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MIT Rocket Team
USLI 2011
Critical Design Review
January 23, 2011
Project Valhalla
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REPORT AUTHORS
Christian Valledor Project Manager
Andrew Wimmer Team Safety Officer Tripoli Rocket Association Level 3
Ben Corbin EHS Representative
Ryan McLinco Launch Vehicle Lead
Jonathan Allen Payload Lead
Eric Peters Recovery Lead
Ben Couchman Avionics Lead
Anna Ho Outreach Coordinator
Jake Bograd-Denton
Michelle Burroughs
Jason Elizalde
Jedediah Storey
Leo Tampkins
3
NAMING CONVENTIONS
NASA and the space industry have typically been very creative with their acronyms for
spacecraft and space instruments. From LEM to SRB, NASA has come up with the
most recognized TLAs the -world has ever known, approaching even greater rhetorical
strength with C.O.L.B.E.R.T.. We in the MIT Rocket Team also share such an affinity for
acronyms, so we decided to put some extra thought into the naming of our rocket
project. We also honor NASA‟s tradition of naming conventions based off mythological
gods from ancient civilizations. However, where NASA chooses its naming inspiration,
we wholeheartedly disagree; the Greek and Roman gods are too overused for us to
name our rocket parts after. Instead, we turned to the Norse gods, whose deeds are so
epic and intense that modern comic book heroes are based off of them.
While most people are familiar with the god Odin, few are familiar with the one-handed
god if single combat, victory, and heroic glory, for which the third day of our modern
week is named after. Scandinavians would have called this god in his human form a
yager, which is German for "hunter" or someone who tracks animals. This is appropriate
considering the goal of the UAV payload is to track targets on the ground and
autonomously report their location. For these reasons, we are naming the entirety of our
rocket the Tactical Yager Rocket – TYR
One of the most famous Norse gods is the Norse god of fire. A brother to Odin and
father of a wolf, this god has inspired many names including a company that produces
rocket engines. We at the MIT Rocket Team couldn't pass up the opportunity to use his
name for a piece of our rocket. This is why we have named the structural section of the
rocket that houses the motor and spits fire the Ludicrously Overpowered Kinetic
Impulser – LOKI
Vikings were inspired to be brave in their fights because only the righteous and brave
who die in battle would enter the gates of Valhalla to sit with Odin and fight by his side
in the end of times. One son of Loki, an eight-legged horse that travels across land, sea,
and air, is responsible for delivering fallen warriors to their glorious fate. This god carries
those valiant warriors who die in battle over land and sea across the Rainbow Bridge
through the gates of Vahalla. For this reason, we have named our motor the Single-
Loaded Electrically-Ignited Propulsively Numerous Integrated Rocket – SLEIPNIR
From the basis of these Norse naming inspirations, there was only one logical choice for
building a machine that would reach the heavens, named after the home of Odin himself
– Project Valhalla
However, not even the Norse gods were powerful enough by modern standards when it
came to naming the crown jewel of our rocket: the UAV payload we are ejecting from
the rocket to perform a search mission. The one series of magnificent beings that exists
today from which we can draw truly incredible names from is none other than the
Transformers series. While casual fans of the series may often bicker about who their
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favorite Transformer is, often arguing between Optimus Prime and Megatron, there is
only one Transformer that truly and unequivocably incites emotions of grandeur, terror,
destruction, power, and might, and it first appeared in the 1986 Transformers theatrical
release as the Transformer that was so powerful it literally ate planets. For this reason,
we have given our unmanned aerial vehicle the name UNmanned Integrated Craft for
Rescue with Onboard Navigation – UNICRON
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CONTENTS
Report Authors ................................................................................................................ 2 Naming Conventions ....................................................................................................... 3 1 Summary of CDR ................................................................................................... 10
1.1 Team Summary ................................................................................................ 10 1.2 Launch Vehicle Summary ................................................................................ 10
1.3 Payload Summary ............................................................................................ 10
2 Changes Since PDR .............................................................................................. 11 2.1 Launch Vehicle Changes ................................................................................. 11 2.2 Payload Changes ............................................................................................. 11
2.3 Activity Plan Changes ...................................................................................... 12
3 Design and Verification of Launch Vehicle ............................................................. 12 3.1 Mission Statement, Requirements, and Mission Success Criteria ................... 12
3.2 Major Vehicle Milestone schedule .................................................................... 13
3.3 System Level Review ....................................................................................... 13 3.3.1 Rocket Design and Subsystems ................................................................ 14
3.3.2 Subsystem Requirements and Descriptions .............................................. 14 3.3.3 Demonstrate design meets system level requirements ............................. 39
3.3.4 Workmanship and relation to success ....................................................... 39 3.3.5 Planned Component, functional, and static testing .................................... 40
3.3.6 Status and plan for remaining Manufacture and Assembly........................ 41
3.3.7 Integrity of design ...................................................................................... 42
3.4 Recovery System ............................................................................................. 43
3.4.1 Parachute choice and testing .................................................................... 43
3.5 Safety and Failure analysis .............................................................................. 44
3.6 Mission Performance Predictions ..................................................................... 46
3.6.1 Mission Performance Criteria .................................................................... 46
3.6.2 Flight Profile Simulation ............................................................................. 46
3.6.3 Scale Model Test ....................................................................................... 50 3.6.4 Stability ...................................................................................................... 51
3.7 Payload Integration Plan .................................................................................. 51
3.7.1 Installation and removal, dimensions, precision fit ..................................... 51
3.7.2 Tasking & Integration Schedule ................................................................. 55 3.7.3 Compatibility of elements ........................................................................... 55
3.7.4 Simplicity of integration procedure ............................................................. 56 3.8 Launch Operation Procedures ......................................................................... 56
3.8.1 Recovery Preparation ................................................................................ 64
3.8.2 Motor Preparation ...................................................................................... 67
3.8.3 Igniter Installation ...................................................................................... 67
3.8.4 Setup on Launcher .................................................................................... 68 3.8.5 Troubleshooting ......................................................................................... 68
3.8.6 Post Flight Inspection ................................................................................ 68 3.9 Vehicle Safety .................................................................................................. 69
3.9.1 Identification of Safety Officers .................................................................. 69
6
3.9.2 Analysis of Failure modes and Mitigations................................................. 69
3.9.3 Potential Hazards ...................................................................................... 71
3.9.4 Environmental Concerns ........................................................................... 76
4 Payload Criteria ...................................................................................................... 76 4.1 Testing and Design of Payload Experiment ..................................................... 76
4.1.1 System Level Review ................................................................................ 76 4.1.2 Demonstrate design meets system level requirements ............................. 97
4.1.3 Workmanship and relation to success ....................................................... 98 4.1.4 Planned Component, functional, and static testing .................................... 99 4.1.5 Status and plan for remaining Manufacture and Assembly...................... 100
4.1.6 Describe Integration Plan ........................................................................ 103 4.1.7 Instrumentation precision and repeatability ............................................. 104
4.2 Payload Concept Features and Definition ...................................................... 105
4.2.1 Creativity & Originality of Payload ........................................................... 105
4.3 Science Value ................................................................................................ 106 4.3.1 Payload Science Objectives .................................................................... 106 4.3.2 Payload Success Criteria ......................................................................... 106
4.3.3 Experimental Logic, Approach, Verification ............................................. 106 4.3.4 Describe test and measurement, variables and controls ......................... 107
4.3.5 Relevance of Expected Data ................................................................... 109 4.3.6 Accuracy and Error Analysis .................................................................... 109
4.3.7 Experiment Process Procedures ............................................................. 109 4.4 Payload Safety ............................................................................................... 110
4.4.1 Identification of Safety Officer .................................................................. 110
4.4.2 Failure Modes .......................................................................................... 110
4.4.3 Potential Hazards .................................................................................... 116
4.4.4 Environmental Concerns ......................................................................... 117
5 Activity Plan ......................................................................................................... 118 5.1 Budget Plan.................................................................................................... 118
5.2 Timeline ......................................................................................................... 119 5.3 Educational Engagement ............................................................................... 119
6 Conclusion ........................................................................................................... 122
7
Table of Figures Figure 3-1: Overall Rocket ............................................................................................ 14
Figure 3-2: Tube Coupler Segment ............................................................................... 16
Figure 3-3: Nose Cone .................................................................................................. 16
Figure 3-4: Nose Cone Coupler .................................................................................... 17
Figure 3-5: Motor Centering and Retention ................................................................... 18
Figure 3-6: Recovery System Bulkhead ........................................................................ 19
Figure 3-7: Sabot Overview ........................................................................................... 20
Figure 3-8: Payload Integration Stacking ...................................................................... 20
Figure 3-9: Recovery Configuration .............................................................................. 21
Figure 3-10: Sabot Hard Point ....................................................................................... 22
Figure 3-11: MAWD Flight Computer ............................................................................ 24
Figure 3-12: ARTS2 Flight Computer ............................................................................ 25
Figure 3-13: ARTS2 Telemetry Transmitter .................................................................. 25
Figure 3-14: ARTS2 Telemetry Receiver ...................................................................... 26
Figure 3-15: ARTS GUI ................................................................................................. 27
Figure 3-16: ARTS Data Analyzer ................................................................................. 27
Figure 3-17: Power Switch ............................................................................................ 28
Figure 3-18: Avionics Package ...................................................................................... 29
Figure 3-19: Axial Case BCs ......................................................................................... 33
Figure 3-20: Lateral Case BCs ...................................................................................... 33
Figure 3-21: Axial Case Results .................................................................................... 34
Figure 3-22: Lateral Case REsults ................................................................................ 35
Figure 3-23: Motor Retention BCs ................................................................................. 37
Figure 3-24: Motor Retention Displacement .................................................................. 37
Figure 3-25: Motor Retention Stress ............................................................................. 38
Figure 3-26: Predicted CM and CP Locations ............................................................... 46
Figure 3-27: Predicted Acceleration and Velocity Profiles ............................................. 48
Figure 3-28: Simulated Altitude Profile .......................................................................... 49
Figure 3-29: Scaled Launch Simulated Results ............................................................ 50
Figure 3-30: Scale Model Rocket .................................................................................. 50
Figure 3-31: Tube-Tube interface .................................................................................. 53
Figure 3-32: Main parachute/shock cord attached to eye bolt and recovery system
bulkhead ........................................................................................................................ 54
Figure 3-33: Nose cone/upper body tube interface ....................................................... 55
Figure 3-34: Integrated avionics assembly, main parachute, sabot and UAV assembly 55
Figure 4-1: Airfoil Comparison ....................................................................................... 82
Figure 4-2: NACA 4412 Polar ........................................................................................ 83
Figure 4-3: Tail .............................................................................................................. 84
Figure 4-4: Wing Rotator Mechanism ............................................................................ 86
Figure 4-5: UAV In Stowed Configuration ..................................................................... 87
Figure 4-6: Stowed tail .................................................................................................. 87
Figure 4-7: Wing Boundary Conditions.......................................................................... 90
8
Figure 4-8: Effective Strain ............................................................................................ 91
Figure 4-9: Wing Deflection ........................................................................................... 92
Figure 4-10: New Wing Boundary Conditions ............................................................... 93
Figure 4-11: New Effective Strain .................................................................................. 94
Figure 4-12: New Wing Deformation ............................................................................. 94
Figure 4-13: Wing Rotator Boundary Conditions ........................................................... 95
Figure 4-14: Stress Distribution ..................................................................................... 96
Figure 4-15: Deformation .............................................................................................. 96
Figure 4-16: Ground Station User Interface ................................................................ 102
Table of Tables
Table 3-1: Rocket Budget Summary ............................................................................. 14
Table 3-2: Hardware Specifications............................................................................... 26
Table 3-3: Launch Loading ............................................................................................ 30
Table 3-4: Recovery Shock Calculations ....................................................................... 31
Table 3-5: Carbon Fiber Properties ............................................................................... 31
Table 3-6: Axial Stress Calculations .............................................................................. 32
Table 3-7: Buckling Calculations ................................................................................... 32
Table 3-8: Bending Calculations ................................................................................... 32
Table 3-9: Payload Bulkhead Bolt Shear Calculations .................................................. 35
Table 3-10: Threaded Rod Sizing ................................................................................. 35
Table 3-11: Motor Retention Sizing ............................................................................... 36
Table 3-12: Stringer Sizing ............................................................................................ 38
Table 3-13: Parachute Descent Rates .......................................................................... 43
Table 3-14: Potential Rocket Failure Modes ................................................................. 44
Table 3-15: Scale Model Dimensions ............................................................................ 50
Table 3-16: Tasking and Integration Schedule .............................................................. 55
Table 3-17: Possible Launch Failure Modes ................................................................. 68
Table 3-18: Potential Failure Modes ............................................................................. 69
Table 3-19: Tool Use Injury Potentials and Mitigations ................................................. 73
Table 4-1: UAV Characteristics ..................................................................................... 77
Table 4-2: Stability Analysis .......................................................................................... 78
Table 4-3: Flight Case Analysis ..................................................................................... 79
Table 4-4: Deployment Loading .................................................................................... 88
Table 4-5: Foam Material Properties ............................................................................. 88
Table 4-6: Fiberglass material Properties ...................................................................... 89
Table 4-7: Fiberglass Layup Properties ......................................................................... 89
Table 4-8: Wing Deflection Calculations ........................................................................ 90
Table 4-9: Carbon Fiber Layup Properties .................................................................... 92
Table 4-10: Carbon Fiber Deflection Calculations ......................................................... 92
Table 4-11: Avionics testing ........................................................................................ 101
Table 4-12: Precision of Transmitted Data .................................................................. 104
Table 4-13: Sensor Precision ...................................................................................... 104
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Table 4-14: Potential UAV Failure Modes ................................................................... 110
Table 4-15: Potential Avionics and ground Station Failure Modes .............................. 114
Table 5-1: Budget ........................................................................................................ 118
Table 5-2: Total Budget ............................................................................................... 118
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1 SUMMARY OF CDR
1.1 TEAM SUMMARY
MIT Rocket Team,
Massachusetts Institute of Technology
Cambridge, MA
Dr. Paulo Lozano
Faculty Advisor
Andrew Wimmer
Safety Officer, Rocket Owner, TRA # 9725 Level 3
John Kane
Local NAR Contact
1.2 LAUNCH VEHICLE SUMMARY
The purpose of the launch vehicle is to reach an apogee of 1 mile and deploy the UAV
payload after descending to an altitude of 2500 feet. Diagrams of the vehicle are
provided below in the rocket section.
The carbon fiber and fiberglass airframe will be 132 inches long, the inner diameter of
the rocket tube is designed to be 6 inches, and the outer diameter of the fins is 16.25
inches. Furthermore, the mass of the rocket is projected to be 44.5 lbs (not including a
payload mass of 7 pounds) and ballasted (in the nose cone) as necessary in order to
reach an apogee of 5280 feet using a single commercial Cesaroni L1115 motor.
Payload deployment will be performed at 2500 feet using two sabot halves that will be
pulled out of the tube by the drogue parachute and separated using the deployable UAV
wings.
1.3 PAYLOAD SUMMARY
The rocket payload will consist of a 3.75 ft long, 7.5 pound UAV that will be launched
from the rocket at an altitude of 2500 ft. It will fit inside the rocket by means of folding
wings, tail, and propeller. The UAV will have a TR 35-30A 1700kv Brushless Outrunner
motor onboard but will function as a glider for the majority of its flight.
The UAV will fly to GPS coordinates supplied by a human operator. The UAV will not
require advanced airplane or flight knowledge, which will make it useful for search and
rescue type missions as well as for scientific research. The UAV will carry GPS tracking,
airspeed sensors, atmospheric sensors, an accelerometer, a video capture device, and
an onboard computer.
11
2 CHANGES SINCE PDR
2.1 LAUNCH VEHICLE CHANGES
A number of changes have been made to the rocket design since PDR in order to solve
problems and improve the design. The changes revolve primarily around review of the
design to simplify fabrication and integration.
Airframe
The length of the rocket was increased to 132 in to allow for an internal coupler
between the two airframe segments
Fin slots in the motor centering rings will extend to the motor tube to improve
strength
Rocket diameter decreased from 6.5” to 6”
Inter-segment coupler changed from external doublers to internal phenolic
coupler
Integration access door added to upper part of intersegment coupler so permit
integration with the new coupler design
Material type changed to Soller Composites braided carbon fiber
Recovery
The recovery simulation was updated to reflect the parachutes being X-form
rather than hemispherical
Drogue and main parachute diameters increased to 4ft and 13ft, respectively, to
account for decreased drag of X-form profile.
Deployment
Standard (non-expanding) foam will be used for the sabot
Propulsion
The sub-scale test motor will be an Aerotech G80
The full-scale test flight motor has been changed to a Cesaroni K1085
Avionics
Two plates are used for mounting avionics components within the avionics
assembly to allow the threaded rod to pass through the middle
2.2 PAYLOAD CHANGES
Avionics
12
Replaced uBlox GPS module with MediaTek to reduce cost
Removed external SCP1000 sensor from flight computer to use pressure sensor
internal to flight computer
Replaced SD Card board with FAT16/32 support to one without to reduce cost.
2.3 ACTIVITY PLAN CHANGES
Since the completion of the preliminary design review, the team has moved into the final
stages of planning for educational outreach events at both the Boston Museum of
Science and the MIT Museum. The event at the Museum of Science will be held on
Saturday February 5th, with two 15-20 minute talks. One discussion will focus on rocket
technology and history, and another will be a discussion of the space industry. Two
hands-on workshops will be conducted before, during, and after the presentations: one
will allow participants will build their own Alka-Seltzer, and another where participants
will build their own parachutes. We are currently working on securing space to set off
the rockets and drop the parachutes. We have set a date for our outreach event at the
MIT Museum will follow a similar format. The MIT Museum event will take place on
Sunday, May 1, from 10am to 1pm.
3 DESIGN AND VERIFICATION OF LAUNCH VEHICLE
3.1 MISSION STATEMENT, REQUIREMENTS, AND MISSION SUCCESS
CRITERIA
Mission Statement
Use a rocket to rapidly deploy a UAV capable of completing search and rescue type
missions with the use of a ground based system requiring little to no UAV flight training.
Constraints
Follow all rules of NASA USLI 2011, including but not limited to:
Rocket apogee shall be closest to but not exceeding 5280ft.
At no time may a vehicle exceed 5600ft
Minimum science payload deployment of 2500ft
Must carry one PerfectFlight MAWD for official altitude recording
Dual deployment recovery must be used
Dual altimeters must be used for all electronic flight systems
Each altimeter must have its own battery and externally located arming switch
Recovery and payload electronics must be independent from each other
At all times the system must remain subsonic
Shear pins must be used in the deployment of both the drogue and main parachute
13
All components of the system must land within 2500ft of the launch site in a wind
speed of 10 mi/hr
Scientific method must be used in the collection, analysis and reporting of all data.
Electronic tracking devices must be used to transmit the location of all components
after landing
Only Commercially available, NAR/TRA certified motors may be used
Full-scale flight model must be flown prior to FRR
Students must do 100% of all work for USLI competition related projects
$5000 maximum value of rocket and science payload as it sits on the launch pad
Requirements
Launch UAV with Rocket
Meet the needs of NASA Science Mission Directorate including:
o Gather atmospheric measurements of: Pressure, Temperature, relative
humidity, solar irradiance, and ultraviolet radiation at a frequency no less than
once every 5 seconds upon decent, and no less than once every minute after
landing.
o Take at least two still photographs during decent, and at least 3 after landing.
All pictures must be in an orientation such that the sky is at the top of the
frame.
o All data must be transmitted to ground station after completion of surface
operations.
o Science payload must carry GPS tracking unit.
Successfully perform model search and rescue/reconnaissance mission
3.2 MAJOR VEHICLE MILESTONE SCHEDULE
The full schedule for rocket development may be found in Appendix 2. Key dates are
presented below for reference:
9/10: Project initiation
11/19: PDR materials due
12/30: Scaled test launch
1/24: CDR materials due
2/15: Balloon deployment test
2/30: Full-scale test launch
3/21: FRR materials due
4/14: Competition launch
3.3 SYSTEM LEVEL REVIEW
14
3.3.1 ROCKET DESIGN AND SUBSYSTEMS
As described in the summary section, the purpose of the rocket is to reach 1 mile and deploy the UAV at an altitude of 2500 feet. This will be accomplished with a Cesaroni L1115 motor and a 132 inch long, 6 inch diameter airframe. The UAV will be contained within a sabot, which will be located just aft of the nosecone. The drogue parachute will be above the sabot, the main parachute below the sabot, and the avionics below the recovery system. The overall rocket can be seen in Figure 3-1.
FIGURE 3-1: OVERALL ROCKET
Furthermore, the rocket budget summary (for mass and cost) can be seen in Table 3-1.
TABLE 3-1: ROCKET BUDGET SUMMARY
System
Mass (kg) Cost (USD)
Rocket
Propulsion 5.46 552.00
Airframe-Body 3.62 455.09
Airframe-Fairing 1.01 27.00
Avionics/Comm 0.99 947.38
Payload Support Equipment 1.82 152.24
Recovery 2.19 434.60
SUBTOTAL 15.09 2568.31
The subsystems, which will be described in greater detail below, are:
Airframe
Recovery
Deployment
Propulsion
Avionics/Communications
3.3.2 SUBSYSTEM REQUIREMENTS AND DESCRIPTIONS
15
Airframe
The airframe is comprised of the following components:
Body Tube
Nose Cone
Fins
Motor Retention System
Recovery System Bulkhead
Each of these will be described in detail below.
The body tube is a carbon fiber laminate tube of inner diameter 6”. The laminate is a 2-
ply layup of Soller Composites 14.5 oz/sqyd biaxial sleeve carbon fiber fabric and
Aeropoxy 2032/3665 matrix. Carbon fiber was chosen as the material for the primary
structure due to its high strength-to-weight ratio, toughness, and ease of manufacture to
customized shapes and dimensions. The biaxial sleeve was chosen due to difficulties in
fabricating wrinkle-free tubes. All layups for the rocket are done in-house using a
custom oven and vacuum bagging equipment in the rocket team lab. For fabrication and
transportation reasons, it would be difficult to make the entire tube in one segment. As a
result, the body tube is split into 2 segments, with a “seam” just below the base of the
sabot, as seen in Figure 3-2. The two segment lengths are 56” for the upper segment
and 48” for the lower segment. The seam between the tubes is accomplished by adding
use of a phenolic coupler that is epoxied to the upper segment. The use of an internal
coupler was accomplished by lengthening the upper segment.
16
FIGURE 3-2: TUBE COUPLER SEGMENT
Additionally, the tube will have 2 pressure relief holes (of 0.25” diameter, unless
otherwise specified) in each of the following locations:
Just above the fins in the propulsion section
Avionics bay: the hole for the switches will double as a pressure relief hole
In the middle of the section between the avionics bay and the sabot
In the nose cone
The nose cone is a 2-ply carbon laminate, just like the body tube. The shape is a
tangent ogive for manufacturing simplicity, and the length was chosen to fit the drogue
parachute and maintain stability of the rocket. The length is 27 in and shape is shown in
Figure 3-3.
FIGURE 3-3: NOSE CONE
The nose cone is mounted to the body tube using 4 nylon 2-56 bolts (MMC 97263A077)
fastened to the inside using Helicoils (MMC 91732A203), which will act as shear pins.
Shear area is provided using doublers on both the top of the tube and the inner portion
of the nose cone, as shown in Figure 3-4. Similarly, the doublers are on the outside of
the tube to allow the sabot to cleanly exit the body tube. Bolts are used because they
can be easily threaded into the inner doubler during integration and will fail at low
loading since they are plastic.
17
FIGURE 3-4: NOSE CONE COUPLER
Four fins were chosen with the dimensions as shown in Appendix 3 for rocket stability
reasons (see Section Error! Reference source not found. and Modeling XX). The fins
are a carbon fiber, 3/16” plywood, carbon fiber sandwich laminate to maximize stiffness
with minimum mass. The fins are located in position and angle relative to the rocket
using slots that are laser-cut into the motor centering rings. Oversized slits added to the
body tube to allow the fins to pass through, but provide no DOF restrictions. Fabrication
of the fins is as follows:
Sand the laser-cut plywood core edges (not tabs) to a taper
Laminate the plywood core with a ply of carbon fiber on each face using standard
plate lamination techniques (see manufacturing plan section)
Obtain body tube with motor tube and centering rings installed
Affix fins to the centering rings/motor tube assembly using 5 minute epoxy and let
cure
Apply another layer of carbon fiber across and between the fins, i.e. “Tip-to-Tip”
The motor mount will consist of a commercial 75mm motor tube and laser-cut, plywood
centering rings. There will be four centering rings in total, one located at each end of the
motor tube and two in the middle. The farthest forward will be made from 1/2” plywood.
The farthest aft centering ring will be made from two rings of 3/16” plywood sandwiched
together; the OD of the forward ring will be the ID of the body tube, and the OD of the
aft ring will be the OD of the body tube. This will transfer some of the thrust load through
compression of the aft centering ring, rather than through shear in the epoxy joints
holding the motor mount in the body tube. The middle centering rings will be made from
½” plywood, with four slots to accept the fin tabs. The fin tabs will also be slotted at the
centering ring locations to allow the fins to contact the motor tube for additional support.
One will be located near the forward edge of the fin tabs and the other near the aft edge
of the fin tabs, close to the aft-most centering ring. Plywood is chosen because it is
relatively cheap, strong, light, and able to withstand the high temperatures of the motor
casing without deforming.
Motor retention will be accomplished as follows. Two 8-32 T-nuts will be mounted to the
aft-most centering ring, 180° apart at a radius of roughly 2.375”. 1.5” 8-32 screws will go
through two small clearance holes in the motor retention plate and screw into the T-nuts
to hold the plate in place. The motor retention plate will be a piece of 1/32” steel sheet
that has a hole cut in it; this hole will be made large enough for the motor‟s nozzle to fit
through, but small enough to keep the motor casing from falling out of the motor tube.
There is a thrust ring on our 75mm hardware that prevents the motor casing from
moving forward during burn.
The mounting and retention system can be seen in Figure 3-5.
18
FIGURE 3-5: MOTOR CENTERING AND RETENTION
The recovery system bulkhead serves as a reaction point for lateral forces from the
payload, which come from two sources: inertial force of payload during boost, and
drogue drag force between drogue parachute deployment and main parachute
deployment. Axial forces will be reacted through a threaded rod to the motor casing, for
reaction via the motor mount assembly. The bulkhead must also be removable to
enable removal of the avionics bay, which sits between the motor retention bulkhead
and the recovery system bulkhead. The threaded rod will be used to affix the payload
support bulkhead to the airframe. No hardware will be used to affix the payload support
bulkhead to the lower body segment directly. The body tube will have a doubler
concentric with the recovery system bulkhead to allow for transfer of shear loads. The
bulkhead will need to attach to both the charge released locking mechanism and the
quick link to the main parachute shock cord. As a result, the bulkhead needs to have an
eye bolt that is capable of transferring the loads to the bulkhead, which will be done
through via an eyenut (MMC 3274T41) and threaded rod (MMC 95412A652). The
design of the bulkhead is shown in Figure 3-6. Furthermore, the bulkhead will be
manufactured from polycarbonate for its relative cheap price (compared to other plastic
rod stock).
19
FIGURE 3-6: RECOVERY SYSTEM BULKHEAD
Recovery
A detailed description of the recovery process can be found in the Section 3.2.
Deployment
Deployment of the UAV and parachutes is as follows.
Initially, the stacking of the rocket above the recovery system bulkhead is as follows (as
seen in Figure 3-7 and Figure 3-8):
Payload Bulkhead attachment quick links
Charge released locking mechanism
Main parachute
Sabot base hardpoint
Sabot halves (cradling UAV)
Sabot top hardpoint
Drogue parachute quick link
Drogue parachute
20
Nose cone ejection charge
Note: There is a redundant igniter in the charge in the nose cone and a redundant
igniter in the charge released locking mechanism.
FIGURE 3-7: SABOT OVERVIEW
FIGURE 3-8: PAYLOAD INTEGRATION STACKING
The deployment then occurs as follows:
Just after apogee, nose cone ejection charge fires
Nose cone separates, but remains attached to the drogue parachute
Drogue parachute deploys
Rocket descends to 2500 feet
At 2500 feet, the charge released locking mechanism fires. Mechanism to be
used is the “FruityChutes L2 Tender Descender”
The drogue parachute pulls the sabot out of the rocket tube
As the sabot leaves the tube, the spring-loaded UAV wings push the sabot
halves apart
The sabot pulls the main parachute bag out behind it
Main parachute deploys and remains attached to the main body tube
After deployment, the rocket will fall to the ground in two sections, as shown in Figure
3-9:
21
Sabot and nose cone, which are attached to the drogue parachute via the upper
hardpoint and a shock cord
Main body tube, which is attached to the main parachute via the recovery system
bulkhead and a shock cord
FIGURE 3-9: RECOVERY CONFIGURATION
Deployment into two pieces (rather than one) is performed in order to minimize the
chance of contact between the sabot/UAV and the body tube after separation. This will
enable the drogue parachute to pull the UAV/sabot away from the rocket to allow clean
separation and minimize the chances of entanglement.
As described above, the UAV is encased within the two sabot halves, which are made
of foam and laminated in a ply of fiberglass so as to maintain shape. Force will be
transferred between the hardpoints using a 4x 10-24 nylon threaded rods (MMC
94435A355), which will mount to the upper and lower hardpoints using clearance holes
and nuts. Finally, plastic hardpoints are glued to the upper and lower ends of the sabot
halves. These hardpoints enable recovery and deployment system fixtures to be
attached to the sabot. One of these hardpoint sets is shown in Figure 3-10. As can be
seen below, the hardpoint halves overlap to ensure force transfer between halves when
in tension. Adhesive is applied as shown in the figure. An eye bolt is threaded into the
lower hardpoint half, which serves as the attachment points for:
Lower hardpoint: the charge released locking mechanism
22
Upper hardpoint: the drogue parachute and upper shock cord (attaches to nose
cone)
It should be noted that the upper hardpoint will require eye bolts in both hard point
halves due to ensure both sabot halves remain attached to the drogue parachute.
FIGURE 3-10: SABOT HARD POINT
Propulsion
The rocket will be powered by a Cesaroni L1115 solid rocket motor. This motor was
chosen because it is commercially available and does not require any modifications in
order to reach the flight altitude requirement of 5280 feet based off the mass estimates
available this early in the design process. The motor is actually more powerful than
required given the current mass estimates, but this will ensure that even with mass
creep over multiple design iterations, the rocket mass can be optimized with ballast
weight to come as close to 5280 feet as the models can predict.
The Cesaroni L1115 is also reloadable and relatively inexpensive compared to its
Aerotech counterparts. It does not require extensive ground support equipment
compared to hybrid motors, which were originally considered for propulsion. The L1115
is 75mm in diameter, 24.5 inches in length, and has a total impulse of 4908 Newton-
seconds over a 4.49 second burn time.
For the full-scale test, the Cesaroni K1085 solid rocket motor will be used. The K1085
has enough power to launch the full system up to an altitude of 2000 feet and still has
the same diameter as the L1115, so minimal changes will have to be made to the motor
housing section for the full scale test launch. The K1085 is 75mm in diameter, 13.8
inches long, and provides 2486 Newton-seconds of thrust over a 4.84 second burn time
Avionics/Communications
The purpose of the rocket avionics is to control parachute deployment while collecting
rocket flight data and relaying it to the ground station.
23
The rocket avionics system is comprised of two flight computers (miniAlt/WD and
ARTS2) and an ARTS2 transmitter. The miniAlt/WD flight computer serves as a backup
altimeter that measures the rockets altitude during launch and stores in on the computer
board and will fire a redundant igniter for the recovery charge after the ARTS is
programmed to. This data can be retrieved after rocket recovery where the miniAlt/WD
flight computer is connected to the ground station computer via a miniAlt/WD to PC
Connect Data Transfer Kit. The ARTS2 flight computer handles primary parachute
deployment as well as determining the rocket state variables and flight states. The
ARTS2 Transmitter transmits the data from the ARTS2 to the ground station receiver.
Rocket Flight data includes:
State Variables:
o Altitude
o Maximum Altitude
o Velocity
o Acceleration
Flight State:
o On Pad
o Thrust
o Coast
o Apogee
o Descent
o Drogue parachute Deployment
o Main parachute Deployment
Power Supply
Three 9 volt batteries will provide power for the flight computers and transmitters. One
of the batteries will be dedicated towards powering the miniAlt/WD while the other two
will power the ARTS2 flight computer and telemetry system to create a power source
redundancy in case one was to fail. On the ARTS2 board one battery powers the two
systems while the other powers the igniters. They will be located inside the removable
rocket avionics section of the rocket, alongside the rest of the avionics system.
Hardware Description
MiniAlt/WD Logging Dual Event Altimeter (PerfectFlite)
This flight computer measures the rocket‟s altitude by sampling the surrounding air
pressure relative to the ground level pressure. The altitude above the launch platform is
calculated every 50 milliseconds. After launch, the device continuously collects data
until landing. Altitude readings are stored in nonvolatile memory and can be
downloaded to a computer through a serial data I/O connector. The miniAlt/WD has two
24
channels for parachute deployment; one for the main parachute and the other for
drogue parachute.
FIGURE 3-11: MAWD FLIGHT COMPUTER
Altimeter Recording and Telemetry System (ARTS2 Flight Computer) (Ozark
Aerospace)
This flight computer calculates the rockets altitude by sampling the surrounding air
pressure relative to the ground level pressure and measuring the rockets acceleration.
The rate at which the altitude above the launch platform is calculated is adjustable and
will be set at 100 samples per second with an overall recording time of 5.4minutes.
Altitude readings are sent to the ground station via the ARTS2 telemetry transmitter.
Also the altitude and other flight data are stored in nonvolatile memory to be
downloaded to a computer through a serial data I/O connector. The ARTS2 has two
channels for parachute deployment; one for the main parachute and the other for
drogue parachute.
1. Terminal Connector
2. GPS Connector
3. Programming header
4. Battery Configuration
5. Main Battery Connection
6. Power Switch Connector
7. 9V Pyro Battery Connection
8. Option Switches
9. Output Channel Terminals. Channel 1 Apogee, Channel 2 Main.
25
FIGURE 3-12: ARTS2 FLIGHT COMPUTER
ARTS2 Telemetry Transmitter (Ozark Aerospace)
100mW 900MHz spread spectrum transmitter
Integrated wire antenna on the board is connect to a larger antenna on the rocket
Works with the ARTS-TT2-W and ARTS-TT2-RPSMA
ARTS flight computer gets connected directly to the transmitter board
Transmits real time flight data to the ARTS telemetry receiver
FIGURE 3-13: ARTS2 TELEMETRY TRANSMITTER
ARTS2 Telemetry Receiver (Ozark Aerospace)
Receives telemetry from the ARTS2 transmitter and sends it to the computer
Connect to the ground station computer via a serial cable
26
FIGURE 3-14: ARTS2 TELEMETRY RECEIVER
TABLE 3-2: HARDWARE SPECIFICATIONS
Hardware Operating
Voltage
Minimum
Current Dimensions Weight
Altitude
Accuracy
Operating
Temperature
Maximum
Altitude
MiniAlt/WD 6-10 volts 10
milliamps
0.90”W,
3.00”L,
0.75”T
20
grams +/- .5%
0C to 70C
25,000
feet
ARTS2 9-25 volts
1.40"W,
3.75"L,
0.75"T
~20
grams
100,000
feet
ARTS2
Transmitter
6.8 -25
volts
~2.50"W,
~7.50"L,
~1.50"T
~200
grams
100,000
feet
Switches
The single power switch is a push-on/push-off switch that delivers power to all of the
avionics components when on. The three arming switches are RBF pull-pin switches.
Software
Telemetry software:
Displays flight data in real time using text, 2-D, and 3-D graphical user interfaces.
27
FIGURE 3-15: ARTS GUI
ARTS Software V1.61 (Data Analyzer):
Used in analyzing the data collected by the ARTS2 and also configuring parachute
deployment and sample rate settings.
FIGURE 3-16: ARTS DATA ANALYZER
Parachute Deployment
Both the ARTS2 and the miniAlt/WD are programmed to deploy the drogue parachute at
apogee, while the main parachute and the UAV are set to deploy after apogee is
reached at an altitude of 2500 feet. This creates system redundancy in case one of the
flight computers fails.
Transmission from Rocket to Ground Station
Since the carbon fiber material of the rocket body tube disrupts any RF signals the wire
on the ARTS2 transmitter will be extended out of the avionics bay. There are two
possible options for doing this.
28
1) The 14 gauge copper wire is connected to the transmitter antenna via a binding
post. It then extends from the avionics bay and warps around the upper section
of the rocket body. Each loop of the helix is spaced 33 centimeters apart to
prevent destructive interference. Kapton tape is placed above and below the wire
to prevent contact with the carbon fiber.
2) The ARTS2 transmitter antenna is extended directly to the fiberglass nose cone
and connects to a 25 inch long (4 millimeter in diameter) RF antenna. It should
be noted that the nose cone is currently carbon fiber, so it would need to be
changed to fiberglass if this modification is made.
Arming and Power Switches
FIGURE 3-17: POWER SWITCH
The avionics bay contains two plunger type switches that connect the power to the
miniAlt/WD flight computer and the ARTS2 altimeter telemetry system and two switches
with Remove Before Flight [RBF] pins to arm the flight computers. The power and
arming switches are used in order to prevent premature firing of ejection charges and
power usage before the rocket is on the launch pad.
Mounting/Placement
Placed in the avionics bay, which is in the lower segment of the rocket as described
below. The flight computers will be mounted in such as way so that their pressure and
acceleration readings are not disturbed. This means that the barometer on both the
ARTS2 and miniAtl/WD would have to have at least a 1 centimeter clearance from any
closest surface parallel to it. Also, the ARTS2 will be mounted with its length parallel to
the rocket‟s length in order for the accelerometer to record proper positive values.
The avionics will be mounted into an avionics integration tube, which is shown below in
Figure 3-18.
29
FIGURE 3-18: AVIONICS PACKAGE
As can be seen in the figure, the boards and battery are mounted to a plate, which will
be mounted vertically in the rocket frame. A series of L brackets will be used to mount
this vertical avionics plate to the avionics tube, which can be integrated with the rocket
in a preassembled form. Additionally an ELT will be mounted to the side of the avionics
assembly. The hole in the top of the avionics package is for wires to reach into the
upper portion of the rocket (a similar hole exists in the recovery system bulkhead). The
switches on the bottom image are used to arm each of the following just before launch:
ARTS2/ARTS2 Transmitter Power Switch
MAWD Power Switch
ARTS2 Arming Switch
MAWD Arming Switch
The power switch will be push-on/push-off. The arming switches will be armed by
removing Remove Before Flight [RBF] tags. The switches will also have the capability of
being flipped without reinsertion of the RBF tag.
The bottom avionics plate is grooved so that the phenolic tube packaging shell can be
attached. The boards and batteries can be mounted to the avionics plate, which is
mounted to the top plate using 4x 6-32 fasteners. Nutplates are glued to the insides of
each of the mounting brackets, such that bolts can be used without standard nuts. After
insertion into the tube, four additional 6-32 fasteners are bolted from the bottom plate
30
into the bottom L brackets, the holes of which are threaded. After this is assembled, the
whole avionics package may be inserted into the rocket as described in the payload
integration plan. This design was chosen to make the avionics assembly as modular as
possible, while still maintaining access just before flight and low mass/cost of the
assembly.
3.3.2.1 DRAWINGS
[See CDR-Drawings document]
3.3.2.2 ANALYSIS RESULTS
In order to verify the design of the rocket, a battery of analysis was applied to the rocket airframe, bulkheads, and mechanisms. The order of analysis is as follows:
Define loading conditions
Design part
Use hand calculations to size the part
Validate hand calculations using finite element method
Re-size as necessary Loading Conditions Determining loading conditions for a vehicle that must withstand a variety of largely unknown dynamic and static loading is a difficult task. Furthermore, the rocket airframe can be significantly overdesigned without applying too significant of a penalty to the mass budget. As a result, the loading conditions were often estimated using significant margin to account for uncertainty. Launch Loading is summarized in Table 3-3.
TABLE 3-3: LAUNCH LOADING
Launch Loading
Aero Loading 90 lbf
Peak Thrust 385 lbf
Payload Mass 15 lbm
Max Accel 8 G
Total Axial 595 lbf
MUF 1.5
Design Axial 892.5 lbf
Total Lateral 150 lbf
MUF 1.5
Design Lateral 225 lbf
Aerodynamic loading is determined from the Rocksim model, peak thrust is determined
from the Cesaroni L1115 Thrust Curve, payload mass is determined from the UAV
31
design and sabot, and maximum acceleration is determined from the Rocksim model.
Although many of these peak loads are applied independently from each other, to
provide for a conservative calculation, the loads are summed to create a total load,
which is then margined by a 1.5 model uncertainty factor, resulting in a design axial load
of 890 lbf. Lateral loading is determined by summing half of the aerodynamic and
payload forces and margining by a 1.5 model uncertainty factor. This is assumed to be
highly margined since as much as half of each of these loads is unlikely to be applied in
the lateral direction. Regardless, the design lateral load is therefore 225 lbf.
Recovery shock calculations are determined by examining the change in momentum of the rocket due to deployment, as shown in Table 3-4.
TABLE 3-4: RECOVERY SHOCK CALCULATIONS
Recovery Shock Calculations
Initial Rate 64 ft/s
Final Rate 18 ft/s
g 32 ft/s^2
t 0.1 s
accel 460 ft/s^2
Gs 14.375
Rocksim Gs 8
MUF 2
Design Gs 28.75
Design Force 431.25 lbf
Recovery calculations show the descent rates of the system prior to deployment (under the drogue) to be 64 ft/s and of the main rocket after deployment (under tha main) to be 18 ft/s. Assuming a deployment time of 0.1s, this results in 14 Gs. Adding a model uncertainty factor of 2 to this results in 29 Gs, which (given the mass of the payload system) results in a design recovery shock force calculation of 430 lbf. Body Tube Analysis As described in the design section, the body tube is made from Soller Composites Biaxial Weave carbon fiber. The modulus and strength are taken from Soller Composites and the resulting strain allowable is derived, as shown in Table 3-5-5.
TABLE 3-5: CARBON FIBER PROPERTIES
Material Properties
E 4675000 psi
v 0.3
E (claimed) 34 Msi
E_lam (claimed)
4.675 Msi
strength 110 ksi
32
strain 0.02352941 strain
FOS 3
strain w/MOS 7843.13725 µstrain
Using these properties, hand calculations could be performed for axial compression, global buckling, and bending. These intermediate calculations as well as the resulting margins of safety can be seen in Table 3-6 through Table 3-8. It should be noted that, for the sake of being conservative, lateral loads are taken to be applied at the top of the rocket and restrained at the base.
TABLE 3-6: AXIAL STRESS CALCULATIONS
Axial Stress Calculations
ID 6 in
OD 6.092 in
Area 0.87372718 In^2
Axial Stress
1021.4859 psi
strain 0.0002185 strain
218.499658 µstrain
MOS 34.8954211
TABLE 3-7: BUCKLING CALCULATIONS
Buckling
r/t 81.5217391
Z 62136.005
Kc 4000
Fcr 3182.89745 psi
L/r 28.2666667
Axial Load 2780.98402 lbf
MoS 2
Axial Load Allowable
1390.49201 lbf
MOS 0.55797424
TABLE 3-8: BENDING CALCULATIONS
Bending
I 7.620816554 in^4
z 3.75 in
M 23850 in-lbf
stress 11735.94711 psi
strain 0.002510363 strain
2510.363018 µstrain
MOS 2.124304015
33
In order to verify these calculations, a finite element model was developed in Femap. The applied boundary conditions of the lateral case are shown in Figure 3-19. The boundary conditions in the axial load case are shown in Figure 3-20.
FIGURE 3-19: AXIAL CASE BCS
FIGURE 3-20: LATERAL CASE BCS
This model was then solved using NEi Nastran, resulting in the ply 1 effective strain and displacement outputs for the axial case, as shown in Figure 3-21. The results for the lateral case are shown in Figure 3-22.
34
FIGURE 3-21: AXIAL CASE RESULTS
35
FIGURE 3-22: LATERAL CASE RESULTS
Payload Bulkhead Analysis As described in the design part of the document, 4x #6 bolts are used in order to fasten the tube segments together. From this, the effective strain may be determined and compared to allowables. This calculation takes into account the allowable shear area, using a 4 ply thick doubler and assuming that only two of the bolts are being used, as shown in Table 3-9.
TABLE 3-9: PAYLOAD BULKHEAD BOLT SHEAR CALCULATIONS
P/L Bulkhead Bolt Hole Shear
n 4
n "used" 2
hole dia 0.0997 in
t 0.092 in
shear area 0.0091724 in^2
stress 23508.0241 psi
strain 0.00502845 strain
5028.45435 µstrain
MoS 0.55975111 µstrain
The 3/8” threaded rod is used to connect the parachute shock loads to the motor retention system, as shown in Table 3-10.
TABLE 3-10: THREADED ROD SIZING
Threaded Rod Sizing
rod dia 0.2983 in
36
A_tot 0.069887 in^2
stress 6170.67566 psi
allowable stress
33000 psi
MoS 4.34787466
Motor Retention Analysis Since the threaded rod transfers shock load to the motor retention system, the motor retention system must be able to react the load of the deployment. It must also, however, be able to react the original axial load. The initial sizing calculations are as shown in Table 3-11.
TABLE 3-11: MOTOR RETENTION SIZING
Motor Retention Sizing
# Used 3
thk 0.5 in
ID 3 in
OD 6 in
avg D 4.5 in
"b" 14.1371669 in
I 0.14726216 in^4
M 669.375 in-lbf
z 0.25 in
stress 1136.36629 psi
failure stress
1500 psi
MoS 0.31999691
This was validated using Ansys Workbench. The loading conditions are shown in Figure 3-23.
37
FIGURE 3-23: MOTOR RETENTION BCS
The resulting displacement is shown in Figure 3-24 and the stress is shown in Figure 3-25.
FIGURE 3-24: MOTOR RETENTION DISPLACEMENT
38
FIGURE 3-25: MOTOR RETENTION STRESS
Payload Support Equipment Sabot Stringer Sizing As described in the design portion of the document, the stringers are 4x #10 rods that connect the top and bottom parts of the sabot together. The size of these aluminum-threaded rods can be verified as shown in Table 3-12.
TABLE 3-12: STRINGER SIZING
Stringer Sizing
stringer dia 0.1318 in
# 4
A_tot 0.05457336 in^2
stress 7902.20724 psi
rho 0.041 lb/in^3
length 50 in
sabot dia 6.5 in
mass 0.11187539 lbm
allowable stress
33000 psi
MoS 3.17604841
3.3.2.3 TEST RESULTS
[See Section 3.1.3]
3.3.2.4 MOTOR SELECTION
[See Section 3.1.3]
39
3.3.3 DEMONSTRATE DESIGN MEETS SYSTEM LEVEL REQUIREMENTS
[See section 3.1.3]
3.3.4 WORKMANSHIP AND RELATION TO SUCCESS
Through past experiences the MIT Rocket Team has identified that the workmanship of
individual components plays an integral role in the final outcome of any project. With
this in mind, the team has set in place schedule of testing and teaching of the various
skills necessary for the fabrication and assembly of all components. Fabrication
methods used by the team are learned from experienced sources, and all methods are
tested at various scales. Team members are taught basic fabrication techniques under
the instruction of senior members, and all components are inspected and tested as
necessary before they are cleared for flight.
The rocket tubes needed for this year‟s USLI competition are the largest composite
tubes that the MIT Rocket Team has constructed to date. As such, the team has spent a
large amount of time evaluating our fabrication techniques and their ability to scale up.
In this pursuit, the team has constructed a series of tube sections using various
materials and methods to test the feasibility of our upcoming fabrication schedule. In the
course of this investigation the team has learned a few important lessons.
First and foremost, the team has learned that techniques normally employed for the
fabrication of smaller diameter body tubes do not easily work for the large diameter
needed for this project. The normal technique employed by the team required plain-
weave fabric to be wrapped around a mandrel during the wet lay-up procedure until the
appropriate number of plies had been achieved. The layup would then be wrapped in a
release film and a layer of breather material. Finally a vacuum bag would be
constructed around the whole layup and when a good vacuum was pulled, the whole
bag would be placed into our custom built curing oven. However when this process was
applied to the large diameter tube sections, we found that because the fabric was
wrapped around the tooling, there was no way to prevent wrinkles from forming within
the layup when a vacuum was pulled. Furthermore because the wrinkles were
throughout the structure of the tubes section, the structural integrity of the part was
weakened. With this lesson we realized we had to move onto a different fabrication
technique.
The team has since developed a relatively simple way of constructing large diameter
tube sections without wrinkles. The new fabrication method uses many of the same
techniques of the old method, but now a tubular braided fabric is being used in placed of
the plain-weave fabric. This change allows for the fabric to adjust diameter to exactly
match what is required. Furthermore, a new bagging method has been developed, and
tested with sample sections, which results with no seam in the finished tube.
40
For the fabrication of other vehicle components, such as the rocket fins and faring, a
similarly high standard of workmanship is held. To this end, the plywood cores of our
fins are first designed in SolidWorks and then the model is used to generate a cut file for
a Computer-Numeric-Controlled laser cutter available on campus through the Edgerton
Center. This method allows for extremely accurate cores, which are then faced with
carbon fiber as described in section [3.1.7]. For the fabrication of the faring, a computer
model is first generated in SolidWorks, and a full sized template is created of the
complex surface. Next, a custom router jig is used to turn down a foam core to the
appropriate shape according to the template exported from SolidWorks. In this way we
are left with a tool, which will allow for the precise fabrication of multiple faring.
3.3.5 PLANNED COMPONENT, FUNCTIONAL, AND STATIC TESTING
The team‟s first priority will be to perform qualification testing on the structural
components of the rocket. The tests to be performed are as follows and will be
completed after the structural test article is completed post-CDR:
The body tube will be tested using a crush test in the axial direction and bending
test in the lateral direction. It will be tested with a variable mass, such as sand, to
determine the stiffness and failure force.
A crush test will also be performed between two tubes to verify the strength of
the tube coupler.
The bulkheads and their attachment to the body tube will be tested with a pull
test, in which the tube will be fixed and variable mass will be used to determine
pullout force.
The fins will also be tested using a series of pull/push tests (also using a variable
mass and gravity) in order to test the fin strength in each of the 3 orthogonal
directions.
In addition to structural testing, several deployment and recovery tests will need to be
performed. These tests will be performed after the UAV prototype is completed post-
CDR.
Deployment altitude will be verified using barometric testing. The team has
constructed a small vacuum chamber, which is capable of roughly simulating
ambient pressure. As a result, the avionics package will be placed into the
vacuum chamber to ensure that it sends charge ignition commands at the right
times.
In order to verify the failure force of the shear pins, a representative tube will be
used with a representative nose cone, with the open side of the tube covered.
The shear pins are mounted into the relevant brackets in flight orientation. The
black powder charge will be ignited at the closed end to validate the mass of
black powder to be used.
41
UAV deployment will also require testing, which can be performed in a couple of
phases: (1) the force of the drogue parachute on the sabot can be simulated to
ensure that the sabot separates from the tube and the UAV deploys and (2)
integrated deployment tests from a balloon platform. This test will be described
further in Section 4.3.4.
A series of avionics tests will also be performed. A summary of the tests is provided
below. Greater detail can be found in Section 4.3.4. These tests will be performed prior
to the FRR.
The emergency locator beacons (transmitters and receiver) operation will be
checked, by searching for the beacons in a representative location.
Each computer will also be checked to see if they downlink properly to the
ground station. This will be performed on the ground in a field and then on a
balloon platform using a representative ground station and rocket.
Finally, these tests will culminate in a representative scaled test launch, which will verify
functionality of all systems, including the UAV.
3.3.6 STATUS AND PLAN FOR REMAINING MANUFACTURE AND ASSEMBLY
To produce components of a high caliber the MIT Rocket Team has decided to use a
four-stage fabrication and assembly process. After the finalization of the designs, the
team moved straight into a fabrication-testing period where possible manufacturing
methods were evaluated for their feasibility. In this period the team decided on the
appropriate technique and materials needed for the fabrication of each component of
the flight vehicle. In this stage, fabrication methods were tested on representative
components sized to the approximate dimensions of the specified design. The testing
phase of production has been completed as of mid-January and the team has now
moved into the prototyping phase for the flight vehicle.
In the prototype phase, full-scale components will be constructed using the methods
determined during the testing phase of development. The resulting components will
then be assembled into a full-scale prototype with extra components being produced for
destructive testing methods. When the components are all fully tested, the any
components needing design changes will be refabricated to the new specifications and
a proto-flight model will be constructed.
The purpose of the proto-flight model is to allow for full-scale flight-testing procedures
with, and without the completed payload. It is expected that the proto-flight model will be
completed in early February to allow for multiple launch attempts before the Flight
Readiness Review. Upon successful flight testing of the proto-flight vehicle, any
necessary design changes and repairs will be made to the airframe for the flight model
to be launched in Huntsville AL. Furthermore spare components will be manufactured to
the specifications of the flight model to mitigate the loss of components in transit to AL.
42
As stated above, the team is currently in the prototype phase of fabrication on track for
completion in early February. The team is has purchased all materials necessary to
construct a complete vehicle as well as a set of spare components. Construction is
currently underway and the components will be tested as listed in section 3.1.6.
3.3.7 INTEGRITY OF DESIGN
[See section 3.3.1]
3.3.7.1 SHAPE AND FIN STYLE
[See section 3.3.1]
3.3.7.2 PROPER USE OF MATERIALS FOR: FINS, BULKHEADS,
STRUCTURE
[See section 3.3.1]
3.3.7.3 PROCEDURES: ASSEMBLY, ATTACHMENT,
ALIGNMENT
[See section 3.3.1]
3.3.7.4 MOTOR MOUNTING AND RETENTION
The motor mount will consist of a commercial 75mm motor tube and laser-cut, plywood
centering rings. There will be four centering rings in total, one located at each end of the
motor tube and two in the middle. The farthest forward will be made from 1/2” plywood.
The farthest aft centering ring will be made from two rings of 3/16” plywood sandwiched
together; the OD of the forward ring will be the ID of the body tube, and the OD of the
aft ring will be the OD of the body tube. This will transfer some of the thrust load through
compression of the aft centering ring, rather than through shear in the epoxy joints
holding the motor mount in the body tube. The middle centering rings will be made from
½” plywood, with four slots to accept the fin tabs. The fin tabs will also be slotted at the
centering ring locations to allow the fins to contact the motor tube for additional support.
One will be located near the forward edge of the fin tabs and the other near the aft edge
of the fin tabs, close to the aft-most centering ring. Plywood is chosen because it is
relatively cheap, strong, light, and able to withstand the high temperatures of the motor
casing without deforming.
Motor retention will be accomplished as follows. Two 8-32 T-nuts will be mounted to the
aft-most centering ring, 180° apart at a radius of roughly 2.375”. 1.5” 8-32 screws will go
through two small clearance holes in the motor retention plate and screw into the T-nuts
to hold the plate in place. The motor retention plate will be a piece of 1/32” steel sheet
that has a hole cut in it; this hole will be made large enough for the motor‟s nozzle to fit
through, but small enough to keep the motor casing from falling out of the motor tube.
43
There is a thrust ring on our 75mm hardware that prevents the motor casing from
moving forward during burn.
3.3.7.5 STATUS OF VERIFICATION
[See section 3.1.6]
3.4 RECOVERY SYSTEM
3.4.1 PARACHUTE CHOICE AND TESTING
When the drogue parachute is deployed at apogee, it will need to support a total system
mass of 23 kg. A 4ft diameter parachute will be used to achieve a descent rate of 75
ft/s.
Once an altitude of 2500 ft AGL is reached, the tether securing the sabot inside the
rocket will release, allowing the drogue parachute to pull the sabot and the main
parachute out of the rocket. At this point, the rocket body will separate from the
sabot/nose/drogue section and free fall as the main parachute deploys. This will allow
for a considerable gap between the rocket body and the sabot, decreasing the risk of
the deployed UAV colliding with the rocket or becoming entangled in the main
parachute.
With the UAV deployed and the sabot separated from the rocket body, the remaining
structure has a mass of 15 kg. With a 14ft diameter parachute, a final descent rate of 19
ft/s can be achieved. Under the 4ft parachute, the nose cone and sabot will have a final
descent rate of 26 ft/s.
TABLE 3-13: PARACHUTE DESCENT RATES
Final Descent Rate
System Under Drogue 75 ft/s
Nose/Sabot Final Descent
Rate 26 ft/s
Rocket Body Under Main 19 ft/s
The drogue parachute and nose cone are directly connected to the sabot. This
assembly is initially connected to the recovery system bulkhead via the explosive tether.
The main parachute is also secured directly to the recovery system bulkhead (not by the
tether). Its deployment is constrained by the sabot.
The calculations for the amount of black powder required to successfully separate the
nose cone from the body tube can be found below.
44
The charge release mechanism will contain 0.2 grams of black powder. This number is
recommended by the manufacturer.1
The drogue deployment charge must provide ample force to break the shear pins,
accelerate the nose cone away from the rocket body, and accelerate the drogue
parachute out of the nose cone. Four #2-56 nylon screws (MMC 94735A177) will be
used as shear pins to retain the nose cone. Nylon 6/6 has a shear strength of 10ksi.2
With this, the maximum shear force can then be calculated by the following equation:
,
where A is the cross-sectional area of the bolt, and τ is the shear strength. For a #2-56
screw, the minimum pitch diameter is 0.0717 in.3 This leads to a shear force of 40 lbf.
With four pins, the charge will have to provide a minimum force of 120 lbf. Adding 25%
margin, the charge will need to provide a total force of 150 lbf. This leads to a required
black powder mass of 2.1 g.4
3.4.1.1 TEST RESULTS FROM EJECTION CHARGES AND
ELECTRONICS
3.5 SAFETY AND FAILURE ANALYSIS
TABLE 3-14: POTENTIAL ROCKET FAILURE MODES
Risk Likelihood Effect on Project Risk Reduction Plan
Catastrophe at
Take-Off Low Total mission failure
To mitigate this risk, we
have detailed setup,
integration, and launch
procedures. We will
conduct safety checks
at every stage to
ensure adherence to all
safety guidelines.
Structural
failure Low Total mission failure
Large safety factors
accounted for during
the design process
reduce the impact that
launch loads will have
on weaker structural
areas
1 Tender Descender User‟s Guide, http://fruitychutes.com/Recovery_Tether_manual.pdf
2http://www.aptllc.net/datasheets/Nylon66.pdf
3http://www.engineersedge.com/screw_threads_chart.htm
4Black Powder Pressure-Force Calculator: http://www.info-central.org/files/303-
Pressure_Force_Calculator_Ver2.xls
45
Birdstrike Low Flight path altered
Follow all NAR launch
rules, holding launch if
any wildlife overhead.
Lack of failure
of shear pins Low
No parachute
deployment;
catastrophic failure
Extensive deployment
testing will be
conducted to validate
the amount of black
powder being used for
deployment is sufficient
to break pins.
Sabot not
deploying Medium
Payload not
deployed; main
parachute not
deployed
Extensive testing. Wing
release locking
mechanism will keep
wings locked until sabot
exits body tube. This
will prevent premature
opening of the sabot,
decreasing the
possibility of the sabot
binding inside the
rocket body.
Drogue
parachute not
deploying
Low
No force available to
pull sabot and main
parachute from
rocket body;
catastrophic failure
Extensive deployment
testing will be
conducted to find
optimal packing method
for drogue parachute.
Entanglement
of main
parachute
Medium
Partial mission
failure. Payload
deployment still
viable. Recovery of
main rocket body
unlikely
Parachute will be
properly packed.
Failure of
recovery
system
attachment
point
Medium
Partial mission
failure. Payload
deployment still
viable. Recovery of
main rocket body
unlikely
Ensure extensive
testing of recovery
system attachment
points to ensure their
ability to meet strength
requirements
Sabot fails to
separate after
ejection from
rocket
Low
UAV unable to
deploy; mission
failure
Extensive testing to
ensure wing rotator
locking mechanism
disengages after sabot
exits rocket body, and
46
that spring force of
deploying wings is
sufficient to separate
sabot halves.
Motor fails to
ignite Low
Unable to launch;
mission failure Replace igniter
3.6 MISSION PERFORMANCE PREDICTIONS
3.6.1 MISSION PERFORMANCE CRITERIA
In order for this mission to be considered a success, the following events must occur:
Achieve an altitude as close to 5280 feet (1 mile) as possible. (It is preferable to
undershoot the target, as the flight score penalty for overshooting is twice as
great.)
Eject nose cone and deploy drogue parachute at apogee
Deploy UAV and main parachute at an altitude of 2500 ft
The UAV must unfold its wings and start the primary science mission objective.
Land safely (intact and reusable with no necessary repairs) on the ground.
3.6.2 FLIGHT PROFILE SIMULATION
For the Preliminary Design Review flight profile simulations, RockSim was used. A
model of the rocket was built in RockSim, and the mass was verified against the
Solidworks model. Parachute descent rates were verified against the MATLAB
parachute sizing model. The RockSim model agreed with the Solidworks model mass to
within 0.1 pounds and with the MATLAB model descent rates to within 3 feet per
second. Figure 3-19 shows the RockSim model.
FIGURE 3-26: PREDICTED CM AND CP LOCATIONS
A battery of simulations was run, taking into account the approximate location and
altitude of the launch site and average temperature, pressure, and humidity conditions.
It was known that the Cesaroni L1115 would be more powerful than necessary and
propel the rocket higher than the target altitude. With no added ballast or winds, the
47
rocket flew over 800 feet above the target altitude. This was expected and desired,
especially considering the mass margin of the payload and other components, the
masses of which have only been measured up to this point. Initially, the RockSim model
had a mass of 19.75 kg and an initial stability margin of 2.26, which is comfortably
stable but makes the rocket susceptible to angling toward gusting winds.
Groups of ten simulations were run to find an optimal mass of the ballast that needed to
be added. Each simulation had variable light winds (3-7 mph), and the ballast mass and
launch rail angle were varied until the desired apogee and landing range distance were
achieved. The optimal value for the ballast weight added to the bottom motor bulkhead
mount is 3.65kg, giving the rocket a total wet mass of 23.4 kg, and the optimal launch
angle is 2 degrees. This gives an average altitude over 10 simulations of 5275 feet
(maximum 5296 feet, minimum 5243 feet) and a distance at landing of no more than
500 feet from the launch location, with an average distance of 100 feet.
At t = 0, the Cesaroni L1115 is ignited. Burnout occurs at 4.49s, and apogee occurs at
approximately 19.4 seconds. At this time, the first charge is ignited to eject the
nosecone and deploy the drogue chute, which pulls the sabot out of the rocket. At an
altitude of 2500 feet, the second charge is ignited. This charge releases the UAV from
the sabot, separates the nosecone, drogue chute, and sabot from the rest of the rocket
body tube, and deploys the main parachute.
Figure 3-20 shows the acceleration and velocity of the rocket during the first 30 seconds
of flight (the remaining flight time was omitted for clarity). The maximum speed occurs
near burnout, and does not exceed Mach 0.5. The maximum predicted acceleration
occurs at the parachute deployment, as expected. While the magnitude of the maximum
acceleration is high compared to what was expected, this is still within the range that the
carbon fiber structure of the rocket can stand. An initial concern was that the parachute
cords could rip the body tube apart during high-speed deployment. Future modeling will
try to reconcile the nearly instantaneous parachute deployment featured in RockSim
and the expected unraveling time of the chute to prevent such high accelerations in
simulations.
48
FIGURE 3-27: PREDICTED ACCELERATION AND VELOCITY PROFILES
Figure 3-27 shows the simulated altitude profile of the rocket. Burnout and apogee are
shown with red and blue dotted lines, respectively, and the main parachute deployment
can be seen as the kink in the altitude line near 50 s.
49
FIGURE 3-28: SIMULATED ALTITUDE PROFILE
Future flight profile modeling will more accurately define the launch conditions, including
launch pad altitude, predicted weather conditions (relative humidity, average wind
speed, etc.), and competition settings. Immediately before the flight, these conditions
will be taken into account and the mass of the ballast will be adjusted according to on-
site simulations to achieve the predicted altitude given the very best initial conditions
simulations the team can generate.
The scale motor used is the Cesaroni K1085, which was chosen to subject the airframe to full-scale launch conditions and deliver the UAV to a deployable altitude. A simulation of these results is as shown in 3-29.
50
FIGURE 3-29: SCALED LAUNCH SIMULATED RESULTS
3.6.3 SCALE MODEL TEST
A 1/2 scale test rocket was constructed and launched to test the aerodynamic
properties of our design. See Figure 3-30 for a picture of the scale rocket. The scale
rocket was not scaled in weight because doing so would have required a much larger
and more expensive -class motor. The goal of the scale test launch was to test rocket
stability, and it was determined that this goal could be achieved using a smaller motor.
FIGURE 3-30: SCALE MODEL ROCKET
An Aerotech G80 was originally picked for the test motor. However, the scale rocket‟s
final mass was larger than anticipated, necessitating the use of a larger motor. An
Aerotech H128 was used as the test motor due to its higher average and initial thrust. A
summary of the specifications of the ½ scale rocket can be found in Table 3-15.
TABLE 3-15: SCALE MODEL DIMENSIONS
Specification Value
51
Diameter 3 in
Length 62 in
Mass (without motor) 4.1 lbs
CG (measured, with motor) 41 in
CP (Rocksim prediction) 45.8 in
Stability Margin 1.6
By adding mass, the CG of the scale test rocket was placed at the scaled location of the
predicted CG of the full scale rocket. This resulted in a very similar stability margin to
that predicted by Rocksim for the full scale rocket.
The scale rocket was outfitted with a PerfectFlite MAWD altimeter for altitude
measurement. The rocket flew straight and stable. The MAWD recorded a final altitude
of 778 ft on the H128, only 50 ft higher than Rocksim predicted, thus validating our
Rocksim model.
3.6.4 STABILITY
The initial static margin of the rocket with all the ballast placed at the bottom of the
bottom motor bulkhead is 1.17. This is an appropriate static margin that makes the
rocket less susceptible to wind gusts during flight that would cause an overstable rocket
to tilt into the wind, but not so close to unstable that unexpected changes in the masses
of some of the components would jeopardize the overall stability. During flight, the static
margin will increase as propellant is burned and the center of mass moves toward the
nose of the rocket. The static margin at burnout is 1.64.
If 0.8 kg of ballast mass is moved from the bottom bulkhead to the sabot, then the initial static margin is 1.56. At burnout, the static margin is 2.08. This does not significantly change the maximum altitude of the rocket.
3.7 PAYLOAD INTEGRATION PLAN
3.7.1 INSTALLATION AND REMOVAL, DIMENSIONS, PRECISION FIT
1) Integrate Avionics Bay a) Integrate avionics boards and ELT onto avionics plate b) Integrate 3 New Batteries c) Test electronics (turn on) d) Attach avionics plate onto top cap with L-brackets and 6-32 bolts, which are
inserted, from the outside, through nutplates on the inside e) Attach avionics plate onto bottom cap with L-brackets and 6-32 bolts, which are
inserted from the outside through nutplates on the insides f) Slide assembly into tube
52
g) Slide recovery system bulkhead into rocket and secure with screws h) Check all connections i) Check pressure holes
2) Make Black Powder Ejection Charge a) Safety Officer will oversee this step b) Connect to avionics
3) Integrate antenna a) Antenna is pre-attached to main body b) Connect the antenna (14 gauge insulated copper wire) to the avionics using a
binding post connection c) Slide sabot (with UAV) in, while routing the wires from the avionics assembly
through the raceway
Figure 3-22: Avionics assembly on plate inside tube, attached with L-brackets
53
FIGURE 3-31: TUBE-TUBE INTERFACE
4) Recovery a) If not pre folded, fold drogue parachute b) Integrate drogue parachute and parachute protector
i) Attach to upper sabot hard point with quick link c) If not already properly packed in parachute bag, fold main parachute d) Integrate payload bulkhead
i) Attach main parachute bag to lower sabot hard point ii) Attach main parachute to shock cord with quick link iii) With the upper body segment still detached from the lower body segment,
use quick links to attach one end of the main shock cord to the payload eyebolt via the charge release locking mechanism, and the other end to the sabot eyebolt. Tie the cord off using a bowline knot.
5) Nose Cone a) Attach secondary shock cord between nose cone and sabot with quick links b) Install ELT c) Nose Cone, Drogue, and Sabot should not be attached to main shock cord
6) Integrate rocket body with sabot/UAV assembly a) Connect leads to avionics b) Attach two rocket body segments together (Fig.3)
i) Thread four 6-32 ½‟‟ screws through doublers ii) Reach through the 5‟‟ by 5‟‟ access door and untie the end of the shock cord
connected to the payload eyebolt iii) Tighten the cord and re-knot (using a bowline knot). iv) Cut off extra length of cord. v) Close access door by bolting plastic piano hinge shut using 3 6-32‟‟ bolts
54
FIGURE 3-32: MAIN PARACHUTE/SHOCK CORD ATTACHED TO EYE BOLT AND RECOVERY
SYSTEM BULKHEAD
7) Integrate Nose Cone a) Slide into upper body tube (Fig.4) and mount using 4 nylon bolts fastened to the
inside using Helicoils, which will act as shear pins
55
FIGURE 3-33: NOSE CONE/UPPER BODY TUBE INTERFACE
FIGURE 3-34: INTEGRATED AVIONICS ASSEMBLY, MAIN PARACHUTE, SABOT AND UAV
ASSEMBLY
3.7.2 TASKING & INTEGRATION SCHEDULE
TABLE 3-16: TASKING AND INTEGRATION SCHEDULE
Overall Task Number of People* Time
Integrate avionics
assembly
3 15 minutes
Assemble UAV 2 5 minutes
Integrate main parachute 2 15 minutes
Integrate UAV assembly
with recovery system
3 5 minutes
Integrate drogue
parachute
2 5 minutes
Integrate nose cone 2 2 minutes
Integrate motor 2 4 minutes
Total time: Approximately 60 minutes
*This includes one person with the checklist who will be supervising
3.7.3 COMPATIBILITY OF ELEMENTS
[See section or 3.7.1]
56
3.7.4 SIMPLICITY OF INTEGRATION PROCEDURE
[See section 3.7.1]
3.8 LAUNCH OPERATION PROCEDURES
Caution Statement
Recall the Hazards Recognition Briefing. Always wear proper clothing and safety gear.
Always review procedures and relevant MSDS before commencing potentially
hazardous work. Always ask a knowledgeable member of the team if unsure about
equipment, tools, procedures, material handling, and/or other concerns. Be cognizant of
your and others’ actions. Keep work station as clutter-free as possible.
Equipment Packing Checklist:
1. Support Equipment and Tools
a. Safety Gear
i. Goggles
ii. Rubber Gloves
iii. Leather/Work Gloves
iv. Face Masks
v. All Safety Documents and References
b. Furniture
i. Tent (1x)
ii. Tables (2x)
iii. Chairs (6x)
iv. Rocket assembly benches
c. Generator
i. Gas
ii. Power Strip(s) (3x)
iii. Extension Cord(s) (3x)
d. Tools
i. Corded Drill
ii. Cordless Drill
1. Cordless Drill Batteries
2. Charger
iii. Drill Bit Index(s)
iv. Wrench Set
v. Pliers
vi. Screwdriver Set
vii. Hex Keys Set
57
viii. Files
ix. Sandpaper
x. Knives
xi. Flashlight
xii. Soldering Iron
1. Solder
2. Solder Wick
3. Sponge
xiii. Wire Cutter/Stripper(s)
xiv. Extra Wire (Black and Red)
xv. Pocket Scale
e. Adhesive
i. 5-minute Epoxy (2 part)
ii. CA and Accelerant
iii. Aeropoxy (2 part)
iv. Epoxy Mixing Cups
v. Popsicle Sticks
vi. Foam (2-part)
vii. Foam (solid)
f. Other supplies
i. Tape
1. Duct Tape
2. Scotch Tape
3. Vacuum Tape
4. Electrical Tape
5. Masking Tape
6. Gaffer‟s Tape
ii. Trash Bags
iii. UAV Camera Port Cleaner
iv. Isopropyl Alcohol (general clean up)
v. Water Bottle
vi. Camera Lens Cleaning Supplies
vii. Paper Towels
viii. Wipes
ix. Spare Hardware
x. Lithium/Silicon Grease (for building reload; other)
xi. Zip-ties
xii. Talcum Powder (for parachutes)
2. Ground Station
a. Antennas
58
i. Rocket (1)
ii. UAV (3)
iii. Antenna Mounts
b. Emergency Locator Transponder (ELT) (UAV and Rocket) (3x)
c. Emergency Locator Receiver
d. UAV Main “Pilot” Computer
e. UAV Secondary Computer
f. Rocket Ground Station Computer
g. UAV Manual R/C Controller
h. Binoculars
i. Monitors
j. Power Adapters for all Computers
k. Mice (3x)
l. Cables
i. Antennas
ii. Monitors
iii. Other
m. Miniature Weather Station (wind speed/direction, temperature)
3. Launching Equipment
a. Launch Pad
b. Launch Rail
c. Stakes for Pad
d. Angle Measuring Tool
e. Electronic Launch System (ELS)
i. Battery
ii. Battery Charger
iii. Controller
iv. Leads
4. Rocket
a. Body
i. Lower Tube Section
ii. Upper Tube Section
iii. Nose Cone
iv. Ballast
v. Shear Pins (10x)
b. Recovery
i. Parachutes
1. Drogue (2x)
2. Main (2x)
3. Nomex Parachute Protectors (3x)
59
ii. Shock Cord
iii. Ejection Charges
1. Black Powder
2. Charge Holders (4x)
3. Igniters (4x)
iv. Charge Released Locking Mechanism (2x)
v. Quick links (10x)
c. Motor
i. Casing
ii. Reload (2x)
iii. Retention
1. Retention Plate
2. Retention Hardware
d. Avionics
i. Avionics Bay
ii. Altimeters
1. ARTS2 (1x)
2. ARTS2 Transmitter Board (1x)
3. MAWD (1x)
iii. Antenna (attached to outside of rocket body)
iv. 9V Batteries (10x)
v. ELTs (one in Bay, one in nose cone) (3x)
vi. Hardware
1. 4-40x1” bolts (10x)
2. 4-40 locknuts (6x)
5. UAV
a. UAV
b. Motor (2x)
c. UAV Propeller (3x)
d. UAV Lithium Polymer Batteries (2x) and Spare Batteries (3x)
e. Lithium Polymer Battery Charger/Balancer
f. Spare Servos (3x)
g. Spare Control Linkages
h. Sabot
i. Avionics
i. Flight Computer
ii. Back up Sensor Logging Board
iii. Sensors
iv. Flight Digital Still Camera
v. Video Board and Video Camera
60
vi. Manual Control Receiver (Back Up: 72MHz)
vii. Antennas (72MHz, 900MHz, 2.4GHz)
viii. ELT
6. Miscellaneous
a. Digital Camera
b. Video Camera
c. Extra Batteries
d. Two-Way Radios
e. Two-Way Radio Chargers
f. Manuals for all Equipment and Gear
Pre-Flight/Final Assembly Checklist:
1. Ground Station
a. Furniture Set Up
b. Generator
i. Full Tank
ii. Extra Gas
iii. Connect Extension Cord(s)/Power Strip(s)
c. Computers
i. Set Up
ii. Plug in Power Adapters
iii. Mice
iv. Set Up Monitors
v. Power Up
d. Antennas
i. Mount and Set Up
1. 2.4GHz
2. 900MHz
ii. Connect to Computers
e. Set Up ELT Receivers
i. Test on each of 3 channels
2. UAV
a. Mechanical
i. Inspect Fuselage (follow detailed checklist)
1. Internal Structure
2. External Structure
3. All Electronics/Avionics Mounts
4. Motor Mounted Securely
5. Kevlar Skid Plate
61
ii. Inspect Wing and Wing Folding Mechanism
iii. Test Wing Folding Mechanism
1. Fold and let Unfold at least twice
2. Adjust as necessary
iv. Inspect all Hinges
v. Test All Folding Hinges
1. Fold and let Unfold
2. Adjust as necessary
vi. Unfold Everything
vii. Inspect All Control Surfaces
1. All should be free and clear to rotate
2. Inspect and Move All Hinges
3. Inspect Control Linkages and Servos
viii. Inspect Camera Dome
1. Clean Dome if necessary
2. Check Connection to Fuselage
3. Check Camera Mount
ix. Inspect UV Sensor Window
1. Clean if necessary
b. Power Systems
i. Inspect Motor
ii. Check if Propeller Secure
iii. Give Motor a Test Spin (by hand)
iv. Inspect Motor Controller
v. Make sure all electronics are Switched Off
vi. Connect and Secure Charged Lithium Polymer Batteries
c. Avionics
i. Install Flight Computer
ii. Install Back up Sensor Logging Board
iii. Install Video Board and Video Camera
iv. Install Digital Camera
v. Install Manual Control 72MHz Receiver (Back Up)
vi. Inspect All Sensors
vii. Install ELT
viii. Connect Everything
ix. Set No-Fly Zones
x. Set Loiter-mode Landing Location
d. Communication/Controls
i. All servos connected to proper channels
ii. All Avionics Connected
62
iii. Power On
iv. Test All Control Surfaces (using standard/manual R/C 72MHz
transmitter)
1. Trim
2. Actuate one direction
3. Actuate other direction
v. Test Motor (using standard/manual R/C 72MHz transmitter)
1. Clear objects/people from the plane of the propeller
2. Throttle Up
3. Throttle Down
vi. Power Motor/Motor Controller Off
vii. Test Flight Computer
1. Communicating with Ground Station
viii. Test Data Feeds (turn UAV avionics on)
1. Temperature
2. Humidity
3. Solar Irradiation
4. UV Irradiation
5. Pressure
ix. Test IMU/GPS
1. Transmitting Telemetry
x. Test Autopilot (Make sure control surfaces respond correctly)
1. Pitch UAV Up
2. Pitch UAV Down
3. Yaw UAV Right
4. Yaw UAV Left
5. Roll UAV Left
6. Roll UAV Right
xi. Test Data Logging
1. Digital Camera Still Shot Recorder
2. Back Up Sensor Data Logging
xii. Test Video Feed
1. Receiving Video
xiii. Test ELT
1. Receiving ELT signal
xiv. Power Up Motor/Motor Controller
xv. Flight Test with Manual R/C Control (no autopilot)
1. Receiving All Data
2. Proper Control Responses
xvi. Ground Test of Point-and-Click Control (with autopilot)
63
1. Receiving All Data
2. Proper Control Responses
xvii. Aerial Test of Point-and-Click Control
1. Trim control surfaces before flight
2. Back Up with Manual R/C Control
e. Switch out Lithium Polymer Batteries
f. Final Overall Inspection
g. Install UAV into Sabot. See Payload Integration Plan.
3. Rocket
a. Lay-out rocket sections in order
b. Check Body Antenna
c. Install Ballast into appropriate sections of sabot and body tube
d. Refer to Payload Integration Plan
i. Follow, then continue with this checklist
e. Install all shear pins
f. Prepare Motor Reload
i. Safety Officer will oversee this step
g. Slide motor casing into rocket
h. Screw on motor retention
i. Make sure the tube-tube and tube-nose cone interfaces are secure
j. Inspect rail guides
k. Do a pre-launch briefing
Launch Checklist:
1. Get approval from event administration to set up pad, ELS, and rocket
2. Set up pad
3. Tip pad over and install rail
4. Check all tube interfaces
5. Slide rocket onto rail down to stop
6. Tip up launch pad
7. Stake pad to ground
8. Arm Electronics
a. Have manuals on-hand
b. Listen for proper beeps
9. Put igniter into motor and secure it
10. Connect launch clips
11. Connect ELS to battery
12. Clear launch area/back up appropriate distance
13. Make sure Ground Station and Pilots are ready
64
14. Get approval from event administration for launch
The following depend on procedures outlined by event administration:
15. Cameraman ready
16. Check to see if range and skies are clear
17. Insert key into ELS check continuity
18. Countdown from 5
19. Launch
20. Remove key from ELS
21. Disconnect ELS from battery
22. Recover Rocket and UAV
3.8.1 RECOVERY PREPARATION
Using a short length of nylon webbing, attach the inside of the deployment bag to the
loop at the top of the main parachute.
Secure the top of the deployment bag to a table leg or other hard point. Stretch out the
parachute and untangle the lines. Place a weight on the outstretched, untangled lines,
to hold them in place. Next, flake out the canopy.
65
Next, fold the canopy width-wise so it can fit inside the deployment bag.
Fold the leader connecting the deployment bag and the parachute in a “figure 8” and
secure with a rubber band. Place inside deployment bag. Begin placing the canopy
inside the bag, folding it over itself in an S pattern.
66
67
After the parachute is in the bag, begin folding the shroud lines, again, in an S-pattern.
Tuck the folded lines into the bottom of the deployment bag. Secure the Velcro flap of
the bag.
3.8.2 MOTOR PREPARATION
One of the team‟s L2 members will supervise motor assembly. All fire hazards, e.g.
people smoking, lighters, potential ignition sources, will be removed from the immediate
surroundings during motor preparation.
See Appendix 6 for official and detailed Pro75 motor preparation instructions.
The assembled motor will be slid into the motor tube and motor retention will be
screwed on.
3.8.3 IGNITER INSTALLATION
Once the rocket is on the pad, tipped vertical, and all electronics are armed, the motor
igniter will be installed. Care will be taken to fully insert the igniter into the motor. The
igniter will be held in with tape and a 1/8” dowel, which will be easily pushed out when
the motor lights. Launch lead clips will be securely attached to the igniter leads at the
appropriate time.
68
3.8.4 SETUP ON LAUNCHER
Refer to Section 3.8 for Launch Checklist. After preparation is completed, the rocket will
be carefully carried out to the launch pad. The launch pad will be oriented such that the
rocket sits on the downwind side of the rail to avoid torque on the rail-buttons. The
launch rail will be tipped over to allow the rocket to be slid on horizontally. Care will be
taken when sliding the rail buttons into the rail slot. Then the rocket and rail will be
carefully tipped up to vertical.
3.8.5 TROUBLESHOOTING
Electronics will be disarmed any time the rocket is approached while on the pad. Table
3-17 summarizes possible problems and solutions with the rocket while it is on the pad.
TABLE 3-17: POSSIBLE LAUNCH FAILURE MODES
Problem Possible Causes Possible Solutions
Launch Button suppressed,
but rocket does not launch.
1. Lead acid launch
battery is dead.
2. Igniter lights, but motor
does not.
3. Igniter does not light,
but lead acid launch
battery is charged.
1. Change out battery.
2. Change out igniter.
3. Check all wire
connections.
Rocket electronics do not
arm when switched on.
1. The electronics‟
batteries were not
properly installed or
connected.
2. The electronics‟
batteries are dead.
1. Take rocket off of pad,
2. Take rocket off of pad and
replace electronics‟
batteries.
Avionics fail to report full
continuity
1. Charges are likely not
properly connected
1. Restart electronics to see
if it fixes the issue. If not,
remove rocket from pad and
troubleshoot situation at
prep table
Rail-button pulls out upon
sliding rocket onto pad.
1. Not enough care was
taken when installing the
rocket on the pad.
1. Scrub launch attempt,
bring rocket back to work
station, and repair.
3.8.6 POST FLIGHT INSPECTION
Given the nature of the USLI Launch field, it is unlikely launch organizers will allow more
than 2 people to recover the rocket and UAV. Given the large size of the vehicle, it will
69
be important to plan ahead to be able to carry it back. A small duffle bag will be carried
by one team member to pack the parachute and other items into after recovery.
Before reaching the rocket and UAV, pictures will be taken for future reference before
disturbing any part of it
Upon approaching the rocket, the two ejection charges will be carefully inspected to
ensure they are no longer live. If one is still live, its wire leads will be cut with a pliers
and it will be placed in a safe location (not to be carried directly by a person). The main
parachute will be disconnected from the shock cord and it will be rolled up and placed in
the bag. The rocket and nosecone section will be repacked for transport. The rocket will
be carried back in once piece by both people. The UAV will be retrieved either
concurrently by the same team if it is determined that one group of people can carry it
back with the rocket, or separately by a different team. Similar photo documentation
techniques will be used.
In the event of a contingency during recovery, plans will be made to adjust to these
situations. Pyrotechnics safety will be a top priority during any situation.
Upon return to the prep area, data will be collected from data logging devices and the
rocket will be prepared for the trip back.
3.9 VEHICLE SAFETY
3.9.1 IDENTIFICATION OF SAFETY OFFICERS
Andrew Wimmer will be the primary rocket safety officer for the team. Ben Corbin is the
team‟s MIT EHS representative and is the assistant safety officer and is in
charge of safety issues not directly related to the rocket. Both team members have
considerable experience in their respective areas.
3.9.2 ANALYSIS OF FAILURE MODES AND MITIGATIONS
The following table provides an updated analysis of the failure modes of the proposed
vehicle design, integration and launch operations.
TABLE 3-18: POTENTIAL FAILURE MODES
Failure Mode Effects Precautions to prevent
result
Precautions to
prevent event
Motor Failure Property Damage,
Injury
Stand up, follow path
of rocket visually,
move if needed.
Follow proper launch
Store and assemble
motor in accordance
with manufacturer‟s
70
safety distances instructions
Recovery
System
Entanglement
Property Damage,
Injury
Follow rocket‟s
descent path visually,
move if needed
Design and rigorously
test recovery system
in accordance with
accepted HPR
standards
Recovery
System
Structural
Failure
(bulkheads,
shockcords, etc)
Property Damage,
Injury
Follow rocket‟s
descent path visually,
move if needed
Perform pull tests on
unrated components
to ensure their
strength. Components
to be tested to 50g
shock loads
Recovery
System failure to
deploy
Property Damage,
Injury
Follow rocket‟s
descent path visually,
move if needed
Ensure rigorous
testing of black
powder charges,
Tether release
mechanisms and
deployment altimeters
and power supplies.
Don‟t forget to arm
altimeters
Recovery Device
deployment on
ground
Property Damage,
Injury (especially
eye)
Avoid placing body in
path of parts if
electronics are armed.
Wear safety glasses if
necessary.
Shunt charges until
they are attached to
recovery electronics.
Do not move the
rocket with armed
electronics.
Unstable Vehicle Property Damage,
Injury
Stand up, follow
rocket‟s path visually,
move if needed.
Confirm vehicle
stability before launch.
Ensure actual CG
position is acceptable
relative to calculated
CP
Brush Fire Fire damage, injury
Have fire protection
equipment and
personnel trained in its
use onsite
Follow NFPA table for
dry brush around pad
area.
Mid-flight vehicle
destruction
Loss of vehicle,
Injury, Property
Follow rocket‟s path
visually and move if
Design, construct and
test vehicle to assure
71
3.9.3 POTENTIAL HAZARDS
A listing of personnel hazards and evidence of understanding of safety hazards
is provided in the sections below.
(excessive
forces on
vehicle)
damage needed if vehicle does
come apart
successful flight. Use
standard construction
procedures for LII-LIII
rockets, including
sufficient bulkheads,
fins, motor retention
and couplers.
Failure of UAV
to deploy
Loss of science
value, potential
failure of main
recovery device
Visually track vehicle
& UAV, and move if
needed
Rigorously test UAV
deployment method
as an integrated
component in the
rocket recovery
system. Ensure all
other aspects of the
rocket flight succeed
Failure to
successfully
integrate vehicle
in allotted 4 hour
time period
Loss of launch
opportunity N/A
Practice integration
techniques under time
constrains to ensure
they are achievable in
allotted time.
Failure of vehicle
to reach desired
altitude
Loss of competition
points and potential
loss of science value
N/A
Use multiple
simulation programs
and data from actual
flight tests to fine tune
rocket mass and
motor selection
UAV Scientific
data is
unrecoverable
Loss of science value N/A
Perform range
communication tests
with all flight hardware
in flight configuration.
Allow electronics to
be quickly
reconfigurable in case
of frequency conflicts.
72
Safety Checklist
In order to assure a safe and successful flight, a checklist must be followed during prep
activities and launch. In order to reduce personnel hazards during the prep of the
vehicle before taking it to the pad, the following precautions must also be taken into
consideration.
Always wear safety glasses when dealing with rocket parts containing small
hardware or pyrotechnic charges.
Never look down a tube with live pyrotechnic charges in it.
Always point rocket and pyrotechnic charges away from body and other people
Avoid carrying devices that have live electrical contacts (radios, cell phones, etc.)
while prepping live pyrotechnic charges.
Never arm electronics when rocket isn‟t on pad unless the area has been cleared
and everyone knows that pyrotechnic continuity checks are being done.
Always follow the NAR/TRA safety codes.
Always follow all applicable local, state and national laws and regulations
Do not allow smoking or open flames within 25 feet of the motor or pyrotechnics.
Respect the launch organizer‟s decisions regarding range safety and weather
conditions
Pay attention to other launches, especially when conditions require a “head‟s up”
Make sure the checklist is followed and all steps are completed properly in a
thorough, workmanlike manner to assure mission success.
To further ensure mission success, considerations must be taken while at the
launch prepping and flying the vehicle to keep all the people around and the vehicle
itself safe.
Important safety related considerations are found in the following list:
Always follow the NAR/TRA safety code.
Adhere to local, state and federal regulations.
Never arm electronics unless rocket is vertical and the criterion for testing
continuity listed above is met.
Never proceed with launch if there are any outstanding technical issues that may
reduce the chances of a safe flight without first consulting both safety officers and
NASA officials if needed.
No smoking or open flames within 25 feet of the vehicle.
Do not put self or others in path of body tube in case of early ejection on
the ground; always be aware of the possibility of ejection charges firing at any time.
Verify that ignition leads are not live before connecting igniter to ground control.
(Touch leads together in the shade and listen and watch for sparks or place against
tongue)
73
Verify rocket will exit launching device vertically with almost no friction from the
launch guides
Handle rocket carefully to avoid injury to self, others or the vehicle
Be observant of other rocket prep operations in the launch pad area
Verify that ground around launch pad is cleared of flammable materials.
TABLE 3-19: TOOL USE INJURY POTENTIALS AND MITIGATIONS
Tool: Injury Potential: Risk mitigation procedure:
Electric Handheld
Sander Burns, cuts, skin abrasion Avoid loose clothing
Rotary Cutter/Dremel
Cuts, skin abrasion, eye
damage from flying debris,
respiratory damage from dust
Always wear a mask, gloves,
and safety goggles when
operating
Soldering Iron Burns Exhibit care not to come in
contact with hot element
Handsaw Cuts, splinters, skin abrasion Wear proper safety gear (gloves
and goggles)
Table Saw Cuts, Limb/appendage
removal
Avoid loose clothing, follow
safety procedures found in
instruction manual.
Wood Lathe Cuts, broken appendages Avoid loose clothing, use proper
tools and safety equipment
Table Router Cuts, Limb/appendage
removal Use proper protective gear.
Drill Press Cuts, abrasion, loss of limbs/
appendages
Use proper protective gear, hold
down work with clamps
Miter Saw Cuts, Limb/appendage
removal
Avoid loose clothing, follow
safety procedures found in
instruction manual.
Band Saw Cuts, loss of
limbs/appendages Use proper protective gear.
Belt Sander Burns, skin abrasion No loose clothes, wear proper
protective gear (gloves)
74
CNC Water Cutter Cuts, loss of
limbs/appendages
Only trained personnel use this
tool. Oversight by shop
manager
CNC Laser Cutter
Burns, eye damage,
respiratory issues, poisonous
off-gassing
A training course is required by
MIT to use the laser cutters on
campus. Many safety measures
exist, including ventilation and
failsafe switches.
Mill Loss of limbs, scarring, eye
damage from flying chips
A training course is required by
all shop managers to use mills.
Safety goggles are always worn.
Metal Lathe Loss of limbs, cuts, eye
damage from flying chips
A training course is required by
all shop managers to use lathes.
Safety goggles are always worn.
Safety Codes
The Tripoli Rocketry Association and the National Association of Rocketry have adopted
NFPA 1127 as their safety code for all rocket operations. A general knowledge of these
codes is needed and will be required by all team members. These codes are found in
the CDR-Safety document.
Hazards Recognition
The Hazards Recognition Briefing PowerPoint Presentation will be given prior to
commencing rocket construction. It will cover accident avoidance and hazard
recognition techniques, as well as general safety.
1) General
a) Always ask a knowledgeable member of the team if unsure about:
i) Equipment
ii) Tools
iii) Procedures
iv) Materials Handling
v) Other concerns
b) Be cognizant of your own actions and those of others
i) Point out risks and mitigate them
ii) Review procedures and relevant MSDS before commencing potentially
hazardous actions
c) Safety Equipment
i) Only close-toed shoes may be worn in lab
75
ii) Always wear goggles where applicable
iii) Always use breathing equipment, i.e. face masks, respirators, etc, where
applicable
iv) Always wear gloves where applicable, e.g. when handling epoxy and other
chemicals
2) Chemicals
a) The following are risks of chemical handling:
i) Irritation of skin, eyes, and respiratory system from contact and/or inhalation
of hazardous fumes.
ii) Secondary exposure from chemical spills
iii) Destruction of lab space
b) Ways to mitigate these risks:
i) Whenever using chemicals, refer to MSDS sheets for proper handling
ii) Always wear appropriate safety gear
iii) Keep work stations clean
iv) Keep ventilation pathways clear
v) Always wear appropriate clothing
3) Equipment and Tools
a) The following are risks of equipment and tool handling:
i) Cuts
ii) Burning
iii) General injury
b) Ways to mitigate these risks:
i) Always wear appropriate clothing, e.g. closed-toed shoes.
ii) Always wear appropriate safety equipment
iii) Always ask if unsure
iv) Err on the side of caution
4) Composites Safety
a) Carbon fiber, fiberglass, epoxy, and other composite materials require
special care when handling.
b) The following are risks composites handling:
i) Respiratory irritation
ii) Skin irritation
iii) Eye irritation
iv) Splinters
v) Secondary exposure
c) Ways to mitigate these risks:
i) Always wear face masks/respirators when sanding, cutting, grinding, etc., lay-
ups.
ii) Always wear gloves when handling pre-cured composites
iii) Always wear puncture-resistant gloves when handling potentially sharp
composites
76
iv) A dust-room has been constructed, as per MIT EHS guidelines,
specifically for the handling of composite materials.
d) No team member will handle carbon fiber until properly trained
3.9.4 ENVIRONMENTAL CONCERNS
All waste materials will be disposed of using proper trash receptacles
Non-polluting recovery system heat protection will be used
Solid rocket motor manufacturers‟ instructions will be followed when disposing of any
rocket motor parts
Consideration of environmental ramifications will be made regarding applicable
activities
Proper blast shields on the launch pad will be used to prevent direct infringement of
rocket motor exhaust on the ground
Waste receptacles (trash bags) will be available for use around the prep area to
encourage proper disposal of waste from rocket prep activities
The following list of materials have been identified as potentially hazardous:
o Aeropoxy 2032 Epoxy Resin
o Aeropoxy 3660 Hardener
o Ammonium Perchlorate Composite Propellant
o Black Powder
See CDR-MSDS document for complete MSDS specifications on these materials.
4 PAYLOAD CRITERIA
4.1 TESTING AND DESIGN OF PAYLOAD EXPERIMENT
4.1.1 SYSTEM LEVEL REVIEW
Note: All UAV Dimensions are in English (IPS) and all the aerodynamics calculations
are in Metric.
77
System Mass (lbs) Cost (USD)
UAV
Airframe 3.06 81.87
Motor 0.73 59.58
Avionics 3.21 ?
TABLE 4-1: UAV CHARACTERISTICS
78
TABLE 4-2: STABILITY ANALYSIS
79
ϒ
TABLE 4-3: FLIGHT CASE ANALYSIS
ρ
μ
80
Fuselage
The UAV fuselage is made of separable top and bottom halves. The fuselage is
rotationally symmetric except for two flat sections which are used to place a skid plate
made of Kevlar-epoxy and a camera window. The UAV does not have landing gear so
the Kevlar skid plate will prevent the fuselage from being damaged during landing. An
internal frame consisting of bulkheads and a plate on the bottom of the fuselage provide
structural support and allow attachment of the tail, wings, and avionics. The bulkheads
are epoxied to the top half of the fuselage and screws are used to secure the bottom
half of the fuselage to the bulkheads. This allows for easy removal of the bottom half of
the fuselage in order to gain access to the avionics.
The motor and a folding propeller are located at the rear of the aircraft which allow an
unobstructed view for the camera at the front of the aircraft and allows the propeller to
automatically fold back when it stops spinning. A folding propeller is used to ensure that
the UAV fits inside the rocket. The propeller will also fold up while gliding to reduce drag
and upon landing to avoid incurring damage.
The center of gravity will be located 25% of the fuselage length from the nose (11.25 in),
which will be accomplished by putting the avionics equipment in front of the wings. The
center of gravity and consequently the wings were placed this close to the nose to allow
the wings to fold up without hitting the vertical stabilizer. A stability margin (S.M.) of 0.1,
a horizontal tail volume coefficient (Vh) of 0.5, and a horizontal stabilizer aspect ratio
(ARh) of 4.5 were used to estimate the location of the neutral point and the aircraft wing.
These values were chosen to increase the stability of the airplane.
xnp
c1
41
2
AR
12
ARh
14
AR 2
Vh
81
The estimate of the neutral point of the aircraft is 2.7 in from the wing leading edge. This
estimate puts the center of gravity of the UAV 2.2 in from the wing leading edge. With
this information, we then calculated the distance from the nose to the wing leading edge
to be 9.1 in (wing chord is 5 in).
Wings
The wings have a chord of 5 in which was chosen because it is the largest chord that
will still allow the UAV to fit inside the rocket. Using this chord we then plotted
coefficients of lift and drag for various angles of attack using Xfoil (Xfoil is CFD
command line program developed at MIT for designing and analyzing airfoils).
Atmospheric data was found using a java applet by DesktopAeronautics
(http://www.desktop.aero/stdatm.php). Standard cruising conditions were considered to
be
Velocity = 25m/s ( )
Altitude = 2500 ft
Reynolds number = 204,000
Using this Reynolds number, nine airfoils were analyzed. The 6 airfoils with the highest
CL are shown in Figure 4-5.
82
FIGURE 4-1: AIRFOIL COMPARISON
The leftmost graph shows the coefficient of lift plotted against the coefficient of drag.
The center graph shows CL vs. α (the steeper graph) as well as the pitching moment
(Cm) vs. α (the flatter graph). After some consideration, the NACA 4412 was chosen
because it has a very high ratio of CL/cd and it has one of the highest CL around an
angle of attack of 3°, which will be the UAV wing‟s angle of attack while cruising. The 3°
angle of attack was chosen because it provides significantly more lift but doesn‟t
produce too much drag.
83
At a 3° angle of attack, the NACA 4412 airfoil has a coefficient of lift of 0.778. The stall
angle of attack at 66.0 ft/sec was determined to be roughly 14° (shown in Figure 4-2:
NACA 4412 Polar). The glide ratio and glide angle at 3° are 0.08 and 4.6° respectively.
FIGURE 4-2: NACA 4412 POLAR
After having chosen an airfoil, calculations for total weight were made with the equation:
[Density (ρ) = 2.21E-3 slugs/ft3]
A wingspan of 54 in was then chosen because it allows the aircraft to lift 7 lbs, roughly
the estimated weight of the airplane, while flying at its cruising speed.
Each wing will be horizontal for 3.5 in on either side of the aircraft and will then have a
5.8° dihedral (Υ) for the rest of the wing in order to increase spiral stability. A dihedral of
5.8° was chosen using:
84
[B = 4; lv= 29.3 in]
In this equation, Υ actually refers to the equivalent dihedral angle but in our
approximations we ignored the contributions of the flat center section of the wing to the
dihedral. The equivalent dihedral will be calculated more accurately in the future when a
more thorough stability analysis is performed.
FIGURE 4-3: TAIL
The UAV tail will consist of a vertical stabilizer and a horizontal stabilizer mounted part
way up the vertical stabilizer. The horizontal stabilizer is mounted away from the
fuselage in order to maximize its area while minimizing the total width of the airplane.
Both the horizontal and vertical tails will be made using the NACA 0008 airfoil. This
airfoil was chosen because it is symmetric and has a slightly smaller thickness/chord
ratio than the wing.
Control surfaces for the aircraft will consist of elevators and a rudder located on the tail.
They will be controlled with pushrods originating from servos located in the fuselage.
85
For size estimates, a horizontal tail volume coefficient of .5 and a vertical tail volume
coefficient of .05 were chosen in order to increase stability which is important for the
autopilot to function. The following equations were used for estimates of tail sizes.
Typical aspect ratios of 4.5 and 1.4 were chosen for the horizontal and vertical tails
respectively. Furthermore, taper ratios of .4 were used on both the horizontal and
vertical tail.
Aerodynamic Analysis
AVL and XFOIL were used to calculate lift and drag for the aircraft. AVL uses the
horshoe vertex lattice method to calculate inviscid lift and drag and aircraft trim. The
UAV wing and tail were modeled in AVL for standard cruising conditions (angle of attack
of UAV was set to 0°). AVL makes its calculations by breaking up each surface into
many small sections. XFOIL was used to calculate viscous lift and drag for each of the
sections using the inviscid lift coefficient calculated by AVL. Total lift and drag was
calculated for each section and the results from each section were summed to give total
lift and drag for the wing and tail.
AVL and XFOIL are not able to model the fuselage so hand calculations were employed
using the following equations:
(
)
All the results can be seen in the Flight Case analysis table. The UAV has an estimated
lift to drag ratio of 19 which means the UAV should glide well.
Folding Mechanisms
86
FIGURE 4-4: WING ROTATOR MECHANISM
In order to fit inside the rocket, the wings, horizontal stabilizer, and horizontal stabilizer
all fold.
The wings are attached to a spring loaded rotating mechanism. The mechanism allows
one wing to lift above the other. Both wings will then fold backwards. When unfolding, a
torsion spring swings both wings out until they achieve a 180° angle with each other. A
compression spring is used to pull the elevated wing down into position. The wings are
locked in place with magnets.
87
FIGURE 4-5: UAV IN STOWED CONFIGURATION
The tail folds by means of spring hinges. Spring hinges were chosen because of their
ease of integration into the existing plan and their relative low cost. Furthermore, the
spring hinges will automatically deploy when the UAV is released from the sabot. A pair
of rare earth magnets will be used to hold the tail in its unfolded position after tail
deployment.
FIGURE 4-6: STOWED TAIL
While inside the rocket, rods will be inserted into the wings and fuselage to hold the
wings in their folded position. This part will be attached to the rocket and will be pulled
out when the UAV falls out of the rocket.
4.1.1.1 DRAWINGS AND SPECIFICATIONS
[See Appendix 4: UAV Drawings]
4.1.1.2 ANALYSIS RESULTS
88
In order to verify the design of the UAV, analysis was performed on the wings and wing
rotator mechanism. The order of analysis is as follows:
Define loading conditions
Design part
Use hand calculations to size the part
Validate hand calculations using finite element method
Re-size as necessary
Loading Conditions
The wings of the UAV must be capable of withstanding two loads: the aerodynamic
forces during deployment and the weight of the fuselage during steady level flight. The
same analysis was performed on the wing rotator mechanism, as questions to its load-
bearing capability were brought up during PDR.
Deployment loading conditions for a single wing are summarized in Table X
TABLE 4-4: DEPLOYMENT LOADING
Deployment
Loading
Aero Loading 1.57 lbf
Fuselage Weight 8 lbf
MUF 1.5
Design Axial 14.4 lbf
Aerodynamic loading was calculated based on sabot speed at time of deployment, as
predicted by the Matlab recovery simulation. After separation from the main rocket
body, the sabot is decelerated to 26 ft/s, at which point the UAV is deployed. In reality, it
is hoped that the UAV will be in a nose dive during deployment. However, for analysis
purposes, a worst-case scenario of a flat fall was used. Additionally, the entire fuselage
weight was applied to each wing, rather than distributed between the two. The loads are
summed to create a total load, which is then margined by a 1.5 model uncertainty
factor, to account for inaccuracies in the recovery model and unpredictable
aerodynamic turbulence.
Wing Analysis
As described in the design section, the wings are a composite laminate of fiberglass
weave over a rigid foam core. The moduli and strength for the fiberglass and foam are
shown in Table X and Table Y. Note that the properties cited were for a unidirectional
fiberglass composite. Because we use a woven fiberglass sheet, the modulus was
divided by two, as approximately half of the fibers are aligned in a given direction.
TABLE 4-5: FOAM MATERIAL PROPERTIES
89
Material Properties - Foam5 6
E11 1.6 ksi
E22 2.8 ksi
E33 6.8 ksi
v 0.3
strength 25 psi
TABLE 4-6: FIBERGLASS MATERIAL PROPERTIES
Material Properties - Fiberglass7
E 3.25 Msi
v 0.3
strength 257 ksi
Material properties for the fiberglass/foam laminate were then calculated using the
Layup creator in Femap. Based on the unique shape of the layup and the limited
information available regarding the structural properties of foam, several assumptions
had to be made. First, the average thickness of the airfoil was used as the thickness of
the foam layer. For a NACA 4412 airfoil with a 5in chord, the maximum thickness is
0.60in. Half of this value, or 0.30in, was used as the foam ply thickness. Second,
because the shear modulus of the foam was unknown, the foam was approximated as
an isotropic material with the modulus of elasticity equal to E11. This was considered a
fair assumption because of the difference in magnitude between E11 and E22 of the
foam was small in relation to the magnitude of the fiberglass modulus.
TABLE 4-7: FIBERGLASS LAYUP PROPERTIES
Bending Properties – Fiberglass Layup
Ex 306 ksi
Ey 306 ksi
Gxy 118 ksi
vxy 0.3
vyx 0.3
Using these properties, hand calculations could be performed to calculate the maximum
deflection of the beam. These intermediate calculations can be seen in Table Z.
5 Moduli: http://web.mit.edu/16.unified/www/SPRING/systems/Lab_Notes/appendix.pdf
6 Strength: http://insulation.owenscorning.com/WorkArea/linkit.aspx?LinkIdentifier=id&ItemID=788
7 E-glass Epoxy Composite; http://www.carbonfibertubeshop.com/tube%20properties.html
90
TABLE 4-8: WING DEFLECTION CALCULATIONS
Wing Deflection
Calculations
Chord 5 in
Mean Thickness 0.3 in
Moment of Inertia 0.01125 In^4
Deflection 10.29 in
In order to verify these calculations, a finite element model was developed in Femap.
The applied boundary conditions are shown in Figure 4-7.
FIGURE 4-7: WING BOUNDARY CONDITIONS
The model was then solved using NEi Nastran, resulting in the ply 1 effective strain and
displacement outputs as shown in Figure 4-8 and Figure 4-9.
91
FIGURE 4-8: EFFECTIVE STRAIN
92
FIGURE 4-9: WING DEFLECTION
Analysis verified the deflection value produced by hand calculations. To reduce the
amount of deflection, the wing layup was redesigned with 1in thick carbon fiber strips
centered along the quarter-chord location on the top and bottom surfaces of the wing.
Properties for this layup, as computed by Femap, are listed in Table 4-9.
TABLE 4-9: CARBON FIBER LAYUP PROPERTIES
Bending Properties – Carbon Fiber
Layup
Ex 3.15 Msi
Ey 3.15 Msi
Gxy 1.21 Msi
vxy 0.3
vyx 0.3
Hand calculations were again performed, this time on the I-beam made by the carbon
fiber strips. These intermediate calculations can be seen in Table 4-10.
TABLE 4-10: CARBON FIBER DEFLECTION CALCULATIONS
93
CF Deflection Calculations
Width, b 1 in
Thickness 0.025 in
Separation
Thickness
0.3 in
Moment of Inertia 0.00132 In^4
Deflection 3.15 in
In order to verify these calculations, the Femap model was refined to incorporate the
carbon fiber strip. The new boundary conditions are shown in Figure 4-10.
FIGURE 4-10: NEW WING BOUNDARY CONDITIONS
The model was then solved using NEi Nastran, resulting in the ply 1 effective strain and
displacement outputs as shown in Figure 4-11 and Figure 4-12.
94
FIGURE 4-11: NEW EFFECTIVE STRAIN
FIGURE 4-12: NEW WING DEFORMATION
95
Wing Rotator Mechanism Analysis
The two major structural components of the wing rotator mechanism are the plates that
connect to the wing halves, and the central shaft that connects the entire mechanism to
the UAV fuselage. These components must be able to withstand the aforementioned
loading condition, as well as the moment generated by the force acting along the span
of the wing. A model of the rotator mechanism was made in ANSYS Workbench. Figure
4-13 shows the boundary conditions placed on the rotator mechanism.
FIGURE 4-13: WING ROTATOR BOUNDARY CONDITIONS
The model was then solved using the ANSYS engine. Stress and deformation
distributions can be seen in Figure 4-14 and Figure 4-15.
96
FIGURE 4-14: STRESS DISTRIBUTION
FIGURE 4-15: DEFORMATION
97
4.1.1.3 TEST RESULTS
To verify the wing analysis, mockup wings were tested with variable mass to determine
if the unidirectional carbon fiber strips were necessary to ensure the wings will be
capable of handling expected forces upon deployment in launch. Wings with fiberglass
only and wings with the carbon fiber strips were both tested with the same restrictions
with variable mass at a single point. In all the tests, the main mode of failure was point-
stress failure at the restraint, which is expected. The fiberglass wings failed with an
average applied torque of 6.3 foot-pounds. The fiberglass wings with unidirectional
carbon fiber imbedded in the layup failed with an average applied torque of 8.25 foot-
pounds. The wing design with the imbedded carbon fiber will be used to ensure that the
wings will be capable of handling the impulse upon deployment, and are more than
suitable to generate the necessary lift of the UAV.
Analysis done on the wing rotator shows that the components are more than capable of
handling any anticipated impulse load. To verify initial analysis, several wing rotators
will be tested post-CDR.
Other components and sub-systems will be tested post-CDR as seen fit. Only the
wings and wing rotator mechanism have been analyzed and will be thoroughly tested as
the failure of either of these two subsystems will result in the UAV becoming
fundamentally unflyable.
4.1.1.4 INTEGRITY OF DESIGN
A high level of integrity is expected of the UAV design as otherwise the science payload
will be left incapable of completing the mission. The integrity of the UAV design can be
noted throughout the manufacturing, testing and analysis sections. The beginning of
the design process included ample use of SolidWorks and team meetings with senior
team members to ensure that all systems and subsystems will be of high enough caliber
and structural reliability to meet the requirements set by the mission.
4.1.2 DEMONSTRATE DESIGN MEETS SYSTEM LEVEL REQUIREMENTS
Simplify UAV flight process
Due to the stability of the UAV, a control system is being employed to automatically
stabilize the UAV without input from the user. Not only does this act to remove the user
from the complexity of flying a UAV, but it also adds resistance to lag in the telemetry
system as the stabilization takes place on the UAV regardless of user inputs.
The ground station provides a simple „point and click‟ user interface for choosing the
location the UAV should fly to and land at. The same interface also allows the user to
easily select a cruising speed and altitude. To increase the ergonomic aspect of the
ground station, it is run on a tablet PC, although it is capable of running on any windows
computer with modification to be ported to Apple or Linux possible.
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Stream telemetry and video to ground station
Telemetry is provided by a pair of XBee 900MHz modules. The telemetry system
provides location, speed, altitude and heading data to the ground station to aid the user
in flying as well as temperature, pressure, humidity, UV index and solar irradiance data
as per SMD requirements).
The video transmission occurs via a separate 2.4GHz link to a secondary computer at
the ground station to provide the user with a first person view from the UAV.
Take "simple" commands from ground station
To simplify the flight process, the UAV is flown using a ground station which converts
the user input to commands which the UAV decodes to control its higher level actions
such as cruising altitude and speed or target location
Employ video tracking systems
Changes in the higher level requirements have reducing this to video tracking only.
Video tracking is performed using a first person video feed operating at 2.4GHz. This
video feed is forward looking and inclined to the horizontal to give the user not only a
view of upcoming terrain, but also a view of the ground aiding in the goal of
reconnaissance.
UAV must be capable to operate for at least 30 min
Based on the power and propulsion system, the UAV should be able to perform a
fly/glide flight profile for over this 30 minute target. However during the USLI competition
the UAV is only expected to fly for 5-10minutes for the purposes of flight demonstration.
4.1.3 WORKMANSHIP AND RELATION TO SUCCESS
Avionics
The avionics fall into two separate components; the flight computer and the backup
board. Since the flight computer is tasked with controlling the UAV, its correct
construction is critical. The majority of the construction is soldering however, which is
neither difficult to perform nor to inspect, so all the solders are inspected after
construction as a part of the verification process. The software on the flight computer is
also critical; to ensure high quality of the software, a branch from the ArduPilot Mega
software repository is being used, with additions such as adjustments to the data
protocol and for sensor data logging.
Although the backup data board is not critical, the majority of construction is also
soldering, verification of correct construction is also not a complex task. Also since there
is significant commonality between the additional flight computer software and the
backup board software should have the consequence of reliable backup board software.
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UAV Hardware
As long as a stable UAV platform is produced, then the only aspect of the UAV that is
critical to the correct operation of the UAV is the correct design and construction of the
wing rotator mechanism. A failure of the wing rotator mechanism could result in the UAV
being fundamentally unflyable, and the control system on the flight computer will be
unable to fly the UAV successfully.
UAV Composite Fabrication
The UAV composite components require the testing of new fabrication techniques for
the wings and fuselage. The wings incorporated a technique new to the rocket team
that produces a completely smooth and wrinkle-free wing layup, incorporating plain-
weave carbon fiber strips. An experienced team member in this new fabrication
technique watched over other team members as several practice layups were made,
until satisfactory results were obtained. Additionally, the positive fuselage mold was
made of wood on a wood lathe, by hand, practicing acquiring an appropriate taper. It is
appropriate that several more practice layups are performed to ensure that the
fiberglass layup can be easily removed from the wood mold, without ruining the
fiberglass as experienced from the initial layup. The team members will continue to
improve their skills and layup techniques, continually improving the quality of all the
fabricated composites so the final UAV will be of superb quality. Extra wings,
stabilizers, and entire UAVs will also be made and be available as backups in Huntsville
so that the payload will be readily repairable and flyable in the event of any
transportation or mission breakages.
4.1.4 PLANNED COMPONENT, FUNCTIONAL, AND STATIC TESTING
The team‟s first priority will be to perform qualification testing on the structural
components of the payload. The tests to be performed are as follows and will be
completed after the structural test article is completed post-CDR:
Wings will be tested further with variable mass, using small weights. As verified
by initial testing, wings will not reach full load capacity, failing prematurely at the
point of contact between the wing and edge of table holding wing.
Horizontal stabilizer will be tested in same manner as wings.
Test the wing rotator several times in three manners:
o That wing rotator will deploy wings from a folded manner.
o A symmetric variable mass on attached wings, and wings are expected to
fail before the aluminum wing rotator. Will apply variable mass until wings
break off rotator.
o Drop test of UAV from tethered weather balloon.
Fuselage will be tested in a manner of ways:
o Crush test from nose to tail
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o Impact test from the top of the body and the bottom of the body. This will
ensure that the fuselage will be able to withstand the impact force
expected upon landing. Will also verify the connecting points between the
two halves of the fuselage will withstand forces upon deployment and
landing.
o Abrasive test on the Kevlar landing strip, testing the resilience of the
Kevlar for the skidding expected upon landing.
After component structural testing is complete, deployment and recovery tests will be
performed. This will occur with the final prototype assembly.
Deployment test from the sabot to ensure that wings will separate sabot upon
deployment from rocket tube. Ensures that the spring components of the wing
rotator mechanism are capable of separating sabot.
Several flight test to ensure that control gains are suitable for the final UAV
design and prototype.
Drop test from tethered weather balloon, testing the avionics system and
prototype UAV work under mission conditions.
See section 4.3.4 for greater detail in testing UAV and avionics. These tests will
occur before the FRR.
4.1.5 STATUS AND PLAN FOR REMAINING MANUFACTURE AND ASSEMBLY
Payload
Fuselage:
A prototype wood mold of the fuselage body has been created and several layups have
been performed to test the method that will give the best fiber glassing results. Using a
standard fiberglass layup process in the rocket teams own oven, using 3 plies of
fiberglass, only half of the fuselage is made at a time. The two halves are combined
after reaching a full cure using a thin strip of fiberglass and epoxy. The fabrication
process and the team members experience level produce a fuselage of high enough
quality to proceed with assembly and integration of other sub-components. The UAV
prototype will be used for all testing, and once the systems are all approved and
working properly, the team will go into the final fabrication process that will include
making another wood mold of the fuselage to required dimensions. The final prototype
will have the entire fuselage made of fiberglass, and a small window will be cut and
clear polycarbonate installed into the nose of the fuselage with large enough
dimensions to allow for the camera to have a clear view of the ground.
Wings:
The airfoils will be cut out of Owen Corning Foamular 250 using a foam cutter, the
custom foam cutter provided by the MIT Department of Aeronautics and Astronautics for
student use. Using the NACA 4412 for the wings, the settings of the foam cutter
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machine were adjusted until the desired shape and finish of the foam is attained. Using
techniques developed my previous years of MIT student trial and error, the wings are
going to be lain up using a single ply of fiberglass and 1” unidirectional carbon fiber
strips across the top and bottom of the wing on the quarter cord to increase the strength
and load capabilities of the wings. This is opposed to simply doing a single ply of
fiberglass and no carbon fiber, which produces wings that will have a much higher
likelihood of failing upon initial deployment. Upon some analysis, the carbon fiber strips
decreases the bending of the wing by 6.5” across one wing with a load of 14 lbs of
force, resulting in only about 2.6” of tip deflection versus the 9.125” tip deflection.
Several layups have been performed and the final wings are of suitable quality. Some of
the internal components, including the hinge components, will be cut with a waterjet for
the aluminum pieces, and a laser cutter for the wood components and will be integrated
when assembling the first prototype for flight. Members of the rocket team have
acquired suitable experience using the machines and have been trained in the safe
operation of this equipment, obtaining fantastic results.
Empennage:
The NACA 0008 airfoils for the horizontal and vertical stabilizers will be cut in same
fashion as the wings, and will undergo the same layup process excluding the
unidirectional carbon fiber strips. After the fiberglassing process is completed, the
stabilizers will be cut and magnets will be installed inside compartments that will be cut
into the foam. This will lock the stabilizers in place upon deployment.
Wing Rotator:
The wing rotator will be constructed of 6061-T6 Aluminum, standard aluminum, as
opposed to the polycarbonate construction in response to anticipated loads upon
deployment. All pieces are machined separately in MIT metal machine shops using
CNC Lathes, CNC Mills, and a water jet. Three complete wing rotator mechanisms will
be made, along with several back up components, the spares will be on hand at
Huntsville in case of any transportation or other accidents leaded to integrated
component failure.
Internal components:
All other internal components will be made using standard machines, along with a
waterjet and laser cutter for appropriate components.
Avionics
TABLE 4-11: AVIONICS TESTING
Sub-System Completed Remaining
Ground station Software Software written Testing
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Flight Computer
Core flight computer built
Updated code for UAV specific needs
Tested to ensure correct servo actuation
Integrated with test aircraft
Tested in manual mode via an RC handset
Integrate with sensor assembly
Test in flight
Test telemetry link
Test sensor data logging
Back Up Logging Board
Sensor software written
Logging software written
Construction started
Complete construction
testing
Visual System Integrate with UAV
Test
User Interface Description
FIGURE 4-16: GROUND STATION USER INTERFACE
1
2
3 4 5
6
7
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1. Top down view – This provides the user with information of the local area. The
zoom of the view varies with altitude of the UAV to maintain a view 800m above
that of the UAV. The map tracks such that the UAV remains central to the map
which also rotates so the UAV is always travelling upwards on the UAV i.e. the
UAV sprite remains static.
2. UAV Icon – This denotes the location of the UAV on the map and remains static
since the map translates, rotates and resizes around this icon
3. Target Icon – This marks the location of the „target waypoint‟ and is modified by
clicking within the top down view. Moving the target does nothing until one of the
waypoints commands (4) and selected.
4. Waypoint Commands – Once a suitable target location has been selected these
are used to determine how to use that location information; whether to fly to that
location and loiter there or whether to land there.
5. Cruise Control Settings – Here the target cruise altitude and speed can be
selected. Once selected the control system on the UAV will attempt to maintain
these values during flight.
6. Cruise State – Displays the current cruising speed and altitude of the UAV
7. Sensor Data – Displays the current sensor data as recorded by the flight
computer in accordance with SMD requirements.
4.1.6 DESCRIBE INTEGRATION PLAN
Integrating the UAV into the sabot:
The following steps should be completed after the parts of the Pre-Flight Checklist
before this step have been completed. In other words, the UAV should be completely
ready to be deployed before integrating it into the sabot.
Separate the sabot halves and lay them next to each other.
Set the UAV on a sturdy working surface
While bracing the body, carefully fold the UAV‟s wings back, making sure to
depress the rotating mechanism such that one wing folds under the other.
Lock the wings in place with the four pronged locking plate (this stays with the
sabot when the UAV is deployed)
Fold the three spring-hinged tail surfaces.
Place the UAV in one sabot half.
Insert the activation pin that pulls out at time of deployment.
Do one final look-over to make sure everything seems ok.
While holding the tail surfaces so they do not spring out, place the UAV in one of
the sabot halves.
Carefully close the sabot with the other half.
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Have someone hold it closed until it‟s ready to be integrated into the rocket.
4.1.7 INSTRUMENTATION PRECISION AND REPEATABILITY
There are two areas which effect precision of the data collected beyond the precision of
the sensors themselves since data is recorded in two ways; transmission to the ground
station and storage on non-volatile memory.
Data transmitted to the ground station has a lower precision than the data collected by
the sensors due to the need to encode the floating point data into integer values for
sake of compatibility. This results in the maximum precision shown in table XXX.
TABLE 4-12: PRECISION OF TRANSMITTED DATA
Data Decimal places
Humidity [%] 1
Solar [Wm-2] 2
UV Index 1
Pressure [Pa] 0
Temperature [oC] 1
The precision of data logged by the backup board is the same as the precision of the
sensors themselves as no conversion from floating point need be done. The precision
for the sensors is shown in table XX and the back board components are highlighted for
clarity.
TABLE 4-13: SENSOR PRECISION
Data Sensor Location Precision
Pressure
SCP1000 Back-Up Board 6Pa
(8.7E-3psi)
BMP085 Flight Computer 100Pa
(1.45E-2psi)
Solar Clairex CLD140
Back-Up Board /
Flight Computer
Humidity Honeywell HIH-4000
Back-Up Board /
Flight Computer 3.5%
UV Index Back-Up Board /
Flight Computer 0.1
Temperature
SCP1000 Back-Up Board 0.03125oC
(0.05625oF)
BMP085 Flight Computer 0.1oC
(0.18oF)
Dallas Back-Up Board / 0.5 oC
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Semiconductor DS18S20
Flight Computer (0.9°F)
It is worth noting that the flight computer also logs data from the sensors connected to it
at their precision as this data is not transmitted and so does not need to be converted to
integer values for compatibility.
Given the changing nature of the atmosphere repeatability of data collection is
synonymous with reusability of the data collection hardware and by extension the UAV
and rocket. To this extent all aspects of the rocket, UAV and avionics have been
designed with reusability in mind.
4.2 PAYLOAD CONCEPT FEATURES AND DEFINITION
4.2.1 CREATIVITY & ORIGINALITY OF PAYLOAD
The idea of a deploying an Unmanned Aerial Vehicle (UAV) with a rocket is not an
entirely original idea; however, the end goal of producing a simplified flight control
interface is a new idea. Current UAV technology requires a classically trained pilot to
remotely fly the craft, or at the very least requires operators to undergo a large amount
of training in the operation of remote controlled equipment. The control interface that the
MIT Rocket Team is developing aims to reduce the amount of training required to
successfully complete a UAV mission, opening this class of technology to a greater
range of potential users. As such, a simple “point-and-click” flight system is being
developed to easily translate user operations into functional flight controls, allowing for
successful operation with little or no operator training. Furthermore, by choosing a
rocket deployment, and keeping to a $5000 budget, it further allows for this technology
to be applied to situations where time and budget are controlling factors. This quick
deployment and relative low cost of operation would ideally suit the needs of search and
rescue operations, reconnaissance missions, and even rapid scientific data gathering
missions.
By completing NASA‟s Science Mission Directorate, the MIT Rocket Team is further
proving the range of applications the UAV is capable of completing. The requirements of
the SMD do not explicitly require the complexities of a UAV. However, by using the UAV
form factor, the MIT Rocket Team will be able to complete all required tasks of the SMD
mission in a more precise manner over several missions, and will also have the
capability to loiter in any airspace, being limited only by the charge on the batteries.
payload. For example, during flight, a UAV will generally maintain the same orientation
with respect to the horizon, allowing for all images taken during and after the flight to
keep the sky and ground in the same location with a very low chance of error.
Furthermore, the use of a controllable payload allows for the investigation of specific
areas, allowing for the gathering of data of greater importance while limiting the need for
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secondary missions. Finally, the use of a UAV allows for a greater amount of data to be
collected due to the extended flight time of a UAV platform compared to other payload
options.
4.3 SCIENCE VALUE
4.3.1 PAYLOAD SCIENCE OBJECTIVES
There are two different aspects to the payload, each with their own objectives; the SMD
payload requirement and decreased complexity in UAV flight.
The payload objectives relating to the SMD payload are to log atmospheric pressure,
temperature and humidity along with solar intensity and UVI data at 5-second intervals
as well as taking at least two still images during flight and three after landing.
The payload objective relating to decreasing the complexity in flying a UAV is to
complete the flight and mission (visually locating the rocket and landing) solely using the
software provided at the ground station without reverting to back up manual control.
4.3.2 PAYLOAD SUCCESS CRITERIA
The data logging and sensors shall be deemed successful if the payload obtains and
logs atmospheric pressure, temperature and humidity along with solar intensity and UVI
data at 5-second intervals as well as taking at least two stills during flight and three after
landing. It shall be deemed a success regardless if the data is collected by the main or
the back-up sensor package.
Fulfilling the SMD payload requirement successfully shall also demonstrate the flexibility
in the UAV design.
If the UAV operator successfully visually locates the rocket and lands in a state fit for
reusability, without resorting to use of the back-up manual flight control then this will
demonstrate successful reduction in complexity of UAV control.
4.3.3 EXPERIMENTAL LOGIC, APPROACH, VERIFICATION
By using a science payload in a descending UAV, atmospheric measurements
presented in section 4.2 will be collected. The science payload will be contained inside
built-in compartments in the fuselage, preventing thrashing of instruments from launch
initiation to landing of UAV.
To obtain such data, all the sensors will be turned on just prior to launch and
measurements will be recorded at 5-second intervals during launch and decent within
target area. Using a UAV to carry a science payload of multiple sensors and accurately
obtain such data will provide a more efficient means for obtaining such data.
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Additionally, the telemetry devices inside the UAV will allow for safe operation of the
vehicle via a pilot at the ground station. A single mission by a UAV with such science
payload gathers data at varying altitudes effectively and efficiently, relative to other
means of acquiring such data.
The goal of the UAV is also to have a simplified flight control interface. This will be
achieved by having stability control on the UAV such that it is able to maintain straight
and level flight, perform controlled turn and land safely with no user input. Combining
this with an appropriately designed user interface, this should be sufficiently automated
that it can be controlled by a person with absolutely no flying experience.
4.3.4 DESCRIBE TEST AND MEASUREMENT, VARIABLES AND CONTROLS
Testing and verification of the avionics occurs in three distinct phases: ground testing,
on a test aircraft and lastly on the final UAV, thus enabling ground testing shall consist
of validating the correct operation of all hardware and sensors in a non-critical
environment. The testing on the test aircraft serves to verify that the subsystems within
the avionics system work as expected in flight case and to validate changes made to
the flight computer hardware and software for the purposes of the competition. The
flight testing on the UAV is to demonstrate the avionics system is able to function
correctly in its intended flight configuration and importantly, that is it capable of recovery
after deployment from the rocket.
Phase One – Ground Testing
First the flight computer, GPS/IMU, and telemetry boards shall be connected to ensure
these systems are functioning. Next this system shall be connected to servos and a
program will be used to actuate the servos in a known fashion to confirm that the servos
can be controlled. Furthermore, an R/C receiver will be attached to the avionics system
to ensure that the equipment accepts commands from an R/C handset. The telemetry
system shall also be tested by communicating the position the GPS reports to a mock-
up ground station. That position will be checked for accuracy, the time taken for the
GPS to get position lock at start up will be noted, and whether or not the GPS is able to
maintain lock when moved at a reasonable speed will be checked.
The back-up board will also be constructed and be tested by exposing it to a variety of
conditions such as varying the altitude under different weather conditions and
comparing to the expected values. This acts to validate that both the sensors and the
data recording systems within the back-up board are working correctly.
The real-time video system will also be tested by varying the distance between the
transmitter and receiver to find the limit of the range the system is able to adequately
transmit the video data.
Phase Two – Test Aircraft
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The initial flight-testing will occur on a aircraft team owned and previously built UAV.
This acts to limit the dependency of avionics testing and refining on the construction of
the UAV, thus expediting the development of the avionics. This also minimizes the risk
to the final UAVs; there will be at least one back up UAV and replacement components
for internal mechanisms in the event that the primary UAV fails or breaks.
The initial flights will be manual, i.e. the R/C aircraft shall be controlled solely by an
operator using a standard R/C controller. These initial flights will be to verify the correct
setup of the flight computer and servos. After that, flight testing shall be undertaken to
tune the control gains of the flight computer for stable flight. It is worth noting that these
gains will not necessarily be those required for the final UAV, but the autonomous flying
ability gained from this is essential for further flight testing. Next flight tests shall be
performed where coordinates of waypoints to be flown to will be uploaded to the
avionics system in-flight to ensure that this functionality works as expected. At this point,
the flight computer hardware will be wired to the primary sensors. Also, the flight
software shall be modified to log the sensor data on the internal volatile memory and
transmit the logged data post landing. This additional functionality shall then be tested in
multiple flights to ensure correct operation. Over the course of the latter flight testing the
back-up sensor board, real-time video transmission and still capturing systems shall be
integrated into the R/C aircraft, tested and refined as necessary.
Phase Three –UAV Testing
After the avionics performance has demonstrated adequate performance on the „Test
Aircraft‟, the avionics system shall then be integrated with the UAV. The first flight
testing shall be to determine the control gains required for stable flight of the UAV. For
the purposes of these tests, the equipment not essential for flying (i.e. everything but
flight computer, telemetry link and GPS/IMU) shall be replaced by appropriate ballasting
to minimize the risk of damage to components.
Once adequate control gains have been determined, a series of flight tests shall be
undertaken to ensure that the sensor systems and data logging systems, as well as the
imaging systems, still function as desired. These flights will also determine if the
propulsion system‟s duration and thrust are sufficient to maintain steady-level flight for
at least 30 minutes. Further testing representative of flight scenarios shall also be
undertaken, including point-to-point flying based on user inputs at a ground station.
Drop tests from a tethered weather balloon shall also be used to simulate UAV
deployment to ensure the UAV/Avionics is capable of recovering from the post-
deployment dive. The UAV will be unpowered (propulsion system off) due to safety
reasons for these tests; the lithium polymer propulsion battery will be replaced by ballast
to mitigate the risk of the lithium polymer battery exploding due to damage if the UAV
were to crash. Gliding should be sufficient to test all avionics. A test section of the
rocket body tube will be hung from a balloon platform attached to the weather balloon.
The UAV will be packed into the sabot, and the sabot will be placed in the body tube
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and connected to a radio controlled tender descender. The balloon will be tethered and
raised to an altitude sufficient enough such that the UAV and sabot will be falling at
speeds identical to those of launch conditions when approaching an altitude of 200 ft,
releasing from the rocket tube at 200 ft; this altitude should be sufficient for full UAV
deployment, while restricting the safety radius needed to be cleared of personnel on the
ground to a reasonable value. The sabot will be dropped under drogue parachute, and
the UAV will deploy. These tests shall be performed with ballast instead of non-essential
electronic components. This ballast will be placed in such proportions and
arrangements to maintain the center of mass of the UAV, providing sufficiently accurate
mission conditions for the UAV.
4.3.5 RELEVANCE OF EXPECTED DATA
The data collected is vital for the analysis of the systems and subsystems in
determining any necessary changes to the design of the UAV, or to any instruments and
power devices. Accuracy of the data is also significant in that, if the lift produced or
propeller propulsion is not enough, the UAV will need to adjust its attitude, which can
potentially lead to unbalanced forces, instability of the vehicle, or even stalling of wings
or horizontal stabilizer.
Effectively, all data on the stability lift and drag forces for the wings, horizontal
stabilizers, and the assembled body must be accurate to determine the necessary
attitude of the vehicle to achieve specific tasks, such as steady-level flight, landing, and
elevating altitude.
The various measurements of the atmosphere will be gathered, organized and analyzed
to study changes in the atmosphere with changes in altitude, changes in amount of
atmosphere between the payload and ground, and changes in level of atmosphere
between the payload and space. This will provide real data, to contrast to theoretical
data predicting such qualities of the atmosphere based on location, altitude, and density
of the air.
4.3.6 ACCURACY AND ERROR ANALYSIS
Accurate data provides information about atmospheric conditions to people, giving
realistic data for the analysis and design of different potential aerial mechanisms. Such
data will also allow for scientific groups to consider the protection necessary for
instruments of varying sensitivity to cosmic electromagnetic radiation, that are planned
on being deployed at varying altitudes. Devices and forces can be greatly affected by
variables such as pressure, temperature, relative humidity, solar irradiance, and UV
radiation; appropriate knowledge of such variables can allow for proper preparation for
objects entering such conditions.
4.3.7 EXPERIMENT PROCESS PROCEDURES
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Individually test all sensors for temperature, pressure, relative humidity, solar irradiance, and UV radiation (primary and back-up sensors).
Pressure can be tested in a wind tunnel with a known dynamic pressure o Temperature probes can be tested at room temperature, and outside ambient
temperature, which will range from 30°F to 60°F, at time of anticipated testing o Relative humidity can be tested in dry rooms and in ambient air, with known
humidity levels known from known weather conditions.
Determine mass of all instruments, avionics, and power devices
Estimate mass of UAV body materials
Identify a suitable propulsion system and battery for device powering
Using computational software, Excel and MATLAB, verify calculations for expected parameters and requirements of the UAV.
Using CAD software, model UAV with appropriate dimensions and parts.
Using hand calculations and analysis software (i.e. AVL, ANYSYS, Nastran) to determine failure loads of wings and internal mechanisms of UAV
Use flight simulation software to determine flight patterns of UAV
Develop mission success criteria o All data accurately acquired and stored properly o Still photographs acquired at SMD prescribed intervals o Communication between payload and ground station seamless o Semi-autonomous navigation capable of navigating to command coordinates o Safe landing of rocket and tethered pieces with use of parachutes o Safe landing of UAV, employing protective underside Kevlar coat
Ensure rocket, UAV and other equipment are reusable after each mission
4.4 PAYLOAD SAFETY
4.4.1 IDENTIFICATION OF SAFETY OFFICER
Andrew Wimmer will be the primary rocket safety officer for the team. Ben Corbin is the
team‟s MIT EHS representative and is the assistant safety officer and is in
charge of safety issues not directly related to the rocket. Both team members have
considerable experience in their respective areas.
4.4.2 FAILURE MODES
UAV
TABLE 4-14: POTENTIAL UAV FAILURE MODES
Risk Consequence Mitigation
UAV is damaged
before launch
weekend
The UAV needs to be
repaired or replaced
quickly to allow for
Avionics testing in a commercial R/C
aircraft and ground testing of the UAV
will be performed before the UAV is
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testing to continue. flown. Extensive flight testing with the
final UAV design will be performed, and
an experienced backup pilot will be
standing by during any flight test with
the R/C controller to mitigate chances
of a crash landing.
UAV is damaged on
launch weekend
before launch
The degree of damage
needs to be assessed.
Quick repair or
complete replacement
with a backup UAV
needs to be done.
Extreme care will be taken when
packing and handling the UAV. A
backup UAV will be taken to the launch
site to prevent total mission failure in
the event of the primary UAV being
damaged.
The UAV is
damaged during
deployment
The UAV may not be
controllable in autopilot
or manual mode, and
part or all of the
mission may fail.
High strength materials, including
fiberglass, carbon fiber, polycarbonate,
aluminum, and steel, are used in the
construction of the UAV to prevent total
failure and part separation.
The UAV‟s autopilot
fails
The backup pilot
engages manual
control and the UAV is
flown like a normal R/C
aircraft. All data can
still be collected and
transmitted.
A backup pilot with manual R/C
controller and a spotter with binoculars
will be on hand at all times, ready to
take over control of the UAV in case of
any autopilot failure.
The UAV becomes
entangled in shock
cords or shroud lines
upon exiting the
sabot.
The UAV will not be
able to fly. The pilot will
immediately go to
manual mode to
prevent the UAV from
trying to correct its
flight and potentially
damage the rocket.
The UAV will still be
able to collect and
stream data. Rocket
parachutes may not
deploy correctly.
The location of the parachutes and the
length of the shock cords are such that
the UAV should avoid entanglement.
Testing will drive adjustments to the
rocket‟s recovery system to mitigate
UAV entanglement.
Control surface(s)
break off during
initial dive pull-out
Depending on the
control surface, the
UAV may no longer be
maneuverable and will
likely crash land. Data
Balloon drop tests will verify that the
UAV can pull out of a dive and fly
without sustaining damage.
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can still be collected.
Wing breaks off
immediately after
deployment
The UAV will no longer
be maneuverable and
will crash. Depending
on the damage
sustained, data may or
may not be able to be
collected.
The UAV‟s wing rotator and wings have
been over-engineered to sustain much
higher loading than could be
experienced throughout the entire
flight. Balloon drop tests will verify that
the UAV can pull out of a dive without
sustaining damage.
High-wind conditions
on launch day
The UAV could be
damaged by wind
gusts. The UAV
experiences turbulent
flying conditions,
making control difficult.
Video may be shaky.
The UAV structure is designed to
handle high-loading situations, such as
those experienced by strong winds.
The UAV autonomous controls will be
able to stabilize the flight of the UAV in
high winds. The UAV design will not
need to take into account flying in very
high winds (15mph+), because it is
likely that the entire launch will be
postponed if high-wind conditions are
present due to the danger of launching
high power rockets in high-winds.
The propulsion
system fails
The UAV will not be
able to sustain or gain
altitude, decreasing
mission time. The UAV
will still be controllable.
Flight testing and pre-flight checks will
ensure that all UAV systems are
working properly before final
integration. The avionics are designed
to not need a propulsion system to
function properly, and will still be
capable of landing the plane safely.
The UAV‟s
propulsion battery
dies before the end
of the mission
The UAV loses the
ability to sustain or
gain altitude,
decreasing mission
time. The UAV will still
be controllable.
The propulsion and avionics batteries
will be separate, so if the propulsion
battery dies, the UAV will act as a
glider. Both batteries, as well as
backups, will be fully charged on
launch day. Flight testing will determine
necessary battery capacities.
The UAV‟s avionics
battery dies before
the end of the
mission
The UAV loses all
control abilities and will
crash land.
Flight testing will determine minimum
avionics battery capacity. For safety
purposes, this capacity will then be
doubled for the final UAV. Ways of
utilizing the propulsion battery as a
backup avionics battery are currently
being investigated.
The UAV is out of
sight at time of
No knowledge of
whether or not the
Not having sight of the UAV should not
be a problem, as the autopilot in the
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deployment UAV has successfully
pulled out of its initial
dive. Backup pilot
cannot engage manual
control without sight of
UAV.
UAV will undergo extensive testing
before launch day, and the tablet pilot
interface should work as planned. A
spotter with binoculars, whose main
responsibility is to track the UAV by
sight, will be with the backup pilot at all
times. The live video feed and GPS
coordinates transmitted from the UAV
should also give information pertaining
to the status of the UAV.
The UAV makes it
out of the sabot, but
does not right itself
(The autopilot and
GPS require the
UAV to be upright.)
The backup pilot will
engage manual control
and hopefully be able
to steady the UAV. If
control is not gained,
the UAV will crash
land, but still be able to
collect and transmit
data.
Upon release from the sabot, the UAV
will immediately and automatically set
its control surfaces to a position that
should pull it out of the initial dive in an
upright position. Wing dihedral should
make being upright a much more
favorable position. If all this fails, the
backup pilot has a long time (2500 ft) to
steady the UAV with manual control.
The UAV‟s propeller
contacts part of the
rocket while under
throttle
Damage to the rocket
and UAV may occur.
The UAV may be put
into a spin due to such
contact.
The autopilot will be programmed not
to engage the motor until a set time
has passed from deployment. If this
fails, the backup pilot can engage
manual control and set the motor
throttle to zero.
An internal
component in the
UAV moves during
flight, shifting the
UAV‟s CG
The UAV becomes
difficult to maneuver.
Autonomous flight
capability may be lost
and manual control
may have to be
engaged.
All components will be securely
mounted. Balloon drop tests will verify
all component mounts are secure
enough.
The UAV deploys
correctly, but flies
over the crowd or
out of range of the
ground station.
Manual control is
engaged and the UAV
is piloted out of the no
fly zone.
A no fly zone will be coded into the
tablet pilot system disallowing the
addition of waypoints above the crowd
or outside the range of the ground
station. The rocket will be launched far
away from the crowd, as per NAR
regulations, so the chance of UAV
deployment over crowd is minimal. The
range of the ground station will be large
enough to cover the UAV at any
possible deployment location.
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UAV contact with
ground station is lost
Contact will try to be
regained and/or
manual control will be
engaged.
The UAV will be programmed with a
“loiter-mode” that will be activated if
ground station signal is lost. The UAV
automatically lands at a set location
after a set amount of time out of
contact with the ground station. In
“loiter-mode”, the UAV will still avoid
the no fly zones.
The UAV requires
an emergency or
forced landing due
to one of the above
or other risks
Manual control is
engaged and the
backup pilot brings the
UAV down as quickly
and safely as possible.
If that is not an option,
an emergency “kill
switch” is engaged that
puts the plane in a
steep downward spiral.
Low cruising/reconnaissance altitude
allows for minimal landing time in case
of an emergency. The mass of the UAV
has been minimized to decrease
damage in case of an emergency. The
addition of an emergency “kill switch” in
the programming (controlled by the
backup pilot) is a last resort to quickly
land the UAV.
Note, in all circumstances where control of the UAV is lost, with the exception of loss of
avionics power, the UAV will be able to collect and transmit data while it is falling.
Avionics and Ground Station
TABLE 4-15: POTENTIAL AVIONICS AND GROUND STATION FAILURE MODES
Failure Mode Description Consequence Mitigation
GPS Failure
GPS hardware
failure
Satellite lock failure
No navigation data
to run UI
Testing to ensure
antenna function
and adequate GPS
lock prior to launch
day
Flight is continued
under mode using
visual data
Telemetry Failure
Telemetry module
hardware failure
Antenna failure
Loss of
communication
between UAV and
Ground Station
Testing to ensure
antenna and
telemetry module
functions
adequately prior to
launch day
Flight is continued
under manual
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mode using visual
data and sensor
data logged by
backup board
Video Camera
Failure
Camera hardware
failure
Antenna failure
Loss of first person
view video stream
Testing to ensure
antenna and video
hardware functions
correctly
Flight is continued
using pilot and
spotter team
Still Camera
Failure
Servo actuation
failure
Loss of capability
to take still photos
Stills taken from
the video stream
during flight and
landing
If combined with a
video camera
failure then results
in a loss of image
gathering capability
Flight Computer
Sensors Failure
Sensor hardware
failure
No sensor data
collected by the
flight computer
No sensor data
transmitted to
ground station
Backup board logs
sensor data using
an independent set
of sensors to an
SD card
Backup Board
Failure
Ardiuno board
failure
Sensor hardware
failure
SD card failure
SD card writer
failure
No sensor data
collected by the
backup board
Flight computer
logs sensor data
and transmits it to
ground station
If combined with
flight computer
sensor failure then
results in a lack of
sensor data
Flight Computer
Failure
Flight computer
hardware failure
Flight computer
software freezes
Flight computer
software crashes
Loss of
communication
between UAV and
Ground station
No sensor data
collected by the
flight computer
Motor throttled to
Manual control and
a pilot spotter
teams glide the
UAV down safely
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idle
Gyroscopic
Sensor Failure
Sensor hardware
failure
Loss of orientation
data
Can be flown under
manual control if
control loops
become unstable
Accelerometer
Failure
Sensor hardware
failure Little consequence
Can be flown under
manual control if
necessary
Loss of Motor
Power
Motor dedicated
battery pack runs
out
Motor hardware
failure
Motor throttled to
idle
UAV can be glide
to landing via the
ground station
Loss of Avionics
Power
Avionics dedicated
battery pack runs
out
Loss of
communication
between UAV and
Ground Station
No sensor data
collected by the
flight computer
Motor throttled to
idle
UAV performs
uncontrolled glide
to ground
Primary Ground
Station Computer
Failure
Computer system
crash
Application crash
Laptop battery runs
out
Loss of telemetry
link
Multiple backup
computers capable
of running the
ground station
software
Employ manual
control until
software restart
Secondary
Ground Station
Computer Failure
Computer system
crash
Application crash
Laptop battery runs
out
Loss of first person
video feed
Multiple backup
computers capable
of displaying video
feed
Use a spotter in
place of video feed
until software
restart
4.4.3 POTENTIAL HAZARDS
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For manufacturing hazards, see “Hazards Recognition” and “Tool Use” in Section XX
(Personnel Hazards in rocket section).
In order to assure safe and successful operations concerning the payload, a checklist
must be followed. In order to reduce personnel hazards the following precautions must
be taken:
Avoid standing in the plane of the propeller when UAV propulsion system is on.
Do not try to catch the UAV during landing.
Make sure all relevant testing (reference checklist) has been completed prior to
attempting a flight test.
Make sure the checklist is followed and all steps are completed properly in a
thorough, workmanlike manner to assure mission success.
Lithium Polymer Battery Hazards and Procedures:
Always charge lithium polymer batteries with a balancer. Out of balance packs
can explode.
Never over-discharge a lithium polymer battery (below 2.7V per series cell).
Always use an electronic speed controller (ESC) with a low voltage cut off
feature.
Never attempt to charge a lithium polymer battery if it looks bloated, damaged,
over discharged (below 2.7V per series cell). Damaged packs can explode.
Never leave a lithium polymer battery unattended while charging.
Always charge lithium polymer batteries on a non-flammable surface and away
from flammables.
Take extreme caution around the UAV in the case of a crash. The battery packs
may explode if damaged.
Never discharge a lithium polymer battery at more than the published discharge rate.
The pack may explode if discharged too quickly.
4.4.4 ENVIRONMENTAL CONCERNS
All waste materials will be disposed of using proper trash receptacles
Consideration of environmental ramifications will be made regarding applicable
activities
The following materials have been identified as potentially hazardous:
o Aeropoxy 2032 Epoxy Resin
o Aeropoxy 3660 Hardener
o Lithium Polymer Batteries
See CDR-MSDS document for complete MSDS specifications on these and other
materials.
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5 ACTIVITY PLAN
5.1 BUDGET PLAN
Since the completion of CDR the MIT Rocket Team is still awaiting final funding awards
from the Massachusetts Institute of Technology department of Aeronautics and
Astronautics, as well as the Massachusetts Institute of Technology Edgerton Center.
Talks with these two sources are ongoing, and final decisions from these two groups
are expected by the first week of February. The MIT Rocket Team is also investigating
possible funding from two new sources. The first source being looked into is the MIT
Gordon Engineering Leadership program; this program is new on campus and is
offering funding to student engineering groups in return for allowing students within their
program to take part in the project throughout the semester. The MIT Rocket team is
currently waiting to submit a proposal as the program is currently assessing their
application process. The team is also interested in approaching Aurora Flight Sciences.
From research and contacts with a senior level design class the team has learned that
Aurora is interested in funding student-engineering projects that involve UAV
development as it is a field they are heavily invested in. The team has initiated contact
with a member of the Aurora Flight Sciences company and we are awaiting a response.
The MIT Rocket Team has also reevaluated the cost of our project taking into account
the changing design features of both the flight vehicle and UAV payload. At the
Preliminary Design Review many of the components of both the vehicle and payload
were priced using estimates with large margins. Since the maturation of the design the
Master Equipment List has been revised and we now have a more accurate measure of
system value as it sits on the launch pad. A table of contributions and a summary of
system costs are shown below in tables XX and XX
TABLE 5-1: BUDGET
Funding Sources
Source Contribution
MIT Aero-Astro ($7,000)
MIT Edgerton Center ($5,000)
NASA SMD Grant $5,000
MIT RT Savings $5,000
Total ($22,000)
Note: values in parenthesis are anticipated funding sources
TABLE 5-2: TOTAL BUDGET
System Summary
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System
Mass (kg) Cost (USD)
Rocket
Propulsion 5.46 552.00
Airframe-Body 3.62 455.09
Airframe-Fairing 1.01 27.00
Avionics/Comm 0.99 947.38
Payload Support Equipment 1.82 152.24
Recovery 2.19 434.60
SUBTOTAL 15.09 2568.31
UAV
Propulsion 1.15 190.00
Airframe 0.00 80.00
ACS 0.16 70.00
Avionics 0.21 647.73
Payload 0.15 622.30
Recovery 0.00 10.00
SUBTOTAL 1.67 1620.03
Support
Ground Station N/A 5,459.90
Testing N/A 2,000.00
Spares N/A 4,000.00
Team Support N/A 4,260.00
TOTALS 16.76 19,908.24
5.2 TIMELINE
The Gantt chart that has been in use since the initial proposal is still being used by the
MIT Rocket Team as the main method of keeping the team on target for completion.
Some small changes have occurred due to the fluid nature of student schedules during
the month of January. However, the team is still on track for completion of all
milestones. The updated Gantt Chart can be found in Appendix 2.
5.3 EDUCATIONAL ENGAGEMENT
We would like to build upon out fall outreach event with three events in the spring, which
will be shaped by the feedback we received.
The team has committed to holding three community outreach events over the next few
months in order to continue to engage the surrounding community in hands-on learning.
These programs will accommodate a range of ages, and will assume no prior
knowledge of rocketry. Our aim is to get the public, particularly middle to high school
students, excited about science, engineering, and the space industry.
MIT Splash and Spark Weekends
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MIT‟s Educational Studies Program is a student group that offers services to student
and community members alike. As part of its community outreach it offers student-
taught classes all weekend long during the months of November (called Splash) and
March (called Spark) on campus to a target group of 7th-12th graders.
At Splash, we taught two classes for a total of 25 high school students called “Rockets
and Composites”. We began with a brief lecture on composites, and then guided the
groups through the constructions of their own balsa wood and carbon fiber fins, which
they were able to take home with them. We then moved on to an introduction to rockets,
and further discussion on composites design, fabrication, and testing, and engine
testing.
We handed out a feedback form at the end of the session, which posed a series of
short-answer questions (what did you like most, what improvements could be made,
etc) as well as some multiple choice questions, to gage how the class impacted their
understanding of rocketry as well as their excitement about the subject.
The students reported that they enjoyed the hands-on part the most (making the fins),
and talking to us personally about their interests and any questions they happened to
have. Almost all of them suggested that the class be made longer. While this is not
possible due to the nature of the event, it did make us think about how we could use our
time better, and for the second group we cut down on our first lecture, which left more
time for questions at the end and made us feel less rushed. Most students reported
knowing “a bit” about rockets before the class, and that the class enhanced their
understanding “a lot”. Most said that the class “definitely” made them want to learn
about rockets in the future. Finally, almost all reported that the math and the equations
were the most confusing part, so we cut down on this for our next group.
For Spark, the corresponding spring event, we plan to use a presentation similar to that
given as Splash. Our survey results from Splash will help us to be more effective.
Boston Museum of Science
The MIT Rocket Team is a subset of the MIT chapter of a national organization called
the Students for the Exploration and Development of Space (SEDS), which aims to get
undergraduates involved in space-related projects. These projects often have a
community service side, and in the past the group has organized highly successful
community workshops and presentations at the Boston Museum of Science, where
undergraduate and graduate students conduct hands-on activities for the purpose of
increasing public interest in math, science and higher education.
We are working with the museum‟s Current Science & Technology group to run an
event on February 5th, which will consist of two presentations with hands-on activities
and demonstrations. The two presentations will last 15-20 minutes; the first will be about
rocket technology and its development over time, and another will be about the space
industry. We will run two hands-on workshops before, during, and after the
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presentations: one where participants will build their own Alka Seltzer bottle rockets,
and another where participants will design build their own parachutes. We will present
them with a set of materials, such as paper and plastic and cloth, and have them attach
a weight and observe the results.
We are currently working on securing space to set off the rockets and drop the
parachutes. This event will accommodate the wide range of age and experience that the
museum attracts.
We will be able to promote this event through various student websites and public radio,
as well as the museum‟s public relations personnel. We also plan to distribute posters
and flyers around the museum.
The learning objectives for this activity will be the following:
Arrive at a basic understanding of the history of rocketry and the space
industry. We believe that understanding rocketry requires learning about its
development, which includes the figures and organizations that have been key to
the field. Topics will include Wernher von Braun, NASA, the Space Race, and
current commercial organizations such as SpaceX.
How does a rocket work? We want to explain how rockets work, and get our
audience interested in math and science through the amazing technology that is
rocketry. This portion of the presentation will introduce the importance of math
and science in developing rockets by explaining the basics principles that allow
us to send rockets into space. We believe in learning by doing, and therefore will
supplement this with a bottle rocket workshop.
The social impact that low-Earth orbit rocketry has brought to our everyday
lives. We want our audience to leave with an understanding of not only how
rockets work, or how the space industry developed over time, but why all of this
is important and relevant to modern society, and to each of their daily lives. We
will explore rocketry‟s contributions to, for example, telecommunication and
accurate weather forecasting.
To evaluate the success of our engagement, we plan to conclude each talk with a
question session, and consider the accuracy and enthusiasm of their engagement. We
also plan to have questionnaires at each of the workshop tables, asking participants for
their feedback on the activity and the event as a whole.
MIT Museum
The MIT Museum is holding an event called the Cambridge Science Festival (CSF) from
April 30 to May 8. The mixture of talks and hands-on activities and workshops appealed
to our team, and we felt that our presentation would be a good fit for the venue. We
have worked with the museum‟s Educational Coordinator to secure a date and finalize
an event description, reproduced below. The event will be held on May 1, and will last
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from 10am to 1pm. The audience will be of all ages. We plan to run the event identically
to our program at the Boston Museum of Science, but will update our talks and
workshops based on the feedback we receive at the Museum of Science.
You Can Be a Rocket Scientist
How do rockets work? How do they affect our daily lives? Explore the history of
rocketry from before NASA to present-day and try your hand at aerospace
engineering with demonstrations, presentations, and hands-on activities by the
MIT Rocket Team.
6 CONCLUSION
From its inception, the MIT Rocket Team has known that project Valhalla would be a
difficult and aggressive challenge. However by taking a structured and thorough
approach to the development of the project, the team is confident that the project will
succeed. Upon its completion, the team will have developed a sophisticated
reconnaissance system capable of being employed in a variety of situations and
mission types. In addition to being rapidly deployable, the system benefits from the
simplified flight control interface already in the final stages of development. Thanks to
the simple “point-and-click” nature of the ground based flight control system, the skill
needed to successfully deploy the system is decreased, leading to a greater success
rate. Furthermore, to demonstrate the adaptability of the UNICRON flight system, we
have chosen to undertake the 2011 USLI Science Mission Directorate, which will
demonstrate the craft‟s performance in a science-gathering mission.
In the time since the completion of the Preliminary Design Review, the MIT Rocket
Team has taken great strides in the design and early testing of project Valhalla.
Furthermore, as the designs of the key components matured, design specifications
were adapted to simplify the manufacture process, while simultaneously increasing the
overall safety of our system. As such the designs of the launch vehicle, TYR, and the
payload, UNICRON, have both undergone a considerable number of revisions. And with
the analysis shown above the MIT Rocket Team is confident that we will be delivering a
system that exceeds the minimum safety requirements of NASA‟s University Student
Launch Initiative, while excelling at the mission for which project Valhalla was designed.
In the coming month, the MIT Rocket Team will continue fabrication and testing for
project Valhalla starting with the proto-flight models of both TYR and UNICRON. By
mid-February both proto-flight structures will be completed and ready to undergo further
testing to confirm the safety and viability of the final designs. Furthermore the
UNICRON avionics subsystem will begin flight-testing in the last week of January using
a standard model aircraft, and transitioning to the prototype versions of UNICRON as
soon as possible. The planned deployment method, MAGIC, will be undergoing testing
in mid-February using a balloon platform drop test. Finally all aspects of project Valhalla
will be tested with a scale launch and UAV deployment by the end of February.
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The MIT Rocket Team is looking forward to the coming month and the challenges it
presents. But more importantly the team is looking forward to April‟s launch weekend
where we look forward to a successful launch and mission completion. Until then, the
team will keep working hard to meet its deadlines, and further refine our system.
