The Perfect Game
Team 11 Project Technical Report for the 2021 LASC
Sanket Zambare1, Tanmay Kanmahale2, Piyusha Patil3, Sanket Pagadpalliwar4, Abhirav Gote 5, Akshay Raghuvanshi 6
, Atharva Pingle 7, Girija Wale8, Pratik Bhusari9, Saakshi Sule10, and Sarthak Zende11
Sinhgad Technical Education Society, Pune-MH, India-411041
Approved by
Dhananjay Khankal12, and Gaurav Sandeep Dave13
Sinhgad Technical Education Society, Pune-MH, India-411041
This project details the design of STES Rocketry’s Level 2 HPR, to compete in the ½ Km
Height Target Solid motor category at 2021 Latin American Space Challenge. This rocket,
Hero, is a 4.7 ft tall fiberglass composite rocket equipped with a J425 38mm Cesaroni Solid
Motor. The rocket is designed to carry a 300g payload to ½ Km height target and get it back
safely with the data used to interpret rockets in flight dynamics. Modeling and analysis were
conducted on the entirety of the rocket. The research was performed on composites,
aerodynamics, avionics, and recovery aspects of Hero. Current models project Hero to reach
an apogee of 525m.
1
Systems Engineering, Operations, sanketzambre1@gmail.com
Aerostructures Lead, Aerostructure and Analysis, kanmahaletanmay@gmail.com
3
Avionics Lead,Avionics, ppiyusha2001@gmail.com
4
Payload Lead, Payload, sanketrp123@gmail.com
5
Recovery Co-Lead, Recovery, anhiravgot133@gmail.com
6
Propulsion Co-Lead, Propulsion,akraghuvanshi07@gmail.com
7
Aerostructures Analyst, Aerostructures and Analysis, atharva99pingale123@gmail.com
8
Recovery Co-Lead, Recovery, girijawale432@gmail.com
9
Avionics Technician, Avionics, pratikbhusari77@gmail.com
10
Avionics Technician, Avionics, saakshi535@gmail.com
11
Aerostructures Co-Lead, Aerostructures, sarthakzende379@gmail.com
12
Faculty Advisor, Department of Production Engineering, dhananjaykhankal@sinhgad.edu
13
Faculty Advisor, Department of Mechanical Engineering, gsdave_skncoe@sinhgad.edu
2
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Latin American Space Challenge
To provide a project-based learning experience for the overall development of aspiring space
pioneers.
To make space for everyone.
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Nomenclature
A
a
Cp
Cx
Cy
c
dt
Fx
Fy
f, g
K
R
ρ
𝛿
SRT
STES
COTS
SRAD
SE
CP
CG
PCB
UI
BOM
CAD
AGL
LV
L3
DMP
RSRM
GFRP
RRC3
DAQ
STEM
HPR
=
=
=
=
=
=
=
=
=
=
=
=
Area
cylinder diameter
pressure coefficient
force coefficient in the x-direction
force coefficient in the y-direction
chord
time step
X component of the resultant pressure force acting on the vehicle
Y component of the resultant pressure force acting on the vehicle
generic functions
trailing-edge (TE) nondimensional angular deflection rate
resistance
= resistivity
air density
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
=
STES Rocketry Team
Sinhgad Technical Education Society
Commercial off the shelf
Student Researched And Developed
Systems Engineering
Centre of Gravity
Centre of Pressure
Printed Circuit Board
User Interface
Bill of Materials
Computer-Aided Design
Above Ground Level
Launch Vehicle (i.e. Hero)
Level 3
Digital Motion Processor
Reloadable Solid Rocket Motor
Glass Fibre Reinforced
Rocket Recovery Controller 3
Data Acquisition System
Science, Technology, Engineering, and Mathematics
High Power Rocket
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Latin American Space Challenge
I. Introduction
About us
STES Rocketry is a student-run rocket engineering project group under Sinhgad Institutes, Pune-India. The
team comprises 14 undergraduate students from various engineering disciplines working as one unit on designing
experimental HPR and research on its various subsystems. The team has been participating in the Intercollegiate
Rocket Engineering Competition at New Mexico- USA for 2 flight events and 1 virtual event designing L3 HPRs.
With The Perfect Game, the team will be participating in the Latin American Space Challenge for the first
time with all new members with their competition rocket, Hero, in the ½ km height target entry-level category. Due
to COVID restrictions, the team wasn’t able to complete the manufacturing but the design was made as complete as
possible on paper to be flight-ready when the time comes. Hero is designed such that it’s scalable and multiple parts
can be reused for bigger projects with little to no change.
The core principle of this team is to develop our members into talented engineers by providing them with
the necessary skills and tools to grow and succeed. Every day we aim to accomplish feats only thought possible in
the industry by incorporating professional design reviews, rigorous testing, and in-depth documentation of
individual systems, hardware, and software.
The bridge between the classroom and the workroom is one of the most significant components the team
gives to students. Students get the ability to apply and expand their engineering education through real-world design
and fabrication difficulties through SRT. Undergraduate student candidates from all Sinhgad Institutes Engineering
campuses build up the SRT. This means that freshmen on our team have the great opportunity to work on a project
with seniors and graduate students, and to participate in real, hands-on aerospace projects. Making rocket science a
more popular subject to work in and promoting STEM discipline among students is one of the team's main duties.
Team Organisation
The team is organized into two sections: management and technical, and is managed by a Project Manager.
Public Relations, Finances, Logistics, and Operations are all part of the management division. Sponsorship and
communication are handled by the PR team, while accounts and budgets are handled by the Finances department.
Test launches, Preliminary and Critical Design Reviews, and on-site arrangements at the launch venue are all
handled by the Logistics team. Finally, the Operations team is responsible for all technical elements of launches,
including buying competitive motors, setting up launch locations, and most importantly, preparing checklists and
procedures to assure a successful rocket launch. The Mission Supervisor is in charge of making sure the checklists
are followed and the rocket is ready to launch on time on launch days. The technical section is led by System
Engineering, which is composed of six sub-teams: Avionics, Payload, Propulsion, Recovery, Analysis, and
Aerostructures. Each one is responsible for their respective departments.
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Figure 1. Team Organisation Chart
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II. System Architecture Overview
1. Rocket Layout
The rocket's design is meant to be basic while still accomplishing the competition objectives. Components
include nose cone, body tubes, fins, and boat tail. The goal was to keep the mass of each component to a minimum
while keeping the strength required for the entire body because the structure's weight directly influences the rocket's
performance. An intensive material investigation was undertaken to attain the best strength-to-weight ratio.
Aluminum alloy, Aluminum 6061 T6, Glass Fiber Reinforced Plastic, and ABS Plastic were used for various parts
based on their design needs. The structures team is then allocated to each of these elements for further research and
development.
Figure 2: Hero Internal Configuration
2. Design Overview & Key Technical Specifications
Rocketeers who wish to evaluate the performance of a model rocket before building and flying it often use
commercially available model rocket simulation software such as an OpenRocket. The software computes and
simulates the aerodynamic properties of rockets, yielding a wide range of technical findings like CG-CP locations,
stability margin calibers, angle of attack, pitch and yaw moment coefficient, roll characteristics, flight position, and
flight duration. The flight objectives were as follows: an apogee of 500m carrying a payload of minimum mass
required 250g. On the basis of the rocket’s CAD model, the shape and mass distributions were replicated as
accurately as feasible. The simulation produced an apogee of 525 meters and a maximum velocity of 98.7 meters per
second (Mach 0.287). The peak drag coefficient was found to be 0.495 (Refer Appendix E-1 for Drag Calculation).
Specification
Value
Airframe Length
1460 mm
Diameter
104 mm
Expected Apogee
525 m
Max Velocity
98.7 m/s
Motor
Cesaroni Pro38 J425
Motor Impulse/ Average Thrust
785Ns/ 423N
Burn-time
1.85 sec
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A. Propulsion Subsystem
The design was made to accommodate large motors, up to 75mm, to reach higher height targets. Since our
KNSB motor project, Damocles, was halted due to COVID, a similar performing (in terms of total impulse) COTS
motor was selected for the virtual event that will help us reach the ½ km height target.
For the 2021-22 academic term the propulsion team has chosen a Cesaroni Pro38 Reloadable Solid Rocket
Motor (RSRM). The goal of the Solid motor is to deliver the vehicle to ½ km apogee. While simulating in
OpenRocket taking into account the weights of certain subcomponents was lighter, thus allowing us to select an
appropriate motor. The Cesaroni Pro38 J425 motor was chosen in this concept based on performance and trajectory
simulation. It is a 38 mm J-class motor with an average thrust of 423N, a max thrust of 720N, a total impulse of 785
N-s, and 1.85 seconds burn time.
Figure 3: Thrust Curve of COTS J425-7 Motor
B. Aero-structures Subsystem
B.1 Nose Cone
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Figure 4: Von-Karman Nosecone
The nose cone is made up of two parts: the airframe and the nose tip. The nose cone is a haack series
designed to minimize aerodynamic drag by employing the L-D Haack equation, also known as the Von-Karman
equation. This shape has a 3.84:1 fineness ratio with a diameter of 104 mm, 400 mm long, and a shape parameter of
0. The nose cone shoulder is a simple tubular design with 50 mm length and 100 mm outer diameter. The thickness
was set to 2 mm for both the airframe and shoulder. The nose cone airframe is made of GFRP.
B.2 Nose Tip
Figure 5: Aluminum Nose Tip
Aluminum is used to make the nose tip. The significant temperature rise during flight is the cause for the
metal nose tip’s debut. Machining composite material nose tip would be difficult hence, aluminum was chosen for
tip material because of its lightweight property, high availability, and ease of manufacturing.
B.3 Composites and epoxy
Hero consists of two composite fiberglass tubes, upper airframe and booster tube acting as the primary
aerodynamics and structural elements. Composite materials such as Carbon Fiber have a good weight-to-strength
ratio, yet are prohibitively expensive. Aluminum is too heavy for our use and costs are also very high due to the
diameter and length of our rocket. Thus, fiberglass was the solution employed to achieve both low economic costs,
and an appropriate strength-to-weight ratio. For the lamination of tubes, 1522 (plain weave) E-Glass fiber with
carbon black composites epoxy laminating resin as the substrate are used.
Material
Composition
E-Glass Fibre
54%SiO2-15%Al2O3-12%CaO
Property
Maximum Value
Density
2.6 Mg/m3
Energy Content
120 MJ/kg
Bulk Modulus
50 Gpa
Compressive Strength
5000 Mpa
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Ductility
0.028
Elastic Limit
2875 Mpa
Endurance Limit
3110 Mpa
Hardness
6000 Mpa
Poisson’s Ratio
0.23
Shear Modulus
36 Gpa
Tensile Strength
2050 Mpa
Young’s Modulus
85 Gpa
Thermal Conductivity
1.35 W/m.k
B.4 Upper Airframe
Figure 6: Upper Airframe
The upper airframe is tubular in design, with a length of 380 mm and an outside diameter of 104 mm,
which matches the outer diameter of all the tubes. The thickness is 2 mm, resulting in a 100 mm inner diameter. The
nose cone shoulder will be installed in the upper airframe. This tube will house the parachute coupler with recovery
subsystem and half avionics coupler.
B.5 Booster tube
Figure 7: Booster Tube
The rocket’s bottom or aft tube is 600 mm long with a 104 mm outside diameter, which matches the outer
diameter of all the tubes. The thickness is 2mm, resulting in a 100 mm inner diameter. This tube will house the other
half of the avionics coupler with avionics subsystems, fins, centering rings, stringers and motor. From the base of the
aft tube upward, three 158.44 mm long slits are made to allow the fins to slide through.
B.6 Boat tail
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Figure 8: Boat-tail
The bottom of the rocket ends in a conical transition called boat tail is used to reduce pressure drag. For
boat tail material GFRP was chosen because of its relatively low weight and good strength. Dimensions were
defined by using open rocket software. The transitional length of the boat tail is 75 mm with a front diameter of 104
mm and aft diameter of 80mm.
B.7 Couplers
Figure 9: Avionics Coupler
Two fiberglass Coupler tubes were used to house the parachute and avionics subsystems. The parachute
coupler is made up of recovery subsystems, and an avionics coupler also serves as a coupler, connecting the upper
airframe to the booster tube. The carbon fiber couplers are not chosen at random; they block signals from the
antennas inside the avionics bay, making live data transmission and system location impossible upon landing.
Because fiberglass allows RF signals to pass through it and also has the ability to resist buckling, it was determined
to be the best material for the specific application.
B.8 Fins
B.8.1 Shape Selection
The open rocket software was used to select the fin shape according to flight stability and material usage.
Back slanted fins were ruled out due to the potential of structural integrity collapse during landing. Round fins
produced the highest apogee, but the team decided against it because there may be a possibility that the rounded fins
will suffer fluttering due to sudden changes of air pressure during flight. The trapezoidal shape of the fins has a
number of benefits over other fin shapes like the trapezoidal shape produces low induced drag. Thus, the trapezoidal
shape of fins was selected for our rocket.
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B.8.2 Fin Assembly
Figure 10: Fin Assembly
The fin is fastened to stringers made of aluminum 6061 T6. An M5 button head screw connects the fins to
the stringers, while an M4 button head screw connects the stringers to the centering rings.
B.8.3 Number of fins
The team decided on three fins to make fabrication easier for the upcoming flight events. The drag
coefficient and stability rise as the number of fins increases. As a result, we need to choose the best quantity so that
the rocket does not become too steady and also reaches the expected apogee. Hence, 3 fins were selected.
B.8.4 Fin Material
The fins will be made of aluminum from the T6 6061 series. After analyzing the findings of fin flutter
simulations, aluminum was chosen. Aluminum T-6061 has the added benefit of being lightweight and resistant to
corrosion, stress, and cracking.
B.8.5 Fin Flutter
Using AERO FINSIM fin flutter analysis to ensure the stability of our fin design, we discovered that the
Hero would need to attain a velocity of 613.76 m/s for aluminum fins of this form to experience fin flutter. (For hand
calculation refer to Appendix E-1).
Flutter
Divergence
Young’s
Modulus
(E)
Shear
Modulus
(G)
Poisson
Ratio
Density
f_bending
613.76 m/s
390.73 m/s
10000000
psi
3500000
psi
0.33
0.098 lb/in3
169.861 HZ
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f_torsion
333.245 HZ
Table: Additional Results from AERO FINSIM Software
B.9 Thrust plate
The thrust plate is 10 mm thick and made of 6061-T6 aluminum. Extensive FEA was carried out using
Ansys as the main tool to verify that the design was strong enough to resist the maximum thrust of nearly 720N
produced by the Pro38 784J425-16A motor.
Figure 11: Thrust plate with MDF Motor Damper
The thrust plate was subjected to a finite element analysis (FEA) to replicate launch conditions. To imitate the motor
thrust applied to the region covered by the lip of the motor casing, a 720 N force was applied from the face diameter
of the thrust plate.
B.9.1 Thrust Plate Total Deformation
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B.9.2 Thrust Plate Equivalent Stress
B.9.3 Thrust Plate Maximum Principal Stress
B.9.4 Result
The FEA results showed that the thrust plate would suffer maximum principal stress at 4.91Mpa, with the
yield strength of 6061-T6 aluminum being 259.2Mpa, giving the thrust plate a maximum safety factor of 15. The
highest displacement experienced by the thrust plate was determined to be 0.0055204 mm, which is significantly
less deflection.
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B.10 Motor Retainer
Figure 12: Motor Retainer
A Motor Retainer is used by Hero to keep the motor in place throughout the flight. A 10 mm thick 6061-T6
aluminum plate is used to make the motor retainer.
B.10.1 Motor Retainer Total Deformation
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B.10.2 Motor Retainer Equivalent Stress
B.10.3 Result
By making the screw placement fixed support and applying a force of 2950N which is equal to the mass of
the empty motor and gravity during descending. The maximum stress that the motor retainer could withstand was
determined to be 43.797Mpa. The motor retainer has a factor of safety of 5.91 which is acceptable as a safe
component since the yield strength of 6061 T6 aluminum is 259.2Mpa.
B.11 Bolt Strength Calculation
To check whether the bolts we are using are strong enough to hold the components in place, shearing
strength calculation was done for M4, M5, and M6 bolts. Additional calculations were done for M5 and M6 bolts
because these bolt sizes were chosen for booster tube components where maximum thrust force will be experienced
by these bolts. M5 size bolt was used for thrust plate and M6 bolt size was used for centering rings. (Refer Appendix
E-1 for Bolt Calculations).
Bolt Size
Core
Diameter
Shearing
Strength
τ max
FOS
Min Proof
Strength
Min Tensile
Strength
M4
3.1412mm
2.321KN
-
-
255N/mm2
400N/mm2
M5
4.0184mm
3.627KN
9.46 N/mm2
23.7
255N/mm2
400N/mm2
M6
4.7732mm
5.223KN
6.7061 N/mm2
33.5
255N/mm2
400N/mm2
C. Recovery Subsystem
The recovery system is composed of a single deployment of the commercially available main parachute.
The rocket body airframes will be separated by a black powder ejection system with the help of an altimeter. The
system deploys the parachute once the target apogee is achieved. The primary method we use for deployment is
using a pneumatic piston. Inspired from bps.space’s Lumineer project, this method uses small black powder charges
inside a bolt which in turn generate high pressure inside the chamber when triggered at the apogee, moving a piston
that shears the nylon bolts and allowing the nose cone to pop open for parachutes to exit. For redundancy. a vinyl BP
charge is triggered 2 seconds after apogee. After the charge goes off, the built-up gas pressurizes the recovery
coupler and deploys the main parachute. A variety of separation methods were researched. Ultimately, we chose
black powder as our separation method because it is the most reliable method.
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Methods used to deploy parachutes ● Pyrotechnic Deployment
● Electronic Hatch Opening
● Pin Release
● Drag Separation
Factors considered while selecting parachutes ● Descent rate of rocket
● Mass of the rocket
● Available packing Volume
● Loads during deployment
● Length of the rocket body
C.1 Pyro-Bolts :
Inspired by MIT Rocket Team and BPS.space’s HPR projects, Hero uses a pneumatic piston to push the
recovery devices out of the rocket with the force being generated by igniting a small amount of black powder at the
aft of the piston. This bolt holds 0.4 grams of black powder with 2 redundant ignitors, one for each altimeter, and
sealed off with adhesives. Either of the 2 altimeters can give the signal and will set off the charge, pushing the shaft
upwards, which pushes the piston and pops out the nose cone.
Figure 13: Pyro-Bolt (Credits: BPS.space)
C.2 Recovery avionics :
RRC3 is primarily used to aid in the deployment of the main parachute along with our SRAD altimeter.
The altimeter calculates our rocket's instantaneous and smoothed velocity throughout the ascent and uses this
information to determine when it has reached apogee (zero vertical velocity). Two separate igniters are placed inside
the pyro-bolt, one wired to RRC3 and another to SRAD Avionics, either one of them or both of them can set out a
black powder charge. This pyro-bolt will be connected to the piston which will, in turn, cause piston ejection when
apogee is detected. This will be our primary method for deployment For redundancy, RRC3 triggers a backup black
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powder vinyl charge inside the recovery section with a 2-second delay after apogee. After detecting touchdown, the
last known coordinates from SRAD GPS are used to locate the rocket for manual recovery.
Figure 14: Recovery E-match Connection diagram
C.3 Parachute:
A single 36” toroidal-shaped parachute, with a Cd of 2.2, is used for recovery of the
rocket. The toroidal shape is selected for its relatively high coefficient of drag of
approximately 2.2 for the main chute. The parachute is composed of ripstop nylon
and sewn together with nylon thread. The inner and outer nylon shroud lines are
attached to a nylon bridle which is made of tubular nylon webbing and a mix of nylon
and kevlar thread. The bridle is attached to the shock cord using a steel swivel link.
The parachute size was mainly chosen according to the descent velocities determined
for the deployment of the parachute. For the main parachute, the descent velocity was
determined to be 8.14 m/s. The basic drag equation was used to determine the descent
velocity of the rocket after the deployment of the chute. Here, m is the dry mass of
the rocket, g is the gravitational constant, Cd is the coefficient of drag, δ is the air
density, A is the canopy area and v is the descent velocity.
Figure 15: Toroidal Parachute by Fruity Chutes
Descent Velocity (m/s) =
2 𝑚𝑔
𝐶𝑑 × δ × 𝐴
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Latin American Space Challenge
=
∼ 8.14m/s
C.4 Pyrotechnic Ejection System:
For recovery pyrotechnics, 2 sets of Black Powder were used, one for the piston ejection (primary) and
another for vinyl charge (backup) Ignited with their e-matches.The black powder charge was determined on the basis
of the adequate amount of force required to push the recovery devices and shear the nylon bolts. To determine the
amount of black powder needed for separation, Ideal Gas Law was used. Here, P is the pressure required to shear the
nylon bolts, V is the volume of the compartment, n is the amount of black powder, R is the gas constant and T is the
combustion temperature of black powder.
n=
𝑃𝑉
𝑅𝑇
×
(
454 (𝑔𝑟𝑎𝑚𝑠)
1 (𝑙𝑏𝑓)
)
For more details on black powder amount calculation, check Appendix E-2.
C.5 Piston Ejection Mechanism:
For this project, when RRC3 (COTS) + SRAD avionics detect that the rocket has reached the apogee it
sends a signal to ignite the black powder charge. When this black powder charge explodes and creates pressure in
the pneumatic piston, it is used to push the shaft which moves the bulkhead sitting on the shaft. The bulkhead sitting
flush with the parachute coupler moves the nose cone to exit the tube and hence pushes the recovery devices off. In
case of an anomaly, the redundant RRC3 jumps into action and sends another signal to the backup black powder
charge. This second BP charge sits right in the parachute coupler which is triggered 2 secs after apogee which is
used to pop the nose cone away to pull the recovery hardware out.
Figure 16: Piston Ejection Placement
C.6 Shock Cords:
Shock cord is used to reduce the impact of the opening of the main chutes on the rocket's connecting
bulkhead. It also reduces wear and tear on the rocket as it descends. According to the 'rule of thumb,' the shock cord
is roughly three times the length of the body tube, which came around 15ft in total. The COTS shock cord used is
made of tubular nylon webbing with kevlar on one end with loops sewn on both ends. The calculated shock load that
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is going to act on the shock cord is 340 N. To determine the shock load, the drag equation given below is used. Here,
Cd is the coefficient of drag, δ is the air density, A is the canopy area and v is the descent velocity.
Shock Load (N) =
1
2
2
𝐶𝑑 δ 𝐴 𝑉
= 340N
C.7 Static Pressure Port:
The primary and backup altimeters will need to sense changes in air pressure to provide altitude
measurements during flight, hence vent holes are required along the avionics coupler and airframe. The holes were
primarily designed for the RRC3 altimeter and the backup altimeter, which use the BMP280 sensor. The Static port
hole sizing calculator was used to determine the required diameter. Three 10 mm holes placed at 120 degrees from
each other in the airframe's avionics bay wall were drilled. Additionally three 4mm holes were made in the
parachute compartment to vent out excess pressure preventing the nose cone pop open prematurely.
D. Avionics Subsystem
The primary function of the avionics subsystem is to trigger recovery events and to track the rocket's
location after its touchdown. Additional functions performed by the avionics subsystem include sending data to the
ground station and logging experimental flight data for post-flight analysis. For these, two independently powered
flight computers, one COTS and one SRAD are used which are placed inside the avionics coupler joining booster
and upper airframe.
D.1 Block Diagram
Figure 17: Avionics Block Diagram
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Latin American Space Challenge
D.2 Choice Justification and proof of design
The avionics were designed to fulfill a number of tasks defined by System Engineering. These tasks were
separated between COTS and SRAD systems. The SRAD and COTS systems are completely independent on the
energetic as well as the informational level for redundancy reasons.
● COTS avionics:
1.
2.
Trigger recovery events.
Log altitude, velocity, temperature, and battery voltage locally.
● SRAD avionics:
1.
2.
3.
Send flight data to Ground Station via onboard telemetry
Log flight data such as altitude, location, and pressure locally.
Trigger Recovery Events.
D.2.1 SRAD Avionics Specifications :
The SRAD Avionics is a custom flight computer that consists of an Arduino Nano microcontroller that is
connected to different COTS sensors. The Arduino Nano was selected for its portability, low pricing, computational
power, ease of use, and prior experience of using it.
1.
2.
Use of Sensors:● GY-BMP280 - Used to detect flight conditions such as temperature and pressure, which are then
used to calculate the altitude.
● Adafruit Ultimate GPS - For triangulating coordinates of the rocket after landing.
● Adafruit Micro SD Card - For logging data provided by sensors.
● Xbee S3B Pro - For telemetry via rocket to ground station
Reasons for choosing them:● GY-BMP280 - smaller footprint, lower noise measurements, with good accuracy.
● Adafruit Ultimate GPS - It can track up to 22 satellites on 66 channels, has an excellent
high-sensitivity receiver, a small form factor, and low power consumption.
● Adafruit Micro SD Card - low latency and high-frequency of data logging over SPI
● Xbee S3B Pro - data rate of up to 20Kbps and has a HAM license-free frequency range of
operation from 902MHz to 928MHz.The SRAD altimeter would be operating at 916 MHz.
D.2.2 COTS Avionics Specification :
Missile Works RRC3, is a barometric dual‐deployment altimeter,
and is used as our primary COTS altimeter. The RRC3 is
recommended to be operated with a standard 9‐volt alkaline
battery. It also logs flight data like altitude, velocity, temperature,
and battery voltage locally on non-volatile storage. We use RRC3
to handle the recovery events on Hero by triggering the piston
ejection and the backup BP charge at apogee.
Figure 18: Missile Works RRC3
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It was selected as it was heard from word of mouth of many amateur rocketeers in terms of reliability and was
configurable enough to set 2 charges off at a delay. It also has great UI and sanity checking onboard to mitigate any
anomalies on the pad if such an event occurs.
D.2.3 Power Budget:
The avionics system is designed to provide power to all systems with enough redundancy to allow the
rocket to fly even after being on the pad for many hours. Two guiding specifications were used when selecting a
power value for the battery: the voltage of the battery must be greater than the required voltage for any one
component, and one of the commonly available voltages must be used. After investigation, it was determined that
11.1V and 9V were among some of the most common voltages for high-power batteries (For power budget
calculations check Appendix E-3). Since a combination of COTS and SRAD electronics are used, for redundancy,
they both were powered individually. Hence the figure below contains the chosen configuration to power the
onboard electronics.
Figure 19: Power Budget
D.2.4 GPS and Telemetry:
Adafruit Ultimate GPS is used as our primary GPS. It is a high-quality
GPS module that can track up to 22 satellites on 66 channels, has an excellent
high-sensitivity receiver, and has a built-in antenna. It can do up to 10 location
updates a second for high speed, high sensitivity logging, or tracking.
Figure 20: Adafruit Ultimate GPS Breakout-66 channel
To receive flight data from SRAD Avionics in real-time, a telemetry system was developed to send flight
data to the ground station. For telemetry Xbee S3B pro was used. Xbee sends data like altitude, GPS coordinates,
velocity, and flight states to the ground station. One of the primary reasons for choosing this module is it delivers a
needed transmission range of up to 500 m Line-of-Sight (LOS) along with low power consumption. The model
21
Latin American Space Challenge
we’re using operates at a 916MHz HAM license-free frequency band and 10Kbps to 20Kbps RF data rate. An
external COTS omnidirectional antenna, ANT-916-CW-RH, is used for transmission.
For our configuration, 2 Xbee modules were used:
● One connected to the microcontroller of the SRAD Avionics unit to send
out packets containing flight data.
● The other is on the receiving end to fetch data from the rocket at the
ground station. The module is connected to a local machine to save those
packets, interpret them and display appropriate data on the software.
The communication between XBee in the flight computer and XBee used at
the ground station will be a simplex communication.
Figure 21: XBee Pro S3B
D.2.5 Arming System:
The arming system is a critical safety feature that prevents accidental deployment of any recovery events
before the rocket is safely on the pad, as well as saving battery life. Two SPST Screw Switches were used, one to
arm SRAD Avionics and the other to arm the RRC3. They both
are powered by separate batteries. When both the systems are
armed, the buzzers on RRC3 as well as SRAD Altimeter go off
which confirms that both the systems are armed successfully.
Additionally, the connection established is also validated by the
ground station by checking the packets received via telemetry
from the SRAD Avionics. A series of validation checks are to be
performed on the launch pad (found in ‘Preflight Launch
Checklist’ in Appendix E-3) to validate all systems are
functioning as planned.
Figure 22: SPST Screw Switches
Figure 23: Arming Mechanism
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Latin American Space Challenge
D.2.6 Control Panel (Ground Station):
As soon as the avionics is in the armed state, the ground station immediately starts receiving packets from
the Xbee telemetry module. This data includes altitude, location, velocity, and flight states from SRAD Altimeter.
The flight state has 4 events namely “On the pad, Apogee, Main Deployment, and Touchdown”. After arming, the
device continuously collects data until landing. Data is stored in a Micro SD Card and can be downloaded to a
computer for post-flight analysis.
D.3 Ground Station
Figure 24: Ground Station Block Diagram
The ground station checks the status of the rocket during its flight. A Yagi-Uda antenna is used for
receiving telemetry data. The Yagi Uda antenna’s connector is directly connected to the XBee module’s RP-SMA
male connector. This Xbee S3B pro module is placed on an adapter that is connected to the GS computer. The
desktop application, by connecting to the serial port, will display the data sent by the SRAD Computer.
When there is little to no change in GPS coordinates, it means that the rocket has landed, those final
touchdown location coordinates are then loaded to maps and the recovery team leaves for the touchdown location
for manual recovery. It does that by downloading the launch site area map beforehand with the help of Google Earth
which is then later used in offline mode for recovery by searching the coordinates we need.
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Latin American Space Challenge
E. Payload Subsystem
Flying inside the nose of Hero is a Data acquisition unit whose main goal is to model the expected flight
behavior; select, design, and build an appropriate sensor package to measure phenomena related to the scientific
mission of their flight, and compare expected behavior to their measured flight results. This single-board design
consists of an array of sensors ranging from IMUs to BMEs and logs this data to a micro SD card.
A data acquisition system (DAQ) is a collection of sensors, circuitry to modify/condition the sensor signal,
analog-to-digital converters, and either onboard storage or a means to interface with a computer for remote data
storage. Modern data acquisition systems (DAQs) include computation capabilities as well as graphical interfaces
for quickly processing and visualizing data. A stand-alone (or traditional) data logger is a type of DAQ that is not
connected to a PC and thus has its own power source as well as onboard data storage (often done using a memory
card).
Figure 25: Payload Enclosure
For our little experiment, we’re using IMU and BMP to model a flight trajectory by understanding roll
characteristics, pitching moment, and vibrations in-flight to model a mockup of our flight trajectory by dead
reckoning. This software-based approach is often used in ICBMs, submarines, and aviation vehicles, but the idea is
to use the data from inertial sensors, propagate forward with the accelerations and rotations by integrating them to
get position and you get some estimate on your location. Since the sensors onboard are MEMS, there are 3 IMUs to
filter out noisy data.
We are using the MPU9250 breakout as our IMU. The MPU9250 is a 9-axis MEMS sensor from
InviSense® designed to be used for impact recognition and logging, motion-activated functions, vibration
monitoring and compensation. MPU9250 9-Axis Attitude Gyro Accelerator Magnetometer Sensor Module features
the MPU-9250, which is a multi-chip module (MCM) consisting of two dies integrated into a single QFN package.
One die houses the 3-Axis accelerometer and gyroscope. The other die houses the AK8963 3-Axis magnetometer.
Hence, the MPU-9250 is a 9-axis Motion Tracking device that combines a 3-axis accelerometer, gyroscope and
magnetometer, and a Digital Motion Processor (DMP). It is based on I2C Address, Address: 0x68. It consists of an
Accelerometer of range ± 2 ± 4 ± 8 ± 16g. The Gyroscope has a range of ± 250 500 1000 2000 ° / s. It also consists
of a Magnetometer measuring magnetic fields ranging ± 4800uT.
The atmospheric sensor used is BME280 Atmospheric Sensor Breakout by Sparkfun. The BME280
Breakout has been designed to be used in navigation and weather forecasting. The on-board BME280 sensor
measures atmospheric pressure from 30kPa to 110kPa as well as relative humidity ranging from 0-100% RH, =-3%
from 20-80% and temperature ranging from -40°C to 85°C. The breakout provides a 3.3V SPI interface takes
measurements at less than 1mA.
24
Latin American Space Challenge
Ⅲ. Mission Concept of Operations
Hero’s objective is to reach an altitude close to 0.5km as per LASC competition guidelines, before
actioning a safe and successful recovery. The mission structure can be split into eleven successive stages :
Assembly, Launchpad Setup, Procedure checks, Arming, Firing, Liftoff, Motor Burnout, Coastphase, Apogee, Main
Deployment, Landing.
Figure 26: Flowchart of Launch Mission Phases
Pre-Launch Phases
Phase
Pre-Flight Assembly
Description
Time
Joined body tubes and airframe with couplers.
Payload inserted in the bay.
Avionics and Motor components assembled and inserted in
their respective bay.
Parachute inserted in rocket.
Ground Station Setup.
T-2 Hours
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Latin American Space Challenge
Launch pad Setup
Rocket placed on the launch rail at a specific angle.
T - 30 Mins
Preflight Checks
Every sub-system checkups.
System Testing checkup
T - 30 Mins
Arming
The arming phase starts when the rocket is attached to the
launch rail.
Altimeters armed.
Igniters inserted.
T - 15 Mins
Launch Mission Phases
Phase
Description
Time
Firing
Motor ignited.
Propulsion System Firing.
Rocket ready to Liftoff.
T-1 Sec
Liftoff
Right after ignition, the motor begins to generate thrust
which pushes the rocket into the air, it is during this phase
when the rocket gets its upward acceleration.
Avionics starts detecting launch and altitude
measurements which are displayed by ground station.
Rocket leaves thes launch rail and takes off.
T - 0 Sec
Rocket Ascent - Motor
Burn
A point where the propellant inside the motor is fully
consumed, the motor is no longer producing thrust.
Avionics starts calculating altitude.
T + 2 Sec
Rocket Ascent - Cruise
After the burntime the rocket enters the coast phase,
during this phase there is no thrust coming from the
motor, and the rocket is moving with inertia.
2 sec< T < 11 Sec
Apogee
Once the vehicle has been through the coast phase it
reaches apogee, Point where the velocity of Rocket
reaches zero.
Avionics fires ejection charges.Recovery system initiates.
T + 11 Sec
Chute Deployment
The Main chute deployment phase is initiated when the
avionics system detects an altitude below the apogee,
At this point the ejection charge is fired which further
pushes off the nose cone, and ejects the parachute out of
the rocket.
Shear pins separate Nose Cone from Upper Airframe.
Deployment event is displayed by the ground station.
T + 13 Sec (worst
case scenario)
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Latin American Space Challenge
Touchdown
Rocket returns to the ground safely.
Ground Station displays landing events.
T + 72 Sec
Launch Mission Phases
Figure 27: Diagram representing rocket phases
Ⅳ. Conclusion and Lessons Learned
Historically, the team has been primarily concerned with payload. This year's rocket was a significant
departure from previous models in that we concentrated heavily on aerostructures and optimizing parts to reduce
manufacturing costs. Because we underestimated how difficult this would be, we set far loftier goals for this rocket,
including custom cameras, a more reliable arming mechanism, a far more advanced ground station, a DAQ system
to understand flight dynamics, and an integrated umbilical connection for liftoff. Despite the difficulties, the team
has gained knowledge about manufacturing composite parts and additive/subtractive manufacturing. We made many
mistakes, but we were able to learn from them and avoid them in future projects.
The pandemic had a significant impact on the team. Throughout the year, we lost a significant number of
skilled team members, resources, funding, and focus, but through the pain, we have set up next year's rocket to be
extremely successful. We hope that the lessons learned from this project will be useful to the team in the future.
27
Latin American Space Challenge
2nd Progress Report Receipt
2021 Latin American Space Challenge
Hello, Team 11
The Latin American Space Challenge (LASC) successfully received your
submission for the 2nd Progress Report. For your information, the data
collected by our judges from the system is available below:
Category: Entry-level: 0.5 km apogee
Mission: The Perfect Game
Predicted Apogee (m):
Max. Velocity (m/s):
551
98.9
Airframe Length (mm):
Airframe Max. Diameter (mm):
1460
104
Liftoff Mass (kg):
Liftoff Thrust-Weight Ratio:
6.36
5.51
Launch Rail Dep. Velocity (m/s):
Min. Static Margin During Boost:
31.2
1.8
Recovery System:
Main Deploy Altitude (m):
500
In terms of rocket recovery a single 36" toroidal parachute with Cd of 2.2 we bring rocket safely to the ground with
a decent velocity of 8.14 m/s. This is done primarily using a pyro-based piston Ejection method Along with vinyl
charges acting as backup.
Latin American Space Challenge | www.lasc.space
2nd Progress Report Receipt
2021 Latin American Space Challenge
Propulsion:
Class J solid rocket motor, the average thrust of 325.29 N, the total impulse of 765.00 Ns, based on APCP. For our
primary design, a 38 mm COTS Cesaroni J330 motor is being used which will help us reached our target apogee.
Alternatively, an SRAD motor bas
Updates
The Team Had no Change in structure however the team has started arranging the Insitu launch site for an
experimental test launch. Since this design is based on Virtual event we make sure that the design can be scaled
up easily once we make our propulsion
General Progress:
The team will soon be regrouping at the university campus to start the actual production of the project and the
static fire round for the SRAD motor. For the sake of authenticity of documentation, the motor selection comes at
par with the designed SRAD mot
Report generated on October 26, 2021 by LASC Organization for Team 11
Latin American Space Challenge | www.lasc.space
Appendix
Appendix A: SYSTEM WEIGHTS, MEASURES, AND PERFORMANCE DATA APPENDIX
1. Propulsion
At Lift off
Component
Mass (g)
1
Propellant grain
405
2
Motor Casing
295
Total
At Motor Burnout
700
Component
Mass (g)
1
Propellant grain
0
2
Motor Casing
295
Total
295
2. Aerostructures
Component
Mass (g)
1
Nosecone Assembly
364
2
Upper Airframe Assembly
425
3
Booster Tube Assembly (Without motor)
665.6
4
Parachute Coupler
108.8
5
Avionics Coupler
317.8
Total
1881.2
Component
Mass (g)
1
Parachute
141.74
2
Shock Cord
520
3
Recovery Hardware
300
Total
961.74
3. Recovery
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Latin American Space Challenge
4. Avionics
Component
Mass (g)
1
COTS Altimeter
17.01
2
11.1V Li-ion battery
172
3
9V Alkaline battery
100
4
SRAD Avionics
130
5
Cables & Connectors
20
Total
439.01
Component
Mass (g)
Payload
300g
5. Payload
1
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Latin American Space Challenge
Figure 28: Hero’s Internal Configuration
Rocket Parameters
Airframe Length (in)
57.48
Airframe Diameter (in)
4.09
Vehicle Weight
6765 g
Propellant Weight
405 g
Payload Weight
300 g
Flight Parameters
Simulation Software
OpenRocket
Launch Rail Length
16.4 ft
Lift-Off Thrust to Weight Ratio
6.37:1
Launch Rail Departure Velocity
31.2 m/s
Maximum Acceleration
90.8 m/s2
Maximum Velocity
98.7 m/s2
Predicted Apogee
525 m
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Latin American Space Challenge
Propulsion Parameters
Propulsion Type
Solid
Manufacturer
Cesaroni Technologies
Class
J
Motor Designation
J425
Total Impulse (N-sec)
785 N-sec
Peak Thrust (N)
720 N
Average Thrust (N)
423 N
Burn time (sec)
1.85
Parachute Recovery Summary
Diameter
36”
Coefficient of Drag (Cd)
2.20
Decent Velocity
8.14 m/s
Material
Nylon
Quick Links
Stainless Steel (3/16”)
Shock Cord Parameters
Material
Tubular nylon
Length (ft)
15 ft
Width (in)
1.25 in
Maximum Breaking load
4200 lbs
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Latin American Space Challenge
Appendix B: HAZARD ANALYSIS APPENDIX
Team 11
#The Perfect Game
Severity
Hazard
Possible Causes
Mitigation Approach
Medium
Person injuries during
manufacturing and testing.
Lack of safety
regulations and
procedures.
Maintaining safe distances and making
sure to follow the safety measures list.
Low
Black powder mishap.
Exposure to heat
sources.
Loading of black powder during
assembly should be done carefully and
should be stored in cold, dry places.
There should be an additional checklist
to approach the rocket when it has active
black powder charges.
Low
Premature motor ignition.
Motor close contact to
heat flame or other
heat sources.
The motor should be stored in cold, dry
places for storage and transportation.
Low
Hazardous material handling.
Personal Protective Equipment must be
worn at all times when handling these
materials.Usage of these materials
should be timely. Active ventilation
system attached to the workspace.
Appendix C: RISK ASSESSMENT APPENDIX
Team 11
#The Perfect Game
Severity
Risk
Possible Causes
Mitigation Approach
Medium
Task Incompletion due to time
constraints which are probably
used on other tasks in hand.
Poor communication
between team
members and leads.
Senior team members and leads should
make sure new members complete their
tasks in hard assigned deadlines, ensure
most of the team members speak
English, and are active whenever
possible.
Medium
Subsystem task incompletion
New inexperienced
Assigning a particular task to every
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Latin American Space Challenge
till date.
team members.
member and arranging weekly general
meetings, wherein respective team leads
get an update on everyone and verify
them.
Low
Team split between tasks
Team members'
ambition to achieve
the perfect rocket.
Prioritize the tasks; assign the team
members to focus completely on the
highest priority ones; terminate
infeasible objectives; discuss
compromised objectives at team
meetings.
Medium
Insufficient money to complete
goal objectives.
Less time assigned or
not proper approach
and interest on the
non-technical side of
the project.
Finding members who have real
experience in publicizing the idea. Other
team members should put more effort
into sponsoring and making strong
relations with other companies and
organizations.
Low
Failure to ignite the propellant.
Improper
installation of
igniter
Reset/installation of new ignitor and
checking proper continuity.
Low
Failure/breakage of the
airframe.
Not properly coupled
or environmental
errors.
Ensure proper assembly and careful
inspection of joining of components,
using simulations for environmental
errors.
Low
Explosion of solid-propellant
rocket motor during launch.
Cracks in propellant
grain or gaps in the
propellant sections.
Inspections of the motor grains for
cracks and gaps during assembly prior to
launch.
Low
Motor ejection and full
structural failure.
Assembly integration
issue.
Ensure proper assembly.
Medium
Failure/breakage of fins.
Assembly integration
problem.
Follow calculations and proper material
use.
Low
Target Apogee not reached.
Error in calculations
and simulations,
unpredicted weather
conditions.
Making proper use of OpenRocket and
other simulation software to make sure
there are solutions for all scenarios.
Medium
Breakage/damage to avionics
bay.
Due to buckling
which causes
structural instability in
the rocket.
Place the avionics coupler half their
length of the body tube, it will increase
their strength 4 times.
Medium
Failure/breakage to payload
and avionics bay.
Physical damage to
the main chute
leading to ballistic
descent.
Check out parachute trajectory
simulations and calculations.
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Latin American Space Challenge
Low
Rocket deviates from nominal
trajectory and delay between
launch and firing.
Incorrect mounting
of the rocket to the
launch rail.
Follow pre-launch procedures during
launch pad setup.
High
Failure to achieve minimum
stability off the rail.
Low thrust and speed
or insufficient
aerodynamic stability
which may cause poor
stability off the launch
rail.
Using OpenRocket simulations to
overcome and verify launch rail stability.
Follow pre-launch procedures.
Medium
Telemetry Failure.
Aerodynamic forces
during flight might
result in loose
electrical connections.
A proper assembly checklist is required.
Make use of high-quality electrical
connectors.
Low
Unsuitable drag force from
simulation.
Commercial Rocket
simulation software
error.
Hand-check using TR-11 Aerodynamics.
Low
Parachute inflation
unsuccessful.
Parachute packing
problem during
assembly.
The parachute must be properly fitted
into the rocket leading to proper
separation of the airframe.
Low
Main Parachute deployment
failure.
Failure of altimeters.
Testing all systems prior to launch.
Check for continuity on all pyro
channels.
References
Books
Dr. Gerald M. Gregorek, Aerodynamic Drag of Model Rocket, ESTES Industries.INC.Box 227, 1970
2
Barrowman J.S. and Barrowman J.A, 1996, Theoretical Prediction of the center of pressure, NARAM
3 Rick Newland, Martin Heywood, Andy Lee, Rocket Vehicle Loads and Airframe Design, Aspirepace technical
book
4
V.B Bhandari, Design of Machine Element, 2nd edition, Tata McGraw Hill 2007
1
Documents
Apogee Newsletter 291, Peak of Flight Fin Flutter Calculation, Doc link:
https://apogeerockets.com/education/downloads/Newsletter291.pdf
6
Sampo Niskanen, Open Rocket Technical Document, Based on the Master’s thesis, Development of an Open
Source model rocket simulation software, OR 13.05, http://openrocket.sourceforge.net/.
7
Missile Work Corporation, RRC3-Rocket Recovery Controller 3, user manual, www.missileworks.com
5
Websites
34
Latin American Space Challenge
8
The U.S. Standard Atmosphere, Engineering ToolBox, (2001), URL: www.EngineeringToolBox.com
AzoMaterials, E Glass Fibre Properties, URL: https://www.azom.com/properties.aspx?ArticleID=764
10
Tom Benson, Rocket Principles, Beginners Guide to Rockets, URL:
https://www.grc.nasa.gov/www/k-12/rocket/TRCRocket/rocket_principles.html
11
Tom Benson, Rocket Thrust, Beginners Guide to Rockets, URL:
https://www.grc.nasa.gov/www/k-12/airplane/rockth.html
12
Richard Nakka, Richard Nakka’s Experimental Rocketry Website, URL: Nakka's Experimental Rocketry Site .
13
Tom Benson, Four Forces on a Rocket, Beginners Guide to Rockets, URL: Four Forces on a Rocket .
14
Thomson Engineering Design, A Short Guide to Metric Nut and Bolt, Proof Load for std pitch bolt, URL:
https://thomsonrail.com/metric-nuts-and-bolts/
15
ASM Aerospace Specification Metals Inc, Aluminum 6061 T6 Properties, URL:
http://asm.matweb.com/search/SpecificMaterial.asp?bassnum=ma6061t6
16
Joe Barnard, Advanced Rocketry Community Platform, URL: https://bps.space/
17
SparkFun Atmospheric Sensor Breakout, BME280, Temperature and Humidity Sensor, https://robu.in/
9
Computer Softwares
Autodesk, Fusion360, 3D CAD Software, Student Version 2.0.10806
19
John Swanson, Ansys Workbench, Mechanical, Ansys Student 2021 R2
20
Sampo Niskanen, Open Rocket designing and simulation, Version OpenRocket 15.03, 2021
21
John Cipolla/AeroRocket, AEROFINSIM, Fin Flutter simulation, Finsimlite Version 5.7
22
Autodesk, EAGLE,Version 9.6.2
18
Data Sheets
Lady ada, Micro SD Card Breakout Board Tutorial, adafruit learning system,
https://learn.adafruit-micro-sd-breakout-board-card-turorial
24
Arduino Nano, User Manual, www.arduino.cc
25
Bosch Sensortec, BME280, Combined Humidity and Pressure Sensor, www.bosch-sensortec.com
26
Multicomp Pro. Carbon Film Fixed Resistors, RoHS Compliant V1.1, Newark.com/multicomp-pro
27
MPS The Future of Analog IC Technology, MP1584, 3A, 1.5MHz, 28V Step Down Converter,
www.MonolithicPower.com
28
InvenSense, MPU-9250 Product Specification 1.1, PS-MPU-9250A-01, www.invense.com
29
Digi International,XBee-PRO® XSC S3B,
https://www.digi.com/products/embedded-systems/digi-xbee/rf-modules/sub-1-ghz-rf-modules/xbee-pro-xsc
30
Lady ada,Adafruit Ultimate GPS, adafruit learning system,https://www.adafruit.com/product/746
31
Bosch Sensortec, BMP280, Combined Humidity and Pressure Sensor, www.bosch-sensortec.com
23
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Latin American Space Challenge
Appendix D: ENGINEERING DRAWINGS APPENDIX
Drawing 1: Nosecone
Drawing 2: Nose Tip
Drawing 3: Noseplate
Drawing 4: Nosecone Endcap
Drawing 5: Fins
Drawing 6: Booster tube
Drawing 7: MDF Motor Damper
Drawing 8: Thrust Plate
Drawing 9: Retention Plate
Drawing 10: Boat Tail
Drawing 11: Upper Airframe
Drawing 12: Fin Assembly: Centering Rings and Stringers
Drawing 13: Shaft Piston
Drawing 14: Avionics
Drawing 15: Avionics Coupler Top
The Perfect Game
LASC Tests Summary
Theoretical Calculations
Date: DD-MM-YYYY
Category: Aerostructures
Version: 1.0
Technical Notes(Appendix E-1)
Title:
Technical Notes: Aerostructure
Aim:
Testing Team:
Approved By:
Location:
Status:
-
Bolt Shearing Strength Calculation
Bolt size: M5 bolt of 4.6 grade plain carbon steel
Nominal Diameter: 5 mm
Clearance Hole: 5.3 mm
Ultimate Strength of bolt (Fub): 400 Mpa
Yield Strength of bolt (Fyb): 240 Mpa
Shearing Strength of bolt (Vdsb):
𝑉𝑛𝑠𝑏
𝑌𝑚𝑏
Where, Vnsb is nominal shear strength and Ymb is partial factor of safety in yielding for the material in bearing
Vdsb=
𝐹𝑢𝑏
3
×
𝑁𝑛.𝐴𝑛𝑏+𝑁𝑠.𝐴𝑠𝑏
𝑌𝑚𝑏
Where, Nn is the number of shear planes passing through the threaded part and Ns is the number of shear planes passing
through the shank part.
INTERNAL USE ONLY
1
Team #11 STES India
The Perfect Game
Vdsb=
400
×
3
π
4
×
25
1.25
=3627.598N
Shearing Strength of M5 bolt of grade 4.6 is 3.627 KN
Similarly, shearing strength of M4 and M6 bolts of grade 4.6 is 2.321 KN and 5.223 KN respectively.
Bolt Safety Factor Calculation
Bolt size: M5 bolt of 4.6 grade plain carbon steel
Core diameter of M5 bolt: 4.0184mm
Minimum proof strength of grade 4.6 bolt: 225 N/mm2
Minimum tensile strength of grade 4.6 bolt: 400 N/mm2
𝑇ℎ𝑟𝑢𝑠𝑡 𝑓𝑜𝑟𝑐𝑒
𝑁𝑜𝑠 𝑜𝑓 𝑏𝑜𝑙𝑡𝑠
Thrust force on one bolt=
Tmax=
FOS=
𝑇ℎ𝑟𝑢𝑠𝑡 𝑓𝑜𝑟𝑐𝑒 𝑜𝑛 𝑜𝑛𝑒 𝑏𝑜𝑙𝑡
𝐴𝑟𝑒𝑎 𝑜𝑓 𝑜𝑛𝑒 𝑏𝑜𝑙𝑡
=
=
720
6
= 120𝑁
120
π
4
= 9. 462
2
×(4.0184)
𝑀𝑖𝑛𝑖𝑚𝑢𝑚 𝑝𝑟𝑜𝑜𝑓 𝑠𝑡𝑟𝑒𝑛𝑔𝑡ℎ 𝑜𝑓 𝑝𝑟𝑜𝑝𝑒𝑟𝑡𝑦 𝑐𝑙𝑎𝑠𝑠 4.6
𝑇𝑚𝑎𝑥
=
225
9.462
𝑁
2
𝑚𝑚
= 23. 779
Drag Coefficient Calculation by TR-11 Aerodynamics
1+1.5
CDN+CDBT=1.02Cf×
CDB=
3
0.029
𝐶𝐷𝑁+𝐶𝐷𝐵𝑇
(
×
𝑑𝑏
𝑑
2𝑡
𝑐
CDOF=C*DOF×
𝑆𝐹
𝑆𝐵𝑇
𝐶𝑅
𝑆𝐵𝑇
𝑆𝑤
𝑆𝐵𝑇
=
0.029
×
0.2
(
= 1. 02×2. 08×
1+1.5
52.5
)×
86.59
34.3
= 0. 2
3
(0.08)
0.104
= 0. 025
) = 0. 595
C*DOF=2Cf 1 +
CDint=CDOF×
×
𝐿 1.5
𝑑
( )
×
= 0. 595×
𝑑
2
8.28
34.3
= 0. 143
×𝑛𝑜𝑠 𝑜𝑓 𝑓𝑖𝑛𝑠 = 0. 143×
0.15
34.3
×
104
2
×3 = 0. 097
CD=CDN+CDBT+CDB+CDOF+CDint=0.2+0.025+0.143+0.097=0.465
Drag coefficient by hand calculation:0.465
Drag coefficient by open rocket simulation:0.495
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Team #11 STES India
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Drag force Calculation
Taking peak value of Drag coefficient from open rocket
Coefficient of drag (CD)= 0.495
Maximum Velocity (V)= 98.7 m/s
Density (ρ)= 1.112 Kg/m3
Ref Area (Aref)=
π
4
2
×𝑑 =
Drag force (D)= 𝐶𝐷×
1
2
π
4
2
2
× (0. 104) = 0. 008494 𝑚
2
2
×ρ×𝐴𝑟𝑒𝑓×𝑉 = 0. 495×0. 5×1. 112×0. 008494×(98. 7)
Drag Force (D)= 22.77 N
where,
CD= Drag coefficient of an aerodynamic shape
CDB= Base drag coefficient
CDN= Drag coefficient of nose cone
CDBT= Drag coefficient of Body tube
CDOF= Drag coefficient of the fins at zero angle of attack
C*DOF= Drag coefficient of the fins at zero angle of attack, based on fin surface area SF
CDint= Drag coefficient due to fin and body interference
Cf= Skin friction coefficient due to boundary layer
D= Drag force
L= Length of Rocket
SF= surface area of all fins
Sw= wetted surface area of entire rocket
SBT= Cross sectional area of the body tube
Aref= Reference area of rocket body
db= Base diameter of rocket body
t/c= thickness ratio of the fin; thickness divided by chord
ρ= Density of air
Axial Force Calculation
Maximum Thrust (T)= 720N
Drag Force (D)= 22.77N
Maximum Acceleration (a x)=90.8m/s2
Inertial Load (Mx)= 3.3825 Kg
𝑥
Axial Force= -T+D+ax∑Mx = -720+22.77+(307.131) = -390.099 N
𝑥0
Note: The thrust is in the opposite direction to the drags so has the opposite sign. (Negative signum only indicates direction)
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Fin Flutter Calculation
Root Chord (Cr)= 5.905 in
Tip chord (Ct)= 3.14961 in
Thickness (t)= 0.11811 in
Semi-span (b)= 4.724 in
Shear Modulus (G)= 3500000 psi
Flutter Boundary Equation:
𝑉𝑓 = 𝑎
𝐺
1.337×(𝐴𝑅)^3×𝑃×(λ+1)
𝑡 3
𝑐
( )
2(𝐴𝑅+2)
S= ½(Cr+Ct) b= ½ (5.905+3.14961) ×4.724= 21.386 in2
AR=
2
𝑏
𝑆
= (4.724)2/21.386= 1.043
λ= Ct/Cr= 3.14961/5.905= 0.533
T= 59-0.00356h= 59-0.00365(1640.42) = 53.16 F
P= 2116((
𝑇+459.7
518.6
)=14.5317 psi
a= 1. 4×1716. 59×(𝑇 + 460)= 1110.513149 ft/sec
Hence, after calculating, Vf was found to be 2494.004 ft/s or 760.17 m/s
Fin Flutter velocity by Aero FINSIM simulation:613.76 m/s
Fin Flutter Velocity by Hand Calculation: 760.17 m/s
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The Perfect Game
LASC Tests Summary
Date: DD-MM-YYYY
Category:Recovery
Version: 1.0
Theoretical Calculations
Technical Notes(Appendix E-2)
Title:
Technical Notes: Recovery
Aim:
Testing Team:
Approved By:
Location:
Status:
-
Descent Velocity Calculation
Gravitational Constant (g) = 9.81 m/𝑠
2
Dry Mass (m) = 6500 g
Coefficient Of Drag (Cd) = 2.2
3
Density of Air ( δ) = 1225 g/𝑚
Parachute Diameter = 36 inch =
Surface Area (A) =
𝑖𝑛𝑐ℎ
39.37
Π×𝑃𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒 𝐷𝑖𝑎𝑚𝑒𝑡𝑒𝑟
4
2
=
2
36
39.37
=
= 0.91 meter
2
Π × 0.91
4
Surface Area (A) = 0.65 𝑚
Descent Velocity (v) =
2×𝑔×𝑚
𝐶𝑑 × δ × 𝐴
=
2 × 9.81 × 6500
2.2 × 1225 × 0.65
Descent Velocity (v) = 8.14 m/s
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Team #11 STES India
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Shock Load Calculation
Drag Coefficient (Cd) = 2.2
3
Density of Air ( δ) = 1.225 kg/𝑚
2
Parachute Surface Area (A) = 0.65 𝑚
Terminal Velocity (V) = 19.6 m/s
Shock Load =
1
2
2
× 𝐶𝑑 × δ × 𝐴 × 𝑉 =
1
2
2
x 2.2 x 1.225 x 0.65 x 19. 6
Shock Load = 340 N
Black Powder Calculation
Pressure (P) = ∼15 psi
2
2
Airframe Volume (V) = Π × 𝑟 × ℎ = Π × 1. 96
× 14. 96
3
Airframe Volume (V) = 184.25 𝑖𝑛𝑐ℎ
Converted Gas Constant (R) = 266 inch.lbf/lbm
Temperature inside airframe during combustion (T) = 3307
PV=nRT
Mass of Black Powder (n) =
𝑃𝑉
𝑅𝑇
×
(
454 (𝑔𝑟𝑎𝑚𝑠)
1 (𝑙𝑏𝑓)
)=(
15 × 184.25
266 × 3307
)× ( )
454
1
Mass of Black Powder (n) = 1.42 g
Note: Pressure assumed as 15 psi
Piston Force Calculation
Pressure = ∼15 psi
Radius of the piston (r) = 1.96 inch
2
2
Area (A) = Π × 𝑟 = Π × 1. 96
2
Area (A) = 12.06 𝑖𝑛𝑐ℎ
Force (F) = P × A = 15 × 12.06
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Team #11 STES India
Force (F) = 180.9 N
Note: Pressure assumed as 15 psi
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The Perfect Game
LASC Tests Summary
Date: DD-MM-YYYY
Category:Avionics
Version: 1.0
Theoretical Calculations
Technical Notes(Appendix E-3)
Title:
Aim:
Testing Team:
Approved By:
Location:
Status:
-
1.On Pad Checklist (Preflight & Launch)
⬜ Carefully slide the rocket onto the launch rail
⬜ Raise the rocket to the desired launch angle
⬜
Arm the COTS and SRAD altimeter by turning ON the arming switch and ensuring proper startup sequence. A continuous beep of
5 secs for RRC3 & of 3 secs for SRAD Avionics to ensure initialization.Three short beeps for RRC3 to indicate continuity on the main
and drogue terminals. Check Connection Establishment for location on the mobile app for SRAD GPS. If either altimeter displays
off-nominal, proceed to section 2.1 “clearing anomalies on the pad”
⬜ Arm payload with the arming switch. Wait for confirmation of arming
⬜ Clear the pad of all non-essential personnel
⬜ Get permission from the RSO to install ignitors
⬜ Insert the dowel with igniters as far up into the engine as it will go
⬜
Strip the leads of the igniters and attach them to the launch control leads. Confirm they are attached in parallel. Confirm they are
laid out so there is no possibility of a short
⬜
Test launch controller continuity at the pad level. If there is no continuity, proceed to section 2.1. Ensure one final time that all
avionics are functional. If not, proceed to section 2.1
⬜ Leave launchpad and maintain a safe distance.
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Team #11 STES India
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1.1 Clearing anomalies on the pad
1.
RRC3
a.
One long continuous beep pattern
i.no continuity on any event terminal
ii. Disarm altimeter
iii. Pull rocket off the rail, the problem is in both bays and will likely take some time to fix
b.
One short beep pattern
i.continuity on only the drogue terminal i.e no continuity on the main
1. Disarm altimeter
2. Lower rocket
3. Unscrew parachute bay
a. If the problem can be determined, fix and restart pad ops procedures.
Otherwise, stand down.
c.
Two short beep pattern
i. continuity on only the main terminal i.e no continuity on the drogue
1. Disarm altimeter
2. Lower rocket
3. Unscrew parachute bay
a. If the problem can be determined, fix and restart pad ops procedures.
Otherwise, stand down.
2. Custom avionics non-functional
a. Stand down, will require taking the rocket more or less fully apart
3. No continuity for the motor igniter
a. Check leads for contact
b. See RSO if other issues
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2.Avionics Assembly Procedure
⬜ Test two batteries, they should each be at or above their rated voltages
⬜
Attach the screw switches and battery leads to the two altimeters and powers them on. Confirm the altimeters enter startup
sequence
⬜ Turn off the switches to conserve battery life.
⬜ Insert arming keys and turn so that the key is retained (off position)
⬜ Connect the igniters connected to the pyro-bolt inside in the pneumatic piston where black powder is kept.
⬜ This is in the avionics bay. Check if the igniters are connected to the SRAD avionics and RRC3’s main terminal.
⬜
Check the wire passing through the avionics bay’s top bulkhead via RRC3’s drogue terminal is connecting vinyl charge that is
present in the recovery bay properly as it is the redundant recovery system team is using.
⬜
Mount the batteries and electronic boards on the Avionics structure. PCB needs to be mounted on top of risers and bolted nicely to
the structure.
⬜ Connect the COTS and SRAD Avionics units to the 2 separate SPST Switches. Keep the electronics turned OFF.
⬜ Integrate Avionics inside the coupler.
3.Power Budget Calculations
1)SRAD AltimeterCurrent consumption ·
For triggering deploymentWe will be using nichrome wire of 22 gauge that hasLength(l) = 5 cm = 0.05m
Radius(r) = 0.3213 mm = 0.0003213m
ρ = 1.1x10-6 Ωm
R=ρx
𝑙
𝐴
−6
= 1.1 x 10 x
−6
= 1.1 x 10 x
−2
(5 × 10 )
2
π𝑟
−2
(5 × 10 )
3.14 × 0.0003213 × 0.0003213
= 16.959Ω
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Team #11 STES India
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Using Ohm’s LawI = V/R
= 11.1/16.959
= 0.654 A=654mA
·
By other components included in SRAD Altimeter-
Sr.no
Components
Current Consumed
1
Bmp280
0.0027mA
2
Adafruit Ultimate GPS
20mA
3
Adafruit Micro SD Card
100mA
4
Xbee S3B Pro
215mA
5
Arduino Nano
19mA
Total current
354.0027mA
Total current consumed = current consumed by deployment + current consumed by sensors
= 654mA + 354.0027 mA
= 1008.0027mA
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