry
McMaster-Carr
Giant Leap Rocketry
Giant Leap Rocketry
Loki
Cesaroni
Cesaroni
Cesaroni
Unit Cost (USD)
16.50
7.00
52.00
12.47
203.45
24.95
160.00
239.54
240.00
240.00
Quantity
1
4
1
1
1
1
1
1
1
1
TOTAL
Total Cost (USD)
16.50
28.00
52.00
12.47
203.45
24.95
160.00
239.54
240.00
240.00
1216.91
Quantity
5
5
5
5
5
5
5
5
5
3
5
3
1
TOTAL
Total Cost (USD)
73.25
31.25
14.25
31.25
104.75
23.75
29.00
14.25
26.25
18.75
89.95
128.91
18.76
604.37
Table 6: Propulsion Subsystem Component Costs
Outreach Project Components
Nose Cone
Payload Body
Tube Coupler
Main Body
Parachute 36"
Motor Mount
Centering ring(2)
Bulkhead assembly
Rail Buttons
Shock Chord
G-10 Fiberglass
F-Class Motors
Rocket Fair Display
Supplier
LOC Precision
LOC Precision
LOC Precision
LOC Precision
LOC Precision
LOC Precision
LOC Precision
LOC Precision
LOC Precision
LOC Precision
Giant Leap Rocketry
Apogee
Home Depot
Unit Cost (USD)
14.65
6.25
2.85
6.25
20.95
4.75
5.80
2.85
5.25
6.25
17.99
42.97
18.76
Table 7: Outreach Project Costs
Travel Expense
Car Rental
Gas
Hotel
Unit Cost (USD)
300.00
0.24
272.22
Quantity
2 vehicles
2000 miles
5 nights
TOTAL
Table 8: Travel Costs
Total Cost (USD)
600.00
482.35
1361.08
2443.43