PROJECT MANAGEMENT & DESIGN II
MAE 435
Rocket Project
March 4th, 2015
Group Members:
Nathan Akers
Chad Bagley
Paul Campbell
Class Supervisors:
Dr. Sebastian Bawab
Dr. Colin Britcher
Mr. Michael Polanco
Charles Juenger
Brian Mahan
Chris Skiba
Michael Weber
Michael Wermer
Project Advisor:
Dr. Thomas Alberts
2
Contents
Nomenclature .................................................................................................................................. 4
1 Abstract ....................................................................................................................................... 5
2 Introduction ................................................................................................................................. 6
3 Structures .................................................................................................................................... 6
3.1 Introduction .......................................................................................................................... 6
3.2 Completed Methods ............................................................................................................. 6
3.3 Proposed Methods ................................................................................................................ 9
4 Propulsion ................................................................................................................................... 9
4.1 Fuel and Combustion Chamber Design .......................................................................... 10
4.1.1 Introduction ................................................................................................................. 10
4.1.2 Completed Methods ..................................................................................................... 11
4.1.3 Proposed Methods ....................................................................................................... 13
4.2 Nozzle Design .................................................................................................................... 14
4.2.1 Introduction ................................................................................................................. 14
4.2.2 Completed Methods ..................................................................................................... 15
5 Aerodynamics ........................................................................................................................... 17
5.1 Introduction ........................................................................................................................ 17
5.2 Completed Methods ........................................................................................................... 17
5.3 Proposed Methods .............................................................................................................. 18
6 Avionics .................................................................................................................................... 19
6.1 Introduction ........................................................................................................................ 19
6.2 Completed Methods ........................................................................................................... 19
3
6.3 Proposed Methods .............................................................................................................. 19
7 Results ....................................................................................................................................... 21
8 Discussion ................................................................................................................................. 22
9 Appendix ................................................................................................................................... 24
9.1 Additional Figures .............................................................................................................. 24
9.2 Gantt Chart ......................................................................................................................... 30
9.3 Budget ................................................................................................................................ 31
10 References ............................................................................................................................... 32
List of Figures
Figure 1: Rocket Components ........................................................................................................ 7
Figure 2: Rocket Assembly............................................................................................................. 8
Figure 3: Thrust vs Time............................................................................................................... 12
Figure 4: Combustion Chamber with Convering-Diverging Nozzle ............................................ 15
Figure 5: Velocity Contour ........................................................................................................... 16
Figure 6: Temperature Contour .................................................................................................... 17
Figure 7: Coefficient of Drag vs Mach Number ........................................................................... 18
Figure 8: Fuel Performance .......................................................................................................... 21
Figure 9: Predicted Flight Performance ........................................................................................ 21
Figure 10: Rocket Nozzle Data ..................................................................................................... 22
Figure 11: Dual Deployment Wiring Diagram ............................................................................. 24
Figure 12: Avionics Bay Arrangement Sketch ............................................................................. 25
Figure 13: Rocket Motor ............................................................................................................... 26
4
Figure 14: Rocket Upper Body ..................................................................................................... 27
Figure 15: Avionics Bay ............................................................................................................... 28
Figure 16: Rocket Lower Body..................................................................................................... 29
Nomenclature
5
1 Abstract
The goal of the project is to design a high-powered rocket capable of taking a cosmic ray
detector payload to an apogee of 10,000 feet, deploying the payload, and being recovered.
Detailed designs of the structure, propulsion system, aerodynamic fins, avionics, and recovery
system were developed using computer aided drafting software (CAD). Structural stability will
be determined through finite element analysis (FEA). The propulsion system will be powered by
a potassium nitrate (KNO3)-dextrose (C6H12O6) solid propellant that will undergo combustion
and expansion through a converging-diverging nozzle, delivering up to 5,120 Newton-seconds of
impulse. The expected thrust for the fuel was calculated and an optimized nozzle was designed
for the fuel. Aerodynamic analysis was performed through modeled computational fluid
dynamics (CFD) to determine coefficient of drag versus Mach number. Scaled rocket supersonic
wind tunnel testing will be performed to determine the accuracy of the CFD analysis. The
avionics system will autonomously record in-flight data and control the deployment of the
payload, the drogue and main parachutes, ensuring safe recovery after flight. Fundamental
vibration analysis will be performed to minimize in-flight vibrational effects on the payload and
avionics system. All analysis and testing performed on the rocket will allow the team to extract
the necessary information to validate the design and ensure a safe and reliable flight.
6
2 Introduction
Rockets are used to carry payloads into orbit, but are expensive and difficult to design.
They need to be designed for the type, size, and weight of the payload to be carried. They need to
be designed for the desired payload deployment height. The design requirements of this project
are to design a high-powered rocket capable of deploying a five pound cosmic ray detector at
10,000 feet. The rocket design was divided into four key areas necessary to satisfy the goal of the
project. These include: structures, propulsions, aerodynamics, and avionics.
3 Structures
3.1 Introduction
A high powered rocket is composed of cylindrical body, nose cone, avionics bay, tail
fins, motor assembly, and recovery system. CAD software is used to create an accurate visual
representation of the rocket and its components. The representation is used to perform analysis
that determine design feasibility prior to construction [1]. FEA is a method for determining stress
and strain properties on complex solid bodies due to applied forces [2-4]. FEA is used on the
rocket assembly to design for structural integrity and to verify material selection [5]. The
analysis is needed to determine if the rocket design is capable of successful flight. Every new
rocket design needs to be tested for structural integrity.
3.2 Completed Methods
A 3D model of the rocket was created using Autodesk Inventor (Autodesk Inc., San
Rafael, CA) and was developed by designing part assemblies of the different rocket components.
These components (Figure 1) were then joined together to produce the main rocket assembly
(Figure 2).
7
Figure 1: Rocket Components
The first part assembly to be designed is the avionics bay (Figure 15), which houses the
electrical equipment and connects the top and lower portions of the rocket. The ends are fitted
with two circular end caps and are secured to the tubular housing by two threaded rods with
accompanying fastener hardware. Eyebolts are fitted to the center of the two end caps by
fasteners. These bolts serve as an anchor point for the shock chords after in flight rocket
separation has occurred. The nose cone (Figure 14) is designed with a hollow cavity surrounded
by the nose cone walls. A circular end cap with a U-bolt and the accompanying fasteners are
fitted to the end of nose cone. The U-bolt serves as another anchor point for the top shock chord.
Connecting the nose cone and the avionics bay is the upper rocket body (Figure 14). The lower
8
rocket body is a tube fitted with six fins (Figure 16). The engine assembly is made up of a
nozzle, outer casing, resin liner and end cap (Figure 13). This is attached to the lower rocket
body using six circular motor mounts attached to inner walls of the bottom rocket body.
Connecting the avionics bay to the nose cone are Kevlar® shock chords. Connecting the nose
cone to the upper body and the avionic bay to the lower body are three shear pins for each
connection. The avionics bay is connected to the upper rocket body using three riveted joints.
Figure 2: Rocket Assembly
9
3.3 Proposed Methods
The FEA will be carried out through the application of NASTRAN In-Cad (Autodesk
Inc., San Rafael, CA). The test to be performed will consider the outer rocket body and the
stress put on it from thrust. The maximum thrust force limit will be applied during the FEA
analysis.
Stress testing will be carried out on the shear pins to determine the force required to cause
shear. A testing fixture will be constructed with two interlocking metal tubes that form a
couple. A bulkhead will be fixed to each end of the couple with an eyebolt running through the
center of each. Three equally spaced 5/64-inch holes will be drilled through both pipes in the
couple. After the three shear pins are inserted into the corresponding holes in the couple, the test
fixture will be secured to the MTS Alliance RF/300 electromechanical load frame in the ODU
materials laboratory. The MTS machine will then be set to expand causing tension in the testing
fixture. The data will be recorded once the shear pins fail and the couple separates. The MTS
machine will be reset and a new set of shear pins will be secured into the rejoined couple. This
testing procedure will be repeated until six tests have been performed.
A test stand is under development that will be used to measure rocket thrust. The stand
will utilize an Interface load cell and be mounted in the vertical position. The test will use the
same motor that was designed for the rocket, but with half of the amount of fuel. The data will
then be used in conjunction with the fuel calculations to predict the amount of fuel that will be
needed to deliver the designed total impulse.
The rocket will be assembled in steps starting first with the motor mount. The fins will be
installed through the outer body of the rocket and be glued to the motor mount. Then the avionics
10
bay will be assembled. After the avionics bay is built, the upper sections of the rocket will be
assembled. The motor will be the last piece to be installed.
4 Propulsion
4.1 Fuel and Combustion Chamber Design
4.1.1 Introduction
Rocket motors are thrust engines, which operate by generating a pressure and momentum
thrust. This is achieved by combusting fuel inside a pressure vessel and expanding the
combustion products through a nozzle. This is typically accomplished in small-scale rockets by
using solid rocket fuels. Commercially available “high-power” hobby rocket motors are typically
made of Ammonium Perchlorate, HTPB and Aluminum, which is similar to the formulations
used in rockets designed for space flight as well as missiles. These formulations, commonly
referred to as APCP (Ammonium Perchlorate Composite Propellant), provide a high specific
impulse relative to other solid formulations. Specific Impulse is a key concept in the design and
of a rocket motor because it is used to describe the efficiency of a fuel based on its weight.
Specific Impulse is the ratio of the amount of thrust a fuel is able to produce to the weight flow
of the propellants [6]. One drawback to using an APCP is the production cost. The material cost
is high and the manufacturing process can be difficult. An alternative to APCP, which is
commonly used in homemade rockets, is Potassium Nitrate-Sugar propellant, sometimes referred
to as “R-Candy”. This formula is easy to manufacture using readily available materials. The
greatest drawback to using these “R-Candy” formulations is the relatively low specific impulse
they produce; meaning to reach a given altitude, more fuel is required. These formulations can be
modified using different sugars such as sucrose, sorbitol, or dextrose, which can impact the
material properties of the casted propellant, and can influence the ease of manufacturing, and the
11
reliability of the motor itself. While general formulations are available, the amounts and types of
fuel need to be modified in order to reach the target payload deployment altitude.
4.1.2 Completed Methods
The rocket fuel chosen is an “experimental” fuel, meaning that the motor is to utilize a
homemade propellant formulation in lieu of a commercially available motor. The formulation
was taken from Richard Nakka’s Experimental Rocketry Website [7]. The need to select a preinvestigated formulation was necessitated by the compressed budget and timeline of this project
as there were not enough resources available to develop a unique formulation which could
function safely and reliably. The propellant selected for this project was a mixture of 65% by
mass Potassium Nitrate (KNO3) and 35% by mass Dextrose (C6H12O6) and is seen below [8].
C6H12O6 + 3.31 KNO3  2.116 CO2 + 2.300 CO + 4.512 H2O + 1.424 H2 + 1.655 N2 + 1.585
K2CO3 + 0.133 KOH (Equation 1)
(theoretical combustion reaction at 68atm [8])
This formulation was tested extensively and is commonly used by high-power rocketry
hobbyists. This relatively inexpensive formula has been proven to be safe and reliable while
being relatively easy to manufacture, and produces results within a reasonable margin of error by
using the published values.
A preliminary motor design was developed using a “freeware” Microsoft Excel
(Microsoft Corporation, Redmond, WA) spreadsheet (“SRM_2014”) published by Richard
Nakka [9] to target a maximum operating pressure within the combustion chamber of 1100psia.
This target pressure was based off a preliminary motor casing design, to be constructed from
2.95” outside diameter, 6061-T6 Aluminum tubing with a wall thickness of 0.11”, giving an
inside diameter of 2.73”. The selected pressure of 1100psia used a conservative factor of safety
12
of 2, this conservative value was selected because of the need to re-use the motor casing multiple
times and the relatively low-impact of the casing weight on the rocket performance and design.
The size of our motor was scaled to deliver as close to 5120 Newton-seconds of impulse as
possible, without exceeding that limit. The desired pressure is achieved by modifying the “Kn”
value, which is the ratio of burning surface area of the propellant to the area at the throat of the
nozzle [10]. This was done by selecting a commercially available propellant casting mold and
modifying the nozzle diameter to meet the required “Kn” value.
Using the “freeware” Microsoft Excel (Microsoft Corporation, Redmond, WA)
spreadsheet (“SRM_2014”) published by Richard Nakka [9] and the fuel amount dictated by the
size of the combustion chamber, plots of Thrust vs Time (Figure 3) were obtained. Additionally,
the predicted altitude, flight time, and max velocity using a conservative coefficient of drag of
0.35 was estimated.
SRM 2014.XLS
Thrust vs time
500
450
400
Thrust (lbf)
350
300
250
200
150
100
50
0
0.0
0.5
1.0
Time (sec.)
Figure 3: Thrust vs Time
1.5
2.0
13
4.1.3 Proposed Methods
Now that the fuel has been selected it will be manufactured. The process of
manufacturing the fuel requires the two components to be thoroughly mixed together and then
heated until the dextrose melts, homogenously binding the two products together. The particles
must be desiccated and finely milled before mixing and heating to ensure the maximum
performance of the propellant. Errors in manufacturing can cause results which are significantly
less than predicted. The particles will be mixed together using a rotating drum or other low-speed
mixture to ensure an evenly distributed mixture. The dextrose will fully melt at 123°C [7], in
order to create the propellant; the particle mixture will be heated between 125-130°C. The sugar
will begin to caramelize at 157°C [7] so it is important to use a calibrated hot plate to precisely
control the temperature; caramelization of the sugar will negatively impact the performance of
the propellant. While heated, the mixture will be packed into a 2.562” inside-diameter,
cylindrical casting tube, placed around a cylindrical core with a diameter of 0.6875”. To form the
motor, six of these castings will be made at a length of 4.83”; each of these castings is known as
a grain. The completed motor will be comprised of six grains, totaling 28.98” long with an
outside diameter of 2.562”, and a core diameter of 0.6875”, for a total volume of 132in3.
After the fuel has been mixed, the rocket motor manufactured, and the test stand
manufactured, the motor will be test fired. Pressure will be measured using a PX-303 load cell
(Omega, Stamford, Connecticut) and the thrust data is measured using a PX-LCJA (Omega,
Stamford, Connecticut). The data will be collected from the static test using a DI-194 (DATAQ
Instruments, Akron, OH). The data from this test will be used to validate the software
simulations and used to adjust the fuel component percentages in order to achieve the required
thrust.
14
4.2 Nozzle Design
4.2.1 Introduction
In order to achieve maximum propulsive efficiency, a rocket motor must sustain a high
flow of heavy particles and elevated combustion temperatures while maximizing the speed of
gases through the nozzle exhaust such that the exit pressure is equal to the atmospheric pressure
[11, 12]. Exponentially increasing temperatures in the combustion chamber can cause rapid
deformation of the materials of the nozzle and chamber [11]. Combustion gas expansion induces
high internal pressures within the combustion chamber, which further exacerbates component
deformation. Gases that travel through the nozzle carry burnt aluminum particles, which cause
ablation to the throat section of the nozzle and agglomeration in the nozzle cone [13-15]. This
results in decreased efficiency by altering the nozzle inner dimensions. Additionally, heat
transfer in the chamber must be controlled, as heat conducted through the chamber walls results
in heat loss to the combustion gases [16]. The inability to accommodate these operating
conditions results in performance inefficiencies as well as potential component failure.
Determination of a nozzle design, which accommodates maximum operating conditions, will
increase performance and longevity while reducing cost. The nozzle dimensions must be
individually designed for every rocket because of unique fuel.
15
4.2.2 Completed Methods
The nozzle (Figure 4) was assumed to be isentropic, and compressible flow equations
were used to determine the thermodynamic properties at the subsonic, sonic, and supersonic
regions. The properties at each station and the constraint of three inches on the converging
section were used to determine the dimensions of the nozzle. To reach maximum propulsive
Figure 4: Combustion Chamber with Convering-Diverging Nozzle
efficiency, the exit gases must be expanded through the diverging section of the nozzle so that
Pexit equals Patm. The mass flow rate in the nozzle was determined by
(Equation 2) where
g -1
2g RTc é Pe ù g
Ve =
ê1- ú
g -1 ë Pc û
(Equation 3). Using the mass flow rate and exit velocity, the throat diameter
(Equation 4) and 𝑋 ∗ is a function of the ratio
was determined using the equation
g +1
æ 2 ö 2(g -1)
of specific heats X * (g ) = g ç
(Equation 5). Next, the nozzle exit diameter was found
÷
è g +1 ø
16
g +1
é
ù
2(g -1)
æ
ö
2 ê Pc
by first finding the Mach number at the exit by M e =
-1úú (Equation 6).
ç ÷
ê
g -1 è Pe ø
êë
úû
Substituting the Mach number into the equation for area ratios provided the diameter for
g +1 ö
æ
2(g -1)
é
ù
æ
ö
æ
ö
4
1
2
g
-1
ç
÷
the exit De = A* ç
M e2 ÷ú
êç
֍1+
÷ (Equation 7). A simulation of the nozzle
øû
p ç M e ëè g +1 øè
2
÷
è
ø
was run at the determined area and pressure ratios using the CD Nozzle Simulator (engApplets,
Blacksburg, VA).
Using Simulation CFD (Autodesk Inc., San Rafael, CA) a CFD analysis was performed
of the nozzle with the inlet conditions found in the fuel analysis. The assumptions were a fluid of
air, compressible flow, sea level standard air conditions, steady flow, and no heat transfer. The
CFD analysis supplied velocity (Figure 5) and temperature (Figure 6).
Figure 5: Velocity Contour
17
Figure 6: Temperature Contour
5 Aerodynamics
5.1 Introduction
Aerodynamics is examined to reduce drag and shock effects. Shapes of nose cones and
fins have been examined in an effort to reduce drag and shock effects [17-22]. Nose cone designs
showed that increased ratios of nose cone length to body length decreased drag, and reduction of
the angle of the nose cone resulted in a smaller shock transition between the nose and the rocket
body during supersonic flight [18, 19]. Fin designs were primarily focused on the flight stability,
but as the size and bluntness of the leading edge increased, the drag and shock effects increased
[17,20-22]. While the basic aerodynamics of fins and nose cones are known, the final fin designs
and nose cone designs are modified for every rocket based upon mass, and length of the rocket in
order to move the center of pressure below the center of mass. This ensures flight stability.
5.2 Completed Methods
The software chosen to perform CFD analyses on the fins and nose cone was Simulation
CFD (Autodesk Inc., San Rafael, CA). The assumptions for the flow to be used in the CFD
18
analyses were a fluid of air, compressible flow, sea level standard air conditions, steady flow,
and no heat transfer. The inputs were Mach numbers from 0.1 to 1.2 in step sizes of 0.1. The
output criteria to judge the different designs were coefficient of drag and center of pressure. The
Coefficient of Drag vs Mach number are plotted below (Figure 7).
Cd vs Mach
0.7
0.6
Cd
0.5
0.4
0.3
0.2
0.1
0
0
0.2
0.4
0.6
0.8
1
1.2
1.4
Mach
Figure 7: Coefficient of Drag vs Mach Number
5.3 Proposed Methods
A model rocket will be used to test in the supersonic wind tunnel at Old Dominion
University. A supersonic wind tunnel test was chosen because it will be inexpensive and easier
to perform. This is because the models are smaller and there are a few model rockets already
available for choosing, but the fins need to be modified. The model will also be threaded for
attachment to the wind tunnel apparatus. A Mach number of 1.8 will be used since that is the
lowest speed this wind tunnel can operate. A CFD analysis will be done at a Mach of 1.8 to
ensure that the CFD program used is accurate. Schlieren photography will be used to see the
flow of fluid and shockwaves across the nose cone.
19
6 Avionics
6.1 Introduction
The avionics system will consist of a redundant, dual-deployment altimeter system capable of
ejecting the drogue and main parachutes. The typical wiring diagram for dual deployment altimeters is
seen in the appendix (Figure 11). The dual-deployment altimeters will be programmed to deploy two
separate parachute ejection charges that will allow failure of the rocket airframe shear pins. The ejection
charges and airframe shear pins will be sized using common rocketry calculators [23] and experimental
bench testing. The first ejection charge will cause shear pin failure of the rocket’s forward airframe and
will release the drogue parachute and cosmic ray detector payload. The drogue parachute will be deployed
at the rocket’s apogee and used to decelerate and stabilize the descent of the rocket. When the drogue
parachute is deployed the payload will be ejected. The second ejection charge will cause shear pin failure
of the rocket’s rear airframe and will deploy the main parachute at approximately 800 feet above the
ground.
6.2 Completed Methods
The electronics, electrical mounting structures, parachutes, shock cords, and rocket
airframe shear pins have been ordered. A conceptual sketch has been developed for the
arrangement of the avionics bay section of the rocket and is in the appendix (Figure 12).
Software Simulation Mechanical (Autodesk Inc., San Rafael, CA) has been acquired and
will enable the team to perform the vibration and stress analysis using the rocket model. To gain
an understanding of how the software can be used, the team has imported the Inventor® model
of the avionics bay of the rocket’s airframe and have ran a preliminary design analysis report.
6.3 Proposed Methods
Once construction of the avionics bay is complete, the bay will be fully modeled in
Autodesk Inventor and incorporated into the rocket model. The rocket model will then be
20
imported into Autodesk Simulation Mechanical (Autodesk Inc., San Rafael, CA) to determine
the effects of vibration and stress on the avionics bay section of the rocket. Using Autodesk
Simulation Mechanical (Autodesk Inc., San Rafael, CA), constant random vibration analysis
(modal superposition) will be conducted. The results provided from the analysis are the root
mean square response of the displacement is equal to the standard deviation. From this analysis,
a design analysis report will be generated and included in the final report. Based on the results
from the analysis, action will be taken to mitigate the effects of vibration on the electronics (e.g.
adding rubber isolators to the electronic mounting points).
Ejection charge and shear pin sizing will be performed, first using commonly available
calculators used in model rocketry [23]. Data obtained from these calculators will be used as a
baseline to size the ejection charges and shear pins for experimental testing. Bench tests will be
performed on the assembled rocket testing the rocket’s separating force. Bench testing will
verify the charges are sized appropriately, allowing failure of the rocket airframe shear pins
during flight.
The manufacturer of the dual-deployment altimeters provides free downloadable software
(mDACS – Missle Works Data Acquision and Configuration Software) that will allow the team
to extract detailed dynamic data captured during flight. Flight data will be used to verify the
performance of the rocket after flight and will allow the team to make adjustments to the rocket
to achieve the rocket’s performance objectives. This software will be downloaded prior to flight,
allowing the team to become familiar with its functions and capabilities.
21
7 Results
A preliminary rocket design was completed based upon commercially available rocket
designs. These include the nose cone, avionics bay, rocket outer body, motor mount, fins and
engine (Figure 1).
The rocket fuel propellant was chosen from a formulation commonly used by rocketry
hobbyists. Using this fuel the combustion chamber pressure reached the desired value of
1100psia given a factor of safety of 2. The fuel performance results are seen in Figure 8.
Fuel Diameter (in)
Max Operating Pressure (psi)
2.493
1050
Max Thrust (lbf)
460
Average Thrust (lbf)
414
Total Impulse (lbf-sec)
Burn Time (s)
Specific Impulse (s)
780
1.89
134
Figure 8: Fuel Performance
Using the fuel performance and a conservative coefficient of drag of 0.35, which is greater than
that predicted by the CFD analysis, the flight performance was estimated (Figure 9).
Peak altitude
Time to peak
altitude
Max velocity
Burnout altitude
Z peak =
9593
feet
t peak =
V max =
or V max
=
Z bo =
23.6
964
sec.
feet/sec.
657
886
MPH
feet
Figure 9: Predicted Flight Performance
Predicted (with
drag)
22
The rocket nozzle dimensions and properties at the inlet throat and exit are seen below
(Figure 10). Additionally no shock formed inside the nozzle because there are no sudden drops in
velocity (Figure 5).
Diameter (in)
Area (in2)
Mach
Velocity (ft/s)
Pressure (psi)
Temperature (F)
Gamma
mdot (lbm/s)
Rocket Nozzle Data
Inlet
2.4
4.53
1131
2465
1.1308
4.453
Throat
0.625
0.310
Exit
2.25
4.00
1
1965
655.4
2285
1.1308
4.453
3.24
5049
14.67
1310
1.1308
4.453
Figure 10: Rocket Nozzle Data
8 Discussion
The purpose of this project was to design high-powered rocket capable of carrying and
deploying a cosmic ray detector at 10,000 feet. This was done by designing the structure,
propulsions, aerodynamics, and avionics sections of the rocket.
The results focused mainly upon the propulsion and fuel design. Analysis of the fuel
resulted in a predicted max altitude of 9593 feet. This was done with a conservative coefficient
of drag which was higher than the coefficient of drag predicted by the CFD analysis, so the
rocket will exceed the maximum predicted altitude. A conservative coefficient of drag was used
due to the inability to determine the accuracy of the program used to perform CFD analysis.
Additionally, the CFD analysis was run assuming sea level standard conditions which will not be
the case as the rocket’s altitude changes. However, a CFD analysis will be run at Mach 1.8 and
compared with supersonic wind tunnel testing to determine the accuracy of the CFD program.
23
In addition to the wind tunnel testing, future work includes manufacturing the fuel,
building the rocket, building a test stand, performing a vibration analysis on the avionics section,
determining the force and amount of black powder needed to shear the shear pins, and
performing test burns with the combustion chamber and nozzle on the test stand to ensure the
accuracy of the thrust calculations, and testing the recovery system.
24
9 Appendix
9.1 Additional Figures
Figure 11: Dual Deployment Wiring Diagram
25
Figure 12: Avionics Bay Arrangement Sketch
26
Figure 13: Rocket Motor
27
Figure 14: Rocket Upper Body
28
Figure 15: Avionics Bay
29
Figure 16: Rocket Lower Body
30
9.2 Gantt Chart
31
9.3 Budget
32
10 References
[1]
W. A. D. Michael T. Heath, "Virtual Prototyping of Solid Propellant Rockets," Computer
Science & Engineering, vol. 2, pp. 21-32, 2002.
[2]
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and Exposition, 2006, June 18, 2006 - June 21, 2006, Chicago, IL, United states, 2006, p.
Dassault Systemes; HP; Lockheed Martin; IBM; Microsoft; et al.
[3]
J. F. Dues Jr, "Stress analysis for novices using autodesk inventor," in 2006 ASME
International Mechanical Engineering Congress and Exposition, IMECE2006, November
5, 2006 - November 10, 2006, Chicago, IL, United states, 2006.
[4]
W. Younis, "CHAPTER 9 - The Stress Analysis Environment," in Up and Running with
Autodesk Inventor Simulation 2011 (Second edition), W. Younis, Ed., ed Oxford:
Butterworth-Heinemann, 2010, pp. 235-275.
[5]
S. Ping, "Dynamic Modeling and Analysis of a Small Solid Launch Vehicle," in
Information Engineering and Computer Science, 2010.
[6]
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November 2014].
[7]
R. Nakka, "KN - Dextrose Propellant," Richard Nakka's Experimental Rocketry Web
Site, 15 July 2006. [Online]. Available: http://www.nakka-rocketry.net/dex.html.
[Accessed 19 November 2014].
[8]
R. Nakka, " KN-Dextrose Propellant Chemistry and Performance Characteristics,"
Richard Nakka's Experimental Rocketry Web Site, 10 December 2000. [Online].
Available: http://www.nakka-rocketry.net/dexchem.html. [Accessed 19 November 2014].
33
[9]
R. Nakka, " Rocketry Software," Richard Nakka's Experimental Rocketry Web Site, 18
October 2014. [Online]. Available: http://www.nakka-rocketry.net/softw.html. [Accessed
19 November 2014].
[10]
J. S. DeMar, "The Propellant:Nozzle Area Ratio - A Practical Guide to Kn,"
ThrustGear.Com, 2 February 2007. [Online]. Available:
http://www.thrustgear.com/topics/Kn_Notes.htm. [Accessed 3 November 2014].
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